Abstract: A compressor assembly for a gas turbine engine includes an outer band, a plurality of variable pitch vanes, and an inner band. The plurality of variable pitch vanes extend radially between the outer band and the inner band. The inner band extends circumferentially partway about an axis and includes a first ring segment and a second ring segment that cooperate to receive the plurality of variable pitch vanes.

Abstract: A turbo machine including a device for ventilating and pressurising a turbine rotor of a turbomachine having an axis, the device including at least one collection pipe suitable for collecting a fraction of the air circulating in a high-pressure compressor of the turbomachine and conveying it to a first inner chamber inside the turbomachine that communicates with a second inner chamber delimited by the turbine rotor, the first and second chambers being at least partially separated by a stationary shroud having the axis, wherein it includes at least one injector passing through the stationary shroud and having a cross-section that varies in response to a pressure difference between the first and second chambers.

Abstract: Counter-rotating turbine (C) of a turbomachine (10) extending about an axis (X) and comprising an inner rotor configured to rotate about the axis of rotation (X), and comprising an inner drum on which an inner movable blading (22) is fixed, an outer rotor configured to rotate about the axis of rotation (X) in a direction opposite to the inner rotor, and comprising an outer drum (50) on which an outer movable blading (20) is fixed, the outer movable blading (20) comprising at least one fixing rod (212) extending through an orifice (51) of the outer drum (50), the outer movable blading (20) being fixed to the outer drum (50) via a clamping means (100) fixed to the fixing rod (212) from an outer face of the outer drum (50), a set ring (80) being disposed around the fixing rod (50) in the orifice (51) of the outer drum (50).

Abstract: An airfoil assembly includes an airfoil fairing, a spar, a seal plate, and a seal. The airfoil fairing has a fairing platform and a hollow airfoil section that extends from the fairing platform. The spar has a spar leg that extends in the hollow airfoil section. The spar leg defines a spar leg periphery. The seal plate is secured with the fairing platform. The seal plate has an opening, and the opening has an opening periphery that is complementary to the spar leg periphery. The spar leg extends through the opening. The seal is between the seal plate and the spar leg. The seal seals around the spar leg periphery.

Abstract: Turbine nozzle for a turbomachine extending about an axis, the nozzle including a plurality of sectors each having at least one vane which extends radially between an internal platform and an external platform, the nozzle including at least two sealing strips each including a main portion configured to ensure the sealing between the nozzle and an element of the turbomachine which is adjacent to the nozzle, and a first portion folded relative to the main portion, the first folded portion being in contact with an outer face of at least one platform, each strip being held to the at least one platform by a clamping device configured to clamp the first folded portion between the outer face of the platform and the clamping device.

Abstract: A geared architecture for a gas turbine engine includes a central gear supported for rotation about the axis, a plurality of intermediate gears engaged with the central gear and a ring gear circumscribing the intermediate gears. A first flexible coupling is provided between an input shaft driven by a turbine section and the sun gear. The geared architecture provides a power density comprising a power measured in horsepower (HP) related to a weight of the geared architecture within a defined range that benefits overall engine weight and efficiency.

Abstract: A vane assembly for use with a gas turbine engine includes an outer wall, an inner wall, and a plurality of airfoils. The outer wall and the inner wall extend at least partway about an axis. At least one of the airfoils is coupled with the outer end wall and the inner end wall to transmit force loads through the vane assembly.

Abstract: An end or intermediate casing for a combustion turbine engines includes prefabricated vanes of a first metal. Ends of the prefabricated vanes are then embedded within cast-in place inner and outer, annular-shaped ring castings, formed from a second metal having a lower melting point than the first metal. The respective ends of the prefabricated vanes include first and second shanks, with respective first and second surface features that are oriented transverse to the central axis of the vane are encapsulated in the molten second metal during the inner and outer ring casting. Once the castings harden, the first and second surface features, such as for example circumferential fillets projecting outwardly from the airfoil portion of the vane, inhibit separation of the vanes from the respective inner and outer rings.

Abstract: An assembly adapted for use in a gas turbine engine has a carrier component and a supported component. The assembly includes a mounting system for coupling the supported component to the carrier component. In an illustrative embodiment, the assembly is a turbine shroud segment for blocking gases from passing over turbine blades included in the gas turbine engine.

Abstract: An airfoil includes an airfoil body that extends between a leading edge and a trailing edge, a suction side and a pressure side, and a first end and a second end, a fitting located at one of the first end or the second end, the fitting including a body portion, a fillet portion, and a neck portion joining the body portion and the neck portion, the neck portion including a shelf, a fastener through the airfoil body, and a shroud having a complementary shape to the shelf such that the shroud nests in the shelf.

Abstract: A data processing method comprises: receiving a data packet related to a service. The method also comprises determining whether predefined configuration information indicates a priority of the service. The method further comprises in response to the configuration information indicating the priority of the service, adding an indication of the priority into the data packet. The above method can enable the underlying network devices to adopt a suitable strategy to process data packets from services of different priorities, so as to improve the transmission efficiency of the data packets and enhance user experience of the services.

Abstract: A seal member assembly structure and assembly method, a seal member, and a gas turbine, wherein, in a state where first flange parts provided on a combustor transition piece are fitted into first fitting parts provided in seal members, protruding parts provided on the first flange parts are inserted into recessed parts provided in the first fitting parts.

Abstract: A stator vane support with anti-rotation features is provided. The stator vane support may comprise an inner diameter surface opposite an outer diameter surface. The stator vane support may comprise an anti-rotation lug defining a protrusion extending inward from the inner diameter surface. The stator vane support may have a first recess defining a first void on the inner diameter surface proximate a first surface of the anti-rotation lug. The stator vane support may have a second recess defining a second void on the inner diameter surface proximate a second surface of the anti-rotation lug. The anti-rotation lug may be configured to interface with a stator vane to at least partially limit circumferential movement, and each recess may be configured to allow the stator vane to thermally expand during gas turbine engine operation.

Abstract: A nozzle assembly is disclosed, including a CMC nozzle shell, a nozzle spar, and an endwall. The CMC nozzle shell includes a CMC composition and an interior cavity. The nozzle spar is partially disposed within the interior cavity and includes a metallic composition, a cross-sectional conformation, a plurality of spacers protruding from the cross-sectional conformation, the plurality of spacers contacting the CMC nozzle shell, and a spar cap. The endwall includes at least one surface in lateral contact with the spar cap and maintains a lateral orientation of the CMC nozzle shell and the nozzle spar relative to the endwall. The lateral orientation maintains a predetermined throat area of the nozzle assembly. A method for forming the nozzle assembly includes inserting the nozzle spar into the interior cavity, rotating the CMC nozzle shell and the nozzle spar laterally relative to the endwall, and maintaining the lateral orientation.

Abstract: A turbine nozzle for a rotary machine including a centerline axis includes an airfoil including a leading edge and a trailing edge. The airfoil defines a throat location proximate the trailing edge. The turbine nozzle also includes an inner band assembly including a platform portion coupled to the airfoil, and a first flange coupled to the platform portion. The first flange is obliquely oriented with respect to the platform portion, and the platform portion and the first flange intersect at a point axially aligned with the throat location.

Abstract: Turbine center frames are provided. For example, a turbine center frame comprises an annular outer case and an annular hub. The hub is defined radially inward of the outer case such that the outer case circumferentially surrounds the hub. The turbine center frame further comprises an annular fairing extending between the outer case and the hub, a ligament extending from the fairing to the outer case to connect the fairing to the outer case, a plurality of struts extending from the hub to the outer case, and a boss structure defined on an outer surface of the outer case. The outer case, hub, fairing, ligament, plurality of struts, and boss structure are integrally formed as a single monolithic component. For instance, the turbine center frame is additively manufactured as an integral structure, and methods for manufacturing turbine center frames also are provided.

Abstract: According to some aspects of the present disclosure, a diffuser for a centrifugal compressor is provided. The diffuser may comprise an outerband casing and an innerband casing. The outerband casing may comprise an annular flowpath boundary member that has a flowpath boundary surface. The flowpath boundary member may define a plurality of vane-receiving pockets spaced about a circumference of the member. The innerband casing may comprise an annular flowpath boundary member that has a flowpath boundary surface. The flowpath boundary member may comprise a plurality of vanes spaced about a circumference of the member. Each of said plurality of vanes may comprise a vane body that extends from the flowpath boundary surface, a platform head that has a lateral dimension normal to the length of the vane body greater than the lateral dimension of the vane body, and a fillet between the platform head and the vane body.

Abstract: The invention is directed to a guide vane ring for a turbomachine, in which a vane bearing is produced in the inner ring via platform plates of the vanes and reliability against disintegration is produced via journals that extend radially inward from the platform plates, as well as a turbomachine.

Abstract: A nozzle box assembly includes steam inlets, through which working steam is supplied, a torus part connected to the steam inlets so as to form an annular steam path and having an opening portion, in which a part of the front surface of the annular steam path is opened, a bridge ring connected to the front surface of the torus part and having a bridge inside, and a steam path ring connected to the bridge ring so as to provide a path, which is connected to a stage, and provided with a plurality of vanes, wherein the bridge ring and the steam path ring are formed of the coupling of a plurality of divisions, which are divided in the circumferential direction thereof, and the divisions are coupled to the front surface of the torus part.

Abstract: A blade row group arrangeable in a main flow path of a fluid-flow machine and including N adjacent member blade rows firmly arranged relative to one another in both the axial direction and the circumferential direction is provided. Here, a front member blade row with front blades as well as a rear member blade row with rear blades are provided in the meridional plane established by the axial direction and the radial direction. The blade row group has two main flow path boundaries. It is provided that the blades of the member blade rows are fixed to the one main flow path boundary in the surrounding structure by means of a blade root structure, and that the blades of the member blade rows on the other main flow path boundary are each firmly connected to a base.

Abstract: A stator includes a plurality of stator vanes and a shroud operably connected to the stator vanes. The shroud includes one or more positioning tabs configured to engage one or more corresponding alignment features of a mating component to radially position the shroud at the mating component, the one or more shroud positioning tabs position the outer shroud to define a radial tip clearance between an outer shroud and an adjacent rotor of the gas turbine engine. A gas turbine engine includes a stator and case assembly in fluid communication with a combustor. The stator and case assembly includes a stator located at a case. The stator has a plurality of stator vanes and a shroud including one or more positioning tabs configured to engage one or more corresponding case alignment features to radially position the shroud at the case.

Abstract: A shroud mounting arrangement comprises a shroud housing and a shroud mounted to the shroud housing. The shroud is configured to surround a stage of rotor blades of a gas turbine engine. A circumferentially segmented vane ring is disposed axially adjacent to the stage of rotor blades. The circumferentially segmented vane ring comprises a plurality of vane segments. The vane segments jointly support the shroud housing.

Abstract: A shroud supports one of an inner and an outer trunnion on a variable vane. The shroud is provided by a first and a second axial half. Each of the axial halves is provided by a plurality of circumferentially spaced segments. Circumferential edges are defined on each of the plurality of circumferentially spaced segments. Edges between adjacent ones of the circumferentially spaced segments on the first half are circumferentially offset from the edges on adjacent ones of circumferentially spaced segments of the second half, such that no direct leakage path exists across an axial width of the shroud.

Abstract: A seal land for a gas turbine engine can include a seal body circumferentially extending about a longitudinal centerline axis. The seal body includes at least one sealing surface that extends in a plane that is transverse to the longitudinal centerline axis.

Abstract: A bundle guiding device includes an inner roller mounted on a front portion of a bundle in an insertion direction where the bundle is to be inserted into a casing and configured to roll on an inner peripheral surface of the casing, an outer roller mounted on a rear portion of the bundle in the insertion direction where the bundle is to be inserted into the casing and configured to travel on a rail member, and a position adjusting mechanism configured to adjust a position of the bundle in a horizontal direction orthogonal to the insertion direction by applying an adjusting force to a receiving portion, one of the position adjusting mechanism and the receiving portion being on the bundle, and the other of the position adjusting mechanism and the receiving portion being on one of a base surface and the rail member.

Abstract: Systems and methods are disclosed for anti-rotation lugs. A stator for a gas turbine engine may comprise an outer shroud, an inner shroud, and a plurality of vanes located between the outer shroud and the inner shroud. A plurality of anti-rotation lugs may be coupled to the inner shroud. The anti-rotation lugs may be configured to contact a diffuser case in order to prevent rotation of the stator. The anti-rotation lugs may comprise a body and a tapered shoulder. The tapered shoulder may distribute stress concentrations in the anti-rotation lugs.

Abstract: A turbine nozzle comprises a plurality of composite material vanes, each vane comprising an inner platform, an outer platform with attachment tabs on the outside, and at least one airfoil extending between the inner and outer platforms and secured thereto. The attachment tabs of the vanes are engaged on a metal assembly ring supporting all of the vanes, extending continuously along the outer platforms of adjacent vanes, and forming a distinct assembly part between the vanes and a turbine casing.

Abstract: An embodiment of the present invention is a turbine nozzle baffle including an annular box structure. The turbine nozzle baffle includes a first diaphragm, a second diaphragm, and a static portion of a rotating seal that form a one-piece annular box structure. The first diaphragm extends from an inner diameter of the box structure to an outer diameter of the box structure to form a first axial side. The second diaphragm includes a first portion and a second portion. The first portion extends from the inner diameter to the outer diameter to form a second axial side. The second portion extends toward the first axial side to form the outer diameter. The second diaphragm connects to the first diaphragm at the outer diameter. The static portion of the rotating seal faces radially inward at the inner diameter between the first axial side and the second axial side.

Abstract: The invention relates to a housing system (10) for an axial turbomachine, in particular for an aircraft turbine, comprising a housing (12), at least one guide blade (14), which is fastened to the housing (12) by means of at least one fastening element (18),a scaling element (20) for reducing a leakage flow between the listening element (18) and the housing (12), wherein the sealing element (20) is arranged between a downstream wall (22) of the fastening element (18) with respect to a flow direction (I) of the axial turbomachine and a wall (24) of the housing (12), and a pressing means (36), by means of which an axial force can be applied to the scaling element (20) And the scaling element can be pressed against the wall (24) of the housing (12), at least during the operation of the axial turbomachine. The invention further relates to an axial turbomachine having such a housing system (10).

Abstract: A fan inlet diffuser housing system includes an air cycle machine, a housing coupled to the air cycle machine, an inlet fan disposed on the air cycle machine and configured to pass inlet air into the fan inlet diffuser housing, a center tube disposed in the housing, a diffuser cone disposed around the center tube, an inboard strut supporting the inner tube within the diffuser cone, and an outboard strut supporting the inner tube within the diffuser cone, the outboard strut being integral with the diffuser cone, wherein the inboard strut is coupled to the inner tube with a first plurality of mechanical fasteners, and the outboard strut is coupled to the inner tube with a second plurality of mechanical fasteners.

Abstract: An exemplary turbomachine securing apparatus includes a coupling member that couples a first component to a second component within a turbomachine. The first component and the second component are both non-rotating components during operation of the turbomachine. The coupling member limiting relative circumferential movement between the first component and the second component, and permitting relative radial movement.

Abstract: A turbomachine comprises a flow duct with coaxially arranged radially inner and outer endwalls, first and second sets of axially spaced rotor stages arranged between the inner and outer endwalls, and a plurality of variable endwall segments arranged along the inner endwall. The first set of rotor stages rotates in a first direction, and the second set rotates in a second direction. The first and second sets alternate in axial series along the flow duct, such that axially adjacent rotor stages rotate in different directions. The variable endwall segments are radially positionable, in order to regulate loading on the first and second sets of rotor stages by changing a cross-sectional flow area between the inner and outer endwalls.

Abstract: A turbine is provided and includes a turbine shell including shrouds at multiple stages thereof, and constraining elements, disposed at least at first through fourth substantially regularly spaced perimetrical locations around the turbine shell, which are configured to concentrically constrain the shrouds of the turbine shell.

Abstract: An article of manufacture has a body formed in part of a first metal alloy and in part of a second metal alloy, the second metal alloy having a thermal coefficient of expansion that is less than the thermal coefficient of expansion of the first metal alloy. A BOAS segment for a gas turbine engine is disclosed wherein the formation of cracks due to thermal mechanical fatigue in the body of the disclosed BOAS segment is minimized, if not eliminated, through a unique construction of the disclosed BOAS segment, whether original equipment manufacture or a repaired blade outer air seal. A method for manufacture of a BOAS segment and a method for modifying a BOAS segment are disclosed.

Abstract: An air metering apparatus for a secondary air system of gas turbine engine includes a metering gap defined between surfaces comprised of one or more metals having one of similar coefficients of thermal expansion or similar thermal expansions at an operating temperature. Provided is an air metering gap which undergoes thermal growth in a way which provides a suitable air metering gap during engine running conditions.

Abstract: In exemplary embodiments, a nozzle can include a first flow wall, a second flow wall and a vane disposed between the first and second flow walls, wherein the vane is mechanically coupled to the first flow wall and in contact with the second flow wall.

Abstract: The present invention relates to an apparatus and a method for calculating a fatigue usage factor with respect to fatigue generated in the material of a mechanical device depending upon changes of an operation environment during operation of the mechanical device. A fatigue usage factor UOP.cyc,transient n, which is based on the operation-transient state n, is calculated by multiplying a fatigue usage factor based on a design-transient state UDSGN.cyc,transient n by the value of a characteristic fatigue usage factor F (k) as shown in the following equation: UOP.cyc,transient n=F(k)×UDSGN.

Abstract: A turbine is provided and includes a turbine shell including shrouds at multiple stages thereof, and constraining elements, disposed at least at first through fourth substantially regularly spaced perimetrical locations around the turbine shell, which are configured to concentrically constrain the shrouds of the turbine shell.

Abstract: An inlet case for a small gas turbine engine, where the inlet case is formed as a hybrid piece with a metallic inner shroud cylinder and plastic guide vanes and plastic outer shroud molded around the metallic inner shroud cylinder. The metallic inner shroud cylinder carries an electric generator and provides for a good convective heat transfer surface to carry heat away from the electric generator and into the inlet air flow entering the engine. The plastic guide vanes and outer shroud allow for a light weight inlet case assembly that is also much easier to form the airfoil shapes than a machined metallic inlet case. The plastic guide vanes and outer shroud are injection molded around the metallic inner shroud cylinder to produce a rigid single inlet case. The metallic cylinder includes axial extending slots that receive the lower ends of the guide vanes during the injection molding process.

Abstract: A vane assembly for a gas turbine engine comprising a number of radial loading elements disposed between lugs of the vane ring and the vane support, such as to generate a radial load force against the vane ring. The radial load force prevents unwanted relative movement between the vane ring and the vane support during operation of the gas turbine engine.

Abstract: Turbine nozzles and methods of manufacturing the turbine nozzles are provided. In an embodiment, by way of example only, a turbine nozzle includes a first ring, a vane, and a first joint. The first ring comprises a single unitary component and having a first opening and including a first metal alloy. The vane includes a first end disposed in the first opening and includes a second metal alloy. The first joint is formed in the first opening between the first ring and the vane and includes a first braze layer and an oxide layer. The first braze layer is disposed adjacent to the oxide layer, and the first braze layer and the oxide layer are disposed between the first ring and the vane.

Abstract: A turbomachine (1), in particular a turbine or compressor, includes a rotor (2) which has at least one moving blade row (4) with a plurality of moving blades (5), and a stator (3) which has at least one guide vane row (6) with a plurality of guide vanes (7), at least in one of the blade or vane rows (4, 6) all the blades or vanes (5, 7) each having a shroud plate (10). In order to simplify the maintenance of the turbo-machine (1), the shroud plates (10) are separate components with respect to the respective blade or vane (5, 7) and are mounted on the respective blade or vane (5, 7), shroud plates (10) adjacent in the circumferential direction (9) being coupled to one another for the transmission of tensile forces in the circumferential direction (9).

Abstract: An example exit guide vane for a gas turbine engine is mounted adjacent to a diffuser case defining an air flow path. A thrust balance seal is attached to the diffuser and the exit guide vane is mounted adjacent the diffuser case for guiding air flow into the air flow path. The exit guide vane is mounted adjacent to the diffuser without interfering with the air flow path or restricting thermal expansion.

Abstract: The present invention provides a mixer assembly including a mixer and a plurality of scoops attached around a circumference of the mixer element via a combination of sliding brackets and fixed brackets. The sliding brackets provide attachment and allow relative movement between the plurality of scoops and the mixer element while the fixed brackets provide only attachment between the mixer element and the plurality of scoops. The first sliding bracket includes a first flange that is fixedly mounted to the mixer element and a second flange that is biased against a surface of the scoop. A second sliding bracket is fixedly mounted to the mixer element and adjustably mounted to the scoop. The second sliding bracket includes an attachment hole and the scoop includes an adjustment slot. A fastener assembly including a spacer is received within the adjustment hole and the adjustment slot to allow relative movement between the mixer element and the scoop.

Abstract: A method of operating a gas turbine engine includes forming at least one stator ring passage within at least a portion of a radially inner surface of a casing. The method also includes providing at least one stator ring segment that is sized to be inserted into the at least one stator ring passage. The method further includes forming at least one radial passage within at least a portion of the at least one stator ring segment such that at least a portion of the at least one radial passage is adjacent to at least one substantially axial stator blade passage.

Abstract: A method of assembling a stator assembly includes coupling at least one stator ring segment to a portion of a casing using at least one circumferential stator ring groove defined in the casing. The method also includes coupling at least one stator blade assembly to a portion of the stator ring segment such that at least one radial passage is at least partially defined by at least one of the stator blade assembly and the stator ring segment. The stator assembly includes at least one first radial passage defined within a portion of the stator ring segment. The assembly also includes at least one second radial passage coupled in flow communication with the first radial passage. The second radial passage is at least partially adjacent to at least one of a portion of the stator blade assembly and a portion of the stator ring segment.

Abstract: One embodiment of a transition-to-turbine seal (300) comprises a transition exit seal (302) adapted for fixed attachment to an exit rail (350) of a transition (109), and a stage-1 seal (340). The stage-1 seal (340) comprises an axial seal slot (348) comprising parallel walls (346, 347) adapted to slidingly engage a first male member (310) of the transition exit seal (302). The stage-1 seal (340) also comprises a radial seal slot (354) comprising parallel walls (351, 353) adapted to slidingly engage a plate (338) of a row 1 vane segment (330). The transition-to-turbine seal (300) thus provides both axial and radial freedom of movement. Materials and coatings for the transition exit seal (302) and the stage-1 seal (340) are selected to promote preferential wear of these components relative to wear of the surface of the row 1 vane segment (330). Wear of the exit rail (350) is reduced or eliminated due to the noted fixed attachment.

Abstract: A seal arrangement 32 between a turbine shroud seal segments 38,40 in a gas turbine engine. The arrangement 32 comprises facing slots 34, 36 provided respectively in the segments 38,40 with offset spaces 42, 44 provided between the segments 38,40. A strip 46 is cemented into the slot 36 such that the flow path 50 passes through the space 42, around the free end of the strip 46, and out through the offset space 44, thereby providing a longer flow path.

Abstract: A seal usable to seal a transition in a can-annular combustion system of a turbine engine to a turbine vane assembly to direct exhaust gases through the turbine vane assembly. The seal may be formed from an elongated body extending along an outer edge of the transition and having first and second edges. The first edge of the seal may be attached to the transition, and the elongated body may extend away from the transition edge and contact a portion of the turbine vane assembly. The elongated body may flex during use without yielding or otherwise deforming.

Abstract: An exhaust diffuser unit for a gas turbine engine includes an outer ring, a diffuser cone and vanes, all made from a ceramic matrix composite material. Each vane has end components which have apertures through which a metallic support strut extends. The support strut is rigidly secured at its radially outer end to a diffuser casing, and is slidably but non-rotatably received at its radially inner end in an aperture in a boss provided on a drum. Structural loads between the casing and the drum are transferred through the strut independently of the vane. Gas loads on the vane are transferred to the strut by way of metallic end components.