Abstract: A method facilitates assembling a diaphragm assembly for a steam turbine. The method includes forming at least one opening within a radially outer band and forming at least one opening within a radially inner band. The method further includes coupling at least one partition to at least one opening within the radially outer band including a radially inner surface and a radially outer surface wherein the at least one opening is at least partially defined by a wall that extends obliquely between the outer band radially inner surface and the outer band radially outer surface. The method additionally includes coupling the at least one partition to the at least one opening within the radially inner band wherein the at least one partition extends between the at least one radially outer band opening and the at least one radially inner band opening.

Abstract: A system and method for configuring a steam turbine to accommodate changing resource conditions, such as may be encountered with geothermal wells. A plurality of sets of nozzle/blade assemblies are provided for installation in a diaphragm structure and on a rotor, respectively. As the condition of steam provided to the turbine changes, a different set of nozzle/blade assemblies may be installed to maintain a preferred thermodynamic efficiency for the turbine.

Abstract: A flow discouraging system includes a stator assembly, fins, and a rotor assembly. The stator assembly includes stationary components forming a side wall including an annular groove defined by an outer axially-extending surface, an inner axially-extending surface, and a radial surface. One or more outer axial fins disposed in the annular groove extend along the outer axially-extending surface of the side wall. One or more inner axial fins disposed in the annular groove extend along the inner axially-extending surface of the side wall. One or more radial fins disposed in the annular groove extend axially from the radial surface of the side wall. The rotor assembly is disposed adjacent to and is spaced apart from the stator assembly to form a portion of a cavity and includes an annular rim extending at least partially into the annular groove and disposed between the outer and inner axial fins.

Abstract: A gas turbine includes a nozzle vane and a sealing unit engaged with the nozzle vane inside a turbine supplied with combustion gases produced by mixing and burning air for combustion and fuel. The nozzle vane and the sealing unit are disposed in a channel of the downward flowing combustion gases on the outlet side of a gas path. A plurality of engagement portions between the sealing unit and the nozzle vane are provided successively from the upstream side toward the downstream side in a direction of flow of the combustion gases, and a downstream one of the plurality of engagement portions has a contact interface formed in a direction across a turbine rotary shaft. A reduction in the thermal efficiency of the gas turbine can be suppressed.

Abstract: A gas turbine engine shroud includes a plurality of shroud segments disposed circumferentially one adjacent to another to form a full, circumferentially segmented ring about the rotor. The radial seal between each pair of adjacent shroud segments is positioned on the back side of the shroud platform. Cooling underneath the radial seals may be done by conduction heath transfer through the platform of the shroud segments.

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: Embodiments of a turbine nozzle assembly are provided for deployment within a gas turbine engine (GTE) including a first GTE-nozzle mounting interface. In one embodiment, the turbine nozzle assembly includes a turbine nozzle flowbody, a first mounting flange configured to be mounted to the first GTE-nozzle mounting interface, and a first radially-compliant spring member coupled between the turbine nozzle flowbody and the first mounting flange. The turbine nozzle flowbody has an inner nozzle endwall and an outer nozzle endwall, which is fixedly coupled to the inner nozzle endwall and which cooperates therewith to define a flow passage through the turbine nozzle flowbody. The first radially-compliant spring member accommodates relative thermal movement between the turbine nozzle flowbody and the first mounting flange to alleviate thermomechanical stress during operation of the GTE.

Abstract: A nozzle assembly for a turbine that may include: (1) a nozzle blade having inner and outer sidewalls and, in part, defining a flowpath upon assembly into the turbine; (2) an outer ring; (3) a flowsplitter having a horizontal extension; (4) an interface between the outer ring and the outer sidewall having at least one of (i) a male/female interface or (ii) a radial interlock; and (5) an interface between the horizontal extension and the inner sidewall having at least one of (i) the male/female interface or (ii) the radial interlock. In some embodiments, one of the interface between the outer ring and the outer sidewall and the interface between the horizontal extension and the inner sidewall comprises a weld and one of the interface between the outer ring and the outer sidewall and the interface between the horizontal extension and the inner sidewall is weld free.

Abstract: A rotary drum of an axial compressor having a web made of composite material has an upstream flange for fastening the fan, a segment having a sealing flange upstream, a segment with a groove adapted to receive the bases of the vane of a first row of vanes, a segment with a groove adapted to receive the bases of the vane of a second row of vanes, and a segment with a groove adapted to receive the bases of the vane of a third row of vanes. These segments are made of metallic material mostly for reasons of mechanical resistance to the centrifugal forces. The segments are connected to each other with web segments made of composite material. The ends of the metallic segments have a dovetail-shaped section in a longitudinal plane that is able to ensure a positive connection with the composite material.

Abstract: A support frame includes a rail engaging a hook of a composite article having plys. The rail included a rail contact surface generally conformal to and contacting a hook contact surface of the hook. The rail and hook contact surfaces being substantially normal to the plys beneath the hook contact surface. One embodiment of the article is a composite nozzle segment including one or more airfoils extending radially between arcuate radially outer and inner band segments having plys curved about a centerline axis and the rail engages a hook of the radially inner band segment. The hook and the rail may be straight or arcuate. The frame includes a central body extending axially aftwardly from the rail to an aft frame flange. A frame pin may be mounted on the frame, engage the inner band segment or one of the airfoils, and may extend into a cavity in the airfoil.

Abstract: A turbine vane attachment system configured to eliminate movement of a turbine vane relative to a turbine vane carrier. The turbine vane attachment system may include a base attached to a turbine airfoil. The base may be configured to contact a wedge support along a plane that is generally nonparallel and nonorthogonal with a longitudinal axis of the airfoil. A bolt may connect the base with the wedge support. The bolt may change a distance between a channel in the base and an outer bearing surface. The turbine vane may be positioned in a vane carrier such that tongues extending from the vane carrier are positioned in the channels. As the bolt is advanced, the wedge support is moved laterally along the support surface of the base and the channels engage the tongues, thereby preventing movement of the turbine vane.

Abstract: A device for guiding a stream of air entering a combustion chamber of a turbomachine is disclosed. The device includes a flow straightener followed by a diffuser. One of the flow straightener shrouds is formed as one with one wall of revolution of the diffuser. The other flow straightener shroud is added and attached to the other wall of revolution of the diffuser. The vanes of the flow straightener are secured by one end to one shroud of the flow straightener and separated by a small clearance from the other shroud at their other end.

Abstract: The invention relates to a novel reworking of a geometric shape of a stator vane in order locally to reduce the stress levels. It proposes a novel type of stator section in which a local reworking of the geometry of the vanes has been performed, this reworking making it possible to reduce the level of stress in the brazed joint connecting the vane to its outer shroud without entailing further aerodynamic engineering. To do that, the invention consists in a progressive and localized increase in the radius at the trailing edge combined with a progressive increase in the thickness of the vane.

Abstract: The present invention relates to a method for manufacturing a turbomachine rectifier comprising an outer ring (5) and a plurality of stator vanes (2), said vanes comprising a platform (9) and a blade (12), wherein said method comprises at least the following steps: a) first folds (7) are draped over a core comprising perforations and having the shape of the aerodynamic flow; b) optionally, the first folds (7) are cut along incisions (8) placed opposite the perforations of the core; c) the blades (12) of the vanes or, alternatively, the blades (12) of the vane preforms, are inserted through the incisions (8) of the first folds (7) and through the perforations of the core; d) last folds (10) are draped over the platforms (9) thus forming a rectifier preform; e) a resin is injected into a closed mold with the preform and the resin-impregnated preform is polymerized; f) a molded piece, essentially having the shape and dimensions of said rectifier, is retrieved from the mold.

Abstract: A turbine nozzle includes radially inner and outer bands integrally joined to opposite ends of an airfoil. The outer band has forward and aft hooks extending radially outwardly at axially opposite ends. And the hooks have different prechorded circumferential curvatures.

Abstract: A compressor having a plurality of vane units each having at least one airfoil projecting from the base. Each base has a groove, so that when the vane units are disposed circumferentially adjacent, the grooves of the respective vane units are circumferentially aligned. A metal strip is disposed in the groove of a plurality of adjacent vane units to link the vane units forming a more rigid ring of vanes that are less susceptible to vane motion, e.g., caused by pressure fluctuations within the compressor of a gas turbine.

Abstract: An outer sidewall retention scheme for a singlet first stage nozzle of a gas turbine. The retention scheme includes a circumferential retaining ring with a main body and a pair of circumferential retaining lands projecting inward radially. A circumferential annular retaining groove is formed between the pair of circumferential retaining lands. A first lug and a second lug mounted on an outer face of the outer sidewall of each nozzle are adapted to fit within the circumferential annular retaining groove of the retaining ring and are supported radially and circumferentially by a first retaining pin and a second retaining pin, each pin passing though the circumferential retaining lands. Each nozzle further includes a chordal hinge rail and seal on the outer sidewall and a chordal hinge rail and seal on the inner sidewall providing axial support for the nozzle.

Abstract: A method of repairing a turbine nozzle segment having at least one vane disposed between outer and inner bands. The method includes using an salvageable part, such as honeycomb backing strip remnant, from a nozzle segment that is unrepairable, and joining the honeycomb backing strip to a replacement casting. The replacement casting includes a tenon formed on the inner band that mates with a mortise formed on the back side of the honeycomb backing strip. A nozzle segment and nozzle segment casting used in connection with the method is also disclosed.

Abstract: An apparatus and method for locking a composite component such as blade or vane of a turbine engine is disclosed herein. The disclosed lock can also be used generally to fix the position of composite structures. The lock includes a first locking member including a first structure operable to at least partially fix the first locking member along an axis of a slot for receiving a blade or a vane. The lock also includes a second locking member including a second structure operable to engage the blade or vane. The first and second locking members are slidably engaged with one another along mating ramped surfaces.

Abstract: A turbine module for a gas turbine engine includes at least an annular distributor and a turbine rotor inside a casing, where the annular distributor includes a plurality of elements in the form of a ring sector, of which a first part supports fixed blades positioned radially towards the turbine axis, and a second part forms a seal with the tips of the turbine rotor blades. The elements in the form of a ring sector are held inside the casing by attachment resources.

Abstract: A mounting arrangement (10) for a multivane segment (12) of ceramic matrix composite (CMC) composition positioned between outer and inner metallic rings (14, 16). Selected ones of the vanes (18a) of the multivane segment surround internal struts (24) joining the outer and inner rings. Spring members (26, 28) accommodate differential thermal growth between the multivane segment and the outer and inner rings, and a compliant material (30) seals against gas leakage around the segments.

Abstract: A gas turbine engine is provided comprising an outer casing and a plurality of circumferentially positioned vane segments. The outer casing is provided with a circumferential casing slot. The plurality of circumferentially positioned vane segments are coupled to the outer casing. Each vane segment comprises at least one vane airfoil, a radially inner shroud coupled to a first end of the airfoil, a radially outer shroud coupled to a second end of the airfoil, and a strongback fixedly coupled to axially spaced-apart portions of the outer shroud such that a gap is provided between the strongback and the outer shroud. The strongback may comprise axially spaced-apart first and second end portions received in the casing slot.

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: A method for assembling a stator assembly for a turbine engine is provided. The method includes providing a blade with a base including an end wall having at least one hole defined therein and providing a shim having at least one aperture extending therethrough. The shim aperture is aligned with the end wall hole, and the shim is secured to the blade base end wall using a fastener. The fastener is inserted through the shim aperture in an interference fit within the end wall hole. The blade and the shim are coupled to a turbine casing.

Abstract: A turbine vane ID support system usable to support the ID of a turbine vane and including a transition seal system and a turbine vane ID forward rail seal system. The transition seal system may seal the turbine vane ID support body to a transition, and the turbine vane ID forward rail seal system may seal the turbine vane ID support body to a turbine vane. Use of the transition seal system and the turbine vane ID forward rail seal system eliminates the problems inherent with conventional seals used to seal transitions directly to turbine vanes. The turbine vane ID support system may also be configured such that the turbine vane ID support system may be removed to facilitate removal of turbine vanes and blades for repair or replacement.

Abstract: A flange for supporting arcuate components comprising at least one arcuate rail, each arcuate rail having an inner radius, a first taper location, a first taper region, a second taper location, a second taper region, wherein the thickness of at least a portion of the first taper region is tapered and wherein the thickness of at least a portion of the second taper region is tapered.

Abstract: A variable stator vane airfoil is mounted on a button centered about a rotational axis and includes circular leading and trailing edges circumscribed about the rotational axis at a button radius. Contoured pressure and suction sides extend from the circular leading edge to the circular trailing edge and are recessed inwardly from a perimeter circumscribed about the rotational axis at the button radius. One of upstream and downstream pressure side portions of the contoured pressure side is straight and another of the upstream and downstream pressure side portions is curved. One of the upstream and downstream suction side portions is straight and another of the upstream and downstream suction side portions is curved. One of the upstream pressure side portion and upstream suction side portion is straight and another of the upstream pressure side portion and upstream suction side portion is curved.

Abstract: In one embodiment, a system includes a turbine engine that includes a rotor including multiple blades. The turbine engine also includes a shroud disposed about the blades. The shroud includes multiple segments engaged with one another via mating teeth. The mating teeth are oriented in an axial direction along a longitudinal axis of the turbine engine.

Abstract: A cooling fan assembly for cooling a heat exchanger includes a shroud having a plurality of stators extending radially to support a hub. A flow-guide including an outer rim and inner rim surrounds the shroud and extends to the heat exchanger. The outer rim extends from the heat exchanger to the stators. The inner rim extends annularly about the axis and curves convexly from the outer rim. A fan unit includes a plurality of fan blades extending radially from a motor supported by the hub. Each fan blade includes a leading edge facing the stators and a trailing edge facing in the opposite direction with a blade tip extending therebetween. A fan section extends from the inner rim and axially about the fan blades. The stators are disposed between the heat exchanger and the fan blades. The stators have a front edge facing the heat exchanger and a back edge facing the fan blades. Each stator is connected to the inner rim with the front edge extending radially outward and farther than the back edge.

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 vane fixing apparatus for removably fixing a vane plate (14) to a web (12) is described. The vane fixing apparatus comprises a mounting bracket (16) attached to the vane plate (14), and a mounting block (20,30) held on the web (12) by at least one fastener (18), wherein the mounting block (20,30) is moveable between a first configuration in which the mounting bracket (16) may engage and disengage the mounting block (20,30) and a second configuration in which the mounting block (20,30) securely retains the mounting bracket (16) relative to the web (12), and wherein the mounting block (20,30) is moveable between the first and second configurations by operation of the fastener (18).

Abstract: A method of assembling a turbine includes positioning a turbine nozzle against a forward inner nozzle support extending from a rotor assembly, and coupling a cover to the forward inner nozzle support. The method also includes inserting a tab that extends from the cover within at least one aperture defined in the turbine nozzle.

Abstract: An annular vane assembly for a gas turbine engine is provided. The assembly includes a vane segment, the vane segment includes an arcuate rail and a vane that extends radially inwardly from the arcuate rail. The assembly also includes a hollow cylindrical casing, the inside curved surface of which an annular groove is formed that receives the arcuate rail. The arcuate rail is secured in the annular groove using a resilient strip interposed between the rail and the groove. The resilient strip includes a planar main body and sprung wings that extend to either side of the main body. The wings are angled with respect to the plane of the main body. The resilient strip is moveable circumferentially between a first position in which the strip exerts a force radially on the arcuate rail and a second position in which the wings occupy recesses in the assembly.

Abstract: An assembly for mounting a ceramic turbine vane ring onto a turbine support casing comprises a first metal clamping ring and a second metal clamping ring. The first metal clamping ring is configured to engage with a first side of a tab member of the ceramic turbine vane ring. The second metal clamping ring is configured to engage with a second side of the tab member such that the tab member is disposed between the first and second metal clamping rings.

Abstract: A mid-turbine frame connected to at least one mount of a gas turbine engine transfers a first load from a first bearing and a second load from a second bearing to the mount. The mid-turbine frame includes a load transfer unit, a torque box rotatably positioned within the load transfer unit, and a plurality of struts. The load transfer unit has a first locking element and combines the first load and the second load into a combined load. The torque box has a second locking element that is engagable with the first locking element of the load transfer unit. The plurality of struts are connected between the torque box and the mount, and transfer the combined load from the torque box to the mount. The first locking element and the second locking element are at least one of a rib or a groove.

Abstract: A support for the first row turbine vanes (12) in a gas turbine engine, the support including a support framework (18). The support framework (18) includes an outer vane carrier (34), and an inner vane carrier (36) connected to the outer vane carrier (34) by a plurality of struts (44, 46). The outer vane carrier (34) is mounted to an inner casing (16) of the engine and the struts (44, 46) support the inner vane carrier (36) in cantilevered relation to the outer vane carrier (34). An aft outer flange (94) of each first row vane (12) is supported on the outer vane carrier (34) and a forward inner flange (114) of the vane (12) is support on the inner vane carrier (36).

Abstract: A method for assembling a turbine nozzle for a gas turbine engine. The method includes providing a turbine nozzle including a plurality of airfoil vanes extending between an inner band and an outer band, wherein the outer band includes a forward hook assembly having a rail and at least one hook, providing at least one scalloped recessed area within the forward hook assembly at least one hook to facilitate reducing stresses induced to the turbine nozzle, and coupling the turbine nozzle into the gas turbine engine using the forward hook assembly such that the turbine nozzle is at least partially supported by the forward hook assembly.

Abstract: A gas turbine CMC shroud plate (48A) with a vane-receiving opening (79) that matches a cross-section profile of a turbine vane airfoil (22). The shroud plate (48A) has first and second curved circumferential sides (73A, 74A) that generally follow the curves of respective first and second curved sides (81, 82) of the vane-receiving opening. Walls (75A, 76A, 77A, 78A, 80, 88) extend perpendicularly from the shroud plate forming a cross-bracing structure for the shroud plate. A vane (22) may be attached to the shroud plate by pins (83) or by hoop-tension rings (106) that clamp tabs (103) of the shroud plate against bosses (105) of the vane. A circular array (20) of shroud plates (48A) may be assembled to form a vane shroud ring in which adjacent shroud plates are separated by compressible ceramic seals (93).

Abstract: A guide vane arrangement 20 for a gas turbine engine (10, FIG. 1) includes a vane member 21 extending between inner and outer platforms 22, 24 which are respectively mounted on inner and outer mounting members 42, 34. One of the inner and outer platforms 22, 24 includes a resilient that abuts the respective inner or outer platform 22, 24 to permit relative movement between the inner or outer platform 22, 24 and the respective inner or outer mounting member 42, 34.

Abstract: Disclosed herein is a compressor case assembly that includes a compressor case section having an assembly track and a set of stator portions that are insertable in the assembly track, the compressor case section and the set of stator portions are configured such that the set of stator portions only assembles to the compressor case section in a single configuration.

Abstract: Vanes are generally hollow with a passage through which service pipes extend. These service pipes must be connected such that they can allow for thermal expansion and contraction but also anti rotation features are needed in order to prevent damage when a supply coupling is connected through a pipe end. This anti-rotation feature is generally provided between a top hat and a pipe end through flanges. In order to allow reduced spacing between the top of the vane and a casing the flanges are arranged to allow at least recessing of a first flange into a passage opening profile whereby those flanges inter-engage with second flanges at a surface interface to prevent rotation.

Abstract: Guide vane stage of a turbine engine comprising a series of fixed blades connecting an inner collar to an outer collar, characterised in that said fixed blades comprise attachment platforms with flat surfaces that co-operate with a plurality of juxtaposed flat facets, said 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 gas turbine includes a turbine housing; a rotatable turbine rotor with turbine blades; and at least one heat shield segment for a stator. The heat shield segment is disposed radially between the rotor and the housing and attached to the housing and has a profile with a curved section in at least one region in an axial direction of the gas turbine and a radial outer surface. The heat shield segment includes a rib, a first boss, and a second boss disposed on the radial outer surface.

Abstract: A method for preventing cracking of a turbine engine case includes the steps of disposing at least two rails upon an exterior surface of a turbine engine case; and securing a first rail to a first means for attaching at least one fan exit guide vane to the turbine engine case and securing a second rail to a second means for attaching the at least one fan exit guide vane to the turbine engine case.

Abstract: A fan frame (10) for a gas turbine is provided. The fan frame (10) has fan exit guide vanes (100) that are connected to an intermediate case (50) for structural support, and are constrained from movement in a radial direction by a first joint (110, 110a) and constrained from movement in an axial and radial direction by a second joint (120a, 110a). The intermediate case (50) provides structural integrity while reducing weight by providing struts (56, 95) with a width that increases in a direction away from a central bearing (55) as the thickness of the struts (56, 95) decreases.

Abstract: The present invention has an object to provide an assembling method of a stator blade ring segment that can prevent the risk of thermal deformation and a reduction in strength, ensure flexibility of shape, and increase compression performance of a compressor. The present invention has another object to provide welding that prevents an object to be welded from rising from a groove or a recess during welding and allows satisfactory welding. A groove 50 formed in each segment piece 41 is formed to have side wall portions 50b on opposite sides inclined with respect to a bottom 50a so that a width of the groove 50 gradually decreases from an outer peripheral side toward the bottom 50a, and a coupling member 42 has a sectional shape so as to have a maximum width at protrusions 42b at an intermediate portion in a thickness direction.

Abstract: An airfoil assembly is described comprising an airfoil, a band coupled to the airfoil for providing at least a portion of support for the airfoil in the airfoil assembly and a slot capable reducing stress in the airfoil wherein the slot is located in the band near a portion of the airfoil. A compression system is described comprising a stator assembly comprising a plurality of arcuate vane sectors arranged circumferentially around a longitudinal axis, each vane sector comprising an airfoil, an arcuate inner band and an arcuate outer band coupled to the airfoil, and a slot capable of reducing stress in the airfoil, the slot being located in the vane sector near a portion of the airfoil.

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: A turbo-engine stator blading is disclosed. The turbo-engine stator blading includes a crown of fixed blades mounted on a ring. Each blade includes a platform and each blade is fastened to the ring. The ring includes a plurality of individual receptacles for the each of the platforms, the receptacles being machined in the thickness of the ring, the form of each receptacle being complementary to that of the corresponding platform. By means of the receptacles, the mounting of the blades within the blading is both simple and accurate.

Abstract: A bling nozzle/carrier interface design for a steam turbine is provided wherein there is limited contact between the carrier (casing, shell) to allow for ease of assembly. There is also limited contact in strategic locations to both reduce the roll, or downstream deflection, of the inner portion and to improve disassembly due to limited contact in areas where corrosion can occur.