Abstract: A blade outer air seal segment including a radially outward surface, a radially inward surface oriented away from the radially outward surface, and a cooling channel located between the radially outward surface and the radially inward surface. The blade outer air seal segment also including a stress-relief boss extending into the cooling channel and an inlet orifice fluidly coupled to the cooling channel through the stress-relief boss. The blade outer air seal segment further including a stress-relief recess. The stress-relief boss being located within the stress relief recess.

Abstract: Disclosed are a flue pipe separation sensing apparatus and a boiler including the same. The flue pipe separation sensing apparatus of a boiler for sensing separation of an exhaust flue pipe that is fitted with and coupled to an exhaust adapter provided in an exhaust opening of a boiler body includes a sensing member located on an outer side of the exhaust flue pipe in a state, in which the exhaust flue pipe is coupled to the exhaust adapter, and that passes through the exhaust adapter by an elastic force and proceed into an interior of the exhaust adapter to sense the separation of the exhaust flue pipe when the exhaust flue pipe is separated from the exhaust adapter.

Abstract: A sealing device according to at least one embodiment includes not less than three arc-shaped fins arranged in an axial direction. The arc-shaped fins include: a first fin which is one of two outermost fins located on an outermost side in the axial direction; a second fin disposed adjacent to the first fin in the axial direction; and at least one third fin disposed opposite to the first fin across the second fin in the axial direction. It is preferable that the third fin is disposed to be inclined with respect to a radial direction such that a tip end portion is located on a side of the first fin in the axial direction relative to a base end portion, and the third fin has a larger inclination angle than the first fin or the second fin with respect to the radial direction.

Abstract: A process for additively controlled surface features of a gas turbine engine casing. The process comprises forming the casing having an inner surface and an outer surface opposite the inner surface; forming a surface feature on the casing proximate the inner surface, wherein the surface feature comprises a structure on the inner surface configured to align or misalign with respect to a flow direction of a working fluid in a flow path of the casing.

Abstract: A steam turbine provided with: a rotor; a plurality of movable blades each of which has a movable blade body and a shroud, and which are circumferentially arranged at intervals; a casing that has a recessed section for accommodating the shroud; fins that form clearances from the outer circumferential surface of the shroud; a stator blade support part disposed on the downstream side of the recessed section; a plurality of stator blades circumferentially arranged at intervals; and an annular partition plate that extends from a recessed section downstream surface toward the upstream side thereof. The casing has formed therein a bypass flow passage that communicates a region radially outside of the partition plate and a region between the stator blades.

Abstract: A centrifugal compressor includes a shaft member and an impeller. The shaft member extends in an axial direction and is rotatable. The impeller includes a hub that is disposed on an end portion of the shaft member, and a plurality of blades disposed on an outer peripheral surface of the hub. The impeller has a shape that tapers outward in the axial direction. A cover member is disposed at least at an end portion of the impeller outside in the axial direction. The cover member is joined to outer peripheries of the plurality of blades in a radial direction and covers at least a part of the plurality of blades from outside in the radial direction. A first bearing is disposed outside in the radial direction of the cover member, the first rotatably supporting the impeller.

Abstract: A ring assembly has a housing made of woven composite material carrying on its radially inner face an acoustic panel with a cellular structure covered with a composite material. The housing includes at least one radial recess in which is engaged at least one radially outwardly projecting part of the radially outer face of the acoustic panel.

Abstract: A turbine assembly (1) comprising: —a plurality of turbine ring sectors (20) made of ceramic-matrix composite material, —a ring support structure (3), comprising an annular shroud (6), and in addition —a plurality of angular spacer sectors (70) together forming an annular spacer (7), said annular spacer (7) being, on the one hand, fixed to the turbine ring (2) and, on the other hand, fixed to said annular shroud (6), characterized in that said turbine assembly (1) comprises at least one air diffuser (8), which is configured to diffuse cooling air onto the radially outer face (212) of at least one of said turbine ring sectors (20), and in that said at least one air diffuser (8) is mounted by being nested on one of said angular spacer sectors (70), in a nested position.

Abstract: A turbomachine turbine casing that extend around an axis and includes an annular wall and a cooling device. The annular wall is provided with a casing hook which extends in radial protrusion from an inside of the annular wall. The casing hook allows a mounting, on the turbomachine turbine casing, of ring segments disposed circumferentially end to end around the axis. The cooling device includes a collector duct intended to convey cooling air, the collector duct extending circumferentially around the annular wall. The collector duct has a cooling air inlet and a cooling air outlet. The collector duct and the annular wall have a common portion, which delimits the collector duct and from which the corresponding casing hook extends.

Abstract: A shroud hanger assembly is provided for hangers and shrouds defining dimensionally incompatible components such as those which are press or frictionally fit to engage one another. The shroud hanger assembly includes a multi-piece hanger and a shroud that is pinned to the hanger assembly by at least one axially extending pin and which locates the shroud relative to the hanger to control motion in one or both of circumferential (tangential) and radial directions relative to the engine.

Abstract: A compressor assembly may include a compressor housing and a compressor impeller disposed within the compressor housing. The compressor housing may have an internal aerodynamic surface that defines a circumferentially extending volute, and the compressor impeller may have an external aerodynamic surface that faces toward at least a portion of the internal aerodynamic surface of the compressor housing. A nonstick coating may be formed on the internal aerodynamic surface of the compressor housing or on the external aerodynamic surface of the compressor impeller. The nonstick coating may prevent foreign material introduced into the compressor assembly from collecting on the internal aerodynamic surface of the compressor housing or on the external aerodynamic surface of the compressor impeller.

Abstract: Disclosed is an outer air seal, having: an axial member, the axial member extending axially from an axial front end to an axial aft end, and extending radially from a radial inner surface to a radial outer surface; a radial flange extending radially from the radial outer surface of the axial member to a radial outer tip, and extending axially from an axial front surface to an axial aft surface; and a first kerf slot defined through the axial member from the axial front end to the axial aft end and from the radial inner surface to the radial outer surface, and through the radial flange from the axial front surface to the axial aft surface, wherein a radial top end of the first kerf slot is radially spaced apart from the radial outer tip of the radial flange.

Abstract: One feature pertains to an acoustic attenuation device. The acoustic attenuation device comprises a fan assembly, the fan assembly including at least one fan module that directs airflow in at least one direction, wherein the at least one fan module includes a top surface and a bottom surface, a plurality of air deflectors mounted to the top surface and the bottom surface of the at least one fan module, wherein the plurality of air deflectors include a plurality of surfaces, and wherein at least one of the plurality of surfaces is an angled surface that redirects the airflow 90-degrees, and acoustic attenuating foam, wherein the acoustic attenuating foam has a two-dimensional (2-D) surface and is applied to interior surfaces of the plurality of surfaces.

Abstract: A turbine blade tip, turbine blade and method where improved cooling is made possible by an improved cooling structure with cooling air holes inside a depression in a blade tip and a special arrangement of multiple cooling air holes which are supplied by a single cooling air channel inside a wall.

Abstract: The invention relates to a rotor blade (10) for a turbomachine, in particular of an aircraft, comprising an airfoil (12) comprising a pressure face (15) and a suction face (17) extending from a leading edge (14) to a trailing edge (16), the airfoil (12) comprising an axis of elongation extending substantially along the leading (14) and trailing (16) edges, the airfoil (12) comprising a radially firmer end for connection to a rotor and a free radially outer end. According to the invention, the airfoil (10) further comprises at least one series of fins (24) situated on said free end, each of these fins (24) comprising a pressure face (26) situated on the suction face (17) side of the airfoil (12), and a suction face (28) situated on the pressure face (15) side of the airfoil (12).

Abstract: A turbine blade includes a blade tip defining pressure side cooling apertures. The turbine blade defines a serpentine cooling passage having a first, second, and third legs, and first and second junction portions. The first leg extends radially and is connected to the second leg by the first junction portion proximate the blade tip. The second leg extends radially between the first and second junction portions. The second junction portion connects the second leg to the third leg which extends radially toward the blade tip and is connected to a trailing edge cooling aperture to exhaust the gas to an exterior of the turbine blade. The turbine blade defines a plenum connected to the first junction portion. At least one tip cooling aperture connects to the plenum and is radially outward of the third leg and axially aftward of at least a portion of the third leg.

Abstract: A flange arrangement of a gas turbine engine includes a first flange of a first component, and a second flange of a second component axially offset from the first flange along an engine central longitudinal axis. The first flange is secured to the second flange. One or more flange flowpaths are defined between the first flange and the second flange to convey a flange airflow from an interior of the second component thereby thermally conditioning the flange arrangement. The flange airflow is driven through the one or more flange flowpaths by a pressure differential.

Abstract: A centrifugal compressor includes a tubular shape casing extending in an axial direction, a motor stator fixed to an inside of the casing in a radial direction, a shaft disposed inside of the motor stator in the radial direction and extending in the axial direction, a motor rotor fixed to the shaft and facing the motor stator with a gap in the radial direction, a compression unit fixed to an end portion of the shaft on one side in the axial direction, a first thrust magnetic bearing disposed between the motor rotor and the impeller in the axial direction, and a second thrust magnetic bearing disposed on an other side in the axial direction relative to the motor rotor. A magnetic attraction force of the second thrust magnetic bearing toward the other side is larger than a magnetic attraction force of the first thrust magnetic bearing toward the one side.

Abstract: A casting method includes: forming a seed, the seed having a first end and a second end, the forming including bending a seed precursor; placing the seed second end in contact or spaced facing relation with a chill plate; contacting the first end with molten material; and cooling and solidifying the molten material so that a crystalline structure of the seed propagates into the solidifying material. The forming further included reducing a thickness of the seed proximate the first end relative to a thickness of the seed proximate the second end.

Abstract: An electric generator includes a process gas inlet on a downstream side of the flow control valve to receive process gas into the electric machine; a rotor shaft including a node position, the node position defining a position of a node of a first bending mode of the rotor shaft; a turbine wheel coupled to the rotor shaft at the node position, the turbine wheel configured to receive process gas and rotate in response to expansion of the process gas flowing into an inlet of the turbine wheel and out of the outlet of the turbine wheel, the rotor shaft configured to rotate with the turbine wheel, and a stationary stator, the electric generator to generate an alternating current upon rotation of the rotor within the stator. The turbine wheel selected from a plurality of turbine wheels based on the operational conditions of the electric generator.

Abstract: A ram air turbine rotor comprises at least one intra-flow path shroud structure coupled between rotor blades, along a radial position between a support disc and an outer rim. The shroud structure includes shroud sectors each coupled between a respective pair of blades. The sectors each include a first edge adjacent to leading edges of the respective pair of blades, the first edge including a first curved segment, and a second edge adjacent to trailing edges of the respective pair of blades, the second edge including a second curved segment. The curved segments are each partially defined by a respective ellipse having a semi-major axis and a semi-minor axis. The semi-major axis is a portion of a spanwise distance between the respective pair of blades. The semi-minor axis is a portion of an axial distance between the leading edge of one blade and the trailing edge of an adjacent blade.

Abstract: An impeller shroud assembly for a gas turbine engine includes an annular impeller shroud disposed about an axial centerline. The impeller shroud includes a shroud inducer portion and a shroud exducer portion disposed radially outward of the shroud inducer portion and extending to an outer radial end of the impeller shroud. The shroud inducer portion and the shroud exducer portion defining an impeller-facing surface of the impeller shroud. The impeller shroud has a pivot point defined between the shroud inducer portion and the shroud exducer portion. The impeller shroud assembly further includes a clearance control device connected to the shroud exducer portion of the impeller shroud proximate the outer radial end. The clearance control device is configured to pivot the shroud exducer portion of the impeller shroud about the pivot point between a first axial position and a second axial position.

Abstract: The invention relates to a turbine assembly comprising: a plurality of ring sectors forming a ring; a support structure comprising a shroud from which there project radial flanges which serve to hold latching tabs of each ring sector; a stator positioned downstream of said ring and comprising a blade provided with a radially outer platform and extending axially opposite said ring; an annular space being defined between the ring, the support structure and the platform, this annular space having passing through it a leakage air current, the assembly comprising a hollow ring which occupies said annular space and which is shaped so as to collect, channel and expel said leakage air current, more specifically to channel and expel it into the radially inner region of said annular space.

Abstract: An assembly adapted for use in a gas turbine engine includes a carrier and a blade track segment. The carrier extends at least partway about an axis. The blade track segment is supported by the carrier radially relative to the axis to define a portion of a gas path of the assembly.

Abstract: A turbine shroud assembly adapted for use with a gas turbine engine includes a carrier, a seal segment, and a mount pin assembly. The carrier is configured to be coupled to a turbine case. The seal segment is shaped to define a gas path boundary of the shroud assembly. The mount pin assembly is configured to couple the seal segment to the carrier.

Abstract: The invention relates to an airfoil as well as to a method for producing an airfoil for a turbomachine, comprising a leading edge and a trailing edge joined to each other by a suction side and a pressure side and which, in at least one region, extends in a curved manner from an airfoil root to an airfoil tip, wherein the airfoil tip has a squealer tip, which is arranged at the airfoil tip.

Abstract: A gas turbine compressor has a flow duct wall disposed radially opposite to an airfoil tip and has a circumferential groove having an upstream groove edge and a downstream groove edge, the circumferential groove having a web having a radial cutback.

Abstract: A shroud assembly including a shroud support and an annular shroud is provided. The shroud assembly includes one or more pins for securing the annular shroud to the shroud support. The pins having a block capable of translating radially to allowing the shroud to expand and contract in the radial direction. A gas turbine engine having a compressor section, a combustion section, a turbine section and a shroud assembly is also provided. The shroud assembly includes one or more pins for securing the continuous shroud to the shroud support. The pins having a block capable of translating radially to allowing the shroud to expand and contract in the radial direction. Methods for assembling a shroud assembly structure in a gas turbine engine are also provided.

Abstract: A method of assembling a blade outer air seal assembly includes engaging a first blade outer air seal with a first attachment surface on a first attachment body. The first blade outer air seal includes a first attachment body passage for accepting the first attachment body. A second blade outer air seal is engaged with a second attachment surface on the first attachment body. The second blade outer air seal includes a second attachment body passage for accepting the first attachment body. Rotation is prevented of the first attachment body relative to the first blade outer air seal with a first post engaging the first blade outer air seal. Rotation is prevented of the first attachment body relative to the second blade outer air seal with a second post engaging the second blade outer air seal.

Abstract: There is provided a system and a method for controlling a tip clearance between a turbine casing and turbine blade tips of an aircraft engine. At least one operational parameter of the aircraft engine is obtained. Based on the at least one operational parameter, a current value of the tip clearance and a target value of the tip clearance are determined. A limiting factor to be applied to the target value of the tip clearance is computed. The limiting factor is applied to the target value of the tip clearance to obtain a tip clearance demand for the aircraft engine. A tip clearance control apparatus of the aircraft engine is controlled based on a difference between the current value of the tip clearance and the tip clearance demand.

Abstract: A tribological and creep resistant system configured to operate at temperatures in excess of 700° C. A seal body extends between a leading edge and a trailing edge. A first component contact surface is adjacent the leading edge and a second component contact surface is adjacent the trailing edge. The seal body is formed from a high entropy alloy.

Abstract: A seal assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a first side plate and a backing plate that extend circumferentially about an assembly axis. The backing plate includes a main body and a first set of abutments that extend outwardly from the main body to abut the first side plate. A first set of pockets are established between adjacent abutments of the first set of abutments. A first brush seal is between the first side plate and the backing plate. Sections of the first brush seal received in respective pockets of the first set of pockets are dimensioned to join together to establish a first sealing relationship with a first gas turbine engine component. A method of assembly is also disclosed.

Abstract: An airfoil for a gas turbine engine defining a radial direction and an axial direction, the airfoil including: a flared portion extending from a suction surface of the airfoil at an outer edge of the airfoil along the radial direction; and a plenum disposed at the outer edge of the airfoil, the plenum having a suction-side sidewall and a pressure-side sidewall, wherein the suction-side sidewall has a first sidewall portion adjacent to a second sidewall portion disposed outside of the first sidewall portion along the radial direction, wherein the first sidewall portion defines a first angle, wherein the second sidewall portion defines a second angle, and wherein the first angle is greater than the second angle.

Abstract: A method for manufacturing a compressor includes a step of installing a bundle having an O-ring on an outer peripheral surface inside a lower half casing so that a position of the O-ring in an axial direction coincides with a position of a lower half relief groove, a step of installing an upper half casing having an upper half relief groove on the lower half casing so that a position of the upper half relief groove in the axial direction coincides with the position of the O-ring, and a step of pressing the bundle in the axial direction to move the O-ring to a position away from the lower half relief groove and the upper half relief groove, and bringing the O-ring into contact with the inner peripheral surface of the lower half casing and the inner peripheral surface of the upper half casing.

Abstract: A ring segment having improved cooling efficiency is provided. The ring segment may include a shield plate mounted to a casing which accommodates a turbine and configured to face an inner wall of the casing, a pair of hooks configured to protrude from the shield plate toward the casing to be coupled to the casing, a cavity defined between the shield plate and the pair of hooks, a plurality of first cooling passages configured to connect the cavity and first side surfaces facing each other of the shield plate, and a plurality of second cooling passages configured to connect the cavity and second side surfaces facing each other of the shield plate, wherein the first cooling passages extend in a longitudinal direction of a central axis of the turbine, and the second cooling passages extend in a circumferential direction of the turbine.

Abstract: An internal casing ferrule for a turbomachine, centered on a longitudinal central axis, which includes: a main body centered on this axis, with two ends, delimited by surfaces that are radially inner and outer relative to this axis; a thermal porous-structure insulation envelope having a volumetric porosity ?50%, which includes: a lateral portion entirely covering the two ends; when viewed in section transversely to this axis, exterior and interior portions entirely covering, respectively, the radially outer and inner surfaces of the main body; and a protective envelope which at least partially covers the envelope and which includes, when viewed in section transversely to this axis, radially outer and inner protective portions, respectively covering, at least in part, the exterior and interior portions.

Abstract: A process for manufacturing a preformed sheet having geometric surface features for a geometrically segmented abradable ceramic thermal barrier coating on a turbine engine component, the process comprising the steps of providing a preformed sheet material. The process includes forming a partially of geometric surface features in the sheet material. The process includes joining the sheet material to a substrate of the turbine engine component. The process includes disposing a thermally insulating topcoat over the geometric surface features and forming segmented portions that are separated by faults extending through the thermally insulating topcoat from the geometric surface features.

Abstract: A ring segment and a turbomachine including the ring segment are provided. The ring segment installed on an inner circumferential surface of a casing and disposed to face an end of a blade disposed inside the casing, the ring segment includes a segment body disposed inside the casing in a radial direction of the casing and including a plurality of cooling channels through which cooling air flows, a pair of segment protrusions protruding outward from the segment body, coupled to the inner circumferential surface of the casing, and spaced apart from each other in a flow direction of fluid flowing through the casing to form an RS cavity into which cooling air is introduced, wherein when the segment body has a cross section along an imaginary plane including a radial straight line of the casing, the cooling channel is formed such that a width in a direction perpendicular to a radial direction of the casing is greater than a width in the radial direction of the casing.

Abstract: A wind wheel including a first air plate, a second air plate, and a plurality of blades disposed between the first air plate and the second air plate. The first air plate includes a central part and an air inlet disposed in the central part; each two adjacent blades form an air duct. The plurality of blades each includes an inlet section next to the air inlet, an intermediate section, and an outlet section. The thickness of the intermediate section is greater than the thickness of the inlet section and the thickness of the outlet section.

Abstract: A cooling assembly comprises a coolant source chamber inside an airfoil that directs coolant inside the airfoil that extends between a hub end and a tip end that includes a tip body and tip rail along a radial length. A first body cooling chamber and a second body cooling chamber are disposed inside the tip body. The second body cooling chamber is positioned between the tip end and the first body cooling chamber. At least one of the first or second body cooling chambers are fluidly coupled with the coolant source chamber. The coolant source chamber directs the coolant into the first or second body cooling chambers. A rail cooling chamber disposed inside of the tip rail is fluidly coupled with the first or second body cooling chambers. The first or second body cooling chambers directs coolant out of the body cooling chambers and into the rail cooling chamber.

Abstract: An example high-performance system includes an example high-performance component including a substrate and a multilayer abradable track adjacent to the substrate. The abradable track includes a plurality of alternating layers along a thickness of the abradable track. The plurality of alternating layers includes at least one relatively porous abradable layer and at least one relatively dense layer. A porosity of the relatively dense layer is lower than that of the at least one relatively porous abradable layer. The example high-performance system may include a rotating component configured to contact and abrade the multilayer abradable track. An example technique for forming the multilayer abradable track includes thermal spraying a first precursor composition toward the substrate to form a relatively porous abradable layer of a layer pair of a plurality of layer pairs of the multilayer abradable track, and a second precursor composition to form a relatively dense layer of the pair.

Abstract: A gas turbine engine may include a fan, a plurality of inlet pre-swirl features disposed upstream of a fan, and an outlet guide vane assembly disposed downstream of the fan. The outlet guide vane assembly includes a plurality of outlet guide vanes that may define an outlet guide vane solidity profile, wherein the outlet guide vane solidity profile is achieves a minimum solidity at a radial position between an inner boundary and seventy percent of an outlet guide vane span. The fan includes a plurality of fan blades that may define a fan solidity profile, wherein the fan solidity profile maintains a solidity of greater than 1.1 between a radial position at seventy percent of the fan blade span and an outer boundary.

Abstract: Rotor systems and methods for improved performance with extended range to stall fabricated through the use of additive manufacturing. A rotor has blades that extend to tips and rotates about an axis. A casing fits over the rotor so that the tips are configured to pass proximate the casing when the rotor rotates. The casing channels a flow stream across the rotor. Grooves are defined in the casing and extend longitudinally at an acute angle relative to the axis. The grooves extend a distance upstream from a leading edge of the blades and over at least a portion of the blade tips so that the blade tips are configured to pass across the grooves when the rotor rotates.

Abstract: The invention is related to the gas turbine stators of the gas turbine engines applied in aviation. The gas turbine stator, in the outer housing of which sectors of the split honeycomb ring (made as double-layer one) are installed with support elements on the front and rear axial ends of the sector. In this invention, the layer of the sector facing the outer housing is made U-shaped in the plane, the support elements are made as separate rotary bodies distributed uniformly along the circumference and the front support elements (on the gas flow direction) are larger than rear ones in terms of geometrical dimensions by factor 1.1 . . . 1.5. Therefore, the implementation of the invention proposed with the characteristic features above, in conjunction with the known features of the invention claimed enables reduction of the gas turbine stator weight and improvement its reliability without compromising the turbine efficiency.

Abstract: A flow path component assembly includes a flow path component having a plurality of segments that extend circumferentially about an axis and mounted in a support structure. At least one of the plurality of segments have a first wall and a second wall that extend radially outward from a base portion. The first wall is axially spaced from the second wall. A coating is on a portion of the first wall and a portion of the second wall. The coating is in contact with a feature on the support structure.

Abstract: A nanocellular single crystal nickel based material is provided having a thermal diffusivity in the range of 0.0002 cm{circumflex over (?)}2/s to 0.02 cm{circumflex over (?)}2/s and a thermal conductivity in the range of 0.024 W/mK to 9.4 W/mK. The nanocellular single crystal nickel based material may be used to form turbine engine components. The nanocellular single crystal nickel based material may be produced by providing a first solution containing a nickel precursor and deionized water, providing a second solution containing a structure controlling polymer/surfactant and an alcohol, mixing the first and second solutions into a solution containing a reducing agent to form a third solution, and processing the third solution to create the nanocellular single crystal based material.

Abstract: A turbine shroud assembly adapted for use with a gas turbine engine includes a carrier, a seal segment, and a mount assembly. The carrier is configured to be coupled to a turbine case. The seal segment is shaped to define a gas path boundary of the shroud assembly. The mounting assembly is configured to couple the seal segment to the carrier.

Abstract: An axial compressor includes: a rotary shaft rotatably supported in a cylindrical casing such that an annular fluid passage is defined therebetween; a rotor blade row including rotor blades provided on an outer circumferential surface of the rotary shaft; a stator blade row including stator blades provided on an inner circumferential surface of the casing at a position adjacent to and behind the rotor blade row; and a recirculation passage provided in the casing and having a suction port and an ejection port on downstream and upstream sides of the fluid passage, respectively. The suction port is located rearward of leading edges of bases of the stator blades, and the ejection port is located at a position forward of centers of tips of the rotor blades and at least partially opposing the tips of the rotor blades.

Abstract: A turbine ring assembly including ring sectors forming a turbine ring and a ring support structure, each sector having, in a section plane defined by an axial direction and a radial direction of the ring, a first and a second attachment tabs extending in the radial direction, and the structure including a central shell ring from which extend as projections a first and a second radial flanges between which are retained the first and second attachment tabs of each sector. Each sector includes rectilinear seatings mounted on the faces of the first and second attachment tabs in contact, respectively, with the second annular flange and the annular ring-flange and including, along a tangent to the circumferential direction, a variable thickness in the axial direction with a minimum thickness at the first and second ends of the sector and a maximum thickness in a median portion of the rectilinear seating.

Abstract: A gas turbine engine includes a shroud with an abradable section and a non-abradable section that cooperatively define a shroud surface. The gas turbine engine also includes a rotor that is supported for rotation within the shroud to generate an aft axial fluid flow. The rotor includes a blade with a blade tip that is crowned and that opposes the abradable section and the non-abradable section of the shroud surface. A crown area of the blade tip opposes the abradable section. A casing treatment feature is provided in the non-abradable section of the shroud to oppose the blade tip of the rotor.