Abstract: A stator vane for a gas turbine engine combustor has a vane segment that includes a vane having a leading edge, a trailing edge, a suction side, a pressure side, a vane inner base, and a vane outer base, an inner shroud segment attached to the vane at the vane inner base, and an outer shroud segment attached to the vane at the vane outer base. The stator vane further includes a support structure attached to the vane at a vane trailing edge, and a first weld shield attached to the support structure by a first connector. The first weld shield is positioned over and spaced away from the suction side. The vane, inner shroud segment, outer shroud segment, support structure, and first weld shield are integrally formed during a single, continuous additive manufacturing process.

Abstract: An assembly including a turbine engine casing, in particular for an aircraft, the turbine engine casing including an annular or cylindrical radially outer shell, provided with a radially inner surface, and an aerody-namic treatment member including a plurality of grooves distributed in a circumferential direction. The radially outer shell includes a circumferential recess in which the aerodynamic treatment member is removably mounted, the treatment member extending along at least one angular sector and having a shape that matches that of the circumferential recess, the treatment member including a circumferential opening into which the plurality of grooves open.

Abstract: A gas turbine nozzle guide vane structure includes a vane shaped as an airfoil and having a vane trailing edge, an endwall including an opening to receive an end of the vane, and an element securing the endwall and the vane to each other. Clearance remaining between the endwall and the vane defines a plenum to feed cooling air to the vane at a location adjacent the vane trailing edge. Certain arrangements may have a purge groove defined in at least one of the endwall and the vane and located between the endwall and the vane to receive cooling fluid supplied. In certain arrangements, the structure may include a cover sheet on the endwall defining a gap with the vane, the purge groove configured to receive cooling fluid that exits through the gap.

Abstract: A bypass turbomachine is disclosed. The turbomachine includes a longitudinal axis having an upstream blower impeller and a downstream air-flow straightening assembly of a secondary annular passage delimited radially on the inside by a radially inner shroud and radially on the outside by a radially outer shroud, blades extending between the radially inner and outer shrouds and being attached at a first end portion to the radially inner shroud and at a second end portion to the radially outer shroud, the blades including a useful portion extending between the first and second end portions and defining a lower face and an upper face. For each blade, in a plane perpendicular to the longitudinal axis, each aerodynamically useful portion is bent into a C-shape in the circumferential direction.

Abstract: A gas turbine is provided, the gas turbine engine including a turbomachine having an inlet splitter defining in part an inlet to a working gas flowpath and a fan duct splitter defining in part an inlet to a fan duct flowpath. The gas turbine engine also includes a primary fan driven by the turbomachine defining a primary fan tip radius R1, a primary fan hub radius R2, and a primary fan specific thrust rating TP; and a secondary fan downstream of the primary fan and driven by the turbomachine, the secondary fan defining a secondary fan tip radius R3, a secondary fan hub radius R4, and a secondary fan specific thrust rating TS; wherein the gas turbine engine defines an Effective Bypass Area, and wherein a ratio of R1 to R3 equals R 1 R 3 = ( EFP ) ? ( 1 - RqR ( Sec . - Fan ) 2 ) ( 1 - RqR Prim . - Fan 2 ) ? ( T P T S ) ? ( EBA ) .

Abstract: A turbine housing can include a bearing housing end and a treatment unit end; a volute wall that defines a volute; a wall that defines at least a portion of a turbine wheel space that defines a turbine wheel space axis and a turbine wheel space diameter, where the wall extends to an axial peak to define an extended space with an extended space outlet having an extended space outlet dimension; and an outlet wall that defines an outlet space with a treatment unit end outlet having an outlet dimension, where the extended space is disposed at least in part axially between the turbine wheel space and the outlet space to increase axial velocity uniformity at the treatment unit end outlet.

Abstract: There is provided an assembly for a turbocharger, the assembly comprising a turbine housing and a nozzle ring, wherein one of the turbine housing and the nozzle ring comprises a tongue and the other of the turbine housing and the nozzle ring comprises a corresponding slot, the tongue and slot being configured to engage and prevent relative rotation of the turbine housing and the nozzle ring. Also provided is a method of manufacturing a turbine housing, assembly for use in a turbocharger, or a turbocharger.

Abstract: Turbine comprising a casing and a nozzle including a metal outer shroud integral with the casing, a metal inner shroud, and a plurality of nozzle sectors of CMC forming a ring extending between the metal outer shroud and the metal inner shroud, each sector including a mast, an inner platform, an outer platform and at least one airfoil having a hollow profile defining an inner recess, each of the inner and outer platforms having an opening communicating with said inner recess, and the mast passing through said openings and the inner recess and being attached to said casing and connected with said nozzle sector. The mast comprises at least one fastening projection having at least one portion extending from a radially outer end of the mast in a direction opposite to the center of the mast and in a plane orthogonal to the radial direction.

Abstract: A compressor, for use alone or as a component part of a turbocharger. The compressor comprises a compressor housing and a compressor wheel mounted in the compressor housing. An internal surface of the compressor housing defines an axial inlet. The axial inlet comprises an annular groove for reducing compressor noise.

Abstract: A vane system includes a plurality of vane assemblies and a continuous support ring. The vane assemblies are arranged circumferentially about an axis and each include a hollow airfoil fairing and a spar. The spar has a spar flange and a spar leg extending radially inwardly from the spar flange and through the hollow airfoil fairing. The continuous support ring has radially inner and outer sides and defines a circumferential row of through-holes between the radially inner and outer sides. The spar legs extend through the through-holes and the spar flanges are affixed at the radially outer side of the continuous support ring.

Abstract: The present application provides a journal bearing assembly for use with a rotor shaft of a turbine engine. The journal bearing assembly may include an outer shell, an internal oil feed port on a first side of the outer shell, an external oil feed port on a second side of the outer shell, and a connecting bore in communication with the internal oil feed port and the external oil feed port.

Abstract: A fan blade (1) has a front inflow edge (2) and a rear outflow edge (3). The fan blade (1) has an at least partially wavy inflow edge (4) that forms a wave (W) having a specific three-dimensional waveform.

Abstract: A method of making a fluid injection component for a gas turbine engine includes depositing material onto a piece of tube stock. The method includes machining an elbow into the deposited material, wherein machining the elbow includes forming a braze joint surface in the deposited material. Depositing can include laser cladding the material onto the piece of tube stock.

Abstract: A nozzle for a turbomachine turbine. The nozzle includes blades made of ceramic matrix composite material, and at least one metal liner passing through a respective blade. The liner is connected in a sealed manner to a metal internal shroud so as to guide a ventilation fluid through the blade. The nozzle can be configured to maintain effective sealing in spite of the different thermal expansions of the blade and of the internal shroud.

Abstract: Systems and methods for attaching a protective liner to a an aircraft engine component having a different coefficient of thermal expansion in comparison to the protective liner include a mounting member configured to be mounted to the aircraft engine component and an attachment member configured to be attached to the protective liner. The mounting member and the attachment member are coupled to each other such that a portion of the mounting member is received within a portion of the attachment member and such that the mounting member is permitted to move with both radial and rotational freedom relative to the attachment member.

Abstract: A device is suitable for inserting sealing pads into one of the sectors of a turbine stator of a turbine, wherein the turbine stator includes a plurality of sectors with adjacent side faces abutting each other and slots arranged opposite each other in the adjacent side faces of two successive sectors. Each stator is configured to receive a predefined sealing pad. The device includes a support for supporting the sector; a robot arm configured to grip the sealing pads, each predefined in accordance with the slot intended to receive it. The robot arm is configured to insert each predefined sealing pad into the slot intended to receive it of one of the side faces of the sector.

Abstract: Airfoil assemblies for gas turbine engines include an airfoil body and a leading edge baffle installed within a leading edge cavity of the airfoil body. The airfoil body includes a plurality of radially extending rails configured to engage with the baffle and define radially extending channels therebetween. First and second forward radially extending rails are segmented in the radial direction and a showerhead radial channel is defined along the leading edge. Pressure and suction side radial flow channels are defined between an interior surface of the airfoil body, an exterior surface of the baffle, and radially extending rails. An aft channel is defined between an interior surface of the airfoil body along pressure and suction side walls, an interior rib, exterior surfaces of the baffle, and the radially extending rails.

Abstract: An engine section stator for a gas turbine engine having a compressor, a combustor, and a turbine. The engine section stator includes an inner band, an outer band spaced radially outwardly from the inner band, and a series of spaced apart aerofoils extending the inner and outer bands. The engine section stator includes a stiffness feature that extends away from one of the inner and outer bands of the engine section stator. The stiffness feature configured to increase the high cycle fatigue strength of the aerofoils without impeding airflow passing between the inner and outer bands.

Abstract: A vane system includes a plurality of vane assemblies and a continuous support ring. The vane assemblies are arranged circumferentially about an axis and each include a hollow airfoil fairing and a spar. The spar has a spar flange and a spar leg extending radially inwardly from the spar flange and through the hollow airfoil fairing. The continuous support ring has radially inner and outer sides and defines a circumferential row of through-holes between the radially inner and outer sides. The spar legs extend through the through-holes and the spar flanges are affixed at the radially outer side of the continuous support ring.

Abstract: Systems and methods for identifying and mitigating gas turbine component misalignment using virtual simulation are disclosed herein. An example method may include capturing data associated with a first nozzle segment and a second nozzle segment of a gas turbine. The method may also include creating, based on the captured data, a virtual representation of the first nozzle segment and the second nozzle segment. The method may also include determining that a misalignment exists in a connection between the virtual representation first nozzle segment and the virtual representation of the second nozzle segment. The method may also include identifying, based on the determination that the misalignment exists, a third nozzle segment. The method may also include determining that a connection between a third nozzle segment and the first nozzle segment includes a smaller misalignment.

Abstract: A turbine stator blade made of ceramic matrix composite material includes a hollow blade profile and has a trailing edge and a leading edge, the blade including a first portion including an extrados face and a second portion distinct from the first portion including an intrados face, the first and second portions being connected to one another by a connection interface present at least on the trailing edge or leading edge, the connecting interface including a region of overlap between the first and second portions present on at least one longitudinal end of the blade profile and intended to be present outside a flow path of a gas stream of the turbine, the blade also including a platform present at a longitudinal end of the blade profile and that includes a first portion integral with the extrados face and a second part integral with the intrados face.

Abstract: A method for inhibiting particulate ingress in a combustion system includes acquiring particulate ingress data, determining a probable particulate ingress location on an air inlet of the combustion system in dependence upon the particulate ingress data, and deploying a filter screen at the determined probable particulate ingress location to inhibit particulate ingress at the determined probable particulate ingress location.

Abstract: A plurality of cooling passages extending in an axial direction are formed in a transition piece so as to be aligned in a circumferential direction and the axial direction. One or more downstream side passages are formed in a downstream side region (Rd) within one circumferential region. One or more upstream side passages are formed in an upstream side region Ru within the circumferential region. The total cross-sectional area per unit circumferential length of the one or more downstream side passages is larger than the total cross-sectional area per unit circumferential length of the one or more upstream side passages.

Abstract: A vane assembly includes a vane that includes an outer platform, an inner platform, and an airfoil. The outer platform defines an outer boundary of a gas path. The inner platform is spaced apart radially from the outer platform relative to an axis and defines an inner boundary of the gas path. The airfoil extends radially between and interconnects the outer platform and the inner platform.

Abstract: A sealing assembly includes a plurality of sealing members configured to be located circumferentially adjacent one another between adjacent rotating blade stages in a turbine. The sealing members include axially extending sealing portions having spline seal slots defined in opposing slash faces thereof. The slots are structured to allow insertion of a spline seal into facing slots on adjacent sealing members when the sealing members are in an axially offset position, and to retain the spline seal in the facing slots when the sealing members are moved to an operative axial position. The slots may include spline seal-retaining seats therein to properly position and retain the spline seal. The sealing members allow easy assembly of the sealing assembly with proper positioning of the spline seal, and without damaging the spline seal or adjacent structure. A method assembling the sealing assembly is also disclosed.

Abstract: A turbine vane assembly adapted for use in a gas turbine engine includes a plurality of turbine vanes, an outer vane support, and an inner vane support. The plurality of turbine vanes comprise ceramic matrix composite material and are adapted to interact with hot gases flowing through a gas path of the gas turbine engine during use of the turbine vane assembly.

Abstract: Shrouds and shroud segments for gas turbine engines are provided. In one embodiment, a shroud segment for a gas turbine engine having a rotor blade stage and a nozzle stage is provided. The shroud segment comprises a forward end defining an outer wall of the rotor blade stage and an aft end defining an outer wall of the nozzle stage. The aft end defines at least a portion of an opening therethrough for receipt of a nozzle, and the forward end and the aft end form a single, continuous component. In another embodiment, a gas turbine engine is provided, having a shroud with a forward end positioned near a leading edge of a plurality of rotor blades of a rotor blade stage and an aft end positioned near a trailing edge of a plurality of nozzles of a nozzle stage. Methods of assembling a gas turbine engine also are provided.

Abstract: A method of manufacturing a variable vane mechanism includes arranging a vane and a spacer between a first support structure and a tool member of a tool. The method also includes abutting a first inner surface of the first support structure against a first side surface of the vane and a second side surface of the vane against an opposing surface of the tool member, leaving a control surface of the spacer projecting from the first support structure at a predetermined distance. The method further includes fixedly attaching the spacer to the first support structure with the control surface of the spacer projecting from the first support structure at the predetermined distance. The method further includes abutting a second support structure on the control surface of the spacer to define a gap between the first support structure and the second support structure with the vane disposed within the gap.

Abstract: An inlet guide vane assembly of gas turbine is disclosed. The inlet guide vane assembly includes: a plurality of inlet guide vanes, each including a fixing shaft disposed at the end away from a fixing part; and a fixing assembly including a support ring, a fixing ring and an abutting member. The fixing ring includes a fixing wall facing towards a support part, the fixing wall forming, in cooperation with a support wall, a plurality of accommodating cavities corresponding to the plurality of inlet guide vanes. The fixing shafts are rotatably inserted into the accommodating cavities, and the abutting member is in interference fit with the inner inlet casing and the fixing ring. The inlet guide vane assembly implements support for the inlet guide vanes without affecting installation of the inlet guide vanes.

Abstract: An airfoil assembly may include a shroud and an airfoil. The shroud may include a first attachment arm, a second attachment arm, and a shroud rail extending from a first surface of the shroud. A first channel may be defined between the first attachment arm, the first surface, and the shroud rail and a second channel may be defined between the second attachment arm, the first surface, and the shroud rail. The airfoil may extend from a second surface of the shroud opposite the first surface. In various embodiments, a height of the shroud rail, as measured from the first surface of the shroud, is non-uniform.

Abstract: A method for manufacturing a cascade for a thrust reverser for a jet engine includes forming a strip assembly. The strip assembly includes a strong back member which includes a length. The strip assembly includes a plurality of first vane members which extend from a first side of the strong back member in a first direction nonparallel relative to the length of strong back member wherein the plurality of the first vane members are spaced apart from one another along the length of the strong back member.

Abstract: An annular array of turning vanes 200 is provided in a duct 100 of a gas turbine engine 10. The annular array of turning vanes 200 comprises aerodynamic vanes 220 and strut-vanes 240. The strut-vanes 240 have greater chord length and extend further axially downstream than the aerodynamic vanes 220. The leading edge of the strut-vanes 240 is upstream of the trailing edge of the aerodynamic vanes 220. The strut-vanes provide flow turning. The arrangement allows the duct 100 to be axially short.

Abstract: An aerofoil member that extends between radially inner and radially outer endwalls which define a gas annulus of the compressor. Aerofoil member has a leading edge, a trailing edge, a pressure surface and a suction surface such that successive cross-sections through the aerofoil member transverse to the radial direction provide respective aerofoil sections. The external shape of the aerofoil member is defined by the stacking of the aerofoil sections on a stacking axis which passes through a reference point common to each aerofoil section. The aerofoil member has lean produced by the projection of the stacking axis onto a plane normal to the engine axis intersecting a first one of the endwalls at an angle of from 5° to 25° to the circumferential direction direction such that the pressure surface faces the first endwall.

Abstract: A fuel injector for a gas turbine combustor includes side wall fuel injection bodies extending between opposite end walls. Each side wall fuel injection body defines a fuel plenum and includes an outer surface defining fuel injection ports in communication with the fuel plenum. One or more fuel injection bodies extending between the end walls are positioned between the side wall fuel injection bodies. Each fuel injection body defines a fuel plenum and includes an outer surface defining fuel injection ports in fluid communication with the fuel plenum. A conduit fitting coupled to the frame is fluidly connected to the respective fuel plenums. The fuel injection ports fluidly communicate with air flow paths defined between the fuel injection bodies and the side wall fuel injection bodies. A combustor for a gas turbine includes a liner and an axial fuel staging system with the present fuel injector.

Abstract: A boost control device for a turbocharger, assembled at room temperature, may not be geometrically optimal at turbocharger operating temperature. The geometric deviation in the boost control device introduced by the heating of the turbocharger to operating temperature, or during welding of the boost control device, is corrected by producing a first turbocharger at room temperature, determining the geometry of the deviation at operating temperature, producing a distance element that will introduce an equal and opposite deviation, introducing the equal and opposite deviation via the distance element in second and subsequent turbochargers, whereby thermal deviation and corrective deviation offset so that the turbocharger is geometrically accurate at operating temperature.

Abstract: A turbo fan engine includes a housing that supports a compressor section. At least one compressor hub is within the compressor section. A compressor shaft is arranged in the compressor section. An epicyclic gear train is coupled to the compressor shaft. At least one compressor hub and the compressor shaft are coupled at a common attachment point. The epicyclic gear train includes a carrier. A sun gear and intermediate gears are arranged about and intermeshing with the sun gear. The intermediate gears are supported by the carrier. A baffle is supported relative to the carrier and includes a lubrication passage near at least one of the sun gear and intermediate gears for directing a lubricant on at least one of the sun gear and the intermediate gears.

Abstract: An assembly for a turbine engine includes a case, a guide vane arrangement and a flexible seal ring. The case extends circumferentially around an axis, and includes a first seal land. The guide vane arrangement is located radially within the case, and includes a second seal land. The seal ring at least partially seals a gap between the first seal land and the second seal land. The seal ring includes a first leg and a second leg that is connected to the first leg at a corner of the seal ring. The first leg is axially engaged with the first seal land. The second leg is axially engaged with the second seal land.

Abstract: A system for an inlet guide vane shroud and baffle assembly includes a shroud. The shroud includes a first end configured to couple to an annular support member of the gas turbine engine, and a second end. The shroud also includes a distal end including an axially aft extending lip, and a seal assembly configured to form a seal with an annular set of rotating teeth on a rotor of the gas turbine engine. The shroud further includes an arcuate shroud body extending between the distal end and the second end. The shroud and baffle assembly also includes a baffle coupled to an aft side of the shroud. The baffle includes a radially outward end coupled to the aft side, a radially inward end coupled to the axially aft extending lip, and an arcuate baffle body extending there between. Both the shroud and the baffle extend circumferentially about a central axis.

Abstract: A gas turbine engine includes a blade track, a support assembly, and a seal system. The blade track is arranged around a central axis of the gas turbine engine to direct gasses through the gas turbine engine. The support assembly is arranged around the blade track to support the blade track relative to the central axis. The seal system is configured to block fluid communication between pressurized cavities formed in the support assembly.

Abstract: An assembly for operating a device, such as a variable vane, has an arm, a stem to which the arm is connected, the stem having a slot, and a member having a tab which fits into the slot.

Abstract: The invention relates to an assembly which includes: two rotationally symmetrical components (10, 20) for supporting the blades of a turbine engine, arranged one inside the other concentrically about a turbine engine axis, and a system (30) for locking the components (10, 20) such as to prevent the relative translation of same in the axial and radial directions relative to said axis, the system including a slotted ring (40) comprising a U-shaped cross-section suitable for receiving one end of the components (10, 20), the assembly being characterized in that the ring (40) and one of the components (10, 20) are shaped such as to allow the ring to be interlocked on the component, and in that the locking system (30) also includes a member for stopping the rotation of the ring relative to the component with which the latter is interlocked. The invention further relates to a method for assembling such an assembly.

Abstract: A combustor wall is provided for a turbine engine and includes a combustor shell and a heat shield. The heat shield is attached to the shell with first and second cavities extending between the shell and the heat shield. The first cavity fluidly couples apertures defined in the shell with the second cavity. The second cavity fluidly couples the first cavity with apertures defined in the heat shield. The shell and the heat shield converge toward one another about the second cavity.

Abstract: A robot system includes a robot and a robot control apparatus. The robot control apparatus is connected to the robot for controlling the robot and includes a casing and a display unit movable with respect to the casing and displaying information. The casing has a plurality of attachment parts to which the display unit is attached. The plurality of attachment parts are provided on first and second surfaces of the casing, and the first surface is not parallel to the second surface.

Abstract: A number of variations may include a product comprising: a vane lever comprising a first end, a second end, a top surface, and a bottom surface; an opening which extends through the second end and which is defined by an inner surface, and wherein a perimeter of the inner surface comprises a plurality of vertical flutes which extend a length of the inner surface and which are constructed and arranged to displace a material of a vane shaft.

Abstract: A guide vane arrangement of a gas turbine and a method of manufacturing a guide vane arrangement of a gas turbine are provided herein. The guide vane arrangement includes a first guide vane device including a first radially inner platform and a first number of first airfoils, and a second guide vane device including a second radially inner platform and a second number of second airfoils. The first guide vane device and the second guide vane device are arranged along a circumferential direction of the turbine, wherein the first number of the first airfoils differs to the second number of the second airfoils. The first guide vane device is designed with a higher heat resistance than the second guide vane device.

Abstract: An axial flow turbine diaphragm is constructed without welding or other metal fusion or melting techniques. Static blade units are attached to inner and outer diaphragm rings by radially inner platform portions that engage the radially inner ring, and radially outer platform portions that engage the radially outer ring, the inner platform portions being elongate in the circumferential direction of the turbine diaphragm and the outer platform portions being elongate in a direction compatible with the stagger angle of the aerofoils. The outer circumference of the radially inner ring has a blade unit retaining feature of complementary shape and orientation to the inner platform portions of the static blade units, and the inner circumference of the radially outer ring is provided with a plurality of blade unit retaining features of complementary shape and orientation to corresponding outer platform portions of the static blade units.

Abstract: An airfoil fairing shell for a gas turbine engine includes an airfoil section between an outer vane endwall and an inner vane endwall, at least one of the outer vane endwall and the inner vane endwall including a radial attachment face, a suction side tangential attachment face, a pressure side tangential attachment face, and an axial attachment face.

Abstract: It is provided with: a variable nozzle mechanism section 23; an opening degree regulating member 43 for regulating an angle of a nozzle vane 19a by bringing an outer peripheral surface 41 of an inner cylinder of a bearing housing 13 to which a nozzle mount is affixed and an inner peripheral surface of a lever plate 33 into contact; and a discharge recess 47 for discharging soot which accumulates in the gap between an inner peripheral end portion of the lever plate 33, the nozzle mount 21, and the bearing housing 13 to the inner peripheral end portion of the lever plate 33.

Abstract: A turbine or compressor stage of a gas turbine, the rotor blade having a radially outer shroud (1) which has a sealing fin array having a first sealing fin (3.1) and a second sealing fin (3.2) which is adjacent to the first sealing fin and connected thereto by a first groove base (10) having a circumferential region (13) of maximum radial height, which is located at a first circumferential position is provided. The sealing fin array has a third sealing fin (3.3) adjacent to the second sealing fin and opposite to the first sealing fin, the third sealing fin being connected to the second sealing fin by a second groove base (20) having a circumferential region (23) of maximum radial height, which is located at a second circumferential position different from the first circumferential position.

Abstract: Three or more attachment pins are arranged on a right lateral surface of a nozzle ring at intervals. All the attachment pins are placed outside support holes in the nozzle ring in radial directions of the nozzle ring. A guide ring is provided across right lateral surfaces of the attachment pins. A first side guide member to support a right lateral surface of a drive ring in away that allows the right lateral surface to be in sliding contact therewith is provided to the right of the guide ring. A second side guide member to support a left lateral surface of the drive ring in a way that allows the left lateral surface to be in sliding contact therewith is provided to the left of the guide ring.