Abstract: A fish mouth seal carrier for a guide vane arrangement of a gas turbine comprises a first half-shell element and a second half-shell element bonded to it, which together form a box profile with two axial arms and two radial arms. A sealing element is arranged on one of the axial arms of the box profile. At least one of the two half-shell elements has an integrally formed axial flange for forming, with a guide vane platform, a fish mouth seal accommodating an axial flange of an adjoining moving blade radially between the guide vane platform and the axial flange of the seal carrier.

Abstract: A flow discouraging system includes a stator assembly, a rotor assembly, and a plurality of fingers. The stator assembly includes one or more stationary components forming a side wall, the side wall including an annular groove defined by an outer axially-extending surface, an inner axially-extending surface, and a radial surface extending between the outer and inner axially-extending surfaces. The rotor assembly is disposed adjacent to and spaced apart from the stator assembly to form a portion of a cavity and includes an annular rim extending at least partially into the annular groove. The plurality of fingers is disposed in the annular groove and extends from the one or more stationary components the annular rim.

Abstract: A seal in a turbine of a combustion turbine engine is described. The seal is formed within a trench cavity defined between a rotor blade and a stator blade. The stator blade includes a sidewall projection and the rotor blade includes an angel wing projection extending toward the stator blade. The side wall projection overhangs the angel wing projection. The seal include: a port disposed on an inboard surface of the stator projection; and deflecting structure disposed on the angel wing projection. The deflecting structure may be configured to receive the fluid expelled from the port and deflect the fluid toward an inlet of the trench cavity.

Abstract: A seal assembly provided between a hot gas path and a disc cavity in a turbine engine includes an annular outer wing member extending from an axially facing side of a rotor structure toward an adjacent non-rotating vane assembly, and a plurality of fins extending radially inwardly from the outer wing member and extending toward the adjacent non-rotating vane assembly. The fins are arranged such that a space having a component in a circumferential direction is defined between adjacent fins. Rotation of the fins during operation of the engine effects a pumping of purge air from the disc cavity toward the hot gas path to assist in limiting hot working gas leakage from the hot gas path to the disc cavity by forcing the hot working gas away from the seal assembly.

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 compressor for a turbine engine includes multiple compressor disks having rotor blades mounted about the circumference of each of the disks. A plurality of stator blades extend between the rotor blades of axially adjacent disks. A knife edge seal is supported and retained by retaining flanges extending from a rim on each disk, and contacts the stator blades to limit the recirculation of air within the compressor. A plurality of lock assemblies are spaced about the circumference of disk backbones formed in each disk, with a plurality of knife edge seals located between each lock assembly. When in the lock position the lock assemblies reduce the slack used for assembly of the final knife edge seal to prevent shifting and rotating during operation.

Abstract: A system includes a multi-stage turbine. The multi-stage turbine has an interstage seal extending axially between a first turbine stage and a second turbine stage. The interstage seal has an upper body that extends from an upstream seating arm to a downstream seating arm. The upstream and downstream seating arms are designed to constrain movement of the interstage seal along a radial direction of the multi-stage turbine. The interstage seal also has a lower body that extends from a seating end to a hook end. The seating end is designed to constrain movement of the interstage seal along the radial direction. The hook end has a protrusion that extends crosswise relative to a base of the lower body. The hook end is designed to constrain movement of the interstage seal along the radial direction and an axial direction of the multi-stage turbine.

Abstract: A turbine engine section has a stationary vane stage and a rotating blade stage. The blade stage is spaced from the vane stage to form an annular chamber therebetween. A manifold delivers a pressurized fluid to the chamber. An outer diameter (OD) sealing system restricts leakage of the pressurized fluid from the chamber. An inner diameter (ID) sealing system restricts leakage of the pressurized fluid from the chamber. A flow guide extends radially between the inner diameter sealing system and the manifold. The flow guide bisects the chamber to form a stationary chamber portion and a rotating chamber portion, the rotating chamber portion at least partially along a disk of the first blade stage.

Abstract: In accordance with one embodiment, a system includes a multi-stage turbine including a first turbine stage including a first wheel having multiple first slots spaced circumferentially about the first wheel, and multiple first blade segments each coupled to at least one of the multiple first slots. The multi-stage turbine also includes a second turbine stage including a second wheel having multiple second slots spaced circumferentially about the second wheel, and multiple second blade segments each coupled to at least one of the multiple second slots. The multi-stage turbine further includes a one-piece interstage seal extending axially between the first and second turbine stages, wherein the one-piece interstage seal comprises a first radial support coupled to the first wheel and a second radial support coupled to the second wheel, and the one-piece interstage seal extends circumferentially across at least two of the multiple first slots or at least two of the multiple second slots.

Abstract: A near flow path seal member for a turbomachine includes a seal body having a seal support member including a first end portion that extends to a second end portion through an intermediate portion. An arm member extends from the first end portion of the seal body. The arm member has a first end that extends to a second end to define an axial dimension of the arm member, a first edge that extends to a second, opposing edge to define a circumferential dimension of the arm member, and a surface having a profile that establishes a thickness variation of the arm member in each of the axial dimension and the circumferential dimension.

Abstract: A seal apparatus in a gas turbine engine between a transition duct and a vane structure in a first row vane assembly. The seal apparatus includes a transition seal structure associated with the transition duct and a vane seal structure associated with the vane structure. The transition seal structure includes a transition base portion that defines a first axially facing surface, a first transition lip member extending axially from the first axially facing surface, and a second transition lip member radially spaced from the first transition lip member and extending axially from the first axially facing surface. The vane seal structure includes a vane base portion that defines a second axially facing surface, a first vane lip member extending axially from the second axially facing surface, and a second vane lip member radially spaced from the first vane lip member and extending axially from the second axially facing surface.

Abstract: A turbine section has a turbine rotor carrying turbine blades. The turbine blades include seal members at a radially inner location. A vane section is formed of a plurality of circumferentially spaced vane components, each of which has an airfoil extending radially outwardly of a platform. A first seal member is fixed to the platform, and is positioned to be adjacent a seal from a blade which is positioned in one axial direction relative to the first seal member. A second seal member extends circumferentially beyond at least a plurality of the vane components and is positioned to be adjacent a seal member of a blade on an opposed axial side from the first blade. A vane component is also disclosed and claimed.

Abstract: A turbomachine includes a motor including a rotor assembly and a stator assembly. A sleeve fluidly separates the rotor assembly from the stator assembly. The sleeve has first and second ends axially spaced from one another. In one example, first and second seals are arranged at the first end and fluidly separate the rotor assembly from the stator assembly. The second seal is arranged downstream from the first seal relative to the rotor assembly. In another example, the sleeve includes wet and dry sides. A first seal is arranged at the first end on the wet side. A drain is provided in the housing and arranged in a region downstream from the first seal. The drain fluidly connects the region to an exterior of the housing.

Abstract: A turbocharger system, in certain embodiments, includes a compressor, a turbine, a shaft of common diameter coupling the compressor to the turbine, and a first fluid film bearing disposed about the shaft at a compressor end portion of the shaft. The system also includes a second fluid film fixed pad bearing disposed about the shaft at a turbine end portion of the shaft, wherein the first and second fluid film fixed pad bearings have different clearance ratios, effective lengths, or both, relative to one another. The system, in some embodiments, includes a compressor fluid film fixed pad bearing and a turbine fluid film fixed pad bearing, wherein the compressor and turbine fluid film fixed pad bearings have different clearance ratios and effective lengths, relative to one another.

Abstract: A turbine airfoil seal system of a turbine engine having a seal base with a plurality of seal strips extending therefrom for sealing gaps between rotational airfoils and adjacent stationary components. The seal strips may overlap each other and may be generally aligned with each other. The seal strips may flex during operation to further reduce the gap between the rotational airfoils and adjacent stationary components.

Abstract: A compressor stator blade 10 includes a seal mechanism 12 configured to hermetically seal a part between a radial inner end on the upstream side and an outer surface of a rotary body, and a remainder part on the downstream side of the seal mechanism includes a hub clearance 14 between the outer surface of the rotary body and itself. Additionally, a compressor rotor blade 20 includes a seal mechanism 22 configured to hermetically seal a part between a radial outer end on the upstream side and an inner surface of a stationary body, and a remainder part on the downstream side of the seal mechanism includes a tip clearance 24 between the inner surface of the stationary body and itself.

Abstract: A seal assembly that limits gas leakage from a hot gas path to one or more disc cavities in a turbine engine. The seal assembly includes a seal apparatus that limits gas leakage from the hot gas path to a respective one of the disc cavities. The seal apparatus comprises a plurality of blade members rotatable with a blade structure. The blade members are associated with the blade structure and extend toward adjacent stationary components. Each blade member includes a leading edge and a trailing edge, the leading edge of each blade member being located circumferentially in front of the blade member's corresponding trailing edge in a direction of rotation of the turbine rotor. The blade members are arranged such that a space having a component in a circumferential direction is defined between adjacent circumferentially spaced blade members.

Abstract: A disc seal assembly for use in a turbine engine. The disc seal assembly includes a plurality of outwardly extending sealing flange members that define a plurality of fluid pockets. The sealing flange members define a labyrinth flow path therebetween to limit leakage between a hot gas path and a disc cavity in the turbine engine.

Abstract: A sealing apparatus in a gas turbine. The sealing apparatus includes a seal housing apparatus coupled to a disc/rotor assembly so as to be rotatable therewith during operation of the gas turbine. The seal housing apparatus comprises a base member, a first leg portion, a second leg portion, and spanning structure. The base member extends generally axially between forward and aft rows of rotatable blades and is positioned adjacent to a row of stationary vanes. The first leg portion extends radially inwardly from the base member and is coupled to the disc/rotor assembly. The second leg portion is axially spaced from the first leg portion, extends radially inwardly from the base member, and is coupled to the disc/rotor assembly. The spanning structure extends between and is rigidly coupled to each of the base member, the first leg portion, and the second leg portion.

Abstract: A disc seal assembly for use in a turbine engine. The disc seal assembly includes a plurality of outwardly extending sealing flange members that define a plurality of fluid pockets. The sealing flange members define a labyrinth flow path therebetween to limit leakage between a hot gas path and a disc cavity in the turbine engine.

Abstract: An exemplary turbomachine seal assembly includes a ring seal that has a radially-facing sealing surface, an axially-facing sealing surface, and a first tapered surface. The turbomachine seal assembly also includes a spacer that has a second tapered surface. The second tapered surface interfaces with the first tapered surface to bias the radially facing sealing surface and the axially facing sealing surface toward a sealing position.

Abstract: A turbomachine seal assembly includes a base member, a rocker arm pivotally mounted to the base member, and a seal element fixedly mounted to the rocker arm. The seal element is configured and disposed to selectively shift relative to the base member.

Abstract: A seal assembly for a gas turbine engine includes an annular body and a flow-through tube that extends through the annular body. The flow-through tube includes an upstream orifice, a downstream orifice and a tube body that extends between the upstream orifice and the downstream orifice. The tube body establishes a gradually increasing cross-sectional area between the downstream orifice and the upstream orifice.

Abstract: A gas turbine includes forward and aft rows of rotatable blades, a row of stationary vanes between the forward and aft rows of rotatable blades, an annular intermediate disc, and a seal housing apparatus. The forward and aft rows of rotatable blades are coupled to respective first and second portions of a disc/rotor assembly. The annular intermediate disc is coupled to the disc/rotor assembly so as to be rotatable with the disc/rotor assembly during operation of the gas turbine. The annular intermediate disc includes a forward side coupled to the first portion of the disc/rotor assembly and an aft side coupled to the second portion of the disc/rotor assembly. The seal housing apparatus is coupled to the annular intermediate disc so as to be rotatable with the annular intermediate disc and the disc/rotor assembly during operation of the gas turbine.

Abstract: A turbine is provided and includes a spacer having an annular body formed with opposing outward and inward surfaces and an orifice extending through the body from the outward to the inward surface, an assembly to secure the spacer around a rotor axially between sequential buckets of a forward turbine stage and an aft turbine stage, the spacer forming an annular passage around the rotor into which a fluid flows through the orifice and a circuit fluidly coupled to the annular passage to deliver the fluid from between the sequential buckets of the forward turbine stage and the aft turbine stage to an axial location forward of the forward turbine stage.

Abstract: A gas turbine inner flow path cover piece for a gas turbine a first turbine wheel and a second turbine wheel is provided is provided. The gas turbine inner flow path cover piece can include a main body having an first surface and a second surface, side pieces disposed on the first surface of the main body and mating pairs disposed on the second surface of the main body.

Abstract: A card seal with an annular arrangement of plates that form the card seal, where the plates are rotated at an angle around a longitudinal axis of the plates so that the gaps between adjacent plates are not parallel to the rotational axis of the rotor shaft, and where an annular disk is rotatably secured to the rotor shaft and located adjacent to card seal on the low pressure side, or on the high pressure side or even on both sides such that rotation of the annular disk pumps fluid into or draws fluid out from the plate gaps to decrease or eliminate any differential pressure across the card seal. Thus, card seal leakage is reduced or eliminated.

Abstract: A seal assembly for a gas turbine engine is provided that includes a seal support member, at least one knife edge seal blade member, and at least one cooling member. The seal support member is configured for rotation within the gas turbine engine around an axial centerline of the engine. The support member has a thickness extending between a first surface and a second surface. The blade member extends outwardly from the first surface of the support member. The blade member has a central portion that extends between a base end and a distal knife edge end. The base end is attached to the first surface of the seal support member. The cooling member extends outwardly from the second surface, and is oppositely aligned with the blade member.

Abstract: Gas turbine engine systems involving rotor bayonet coverplates and tools for installing such coverplates are provided. In this regard, a representative turbine assembly for a gas turbine engine includes: a turbine disk operative to mount a set of turbine blades; and a coverplate having an annular main body portion and a spaced annular arrangement of tabs extending radially inwardly from the main body portion with open-ended gaps being located between the tabs, the tabs being operative to secure an inner diameter of the coverplate to the turbine disk.

Abstract: A seal formed between at least two blades in the turbine of a turbine engine, a first turbine blade and a second turbine blade, wherein one of the turbine blades comprises a turbine rotor blade and the other turbine blade comprises a turbine stator blade, and wherein a trench cavity and the seal is formed between the first turbine blade and the second turbine blade when first turbine blade is circumferentially aligned with the second turbine blade, the seal comprising: a cutter tooth and a honeycomb; wherein: the cutter tooth comprises an axially extending rigid tooth that is positioned on one of the first turbine blade and the second turbine blade and the honeycomb comprises an abradable material that is positioned on the other of the first turbine blade and the second turbine blade; and the cutter tooth and the honeycomb are positioned such that each opposes the other across the trench cavity when the first turbine blade is circumferentially aligned with the second turbine blade.

Abstract: A gas-turbine axial compressor has a casing 3 and a rotatable shaft 6, which form an annular duct, in which at least one stator 2 and one rotor 4 are arranged. A shroud 21 is arranged at a free end of the stator 2 vanes, which extends over part of the axial length of the stator 2. A rotor platform 23 of at least one rotor 4 extends axially beneath a further part of the axial length of the stator 2, at which no shroud 21 is arranged.

Abstract: In one embodiment, a catenary seal structure may be mechanically supported only by two adjacent rotor wheels to isolate a volume between the two adjacent rotor wheels. The catenary seal structure may be annularly disposed about the rotational axis of the adjacent rotor wheels.

Abstract: A device for reducing secondary airflow in a gas turbine is disclosed. The device includes an inter-stage sealing member located between a plurality of first turbine buckets attached to a first rotor disk, and a plurality of second turbine buckets attached to a second rotor disk. The first rotor disk and the second rotor disk are rotatable about a central axis. The inter-stage sealing member is configured to be attached in a fixed position relative to the first rotor disk and the second rotor disk, and to contact the plurality of first buckets and the plurality of second buckets in a sealing engagement.

Abstract: An air seal assembly for a gas turbine engine includes a first cover plate with a radially extending knife edge seal defined about and axis of rotation. The first cover plate is mountable to a first rotor disk for rotation therewith, the first radially extending knife edge seal interfaces with a vane structure. A second cover plate with a second radially extending knife edge seal defined about the axis of rotation, the second cover plate mountable to the second rotor disk for rotation therewith. The second radially extending knife edge seal interfaces with the vane structure.

Abstract: There is described a gas turbine with a channel wall that limits a flow, said channel wall comprising first and second wall sections, one of the wall sections being part of a combustion chamber of the gas turbine and the other wall section being part of an annular hot gas channel of a turbine unit, and the two adjacent wall sections facing each other with a gap there between. In order to reduce the leakage quantities of blocking gas that protect the gap from the entry of hot gas, it is proposed that blocking element projects from the side of the first wall section into the gap and can be displaced transversely thereto, said blocking element comprising a piston face disposed in the first wall section and a sealing structure lying opposite the piston face, it being possible to press said sealing means against an end sealing section of the second wall section by means of a pressure medium that acts on the piston face.

Abstract: A method of repairing a gas turbine knife edge seal assembly in which assembly is included a used honeycomb ring seal and a knife edge. The method includes removing the used honeycomb ring seal, removing material from the knife edge to a first knife edge radial height, installing a replacement honeycomb ring seal with a stock radial thickness, and removing material from the replacement honeycomb ring seal to a finished radial thickness for sealable engagement with the knife edge. The finished radial thickness of the replacement honeycomb ring seal establishes a first radial gap between an inner annular surface of the replacement honeycomb ring seal and the knife edge before the knife edge seal assembly begins rotating in the gas turbine.

Abstract: A fluid-flow machine has a main flow path, in which at least one row of blades (1) is arranged, and a shroud (2) embedded in a cavity (3) of a component, with the component and the blades being in rotary movement relative to each other. The shroud (2), as referred to an axial direction of the fluid-flow machine is arranged between two reference points (A) and (B) at the periphery of the main flow path of the fluid-flow machine and projects beyond a rectilinear connection of the reference points (A) and (B) into the flow path in at least one location of the circumference and in at least one of the areas of the leading or trailing edge of the appertaining blade row.

Abstract: The present invention relates to a pressure seal (7) of a turbomachine stator, said seal (7) comprising a first abradable surface opposite a rotor portion of the turbomachine and a second surface in contact with an inner shroud (3) of the stator, said seal (7) comprising a plurality of component units (10), each component unit (10) having, on its first abradable surface, a circumferential step (9) creating an obstacle in the circumferential direction of the inner shroud (3).

Abstract: A gas turbine engine comprising a rotor and a stator which define first, second and third cavities; the rotor and stator define a seal therebetween and which is located for sealing between the second and third cavities, the rotor comprises an aperture through which a gas flow passes from the first cavity to the second cavity characterized in that the seal comprises a deflector that extends axially over at least a portion of the aperture to deflect at least a part of the gas towards the rotor.

Abstract: A turbomachine (1), in particular a gas turbine, has a stator (2) and a rotor (3) and with at least one axial sealing device (6) arranged between them. On the stator side, the sealing device (6) has a radially stepped sealing contour (8) with regions (9, 10) projecting and retracting in the direction of the rotor (3), on the rotor side a plurality of sealing fins (11) projecting in the direction of the stator (2) being arranged, which engage in each case into adjacent retracting regions (10) of the stator-side sealing contour (8). On the rotor side, at least one additional fin (12) projecting in the direction of the stator (2) is provided, which is positioned between two adjacent sealing fins (11) arranged on the rotor side and which lies opposite a projecting region of the stator-side sealing contour (8).

Abstract: An exemplary embodiment of the invention is directed to a turbine engine having a rotatable turbine rotor assembly, a stationary nozzle assembly disposed adjacent thereto and a wheel space defined therebetween. The wheel space receives cooling air therein and includes a sealing feature located on the first, rotatable turbine rotor assembly and extending axially into the wheel space and a sealing land assembly having a sealing land associated with a moveable member installed in an opening in the second, stationary nozzle assembly. A biasing member constructed of shape memory alloy is associated with the moveable member and operates to bias the moveable member, and associated sealing land, axially into the wheel space towards the sealing feature as the turbine engine transitions from a cold state to a hot state to reduce the release of cooling air from within the wheel space.

Abstract: A stage of a turbine or compressor for a turbomachine including at least a vane disk surrounded with a sectorized ring supported by a casing and including a circumferential flange radially clamped on a rail of the casing by substantially C-shaped cross-section locks which are axially engaged on the casing rail and the flange of the ring is disclosed. The inner and outer walls of each lock includes radial bearing surfaces on the flange of the ring and the casing rail, respectively. The bearing surfaces only extend on a part of the width of the lock.

Abstract: A turbine engine blade outer air seal segment has a body having a base portion. The base portion has a transversely concave ID face, a forward end, an aft end, and first and second circumferential edges. The body has at least one mounting hook. At least one cover plate is secured to the body to define at least one cavity. The cover plate has a plurality of feed holes. A plurality of outlet holes extend through the base portion to the ID face. At least one of the base portion and cover plate comprises a protruding portion protruding into the cavity to form a partial restriction separating circumferentially and fore-aft offset cavity portions.

Abstract: A knife edge seal assembly includes at least one knife edge and at least one honeycomb ring seal. The at least one knife edge has material removed to a first knife edge radial height and the least one honeycomb ring seal has material removed to a finished radial thickness for sealable engagement with the at least one knife edge. The finished radial thickness of the replacement honeycomb ring seal establishes a first radial gap between an inner annular surface of the replacement honeycomb ring seal and the knife edge before the knife edge seal assembly begins rotating in the gas turbine.

Abstract: A method for assembling a rotating machine includes providing a rotating element including a plurality of rotor wheels. The method also includes positioning the rotating element such that at least a portion of a stationary portion extends at least partially about the rotating element. The method further includes assembling an interstage seal mechanism including coupling at least a portion of a first hook device to the rotating element, and also including coupling at least a portion of a second hook device to the first hook device. The first hook device and the second hook device are radially inboard of at least a portion of the stationary portion.

Abstract: The present application provides an angel wing seal for a turbine bucket. The angel wing seal may include a first wing with a sinusoidally-shaped outer edge and a number of wing teeth positioned thereon.

Abstract: A turbine seal member for use in a gas turbine engine includes a turbine seal substrate having a gas-path side and a ceramic layer disposed on the gas-path side that includes a plurality of mechanical indentations.

Abstract: A seal assembly limits gas leakage from a hot gas path to one or more disc cavities in a gas turbine engine. The seal assembly includes a seal apparatus associated with a blade structure including a row of airfoils. The seal apparatus includes an annular inner shroud associated with adjacent stationary components, a wing member, and a first wing flange. The wing member extends axially from the blade structure toward the annular inner shroud. The first wing flange extends radially outwardly from the wing member toward the annular inner shroud. A plurality of regions including one or more recirculation zones are defined between the blade structure and the annular inner shroud that recirculate working gas therein back toward the hot gas path.

Abstract: Disclosed is an annular flow duct for a turbomachine, the flow duct arranged concentrically about a machine axis running in the axial direction and defined by a boundary wall of circular cross section for directing a main flow, wherein the boundary wall has a plurality of return flow passages distributed over its circumference and through which in each case a partial flow is detached from the main flow at a bleed position and returned to the main flow at a feed position situated upstream of the bleed position, and having aerofoils, arranged radically in the flow duct, of a blade ring, the aerofoil tips of which lie opposite the boundary wall, with a gap formed in each case, wherein the aerofoils are movable in a predetermined rotation direction along the circumference of the boundary wall. Furthermore, a compressor which is insensitive to pumping and flow separations, as viewed in rotation direction, the bleed position of each return flow passage lies upstream of the corresponding feed position.

Abstract: A stator-rotor assembly which includes at least one interface region between the stator and rotor is described. At least one stator or rotor surface in the interface region includes a pattern of concavities. The concavities restrict gas flow through a gap between the stator and the rotor. Various turbomachines which can contain such a stator-rotor assembly are also described. The disclosure also discusses methods to restrict gas flow through gaps in a stator-rotor assembly, utilizing the concavities.