Abstract: A turbomachine sealing ring has an axis of revolution and includes an annular support, an annular coating made of abradable material which is carried by the support, and an annular thermal protection plate which is carried by the support. The ring is divided into sectors and has a plurality of ring sectors disposed circumferentially next to one another about the axis. Each ring sector has a support sector, a coating sector, and a plate sector. Each plate sector ling includes a flat tab that is pressed against a face of the corresponding support sector and is fixed to this face by brazing. The tab is inserted into a through-slot in the support sector in order to improve its integrity in operation.

Abstract: A stator configuration for a gas turbine engine including: a splitter segment, the splitter segment extending from a forward end to a rearward end; a forward most first stator extending radially outwardly from the splitter segment, the forward most first stator being completely located downstream from the forward end of the splitter segment; and a forward most second stator extending radially inwardly from the splitter segment, the splitter segment, the forward most first stator and the forward most second stator being formed as a single, integrally formed structure, the forward most first stator positioned closer to the forward end relative to a distance between the forward most second stator and the forward end and the forward most second stator positioned closer to the rearward end relative to a distance between the forward most first stator and the rearward end.

Abstract: A method of cleaning a component includes providing the component following the operation of the component in a high temperature environment, the component including a thermal barrier coating (TBC), cleaning the TBC of the component using a sponge jet blasting process, and measuring a cleaned thickness of the TBC to verify that the cleaned thickness exceeds a predetermined minimum value that will allow the return of the component to the high temperature environment.

Abstract: A gas turbine engine article includes a substrate that has a pre-bond surface that includes a serpentine groove. A coating is disposed on the pre-bond surface and mechanically interlocks with the serpentine groove.

Abstract: A fan track liner for a fan containment arrangement for a gas turbine engine comprises a cellular impact structure and a supporting sub-laminate integrally formed with each other from a fibre-reinforced polymer material.

Abstract: A turbine shroud segment for use in a gas turbine engine includes a ceramic shroud segment formed to define a circumferentially extending channel that opens radially inwardly and a layer of abradable material that extends axially along a radial inner surface of the ceramic shroud segment to provide a flow path surface of the turbine shroud segment.

Abstract: An actuator includes a primary electrode, a secondary electrode overlapping at least a portion of the primary electrode, a nanovoided polymer layer disposed between and abutting the primary electrode and the secondary electrode, the nanovoided polymer layer having a plurality of nanovoids dispersed throughout a polymer matrix, and a sealing layer at least partially encapsulating the nanovoided polymer layer, where the nanovoids include a fill gas.

Abstract: The invention relates to a ceramic sealing system between a rotor blade (120) and a housing (I). A small porous zirconium oxide layer (II) on a turbine rotor blade, which zirconium oxide layer faces a ceramic layer system (15?, 15?) of higher porosity, achieves durable sealing systems. The housing (I?) has a metal substrate (7), a metal adhesion-promoting layer (10), and a thick, outer, ceramic layer (15?, 15?) based on zirconium oxide, in particular having a porosity 2?14%.

Abstract: A flow-conducting component having at least one functional region for contact with a flowing medium and at least one functional region having supporting characteristics is provided. The two functional regions are produced from a material by successively solidifying layers using radiation in a manner that provides different material characteristics in the different functional regions.

Abstract: Shroud assembly of an axial turbomachine (2) having a rotor (12), such as a turbomachine compressor air-intake shroud (28). The assembly comprises an annular layer (32) of abradable material that is able to cooperate by abrasion with lips (44) of the turbomachine in order to ensure tightness, and a heating module (38) for the abradable layer (32). The module (38) comprises electrical heating resistors (42) within the abradable material in order to prevent the presence of ice on the shroud (28). The abradable material is thus used as a heating body which receives, transports and distributes the heat used for de-icing.

Abstract: Apparatus and method for volatilizing a source reagent susceptible to particle generation or presence of particles in the corresponding source reagent vapor, in which such particle generation or presence is suppressed by structural or processing features of the vapor generation system. Such apparatus and method are applicable to liquid and solid source reagents, particularly solid source reagents such as metal halides, e.g., hafnium chloride. The source reagent in one specific implementation is constituted by a porous monolithic bulk form of the source reagent material. The apparatus and method of the invention are usefully employed to provide source reagent vapor for applications such as atomic layer deposition (ALD) and ion implantation.

Abstract: A fan is mounted on a fan shaft which constitutes part of a mainline shaft assembly centered for rotation on an engine axis X-X of a gas turbine engine. When in operation, the fan rotates about the engine axis X-X. A stub-shaft is arranged radially outwardly and concentrically with the fan shaft and adjoins an upstream facing face of a bearing sleeve. A fan catcher arm extends from the stub-shaft through the bearing sleeve and joins a fan catcher ring arranged downstream of the bearing sleeve and having an upstream facing face arranged to abut against a downstream facing face of the bearing sleeve in the event of a fan blade-off. Immediately downstream of the fan catcher ring is formed a shaft end which is configured to couple with an axially adjacent end of a low pressure compressor shaft. Immediately downstream of the bearing sleeve is a shaft end configured to connect with a turbine shaft.

Abstract: An assembly for mounting a circumferential shroud to a casing, including a plurality of arcuate shroud segments collectively forming the circumferential shroud. Each segment includes a circumferentially extending wall and a channel extending from the wall. The channel is receivable in a hole passing through a circumferential casing and has a first connector. A fastener has a head and a shank, which is receivable in the channel and has a second connector that engages with the first connector. A first spacer receives the channel and a second spacer receives the shank and sit adjacent the head of the fastener on either side of the casing. The spacers can be selected to adjust for optimal radial positioning of the shroud and of probes received within a space defined between the casing and the shroud for monitoring rotor speed and the tip air gap of a turbine rotor contained in the casing.

Abstract: A ceramic honeycomb structure comprising porous cell walls defining large numbers of flow paths, the cell walls having (a) porosity of 55% or more and less than 65%, and (b) 35,000/mm3 or more of substrate branches, wherein the number of substrate branches is defined by the number of branch points (including connecting points of 3 or more branches and connecting points of different-width branches) per a unit volume, in a network structure obtained by the skeletonization of the three-dimensional structure of cell wall substrates determined by X-ray CT.

Abstract: A gas turbine engine includes an inlet duct, a compressor section, a combustor section, and a turbine section connected to drive the compressor section. The compressor section includes circumferentially-spaced blades having abrasive blade tips. A seal is disposed radially outwards of the blades. The seal includes a substrate that has a substrate hardness, an abradable layer that has an abradable layer hardness, and a hard interlayer between the substrate and the abradable layer. The hard interlayer has an interlayer hardness that is higher than the abradable layer hardness and higher than the substrate hardness.

Abstract: A corrosion resistant aluminum alloy abradable coating for use as a seal material consists of a porous base metal alloy layer containing corrosion inhibiting metal compounds dispersed throughout the porous base metal alloy layer. A method of forming a corrosion resistant aluminum alloy abradable coating consists of co-thermal spraying aluminum alloy powder plus polymer powder and particles containing corrosion inhibiting metal compounds.

Abstract: A stator seal for a turbine assembly includes a seal base securable to a turbine stator and including an annular inner surface, and an abradable coating disposed on the annular inner surface. The abradable coating and the annular inner surface have a predefined cross-sectional profile including a transient operation section that facilitates axial expansion and a steady state operation section that facilitates a tighter clearance.

Abstract: A seal segment according to an exemplary aspect of the present disclosure includes, among other things, a first axial wall, a second axial wall radially spaced from the first axial wall and a radially outer wall that interconnects the first axial wall and the second axial wall. At least one curved member is radially inwardly offset from the radially outer wall and extending between the first and second axial walls.

Abstract: A gas turbine engine component (50, 100, 150, 160, 174, 206, 236), including: a surface (54) subject to loss caused by a wear instrument during operation of the component in a gas turbine engine and a performance feature (80, 82, 102, 152, 162, 172, 200, 230) associated with the surface. The surface and the performance feature interact in a manner that changes with the loss such that a change in performance of the gas turbine engine resulting from the loss is mitigated.

Abstract: The invention relates to a composition for an abradable seal of a turbomachine, the composition comprising an aluminum base, nickel powder, polyester powder. The invention also relates to an outer casing of a low-pressure compressor of an axial-flow turbomachine with an abradable seal surrounding an annular row of rotor blades. The seal comprises a rounded support covered with a layer of abradable material comprising a metallic phase mainly made of aluminum and with nickel in a lesser proportion. The abradable material additionally comprises from 25% to 55% of additive, such as polyester, methyl methacrylate, hexagonal boron nitride, calcium fluoride. The support is segmented, and forms an organic matrix composite outer casing of the compressor. The invention also proposes a process for producing an abradable seal by plasma spraying an Al—Ni-polyester powder.

Abstract: A stiffness controlled abradeable seal system for a gas turbine engine includes a cantilevered arm that supports one of a rotating seal surface and a static seal surface, a stiffness of the cantilevered arm controlled to achieve a desired operational temperature at a seal interface.

Abstract: An article including a substrate and a plurality of coatings disposed on the substrate is presented. The plurality of coatings includes a thermal barrier coating disposed on the substrate; and a protective coating including a calcium-magnesium-aluminum-silicon-oxide (CMAS)-reactive material disposed on the thermal barrier coating. The CMAS-reactive material includes an NZP-type material. The CMAS-reactive material is present in the plurality of coatings in an effective amount to react with a CMAS composition at an operating temperature of the thermal barrier coating, thereby forming a reaction product having one or both of melting temperature and viscosity greater than that of the CMAS composition. A method of making the article and a related turbine engine component are also presented.

Abstract: The invention relates to a guide vane arrangement and a method for mounting a guide vane. The guide vane arrangement has at least one guide vane having a radially inner journal, which engages in a borehole of a bushing. The bushing is disposed in a borehole of an inner ring segment. The journal is connected to the bushing in a rotation-resistant manner and a slide mounting is provided between the bushing and the inner ring borehole. For mounting the guide vane in an inner ring segment, the journal of the guide vane is introduced into the borehole of the bushing. Further, the journal is fastened to the bushing in a rotation-resistant manner, and the bushing is slide-mounted with the inner ring segment. The slide mounting is conducted prior to or simultaneously with or after the fastening of the journal to the bushing in a rotation-resistant manner.

Abstract: A rotatable engine component, coating and method of coating the rotatable component includes a substrate, a transitional zone applied to the substrate, and an abradable zone applied to the transitional zone. During operation of an engine in a turbine, the abradable zone is consumed upon contact with a static portion of the engine preventing damage to the rotatable component.

Abstract: The invention relates to a blower (99), in particular for an engine cooling blower in a motor vehicle, comprising a housing (10) having a cooling air guide device (200), wherein the housing (10) is designed for mounting a drive unit (102, 103) and the cooling air guide device (200) in the motor vehicle, and for guiding a cooling air flow (20, 30, 31, 32) to the drive unit (102, 103), wherein the cooling air guide device (200) is designed as one piece with the housing (10).

Abstract: A gas turbine engine includes a compressor, a combustor section, and a turbine. The turbine includes an integrated case/stator segment that is comprised of a ceramic matrix composite material.

Abstract: The present application provides a seal carrier for use about a number of flow orifices of a platform of a turbine nozzle. The seal carrier may include an inner surface facing the platform with the inner surface having a number of slots therein aligning with the flow orifices of the platform and an opposed outer surface with a seal positioned about the outer surface.

Abstract: The present invention creates a layer system for the rotor/stator seal of a turbomachine, in particular a compressor, which is disposed between components of the turbomachine and can be run in with a movement of the components relative to one another, in such a way that at least one of the components is run into the layer system, having: a first adhesive layer disposed on at least one of the components; a protective layer disposed on the first adhesive layer; a second adhesive layer disposed on the protective layer; and a running-in layer, which is formed softer than the protective layer and which is disposed on the second adhesive layer. The present invention further provides a method for producing a layer system for the rotor/stator seal of a turbomachine, as well as a turbomachine.

Abstract: A composite fan case includes a composite fan case body having an outer surface. A bolt attachment is attached to the outer surface of the fan case and has an internal composite boss with at least one threaded metallic insert. The boss is attached to the outer surface of the fan case with a combination of fiber reinforced ply and adhesive. A bolt attachment and a method of attaching components to the outer periphery of the composite fan case are also disclosed.

Abstract: A turbine blade is provided and includes a tip end carrying a shroud and at least one fin, which extends radially away from the shroud. The fin includes a first sidewall and a second sidewall, which are spaced apart, arranged parallel to each other, and are connected to the shroud, and a cutting edge, which is connected to the first and second sidewalls. The cutting edge thereby creates a hollow space between the sidewalls, the shroud, and the cutting edge, and further extends radially away from the first and second sidewalls. Also provided is a method of manufacturing the blade by casting the blade as single piece with the hollow fin or by forging the blade; and machining the fin to create the first and second sidewalls and cutting edge thereby opening the hollow space between said sidewalls and the cutting edge.

Abstract: A low leakage, integral fixed vane system for a gas turbine engine compression system which can accommodate the latest 3-D aerodynamic airfoil geometries and provide endwall ovalization control. A unitized construction wherein the vanes are integral with the case wall in the compression system.

Abstract: Turbine and compressor casing abradable component embodiments for turbine engines, with nested loop pattern abradable surface ridges and grooves. The nested loops comprise nested fully closed loops or a spiraling maze pattern loop. Some embodiments include distinct forward upstream and aft downstream composite multi orientation groove and vertically projecting ridges planform patterns, to reduce, redirect and/or block blade tip airflow leakage downstream into the grooves rather than from turbine blade airfoil high to low pressure sides. Ridge or rib embodiments have first lower and second upper wear zones. The lower zone optimizes engine airflow characteristics while the upper zone is optimized to minimize blade tip gap and wear by being more easily abradable than the lower zone.

Abstract: Turbine and compressor casing abradable component embodiments for turbine engines, with composite grooves and vertically projecting rows of ridges in planform patterns, establishing upper and lower wear zones. The lower wear zone reduces, redirects and/or blocks blade tip downstream airflow leakage, while the upper wear zone is optimized to minimize blade tip gap and wear by being more easily abradable than the lower zone. An elongated first ridge in the lower wear zone terminates in a continuous surface plateau. A plurality of second ridges or nibs, separated by grooves, project from the plateau, forming the upper wear zone. Each of the second ridges has a planform cross section smaller than the plateau planform cross section and a height smaller than the first ridge height. Some embodiments of the second ridges have spacing, planform cross sections, heights and separating groove dimensions selected for shearing when contacted by turbine blade tips.

Abstract: A seal assembly between a hot gas path and a disc cavity in a turbine engine includes a non-rotatable vane assembly including a row of vanes and an inner shroud, a rotatable blade assembly adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor, and an annular wing member located radially between the hot gas path and the disc cavity. The wing member extends generally axially from the blade assembly toward the vane assembly and includes a plurality of circumferentially spaced apart flow passages extending therethrough from a radially inner surface thereof to a radially outer surface thereof. The flow passages effect a pumping of cooling fluid from the disc cavity toward the hot gas path during operation of the engine.

Abstract: Turbine and compressor casing abradable component embodiments for turbine engines, with composite grooves and vertically projecting rows of first ridges in planform patterns, to reduce, redirect and/or block blade tip airflow leakage downstream into the grooves rather than from turbine blade airfoil high to low pressure sides. The abradable component embodiments have a multi-depth grooves formed in the abradable substrate surface between pairs of elongated continuous tip surface first ridges. These ridge and multi-depth groove embodiments have first lower and second upper wear zones. The lower zone, in the deeper grooves, optimizes engine airflow characteristics, while the upper zone, in the shallower grooves, is optimized to minimize blade tip gap and wear by being more easily abradable than the lower zone.

Abstract: Turbine and compressor casing abradable component embodiments for turbine engines, with composite grooves and vertically projecting rows of stepped first ridges in planform patterns, to reduce, redirect and/or block blade tip airflow leakage downstream into the grooves rather than from turbine blade airfoil high to low pressure sides. Each stepped first ridge has a first portion proximal the substrate surface with a pair of first opposed lateral walls terminating in a plateau, and a second portion terminating in a ridge tip. These ridge or rib embodiments have first lower and second upper wear zones. The lower zone, which at and below first portion height, optimizes engine airflow characteristics, while the upper zone, between the plateau and the second portion ridge is optimized to minimize blade tip gap and wear by being more easily abradable than the lower zone.

Abstract: A low density abradable coating for use as a seal material is formed by depositing a metal bond coat followed by depositing a low density, porous, abradable metal seal layer on the bond coat. The seal layer is formed by co-depositing metallic sponge particles and metal precursor particles. The metal precursor particles decompose during heat treatment leaving fine pores distributed throughout the microstructure.

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

Abstract: A turbomachine casing assembly including: a first casing element arrangeable adjacent to one or more rotating aerofoil structures of a turbomachine; a second casing element provided at a radially outer position with respect to the first casing element; and an infill member disposed between the first and second casing elements, wherein an end of the infill member is set back from an end of the first casing element such that a void is provided between the first and second casing elements.

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

Abstract: A running-in coating for a compressor housing is disclosed. The running-in coating has at least two layers where a first layer is dimensionally stable and at least one additional layer has run-in capability.

Abstract: A turbomachine casing assembly including a first casing element locatable radially outward of one or more rotating aerofoil elements of a turbomachine, a second casing element located radially distal to the first casing element is provided. A void is provided between a radially proximal face of the first casing element and a radially proximal face of the second casing element, at the first end of the first casing element. The first end of the first casing element is located against a corresponding first end of the second casing element by an energy absorbing element, with the first casing element; including a cantilever. The cantilever is arranged in a region extending along the first casing element from the first end; and having a pivot axis at a mid portion thereof.

Abstract: A turbomachine rotor includes: a rotor body having an outer surface; and a patterned abrasive or abradable coating formed over the outer surface of the rotor body, the patterned abrasive or abradable coating for directing a flow of a working fluid across the turbomachine rotor.

Abstract: An abrasive coating for a rotor shaft, the abrasive coating having a metal bond coat layer on the rotor shaft and an abrasive layer over the bond layer for contact with vanes during operation of the rotor shaft. The abrasive coating includes a matrix containing a low strength filler having an increased concentration radially outward from the bond layer to the outer surface of the abrasive layer. The matrix may be ceramic or metal.

Abstract: A sealing structure in which free-cutting spacers on a sealing base plate and sealing fins on a rotor are opposed to each other is provided. The sealing base plate is disposed so as to be movable in an axial direction of the rotor. When a steam turbine increases in load, a pressure head moves in an axial steam-pressure loading direction of the rotor and the sealing base plate connected to the pressure head moves in the axial steam-pressure loading direction. Since the free-cutting spacers at positions opposed to those of the sealing fins existing after the movement of the sealing base plate are maintained in a non-contact state, clearances between the sealing fins and the free-cutting spacers are dimensionally minimized.

Abstract: Shroud device thermally protecting a gas turbine blade, having a ceramic layer and a metallic layer, the metallic layer being thermally protected by the ceramic layer, the ceramic layer being mechanically joined to the metallic layer by a fixation device having a plurality of protrusions located in the metallic layer designed so as to engage with a plurality of cavities located in the ceramic layer, such that there exists a gap between the cavities and the protrusions at ambient temperature, the gap disappearing at high temperature operation of the gas turbine, the protrusions being then locked into the cavities.

Abstract: A labyrinth seal is provided for forming a seal between a first and a second component which rotate relative to each other. The seal has an abradable lining mounted to the first component, and a plurality of fins projecting from the second component. The fins are arranged in abutment with the abradable lining to form a labyrinthal path for a flow of air through the seal. The seal further has a bypass passage which extends through the abradable lining. The bypass passage allows air to flow through the seal and bypass the labyrinthal path.

Abstract: A gas turbine engine component has an airfoil and a squealer tip. The airfoil has a pressure side and a suction side. The squealer tip is located at one end of the airfoil to engage with an adjacent surface and thereby form a seal. The squealer tip terminates in a squealer tip apex with an arched cross-sectional profile in a plane extending from the pressure side to the suction side of the airfoil. A method for producing an airfoil seal for the gas turbine engine component is also provided.

Abstract: A fragment containment assembly for a turbine is provided. The fragment containment assembly includes a plurality of bands disposed around a shroud of the turbine and positioned such that the shroud is disposed between blades of the turbine and the bands along radial directions outwardly extending from a shaft of the turbine. The bands include a material having a first modulus of toughness parameter that is greater than a second modulus of toughness parameter of the shroud at temperatures of at least 260 degrees Celsius. The bands are disposed around the shroud to prevent debris of the turbine from being released outside of the bands along the radial directions caused by failure of the turbine.

Abstract: A metal sealing material used in a sealing device which can reduce a gap between a stator and a rotor of a turbine. The metal sealing material used in a sealing device for a stator and a rotor of a turbine includes a surface layer and a lower layer composed of a porous metal layer, wherein the porosity of the surface layer is smaller than the porosity of the lower layer; the porosity of the surface layer is 60 to 65% and the porosity of the lower layer is 67 to 75% or less; and the porous metal layer has a thickness of 0.3 to 3.0 mm and may include, as a main component, an MCrAlY alloy where M is either one of Ni and Co or both thereof, and h-BN as a solid lubricant.