Abstract: A stationary blade cascade 29 of an embodiment includes stationary blade structures 50 and a ring-shaped support structure 40 supporting the stationary blade structures 50. The stationary blade structures 50 each include: a stationary blade part 51 where steam passes; and an outer circumference side constituent part 52 formed on an outer circumference side of the stationary blade part 51 and having a fitting groove 56 which penetrates all along a circumferential direction and which has an opening 55 all along the circumferential direction in a downstream end surface 54 of the outer circumference side constituent part 52. The support structure 40 includes a ring-shaped support part 42 having a fitting portion 41 fitted in the fitting grooves 56 of the outer circumference side constituent parts 52. The plural stationary blade structures 50 are supported along the circumferential direction by the ring-shaped support part 42.

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 non-metallic vertical turbine pump and a method of manufacturing same are provided. In accordance with one embodiment, the vertical turbine pump can include a motor, and a drive shaft coupled to the motor for rotation, a casing with a mixed flow diffuser including a diffuser hub and diffuser vanes that project out from the diffuser hub, and a mixed flow impeller configured to rotatably fit inside the casing, having an impeller hub, impeller vanes, extending from the hub, and a front and a back shroud connected to opposing ends of the impeller vanes. At least one of the mixed flow impeller and the mixed flow diffuser can be made of a non-metallic material and is a singular homogenous component.

Abstract: A quick engine change assembly, comprises a first circular frame member, a plurality of doublet supports spaced about the first circular frame member, the doublet supports being contoured along the axial direction, a flow surface defined between the plurality of doublet supports and, a plurality of cradles, each of the cradles including the doublet supports, the doublet supports from a lower portion of the cradle to the flow surface.

Abstract: An integrally bladed rotor has an outer rim with a plurality of blades extending radially outwardly of the outer rim. A plurality of channels are formed radially inwardly of the outer rim. A discontinuity formed at a radially outer surface of the outer rim includes a first thin slot at a radially outer face of the outer rim with an enlarged seal holding area. A second thin slot is positioned radially inwardly of the seal holding. The first and second thin slots are thinner circumferentially than the enlarged seal holding area. A seal is inserted into the seal holding area. The seal does not extend into the first and second thin slots, nor into the channels.

Abstract: A damped assembly within a gas turbine, comprising a first member and a second member mechanically connected via an intermediate vibration dampener which has at least first and second surfaces for contacting the first and second members respectively, wherein either or both of the first and second surfaces include a plurality of raised portions in direct contact with the respective member and a plurality of recessed portions which contain adhesive for bonding the intermediate vibration dampener to the respective member.

Abstract: Methods of forming ceramic matrix composite structures include joining at least two lamina together to form a flexible ceramic matrix composite structure. Ceramic matrix composite structures include at least one region of reduced inter-laminar bonding at a selected location between lamina thereof. Thermal protection systems include at least one seal comprising a ceramic matrix composite material and have at least one region of reduced inter-laminar bonding at a selected location between lamina used to form the seal. Methods of forming thermal protection systems include providing one or more such seals between adjacent panels of a thermal protection system.

Abstract: An exhaust plenum for connection to an exhaust outlet portion of a gas turbine for receiving turbine exhaust gas. The exhaust plenum includes a first wall. The first wall includes a seal attachment component in thermal contact with an exterior shell of the first wall and is adapted for attaching an annular flex seal to the first wall. The seal attachment component extends generally around a perimeter of the inlet of the exhaust plenum, and is thermally insulated from an interior space of the exhaust plenum.

Abstract: A seal assembly for an exhaust duct section of a gas turbine engine includes a multiple of springs that axially extend from a split hoop opposite a seal surface.

Abstract: One aspect of present application provides an intermediate structure in a gas turbine engine. The intermediate structure is positioned between a first component and another component. The first component may be a composite component. The components may be interlocking. The intermediate structure may be load bearing. Also disclosed is a method using the intermediate structure.

Abstract: A first ceramic matrix composite structure having a first flange with a first aperture, and a second structure having a second flange with a second aperture are clamped together by a fastener and spring assembly. The second flange is in contact with the first flange and the apertures of the first and second flanges align with each other. A threaded fastener having a thermal expansion rate that is different than the thermal expansion rate of at least one of the first ceramic matrix composite structure and second structure extends through the aligned apertures. The fastener includes a bolt, a nut, and a spring assembly that maintains a constant clamping pressure irrespective of the thermal expansion differences between the bolt and the first ceramic matrix composite structure and second structure.

Abstract: A sealing assembly is provided for sealing against a component. The sealing assembly comprises a seal carrier and a continuous seal element that is attached to the seal carrier at at least two discrete positions of attachment. The seal element is provided with a sealing surface that is arranged to seal against the component. The seal element is further provided with a thermal expansion slot that is located between the discrete positions of attachment and which extends from a first edge of the seal element over a portion of the width of the seal element. This allows the seal carrier and seal element to thermally expand in the length direction at different rates.

Abstract: A system includes a gas turbine engine, a nozzle structure located adjacent to the gas turbine engine, a flexible annular seal linking the gas turbine engine and the nozzle structure, the flexible annular seal having a plurality of corrugate sections, and a reinforcement member positioned between two of the plurality of corrugate sections.

Abstract: To provide a thermal insulation structure for a structural member and a scroll structure which can be easily used even in integrally constructed structural members and which can ensure reliability and thermal insulation properties. Provided are securing portions (52 and 53) extending outward from an outer wall of the structural member; a press plate (55) attached to the securing portions (52 and 53), with a gap between the press plate and the outer wall; and a thermal insulation member (56) disposed between the outer wall and the press plate (55).

Abstract: An arrangement for optimizing the running clearance for turbomachines of the axial type, such as turbocompressors, gas turbines, and steam turbines, in particular for compressors of stationary gas turbines, by controlling the inner diameter, which is relevant to the running clearance, of at least one stator structure that surrounds a rotor blade ring, including: the stator structure has a closed, circular inner ring, a circular outer ring that is situated concentric to the inner ring at a radial distance therefrom, and a plurality of links that integrally connect the inner ring to the outer ring, the links being circumferentially inclined at a defined angle (?) to the radial direction and distributed around the circumference of the stator structure, and the arrangement includes an adjustment device for rotating the inner ring relative to the outer ring with elastic modification of the running clearance-relevant inner diameter (D) of the inner ring.

Abstract: An expansion joint that is provided between a steam turbine outlet and a condenser inlet, including: an upstream baffle tube that has one end secured to an upstream fixing end and the other end as a free end, and forms a steam channel; a downstream baffle tube that has one end secured to a downstream fixed end and the other end as a free end, the free end being relatively movable outside the free end of the upstream baffle tube; and a flexible cylinder that is formed of a non-metal material using a resin sheet, and deformable while airtightly surrounding the baffle tubes.

Abstract: A blade outer air seal with ring segments forming a mate face gap in which a riffle seal is placed to seal the axial gap. The riffle seal includes a horizontal plate and a vertical plate extending from a bottom surface of the horizontal plate and occupies the axial gap space. The bottom end of the vertical extending plate is angled in a direction of rotation of rotor blades and includes ribs that form open slots in the riffle seal. The ring segments include metering holes that discharge cooling air into the slots to discharge film cooling air onto the hot gas surface of the BOAS and cool the ring segments.

Abstract: A flexible seal having an X-shaped cross section that forms four contact points on four contact surfaces of two opposed seal slots. The flexible seal is used for a component in which the two seal slots undergo a large deflection such that the opposed slots are not aligned and a rigid seal will not form an adequate seal. The flexible seal can be used in a component of a combustor or a turbine in a gas turbine engine where opposed seal slots undergo the large deflection during operation.

Abstract: A method of assembling a heat exchanger in a gas turbine engine assembly is provided. The method includes providing a structural body including a wall having an elongated slot, providing a heat exchanger operably associated with a bracket, and providing a slip joint for coupling the bracket to the structural body. The slip joint includes a standoff, an elongated member, and a biasing member disposed about the elongated member. The standoff is inserted through the wall so that the biasing member imparts a predetermined biasing force between the bracket and the structural body. The slip joint allows relative sliding movement between the bracket and the structural body when a sliding force due to thermal expansion or contraction is greater than the biasing force.

Abstract: A system includes a gas turbine engine, a nozzle structure located adjacent to the gas turbine engine, a flexible annular seal linking the gas turbine engine and the nozzle structure, the flexible annular seal having a plurality of corrugate sections arranged about a centerline, and a connection device secured to the gas turbine engine and to the nozzle structure. The connection device is adjustable to select a length of the flexible annular seal measured along the centerline.

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: 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: In one embodiment, a system includes a seal having a first sealing end portion, a second sealing end portion, and an intermediate portion between the first and second sealing end portions, wherein the seal has at least one metering hole configured to control a leakage flow across the seal.

Abstract: The invention relates to a compressor, particularly a high pressure compressor, of a gas turbine, particularly of a gas turbine aircraft engine, having at least one rotating blade ring on the rotor side, and having at least one guide blade ring on the stator side, wherein the, or each, guide blade ring is formed of a plurality of guide blade segments, and wherein each guide blade segment is formed of a plurality of individual blades. According to the invention, adjacent individual blades (11, 12; 12, 13; 13, 14) are permanently connected to each other within each guide blade segment (10) of at least one guide blade ring on opposing surfaces positioned radially outward, whereas the same are not connected to each other on opposing surfaces positioned radially inward.

Abstract: A turbine includes a static member, a moveable member fluidly connected to the static member, and a flexible seal. The flexible seal includes a spring component having a base section mounted to one of the static member and the moveable member of the turbine and a plurality of biasing members. Each of the plurality of biasing members includes a first portion extending from the base section and a cantilevered portion. The flexible seal further includes a seal component provided on the cantilevered portion of each of the plurality of biasing members. The seal component is biased into contact with the other of the static member and the moveable member.

Abstract: A turbine ring assembly for a gas turbine includes a one-piece split ring (10) of ceramic matrix composite (CMC) material, a CMC wedge-shaped part (20) having flanks in contact with the ends of the ring, on either side of the split, so as to close the ring, and an annular metal support structure (40) surrounding the CMC ring and in contact therewith over the major fraction of its outline, the CMC ring being mounted with prestress in the metal structure, at least one element (26) exerting a resilient return force on the wedge-shaped part to keep it in contact with the ends of the CMC ring when the split opens under the effect of differential expansion between the annular metal structure and the CMC ring, and at least one element for preventing the CMC ring from turning about its axis.

Abstract: A turbine nozzle segment includes a band having a plurality of tabs, an airfoil extending from the band and a support structure attached to the tabs. The support structure has a plurality of biasing structures.

Abstract: A bi-metallic strip seal used to seal the gap between adjacent shroud segments in a gas turbine engine. The bi-metallic strip seal is formed from one or more layers of bi-metallic materials with different coefficients of thermal expansion. In a cold state the bi metallic strip seal will fit within the slot easily, while in the hot state the strip seal will expand to fit tightly within the slot and block leakage flow. The strip seal can be formed with the bi-metallic layers oriented similarly or oppositely in order to control the seal force and deflection properties. In one embodiment, the strip seal is formed from four bi-metallic layers in which the middle is bonded while the adjacent layers are free from shear.

Abstract: A guide vane for a gas turbine includes a vane body having a leading edge and a trailing edge and a shroud extending at least between the leading edge and the trailing edge. The shroud has a first side wall and a second side wall extending essentially radially and in a longitudinal direction of the gas turbine. The first side wall has a groove disposed in a region of the trailing edge extending in a longitudinal direction of the shroud and is configured to receive a sealing plate. The first side wall also has a clearance extending from the groove in a region of the trailing edge, wherein a depth of the clearance in a circumferential direction of the gas turbine is equal to a depth of the groove, and wherein a width of the clearance in the longitudinal direction of the shroud is between one to three times the depth of the clearance.

Abstract: An impeller with an axial inlet in relation to an impeller axis for pressurizing a fluid has: a nose section that has a nose section central hub and multiple nose vane sections, each nose vane section extending from a leading edge to a generally radial nose section interface plane; a tail section that has a tail section central hub and multiple tail vane sections, each tail vane section extending from a generally radial tail section interface plane to a vane tip; a coupling that joins the nose section central hub to the tail section central hub and aligns each one of the nose vane sections with a corresponding one of the tail vane sections with an axial gap between the radial nose section interface plane and the radial tail section interface plane.

Abstract: It is an object of the present invention to provide a sealing apparatus having high sealing performance, entailing a low cost, and satisfactory in assembly characteristics. The sealing apparatus is a differential pressure-responsive sealing apparatus which is inserted between seal grooves formed in the side surfaces of outlet flange portions of adjacent combustor transition pipes to seal clearance between the side surfaces, comprising: a minimum required number of seal pieces connected in the longitudinal direction for imparting flexibility, the seal piece comprising a heat resistant, wear resistant material, and having a pair of sealing ridges each having an arcuate surface contacting the wall surface of each seal groove.

Abstract: A vane assembly for a gas turbine engine comprising a number of radial loading elements disposed between lugs of the vane ring and the vane support, such as to generate a radial load force against the vane ring. The radial load force prevents unwanted relative movement between the vane ring and the vane support during operation of the gas turbine engine.

Abstract: A turbomachine system includes a compressor having a compressor intake and a compressor extraction outlet, and an inlet system fluidly connected to the compressor intake and the compressor extraction outlet. The inlet system includes a plenum having a first end portion that extends to a second end portion through an intermediate portion. The inlet system also includes a heating system having a plurality of conduits extending horizontally through the intermediate portion of the plenum and arranged in a vertical relationship. Heated air from the compressor extraction outlet passes through the plurality of conduits and raises a temperature of ambient air passing through the plenum and into the compressor intake.

Abstract: An inspection port plug device is provided for sealing three or more oppositely placed inspection ports in a multi-chambered gas turbine engine. An embodiment of the inspection port plug device includes a cap and at least two shafts coupled end-to-end. A first shaft includes a first end coupled to the cap and a second end having a first sealing plug, wherein the first sealing plug includes a recess. A second shaft includes a third end coupled to the first sealing plug within the recess and a fourth end having a second sealing plug. The first shaft and the second shaft also include a first biasing mechanism and a second biasing mechanism, respectively, for maintaining the seals at the first and second sealing plugs when the inspection port plug device is installed.

Abstract: A compressor has a rotor (13), which rotor (13) has a number of rows of rotor blades (14, 14?) which are at a distance from one another, one behind the other in the axial direction, with a number of thermal barrier segments (10), which are detachably attached to the rotor (13) and are mounted such that they can move in the circumferential direction, being arranged one behind the other in the circumferential direction on the circumference of the rotor (13), and with securing devices (15) being provided on the thermal barrier segments (10), which secure the thermal barrier segments (10) against being moved in the circumferential direction. In the case of such a rotor (13), production and assembly are simplified in that only some of the thermal barrier segments (10) are equipped with the securing devices (15).

Abstract: A transition channel between a first turbine section and a second turbine section for a gas turbine engine is disclosed. The channel includes a first radially external annular wall including ring sector elements housed inside an annular ring and orifices for injecting a fluid into the channel in order to re-energize the boundary layer thereof, a second radially internal annular wall, and a fluid supply device arranged between the space outside the ring and the injection orifices.

Abstract: A seal assembly (60) for sealing first and second static parts (40A, 40B) with opposed slots (44A, 44B) together forming a pocket (46). The seal assembly (60) comprises a floor strip (70) positioned adjacent the pocket's floor (48) and a ceiling strip (80) positioned adjacent the pocket's ceiling (50). A connection (90) of the strips' central portions (72, 82) provides a flat hinge between cooperating hinge portions (76, 86). A first set of the hinge portions (76A-86A) converge for insertion into the first slot (44A) and a second set of hinge portions (76B-96B) converge for insertion into the second slot (44B). After such slot insertion and pocket installation, the converged hinge portions (76A-86A, 76B-86B) resiliently diverge within the pocket to seat the floor strip's end portions (74A, 74B) against the pocket's floor (48) and to seat the ceiling strip's end portions (84A, 84B) against the pocket's ceiling (50).

Abstract: The invention relates to a charging device, such as an exhaust gas turbocharger for a motor vehicle, with a shaft carrying a compressor wheel, with a fixing device for fixing the compressor wheel on the shaft, with at least one component fixed on the shaft together with the compressor wheel by the fixing device. It can be achieved that heat expansion of the compressor wheel through the expansion coefficient (?3) of at least one component can be at least partially offset. Because of this it is possible to increase the initial clamping force with which the components are positioned or clamped on the shaft without the permissible surface pressure for the clamp components being exceeded.

Abstract: The invention relates to a pump module (2) comprising a housing part (4) with precisely determined force introduction points (25) in relation to adjacent components. Supporting elements (24, 27) and a groove (26) are located between the force introduction points (25) and a disc element (20) that comprises a partially annular channel (10). Forces that are introduced into the housing parts (4) thus lead to negligible deformations in the areas of said housing parts (4) that are adjacent to the impeller.

Abstract: The interturbine duct (ITD) directs hot combustion gases from a high pressure turbine stage outlet to a low pressure turbine stage inlet in a gas turbine engine. The ITD comprises an outer wall and a heat resistant non-metallic resilient annular seal provided around an end of the outer wall that is adjacent to the high pressure turbine stage outlet.

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: The present invention provides a mixer assembly including a mixer and a plurality of scoops attached around a circumference of the mixer element via a combination of sliding brackets and fixed brackets. The sliding brackets provide attachment and allow relative movement between the plurality of scoops and the mixer element while the fixed brackets provide only attachment between the mixer element and the plurality of scoops. The first sliding bracket includes a first flange that is fixedly mounted to the mixer element and a second flange that is biased against a surface of the scoop. A second sliding bracket is fixedly mounted to the mixer element and adjustably mounted to the scoop. The second sliding bracket includes an attachment hole and the scoop includes an adjustment slot. A fastener assembly including a spacer is received within the adjustment hole and the adjustment slot to allow relative movement between the mixer element and the scoop.

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 ceramic seal element (200) for use in a turbo-machine comprises a first, rigid portion (210) comprising ceramic fibers (212) bound within a ceramic matrix binder (214), a second, flexible portion (220) comprising ceramic fibers (222), and a third, rigid portion (230) comprising ceramic fibers (232) bound within a ceramic matrix binder (234). Ceramic fibers (222) retain a desired flexibility because they are not bound in ceramic matrix binder. In some embodiments the ceramic fibers (212, 222, 232) are stacked as horizontally disposed layers (225). Also, the fibers (212, 222, 232) of any layer (225) comprise bundles of fibers wherein some of the bundles extend continuously across portions (210, 220, 230). An alternative sealing element (300) comprises a first, rigid portion (310) comprising ceramic fibers (312) that are bound within a ceramic matrix binder (314), and a second, flexible portion (320) that comprises ceramic fibers (313) that retain a desired flexibility. Methods of manufacture are disclosed.

Abstract: A high temperature seal that may include an interior region comprising a material that swells when exposed to a high temperature environment and at least one projection that is able to be inserted into a slot, and a coating that substantially prevents swelling to coated areas of the interior region. The coating may be disposed upon an exterior surface of the interior region such that a distal end of at least one projection remains at least partially uncoated and the remainder of the interior region is substantially covered by the coating. The interior region may include a ceramic matrix composite.

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: Primary objective of the present invention is to effect a seal against a variable clearance formed between assembly components which admit flows of high temperature fluids or are subjected to vibrations. Another objective is to reduce manufacture cost thereof.

Abstract: A fuel pump unit comprising a housing part (4) arranged between an electric motor (2) and a fuel pump (3). The housing part is provided with expansion joints for compensating thermal expansions or swelling of the material of the housing part (4). The gap between the impeller (6) and the housing parts (4, 5) of the fuel pump (3) can be kept constant by virture of the expansion joints, whereby the fuel pump (3) is essentially prevented from jamming especially when it runs dry.

Abstract: A blade gap control system configured to move a blade ring of a turbine engine relative to a blade assembly to reduce the gaps between the tips of the blades and the blade rings to increase the efficiency of the turbine engine is provided. The blade rings can be at an acute angle with respect to the rotational axis of the blade assembly. The axial movement of the blade ring can be done by a pressure differential supplied across the blade ring, the thermal expansion and/or contraction of a linkage or by a piston.

Abstract: A method for assembling a gas turbine engine is provided. The method comprises coupling a first turbine nozzle within the engine, coupling a second turbine nozzle circumferentially adjacent the first turbine nozzle such that a gap is defined between the first and second turbine nozzles and providing at least one spline seal including a substantially planar body. The method also comprises forming at least one catch to extend outward from the body portion of the at least one spline seal, and inserting the at least one spline seal into a slot defined in at least one of the first and second turbine nozzles to facilitate reducing leakage through said gap, such that a portion of the at least one spline seal is received within a recess defined within the turbine nozzle slot to facilitate retaining the at least one spline seal within the turbine nozzle slot.