Abstract: A static gas turbine component, especially for an aircraft engine, is formed at least partially region-wise of metal foam. An abradable shroud lining of metal foam and a carrier allow a radial through-flow of gas.

Abstract: A turbine blade tip shroud. The tip shroud may include a platform, a number of Z-notched positioned within the platform, a number of seal rails positioned on the platform, and a number of cavities positioned within the platform and adjacent to seal rails and the Z-notches.

Abstract: A ceramic ring segment for a turbine engine that may be used as a replacement for one or more metal components. The ceramic ring segment may be formed from a plurality of ceramic plates, such as ceramic matrix composite plates, that are joined together using a strengthening mechanism to reinforce the ceramic plates while permitting the resulting ceramic article to be used as a replacement for components for turbine systems that are typically metal, thereby taking advantage of the properties provided by ceramic materials. The strengthening mechanism may include a bolt or a plurality of bolts designed to prevent delamination of the ceramic plates when in use by keeping the ceramic plates in compression.

Abstract: The invention provides a turbine machine rotor having blades of composite material, the rotor comprising an assembly disk, a plurality of composite material blades mounted on the assembly disk via their roots and each provided at its free end with a platform, and a metal shroud positioned on the blade platforms concentrically around the assembly disk and including labyrinth teeth that project radially outwards in order to abrade an abradable material carried by a casing surrounding the rotor, the casing being held stationary both axially and circumferentially relative to the blade platforms and being suitable for moving radially outwards in operation relative to the blade platforms.

Abstract: A seal (48) between a compressor rotor blade (26) and compressor casing (28) comprises an abradable structure (59) on the compressor casing (28). The abradable structure (59) comprises a metallic foam (60) having pores (66, 68). The metallic foam (60) has a first a region (62) and a second region (64). The first region (62) of the metallic foam (60) is arranged adjacent to the compressor casing (28) and the second region (64) of the metallic foam (60) is spaced from the compressor casing (28) by the first region (62) of the metallic foam (60). The pores (68) of the second region (64) of the metallic foam (60) contain an abradable material (70) and the pores (66) of the first region (62) of the metallic foam (60) do not contain an abradable material.

Abstract: A housing for a supercharger assembly is provided having an inner wall at least partially defining a rotor cavity. A layer is formed from a sacrificial polymeric material and is provided on at least a portion of the inner wall. The layer is operable to provide approximately zero running clearance and improve scuff resistance between the first and second rotors and the inner wall. The sacrificial polymeric material is applied to the inner wall by insert molding to form the layer. A method of forming the housing is also provided.

Abstract: Aspects of the invention relate to a ring seal for a turbine engine. The ring seal can be made up of a plurality of circumferentially abutted ring seal segments. Each ring seal segment can comprise a plurality of individual channels. The channels can be generally U-shaped in cross-section with a forward span, and aft span and an extension connecting therebetween. The channels can be positioned such that the aft span of one channel can substantially abut the forward span of another channel. The plurality of separate channels can be detachably coupled to each other by, for example, a plurality of pins. The ring seal segment according to aspects of the invention can facilitate numerous advantageous characteristics including greater material selection, selective cooling, improved serviceability, and reduced blade tip leakage. Moreover, the configuration is well suited to handle the operational loads of the turbine.

Abstract: A fan containment casing includes a core comprising a number of tows. The core has a central region and two end regions. The tows are grouped into a number of bands and form a number of filament-wound radially disposed layers. The bands of tows are interleaved as a number of direct bands and indirect bands along at least a portion of one or more of the layers. A method of forming a fan containment casing includes winding a number of tows on a liner to form a number of radially disposed layers, where each of the layers is continuously connected to at least one adjacent one of the layers across one of the end regions via at least one of the tows.

Abstract: This invention relates to an abradable coating system for use in axial turbine engines. When coated onto a turbine ring seal segment the coating system may allow formation of an individualized seal between turbine blade disks and the surrounding ring seal without causing excessive wear to the blade tips. The abradable coating system includes columns of an abradable material. Thus, interference between the blades and the abradable coating system causes the individual columns to break off at the base. This abrasion mechanism may reduce blade wear and spalling of the coating system when compared to conventional coatings.

Abstract: A ceramic ring segment for a turbine engine that may be used as a replacement for one or more metal components. The ceramic ring segment may be formed from a plurality of ceramic plates, such as ceramic matrix composite plates, that are joined together using a strengthening mechanism to reinforce the ceramic plates while permitting the resulting ceramic article to be used as a replacement for components for turbine systems that are typically metal, thereby taking advantage of the properties provided by ceramic materials. The strengthening mechanism may include a ceramic matrix composite overwrap or plurality of overwraps designed to help prevent delamination of the ceramic plates when the ceramic article is in use by placing the plates in compression.

Abstract: A method for fabricating a turbine bucket and an apparatus facilitate reducing tip shroud creep. The method includes providing a turbine bucket that includes a tip shroud including at least one seal rail. The method also includes coupling at least one cutter tooth to the tip shroud, wherein the at least one cutter tooth is fabricated from an abradable material that enables the at least one cutter tooth to be removed from the tip shroud during operation of the turbine engine.

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 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 gas turbine comprises forward and aft rows of rotatable blades coupled to a disc/rotor assembly, a row of stationary vanes positioned between the forward and aft rows of rotatable blades, and rotatable sealing apparatus. Each of the stationary vanes comprises an inner diameter platform having first sealing structure. The rotatable sealing apparatus comprises seal housing apparatus coupled to the disc/rotor assembly and has second sealing structure adapted to engage with the first sealing structure.

Abstract: An air seal unit to be positioned adjacent blade structure in a gas turbine is provided. The air seal unit comprises a main body including a plurality of first impingement cavities and at least one first interconnecting passage extending between and communicating with a first pair of the first impingement cavities. The first interconnecting passage is nonparallel to the first pair of the first impingement cavities and defines a path for cooling air to pass from one impingement cavity of the first pair of the first impingement cavities to another impingement cavity of the first pair of the first impingement cavities so as to strike a wall defining at least a part of the other impingement cavity of the first pair of the first impingement cavities.

Abstract: A turbine outer air seal includes, in a radially inner major surface thereof, at least one circumferential groove accommodating a plurality of holes which supply air for film cooling the radially inner surface of the air seal.

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 CMC wall (22) with a front surface (21) heated (24) by a working fluid in a gas turbine. A back CMC surface (23) is coated with a layer (42) of a thermally conductive material to accelerate heat transfer in the plane of the CMC wall (22), reducing thermal gradients (32-40) on the back CMC surface (23) caused by cold spots (32) resulting from impingement cooling flows (26). The conductive material (42) may have a coefficient of thermal conductivity at least 10 times greater than that of the CMC material (22), to provide a minimal thickness conductive layer (42). This reduces thermal gradient stresses within the CMC material (22), and minimizes differential thermal expansion stresses between the CMC material (22) and the thin conductive layer (42).

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 pattern for improving aerodynamic performance of a turbine includes a material disposed in a pattern at a base surface of a turbine shroud such that the material is capable of abradable contact with a tip portion of a turbine bucket. The pattern includes a first plurality of ridges disposed at the base surface such that a first portion of the first plurality of ridges corresponding to a back portion of the turbine bucket is oriented at a first angle with respect to an axis of rotation of the turbine bucket. Each ridge of the first plurality of ridges has a first sidewall and a second sidewall having a first end and a second end. The first ends of the first and second sidewalls extend from the base surface. The first and second sidewalls slope toward each other with substantially equal but opposite slopes until meeting at the second ends of respective first and second sidewalls defining a centerline and a top portion of the ridge.

Abstract: Aspects of the invention are directed to a system for improving engine performance by reducing vane tip clearances in the compressor section of a turbine engine. According to aspects of the invention, an abrasive material can be attached to a rotor disc. Thus, if a vane tip contacts the rotor disc during engine operation, the vane tip is worn away by the abrasive material. The system can reduce the risk of substantial component damage resulting from a vane tip rubbing event. The abrasive material can be provided in the form of an insert that can be removably attached to the rotor disc. Thus, the insert can be readily removed and replaced when necessary.

Abstract: An article of manufacture pattern for improving aerodynamic performance of a turbine including an abradable material capable of abradable contact. The abradable material is disposed in a pattern. The pattern includes a first plurality of ridges disposed at a base surface of the turbine. Each ridge of the first plurality of ridges has a first sidewall and a second sidewall having a first end and a second end. The first ends of the first and second sidewalls extend from the base surface. The first and second sidewalls slope toward each other with substantially equal but opposite slopes until meeting at the second ends of respective first and second sidewalls defining a centerline and a top portion of the ridge.

Abstract: A blade outer air seal used in a gas turbine engine, the BOAS including a metering plate with metering holes and an impingement plate with impingement holes, the metering plate and impingement plate forming a plurality of separate diffusion cavities forming a grid. A porous metallic plate is bonded to the underside of the impingement plate and has a plurality of cooling channels extending from the leading edge to the trailing edge of the BOAS. Cooling air from the blade ring carrier is metered through the metering holes and into the diffusion cavities, and then passes through a plurality of impingement holes and into a cooling channel, to be discharged out the trailing edge side of the BOAS. Inter-segment cooling holes also pass cooling air out to the sides of the BOAS.

Abstract: A gas turbine with a compressor includes at least one row of blades, with the blades having a free end each, with a self-healing abradable coating being provided adjacent to the free end of the blades on an annular casing area and/or an annular drum area.

Abstract: This invention relates to an abradable coating system for use in axial turbine engines. When coated onto a turbine ring seal segment the coating system may allow formation of an individualized seal between turbine blade disks and the surrounding ring seal without causing excessive wear to the blade tips. The abradable coating system includes columns of an abradable material. Thus, interference between the blades and the abradable coating system causes the individual columns to break off at the base. This abrasion mechanism may reduce blade wear and spalling of the coating system when compared to conventional coatings.

Abstract: An abradable rim seal arrangement for a turbine stage of a gas turbine engine, which includes an abradable rim seat attached to a leading edge of an annular inner duct of a turbine exhaust case, causing a trailing edge platform of each rotor blade to overlap the abradable rim seal, and allowing rubbing with the trailing edge platform during engine operation.

Abstract: A method for producing a contoured gap between a rotor and a stator of a turbo-engine is provided. A first turbo-engine drives a second turbo-engine by means of a common shaft, the shaft being mounted in a bearing housing by means of a bearing. The aim of the invention is to enable the contoured gap to be created in a simple and reliable manner. To this end, the contoured gap is formed between a contoured surface of the rotor and a contoured surface of an engine housing associated with the rotor, by grinding the contoured surfaces against each other, using the axial play of the bearing.

Abstract: A blade outer air seal for a gas turbine has an abradable coating with a substantially uniform thickness during service use. Prior to being placed in service, the abradable coating has a central portion with a greater thickness than leading and trailing portions. When placed in service, the blade outer air seal is positioned so that the central portion of the abradable coating with the greater thickness is located in a blade rub path of the turbine blades. As a result of abrasion in the blade rub path, the thickness of the central portion of the abradable coating is reduced so that it becomes substantially similar to the thickness of the leading and trailing portions.

Abstract: A compressor includes a compressor wheel, a diffuser and a device for the wet cleaning of the diffuser of a compressor, the diffuser being arranged downstream of the compressor wheel in the flow direction of the medium to be compressed. The device includes an arrangement for injecting a liquid, which is arranged in such a way that the liquid can be introduced, downstream of the compressor wheel, to the diffuser. The diffuser has at least partially a non-stick coating. A device of this type is suitable for removing deep-seated deposits from the surface of the diffuser.

Abstract: Systems and methods involving air seals are provided. In this regard, a representative air seal for a gas turbine engine comprises a layer of ceramic exhibiting patterned surface features, wherein a substrate supporting the layer of ceramic lacks the patterned surface features.

Abstract: A gas turbine engine blade containment system that includes a laminate material including first and second layers. The first layer is an inner layer within the second layer. The second layer includes a plurality of resiliently compressive springs extending transversely relative to the first layer, and each spring is encased in a crushable support material. The first layer includes a hard material to blunt a broken blade. The springs in the second layer are compressed and the crushable support material of the second layer is crushed to absorb energy of a broken blade.

Abstract: A turbine blade tip and shroud clearance control coating system comprising an abrasive blade tip coating and an abradable shroud coating are provided. The abrasive layer may comprise abrasive particles of cubic zirconia, cubic hafnia or mixtures thereof, and the abradable layer may be a nanolaminate thermal barrier coating that is softer than the abrasive layer. The invention further provides an alternate coating system comprising an abradable blade tip coating and an abrasive shroud coating.

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: The invention relates to a turbine blade with a cover plate shaped onto the pan of the blade. The aim of the invention is to provide a turbine blade of this type that, while having a high level of efficiency, is designed for a particularly reliable and safe operation in a turbine, particularly of a steam turbine. To this end, the invention provides that a protective layer made on an alternative material is applied to the surface of the cover plate facing away from the pan of the blade. The friction behavior with regard to a turbine component, particularly a sealing strip, which is opposite the protective layer, can be specifically influenced whereby enabling favorable emergency running properties to be provided in the event of rubbing.

Abstract: On an aircraft engine with a shroudless rotor wheel arranged in a flow duct, the flow duct includes a shroud segment (1) formed by a ceramic rubbing coating (3) which, while having good thermal conductivity, is highly-temperature resistant, attaches firmly to the metallic substrate (2) and is abradable by the tips of the rotor wheel to form a sealing gap as narrow as possible, this rubbing coating, owing to the lack of self-insulation, being coolable from a free side of the metallic substrate and, therefore, permitting for working gas temperatures occurring in high-pressure turbines and coating thicknesses sufficient for abrasion.

Abstract: A turbine blade tip and shroud clearance control coating system comprising a dense abrasive blade tip layer and an abradable shroud layer are provided. The dense abrasive coating may comprise cubic zirconia, hafnia or mixtures thereof and the abradable layer may be a nanolaminate thermal barrier coating that is softer than the dense abrasive layer.

Abstract: A housing for a supercharger assembly is provided having an inner wall at least partially defining a rotor cavity. A layer is formed from a sacrificial polymeric material and is provided on at least a portion of the inner wall. The layer is operable to provide approximately zero running clearance and improve scuff resistance between the first and second rotors and the inner wall. The sacrificial polymeric material is applied to the inner wall by insert molding to form the layer. A method of forming the housing is also provided.

Abstract: A pattern of depressions (36) in a sealing surface (34) on a CMC wall (32) of gas turbine ring segment (30) allows minimum clearance against turbine blades tips, and thus maximizes working gas sealing. An array of depressions (36) on the surface (34) increases abradability of the surface (34) by blade tip contact during zero clearance conditions and reduces blade tip damage. The depressions (36) are unconnected, preventing bypass of the working gas around the blade tips. A desired abradable surface geometry may be formed in a stacked laminate wall construction (40-43, 52) by staggered laminate edge profiles (50, 52) or by machining of depressions (36, 54) after construction.

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: An abradable sealing land for a gas turbine engine includes a mount plate and an open cell portion which is to be abraded. Both portions are formed integrally from a single piece of material.

Abstract: An abradable land assembly for a gas turbine engine includes a mount portion and an open cell portion which is to be abraded. Both portions are formed integrally from a single piece of material.

Abstract: A thermal turbomachine is disclosed having at least one row of rotor blades. At least one first rotor blade has a greater radial length than the others and at the blade tip is equipped with a first abrasive layer. At least one rotor blade which has a shorter radial length than the first rotor blade is equipped with a second abrasive layer at the blade tip. The first abrasive layer has a better cutting capacity and a lower thermal stability than the second abrasive layer. During commissioning of the thermal turbomachine, the first abrasive layer is in contact with the abradable layer of the stator, and during continuous operation of the thermal turbomachine the second abrasive layer is in contact with the abradable layer of the stator.

Abstract: A gas turbine engine component includes an airfoil having a radial outward end and a radial inward end. A seal member is positioned adjacent to the radial inward end of the airfoil. A tip of the radial inward end of the airfoil is coated with an abradable material. The seal member is coated with an abrasive material.

Abstract: Described is reinforcement of a fan case in a gas turbine jet engine. In one embodiment, a containment ring and a heat resistance ring are shrink interference fit on the inside diameter of the fan case, the containment ring where the large fan blades turn, and the heat resistance ring where heated air from backfiring heats up the fan case. In one example, the containment ring is made of a super alloy to provide added strength to the fan case should a fan blade break, containing the fan blade within the fan case. The heat resistance ring is made of titanium or other suitable material. Additionally, one or more stiffener rings may be shrink interference fit on the outside diameter of the fan case. The containment ring and stiffener rings can reduce the flight weight of the fan case and lower the material costs, while increasing the containment strength of the fan case. Other embodiments are described and claimed.

Abstract: A steam turbine has: a casing; a rotor rotatably arranged in the casing; a nozzle diaphragm concentrically arranged with respect to the rotor, the nozzle diaphragm being engaged with the casing; moving blades arranged in circumferential direction on outer circumference of the rotor at positions adjacent to the nozzle diaphragm; seal strips circumferentially extending on tips of the moving blades, the seal strips protruding in radial outward direction; and an abradable structure rigidly connected to the nozzle diaphragm. The abradable structure faces the seal strips in radial direction at a facing surface, and has an abradable part made of an abradable material arranged at the facing surface.

Abstract: A turbine blade tip and shroud clearance control coating system comprising an abrasive blade tip coating and an abradable shroud coating are provided. The abrasive layer may comprise abrasive particles of cubic zirconia, cubic hafnia or mixtures thereof, and the abradable layer may be a nanolaminate thermal barrier coating that is softer than the abrasive layer. The invention further provides an alternate coating system comprising an abradable blade tip coating and an abrasive shroud coating.

Abstract: A cooled turbine shroud includes an arcuate flow path surface adapted to surround a row of rotating turbine blades, and an opposed interior surface; a forward overhang defining an axially-facing leading edge, an outwardly-extending forward wall and an outwardly-extending aft wall; opposed first and second sidewalls, wherein the forward and aft walls and the sidewalls define an open shroud plenum; at least one leading edge cooling hole extending from the shroud plenum to the leading edge; and at least one sidewall cooling hole extending from the plenum to one of the sidewalls. The flow path surface is free of cooling holes and may include a protective coating applied thereto.

Abstract: Fabricating a refractory component for a gas turbine engine, such as a turbine shroud ring segment, by arranging refractory fiber tows (24) in a flaired tubular geometry (20) comprising a stem portion (21) and a funnel-shaped portion (22); impregnating the refractory fibers (24) with a ceramic matrix to form a flaired tube (20) of ceramic composite matrix material; at least partially filling the funnel-shaped portion (22) with a ceramic core (30) extending beyond the end of the funnel-shaped portion to provide a working gas containment surface (31); curing the flaired tube (20) and the ceramic core (30) together; cutting the funnel-shaped portion (22) to provide rectangular edges (27); and providing an attachment mechanism (34, 36, 38, 40) on the stem portion (21) for attaching the component to a surrounding support structure. Additional tows (24) may be introduced at intermediate stages to maintain a desired fabric density.

Abstract: A gas turbine engine is provided that includes one or more rotor assemblies, one or more shroud segments, and an actuator. The rotor assemblies are rotatable around an axially extending rotational axis, and each rotor assembly has a plurality of blades. Each blade has a blade tip. The one or more shroud segments each has a blade seal surface disposed radially outside of the one or more rotor assemblies. The blade tips and blade seal surfaces (collectively referred to as the blade seal surface) have mating conical geometries and a clearance distance extending between the blade tips and blade seal surface. The blade tips and blade seal surface are disposed at an angle relative to the axial centerline of the engine, which angle is greater than zero. The actuator is selectively operable to axially move one or both of the shroud segments and rotor assemblies relative to the other of the shroud segments and rotor assemblies to alter the clearance distance.

Abstract: Aspects of the invention are directed to a ceramic matrix composite ring seal segment. The ring seal segment according to aspects of the invention includes a relatively simple body that is circumferentially curved. At least a portion of the hot gas path surface of the ring seal segment can be coated with a thermal insulating. material. In one embodiment, each ring seal segment can be operatively connected to a stationary support structure, such as by way of isolation rings. The ring seal segments and/or the isolation rings can be configured so as to restrain the ring seal segments in the axial, radial and/or circumferential directions. The ring seal segments can be attached to the isolation rings so that the support points act opposite the operating pressure loads. Thus, the ring seal segments carry these loads in compression, a strong direction of the CMC fibers.