Abstract: A rotary machine, such as a turbomachine of a gas turbine engine, for example, including a rotor and a rotor seal assembly. The rotor is rotatable about a rotational axis and has a rotor seal face. The rotor seal assembly includes a seal body, one or more primary fluid conduits formed in the seal body, and a retraction assembly. The seal body has a seal face and is positionable to form a fluid-bearing gap between the seal face of the seal body and the rotor seal face. The one or more primary fluid conduits are fluidly connected to the fluid-bearing gap to supply a fluid to the fluid-bearing gap. The retraction assembly is connected to the seal body to move the seal body in a retraction direction away from the rotor seal face.

Abstract: A blade outer airseal (BOAS) assembly of a gas turbine engine includes a BOAS segment, and a BOAS carrier located radially outboard of the BOAS segment relative to an engine central longitudinal axis. The BOAS segment secured to the BOAS carrier. The BOAS assembly is selectably radially movable during operation of the gas turbine engine. One or more secondary retention features are configured to limit the radial movement of the BOAS assembly.

Abstract: An abradable member for a turbomachine turbine, including a cellular structure including walls defining wells which open through a wear face. The cellular structure includes at least one flow straightener jutting from the wear face.

Abstract: A centrifugal compressor, has: an impeller; and a housing including: a shroud extending between a first end and a second end; a structural member having an outer end securable to a casing of the gas turbine engine, an inner end of the structural member intersecting the rear side of the shroud; and a reinforced region at the location where the structural member and the rear side of the shroud intersect, a thickness of the reinforced region in a direction normal to the gaspath side greater than a nominal thickness of the shroud, the reinforced region defining a curved surface extending from a first location on the structural member to a second location on the rear side of the shroud, a portion of the curved surface having a radius that increases from a first radius to a second radius at the second location.

Abstract: A blade outer air seal assembly includes a support structure. A blade outer air seal has a plurality of segments that extend circumferentially about an axis and mounted in the support structure. At least two of the segments have a first wall circumferentially spaced from a second wall. A base portion extends from the first wall to the second wall. A recess is arranged in at least one of the first wall and the second wall. An intersegment seal is arranged between the at least two segments. The intersegment seal has a circumferentially extending portion and a radially extending tab. The radially extending tab extends into the recess to define at least two cavities within the recess.

Abstract: A seal segment for a turbine and an assembly for sealing the gaps between seal segments and stator vanes of a turbine. The seal segments have a plate-shaped wall, the first lateral surface of which faces the vane tips in the assembled state of the seal segments, is surrounded by a closed circumferential edge, and can be divided into four lateral wall sections, and the plate-shaped wall has a seal element which is arranged over the entire surface of the lateral surface. A number of seal lamellae which are secured on one side are provided on at least one of the lateral wall sections and/or on at least one of the seal lateral wall sections facing adjacent seal segments when the seal segments are assembled in a turbine so as to form a ring in order to reduce a flow along the corresponding lateral wall section.

Abstract: A shroud attaching structure of a gas turbine includes: a support provided around an axis of the gas turbine and inside a casing of the gas turbine in a radial direction; and a shroud assembly attached to the support so as to cover an inner peripheral surface of the support, the shroud assembly being formed by laminating a large number of plate-shaped shroud elements containing a ceramic matrix composite. The shroud elements are lined up in a circumferential direction of the support. The adjacent shroud elements are arranged so as to be slidable on each other.

Abstract: A shroud assembly of a gas turbine includes a holder including an insertion hole, a plate-shaped shroud main body covering the holder from a radially inner side and made of a ceramic matrix composite, and a plate-shaped insertion member made of the ceramic matrix composite. The shroud main body includes a main body portion, a flange portion, and an insertion hole formed at the main body portion. The insertion member includes an insertion portion inserted into the insertion hole of the shroud main body and the insertion hole of the holder in an axial direction, and a head portion arranged at the radially inner side of the flange portion of the shroud main body and configured to be brought into contact with the flange portion from the radially inner side.

Abstract: A turbine shroud assembly of a gas turbine engine includes a carrier, a blade track segment and a mounting assembly. The carrier is made from metallic materials. The blade track segment is made from ceramic matrix composite materials. The mounting assembly is configured to couple the blade track segment to the carrier.

Abstract: A blade outer air seal includes a base portion that extends between a first circumferential side and a second circumferential side and from a first axial side to a second axial side. A first wall is axially spaced from a second wall. The first and second walls extend from the base portion. The second wall has at least one wall window configured to engage with a support structure. An outer wall radially spaced from the base portion between the first and second walls. The outer wall has at least one outer wall window configured to engage with the support structure.

Abstract: A seal assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a support mountable to an engine static structure. The support has a main body extending axially between leading and trailing edge portions and circumferentially between opposed mate faces, and the main body includes a pair of opposed retention hooks extending inwardly from the leading and trailing edge portions to bound an elongated groove extending circumferentially between the opposed mate faces. A seal has a sealing portion that extends from an engagement portion. The sealing portion has a sealing face that extends circumferentially between first and second mate faces. An overwrap has one or more plies that follow a perimeter of the engagement portion to define an interface between the retention hooks and the engagement portion. A method of sealing is also disclosed.

Abstract: A sealing assembly for a turbine is provided. The turbine includes stages each including a nozzle blade extending between radially inner and outer platforms, the outer platform including a spoiler which delimits a groove. The assembly includes a ring, a tab, and an anti-wear foil. The ring is to be attached to an outer casing of the turbine. A part of the sealing tab is engaged in a first housing of a ring sector of the ring and another part in a second housing of an adjacent ring sector. The anti-wear foil, open upstream, is mounted in the groove. The downstream end of the ring sectors are to be engaged in the anti-wear foil. Each of the first and second housings is formed by a slot. The tab is bent and disposed in the slots so that a downstream end of the tab extends out of the slots.

Abstract: A turbine engine component has a first portion and a hotspot portion with both portions being substantially covered with a thermal barrier coating (TBC). The hotspot portion may be additively manufactured to the first portion. Furthermore, the hotspot portion may have a single crystal microstructure to resist high temperatures and the first portion may not have such a microstructure and may further be cast. A method of forming the component includes the steps of casting the first portion and additively manufacturing the hotspot portion to the first portion, then covering the first and hotspot portion with the TBC.

Abstract: A filled abradable seal component, an associated method of manufacturing, and a turbomachine including the filled abradable seal component are disclosed. The method includes positioning the abradable seal component including a plurality of honeycomb cells, applying a filler material on the abradable seal component to fill the plurality of honeycomb cells, and curing the filler material at a temperature below 250 degrees Celsius to produce the filled abradable seal component. The filler material includes an abradable material, a binder material, and a fluid catalyst. The abradable material includes at least one of nickel chromium aluminum-bentonite, cobalt nickel chromium aluminum yttrium-polyester, cobalt nickel chromium aluminum yttrium-boron nitride, aluminum silicon-bentonite, aluminum bronze-polyester, nickel graphite, or aluminum silicon-boron nitride. The binder material includes at least one of aluminum, nickel-aluminum, aluminum thiophosphate, or aluminum thiosulfate.

Abstract: A seal assembly includes a seal arc segment defining first and second seal supports with a carriage defining first and second support members. The first support member supports the seal arc segment in a first ramped interface, and the second support member supports the seal arc segment in a second ramped interface. A spring is configured to bias the seal arc segment axially. A heat shield is radially inward of the spring.

Abstract: Various embodiments include gas turbine seals and methods of forming such seals for advanced gas path (AGP) heat transfer designed gas turbine engines. In some cases, a turbine includes: a first arcuate component adjacent to a second arcuate component, each arcuate component formed of a ceramic matrix composite and including a seal slot located in an end face, each seal slot having a plurality of axial surfaces and radially facing surfaces extending from opposite ends of the axial surfaces and an overhanging seal assembly disposed in the seal slot. The overhanging seal assembly including an intersegment seal including a plurality of seal segments defining two vertical seal segments and a plurality of horizontal seal segments and a first overhanging positioner shim seal and a second overhanging positioner shim seal, each including an extended portion. The overhanging seal assembly disposed in the seal slot so as to provide position constrainment of the two vertical seal segments.

Abstract: A blade track system includes an outer ring coupled with a blade track, and an inner ring positioned concentrically between the outer ring and the blade track. The outer ring designed to radially expand, at a first predetermined rate, in response to acceleration of a gas turbine engine. The segments of the blade track radially outwardly expanded away from a hub of the gas turbine engine by the radial expansion of the outer ring. The inner ring designed to radially contract inwardly toward the hub of the gas turbine engine at a second predetermined rate in response to deceleration of the gas turbine. The first predetermined rate being greater than the second predetermined rate, and radial contraction of the outer ring and the segments of the blade track constrained to the predetermined rate of contraction of the inner ring.

Abstract: A shroud hanger assembly or shroud assembly is provided for components which may be formed of materials having differing coefficient thermal expansion. The assembly includes a multi-piece hanger including a shroud positioned in a cavity between a first hanger portion and a second hanger portion. A shroud may be formed of a low coefficient of thermal expansion material which may have a differing coefficient thermal expansion than the material defining the shroud hanger. The shroud is deflected by an axial force acting between the hanger and the shroud which also forms a seal. The seal compensates for differing rates of thermal growth between the shroud and the hanger throughout the engine operating envelope.

Abstract: A blade outer air seal support includes, at least one arc body having a first portion and a second portion, a blade outer air seal mounting region defined at least partially between the first portion and the second portion, and an interface feature interfacing the first portion and the second portion. The interface feature is configured such that axially aligned forces are communicated between the first and second portions through the interface feature, bypassing the blade outer air seal mounting region.

Abstract: A shroud support system with load spreading comprises a shroud hanger having a first wall and a second wall spaced apart in an axial direction by a retainer support wall, a ceramic matrix composite shroud segment disposed in the shroud hanger between the first and second walls and the retainer support wall, a retainer having circumferentially spaced first and second bolt holes, the retainer passing through the shroud, first and second bolts passing through the shroud hanger and engaging the first and second bolt holes of the retainer.

Abstract: A blades outer air seal includes a seal arc segment that defines radially inner and outer sides. The radially outer side includes radially-extending sidewalls and a radially inner surface that joins the radially-extending sidewalls. The radially-extending sidewalls and the radially inner surface define a pocket. A manifold is disposed at least partially in the pocket. The manifold subdivides the pocket such that there is a manifold chamber bounded by the manifold and the radially inner surface. The manifold includes at least one inlet and a plurality of outlets.

Abstract: A turbine component including an outer shroud arranged within a turbine and including opposed extending portions. An inner shroud shields the outer shroud from a gas flowing along a gas path within the turbine during its operation and including opposed first and second arcuate portions extending around and in direct contact with a corresponding extending portion of the outer shroud. A first pin and second pin have a respective first end and second end. The first arcuate portion having a first engagement region for engaging the first end, and second arcuate portion having a second engagement region for engaging the second end. In response to engagement of the first engagement region and the first end of the first pin, and engagement of the second engagement region and the second end of the second pin, the inner shroud is prevented from twisting relative to the outer shroud.

Abstract: In one aspect the present subject matter is directed to a system for thermally shielding a portion of a shroud assembly for a gas turbine. The system includes a shroud support having a forward wall that includes a front side that is axially spaced from a back side and a radially inner surface that extends axially between the front and back sides. A shroud is mounted to the shroud support. The shroud includes a leading edge portion that extends towards the forward wall of the shroud support and a trailing edge portion that extends towards the aft wall of the shroud support. A radial gap is defined between a top surface of the leading edge portion and the radially inner surface of the forward wall. A thermal shield is disposed along a bottom portion of the forward wall and is oriented to face towards a flow of combustion gases.

Abstract: A turbine assembly can include a turbine wheel, a shroud component, a turbine housing and a seal that includes a wall and a lower lip that that extends radially outwardly from the wall where the seal is disposed, at least in part, between an outer surface of the shroud component and an inner surface of the turbine housing and where the lower lip is in contact with the turbine housing.

Abstract: A gas turbine includes an insert element for fastening to a ring segment body of a turbine of the gas turbine. The ring segment body has a recess on a hot gas side. The insert element is designed to cover the recess. The insert element includes a cover plate, which has a concavely shaped front side and a back side, and at least one foot arranged on the back side for positioning on the ring segment body. A mounting method includes mounting the insert element where the insert element is fixed to a recess of a ring segment body of a turbine of a gas turbine.

Abstract: A component including a porous portion which may be permeable or impermeable to air, and a method for making. In one example, the component is a cooled wall segment for a gas turbine engine, including a body defining a contact surface configured to be in contact with circulating hot gas and an outer surface configured to be in contact with cooling air. The body includes a first portion with at least one retention element, and a porous second portion made of a porous material permeable to air, containing a plurality of interconnected pores, and having a porosity greater than that of the first portion. The second portion is engaged to the first portion, defines at least part of the contact surface, and defines at least part of a fluid communication between the outer surface and the contact surface through the interconnected pores. The wall segment may be for example a heat shield or shroud segment. Methods of forming components are also discussed.

Abstract: A shroud and hanger assembly for a gas turbine engine and a method of transferring load from a ceramic matrix composite (CMC) shroud to a CMC shroud hanger assembly are provided. The shroud and hanger assembly includes a shroud hanger assembly formed of a first material having a first coefficient of thermal expansion, the shroud hanger assembly having a forward hanger portion and a separate aftward hanger portion and a shroud formed of a second material having a second coefficient of thermal expansion, the forward and aftward hanger portions configured to couple together to clamp across a forward radially extending wall of the shroud.

Abstract: A blade for a gas turbine engine includes a fixed length member and a floating blade seal that is movable relative to the floating blade seal to change the length of the blade and vary the gap between the blade and an engine housing component.

Abstract: A fan of a motor vehicle is particularly suited as a main fan of an internal combustion engine. The fan has a fan wheel with an outer ring which has a substantially L-shaped ring cross section. The L-shape is defined with a radial limb and an axial limb. The radial limb has a cross-sectional enlargement on a free-end side.

Abstract: A gas turbine engine is disclosed with a turbine section having at least one turbine rotor with a plurality of turbine blades, a plurality of blade tracks positioned circumferentially around the turbine blades, at least one dovetail shaped connecting member extending radially outward from each blade track, and a hanger connected to a structural member of the gas turbine engine and configured to releasably couple with the at least one dovetail shaped connecting member of a corresponding blade track.

Abstract: A turbine shroud for a gas turbine engine includes an annular metallic carrier, a ceramic blade track, and a cross-key connection formed between the annular metallic carrier and the ceramic blade track. The cross-key connection is formed between the annular metallic carrier and the ceramic blade track to locate the ceramic blade track relative to the annular metallic carrier. The cross-key connection includes a plurality of keys and a plurality of corresponding keyways.

Abstract: A sealing band arrangement for a gas turbine including first and second adjoining rotor disks separated by a gap wherein the first rotor disk includes an aperture. The sealing band arrangement includes at least one seal strip segment located within the gap, wherein the seal strip segment includes a raised portion having a first mating surface. The sealing arrangement further includes a locking pin having a planar section for receiving the first raised surface. The locking pin also includes a pin section having a second mating surface that abuts against the first mating surface to thereby lock the locking pin and the seal strip segment together. Further, the pin section is located within the aperture to stop circumferential movement of the seal strip segment relative to first and second disks.

Abstract: Turbine and compressor casing abradable component embodiments for turbine engines, with zig-zag pattern abradable surface ridges and grooves. 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: A seal assembly for sealing between a rotating component and a stationary component in a turbomachine. The seal assembly includes a plurality of radially inwardly projecting, axially spaced teeth extending from the stationary component, wherein at least one of the plurality of teeth has at least one axially extending hole therethrough. Axial flow of an operating fluid through the holes acts as an air-curtain to interrupt swirl flow in a seal cavity, therefore reducing steam force that could act to destabilize rotordynamics.

Abstract: A seal is positioned radially outwardly of a compressor blade. The seal has a seal hook facing in a downstream direction. A seal mount has a mount hook for receiving the seal hook in a gap. The hook face in an upstream direction. The seal mount has an upstream portion extending from an upstream end into the gap. The mount hook extends into a web which extends in a downstream direction. A thermal barrier coating on the upstream portion of the seal mount extends to a downstream end. There is a thermal barrier coating on a radially inner surface of the mount hook, and along the web to a downstream coating end. An upstream end of the coating on the inner surface of the hook being formed axially upstream of the downstream end of the coating on the upstream portion.

Abstract: A shroud segment for a gas turbine engine, the shroud segment constructed from a composite material including reinforcing fibers embedded in a matrix, and having a cross-sectional shape defined by opposed forward and aft walls, and opposed inner and outer walls, the walls extending between opposed first and second end faces, wherein the inner wall defines an arcuate inner flowpath surface; and wherein a compound fillet is disposed at a junction between first and second ones of the walls, the compound fillet including first and second portions, the second portion having a concave curvature extending into the first one of the walls.

Abstract: A turbine shroud has a plurality of shroud segments disposed circumferentially one adjacent to another. Each segment has a flow restrictor projecting integrally from one end face thereof and overlapping a corresponding end face of a circumferentially adjacent segment. The overlap between the circumferentially adjacent segments restricts gas leakage through the inter-segment gap between adjacent shroud segments.

Abstract: A blade outer air seal system includes a body that extends between two circumferential sides, a leading edge and a trailing edge, and a radially inner side and a radially outer side. An attachment section associated with the body and includes at least one engagement surface that is transverse to the radially outer side.

Abstract: An annular seal positionable in a gap is disclosed herein. The annular seal includes a plurality of leaves positioned about a center axis. Each leaf extends to a distal end that is elastically deflectable relative to the center axis to accommodate changes in a size of the gap. The annular seal also includes a ring encircling the center axis and at least partially interconnecting the plurality of leaves together for concurrent installation and removal. The annular seal also includes a plurality of tabs projecting from the ring along the center axis. Each of the tabs is directly connected to at least one of the plurality of leaves such that elastic deformation of one of the plurality of leaves induces bending in the corresponding tab before inducing bending in the ring.

Abstract: A seal arrangement for a gas turbine is disclosed. The seal arrangement is used for sealing a gap between radially internally located ends of guide vanes of a guide vane ring and a rotor, in which case the rotor has at least two seal projections positioned at an axial distance relative to each other in a circumferential direction of the rotor. The seal projections effecting a seal of the gap in combination with intake linings associated with the radially internally located ends of the guide vanes. The seal projections are inclined or tilted in the axial direction toward a side of higher pressure, where, in a space limited by the minimum of two seal projections and the corresponding intake linings, at least one recirculation structure is provided. The recirculation structure, or each recirculation structure, is oriented toward the side of the higher pressure.

Abstract: A method and system for sealing between components within a gas turbine is provided. A first recess defined in a first component receives a seal member. A second recess defined in a second component adjacent the first component also receives the seal member. The first and second recesses are located proximate a hot gas path defined through the gas turbine, and define circumferential paths about the turbine axis. The seal member includes a sealing face that extends in a direction substantially parallel to the turbine axis. The seal member also includes a plurality of seal layers, wherein at least one of the seal layers includes at least one stress relief region for facilitating flexing of the first seal member.

Abstract: A turbine flowpath apparatus for a gas turbine engine includes: a flowpath component exposed at least partially to a primary combustion gas flowpath of the engine, the flowpath component comprising low-ductility material; a metallic annular stationary structure surrounding the flowpath component, including a bearing surface which bears against the flowpath component, so as to restrain the flowpath component from axial movement in a first direction; and a spring element disposed between the flowpath component and the stationary structure which resilient urges the flowpath component in the first direction against the bearing surface.

Abstract: The present application relates to an axial turbine engine compressor having a stator with an annular row of stator blades extending radially, an inner ferrule disposed at the inner ends of the stator blades, a downstream brush seal disposed on the inner ferrule which includes a profile of revolution extending substantially axially. The brush seal comprises bristles which extend mainly axially. The compressor also includes a rotor having an annular sealing surface, generally cylindrical or frusto-conical, disposed on the downstream side of the stator blades and which cooperates with the brush seal, so as to provide sealing between the internal ferrule and the rotor. The annular surface surrounds the brush seal, such that the pressure downstream of the brush seal tends to push said seal away from the annular surface. The brush seal wears less in contact with the annular surface.

Abstract: A gas turbine engine section has a rotor carrying a plurality of blades. The blades have airfoils which define a radially outer tip. A blade outer air seal is positioned radially outwardly of the tips of the blades. The blade outer air seal is provided by at least a plurality of circumferentially spaced segments, which slide circumferentially relative to each other to adjust an inner diameter of an inner surface of the blade outer air seal segments. An actuator actuates the blade outer air seal segments to slide towards each other to control a clearance between the inner periphery of the blade outer air seal segments and the radially outer tip of the blade airfoils. A gas turbine engine is also disclosed.

Abstract: A gas turbine engine that includes: a flowpath defined through one of a compressor and a turbine; an inner casing defining an axially tilted outboard boundary of the flowpath, which, relative to the axial tilt, defines a converging direction in which the flowpath converges and a diverging direction in which the flowpath diverges; a row of rotor blades having outer tips that oppose the outboard boundary across a gap clearance defined therebetween; an outer casing concentrically arranged about the inner casing so to form an annulus therebetween; and a connection assembly that slidably connects the inner casing to the outer casing and includes a biasing means for axially preloading the inner casing in the converging direction.

Abstract: A system for passively varying an axial position of an inner casing of a gas turbine pursuant to changing pressure in a flowpath during transient engine operation. The system may include: a connection assembly slidably connecting the inner casing to the outer casing for axial movement of the inner casing between a first position and a second position; means for pressurizing the annulus relative to a flowpath pressure; biasing means for axially preloading the inner casing toward the first position; and an inner casing receiving surface configured to receive a pressure in the annulus for axially loading the inner casing in opposition to the axial preload of the biasing means.

Abstract: The invention relates to a housing for covering the ends of a row of rotor blades of an axial turbomachine compressor, the housing being provided with a sealing device between the blade tips and the housing. It comprises a shell segmented along its circumference, each segment being fixed to the housing by a series of elastomeric elements in a recess in the shape of a channel cut into the inner surface of the housing. In this way, in the event of misalignment of the rotor relative to the stator, the rotor blades coming into contact with sections of the shell will be able to move to compensate for this misalignment while reducing the frictional forces generated by contact between the blades and the shell. In the event of alignment being re-established the segments of the shell will be able to resume their initial position because of the elastic behavior of the elements.

Abstract: A system for operating a turbine includes a rotating component and a non-rotating component separated from the rotating component by a clearance. A first actuator is connected to the non-rotating component, and the first actuator comprises a shape-memory alloy. A method for operating a turbine includes sensing a parameter reflective of a clearance between a non-rotating component and a rotating component and generating a parameter signal reflective of the clearance. The method further includes generating a control signal to at least one actuator based on the parameter signal and moving at least a portion of the non-rotating component relative to the rotating component to change the clearance.

Abstract: Certain embodiments of the invention may include systems, methods and apparatus for providing a labyrinth seal. In an example embodiment, a method is provided for sealing a flow path between a stationary element and a rotating element of a turbomachine. The method can include disposing at least one fixture on an inner surface of a stationary element associated with the turbomachine; disposing a packing ring linked via a spring element to the fixture wherein the packing ring comprises at least one bore; and disposing, according to a predetermined profile, a plurality of interdigitated packing ring teeth and rotor teeth intermediate to the packing ring and the rotating element; wherein the bore, packing ring teeth, and rotor teeth, cooperate to counter a moment associated with one or more axial forces.

Abstract: A turbomachine rotor wheel including blades made of composite material; an annular mounting plate; a plurality of composite material blades mounted on the mounting plate, each blade also including a top platform at its tip carrying radial labyrinth teeth; an orientation disk fastened coaxially around the mounting plate and including at its outer periphery a plurality of axial slots that are open at one end, each slot presenting a cross-section that matches the cross-section of a blade root so as to hold it angularly in position by co-operation of shapes; and a variable-diameter spring ring housed in part in an axial notch formed in each blade root and projecting axially relative thereto, the spring ring to come into radial abutment against an inside surface of the orientation disk during radial outward movement of the blades relative to the mounting plate.