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Небесная энциклопедия

Космические корабли и станции, автоматические КА и методы их проектирования, бортовые комплексы управления, системы и средства жизнеобеспечения, особенности технологии производства ракетно-космических систем

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Мониторинг СМИ

Мониторинг СМИ и социальных сетей. Сканирование интернета, новостных сайтов, специализированных контентных площадок на базе мессенджеров. Гибкие настройки фильтров и первоначальных источников.

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Поддерживает ввод нескольких поисковых фраз (по одной на строку). При поиске обеспечивает поддержку морфологии русского и английского языка
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Применить Всего найдено 5342. Отображено 100.
01-03-2012 дата публикации

Turbine nozzle with contoured band

Номер: US20120051900A1
Автор: Ching-Pang Lee, Jeffery Donald Clements, Vidhu Shekhar Pandey
Принадлежит: General Electric Co

A turbine nozzle includes an array of turbine vanes between inner and outer bands. Each vane includes opposed pressure and suction sides extending between opposed leading and trailing edges. The vanes define a plurality of flow passages each of which is bounded between the inner band, the outer band, and adjacent first and second vanes. A surface of the inner band in each of the passages is contoured in a non-axisymmetric shape including a peak of relatively higher radial height adjoining the pressure side of the first vane adjacent its leading edge, and a trough of relatively lower radial height is disposed parallel to and spaced-away from the suction side of the second vane aft of its leading edge. The peak and trough define cooperatively define an arcuate channel extending axially along the inner band between the first and second vanes.

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Номер записи: 1
03-05-2012 дата публикации

Blade arrangement, especially stator blade arrangement

Номер: US20120107111A1
Автор: Sejko Kolev, Ulrich Wellenkamp
Принадлежит: Alstom Technology AG

A blade arrangement on a wall of a fluid flow path of a turbomachine includes a locating channel disposed in the wall and having undercut flanks. A plurality of glades each having a blade-side root section corresponding to the undercut flanks a plurality of intermediate filling pieces correspond to the undercut flanks and separate each of the blade-side root sections. The plurality of blade-side root sections and intermediate filling pieces are seated in a form-fitting manner and fixed in a force-locking manner in the locating channel. The plurality of intermediate filling pieces fill out the locating channel so as to form a gas flow-side surface essentially free of a disturbing contour between the blade-side root sections. Each of the plurality of intermediate filling pieces includes a threaded hole disposed essentially perpendicular to a base of the channel. A screw is disposed in the threaded hole.

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Номер записи: 2
07-06-2012 дата публикации

Steam turbine singlet interface for margin stage nozzles with pinned or bolted inner ring

Номер: US20120141260A1
Автор: Dominick Joseph Werther, Steven Sebastian Burdgick
Принадлежит: General Electric Co

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.

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Номер записи: 3
12-07-2012 дата публикации

Gas Turbine Nozzle Arrangement and Gas Turbine

Номер: US20120177489A1
Автор: Stephen Batt
Принадлежит: SIEMENS AG

A sealing element is provided for sealing a leak path between a radial outer platform of a turbine nozzle and a carrier ring for carrying said radial outer platform. The carrier ring has an axially facing carrier ring surface and the radial outer platform has an axially facing platform surface. The carrier ring surface forms a first sealing surface and the platform surface forming a second sealing surface. The first and second sealing surfaces is aligned in a plane with a radial gap between them. The sealing element includes a leaf seal adapted to cover the gap between the first and second sealing surfaces, and an impingement plate for allowing impingement cooling of a radial outer surface of the radial outer platform. The impingement plate is adapted to be fixed to the turbine nozzle. The sealing element may be part of a nozzle arrangement of a gas turbine.

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Номер записи: 4
19-07-2012 дата публикации

Aerofoil blade for an axial flow turbomachine

Номер: US20120183411A1
Автор: Brian Robert Haller
Принадлежит: Alstom Technology AG

An exemplary aerofoil blade for an axial flow turbomachine has a radially inner platform region, a radially outer tip region, an axially forward leading edge, and an axially rearward trailing edge. The aerofoil blade has a pressure surface which is convex in a radial direction, and a suction surface which is concave in the radial direction. The axial width (W) of the aerofoil blade can vary parabolically between maximum axial widths (W max ) at the platform and tip regions, respectively and a minimum axial width (W min ) at a position between the platform region and the tip region.

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Номер записи: 5
26-07-2012 дата публикации

Axial flow turbine

Номер: US20120189441A1
Автор: Brian Robert Haller, Gursharanjit Singh
Принадлежит: Alstom Technology AG

An axial flow turbine includes in axial flow series a low pressure turbine section and a turbine exhaust system. The low pressure turbine section includes a final low pressure turbine stage having a circumferential row of static aerofoil blades followed in axial succession by a circumferential row of rotating aerofoil blades. Each aerofoil blade has a radially inner hub region and a radially outer tip region. The K value, being equal to the ratio of the throat dimension (t) to the pitch dimension (p), of each static aerofoil blade of the final low pressure turbine stage varies along the height of the static aerofoil blade, between the hub region and the tip region, according to a substantially W-shaped distribution.

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Номер записи: 6
02-08-2012 дата публикации

compressor nozzle stage for a turbine engine

Номер: US20120195745A1
Автор: Patrick Edmond Kapala
Принадлежит: SNECMA SAS

A single-piece compressor nozzle stage for a turbine engine, the stage comprising two coaxial rings, connected together by radial vanes, the inner ring including an annular cavity for housing damper means for damping vibration by friction, which damper means are secured to an annular abradable-material support.

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Номер записи: 7
23-08-2012 дата публикации

Gas turbine engine component

Номер: US20120213634A1
Автор: Rickard Samuelsson, Robert Reimers
Принадлежит: Volvo Aero Corp

In a gas turbine engine at least one vanes is made of a composite material and an end part of the vane is fastened to the ring element by at least a first bracket member arranged in a corner region formed by the vane and a ring element. The first bracket member is fastened to the vane by a first fastening device arranged in association with a first hole extending through the vane. The first bracket member is fastened to the ring element by a second fastening device arranged in association with a second hole extending through the ring element. An opening is arranged in the ring element. A second bracket member at least partly has an angular cross section. The second, angled bracket member is arranged through the opening such as to be positioned on an opposite side of both the vane and the ring element in relation to the first bracket member. The second, angled bracket member is fastened to the first bracket member by the first and second fastening devices.

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Номер записи: 8
30-08-2012 дата публикации

Stationary vane unit of rotary machine, method of producing the same, and method of connecting the same

Номер: US20120219412A1
Автор: Takeki NAKAYAMA, Yuki Yamamoto
Принадлежит: Individual

The stationary vane unit of a rotary machine includes: a first band member that extends in the circumferential direction and comes into contact with the outer shrouds of the plurality of stationary vane members from one side thereof in the main axial direction in which a central axis extends; a second band member that extends in the circumferential direction and comes into contact with the outer shrouds of the plurality of stationary vane members from the other side thereof in the main axial direction; and a fastening member that fastens the first band member and the second band member to each other so that the outer shrouds of the plurality of stationary vane members are connected to each other.

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Номер записи: 9
06-09-2012 дата публикации

Airfoil core shape for a turbomachine component

Номер: US20120224954A1
Автор: Julianne Edna Baer, Umesh Garg
Принадлежит: General Electric Co

A turbomachine component includes a compressor stator vane having an airfoil core shape. The airfoil core shape includes a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in TABLE 1, and wherein X and Y are distances in inches which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z in inches. The profile sections at the Z distances are joined smoothly with one another to form a complete airfoil core shape.

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Номер записи: 10
01-11-2012 дата публикации

High area ratio turbine vane

Номер: US20120275922A1
Автор: Eunice Allen-Bradley, Norbert Huebner, Thomas J. Praisner
Принадлежит: Individual

A vane for a turbine engine comprises an airfoil section, an inner platform and an outer platform. The airfoil section comprises pressure and suction surfaces extending from a leading edge to a trailing edge. The inner platform is attached to the airfoil section along an inner flow boundary, where the inner flow boundary extends from an upstream inlet region of the vane to a downstream outlet region of the vane. The outer platform is attached to the airfoil section along an outer flow boundary, where the outer flow boundary extends from the inlet region to the outlet region. An area ratio of the outlet region to the inlet region is greater than 2.4.

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Номер записи: 11
14-02-2013 дата публикации

Turbomachine component having an airfoil core shape

Номер: US20130039771A1
Автор: Benjamin Michael Check, Craig Allen Bielek, Hal Feaster Pollard, Sivaraman Vedhagiri
Принадлежит: General Electric Co

A turbomachine component includes a turbine stator nozzle member having an airfoil core shape. The airfoil core shape includes a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in TABLE 1, and wherein X and Y are distances in inches which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z in inches, the profile sections at the Z distances being joined smoothly with one another to form a complete airfoil core shape.

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Номер записи: 12
28-02-2013 дата публикации

Transition channel of a turbine unit

Номер: US20130051996A1
Автор: Kai Koerber, Karl Engel, Martin Hoeger
Принадлежит: MTU AERO ENGINES GMBH

A transition channel for a turbine unit with at least two components is configured as a flow channel from one component of a first pressure to a component of a second pressure. The transition channel has support ribs, extending between envelope surfaces of the transition channel and having a profile that is configured for the deflecting of a flow from an inlet cross section to an outlet cross section of the transition channel. Flow splitter blades are arranged between the support ribs, having a smaller relative profile thickness than the support ribs and/or a shorter axial design depth or profile chord length than the support ribs. Thanks to the integration of the slim and/or short flow splitter blades (tandem blades), it is possible to largely dissipate parasite secondary flows.

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Номер записи: 13
06-06-2013 дата публикации

Alternate shroud width to provide mistuning on compressor stator clusters

Номер: US20130142640A1
Автор: David P. Houston, Richard A. Lomenzo, Robert J. Morris
Принадлежит: United Technologies Corp

A stator for a turbo-machine having a plurality of airfoils extending radially therefrom has a base from which the airfoils depend, and slits disposed in the base, each slit disposed adjacent a pair of airfoils, wherein a first set of adjacent slits and a distance between a second set of adjacent slits varies

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Номер записи: 14
13-06-2013 дата публикации

STEAM TURBINE, BLADE, AND METHOD

Номер: US20130149106A1
Автор: Giusti Enrico, Grilli Marco, Imparato Enzo
Принадлежит: NUOVO PIGNONE S.P.A

A stator blade ring comprising a plurality of stator blade modules defining an annular chamber is provided. The plurality of stator blade modules comprises an elongated blade portion comprising a first and a second blade shell portion, a longitudinal passageway, and at least one opening extending through at least one of the first and the second blade shell portion to the longitudinal passageway, an inner portion brazed to a first longitudinal end of the elongated blade portion, wherein the inner portion comprises a through hole forming a portion of the annular chamber, and an inner passageway extending from the through hole to the longitudinal passageway, and an outer portion brazed to a second longitudinal end of the elongated blade portion and engaged to a steam turbine, the outer portion comprising an outer passageway open to a surface of the steam turbine and the longitudinal passageway. 1. A stator blade ring for a steam turbine , the stator blade ring comprising: an elongated blade portion comprising a first blade shell portion, a second blade shell portion brazed to the first blade shell portion, a longitudinal passageway; and at least one opening extending through at least one of the first blade shell portion and the second blade shell portion to the longitudinal passageway;', 'an inner portion brazed to a first longitudinal end of the elongated blade portion, wherein the inner portion comprises a through hole forming a portion of the annular chamber, and an inner passageway extending from the through hole to the longitudinal passageway; and', 'an outer portion brazed to a second longitudinal end of the elongated blade portion and engaged to the steam turbine, wherein the outer portion comprises an outer passageway open to a surface of the steam turbine and the longitudinal passageway., 'a plurality of stator blade modules defining an annular chamber, wherein each of the plurality of stator blade modules comprises2. The stator blade ring of claim 1 , wherein ...

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Номер записи: 15
13-06-2013 дата публикации

Fan Hub Frame for Double Outlet Guide Vane

Номер: US20130149130A1
Автор: Courtney James Tudor, William Howard Hasting
Принадлежит: General Electric Co

A fan hub frame comprises a circular hub having an opening extending axially wherein an engine core is capable of being positioned, the circular hub having a radially outer surface, the radial outer surface having a plurality of cradles, each of the cradles having a lower surface and fillets disposed between the lower surface and upwardly extending sidewalls, the cradles capable of receiving a double outlet guide vane.

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Номер записи: 16
13-06-2013 дата публикации

STATOR VANE ARRAY

Номер: US20130149135A1
Автор: HIELD Paul Michael
Принадлежит: ROLLS-ROYCE PLC

A stator vane assembly for a gas turbine engine includes circumferentially spaced vanes about a common axis. The array of vanes further includes three or more sub-arrays, which are configured such that the vane spacing in one sub-array is different from the vane spacing in the other sub-arrays. 116-. (canceled)17. A vane assembly for a gas flow machine , the vane assembly comprising an array of vanes circumferentially spaced about a common axis , wherein the array of vanes comprises three or more sub-arrays , wherein the vane spacing within one sub-array is different from the vane spacing within the other sub-arrays; wherein each sub-array comprises a plurality of adjacent vanes , each vane of a sub-array being substantially equally spaced from an adjacent vane in that sub-array.18. A vane assembly according to claim 17 , wherein each sub-array extends through a portion of a revolution about said axis and the sub-arrays are arranged in an end to end arrangement such that the array forms a complete revolution about the axis.19. A vane assembly according to wherein a step change in vane spacing occurs upon passage from one sub-array to an adjacent sub-array.20. A vane assembly according to claim 17 , wherein the vane spacing varies in a non-cyclic manner through a single revolution of the array about the axis.21. A vane assembly according to claim 17 , wherein at least a pair of sub-arrays have a vane spacing which differs from an average vane spacing for the array by an equal magnitude.22. A vane assembly according to claim 17 , comprising a casing disposed about the axis claim 17 , wherein each vane of the array depends inwardly from the casing towards the axis from the casing.23. A vane assembly according to claim 17 , comprising a stator vane assembly for a gas turbine engine.24. A gas flow machine comprising:a plurality of rows of rotor blades arranged in serial flow arrangement; anda stator vane assembly disposed in the flow path between said rows of rotor ...

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Номер записи: 17
13-06-2013 дата публикации

Stationary blade cascade, assembling method of stationary blade cascade, and steam turbine

Номер: US20130149136A1
Автор: Itaru Murakami, Motoki Yoshikawa, Tsuguhisa Tashima
Принадлежит: Individual

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.

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Номер записи: 18
11-07-2013 дата публикации

Turbomachine component including a cover plate

Номер: US20130177408A1
Автор: Aaron Gregory Winn, Michael Gordon Sherman, Niranjan Gokuldas Pai
Принадлежит: General Electric Co

A turbomachine component includes a body having a first end that extends to a second end. One of the first and second ends includes a mounting element, and a mounting component. A cover plate is arranged at the one of the first and second ends to establish an interface region. The cover plate includes a mounting member configured to align with the mounting element, and a mounting portion configured to align with the mounting element. A fastener member is configured and disposed to cooperate with the mounting element and the mounting member to constrain the cover plate to the body along at least two axes with the interface region being devoid of a metallurgical bond.

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Номер записи: 19
01-08-2013 дата публикации

NOISE-REDUCED TURBOMACHINE

Номер: US20130195610A1
Автор: ROSE Marco, Willer Lars
Принадлежит: ROLLS-ROYCE DEUTSCHLAND LTD & CO KG

Turbomachine with an annular main flow duct () through which passes a flow, and in which is arranged at least one stator provided with stator vanes (), characterized in that the stator vanes () each have at least one recess () issuing into a flow duct () inside the stator vanes () and that the flow duct () issues into a bypass duct () of the turbomachine via at least one outflow duct () provided with a shut-off element (). 1. A turbomachine , comprising:an annular main flow duct through which passes a flow;a stator having a plurality of stator vanes arranged in the main flow duct, the stator vanes each having a recess;a flow duct inside the stator vane, the recess issuing into the flow duct;an outflow duct, the flow duct issuing into a bypass duct of the turbomachine via the outflow duct;a shut-off element for shutting-off flow through the outflow duct.2. The turbomachine of claim 1 , wherein the stator is a guide vane.3. The turbomachie of claim 2 , wherein the recess is has a slot form.4. The turbomachine of claim 2 , wherein the recess is formed as a line-type row of holes.5. The turbomachine of claim 2 , wherein the recess s located on a suction side of the stator vane.6. The turbomachine of claim 5 , wherein the recess is arranged adjacent to a stator profile trailing edge of the stator vane.7. The turbomachine of claim 6 , wherein the recess extends substantially over an entire radial length of the stator vane.8. The turbomachine of claim 7 , and further comprising a plurality of recesses arranged at at least one chosen from equal and varying spacing to one another.9. The turbomachine of claim 8 , wherein the at least one outflow duct is connected to an annular duct extending in the circumferential direction relative to a machine axis claim 8 , and the flow ducts issue into the annular duct10. The turbomachine of claim 9 , and further comprising a diffuser claim 9 , the outflow duct issuing into the bypass duct via the diffuser.11. The turbomachine of claim 10 ...

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Номер записи: 20
01-08-2013 дата публикации

Stress relieving slots for turbine vane ring

Номер: US20130195643A1
Автор: Douglas MacCaul, John Pietrobon, Keppel Nyron Bharath, Vincent Paradis
Принадлежит: Pratt and Whitney Canada Corp

A turbine vane ring has a 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.

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Номер записи: 21
15-08-2013 дата публикации

Cooled vane of a turbine and corresponding turbine

Номер: US20130209230A1
Автор: Andy Pacey, Stephen Batt
Принадлежит: SIEMENS AG

A vane is provided for use in a fluid flow of a turbine engine. The vane includes a thin-walled radially extending aerodynamic vane body having axially spaced leading and trailing edges, and a radially outer platform. The wall of the vane body includes an outer shell and an inner shell and defines an interior cavity therein for flowing a cooling medium. A radially extending load strut is arranged at the inner shell of the wall of the leading edge of the vane body.

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Номер записи: 22
15-08-2013 дата публикации

Nozzle guide vane with cooled platform for a gas turbine

Номер: US20130209231A1
Автор: Anthony Davis, Paul Mathew Walker
Принадлежит: SIEMENS AG

A platform for supporting a nozzle guide vane for a gas turbine is provided. The platform has a gas passage surface arranged to be in contact with a streaming operation gas, and a cooling channel for guiding a cooling fluid within the cooling channel formed in an inside of the platform. A cooling portion of an inner surface of the cooling channel is in thermal contact with the gas passage surface. The platform is an integrally formed part representing a segment in a circumferential direction of the gas turbine. The cooling channel has a first cooling channel portion and a second cooling channel portion arranged downstream of the first cooling channel portion with respect to a streaming direction of the operation gas. The first cooling channel portion and the second cooling channel portion are interconnected.

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Номер записи: 23
15-08-2013 дата публикации

Anti-Rotation Stator Segments

Номер: US20130209248A1
Автор: Jonathan J. Earl, Mark David RING
Принадлежит: Pratt and Whitney Co Inc

A stator assembly for a turbofan gas turbine engine is disclosed. The stator assembly is coupled to a shroud of the engine. The stator assembly includes an endless case fixedly coupled to the engine shroud. The case includes a forward portion, an aft portion and a central portion disposed therebetween. The case extends about an axis of the engine. The forward and aft portions of the case include rails that extend towards each other and form forward and aft pockets with the central portion respectfully. The stator assembly also includes a locking stator segment. The locking stator segment includes a shroud that includes a forward hook and a pair of aft hooks with a platform disposed therebetween. The forward and aft hooks are retained in the forward and aft pockets of the case respectfully.

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Номер записи: 24
26-09-2013 дата публикации

TURBOJET CASING AND TURBOJET RECEIVING SUCH CASINGS

Номер: US20130251519A1
Автор: Desjoyeaux Bertrand, Dunleavy Patrick, Masson Richard
Принадлежит:

A turbojet casing adapted to receive a plurality of vanes, the casing including attachment means () for attaching one end of each vane to the casing, the casing being characterized in that the attachment means extend on a face of the casing facing away from the vanes, the casing including orifices () for passing the ends of the vanes so that they can co-operate with the attachment means of the casing. A turbojet including such a casing. 1213026. A turbojet casing adapted to receive a plurality of vanes , the casing including attachment means ( , ) for attaching one end of each vane to the casing , the casing being characterized in that the attachment means extend on a face of the casing facing away from the vanes , the casing including orifices () for passing the ends of the vanes so that they can co-operate with the attachment means of the casing.22130. A casing according to claim 1 , wherein the attachment means comprise an annular member ( claim 1 , ) extending around the casing.3212423. A casing according to claim 2 , wherein the annular member comprises at least one peripheral rail () having the ends () of fastener elements () for fastening to the ends of the vanes inserted therein.430. A casing according to claim 2 , wherein the annular member comprises a peripheral angle bar () to which the ends of the vanes are fastened directly.52130. A casing according to claim 2 , the casing being made of long fibers associated with a thermoplastic resin claim 2 , while the annular member ( claim 2 , ) is obtained by pultrusion and impregnated with a thermoplastic resin that is heat-sealable with the thermoplastic resin of the casing claim 2 , the assembly being joined together by hot compaction.6. A turbojet including at least one casing according to claim 1 , and a plurality of vanes claim 1 , each having one end connected to the casing.7. A turbojet according to claim 6 , wherein each of the vanes comprises:{'b': '2', 'an elongate one-piece front portion () cut from a ...

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Номер записи: 25
03-10-2013 дата публикации

Turbine nozzle

Номер: US20130259703A1
Автор: Hee Koo Moon, Juan Yin, Luzeng ZHANG
Принадлежит: Solar Turbines Inc

A nozzle arrangement for a gas turbine engine comprising a first housing member and a second housing member. The nozzle arrangement may further include a first nozzle and a second nozzle. Each of the first nozzle and second nozzle may extend between the first housing member and the second housing member so as to form a doublet. A plurality of cooling apertures may be arranged on at least one of the first nozzle, the second nozzle, the first housing member, or the second housing member so as to provide a different degree of first order cooling across the doublet.

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Номер записи: 26
24-10-2013 дата публикации

Turbine engine stator and method of assembly of the same

Номер: US20130280054A1
Автор: Lewis J. HOLMES
Принадлежит: Rolls Royce PLC

A turbine engine stator stage includes a plurality of vanes with each of the plurality of vanes having a camber angle. The plurality of vanes is arranged in a plurality of groups with each group including a pre-determined sequence of vanes. The ordering of vanes within each group is determined by the camber of the individual vanes. This results in an arrangement of vanes within the stator stage which can modify the flow characteristics of the air entering the stator stage to reduce the circumferential pressure variation in the flow region immediately downstream of the stator stage.

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Номер записи: 27
31-10-2013 дата публикации

TURBINE DIAPHRAGM CONSTRUCTION

Номер: US20130287563A1
Автор: Brummitt-Brown Angus Robert, Lord Adrian Clifford, Mistry Kanu
Принадлежит: ALSTOM Technology Ltd

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. 2. An axial flow turbine diaphragm according to claim 1 , in which a radially outer port wall of the diaphragm comprises the radially outer elongate platform portions of the blade units claim 1 , alternating in the circumferential direction with exposed portions of the inner circumference of the outer diaphragm ring.3. An axial flow turbine diaphragm according to claim 1 , in which the radially inner platform portions of the blade units are elongate in the circumferential direction of the turbine diaphragm and an outer circumference of the radially inner diaphragm ring is provided with a blade unit retaining feature of complementary shape and orientation to the inner platform portions of the static blade units claim 1 , whereby the inner platform portions are retained to the radially inner ring.4. An axial flow turbine diaphragm according to claim 3 , in which a radially inner port wall of the diaphragm comprises the radially inner elongate platform portions of the blade units claim 3 , flanked on their axially opposed sides by portions of ...

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Номер записи: 28
02-01-2014 дата публикации

Gas turbine engine turbine vane airfoil profile

Номер: US20140000287A1
Автор: Brandon W. Spangler, Russell J. Bergman
Принадлежит: Individual

A turbine vane for a gas turbine engine includes inner and outer platforms joined by a radially extending airfoil. The airfoil includes leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface. The inner and outer platforms respectively include inner and outer sets of film cooling holes, wherein one of the inner and outer sets of film cooling holes are formed in substantial conformance with platform cooling hole locations described by one of the sets of Cartesian coordinates set forth in Tables 1 and 2. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate. The cooling holes with Cartesian coordinates in Tables 1 and 2 have a diametrical surface tolerance relative to the specified coordinates of 0.200 inches (5.08 mm).

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Номер записи: 29
02-01-2014 дата публикации

COMPRESSOR FOR A GAS TURBINE AND METHOD FOR REPAIRING AND/OR CHANGING THE GEOMETRY OF AND/OR SERVICING SAID COMPRESSOR

Номер: US20140003926A1
Автор: CORRADI Mauro, RAUCH Marc, STEFFEN Tobias, Veith Bjoern
Принадлежит:

A compressor for a gas turbine having a rotor rotating about an axis and a housing coaxially surrounding the rotor, thereby defining a flow channel between the rotor and the housing for a fluid to be compressed. Alternating rows of vanes and blades are arranged within the flow channel, with the vanes being mounted in the inner wall of the housing, and the blades being mounted in the outer wall of the rotor. 1. A compressor for a gas turbine , said compressor comprising;a rotor rotating about an axis;a housing coaxially surrounding said rotor, thereby defining a flow channel between rotor and housing for a fluid to be compressed;alternating rows of vanes and blades arranged within the flow channel, the vanes being mounted in the inner wall of the housing, and the blades being mounted in the outer wall of the rotor; andintermediate rings provided in the inner wall of the housing between neighbouring rows of vanes such that the outer wall of the flow channel is formed by said vanes and intermediate rings.2. The compressor according to claim 1 , wherein said intermediate rings are adjacent to the tips of the blades.3. The compressor according to claim 2 , wherein said intermediate rings bear a coating on their inner side just opposite to the tips of the blades.4. The compressor according to claim 3 , wherein said coating is an abradable coating.5. The compressor according to claim 4 , wherein the inner side of the intermediate rings is completely covered by said abradable coating.6. The compressor according to claim 4 , wherein the inner side of the intermediate rings is only covered by said abradable coating only in an area opposite to the tips of the respective blades.7. The compressor according to claim 4 , wherein said abradable coating comprises a honeycomb structure filled with an abradable material.8. The compressor according to claim 1 , wherein the intermediate rings are mounted in circumferential mounting grooves provided in the inner wall of the housing.9. ...

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Номер записи: 30
16-01-2014 дата публикации

STATIC VANE ASSEMBLY FOR AN AXIAL FLOW TURBINE

Номер: US20140017071A1
Автор: Havakechian Said, MCBEAN Ivan William, MEGERLE Benjamin, RICE Timothy Stephen
Принадлежит:

An axial flow turbine is described having a casing defining a flow path for a working fluid therein, a rotor co-axial to the casing, a plurality of stages, each including a stationary row of vanes circumferentially mounted on the casing a rotating row blades, circumferentially mounted on the rotor, with within a stage n vanes have an extension such that at least a part of the trailing edge of each of the n vanes reaches into the annular space defined by the trailing edges of the remaining N-n vanes and the leading edges of rotating blades of the same stage. 1. An axial flow turbine comprising:a casing defining a flow path for a working fluid therein;a rotor co-axial to the casing; a row of N stationary vanes circumferentially mounted on the casing; and', 'a row of rotating blades circumferentially mounted on the rotor,, 'a plurality of stages, each comprisingwherein within a stage, n vanes have an extension such that at least a part of the trailing edge of each of the n vanes reaches into the annular space limited by the rotor and the casing and the trailing edges of the remaining N-n vanes and the leading edges of rotating blades of the same stage,wherein the number n of extended vanes is larger than zero but less than half the total number N of vanes in the stage.2. The turbine according to wherein the stage is a last stage of a low pressure steam turbine.3. The turbine according to wherein the number n is selected to be 0 Подробнее

Номер записи: 31
06-02-2014 дата публикации

Anti-rotation lug for a gas turbine engine stator assembly

Номер: US20140037442A1
Автор: Neil L. Tatman, Richard K. Hayford
Принадлежит: Individual

A stator assembly includes a case including an arcuate wall having an aperture with circumferentially spaced first lateral surfaces. A stator vane has an outer platform with a notch. An anti-rotation lug has a base that is received in the notch and a boss extends from the base. The boss is received in the aperture. The boss has second lateral surfaces that engage the first lateral surfaces in an interference fit relationship.

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Номер записи: 32
06-03-2014 дата публикации

SINGLET VANE CLUSTER ASSEMBLY

Номер: US20140060081A1
Автор: Earl Jonathan J., Kuehne Eric A.
Принадлежит:

A vane cluster for a gas turbine engine includes multiple singlet vanes and a forward wear liner connecting a forward edge of each singlet vane, thereby allowing the vane cluster to be manipulated as a single component. 1. A gas turbine engine vane cluster comprising:a plurality of singlet vanes including an anti-rotation singlet vane and an end vane singlet vane; anda forward wear liner connecting a forward edge of each singlet vane in said plurality of singlet vanes, such that said plurality of singlet vanes is operable to be manipulated as a single component.2. The gas turbine engine vane cluster of claim 1 , wherein said plurality of singlet vanes includes said anti-rotation singlet vane on a first end claim 1 , said end vane singlet vane on an opposite end claim 1 , and at least one intermediate vane between said anti-rotation singlet vane and said end vane singlet vane claim 1 , and wherein each of said singlet vanes interfaces with each adjacent singlet vane in said vane cluster.3. The gas turbine engine vane cluster of claim 1 , further comprising an aft wear liner connecting an aft edge of each singlet vane in said plurality of singlet vanes.4130. The gas turbine engine vane cluster of claim 3 , wherein said anti-rotation singlet vane includes an anti-rotation notch and a forward wear liner retention notch.5430. The gas turbine engine vane cluster of claim 3 , wherein said end vane singlet vane includes a forward retention lug interfaced with said forward wear liner and an aft retention lug interfaced with said aft wear liner.6. The gas turbine engine vane cluster of claim 3 , wherein said forward wear liner comprises an anti-rotation interface feature operable to interface with said anti-rotation singlet vane on a first end of said forward wear liner claim 3 , and an end vane singlet vane interface feature on a second end of said forward wear liner.7. The gas turbine engine vane cluster of claim 6 , wherein said anti-rotation interface feature comprises a ...

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Номер записи: 33
06-03-2014 дата публикации

GUIDE VANE ASSEMBLY

Номер: US20140064956A1
Автор: DRANE Andrew James
Принадлежит: ROLLS-ROYCE PLC

A guide vane assembly for a stator vane stage comprises at least two aerofoil members joined together using a vane assembly attachment web. The vane assembly attachment web may be provided at the inner and/or outer radius of the aerofoil members. The vane assembly attachment web allows the guide vane assembly to be attached to an inner and/or outer attachment ring, thereby forming a stator vane stage. Such an arrangement may allow composite fibre reinforced guide vane assemblies to be readily assembled together to form stator vane stages. Stator vane stages that comprise such composite fibre reinforced guide vane assemblies may be lighter than conventional metallic stator vane stages. 1. A guide vane assembly for a gas turbine engine stator vane stage comprising:at least two adjacent aerofoil members, each extending from a root to a tip, wherein:the adjacent aerofoil members are integral with a vane assembly attachment web formed either between the tips or between the roots of the adjacent aerofoil members, the vane assembly attachment web being arranged for use in fixing the guide vane assembly in the stator vane stage in use;said adjacent aerofoil members and vane assembly attachment web are formed using a fibre-reinforced composite material; andthe guide vane assembly extends from a first end to a second end.2. A guide vane assembly according to claim 1 , further comprising:a first attachment flange extending from the first end to the tip or root of one of the aerofoil members; anda second attachment flange extending from the second end to the tip or root of another one of the aerofoil members.3. A guide vane assembly according to claim 2 , wherein the circumferential extent of the guide vane assembly claim 2 , with respect to a rotational axis of the gas turbine engine in use claim 2 , is defined by the first and second ends.4. A guide vane assembly according to claim 2 , wherein the first and second ends are within the circumferential extent of the guide vane ...

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Номер записи: 34
03-04-2014 дата публикации

LINER AND METHOD OF ASSEMBLY

Номер: US20140093363A1
Автор: Cassella David, Earl Jonathan, Kuehne Eric, Ring Mark David, Warner Charles
Принадлежит: UNITED TECHNOLOGIES CORPORATION

An assembly includes a plurality of vanes, a forward liner segment, and an aft liner segment. The forward liner segment and the aft liner segment are mounted to the plurality of vanes and each segment comprises an arc of less than 360° in length. 1. An assembly comprising:a plurality of vanes each vane having an airfoil, a platform, and forward and aft mounting hooks; anda forward liner segment mounted on the forward mounting hooks of the plurality of vanes; andan aft liner segment mounted to the aft mounting hooks of the plurality of vanes, wherein the forward liner segment and the aft liner segment are arcs of less than 360°.2. The assembly of claim 1 , wherein each liner segment comprises a single-piece segment less than a complete circular ring.3. The assembly of claim 1 , wherein the plurality of vanes are mounted adjacent one another to form a vane pack that comprises an arc that extends substantially 45° about a centerline axis of a gas turbine engine.4. The assembly of claim 1 , wherein the plurality of vanes comprise cantilevered vanes.5. The assembly of claim 1 , wherein the plurality of vanes are mounted adjacent one another to form a vane pack claim 1 , and wherein the vane pack has a first end vane at a first end and an second end vane at a second end.6. The assembly of claim 5 , wherein each liner segment includes one or more slots adapted to receive one or more standups of the first end vane and/or second end vane.7. The assembly of claim 6 , wherein the one or more slots allows at least one of the first end vane or second end vane to be inserted therethrough.8. The assembly of claim 5 , wherein at least one of the forward liner segment and the aft liner segment is disposed at a distance from the first end vane and/or the second end vane.9. The assembly of claim 1 , wherein a first end vane of a first vane pack is adapted to interface with a second end vane of a second vane pack.10. A gas turbine engine comprising:a casing with first and second ...

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Номер записи: 35
03-04-2014 дата публикации

DOUBLE FLOW TURBINE HOUSING TURBOCHARGER

Номер: US20140093364A1
Автор: Alajbegovic Vahidin, Narehood Abraham C., Watson John P.
Принадлежит: BORGWARNER INC.

Implementations of the present disclosure are directed to turbine assemblies for turbocharger systems. In some implementations, turbine housings include a body that defines an inlet for fluid communication with a fluid source, and a wall, the wall dividing the inlet into an inner inlet and an outer inlet, and a fluid guide assembly disposed within the housing, the fluid guide assembly including a plurality of vanes that demarcate an inner volute and an outer volute within the housing, the inner volute being in fluid communication with the inner inlet and the outer volute being in fluid communication with the outer inlet, each vane of the plurality of vanes being fixed at a respective angle relative to a radial direction, the plurality of vanes guiding fluid flow from the outer volute to the inner volute. 1. A turbine housing for a turbocharger , the housing comprising:a body that defines an inlet for fluid communication with a fluid source, and a wall, the wall dividing the inlet into an inner inlet and an outer inlet; anda fluid guide assembly disposed within the housing, the fluid guide assembly comprising a plurality of vanes that demarcate an inner volute and an outer volute within the housing, the inner volute being in fluid communication with the inner inlet and the outer volute being in fluid communication with the outer inlet, each vane of the plurality of vanes being fixed at a respective angle relative to a radial direction, the plurality of vanes guiding fluid flow from the outer volute to the inner volute.2. The turbine housing of claim 1 , wherein the fluid guide assembly further comprises a guide plate that is secured to the body claim 1 , the plurality of vanes being secured to the guide plate.3. The turbine housing of claim 1 , wherein at least one of the vanes is positioned at a selected angle relative to the radial center of the turbine wheel.4. The turbine housing of claim 3 , wherein the selected angle is between approximately 30° and ...

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Номер записи: 36
05-01-2017 дата публикации

GAS TURBINE BLADE

Номер: US20170002665A1
Автор: Facchinetti Emanuele, HENZE Marc, WAGNER Guillaume
Принадлежит: ANSALDO ENERGIA SWITZERLAND AG

A gas turbine blade includes a blade root and a blade aerofoil, a cooling fluid plenum extending inside the gas turbine blade through the blade root, the blade aerofoil and the blade tip, a blade root impingement plate in the cooling fluid plenum inside the blade root and a blade tip impingement plate in the cooling fluid plenum inside the blade tip, the blade tip impingement plate having at least one cooling fluid hole configured and arranged to enable a cooling fluid to flow from the blade tip into the blade aerofoil via the cooling fluid hole or holes, and a pipe extending in the cooling fluid plenum from the blade root impingement plate to the blade tip impingement plate. The blade root impingement plate can direct the cooling fluid from the blade root to the pipe. 1. A gas turbine blade comprising:a blade root and a blade aerofoil, the blade root being attached to a first end of the blade aerofoil;a blade tip attached to a second end of the blade aerofoil;a cooling fluid plenum extending inside the gas turbine blade through the blade root, the blade aerofoil and the blade tip;a blade root impingement plate in the cooling fluid plenum inside the blade root and a blade tip impingement plate in the cooling fluid plenum inside the blade tip, the blade tip impingement plate having at least one cooling fluid hole configured and arranged to enable a cooling fluid to flow from the blade tip into the blade aerofoil via the cooling fluid hole or holes; anda pipe extending in the cooling fluid plenum from the blade root impingement plate to the blade tip impingement plate, and the pipe being configured and arranged to transport the cooling fluid from the blade root to the blade tip; andthe blade root impingement plate being configured and arranged to direct the cooling fluid from the blade root to the pipe.2. The gas turbine blade of claim 1 , wherein the pipe is attached to the blade tip impingement plate and slidably attached to the blade root impingement plate claim 1 ...

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Номер записи: 37
05-01-2017 дата публикации

Mechanical fastening system for rotating or stationary components

Номер: US20170002669A1
Автор: Andreas Bauer, Jaroslaw Leszek Szwedowicz, Jonas Boettcher, Marcus NEUBAUER
Принадлежит: General Electric Technology GmbH

A mechanical fastening system for rotating or stationary components such as turbine or compressor blades on a rotor or a shaft or a casing, respectively, is disclosed. The system includes a circumferential mounting groove adapted for receiving root sections of the rotating or stationary components as well as intermediate fastening components for fixation of the components in equidistance positions. The intermediate fastening components comprise at least an upper platform and at least a side plate having a groove for mounting on said rotor. The intermediate fastening components are made of a plurality of distinct parts of different materials from which at least one clamping part is made of or comprises a shape memory alloy or similar material having a pseudo-elasticity behavior.

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Номер записи: 38
05-01-2017 дата публикации

SEAL FOR A GAS TURBINE ENGINE

Номер: US20170002675A1
Автор: Griffin David Richard, Lutjen Paul M., Moore Christopher William
Принадлежит:

A gas turbine engine assembly includes a shield that has a first portion and a second portion. The first portion extends radially from an axial end portion of the shield and includes a blade outer air seal contact surface. The second portion extends axially from a radially outer end of the first portion and includes a vane contact surface. 1. A gas turbine engine assembly comprising:a shield having a first portion and a second portion, the first portion extending radially from an axial end portion of the shield including a blade outer air seal contact surface and the second portion extending axially from a radially outer end of the first portion including a vane contact surface.2. The gas turbine engine assembly of claim 1 , wherein the shield forms a complete unitary circumferential hoop.3. The gas turbine engine assembly of claim 1 , wherein the shield forms a circumferential hoop with a single discontinuity.4. The gas turbine engine assembly of claim 1 , comprisinga seal in contact with the shield.5. The gas turbine engine assembly of claim 4 , wherein the second portion of the shield is located radially outward from the seal and the first portion of the shield is located axially upstream from the seal.6. The gas turbine engine assembly of claim 4 , wherein the seal includes a “W” shaped cross section pointing radially outward.7. The gas turbine engine assembly of claim 4 , wherein the second portion of the shield is located radially outward from the seal and the first portion of the shield is located axially downstream from the seal.8. A gas turbine engine comprising:at least one vane;at least one blade outer air seal adjacent the at least one vane;a shield located axially between the at least one vane and the at least one blade outer air seal; anda seal located radially inward from the shield.9. The gas turbine engine of claim 8 , wherein the shield comprises a first portion extending radially on an axially upstream end of the shield.10. The gas turbine engine ...

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Номер записи: 39
05-01-2017 дата публикации

ROTOR OFF-TAKE ASSEMBLY

Номер: US20170002678A1
Автор: N. Chandrasekaran, Pattnaik Santosh Kumar, Subramanian Sesha, Sundaran Swaroop
Принадлежит:

According to exemplary embodiments, a rotor off-take assembly is provided by positioning an angled hole or aperture in a stator assembly. This angled hole provides improved pressure recovery and utilizes higher dynamic pressure to drive the bleed air flow into the off-take cavity. 1. A rotor off-take assembly for improved pressure recovery , comprising:a first rotor disk, including at least one first blade connected to the first rotor disk and extending radially outwardly;a second rotor disk, including at least one second blade connected to the second rotor disk and extending radially outwardly;an at least one stator assembly disposed between the first rotor disk and the second rotor disk;the stator assembly including a flow surface generally extending from adjacent the first rotor disk toward the second rotor disk;the stator assembly including an off-take aperture extending downwardly at a non-perpendicular angle through the flow surface;wherein air passes through the off-take aperture of the stator assembly reducing swirl.2. The rotor off-take assembly of further comprising a bleed air passage in rotor structure.3. The rotor off-take assembly for improved pressure recovery of further comprising an impeller tube disposed radially inward of the stator.4. The rotor off-take assembly for improved pressure recovery of claim 3 , wherein the impeller tube is of reduced weight due to reduced height.5. The rotor off-take assembly for improved pressure recovery of claim 4 , wherein the offtake aperture and the impeller tube arrangement results in decreased pressure drop between the off-take aperture and impeller tube.6. The rotor off-take assembly for improved pressure recovery of claim 1 , wherein the offtake aperture is circular in cross-sectional shape.7. The rotor off-take assembly for improved pressure recovery of claim 1 , wherein the offtake aperture is oval in cross-sectional shape.8. The rotor off-take assembly for improved pressure recovery of claim 1 , wherein ...

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Номер записи: 40
07-01-2016 дата публикации

GAS TURBINE ENGINE STATOR VANE BAFFLE ARRANGEMENT

Номер: US20160003071A1
Автор: Devore Matthew A., Ghigliotty Jaime G.
Принадлежит:

A stator vane for a gas turbine engine includes an airfoil that has an exterior wall that provides a cooling cavity. The exterior surface has an interior surface that has multiple pin fins extending therefrom. A baffle is arranged in the cooling cavity and is supported by the pin fins. 1. A stator vane for a gas turbine engine comprising:an airfoil having an exterior wall providing a cooling cavity, the exterior surface has an interior surface having multiple pin fins extending therefrom; anda baffle arranged in the cooling cavity and supported by the pin fins.2. The stator vane according to claim 1 , wherein the baffle is sheet steel.3. The stator vane according to claim 2 , wherein the exterior wall provides pressure and suction sides joined at leading and trailing edges claim 2 , and the baffle includes impingement holes configured to provide impingement cooling fluid onto the exterior wall at the leading edge.4. The stator vane according to claim 2 , wherein the baffle includes a generally smooth outer contour free of protrusions.5. The stator vane according to claim 4 , wherein the outer contour is provided by plastically deformation.6. The stator vane according to claim 4 , wherein cooling holes are provided by at least one of drilling claim 4 , laser drilling claim 4 , or electro discharge machining.7. The stator vane according to claim 1 , wherein a perimeter cavity is provided between the baffle and the exterior wall claim 1 , the pin fins arranged in the perimeter cavity.8. The stator vane according to claim 7 , wherein the perimeter cavity circumscribes the baffle.9. The stator vane according to claim 8 , wherein the pin fins provide the sole support for the baffle in the perimeter cavity.10. The stator vane according to claim 1 , wherein the pin fins are arranged in rows.11. The stator vane according to claim 1 , wherein the pin fins are radially spaced from one another.12. The stator vane according to claim 1 , wherein a rib separates the cooling cavity ...

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Номер записи: 41
07-01-2016 дата публикации

GAS TURBINE ENGINE THIN WALL COMPOSITE VANE AIRFOIL

Номер: US20160003072A1
Автор: Chang Hoyt Y., Duelm Shelton O.
Принадлежит:

An airfoil for a gas turbine engine has a first layer forming a cavity having transitioning from a first thickness to a second thickness through a ply drop region. A second layer is secured to the first layer. 1. An airfoil for a gas turbine engine comprising:a first layer forming a cavity having transitioning from a first thickness to a second thickness through a ply drop region; anda second layer secured to the first layer.2. The airfoil according to claim 1 , comprising a space arranged between the first and second layers claim 1 , and a filler is provided in the space.3. The airfoil according to claim 2 , wherein the second layer terminates in ends forming a V-shape at a trailing edge of the airfoil claim 2 , and the filler is provided between the first layer and second layer.4. The airfoil according to claim 3 , wherein the second thickness is provided at a location between the first thickness and the filler.5. The airfoil according to claim 2 , wherein the filler is provided near a leading edge of the airfoil.6. The airfoil according to claim 1 , wherein each layer includes multiple plies.7. The airfoil according to claim 6 , wherein the plies are constructed from a ceramic matrix composite bonded to one another by a resin.8. The airfoil according to claim 7 , wherein the ceramic matrix composite is a silicon carbide material.9. The airfoil according to claim 1 , wherein the airfoil is a vane.10. The airfoil according to claim 9 , wherein the vane is a mid turbine frame vane.11. The airfoil according to claim 9 , comprising a component passing through the cavity of the vane claim 9 , the component adjacent to the first thickness.12. The airfoil according to claim 9 , wherein a single cavity is provided in the airfoil.13. A method of forming an airfoil comprising:wrapping a first layer about a mandrel and building a thickened area with the first layer relative to an adjacent area of the first layer;applying a filler over the thickened area; andwrapping second ...

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Номер записи: 42
07-01-2016 дата публикации

GUIDE VANE ASSEMBLY VANE BOX OF AN AXIAL TURBINE ENGINE COMPRESSOR

Номер: US20160003073A1
Автор: Derclaye Alain
Принадлежит:

The invention relates to an angular sector of a bladed stator of a low-pressure compressor of an axial turbine engine. The sector comprises an outer shroud and an inner shroud in the form of circular arcs intended to be mounted in a concentric manner on the outer casing of the turbine engine compressor. The sector likewise comprises a row of stator vanes extending radially and anchored in the shrouds in such a manner as to form a bladed box. The vanes of the box comprise anchoring lugs at their outer ends, the lugs being disposed in the thickness of the outer shroud. The inner shroud comprises stubs for anchoring vanes. 1. An angular sector of a bladed stator of an axial turbine engine , said sector comprising:an arcuate segment of an outer shroud intended to be mounted on a casing of the turbine engine;an arcuate segment of inner shroud; and 'at least one anchoring portion of a box vane comprises an anchoring lug which mainly extends in the circumferential direction, and which is disposed in the thickness of one of the shrouds in such a manner as to anchor the vane to the shroud to make the box rigid.', 'a row of stator vanes extending radially from the outer shroud to the inner shroud, each of the stator vanes comprising an inner anchoring portion anchored to the inner shroud and an outer anchoring portion anchored to the outer shroud in such a manner that the stator vanes, the inner shroud and the outer shroud form a bladed box, wherein'}2. The angular sector in accordance with claim 1 , wherein each box vane comprises an airfoil extending between the shrouds in the radial direction claim 1 , the anchoring lugs extending perpendicularly to the radial direction and generally perpendicularly in respect of a chord of the associated vane.3. The angular sector in accordance with claim 1 , wherein at least one box vane comprises two lugs disposed at a same end claim 1 , the lugs being generally curved.4. The angular sector in accordance with claim 1 , wherein at least ...

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Номер записи: 43
07-01-2016 дата публикации

GAS TURBINE ENGINE STATOR VANE PLATFORM COOLING

Номер: US20160003074A1
Автор: Chlus Wieslaw A., Thomen Seth J.
Принадлежит:

An airfoil component for a gas turbine engine includes a platform joined to an airfoil. The platform includes a flow path surface that extends between spaced apart lateral surfaces. The airfoil extends from the flow path surface. A contoured surface adjoins the flow path surface and one of the lateral surfaces. 1. An airfoil component for a gas turbine engine comprising:a platform joined to an airfoil, the platform includes a flow path surface extending between spaced apart lateral surfaces, the airfoil extends from the flow path surface, and a contoured surface adjoins the flow path surface and one of the lateral surfaces.2. The airfoil component according to claim 1 , comprising inner and outer platforms joined by the airfoil claim 1 , one of the inner and outer platforms providing the platform.3. The airfoil component according to claim 2 , wherein the platform is provided by an inner platform.4. The airfoil component according to claim 1 , comprising a cooling passage claim 1 , and cooling holes extend through the contoured surface and are in fluid communication with the cooling passage.5. The airfoil component according to claim 4 , wherein the contoured surface is at first and second angles with respect to the flow path surface and the lateral surface claim 4 , respectively claim 4 , the first and second angles are in the range of greater than 0° to 65°.6. The airfoil component according to claim 5 , wherein the contoured surface is curved.7. The airfoil component according to claim 5 , wherein the first and second angles are about 45°.8. The airfoil component according to claim 4 , wherein the exit of the cooling holes are directed aftward toward a trailing edge of the airfoil.9. The airfoil component according to claim 4 , comprising a thermal barrier coating provided on the inner flow path surface and the contoured surface claim 4 , the cooling holes extend through the thermal barrier coating.10. The airfoil component according to claim 1 , wherein a slot ...

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Номер записи: 44
07-01-2016 дата публикации

Gas turbine engine component having variable width feather seal slot

Номер: US20160003079A1
Автор: Allison Mainelli, Jeffrey J. Degray, Kevin RAJCHEL, Mark A. Boeke, Richard M. Salzillo
Принадлежит: United Technologies Corp

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a mate face and a feather seal slot axially extending along the mate face, the feather seal slot having a variable width along a portion of its axial length.

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Номер записи: 45
07-01-2016 дата публикации

CONTOURED BLADE OUTER AIR SEAL FOR A GAS TURBINE ENGINE

Номер: US20160003082A1
Автор: Bruskotter Michael J., Lutjen Paul M.
Принадлежит:

A blade outer air seal (BOAS) segment according to an exemplary aspect of the present disclosure includes, among other things, a seal body having a radially inner face that circumferentially extend between a first mate face and a second mate face and axially extend between a leading edge face and a trailing edge face, wherein a radial position of the radially inner face varies at a given axial position. 1. A blade outer air seal (BOAS) segment , comprising:a seal body having a radially inner face that circumferentially extend between a first mate face and a second mate face and axially extend between a leading edge face and a trailing edge face, wherein a radial position of the radially inner face varies at a given axial position.2. The BOAS segment of claim 1 , wherein the given axial position is upstream from a rub track of the radially inner face.3. The BOAS segment of claim 2 , wherein the given axial position is a first given axial position claim 2 , and a radial position of the radially inner face varies at a second given axial position that is downstream from the rub track of the radially inner face.4. The BOAS segment of claim 1 , wherein the radial position of the radially inner face smoothly varies at the given axial position.5. The BOAS segment of claim 1 , wherein the radial position of the radially inner face undulates at the given axial position between positions that are radially closer to the a central axis and positions that are radially further from the central axis.6. The BOAS segment of claim 1 , wherein the radial position of the radially inner face is contoured.7. The BOAS segment of claim 1 , wherein the BOAS includes at least a layer of an additive manufacturing material.8. A blade outer air seal (BOAS) assembly claim 1 , comprising:a BOAS segment including a radial inner face that circumferentially extends between a first mate face and a second mate face and axially extends between a leading edge face and a trailing edge face; andat least ...

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Номер записи: 46
07-01-2016 дата публикации

UNDULATING STATOR FOR REDUCING THE NOISE PRODUCED BY INTERACTION WITH A ROTOR

Номер: US20160003095A1
Автор: JODET Norman Bruno André, Romano Pascal Jérôme
Принадлежит: SNECMA

A stator designed to be placed radially in a flow which passes through one or more rotors which share the same axis of rotation, with a leading edge and a trailing edge. The leading edge and trailing edge are connected by a lower face and an upper face, wherein at least one of the faces of the stator has radial undulations which extend axially from the leading edge to the trailing edge. The radial undulations can have at least two bosses in the same azimuth direction, the amplitude of which is at least one centimeter on at least part of the axial length of the stator. A propulsion assembly formed by the rotor and the stator, and to a turbine engine comprising such assembly is also provided. 1. Assembly comprising one or more rotors which share the same axis of rotation , and at least one stator which is designed to be placed radially in a flow which passes through said rotor(s) upstream or downstream thereof , said stator having a leading edge and a trailing edge , said leading edge and trailing edge being connected by a lower face and an upper face , wherein at least one of the faces of said stator has radial undulations which extend axially from the leading edge to the trailing edge , said radial undulations having at least two bosses in the same azimuth direction , the amplitude of which is at least one centimeter on at least part of the axial length of the stator , and in that , with the assembly being designed such that the crossing of said flow by the stator creates on said undulating surface pressure fluctuations with oscillations of the temporal phase according to the radial position , the radial undulations of said face have azimuth maximums and/or minimums in the vicinity of the zero mean dephasing regions for the pressure on the undulating face.2. Assembly according to claim 1 , wherein the radial undulations have a wavelength which is substantially constant along the radial extension of the stator.3. Assembly according to claim 1 , wherein the amplitude ...

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Номер записи: 47
07-01-2016 дата публикации

AXIAL RETAINING RING FOR TURBINE VANES

Номер: US20160003102A1
Автор: SYNNOTT Remy
Принадлежит:

A gas turbine engine is described which has first and second turbine vane assemblies with multiple turbine vanes within respective first and second circumferential outer shrouds. The first outer shroud has a first radially extending flange and the second outer shroud has a second radially extending flange. The radially extending first and second flanges each defining an upstream mating surface and a downstream mating surface relative to a direction of air flow through the engine in use. The downstream mating surface of the first flange mates with the upstream mating surface of the second flange. An axial retaining ring axially retains together the first and second flanges, and has an annular body extending between an upstream portion of the body abutted against the upstream mating surface of the first flange and a downstream portion of the body abutted against the downstream mating surface of the second flange. 1. A gas turbine engine having a center axis of rotation , the engine comprising:first and second turbine vane assemblies having multiple turbine vanes within respective first and second circumferential outer shrouds, the first outer shroud having a first radially extending flange and the second outer shroud having a second radially extending flange, the radially extending first and second flanges each defining an upstream mating surface and a downstream mating surface relative to a direction of air flow through the engine in use, the downstream mating surface of the first flange mating with the upstream mating surface of the second flange; andan axial retaining ring axially retaining together the first and second flanges, the axial retaining ring having an annular body extending between an upstream portion of the body abutted against the upstream mating surface of the first flange and a downstream portion of the body abutted against the downstream mating surface of the second flange.2. The turbine section of claim 1 , wherein the first and second flanges ...

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Номер записи: 48
01-01-2015 дата публикации

GAS TURBINE ENGINE VANE END DEVICES

Номер: US20150003963A1
Автор: Freeman Ted J., Munson John, Taketa Linnea
Принадлежит:

A turbomachinery component of a gas turbine engine is disclosed having a number of techniques of reducing the effects of a gap flow between an airfoil member of the gas turbine engine and a wall of the gas turbine engine. The airfoil member can be a variable and in one form is a variable turbine vane. In one embodiment a brush seal is included between the vane and the wall. In another form a wear surface is disposed between the vane and the wall. In yet another form a moveable member capable of being actuated to change position can be disposed between the vane and the wall to alter the size of a gap between the two. 1. An apparatus comprising:a moveable airfoil member structured for use in a working fluid flow path of a gas turbine engine; anda brush seal disposed at an end of the moveable airfoil member, the brush seal having a plurality of extensions projecting outwardly and configured to discourage a flow of working fluid through the extensions as the working fluid traverses the working fluid flow path.2. The apparatus of claim 1 , which further includes the gas turbine engine claim 1 , the engine including a plurality of the moveable airfoil members claim 1 , wherein each of the plurality of moveable airfoil members is a rotatable vane that includes a range of travel claim 1 , and wherein the extensions contact the wall in the range of travel.3. The apparatus of claim 1 , wherein the plurality of extensions each have a first end and a second end both disposed toward a distal side of the plurality of extensions claim 1 , the first end and second end connected via a body that is looped around a central member.4. The apparatus of claim 3 , which further includes a crimp to couple the plurality of extensions to the central member claim 3 , the central member extending along a chord of the moveable airfoil member claim 3 , and wherein the moveable airfoil member is a rotatable turbine vane.5. The apparatus of claim 1 , wherein the plurality of extensions cover a ...

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Номер записи: 49
01-01-2015 дата публикации

GAS TURBINE ENGINE WITH ATTACHED NOSECONE

Номер: US20150003968A1
Автор: Dumavath Rakesh Sharma N., Jayasingh Arun Kumar, Little Jonathan D., Maliniak David, Mayachari Veerachari J., Monahan Daniel J., Walker Herbert L.
Принадлежит:

A gas turbine engine includes a compressor section and a nosecone assembly. The compressor section includes an inlet guide vane assembly including an inner shroud, an outer shroud, and an inlet guide vane extending from the inner shroud to the outer shroud. The nosecone assembly is attached to the inner shroud. 1. A gas turbine engine comprising: an inner shroud;', 'an outer shroud; and', 'an inlet guide vane extending from the inner shroud to the outer shroud; and, 'an inlet guide vane assembly comprising, 'a compressor section comprisinga nosecone assembly attached to the inner shroud.2. The gas turbine engine of claim 1 , wherein the nosecone assembly comprises:a nosecone; anda nosecone support ring connected to an aft end of the nosecone.3. The gas turbine engine of claim 2 , wherein the nosecone is bolted to the nosecone support ring claim 2 , which is bolted to the inner shroud.4. The gas turbine engine of claim 2 , wherein the nosecone comprises a plurality of tabs extending axially aft of the aft end of the nosecone and wherein a plurality of bolts extend radially inward through the nosecone support ring and into the plurality of tabs.5. The gas turbine engine of claim 2 , wherein the plurality of bolts are countersunk into the nosecone support ring such that heads of the plurality of bolts are positioned flush with or below an outer surface of the nosecone support ring.6. The gas turbine engine of claim 2 , wherein a first set of bolts extend substantially radially to connect the nosecone to the nosecone support ring and wherein a second set of bolts extend substantially axially to connect the nosecone support ring to the inner shroud.7. The gas turbine engine of claim 2 , wherein the nosecone support ring comprises:a rim; anda flange ending radially inward of the rim.8. The gas turbine engine of claim 7 , wherein the flange is bolted to the inner shroud.9. The gas turbine engine of claim 7 , wherein the nosecone is bolted to the rim.10. The gas turbine ...

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Номер записи: 50
01-01-2015 дата публикации

Steam turbine

Номер: US20150003969A1
Автор: Junichi Tominaga, Naoki Shibukawa, Shinichiro Ohashi, Tomohiko Tsukuda
Принадлежит: Toshiba Corp

A steam turbine 10 according to an embodiment includes rotor blades 22 implanted to a turbine rotor 21, stationary blades 26 making up a turbine stage together with the rotor blades 22, diaphragm outer rings 23 including an annular extending part 24 surrounding a periphery of the rotor blades 22, and supporting the stationary blades 26, and diaphragm inner rings 25 supporting the stationary blades 26. The steam turbine 10 further includes an annular slit 40 formed at an inner surface of the diaphragm outer ring 23 between the stationary blades 26 and the rotor blades 22 along a circumferential direction, and communication holes 50 provided in plural at an outer surface of the diaphragm outer ring 23 along the circumferential direction, communicated to the annular slit 40 from the outer surface side, and communicated to an exhaust chamber sucking water films via the annular slit 40.

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Номер записи: 51
04-01-2018 дата публикации

TURBINE ENGINE WHEEL

Номер: US20180003065A1
Автор: COLLADO MORATA Elena
Принадлежит: SAFRAN HELICOPTER ENGINES

The invention relates to a turbine wheel () comprising a plurality of vanes connected to an annular platform () carrying annular lips (). According to the invention, one of the upstream lip () and the downstream lip () is of a first type or of a second type, with the first type corresponding to one lip () having the upstream face () which is concave curved and the downstream face () which is convex curved and the second type corresponding to a lip () having the upstream () and downstream () faces which are substantially flat and mutually parallel. 1. A wheel of a turbine engine comprising a plurality of radially extending vanes , one radially internal or external end of which is connected to an annular platform carrying annular lips extending from said platform in a direction opposite the vane between a first radial end connected to the platform and a second opposed radial free end , in order to sealingly cooperate with a radially facing ring , wherein one of the upstream lip and of the downstream lip is of a first type or of a second type , with the first type corresponding to one lip having the upstream face which is concave curved and the downstream face which is convex curved and the second type corresponding to a lip having the upstream and downstream faces which are substantially flat and mutually parallel , with the first end of said lip being arranged downstream of the second end for both the first type and the second type.2. A wheel according to claim 1 , wherein said lip is arranged at the downstream end of said platform.3. A wheel according to claim 2 , wherein said lip is of the first type.4. A wheel according to claim 3 , further comprising another annular lip arranged at the upstream end of the platform claim 3 , with the other lip being of the second type.5. A wheel according to claim 1 , wherein the generatrix of the cone of revolution going through the first end and the second end of the lip is inclined relative to a plane perpendicular to the axis ...

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Номер записи: 52
04-01-2018 дата публикации

SEGMENTED FACE SEAL ASSEMBLY AND AN ASSOCIATED METHOD THEREOF

Номер: US20180003067A1
Автор: Bidkar Rahul Anil, SEVINCER Edip, Thatte Azam Mihir, Wang Jifeng, Zheng Xiaoqing
Принадлежит:

A turbomachine and a method of operating the turbomachine are disclosed. The turbomachine includes a stator, a rotor including a rotor bearing face, and a face seal assembly including a first segmented seal ring and a second segmented seal ring. The first segmented seal ring includes a plurality of joints and a first flat-contact surface and the second segmented seal ring includes a plurality of segment ends and a second flat-contact surface. One of the first and second segmented seal rings includes a seal bearing face. The second segmented seal ring is coupled to the first segmented seal ring such that the second flat-contact surface is in contact with the first flat-contact surface. The plurality of segment ends is circumferentially offset from the plurality of joints. The first segmented seal ring is slidably coupled to the stator and defines a face seal clearance between the rotor and seal bearing faces. 1. A turbomachine comprising:a stator;a rotor comprising a rotor bearing face; and a first segmented seal ring comprising a plurality of joints and a first flat-contact surface; and', 'a second segmented seal ring comprising a plurality of segment ends and a second flat-contact surface,, 'a face seal assembly comprisingwherein one of the first segmented seal ring and the second segmented seal ring comprises a seal bearing face, wherein the second segmented seal ring is coupled to the first segmented seal ring such that the second flat-contact surface is in contact with the first flat-contact surface, wherein the plurality of segment ends is circumferentially offset from the plurality of joints, and wherein the first segmented seal ring is slidably coupled to the stator and defines a face seal clearance between the rotor bearing face and the seal bearing face.2. The turbomachine of claim 1 , wherein the first segmented seal ring further comprises a circumferential slot extending inwards from a first peripheral side towards a second peripheral side of the first ...

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Номер записи: 53
02-01-2020 дата публикации

VANE SYSTEM WITH CONNECTORS OF DIFFERENT LENGTH

Номер: US20200003064A1
Автор: Kline Bradley P., Shahab Farruqh
Принадлежит:

A vane system includes vane segments that each have a platform, a connector box, and at least one airfoil extending between the platform and the connector box. The connector box has a first circumferential side in the form of a male connector and a second circumferential side in the form of a female socket. The vane segments are connected together in a circumferential row with the male connector of each said vane segment being received in the female socket of the next vane segment in the circumferential row. A majority of the male connectors are of a first, common connector length, and at least one of the male connectors is of a second connector length that is different than the common connector length. 1. A vane system comprising:a plurality of vane segments, each vane segment having a platform, a connector box, and at least one airfoil extending between the platform and the connector box, the connector box having a first circumferential side in the form of a male connector and a second circumferential side in the form of a female socket, the vane segments being connected together in a circumferential row with the male connector of each said vane segment being received in the female socket of the next vane segment in the circumferential row, a majority of the male connectors being of a first, common connector length, and at least one of the male connectors being of a second connector length that is different than the common connector length.2. The vane system as recited in claim 1 , wherein the first connector length and the second connector length are the distance from a base of the male connector to a tip of the male connector.3. The vane system as recited in claim 1 , wherein the second connector length is greater than the first connector length.4. The vane system as recited in claim 1 , wherein each said male connector claim 1 , inclusive of the male connectors that have the first connector length and the at least one male connector that has the second ...

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Номер записи: 54
02-01-2020 дата публикации

GUIDE VANE ARRANGEMENT FOR TURBOMACHINE

Номер: US20200003065A1
Автор: Feldmann Manfred
Принадлежит: MTU Aero Engines AG

The present invention relates to a guide vane arrangement for a turbomachine, comprising a guide vane segment and a housing part that are fastened to one another, for which a back guide vane hook that rises radially toward the outside from an outer shroud of the guide vane segment, when referred to a longitudinal axis of the turbomachine, and a housing hook, which is arranged radially inside circumferentially at the housing part, engage in one another in form-fitting manner, wherein, in a first peripheral segment of the guide vane arrangement, the guide vane hook has a front wall and a back wall, and therewith forms a groove open radially toward the outside, in which a ring section of the housing hook is arranged and held axially. 1. A guide-vane arrangement for a turbomachine , comprising a guide vane segment and a housing part that are fastened to one another , for which a back guide vane hook that rises radially toward the outside from an outer shroud of the guide vane segment , referred to a longitudinal axis of the turbomachine , and a housing hook , which is arranged radially inside circumferentially at the housing part , engage in one another in form-fitting manner , wherein , in a first peripheral segment of the guide vane arrangement , the guide vane hook has a front wall and a back wall , and therewith forms a groove open radially toward the outside , in which a ring section of the housing hook is arranged and held axially ,and wherein, in a second peripheral segment of the guide vane arrangement, the back wall of the guide vane hook is provided with a discontinuity, and a cam projecting axially toward the back is arranged at the ring section of the housing hook,the cam extending axially toward the back in the discontinuity of the back wall and is held therein for circumferential fixing in place.2. The guide vane arrangement according to claim 1 , wherein claim 1 , in a third peripheral segment of the guide vane arrangement claim 1 , which follows the ...

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Номер записи: 55
07-01-2021 дата публикации

VANE ANGLE SYSTEM ACCURACY IMPROVEMENT

Номер: US20210003029A1
Автор: Ward Thomas W.
Принадлежит:

A stator vane angle system includes an engine case, a plurality of stator vanes located at an interior of the engine case. Each stator vane is rotatable about a stator vane axis. A synchronization ring is located at an exterior of the engine case. The synchronization ring is operably connected to each stator vane of the plurality of stator vanes such that movement of the synchronization ring urges rotation of each stator vane of the plurality of stator vanes about their respective stator vane axes. A plurality of impingement openings extend through the engine case from the interior of the engine case to the exterior of the engine case. The plurality of impingement openings are configured to direct flowpath gases from the interior of the engine case to impinge on the synchronization ring, thereby reducing a thermal mismatch between the engine case and the synchronization ring. 1. A stator vane angle system , comprising:an engine case;a plurality of stator vanes disposed at an interior of the engine case, each stator vane rotatable about a stator vane axis;a synchronization ring disposed at an exterior of the engine case, the synchronization ring operably connected to each stator vane of the plurality of stator vanes such that movement of the synchronization ring urges rotation of each stator vane of the plurality of stator vanes about their respective stator vane axes; anda plurality of impingement openings extending through the engine case from the interior of the engine case to the exterior of the engine case, the plurality of impingement openings configured to direct flowpath gases from the interior of the engine case to impinge on the synchronization ring.2. The stator vane angle system of claim 1 , wherein the plurality of impingement openings each have an impingement opening outlet disposed at a same axial location as the synchronization ring.3. The stator vane angle system of claim 1 , wherein the plurality of impingement openings each extend perpendicular to ...

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Номер записи: 56
03-01-2019 дата публикации

Blade removal device and method

Номер: US20190003310A1
Автор: Ryota CHAKI, Tadakazu Sakaguchi
Принадлежит: Mitsubishi Hitachi Power Systems Ltd

A blade removal device for removing a blade in a circumferential direction of a blade ring along a groove, the blade being engaged with the groove extending in the circumferential direction on an inner peripheral side of the blade ring, is provided with a towing part for towing the blade, a string member connecting the towing part and the blade, and a first turning part attached to the blade ring so as to be in contact with a portion of the string member between the towing part and the blade to change a direction of a towing force transmitted via the string member.

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Номер записи: 57
03-01-2019 дата публикации

AIRFOIL ASSEMBLY WITH A SCALLOPED FLOW SURFACE

Номер: US20190003323A1
Автор: Hura Harjit Singh, Shirley Connor Marie, Vitt Paul Hadley
Принадлежит:

A stage for a compressor or a turbine in a turbine engine can include an annular row of airfoils radially extending from corresponding platforms, where each platform can include a fore edge and aft edge and each airfoil can include a leading edge and trailing edge. At least one of the platforms can have a scalloped flow surface including a bulge and a trough. 1. A stage for at least one of a compressor or a turbine , the stage comprising:an annular row of airfoils radially extending from corresponding platforms, the airfoils circumferentially spaced apart to define intervening flow passages;each platform having a fore edge and an aft edge;each airfoil having an outer wall defining a pressure side and a suction side opposite the pressure side, the outer wall extending axially between a leading edge and a trailing edge defining a chord-wise direction, and the outer wall extending radially between a root and a tip defining a span-wise direction, with the root adjacent the platform and the leading edge aft of the fore edge of the platform; and the bulge having a portion extending forward of the fore edge and a local maximum located aft of the fore edge and spaced from the pressure side to define a bulge flow channel between the bulge and the pressure side, and', 'the trough extending adjacent at least a portion of the suction side with a fore portion of the trough located in front of the leading edge., 'at least one of the platforms having a scalloped flow surface including a bulge adjacent the pressure side and a trough adjacent the suction side,'}2. The stage of further comprising a fillet extending between the pressure side and the platform and located between the pressure side and the bulge.3. The stage of wherein the fillet extends between the suction side and the platform and is located between the suction side and the trough.4. The stage of wherein the fillet extends about the periphery of the outer wall.5. The stage of wherein the fore portion of the trough is ...

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Номер записи: 58
20-01-2022 дата публикации

Steam Turbine Hollow Blade

Номер: US20220018265A1
Автор: SASAO Yasuhiro, TABATA Soichiro, TAKATA Ryo, TANIGUCHI Nao
Принадлежит:

A steam turbine hollow stationary blade is able to reduce the amount of water droplets captured on a blade surface. The steam turbine hollow stationary blade, which has a cavity therein, includes a partition wall dividing the cavity into a pressure chamber on a leading edge side and an exhaust chamber on a trailing edge side, at least one steam inlet hole connecting the pressure chamber and an outside of the stationary blade to each other, and at least one pressure conditioning hole connecting the pressure chamber and the exhaust chamber. Total opening area of the pressure conditioning hole is smaller than total opening area of the steam inlet hole. 1. A steam turbine hollow stationary blade , that has a cavity therein , comprising:a partition wall dividing the cavity into a pressure chamber on a leading edge side and an exhaust chamber on a trailing edge side;at least one steam inlet hole connecting the pressure chamber and an outside of the stationary blade to each other; andat least one pressure conditioning hole connecting the pressure chamber and the exhaust chamber, whereintotal opening area of the pressure conditioning hole is smaller than total opening area of the steam inlet hole.2. The steam turbine hollow stationary blade according to claim 1 , Whereinthe steam inlet hole is positioned on a leading edge of the stationary blade.3. The steam turbine hollow stationary blade according to claim 1 , whereinthe pressure conditioning hole is provided to the partition wall at both inner and outer circumferential sides of the partition wall.4. The steam turbine hollow stationary blade according to claim 1 , further comprising:a slit connecting the exhaust chamber to the outside of the stationary blade.5. The steam turbine hollow stationary blade according to claim 1 , further comprising:an exhaust hole connecting the exhaust chamber to a steam condenser. The present invention relates to a steam turbine hollow stationary blade.In steam turbines, during the process ...

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Номер записи: 59
08-01-2015 дата публикации

Stator Blade Sector for an Axial Turbomachine with a Dual Means of Fixing

Номер: US20150010395A1
Автор: Cortequisse Jean-Francois
Принадлежит:

The present application relates to a stator blade sector configured to be fixed to a housing of an axial turbomachine, the sector having a plurality of blades with platforms juxtaposed, so as to describe an arc of a circle. At least one of the platforms comprises on its outer face a fixing screw and at least one other platform has no fixing screws, the platforms being fixed together at their adjacent edges. The application also relates to a stator or portion of stator having a housing forming a generally circular wall and blade sectors arranged along the wall. The housing includes several parts connected to each other by longitudinal flanges. Platforms with no fixing screws are located opposite the flanges. 1. A stator blade sector for attachment to a housing of an axial turbomachine , comprising:a plurality of blades with platforms juxtaposed, so as to describe the arc of a circle, and with an airfoil projecting from the inner face of each platform and directed towards the center of the circular arc described by the platforms;wherein at least one of the platforms comprises:a fixing screw on the outer face thereof and at least one other platform having no fixing screws, the platforms being fixed together at their adjacent edges.2. The stator blade sector in accordance with claim 1 , further comprising:three blades with a central blade and two lateral blades on either side of the central blade, the platform of the central blade being the platform that has no fixing screw, the two platforms of the lateral blades being the platforms having fixing screws.3. The stator blade sector in accordance with claim 2 , wherein at least one of the edges of the platforms with fixing screws forms one end of the sector and includes a shoulder configured to overlap an adjacent edge of an adjacent sector.4. The stator blade sector in accordance with claim 1 , further comprising:three blades with a central blade and two lateral blades on either side of the central blade, the platform of ...

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Номер записи: 60
08-01-2015 дата публикации

BLADE ROW POISITIONING DEVICE, BLADE-DEVICE COMBINATION, METHOD AND TURBOMACHINE

Номер: US20150010396A1
Автор: Bokhorst Willem, Rauscher Stefan, Stricker Hans
Принадлежит:

A device such as a drum for positioning at least one guide blade row from a plurality of guide blade groups in a turbomachine, the device on the outer circumferential side including at least one flange for attachment to a housing section of the turbomachine and on the inner circumferential side including a plurality of uniformly distributed receptacles for accommodating holding elements of the guide blade groups and a plurality of recesses, the device having a depth-reduced inner circumferential section, relative to the accommodating grooves, between at least two adjoining receptacles, which extends in each case from the one receptacle to the other receptacle; a blade-device combination; a method for assembling such a blade-device combination; and a turbomachine. 1. A device for positioning at least one guide blade row in a turbomachine in the circumferential direction , the guide blade row including guide blade groups , each guide blade group including at least two guide blades and at least two holding elements spaced apart from each other in the circumferential direction for connection to the device , the device comprising:at least one flange on the outer circumferential side for attachment to a housing section of the turbomachine and on the inner circumferential side having a plurality of uniformly distributed receptacles for accommodating the holding elements and a plurality of recesses, anda depth-reduced inner circumferential section, relative to the receptacles, between at least two adjoining receptacles, and extending from one of the two adjoining receptacles to the other adjoining receptacle.2. The device as recited in wherein the device is a drum for accommodating multiple guide blade rows successively.3. The device as recited in further comprising a further depth-reduced inner circumference section situated in the circumferential direction alternatingly with the recesses.4. The device as recited in wherein the receptacles claim 1 , recesses and depth- ...

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Номер записи: 61
12-01-2017 дата публикации

METHOD FOR GENERATING AN AIRFOIL INCLUDING AN AERODYNAMICALLY-SHAPED FILLET AND AIRFOILS INCLUDING THE AERODYNAMICALLY-SHAPED FILLET

Номер: US20170009587A1
Автор: Szymanski Stanley J.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

Methods are provided for generating an airfoil including an aerodynamically-shaped fillet. 2D fillet curve offset values are defined by obtaining a 2D airfoil image having airfoil surface line and transition region interconnecting airfoil surface line to flowpath surface line with 2D fillet curve. Flowpath offset lines are generated on image at predetermined flowpath offset values and, in the transition region, graphically represent a flowpath offset surface. Intersection point between 2D fillet curve and each flowpath offset line is determined Distance between airfoil surface line and each intersection point generates airfoil offset values. Airfoil and flowpath offset surfaces according to airfoil offset values and flowpath offset values respectively and 3D fillet streamline curves are generated on computer model to define the aerodynamically-shaped fillet. Each 3D fillet streamline curve is an intersection between the airfoil and flowpath offset surfaces. 1. A method for generating an airfoil including an aerodynamically-shaped fillet , the method comprising: obtaining a 2D image of an airfoil having an airfoil surface line and including a transition region interconnecting the airfoil surface line to a flowpath surface line with a 2D fillet curve;', 'generating a plurality of flowpath offset lines on the 2D image at pre-determined flowpath offset values, each flowpath offset line of the plurality of flowpath offset lines in the transition region graphically representing a flowpath offset surface of a plurality of flowpath offset surfaces;', 'determining an intersection point between the 2D fillet curve and each flowpath offset line;', 'computing a distance between the airfoil surface line and each intersection point to generate a plurality of airfoil offset values;, 'defining a plurality of two-dimensional (2D) fillet curve offset values bygenerating a plurality of airfoil offset surfaces according to each airfoil offset value of the plurality of airfoil offset ...

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Номер записи: 62
14-01-2016 дата публикации

COMPONENTS WITH COOLING CHANNELS AND METHODS OF MANUFACTURE

Номер: US20160010464A1
Автор: Bonini Eric Richard, BUNKER Ronald Scott
Принадлежит:

A component is provided and includes a substrate comprising an outer and an inner surface, where the inner surface defines at least one hollow, interior space. The component defines one or more grooves, where each groove extends at least partially along the outer surface of the substrate and has a base and a top. The base is wider than the top, such that each groove comprises a re-entrant shaped groove. One or more access holes are formed through the base of a respective groove, to connect the groove in fluid communication with the respective hollow interior space. Each access hole has an exit diameter D that exceeds the opening width d of the top of the respective groove. The diameter D is an effective diameter based on the area enclosed. The component further includes at least one coating disposed over at least a portion of the surface of the substrate, wherein the groove(s) and the coating together define one or more re-entrant shaped channels for cooling the component. A method for manufacturing the component is also provided. A method for manufacturing a component is also provided, where the groove and the access hole(s) are machined as a single continuous process, such that the groove and the access hole(s) form a continuous cooling passage. 1. A component comprising:a substrate comprising an outer surface and an inner surface, wherein the inner surface defines at least one hollow, interior space, wherein the component defines one or more grooves, wherein each groove extends at least partially along the substrate and has a base and a top, wherein the base is wider than the top, such that each groove comprises a re-entrant shaped groove, wherein one or more access holes are formed through the base of a respective groove, to connect the groove in fluid communication with the respective hollow interior space, wherein each access hole has an exit diameter D that exceeds an opening width d of the top of the respective groove, wherein the diameter D is an effective ...

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Номер записи: 63
14-01-2016 дата публикации

Gas turbine engine airfoil leading edge cooling

Номер: US20160010465A1
Автор: Dominic J. Mongillo, Jr., Matthew A. Devore, Thomas N. Slavens
Принадлежит: United Technologies Corp

An example gas turbine engine component includes an airfoil having a leading edge area, a first circuit to cool a first section of the leading edge area, and a second circuit to cool a second section of the leading edge area. The first circuit separate and distinct from the second circuit within the airfoil.

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Номер записи: 64
14-01-2016 дата публикации

COMPOSITE AIRFOIL METAL LEADING EDGE ASSEMBLY

Номер: US20160010468A1
Автор: Albrecht, JR. Richard W., Kray Nicholas Joseph, Li Qiang
Принадлежит: GENERAL ELECTRIC COMPANY

An airfoil assembly () comprises a composite airfoil () having a leading edge () and a trailing edge (), a pressure side () extending between the leading edge and the trailing edge, a suction side () extending between the leading edge and the trailing edge, opposite the leading edge, a metallic leading edge assembly () disposed over the composite air-foil, the metallic leading edge assembly including a high density base (), the metallic leading edge assembly also including a nose () disposed over the base, an adhesive bond layer disposed between the composite airfoil and the metallic leading edge assembly. 1. An airfoil assembly , comprising: a leading edge and a trailing edge;', 'a pressure side extending between said leading edge and said trailing edge;', 'a suction side extending between said leading edge and said trailing edge, opposite said leading edge;, 'a composite foil havinga metallic leading edge assembly disposed over said composite foil;said metallic leading edge assembly including a high density base;said metallic leading edge assembly also including a nose disposed one of over or under said base;an adhesive bond layer disposed between the composite foil and the metallic leading edge assembly.2. The airfoil assembly of claim 1 , wherein said high density base is formed of a uniform thickness.3. The airfoil assembly of claim 1 , wherein said high density base is formed of a varying thickness.4. The airfoil assembly of claim 1 , said base being welded to said nose.5. The airfoil assembly of claim 1 , said base being bonded to said nose.6. The airfoil of claim 1 , said base having first and second legs which are longer than side walls of said nose.7. The airfoil of claim 1 , wherein said metal leading edge assembly is formed of a single construction in a radial direction.8. The airfoil of claim 1 , wherein said metal leading edge assembly is formed of multiple segments in a radial direction.9. The airfoil of claim 1 , wherein said nose is bonded to said ...

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Номер записи: 65
14-01-2016 дата публикации

TURBINE NOZZLE COMPONENTS HAVING REDUCED FLOW AREAS

Номер: US20160010474A1
Автор: Macelroy Bill
Принадлежит: HONEYWELL INTERNATIONAL INC.

Embodiments of a method for controllably reducing of the flow area of a turbine nozzle component are provided, as are embodiments of turbine nozzle components having reduced flow areas. In one embodiment, the method includes the steps of obtaining a turbine nozzle component having a plurality of turbine nozzle flow paths therethrough, positioning braze preforms in the plurality of turbine nozzle flow paths and against a surface of the turbine nozzle component, and bonding the braze preforms to the turbine nozzle component to achieve a controlled reduction in the flow area of the turbine nozzle flow paths. 1. A turbine nozzle component , comprising:an inner endwall;an outer endwall radially spaced from the inner endwall;a plurality of nozzle vanes extending between the inner and outer endwalls;a plurality of turbine nozzle flow paths extending through the turbine nozzle component and generally defined by the inner endwall, the outer endwall, and the plurality of nozzle vanes; andbraze preforms positioned in the turbine nozzle flow paths and bonded to at least one of the inner endwall and outer endwall reducing the flow area of the turbine nozzle flow paths.2. The turbine nozzle component of wherein the plurality of nozzle vanes have leading and trailing edges around which the braze preforms wrap.3. The turbine nozzle component of wherein the braze preforms are further welded to at least one of the inner endwall and outer endwall.4. The turbine nozzle component of further comprising inner inter-blade flow areas provided on the inner endwall and bounding the plurality of flow paths claim 1 , the plurality of braze preforms having planform geometries substantially conformal with the inner inter-blade flow areas.5. The turbine nozzle component of further comprising outer inter-blade flow areas provided on the outer endwall and bounding the plurality of flow paths claim 1 , the plurality of braze preforms having planform geometries substantially conformal with the outer ...

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Номер записи: 66
14-01-2016 дата публикации

SHAPED RIM CAVITY WING SURFACE

Номер: US20160010476A1
Автор: Grover Eric A.
Принадлежит:

A shaped rim cavity wing includes an upper surface and a lower surface. The lower surface has a geometric shape to control the separation of airflow as it passes around the lower surface to the top surface. A point of maximum extent defines the boundary between the upper surface and the lower surface, wherein the point of maximum extent defines a corner that that separates airflow from the shaped rim cavity rim and creates a flow re-circulation adjacent to the top surface of the shaped rim cavity wing. 1. A shaped rim cavity wing comprising:a body configured to extend from one of a rotating component and a stationary component of a turbomachine to inhibit airflow through a gas path between the rotating component and the stationary component;an upper surface of the body; a first concave portion;', 'a convex portion adjacent the first concave portion;', 'a first inflection point between the convex portion and the first concave portion;', 'a second concave portion adjacent the convex portion; and', 'a second inflection point between the second concave portion and the convex portion; and, 'a lower surface of the body, the lower surface having a geometric shape to control the separation of airflow as it passes around the lower surface, the geometric shape includinga point of maximum extent that defines a boundary between the upper surface and the lower surface, wherein the point of maximum extent defines a corner that separates airflow from the shaped rim cavity wing and creates flow re-circulation adjacent to the upper surface of the shaped rim cavity wing;wherein the first concave portion is located adjacent the point of maximum extent.2. The shaped rim cavity wing of claim 17 , further including:a flat portion located between the point of maximum extent and the lower surface.3. The shaped rim cavity wing of claim 18 , wherein the flat portion is vertical.4. The shaped rim cavity wing of claim 17 , wherein the body extends from the rotating component of the ...

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Номер записи: 67
14-01-2016 дата публикации

Intercooled Compressor for a Gas Turbine Engine

Номер: US20160010498A1
Автор: TAYLOR Jack R.
Принадлежит:

A multi-stage intercooled compressor for a gas turbine engine, including multiple stages of rotating blades and cooling stator vanes, a cooling stator vane including an outer wall that defines an internal coolant fluid passage and has a length along a centerline from a leading edge to a trailing edge of the outer wall, and an internal flow divider wall disposed within the internal passage and extending along the centerline to divide the internal coolant fluid passage into an inflow pathway and an outflow pathway. 1. A multi-stage intercooled compressor for a gas turbine engine , including multiple stages of rotating blades and cooling stator vanes , the cooling stator vane including an outer wall that defines an internal coolant fluid passage and has a length along a centerline from a leading edge to a trailing edge of the outer wall , and an internal flow divider wall disposed within the internal passage and extending along the centerline to divide the internal coolant fluid passage into an inflow pathway and an outflow pathway.2. The multi-stage intercooled compressor of claim 1 , wherein the outer surface of the stator vane is substantially free of an extending cooling fin.3. The multi-stage intercooled compressor of claim 1 , wherein the surface area of an interior surface of the outer wall claim 1 , exposed to cooling fluid claim 1 , is at least about 90% of a surface area of an outer surface of the outer wall claim 1 , exposed to compression air.4. The multi-stage intercooled compressor of claim 1 , wherein a leading edge of the internal flow divider wall is connected to the leading edge of the outer wall claim 1 , and a trailing edge of the internal flow divider wall is connected to the trailing edge of the outer wall.5. The multi-stage intercooled compressor of claim 1 , wherein an outer surface of the internal flow divider is completely separated from an interior surface of the outer wall.6. The multi-stage intercooled compressor of claim 5 , wherein an ...

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Номер записи: 68
11-01-2018 дата публикации

RING STATOR

Номер: US20180010470A1
Автор: Amadon Colin G., Baumann Paul W., Ford Steven J.
Принадлежит:

A stator assembly for a gas turbine engine includes an annular outer shroud, an annular inner shroud radially spaced from the outer shroud and a plurality of stator vanes extending from the outer shroud to the inner shroud. A volume of potting is located at the inner shroud and at the outer shroud to retain the plurality of stator vanes thereat. A stator and case assembly for a gas turbine engine includes a case defining a working fluid flowpath for the gas turbine engine and a stator assembly located at the case. The stator assembly includes an annular outer shroud secured to the case, an annular inner shroud secured to the case and a plurality of stator vanes extending from the outer to the inner shroud. A volume of potting is located at the inner shroud and at the outer shroud to retain the plurality of stator vanes thereat. 1. A stator assembly for a gas turbine engine , comprising:an annular outer shroud;an annular inner shroud radially spaced from the outer shroud;a plurality of stator vanes extending from the outer shroud to the inner shroud; anda volume of potting disposed at the inner shroud and at the outer shroud to retain the plurality of stator vanes thereat.2. The stator assembly of claim 1 , wherein each stator vane of the plurality of stator vanes includes:an airfoil portion;an outer leg extending radially outwardly from the airfoil portion; andan inner leg extending radially inwardly from the airfoil portion.3. The stator assembly of claim 2 , wherein:the outer leg is installed into an outer shroud opening in the outer shroud; andthe inner leg is installed into an inner shroud opening in the inner shroud.4. The stator assembly of claim 3 , wherein the potting comprises:an outer grommet disposed at each outer shroud opening; andan inner grommet disposed at each inner shroud opening to retain each stator vane thereat.5. The stator assembly of claim 2 , wherein each stator vane further includes:an outer leg opening; andan inner leg opening;wherein a ...

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Номер записи: 69
11-01-2018 дата публикации

Spall break for turbine component coatings

Номер: US20180010471A1
Автор: Seth Thomen
Принадлежит: PW Power Systems LLC

A turbine engine component can include a surface comprising at least one edge and a coating disposed upon the surface that can extend to the edge. A spall break can be disposed along a line upon the surface adjacent the edge to prevent spallation of the coating from spreading from the edge onto the surface beyond the spall break. The spall break can comprise a discontinuity of the coating. A method of coating a turbine component can include preparing a substrate to receive a coating and selecting a fail location along the substrate for a coating. One or more coating can be applied to the substrate and a spall break can be incorporated into the one or more coatings. The spall break can comprise a line of discontinuity in the one or more coatings along the fail location.

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Номер записи: 70
11-01-2018 дата публикации

Segmented Stator Assembly

Номер: US20180010472A1
Автор: Baumann Paul W.
Принадлежит:

A stator assembly for a gas turbine engine includes an arcuate outer shroud, an arcuate inner shroud radially spaced from the outer shroud and a plurality of stator vanes extending from the outer shroud to the inner shroud. A volume of potting is located at the inner shroud and at the outer shroud to retain the plurality of stator vanes thereat. A stator and case assembly includes a case defining a working fluid flowpath and a stator assembly positioned at the case. The stator assembly includes a plurality of stator segments arranged circumferentially about an engine axis, each stator segment including an arcuate outer shroud secured to the case, an arcuate inner shroud, and a plurality of stator vanes extending from the outer to inner shroud. A volume of potting is located at the inner shroud and at the outer shroud to retain the plurality of stator vanes thereat. 1. A stator assembly for a gas turbine engine , comprising:an arcuate outer shroud;an arcuate inner shroud radially spaced from the outer shroud;a plurality of stator vanes extending from the outer shroud to the inner shroud; anda volume of potting disposed at the inner shroud and at the outer shroud to retain the plurality of stator vanes thereat.2. The stator assembly of claim 1 , wherein each stator vane of the plurality of stator vanes includes:an airfoil portion;an outer leg extending radially outwardly from the airfoil portion; andan inner leg extending radially inwardly from the airfoil portion.3. The stator assembly of claim 2 , wherein:the outer leg is installed into an outer shroud opening in the outer shroud; andthe inner leg is installed into an inner shroud opening in the inner shroud.4. The stator assembly of claim 3 , wherein the potting comprises:an outer grommet disposed at each outer shroud opening; andan inner grommet disposed at each inner shroud opening to retain each stator vane thereat.5. The stator assembly of claim 2 , wherein each stator vane further includes:an outer leg opening ...

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Номер записи: 71
11-01-2018 дата публикации

Attachment Faces for Clamped Turbine Stator of a Gas Turbine Engine

Номер: US20180010473A1
Автор: Carr Jesse M., Dube Bryan P.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

An airfoil fairing shell for a gas turbine engine includes an airfoil section between an outer vane endwall and an inner vane endwall, at least one of the outer vane endwall and the inner vane endwall including a radial attachment face, a suction side tangential attachment face, a pressure side tangential attachment face, and an axial attachment face. 1. An airfoil fairing shell for a gas turbine engine comprising:an airfoil section between an outer vane endwall and an inner vane endwall, at least one of said outer vane endwall and said inner vane endwall including a radial attachment face, a suction side tangential attachment face, a pressure side tangential attachment face, and an axial attachment face, said suction side tangential attachment face transverse to a resultant aerodynamic load generated by said airfoil.2. The airfoil fairing shell as recited in claim 1 , wherein said radial attachment face claim 1 , said suction side tangential attachment face claim 1 , said pressure side tangential attachment face claim 1 , and said axial attachment face are formed by a thickened region of at least one of said outer vane endwall and said inner vane endwall.3. The airfoil fairing shell as recited in claim 1 , wherein said radial attachment face claim 1 , said suction side tangential attachment face claim 1 , said pressure side tangential attachment face claim 1 , and said axial attachment face are formed by a thickened region of said inner vane endwall.4. The airfoil fairing shell as recited in claim 1 , wherein said suction side tangential attachment face is parallel to said pressure side tangential attachment face.5. The airfoil fairing shell as recited in claim 1 , wherein said suction side tangential attachment face and said pressure side tangential attachment face are non-parallel to said inner vane endwall.6. The airfoil fairing shell as recited in claim 1 , wherein said suction side tangential attachment face and said pressure side tangential attachment face ...

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Номер записи: 72
14-01-2021 дата публикации

STATOR VANE FOR STEAM TURBINE, STEAM TURBINE, AND METHOD FOR HEATING STATOR VANE FOR STEAM TURBINE

Номер: US20210010379A1
Автор: Hirano Yuichiro, Kitamura Tsuyoshi, TAKATA Ryo, TANIGUCHI Nao
Принадлежит: MITSUBISHI HEAVY INDUSTRIES, LTD.

A stator vane for a steam turbine includes: a vane body having an airfoil cross section including a pressure-side partition wall having a concave surface shape and a suction-side partition wall having a convex surface shape, the vane body having a hollow section formed between an inner surface of the pressure-side partition wall and an inner surface of the suction-side partition wall; and a first division wall dividing the hollow section into a first hollow section positioned at a leading edge side and a second hollow section positioned at a trailing edge side. The first hollow section is configured to be supplied with a fluid, or as a sealed space, and a slit is formed on at least one of the pressure-side partition wall or the suction-side partition wall, the slit being in communication with the second hollow section. 1. A stator vane for a steam turbine , comprising:a vane body having an airfoil cross section including a pressure-side partition wall having a concave surface shape and a suction-side partition wall having a convex surface shape, the vane body having a hollow section formed between an inner surface of the pressure-side partition wall and an inner surface of the suction-side partition wall; anda first division wall dividing the hollow section into a first hollow section positioned at a leading edge side and a second hollow section positioned at a trailing edge side,wherein the first hollow section is configured to be supplied with a fluid, and a slit is formed on at least one of the pressure-side partition wall or the suction-side partition wall, the slit being in communication with the second hollow section.2. The stator vane for a steam turbine according to claim 1 , further comprising:a second division wall dividing the first hollow section into a pressure-side space closer to the pressure-side partition wall and a suction-side space closer to the suction-side partition wall,wherein the pressure-side space is configured to be an outgoing passage of ...

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Номер записи: 73
10-01-2019 дата публикации

TURBINE NOZZLE

Номер: US20190010815A1
Автор: WATANABE Fumiaki
Принадлежит: IHI CORPORATION

A tapered surface section is formed on an inner surface of a base facing a flow path in an outer band section or inner band section of a turbine stator vane of one of two vane segments disposed adjacent to each other with a space therebetween, the turbine stator vane being formed of a ceramic matrix composite. The tapered surface section has a tapered shape which approaches an outer surface as extending toward the tip (i.e., toward an end face) of the base. 1. A turbine nozzle comprising:a plurality of turbine stator vanes each being formed by combining ceramics with a fiber fabric, an end of the turbine stator vane being bent and being integrally molded into a shape corresponding to an airfoil portion and to a band section connecting to the airfoil portion;a flow path of gas between airfoil portions of adjacent two turbine stator vanes;a seal member extending across a bent section connecting to the airfoil portion of the band section of one turbine stator vane of the adjacent two turbine stator vanes, and a tip part of the band section, which is spaced from the bent section of the one turbine stator vane, of the other turbine stator vane of the adjacent two turbine stator vanes; anda thin-walled part formed in an inner surface facing the flow path of the band section of the other turbine stator vane, wherein a thickness between the inner surface of the band section and an outer surface opposite to the inner surface is, at a tip, thinner than the bent section.2. The turbine nozzle according to claim 1 , whereina locking piece of the seal member is provided in the outer surfaces of the both band sections, respectively, and wherein the seal member is abutted against the outer surfaces of the both band sections by inserting the seal member into a locking groove formed from the locking piece of the each band section and the outer surface.3. The turbine nozzle according to claim 1 , whereinthe thin-walled part is formed in a portion corresponding to a blade width of the ...

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Номер записи: 74
10-01-2019 дата публикации

Stator vane assembly for a gas turbine engine

Номер: US20190010816A1
Автор: Brian BARAINCA, Paul W. Baumann
Принадлежит: United Technologies Corp

A gas turbine engine has a stator vane assembly. The stator vane assembly includes an inner diameter shroud, an outer diameter shroud located radially outward from the inner diameter shroud, a vane extending radially outward from the first inner diameter shroud to the outer diameter shroud. The wedge clip is positioned horizontally through the vane to prevent the vane from being dislodged from the stator vane assembly.

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Номер записи: 75
15-01-2015 дата публикации

SYSTEM HAVING DUAL-VOLUTE AXIAL TURBINE TURBOCHARGER

Номер: US20150013332A1
Автор: NASIR Shakeel, SCHUELER Joshua David
Принадлежит:

A turbocharger is disclosed for use with an engine system. The turbocharger may have a housing at least partially defining a compressor shroud and a turbine shroud. The turbine shroud may form a first volute and a second volute, each having an inlet configured to receive exhaust from an exhaust manifold of the engine in a tangential direction, and an axial channel disposed downstream of the inlet. The turbocharger may also have a turbine wheel disposed within the turbine shroud and configured to receive exhaust from the axial channels of the first and second volutes, a compressor wheel disposed within the compressor shroud, and a shaft connecting the turbine wheel to the compressor wheel. The turbocharger may further have a nozzle ring in fluid communication with the axial channels of the first and second volutes at a location upstream of the turbine wheel. 1. A nozzle ring for a turbocharger , comprising:an inner annular band;an outer annular band;an intermediate annular band located radially between the inner and outer annular bands;a first plurality of vanes extending between the inner and intermediate annular bands; anda second plurality of vanes extending between the intermediate and outer annular bands.2. The nozzle ring of claim 1 , wherein a number of the first plurality of vanes is different than a number of the second plurality of vanes.3. The nozzle ring of claim 1 , wherein an aerodynamic design and configuration of the first plurality of vanes is different than an aerodynamic design and configuration of the second plurality of vanes.4. The nozzle ring of claim 1 , further including at least one tab extending radially inward from the inner annular band.5. The nozzle ring of claim 1 , further including a flange extending radially outward from a periphery of the outer annular band.6. A turbocharger for an engine claim 1 , comprising: an inlet configured to receive exhaust from an exhaust manifold of the engine in a tangential direction; and', 'an axial ...

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Номер записи: 76
15-01-2015 дата публикации

BI-CAST TURBINE VANE

Номер: US20150016972A1
Автор: Chamberlain Adam L., Freeman Ted J., Landwehr Sean E., Proctor Robert F.
Принадлежит:

One aspect of the present disclosure includes a turbine vane assembly comprising a vane made from ceramic matrix composite material having an outer wall extending between a leading edge and a trailing edge and between a first end and an opposing second end; an endwall made at least partially from a ceramic matrix composite material configured to engage the first end of the vane; and a retaining region including corresponding bi-cast grooves formed adjacent the first end of the vane and a receiving aperture formed in the endwall; wherein a bond is formed in the retaining region to join the vane and endwall together. 1. A turbine vane assembly comprising:a vane made from ceramic matrix composite material having an outer wall extending between a leading edge and a trailing edge and between a first end and an opposing second end;an endwall made at least partially from a ceramic matrix composite material configured to engage the first end of the vane; anda retaining region including corresponding bi-cast grooves formed adjacent the first end of the vane and a receiving aperture formed in the endwall; wherein a bond is formed in the retaining region to join the vane and endwall together.2. The turbine vane assembly of claim 1 , further comprising:a second endwall made at least partially from ceramic matrix composite material configured to engage the second end of the vane.3. The turbine vane assembly of claim 1 , wherein the vane includes internal open regions for receiving cooling fluid therein.4. The turbine vane assembly of claim 3 , wherein the endwall includes internal open regions fluidly connected to the open regions of the vane.5. The turbine vane assembly of claim 1 , wherein the vane includes a plurality of cooling holes formed through an outer wall in fluid communication with the open regions.6. The turbine vane assembly of claim 1 , wherein the vane includes one or more coating layers applied to an outer surface thereof.7. The turbine vane assembly of claim 6 ...

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Номер записи: 77
03-02-2022 дата публикации

Ceramic matrix composite turbine vane and method for making

Номер: US20220034233A1
Автор: Kevin Lukhard, Robert J. Shinavski, Steffan Brown, Ted J. Freeman, Thomas Tran
Принадлежит: Rolls Royce High Temperature Composites Inc, Rolls Royce North American Technologies Inc

A turbine vane comprising ceramic matrix composite materials includes a vane support core, an airfoil, and an end wall that at least partially defines a gas path. The turbine vane is formed from a plurality of ceramic plies or preforms that are infiltrated with ceramic matrix material to form a one-piece ceramic matrix composite turbine vane.

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Номер записи: 78
19-01-2017 дата публикации

COOLING STRUCTURE FOR STATIONARY BLADE

Номер: US20170016338A1
Автор: Golden Christopher Lee, Porter Christopher Donald
Принадлежит:

Embodiments of the present disclosure provide a cooling structure for a stationary blade. The cooling structure can include: an airfoil having a cooling circuit therein; an endwall coupled to a radial end of the airfoil; a chamber positioned within the endwall for receiving a cooling fluid from the cooling circuit, wherein the cooling fluid absorbs heat from the endwall, and a temperature of the cooling fluid in an upstream region is lower than a temperature of the cooling fluid in a downstream region; a first passage within the endwall fluidly connecting the upstream region of the chamber to a wheel space positioned between the endwall and the turbine wheel; and a second passage within the endwall fluidly connecting the downstream region of the chamber to the wheel space. 1. A cooling structure for a stationary blade , the cooling structure comprising:an airfoil having a cooling circuit therein;an endwall coupled to a radial end of the airfoil, relative to a rotor axis of a turbomachine;a chamber positioned within the endwall for receiving a cooling fluid from the cooling circuit and including an upstream region and a downstream region therein, wherein the cooling fluid absorbs heat from the endwall, and a temperature of the cooling fluid in the upstream region is lower than a temperature of the cooling fluid in the downstream region;a first passage within the endwall fluidly connecting the upstream region of the chamber to a wheel space positioned between the endwall and the turbine wheel, wherein a first portion of the cooling fluid in the upstream region passes through the first passage; anda second passage within the endwall fluidly connecting the downstream region of the chamber to the wheel space, wherein a second portion of the cooling fluid in the downstream region passes through the second passage, and a remainder portion of the cooling fluid bypasses the first passage and the second passage without entering the wheel space.2. The cooling structure of ...

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Номер записи: 79
19-01-2017 дата публикации

COOLING STRUCTURE FOR STATIONARY BLADE

Номер: US20170016339A1
Автор: Gergely George Andrew, Holloway Mary Virginia, Smith Paul Kendall, Winn Aaron Gregory
Принадлежит:

Embodiments of the present disclosure provide a cooling structure for a stationary blade, including: an endwall coupled to a radial end of an airfoil; a chamber positioned within the endwall and radially displaced from a radially outer end of the trailing edge of the airfoil, wherein the chamber includes a pair of opposing chamber walls, one of the pair of opposing chamber walls being positioned proximal to the pressure side surface of the airfoil and the other of the pair of opposing chamber walls being positioned proximal to the suction side surface and the trailing edge of the airfoil, and wherein the cooling fluid in the chamber is in thermal communication with least a portion of the endwall positioned proximal to the pressure side surface and the trailing edge of the airfoil; and a plurality of thermally conductive fixtures positioned within the chamber. 1. A cooling structure for a stationary blade , comprising:an endwall coupled to a radial end of an airfoil relative to a rotor axis of a turbomachine, the airfoil including a pressure side surface, a suction side surface, a leading edge, and a trailing edge;a chamber positioned within the endwall and radially displaced from the radial end of the trailing edge of the airfoil, the chamber receiving a cooling fluid from a cooling fluid source, wherein the chamber includes a pair of opposing chamber walls, one of the pair of opposing chamber walls being positioned proximal to the pressure side surface of the airfoil and the other of the pair of opposing chamber walls being positioned proximal to the suction side surface and the trailing edge of the airfoil, and wherein the cooling fluid in the chamber is in thermal communication with least a portion of the endwall positioned proximal to the pressure side surface and the trailing edge of the airfoil; anda plurality of thermally conductive fixtures positioned within the chamber and distributed substantially uniformly throughout the chamber.2. The cooling structure ...

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Номер записи: 80
19-01-2017 дата публикации

SHROUD ASSEMBLY FOR GAS TURBINE ENGINE

Номер: US20170016341A1
Автор: Armanios Laura Isabelle, HEITMAN Bryce Loring, Stapleton David Scott
Принадлежит:

Shroud assemblies for gas turbine engines are provided. A shroud assembly includes a hanger having a forward hanger arm, a rear hanger arm, and a hanger body extending between the forward hanger arm and the rear hanger arm. The shroud assembly further includes a shroud having a forward surface, a rear surface, and an inner surface and outer surface extending between the forward surface and the rear surface, the outer surface radially spaced from the inner surface, the shroud connected to the hanger. The shroud assembly further includes a support member positioned axially forward of the forward hanger arm, the support member having a radially outer portion connected to the forward hanger arm and a radially inner portion axially spaced from the shroud such that a gap is defined between the radially inner portion and an axially adjacent surface of the shroud. 1. A shroud assembly for a gas turbine engine , the shroud assembly comprising:a hanger, the hanger comprising a forward hanger arm, a rear hanger arm axially spaced from the forward hanger arm, and a hanger body extending between the forward hanger arm and the rear hanger arm;a shroud, the shroud comprising a forward surface, a rear surface axially spaced from the forward surface, an inner surface extending between the forward surface and the rear surface, and an outer surface extending between the forward surface and the rear surface and radially spaced from the inner surface, the shroud connected to the hanger; anda support member positioned axially forward of the forward hanger arm, the support member comprising a radially outer portion connected to the forward hanger arm and a radially inner portion axially spaced from the shroud such that a gap is defined between the radially inner portion and an axially adjacent surface of the shroud.2. The shroud assembly of claim 1 , wherein the radially outer portion contacts the forward hanger arm.3. The shroud assembly of claim 1 , further comprising a hanger plate ...

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Номер записи: 81
19-01-2017 дата публикации

Axial Flow Turbine

Номер: US20170016342A1
Автор: FUKUSHIMA Hisataka, LEE Goingwon, NISHIJIMA Noriyo, SEGAWA Kiyoshi, SHIBATA Takanori
Принадлежит:

An axial flow turbine that can enhance an effect of reducing a mixing loss is disclosed. The axial flow turbine includes a plurality of stator blades provided on the inner circumferential side of a diaphragm outer ring; a plurality of rotor blades provided on the outer circumferential side of a rotor; a shroud provided on the outer circumferential side of the plurality of rotor blades; an annular groove portion formed in the diaphragm outer ring and housing the shroud therein; a clearance passage defined between the groove portion and the shroud, into which a portion of working fluid flows from the downstream side of the stator blades in a main passage; seal fins provided in the clearance passage; a circulation flow generating chamber defined on the downstream side of the clearance passage; and a plurality of shielding plates secured to the diaphragm outer ring. 1. A stationary body for a steam turbine comprising:an inner circumferential surface constituting a main passage through which steam flows;an annular groove portion housing a shroud therein, the shroud being provided on the outer circumferential side of rotor blades;a projecting portion projecting from a downstream-side lateral surface of the groove portion toward a downstream-side end face of the shroud, the downstream-side end face of the shroud being located on the radial inside of an outer circumferential surface of the shroud; anda plurality of shielding plates arranged at given intervals in the circumferential direction in a space, the space being defined by an inner circumferential surface of the groove portion, the downstream-side lateral surface of the groove portion, and an outer circumferential surface of the projecting portion.2. The stationary body according to claim 1 ,wherein the shielding plates are located on the downstream side of seal fins arranged between the shroud and the inner circumferential surface of the groove portion.3. A steam turbine comprising:{'claim-ref': {'@idref': 'CLM- ...

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Номер записи: 82
19-01-2017 дата публикации

COOLING STRUCTURE FOR STATIONARY BLADE

Номер: US20170016348A1
Автор: Earnhardt Dustin Michael, Golden Christopher Lee, Iduate Michelle Jessica, Lewis Bryan David, Porter Christopher Donald, Weber David Wayne
Принадлежит:

Embodiments of the present disclosure provide a cooling structure for a stationary blade, which can include: an endwall coupled to a radial end of an airfoil, relative to a rotor axis of a turbomachine; and a substantially crescent-shaped chamber positioned within the endwall and radially displaced from a trailing edge of the airfoil, the substantially crescent-shaped chamber receiving a cooling fluid from a cooling circuit, wherein the substantially crescent-shaped chamber extends from a fore section positioned proximal to one of a pressure side surface and a suction side surface of the airfoil to an aft section positioned proximal to the trailing edge of the airfoil and the other of the pressure side surface and the suction side surface of the airfoil, wherein the aft section of the substantially crescent-shaped chamber is in fluid communication with the fore section of the substantially crescent-shaped chamber. 1. A cooling structure for a stationary blade , comprising:an endwall coupled to a radial end of an airfoil, relative to a rotor axis of a turbomachine, the airfoil including a pressure side surface, a suction side surface, a leading edge, and a trailing edge; anda substantially crescent-shaped chamber positioned within the endwall and radially displaced from the trailing edge of the airfoil, the substantially crescent-shaped chamber receiving a cooling fluid from a cooling circuit, wherein the substantially crescent-shaped chamber extends from a fore section positioned proximal to one of the pressure side surface and the suction side surface of the airfoil to an aft section positioned proximal to the trailing edge of the airfoil and the other of the pressure side surface and the suction side surface of the airfoil,wherein the cooling fluid in the fore section is in thermal communication with a portion of the endwall proximal to one of the pressure side surface and the suction side surface of the airfoil, the cooling fluid in the aft section is in thermal ...

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Номер записи: 83
21-01-2016 дата публикации

A TURBOMACHINE COMPONENT WITH A STRESS RELIEF CAVITY

Номер: US20160017716A1
Автор: Haggmark Anders, Szijarto Janos
Принадлежит: SIEMENS AKTIENGESELLSCHAFT

A turbomachine component with a stress relief cavity includes an airfoil and a platform. The airfoil has a trailing edge. The platform has a trailing edge region, a seal strip slot and a stress relief cavity. The trailing edge region supports at least a part of the trailing edge. The stress relief cavity extends inside the platform into the trailing edge region and is an extension of the seal strip slot.

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Номер записи: 84
21-01-2016 дата публикации

VANE ASSEMBLY

Номер: US20160017731A1
Автор: Becker Jeffrey M., HODGSON BENEDICT N., Johnson Rebecca M., Margetts Michael C., McAdams Danny E.
Принадлежит:

A compressor for a gas turbine engine includes a plurality of rotating wheel assemblies, a plurality of static vane assemblies, and a case extending around the rotating wheel assemblies and static vane assemblies. The static vane assemblies include an inner band, an outer band, and a plurality of vanes extending between the inner and outer bands. 1. A vane ring segment for use in a gas turbine engine , the vane ring segment comprisingan inner band that extends around a portion of a central axis, the inner band including a radially-inner surface facing toward the central axis, a radially-outer surface facing away from the central axis, and a plurality of inner-band vane apertures extending through the radially-inner and radially-outer surfaces of the inner band,an outer band that extends around a portion of a central axis and that is radially spaced apart from the inner band, the outer band including a radially-inner surface facing toward the central axis, a radially-outer surface facing away from the central axis, and a plurality of outer-band vane apertures extending through the radially-inner and radially-outer surfaces of the outer band, anda plurality of vanes coupled to the inner and outer bands, wherein each vane extends radially outward through one of the plurality of outer-band vane apertures beyond the radially-outer surface of the outer band and each vane is bonded to the outer band by a first layer of braze.2. The vane ring segment of claim 1 , wherein each vane includes a body portion that extends from the radially-outer surface of the inner band to the radially-inner surface of the outer band and an outer attachment portion that extends radially outward from the body portion through one of the plurality of outer-band vane apertures beyond the radially-outer surface of the outer band and the first layer of braze is located between the outer attachment portion and the outer band.3. The vane ring segment of claim 2 , wherein a radial cross-section of the ...

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Номер записи: 85
21-01-2016 дата публикации

Off-Cambered Vanes for Gas Turbine Engines

Номер: US20160017732A1
Автор: THOMAS Flavien L.
Принадлежит:

An off-cambered vane for a guide vane assembly in a gas turbine engine is described. The guide vane assembly may comprise a nominal vane having a tip portion, a mid-span portion, and a hub portion. The mid-span portion of the nominal vane may adopt a nominal geometry and the hub portion of the nominal vane may adopt a common geometry. The guide vane assembly may further comprise an off-cambered vane having a tip portion, a mid-span portion, and a hub portion. The mid-span portion of the off-cambered vane may deviate variably with respect to the nominal geometry and at least one of the hub portion and the tip portion may adopt the common geometry.

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Номер записи: 86
21-01-2016 дата публикации

MULTIPLE COATING CONFIGURATION

Номер: US20160017733A1
Автор: Boeke Mark A., Levine Jeffrey R., Riley Sarah
Принадлежит:

An article includes a body that has a coating thereon. The coating has a first portion disposed on a first section of the body and a second portion disposed on a second, different section of the body. The first portion has a first microstructure and the second portion has a second, different microstructure.

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Номер записи: 87
21-01-2016 дата публикации

AXIAL FLOW ROTATING MACHINE AND DIFFUSER

Номер: US20160017734A1
Автор: AKIMOTO Kentaro, Iida Koichiro, Ito Eisaku, NISHIMURA Kazuya
Принадлежит:

An axial flow rotating machine having: a rotor that is provided with a plurality of rotor blades; a stator provided with a plurality of stator blades; an axial flow rotating portion formed by the rotor and the stator; and a diffuser connected to the axial flow rotating portion on the downstream side of the axial flow rotating portion. The final blade portion inner-circumferential inner wall, which is a portion of the inner-circumferential inner wall of the axial flow rotating portion, is formed so that the diameter thereof at the trailing edge position of the final blade is smaller than the diameter at the leading edge position of the final blade. In addition, the diameter of all or a portion of the diffuser inner-circumferential inner wall decreases in a direction of a first side in the axial direction, the first side being the downstream side.

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Номер записи: 88
21-01-2016 дата публикации

VANE TIP MACHINING FIXTURE ASSEMBLY

Номер: US20160017735A1
Автор: Richardson Carl S., Szymbor John A.
Принадлежит:

A vane tip machining fixture assembly includes a first clamp ring mountable to a fixture, the first clamp ring includes a rigid back ring mounted to a resilient ring.

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Номер записи: 89
21-01-2016 дата публикации

SEAL ASSEMBLY FOR A GUIDE VANE ASSEMBLY

Номер: US20160017739A1
Автор: DiBenedetto Enzo, Firnhaber Michael C.
Принадлежит:

The present disclosure relates generally to a guide vane assembly including a first airfoil, including a first airfoil trailing edge, a second airfoil, including a second airfoil leading edge, positioned aft the first airfoil to create a gap therebetween, and a seal assembly disposed within the gap to engage the first airfoil trailing edge and the second airfoil leading edge.

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Номер записи: 90
21-01-2016 дата публикации

RIFFLED SEAL FOR A TURBOMACHINE, TURBOMACHINE AND METHOD OF MANUFACTURING A RIFFLED SEAL FOR A TURBOMACHINE

Номер: US20160017740A1
Автор: McKenna Mike
Принадлежит: SIEMENS AKTIENGESELLSCHAFT

A seal of a turbomachine reduces a leakage flow between a first and second component of the turbomachine. The first component has a first surface and the second component has a second surface, wherein the first component is stiff with regard to a first force exerted perpendicularly thereto and the second component is stiff with regard to a second force exerted perpendicularly thereto. The first surface is opposite the second surface, together defining boundaries of a fluid passage for the leakage flow. The first surface has a first surface riffle. A turbomachine has a seal described above, wherein the turbomachine is a gas turbine engine. A method of manufacturing a first component of a turbomachine with a reduced leakage flow between the first component and a second component of the turbomachine includes fabrication of a first surface riffle, in particular by grinding and/or by electrical discharge machining.

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Номер записи: 91
21-01-2016 дата публикации

Box Rim Cavity for a Gas Turbine Engine

Номер: US20160017741A1
Автор: Ebert Todd A.
Принадлежит:

A gas turbine engine having a rotor with blades and a stationary vane, a platform seal is formed between the blade and vane for inhibiting ingestion of hot gas from a hot gas flow through the turbine into turbine wheel spaces, the platform seal including axial extending platforms on the blade and vane, and radial extending fingers extending from the platforms and forming restrictions between the fingers and the platforms, and a buffer cavity formed between the restrictions, where the fingers are so arranged in a generally radial direction that the vane can be removed from the turbine engine in a radial direction without having to remove the blades first. In additional embodiments, the platform seal assembly can have two or three buffer cavities formed between additional restrictions.

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Номер записи: 92
18-01-2018 дата публикации

SHROUD FOR A GAS TURBINE ENGINE

Номер: US20180016924A1
Автор: Craig, Gallier Kirk Douglas, III Charles William
Принадлежит:

Shrouds and shroud segments for gas turbine engines are provided. In one embodiment, a shroud segment for a gas turbine engine having a rotor blade stage and a nozzle stage is provided. The shroud segment comprises a forward end defining an outer wall of the rotor blade stage and an aft end defining an outer wall of the nozzle stage. The aft end defines at least a portion of an opening therethrough for receipt of a nozzle, and the forward end and the aft end form a single, continuous component. In another embodiment, a gas turbine engine is provided, having a shroud with a forward end positioned near a leading edge of a plurality of rotor blades of a rotor blade stage and an aft end positioned near a trailing edge of a plurality of nozzles of a nozzle stage. Methods of assembling a gas turbine engine also are provided. 1. A shroud segment for a gas turbine engine , the gas turbine engine including a rotor blade stage and a nozzle stage , the shroud segment comprising:a forward end, the forward end defining an outer wall of the rotor blade stage; andan aft end, the aft end defining an outer wall of the nozzle stage, the aft end defining at least a portion of an opening therethrough for receipt of a nozzle,wherein the forward end and the aft end form a single, continuous component.2. The shroud segment of claim 1 , further comprising a first side extending axially from the forward end to the aft end claim 1 , wherein the first side defines a portion of the opening and a side of an adjacent shroud segment defines the remainder of the opening.3. The shroud segment of claim 1 , wherein the aft end defines two openings therethrough claim 1 , and wherein each opening is configured for receipt of a nozzle.4. The shroud segment of claim 1 , further comprising a radially inner surface and a radially outer surface claim 1 , and wherein the nozzle comprises a radially outer end and a radially inner end claim 1 , the radially outer end extending radially beyond the outer surface ...

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Номер записи: 93
18-01-2018 дата публикации

OPTIMIZED AERODYNAMIC PROFILE FOR A TURBINE VANE, IN PARTICULAR FOR A NOZZLE OF THE FOURTH STAGE OF A TURBINE

Номер: US20180016925A1
Автор: COUDERT Laurent Patrick, DESTOUCHES Jean-Armand Marc, Man Daniel Marius, MERVANT Thomas Michel
Принадлежит: SAFRAN AIRCRAFT ENGINES

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine. 1. An aerodynamic profile for a turbine vane , the profile being , when cold and in a non-coated state , substantially identical to a nominal profile determined by the Cartesian coordinates X ,Y ,Zadim given in Table 1 , in which the coordinate Zadim is the quotient D/H , where D is the distance of the point under consideration from a reference X ,Y plane situated at the base of the nominal profile , and H is the height of said profile measured from said reference plane out to the end of the vane , the measurements D and H being taken radially relative to the axis of the turbine , while the coordinate X is measured in the axial direction of the turbine.2. The aerodynamic profile as claimed in claim 1 , wherein said profile is defined within an envelope of ±1 mm in a direction normal to the surface of the nominal profile.3. The aerodynamic profile as claimed in claim 1 , wherein the coordinates X claim 1 ,Y of said profile lie within a range of ±5% relative to the coordinates X claim 1 ,Y of the nominal profile.4. The aerodynamic profile as claimed in claim 1 , wherein the vane is a nozzle vane forming a part of a stator of a turbine.5. The aerodynamic profile as claimed in claim 4 , wherein the vane is a nozzle vane of the fourth stage of the turbine.6. The aerodynamic profile as claimed in claim 4 , wherein the vane is a vane of the fourth stage nozzle ...

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Номер записи: 94
18-01-2018 дата публикации

OPTIMIZED AERODYNAMIC PROFILE FOR AN ARM OF A STRUCTURAL CASING OF A TURBINE, AND STRUCTURAL CASING HAVING SUCH AN ARM

Номер: US20180016926A1
Автор: Leonardon Vincent Nicolas, MARCHE Pierre Hervé, MERVANT Thomas Michel, ROYAN Renaud Gabriel
Принадлежит: SAFRAN AIRCRAFT ENGINES

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine. 1. An aerodynamic profile for an arm of a structural casing of a turbine having a central hub and a shroud , the arm connecting the central hub and the shroud , the profile being , when cold and in a non-coated state , substantially identical to a nominal profile determined by the Cartesian coordinates X ,Y ,Zadim given in Table 1 , in which the coordinate Zadim is the quotient D/H , where D is the distance of the point under consideration from a first reference X ,Y plane situated at the base of the nominal profile , and H is the height of said profile measured from said first reference plane that is the intersection of the stacking axis of the arm and the axisymmetric surface of the hub , out to a second reference plane that is the intersection of said stacking axis with the axisymmetric surface of the shroud , the measurements D and H being taken radially relative to the axis of the turbine , while the coordinate X is measured in the axial direction of the turbine.2. The aerodynamic profile as claimed in claim 1 , wherein said profile is defined within an envelope of ±1 mm in a direction normal to the surface of the nominal profile.3. The aerodynamic profile as claimed in claim 1 , wherein the coordinates X claim 1 ,Y of said profile lie within a range of ±5% relative to the coordinates X claim 1 ,Y of the nominal profile.4. The aerodynamic profile ...

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Номер записи: 95
18-01-2018 дата публикации

TURBINE

Номер: US20180016928A1
Автор: KUWAMURA Yoshihiro, OYAMA Hiroharu, SUGISHITA Hideaki, TANAKA Yoshinori
Принадлежит: Mitsubishi Hitachi Power Systems, Ltd.

A turbine is provided with a seal device. The seal device includes: at least one step surface disposed in a region of an outer peripheral surface of a rotor facing a shroud of a stationary vane in the radial direction of the rotor, the at least one step surface facing upstream in a flow direction of the fluid and dividing the region of the outer peripheral surface into at least two sections in an axial direction of the rotor: at least two seal fins protruding toward the at least two sections from the stationary vane and facing the at least two sections via a seal gap; and a swirling-component application portion disposed on an end side of the shroud of the stationary vane with respect to the axial direction of the rotor and configured to be capable of applying a swirling component to the fluid flowing toward the seal gap. 1. A turbine , comprising:a casing;a rotor extending inside the casing;a plurality of rotor blades fixed to the rotor and arranged in a circumferential direction of the rotor:a plurality of stationary vanes fixed to the casing and arranged in the circumferential direction of the rotor, each of the stationary vanes having a vane body and a shroud which is connected to the vane body and which faces an outer peripheral surface of the rotor via a clearance in a radial direction of the rotor; anda seal device capable of restricting a flow of a fluid in he clearance, at least one step surface disposed in a region of the outer peripheral surface of the rotor facing the shroud of the stationary vane in the radial direction of the rotor, the at least one step surface facing upstream in a flow direction of the fluid and dividing the region of the outer peripheral surface into at least two sections in an axial direction of the rotor;', 'at least two seal fins protruding toward the at least two sections from the stationary vane and facing the at least two sections via a seal gap; and', 'a swirling-component application portion disposed on an end side of the ...

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Номер записи: 96
18-01-2018 дата публикации

ASSEMBLY FOR CONTROLLING VARIABLE PITCH VANES IN A TURBINE ENGINE

Номер: US20180016931A1
Автор: BENDERRADJI Kamel, Bergan Blaise, Bromann Alain Marc Lucien, COUSTILLAS Suzanne Madeleine, DUMAS Lilian Yann
Принадлежит: SAFRAN AIRCRAFT ENGINES

An assembly, in particular for controlling variable pitch vanes in a turbine engine, comprising an actuating ring surrounding a casing of the turbine engine and connected by rods to variable pitch vanes, in addition to a driving means for rotating the actuating ring around the casing. The assembly includes a slidingly connected passive element, one end of which is connected by a sliding pivoting link on the actuating ring and a second end is connected by a ball-joint link to the casing. 1. Assembly for controlling variable pitch vanes in a turbine engine comprising an actuating ring surrounding a casing of the turbine engine and connected by rods to variable pitch vanes in addition to a driving means for rotating the actuating ring around the casing , the assembly comprising a slidingly connected passive element , one end of which is connected by a sliding pivoting link on the actuating ring and a second end is connected by a ball-joint link to the casing.2. Assembly according to claim 1 , wherein the passive element is arranged circumferentially substantially opposite the driving means of the ring.3. Assembly according to claim 1 , wherein the passive element comprises a body bearing the first end and in which a pin bearing the second end is mounted for translational movement.4. Assembly according to claim 1 , wherein the first end of the passive element is mounted for rotation and translational movement around and according to a radial axis in a yoke of a first support element integral with the actuating ring.5. Assembly according to claim 1 , wherein the second end of the passive element is connected via a ball-joint link to a second support element integral with the casing.6. Assembly according to claim 4 , wherein the first support element is arranged axially downstream claim 4 , or upstream respectively claim 4 , from the rods and the second support element is arranged upstream claim 4 , or downstream respectively claim 4 , from the rods.7. Assembly according ...

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Номер записи: 97
18-01-2018 дата публикации

OPTIMIZED AERODYNAMIC PROFILE FOR AN ARM OF A STRUCTURAL CASING OF A TURBINE, AND STRUCTURAL CASING HAVING SUCH AN ARM

Номер: US20180016940A1
Автор: DELABRIERE Maxime Didier, Leonardon Vincent Nicolas, MARCHE Pierre Hervé, Routier Pascal Pierre
Принадлежит: SAFRAN AIRCRAFT ENGINES

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane Psituated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine. 1. An aerodynamic profile for an arm of a structural casing of a turbine having a central hub and a shroud , the arm connecting the central hub and the shroud , the profile being , when cold and in a non-coated state , substantially identical to a nominal profile determined by the Cartesian coordinates X ,Y ,Zadim given in Table 1 , in which the coordinate Zadim is the quotient D/H , where D is the distance of the point under consideration from a first reference X ,Y plane situated at the base of the nominal profile , and H is the height of said profile measured from said first reference plane that is the intersection of the stacking axis of the arm and the axisymmetric surface of the hub , out to a second reference plane that is the intersection of said stacking axis with the axisymmetric surface of the shroud , the measurements D and H being taken radially relative to the axis of the turbine , while the coordinate X is measured in the axial direction of the turbine.2. The aerodynamic profile as claimed in claim 1 , wherein said profile is defined within an envelope of ±1 mm in a direction normal to the surface of the nominal profile.3. The aerodynamic profile as claimed in claim 1 , wherein the coordinates X claim 1 ,Y of said profile lie within a range of ±5% relative to the coordinates X claim 1 ,Y of the nominal profile.4. The aerodynamic profile as ...

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Номер записи: 98
18-01-2018 дата публикации

SHROUD HOUSING SUPPORTED BY VANE SEGMENTS

Номер: US20180016943A1
Автор: GRIVAS NICOLAS, Lefebvre Guy, PARADIS Vincent
Принадлежит:

A shroud mounting arrangement comprises a shroud housing and a shroud mounted to the shroud housing. The shroud is configured to surround a stage of rotor blades of a gas turbine engine. A circumferentially segmented vane ring is disposed axially adjacent to the stage of rotor blades. The circumferentially segmented vane ring comprises a plurality of vane segments. The vane segments jointly support the shroud housing. 1. A turbine assembly for a gas turbine engine , the turbine assembly comprising: a shroud housing supporting a circumferential array of shroud segments about a tip of a circumferential array of turbine blades mounted for rotation about an engine axis , and a circumferentially segmented vane ring mounted to an internal structure of the engine axially adjacent to the circumferential array of turbine blades , the circumferentially segmented vane ring including a plurality of vane segments , the vane segments jointly supporting the shroud housing , the shroud housing being axially restrained on the vane segments by a retaining ring.2. The turbine assembly defined in claim 1 , wherein each vane segment has at least one vane extending between inner and outer platforms claim 1 , and wherein the shroud housing is axially clamped to a mounting structure extending radially outwardly from the outer platform of the vane segments.3. The turbine assembly defined in claim 2 , wherein the shroud housing has an annular body claim 2 , the retaining ring being mounted to an end portion of the annular body claim 2 , and wherein the mounting structure comprises a flange and at least one lug extending radially outwardly from the outer platform of each vane segment claim 2 , the flange being axially clamped between the annular body and the retaining ring claim 2 , the at least one lug being received in a radial slot between the annular body and the retaining ring.4. The turbine assembly defined in claim 1 , wherein each vane segment has at least one vane extending between ...

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Номер записи: 99
17-01-2019 дата публикации

TURBOMACHINE AIRFOIL ARRAY

Номер: US20190017384A1
Автор: BRETTSCHNEIDER Markus, MAATOUK Fadi, Mahle Inga
Принадлежит:

An airfoil array for a turbomachine, in particular a turbine or compressor stage of a gas turbine. The airfoil array includes at least two airfoils and at least one contoured circumferential surface which connects a pressure side of one airfoil to a suction side of the other airfoil and includes an upstream first section and a downstream second section which adjoins the first section along an elevation contour line; the first section being depressed relative to a rotationally symmetric reference surface containing this elevation contour line away from the airfoils, and the second section not being depressed relative to this reference surface away from the airfoils; this elevation contour line lying in an axial area which terminates at most 30% of an axial chord length of one of the airfoils downstream of its leading edge; and an axial distance of this elevation contour line increasing toward the pressure side and toward the suction side, starting at a point between the pressure and suction sides that is closest to the leading edge; and the first section extending over at least 90% of the space between the pressure side and the suction side. 1. An airfoil array for a turbomachine , the airfoil array comprising:at least a first and a second airfoil and at least one contoured circumferential surface connecting a pressure side of the first airfoil to a suction side of the second airfoil, the contoured circumferential surface including an upstream first section and a downstream second section adjoining the first section along an elevation contour line, the first section being depressed relative to a rotationally symmetric reference surface containing the elevation contour line away from the first and second airfoils, and the second section not being depressed relative to the reference surface away from the first and second airfoils, the elevation contour line lying in an axial area terminating at most 30% of an axial chord length of one of the first and second airfoils ...

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Номер записи: 100