Abstract: The method for image segmentation includes: acquiring, according to an image to be segmented including a background, a mediastinum, an artery and a vein, a first segmentation result of the mediastinum, the artery, the vein and the background in a mediastinum region of the image to be segmented; acquiring, according to the image to be segmented, a second segmentation result of a blood vessel and the background in an epitaxial region of the image to be segmented; and acquiring, according to the first segmentation result and the second segmentation result, a segmentation result of the mediastinum, the artery, the vein and the background of the image to be segmented, so that the segmentation accuracy and the segmentation efficiency of the artery and the vein may be improved.

Abstract: The turbine (12) for a turbine engine (100) extends around a major axis (X-X) and comprises: —a casing (4) comprising an annular hook (38), —a movably mounted impeller (16), —a ring (30) extending opposite the impeller in a direction radial to the major axis (XX), —a distributor (14) comprising a blade provided with a platform (11), the platform being extended radially outwards by a spoiler (22, 24), the spoiler (22, 24) being radially mounted on the hook (38), and —foils (40) each having a profiled trough shape in a direction circumferential to the axis, the foils extending in succession in the circumferential direction, each foil (40) extending between the spoiler (22, 24) and the hook (38). The turbine comprises stops (162) to prevent the foils (40) from moving in the circumferential direction. Each foil (40) comprises a protrusion (162) arranged so as to extend circumferentially opposite another protrusion (162) of another adjacent foil (40).

Abstract: An assembly for a gas turbine engine includes a plurality of vanes having a radially outer platform with a flange that extends radially outward therefrom and a plurality of notches in the flange. A ring is located radially outward from the radially outer platform of the plurality of vanes. An axially extending projection extends through and corresponds with each of the plurality of notches on the flange.

Abstract: The present disclosure is directed to a composite component defining a component aperture extending between a first surface and a second surface. The composite component includes an insert having an insert annular wall positioned in the component aperture. The insert annular wall defines an insert aperture therethrough. An insert flange extends radially outwardly from the insert annular wall and contacts the first surface of the composite component. The insert flange includes a diameter about 1.5 times to about 5 times greater than a smallest diameter of the component aperture defined by the composite component.

Abstract: A stator vane assembly of a gas turbine engine includes a stator shroud having a plurality of shroud openings formed therein. The stator shroud defines a flowpath surface. The assembly includes a plurality of stator vanes, one stator vane of the plurality of stator vanes installed in each shroud opening of the stator shroud. A volume of potting material is located in the plurality of shroud openings to retain the stator vanes thereat. One or more retainers are located at a back surface of the stator shroud opposite the flowpath surface to restrict circumferential movement of the plurality of stator vanes relative to the stator shroud.

Abstract: A nozzle assembly for a gas turbine engine and methods for assembling a nozzle assembly are provided. In one example aspect, the nozzle assembly includes an outer wall and an inner wall radially spaced from the outer wall. The outer wall defines a plurality of mounting openings spaced circumferentially from one another. The inner wall defines a plurality of mounting openings spaced circumferentially from one another. The mounting openings defined by the inner wall are positioned circumferentially between adjacent mounting openings defined by the outer wall. The nozzle assembly includes vanes that are inserted through the mounting openings of the outer wall in a radially inward direction and vanes that are inserted through the mounting openings of the inner wall in a radially outward direction in an alternating manner.

Abstract: A vane for a gas turbine engine, the vane including a platform with an airfoil extending radially from the upper surface of the platform. The platform includes a joint portion which includes a circumferentially extending flange and a recessed surface both formed on either the upper or lower surface of the platform. The flange and the recessed surface extend from opposing circumferential edges of the joint portion and each include a substantially radially-extending through hole.

Abstract: An assembly for a gas turbine engine according to an example of the present disclosure includes at least one platform having a main body extending between a first mate face and a second mate face to establish a gas path surface. The main body has an internal passage extending circumferentially between a first opening along the first mate face and a second opening along the second mate face relative to an assembly axis. A perimeter of the first mate face establishes a first area, a perimeter of the second opening establishes a second area, and the second area is greater than the first area. An airfoil section extends radially from the at least one platform.

Abstract: A component for a gas turbine engine according to an example of the present disclosure includes, among other things, a body including a cold side surface adjacent to a mate face. A plurality of ridges extends from the cold side surface. A seal member abuts the plurality of ridges to define a plurality of cooling passages. The seal member is configured to move between a first position and a second position relative to the plurality of ridges. Each of the plurality of cooling passages includes a first inlet defined at the first position and a second, different inlet defined at the second position. A method of sealing between adjacent components of a gas turbine engine is also disclosed.

Abstract: An apparatus configured to support a diffuser segment during removal from and installation into an assembly. The apparatus includes a body having two or more supports and at least two clamp rods. The clamp rods each extend from a first end to a second end. The body defines a pocket extending between the supports on a first side of the clamp rods. An end clamp releasably mates with the second end of each clamp rod. The end clamp has a foot extending into the first side of the clamp rod. A wheel is received within each support. The pocket and the foot of each end clamp releasably retain a diffuser segment and the wheels rotate to allow the diffuser segment to be removed without damaging a surface coating of the diffuser segment.

Abstract: A turbine exhaust case (TEC) has an outer case and an inner case structurally interconnected by a plurality of circumferentially spaced-apart struts. At least one of the struts has an airfoil body with a hollow core. The airfoil body has opposed pressure and suction sides extending chordwise from a leading edge to a trailing edge and spanwise from a radially inner end to a radially outer end. The radially inner end of the strut has a strut wall extension that extends through the inner case to a location radially inward of the inner case relative to the central axis.

Abstract: An assembly for a turbomachine turbine includes at least one ring sector made of CMC material and a support casing including an upstream flange and a downstream flange between which each ring sector is disposed, each ring sector including a base that has a radially external face from which radially extend two lugs, the lugs of each ring sector being retained between the two flanges of the support casing by axial pins each engaged in one of the flanges of the ring support casing and in the lug of the ring sector facing said flange, the assembly further including, for each ring sector, at least one radial retaining pin screwed into the support casing and coming to radially bear against a lug of the at least one ring sector to retain it in position, and anti-rotation system for rotationally locking the radial retaining pin.

Abstract: A turbine nozzle for a gas turbine engine includes an outer endwall, and an inner endwall spaced apart from the outer endwall. The turbine nozzle includes at least one airfoil coupled between the inner endwall and the outer endwall, and a compliant joint defined in at least one of the outer endwall and the inner endwall. The compliant joint includes at least one partially-fused seal that is configured to restrict a flow of fluid through the at least one of the outer endwall and the inner endwall and to form a slip-joint above a predetermined threshold stress. The slip-joint is configured to maintain a radial position of the at least one of the outer endwall and the inner endwall.

Abstract: A distributor sector of a turbine of a turbomachine, including an outer platform including a downstream cylindrical hook adapted to engage with the turbomachine casing, the hook including a part in which is formed a U-shaped anti-rotation notch including a bottom and two lateral walls extending from the bottom, the notch being at least partially covered by a wear insert inserted in the notch, the insert being removeable.

Abstract: In one aspect, centrifuge feed pipes are described herein having design and architecture for mitigating wear during centrifuging operations. In some embodiments, a centrifuge feed pipe comprises a feed material inlet, a feed material outlet, and a conduit body extending along an axis between the feed material inlet and outlet, the conduit body having a variable inner diameter at one or more locations along the axis. In being variable, the inner diameter is not uniform over all radial positions of the inner diameter.

Abstract: A stator assembly for use in a gas turbine engine is provided. The stator assembly including: a conical stator shroud including a radially inward surface and a radially outward surface opposite the radially inward surface; and a plurality of stator vanes integrally attached to the conical stator shroud, each of the plurality of stator vanes being integrally attached to the conical stator shroud at a base of the stator vane, wherein radially outward surface of the base of the stator vane mates flush with the radially inward surface of the conical stator shroud.

Abstract: A nozzle assembly for a gas turbine engine and methods for assembling a nozzle assembly are provided. In one example aspect, the nozzle assembly includes an outer wall and an inner wall radially spaced from the outer wall. The outer wall defines a plurality of mounting openings spaced circumferentially from one another. The inner wall defines a plurality of mounting openings spaced circumferentially from one another. The mounting openings defined by the inner wall are positioned circumferentially between adjacent mounting openings defined by the outer wall. The nozzle assembly includes vanes that are inserted through the mounting openings of the outer wall in a radially inward direction and vanes that are inserted through the mounting openings of the inner wall in a radially outward direction in an alternating manner.

Abstract: A retention device includes a retention block including an opening extending through the retention block in a first direction. A tab also extends from the retention block in the first direction. A fastener extends through the opening and includes a head configured to engage the retention block. A clinch nut is coupled to the fastener with an interface surface of the clinch nut configured to engage the tab of the retention block.

Abstract: A turbine engine turbine including a nozzle stage made of ceramic matrix composite material and including a plurality of annular sectors forming an annulus presenting an inner shroud and an outer shroud, each sector having an inner platform forming a portion of the inner shroud, an outer platform forming a portion of the outer shroud, and at least one airfoil extending between the outer and inner platforms and secured thereto. A metal ring includes at least one annular sector, and presents an outer surface in contact with the surface of the inner shroud opposite from the surface from which the airfoils extend, the metal ring presenting an outside diameter at its outer surface that is greater than the diameter of the inner shroud such that the nozzle stage is held in compression between a casing and the metal ring.

Abstract: A pre-diffuser for a gas turbine engine includes an exit guide vane ring having a multiple of exit guide vanes defined around an engine longitudinal axis; a hot fairing structure adjacent to the exit guide vane ring to define a multiple of diffusion passages around the engine longitudinal axis; an outer radial interface between a radial outer surface of the hot fairing structure and the exit guide vane ring, the outer radial interface being a full hoop structure; and an anti-rotation feature between the hot fairing structure and the exit guide vane ring, the anti-rotation features inboard of the multiple of diffusion passages.

Abstract: A nozzle assembly is provided which is, in part, formed of a low coefficient of thermal expansion material. The assembly includes a nozzle fairing formed of the low coefficient of thermal expansion material and includes a metallic strut extending radially through the nozzle fairing. Load is transferred from the nozzle fairing to a static structure in either of two ways: first, the strut may receive the load directly and/or second, load may be transferred from the nozzle fairing to at least one of the inner and outer support rings. Further, the nozzle fairing and strut may allow for internal airflow for cooling.

Abstract: A method of replacing a module in a modular gas turbine engine having a first fan module; a first propulsor module including an engine core and a gearbox; and a first fan case module having a fan case; includes the steps of: disassembling the gas turbine engine, replacing one of the fan module, propulsor module or fan case module with a replacement fan module, a replacement propulsor module or a replacement fan case module, the replacement module, having the same configuration as the first module; and reassembling the gas turbine engine using the replacement module.

Abstract: A case assembly for a turbine section has a structural outer case structurally connected to a gas generator upstream of the outer case and to an exhaust case downstream of the outer case. The assembly further comprises a containment ring mounted within the outer case and surrounding a plurality of axially spaced-apart turbine stages. An inner surface of the containment ring defines a plurality of shroud receiving portions. An annular plenum is defined between the outer case and the containment ring. The plenum has an inlet connected to a source of pressurized coolant. Outlets are circumferentially and axially distributed and defined through the containment ring. The outlets provides flow communication between the annular plenum and the plurality of axially spaced-apart turbine stages.

Abstract: A nozzle segment for a gas turbine engine includes an arcuate outer vane platform segment. An arcuate inner vane platform segment spaced from the arcuate outer vane platform segment. A multiple of airfoils between the arcuate inner vane platform segment and the arcuate outer vane platform segment, the arcuate outer vane platform segment includes a scallop slot and a seal that seals the scallop slot.

Abstract: A vane assembly adapted to be disposed through an annular gas flow path defined between a fan case and an engine core of a gas turbine engine. The assembly comprises a vane having a vane body configured for extending between the engine core and the fan case, a vane head disposed at one end of the vane body, the vane head being adapted to be disposed outside the annular gas flow path, the vane head having an abutting surface configured for contacting an outer surface of the fan case, and a recess extending within the vane head and opening to the abutting surface. The recess is configured for circumferentially surrounding a longitudinal axis of the vane and forming a closed figure. The assembly also comprises a sealing member disposed within the recess.

Abstract: Turbine vane assemblies incorporating both metallic and ceramic matrix composite materials are provided in the present disclosure. The turbine vane assemblies further include interface components that allow for differing rates of thermal expansion in the ceramic matrix composite components and the metallic components.

Abstract: A fan exit stator assembly of a gas turbine engine may include a radially inward shroud defining a first slot, a radially outward shroud, and a vane. The vane may include a base portion extending through the first slot and a tip portion coupled to the radially outward shroud. The vane may also include a retainer plate disposed radially inward of the radially inward shroud that projects in a circumferential direction from the base portion of the vane. A circumferential plate dimension of the retainer plate in the circumferential direction is greater than a first slot radius in the circumferential direction of the first slot, thus preventing radially outward movement of the vane relative to the radially inward shroud.

Abstract: A stator assembly of a gas turbine engine according to an example of the present disclosure includes, among other things, a first shroud extending about an assembly axis to bound a flow path. The first shroud defines a first shroud opening. An airfoil has an airfoil body extending from a first end portion. The first end portion is received in the first shroud opening and defines a pair of airfoil openings. A retention clip has an intermediate portion connecting a first leg portion and a second leg portion. The intermediate portion is compressible to position the first and second leg portions in the pair of airfoil openings such that the retention clip limits movement of the airfoil relative to the first shroud.

Abstract: Flow path assemblies and methods for assembling such assemblies are provided. For example, a flow path assembly comprises a turbine nozzle segment including an airfoil extending axially between leading and trailing edges, an inner band defining a portion of a flow path inner boundary, and an outer band defining a portion of a flow path outer boundary. The airfoil includes a trailing edge portion defining the trailing edge and extending axially beyond the inner and outer bands such that the trailing edge is defined axially aft of the inner and outer band aft ends. The flow path assembly further comprises a shroud having a forward portion that extends axially along the trailing edge portion such that the forward portion defines the flow path outer boundary at the trailing edge portion. Methods of assembling flow path assemblies having a turbine nozzle assembly and inner and outer members also are provided.

Abstract: A stator assembly of a gas turbine engine according to an example of the present disclosure includes, among other things, a first shroud that extends about an axis to bound a flow path. The first shroud defines a first shroud opening. An airfoil has an airfoil body that extends from a first end portion. The first end portion is received in the first shroud opening and defines a pair of airfoil openings. At least one retention clip has an intermediate portion connecting a pair of elongated leg portions. The pair of elongated leg portions are received in the pair of airfoil openings such that the at least one retention clip limits movement of the airfoil relative to the first shroud. A method of assembling a stator assembly is also disclosed.

Abstract: When a combination of a vane type of a stator vane and a vane type of a stator vane for one straightening plate is the same as a combination of a vane type of a stator vane and a vane type of a stator vane for the other straightening plate, positions of the first straightening plate side connection portions and the second straightening plate side connection portions of the one straightening plate and the other straightening plate are the same as each other. When the combinations are different from each other, at least one of the positions of the first straightening plate side connection portions of the one straightening plate and the other straightening plate and the positions of the second straightening plate side connection portions of the one straightening plate and the other straightening plate are different from each other.

Abstract: A stator assembly for a gas turbine engine includes an arcuate outer shroud, an arcuate inner shroud radially spaced from the outer shroud and a plurality of stator vanes extending from the outer shroud to the inner shroud. A volume of potting is located at the inner shroud and at the outer shroud to retain the plurality of stator vanes thereat. A stator and case assembly includes a case defining a working fluid flowpath and a stator assembly positioned at the case. The stator assembly includes a plurality of stator segments arranged circumferentially about an engine axis, each stator segment including an arcuate outer shroud secured to the case, an arcuate inner shroud, and a plurality of stator vanes extending from the outer to inner shroud. A volume of potting is located at the inner shroud and at the outer shroud to retain the plurality of stator vanes thereat.

Abstract: Composite airfoil singlet, includes airfoil extending from base to tip of airfoil, integrally formed with no more than one outer platform at tip and/or no more than one inner platform at base. Parallel composite plies or woven fibers extend through airfoil and through outer and/or inner platforms. Outer and/or inner curved sections extend between outer and/or inner platforms and airfoil respectively. Assembly includes circular row of the composite airfoil singlets depending radially inwardly from and mounted to an outer shroud or casing. Outer and/or inner fasteners may secure outer and inner platforms to outer shroud or casing and an inner shroud respectively and include shanks extending substantially perpendicularly from outer and inner fastening plates though platform holes in outer and inner platforms and through outer and inner holes in outer shroud or casing and inner shroud respectively. Nuts are screwed on threaded ends of shanks.

Abstract: A shroud assembly of a turbine system is presented. The shroud assembly includes a plurality of outer shroud segments disposed annularly and defining a first gap between a pair of outer shroud segments of the plurality of outer shroud segments; a plurality of inner shroud segments disposed annularly and radially inward of the plurality of outer shroud segments, and defining a second gap between a pair of inner shroud segments of the plurality of inner shroud segments, wherein the plurality of inner shroud segments is same in number to the plurality of outer shroud segments, and wherein the first gap and the second gap are offset; and an impingement plate coupled to an outer shroud segment of the pair of outer shroud segments and interposed between the outer shroud segment and the pair of inner shroud segments.

Abstract: The invention relates to a method for manufacturing a turbine engine casing (1), characterized in that said method includes the steps of: manufacturing (E1) a plurality of sectors (2), at least one portion of the sectors (2) being manufactured by casting and including, on the surface thereof, fastening elements (3) produced during the casting step, assembly bands (8) being produced at the ends of the sectors (2) during the step of manufacturing the sectors (2) by casting, by means of which the sectors (2) can be assembled; and assembling (E2) the sectors (2) end-to-end such as to form a ring (5) of the casing (1). The invention also relates to a turbine engine casing (1).

Abstract: A nozzle guide vane for a gas turbine engine is disclosed herein. The nozzle guide vane includes an inner endcap, an outer endcap, and at least one airfoil that extends from the inner endcap to the outer endcap. The nozzle guide vane further includes at least one composite heat shield component adapted to shield metallic components from high temperature gasses.

Abstract: A pump intake device comprising a main body which includes a side wall section having an inner side and an outer side, an intake section extending from the outer side of the side wall section and an intake passage extending through the intake section, the intake passage having an inner surface and an entry end and an exit end with a central axis extending between the entry and exit ends, a first portion of the inner surface having one or more first guides thereon for directing fluid passing through the intake passage so that in use said fluid leaves the exit end at the first portion with an exit angle which is inclined relative to the central axis.

Abstract: A turbine exhaust case has an outer housing to be secured within a gas turbine engine and a central hub. Struts extend between the outer housing and the central hub. The struts are formed at least in part of a first material. The central hub is formed at least in part of a second material.

Abstract: A turbine nozzle box includes an annular outer ring plate, an annular nozzle unit, and an annular inner ring plate. The annular outer ring plate is operable to be disposed in a turbine such that an outer peripheral surface thereof is adjacent to an inner peripheral surface of an inner casing of the turbine. The annular nozzle unit is disposed adjacent to an inner peripheral surface of the outer ring plate. The annular inner ring plate is disposed such that an outer peripheral surface thereof is adjacent to an inner peripheral surface of the nozzle unit. The nozzle unit includes a plurality of segment blocks circumferentially installed at an arc angle. The segment blocks are disposed at an inlet end of a first-stage nozzle unit, and an outlet end of the first-stage nozzle unit is open.

Abstract: A fan salt-fog-resistant structure and a fan frame thereof. The fan salt-fog-resistant structure includes a base seat having a bearing cup and at least one flow guide body. Multiple connection bodies outward extend from the base seat. The flow guide body is disposed at a junction between the base seat and the connection bodies. The fan frame includes a main body having a first open side and a second open side and a flow passage defined therebetween. One end of the connection body is connected with the second open side. When airflow entraining moisture and salt fog flows through the junction between the base seat and the connection bodies, the flow guide body serves to guide the airflow and prevent the airflow from forming eddy, whereby the salt fog entrained by the airflow will not crystallize to form sediments at the junction between the base seat and the connection bodies.

Abstract: Various embodiments of the invention include turbine nozzles and systems employing such nozzles. Various particular embodiments include a turbine nozzle having: an airfoil having: a suction side; a pressure side opposing the suction side; a leading edge spanning between the pressure side and the suction side; and a trailing edge opposing the leading edge and spanning between the pressure side and the suction side; and at least one endwall connected with the airfoil along the suction side, pressure side, trailing edge and the leading edge, the at least one endwall including an axisymmetric contour.

Abstract: A rotor assembly for use in a turbofan engine is provided. The rotor assembly includes an annular spool including a first blade opening defined therein, a first rotor blade configured to be radially inserted through the first blade opening, and a fairing positioned on a radially outer side of the annular spool. The first rotor blade includes a blade portion and a flange portion that extends substantially perpendicularly relative to the blade portion such that the flange portion is positioned on a radially inner side of the annular spool. The fairing is configured to receive a fastener radially inserted through the flange portion and the annular spool such that the first rotor blade is secured to the annular spool.

Abstract: A component according to an exemplary aspect of the present disclosure includes, among other things, a shell defining an interior, a spar extending into the interior and a first flange attached to the spar. The spar is configured to pivot to change a positioning of the shell.

Abstract: A gas turbine having a guide vane assembly (2) and a securing ring (1) for axially locking the guide vane assembly in position that has a radially outer rim (1A) which is configured in a housing-side groove (3.1), a radially inner rim (1B) which is configured outside of the groove, and a slot (10) which extends from the radially outer rim to the radially inner rim, and a first flank having a radially inner rim section (12) and an undercut portion (11), the radially inner rim section forming an angle (?) of at least 50° with the radially inner rim.

Abstract: An airfoil for a gas turbine engine includes a first airfoil. A first chordal seal is located adjacent a first end of the airfoil. A second chordal seal is located adjacent a second end of the airfoil. The first chordal seal includes a first edge parallel to a first edge on the second chordal seal.

Abstract: An inlet duct, an induction system, and a system are disclosed for directing an inlet flow into an inlet compressor for use in an internal combustion engine. An example inlet duct may include one or more relief features disposed on an inner surface of the inlet duct. The one or more relief features may be made integral with the inlet duct. The one or more relief features may be disposed to protrude into the inlet flow to cause the inlet flow to swirl before reaching the inlet compressor.

Abstract: An assembly for a gas turbine engine, comprises a fairing assembly extending generally axially through a portion of the gas turbine engine, a first component disposed radially inward of the fairing assembly, a second component disposed axially adjacent to the first component, a flow diverter, and a finger seal. The flow diverter includes a mounting portion securing the flow diverter to the first component, and an arm extending in an axial direction from the mounting portion toward the second component. The finger seal includes a fixed end secured between the flow diverter mounting portion and the first component.

Abstract: A turbine component assembly for a gas turbine engine includes: a first component having a first coefficient of thermal expansion and including an end face; a second component including a mating surface abutting the end face; and a fastener having a second coefficient of thermal expansion different from the first coefficient of thermal expansion, the fastener including a shank engaging the second component and an enlarged head engaging a mounting slot in the first component; wherein the mating surface and the end face shaped to permit relative pivoting movement between the first and second components.

Abstract: The present invention relates to a system and method for joining adjacent sections of a compressor diaphragm. A splitline jumper is positioned within a recessed portion of a seal ring segment of the adjacent compressor diaphragms. The splitline jumper is sized and oriented so as to prevent axial movement of the compressor diaphragms relative to each other and also to prevent rotational movement of the seal ring relative to the diaphragm.

Abstract: An article of manufacture having a nominal airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in a scalable TABLE 1, wherein the Cartesian coordinate values of X, Y, and Z are non-dimensional values convertible to dimensional distances by multiplying the Cartesian coordinate values of X, Y, and Z by a number, and wherein X and Y are coordinates which, when connected by continuing arcs, define airfoil profile sections at each Z height, the airfoil profile sections at each Z height being joined with one another to form a complete airfoil shape.