Abstract: A vane multiplet includes first and second ceramic matrix composite (CMC) singlet vanes that are arranged circumferentially adjacent each other. Each of the CMC singlet vanes includes an airfoil section and a platform at one end of the airfoil section. The platform defines forward and trailing platform edges and first and second circumferential side edges. A CMC overwrap conjoins the CMC singlet vanes. The CMC overwrap includes fiber plies that are fused to the platforms of the CMC singlet vanes.

Abstract: A sealing assembly for a gas turbine engine. The sealing assembly includes first and second gas turbine walls defining a channel therebetween. Additionally, the second gas turbine wall further defines a passage extending therethrough. Furthermore, the sealing assembly includes a leaf seal partially positioned within the channel and a seal holder coupled to the second gas turbine wall. Moreover, the sealing assembly includes a spring compressed between the seal holder and the leaf seal such that the leaf seal is in sealing engagement with the first gas turbine wall. In addition, the sealing assembly includes a pin extending through the passage defined by the second gas turbine wall to couple the seal holder and the leaf seal such that the pin is thermally unconstrained by the second wall during operation of the gas turbine engine.

Abstract: A turbine section has: a rotor rotatable about a central axis, the rotor having blades each protruding radially outwardly from a platform relative to the central axis; a stator having vanes each protruding radially outwardly from a shroud; a rim seal between the platform and the shroud, the rim seal having: an axial overlap between the platform and the shroud, and a lip protruding in a direction having a radial component relative to the central axis from one of the platform and the shroud toward the other of the platform and the shroud, the lip axially overlapping the other of the platform and the shroud, the lip having a radial height such that a radial gap remains between the lip and the other of the platform and the shroud when the turbine section is in operation.

Abstract: An airfoil assembly for a turbine engine defines an axial direction, a radial direction, and a circumferential direction, and includes a first airfoil defining a first end along the radial direction, a first hub disposed on the first end of the first airfoil and having a first extension member extending at least partially in the radial direction, and a second airfoil adjacent to the first airfoil, the second airfoil defining a first end along the radial direction, a second hub disposed on the first end of the second airfoil and comprising a second extension member extending at least partially in the radial direction, and a circumferential bias assembly operable with the first extension member, the second extension member, or both for exerting a circumferential force on the first extension member, the second extension member, or both.

Abstract: Structures, such as compressor casings, for reducing a thermal gradient are provided. For example, a compressor case is split such that it includes first and second case segments. The first case segment extends over a first portion of the compressor case circumference and comprises a first inner structure, a first outer structure, and a first porous structure integrally formed as a monolithic component. The first porous structure is defined between the first inner structure and the first outer structure. The second case segment extends over a second portion of the compressor case circumference and comprises a second inner structure, a second outer structure, and a second porous structure integrally formed as a monolithic component. The second porous structure is defined between the second inner structure and the second outer structure. Methods of cooling structures such as compressor casings also are provided.

Abstract: A turbine assembly for use with a gas turbine engine includes a bladed wheel assembly, a vane assembly, and an inner seal. The bladed wheel assembly is adapted to interact with gases flowing through a gas path of the gas turbine engine. The vane assembly is located upstream of the bladed wheel assembly and adapted to direct the gases at the bladed wheel assembly. The inner seal is configured to block gases from passing around the vane assembly.

Abstract: The invention concerns a turbine nozzle, comprising a plurality of angular nozzle sectors (6) each angular sector (6) comprising two inner and outer platform sectors, connected together by a plurality of radial blades (63), each inner platform sector (62) being rigidly attached to a radially inner foot (621), this nozzle comprising an annular collar (5) on which the nozzle angular sectors are fastened end-to-end circumferentially, this collar (5) comprising a cylindrical ring (50) the radial inner face (51) of which comprises an abradable material (53). This nozzle is characterised in that the radially inner foot (621) of each inner platform sector comprises a radial tab (622) and in that the annular ring (50) comprises a plurality of L-shaped pads (55), each pad delimiting a slot (553) for receiving the tab (622), so as to ensure the fastening by coupling of the collar on each nozzle angular sector.

Abstract: A gas turbine spacer disk includes a disk portion, a rim portion, a first fillet, and a second fillet. The disk portion is disposed about a rotational axis. The rim portion is disposed about the disk portion. An outer face of the rim portion defines a plurality grooves extending circumferentially about the rotational axis. The first fillet transitions from the rim portion to a first side of the disk portion. The second fillet transitions from the rim portion to a second side of the disk portion. The plurality of grooves includes a pair of first grooves having a first diameter and a pair of second grooves having a second diameter that is less than the first diameter. A first one of the first grooves overlaps in an axial direction with the first fillet. A second one of the first grooves overlaps in the axial direction with the second fillet.

Abstract: An airfoil assembly includes an airfoil. The airfoil has first and second platforms and an airfoil between the first and second platforms. A support ring is configured to retain the first platform. The support ring has a lip which extends radially inward from the support ring. The lip is configured to engage an axial face of the first platform. The lip has a primary retention feature and a secondary retention feature. The primary and secondary retention features are configured to retain the first face of the first platform with respect to the annular ring. A support structure for an airfoil and a method of retaining an airfoil assembly are also disclosed.

Abstract: A rotor assembly of a gas-turbine engine may comprise a first rotor blade, a second rotor blade, a third rotor blade, a first platform sealing assembly and a second platform sealing assembly. The first platform sealing assembly may be disposed between a first platform of the first rotor blade and a second platform of the second rotor blade. The second platform sealing assembly may be disposed between the second platform and a third platform of the third rotor blade.

Abstract: A ring for guiding variable-pitch blades and supporting an abradable coating for an aircraft turbomachine ring extends about an axis (X) and includes substantially radial orifices for mounting guiding sleeves intended each to receive a pivot of one of the blades. The ring may include a radially inner crown divided into sectors and supports the abradable coating. A radially outer annular ferrule may be divided into sectors and include the orifices for mounting the sleeves. In some embodiments, each ferrule sector is mounted about a crown sector by a circumferential sliding connection. The crown sector may include an immobilizer for rotationally immobilizing the sleeves of this ferrule sector about the axis.

Abstract: The present application provides a rotor for use in a turbine engine. The rotor may include a first rotor disc, a second rotor disc adjacent to the first rotor disc, and a belly band seal positioned between the first rotor disc and the second rotor disc. The belly band seal may include a band seal with a locking tab brazed thereon.

Abstract: A rotor assembly is disclosed. In various embodiments, the rotor assembly includes a first rotor having a first plurality of tabs; a second rotor having a second plurality of tabs; and a rotor coupling ring having an anti-rotation tab configured for disposition between an adjacent pair of tabs from one of the first plurality of tabs or the second plurality of tabs.

Abstract: A steam turbine diaphragm nozzle segment, related assembly and steam turbine. Various embodiments include a steam turbine diaphragm nozzle segment having: a pair of opposing sidewalls; an airfoil extending between the pair of opposing sidewalls and integral with each of the pair of sidewalls, the airfoil having a single contact surface for directing a flow of working fluid through a flow channel; and a fill region integral with the airfoil and the pair of opposing sides, the fill region extending between the pair of opposing sides along an entirety of a length of the airfoil, the fill region for completely obstructing the flow of working fluid.

Abstract: A turbomachine assembly and, in particular, a low-pressure compressor of an aircraft turbojet engine includes an annular row of upstream vanes with trailing edges extending radially from an upstream support; an annular row of downstream vanes with leading edges axially facing the trailing edges and extending radially from a downstream support; an annular passageway delimited by the upstream support and the downstream support. The downstream support has a profile with: an upstream portion delimiting the annular passageway forming an annular slide, a downstream portion axially at the level of downstream vanes, and a connecting arc connecting the upstream portion to the downstream portion. The connecting arc is arranged downstream of the leading edges.

Abstract: The invention concerns a turbine for a gas turbine comprising a blade, a vane and an abradable lip attached to the blade or to the vane, wherein the blade and the vane are separated by a gap and the abradable lip extends part of the distance across the gap. Embodiments include the addition of an abrasive layer attached on the other side of the gap from the abradable lip. A method of manufacturing is also described.

Abstract: A disk of a rotor including an annular radial web, a radially central hub located at the inner radial end of the web and a rim located at the outer radial end of the web, the web including an upstream face and a downstream face, and a plurality of orifices through which bolts pass for the attachment of at least one annular flange forming part of another adjacent rotor disk on either the upstream face or the downstream face of the web, or on both faces. The upstream face and/or the downstream face of the web includes a globally annular shaped indentation, with a bottom set back along the axial direction inwards into the web, and that extends radially outwards from the hub of the disk towards the rim, and that surrounds a radially inner part of each of the orifices of the web, at a distance.

Abstract: The present disclosure is directed to a rotor blade and method for forming the rotor blade. The rotor blade includes a platform having a bottom side radially spaced from a top side and a leading edge portion axially spaced from a trailing edge portion. An airfoil extends radially outwardly from the top side of the platform and a shank extends radially inwardly from the bottom side of the platform. The shank includes a lip that extends axially outwardly from a forward wall of the shank. The lip defines a radially inward surface and a radially outward surface and a plurality of slots. Swirler vane inserts are disposed within respective slots of the plurality of slots. Each swirler vane insert extends radially inwardly from the inward surface of the lip and axially outwardly from the forward wall of the shank.

Abstract: A component joint including a turbomachine component (10) having at least one component flange (11); a mating turbomachine component (20) connected thereto and having at least one mating component flange (21); and at least one seal (30) to provide a seal against an overpressure (p2?p1) on the component side, the at least one seal (30) having a sealing surface (31) that engages into a groove (12) of the turbomachine component (10), the groove (12) being disposed on a mating-component-flange-facing side of the component flange (11) and having a contact surface (13) for sealingly axially supporting the sealing surface (31).

Abstract: A gas turbine engine includes a coupling and an air seal attachment. The annular coupling extends along a centerline and has coupling teeth that extend axially rearward. The annular air seal attachment includes an air seal at an axially rearward end and air seal attachment teeth at an axially forward end that extend axially forward to interlock with the coupling teeth such that the air seal attachment teeth and the coupling teeth alternate in a circumferential direction.

Abstract: A turbine for a turbine engine, the turbine comprising a rotor including blades the radially external periphery of which includes at least one first wiper radially extending outwards, sealing means radially extending about the blades and including a ring made of abradable material, with the radially external ends of the wipers cooperating with said ring made of abradable material so as to form a labyrinth-type seal, wherein said ring includes at least one first portion axially extending upstream of the first wiper and a second portion, different from the first portion, axially extending downstream of the first wiper, with the first portion and/or the second portion including a groove, wherein the first wiper has been inserted, with said groove being defined by the first portion and by the second portion.

Abstract: A rim seal located between a stator and a rotor in a gas turbine engine includes an axial overlap between platform rims of the stator and the rotor. The axial length of the platform rim of one of the stator and the rotor is accommodated within a radially aligned trench defined in the other of the stator and the rotor.

Abstract: A rotor includes a rotor hub, a plurality of blades mounted in a circumferentially-spaced arrangement on the rotor hub and a plurality of platform segments circumferentially arranged, respectively, between neighboring ones of the blades. The platform segments include core gas-path defining surfaces and are mounted with a freedom to circumferentially move relative to the neighboring ones of the blades.

Abstract: A rotor is provided for a gas turbine engine. The rotor includes a rim that extends transverse to the web. The rim has a rotor spacer arm which defines a coating pocket and a stress relief protrusion opposite the coating pocket.

Abstract: Turbomachinery hardware, used in a rotor assembly and a stator assembly, including an airfoil portion including a leading edge, a trailing edge, a pressure side, and a suction side, and a platform on which the airfoil portion is disposed. The platform including a platform axis, a pressure side mateface located adjacent to the pressure side of the airfoil portion and a suction side mateface located adjacent to the suction side airfoil portion, wherein a portion of a pressure side mateface includes a first geometry, and a portion of a suction side mateface includes a second geometry. The first geometry is selected from a group consisting of: oblique to a platform axis, and a first curved portion. The second geometry is selected from a group consisting of: oblique to the platform axis and a second curved portion.

Abstract: A turbine rotor for a gas turbine engine, includes an upstream turbine disk; a downstream turbine disk; an annular flange; a first ferrule connecting the upstream turbine disk to the annular flange; a second ferrule connecting the downstream turbine disk to the annular flange; an air flow separator device including: a first part, forming a first ring, arranged between the upstream turbine disk and the downstream turbine disk; a second part, forming a second ring, having a first portion facing the downstream turbine disk, and a second portion arranged between the first ferrule and the second ferrule; and a thermal insulation area arranged between the first part and the second part.

Abstract: A gas turbine engine includes a turbine section having static turbine vane rings and turbine wheel assemblies. The vane rings are arranged to direct combustion products toward blades included in the turbine wheel assemblies to cause the turbine wheel assemblies to rotate.

Abstract: An air seal may comprise an annular ring defined by at least a proximal surface, a distal surface, an aft side and a forward side. A channel may be disposed in the forward side of the air seal and/or the aft side of the air seal and may extend between the proximal surface and the distal surface. An additional channel extending from at least one of the forward side or the aft side may be disposed in the distal surface. The channel and the additional channel may be circumferentially in line. The channels may define a flow path for direction cooling air from a proximal side of the air seal to a distal side of the air seal.

Abstract: A disc of a gas turbine engine system with particular retaining ring placement and engagement implementation are provided. For example, the disc includes a disc bore including, a groove formed on an axially extending surface of the disc bore, wherein the groove includes a forward surface that extends radially into the disc bore to a groove floor that is cut into the disc bore, and wherein the groove extends axially to an aft surface that extends radially outward to at least the axially extending surface of the disc bore, and a disc web that extends radially outward from the disc bore, relative to an axis of rotation of the gas turbine engine.

Abstract: A disc of a gas turbine engine system is provided. The disc includes a disc bore including, an interstage coupling disposed on an axially extending surface at a peripheral edge of the disc bore that includes a protrusion that extends axially beyond an outer surface of the disc bore, and a groove formed on the axially extending surface of the disc bore wherein an aft surface of the groove is also a forward surface of the interstage coupling, wherein the groove includes a forward surface that extends radially into the disc bore to a groove floor that is cut into the disc bore and extends axially to the aft surface that extends radially outward to at least the axially extending surface of the disc bore, and a disc web that extends radially outward from the disc bore, relative to an axis of rotation of the gas turbine engine.

Abstract: An integrally bladed fan (IBF) rotor of a gas turbine engine. The IBF rotor includes a hub and a plurality of fan blades extending radially outwardly from the hub and integral therewith. The hub has a fan attachment flange disposed at an end of the hub on a trailing edge side thereof for mounting a booster rotor to a trailing edge side of the fan. The fan attachment flange is disposed at a radial distance from a longitudinal center axis of the integrally bladed fan rotor. The hub has an outer hub surface disposed radially inward from the radial distance of the fan attachment flange.

Abstract: A sealing system for a multi-stage turbine includes multiple interstage seal segments disposed circumferentially about a turbine rotor wheel assembly and extending axially between a forward turbine stage and an aft turbine stage. Each of the interstage seal segments includes a forward end portion including an outer seal surface and an inner support face, an aft end portion, including an outer seal surface and an inner support face and a main body portion extending axially from the forward end portion to the aft end. The main body portion includes at least two support webs coupling the outer seal surfaces and the inner support faces. The outer seal surfaces are configured to be retained in a radial direction by a land support on each of a forward and aft stage turbine buckets, such that substantially all the centrifugal load from the multiple interstage seal segments is transferred to the forward and aft stage turbine buckets. A method of assembling the sealing system is disclosed.

Abstract: An integrally bladed rotor includes a rotor that has a rim that provides an inner flow surface. Circumferentially spaced apart radially extending airfoils integrally with and from the rotor and joined by an airfoil fillet. An asymmetrical trench is provided in the rim between adjacent airfoils.

Abstract: One exemplary embodiment of this disclosure relates to a gas turbine engine. The engine includes a first rotor disk, a second rotor disk, and a circumferentially segmented seal. The segmented seal engages the first rotor disk and the second rotor disk. The segmented seal further includes a fore surface contacting the first disk, an aft surface contacting the second disk, and a radially outer surface. Further, (1) the aft surface and (2) one of the fore surface and the radially outer surface include perforations to allow fluid to flow through the interior of the segmented seal.

Abstract: A blade cascade for a turbomachine, in particular a gas turbine, including multiple identical blade groups which are situated next to each other and which each include a first individual blade (10) having a vane (12) having a blade profile and a first side wall (11), and a second individual blade (20), which differs from the first, having a vane (22) having a blade profile and a second side wall (21), the first and second side walls (11, 21) having different contourings.

Abstract: The present disclosure generally relates to a rotor assembly, and in particular relates to an improved rotor heat shield which provides an innovative configuration for securing the same to the rotor assembly. The rotor heat shield element is secured to the rotor assembly in correspondence of the groove in which it is inserted. Embodiments of the present disclosure can allow the removal of current fixation features on heat shields and blades. Furthermore, since the heat shield is no longer connected to a blade but directly to the rotor assembly, there is more freedom in selecting the number of heat shield elements to be provided to form the circumferential heat shield.

Abstract: A sealing system in a cavity under the stator of a turbomachine flow path, the cavity being located between a root of a vane of the stator and a complementary rotor device, the root including two surfaces each provided with an abradable coating, the rotor device being provided with a first and a second sealing element, arranged facing the first and the second surface respectively, the first surface and the first sealing element forming a first sealing pair and delimiting a first leakage section between them, the second surface and the second sealing element forming a second sealing pair and delimiting a second leakage section between them, one of the two pairs tending towards a minimum leakage section when the other tends towards a maximum leakage section, and vice versa.

Abstract: A gas turbine engine including an axial high pressure compressor having expansion slots in the outer rim of the rotor section. The expansion slots may be positioned between blades of a rotor segment. The fore end of the slots may have an axial seal which is coupled to the inner surface of the outer rim in the first rotor segment, and may comprise a fin configuration. The axial seal may be integral to the inner surface of the outer rim. The compressor may comprise a plurality of expansion slots and axial seals, including in a plurality of rotor segments.

Abstract: A device with a torque-proof first structural component and a second structural component that is connected at least in certain areas in a rotatable manner to the first structural component, wherein hydraulic fluid can be guided to lubrication points via the first structural component and the second structural component. The second structural component is embodied with blade areas which are extending substantially at a defined angle in the radial direction inside the second structural component and between which transmission areas for hydraulic fluid of the second structural component are provided, with their flow cross sections decreasing in the transmission areas in the flow direction of the hydraulic fluid.

Abstract: An assembly for a gas turbine engine includes a first module, a second module interconnected with the first module, a first gas turbine component, and a seal assembly. The first gas turbine component includes a body extending generally radially between an inner mounting portion and an outer seal carrier portion. The inner mounting portion is fastened to the first module such that the first gas turbine component forms a boundary between a first annular cavity in the first module, and a second annular cavity in the second module. The seal assembly is mounted to the seal carrier portion of the first gas turbine component, and includes a flow diverter assembly having a flow diverter and a finger seal commonly fastened to a first axial side of the seal carrier portion.

Abstract: The invention relates to a fixed diffuser vanes assembly (10) for guiding flow through a turbomachine, comprising an internal annular platform (12) and a plurality of fixed vanes (11) which are mounted on this platform, the internal platform comprising a support plate (121) forming the base of said vanes, a radial annular partition (120) extending from the support plate toward an axis of the vanes assembly, and an internal ring (122) attached to the radial annular partition and having an internal surface on which an abradable material (123) is arranged, the vanes assembly being characterized in that the internal ring comprises at least one cut-out (1223) delimiting a tongue (1226), the tongue being bent to press against the radial annular partition (120). The invention also relates to a method of manufacturing such a vanes assembly and to a turbomachine incorporating such vanes.

Abstract: Anti-rotation knife edge seals and gas turbine engines including the same are provided. The anti-rotation knife-edge seal includes a first end portion having at least one knife edge, an opposing second end portion including an outside peripheral edge having a nut anti-rotation feature and an inside peripheral edge having a shaft anti-rotation feature, and an intermediate portion extending between the first end portion and the opposing second end portion. The first end portion, the opposing second end portion, and the intermediate portion form a one-piece annular seal member. A method is also provided for substantially preventing rotation of a stack nut against an adjacent rotatable component.

Abstract: A fan section for a gas turbine engine is provided. The fan section having a fan hub with a slot and a fan blade with an airfoil extending from a root to a tip, the airfoil having a leading edge and the root is received in the slot, wherein a first platform is secured to the fan hub and arranged between adjacent fan blades of the fan section, the first platform having a first platform seal including a platform seal leading edge and a base is secured to a side of the first platform and a first winglet extends from the platform seal leading edge and contacts the airfoil leading edge.

Abstract: A shaped rim cavity wing includes an upper surface and a lower surface. The lower surface has a geometric shape to control the separation of airflow as it passes around the lower surface to the top surface. A point of maximum extent defines the boundary between the upper surface and the lower surface, wherein the point of maximum extent defines a corner that that separates airflow from the shaped rim cavity rim and creates a flow re-circulation adjacent to the top surface of the shaped rim cavity wing.

Abstract: A turbine shroud for a turbine of a gas turbine engine is disclosed. The turbine shroud is configured to direct products of a combustion reaction in a combustor of the gas turbine engine toward a plurality of rotatable turbine blades of the turbine to cause the plurality of turbine blades to rotate.

Abstract: A compressor has a plurality of stages, and each of the stages includes a wheel. Each wheel is configured to receive one or more rotor blades. The compressor includes a patch ring mounted on a first rotor wheel. The first rotor wheel has a first rabbet machined therein, and the patch ring is located on the first rabbet. A second rotor wheel is located upstream from the first rotor wheel. The second rotor wheel has a second rabbet machined therein, and the second rabbet is configured to be located opposite to the first rabbet. The second rabbet has a second fillet located between axial and radial surfaces of the second rabbet. The patch ring has a radial height greater than or equal to a radial height of the second fillet.

Abstract: An interface within a gas turbine engine includes a multiple of segmented components, each with a segment flange with a multiple of apertures, at least one of the multiple of apertures a first slot aperture. A full ring component with a ring flange that defines a multiple ring of apertures, at least one of the multiple of ring apertures a second slot aperture, the second slot aperture transverse to the first slot aperture.

Abstract: A hybrid metal and composite spool includes metal rings on an outer diameter of a composite spool shell. Metal rings may include features such as annular or axial dovetail slots. Adhesive layers may be between the metal rings and composite shell which may be connected by a shrink bonded joint. The metal rings may include a single seal tooth ring with an annular radially extending seal tooth. A method for fabricating the spool may include fabricating one or more metal rings with the features therein, positioning the metal rings in place on an outer surface of an uncured composite spool shell of the spool before curing the shell, and curing the shell with the one or more metal rings positioned in place. Alternatively, rings may be heated to a temperature at least sufficient to slide rings over a cured composite shell, and allowed to cool and shrink onto shell.

Abstract: A turbine in a gas turbine engine that includes a stator blade and a rotor blade having a seal formed in a trench cavity. The trench cavity may include an axial gap defined between opposing inboard faces of the stator blade and rotor blade. The seal may include: a stator overhang extending from the stator blade toward the rotor blade so to include an outboard edge and an inboard edge and, defined therebetween, an overhang face; a rotor outboard face extending radially inboard from a platform edge, the rotor outboard face opposing at least a portion of the overhang face across the axial gap of the trench cavity; and a first axial projection extending from the rotor outboard face toward the stator blade. The stator overhang and the first axial projection of the rotor blade may be configured so to axially overlap.

Abstract: A turbomachine including a rotor having an axis and a plurality of disks positioned adjacent to each other in the axial direction, each disk including opposing axially facing surfaces and a circumferentially extending radially facing surface located between the axially facing surfaces. At least one row of blades is positioned on each of the disks, and the blades include an airfoil extending radially outward from the disk A non-segmented circumferentially continuous ring structure includes an outer rim defining a thermal barrier extending axially in overlapping relation over a portion of the radially facing surface of at least one disk, and extending to a location adjacent to a blade on the disk A compliant element is located between a radially inner circumferential portion of the ring structure and a flange structure that extends axially from an axially facing surface of the disk.