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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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16-09-2019 дата публикации

Устройство для регулирования радиального зазора

Номер: RU0000192393U1
Автор: Елена Владимировна Щербакова, Сергей Викторович Белов, Юрий Юрьевич Рыкачев
Принадлежит: Федеральное государственное унитарное предприятие "Центральный институт авиационного моторостроения им. П.И. Баранова"

Полезная модель относится к устройствам для регулирования радиального зазора между ротором и статором турбомашины. Устройство для регулирования радиального зазора между ротором и статором турбомашины содержит надроторную вставку, управляющее кольцо, кинематически связанное с надроторной вставкой, и по меньшей мере три силовых агрегата, шарнирно связанных с управляющим кольцом, при этом надроторная вставка состоит из секций, частично входящих друг в друга и последовательно расположенных таким образом, что их внутренняя поверхность образует поверхность, концентричную поверхности, описываемой торцами лопаток, а каждый силовой агрегат содержит шток. Управляющее кольцо выполнено неразъемным и установлено с возможностью перемещения и наклона относительно продольной оси турбомашины под действием штоков силовых агрегатов, при этом управляющее кольцо связано с каждой секцией надроторной вставки и штоком каждого силового агрегата с помощью шаровых шарниров, расположенных с противоположных сторон управляющего кольца. Внутренняя поверхность секций надроторной вставки образует коническую поверхность. Технический результат - уменьшение усилий, затрачиваемых на перемещение надроторной вставки, регулирующей радиальный зазор. 7 ил. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 192 393 U1 (51) МПК F01D 11/22 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ (52) СПК F01D 11/22 (2019.08) (21)(22) Заявка: 2019119228, 20.06.2019 (24) Дата начала отсчета срока действия патента: Дата регистрации: Приоритет(ы): (22) Дата подачи заявки: 20.06.2019 (45) Опубликовано: 16.09.2019 Бюл. № 26 1 9 2 3 9 3 R U (56) Список документов, цитированных в отчете о поиске: RU 2499891 C1, 27.11.2013. RU 87213 U1, 27.09.2009. WO 2010112421 A1, 07.10.2010. EP 1741880 A2, 10.01.2007. JPS 57195803 A, 01.12.1982. SU 966247 A1, 15.10.1982. (54) Устройство для регулирования радиального зазора (57) Реферат: Полезная модель относится к устройствам для агрегат содержит ...

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Номер записи: 1
26-01-2012 дата публикации

Seal assembly for controlling fluid flow

Номер: US20120017594A1
Автор: Burkhard Voss, Christian Kowalski, Fan Zhang, Robert W. Sunshine, Uwe Lohse
Принадлежит: SIEMENS AG, Siemens Energy Inc

A seal assembly ( 50, 60 ) for a gas turbine engine for controlling air flow between a diffuser ( 48 ) and rotor disks comprising first and second annular flange ends ( 52, 54 ) and an annular seal mid-section ( 56 ) between and operatively connected to the flange ends ( 52, 54 ). The first and second annular flange ends ( 52, 54 ) abut respective outer frame members ( 46 ) of the diffuser, whereby a fluid flow path is formed between the seal assembly ( 50, 60 ) and the rotor disks ( 42 ). The first and second end flanges ( 52, 54 ) are composed of a material having a coefficient of thermal expansion that is substantially the same as a coefficient of thermal expansion of the material of the outer frame members ( 46 ). In addition, the material of the seal mid-section ( 56 ) has a coefficient of thermal expansion that is different than that of the materials of the annular flange ends ( 52, 54 ) and outer frame members ( 46 ).

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Номер записи: 2
02-02-2012 дата публикации

Blade outer air seal and repair method

Номер: US20120027574A1
Автор: Philip R. Belanger, Susan M. Tholen
Принадлежит: United Technologies Corp

An article of manufacture has a body formed in part of a first metal alloy and in part of a second metal alloy, the second metal alloy having a thermal coefficient of expansion that is less than the thermal coefficient of expansion of the first metal alloy. A BOAS segment for a gas turbine engine is disclosed wherein the formation of cracks due to thermal mechanical fatigue in the body of the disclosed BOAS segment is minimized, if not eliminated, through a unique construction of the disclosed BOAS segment, whether original equipment manufacture or a repaired blade outer air seal. A method for manufacture of a BOAS segment and a method for modifying a BOAS segment are disclosed.

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Номер записи: 3
01-03-2012 дата публикации

Electroformed conforming rubstrip

Номер: US20120051922A1
Автор: Gary M. Lomasney, James O. Hansen, Joseph Parkos, Thomas M. Morin
Принадлежит: Individual

A disk made of a first material has a groove in which a blade made of a second material is retained. A strip is placed between the blade and the disk to minimize rubbing damage to the blade and the disk and an insulating material is place between the rub strip and the blade for minimizing damaging responses of the blade to galvanic forces created by rubbing of the first material and the second material.

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Номер записи: 4
15-03-2012 дата публикации

Abradable bucket shroud

Номер: US20120063881A1
Автор: James Albert Tallman
Принадлежит: General Electric Co

The present application provides an abradable bucket shroud for use with a bucket tip so as to limit a leakage flow therethrough and reduce heat loads thereon. The abradable bucket shroud may include a base and a number of ridges positioned thereon. The ridges may be made from an abradable material. The ridges may form a pattern. The ridges may have a number of curves with at least a first curve and a second curve and with the second curve having a reverse camber shape.

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Номер записи: 5
22-03-2012 дата публикации

Turbine shroud thermal distortion control

Номер: US20120070276A1
Автор: Gajawalli V. Srinivasan, Glenn N. Levasseur, Jun Shi, Kevin E. Green, Shaoluo L. Butler
Принадлежит: United Technologies Corp

A shroud suitable for use in combination with an adjacent rotor blade includes a leading portion comprising a first thickness, and a trailing portion adjacent to the leading portion and comprising a second thickness, wherein the second thickness is less than the first thickness so that stiffness in the leading portion is greater than in the trailing portion and radial expansion by thermal growth of the leading portion of the shroud is constrained.

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Номер записи: 6
19-04-2012 дата публикации

Bonded turbine bucket tip shroud and related method

Номер: US20120093634A1
Автор: Gary Charles Liotta
Принадлежит: General Electric Co

A turbine bucket includes an airfoil portion and a tip shroud at a radially outer end of the airfoil portion. The tip shroud includes a first radially inner tip shroud component formed integrally with the airfoil portion and composed of a first metal material, and a second radially outer structural tip shroud component composed of a second metal material bonded to the first radially inner tip shroud component.

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Номер записи: 7
31-05-2012 дата публикации

Gas turbine of the axial flow type

Номер: US20120134779A1
Автор: Alexander Anatolievich Khanin, Valery Kostege
Принадлежит: Individual

In an axial flow gas turbine ( 30 ), a reduction in cooling air mass flow and leakage in combination with an improved cooling and effective thermal protection of critical parts within the turbine stages of the turbine is achieved by providing, within a turbine stage (TS), devices ( 43 - 48 ) to direct cooling air that has already been used to cool, especially the airfoils of the vanes ( 31 ) of the turbine stage (TS), into a first cavity ( 41 ) located between the outer blade platforms ( 34 ) and the opposed stator heat shields ( 36 ) for protecting the stator heat shields ( 36 ) against the hot gas and for cooling the outer blade platforms ( 34 ).

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Номер записи: 8
28-06-2012 дата публикации

System and method for rotary machine online monitoring

Номер: US20120162192A1
Автор: Guanghua Wang, James Peter Delancey, Jeremy Clyde Bailey, Lucy Joelle Summerville, Nirm Velumylum Nirmalan
Принадлежит: General Electric Co

In one embodiment, a system includes an optical monitoring system configured to optically communicate with an interior of a rotary machine. The optical monitoring system is configured to redirect a field of view toward different regions of a component within the interior of the rotary machine while the rotary machine is in operation, and to capture an image of each region.

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Номер записи: 9
02-08-2012 дата публикации

Turbine bucket for use in gas turbine engines and methods for fabricating the same

Номер: US20120195742A1
Автор: Matthew Durham Collier, Rajnikumar Nandalal Suthar, Sanjeev Kumar Jain
Принадлежит: General Electric Co

A turbine bucket for use with a turbine engine. The turbine bucket includes a dovetail that is coupled to a rotor assembly that is positioned within a turbine casing. A platform extends from the dovetail. An airfoil extends from the platform. The airfoil includes a root end and a tip end. The tip end extends outwardly from the root end towards the turbine casing. A tip shroud extends from the tip end. The tip shroud includes a shroud plate. A first shroud rail extending a first radial distance from the shroud plate towards the turbine casing. A second shroud rail extends a second radial distance from the shroud plate towards the turbine casing that is different than the first radial distance.

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Номер записи: 10
09-08-2012 дата публикации

Passive cooling system for a turbomachine

Номер: US20120201650A1
Автор: Joshy John, Rajnikumar Nandalal Suthar, Sanjeev Kumar Jain
Принадлежит: General Electric Co

A turbomachine includes a housing having an outer surface and an inner surface that defines an interior portion. The housing includes a fluid plenum. A rotating member is arranged within the housing. The rotating member includes at least one bucket having a base portion and a tip portion. A stationary member is mounted to the inner surface of the housing adjacent the tip portion of the at least one bucket. At least one fluid passage passes through at least a portion of the stationary member. The at least one fluid passage includes a fluid inlet fluidly coupled to the fluid plenum and a fluid outlet exposed to the interior portion. The fluid outlet being configured and disposed to direct a flow of fluid toward the tip portion of the at least one bucket.

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Номер записи: 11
18-10-2012 дата публикации

Apparatus to seal with a turbine blade stage in a gas turbine

Номер: US20120260670A1
Автор: Andres Jose Garcia-Crespo, Gregory Thomas Foster
Принадлежит: General Electric Co

Disclosed is an apparatus configured to seal with a turbine blade stage of a gas turbine. The apparatus includes an outer shroud coupled to an inner shroud and configured to circumferentially surround the turbine blade stage. The inner shroud is configured to circumferentially surround the turbine blade stage to seal with the turbine blade stage and includes an attachment element configured to be inserted into the outer shroud to couple the inner shroud to the outer shroud.

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Номер записи: 12
20-12-2012 дата публикации

Plug assembly for blade outer air seal

Номер: US20120319360A1
Автор: Ken F. Blaney, Paul M. Lutjen
Принадлежит: Individual

A plug assembly includes a cup which defines a cup portion and a cup anti-liberation portion along an axis, a wedge is mountable within the cup portion to at least partially radially expand the cup anti-liberation portion.

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Номер записи: 13
10-01-2013 дата публикации

Gas turbine shroud arrangement

Номер: US20130008176A1
Автор: Jun Shi, Kevin E. Green, Paul F. Croteau
Принадлежит: United Technologies Corp

A system for supporting a shroud used in an engine has a shroud positioned radially outboard of a rotor, which shroud has a plurality of circumferentially spaced slots; a forward support ring for supporting the shroud; the forward support ring having a plurality of spaced apart first tabs on a first side for functioning as anti-rotation devices; the forward support ring having a plurality of spaced apart second tabs on a second side; and the second tabs engaging the slots in the shroud and circumferentially supporting the shroud.

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Номер записи: 14
10-01-2013 дата публикации

Reduction in thermal stresses in monolithic ceramic or ceramic matrix composite shroud

Номер: US20130011248A1
Автор: Jun Shi, Kevin E. Green, Paul F. Croteau
Принадлежит: United Technologies Corp

A shroud for use in a gas turbine engine has a ring with a plurality of protrusions and a plurality of slots and each of the protrusions having an arc length and parallel sides.

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Номер записи: 15
18-04-2013 дата публикации

SEAL STRUCTURE, TURBINE MACHINE HAVING THE SAME, AND POWER GENERATING PLANT EQUIPPED WITH THE SAME

Номер: US20130094945A1
Автор: KUWAMURA Yoshihiro, MATSUMOTO Kazuyuki, Matsuo Asaharu, OYAMA Hiroharu, TANAKA Yoshinori
Принадлежит: MITSUBISHI HEAVY INDUSTRIES, LTD.

A seal structure includes shrouds that are provided at tips of a plurality of blades on which a main flow of fluid impinges and that come into contact with adjacent shrouds to form a cylindrical shape; a sealing part connected to a structure body facing the shrouds; and a space that is defined by downstream sides of the shrouds, a downstream side of the sealing part, and a wall on a downstream side of the structure body, and that communicates with the main flow of fluid that has impinged on the blades. An end of the wall on the downstream side of the structure body is formed so as to be disposed a predetermined distance away from a shroud extension and away from a sealing-part extension and so as to be located on the opposite side of the shroud extension with respect to the sealing-part extension. 1. A seal structure comprising:shrouds that are provided at tips of a plurality of blades on which a main flow of fluid impinges and that come into contact with adjacent shrouds to form a cylindrical shape;a sealing part connected to a structure body facing the shrouds; anda space that is defined by downstream sides of the shrouds, a downstream side of the sealing part, and a wall on a downstream side of the structure body and that communicates with the main flow of fluid that has impinged on the blades,wherein an end of the wall on the downstream side of the structure body is formed so as to be disposed a predetermined distance away from a shroud extension extending toward the downstream side of the shrouds and away from a sealing-part extension provided substantially parallel to the shroud extension and extending toward the downstream side of the sealing part and so as to be located on an opposite side of the shroud extension with respect to the sealing part extension.2. The seal structure according to claim 1 , wherein surfaces of the shrouds facing the structure body are provided substantially parallel to a flow direction of the main flow of fluid that has impinged on ...

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Номер записи: 16
18-04-2013 дата публикации

TURBINE SHROUD THERMAL DISTORTION CONTROL

Номер: US20130094946A1
Автор: Butler Shaoluo L., Green Kevin E., Levasseur Glenn N., Shi Jun, Srinivasan Gajawalli V.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A shroud for a gas turbine engine includes a leading portion having a leading edge and a first set of circumferentially spaced slots at the leading edge that break up the leading portion into circumferentially spaced segments separated by the first set of slots, and a trailing portion adjacent to the leading portion. The trailing portion has a trailing edge. 1. A shroud for a gas turbine engine , the shroud comprising:a leading portion having a leading edge and a first set of circumferentially spaced slots at the leading edge that break up the leading portion into circumferentially spaced segments separated by the first set of slots; anda trailing portion adjacent to the leading portion, the trailing portion having a trailing edge.2. The shroud of claim 1 , wherein the first set of slots have an open end at the leading edge and extend towards the trailing edge to a closed end within the shroud claim 1 , and extend radially through a full thickness of the leading portion of the shroud.3. The shroud of claim 2 , wherein the first set of slots extend in an axial direction.4. The shroud of claim 1 , wherein the trailing portion further comprises a second set of circumferentially spaced slots at the trailing edge that break up the trailing portion into circumferentially spaced segments separated by the second set of slots.5. The shroud of claim 4 , wherein the first set of slots and the second set of slots are staggered with respect to each other.6. The shroud of claim 4 , wherein the second set of slots extend in an axial direction.7. The shroud of claim 1 , wherein each slot has a length approximately 40% of an axial length of the shroud.8. The shroud of claim 1 , wherein at least one slot has an open end at the leading edge and extend towards the trailing edge to a closed end within the shroud claim 1 , where opposite edges of the slot proximate the open are arranged substantially parallel to each other claim 1 , and wherein the closed end defines a bulbous portion.9. ...

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Номер записи: 17
23-05-2013 дата публикации

Systems and Methods for Adjusting Clearances in Turbines

Номер: US20130129470A1
Автор: Erwing Calleros, Ezio Pena, Jose-Quintino Da-Costa, Nicolas Antoine, Prabhakaran Saraswathi Rajesh, Rahul J. Chillar
Принадлежит: General Electric Co

Embodiments of the invention can provide systems and methods for adjusting clearances in a turbine. According to one embodiment of the invention, there is disclosed a turbine system. The system may include one or more turbine blades; a turbine casing encompassing the one or more turbine blades; and a thermoelectric element disposed at least partially about the turbine casing, wherein the thermoelectric element expands or contracts the turbine casing by heating or cooling at least a portion of the turbine casing, thereby adjusting a clearance between the one or more turbine blades and the turbine casing.

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Номер записи: 18
30-05-2013 дата публикации

SHROUD SEGMENT PRODUCING METHOD AND SHROUD SEGMENT

Номер: US20130136582A1
Автор: Mizokami Yousuke, Tamura Takashi, Tao Nobuya
Принадлежит:

Disclosed herein is a production method of a shroud segment that includes a forming process of molding a cylindrical fiber fabric () into a shroud segment shape by pressing a cylindrical surface of the fiber fabric, and a matrix forming process of impregnating the fiber fabric molded into the shroud segment shape with a matrix. 1. A production method of a shroud segment made of a fiber-reinforced composite material which is arranged between a casing enclosing a rotor blade and the rotor blade by locking a hook portion in a gas turbine engine , the production method of a shroud segment comprising:a forming process of molding a cylindrical fiber fabric into a shroud segment shape by pressing a cylindrical surface of the fiber fabric; anda matrix forming process of impregnating the fiber fabric molded into the shroud segment shape with a matrix.2. The production method of a shroud segment according to claim 1 , wherein when the cylindrical surface of the fiber fabric is pressed at the forming process claim 1 , a gap to allow excessive deformation of the fiber fabric is provided at a part other than a part corresponding to the hook portion.3. The production method of a shroud segment according to claim 1 , wherein a reinforcement member is arranged and accommodated in the cylindrical fiber fabric and the fiber fabric is molded claim 1 , together with the reinforcement member claim 1 , at the forming process.4. The production method of a shroud segment according to claim 2 , wherein a reinforcement member is arranged and accommodated in the cylindrical fiber fabric and the fiber fabric is molded claim 2 , together with the reinforcement member claim 2 , at the forming process.5. A shroud segment made of a fiber-reinforced composite material which is arranged between a casing enclosing a rotor blade and the rotor blade by locking a hook portion in a gas turbine engine claim 2 ,wherein the shroud segment is made of the fiber-reinforced composite material including a ...

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Номер записи: 19
06-06-2013 дата публикации

Abradable seal with axial offset

Номер: US20130142628A1
Автор: Nuo Sheng, Roderick Mark Lusted, Vittorio Michelassi
Принадлежит: Individual

A sealing system for a centrifugal compressor includes a stator having a seal, a seal disposed in the seal housing and having an abradable portion along an inner circumference, a rotor having a plurality of rotor teeth configured to rotate within the inner circumference of the seal and configured to create rub grooves within the abradable portion, and a first spring disposed between the stator and the seal and configured to facilitate axial movement of the seal relative to the seal housing.

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Номер записи: 20
06-06-2013 дата публикации

Preforms and Related Methods for Repairing Abradable Seals of Gas Turbine Engines

Номер: US20130142629A1
Автор: Hall Jacob A., Minor Michael
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A preform, for repairing an abradable seal component of a gas turbine engine, includes a multilayer stack that has a first layer and a second layer. The first layer is operative to bond the multilayer stack to a structural substrate of an abradable seal. The first layer includes structural material that corresponds to a material of the structural substrate and braze material that is compatible with the structural material. The second layer includes an abradable material. The multilayer stack is operative to bond to the structural substrate of the abradable seal component during a brazing process such that the second layer forms a replacement abradable layer of the abradable seal component. 1. A preform for repairing an abradable seal component of a gas turbine engine , the seal component having a structural substrate and an abradable layer , said preform comprising:a multilayer stack having a first layer and a second layer;the first layer being operative to bond the multilayer stack to the structural substrate, the first layer comprising structural material corresponding to material of the structural substrate and braze material compatible with the structural material;the second layer comprising abradable material;the multilayer stack being operative to bond to the structural substrate of the abradable seal component during a brazing process such that the second layer forms a replacement abradable layer of the abradable seal component.2. The preform of claim 1 , wherein each of the first layer and the second layer comprises a cobalt-based claim 1 , silicon-depressed braze alloy with less than approximately 1% boron.3. The preform of claim 1 , wherein the preform is arcuate such that the abradable layer is positioned at a concave radially inner diameter surface of the multilayer stack.4. The preform of claim 1 , wherein an axial outer periphery of the second layer extends outwardly beyond an axial outer periphery of the second layer prior to brazing the preform to ...

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Номер записи: 21
06-06-2013 дата публикации

TURBOMACHINE

Номер: US20130142641A1
Автор: Landi Giacomo
Принадлежит: NUOVO PIGNONE S.P.A.

A turbo machine is provided. The turbo machine comprises a turbo stator having a shroud, a turbo rotor having an impeller within the shroud, a brush seal between the impeller and the shroud, and at least one vane extending from the shroud toward the impeller upstream of the brush seal. 1. A turbo machine comprising:a turbo stator having a shroud;a turbo rotor having an impeller within the shroud;a brush seal between the impeller and the shroud; andat least one vane extending from the shroud toward the impeller upstream of the brush seal.2. The turbo machine of wherein the at least one vane comprises:an upstream end;a downstream end;a first side extending between the upstream end and the downstream end; anda second side extending between the upstream end and the downstream end.3. The turbo machine of wherein the shroud has a surface facing the impeller and the at least one vane comprises an impeller facing surface having an upstream end intersecting the shroud surface and a downstream end intersecting the shroud surface claim 1 , the impeller facing surface being substantially congruent to the impeller from the upstream end to the downstream end.4. The turbo machine of wherein the at least one vane defines a plane coincident with the rotor axis.5. The turbo machine of wherein the impeller has an outer diameter and the at least one vane extends radially outwardly beyond the outer diameter of the impeller.6. The turbo machine of further comprising a main cavity portion upstream of the brush seal claim 1 , the main cavity portion being defined by a recessed surface of the shroud and a surface of the impeller claim 1 , the at least one vane being disposed at least partially within the main cavity.7. The turbo machine of wherein the main cavity shroud surface comprises a step defined by a planar shroud surface normal to the rotor axis and a cylindrical shroud surface upstream of the planar surface claim 6 , the cylindrical surface intersecting the planar shroud surface.8. ...

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Номер записи: 22
20-06-2013 дата публикации

CONTROLLER

Номер: US20130152601A1
Автор: BACIC Marko
Принадлежит: ROLLS-ROYCE PLC

A gas turbine engine has, in flow series, a compressor section, a combustor, and a turbine section. The gas turbine engine further has a system for cooling the turbine section and providing tip clearance control between turbine blades and circumferentially distributed segments forming an annular shroud surrounding the blades outer tips. The turbine section cooling sub-system diverts a first cooling air flow, regulated by a first valve arrangement, from the compressor section to a heat exchanger and then to the turbine section to cool its components. The tip clearance control sub-system supplies a second cooling air flow, regulated by a second valve arrangement, to an engine case where the segments are mounted, which regulates thermal expansion of the case and controls the clearance between the segments and outer tips. The system further includes a closed-loop controller which issues demand signals to the first and second valve arrangements. 1. A gas turbine engine having , in flow series , a compressor section , a combustor , and a turbine section , and further having a system (i) for cooling the turbine section and (ii) for providing tip clearance control between turbine blades of the turbine section and a plurality of circumferentially distributed segments which form an annular shroud surrounding the outer tips of the turbine blades , the system including:a turbine section cooling sub-system which diverts a first cooling air flow received from the compressor section to a heat exchanger and then to the turbine section to cool components thereof, the first cooling air flow by-passing the combustor and being cooled in the heat exchanger, and the turbine section cooling subsystem having a first valve arrangement which regulates the first cooling air flow;a tip clearance control sub-system which supplies a second cooling air flow to an engine case to which the segments are mounted, the second cooling air flow regulating thermal expansion of the case and thereby ...

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Номер записи: 23
22-08-2013 дата публикации

SEAL STRUCTURE AND ROTATING MACHINE EQUIPPED THEREWITH

Номер: US20130216362A1
Автор: NISHIMOTO Shin, ONISHI Tomoyuki
Принадлежит: MITSUBISHI HEAVY INDUSTRIES, LTD.

Provided is a seal structure, which includes fins configured to protrude from an outer circumferential surface of a rotor in a circumferential direction, and stator blades having an abradable coating formed on an inner circumferential surface of an inner shroud so as to face the fins. The inner circumferential surface of the inner shroud is formed in an uneven shape, and the abradable coating is formed along the uneven shape. 1. A seal structure comprising:a fin configured to protrude from an outer circumferential surface of a rotor in a circumferential direction; anda stator blade having an abradable coating formed on an inner circumferential surface of an inner shroud so as to face the fins,wherein the inner circumferential surface of the inner shroud is formed in an uneven shape, and the abradable coating is formed along the uneven shape.2. The seal structure according to claim 1 , wherein the uneven shape is configured by a concave portion formed from one of the inner circumferential surface of the inner shroud and an outer circumferential surface of the abradable coating toward an interior thereof.3. The seal structure according to claim 2 , wherein the concave portion is formed so as to extend in the circumferential direction.4. The seal structure according to claim 2 , wherein the concave portion is formed so as to extend in an axial direction of the rotor.5. The seal structure according to claim 3 , wherein the concave portion is formed on a boundary line between the inner shrouds adjacent in the circumferential direction.6. The seal structure according to claim 4 , wherein:the one of the inner circumferential surface of the inner shroud and the outer circumferential surface of the abradable coating is formed with a second concave portion so as to be opposite to the concave portion formed in the other of the inner circumferential surface of the inner shroud and the outer circumferential surface of the abradable coating; andthe seal structure includes a pin ...

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Номер записи: 24
10-10-2013 дата публикации

SEALING SYSTEM FOR A TURBOMACHINE

Номер: US20130266426A1
Автор: BOECK ALEXANDER, Mahle Inga
Принадлежит: MTU AERO ENGINES GMBH

Disclosed is a sealing system for a turbomachine, in particular for a gas turbine, the sealing system being formed in an annular space between a flow-limiting wall of the turbomachine and at least one rotor blade tip of a rotor blade or an outer shroud arranged on the rotor blade tip, and comprising at least one sealing point. The sealing point comprises at least one run-in coating arranged on the rotor blade tip or on the outer shroud in the direction of the flow-limiting wall of the turbomachine. The invention furthermore encompasses a gas turbine comprising the sealing system. 1. A sealing system for a turbomachine , wherein the sealing system is formed in an annular space between a flow-limiting wall of the turbomachine and at least one rotor blade tip of a rotor blade or an outer shroud arranged on the rotor blade tip , and comprises at least one sealing point , which sealing point comprises at least one run-in coating arranged on the rotor blade tip or on the outer shroud in a direction of the flow-limiting wall of the turbomachine.2. The sealing system of claim 1 , wherein the at least one sealing point comprises at least one sealing tip which is opposite the run-in coating and is arranged on an inner side of the wall.3. The sealing system of claim 1 , wherein the sealing system comprises at least two sealing points claim 1 , the sealing points consisting of in each case one run-in coating arranged on the rotor blade tip or on the outer shroud and of in each case at least one sealing tip which is opposite a particular run-in coating and is arranged on the inner side of the wall claim 1 , and wherein the run-in coatings are arranged one behind another in a flow direction.4. The sealing system of claim 1 , wherein the sealing system has at least two sealing points claim 1 , a first sealing point comprising a sealing tip arranged on the rotor blade tip or on the outer shroud claim 1 , the sealing tip being arranged opposite a run-in coating arranged on an inner ...

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Номер записи: 25
10-10-2013 дата публикации

SEALING SYSTEM FOR A TURBOMACHINE

Номер: US20130266427A1
Автор: BOECK ALEXANDER, Mahle Inga
Принадлежит: MTU AERO ENGINES GMBH

A sealing system for a turbomachine, in particular, for a gas turbine, having an annular space between a flow-restricting wall and at least one series of rotor blades including a plurality of rotor blades. The sealing system comprises at least one first sealing fin disposed on an end of the rotor blade facing the wall, at least one second abradable lining disposed following the first sealing fin in the flow direction on the end of the rotor blade facing the wall, at least one first abradable lining disposed on the inside of the wall and opposite the first sealing fin, and at least one second sealing fin disposed in the flow direction following the first abradable lining on the inside of the wall and opposite the second abradable lining. The invention furthermore relates to a gas turbine comprising at least one sealing system. 1. A sealing system for a turbomachine having a flow-restricting wall and at least one series of rotor blades including a plurality of rotor blades , the flow-restricting wall and the at least one series of rotor blades defining an annular space therebetween having a flow direction , the sealing system comprising:a first seal and a second seal disposed in the annular space between the flow-restricting wall and the at least one series of rotor blades; a first sealing fin that is disposed on an end of a rotor blade facing the flow-restricting wall, the rotor blade being from the at least one series of rotor blades, and', 'a first abradable lining that is disposed on an inside of the flow-restricting wall and is opposite the first sealing fin; and, 'wherein the first seal includes'} a second abradable lining that is disposed on the end of the rotor blade facing the flow-restricting wall, and', 'a second sealing fin that is disposed on the inside of the flow-restricting wall and is opposite the second abradable lining., 'wherein the second seal is disposed following the first seal in the flow direction and includes'}2. A sealing system in ...

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Номер записи: 26
28-11-2013 дата публикации

METHOD FOR ADJUSTING THE RADIAL GAPS WHICH EXIST BETWEEN BLADE AIRFOIL TIPS OR ROTOR BLADES AND A PASSAGE WALL

Номер: US20130312249A1
Автор: Buchal Tobias, Buchholz Bjorn, HILLE THOMAS, Kapustina Anna, Milazar Mirko, Schwerin Eric
Принадлежит:

A method for measuring and adjusting gaps between a rotor and a stator of a machine using a sensor is provided. The measuring is conducted when the rotor is operated at a rotational speed below a nominal rotational speed of the machine and without the machine being in operation. The adjusting of the gap is carried out as a function of at least one gap dimension of the gap. The sensor is not resistant to an operating temperature of the machine occurring in a region where the sensor is located. After completion of the measuring the machine is operated with the sensor. 15.-. (canceled)6. A method for measuring and adjusting gaps between a rotor and a stator of a machine , comprising:providing a sensor;measuring a gap between a rotor and a stator using the sensor, wherein the measuring is conducted when the rotor is operated at a rotational speed below a nominal rotational speed of the machine and without the machine being in operation; andadjusting the gap, wherein the adjusting of the gap is carried out as a function of at least one gap dimension of the gap,wherein the sensor is not resistant to an operating temperature of the machine occurring in a region where the sensor is located, andwherein after completion of the measuring the machine is operated with the sensor.7. The method as claimed in claim 6 , wherein a radial gap in a turbine and/or in a compressor is measured.8. The method as claimed in claim 6 , wherein the method is carried out before commissioning of the machine and/or before an initial startup of the machine after maintenance.9. The method as claimed in claim 6 , wherein the rotational speed of the rotor is equal to or less than 120 min.10. The method as claimed in claim 6 , wherein the machine is an axial throughflow turbomachine and the gap between the rotor and the stator is located between rotor blades and a passage wall. This application is the US National Stage of International Application No. PCT/EP2011/059583 filed Jun. 9, 2011, and claims ...

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Номер записи: 27
28-11-2013 дата публикации

Turbomachine having clearance control capability and system therefor

Номер: US20130315716A1
Автор: Joseph Anthony Cotroneo
Принадлежит: General Electric Co

A turbomachine having clearance control capability is provided and includes a turbine stage including a blade configured to rotate around a centerline, a movable portion of a casing circumferentially surrounding the turbine stage and a rotatable cam operably coupled to the movable portion and thereby configured to control an axial position of the movable portion. A radially outermost tip of the blade and an interior surface of the movable portion are sloped with respect to the centerline such that the controlled axial position of the movable portion is determinative of a clearance between the blade and the movable portion.

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Номер записи: 28
05-12-2013 дата публикации

ROTARY MACHINE

Номер: US20130323034A1
Автор: Isogai Tomoyuki
Принадлежит: TOYOTA JIDOSHA KABUSHIKI KAISHA

In a compressor () of a turbocharger (), a compressor wheel () is provided in a housing () to be capable of rotating. When the wheel () rotates, air suctioned through an inlet of the housing () is compressed and then discharged through an outlet of the housing, (). Further, an abradable seal layer () formed on an inner surface of the housing () is abraded by a vane () of the rotating wheel () such that a tip clearance between the vane () and a part of the inner surface of the housing () that opposes the vane () is adjusted. A corner portion () of the vane () on the outlet side of the housing () is shaped to move gradually further away from a shroud curve (Lc) of the seal layer () toward an end portion of the vane () on the outlet side of the housing (). 1. A rotary machine comprising:a housing; andan impeller having a plurality of vanes and provided in the housing to be rotatable about a shaft,wherein a fluid that flows into the housing passes between the vanes of the impeller and then flows out of the housing,an abradable seal layer, which is abraded by the vanes to adjust a tip clearance between the abradable seal layer and the vanes when the impeller rotates, is provided on an inner surface of the housing such that a surface of the vane and a surface of the abradable seal layer, which oppose each other, are shaped to follow a predetermined shroud curve, anda corner portion of each of the vanes on an outlet side of the housing is shaped such that a distance between the corner portion and the shroud curve of the abradable seal layer gradually increases toward an end portion of the vanes on the outlet side of the housing.2. The rotary machine according to claim 1 , wherein the corner portion of each of the vanes on the outlet side of the housing is shaped such that an end of the corner portion on the outlet side of the housing is withdrawn to a position removed from the shroud curve of the abradable seal layer by a predetermined distance claim 1 , and so as to ...

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Номер записи: 29
19-12-2013 дата публикации

Turbine compressor blade tip resistant to metal transfer

Номер: US20130333392A1
Автор: Christopher W. STROCK, Matthew E. Bintz
Принадлежит: United Technologies Corp

A gas turbine engine having an engine casing extending circumferentially about an engine centerline axis; and a compressor section, a combustor section, and a turbine section within said engine casing. At least one of said compressor section and said turbine section includes at least one airfoil and at least one seal member adjacent to the at least one airfoil, wherein a tip of the at least one airfoil is metal having a thin film ceramic coating and the at least one seal member is coated with an abrasive.

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Номер записи: 30
16-01-2014 дата публикации

INTER STAGE SEAL HOUSING HAVING A REPLACEABLE WEAR STRIP

Номер: US20140015200A1
Автор: Garner Chad, Scimeca Santo F.
Принадлежит: MITSUBISHI POWER SYSTEMS AMERICAS, INC.

An inter stage seal housing for a turbine engine having upper and lower half inter stage seal housings in which a contact sealing surface of the seal housing is restored after an interval of engine operation. The contact sealing surface is restored by fitting a replaceable wear strip on the downstream sealing surface of the seal housing. In order to fit the replaceable wear strip, a circumferential groove is machined along an outer peripheral edge of the seal housing. The groove is machined to include axial location and radial retention such that the wear strips can be slid into the upper half and lower half inter stage seal housing circumferentially from the horizontal joint. The groove includes through holes and the wear strips include corresponding threaded holes such that the wear strips can be fastened in the groove by fasteners and fastener retention hardware. 1. A seal assembly for a turbine engine , comprising:a seal housing having a circumferential groove located along an edge of said seal housing;at least one segment strip, each having an upstream sealing surface, a downstream sealing surface, a right circumferential sealing surface and a left circumferential sealing surface,wherein said circumferential groove is configured to accept the geometry of the said at least one segment strip.2. The seal assembly according to claim 1 , wherein said at least one segment strip does not include any threaded holes.3. The seal assembly according to claim 2 , wherein said seal housing does not include any through holes.4. The seal assembly according to claim 1 , wherein said seal housing further comprises:a downstream surface;wherein said downstream sealing surface of said secured segment strips forms a substantially planar surface with said downstream surface of said seal housing and serves as a replaceable contact surface strip for said seal housing.5. The seal assembly according to claim 1 , wherein said upstream sealing surface of said secured segment strip forms an ...

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Номер записи: 31
23-01-2014 дата публикации

SEAL APPARATUS OF TURBINE AND THERMAL POWER SYSTEM

Номер: US20140020359A1
Автор: Hirano Toshio, MIMURA Yuki, ONODA Akihiro, Sato Iwataro, SHIBUKAWA Naoki, TSUKUDA Tomohiko, TSURUTA Kazutaka
Принадлежит: KABUSHIKI KAISHA TOSHIBA

A sealing device for a turbine has a sealing member provided in a gap between a rotor and a stator arranged to surround the rotor, and a fluid path provided within the stator, to introduce, into the stator, a cooling medium used to cool stationary blades extending radially inward from the stator, and to flow the cooling medium at least to an upstream side of the sealing member. 1. A sealing device for a turbine comprising:a sealing member provided in a gap between a rotor and a stator arranged to surround the rotor; anda fluid path provided within the stator, to introduce, into the stator, a cooling medium used to cool stationary blades extending radially inward from the stator, and to flow the cooling medium at least to an upstream side of the sealing member.2. The sealing device of claim 1 ,wherein the fluid path comprises:a first hole configured to take in the cooling medium used to cool the corresponding stationary blade; anda second hole configured to flow the cooling medium at least into the upstream side of the sealing member.3. The sealing device of claim 2 ,wherein the sealing member has a plurality of sealing fins arranged in an axial direction, andthe second hole is provided between the sealing fins in first and second stages on the upstream side of the sealing fins.4. The sealing device of claim 3 ,wherein a plurality of second holes are provided, and a part of the holes is provided between the sealing fins in stages following the second stage on the upstream side of the sealing fins.5. The sealing device of claim 3 ,wherein the second hole is provided between two sealing fins adjacent to each other, andan interval between these two sealing fins is narrowed around the second hole.6. The sealing device of claim 3 ,wherein the sealing fins are provided on an outer circumferential surface of the stationary blades, and on a surface of the stator facing rotor blades extending radially outward from an outer circumferential surface of the rotor.7. The sealing ...

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Номер записи: 32
23-01-2014 дата публикации

CLEARANCE CONTROL FOR GAS TURBINE ENGINE SEAL

Номер: US20140020390A1
Автор: Jalbert Peter L., Rhoden William E.
Принадлежит:

A gas turbine engine section has a rotor carrying a plurality of blades. The blades have airfoils which define a radially outer tip. A blade outer air seal is positioned radially outwardly of the tips of the blades. The blade outer air seal is provided by at least a plurality of circumferentially spaced segments, which slide circumferentially relative to each other to adjust an inner diameter of an inner surface of the blade outer air seal segments. An actuator actuates the blade outer air seal segments to slide towards each other to control a clearance between the inner periphery of the blade outer air seal segments and the radially outer tip of the blade airfoils. A gas turbine engine is also disclosed. 1. A gas turbine engine section comprising:a rotor carrying a plurality of blades, said blades each having a radially outer tip;a blade outer air seal positioned radially outwardly of said tips of said blades, said blade outer air seal being provided by at least a plurality of circumferentially spaced segments, said circumferentially spaced segments being operable to slide circumferentially relative to each other to adjust an inner diameter of an inner surface of said blade outer air seal segments; andan actuator for actuating said blade outer air seal segments to slide relative to each other to control a clearance between the inner periphery of said blade outer air seal segments and an outer periphery of the tips.2. The gas turbine engine section as set forth in claim 1 , wherein there are at least four of said blade outer air seal segments.3. The gas turbine engine section as set forth in claim 1 , wherein a sensor senses the amount of clearance between the inner periphery of the blade outer air seal segments and the outer periphery of a tip claim 1 , and communicates to a control for said actuator to control the clearance.4. The gas turbine engine section as set forth in claim 1 , wherein said blade outer air seal segments have a tongue at one circumferential ...

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Номер записи: 33
30-01-2014 дата публикации

FLUID TURBINE WITH VARIABLE PITCH SHROUD SEGMENTS

Номер: US20140030059A1
Автор: Dalsgaard Søren, Hjort Søren, JR. Walter M., Lovmand Bo, Presz, Rubak Rune, Werle Michael J.
Принадлежит: FLODESIGN WIND TURBINE CORP.

One or more variable pitch airfoils in fluid communication with a rotor of a fluid turbine can control the amount of energy directed to the rotor, and further control the amount of energy generated by the turbine. Varying the pitch of the airfoils may provide a means of controlling the power output of a fluid turbine without the need to control the pitch of the rotor blades, and may further provide a means of mitigating the effects of wind shear on the rotor. Variable pitch airfoils may also include a means of controlling the active power, reactive power and SCADA, of a group of fluid turbines. 1. A shrouded fluid turbine comprising:a rotor; anda ringed airfoil comprising a plurality of pivotable airfoil segments, each pivotable airfoil segment having a low pressure surface in fluid communication with the rotor.2. The shrouded fluid turbine of claim 1 , wherein each pivotable airfoil segment is rotatable about an axis to change a pitch of the pivotable airfoil segment.3. The shrouded fluid turbine of claim 1 , further comprising a pitch control mechanism that alters the pitch of at least a portion of the ringed airfoil.4. The shrouded fluid turbine of claim 3 , wherein the pitch control mechanism is configured to continuously change a pitch of at least a portion of the ringed airfoil while the shrouded fluid turbine is in use.5. The shrouded fluid turbine of claim 1 , wherein a pitch of each of the plurality of pivotable airfoil segments is individually adjustable.6. The shrouded fluid turbine of claim 1 , wherein the plurality of pivotable airfoil segments includes a plurality of outwardly curving airfoil segments claim 1 , and wherein the ringed airfoil further comprises a plurality of inwardly curving airfoil segments.7. The shrouded fluid turbine of claim 1 , wherein the ringed airfoil further comprises a frame claim 1 , and each of the plurality of pivotable airfoil segments is pivotably coupled to the frame.8. The shrouded fluid turbine of claim 1 , wherein ...

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Номер записи: 34
30-01-2014 дата публикации

ACTIVE CLEARANCE CONTROL MANIFOLD SYSTEM

Номер: US20140030066A1
Автор: Fusinato Daniel John, Glessner John Carl, Mirkovich Dan Nicholas, Proctor Robert, Schimmels Scott Alan
Принадлежит: GENERAL ELECTRIC COMPANY

Active clearance control systems for gas turbine engines are disclosed. An example active clearance control system may include a generally circumferentially mounted spray tube comprising a plurality of impingement holes arranged to impinge thermal control air on a clearance control component of a case; a rigid mounting assembly substantially rigidly coupling the spray tube to the case; and/or a sliding mounting assembly coupling the spray tube to the case while permitting limited relative movement between the spray tube and the case in a direction generally parallel with an engine axis. The sliding mount may be coupled to the case generally axially forward of the rigid mount. A ratio of the stand-off distance to the impingement hole diameter may be less than about 8. A ratio of the arc spacing to the impingement hole diameter may be less than about 15. 1. An active clearance control system for a gas turbine engine , the active clearance control system comprising:a generally circumferentially mounted spray tube comprising a plurality of impingement holes arranged to impinge thermal control air on a clearance control component of a case;wherein an individual impingement hole has an impingement hole diameter;wherein the individual impingement hole of the spray tube is spaced apart from the clearance control component by a stand-off distance; andwherein a ratio of the stand-off distance to the impingement hole diameter is less than about 8.2. The active clearance control system of claim 1 , wherein the ratio of the stand-off distance to the impingement hole diameter is less than about 5.3. The active clearance control system of claim 1 , wherein the ratio of the stand-off distance to the impingement hole diameter is less than about 3.4. The active clearance control system of claim 1 ,wherein the first individual impingement hole of the spray tube is spaced apart from a circumferentially adjacent second individual impingement hole by an arc spacing; andwherein a ratio of ...

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Номер записи: 35
30-01-2014 дата публикации

Blade outer air seal for a gas turbine engine

Номер: US20140030071A1
Автор: Fadi S. Maalouf, Nicholas R. Leslie
Принадлежит: Individual

A blade outer air seal (BOAS) for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a seal body having a radially inner face and a radially outer face that axially extend between a leading edge portion and a trailing edge portion. The BOAS includes a trough disposed on the radially inner face and an abradable seal received within the trough. The trough is open to expose a leading edge of the abradable seal to a core flow path of the gas turbine engine.

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Номер записи: 36
06-02-2014 дата публикации

Sealing structure in steam turbine

Номер: US20140037431A1
Автор: Itaru Murakami, Kenichi Okuno, Shinichiro Ohashi, Toshio Hirano, Yoriharu Murata, Yoshifumi Iwasaki
Принадлежит: Toshiba Corp

According to an embodiment, a rotor blade cover section is integrated with the rotor blades at leading ends thereof. A plurality of sealing fins is disposed at the rotor blade cover section, the sealing fins forming a predetermined clearance relative to an inner peripheral portion of the nozzle outer ring. An annular solid particle trapping space is disposed at the inner peripheral portion of the nozzle outer ring, the solid particle trapping space communicating with an inlet of a steam leak and trapping solid particles that flow in with steam. In the sealing structure, the nozzle outer ring has a through hole through which the solid particles are to be discharged from the solid particle trapping space toward a downstream stage of the steam turbine.

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Номер записи: 37
13-02-2014 дата публикации

Blade outer air seal having anti-rotation feature

Номер: US20140044528A1
Автор: Brian Ellis Clouse
Принадлежит: United Technologies Corp

A blade outer air seal (BOAS) for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a seal body having a radially inner face and a radially outer face that axially extend between a leading edge portion and a trailing edge portion. At least one of the leading edge portion and the trailing edge portion includes a solid wall and a perforated wall. At least a portion of the perforated wall is spaced from the solid wall such that a passage extends between the solid wall and the perforated wall.

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Номер записи: 38
20-03-2014 дата публикации

REPLACEABLE SEALS FOR TURBINE ENGINE COMPONENTS AND METHODS FOR INSTALLING THE SAME

Номер: US20140079538A1
Автор: Ekanayake Sanji, Goetze Raymond Henry, King Peter Samuel, Scipio Alston Ilford, Warwick Douglas Corbin
Принадлежит:

An assembly and method are provided for sealing a compressor in a gas turbine engine. The method comprises forming an annular groove in a compressor casing such that the groove extends circumferentially about a rotor that is housed within the casing. The compressor casing is then coupled to the rotor such that the compressor casing extends circumferentially about the rotor. A plurality of arcuate seal segments are then inserted into the annular groove without removing the rotor from the compressor casing such that the plurality of seal segments extend circumferentially about the rotor to facilitate sealing a gap that is defined between the rotor and the compressor casing. 1. A seal assembly for use in a turbine engine including a compressor casing that at least partially circumscribes a rotor , said seal assembly comprising a plurality of arcuate seal segments configured to be at least partially inserted into a groove defined in the compressor casing to substantially seal a gap defined between the casing and the rotor , wherein each of said plurality of seal segments comprises a radially inner projection , a radially outer projection , and a neck portion extending therebetween.2. An assembly in accordance with claim 1 , wherein said groove includes at least one hook portion configured to retain each of said plurality of arcuate seal segments within said groove.3. An assembly in accordance with claim 1 , wherein said seal assembly further comprises an anti-rotation device coupled to the compressor casing claim 1 , said anti-rotation device configured to prevent said plurality of arcuate seal segments from shifting circumferentially within said groove.4. An assembly in accordance with claim 1 , wherein each of said plurality of seal segments further comprises at least one cutout sized to receive at least a portion of a biasing mechanism therein.5. An assembly in accordance with claim 4 , wherein said biasing mechanism comprises one of a coil spring and a wave spring.6 ...

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Номер записи: 39
06-01-2022 дата публикации

SEAL ASSEMBLY WITH REDUCED PRESSURE LOAD ARRANGEMENT

Номер: US20220003126A1
Автор: Arbona Jaime A., Bleil Justin K., Keene Andrew D., Mahoney Danielle, Roy Thill Carson A.
Принадлежит:

A seal assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a seal arc segment including a sealing portion, and a first rail and a second rail opposed to the first rail. The sealing portion extends in a circumferential direction between opposed mate faces and extends in an axial direction between a leading edge and a trailing edge. Each of the first and second rails extend outwardly in a radial direction from the sealing portion to respective first and second edge faces, and the sealing portion has a sealing face dimensioned to bound a gas path and includes a backside face opposed to the sealing face. The first and second rails include respective first and seconds pairs of hooks dimensioned to mount the seal arc segment to an engine static structure in an installed position. The seal arc segment is radially opposed to the sealing face between the first and second edge faces establishing a first region. The seal arc segment is radially opposed to the sealing face between the leading and trailing edges establishing a second region. A method of sealing for a gas turbine engine is also disclosed. 1. A seal assembly for a gas turbine engine comprising:a seal arc segment including a sealing portion, a first rail and a second rail opposed to the first rail, the sealing portion extending in a circumferential direction between opposed mate faces and extending in an axial direction between a leading edge and a trailing edge, each of the first and second rails extending outwardly in a radial direction from the sealing portion to respective first and second edge faces, and the sealing portion including a sealing face dimensioned to bound a gas path and including a backside face opposed to the sealing face;wherein the first and second rails includes respective first and seconds pairs of hooks dimensioned to mount the seal arc segment to an engine static structure in an installed position; andwherein the seal arc ...

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Номер записи: 40
06-01-2022 дата публикации

SEALING BETWEEN A ROTOR DISC AND A STATOR OF A TURBOMACHINE

Номер: US20220003127A1
Автор: AMORIM Joao Antonio, Rotenberg Maxime Aurélien, TRAPPIER Nicolas Xavier
Принадлежит:

Assembly including a rotor disc, an adjacent stator and a plurality of sealing elements secured to the rotor disc, the stator including an inner platform and a root bearing at least one abradable element configured to cooperate with the sealing elements, the sealing elements being placed in an enclosure formed by the abradable element, the enclosure being open to the inside and delimited axially by an upstream abradable edge and a downstream abradable edge, the enclosure being delimited radially by an outer abradable edge, at least one of the sealing elements including a first lip configured to cooperate with the upstream abradable edge or the downstream abradable edge, and a second, separate lip configured to cooperate with the outer abradable edge. 1. An assembly for a turbomachine comprising a first mobile wheel extending around an axis and an adjacent bladed turbine stator , said bladed turbine stator being coaxial with said axis and axially offset from said first mobile wheel , said assembly comprising a plurality of sealing elements , each sealing element being secured to said first mobile wheel and projecting radially from said first mobile wheel , said bladed turbine stator comprising an inner platform intended to delimit a gas flow channel in the turbomachine and a root extending radially below the inner platform , said root bearing at a radially inner end at least one abradable element configured to cooperate with the sealing elements , characterised in that the sealing elements are placed in an enclosure formed by said at least one abradable element , said enclosure being open inwards and delimited axially by an upstream abradable edge and a downstream abradable edge , said enclosure being radially delimited by an outer abradable edge , and in that at least one of the sealing elements comprises a first lip configured to cooperate with the upstream abradable edge or the downstream abradable edge , and a second lip separate from the first lip and configured ...

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Номер записи: 41
03-01-2019 дата публикации

Coating for Lining a Compressor Case

Номер: US20190001372A1
Автор: CRAWFORD Graeme Edward, KAATZ Chad Christian, ROBERGE Douglas Michael, THOMAS Wayne Wilfred, WITKOWSKA Izabela
Принадлежит:

A lining for a compressor case is provided. The lining comprises a layer of primer applied to an interior surface of the compressor case, and at least one layer of a seal compound applied to the layer of primer. The primer has been found to significantly improve the bond strength between the base metal and the coating. Also, the primer provides a waterproof coating at the bond interface, which prevents any moisture from creeping under the plastic and causing material break-out. The waterproof coating also provides an increased corrosion resistance to the base metal and stator vanes where the plastic lining has been applied. A method of lining a compressor case is also provided, comprising applying a first layer comprising a primer to an interior surface of the compressor case, and applying at least one second layer over the first layer of primer, the at least one second layer comprising a seal compound to line the compressor case. 1. A method of lining a compressor case , the method comprising:applying a first layer comprising a primer to an interior surface of the compressor case; andapplying at least one second layer over the first layer of primer, the at least one second layer comprising a seal compound to line the compressor case.2. The method of claim 1 , wherein applying the first layer comprises:grit blasting an area of the interior surface of the compressor case;spray washing the area of the interior surface;drying the compressor case;applying the primer; andcuring the primer.3. The method of claim 1 , wherein the primer is a silane-based primer.4. The method of claim 2 , wherein the primer is applied using a brush claim 2 , spray claim 2 , or dip application process.5. The method of claim 2 , wherein the primer is cured using a moisture curing process claim 2 , or an oven curing process.6. The method of claim 1 , further comprising installing the compressor case in a casting machine prior to applying the at least one second layer to line the compressor case ...

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Номер записи: 42
03-01-2019 дата публикации

FASTENER REMOVAL TOOLS AND METHODS

Номер: US20190001472A1
Автор: PERKOWSKI Szymon, WOJCIECHOWSKI Piotr Hubert
Принадлежит:

A fastener removal tool is provided. The fastener removal tool includes a body having a cylinder and a puller coupled to the body. The puller includes an arm for engaging an installed fastener and a piston inserted into the cylinder of the body such that, when the cylinder is pressurized, the piston is displaced within the cylinder to displace the arm relative to the body to cause removal of the fastener. 1. A fastener removal tool comprising:a body comprising a cylinder; and an arm for engaging an installed fastener; and', 'a piston inserted into said cylinder of said body such that, when said cylinder is pressurized, said piston is displaced within said cylinder to displace said arm relative to said body to cause removal of the fastener., 'a puller coupled to said body, said puller comprises2. A fastener removal tool in accordance with claim 1 , wherein said body comprises a pair of cylinders claim 1 , said puller comprises a pair of pistons each inserted into a respective one of said cylinders.3. A fastener removal tool in accordance with claim 1 , further comprising a return spring biasing said puller towards said body.4. A fastener removal tool in accordance with claim 3 , wherein said body comprises a sleeve claim 3 , said return spring inserted into said sleeve of said body.5. A fastener removal tool in accordance with claim 1 , wherein said arm defines an open-ended slot for slidably engaging the fastener.6. A fastener removal tool in accordance with claim 1 , wherein said body is generally U-shaped and comprises a first leg member claim 1 , a second leg member claim 1 , and a bridge member coupling said first leg member to said second leg member such that a passage is defined between said first and second leg members.7. A fastener removal tool in accordance with claim 1 , further comprising a shield coupled to said body and at least partially surrounding said puller.8. A fastener removal tool in accordance with claim 1 , wherein said tool is sized for ...

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Номер записи: 43
05-01-2017 дата публикации

TURBINE SHROUD WITH MOVABLE ATTACHMENT FEATURES

Номер: US20170002676A1
Автор: Cole Zacheriah A., Freeman Ted J., Shinavski Robert J., Sippel Aaron D., Thomas David J.
Принадлежит:

A turbine shroud for positioning radially outside of blades of a turbine rotor includes a carrier, a blade track, and a track attachment system. The blade track is moved radially outwardly into a cavity of the carrier, and the track attachment system is adjusted to block radially inward movement of the blade track out of the cavity. 1. A segmented turbine shroud that extends around a central axis , the segmented turbine shroud comprisinga carrier segment that extends partway around the central axis and that forms a radially inwardly-opening cavity,a blade track segment comprising ceramic-containing materials, the blade track segment formed to include a runner that extends partway around the central axis and a positioner attachment post that extends radially outward from the runner into the radially inwardly-opening cavity of the carrier segment, the positioner attachment post formed to include a track-positioning surface that extends both radially and axially, anda track attachment system adapted to couple the blade track segment to the carrier segment, and the track attachment system including a positioner coupled to the carrier segment to move axially from a disengaged position out of contact with the positioner attachment post to an engaged position contacting the positioner attachment post to engage the track-positioning surface of the positioner attachment post with a position-setting surface that extends both radially and axially at an angle corresponding to that of the track-positioning surface.2. The segmented turbine shroud of claim 1 , wherein the position-setting surface is formed by the positioner.3. The segmented turbine shroud of claim 2 , wherein the positioner is formed to include positioner threads that engage corresponding threads formed in the carrier segment.4. The segmented turbine shroud of claim 1 , wherein the position-setting surface is formed by the carrier segment.5. The segmented turbine shroud of claim 1 , wherein the blade track segment ...

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Номер записи: 44
05-01-2017 дата публикации

Break-in system for gapping and leakage control

Номер: US20170002677A1
Автор: Paul M. Lutjen, Richard K. Hayford
Принадлежит: United Technologies Corp

A blade outer air seal for use in a gas turbine engine having an axis of rotation includes a main body having a mating face configured to face, be positioned radially outward from, and be positioned adjacent to a rotor blade of the gas turbine engine. The blade outer air seal also includes an axial member extending aft from the main body, having a first radial face configured to face a second radial face of an outer diameter platform of a stator of the gas turbine engine, and having a first abradable material coupled to the first radial face.

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Номер записи: 45
07-01-2016 дата публикации

GAS TURBINE ENGINE SEAL ASSEMBLY

Номер: US20160003080A1
Автор: McGARRAH CRAIG R.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A seal assembly is positioned within a cavity that extends circumferentially about an axial centerline of a gas turbine engine. The cavity includes a cavity wall. The seal assembly includes a seal and a seal protector. The seal extends circumferentially within the cavity. The seal protector extends circumferentially within the cavity. The seal protector is positioned between the seal and the cavity wall. The seal protector includes a locating feature that is operative to contact the seal to aid in axially positioning the seal protector relative to the seal. 1. A seal assembly positioned within a cavity that extends circumferentially about an axial centerline of a gas turbine engine , which cavity includes a cavity wall , which seal assembly comprises:a seal that extends circumferentially within the cavity;a seal protector that extends circumferentially within the cavity, which seal protector is positioned between the seal and the cavity wall, and which seal protector includes a radially-extending locating feature that is operative to contact the seal to aid in axially positioning the seal protector relative to the seal.2. The seal assembly of claim 1 , wherein the seal is at least substantially annularly-shaped claim 1 , and wherein the seal protector is at least substantially annularly-shaped.3. The seal assembly of claim 2 , wherein the seal forms a split ring claim 2 , and wherein the seal protector forms a split ring.4. The seal assembly of claim 1 , wherein the cavity includes a forward cavity wall claim 1 , an aft cavity wall claim 1 , a radially inner cavity wall claim 1 , and a radially outer cavity wall claim 1 , and wherein the seal protector is positioned between a positioning contact surface of the seal and the radially inner cavity wall.5. The seal assembly of claim 4 , wherein the locating feature is positioned between a forward sealing contact surface of the seal and the forward cavity wall.6. The seal assembly of claim 5 , wherein the seal protector ...

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Номер записи: 46
07-01-2016 дата публикации

FLEXIBLE FINGER SEAL FOR SEALING A GAP BETWEEN TURBINE ENGINE COMPONENTS

Номер: US20160003081A1
Автор: Broomer Mark, Cherolis Anthony P., Choung Conway, Davis Timothy M., Hazzard Robert L., Keene Russell E., McGARRAH CRAIG R., Roy Thill Carson A.
Принадлежит:

An assembly for a turbine engine includes a turbine engine first component, a turbine engine second component and a flexible seal that is attached to the first component. The flexible seal at least partially seals a gap between the first component and the second component. The flexible seal includes a mount and a finger seal that sealingly engages the second component. The mount includes a boss that sealingly engages the first component. 1. An assembly for a turbine engine , comprising:a turbine engine first component;a turbine engine second component; anda flexible seal attached to the first component, the flexible seal at least partially sealing a gap between the first component and the second component;wherein the flexible seal includes a mount and a finger seal that sealingly engages the second component, and the mount includes a boss that sealingly engages the first component.2. The assembly of claim 1 , further comprising a fastener that attaches the mount to the first component claim 1 , and extends through the boss.3. The assembly of claim 2 , whereinthe boss comprises a first boss, and the mount further includes a base and a second boss;the first boss and the second boss are arranged on opposing sides of the base; andthe fastener further extends through the base and the second boss.4. The assembly of claim 2 , wherein the mount further includes a base claim 2 , and the boss comprises a washer that is bonded to the base.5. The assembly of claim 2 , further comprising:a washer sealingly engaged between a flange and a first shoulder;wherein the first component includes the flange, and the fastener includes the first shoulder and a second shoulder that sealingly engages the mount; andwherein the fastener attaches the mount to the flange, and extends through mount, the flange and the washer between the first shoulder and the second shoulder.6. The assembly of claim 2 , further comprising:a second fastener that attaches the mount to the first component;wherein ...

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Номер записи: 47
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: 48
07-01-2016 дата публикации

ABRADABLE SEAL INCLUDING AN ABRADABILITY CHARACTERISTIC THAT VARIES BY LOCALITY

Номер: US20160003083A1
Автор: "ONeill Christopher F.", Boyer Jesse R., Delisle Robert P., Faughnan, JR. Paul R.
Принадлежит:

An abradable seal for a gas turbine engine includes a seal body that has a seal side and a non-seal side. The seal body includes an abradability characteristic that varies by locality. 1. An abradable seal for a gas turbine engine , comprising:a seal body having a seal side and a non-seal side, the seal body including an abradability characteristic that varies by locality.2. The abradable seal as recited in claim 1 , wherein the abradability characteristic is selected from the group consisting of a graded composition claim 1 , a graded porosity claim 1 , a non-uniform geometric cell structure and combinations thereof.3. The abradable seal as recited in claim 1 , wherein the abradability characteristic is a graded composition.4. The abradable seal as recited in claim 1 , wherein the abradability characteristic is a graded porosity.5. The abradable seal as recited in claim 1 , wherein the abradability characteristic is a non-uniform geometric cell structure.6. The abradable seal as recited in claim 1 , wherein the abradability characteristic is a graded composition that varies in an amount of nickel between the seal side and the non-seal side.7. The abradable seal as recited in claim 1 , wherein the abradability characteristic is a graded porosity that varies in a percentage of porosity from a relatively low porosity at the non-seal side to a relatively high porosity at the seal side.8. The abradable seal as recited in claim 7 , wherein the relatively low porosity is 0-5% and the relatively high porosity is 40-60%.9. The abradable seal as recited in claim 1 , wherein the abradability characteristic is a non-uniform geometric cell structure including a cell center-to-center dimension that varies by locality.10. The abradable seal as recited in claim 1 , wherein the seal body includes a plurality of cells defined between cell walls claim 1 , the cell walls being made of a first material claim 1 , and cell cores in the cells claim 1 , the cell cores being made of a ...

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Номер записи: 49
07-01-2016 дата публикации

THERMALLY CONFORMABLE LINER FOR REDUCING SYSTEM LEVEL FAN BLADE OUT LOADS

Номер: US20160003084A1
Автор: Husband Jason, Lancaster Drew W.
Принадлежит:

A fan case for a gas turbine engine includes a fan case surrounding a fan with fan blades. A liner is disposed between the fan case and the fan and is spaced a radial distance from the fan case. A torque stop is arranged between the fan case and the liner. A method for reducing fan case liner loads is also disclosed. 1. A fan section for a gas turbine engine , comprising:a fan case surrounding a fan with fan blades;a liner disposed between the fan case and the fan, spaced a radial distance from the fan case; andat least one torque stop arranged between the fan case and the liner.2. The fan section of claim 1 , wherein the ratio of the radial distance to a radius of the fan blades is 0.25:40.3. The fan section of claim 1 , wherein at least one of a radially inner surface of the fan case and a radially outer surface of the liner includes a low-friction coating.4. The fan section of claim 3 , wherein the low-friction coating is a Teflon® spray.5. The fan section of claim 1 , wherein the at least one torque stop is frangible.6. The fan section of claim 1 , wherein the at least one torque stop prevents rotation of the liner relative to the fan case.7. The fan section of claim 1 , wherein at least a portion of the liner is bonded to the fan case with an adhesive.8. The fan section of claim 7 , wherein the at least one torque stop and the adhesive can withstand a tangential load given by the equation (1+S)*T/R/Nwhere S is a safety factor claim 7 , Tis a rub torque generated by a 2.5 lb (1.3 kg) bird strike claim 7 , Ris a radius of the fan case claim 7 , and Nis the number of torque stops.9. The fan section of claim 8 , wherein S is 0.35.10. The fan section of claim 1 , wherein the liner includes one or more rails which contact the fan case.11. The fan section of claim 1 , wherein the liner includes an abradable rubstrip adjacent to tips of the fan blades.12. The fan section of claim 11 , wherein the abradable rubstrip is arranged radially inward from the at least one ...

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Номер записи: 50
07-01-2016 дата публикации

GAS TURBINE ENGINE SPRING MOUNTED MANIFOLD

Номер: US20160003086A1
Автор: Day Christophe Jude, Laubenthal Jeffrey Scott, MAZUK James L., Poranski Christopher Francis, SMITH Vernon Leroy, Thompson James Edward
Принадлежит:

An arcuate panel for a thermal control assembly includes an arcuate panel base, at least one axially extending panel header sealingly attached to a radially outwardly facing surface of the panel base, a plenum therebetween, one or more spray tubes or channels depending radially inwardly from the panel base and in fluid communication with the plenum, and radially outwardly biasing spring means mounted on or attached to the radially inwardly facing surface of the panel base. The panel may include at least one set of clockwise and counter-clockwise hinge wings attached to clockwise and counter-clockwise ends of the panel base, axial positioning means, and circumferential positioning means. The axial positioning means may include a circular row of spring clips mounted on the panel base. A hoop of arcuate panels pivotably attached by hinges may encircle a portion of a casing compressing leaf springs against rings of the casing. 1525152. An arcuate panel () of a spring mounted air distribution manifold () , the arcuate panel () comprising:{'b': '58', 'an arcuate panel base (),'}{'b': 54', '62', '58, 'at least one axially extending panel header () sealingly attached to a radially outwardly facing surface () of the panel base (),'}{'b': 56', '54', '58, 'a plenum () between the panel header () and the panel base (),'}{'b': 60', '182', '58, 'one or more spray tubes or channels () depending radially inwardly from and mounted on or attached to a radially inwardly facing surface () of the panel base (),'}{'b': 60', '56, 'the spray tubes or channels () in fluid communication with the plenum (), and'}{'b': 110', '182', '58, 'radially outwardly biasing springs () mounted on or attached to the radially inwardly facing surface () of the panel base ().'}252110111. The arcuate panel () as claimed in claim 1 , further comprising the radially outwardly biasing springs () including radially bent leaf springs ().35212212413213458. The arcuate panel () as claimed in claim 1 , further ...

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Номер записи: 51
07-01-2016 дата публикации

EDGE TREATMENT FOR GAS TURBINE ENGINE COMPONENT

Номер: US20160003087A1
Автор: Farris John R., Lutjen Paul M., Strock Christopher W.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A gas turbine engine component according to an exemplary aspect of the present disclosure includes, among other things, a body having a first outer face meeting a second outer face at an intersection, the body having a plurality of apertures extending from an opening in the first outer face to an opening on the second outer face; and a coating filling at least a portion of the plurality of apertures. 1. A gas turbine engine component , comprising:a body having a first outer face meeting a second outer face at an intersection, the body having a plurality of apertures extending from an opening in the first outer face to an opening on the second outer face; anda coating filling at least a portion of the plurality of apertures.2. The gas turbine engine component of claim 1 , wherein the plurality of apertures extend across the intersection.3. The gas turbine engine component of claim 1 , wherein the plurality of apertures comprises more than six apertures.4. The gas turbine engine component of claim 1 , wherein at least one aperture of the plurality of apertures is completely filled with the coating.5. The gas turbine engine component of claim 1 , wherein at least one aperture of the plurality of apertures is defined by a pair of opposing walls and a curved floor.6. The gas turbine engine component of wherein at least one aperture of the plurality of apertures is defined by a pair of opposing walls and at least two planar floors.7. The gas turbine engine component of claim 1 , wherein an axial distance between axially adjacent openings of the plurality of apertures is about the same as an axial width of each of the adjacent openings.8. The gas turbine engine component of claim 1 , wherein the coating comprises Yttria-Stabilized Zirconia.9. The gas turbine engine component of claim 1 , wherein the body comprises a seal body of a blade outer air seal segment claim 1 , the first outer face comprises a first mate face of the seal body claim 1 , and the second outer face ...

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Номер записи: 52
07-01-2016 дата публикации

INTEGRAL SEGMENTED CMC SHROUD HANGER AND RETAINER SYSTEM

Номер: US20160003103A1
Автор: BALDIGA Jonathan David, Shapiro Jason David
Принадлежит: GENERAL ELECTRIC COMPANY

A shroud hanger with integral retainer assembly comprises a ceramic matrix composite shroud hanger a first wall and a second wall, the hanger having a support wall extending between the first and second walls, the support wall having a shoulder near circumferential ends, a retainer depending from the support wall having a first lower leg and a second lower leg extending in the circumferential direction, a first shroud supported by the first lower leg and a second shroud supported by the second lower leg. 1. A shroud hanger with integral retainer assembly , the shroud hanger comprising:a ceramic matrix composite shroud hanger comprising a first wall and a second wall;a support wall extending between said first wall and said second wall, said support wall comprising a shoulder near each of circumferential ends of the shroud hanger;a retainer depending from said support wall having and comprising a first lower leg and a second lower leg, wherein each of the first lower leg and the second lower leg extend towards a corresponding one of said circumferential ends;a first shroud supported by said first lower leg; anda second shroud supported by said second lower leg.2. The shroud hanger of claim 1 , wherein ends of said shrouds are offset from said circumferential ends of said shroud hanger.3. The shroud hanger with integral retainer assembly of claim 1 , wherein said shroud retainer is substantially T-shaped.4. The shroud hanger of claim 1 , wherein one of said first wall and said second wall is a forward wall and the other of said first wall and said second wall is an aft wall.5. The shroud hanger with of claim 1 , wherein said retainer further comprises a first depending member.6. The shroud hanger of claim 5 , wherein said retainer further comprises a second depending member spaced from said first depending member.7. The shroud hanger of claim 1 , wherein said support wall defines a first arm and a second arm extending circumferentially.8. The shroud hanger of claim 1 ...

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Номер записи: 53
01-01-2015 дата публикации

BLADE TRACK ASSEMBLY WITH TURBINE TIP CLEARANCE CONTROL

Номер: US20150003958A1
Автор: Lamusga Joseph P., Sippel Aaron D., Uskert Richard C.
Принадлежит:

An apparatus and method for controlling turbine blade tip clearance is disclosed herein. A blade track assembly can include a blade track carrier having a plurality of slots and rails defined by paths that vary in radial position as a function of circumferential location. An expansion ring can be operably coupled with the slots of the blade track carrier. A plurality of blade track segments can be operably coupled with the expansion ring and engageable with the rails of the blade track carrier such that expansion and contraction of the expansion ring causes radially outward movement and radially inward movement, respectively, of the blade track segments. 1. A turbine blade track assembly comprisinga blade track carrier having a top wall and a pair of sidewalls extending radially inward therefrom,a plurality of slots formed in each sidewall of the blade track carrier, wherein each slot includes a first end located at a first radial position relative to an axis of rotation and a second end located at a second radial position outward of the first radial position,a plurality of rails connected to each sidewall of the blade track carrier, each rail having a bearing surface extending between a first end and a second end, wherein the second end of the rail is positioned radially outward of the first end of the rail,an expansion ring operably coupled with the slots of the blade track carrier, anda plurality of blade track segments operably coupled with the expansion ring and engageable with the rails of the blade track carrier.2. The turbine blade track assembly of claim 1 , wherein the expansion ring moves in a first circumferential direction during expansion and in an opposite circumferential direction during contraction.3. The turbine blade track assembly of claim 2 , wherein each blade track moves in a circumferential direction in response to circumferential movement of the expansion ring.4. The turbine blade track assembly of claim 2 , wherein each blade track segment ...

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Номер записи: 54
01-01-2015 дата публикации

TURBINE SEAL ASSEMBLY AND TURBINE APPARATUS COMPRISING THE TURBINE SEAL ASSEMBLY

Номер: US20150003972A1
Автор: LIM Chan-sun
Принадлежит: SAMSUNG TECHWIN CO., LTD.

According to an aspect of an exemplary embodiment, there is provided a turbine seal assembly comprising: a seal installation groove formed inside a casing; at least one seal member that has at least one tip portion formed in a blade direction and is installed in the seal installation groove; at least one elastic member for elastically connecting the casing to the seal member, wherein an inlet for an inflow of compressed gas is formed in the seal installation groove, and wherein a first space which the compressed gas enters is formed in the seal member, and at least one flow pathway which connects the first space to a space between a blade and the seal member is formed in the seal member. 1. A turbine seal assembly comprising:a seal installation groove formed inside a casing;at least one seal member that has at least one tip portion formed in a blade direction and is installed in the seal installation groove; andat least one elastic member for elastically connecting the casing to the seal member,wherein an inlet for an inflow of compressed gas is formed in the seal installation groove, andwherein a first space which the compressed gas enters is formed in the seal member, and at least one flow pathway which connects the first space to a space between a blade and the seal member is formed in the seal member.2. The turbine seal assembly of claim 1 , wherein the elastic member is a plate spring.3. The turbine seal assembly of claim 1 , wherein a thermal expansivity of a material of the elastic member is different from a thermal expansivity of a material of the casing.4. The turbine seal assembly of claim 1 , wherein the first space is closed in a front direction of the blade.5. The turbine seal assembly of claim 1 , wherein the inlet is connected to a compressed gas tube connected to a compressor.6. A turbine apparatus comprising:a casing in which a seal installation groove is formed;a rotor that is installed inside the casing and has a plurality of blades;at least one ...

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Номер записи: 55
02-01-2020 дата публикации

GAS TURBINE ENGINE COMPONENT

Номер: US20200003066A1
Автор: Clark Thomas E., Whitney Daniel J.
Принадлежит:

A blade outer air seal includes a base portion that extends between a leading edge and a trailing edge. A forward wall and an aft wall extend radially outward from the base portion to a radially outer portion. The radially outer portion is spaced from the base portion. 1. A blade outer air seal comprising:a base portion extending between a leading edge and a trailing edge; anda forward wall and an aft wall extending radially outward from the base portion to a radially outer portion, wherein the radially outer portion is spaced from the base portion.2. The blade outer air seal of claim 1 , wherein the radially outer portion is spaced inward from circumferential edges of the base portion.3. The blade outer air seal of claim 1 , wherein a radially outer edge of the forward wall is spaced a first distance from the base portion and a radially outer edge of the aft wall is spaced a second distance from the base portion and the second distance is greater than the first distance.4. The blade outer air seal of claim 1 , wherein the blade outer air seal is made entirely from a composite matrix composite.5. The blade outer air seal of claim 1 , wherein the radially outer portion is centered between circumferential edges of the base portion.6. The blade outer air seal of claim 1 , wherein the radially outer portion is closer to a first circumferential edge of the base portion than a second circumferential edge.7. The blade outer air seal of claim 1 , wherein the forward wall is spaced a first distance from the leading edge and the after wall is spaced a second distance from the trailing edge and the first distance is greater than the second distance.8. A seal assembly comprising: a base portion extending between a leading edge and a trailing edge; and', 'a forward wall and an aft wall extending radially outward from the base portion to a radially outer portion, wherein the radially outer portion is spaced from the base portion; and, 'at least one blade outer air seal ...

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Номер записи: 56
02-01-2020 дата публикации

SEGMENTED RING FOR INSTALLATION IN A TURBOMACHINE

Номер: US20200003067A1
Автор: Feldmann Manfred
Принадлежит:

A segmented ring for installation in a turbomachine, circumferentially divided into segments, considered with respect to a ring axis of the segmented ring, wherein the segmented ring is adapted for installation from radially inside; i.e., the segments can be assembled radially outwardly to form the segmented ring. At least two immediately circumferentially adjacent segments meet in a joint at which a sealing insert is provided. A pocket which is open toward the joint is formed in each of the at least two immediately circumferentially adjacent segments; i.e., two circumferentially mutually facing pockets are provided at the joint. The sealing insert is disposed in the two mutually facing pockets and axially retained therein and extends circumferentially across the joint, and a pocket of the two mutually facing pockets is additionally also open in a radial direction such that this segment can be radially slid onto the sealing insert. 115-. (canceled)16. A segmented ring for installation in a turbomachine , the segmented ring comprising:segments circumferentially dividing the ring, considered with respect to a ring axis of the segmented ring, the segmented ring being adapted for installation from radially inside so that the segments are assemblable radially outwardly to form the segmented ring;at least two immediately circumferentially adjacent segments of the segments meeting in a joint;a sealing insert provided at the joint, a pocket open toward the joint being formed in each of the at least two immediately circumferentially adjacent segments so that two circumferentially mutually facing pockets are provided at the joint,the sealing insert being disposed in the two mutually facing pockets and axially retained in the two mutually facing pockets and extending circumferentially across the joint;one pocket of the two mutually facing pockets being additionally also open in a radial direction such that the segment having the one pocket is radially slidable onto the sealing ...

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Номер записи: 57
02-01-2020 дата публикации

Jacket ring assembly for a turbomachine

Номер: US20200003076A1
Автор: Feldmann Manfred
Принадлежит:

A jacket ring assembly for a turbomachine, the jacket ring assembly including a casing part, a jacket ring segment which is adapted to radially outwardly surround a rotor blade ring and to this end is disposed radially inwardly of the casing part, as considered with respect to a longitudinal axis of the turbomachine, and a segmented ring which is circumferentially divided into segments and by which the jacket ring segment is mounted to the casing part, the segmented ring being axially form-fittingly disposed on a form-fitting element of the casing part, for which purpose each of the respective segments of the segmented ring is radially outwardly assembled with the form-fitting element, and the segmented ring forming a supporting seat on which the jacket ring segment is seated and radially inwardly supported with an axially forward end. 115-. (canceled)16. A jacket ring assembly for a turbomachine , comprising:a casing part;a jacket ring segment adapted to radially outwardly surround a rotor blade ring and disposed radially inwardly of the casing part, as considered with respect to a longitudinal axis of the turbomachine; anda segmented ring circumferentially divided into segments, the jacket ring segment mounted to the casing part by the segmented ring, the segmented ring being axially form-fittingly disposed on a form-fitting element of the casing part, so that each of the respective segments of the segmented ring is radially outwardly assemblable with the form-fitting element, the segmented ring forming a supporting seat, the jacket ring segment being seated on the supporting seat and radially inwardly supported with an axially forward end.17. The jacket ring assembly as recited in wherein the form-fitting element of the casing part is a radially inwardly projecting web disposed in a radially outwardly open receptacle of the segmented ring.18. The jacket ring assembly as recited in wherein when viewed in an axial section claim 17 , the web extends at an angle of ...

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Номер записи: 58
02-01-2020 дата публикации

GAS TURBINE ENGINE COMPONENT

Номер: US20200003077A1
Автор: Blaney Ken F., Clark Thomas E.
Принадлежит:

An attachment body for a blade outer air seal includes a leading edge connected to a trialing edge by a radially inner wall and a radially outer wall. At least one forward hook extends from the radially outer wall. At least one aft hook extends from the radially outer wall. At least one post extends from the radially outer surface and has a blade outer air seal (BOAS) guide surface. 1. An attachment body for a blade outer air seal comprising:a leading edge connected to a trialing edge by a radially inner wall and a radially outer wall;at least one forward hook extending from the radially outer wall;at least one aft hook extending from the radially outer wall; andat least one post extending from the radially outer surface having a blade outer air seal (BOAS) guide surface.2. The attachment body of claim 1 , wherein the radially outer surface includes at least one BOAS attachment surface.3. The attachment body of claim 2 , wherein the at least one BOAS attachment surface includes a pair BOAS attachment surfaces each located adjacent an opposing circumferential side of the attachment body.4. The attachment body of claim 3 , wherein each of the pair of BOAS attachment surfaces define an arced surface.5. The attachment body of claim 4 , wherein the arced surface includes a varying radius of curvature in an axial direction.6. The attachment body of claim 4 , wherein the arced surface includes a constant radius of curvature in the axial direction.7. The attachment body of claim 4 , wherein the at least one post includes a pair of posts each having the BOAS guide surface facing a circumferential side of the attachment body.8. The attachment body of claim 4 , wherein the at least one aft hook includes a pair of aft hooks each including an anti-rotation tab.9. The attachment body of claim 8 , wherein the at least one post includes a pair of posts each having the BOAS guide surface facing a circumferential side of the attachment body.10. A seal assembly comprising: a base ...

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Номер записи: 59
02-01-2020 дата публикации

Chemistry based methods of manufacture for maxmet composite powders

Номер: US20200003125A1
Автор: Christopher W. STROCK, Shahram Amini, Weina Li
Принадлежит: United Technologies Corp

A method of manufacturing a gas turbine engine air seal comprising forming at least one MAX phase particle. The method includes coating the at least one MAX phase particle with a metallic shell. The method includes applying the at least one MAX phase metallic coated particle to a surface of a substrate of the air seal to form an abradable layer of a MAXMET composite abradable material from the at least on MAX phase metallic coated particle.

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Номер записи: 60
02-01-2020 дата публикации

Fan blades with abrasive tips

Номер: US20200003225A1
Автор: Changsheng Guo, Christopher W. STROCK
Принадлежит: United Technologies Corp

A fan blade for a gas turbine engine is disclosed. The disclosed fan blade includes an airfoil having a leading edge, a trailing ling edge, a convex side, a concave side and a distal tip. The leading edge, trailing edge, convex side and concave side of the airfoil is at least partially coated with an erosion resistant coating. The distal tip of the airfoil is coated with a bonded abrasive coating. The bonded abrasive coating engages the abradable coating disposed on the fan liner and, because of its low thermal conductivity, reduces heat transfer to the distal tip of the fan blade. The reduction in heat transfer to the distal tip of the fan blade preserves the integrity of erosion resistant coatings that may be applied to the body or the airfoil of the fan blade.

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Номер записи: 61
07-01-2021 дата публикации

TURBINE TIP SHROUD ASSEMBLY WITH PLURAL SHROUD SEGMENTS HAVING INTER-SEGMENT SEAL ARRANGEMENT

Номер: US20210003025A1
Автор: Kime Kent L., Petersen Blake, Smoke Jason, Zurmehly Luke
Принадлежит: HONEYWELL INTERNATIONAL INC.

A shroud assembly for a gas turbine engine includes a plurality of shroud segments that are attached to a shroud support with an inter-segment joint defined between shroud segments. The shroud assembly also includes a cooling flow path cooperatively defined by the shroud support and the first shroud segment. The cooling flow path includes an internal cooling passage within the shroud segments. The cooling flow path includes an outlet chamber configured to receive flow from the internal cooling passage. The shroud assembly additionally includes a seal arrangement that extends across the inter-segment joint. The seal arrangement, the first shroud segment, and the second shroud segment cooperatively define a seal chamber that is enclosed. 1. A shroud assembly for a gas turbine engine comprising:a shroud support that extends arcuately about an axis;a plurality of shroud segments that are attached to the shroud support and that are arranged annularly about the axis at different circumferential positions with respect to the axis, the plurality of shroud segments including a first shroud segment and a second shroud segment, an inter-segment joint defined circumferentially between the first and second shroud segments;a seal arrangement that extends circumferentially across the inter-segment joint; andthe seal arrangement, the first shroud segment, and the second shroud segment cooperatively defining a seal chamber that is enclosed.2. The shroud assembly of claim 1 , wherein the intersegment joint includes a leading edge and a trailing edge that are separated apart at a distance along the axis; andwherein the seal chamber is disposed proximate the trailing edge and is spaced apart at a distance from the leading edge.3. The shroud assembly of claim 1 , wherein the seal arrangement includes a first sealing member and a second sealing member that are arranged in-series with the seal chamber separating the first sealing member and the sealing member apart at a distance.4. The ...

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Номер записи: 62
07-01-2021 дата публикации

DOUBLE BOX BOAS AND CARRIER SYSTEM

Номер: US20210003026A1
Автор: Barker William M., Clark Thomas E., III William, Mahoney Danielle, Wangard, Whitney Daniel J.
Принадлежит:

A blade outer air seal assembly includes a support structure. A blade outer air seal has a plurality of seal segments that extend circumferentially about an axis and mounted in the support structure via a carrier. The carrier has a plurality of carrier segments that extend circumferentially about the axis. At least one of the seal segments have a base portion that extends between a first circumferential side and a second circumferential side and from a first axial side to a second axial side. A first wall axially spaced from a second wall. The first and second walls extend from the base portion to a first outer portion to form a first passage. The first wall has at least one slot engaged with a first carrier hook on one of the plurality of carrier segments. At least one of the carrier segments have a carrier window engaged with a support structure hook on the support structure. 1. A blade outer air seal assembly , comprising:a support structure;a blade outer air seal having a plurality of seal segments extending circumferentially about an axis and mounted in the support structure via a carrier, the carrier having a plurality of carrier segments extending circumferentially about the axis;at least one of the seal segments having a base portion extending between a first circumferential side and a second circumferential side and from a first axial side to a second axial side, a first wall axially spaced from a second wall, the first and second walls extending from the base portion to a first outer portion to form a first passage, the first wall having at least one slot engaged with a first carrier hook on one of the plurality of carrier segments; andat least one of the carrier segments having a carrier window engaged with a support structure hook on the support structure.2. The blade outer air seal assembly of claim 1 , wherein the first carrier hook extends into the first passage.3. The blade outer air seal assembly of claim 1 , wherein the first carrier hook extends ...

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Номер записи: 63
07-01-2021 дата публикации

METHOD FOR CONTROLLING A GAP MINIMIZATION OF A GAS TURBINE

Номер: US20210003027A1
Автор: Gamm Hans-Georg, Hüning Marcus, Kahlstorf Uwe
Принадлежит: SIEMENS AKTIENGESELLSCHAFT

A method for controlling a gap minimization for an adjustable gap between a rotor and a housing of a gas turbine carried out on the basis of a correlation extracted from simulation data. If the actual value (P) lies below the lower threshold (P), the gap minimization is deactivated, whereas if the actual value lies above the upper threshold (P), the gap minimization is activated. The gap minimization is activated between the thresholds (P, P) if the actual value lies above the threshold (P) but is deactivated if the actual value (P) lies below the threshold (P). 1. A method for controlling a gap minimization of an adjustable gap between a rotor and a housing of a gas turbine , wherein the gas turbine comprises a gap-adjusting device , in particular a hydraulic gap-adjusting device , the method comprising:with the aid of a simulation program, the operation of the gas turbine with different parameter settings is modeled and a simulation data set is prepared which contains the dependence of the gap size on an operating parameter,{'sub': U', 'O, 'on the basis of the simulation data set, a lower threshold (P) and an upper threshold (P) for the operating parameter are specified,'}{'sub': U', 'O', 'MAX, 'furthermore, for a transition region (M) between the lower threshold (P) and the upper threshold (P), a correlation (F) between the operating parameter and a maximum value (P) of the operating parameter is extracted from the simulation data set,'}{'sub': I', 'U', 'O, 'during operation of the gas turbine, an actual value (P) of the operating parameter is continuously determined and compared with the lower threshold (P) and the upper threshold (P),'}{'sub': MAX', 'I, 'and the maximum value (P) of the actual value (P) is determined over a specified time period,'}{'sub': I', 'U', 'O', 'I, 'wherein, in the comparison of the actual value (P) with the lower threshold (P) and the upper threshold (P), if the actual value (P){'sub': 'U', 'lies below the lower threshold (P), the gap ...

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Номер записи: 64
07-01-2021 дата публикации

TURBINE VANE AND GAS TURBINE INCLUDING THE SAME

Номер: US20210003036A1
Автор: JOO Hyun Woo, Jung Sung Chul
Принадлежит:

A turbine vane and a gas turbine including the same are provided. The turbine vane including an airfoil; an outer shroud formed at a top of the airfoil; and an inner shroud including a stress canceling part formed at a bottom of the airfoil and configured to cancel a stress applied to the airfoil by flowing combustion gas. 1. An inner shroud of a turbine vane comprising:a platform part configured to support an airfoil;a root part configured to be connected to a bottom surface of the platform part; anda stress canceling part formed at a bottom of the airfoil and configured to cancel a stress applied to the airfoil by flowing combustion gas.2. The inner shroud of the turbine vane of claim 1 ,wherein the stress canceling part comprises a protrusion configured to protrude from a bottom of one surface of the root part and a recess configured to be recessed from a bottom of the other surface of the root part.3. The inner shroud of the turbine vane of claim 2 ,wherein the protrusion and the recess include inclined surfaces at predetermined angles.4. The inner shroud of the turbine vane of claim 3 ,wherein the angles of the inclined surfaces are 5° to 45°.5. The inner shroud of the turbine vane of claim 2 ,wherein if a length of the root part is 100, lengths of the protrusion and the recess are 5 to 30.6. The inner shroud of the turbine vane of claim 2 ,wherein if a height of the root part is 100, heights of the protrusion and the recess are 10 to 40.7. A turbine vane comprising:an airfoil;an outer shroud formed at a top of the airfoil; andan inner shroud including a stress canceling part formed at a bottom of the airfoil and configured to cancel a stress applied to the airfoil by flowing combustion gas.8. The turbine vane of claim 7 ,wherein the inner shroud comprises a platform part configured to support the airfoil and a root part configured to be connected to a bottom surface of the platform part, andwherein the stress canceling part comprises a protrusion configured to ...

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Номер записи: 65
03-01-2019 дата публикации

NON-CONTACT SEAL ASSEMBLY FOR ROTATIONAL EQUIPMENT

Номер: US20190003327A1
Автор: Ackermann William K., Crawley, JR. Clifton J., Schwarz Frederick M.
Принадлежит:

Assemblies are provided for rotational equipment. One of these assemblies includes a first bladed rotor assembly, a second bladed rotor assembly, a stator vane assembly, a stator structure and a seal assembly. The second bladed rotor assembly includes a rotor disk structure. The stator vane assembly is axially between the first and the second bladed rotor assemblies. The stator structure is mated with and radially within the stator vane assembly. The seal assembly is configured for sealing a gap between the stator structure and the rotor disk structure, wherein the seal assembly includes a non-contact seal. 1. An assembly for rotational equipment , the assembly comprising:a first bladed rotor assembly;a second bladed rotor assembly including a rotor disk structure;a stator vane assembly axially between the first and the second bladed rotor assemblies;a stator structure mated with and radially within the stator vane assembly; anda seal assembly configured for sealing a gap between the stator structure and a seal land of the rotor disk structure, wherein the seal assembly includes a non-contact seal, and the seal land is configured as a cantilevered tubular body.2. The assembly of claim 1 , wherein the non-contact seal is a hydrostatic non-contact seal.3. The assembly of claim 1 , wherein the non-contact seal comprises:an annular base;a plurality of shoes arranged around and radially adjacent the rotor disk structure; anda plurality of spring elements, each of the spring elements radially between and connecting a respective one of the shoes to the base.4. The assembly of claim 3 , wherein the base is configured with a monolithic full hoop body.5. The assembly of claim 1 , wherein the stator structure is floating radially within the stator vane assembly.6. The assembly of claim 1 , whereinthe first and the second bladed rotor assemblies are turbine rotor assemblies; andthe first bladed rotor assembly is upstream of the second bladed rotor assembly.7. The assembly of ...

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Номер записи: 66
03-01-2019 дата публикации

SEALING DEVICE

Номер: US20190003328A1
Автор: HAMADA Katsuhisa, KUWAMURA Yoshihiro, NISHIKAWA Toyoharu
Принадлежит:

A sealing device includes a fin, a through hole, and a high pressure fluid supplying unit. The fin extends from a stationary body toward a rotating body in a gap between the stationary and rotating bodies. The fin is not in contact with the rotating body. The through hole is formed to be opened in at least one of the stationary body and the rotating body on an upstream side of the fin in a flow direction of a fluid to flow into the gap between the stationary body and the rotating body. The through hole is opened toward an upstream side of the fluid to flow in the gap between the stationary body and the rotating body. The high pressure fluid supplying unit is configured to supply a high pressure fluid to the gap from the through hole. The high pressure fluid has a higher pressure than the fluid. 1. A sealing device comprising:a fin extending from a stationary body toward a rotating body in a gap between the stationary body and the rotating body, the fin being not in contact with the rotating body;a through hole formed to be opened in at least one of the stationary body and the rotating body on an upstream side of the fin in a flow direction of a fluid to flow into the gap between the stationary body and the rotating body, the through hole being opened toward an upstream side of the fluid to flow in the gap between the stationary body and the rotating body; anda high pressure fluid supplying unit configured to supply a high pressure fluid to the gap from the through hole, the high pressure fluid having a higher pressure than the fluid to flow into the gap between the stationary body and the rotating body, whereinthe high pressure fluid supplied from the through hold to the gap boosts a vortex to be stronger, the vortex being generated when the fluid collides with the fin on the upstream side of the flow direction of the fluid at the fin.2. The sealing device according to claim 1 , wherein a plurality of the fins are arranged in the flow direction of the fluid claim 1 ...

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Номер записи: 67
03-01-2019 дата публикации

STEAM TURBINE

Номер: US20190003338A1
Автор: Irikawa Hiroaki, Sasaki Yuichi, Terada Katsumi
Принадлежит: MITSUBISHI HEAVY INDUSTRIES COMPRESSOR CORPORATION

A steam turbine includes: a stator vane disposed in a cylinder inside a casing through which steam flows from an upstream side toward a downstream side; and a stationary support that supports the stator vane relative to the casing. The stationary support includes: a first supporting body fixed to the casing, a second supporting body that connects the stator vane to the first supporting body; and a replacement body detachably disposed between the first supporting body and the second supporting body on the upstream side. 1. A steam turbine , comprising:a stator vane disposed in a cylinder inside a casing through which steam flows from an upstream side toward a downstream side; anda stationary support that supports the stator vane relative to the casing, wherein a first supporting body fixed to the casing;', 'a second supporting body that connects the stator vane to the first supporting body; and', 'a replacement body detachably disposed between the first supporting body and the second supporting body, on the upstream side., 'the stationary support comprises2. The steam turbine according to claim 1 , whereinthe replacement body has an erosion resistance higher than an erosion resistance of both of the first supporting body and the second supporting body.3. The steam turbine according to claim 1 , whereinthe replacement body is detachably disposed in an accommodation region of the second supporting body.4. The steam turbine according to claim 3 , whereinan insertion end of the replacement body is inserted into a holding groove of the second supporting body in the accommodating accommodation region.5. The steam turbine according to claim 1 , whereinthe replacement body is supported relative to the second supporting body by a restraining body that is detachably attached to the first supporting body.6. The steam turbine according to claim 5 , whereinone or both of the replacement body and the restraining body are detachably attached by a fastener.7. The steam turbine ...

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Номер записи: 68
14-01-2021 дата публикации

Abrasive Preforms and Manufacture and Use Methods

Номер: US20210008669A1
Автор: Minor Michael J., Stratton Eric W.
Принадлежит: Raytheon Technologies Corporation

A method for applying an abrasive comprises: applying, to a substrate, the integral combination of: a self-braze material; and an abrasive embedded in the self-braze material; and securing the combination to the substrate. 1. A braze preform comprising the integral combination of:a self-braze material; andan abrasive embedded in the self-braze material.2. The braze preform of wherein the self-braze material comprises a sintered mixture of:at least one first alloy; andat least one second alloy of high melting point relative to the first alloy.3. The braze preform of further comprising:an additional braze material layer without abrasive.4. The braze preform of further comprising:a Ni-based superalloy layer between the additional braze material layer and the combination.5. The braze preform of wherein:the Ni-based superalloy layer is a cast layer.6. An abrasive braze preform comprising:a self-braze layer; anda matrix at least partially embedding an abrasive.7. The preform of further comprising:an intermediate layer between the matrix and the self-braze layer.8. The preform of wherein:the intermediate layer is a pre-cast layer9. The preform of wherein:the intermediate layer is diffusion brazed to the self-braze layer.10. The preform of wherein: at least one first alloy; and', 'at least one second alloy of high melting point relative to the first alloy;, 'the self-braze layer comprises a sintered sheet of11. The preform of wherein:the at least one first alloy comprises about 21.25-22.75 chromium, about 5.7-6.3 aluminum, about 11.5-12.5 cobalt, about 5.7-6.3 silicon, boron in an amount no greater than 1.0 weight percent, and a balance of nickel plus impurities if any; andthe at least one second alloy comprises about 4.75-10.5 chromium, about 5.5-6.7 aluminum, up to about 13 weight percent cobalt, about 3.75-9.0 tantalum, about 1.3-2.25 molybdenum, about 3.0-6.8 tungsten, about 2.6-3.25 rhenium, up to about 0.02 boron, about 0.05-2.0 hafnium, up to about 0.14 carbon, up to ...

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Номер записи: 69
14-01-2016 дата публикации

Low Leakage Multi-Directional Interface for a Gas Turbine Engine

Номер: US20160010482A1
Автор: Broomer Mark, Davis Timothy M, Dunbar Timothy A., Keene Russell E, McGarrah Craig R, Rogers Mark J, Roy Thill Carson A., Swift Anthony B
Принадлежит:

An interface within a gas turbine engine includes a multiple of segmented components, each with a segment flange with a multiple of apertures, at least one of the multiple of apertures a first slot aperture. A full ring component with a ring flange that defines a multiple ring of apertures, at least one of the multiple of ring apertures a second slot aperture, the second slot aperture transverse to the first slot aperture. 1. An interface within a gas turbine engine , comprising:a multiple of segmented components, each with a segment flange with a multiple of segment apertures, at least one of said multiple of segment apertures includes a first slot aperture; anda full ring component with a ring flange that defines a multiple of ring apertures, at least one of said multiple of ring apertures includes a second slot aperture, said second slot aperture transverse to said first slot aperture.2. The interface as recited in claim 1 , wherein said second slot aperture is perpendicular to said first slot aperture.3. The interface as recited in claim 1 , wherein said first slot aperture is circumferentially oriented.4. The interface as recited in claim 1 , wherein said second slot aperture is radially oriented.5. The interface as recited in claim 1 , wherein said first slot aperture is circumferentially oriented and said second slot aperture is radially oriented.6. The interface as recited in claim 1 , wherein each of said multiple of segmented components are Blade Outer Air Seal (BOAS) segments.7. The interface as recited in claim 1 , wherein said full ring component is a full ring seal support.8. The interface as recited in claim 1 , further comprising a case flange adjacent to said segment flange of each of said multiple of segmented components.9. The interface as recited in claim 8 , further comprising a multiple of fastener assemblies mounted through said case flange claim 8 , said multiple of segment apertures and said multiple of ring apertures.10. The interface as ...

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Номер записи: 70
14-01-2016 дата публикации

SEAL ASSEMBLY INCLUDING A NOTCHED SEAL ELEMENT FOR ARRANGING BETWEEN A STATOR AND A ROTOR

Номер: US20160010483A1
Автор: MILLER Jonathan L.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A seal assembly with an axis is provided that includes a rotor and a seal element that is connected to a stator with an anti-rotation element. The seal element includes a first element surface, a second element surface and an annular notch. The first element surface axially engages the stator. The second element surface radially engages the rotor. The notch extends into the seal element between the first and the second element surfaces. The stator radially overlaps a portion of the notch. 1. A seal assembly with an axis , comprising:a stator;a rotor; andan annular seal element connected to the stator with an anti-rotation element, the seal element including a first element surface, a second element surface and an annular notch, the first element surface axially engaging the stator, the second element surface radially engaging the rotor, and the notch extending into a corner of the seal element between the first and the second element surfaces;wherein the stator radially overlaps a portion of the notch.2. The seal assembly of claim 1 , wherein the stator includes a stator surface that axially engages the first element surface and extends radially beyond ends of the first element surface.3. The seal assembly of claim 1 , whereinthe stator includes a stator surface that extends radially inwards to an inner end having a first radius; andthe first element surface axially engages the stator surface, and extends radially inwards to an inner end having a second radius that is greater than the first radius.4. The seal assembly of claim 1 , wherein the notch is at least partially formed by a first notch surface that extends axially from an end of the first element surface towards the second element surface.5. The seal assembly of claim 4 , wherein the notch is further formed by a second notch surface that extends radially from an end of the second element surface towards the first notch surface.6. The seal assembly of claim 5 , wherein the first and the second notch surfaces ...

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Номер записи: 71
14-01-2016 дата публикации

RING-SHAPED COMPLIANT SUPPORT

Номер: US20160010484A1
Автор: Romanov Dmitriy A.
Принадлежит:

A ring-shaped compliant support for a gas turbine engine includes, among other things, an annular case, and an adjustment member that will turn relative to the annular case if exposed to thermal energy. 1. A ring-shaped compliant support for a gas turbine engine , comprising:an annular case; andan adjustment member that will turn relative to the annular case if exposed to thermal energy.2. The ring-shaped compliant support of claim 1 , wherein the annular case defines a bore and the adjustment member is received within the bore.3. The ring-shaped compliant support of claim 1 , wherein the adjustment member comprises a plurality of wings extending radially and circumferentially from a ring.4. The ring-shaped compliant support of claim 3 , wherein slots in the adjustment member separate the plurality of wings from each other.5. The ring-shaped compliant support of claim 3 , wherein the plurality of wings comprises a first array of wings extending radially and circumferentially from the ring claim 3 , and a second array of wings extending radially and circumferentially from the ring.6. The ring-shaped compliant support of claim 5 , wherein the first array of wings are axially sequentially aligned with the second array of wings.7. The ring-shaped compliant support of claim 5 , wherein the wings in the first array are axially thinner than the wings in the second array.8. The ring-shaped compliant support of claim 7 , wherein the wings in the first array are upstream from the wings in the second array.9. The ring-shaped compliant support of claim 5 , wherein the wings in the first array are relatively insulated and the wings in the second array are relatively uninsulated.10. The ring-shaped annular compliant support of claim 7 , wherein the wings in the first array are upstream from the wings in the second array.11. The ring-shaped compliant support of claim 5 , wherein the wings in the first array are axially thinner than the wings in the second array.12. The ring-shaped ...

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Номер записи: 72
11-01-2018 дата публикации

TURBINE COMPONENT THERMAL BARRIER COATING WITH CRACK ISOLATING, CASCADING, MULTIFURCATED ENGINEERED GROOVE FEATURES

Номер: US20180010469A1
Автор: Kadau Kai, Pascual-Gutierrez Jose Antonio, Sharma Atin
Принадлежит:

Turbine engine () components, such as blades (), vanes (), ring segment abradable surfaces or transitions (), have furcated engineered groove features (EGFs) () that cut into the outer surface of the component's thermal barrier coating (TBC). In some embodiments, the EGF planform pattern defines adjoining outer hexagons (). In some embodiments, the EGF pattern further defines within each outer hexagon () a planform pattern of adjoining inner polygons (). At least three respective groove segments () within the EGF pattern () converge at each respective outer hexagonal vertex () or inner polygonal vertex () in a multifurcated pattern, so that crack-inducing stresses are attenuated in cascading fashion, as the stress (σ) is furcated (σ, σ) at each successive vertex juncture. 1. A combustion turbine engine blade , vane , transition , or ring segment abradable component having a heat insulating outer surface for exposure to combustion gas , comprising:a metallic substrate having a substrate surface;an anchoring layer built upon the substrate surface;a thermally sprayed or vapor deposited or solution/suspension plasma sprayed, single- or multi-layer thermal barrier coat (TBC) having a TBC inner surface applied over and coupled to the anchoring layer and a TBC outer surface for exposure to combustion gas; and the EGF pattern defining a planform pattern of adjoining outer hexagons respectively having six hexagonal vertices, with each respective pair of adjoining outer hexagons sharing a common groove segment,', 'the EGF pattern further defining within each outer hexagon a planform pattern of adjoining inner polygons, the adjoining inner polygons respectively sharing at least a common inner polygonal vertex and respectively fully circumscribed within its respective outer hexagon,', 'at least three respective groove segments within the EGF pattern converging at each respective outer hexagonal or inner polygonal vertex in a multifurcated pattern, so that each converging groove ...

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Номер записи: 73
14-01-2021 дата публикации

SEALING ARRANGEMENT BETWEEN TURBINE SHROUD SEGMENTS

Номер: US20210010381A1
Автор: AZAD Gm Salam, Chen Runzhong, Lee Ching-Pang
Принадлежит:

A shroud assembly for a turbine engine includes a seal for sealing a gap between a first mate face of a first shroud segment and a second mate face of a circumferentially adjacent second shroud segment. The seal is received in first and second slots formed respectively on the first and second mate faces. The first and second slots extend axially between a leading edge and a trailing edge of the respective shroud segment. The first slot is open at the leading and the trailing edges while the second slot is open at the leading edge and closed at the trailing edge. The seal has axially extending first and second sides which are receivable respectively within the first and second slots. The seal has an axial length substantially equal to tan axial length of the shroud segments and has a cutout on the second side at a trailing edge end of the seal. 1. A shroud for a turbine engine , comprising:a first shroud segment having a first mate face and a second shroud segment having a second mate face, the first mate face being positioned circumferentially adjacent to the second mate face,a seal for sealing a gap between the first and second mate faces,wherein the seal is received, at least in part, in a first slot formed on the first mate face and a second slot formed on the second mate face,wherein the first and second slots extend axially between a leading edge and a trailing edge of the respective shroud segment, the first slot being open at the leading edge and at the trailing edge, the second slot being open at the leading edge and closed at the trailing edge,wherein the seal comprises axially extending first and second sides which are receivable respectively within the first slot and the second slot, the seal having an axial length substantially equal to an axial length of the shroud segments and having a cutout on the second side at a trailing edge end of the seal.2. The shroud according to claim 1 , wherein an axial length of the cutout is equal to or greater than an ...

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Номер записи: 74
10-01-2019 дата публикации

TURBINE FOR A TURBINE ENGINE

Номер: US20190010818A1
Автор: Drouet Franck Robert, Evain Gaël Frédéric Claude Cyri, Lambert Olivier Arnaud Fabien
Принадлежит: SAFRAN AIRCRAFT ENGINES

The invention relates to a turbine for a turbine engine, having a stator and a rotor comprising a rotor wheel having vanes the radially external periphery of which comprises at least one lip which radially extends outwards, with sealing means radially extending about the vanes and comprising a ring. The radially external end of the lip cooperates with said ring so as to form a seal of the labyrinth type. 1. A turbine for a turbine engine , the turbine having a stator and a rotor comprising a rotor wheel having vanes the radially external periphery of which comprises at least one lip which radially extends outwards , with sealing means radially extending about the vanes and comprising a sealing ring; with the radially external end of the lip cooperating with said sealing ring so as to form a seal of the labyrinth type , wherein said sealing ring comprises at least one first portion and one second portion radially offset relative to one another , with the first portion and/or the second portion defining a groove wherein the lip is inserted , with the first portion and/or the second portion each cooperating with at least one lip of the vanes axially located opposite said first and second portions , with the first portion comprising a first protruding zone engaged in a form-fitting manner in the axial direction into a first recessed zone of the second portion , with the stator comprising means for holding the first and second portions in position relative to the stator.2. The turbine according to claim 1 , wherein the position-holding means comprise a stop for radially bearing the first portion and means for tightening the second portion against the stator claim 1 , with the first portion being positioned upstream of the second portion.3. The turbine according to claim 1 , wherein the first portion is radially held claim 1 , upstream claim 1 , by the stator claim 1 , with the first portion being radially held claim 1 , downstream claim 1 , by the second portion.4. The ...

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Номер записи: 75
10-01-2019 дата публикации

TURBOMACHINE SEALING RING

Номер: US20190010820A1
Автор: EIBELSHAEUSER Peter, Jakimov Andreas, Kraus Juergen, SCHERER Christian, SCHLOTHAUER Steffen
Принадлежит:

A sealing ring for a turbomachine, in particular a compressor or turbine stage of a gas turbine is provided, the sealing ring having an, in particular at least partially honeycomb-like and/or integral, seal, in particular an abradable coating (); a profile cross section of the sealing ring, which is in particular at least partially manufactured by a generative manufacturing process, varying in the circumferential direction, in particular at least in some regions uniformly and/or non-uniformly at one or more circumferential positions, in particular separation points. 113-. (canceled)14. A sealing ring for a turbomachine , the sealing ring comprising:a seal;a profile cross section of the sealing ring varying in a circumferential direction.15. The sealing ring as further comprising at least one radial rib extending circumferentially at least in some regions or at least one axial rib extending circumferentially at least in some regions.16. The sealing ring as recited in wherein the axial rib is disposed on a radial flange or varies in dimension in the circumferential direction dimension claim 15 , or the radial rib is disposed on a circumferential surface or varies in dimension in the circumferential direction.17. The sealing ring as recited in wherein the axial rib varies in the axial or radial dimension.18. The sealing ring as recited in wherein the radial rib is disposed on a circumferential surface in a circumferential groove for attachment of the sealing ring.19. The sealing ring as recited in wherein the circumferential groove includes an undercut.20. The sealing ring as recited in wherein the radial rib varies in the circumferential direction.21. The sealing ring as recited in further comprising circumferentially spaced-apart supports disposed on a circumferential surface and adapted to radially retain bushings of a carrier of the sealing ring.22. The sealing ring as recited in wherein the circumferential surface is in a circumferential groove for attachment of ...

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Номер записи: 76
10-01-2019 дата публикации

SYSTEM FOR MODULATING TURBINE BLADE TIP CLEARANCE

Номер: US20190010822A1
Автор: Chaudhry Zaffir A., McCaffrey Michael G., SUN FANPING
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A system for modulating turbine blade tip clearance is provided. The system may comprise an actuation control system having at least one actuator configured to modulate turbine blade tip clearance. Each actuator may comprise a solid-state motion amplification device such as a flextensional actuator. Each actuator may be in operable communication with a blade outer air seal (BOAS) segment or a BOAS mounting block. The actuators may be configured to move the BOAS segment and/or the BOAS mounting block in a radial direction from a first position to a second position to control tip clearance. 1. An actuation control system , comprising:an actuator in operable communication with a blade outer air seal (BOAS) mounting block; anda BOAS segment coupled to the BOAS mounting block, wherein the BOAS mounting block is configured to move from a first position to a second position in response to an actuation from the actuator.2. The actuation control system of claim 1 , wherein the actuator comprises a control wire configured to control the actuation of the actuator.3. The actuation control system of claim 2 , wherein the control wire comprises a shape memory alloy (SMA) wire configured to contract in response to receiving an electrical current claim 2 , and wherein the contraction of the SMA wire controls the actuation of the actuator.4. The actuation control system of claim 3 , further comprising a power supply having a controller claim 3 , wherein the power supply is configured to transmit the electrical current to the control wire in response to the controller determining a tip clearance actuation event.5. The actuation control system of claim 1 , wherein the actuator comprises a control rod coupled at a first rod end to the actuator and at a second rod end to the BOAS mounting block claim 1 , wherein the control rod is configured to move the BOAS mounting block from the first position to the second position in response to the actuation from the actuator.6. The actuation ...

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Номер записи: 77
10-01-2019 дата публикации

System for modulating turbine blade tip clearance

Номер: US20190010823A1
Автор: Fanping Sun, Hailing Wu, Joel H. Wagner, Lee A. Hoffman, Michael G. Mccaffrey, Seung Bum Kim, Zaffir A. Chaudhry
Принадлежит: United Technologies Corp

A system for modulating turbine blade tip clearance is provided. The system may comprise an actuation control system having at least one actuator configured to modulate turbine blade tip clearance between a turbine blade tip and a blade outer air seal (BOAS). Each actuator may be coupled to the BOAS. Each actuator may comprise a solid-state motion amplification device such as a flextensional actuator. The actuators may be configured to move the BOAS in a radial direction from a first position to a second position to control tip clearance.

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Номер записи: 78
09-01-2020 дата публикации

ROTOR FOR A TURBOMACHINE, AND TURBOMACHINE HAVING SUCH A ROTOR

Номер: US20200011193A1
Автор: ALBERS Lothar
Принадлежит:

A rotor () for a turbomachine, in particular for an aircraft engine, having a rotor base body (), on which at least one sealing fin (), which is disposed on a base (), is provided for cooperating with an associated sealing element () of the turbomachine; relative to an axial direction of the rotor (), the base () having a base portion () disposed upstream of the sealing fin () and a base portion () disposed downstream thereof, for supporting masks during the coating of sealing fins; the upstream base portion () and the downstream base portion () having different radial distances (A, A) to a radially outer sealing tip () of the sealing fin (). Also, a turbomachine having at least one such rotor (). 112-. (canceled)13. A rotor for a turbomachine , the rotor comprising:a rotor base body having at least one sealing fin disposed on a base, the sealing fin for cooperating with an associated sealing element of the turbomachine; and, relative to an axial direction of the rotor, the base having a base portion upstream of the sealing fin, and a base portion downstream of the sealing fin, for supporting masks during coating of the sealing fin, wherein the upstream base portion and the downstream base portion have different radial distances to a radially outer sealing tip of the sealing fin.14. The rotor as recited in wherein a ratio between the radial distance of the upstream base portion and the radial distance of the downstream base portion is between 0.25 and 4 claim 13 , the ratio not being 1.15. The rotor as recited in wherein the rotor is a compressor rotor claim 13 , and the upstream base portion has a larger distance to the radially outer sealing tip of the sealing fin than the downstream base portion.16. The rotor as recited in wherein the rotor is a turbine rotor claim 13 , and the upstream base portion has a smaller distance to the radially outer sealing tip of the sealing fin than the downstream base portion.17. The rotor as recited in wherein the upstream base ...

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Номер записи: 79
09-01-2020 дата публикации

AIRCRAFT TURBINE ENGINE SEALING MODULE

Номер: US20200011194A1
Автор: BEAUQUIN Nicolas Jean-Marc Marcel, Martet Renaud James, MILLIER Vincent François Georges, MORLIERE Simon Nicolas
Принадлежит:

Turbine engine turbine sealing module extending about an axis and including a distributor fixed to a casing, and at least one blade connected to an outer platform. The turbine module further includes an impeller mounted rotating inside the casing and surrounded by a sealing ring fastened to this casing. The sealing ring includes an annular row of ring sectors arranged such that the circumferential end edges of two adjacent sectors are facing one another. Each sector includes a body configured to engage with at least one seal lip carried by the impeller and a hook which extends circumferentially which is configured to engage with a fastening rail of the casing. Each ring sector further includes a deflector which extends radially inwards and upstream with respect to the axis, such that the radially inward end thereof extends around a downstream end of the outer platform of the distributor. 1. Turbine engine turbine sealing module , in particular for aircraft , this sealing module extending about an axis and comprising a distributor fixed to a casing , the distributor having at least one blade connected to an outer platform , the outer platform comprising a spoiler for fixing to the casing , the sealing module further comprising an impeller mounted rotating inside the casing and surrounded by a sealing ring fastened to this casing , this sealing ring being sectored and comprising an annular row of ring sectors arranged such that the circumferential end edges of two adjacent sectors are facing each other , each ring sector comprising a body carrying an abradable coating configured to engage with at least one seal lip carried by the impeller and a hook which extends circumferentially by being located upstream of said abradable coating and which is configured to engage with a fastening rail of the casing , this hook having , in the cross-section , a general C-shape , of which the opening is axially oriented upstream and intended to receive said rail , wherein each ring ...

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Номер записи: 80
09-01-2020 дата публикации

A TIP MACHINING METHOD AND SYSTEM

Номер: US20200011195A1
Автор: Milne Trevor, Tatton Kieran
Принадлежит: SIEMENS AKTIENGESELLSCHAFT

A method of machining a tip profile of a blade for a turbomachine includes coupling the blade to a component of the turbomachine; supporting the component on a machining apparatus, the machining apparatus being configured to remove material from the blade according to a cutting path defined within a coordinate system of the machining apparatus, wherein the component is supported such that a datum axial end face of the component is aligned with a datum of the coordinate system of the machining apparatus; and machining the blade according to the cutting path. A system for machining a tip profile of a blade for a turbomachine accomplishes the method. 1. A method of machining a tip profile of a blade for a turbomachine , the method comprising:coupling the blade to a component of the turbomachine;supporting the component on a machining apparatus, the machining apparatus being configured to remove material from the blade according to a cutting path defined within a coordinate system of the machining apparatus, wherein the component is supported such that a datum axial end face of the component is aligned with a datum of the coordinate system of the machining apparatus;defining the cutting path relative to the datum of the coordinate system; andmachining the blade according to the cutting path.2. The method of claim 1 , wherein the method further comprises:defining the cutting path by offsetting a portion of a gas path of the turbomachine by a desired tip clearance amount, wherein the gas path is defined relative to the datum axial end face of the component.3. The method of claim 1 ,wherein the cutting path extends between an upstream cutting point and a downstream cutting point, wherein the upstream cutting point is upstream of a leading edge of the blade and the downstream cutting point is downstream of a trailing edge of the blade.4. The method of claim 1 , wherein the method further comprises:measuring the tip of the blade at upstream and downstream measurement points, ...

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Номер записи: 81
09-01-2020 дата публикации

Method and Tooling for Manufacturing a Fan Case Assembly

Номер: US20200011198A1
Автор: Eggleton Robert J., Hall Steven M.
Принадлежит:

A method of manufacturing a fan case assembly for a gas turbine engine, the fan case assembly comprising a fan case and a fan liner, wherein the method comprises: providing a mounting ring configured to extend about an inner circumference of the fan case; providing a gasket at an axial end of the mounting ring, wherein the gasket extends around the inner circumference of the fan case; providing the fan liner at the axial end of the mounting ring with the gasket, wherein the fan liner extends around the inner circumference of the fan case; and heating the fan case assembly so as to cure a resin provided between the fan case and fan liner, wherein the heating causes the mounting ring to expand radially relative to the fan case such that the gasket is brought into engagement with the fan case and unwanted migration of resin away from between the fan case and fan liner is restricted. 1. A method of manufacturing a fan case assembly for a gas turbine engine , the fan case assembly comprising a fan case and a fan liner , wherein the method comprises:providing a mounting ring configured to extend about an inner circumference of the fan case;providing a gasket at an axial end of the mounting ring, wherein the gasket extends around the inner circumference of the fan case;providing the fan liner at the axial end of the mounting ring with the gasket, wherein the fan liner extends around the inner circumference of the fan case; andheating the fan case assembly so as to cure a resin provided between the fan case and the fan liner, wherein the heating causes the mounting ring to expand radially relative to the fan case such that the gasket is brought into engagement with the fan case and unwanted migration of resin away from between the fan case and the fan liner is restricted.2. The method of claim 1 , wherein a first end of the fan liner is provided at the axial end of the mounting ring claim 1 , and wherein the method further comprises:providing a further gasket at a second ...

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Номер записи: 82
15-01-2015 дата публикации

BLADE CLEARANCE CONTROL FOR GAS TURBINE ENGINE

Номер: US20150016946A1
Автор: Ottow Nathan W.
Принадлежит:

An apparatus and method for controlling a clearance between the blades of a turbomachinery component and flow forming surface are disclosed herein, and includes controlling the clearance by moving the surface axially relative to the turbomachinery component. In one embodiment the apparatus includes an impeller rotatable about a first axis, a shroud encircling the impeller, and a first ring encircling the first axis. An actuator is operably engaged with the first ring to pivot the first ring about the first axis. The apparatus also includes at least one cam engaged with the first ring and at least one cam follower engaged with the shroud. Pivoting movement of the first ring about the first axis results in the at least one cam urging the at least one cam follower and the shroud along the first axis to vary a distance between the plurality of blades and the shroud. 1. An apparatus comprising:a gas turbine engine bladed turbomachinery component centered on a first axis and operable to rotate about said first axis, said bladed turbomachinery component including an inner base and a plurality of blades extending radially outward from said inner base and also extending along said first axis, wherein a plurality of fluid channels are respectively defined between adjacent pairs of said plurality of blades;a flow path forming component encircling said bladed turbomachinery component and substantially enclosing a radially outward side of said blades along said first axis;a pivoting member circumferentially extending about said first axis and adjacent to at least part of said flow path forming component along said first axis;an actuator operably engaged with said pivoting member to pivot said pivoting member about said first axis;at least one cam engaged with said pivoting member; andat least one cam follower engaged with said flow path forming component, wherein pivoting movement of said pivoting member about said first axis results in said at least one cam urging said at least ...

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Номер записи: 83
15-01-2015 дата публикации

Blade outer air seal assembly and support

Номер: US20150016954A1
Автор: Anne-Marie B. Thibodeau, Bruce E. Chick, Dmitriy A. Romanov, James N. Knapp, Jeffrey Vincent Anastas, Russell E. Keene, Thurman Carlo Dabbs
Принадлежит: United Technologies Corp

An blade outer air seal support assembly includes a main support member configured to support a blade outer air seal. The main support member extends generally axially between a leading edge portion and a trailing edge portion. The leading edge portion is configured to be slidably received within a groove established by the blade outer air seal. A support tab extends radially inward from the main support member toward the blade outer air seal. The support tab configured to contact an extension of the blade outer air seal to limit relative axial movement of the blade outer air seal. A gusset spans between the support tab and the main support member.

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Номер записи: 84
15-01-2015 дата публикации

BLADE TRACK ASSEMBLY, COMPONENTS, AND METHODS

Номер: US20150016970A1
Автор: Freeman Ted Joseph, Griffiths George F., Smith Clayton C., Westphal William
Принадлежит:

A blade track assembly is disclosed having a variety of features. The assembly can have annular or segmented components, or a combination of the two. In one form the assembly includes blade tracks having a forward and aft edge that can be received in an opening of respective hangers. The hangers can include anti-movement features to discourage movement of a blade track. A rib can extend between hangers and in one form can be used as part of a seal assembly. Clips can be used to secure the blade track in openings of the respective hangers, as well as to discourage movement of the blade track. 1. An apparatus comprising:a gas turbine engine having a blade track that includes a main body and fore and aft edges that extend axially. and circumferentially, the blade track supported by a gas turbine engine structure having a receiving portion into which the fore and aft edges can be slindingly inserted.2. The apparatus of claim 1 , wherein the receiving portion includes a radially inner member and a radially outer member for receiving into the receiving portion claim 1 , and wherein the gas turbine engine structure includes a plurality of hangers having the receiving portion.3. The apparatus of claim 2 , wherein the gas turbine engine blade track includes a plurality of segmented blade tracks.4. The apparatus of claim 1 , wherein at least one of the circumferentially extending fore and aft edges extends a substantial length of the blade track.5. The apparatus of claim 3 , wherein the gas turbine engine structure includes a plurality of segmented hangers each having the receiving portion.6. The apparatus of claim 1 , which further includes a u-shaped clip structured to grip the blade track and shaped to fit within the receiving portion of the gas turbine engine structure.7. The apparatus of claim 1 , wherein the gas turbine engine structure is an intermediate structure having the receiving portion for receiving one of the fore and aft circumferentially extending edges and a ...

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Номер записи: 85
03-02-2022 дата публикации

MOVING BLADE FOR A WHEEL OF A TURBINE ENGINE

Номер: US20220034231A1
Автор: BASSERY Josserand Jacques Andre, DUPEYRE Raphael Jean Philippe, FANELLI Jeremy Jacques Attilio, FRANCOIS Etienne Leon
Принадлежит:

A moving blade for a wheel of an aircraft turbine engine, including an aerodynamic aerofoil and an outer heel defining the aerofoil. The heel includes a platform and a first lip that projects from the platform. The first lip is inclined upstream and peripherally along an axis of elongation. The heel includes a row of ribs that are arranged at a distance from each other. The row of ribs extends along the axis of elongation and from the platform up to the first lip. The ribs are arranged upstream of the first lip in such a way as to generate turbulence upstream of first lip. 1. A movable vane for a wheel of an aircraft turbomachine , said vane comprising an aerodynamic blade extending along a stacking axis and an outer heel delimiting said blade along said stacking axis , said heel comprising a platform and a first lip protruding from said platform , the first lip being inclined upstream at an acute angle to said stacking axis , said first lip extending circumferentially along an axis of elongation , wherein the heel comprises a row of ribs spaced from each other , said row comprising at least two ribs , said row of ribs extending along said axis of elongation , each rib extending along said stacking axis from said platform to said first lip , each rib being arranged upstream of the first lip according to the direction of gas flow around said blade so as to generate turbulence upstream of said first lip.2. The vane according to claim 1 , wherein at least one of the ribs is inclined at an acute angle with respect to said axis of elongation claim 1 , said acute angle being defined from the axis of elongation to the corresponding rib in a trigonometric direction claim 1 , said acute angle being measured in a plane perpendicular to said stacking axis.3. The vane according to claim 2 , wherein said acute angle is greater than or equal to 30° and less than 90°.4. The vane according to claim 1 , wherein at least one of the ribs has a parallelogram-shaped profile cross- ...

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Номер записи: 86
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: 87
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: 88
21-01-2016 дата публикации

BRUSH SEAL REPAIR METHOD

Номер: US20160017742A1
Автор: Cote Steven M., Walker David E.
Принадлежит:

A method of working a brush seal assembly for a turbine engine is disclosed and includes the steps of removing an alignment tab from a used brush seal assembly, restoring at least one dimension of the alignment tab to a desired condition and attaching the restored alignment tab to a new brush seal assembly. 1. A method of repairing a brush seal assembly for a turbine engine , the method comprising;removing an alignment tab from a used brush seal assembly;restoring at least one dimension of the alignment tab to a desired condition; andattaching the restored alignment tab to a new brush seal assembly.2. The method of repairing as recited in claim 1 , wherein removing the alignment tab comprises cutting a portion of the used brush seal assembly proximate the alignment tab from the used brush seal assembly.3. The method of repairing as recited in claim 2 , wherein the alignment tab includes an L-shaped cross section with a first part fixed to a first surface of the used brush seal assembly and a second part extending transverse from the first part.4. The method of repairing as recited in claim 3 , wherein the first part of the alignment tab includes an opening for receiving a portion of a weld for attaching the alignment tab.5. The method of repairing as recited in claim 1 , wherein the alignment tab comprises a rectangular section attached to an annular ring of the brush seal assembly.6. The method of repairing as recited in claim 2 , wherein restoring the at least one dimension of the alignment tab includes adding material to the alignment tab.7. The method of repairing as recited in claim 6 , including the step of machining the added material restore the at least one dimensions to the desired condition.8. The method of repairing as recited in claim 1 , wherein attaching the alignment tab to the new brush seal assembly comprise welding the restored alignment tab to a first surface of the new brush seal assembly in a location common with the location of the alignment ...

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Номер записи: 89
18-01-2018 дата публикации

ASSEMBLY FOR SUPPORTING AN ANNULUS

Номер: US20180016941A1
Автор: MCDONAGH Stephen
Принадлежит: ROLLS-ROYCE PLC

The annulus is bound by an inner hub wall and an outer casing and includes a support structure, the support structure bearing the inner hub wall; at least one spigot passing through the hub wall and at least one strut arranged to pass through the spigot of the inner hub wall and across the annulus. The strut has a first end having an abutment arm extending to form an abutment shoulder. Alignable holes pass through the abutment arm and the spigot and these holes are configured to receive a cross pin which is in turn configured to fit snugly through the holes. The configuration is such that, the abutment rim and abutment shoulder are located radially inwardly of the holes and cross pin and outside of the annulus. 1. An assembly for supporting an annulus , the annulus bound by an inner hub wall and an outer casing , the assembly comprising; a support structure , the support structure bearing the hub wall; at least one spigot passing through the hub wall , the spigot incorporating an abutment rim andat least one strut arranged to pass through the spigot of the hub wall and across the annulus, the strut comprising;a first end having an abutment arm and alignable holes passing through the abutment arm and the spigot, the holes configured to receive a cross pin which is in turn configured to fit snugly through the holes;an abutment shoulder of the abutment arm for engaging the abutment rim the configuration being such that, the abutment dm and abutment shoulder are located radially inwardly of the holes and cross pin and outside of the annulus.2. An assembly as claimed in wherein the spigot is formed integrally with the hub wall.3. An assembly as claimed in wherein the spigot is provided with an external profile shaped to blend with a profile of the strut diameter which protrudes into the annulus when the strut is arranged in the annulus.4. An assembly as claimed in further comprising; in a second end of the strut claim 1 , a threaded hole claim 1 , a bolt configured for ...

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Номер записи: 90
18-01-2018 дата публикации

Clearance control between rotating and stationary structures

Номер: US20180017067A1
Автор: Louis R. Mouza, Philip R. Rioux
Принадлежит: United Technologies Corp

Aspects of the disclosure are directed to a system of an engine, comprising: a clearance control thermal ring, and a seal ring, where a radial gap with respect to an axial centerline of the engine is formed between a radial end of the clearance control thermal ring and a facing radial surface of the seal ring, where the clearance control thermal ring is made of a first material and the seal ring is made of a second material that is different from the first material, and where a first coefficient of thermal expansion of the first material is less than a second coefficient of thermal expansion of the second material.

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Номер записи: 91
17-01-2019 дата публикации

METHOD AND APPARATUS FOR SEALING COMPONENTS OF A GAS TURBINE ENGINE WITH A DIELECTRIC BARRIER DISCHARGE PLASMA ACTUATOR

Номер: US20190017406A1
Автор: Bagdonis Guthrie
Принадлежит:

A system and method for aerodynamically sealing rotating and fixed components of a gas turbine engine. The system includes a gas turbine engine having a casing and a rotating portion, a plasma actuator having a first and a second electrodes, the first electrode including at least one section of substantially flat conductive material encased in a dielectric material forming at least a portion of a cylinder disposed circumferentially on the casing. The system also includes the rotating portion operably configured as the second electrode, and an excitation source operably connected between the first electrode and the second electrode, the excitation source generating an excitation signal and applying it to the first and second electrodes to cause the actuator to form a plasma between the first and second electrodes, the plasma inducing an airflow between the casing and the rotating portion. 1. A non-contacting method for aerodynamically sealing rotating and fixed components of a gas turbine engine with a dielectric barrier discharge plasma actuator having a first electrode and a second electrode , the method comprising:disposing the first electrode circumferentially on a casing of the gas turbine engine, the first electrode comprising at least one section of substantially flat conductive material encased in a dielectric material generally forming a cylinder;configuring a rotating portion of the gas turbine engine as the second electrode;operably connecting an excitation source between the first electrode and the second electrode; andgenerating an excitation signal with the excitation source and applying it to the first and second electrode to cause the actuator to form plasma between the first and second electrode, the plasma inducing an airflow between the casing and the rotating portion.2. The method of claim 1 , wherein the disposing includes applying the first electrode to at least a portion of an inner circumference of the casing.3. The method of claim 1 , wherein ...

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Номер записи: 92
17-01-2019 дата публикации

FLOW METERING AND RETENTION SYSTEM

Номер: US20190017412A1
Автор: Juh Christopher M., Petrillo Dalton C.
Принадлежит:

A flow metering and retention system includes a first disk that is annular in shape surrounding a centerline and extending axially along the centerline, a first coverplate axially rearward of the first disk with the first coverplate having an axially rearward extending arm, a second disk that is annular in shape surrounding the centerline and rearward of the first coverplate, a second coverplate at least partially between the first coverplate and the second disk; and a ring adjacent to the radially outer side of the slot of the second disk. The second disk has a slot into which the arm of the first coverplate extends with the slot having a radially outer side and a radially inner side, and the ring is configured to meter air flowing between the radially outer side of the slot and the arm of the first coverplate. 1. A system for a gas turbine engine extending along a centerline , the system comprising:a high pressure turbine disk;a high pressure turbine coverplate axially rearward of the high pressure turbine disk and having an arm that extends axially rearward;a low pressure turbine disk axially rearward of the high pressure turbine coverplate, the low pressure turbine disk having a slot into which the arm of the high pressure turbine coverplate extends;a low pressure turbine coverplate at least partially between the high pressure turbine coverplate and the low pressure turbine disk;an interface between the low pressure turbine disk and the low pressure turbine coverplate, the interface being radially outward from the arm of the high pressure turbine coverplate and having a groove; anda ring within the groove at the interface and extending radially inward towards the arm of the high pressure turbine coverplate, the ring being configured to prevent axial movement of the low pressure turbine coverplate relative to the low pressure turbine and to form a first metering point to meter air flowing between the high pressure turbine coverplate and the low pressure turbine ...

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Номер записи: 93
17-01-2019 дата публикации

NON-CONTACT SEAL WITH PROGRESSIVE RADIAL STOP(S)

Номер: US20190017607A1
Автор: Chuong Conway, Wilson Ross
Принадлежит:

An assembly for rotational equipment includes a plurality of seal shoes, a seal base, a plurality of spring elements and a frangible element. The seal shoes are arranged around an axis in an annular array. The seal base circumscribes the annular array of the seal shoes. Each of the spring elements is radially between and connects a respective one of the seal shoes and the seal base. A first of the spring elements includes a first mount, a second mount and a spring beam. The first mount is connected to a first of the seal shoes. The second mount is connected to the seal base. The spring beam extends longitudinally between and connects the first mount and the second mount. The frangible element is configured to restrict radial outward movement of the first of the seal shoes. 1. An assembly for rotational equipment , comprising:a plurality of seal shoes arranged around an axis in an annular array;a seal base circumscribing the annular array of the seal shoes;a plurality of spring elements, each of the spring elements radially between and connecting a respective one of the seal shoes and the seal base, a first of the spring elements including a first mount, a second mount and a spring beam, the first mount connected to a first of the seal shoes, the second mount connected to the seal base, and the spring beam extending longitudinally between and connecting the first mount and the second mount; anda frangible element configured to restrict radial outward movement of the first of the seal shoes.2. The assembly of claim 1 , wherein the frangible element is configured to progressively restrict the radial outward movement of the first of the seal shoes.3. The assembly of claim 1 , whereinthe frangible element is adapted to enable a first magnitude of the radial outward movement of the first of the seal shoes during a first mode of operation;the frangible element is adapted to enable a second magnitude of the radial outward movement of the first of the seal shoes during a ...

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Номер записи: 94
21-01-2021 дата публикации

CMC BOAS ARRANGEMENT

Номер: US20210017871A1
Автор: Blaney Ken F., Clark Thomas E.
Принадлежит:

A blade outer air seal assembly includes a blade outer air seal that has a plurality of segments that extend circumferentially about an axis and are mounted in a support structure via a carrier. At least one of the segments have a first wall circumferentially spaced from a second wall. A base portion extends from the first wall to the second wall. The first and second walls extend at an angle less than 90° from the base portion. The carrier has a flat portion that extends from the first wall to the second wall. 1. A blade outer air seal assembly , comprising:a blade outer air seal having a plurality of segments extending circumferentially about an axis and mounted in a support structure via a carrier, at least one of the segments having a first wall circumferentially spaced from a second wall, a base portion extending from the first wall to the second wall, the first and second walls extending at an angle less than 90° from the base portion; andthe carrier having a flat portion extending from the first wall to the second wall.2. The blade outer air seal assembly of claim 1 , wherein the carrier has a first angled surface and a second angled surface claim 1 , the first and second angled surfaces in engagement with the first and second walls.3. The blade outer air seal assembly of claim 2 , wherein an aft wall extends circumferentially outward past the first and second angled surfaces at an aft end of the carrier.4. The blade outer air seal assembly of claim 1 , wherein a first hook extends from the first wall and a second hook extends from the second wall.5. The blade outer air seal assembly of claim 4 , wherein the first and second hooks extend away from one another.6. The blade outer air seal assembly of claim 1 , wherein the carrier includes a carrier hook extending radially outward and engaging the support structure.7. The blade outer air seal assembly of claim 6 , wherein the carrier hook extends across a circumferential width of the carrier.8. The blade outer ...

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Номер записи: 95
21-01-2021 дата публикации

CMC BOAS ARRANGEMENT

Номер: US20210017872A1
Автор: Blaney Ken F., Clark Thomas E.
Принадлежит:

A blade outer air seal assembly includes a blade outer air seal that has a plurality of segments that extend circumferentially about an axis and are mounted in a support structure. At least two of the segments have a first wall circumferentially spaced from a second wall. A base portion extends from the first wall to the second wall. A first hook extends from the first wall and a second hook extends from the second wall. A wedge seal is arranged between the at least two adjacent seal segments. A clip is configured to bias the wedge seal radially inward. 1. A blade outer air seal assembly , comprising:a blade outer air seal having a plurality of segments extending circumferentially about an axis and mounted in a support structure, at least two of the segments having a first wall circumferentially spaced from a second wall, and a base portion extending from the first wall to the second wall, a first hook extending from the first wall and a second hook extending from the second wall;a wedge seal arranged between the at least two adjacent seal segments; anda clip configured to bias the wedge seal radially inward.280. The blade outer air seal assembly of claim 1 , wherein the wedge seal extends at least % of an axial length of the at least two adjacent seal segments.3. The blade outer air seal assembly of claim 1 , wherein the wedge seal has a generally triangular cross section.4. The blade outer air seal assembly of claim 1 , wherein the wedge seal has a first angled surface and a second angled surface claim 1 , the first and second angled surfaces configured to mate with the first and second walls on the at least two adjacent seal segments.5. The blade outer air seal assembly of claim 1 , wherein the clip has a first axially extending portion and a second axially extending portion spaced radially from the first axially extending portion by a radially extending portion.6. The blade outer air seal assembly of claim 5 , wherein the radially extending portion is arranged ...

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Номер записи: 96
21-01-2021 дата публикации

CMC BOAS ARRANGEMENT

Номер: US20210017874A1
Автор: Blaney Ken F., Clark Thomas E.
Принадлежит:

A blade outer air seal assembly includes a blade outer air seal that has a plurality of segments that extend circumferentially about an axis and are mounted in a support structure via a carrier. At least one of the segments have a first hook circumferentially spaced from a second hook. A base portion extends from the first hook to the second hook. The carrier has a cavity on a radially inner surface between the carrier and the base portion. 1. A blade outer air seal assembly , comprising:a blade outer air seal having a plurality of segments extending circumferentially about an axis and mounted in a support structure via a carrier, at least one of the segments having a first hook circumferentially spaced from a second hook, and a base portion extending from the first hook to the second hook; andthe carrier having a cavity on a radially inner surface between the carrier and the base portion.2. The blade outer air seal assembly of claim 1 , wherein the cavity extends at least 50% of a circumferential width of the base portion.3. The blade outer air seal assembly of claim 1 , wherein a plurality of grooves are arranged in the cavity.4. The blade outer air seal assembly of claim 1 , wherein an impingement plate is arranged between the carrier and the base portion.5. The blade outer air seal assembly of claim 4 , wherein the impingement plate has a plurality of orifices.6. The blade outer air seal assembly of claim 5 , wherein the plurality of orifices are configured to communicate cooling air to the at least one segment.7. The blade outer air seal assembly of claim 4 , wherein the impingement plate wraps around a flat portion of the carrier.8. The blade outer air seal assembly of claim 4 , wherein a coating is on a surface between the carrier and the impingement plate.9. The blade outer air seal assembly of claim 4 , wherein the impingement plate is a metallic material.10. The blade outer air seal assembly of claim 4 , wherein the impingement plate is welded to the ...

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Номер записи: 97
21-01-2021 дата публикации

TURBOMACHINERY

Номер: US20210017875A1
Автор: FUJITA Yutaka, TAKASHIMA Reiko, Tomita Isao
Принадлежит: MITSUBISHI HEAVY INDUSTRIES ENGINE & TURBOCHARGER, LTD.

A turbomachinery according to an embodiment includes an impeller including at least one blade, and a casing for housing the impeller rotatably. A size of a gap between a tip of the blade and an inner surface of the casing during a stop of the impeller is formed non-uniformly over a circumferential direction of the impeller. 1. A turbomachinery comprising:an impeller including at least one blade; anda casing for housing the impeller rotatably,wherein a size of a gap between a tip of the blade and an inner surface of the casing during a stop of the impeller is formed non-uniformly over a circumferential direction of the impeller.2. The turbomachinery according to claim 1 ,wherein a difference between a maximum value and a minimum value of the gap during the stop of the impeller is not less than 10% of an average value of the gap in the circumferential direction.3. The turbomachinery according to claim 1 ,wherein the casing has an inner circumferential edge formed into an elliptical shape.4. The turbomachinery according to claim 1 ,wherein, during the stop of the impeller, a center axis of the casing is parallel to a rotational axis of the impeller and is displaced from the rotational axis of the impeller to a radial direction.5. The turbomachinery according to claim 1 ,wherein, during the stop of the impeller, a center axis of the casing is not parallel to a rotational axis of the impeller.6. The turbomachinery according to claim 1 ,wherein the impeller is a radial flow impeller, andwherein the casing is rotationally asymmetric about a center axis of the casing.7. The turbomachinery according to claim 6 , a scroll part internally including a scroll flow passage where a fluid flows in the circumferential direction on a radially outer side of the impeller; and', 'a tongue part for separating the scroll flow passage from a flow passage on a radially outer side of the scroll flow passage, and, 'wherein the casing includeswherein, regarding the gap during the stop of the ...

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Номер записи: 98
21-01-2021 дата публикации

SHROUD ASSEMBLY FOR CENTRIFUGAL COMPRESSOR AND METHOD

Номер: US20210017876A1
Автор: MENHEERE Dave
Принадлежит:

The centrifugal compressor can have a shroud engaged to a case via a plurality of circumferentially interspaced slots and lugs, the slots extending in at least one of a radial direction and an axial direction relative to a rotation axis of the compressor, the lugs slidingly received in a corresponding slot and configured for sliding in the slot in response to thermal growth of the case relative to the shroud. 1. A gas turbine centrifugal compressor comprising: a shroud engaged to a case via a plurality of circumferentially interspaced slots and lugs , the slots extending in at least one of a radial direction and an axial direction relative to a rotation axis of the compressor , the lugs slidingly received in a corresponding slot and configured for sliding in the slot in response to thermal growth of the case relative to the shroud.2. The shroud assembly of claim 1 , wherein at least some of the slots are defined in the case portion and at least some of the lugs are attached to the shroud claim 1 , and the dimension is a length extending in the radial direction.3. The shroud assembly of claim 2 , wherein the at least some of the slots are open radially toward the shroud.4. The shroud assembly of claim 3 , wherein the at least some of the lugs include at least one of a bolt connected to the shroud claim 3 , and a lug integral to the shroud.5. The shroud assembly of claim 4 , wherein the shroud includes a low pressure portion defining at least in part an inlet to the centrifugal compressor and a high pressure portion defining at least in part an outlet from the centrifugal compressor claim 4 , and the at least some of the lugs extend from the high pressure portion.6. The shroud assembly of claim 5 , wherein the shroud is dimensioned such that during at least part of a time during which the shroud assembly is in use claim 5 , each lug of the at least some of the lugs and a corresponding slot of the at least some of the slots define a gap between that lug and that slot ...

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Номер записи: 99
16-01-2020 дата публикации

FAN CASE ASSEMBLY FOR GAS TURBINE ENGINE

Номер: US20200018185A1
Автор: Costa Mark W.
Принадлежит:

A fan case assembly for a gas turbine engine includes a honeycomb structure. The fan case assembly also includes a septum operatively coupled to the honeycomb structure and located radially inward of the honeycomb structure. The fan case assembly further includes a rubstrip in contact with the septum and located radially inward of the septum. 1. A fan case assembly for a gas turbine engine comprising:a honeycomb structure;a septum operatively coupled to the honeycomb structure and located radially inward of the honeycomb structure; anda rubstrip in contact with the septum and located radially inward of the septum.2. The fan case assembly of claim 1 , wherein the rubstrip has a uniform circular inner surface for interaction with a rotatable fan blade located within the fan case assembly.3. The fan case assembly of claim 1 , wherein the rubstrip includes a first axial section claim 1 , a second axial section claim 1 , and a third axial section claim 1 , the first axial section and the third axial section having a radial thickness greater than a radial thickness of the second axial section claim 1 , the second axial section positioned for contact with an outer tip of a rotatable fan blade located within the fan case assembly.4. The fan case assembly of claim 3 , wherein the septum includes a bumper portion having a radial thickness greater than a radial thickness of other portions of the septum.5. The fan case assembly of claim 4 , wherein the bumper portion is axially aligned with the second axial section of the rubstrip.6. The fan case assembly of claim 4 , wherein the bumper portion defines at least one circumferentially extending slot.7. The fan case assembly of claim 6 , wherein the rubstrip includes at least one protrusion extending radially outward from the second axial section of the rubstrip claim 6 , the at least one protrusion filling the at least one circumferentially extending slot of the bumper portion.8. The fan case assembly of claim 4 , wherein axial ...

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Номер записи: 100