Abstract: A rim seal assembly includes an annular seal element circumscribing an engine centerline or axis and mounted on an annular platform including a radially inwardly extending annular platform flange disposed between and connected to forward and aft flanges at distal ends of forward and aft annular elements respectively. Annular forward and aft outer rim cavities radially disposed between the forward and aft annular elements and the platform are axially separated by the platform flange. Cooling slots extend radially across axially facing forward and aft surfaces of the forward and aft flanges. The platform flange and the forward and aft flanges are bolted together. The annular seal element may include seal teeth in sealing relationship with an annular seal land such as in a labyrinth seal. The rim seal assembly may be incorporated in a low pressure turbine use a compressor as a source of cooling air.

Abstract: A seal plate assembly for use with a rotor wheel assembly of a turbine engine. The seal plate assembly may include a seal plate positioned about a rim of a rotor wheel. The seal plate may include a number of seal plate segments and a number of locking pins extending through the seal plate segments and into the rim of the rotor wheel.

Abstract: The present invention relates to a sealing arrangement between a stator assembly and a rotor 12 of a gas turbine engine, the rotor 12 being rotatable about an engine axis, the stator assembly comprising vanes 22 provided at their radially inner ends with a vane support structure 28 which comprises an annular platform 30 supporting the vanes and a radial flange 32 which extends radially inwardly from the platform, the sealing arrangement comprising an array of leaves 36 which extend inwardly from the radially inner periphery of the flange 32, towards a circumferential sealing surface 13 of the rotor 12.

Abstract: Apparatus for sealing and purging a bearing region includes a bearing housing having at least one hole that allows a flow of air to flow through the at least one hole, and at least one bearing seal that allows the flow of air to flow through the at least one bearing seal, thereby creating a pressure difference across the bearing seal. The apparatus also includes an air directional device having at least a portion of the flow of air to flow adjacent to the air directional device, and a pair of components having a gap between the pair of components that allows the at least a portion of the flow of air to flow through the gap, thereby purging the gap.

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 inverted turbulators. The inverted turbulators 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 inverted turbulators.

Abstract: The hybrid seal assembly includes a seal for use with rotating surfaces which may suffer significant deflections such as for sealing between flanges and mating surfaces typical of annular seal face surface of a diffuser section of a tip turbine engine.

Abstract: A method of attaching a labyrinth seal support to a turbine engine shaft is disclosed. The method includes inserting axial bolts to secure the labyrinth seal support, the shaft and a device for driving the turbine module with the shaft, and screwing tightening screws in extracting holes, for extracting the drive device with respect to the shaft, provided on the shaft to tighten the labyrinth seal support onto the shaft.

Abstract: A rotor and a stator assembly, wherein the rotor (30) comprises at least one stage of rotor blades (32) and the stator (34) comprises at least one stage of turbine stator vanes (36). The rotor blades (32) have aerofoils and platforms (42) and the turbine stator vanes (36) having aerofoils and platforms (48). A seal (43,45) is defined between the rotor blade (32) platforms (42) and the stator vane (36) platforms (48) wherein a portion (42B) of the rotor blade platforms (42) and/or a portion (48B) of the stator vane (36) platforms (48) comprise a shape memory alloy member or a bimetallic member. The shape memory alloy member, or bimetallic member, controls the flow of cooling air through the seals (43,45) allowing a greater cooling flow at higher temperatures than at lower temperatures.

Abstract: A rotor blade assembly includes a slot for mounting a rotor blade that extends partially through a rotor disk and a continuous uninterrupted hoop seal that extends about the circumference of the rotor disk. The rotor disk includes an integrally formed seal that includes at least one knife edge disposed concentrically about the rotor disk. The slot receives a root portion of a rotor blade but does not extend through the seal. As the seal remains an uninterrupted full hoop about the circumference of the rotor disk, additional rigidity and strength are realized. The increased rigidity and strength provided by the full hoop structure provides for the reduction in material and physical dimensions without compromising desired performance of the seal.

Abstract: A compressor of a gas turbine engine comprises blades 2, 4 provided on blade platforms 8, 12 of a rotor. Stator vanes 20 lie between the blades 2, 4 and are connected to a vane support structure 22. Sealing rings 38 are secured to the vane support structure 22 to restrict air flow into stator wells 36 on each side of the vane support structure 22. The sealing rings 38 are made from a flexible material so that they deflect on installation of the vane support structure 22 in a direction radially inwards between the adjacent blade platforms 8, 12.

Abstract: It is known with regard to particularly cavities below high pressure turbine discs that mixing of hot gas leakage flows through an inner seal with cooling flows can diminish the effectiveness of that cooling flow when presented to other parts for cooling. By providing a path within a wall which is particularly shaped in portions it is possible to provide entrainment of a hot leakage gas flow away from entry into the cavity. Thus, the cooling flow retains a higher cooling effect and maintains its swirling nature in comparison with prior arrangements where mixing with the hot leakage flow occurred.

Abstract: The invention relates to a segmented inner ring for holding guide blades. According to the invention, a lateral wall opposing the front side of the inner ring and pertaining to a shaft shoulder formed on the rotor shaft extends radially, and respectively one half of a labyrinth seal is formed on the front side of the inner ring and on the shaft shoulder. The aim of the invention is to apply an arrangement of stacked labyrinth seals, known from aeroplane turbines, to a stationary gas turbine having a separation plane. To this end, a method is used to mount an inner ring of a gas turbine. The invention also relates to a stationary gas turbine comprising a segmented inner ring.

Abstract: A seal assembly (52, 54) for a gas turbine engine (10). The seal assembly (52, 54) includes a first seal member (40) including an abradable structure having at least first and second seal portions (58, 64) radially spaced from each other, and a second seal member (44) including first and second seal surfaces (56, 62) located in facing relationship to the first and second seal portions (58, 64), respectively. The first seal member (40) is formed of a honeycomb structure, and the first seal portion (58) of the first seal member (40) is a thin fin-like structure having a width in the axial direction of the turbine engine (10) that is less than the width of the second seal portion (64) in the axial direction.

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: 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: 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: 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: 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 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: An impulse turbine is provided. By providing brush seals on the free ends of the guide vanes, which form an acute angle with the radial line in relation to the radial line in the direction of rotation of the rotor, leakage losses may be reduced and the rotor friction between the guide surfaces may be reduced.

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 sealing element for a gas turbine, which is provided with at least one rotor ring and at least a plurality of rotor blades radially arranged about the rotor ring and having an end portion fixed to the rotor ring, is provided with a wall, which is adapted to be coupled to the rotor ring and to at least one rotor blade, has an external face adapted to be arranged in contact with a hot working fluid in the gas turbine and an internal face adapted to be arranged in contact with a cooling fluid of the gas turbine, and a gap, which is adapted to be travelled through by the cooling fluid.

Abstract: A rotary machine includes a rotor, a stationary machine casing extending around the rotor, and a bling assembly extending between the casing and the rotor. The machine also includes at least one rotor tip seal assembly and at least one shaft seal assembly. The seal assemblies have a groove configured to receive at least one seal ring band. A method of assembling a rotary machine is also provided. The method includes fabricating the bling assembly by providing two identical members comprising a mating surface and having a semi-circular profile. The method also includes coupling the two members together at their mating surfaces such that a circular ring is formed and such that the mating surfaces define a horizontal joint. The method further includes machining concentric, circular and annular radially inner and outer and airfoil portions within predetermined radial portions of the bling assembly.

Abstract: A compressor of a gas turbine engine comprises blades 2, 4 provided on blade platforms 8, 12 of a rotor. Stator vanes 20 lie between the blades 2, 4 and are connected to a vane support structure 22. Sealing rings 38 are secured to the vane support structure 22 to restrict air flow into stator wells 36 on each side of the vane support structure 22. The sealing rings 38 are made from a flexible material so that they deflect on installation of the vane support structure 22 in a direction radially inwards between the adjacent blade platforms 8, 12.

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 stator assembly having a windage cover for structure adjacent to a cavity bounded by rotating elements in a rotary machine is disclosed. Various construction details are developed for damping vibrations in the windage cover as the windage cover bounds a cavity having swirling high velocity gases that are capable of transmitting acoustic energy and kinetic energy to adjacent structure. In one detailed embodiment, the windage cover in the uninstalled condition has diverging arms that are resiliently compressed during installation to exert a frictional force on the adjacent structure.

Abstract: On a two or multi-stage turbine, expansion rings are provided in all stages at the sides of the rotors for passive, continuous running gap control whose thermal expansion and contraction behavior corresponds to that of the rotors and which are connected to radially moveable upstream and downstream stator vanes (5, 7). The downstream stator vanes are assembled to the upstream stator vanes via a bridge (16; 17, 18) which is axially and circumferentially fixed and located flexibly in the radial direction on the outer casing (10) of the turbine. The stator vanes (6) arranged between the rotor disks are integrally connected to the bridge or, as separate components, held axially and circumferentially on the bridge to take up rolling and tilting moments.

Abstract: A sealing interface for a component in a turbine machine, wherein a plug is mechanically affixed within an entrance passage of an aperture. The plug is secured in place within the entrance passage through deformation of an inner end of the plug to cause the plug to mechanically engage a groove formed adjacent the entrance passage.

Abstract: A stator vane segment for a gas turbine engine includes at least one airfoil joined to an outer shroud and an inner platform. A sealing element having a first platform radially inward of the inner platform and an abradable material covering at least a portion of the first platform is integrally joined to the inner platform.

Abstract: A pump/flow guide is positioned in a gas turbine engine and attached to a transition duct to face a turbine disk. The pump/flow guide provides a smooth path for guiding purge flow to resist the ingestion of hot gas. One leg of the pump/flow guide extends radially outwardly and contacts the radially inner section of the transition duct that forms a “fish mouth.” Spaced tabs on the pump/flow guide receive bolts to secure the pump/flow guide to the transition duct. The pump/flow guide thus provides a relatively smooth flow passage for purge gas flow.

Abstract: An air circulation method in a turbomachine compressor includes minimizing an inner flowpath between two adjacent blades by providing labyrinths on an inner casing between the two adjacent blades, and sucking in air from the minimized inner flowpath through a radially inner orifice of a vane faced by the labyrinths. The method further includes directing the sucked in air through the vane, bleeding the air from the vane through a radially outer opening of the vane and through an orifice of an outer casing. The air is collected from the vane into a manifold outside the compression stage. A compressor arrangement for circulating air includes a suction device for sucking in air present in the inner flowpath between labyrinths and the vane, and for sending the air into the vane of the stator. The arrangement also includes a device for bleeding air in the vane of the stator and for sending the air outside the compressor.

Abstract: A turbine stage includes a stator nozzle having a row of vanes mounted between inner and outer bands. The inner band terminates in an ovate ledge converging aft from the vanes with radially outer and inner convex surfaces joined at a convex apex. The ovate ledge reduces aerodynamic losses at the rotary seal with a row of following turbine rotor blades.

Abstract: The turbine includes a plurality of successive stages each having a rotatable disk and blades carried thereby. A pair of adjacent rotatable disks define an annular gap therebetween and have respective opposing sealing band receiving slots aligned with the annular gap. At least one of the rotatable disks has at least one notch therein coupled to the respective sealing band receiving slot, and a sealing band is in the opposing sealing band receiving slots to seal the corresponding annular gap. The sealing band includes at least one sealing strip and at least one movable locking tang carried thereby and extending into the at least one notch to define an anti-rotational feature for the sealing band.

Abstract: A seal assembly (52, 54) for a gas turbine engine (10). The seal assembly (52, 54) includes a first seal member (40) including an abradable structure having at least first and second seal portions (58, 64) radially spaced from each other, and a second seal member (44) including first and second seal surfaces (56, 62) located in facing relationship to the first and second seal portions (58, 64), respectively. The first seal member (40) is formed of a honeycomb structure, and the first seal portion (58) of the first seal member (40) is a thin fin-like structure having a width in the axial direction of the turbine engine (10) that is less than the width of the second seal portion (64) in the axial direction.

Abstract: A gas turbine engine having a rotor with blades and a stationary vane, a platform seal is formed between the blade and vane for inhibiting ingestion of hot gas from a hot gas flow through the turbine into turbine wheel spaces, the platform seal including axial extending platforms on the blade and vane, and radial extending fingers extending from the platforms and forming restrictions between the fingers and the platforms, and a buffer cavity formed between the restrictions, where the fingers are so arranged in a generally radial direction that the vane can be removed from the turbine engine in a radial direction without having to remove the blades first. In additional embodiments, the platform seal assembly can have two or three buffer cavities formed between additional restrictions.

Abstract: A rotor for a rotating machine, in particular a steam turbine includes at least one subregion composed of a metal structure having a reduced density and including a multiplicity of finely distributed cavities.

Abstract: A seal plate system adapted to fit between axially adjacent turbine blades and turbine vanes is disclosed. The seal plate system may be formed from a plate adapted to position a turbine blade in a rotor assembly and lock the turbine blade into the rotor assembly. The plate may also function to seal a cooling system in the turbine blade. In at least one embodiment, the plate may be configured to accomplish both of these tasks together once installed into a mini disc of the rotor assembly.

Abstract: A seal assembly includes a body having two circumferential sides, a leading end, and a trailing end. At least one of the circumferential sides includes a first channel sidewall, a second channel sidewall, and a channel bottom wall that together define a seal channel for receiving a seal. The seal channel includes a slot that cooperates with a tab on the seal to facilitate securing the seal within the seal channel.

Abstract: A gas turbine engine is provided with turbine sealing structures including knife edge seals which extend at an angle relative to an axial center line of the engine. Each knife edge seal is associated with a control pocket defined between a radially inner surface and a spaced radially outer surface. The control pockets and their associated knife edge seals create a difficult flow path to prevent leakage into radially inner portions of the turbine section.

Abstract: Gas turbine engine systems involving hydrostatic face seals with back-up seals are provided. In this regard, a representative seal assembly for a gas turbine engine includes: a hydrostatic seal having a seal face and a seal runner; and a back-up seal; wherein, in a normal mode of operation of the hydrostatic seal, interaction of the seal face and the seal runner maintains a pressure differential within the gas turbine engine and, in a failure mode of operation of the hydrostatic seal, the back-up seal maintains a pressure differential within the gas turbine engine.

Abstract: An inner air seal carrier for use in a gas turbine engine having an inlet guide vane surge retainer comprises a body, a stationary sealing element and an outcropping. The body secures around an inlet guide vane inner diameter shroud. The stationary sealing element is disposed on a radially inward face of the body for engaging with a rotatable sealing element of a compressor rotor. The outcropping is positioned on the radially inward face of the body forward of the stationary sealing element for engaging with the surge retainer.

Abstract: A seal arrangement 30 presents a barrier in the form of a combined seal and anti-fret liner 36 across a shroud ring cavity 32 in order to inhibit air leakage in the platform gap between platforms 33. It will be understood that these gaps are provided for thermal expansion and vibration of vane assemblies but create a leakage path between passage 35 and cavity 32 which can increase aerodynamic losses and therefore reduce engine compressor efficiency. By provision of a combined seal and anti-fret liner 36 such air losses are reduced whilst combining the function of an air leakage seal and an anti-fret liner to prevent abrasive degradation of the mountings 37, 39 and platform section 33 ensures that there is a reduced component count as well as cost, weight and assembly time.

Abstract: A gap seal (9) radially seals a gap (8) which extends axially and radially between two blades (1, 2) of a turbomachine which are adjacent in the circumferential direction (3). The two blades (1, 2) have, in each case, an axially extending longitudinal slot (10, 11), which is open towards the gap (8), on its respective blade root (6, 7). A band-form or strip-form sealing element (12) engages with its longitudinal sides (13, 14) in the two longitudinal slots (10, 11) and bridges the gap (8). The one blade (1) has a projection (15) on its blade root (6), which projects from the blade root (6) in the circumferential direction (3) and extends in the circumferential direction (3) and radially, at least in the region of the respective longitudinal slot (10), and bridges an axial longitudinal end of the gap (8) in the process. The other blade (2) has a step-shaped recess (16) on its blade root (7), complementary to the projection (15) of the one blade (1) and in which the projection engages.

Abstract: The invention relates to a segmented inner ring for holding guide blades. According to the invention, a lateral wall opposing the front side of the inner ring and pertaining to a shaft shoulder formed on the rotor shaft extends radially, and respectively one half of a labyrinth seal is formed on the front side of the inner ring and on the shaft shoulder. The aim of the invention is to apply an arrangement of stacked labyrinth seals, known from airplane turbines, to a stationary gas turbine having a separation plane. To this end, a method is used to mount an inner ring of a gas turbine. The invention also relates to a stationary gas turbine comprising a segmented inner ring.

Abstract: A turbine stage includes a stator nozzle cooperating with a row of turbine blades. The nozzle includes a row of vanes mounted between inner and outer bands, with the inner band adjoining the blades at inner platforms thereof at a rotary seal. The inner band includes a stator bullnose at the trailing edge thereof cooperating with the rotary seal for reducing aerodynamic losses thereat.

Abstract: A gas turbine engine is provided with turbine sealing structures including knife edge seals which extend at an angle relative to an axial center line of the engine. Each knife edge seal is associated with a concave pocket defined between a radially inner surface and a spaced radially outer surface. The concave pockets and their associated knife edge seals create a pair of vortices which prevent leakage into radially inner portions of the turbine section.

Abstract: A rotor blade assembly includes a slot for mounting a rotor blade that extends partially through a rotor disk and a continuous uninterrupted hoop seal that extends about the circumference of the rotor disk. The rotor disk includes an integrally formed seal that includes at least one knife edge disposed concentrically about the rotor disk. The slot receives a root portion of a rotor blade but does not extend through the seal. As the seal remains an uninterrupted full hoop about the circumference of the rotor disk, additional rigidity and strength are realized. The increased rigidity and strength provided by the full hoop structure provides for the reduction in material and physical dimensions without compromising desired performance of the seal.

Abstract: This invention relates to a seal plate system adapted to fit between axially adjacent turbine blades and turbine vanes to seal a rim cavity. The seal plate system may be formed from a seal plate supported by a plurality of extended disk lugs extending from a disk. The seal plate system may also include one or more extended disk lug receiving cavities adapted to receive protrusions extending from the disc. One or more ribs may extend from the seal plate for increasing the structural integrity of the seal plate.

Abstract: A turbine rotor for a turbine engine includes multiple rotor disks having rotor blades mounted about the circumference of each of the rotor disks. A fluid seal extends about the circumference of each rotor disk in close proximity to a stationary component of the rotor to separate the space between the rotor blades and a stationary component into separate cavities. The fluid seal includes a seal land on the rotor disk and a brush seal extending from the stationary component. A plurality of disk land segments and a plurality of blade land segments form the seal land. The disk seal segments are located on the rotor disk between the rotor blades. The blade land segments are located on the rotor blades. After the rotor blades are assembled, the blade land segments and the disk land segments fit together to form a segmented ring-like seal land around the circumference of the rotor disk.