Abstract: An airfoil for a gas turbine engine includes pressure and suction walls spaced apart from one another and joined at leading and trailing edges to provide an airfoil having an exterior surface that extends in a radial direction to a tip. A camber line at the tip extends from the leading edge to the trailing edge. Pressure and suction side shelves are arranged in the exterior surface on opposing sides of the camber line respectively in the pressure and suction side walls. A plateau is proud of and separates the pressure and suction side shelves. The plateau is arranged along the camber line and extends to the leading edge.

Abstract: A turbine airfoil usable in a turbine engine and having at least one snubber with a snubber cooling system positioned therein and in communication with an airfoil cooling system is disclosed. The snubber may extend from the outer housing of the airfoil toward an adjacent turbine airfoil positioned within a row of airfoils. The snubber cooling system may include an inner cooling channel separated from an outer cooling channel by an inner wall. The inner wall may include a plurality of impingement cooling orifices that direct impingement fluid against an outer wall defining the outer cooling channel. In one embodiment, the cooling fluids may be exhausted from the snubber, and in another embodiment, the cooling fluids may be returned to the airfoil cooling system. Flow guides may be positioned in the outer cooling channel, which may reduce cross-flow by the impingement orifices, thereby increasing effectiveness.

Abstract: A bimetallic ring for use as a turbine shroud in a gas turbine engine. The bimetallic ring forms a sealing surface as a hot gas flow path boundary in the engine. The ring is comprised of two materials. The first material, a wrought, oxidation resistant metal alloy comprises a first portion, which is the hot gas flow path sealing surface. The second material, a low cost low alloy steel, comprises a second portion that may be at least a pair of supporting side plates. A dissimilar weld joint joins the sealing surface to the second portion, the at least pair of supporting side plates.

Abstract: The present invention relates to a gas turbine having a housing (1), an outer sealing ring (2) that can be fastened detachably to the housing, a clamping member (3) for clamping the outer sealing ring and the housing together radially, and a rotation locking member that has at least one housing groove (10) and a radial flange (20) of the outer sealing ring that can be locked against rotation in the housing groove in form-fitting manner with play (sa) in the axial and/or peripheral direction.

Abstract: Embodiments of a turbomachine, such as a gas turbine engine, are provided. In one embodiment, the turbomachine includes an impeller, a main intake plenum in fluid communication with the inlet of the impeller, and an impeller shroud recirculation system. The impeller shroud recirculation system includes an impeller shroud extending around at least a portion of the impeller and having a shroud port therein. A shroud port cover circumscribes at least a portion of the shroud port and cooperates therewith to at least partially define an impeller recirculation flow path. The impeller recirculation flow path has an outlet positioned to discharge airflow into the main intake plenum at a location radially outboard of the shroud port when pressurized air flows from the impeller, through the shroud port, and into the impeller recirculation flow path during operation of the turbomachine.

Abstract: Systems and programs for controlling thrust in a turbine by adjusting the clearance between a packing seal and a rotating component are disclosed. In one embodiment, a system includes a controller configured to move a packing seal to a selected clearance position in response to a thrust pressure exceeding a target pressure, and a sensor coupled to the controller and configured to detect the thrust pressure.

Abstract: A shroud segment to be arranged on a gas turbine blade is disclosed. The shroud segment includes a shroud segment surface and a stiffening structure that is raised relative to the shroud segment surface. The stiffening structure is cross-shaped at least in some areas.

Abstract: A system and method of controlling turbine blade tip-to-static structure clearance in a gas turbine engine includes processing aircraft data to determine when the aircraft is flying at altitude cruise conditions. When the aircraft is not flying at altitude cruise conditions, then compressor discharge air is directed to impinge upon at least a portion of the turbine shroud. Upon determining that the aircraft is flying at altitude cruise conditions, the compressor discharge air is directed to flow across at least a portion of the turbine case and to impinge upon at least a portion of the turbine shroud.

Abstract: A turbine shroud apparatus for a gas turbine engine having a centerline axis includes: a shroud segment having: an arcuate body extending axially between forward and aft ends and laterally between opposed end faces, wherein each of the end faces includes seal slots formed therein; and an arcuate stationary seal member mounted to the body; a turbine vane disposed axially aft of the shroud segment; and a casing surrounding the shroud segment and the turbine vane; wherein the turbine vane is mounted to the case so as to bear against the stationary seal member, compressing it and forcing the shroud segment radially outward against the casing.

Abstract: A sealing plate for sealing a gap between two primary sealing plates is provided. The sealing plate has a sealing portion and a connector portion, wherein the sealing portion has a width which is greater than a width of the connector portion. An apparatus for sealing a gap between two structural components of a turbine, the apparatus including a plurality of primary sealing components, and at least one sealing plate is also provided, as is a turbine having a plurality of structural components with a gap between said structural components, wherein the gap is sealed by a plurality of sealing components, the sealing components comprising a plurality of primary sealing components and at least one sealing plate.

Abstract: An impingement plate is cooperable with a shroud assembly. The shroud assembly includes an outer shroud and plural inner shrouds with seals between the plural inner shrouds, respectively. The impingement plate includes a trailing edge portion, a leading edge portion and a mid portion between the trailing edge portion and the leading edge portion. A plurality of impingement holes are formed across an area of the impingement plate, and a cooling and damping section includes at least one channel that is shaped to accelerate cooling flow through the impingement plate.

Abstract: There is provided a shroud structure for gas turbines capable of suppressing a drop in the amount of cooling air for cooling the inner shroud by reducing the amount of cooling air leakage that occurs along the cooling air path when feeding cooling air from the one-piece outer shroud to the inner shroud of the gas turbine and ensure more reliable cooling of the inner shroud. The gas turbine shroud structure contains a one-piece outer shroud, and an inner shroud retained on the inner circumferential side of the outer shroud in a structure divided into multiple inner shrouds along the periphery. An inner seal plate groove is formed on the outer circumference of the hook formed on the inner shroud, a seal plate is inserted in the inner seal plate groove, and the seal plate is mounted so that a section of the seal plate protrudes in the gap between the hook mechanism of the outer shroud and the inner shroud.

Abstract: The present invention relates to a run-in coating for a turbomachine, in particular for an aero engine, to be attached to a seal between a rotor and a stator, wherein the run-in coating comprises a metallic scaffold having interstices which are filled with an inorganic-nonmetallic filler material, and wherein the metallic scaffold is formed from an arrangement of metal fibers (9). The invention also relates to a corresponding turbomachine.

Abstract: A system for a turbofan has a variable outer air seal, an actuator coupled to the variable outer air seal, a sensor positioned inside the turbofan for sensing an engine parameter, and a flight controller connected to the turbofan for receiving the parameter from the sensor. The flight controller directs the actuator to regulate the position of the actuator based on the flight condition resulting form the received parameter.

Abstract: The various embodiments of the present invention provide a method for installing a tip shroud on a turbine disk in a steam turbine. A tip shroud in the form of a ring is installed on the outer periphery of the turbine disk to prevent a leakage of the driving fluid. The method comprises the steps of mounting turbine disk on to a rotor shaft, tapering the outer diameter of the turbine disk and the inner diameter of the tip shroud, aligning and mounting the tapered surface of the turbine disk with the tapered surface of the tip shroud and fastening the tip shroud on the turbine disk with the help of fasteners. The bolted tip shroud is finally machined to reduce the outer peripheral diameter to appropriate dimensions.

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 supercharger assembly has a rotor housing defining a chamber. A rotor is within the chamber and has an end with an end face. A seal has a seal face adjacent the end face. The seal face and the end face have complex topographies configured to be complementary to define a gap therebetween. The complex topographies can be, but are not limited to, interfitting concentric annular channels. The gap functions as a tortuous flow path to inhibit fluid flow past the end face. A method of manufacturing a supercharger assembly is also provided.

Abstract: A turbocharger has a casing system which houses a centrifugal impeller connected to a turbine by a shaft. The casing system includes an insert casing which forms a duct for feeding air to and through the impeller, and a volute casing which forms a volute for receiving compressed air from the impeller, the insert casing inserting into the volute casing. The casing system further includes a main casing which forms a housing for the shaft and for the shaft-side end of the impeller. The impeller includes a hub which has a radially outer annular rim, and has front and rear faces which converge towards the rim from respectively the inlet side and the shaft side of the hub. The impeller further includes a plurality of circumferentially arranged vanes on the front face of the hub. The impeller further includes a seal formation formed on the rear face at the rim.

Abstract: There is presented a casing arrangement for a gas turbine engine. The casing arrangement includes a plurality of tip treatment bars extending between a pair of spaced-apart annular supports, with each tip treatment bar being elongate and supported at each end by a respective support such that each end of the bar is received within an opening formed in the respective support. At least the ends of each tip treatment bar are tapered in transverse cross-section so as to have a wedge-shaped profile, and the openings in the supports have a complementary tapered shape for receipt and retention of the bar ends therein.

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 blade outer air seal (BOAS) segment according to an exemplary embodiment of the present disclosure can include a seal body having a radially inner face and a radially outer face that circumferentially extend between a first mate face and a second mate face and axially extend between a leading edge face and a trailing edge face. At least one of the first mate face, the second mate face, the radially outer face and the radially inner face can include an inner wall having a ledge and an undercut passage radially inward from the ledge that extends uninterrupted along a length of the ledge.

Abstract: In certain embodiments of the present disclosure a turbine casing assembly is described. The turbine casing assembly includes an inner casing and an outer casing surrounding the inner casing. The outer casing includes a first outer casing section and a second outer casing section that join together along a flange. Two bolts extend through the flange and join together the first outer casing section and the second outer casing section. A pin having a first segment having a first diameter and a second segment having a second diameter extends through the inner casing and the outer casing and supports the inner casing relative to the outer casing. The pin has a first diameter that is greater than the second diameter and is located between the two bolts along the axis of the flange.

Abstract: A composite annular shroud supported by a support assembly including at least two single piece full 360 degree rings and at least partially disposed within an innermost one of the rings. The shroud is biased against and in sealing engagement with an inner flange of the innermost ring. A three ring assembly includes the inner ring disposed radially inwardly of a middle ring disposed radially inwardly of an outer ring and the shroud at least partially disposed within the inner ring. At least three clocking pins extend radially inwardly from the middle ring through slots in the inner ring into notches in the shroud. The middle ring may be an aft end of a support ring fixedly connected to an engine backbone. Mounting pins may be press fitted into pin holes in the middle ring and extend radially outwardly from the middle ring through radial holes in the outer ring.

Abstract: A method and apparatus are disclosed for a gas turbine spool design combining metallic and ceramic components in a way that controls clearances between critical components over a range of engine operating temperatures and pressures. In a first embodiment, a ceramic turbine rotor rotates just inside a ceramic shroud and separated by a small clearance gap. The ceramic rotor is connected to a metallic volute. In order to accommodate the differential rates of thermal expansion between the ceramic rotor and metallic volute, an active clearance control system is used to maintain the desired axial clearance between ceramic rotor and the ceramic shroud over the range of engine operating temperatures. In a second embodiment, a ceramic turbine rotor rotates just inside a ceramic shroud which is part of a single piece ceramic volute/shroud assembly.

Abstract: The invention relates to a low pressure turbine (1) of a turbomachine comprising a nozzle (62) belonging to an upstream stage (20B) and a nozzle (162) belonging to a downstream stage (20C), the nozzle (62) forming segments (62a) having an external structure (66) comprising: a tab (76) bearing radially on a hook (79) fixed to the case, a tab (85) bearing radially on a reaming (87) in the case, the tabs (76, 85) delimiting an annular upstream cooling air circulation cavity (91), a contact surface (84) extending in the downstream direction from the tab (85) and lined with an abradable element (86) on the inside, the contact surface comprising a downstream end (92) bearing axially on the nozzle (162) and bearing radially on a stop (93) of the case, the contact surfaces (84) delimiting an annular downstream cooling air circulation cavity (94) communicating with cavity (91).

Abstract: A method for disassembly of a rotor, in particular the front rotor of a gas turbine with a housing and a channel which diverges in a direction of flow and in which the rotor is arranged, is disclosed. An embodiment of the method includes axially displacing an outer sealing ring that is radially opposite the rotor, and whose minimum inside diameter is smaller than the maximum outside diameter of the rotor, against the direction of flow. Then, the rotor is axially displaced against the direction of flow, in particular out of the housing. A method for assembly of such a rotor, as well as a tool for the same, is also disclosed.

Abstract: The main object of the invention is an angular stator sector (10) for a turbomachine compressor, comprising an outer shroud and an inner shroud (S) arranged coaxially one inside the other, and at least one vane (P) extending radially between the outer shroud and the inner shroud (S) and connected to the latter by its radial ends, characterised in that the inner shroud (S) comprises at least one brush seal (1, 2, 3) forming an obstacle to the recirculation of the downstream gases upstream of the inner shroud (S).

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 circumferentially extending sealing band is located within sealing band receiving slots formed in adjacent turbine engine disks. The sealing band includes a plurality of seal strips forming overlap joints defined by overlapping end portions, each formed by a tongue portion extending from a seal face of one seal strip past a seal face of the adjacent seal strip, along a radially inward facing side of the adjacent seal strip. A joint gap is defined within at least one overlap joint between the seal face of the one seal strip and the seal face of the adjacent strip, and a spacer is affixed to the one seal strip and is located at a position within the joint gap between the seal faces to limit circumferential movement of the seal faces toward each other.

Abstract: A fan includes a frame, rotor supported by this frame and having a central hub and a number of blades and a drive mechanism for rotatably driving the rotor, in addition to a ring to which the end zones of the blades are connected. The ring is assembled from two part-rings of the same form which each include a circular strip of sheet material, the free ends of which are mutually connected to form the outer surface of a truncated cone, and a third part-ring which mutually connects the inner edges of the first two part-rings. The end zones of the blades are connected to the third part-ring. The ring has a diameter of more than about 1.50 m. The number of blades amounts to at least eight. The blades are hollow and include a framework structure which is provided with a skin connected thereto.

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 bucket assembly is disclosed. The bucket assembly includes an airfoil having a generally aerodynamic contour and defining a tip, and a lower body portion extending generally radially inward from the airfoil. The bucket assembly further includes a tip shroud disposed on the tip of the airfoil and comprising a main body and a rail. The rail includes an exterior surface. The exterior surface defines a microchannel. The bucket assembly further includes a cover layer configured on the exterior surface.

Abstract: The invention relates to a compressor blade of a compressor which, along a main axis, comprises a blade base, a platform area and an adjacent blade profile having a profile tip. The blade profile is configured by a convex wall at the suction end and a concave wall at the pressure end opposite the wall at the suction end. These surfaces extend, with respect to a flow medium, from a leading edge to a trailing edge, a profile center line extending in the center between the two. A front face is arranged on the profile tip at an angle to the main axis, a sealing lip at least partially extending from the leading edge to the trailing edge and the blade profile including the sealing lip having a blade profile height extending in the direction of the main axis. In order to allow for an inexpensive compressor blade having improved aerodynamic properties and a modified sealing lip while having the same sealing properties, the height of the sealing lip is less than 2 percent of the height of the blade profile.

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 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: 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: 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: An blade outer air seal support assembly includes a main support member configured to support a blade outer air seal. The main support member extends generally axially between a leading edge portion and a trailing edge portion. The leading edge portion is configured to be slidably received within a groove established by the blade outer air seal. A support tab extends radially inward from the main support member toward the blade outer air seal. The support tab configured to contact an extension of the blade outer air seal to limit relative axial movement of the blade outer air seal. A gusset spans between the support tab and the main support member.

Abstract: A turbomachine includes a sealing segment ring that is provided between a front guide vane row and a back guide vane row for sealing a radial gap between a casing section and a rotor blade row rotating between the guide vane rows, wherein the sealing segment ring has a plurality of identical sealing segments and at least one of the guide vane rows has a plurality of identical guide vane segments, wherein the sealing segments each have a plurality of engagement sites lying adjacent to one another in the peripheral direction for interaction with securing elements of this guide vane row, wherein the engagement sites and securing elements are distributed uniformly over the periphery and the engagement sites are a multiple of the securing elements, a sealing element, and a guide vane segment.

Abstract: A blade track assembly is disclosed having a variety of features. The assembly can have annular or segmented components, or a combination of the two. In one form the assembly includes blade tracks having a forward and aft edge that can be received in an opening of respective hangers. The hangers can include anti-movement features to discourage movement of a blade track. A rib can extend between hangers and in one form can be used as part of a seal assembly. Clips can be used to secure the blade track in openings of the respective hangers, as well as to discourage movement of the blade track.

Abstract: A device for supporting a ring of a gas turbine, the ring configured to surround mobile blades of the turbine which are driven by a gas stream flowing upstream to downstream. The device includes at least one upstream hook facing upstream, to be housed in an upstream groove of the ring, opened toward the downstream direction, and at least one downstream hook facing downstream, to be housed in a downstream groove of the ring, opened toward the upstream direction. The hooks are protected and clearances at apexes of the blades are more easily controlled. The device further includes, upstream from the upstream hook, a mechanism injecting cooling gas to cool the upstream hook and/or includes, downstream from the downstream hook, a mechanism injecting a cooling gas to cool the downstream hook.

Abstract: A system includes a multi-stage turbine. The multi-stage turbine includes a first turbine stage with a first wheel having a plurality of first blade segments spaced circumferentially about the first wheel, a second turbine stage with a second wheel having a plurality of second blade segments spaced circumferentially about the second wheel, and an interstage seal assembly extending axially between the first and second turbine stages. The interstage seal assembly includes a first coverplate coupled to the first turbine stage. The first coverplate includes a first seal. The interstage seal assembly also includes a second coverplate coupled to the second turbine stage. The second coverplate includes a second seal. The interstage seal assembly also includes a spacer wheel extending from a rotor shaft and configured to engage with the first coverplate and the second coverplate.

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 turbine flowpath apparatus is provided for a gas turbine engine having a centerline axis. The apparatus includes: an annular flowpath member of low-ductility material, the flowpath member having a flowpath surface and an opposed back surface, and having a cross-sectional shape comprising a generally cylindrical forward section and an aft section that extends aft and radially outward at a non-perpendicular, non-parallel angle to the centerline axis; an annular stationary structure surrounding the flowpath member; and an annular centering spring disposed between the stationary structure and the flowpath member, the centering spring urging the flowpath member towards a centered position within the stationary structure.

Abstract: A sealing arrangement (40) is provided for sealing an interface between a casing main body (12) and a casing end cover (16). A circumferential lip (42) on the casing main body (12) defines a first radially facing circumferential surface (46) and the end cover (16) has a circumferential lip (50) defining a second radially facing circumferential surface (52). The sealing arrangement (40) includes a first circumferentially extending groove (55) formed in one of the first and second circumferential surfaces (46,52) and a sealing member (56) disposed within the first groove (55) in sealing contact with the one of said first groove (55) and in sealing contact with the other one of the first and second circumferentially extending surfaces (46,52).

Abstract: The invention relates to a sealing arrangement 15, a guide vane arrangement, and a turbomachine 11 with such a sealing arrangement 15, wherein the sealing arrangement 15 is designed for a guide vane ring 60 of a turbomachine 11, wherein the sealing arrangement 15 comprises a thin-walled annular structure 80 that is substantially closed on all sides, and wherein the annular structure 80 delimits an annular interior space 105, wherein a hollow cell structure 109, which is designed so as to mechanically support the annular structure 80, is provided in the annular interior space 105.

Abstract: A support, for example for a turboshaft engine, or for at least one bearing for a hot section of a turboshaft engine. The support includes at least a central hub incorporating an outer bearing seat for directly receiving the bearing, an annular casing segment around the central hub, and a plurality of radial arms connecting the central hub to the casing segment. The radial arms are inclined in an axial direction and in a tangential direction and are integrated as a single piece with the central hub and with the annular casing segment.

Abstract: A cover plate adapted to axially overlie root or shank portions of one or more buckets or blades secured to a turbomachine wheel, includes an arcuate cover plate body adapted to be secured to the turbomachine wheel so as to cover a root portion of at least one of the turbine buckets or blades; and at least one arcuate angel-wing seal segment detachably secured to one side of the arcuate cover plate body.

Abstract: According to one aspect of the invention, a turbine assembly includes a first component, a second component circumferentially adjacent to the first component, wherein the first and second components each have a surface proximate a hot gas path and a first side surface of the first component to abut a second side surface of the second component. The assembly also includes a first slot formed longitudinally in the first side surface, a second slot formed longitudinally in the second side surface, wherein the first and second slots are configured to receive a sealing member, and a first groove formed in a hot side surface of the first slot, the first groove extending axially from a leading edge to a trailing edge of the first component.

Abstract: A turbine shroud apparatus for a gas turbine engine includes: a plurality of arcuate shroud segments arranged as an annular shroud, each of the shroud segments comprising low-ductility material 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 an open channel is formed through the outer wall of each shroud segment; an annular stationary structure surrounding the shroud segments; and a hanger received in the open channel of each shroud segment and mechanically coupled to the stationary structure, each of the hangers passing through the respective open channel and including an enlarged portion having greater cross-sectional area than the open channel, the enlarged portion engaging the outer wall of the respective shroud segment, so as to retain the shroud segment radially relative to the stationary structure.