Abstract: A flowpath assembly for a gas turbine engine includes a plurality of flowpath insert ducts arranged in a cascade configuration. Each flowpath duct includes a radially inward wall, a radially outward wall, a first side wall, and a second side wall. A flowpath volume is defined between the inward, outward, first side and second side walls. The first side wall of a given one of the plurality of flowpath insert ducts is positioned adjacent to the second sidewall of an adjacent one of the plurality of flowpath insert ducts in the cascade configuration.

Abstract: A turbine vane ring has radially outer and inner annular shrouds defining therebetween an annular gaspath. Circumferentially spaced-apart airfoil vanes extend radially across the gaspath between the outer and the inner shrouds. The radially outer shroud has a circumferentially continuous cylindrical wall extending axially from a leading edge to a trailing edge. A set of circumferentially distributed stress relieving slots is defined in the leading edge of the cylindrical wall at locations adjacent to the leading edge of at least some of said airfoil vanes. The stress relieving slots extend radially through the cylindrical wall from the radially inner surface to the opposed radially outer surface thereof.

Abstract: A method for producing a vane for application in a component in a gas turbine engine includes casting at least one of a vane leading edge part and a vane trailing edge part and connecting the leading edge part and the trailing edge part.

Abstract: A variable vane assembly for a gas turbine engine is disclosed herein. The variable vane assembly includes a vane operable to pivot about a pivot axis. The vane includes a flap portion positionable in a fluid stream for controlling a flow of fluid. The vane also includes a stem portion for mounting the flap portion. The vane also includes a button portion positioned between the flap portion and the stem portion along the pivot axis. The button portion is wider than the stem portion in at least one plane containing the pivot axis. The variable vane assembly also includes a bushing encircling the stem portion and abutting the button portion. The variable vane assembly also includes a labyrinth seal positioned between the bushing and the button portion.

Abstract: An assembling of a guide vane assembly of a turbomachine turbine, formed by an annular row of stationary flow-stabilizing vanes. An outer edge of the guide vane assembly is axially bearing on a line of an external casing of the turbine and an inner edge of the guide vane assembly is in axial sliding connection with the line of the internal casing of the turbine, the axial sliding connection allowing the inner edge to be free along the motor axis with the axial stop being achieved by the axial bearing of the outer edge of the vane.

Abstract: A sectorized turbomachine compressor guide vanes assembly includes assembled sectors forming two concentric shell rings, one outer and one inner, between which vanes are arranged with their leading and trailing edges near transverse faces of the shell rings, and of which an outer shell ring externally includes an attachment mechanism to an external casing that houses the sectors. The attachment mechanism is axially offset from the rear transverse face of the shell ring to be located, in projection, in alignment with the vanes between the leading and trailing edges thereof.

Abstract: A vane assembly for a gas turbine engine comprises at least one vane, a first platform and a second platform. The at least one vane has a first end and an opposite second end with the first end of the at least one vane being secured in a corresponding first aperture in the first platform, and the second end of the at least one vane being secured in a corresponding second aperture in the second platform. The at least one vane intersects the first platform at an acute angle, and the at least one vane intersects the second platform at an obtuse angle with the acute angle and the obtuse angle being supplementary to one another.

Abstract: A wear liner for a stator vane includes a shaped member, comprising a U-shaped portion that receives a foot of a stator vane. A first end of the wear liner is biased into contact with the stator vane. A spring seal extends from the U-shaped portion and into bias contact with an inner surface of the support structure to further prevent leakage between the stator foot and case support structure.

Abstract: A mounting apparatus includes a mounting pin having first and second ends, and a support having first and second sides. The support includes a passage between the first and second sides for receiving the mounting pin so that the first end of the pin protrudes from the first side and the second end of the pin protrudes from the second side. The second end of the pin engages an article to be mounted by the mounting apparatus. The first end rotates the pin about the axis of the pin to vary the position of a mounted article. A clamping plate clamps the first end of the mounting pin to the first side of the support for securing the rotational position of the pin with respect to the support.

Abstract: A turbomachine stator stage including two coaxial shrouds, respectively an inner shroud and an outer shroud with substantially radial vanes extending therebetween, the radially outer ends of the vanes being welded to the outer shroud and the radially inner ends of the vanes passing through orifices in an outer ring of the inner shroud, the inner ends being held by elastomer spacers in an annular cavity formed between the outer ring and an inner ring that is fastened to the outer ring.

Abstract: An aggregate vane assembly is disclosed herein. The aggregate vane assembly includes a core vane assembly encircling a central longitudinal axis. The core vane assembly has a plurality of core vanes each extending radially between an inner hub and an outer band. The core vane assembly extends along the central longitudinal axis between a first forward end and a first aft end. The aggregate vane assembly also includes a bypass vane assembly disposed on a radially opposite side of the outer band relative to the plurality of core vanes. The bypass vane assembly includes at least one bypass vane extending radially outward from a platform. The bypass vane assembly extends along the central longitudinal axis between a second forward end and a second aft end. The aggregate vane assembly also includes at least one boss fixed with the outer band and operable to engage the bypass vane assembly proximate to the second forward end.

Abstract: A vane structure for a gas turbine engine includes a multiple of CMC airfoil sections integrated between a CMC outer ring and a metal alloy inner ring.

Abstract: A rectifier stage of an axial turbine engine compressor is placed directly downstream of the input blower of a turbo-reactor or turboshaft engine compressor. The blade roots of the rectifier stage are housed inside the inner shroud by insertion into respective slots. The blade roots include at least one opening protruding from the slot of the shroud, so as to receive a spiral ring having at least two loops so as to prevent the blade roots from escaping from the respective slots of the inner shroud, and so as to form a connection between the blade roots.

Abstract: A turbocharger with a variable turbine geometry device has one or two blade bearing rings. At least one or more fastening elements are provided for fastening the one or two blade bearing rings. The respective fastening element has at least one section formed with a knurl and/or at least one section formed with at least one or several notches.

Abstract: A structural and aerodynamic module including two blades each extending along a longitudinal axis, in which each blade includes a leading edge, a trailing edge, a radially inner edge, and a radially outer edge, is provided. The module also includes a radially inner platform mechanically connecting the radially inner edges of the two blades, and a radially outer platform mechanically connecting the radially outer edges of the two blades such that the two blades and the radially outer and inner platforms delimit a flow stream. The two blades and the platforms are made from a composite material and the two blades are fixed on the radially outer and inner platforms using screw-nut assemblies.

Abstract: A strut configuration of a static vane ring used in a gas turbine engine having an enlarged end section at least at one of the opposed ends thereof to be welded or brazed to either an outer or inner duct wall of the vane ring. The enlarged end section provides a inner corner curve with a predetermined fillet radius between the strut and the duct wall.

Abstract: The invention relates to a nozzle stage for a turbomachine compressor, the stage comprising two coaxial shrouds, a radially inner shroud and a radially outer shroud, and radial vanes extending between the shrouds and being connected via their ends to the shrouds, together with a support for abradable material that is mounted by crimping onto the upstream and downstream edges of the inner shroud, the inner shroud presenting openings in which the radially inner ends of the vanes are inserted and fastened. The radially inner ends of the vanes are plane and do not have any projecting tenons.

Abstract: A vane includes a composite airfoil having a single dovetail root extending chordwise between or near leading and trailing edges of the airfoil. The root is connected to the airfoil and is slidably received in a dovetail slot in an inner vane mount. The dovetail slot has a triangular slot cross section and a flat slot bottom and flat slot pressure and suction side crush faces corresponding to pressure and suction sides of the airfoil. The root has a triangular root cross section, a flat root bottom, and flat dovetail root pressure and suction side crush faces. A biasing means urges the root bottom of the dovetail root away from the slot bottom and presses and preloads the dovetail root pressure and suction side crush faces against the slot pressure and suction side slot crush faces.

Abstract: A vane includes a composite airfoil having one or more pockets and one or more spars received therein respectively. The spars extending radially away from a bridge of a vane mount supporting airfoil. The spars may be integrally formed with the bridge. The spars may be adhesively bonded to the composite airfoil within the pockets. A passage may extend between an airfoil base and an airfoil tip. A gas turbine engine annular fan frame includes an annular row of the composite outlet guide vanes extending radially between and connecting radially inner and outer rings of the fan frame.

Abstract: The guide vane of a gas turbine comprises a blade having an outer platform connectable to a guide vane carrier or gas turbine casing, and at the opposite end an inner platform that carries a cover plate. The inner platform has a dovetail housing to which the cover plate is connected. The method for replacing the cover plate consists in separating the cover plate from the inner platform by cutting along the profile of the dovetail housing, then making the cover plate to slide out of the dovetail housing, thus inserting a new or refurbished cover plate into the dovetail housing.

Abstract: A nozzle blade and nozzle ring assembly includes a nozzle blade having radially inner and outer sidewalls with an airfoil portion extending therebetween; the inner and outer sidewalls walls formed with axially-extending first surface features along forward and aft marginal edges of the inner and outer sidewalls, respectively; and radially inner and outer nozzle rings formed with corresponding axially-extending second surface features mated with the first surface features, wherein the radially inner and outer sidewalls are welded to the radially inner and outer nozzle rings only along the mated first and second surface features.

Abstract: A vane assembly of a gas turbine engine has a plurality of vanes extending radially and inwardly from an annular casing. An outer end of each vane radially and outwardly projects from the casing and is received in one of apertures defined in an elastomeric ring which is placed around the casing. A strap of a non-metallic material is placed in a pre-tensioned condition around the elastomeric ring to compress the elastomeric ring and to radially retain the outer end of each vane.

Abstract: A vane includes a composite airfoil having one or more outer pockets and one or more outer spars received therein respectively. The outer spars extending radially inwardly from an outer bridge of the outer vane mount outwardly supporting airfoil. The spars may be integrally formed with the bridge and adhesively bonded to the composite airfoil within the pockets with a passage extending between an airfoil base and an airfoil tip. A metallic leading edge tip may be disposed along the leading edge of the composite airfoil. A gas turbine engine annular fan frame incorporates the vanes in an annular row of composite outlet guide vanes extending radially between and connecting radially inner and outer rings of the fan frame.

Abstract: A stator vane assembly includes: inner and outer vane supports, and vanes extending radially between the supports. The supports have openings for receiving the ends of the vanes, and each vane is connected to the supports by a vane boot. Each vane boot has: a rim portion shaped to locate the boot within the corresponding opening in the vane support, the rim portion engaging with the inner surface defining the opening, an abutting portion for abutting against the vane support in the region surrounding the opening, and a vane pocket for accommodating the end of the vane. The vane pockets of the vane boots are shaped so as to position the vane with a predetermined stagger angle.

Abstract: An example airfoil retention arrangement includes a retention assembly having a first retention segment and a second retention segment. Each of the retention segments is separately moveable to an installed position relative to an airfoil assembly and a support structure. The retention segments each have a portion position between a lip of the airfoil assembly and a collar of the support structure when the retention segments are in the installed position. The retention assembly is configured to limit radial movement of an airfoil relative to the support structure when in the installed position.

Abstract: The present invention relates to a method for replacing an inner disk element of a blade integrated disk wherein the blade integrated disk has a plurality of recesses in the form of axial through-bores in the joining area. The method begins with filling the axial through-bores, after which the inner disk element of the blade integrated disk is separated from the remaining blade ring. Next a new inner disk element is inserted into the remaining blade ring and the new inner disk element and the remaining blade ring are joined. Finally the axial through-bores are restored.

Abstract: A method for assembling a rotary machine includes providing a rotor including a plurality of rotor wheels. The method also includes positioning the rotor such that at least a portion of a stationary portion of the rotary machine extends at least partially about the rotor. The method further includes providing a blade that includes a blade platform that is formed with a substantially double-C shape. The method also includes coupling the blade to the rotor.

Abstract: A ferrule of a stator with blades of an axial turbomachine consists of at least two segments having the general form of an arc of a circle and intended to be arranged end to end, so as to form the ferrule, where each segment contains on its outside a plurality of openings arranged at regular distances along the general arc of the circle of the segment, into each of which fits the end of a blade of the stator. Each end of a segment possesses a section of which the profile essentially corresponds with that of the end of the corresponding blade when the segment is placed on the blades of the stator, such that the joint between two adjacent segments corresponds with the profile of the corresponding blade.

Abstract: A high temperature resistant turbine stator vane with an airfoil section made from oxide dispersion strengthened (ODS) material extruded through a die, an internal cooling circuit is machined into the airfoil section after the extrusion, and then two endwalls are formed separately and then secured to the airfoil section ends to form the stator vane. Film cooling holes can be drilled into the airfoil section and the endwalls.

Abstract: A system, in certain embodiments, includes a plurality of detachable, three-dimensional diffuser vanes attached to a diffuser plate of a centrifugal compressor. In certain embodiments, the detachable, three-dimensional diffuser vanes may be attached to the diffuser plate using threaded fasteners. In addition, dowel pins may be used to align the detachable, three-dimensional diffuser vanes with respect to the diffuser plate. However, in other embodiments, the detachable, three-dimensional diffuser vanes may include a tab configured to fit securely within a groove in the diffuser plate. In addition, the tabs of the detachable, three-dimensional diffuser vanes may include indentions that mate with extensions extending from the diffuser plate, wherein the tabs may slide into slots between the extensions and the grooves of the diffuser plate.

Abstract: A columnar air moving device can comprise separately formed modular stator vanes in a stator vane assembly. The stator vanes can be arranged in a radial pattern, and can direct air in an axial direction. The modular stator vanes, as well as other components of the stator vane assembly, can be replaced, adjusted, and/or removed from the columnar air moving device.

Abstract: An assembly for a turbomachine comprising: a plurality of components (16, 26) which, in use, are assembled into a circumferentially extending annular array; each component having a radially outer platform (20, 28), the outer platforms (20, 28) of the components (16, 26) forming a ring in the assembled array; wherein adjacent components (16, 26) interface at adjacent surfaces (22, 30) of their outer platforms (20, 28); wherein the outer platforms (20, 28) each comprise at least one lug (24) which extends from the outer platform (20) over a portion of the outer platform (28) of an adjacent component (26), and in use prevents relative radial displacement of the outer platform (20, 28) of the adjacent component (16, 26).

Abstract: The present invention relates to a strut for a turbine center frame of a jet engine, in particular a gas turbine, with a first end for disposition on an internal structure of the turbine center frame and a second end opposite the first end for attaching the strut on a housing or an internal boundary wall of the turbine center frame, and on the second end at least one flange-like element with at least one first aperture is removably arranged with at least one first aperture for accommodating or passing through an attachment means. The invention further relates to a turbine center frame for a jet engine, in particular a gas turbine as well as a method for producing a turbine center frame of a jet engine, in particular a gas turbine.

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 portion s 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: A steam turbine singlet nozzle airfoil with integral outer sidewall is engaged with an inner ring and an outer ring in a nozzle assembly. The interface of the outer sidewall with the outer ring may include a plurality of mechanical hooks on one or both of the upstream face and the outer radial face of the outer sidewall that engage with complimentary structures on the outer ring. The outer interface may further include low energy welds along limited distances of one or both of the upstream or downstream interface of the outer sidewall and the outer ring. An inner radial end of the singlet nozzle airfoil is pinned into position and fastened to the inner ring. Without a need for high heat welds, distortion of the airfoil and the steam flow path and the associated rework is eliminated and stage performance is improved.

Abstract: A stator vane assembly comprising inner and outer support rings connected by an annular array of stator vanes at locations around the supporting rings, at least one of said stator vanes in said array being a separate, pre-formed stator vane which is slidably connected to one or both of the supporting rings for radial movement relative to the respective support ring thereby to accommodate relative radial expansion of the support rings.

Abstract: A vane includes a composite airfoil having chordwise spaced apart forward and aft dovetail feet at or near leading and trailing edges of the airfoil. Each foot includes a shank connecting a dovetail root to an airfoil base and each root is slidably received in a dovetail slot in an inner vane mount. The dovetail slot has a triangular slot cross section and a flat slot bottom and flat slot pressure and suction side crush faces corresponding to pressure and suction sides of the airfoil. The dovetail root having a triangular root cross section and a flat root bottom and flat dovetail root pressure and suction side crush faces. A biasing means urges the root bottom of the dovetail roots away from the slot bottom and presses and preloads the dovetail root pressure and suction side crush faces against the slot pressure and suction side slot crush faces.

Abstract: Systems, methods, and apparatus for linking machine stators are provided. A bracket may be utilized to link adjacent stators within a machine, for example, a turbine. The bracket may include a body portion, a first plurality of extensions, and a second plurality of extensions. The body portion may include a first edge operable to align with a base of a first stator and a second edge opposite the first edge and operable to align with a base of a second stator. The first plurality of extensions may extend from the first edge and may be operable to align with at least one corresponding groove on the base of the first stator. The second plurality of extensions may extend from the second edge and may be operable to align with at least one corresponding groove on the base of the second stator. When utilized to link the first stator and the second stator, the bracket may facilitate a reduction of vibrations in the first stator and the second stator.

Abstract: A compressor diaphragm for a gas turbine engine having improved wear capability, manufacturability, and assembly techniques is disclosed. The diaphragm includes a shiplap-type joint at an outer vane platform for connecting to adjacent vane assemblies and a clamshell-like assembly of a seal box secures and seals regions around the inner vane platform of the compressor diaphragm so as to reduce wear between the seal box and the vane assemblies. The inner platform of the diaphragm segments are fastened to each other through circumferentially-oriented fasteners at the inner diameter platform.

Abstract: Guide vane stage of a turbine engine comprising a series of fixed blades connecting an inner collar to an outer collar, characterized in that said fixed blades comprise attachment platforms with flat surfaces that co-operate with a plurality of juxtaposed flat facets, the flat facets being located on the inner face of the outer collar in order to ensure attachment with a contact of a flat/flat type between the fixed blades and the outer collar.

Abstract: A static vane ring for gas turbine engines includes a plurality of radial struts extending between and interconnecting outer and inner duct walls which define an annular duct therebetween. A load transfer apparatus is attached to at least one of the outer and inner duct walls to transfer load from vane to ring, and between vanes.

Abstract: A mechanical joint for a gas turbine engine includes:(a) an annular first component having an annular, radially-extending first flange; (b) an annular second component having an annular, radially-extending second flange abutting the first flange; (c) a plurality of generally radially-extending radial channels passing through at least one of the first and second flanges; (d) a plurality of generally axially-extending channels extending through the first flange and communicating with respective ones of the radial channels; and (e) a plurality of fasteners clamping the first and second flanges together.

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

Abstract: Bearing support structure for turbines comprising an inner ring (1) and an outer ring (2) radially connected by means of structural vanes (5) and aerodynamic vanes (6) in a circumference-like arrangement between both rings (1, 2). The aerodynamic vanes (6) are thinner and lighter than the structural vanes (5), and the number of structural vanes (5) depends exclusively on the loads of the bearing (3) housed in the structure to be transmitted to the anchoring points (7) of the engine assembly located in the outer ring (2), and on the amount of service fluids which must go through between the inner ring (1) and the outer ring (2), while the number of aerodynamic vanes (6) and their section depends exclusively on the aerodynamic requirements demanded from the support structure for the straightening of the turbine main flow. Thus, the structural vanes (5) fulfill only structural functions and the aerodynamic vanes (6) fulfill only aerodynamic functions.

Abstract: A section of a gas turbine engine including a radial spoke is provided. The spoke includes an aerodynamic shape with a leading side and a trailing side and, extending from the leading side to the trailing side, a first side and a second side, opposite the first side. The spoke also includes at least one flow guiding element arranged on at least the first side.

Abstract: A retaining seal assembly for use in a gas turbine engine is provided. The retaining seal assembly comprises an outer retaining ring coupled to an aft end of a gas turbine engine combustor. A turbine nozzle is coupled to the outer retaining ring and comprises an outer band. The outer band comprises a leading edge and an opposing trailing edge that defines a slot. A retention seal comprises a first end positioned within the slot, a generally opposing second end that contacts the outer retaining ring, and a body extending between the first end and the second end. The body further comprises an insertion portion positioned within a passage formed in the outer band. The retention seal is fabricated from a resilient material and is configured to facilitate coupling the turbine nozzle to the outer retaining ring.

Abstract: An inner shroud assembly includes an aft core segment mountable to a forward core segment to support a multiple of vanes for rotational movement relative thereto. A shroud backing plate segment engageable with the aft core segment and at least one fastener which passes through the shroud backing plate, the aft core segment and the forward core segment.

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

Abstract: Assembly of the connection of an arm to a ferrule is done by imbricating a side wing protruding from the end of the arm and keys bolted to the ferrule. The assembly may be supplemented with positioning pins.

Abstract: A cooling fan having a fan housing, a motor which is positioned within the fan housing and a plurality of electronics components which are associated with the motor includes at least one strut which extends radially between the motor and the fan housing and on which the electronics components are mounted.