Abstract: A turbomachine compressor rotor including at least two blade-carrying disks mounted on a common axis and connected together by a wall forming a substantially cylindrical surface of revolution, and a centripetal air bleed mechanism including air passages passing through the wall and radial fins for deflecting a stream of air leaving the passages in the wall, the fins being carried by one of the disks and being substantially in radial alignment with the passages through the wall.

Abstract: A seal assembly (116) mounted within a rotationally fixed static outer support structure (14) for engagement with an annular seal face surface (172) located upon a fan-turbine rotor assembly. The seal assembly provides minimal leakage of core airflow when the airflow is turned and diffused by the diffuser section (74) to increase the engine operating efficiency.

Abstract: A seal assembly disposed between a rotor and a stator especially of a turbo-machinery is provided with a close clearance gap there between. The clearance gap is bordered by a rotor surface at the side of the rotor and by a stator surface at the side of the stator. The rotor surface or the stator surface or both are provided with openings which open into at least one chamber arranged at the side of the openings turned away from the clearance gap.

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: A turbine engine ring seal for sealing gaps between turbine engine outer seal segments and turbine blade tips. The turbine engine ring segment may have an inner radial surface that defines a portion of a gap gas flow path where the inner radial surface may be formed of an abradable ceramic coating and includes a plurality of gas flow protrusions that are oriented transverse to the gap gas flow path. The gas flow protrusions may induce vortices in the gas flow in the gap gas flow path. Additionally, the gas flow protrusions may be series of peaks and depressions between two adjacent peaks, where the depressions have an approximate semicircular shape. The distance between two adjacent peaks may be equal or greater than a width of the depression and the height of a single peak may be six percent or greater than the distance between two adjacent peaks.

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: The invention provides a thermomechanical turbomachine part that is circularly symmetrical about a longitudinal axis, in particular a rotor, the part including at least one annular wiper for a sealing labyrinth. In characteristic manner, the wiper presents height in the radial direction that varies around its circumference by forming a plurality of projecting portions.

Abstract: A rotor mechanism of a centrifugal compressor connectable to a power output device and configured to block leakage of lubricants and to recycle lubricants to an oil sump is provided, which includes a rotary shaft with a rolling bearing and a screw nut, a bearing housing for receiving the rolling bearing, and a bearing plate connected to the bearing housing. The screw nut has a surface provided with an oil throw seal, and the bearing housing. As the bearing plate is provided with an oil channel in contact with the oil-returning hole and either one of the bearing housing and the bearing plate is provided with a first labyrinth seal in contact with the screw nut, lubricants leaking out of the rolling bearing is allowed to go to the oil sump via the oil channel and oil-returning hole. The oil slinger thus blocks the lubricants from leaking out of the first labyrinth seal.

Abstract: Disclosed herein is a turbomachinery including a labyrinth seal easy to fabricate and adapted to prevent unstable vibration of a rotor. A turbomachinery according to this invention includes a labyrinth seal 2 formed by dividing the seal into segments 7 in a circumferential direction and combining the segments 7 whose comb teeth 5 differ in height. The labyrinth seal 2 preferably includes two kinds of segments 7 whose comb teeth 5 differ in height for each kind of segment, and the segments 7a of the greater comb tooth height are arranged in a certain reference direction on a cross section right-angled to a rotational axis, and in a direction opposite to the reference direction, and the segments 7b of the smaller comb tool height are arranged in a direction right-angled to the reference direction, and in a direction opposite to that right-angled to the reference direction.

Abstract: Methods and apparatus for a packing ring segment assembly for sealing a turbine gas path are provided. The assembly includes a sealing portion including radially inwardly directed teeth, and a dovetail mounting portion including a hollowed out portion including a biasing member internal to the packing ring segment. The biasing member is coupled to the packing ring segment and configured to engage a pair of annular shoulders of a dovetail groove in the turbine such that the packing ring segment is biasing radially outward at relatively low turbine loads and low working pressure, and at relatively high turbine loads and high working pressure a working fluid overcomes the radial biasing member forces and urge the packing ring segment radially inward.

Abstract: A method for repairing damaged fins in a labyrinth seal without the use of welding or metallurgical bonding by removing a section of the shaft containing the damaged fins, and providing a sleeve with replacement fins.

Abstract: A pulsed discharge is generated between tip ends of a rotating member such as a blade and a discharge electrode including a hard material such as cBN in dielectric liquid or gas by a power supply for discharge to melt the discharge electrode, and a part of the discharge electrode is attached to the tip end of the rotating member to form an abrasive coating film including the hard materials such as cBN.

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 packing ring assembly is provided and includes an arcuate carrier ring segment, an arcuate packing ring segment movable relative to the carrier ring segment, an attachment component disposed in association with the packing ring segment and in association with the carrier ring segment, and an actuator component configured to maintain the packing ring segment in a first position and to allow a movement of the packing ring segment to a second position when the packing ring assembly is exposed to a pressure condition.

Abstract: A vane for a gas turbine engine includes an aerofoil part and a shroud that forms a sealing part at one end of the aerofoil part. The sealing part defines a cavity and an opening to the cavity. The sealing part may include a pair of opposed side walls extending from a radially outer wall of the shroud to a pair of radially inner walls of the shroud to define a cavity. The pair of radially inner walls may be substantially parallel to the radially outer wall and may extend in a substantially circumferential direction to define a cavity opening.

Abstract: A rotating machine (10), in particular a compressor or turbine, has a rotor shaft (11) which is rotatable about an axis (20) and is provided with at least one slot (14). The mechanical stresses in the region of the slot are reduced in that the at least one slot (14) has a slot bottom (17) with an elliptic cross-sectional contour (24).

Abstract: An abradable member, such as gas turbine engine seal for rotary components, is restrained against axial movement relative to an adjacent component via a lip retention engagement.

Abstract: A labyrinth seal for a bypass discharge-type motor-fan unit includes a motor unit from which extends a rotating shaft that passes through an aperture in an end plate to drive a fan unit. Disposed within the aperture is a labyrinth seal configured to prevent liquid entrained in a working airflow from entering the motor unit. The rotating shaft also includes a vacuum bore, such that the negative pressure generated by the operation of the fan unit draws any liquid out of an evacuation zone maintained between the motor unit and the end plate. As such, the components of the motor unit are protected from being exposed to the liquid, thereby extending the operating life motor-fan unit.

Abstract: In the method, a support is provided that presents a base for an annular wiper, and a wiper is built up that presents height that varies around its circumference, by forming a plurality of projecting portions by performing following successive steps: a laser source connected to an optical head focused on a point on the top surface of the base is activated, as is a source of powder connected to the spray nozzle, thereby forming a localized melt at said point, into which the powder is injected, thus forming a localized region of extra thickness; and the optical head and the nozzle are aimed at another point adjacent to said region of extra thickness and the method returns to the preceding step until the layer of material is completed in the angular sector in question.

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: A turbine engine includes a first turbine structure that supports a seal. The seal is movable within a recess of the first turbine structure. The seal is arranged in close proximity to a seal land of a second turbine structure for preventing a fluid from leaking past the seal and seal land. A thermal expansion member interconnects the first turbine structure and the seal. The thermal expansion member expands in response to an increase in temperature to move the seal toward the seal land preventing the typical enlarged gap between the seal and seal land resulting from thermal growth. In one example, the thermal expansion member, which is arranged at each opposing end of a seal segment, is a bimetallic coil spring supported on the first turbine structure by a cage. A free end of the coil spring is secured to the seal at the opposing end portions.

Abstract: A sealing structure includes a labyrinth seal provided between a housing and an outer periphery of an end portion of a turbine shaft on a turbine wheel side, and a piston ring seal disposed on a side toward a center of the turbine shaft in an axial direction thereof from a position on the turbine shaft at which the labyrinth seal is provided, and a ring groove is formed along a circumferential direction of the housing in a middle portion of the labyrinth seal to reverse a flow direction of exhaust gas entering from the exhaust passage.

Abstract: A turbomachine including a wheel with hollow blades that are cooled internally by a forced flow of cooling air. According to an important characteristic of the invention, a hook-annulus (70) is interposed between the outer edge of the labyrinth-disk (52) and the rotor disk (36), the hook-annulus including a retaining shoulder (72) constituting an outer annular hook in which the outer edge of the labyrinth-disk is engaged, and a bearing shoulder (74) constituting an inner hook engaged in a groove around the circumference of the rotor disk.

Abstract: A system and method for extending seal life and for reducing leakage flow in a turbine engine is directed to an interface defined between a stationary component and a rotating component. A seal, which can be a flexible seal such as a brush seal, can be operatively attached to the stationary component. The rotating component has a first region at a first radius relative to the axis of rotation that transitions into a second region at a second, larger radius relative to the axis of rotation. In one embodiment, the rotating component can be selectively axially moved between a first position and a second position. In the first position, the seal is disposed over the first region so as to define a clearance. In the second position, the seal is disposed over the second region so as to decrease the size of the clearance.

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 method of assembling a seal assembly for a turbine engine is provided, wherein the method includes providing a seal ring having an arcuate inner ring portion, an arcuate outer ring portion, and a neck portion extending therebetween, and forming at least one recess within at least one of the outer ring portion and the neck portion. The method also includes extending a biasing mechanism across the seal ring such that the biasing mechanism is positively retained within the at least one recess.

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 assembly for a gas turbine engine compressor includes a rotor disc for mounting a plurality of radially extending rotor blades for rotation thereon, an annulus filler for securing to the disc to provide a radially inner airflow surface for air passing through the rotor assembly and a securing arrangement for securing the annulus filler to the rotor disc, the annulus filler being configured such that it is biased into a condition in which the securing arrangement mechanically retains the annulus filler on the disc, but is movable against its bias to allow the annulus filler to be removed from the disc.

Abstract: This invention relates to a turbine stator (5) for an aircraft turbine engine, comprising: a stator wall (46) supporting a high pressure air injector (50) fixed on the downstream side; an annular abradable element support (54) the lower radial end (54a) of which is fixed to said high pressure air injector and its upper radial end (54b) is fixed to said stator wall, such that they jointly delimit a cavity (58) under pressure. According to the invention, it also comprises at least one vibration damping device (74) located in said cavity (58) under pressure, said vibration damping device being in contact with each of the two elements taken among said stator wall (46) and said annular support (54).

Abstract: Labyrinth seals and methods of manufacturing the seals are provided. In an embodiment, the seal includes a seal base and a first annular tooth. The seal base has an outer peripheral surface. The first annular tooth extends radially from the seal base outer peripheral surface. The first annular tooth includes a forward wall, an aft wall, and a tip wall, and the forward wall and the aft wall each include bottom sections that contact the seal base and angle upwardly toward or substantially parallel to each other. The forward wall has a notch formed therein defined by a shelf wall and a side wall, the shelf wall and side wall join the forward wall to the tip wall, and the aft wall is formed without a notch. The side wall and the aft wall extend substantially parallel to each other.

Abstract: An abradable seal is provided to improve turbine performance by physically reducing the clearance between a flange portion of the nozzle and an opposed angel wing/seal plate member of the bucket. The provision of an abradable seal also mitigates angel wing/seal plate tooth or fin wear by providing for abradable contact without metal to metal hard rub.

Abstract: Gas turbine engine systems involving hydrostatic face seals with anti-fouling provisioning are provided. In this regard, a representative turbine assembly for a gas turbine engine comprises: a turbine having a hydrostatic seal; the hydrostatic seal having a seal face, a seal runner, a carrier, and a biasing member; the seal face and the seal runner defining a high-pressure side and a lower-pressure side of the seal; the carrier being operative to position the seal face relative to the seal runner; and the biasing member being located on the lower-pressure side of the seal and being operative to bias the carrier such that interaction of the biasing member and gas pressure across the seal causes the carrier to position the seal face relative to the seal runner.

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 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 vanes extend between the rotor blades of axially adjacent disks. A knife edge seal segment is supported by each disk backbone extending from the disks and contacts the stator vanes to restrict leakage of compressed air from between the stator vane and the compressor rotor to limit the recirculation of air. Retaining flanges extend from each disk rim to retain the knife edge seal segments to the disk backbone and spacer bridges integral to the knife edge seal segments prevent axial movement of the knife edge seal segments. A plurality of lock assemblies are spaced about the circumference of the disk backbone to prevent circumferential shifting and rotating of the knife edge seal segments during operation.

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 shaft seal for a transmission expander or compressor having at least one rotor which rotates in a housing and is arranged in a floating manner on a shaft, the shaft seal having a seal arrangement with seal tips attached on the hub of the rotor. A stationary sealing ring surrounds the seal tips which are arranged in three seal sections. An annular chamber is arranged between in each adjacent pair of seal sections. The annular chamber which faces the interior of the transmission expander or compressor is equipped with a feed means for a sealing gas, the pressure of which is higher than the pressure in the interior of the transmission expander or compressor. The annular chamber which faces away from the interior of the transmission expander or compressor is provided with an extraction means for the sealing gas.

Abstract: A packing ring assembly is provided and includes an arcuate carrier ring segment having an aperture formed therethrough, an arcuate packing ring segment including an attachment component extending through the aperture of the carrier ring segment and movably associating with the carrier ring segment and the packing ring segment, and an actuator component disposed between the carrier ring segment and the packing ring segment and extending at least partially along the aperture. The actuator component is configured to maintain the packing ring segment in a first position relative to the carrier ring segment and to allow a movement of the packing ring segment to a second position relative to the carrier ring segment when the packing ring assembly is exposed to a pressure condition.

Abstract: A method of assembling a gas turbine engine includes providing a gas turbine engine including a compressor, a combustor downstream from the compressor, and a turbine coupled to the compressor, and coupling a seal assembly aft of the compressor such that air discharged through the seal assembly facilitates reducing the cavity pressure, and therefore reducing the axial forces induced to an aft side of the compressor.

Abstract: A turbine module for a gas turbine engine includes at least one turbine disk and a disk-shaped component mounted on the turbine disk upstream relative to the gas flow. The module can be assembled to the compressor of the engine, wherein the component is preassembled to the turbine disk, before the module is assembled to the compressor, by bolting to an attachment flange fixedly attached to the turbine disk.

Abstract: A method and apparatus for repairing damaged fins in a labyrinth seal, the seal without the use of welding or metallurgical bonding. The method comprises the steps of removing a section of the shaft containing the damaged fins, and providing a sleeve with replacement fins.

Abstract: A turbomachine is described, which includes a rotating member having an interface region with a stationary member. The interface region includes a pattern of concavities. A method for restricting the flow of a fluid through a gap between a stationary member and a rotating member is also described. The method includes the step of forming a pattern of concavities on at least one surface of the stationary member or the rotating member. The concavities have a size and shape sufficient to impede fluid flow.

Abstract: A method for assembling at least one seal assembly to a turbine having a casing extending at least partially around a rotor is provided. The method includes coupling the seal assembly to the casing such that the seal assembly extends into a fluid flow passage defined between the rotor and the casing and coupling at least one spring between the seal assembly and the casing to bias each seal assembly.

Abstract: A turbo-type fluid machine comprises a mouth ring part provided between a shroud of an impeller and a pump casing and having a seal function. The mouth ring part comprises an impeller ring on a side of the shroud of the impeller and a casing ring provided on a side of a non-rotating body wall surface of the casing. The impeller ring is structured to have a stepped portion that is in the form of a staircase with at least two steps to be small in diameter toward a suction side of the impeller and large in diameter toward an outlet of the impeller. The casing ring is labyrinth-shaped to comprise a projection and a recess. A narrow clearance part having a narrow radial clearance is formed between the projection and the impeller ring. A flow passage spatial portion defining an enlarged portion of the narrow radial clearance is formed by the recess and the stepped portion of the impeller ring.

Abstract: The turbine includes a diaphragm having a seal carrier mounted in opposition to seal teeth carried by the rotary component. The seal carrier includes a seal face having a coating of abradable material enabling the rotary component to abrade the material from the seal face. The seal carrier is removable from the diaphragm and is carried by an axial extension integral with or removable from the outer diaphragm ring. A spring may be interposed between the seal carrier and the stationary component.

Abstract: A combustion chamber inner casing (24) has a flange (28) on its downstream end. A retaining seal ring (40) is bolted to the flange (28) and traps a flanged seal ring (46) against the flange (28) in a manner that allows the seal ring (46) to move by expansion, relative to seal ring (40).

Abstract: A sealing arrangement for a compressor that compresses a gaseous mixture of air and fuel includes a pressurized air supply duct for supplying pressurized air into a leakage pathway defined between a compressor wheel and a housing of the compressor, which leakage pathway leads from the main gas flow path into a bearing area of the compressor, the pressurized air being supplied at a pressure sufficient to ensure that air and gaseous fuel cannot flow from the main gas flow path through the leakage pathway into the bearing area. A portion of the pressurized air flows into the bearing area, while the remainder of the air flows outwardly and is either fed back into the main gas flow path or is combined with air and gaseous fuel that leaks past an outer seal and is recirculated back to the compressor inlet.

Abstract: A seal system comprises a rotor and a stator. First and second cavities are defined between the rotor and the stator. A plurality of seals for inhibiting a flow of gas through the cavities are provided. Relative motion between the rotor and the stator can cause a flow of gas through the cavities via the seals. The. seals are arranged such that an increase in pressure in one of the first and second cavities is offset by a decrease in pressure in the other of the first and second cavities.

Abstract: Embodiments of the invention relate to a system and method for controlling the size of gaps in a turbine engine. In many instances, it is desirable to minimize the size of the gaps between neighboring rotating and stationary components in a turbine engine, such as between a disc cover plate and a proximate pre-swirler. According to embodiments of the invention, each component can be provided with a sealing surface. The sealing surfaces can be angled relative to the axis of rotation. The sealing surfaces are spaced from each other so as to form a gap therebetween. The sealing surfaces may or may not be substantially parallel. As a result of such configuration, the size of the gap can be controlled by axial and radial movement of the components. For example, the gap between the cover plate and the pre-swirler can be adjusted by axially movement of the rotor.