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Небесная энциклопедия

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

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Мониторинг СМИ

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

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Поддерживает ввод нескольких поисковых фраз (по одной на строку). При поиске обеспечивает поддержку морфологии русского и английского языка
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Применить Всего найдено 7111. Отображено 100.
19-01-2012 дата публикации

High pressure turbine vane cooling hole distrubution

Номер: US20120014809A1
Автор: Franco Di Paola, Francois Doyon, Larry Lebel
Принадлежит: Pratt and Whitney Canada Corp

A turbine vane for a gas turbine engine with an airfoil portion including a perimeter wall having first, second, third and fourth sets of cooling holes defined therethrough, including the holes numbered HA-1 to HA-13, HB-1 to HB-13, PA-1 to PA-9, and SA-1 to SA-3, respectively, and located such that a central axis thereof extends through the respective point 1 and point 2 having a nominal location in accordance with the X, Y Cartesian coordinate values set forth in Table 3.

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Номер записи: 1
29-03-2012 дата публикации

Cooled turbine blades for a gas-turbine engine

Номер: US20120076665A1
Автор: Erik Janke, Heinz-Peter SCHIFFER, Jens Taege, Knut Lehmann, Nils WINTER, Sebastian Meier, Thomas SCHIESSL
Принадлежит: Rolls Royce Deutschland Ltd and Co KG

The present invention relates to a cooled turbine blade for a gas-turbine engine having at least one cooling duct ( 14 ) extending radially, relative to a rotary axis of the gas-turbine engine, inside the airfoil and air-supply ducts ( 12 ) issuing into said cooling duct, characterized in that the cooling duct ( 14 ) extends into the blade root ( 6 ) in order to generate close to the wall a cooling airflow moved at high circumferential velocity and radially in helical form and that in the area of the blade root ( 6 ) at least one nozzle-shaped air-supply duct ( 12 ) issues into the cooling duct ( 14 ) tangentially or with a tangential velocity component.

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Номер записи: 2
05-04-2012 дата публикации

Apparatus and methods for cooling platform regions of turbine rotor blades

Номер: US20120082550A1
Автор: Adrian Lional Scott, John Wesley Harris, JR., Scott Edmond Ellis, Xiaoyong Fu
Принадлежит: General Electric Co

A platform cooling configuration in a turbine rotor blade that includes platform slot formed through at least one of the pressure side slashface and the suction side slashface; a removably-engaged impingement insert that separates the platform into two radially stacked plenums, a first plenum that resides inboard of a second plenum; a high-pressure connector that connects the first plenum to the high-pressure coolant region of the interior cooling passage; a low-pressure connector that connects the second plenum to the low-pressure coolant region of the interior cooling passage.

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Номер записи: 3
05-04-2012 дата публикации

Cooed IBR for a micro-turbine

Номер: US20120082563A1
Автор: Jack W. Wilson, Jr., James P. Downs, John E. Ryznic
Принадлежит: Florida Turbine Technologies Inc

A micro gas turbine engine in which the turbine rotor blades are formed as an integral bladed rotor with cooling air passages formed within the blades and the rotor disk by an EDM process. an adjacent stator vane includes an air riding seal with an air cushion supplied through the vanes to provide cooling, and where the air cushion is then passed into the turbine blades and rotor disk to provide cooling for the turbine blades. With cooling of the turbine blades, higher turbine inlet temperatures for micro gas turbine engines can be produced.

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Номер записи: 4
05-04-2012 дата публикации

Cooled rotor blade

Номер: US20120082567A1
Автор: Ian Tibbott, Oliver C.T. TIBBOTT
Принадлежит: Rolls Royce PLC

A cooled turbine rotor blade for a gas turbine engine is provided. The engine has an annular flow path for conducting working fluid though the engine. The blade has an aerofoil section for extending across the annular flow path. The blade further has a root portion radially inward of the aerofoil section for joining the blade to a rotor disc of the engine. The blade further has a platform between the aerofoil section and the root portion. The platform extends laterally relative to the radial direction of the engine to form an inner boundary of the annular flow path and to provide a rear overhang portion which projects in use towards a corresponding platform of a downstream nozzle guide vane. The platform contains at least one internal elongate plenum chamber for receiving cooling air. The longitudinal axis of the plenum chamber is substantially aligned with the circumferential direction of the engine. The plenum chamber supplies the cooling air to a plurality of exit holes formed in the external surface of the rear overhang portion to cool that portion.

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Номер записи: 5
12-04-2012 дата публикации

Curved film cooling holes for turbine airfoil and related method

Номер: US20120087803A1
Автор: Benjamin Lacy, Jesse Blair Butler
Принадлежит: General Electric Co

A turbine bucket includes an airfoil portion at one end thereof; a root portion at an opposite end thereof; a platform portion between the airfoil portion and the root portion; at least one internal cavity within or radially inward of the platform portion having at least one film cooling hole extending between the at least one cavity and an external surface of the platform portion. The film cooling hole is curved along a length dimension of the film cooling hole.

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Номер записи: 6
17-05-2012 дата публикации

Rotor and method for manufacturing a rotor for a turbo machine

Номер: US20120121384A1
Автор: Frank Stiehler, Hans Peter Borufka
Принадлежит: MTU AERO ENGINES GMBH

The invention relates to a rotor for a turbo machine, having rotating blades ( 12 ), which are joined to a basic rotor body ( 14 ), whereby a damping element ( 24 ) for damping blade vibrations is provided between blade platforms ( 10 ) of at least two adjacent rotating blades ( 12 ), damping element ( 24 ) being arched radially upward along its axial extent relative to an axis of rotation of the rotor. In addition, the invention relates to a method for manufacturing, repairing, and/or overhauling a rotor for a turbo machine, in which rotating blades ( 12 ) are joined to a basic rotor body ( 14 ), whereby a damping element ( 24 ) for damping blade vibrations is disposed between at least two adjacent rotating blades ( 12 ). In this case, damping element ( 24 ) is arched radially upward along its axial extent relative to an axis of rotation of the rotor.

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Номер записи: 7
17-05-2012 дата публикации

Rotor for a turbo machine

Номер: US20120121435A1
Автор: Frank Stiehler, Hans Peter Borufka
Принадлежит: MTU AERO ENGINES GMBH

The invention relates to a rotor ( 10 ) for a turbo machine, in particular for an aircraft turbine, with rotating blades ( 12 ) that are joined to a basic rotor body ( 16 ), whereby at least one channel ( 22 ) extending between the high-pressure side (HD) and the low-pressure side (ND) of rotor ( 10 ) radially underneath a blade platform ( 18 ) of at least one rotating blade ( 12 ) is provided, whereby a slope of a principal axis of extension (H) of channel ( 22 ) relative to an axis of rotation (D) of rotor ( 10 ) has the same sign as a slope of a principal axis of extension (R) of a radially inner boundary of the flow channel of rotor 10. In addition, the invention relates to a turbo machine having a rotor ( 10 ) as well as to a method for manufacturing a rotor ( 10 ) for a turbo machine.

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Номер записи: 8
17-05-2012 дата публикации

Rotor for a turbo machine

Номер: US20120121436A1
Автор: Frank Stiehler, Hans Peter Borufka
Принадлежит: MTU AERO ENGINES GMBH

The invention relates to a rotor ( 13 ) for a turbo machine, in particular for an aircraft turbine, having rotating blades ( 12 ), which are joined to a basic rotor body ( 14 ), in which at least one rotating blade ( 12 ) has at least one inner cooling channel, which extends at least along a predominant region of a blade element ( 18 ) of rotating blade ( 12 ) and has an inlet opening ( 23 ) for introducing cooling air into blade element ( 18 ), in which the inlet opening ( 23 ) is formed in a blade neck ( 16 ) lying between a blade foot ( 32 ) and a blade platform ( 10 ) of rotating blade ( 12 ). Inlet opening ( 23 ) is disposed in a cavity ( 22 ) formed radially underneath blade platform ( 10 ). In addition, the invention relates to a blade or vane, in particular a rotating blade ( 12 ) for a rotor ( 13 ) of a turbo machine, as well as a turbo machine.

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Номер записи: 9
07-06-2012 дата публикации

Airfoil with wrapped leading edge cooling passage

Номер: US20120141289A1
Автор: Justin D. Piggush, William Abdel-Messeh
Принадлежит: Individual

A turbine engine airfoil includes an airfoil structure having an exterior surface providing a leading edge. A radially extending first cooling passage is arranged near the leading edge and includes first and second portions. The first portion extends to the exterior surface and forms a radially extending trench in the leading edge. The second portion is in fluid communication with a second cooling passage. In one example, the second cooling passage extends radially, and the first cooling passage wraps around a portion of the second cooling passage from a pressure side to a suction side between the second cooling passage and the exterior surface. In the example, the first portion is arranged between the pressure and suction sides. In one example, the first cooling passage is formed by arranging a core in an airfoil mold. The trench is formed by the core in one example.

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Номер записи: 10
14-06-2012 дата публикации

Method of fabricating a component using a two-layer structural coating

Номер: US20120148769A1
Автор: Don Mark Lipkin, Ronald Scott Bunker
Принадлежит: General Electric Co

A method of fabricating a component is provided. The fabrication method includes depositing a first layer of a structural coating on an outer surface of a substrate. The substrate has at least one hollow interior space. The fabrication method further includes machining the substrate through the first layer of the structural coating, to define one or more openings in the first layer of the structural coating and to form respective one or more grooves in the outer surface of the substrate. Each groove has a respective base and extends at least partially along the surface of the substrate. The fabrication method further includes depositing a second layer of the structural coating over the first layer of the structural coating and over the groove(s), such that the groove(s) and the second layer of the structural coating together define one or more channels for cooling the component. A component is also disclosed.

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Номер записи: 11
05-07-2012 дата публикации

Apparatus and methods for cooling platform regions of turbine rotor blades

Номер: US20120171046A1
Автор: Bradley Taylor Boyer
Принадлежит: General Electric Co

A configuration of cooling channels through the interior of a turbine rotor blade having a platform, wherein the rotor blade includes an airfoil cooling channel that includes a cooling channel formed within the airfoil and an outboard airfoil supply channel. The configuration of cooling channels may include: a platform cooling channel that comprises a cooling channel that traverses at least a portion of the platform, the platform cooling channel having an upstream end and a downstream end; an outboard platform supply channel, which comprises a cooling channel that stretches from a second coolant inlet formed in the root to the upstream end of the platform cooling channel; and an inboard platform return channel, which comprises a cooling channel that stretches from the downstream end of the platform cooling channel to a termination point formed in the root.

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Номер записи: 12
16-08-2012 дата публикации

Cooling system having reduced mass pin fins for components in a gas turbine engine

Номер: US20120207591A1
Автор: Ching-Pang Lee, John J. Marra, Nan Jiang
Принадлежит: Siemens Energy Inc

A cooling system having one or more pin fins with reduced mass for a gas turbine engine is disclosed. The cooling system may include one or more first surfaces defining at least a portion of the cooling system. The pin fin may extend from the surface defining the cooling system and may have a noncircular cross-section taken generally parallel to the surface and at least part of an outer surface of the cross-section forms at least a quartercircle. A downstream side of the pin fin may have a cavity to reduce mass, thereby creating a more efficient turbine airfoil.

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Номер записи: 13
04-10-2012 дата публикации

Turbine Blade Platform Undercut

Номер: US20120251331A1
Автор: Douglas James Dietrich, Gregory Edwin Vogel, Stephen Wayne Fiebiger
Принадлежит: Alstom Technology AG

A system and method of extending the useable life of a gas turbine blade is disclosed in which the gas turbine blade includes an undercut configuration designed to relieve mechanical and thermal stress imparted into the pedestal region of the airfoil trailing edge. The embodiments of the present invention include turbine blade configurations having different trailing edge undercut configurations as well as additional cooling supplied to the internal passages of the trailing edge region of the turbine blade.

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Номер записи: 14
08-11-2012 дата публикации

Blade, Integrally Bladed Rotor Base Body and Turbomachine

Номер: US20120282109A1
Автор: Frank Stiehler
Принадлежит: MTU AERO ENGINES GMBH

A blade for a turbomachine is disclosed. At least one cooling channel is configured in the blade neck, whose inlet is disposed near a platform projection on the high-pressure side and whose outlet is disposed in the region of a platform projection on the low-pressure side. An integrally bladed rotor base body having a plurality of these types of blades as well as turbomachine with such a rotor base body is also disclosed.

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Номер записи: 15
31-01-2013 дата публикации

Platform interconnected with mid-body core interface for molding airfoil platforms

Номер: US20130025812A1
Автор: Steven Taffet, Tracy A. Propheter-Hinckley
Принадлежит: Individual

A method of molding a platform opening includes the steps of providing a main body core and a platform core, with the main body core having a portion that forms a portion of the platform. The platform core has at least one side portion that will form a side opening. Molten metal is directed around the cores within a mold and solidifies. The cores are removed, leaving cavities where the cores were within the molten metal, and includes an opening in a side face formed by the side portion of the platform body core. Lost core components are also disclosed and claimed.

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Номер записи: 16
21-02-2013 дата публикации

Angled trench diffuser

Номер: US20130045106A1
Автор: Benjamin Paul Lacy
Принадлежит: General Electric Co

An article is disclosed that includes a substrate having a first surface and a second surface and a coating disposed on the second surface. In addition, the article includes an angled trench at least partially defined in the coating. The angled trench may include a bottom surface, a first sidewall and a second sidewall disposed downstream of the first sidewall. The first and second sidewalls may extend from the bottom surface at an angle of less than about 60 degrees. Moreover, the article may include a plurality of holes defined between the first surface and the bottom surface.

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Номер записи: 17
21-02-2013 дата публикации

Turbine blade cooling system with bifurcated mid-chord cooling chamber

Номер: US20130045111A1
Автор: Ching-Pang Lee
Принадлежит: Siemens Energy Inc

A cooling system for a turbine blade of a turbine engine having a bifurcated mid-chord cooling chamber for reducing the temperature of the blade. The bifurcated mid-chord cooling chamber may be formed from a pressure side serpentine cooling channel and a suction side serpentine cooling channel with cooling fluids passing through the pressure side serpentine cooling channel in a direction from the trailing edge toward the leading edge and in an opposite direction through the suction side serpentine cooling channel. The pressure side and suction side serpentine cooling channels may flow counter to each other, thereby yielding a more uniform temperature distribution than conventional serpentine cooling channels.

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Номер записи: 18
28-02-2013 дата публикации

Pin-fin array

Номер: US20130052036A1
Автор: Aaron Ezekiel Smith
Принадлежит: General Electric Co

The present application provides an airfoil with a cooling flow therein. The airfoil may include an internal cooling passage, a number of cooling holes in communication with the internal cooling passage, and a number of pin-fins positioned within the internal cooling passage. The pin-fins are arranged with one or more turning openings and one or more guiding openings so as to direct the cooling flow towards the cooling holes.

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Номер записи: 19
21-03-2013 дата публикации

TURBINE BLADES AND METHODS OF FORMING MODIFIED TURBINE BLADES AND TURBINE ROTORS

Номер: US20130071254A1
Автор: Halfmann Steve, Mitlin Bob, Morris Mark C., Riahi Ardeshir
Принадлежит: HONEYWELL INTERNATIONAL INC.

Turbine blades and methods of forming modified turbine blades and turbine rotors for use in an engine are provided. In an embodiment, by way of example only, a turbine blade includes a platform and an airfoil. The platform includes a surface configured to define a portion of a flowpath, and the surface includes an initial contour configured to plastically deform into an intended final contour after an initial exposure of the blade to an operation of the engine. The airfoil extends from the platform. 1. A turbine blade for use in an engine , the turbine blade comprising:a platform including a surface configured to define a portion of a flowpath, the surface including an initial contour configured to plastically deform into an intended final contour after an initial exposure of the blade to an operation of the engine; andan airfoil extending from the platform.2. The turbine blade of claim 1 , wherein the operation of the engine includes a predetermined maximum speed and a predetermined maximum temperature.3. The turbine blade of claim 1 , wherein:the airfoil has an airfoil shape defined by a convex suction side wall, a concave pressure side wall, a leading edge, a trailing edge, a root and a tip, the convex suction side wall, the concave pressure side wall, and the tip each including an interior surface that defines an interior with the root, the trailing edge including a plurality of slots formed thereon.4. The turbine blade of claim 1 , wherein the initial contour is a depression.5. The turbine blade of claim 4 , wherein the depression is located adjacent to the concave pressure side wall.6. The turbine blade of claim 4 , wherein the depression is located adjacent to the convex suction side wall. The present application is a divisional application of U.S. Pat. No. 8,297,935 issued Oct. 30, 2012 (application Ser. No. 12/273,108, filed Nov. 18, 2008), the contents of which are herein incorporated by reference in their entirety.This invention was made with Government ...

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Номер записи: 20
02-05-2013 дата публикации

Turbine blade and engine component

Номер: US20130108471A1
Автор: Shu Fujimoto
Принадлежит: IHI Corp

A turbine blade that is used in a turbine of a gas turbine engine and cooled by cooling air, and includes a cooling channel that is formed within the turbine blade and in which the cooling air flows, plural bottomed recesses that are formed on a blade surface of the turbine blade and of which each downstream-side inner wall is inclined, and an ejection hole that is formed on each bottom of the plural bottomed recesses and communicates with the cooling channel to eject the cooling air. The ejection hole is formed so that a central line of the ejection hole extends along the downstream-side inner wall. The above turbine blade can improve cooling without reducing efficiency.

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Номер записи: 21
06-06-2013 дата публикации

HYBRID VAPOR AND FILM COOLED TURBINE BLADE

Номер: US20130142665A1
Автор: Hill James D., Norris James W.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A cooling system for cooling a fluid reaction apparatus of a gas turbine engine includes a vapor cooling subsystem and a film cooling subsystem. The vapor cooling subsystem has a vaporization section and a condenser section for cooling a portion of the fluid reaction apparatus. The condenser section is cooled by a fluid. The film cooling subsystem is configured for cooling a portion of the fluid reaction apparatus by discharging fluid out of openings defined in the fluid reaction apparatus. At least a portion of the fluid used to cool the condenser section of the vapor cooling subsystem is discharged out of the openings of the film cooling subsystem. 1. A cooling system for cooling a fluid reaction apparatus of a gas turbine engine , the system comprising:a vapor cooling subsystem having a vaporization section and a condenser section for cooling a portion of the fluid reaction apparatus, wherein the condenser section is cooled by a fluid;a film cooling subsystem for cooling a portion of the fluid reaction apparatus by discharging fluid out of openings defined in the fluid reaction apparatus, wherein at least a portion of the fluid used to cool the condenser section of the vapor cooling subsystem is discharged out of the openings of the film cooling subsystem; anda flow deflector located at or near a downstream portion of the condenser section for directing the fluid used to cool the condenser section to the film cooling subsystem.23-. (canceled)4. The system of claim 1 , wherein the fluid reaction apparatus is a turbine blade.5. The system of claim 4 , wherein the vapor cooling subsystem provides cooling to a leading edge portion of the turbine blade.6. The system of claim 4 , wherein the turbine blade includes an airfoil and a root claim 4 , and wherein the vaporization section of the vapor cooling subsystem is defined within the airfoil and the condenser section of the vapor cooling subsystem is defined within the root.7. The system of claim 6 , wherein the flow ...

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Номер записи: 22
06-06-2013 дата публикации

Turbine blade incorporating trailing edge cooling design

Номер: US20130142666A1
Автор: Benjamin E. Heneveld, Ching-Pang Lee, Glenn E. Brown
Принадлежит: Mikro Systems Inc, Siemens Energy Inc

A turbine blade ( 10 ) including an airfoil ( 12 ) having multiple interior wall portions ( 70 ) each separating at least one chamber from another one of multiple chambers ( 46, 48, 50, 58, 60 ). In one embodiment a first wall portion ( 70 - 2 ) between first and second chambers ( 60, 52 ) includes first and second pluralities of flow paths ( 86 P, 86 S) extending through the first wall portion. The first wall portion includes a first region R 1 having a first thickness, t, measurable as a distance between the chambers. One of the paths extends a first path distance, d, as measured from an associated path opening ( 78 ) in the first chamber ( 60 ), through the first region and to an exit opening ( 82 ) in the second chamber ( 52 ) which path distance is greater than the first thickness.

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Номер записи: 23
06-06-2013 дата публикации

Cooled turbine blade for gas turbine engine

Номер: US20130142668A1
Автор: Emmanuel Berche, Jean-Luc Bacha, Tangi Brusq
Принадлежит: SNECMA SAS

A cooled turbine blade for a gas turbine engine including a pressure surface wall, a suction surface wall and a distal wall connecting the pressure surface wall and the suction surface wall, arranged so as to create in the region of the distal end of the blade at least one external cavity forming a bathtub-shaped cavity and at least one internal cavity separated by the distal wall, the blade having at least one opening for the introduction of a flow of cooling air into the external cavity, wherein, on the one hand, at least one part of the distal wall is inclined relative to the verticals of the pressure surface wall and, on the other hand, the opening is created in the vicinity of the distal wall so that the flow of cooling air is directed towards the distal end of the pressure surface wall.

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Номер записи: 24
20-06-2013 дата публикации

AMBIENT AIR COOLING ARRANGEMENT HAVING A PRE-SWIRLER FOR GAS TURBINE ENGINE BLADE COOLING

Номер: US20130156579A1
Автор: JR. Samuel R., Lee Ching-Pang, Marra John J., Meeroff Jamie, MILLER, Schroeder Eric, Tham Kok-Mun
Принадлежит:

A gas turbine engine including: an ambient-air cooling circuit () having a cooling channel () disposed in a turbine blade () and in fluid communication with a source () of ambient air: and an pre-swirler (), the pre-swirler having: an inner shroud (); an outer shroud (); and a plurality of guide vanes (), each spanning from the inner shroud to the outer shroud. Circumferentially adjacent guide vanes () define respective nozzles () there between. Forces created by a rotation of the turbine blade motivate ambient air through the cooling circuit. The pre-swirler is configured to impart swirl to ambient air drawn through the nozzles and to direct the swirled ambient air toward a base of the turbine blade. The end walls () of the pre-swirler may be contoured. 1. A gas turbine engine , comprising:an ambient-air cooling circuit comprising a cooling channel disposed in a turbine blade and in fluid communication with a source of ambient air that provides cooling fluid: and an inner shroud;', 'an outer shroud; and', 'a plurality of guide vanes, each spanning from the inner shroud to the outer shroud,', 'wherein circumferentially adjacent guide vanes define respective nozzles there between, the nozzles defining a portion of the cooling circuit, each nozzle defined by a pressure side of a first guide vane, a suction side of the adjacent guide vane, an outer end wall defined by the outer shroud, and an inner end wall defined by the inner shroud;, 'a pre-swirler, comprisingwherein forces created by a rotation of the turbine blade motivate the cooling fluid through the cooling circuit; andwherein the pre-swirler is configured to impart swirl to the cooling fluid drawn through the nozzles and to direct the swirled cooling fluid toward a base of the turbine blade.2. The gas turbine engine of claim 1 , wherein the inner shroud is formed as a monolithic body.3. The gas turbine engine of claim 1 , wherein the outer shroud is formed as a monolithic body.4. The gas turbine engine of ...

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Номер записи: 25
11-07-2013 дата публикации

Double-jet type film cooling structure

Номер: US20130175015A1
Автор: Anas Elyas, Dieter Bohn, Karsten Kusterer, Masahide Kazari, Ryozo Tanaka, Takao Sugimoto
Принадлежит: B&B AGEMA GmbH, Kawasaki Jukogyo KK

Provided is a film cooling structure capable of suppressing a cooling medium film from being separated from a wall surface, to increase a film efficiency on the wall surface and thereby-cool the wall surface effectively. One or more pairs of injection holes are formed on a wall surface facing a passage of high-temperature gas to inject a cooling medium to the passage. A single supply passage is formed inside the wall to supply the cooling medium to the injection holes. A separating section is provided between the injection holes in a location forward relative to rear ends of the injection holes to separate the cooling medium into components flowing to the injection holes. An injection direction of the cooling medium is inclined relative to a gas flow direction so that the cooling medium forms swirl flows that push the cooling medium against the wall surface.

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Номер записи: 26
11-07-2013 дата публикации

Impingement Cooling System for Use with Contoured Surfaces

Номер: US20130177396A1
Автор: Aaron Gregory Winn
Принадлежит: General Electric Co

The present application provides an impingement cooling system for use with a contoured surface. The impingement cooling system may include an impingement plenum and an impingement plate with a linear shape facing the contoured surface. The impingement surface may include a number of projected area thereon with a number of impingement holes having varying sizes and varying spacings.

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Номер записи: 27
15-08-2013 дата публикации

GAS TURBINE BLADE, MANUFACTURING METHOD THEREFOR, AND GAS TURBINE USING TURBINE BLADE

Номер: US20130209271A1
Автор: Asano Hiroshi, HADA Satoshi, HIRATA Norifumi, Imada Takahiko, Ooe Katsutoshi, SHINTANI Tomofumi
Принадлежит: MITSUBISHI HEAVY INDUSTRIES, LTD.

Gas turbine blades which simplify the formation of cooling channels provided inside the turbine blades while simultaneously avoiding loss of turbine blade strength and rigidity due to forming of the cooling channels. Cooling channels provided in the interior of a gas turbine blade include a plurality of straight channel-like base-side elongated holes extending in a longitudinal direction at a base side of the turbine blade, a plurality of straight channel-like tip-side elongated holes extending in a longitudinal direction at a tip side of the turbine blade, and a plurality of communicating hollow portions interposed at connection portions between the two types of elongated holes to allow the two types of elongated holes to communicate with each other, and have larger cross-sectional areas than the channel cross-sectional areas of both elongated holes. The communicating hollow portions are formed to match the position of a platform portion of the turbine blade. 1. A gas turbine blade in which cooling channels are formed inside the turbine blade , and the turbine blade is cooled by causing cooling air to circulate through the cooling channels , whereinthe cooling channels comprise:a plurality of straight channel-like base-side elongated holes that extend in a longitudinal direction at a base side of the turbine blade,a plurality of straight channel-like tip-side elongated holes that extend in a longitudinal direction at a tip side of the turbine blade, anda plurality of communicating hollow portions each having a spherical or spheroidal shape and interposed at respective connection portions that connect the base-side elongated holes and respective corresponding tip-side elongated holes to individually allow the two types of elongated holes to communicate with each other, and the respective communicating hollow portions having larger cross-sectional areas than the channel cross-sectional areas of the two types of elongated holes.2. A gas turbine blade according to ...

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Номер записи: 28
22-08-2013 дата публикации

GAS TURBINE BLADE

Номер: US20130216395A1
Автор: Bregman Vitaly, Petukhovskiy Mikhail
Принадлежит:

A gas turbine blade including a root, an airfoil with a leading edge and a trailing edge, a radial outer tip, and a pressure side and a suction side between the leading edge and the trailing edge, and a cooling air channel system extending from an air inlet opening in the root throughout the airfoil to a plurality of air outlets at the pressure side and the leading edge of the top of the tip of the airfoil. 113-. (canceled)14. A gas turbine blade , comprising:a root;an airfoil with a leading edge and a trailing edge, a radial outer tip, and a pressure side and a suction side between the leading edge and the trailing edge; anda cooling air channel system extending from an air inlet opening in the root throughout the airfoil to a plurality of air outlets at the pressure side and the leading edge of the top of the tip of the airfoil,wherein a number of air outlets per area near the leading edge of the tip is higher than the average number of air outlets per area in the top of the tip,wherein a concentration of air outlets at the top of the tip of the airfoil is higher on the pressure side than on the suction side, andwherein the plurality of air outlets closest to the trailing edge are larger in air cross section than the plurality of air outlets in the middle between the leading edge and the trailing edge.15. The gas turbine blade according to claim 14 , wherein the plurality of air outlets at the leading edge form a group of air outlets arranged at the leading edge of the tip.16. The gas turbine blade according to claim 15 , wherein the shortest distance between the group and an air outlet on the pressure side closest to the group is larger than the diameter of the group.17. The gas turbine blade according to claim 14 , wherein the plurality of air outlets at the pressure side of the top of the tip are arranged in a row completely inside a rib at the pressure side of the tip leaving the thickness of the rib untouched.18. The gas turbine blade according to claim 14 , ...

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Номер записи: 29
12-09-2013 дата публикации

Airfoil with improved internal cooling channel pedestals

Номер: US20130232991A1
Автор: Edwin Otero
Принадлежит: United Technologies Corp

An airfoil for a turbine engine, the airfoil including a first side wall, a second side wall spaced apart from the first side wall, and an internal cooling channel formed between the first side wall and the second side wall. The internal cooling channel includes at least one pedestal having a first pedestal end connected to the first side wall and a second pedestal end connected to the second side wall. The internal cooling channel also includes a first fillet disposed around the periphery of the first pedestal end between the first side wall and the first pedestal end; and a second fillet disposed around the periphery of the second pedestal end between the second side wall and the second pedestal end. At least one of the first fillet and the second fillet includes a profile that is non-uniform around the periphery of the corresponding pedestal end.

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Номер записи: 30
03-10-2013 дата публикации

TURBINE AIRFOIL TRAILING EDGE COOLING SLOTS

Номер: US20130259645A1
Автор: Bergholz, Durstock Daniel Lee, JR. Robert Frederick
Принадлежит:

A turbine airfoil includes pressure and suction sidewalls extending along a span from a base to a tip. Spanwise spaced apart trailing edge cooling holes in the pressure sidewall end at corresponding spanwise spaced apart trailing edge cooling slots extending chordally substantially to the trailing edge. Each cooling hole includes in downstream serial cooling flow relationship, a curved inlet, a constant area and constant width metering section, and a spanwise diverging section leading into the trailing edge cooling slot, and a spanwise height substantially greater than a hole width through the cooling hole. A pressure sidewall surface of the pressure sidewall may be planar through the metering and diverging sections. The width may be constant through the metering and diverging sections. A raised floor may include a flat up ramp in the diverging section, a flat down ramp in the slot. 1. A gas turbine engine turbine airfoil comprising:widthwise spaced apart pressure and suction sidewalls extending outwardly along a span from an airfoil base to an airfoil tip;the pressure and suction sidewalls extending chordwise between opposite leading and trailing edges;a spanwise row of spanwise spaced apart trailing edge cooling holes encased in the pressure sidewall and ending at corresponding spanwise spaced apart trailing edge cooling slots extending chordally substantially to the trailing edge;the cooling hole including in downstream serial cooling flow relationship, a curved inlet, a metering section with a constant area and constant width flow cross section, and a spanwise diverging section leading into the trailing edge cooling slot; anda spanwise height substantially greater than a hole width through the cooling hole.2. The airfoil as claimed in further comprising pressure and suction sidewall surfaces of the pressure and suction sidewalls respectively in the hole and the pressure sidewall surface being planar through the entire metering and diverging sections.3. The ...

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Номер записи: 31
03-10-2013 дата публикации

Turbine nozzle

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

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

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Номер записи: 32
10-10-2013 дата публикации

Surface analysis for detecting closed holes and method for reopening

Номер: US20130268107A1
Автор: Andreas Oppert, Dimitrios Thomaidis, Georg Bostanjoglo, Torsten Melzer-Jokisch
Принадлежит: SIEMENS AG

By means of laser triangulation measurements of an uncoated and a coated component having holes, the exact positions of the holes can be detected for reopening after coating. A method for the surface analysis of at least partially closed holes which are to be opened is provided. A process for reopening coated holes is also provided.

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Номер записи: 33
17-10-2013 дата публикации

Components with microchannel cooling

Номер: US20130272850A1
Автор: Ronald Scott Bunker
Принадлежит: General Electric Co

A component includes a substrate having outer and inner surfaces, where the inner surface defines at least one hollow, interior space. The outer surface defines pressure side and suction side walls, which are joined together at leading and trailing edges of the component. The outer surface defines one or more grooves that extend at least partially along the pressure or suction side walls in a vicinity of the trailing edge. Each groove is in fluid communication with a respective hollow, interior space. The component further includes a coating disposed over at least a portion of the outer substrate surface. The coating comprises at least a structural coating that extends over the groove(s), such that the groove(s) and the structural coating together define one or more channels for cooling the trailing edge. A method of forming cooling channels in the vicinity of the trailing edge is also provided.

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Номер записи: 34
17-10-2013 дата публикации

Axially-split radial turbines and methods for the manufacture thereof

Номер: US20130272882A1
Автор: Amandine Miner, David K. Jan, Don Mittendorf, Jason Smoke
Принадлежит: Honeywell International Inc

Embodiments of an axially-split radial turbine, as are embodiments of a method for manufacturing an axially-split radial turbine. In one embodiment, the method includes the steps of joining a forward bladed ring to a forward disk to produce a forward turbine rotor, fabricating an aft turbine rotor, and disposing the forward turbine rotor and the aft turbine rotor in an axially-abutting, rotationally-fixed relationship to produce the axially-split radial turbine.

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Номер записи: 35
28-11-2013 дата публикации

COMPONENTS WITH COOLING CHANNELS AND METHODS OF MANUFACTURE

Номер: US20130316100A1
Автор: BUNKER Ronald Scott
Принадлежит: GENERAL ELECTRIC COMPANY

A method of manufacturing a component is provided. The method includes forming one or more grooves in an outer surface of a substrate. Each groove extends at least partially along the surface of the substrate and has a base, a top and at least one discharge point. The method further includes forming a run-out region adjacent to the discharge point for each groove and disposing a coating over at least a portion of the surface of the substrate. The groove(s) and the coating define one or more channels for cooling the component. Components with cooling channels are also provided. 1. A component comprising:a substrate comprising an outer surface and an inner surface, wherein the outer surface defines one or more grooves and one or more run-out regions, wherein each groove extends at least partially along the outer surface of the substrate and has a base and at least one discharge point, and wherein each run-out region is adjacent to the respective discharge point for a respective groove; anda coating disposed over at least a portion of the outer surface of the substrate, such that the one or more grooves and the coating together define one or more channels for cooling the component.2. The component of claim 1 , wherein the inner surface defines at least one hollow claim 1 , interior space claim 1 , and wherein one or more access holes extend through the base of a respective one of the one or more grooves to place the groove in fluid communication with respective ones of the at least one hollow interior space.3. The component of claim 1 , wherein the run-out region is wider than a top of the respective groove.4. The component of claim 3 , wherein the coating does not bridge the one or more run-out regions claim 3 , such that each run-out region forms a film hole for the respective groove.5. The component of claim 1 , wherein the base of each groove is wider than the top claim 1 , such that each groove comprises a re-entrant shaped groove.6. The component of claim 5 , ...

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Номер записи: 36
12-12-2013 дата публикации

Cmc blade with pressurized internal cavity for erosion control

Номер: US20130330189A1
Автор: Andres Jose Garcia-Crespo, Jerome Walter Goike
Принадлежит: Individual

A ceramic matrix composite blade for use in a gas turbine engine having an airfoil with leading and trailing edges and pressure and suction side surfaces, a blade shank secured to the lower end of each airfoil, one or more interior fluid cavities within the airfoil having inlet flow passages at the lower end which are in fluid communication with the blade shank, one or more passageways in the blade shank corresponding to each one of the interior fluid cavities and a fluid pump (or compressor) that provides pressurized fluid (nominally cool, dry air) to each one of the interior fluid cavities in each airfoil. The fluid (e.g., air) is sufficient in pressure and volume to maintain a minimum fluid flow to each of the interior fluid cavities in the event of a breach due to foreign object damage.

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Номер записи: 37
26-12-2013 дата публикации

COOLED COMPONENT FOR THE TURBINE OF A GAS TURBINE ENGINE

Номер: US20130343872A1
Автор: JACKSON DOUGAL R., TIBBOTT Ian
Принадлежит: ROLLS-ROYCE PLC

A component for the turbine of a gas turbine engine is provided. The component two facing walls interconnected by one or more generally elongate divider members to partially define side-by-side, generally elongate, cooling fluid passage portions which form a multi-pass cooling passage within the component. The passage portions are connected in series fluid flow relationship by respective bends formed by joined ends of neighbouring of the passage portions. The component further includes one or more core tie linking passages formed in the divider members. One or more differential pressure reducing arrangements are formed in the multi-pass cooling passage adjacent respective of the core tie linking passages. 1. A component for the turbine of a gas turbine engine , the component including:two facing walls interconnected by one or more generally elongate divider members to partially define side-by-side, generally elongate, cooling fluid passage portions which form a multi-pass cooling passage within the component, the passage portions being connected in series fluid flow relationship by respective bends formed by joined ends of neighbouring of the passage portions, andone or more core tie linking passages formed in the divider members, the or each core tie linking passage having an entrance at an upstream passage portion and an exit at a neighbouring downstream passage portion, and allowing cooling fluid to leak therethrough to bypass the bend formed by the joined ends of the neighbouring passage portions,wherein one or more differential pressure reducing arrangements are formed in the multi-pass cooling passage adjacent respective of the core tie linking passages, the or each arrangement reducing the difference in the static pressure of the cooling fluid between the entrance of the respective core tie linking passage and the exit of that core tie linking passage.2. A component according to claim 1 , wherein the or each differential pressure reducing arrangement includes ...

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Номер записи: 38
26-12-2013 дата публикации

ROTARY DEGASSERS AND COMPONENTS THEREFOR

Номер: US20130343904A1
Автор: Cooper Paul V.
Принадлежит:

Disclosed are degassers, couplings, impeller shafts and impellers for use in molten metal. One such coupling transfers gas into an impeller shaft, the coupling having a smooth, tapered internal surface to align with a corresponding surface on the impeller shaft and help prevent gas leakage and to assist in preventing damage to the impeller shaft. Improved impellers for shearing and mixing gas are also disclosed, as is a degasser including one or more of these components. 1. An impeller having a top surface and a bottom surface , the impeller for dispersing gas into molten metal and comprising:an opening in the bottom surface through which gas can be released; andat least one open channel in communication with the opening, the channel for directing the gas; and at least one cavity, the channel directing gas to the cavity, the cavity being defined by a curved side surface and a top surface and the channel extending from the center of the impeller to the center of the curved side surface of the cavity.2. The impeller of further comprising a plurality of cavities and a plurality of channels claim 1 , wherein each of the plurality of channels leads to one of the cavities.3. The impeller of wherein the cavity is defined by a curved side surface and a top surface and the channel extends from the center of the impeller to the center of the curved side surface.4. The impeller of comprising four channels claim 1 , wherein each channel extends from a center of the impeller to a cavity.5. The impeller of that further includes a shearing structure juxtaposed the at least one cavity.6. The impeller of that is comprised of graphite.7. The impeller of that has a top surface claim 1 , a bottom surface and a plurality of cavities.8. The impeller of claim 7 , wherein the top surface has an outer perimeter and at least part of each of each curved side surface is inside the outer perimeter of the top surface.9. The impeller of that has four channels and each channel leads to the center ...

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Номер записи: 39
02-01-2014 дата публикации

GAS TURBINE ENGINE COMPONENT HAVING PLATFORM COOLING CHANNEL

Номер: US20140003961A1
Автор: Deibel Russell, Gleiner Matthew S., Willey Lawrence J.
Принадлежит:

A component for a gas turbine engine according to an exemplary embodiment of the present disclosure can include a platform having an outer surface and an inner surface. A cover plate can be positioned adjacent to the outer surface of the platform. The outer surface of the platform can include a pocket and the cover plate is positioned relative to the pocket to establish a platform cooling channel therebetween. 1. A component for a gas turbine engine , comprising:a platform having an outer surface and an inner surface; anda cover plate positioned adjacent to said outer surface of said platform, wherein said outer surface of said platform includes a pocket and said cover plate is positioned relative to said pocket to establish a platform cooling channel therebetween.2. The component as recited in claim 1 , wherein said platform is an inner diameter platform.3. The component as recited in claim 1 , wherein the component is a turbine vane.4. The component as recited in claim 1 , wherein at least a portion of said pocket is exposed to establish said platform cooling channel.5. The component as recited in claim 4 , wherein said portion of said pocket is a side opening of said pocket that faces a mate face of said platform.6. The component as recited in claim 1 , wherein said pocket is a cast feature of said platform.7. The component as recited in claim 1 , wherein said platform cooling channel is bound by said cover plate and said pocket on all but a single side.8. The component as recited in claim 1 , wherein said platform cooling channel extends adjacent to a pressure side of an airfoil that extends from the platform.9. The component as recited in claim 1 , wherein a pocket wall extends between said pocket and a slot of a mate face of said platform.10. The component as recited in claim 1 , wherein said pocket is enclosed by said cover plate to establish said platform cooling channel.11. The component as recited in claim 1 , wherein said platform cooling channel is a ...

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Номер записи: 40
02-01-2014 дата публикации

TURBINE BLADE

Номер: US20140003962A1
Автор: NAKAMATA Chiyuki, OKITA Yoji
Принадлежит:

A turbine blade () includes a plurality of rows of cooling portions (A, B) that have notch portions () that discharge cooling gas that has been introduced into an internal portion () of a blade portion () onto a ventral side blade surface (), and that are formed in rows that are stacked in a direction between the blade front edge and the blade rear edge. This turbine blade () is provided with: turbulence promoting cooling portions (B) that are provided in the row located furthest to the downstream side from among the plurality of rows, and that have turbulence promoting devices () in areas exposed by the notch portions (); and film cooling portions (A) that are provided in at least one of other rows from among the plurality of rows, and that form a film cooling layer in the areas exposed by the notch portions (). 1. A turbine blade that is provided with a plurality of rows of cooling portions that have notch portions that discharge cooling gas that has been introduced into an internal portion of a blade portion onto a ventral side blade surface , and that are formed in rows that are stacked in a direction between the blade front edge and the blade rear edge , comprising:turbulence promoting cooling portions that are provided in the row located closest to the downstream side from among the plurality of rows, and that have turbulence promoting devices in areas exposed by the notch portions; andfilm cooling portions that are provided in at least one of other rows from among the plurality of rows, and that form a film cooling layer in the areas exposed by the notch portions.2. The turbine blade according to claim 1 , wherein claim 1 , in at least one of the plurality of rows claim 1 , the plurality of cooling portions are placed discretely from each other in the height direction of the blade portion.3. The turbine blade according to claim 1 , wherein claim 1 , in at least one of the plurality of rows claim 1 , the cooling portions are provided continuously in an ...

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Номер записи: 41
09-01-2014 дата публикации

AIRFOIL COOLING ARRANGEMENT

Номер: US20140010632A1
Автор: Bergman Russell J., Spangler Brandon W.
Принадлежит:

An airfoil according to an exemplary embodiment of the present disclosure can include an airfoil body having a plurality of film cooling holes that extend through an exterior surface of the airfoil body. Each of the plurality of film cooling holes break through the exterior surface at geometric coordinates in accordance with Cartesian coordinate values of X, Y and Z as set forth in Table 1. Each of the geometric coordinates is measured from a reference point on a leading edge rail of a platform of the airfoil. 1. An airfoil , comprising:an airfoil body having a plurality of film cooling holes that extend through an exterior surface of said airfoil body, wherein each of said plurality of film cooling holes break through said exterior surface at geometric coordinates in accordance with Cartesian coordinate values of X, Y and Z as set forth in Table 1, wherein each of said geometric coordinates is measured from a reference point on a leading edge rail of a platform of the airfoil.2. The airfoil as recited in claim 1 , wherein said airfoil is a first stage turbine vane.3. The airfoil as recited in claim 1 , wherein said Cartesian coordinate values of Table 1 are expressed in inches.4. The airfoil as recited in claim 1 , wherein said reference point includes a pin hole of said platform.5. The airfoil as recited in claim 1 , wherein said plurality of film cooling holes are spaced along a span of said airfoil body in multiple collinearly aligned rows.6. The airfoil as recited in claim 1 , wherein a first portion of said plurality of film cooling holes are cone shaped and a second portion of said plurality of film cooling holes are round shaped.7. The airfoil as recited in claim 1 , wherein at least a portion of said plurality of film cooling holes are arranged in a herringbone configuration.8. The airfoil as recited in claim 1 , wherein said plurality of film cooling holes are disposed on a pressure side claim 1 , a suction side and a leading edge of said airfoil body.9. ...

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Номер записи: 42
23-01-2014 дата публикации

AIRFOIL ASSEMBLY INCLUDING VORTEX REDUCING AT AN AIRFOIL LEADING EDGE

Номер: US20140023483A1
Автор: Montgomery Matthew D., Smith Bruce L., Wiebe David J.
Принадлежит:

An airfoil assembly including an endwall and an airfoil extending from the into a gas flow path. The endwall includes upstream and downstream edges, and is defined on a platform structure having a front surface extending radially in a direction of a thickness of the platform structure. At least one fluid injection passage extends through the platform structure in a direction from the upstream edge toward the downstream edge of the endwall. The fluid injection passage has an outlet opening defined at the endwall and an inlet opening in fluid communication with a pressurized fluid source. The fluid injection passage extends at a shallow angle relative to a plane of the endwall wherein the fluid injection passage defines a passage axis passing through the front surface and the endwall for effecting energization of a boundary layer between the outlet opening and the airfoil leading edge. 1. An airfoil assembly for an axial flow gas turbine engine , the gas turbine engine including an axially directed flow path defining a passage for a working fluid and a source of pressurized fluid , the airfoil assembly including:an endwall having an upstream edge and a downstream edge axially spaced from the upstream edge;an airfoil extending from the endwall into the flow path, the airfoil having a leading edge and a trailing edge, and a pressure side and a suction side extending between the leading and trailing edges;the endwall is defined on a platform structure having a front surface adjacent to the upstream edge, the front surface extending radially in a direction of a thickness of the platform structure;at least one fluid injection passage extending through the platform structure in a direction from the upstream edge toward the downstream edge, and the fluid injection passage having an outlet opening defined at the endwall and an inlet opening in fluid communication with a pressurized fluid source; andthe fluid injection passage extending at a shallow angle relative to a plane ...

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Номер записи: 43
06-02-2014 дата публикации

TURBINE VANE

Номер: US20140037429A1
Автор: OKITA Yoji
Принадлежит:

A plurality of film cooling holes are formed, so as to communicate with a front cooling passage, in a vane surface on the front-edge side of a stator vane body of a turbine stator vane. The hole cross-section of each of the film cooling holes has a rectangular long-hole shape extending in a direction parallel to the cross-section along the span direction and having a rounded corner. The hole-center line of each of the film cooling holes is inclined with respect to the thickness direction in the cross-section along the span direction. The exit-side portion of the hole wall surface of each of the film cooling holes is inclined with respect to the thickness direction by a greater degree than that of the hole-center line. 1. A turbine vane for a turbine of a gas turbine engine , and capable of being cooled by cooling air , the turbine vane comprising: a vane surface;', 'a cooling passage allowing the cooling air to flow into the vane body, and', 'a plurality of film cooling holes formed in the vane surface on a front edge side of the vane body so as to communicate with the cooling passage to jet out the cooling air through the plurality of film cooling holes, a hole cross section of each film cooling hole having a long-hole shape extending in a direction parallel to a cross section along a span direction of the vane body, a hole-center line of each film cooling hole tilting from a thickness direction of the vane body on a cross section of the vane surface along the span direction, and an exit-side and obtuse angle-side portion of a hole wall surface of each film cooling hole tilting further from the thickness direction than the hole-center line on the cross section along the span direction., 'a vane body including2. The turbine vane according to claim 1 , whereinan aspect ratio of the hole cross section of each film cooling hole is set in a range of 1.1 to 3.0.3. The turbine vane according to claim 1 , whereina tilt angle of the hole-center line of each film cooling ...

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Номер записи: 44
06-02-2014 дата публикации

AIRFOIL DESIGN HAVING LOCALIZED SUCTION SIDE CURVATURES

Номер: US20140037459A1
Автор: Aggarwala Andrew S., Allen-Bradley Eunice, Grover Eric A., Trindade Ricardo
Принадлежит: UNITED TECHNOLOGIES CORPORATION

An airfoil for a gas turbine engine comprises a radially extending body having a transverse cross-section. The transverse cross-section comprises a leading edge, a trailing edge, a pressure side and a suction side. The pressure side extends between the leading edge and the trailing edge with a predominantly concave curvature. The suction side extends between the leading edge and the trailing edge with a predominantly convex curvature. The suction side includes an approximately flat portion flanked by forward and aft convex portions. In another embodiment, the suction side includes a series of local curvature changes that produce inflection points in the convex curvature of the suction side spaced from the trailing edge. 1. An airfoil for a gas turbine engine , the airfoil comprising: a leading edge;', 'a trailing edge;', 'a pressure side extending between the leading edge and the trailing edge with a predominantly concave curvature; and', 'a suction side extending between the leading edge and the trailing edge with a predominantly convex curvature that includes an approximately flat portion flanked by forward and aft convex portions., 'a radially extending body having a transverse cross-section comprising2. The airfoil of wherein the approximately flat portion is formed by a series of local curvature changes that produce inflection points in the convex curvature of the suction side.3. The airfoil of wherein the approximately flat portion is defined by a plurality of changes in sign of the second derivative of a curve defining the suction side.4. The airfoil of wherein the radially extending body defines a chord length and wherein the approximately flat portion is located within a mid-chord region of the airfoil on the suction side.5. The airfoil of wherein the approximately flat portion joins the forward convex portion at a mid-chord point of the airfoil.6. The airfoil of wherein the approximately flat portion joins the forward convex portion at a throat of the ...

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Номер записи: 45
06-02-2014 дата публикации

Cooled blade for a gas turbine

Номер: US20140037460A1
Автор: Helen Marie SAXER-FELICI, Martin Schnieder, Shailendra Naik
Принадлежит: Alstom Technology AG

The invention relates to a cooled blade for a gas turbine that includes a radially extending aerofoil with a leading edge, a trailing edge, a suction side and a pressure side, and wherein a lip overhang is provided on the suction side of the trailing edge The blade also includes a plurality of radial internal flow channels connected via flow bends to form a multi-pass serpentine for a coolant flow, whereby a trailing edge ejection region is provided for cooling said trailing edge, said trailing edge ejection region comprising a trailing edge passage of said multi-pass serpentine running essentially parallel to said trailing edge and being connected over its entire length with a pressure side bleed. An optimized cooling is achieved by mainly determining the cooling flow from the trailing edge passage to the pressure side bleed by means of a staggered field of pins, which is provided between said pressure side bleed and said trailing edge passage, with the lateral dimension of said pins increasing in coolant flow direction.

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Номер записи: 46
06-02-2014 дата публикации

COOLING SYSTEM IN A TURBINE AIRFOIL ASSEMBLY INCLUDING ZIGZAG COOLING PASSAGES INTERCONNECTED WITH RADIAL PASSAGEWAYS

Номер: US20140037461A1
Автор: Lee Ching-Pang
Принадлежит:

An airfoil in a gas turbine engine includes an outer wall, a cooling fluid cavity, a plurality of cooling fluid passages, and a plurality of radial passageways. The outer wall has leading and trailing edges, pressure and suction sides, and radially inner and outer ends. The cooling fluid cavity is defined in the outer wall and receives cooling fluid for cooling the outer wall. The cooling fluid passages are in fluid communication with the cooling fluid cavity and include alternating angled sections, each section having both a radial component and a chordal component. The cooling fluid passages extend from the cooling fluid cavity toward the trailing edge of the outer wall and receive cooling fluid from the cooling fluid cavity for cooling the outer wall near the trailing edge. The radial passageways interconnect radially adjacent cooling fluid passages. 1. An airfoil in a gas turbine engine comprising:an outer wall including a leading edge, a trailing edge, a pressure side, a suction side, a radially inner end, and a radially outer end, wherein a chordal direction is defined between the leading and trailing edges;a cooling fluid cavity defined in the outer wall and extending generally radially between the inner and outer ends of the outer wall, the cooling fluid cavity receiving cooling fluid for cooling the outer wall;a plurality of cooling fluid passages in fluid communication with the cooling fluid cavity, the cooling fluid passages comprising alternating angled sections, each section having both a radial component and a chordal component, the cooling fluid passages extending from the cooling fluid cavity toward the trailing edge of the outer wall and receiving cooling fluid from the cooling fluid cavity for cooling the outer wall near the trailing edge; anda plurality of radial passageways interconnecting radially adjacent cooling fluid passages.2. The airfoil according to claim 1 , wherein the radial passageways between radially adjacent cooling fluid passages ...

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Номер записи: 47
13-02-2014 дата публикации

TRAILING EDGE COOLING CONFIGURATION FOR A GAS TURBINE ENGINE AIRFOIL

Номер: US20140044555A1
Автор: Lewis Scott D., Zess Gary A.
Принадлежит:

An airfoil for a gas turbine engine includes pressure and suction surfaces provided by pressure and suction walls that extend in a radial direction and are joined at a leading edge and a trailing edge. A cooling passage is arranged between the pressure and suction walls and extends to the trailing edge. Elongated pedestals are arranged in the cooling passage and interconnect the pressure and suction walls. The elongated pedestals are spaced apart from one another in the radial direction and extend from a plane to the trailing edge. A metering pedestal includes at least a portion that is arranged between the plane and the trailing edge. The portion is provided between adjacent elongated pedestals in the radial direction. 1. An airfoil for a gas turbine engine comprising:pressure and suction surfaces provided by pressure and suction walls extending in a radial direction and joined at a leading edge and a trailing edge;a cooling passage arranged between the pressure and suction walls and extending to the trailing edge;elongated pedestals arranged in the cooling passage and interconnecting the pressure and suction walls, the elongated pedestals spaced apart from one another in the radial direction and extending from a plane to the trailing edge; anda metering pedestal including at least a portion arranged between the plane and the trailing edge, the portion provided between adjacent elongated pedestals in the radial direction.2. The airfoil according to claim 1 , comprising first claim 1 , second claim 1 , third and fourth rows of pedestals spaced from one another in a chord-wise direction claim 1 , and the first claim 1 , second and third rows respectively include radially spaced first claim 1 , second and third pedestals claim 1 , wherein the fourth row includes the elongated pedestals and the metering pedestals.3. The airfoil according to claim 2 , wherein at least one the first and second pedestals are larger than the third and metering pedestals.4. The airfoil ...

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Номер записи: 48
20-02-2014 дата публикации

Gas turbine engine component having platform trench

Номер: US20140047844A1
Автор: Brandon W. Spangler, Bret M. TELLER, Mark F. Zelesky, Ricardo Trindade, Scott D. Lewis
Принадлежит: Individual

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform that axially extends between a leading edge and a trailing edge and circumferentially extends between a first mate face and a second mate face and a trench disposed on at least one of the first mate face and the second mate face. A plurality of cooling holes are axially disposed within the trench.

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Номер записи: 49
27-02-2014 дата публикации

Gas turbine engine airfoil internal cooling features

Номер: US20140056717A1
Автор: Daniel C. Nadeau, Dominic J. Mongillo, Jr., Jeffrey R. Levine
Принадлежит: Individual

An airfoil for a gas turbine engine includes spaced apart pressure and suction walls joined at leading and trailing edges to provide an airfoil. Intermediate walls interconnect the pressure and suction walls to provide cooling passageways. The cooing passageways have interior pressure and suction surfaces that are respectively provided on the pressure and suction walls. Trip strips include a chevron-shaped trip strip that is provided on at least one of the interior pressure and suction surfaces.

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Номер записи: 50
27-02-2014 дата публикации

Gas Turbine, Gas Turbine Blade, and Manufacturing Method of Gas Turbine Blade

Номер: US20140056719A1
Автор: Hisato Tagawa, Manabu Yagi, Shinichi Higuchi, Tetsuro Morisaki, Yasuhiro Horiuchi
Принадлежит: HITACHI LTD

A gas turbine blade includes a hollow-blade-form portion formed by a leading edge on an upstream side of an working fluid of a gas turbine in a flow direction, a trailing edge on a downstream side of the working fluid in the flow direction, and a suction surface and a pressure surface reaching the trailing edge from the leading edge, and a shank portion for supporting the blade form portion. The blade also includes a partition for connecting the suction surface and the pressure surface in a hollow region of the blade-form portion, coolant paths formed by the partition, the suction surface, and the pressure surface, an impingement cooling hole formed in the partition for dividing the first path that is a flow path nearest to the leading edge side among the coolant paths and the second path adjacent to the first path, and first and second converter portions.

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Номер записи: 51
06-03-2014 дата публикации

TURBINE ROTOR BLADE PLATFORM COOLING

Номер: US20140064942A1
Автор: Itzel Gary Michael, Porter Christopher Donald
Принадлежит:

A cooling arrangement in a platform in a rotor blade or a sidewall in a stator blade in a turbine of a combustion turbine engine is described. The cooling arrangement may include: a cooling chamber configured to pass coolant from an inlet to an outlet; and a rib positioned within the cooling chamber. The rib may partially divide the cooling chamber so to form a switchback. The rib may be canted with respect to the cooling chamber such that the switchback has an ever narrowing channel. 1. A cooling arrangement in one of a sidewall of a stator blade and a platform in a rotor blade in a turbine of a combustion turbine engine , the cooling arrangement comprising:a cooling chamber configured to pass coolant from an inlet to an outlet; anda rib positioned within the cooling chamber, the rib partially dividing the cooling chamber so to form a switchback;wherein the rib is canted with respect to the cooling chamber such that the switchback comprises an ever narrowing channel.2. The cooling arrangement according to claim 1 , wherein ever narrowing channel comprises a channel that narrows at a constant rate as the channel extends from the inlet to the outlet of the cooling chamber.3. The cooling arrangement according to claim 1 , wherein the ever narrowing channel comprises a channel that narrows at a constant rate along both flanks of the rib.4. The cooling arrangement according to claim 3 , wherein the switchback comprises a pass positioned on each flank of the rib: an upstream pass disposed on a flank of the rib that coincides with the inlet; and a downstream pass disposed on a flank of the rib that coincides with the outlet; andwherein, between the upstream pass and the downstream pass, the switchback comprises a turn section that defines a turn of approximately 180°.5. The cooling arrangement according to claim 4 , wherein the cooling chamber comprises a first edge and a second edge claim 4 , wherein the second edge opposes the first edge across the cooling chamber; ...

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Номер записи: 52
06-03-2014 дата публикации

AEROFOIL COOLING ARRANGEMENT

Номер: US20140064967A1
Автор: HARDING Adrian Lewis
Принадлежит: ROLLS-ROYCE PLC

An aerofoil typically for a blade or vane for a gas turbine engine comprises a pressure wall, a suction wall and a web that extends therebetween. One of the walls comprises an inner leaf and an outer leaf which defines a cavity, the other wall defines a first outlet, the aerofoil defines at least one first passage extending from the cavity through the web and to the outlet. Coolant in the cavity on one side wall of the aerofoil is thereby routed across the web to be used to cool the other wall by convention and formation of a coolant film. 1. An aerofoil for a gas turbine engine ,the aerofoil comprises a pressure wall, a suction wall and a web,the web extends between the pressure wall and the suction wall,at least one of the pressure and suction walls comprise an inner leaf and an outer leaf that define a first cavity,the other of the pressure and suction walls defines a first outlet,the aerofoil defines at least one first passage extending from the cavity through the web and to the outlet.2. The aerofoil of wherein the pressure wall claim 1 , the suction wall and the web define at least two main coolant passages.3. The aerofoil of wherein the inner leaf is an impingement plate having an array of impingement holes through which coolant can pass from at least one of the two main coolant passages and impinge on the outer leaf.4. The aerofoil of wherein the inner leaf and the outer leaf define a second cavity and the pressure wall or suction wall defines a second outlet claim 1 ,the aerofoil defines at least one second passage extending from the second cavity through the web and to the second outlet.5. The aerofoil of wherein the suction wall comprises the inner leaf and the outer leaf.6. The aerofoil of wherein the pressure wall comprises the inner leaf and the outer leaf.7. The aerofoil of wherein the aerofoil has an array of first and/or second passages and an array of first and/or second outlets.8. The aerofoil of wherein the arrays of first and second passages ...

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Номер записи: 53
06-03-2014 дата публикации

Cooling arrangement for platform region of turbine rotor blade

Номер: US20140064984A1
Автор: Adebukola Oluwaseun Benson, Camilo Andres Sampayo, Christopher William Kester, Xiuzhang James Zhang
Принадлежит: General Electric Co

A platform cooling arrangement in a turbine rotor blade having a platform at an interface between an airfoil and a root. The platform may include a pressure side slashface and a suction side slashface. The platform cooling arrangement may include: a cooling channel formed within the interior of the platform, the cooling channel extending from a first end toward one of the pressure side slashface and the suction side slashface. At a second end, the cooling channel may include a pocket. The pocket may include an abrupt increase in cross-sectional flow area just before the cooling channel reaches the slashface.

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Номер записи: 54
13-03-2014 дата публикации

BUCKET ASSEMBLY FOR TURBOMACHINE

Номер: US20140069108A1
Автор: Good Randall Richard, Itzel Gary Michael, Smith Aaron Ezekiel
Принадлежит: GENERAL ELECTRIC COMPANY

Bucket assemblies are provided. The bucket assembly includes a shank, and an airfoil positioned radially outward of the shank. The bucket assembly further includes a main cooling circuit defined in the airfoil and the shank, the main cooling circuit comprising seven passages, each of the seven passages fluidly connected with an adjacent one of the seven passages. A maximum rotation number in each of the seven passages is less than or equal to approximately 0.4. 1. A bucket assembly , comprising:a shank;an airfoil positioned radially outward of the shank;a main cooling circuit defined in the airfoil and the shank, the main cooling circuit comprising seven passages, each of the seven passages fluidly connected with an adjacent one of the seven passages,wherein a maximum rotation number in each of the seven passages is less than or equal to approximately 0.4.2. The bucket assembly of claim 1 , wherein the rotation number in each of the seven passages is in a range between approximately 0.01 and approximately 0.35.3. The bucket assembly of claim 1 , wherein a first passage of the seven passages has a maximum cross-sectional area in the airfoil that is in a range between approximately 10% and approximately 20% less than a maximum cross-sectional area in the airfoil of another passage of the seven passages.4. The bucket assembly of claim 1 , further comprising an exhaust passage defined in the airfoil and in fluid communication with a seventh passage of the seven passages.5. The bucket assembly of claim 1 , further comprising a forward cooling circuit and an aft cooling circuit are each defined in the airfoil and the shank.6. The bucket assembly of claim 1 , wherein the main cooling circuit consists of seven passages.7. The bucket assembly of claim 1 , further comprising a platform positioned radially between the shank and the airfoil.8. The bucket assembly of claim 1 , wherein the airfoil has exterior surfaces defining a pressure side and a suction side extending between ...

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Номер записи: 55
20-03-2014 дата публикации

STEAM TURBINE STATIONARY BLADE AND STEAM TURBINE

Номер: US20140079556A1
Автор: Ishibashi Koji, KUDO Takeshi, MIZUMI Shunsuke, Nakano Susumu
Принадлежит: Hitachi, Ltd.

A plurality of slots with different widths are provided in a plurality of line on a stationary blade surface. More specifically, the steam turbine stationary blade has a hollow nozzle with a penetrating space, which is connected with a diaphragm outer ring or inner ring, and a plurality of suction slots extending radially which are arranged on the blade surface. At a position where a water film deposited to the blade surface is thick, the width of a slot is smaller and at a position where the water film is thin, the width of a slot is larger. 1. A steam turbine stationary blade with a hollow inner portion , including:a plurality of slots formed on a blade surface of the steam turbine stationary blade, extending in a blade height direction and provided in a plurality of lines,wherein widths of the slots differ according to a position on the stationary blade surface in a turbine axial direction.2. The steam turbine stationary blade according to claim 1 ,wherein the slots include a first slot located upstream in the turbine axial direction and a second slot located downstream of the first slot and having a slot width larger than the slot width of the first slot.3. The steam turbine stationary blade according to claim 2 ,wherein the slots are located on a pressure side of the stationary blade.4. The steam turbine stationary blade according to claim 2 ,wherein the slots are located on a suction side of the stationary blade.5. The steam turbine stationary blade according to claim 3 ,wherein the first slot is located at a position where a water film deposited to the stationary blade surface is thick, and the second slot is located at a position where the water film is thinner than at the position where the first slot is located.6. The steam turbine stationary blade according to claim 5 ,wherein the first slot is located at a position within a range of l/L ratio from 0.6 to 0.8 where l represents a distance along the blade surface from a leading edge of the stationary blade ...

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Номер записи: 56
27-03-2014 дата публикации

TURBINE BLADE ROOT PROFILE

Номер: US20140083114A1
Автор: Beattie Jeffrey S., Jacques Jeffrey Michael
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A turbine blade for a gas turbine engine includes an airfoil that extends in a first radial direction from a platform. A root extends from the platform in a second radial direction and has opposing lateral sides that provide a firtree-shaped contour. The contour includes first, second and third lobes on each of the lateral sides and that tapers relative to the radial direction away from the platform. The first, second and third lobes each provide contact surfaces arranged at about 45° relative to the radial direction. A contact plane on each lateral side at an angle of about 11° relative to the radial direction defining a contact point on each of the contact surfaces. The first, second and third lobes each include first, second and third grooves that are substantially aligned with one another along an offset plane spaced a uniform offset distance from the contact plane. 1. A turbine blade for a gas turbine engine comprising:an airfoil extending in a first radial direction from a platform; anda root extending from the platform in a second radial direction and having opposing lateral sides providing a firtree-shaped contour, the contour including first, second and third lobes on each of the lateral sides and that tapers relative to the radial direction away from the platform, the first, second and third lobes each provide contact surfaces arranged at about 45° relative to the radial direction, and a contact plane on each lateral side at an angle of about 11° relative to the radial direction defining a contact point on each of the contact surfaces, the first, second and third lobes each include first, second and third grooves that are substantially aligned with one another along an offset plane spaced a uniform offset distance from the contact plane.2. The turbine blade according to claim 1 , wherein the second lobe is arranged radially between the first and third lobes claim 1 , the contact points on the second lobe align in an intersecting plane spaced apart from a ...

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Номер записи: 57
27-03-2014 дата публикации

GAS TURBINE ENGINE COMPONENT

Номер: US20140086724A1
Автор: JACKSON Dougal Richard, TIBBOTT Ian
Принадлежит: ROLLS-ROYCE PLC

An internally cooled gas turbine engine component has a line of cooling air discharge holes, an internal cooling channel, an internal feed cavity for feeding cooling air from the channel to the discharge holes, and flow disrupting pedestals arranged in rows. A method of configuring the component includes: 1. A method of configuring an internally cooled gas turbine engine component , the component having a line of cooling air discharge holes , an internal cooling channel forward of and extending substantially parallel to the line of discharge holes , and an internal feed cavity between the channel and the line of discharge holes for feeding cooling air from the channel to the discharge holes , the component further having a plurality of flow disrupting pedestals extending between opposing sides of the feed cavity , the pedestals being arranged in a number N of rows which extend substantially parallel to the line of discharge holes , the first row being at the entrance from the channel to the feed cavity , the Nrow being at the exit from the feed cavity to the discharge holes , the remaining rows being spaced therebetween , and the pedestals being spaced apart from each other within each row , the method including:determining an angle α of the direction of cooling air flow into the first row;{'sup': 'th', 'determining an angle β of the direction of cooling air flow from the Nrow;'}defining a change in angle φ of the direction of cooling air flow between rows as φ=(β−α)/N; and{'sup': th', 'th', 'th, 'positioning the pedestals such that a line extending forward from the centre of each pedestal in the irow at an angle {α+φ(i−1)} intersects the (i−1)row at a location which is midway between two neighbouring pedestals of the (i−1)row, i being an integer from 2 to N.'}2. A method according to claim 1 , wherein the rows are spaced substantially equal distances apart.3. A method according to claim 1 , wherein the method further includes:{'sup': th', 'th, 'determining the ...

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Номер записи: 58
27-03-2014 дата публикации

METHOD AND COOLING SYSTEM FOR COOLING BLADES OF AT LEAST ONE BLADE ROW IN A ROTARY FLOW MACHINE

Номер: US20140086743A1
Автор: Heidecke Axel, JUSTL Sascha, OLMES Sven, Simon-Delgado Carlos
Принадлежит: ALSTOM Technology Ltd

A method and a cooling system for cooling blades of at least one blade row in a rotary flow machine includes an axial flow channel which is radially limited on the inside by a rotor unit and at the outside by at least one stationary component, the blades are arranged at the rotary unit and provide a shrouded blade tip facing radially to said stationary component. Pressurized cooling air is fed through from radially outside towards the tip of each of said blades in the at least one blade row, and the pressurized cooling air enters the blades through at least one opening at the shrouded blades' tip. 1. A method for cooling blades of at least one blade row in a rotary flow machine , comprising an axial flow channel which is radially limited on the inside by a rotor unit and at the outside by at least one stationary component , said blades are arranged at the rotary unit and provide a shrouded blade tip facing radially to said stationary component , wherein the pressurized cooling air is fed through from radially outside towards the tip of each of said blades in the at least one blade row , and said pressurized cooling air enters the blades through at least one opening at the shrouded blades' tip.2. The method according to claim 1 , wherein the pressurized cooling air is fed through the stationary component surrounding said at least one blade row radially and entering a cavity enclosed by the stationary component and shrouded tips of the blades in the at least one blade row.3. The method according to claim 2 , wherein the pressurized cooling air is fed into the cavity through at least one claim 2 , stationary component sided opening such that a static pressure prevails within said cavity which is higher than a total relative pressure of a flow in the axial flow channel at a leading edge of the blades in the at least one blade row.4. A cooling system for cooling blades of at least one blade row in a rotary flow machine comprising an axial flow channel which is radially ...

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Номер записи: 59
03-04-2014 дата публикации

AIRFOIL WITH VARIABLE TRIP STRIP HEIGHT

Номер: US20140093361A1
Автор: Boeke Mark A., DeGray Jeffrey J., Gregg Shawn J., Riley Sarah
Принадлежит: UNITED TECHNOLOGIES CORPORATION

An airfoil component for a gas turbine engine includes an airfoil extending from a platform. At least one of the airfoil and the platform includes a cooling passage defined by a surface. A chevron-shaped trip strip extends from the surface into the cooling passage at a trip strip height along a length. The trip strip height varies along the length. A turbine vane for a gas turbine engine includes inner and outer platforms. A cooling passage is provided in the inner platform. The cooling passage is provided by first and second radially extending legs spaced circumferentially apart from one another and joined to one another by a circumferential passage. A pair of airfoils extend radially from the same inner platform. A trip strip extends from the surface into the circumferential passage at a trip strip height along a length. The trip strip height varying along the length. 1. An airfoil component for a gas turbine engine comprising:an airfoil extending from a platform, at least one of the airfoil and the platform including a cooling passage defined by a surface; anda chevron-shaped trip strip extending from the surface into the cooling passage at a trip strip height along a length, the trip strip height varying along the length.2. The airfoil component according to claim 1 , wherein the length is provided by multiple zones claim 1 , the height varying between the zones.3. The airfoil component according to claim 2 , wherein the multiple zones include first claim 2 , second and third cooling passage heights claim 2 , and the trip strip includes first claim 2 , second and third trip strip heights respectively within the first claim 2 , second and third zones.4. The airfoil component according to claim 1 , wherein the chevrons are provided by first and second legs joined to one another at an apex to provide the chevron-shape.5. The airfoil component according to claim 3 , wherein a trip strip portion within each of the multiple zone includes a p/e ratio claim 3 , wherein ...

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Номер записи: 60
03-04-2014 дата публикации

Cooled turbine airfoil structures

Номер: US20140093389A1
Автор: Deane Krmpotich, Jon Kettinger, Malak Fouad Malak, Mark C. Morris, Stuart Andrew Harman
Принадлежит: Honeywell International Inc

In accordance with an exemplary embodiment, disclosed is an air-cooled turbine blade having an airfoil shape, including a convex suction side wall, a concave pressure side wall, the walls including an interior surface that defines an interior with the blade, a suction side flow circuit formed within the blade interior, a pressure side flow circuit formed within the blade interior; and a trailing edge pin bank positioned aft of the suction side and pressure side flow circuits. The turbine blade includes a wishbone-shaped architecture at a transition point between the suction side flow circuit and the pressure side flow circuit and the trailing edge pin bank.

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Номер записи: 61
10-04-2014 дата публикации

GAS TURBINE ENGINE COMPONENTS WITH LATERAL AND FORWARD SWEEP FILM COOLING HOLES

Номер: US20140099189A1
Автор: Malak Malak Fouad, Morris Mark C., Tapia Luis, Waldman David
Принадлежит: HONEYWELL INTERNATIONAL INC.

An engine component includes a body having an internal surface and an external surface, the internal surface at least partially defining an internal cooling circuit. The engine component further includes a plurality of cooling holes formed in the body and extending between the internal cooling circuit and the external surface of the body. The plurality of cooling holes includes a first cooling hole with forward diffusion and lateral diffusion. 1. An engine component , comprising:a body having an internal surface and an external surface, the internal surface at least partially defining an internal cooling circuit; anda plurality of cooling holes formed in the body and extending between the internal cooling circuit and the external surface of the body, the plurality of cooling holes including a first cooling hole with forward diffusion and lateral diffusion.2. The engine component of claim 1 , wherein the first cooling hole includes an inlet at the internal cooling circuit claim 1 , a metering section extending from the inlet claim 1 , a first exit portion extending from the metering section claim 1 , a second exit portion extending from the first exit portion claim 1 , and an outlet defined on the external surface and fluidly coupled to the second exit portion.3. The engine component of claim 2 , wherein the first exit portion extends at a first angle relative to the metering section and the second exit portion extends at a second angle relative to the metering section claim 2 , the second angle being greater than the first angle.4. The engine component of claim 2 , wherein the first exit portion is curved with a first radius of curvature and the second exit portion is curved with a second radius of curvature claim 2 , the second radius of curvature being less than the first radius of curvature.5. The engine component of claim 2 , wherein the first exit portion is curved with a first radius of curvature and the second exit portion is curved with a second radius of ...

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Номер записи: 62
06-01-2022 дата публикации

FILM COOLING DIFFUSER HOLE

Номер: US20220003119A1
Автор: Candelori David J., Ghigliotty Jaime G., QUACH San
Принадлежит: Raytheon Technologies Corporation

An airfoil for a gas turbine engine is disclosed. In various embodiments, the airfoil includes a cooling passage; an outer wall separating a core flow path from the cooling passage; a diffuser in fluid communication with the cooling passage and opening into the core flow path, the diffuser being characterized by a linear ridge on a downstream end of the diffuser; and a thermal barrier coating covering the outer wall and the linear ridge. 1. An airfoil for a gas turbine engine , comprising:a cooling passage;an outer wall separating a core flow path from the cooling passage;a diffuser in fluid communication with the cooling passage and opening into the core flow path, the diffuser being characterized by a linear ridge on a downstream end of the diffuser; and 'wherein the linear ridge includes an upstream facing side that is characterized by a height extending a first distance in a direction normal to the outer wall, the height having a value within a range of between five one-hundredths and seventy-five one-hundredths of a depth of the cooling passage, the depth extending between a first wall and a second wall that define the cooling passage and in the direction normal to the outer wall.', 'a thermal barrier coating covering the outer wall and the linear ridge,'}2. The airfoil of claim 1 , wherein the diffuser defines a rectangular shape in the direction normal to the outer wall.3. The airfoil of claim 2 , wherein the linear ridge extends perpendicular to the cooling passage along the downstream end of the diffuser.4. The airfoil of claim 3 , wherein the thermal barrier coating includes a first portion upstream of the linear ridge claim 3 , the first portion extending from the cooling passage and being characterized by a first radius of curvature.5. The airfoil of claim 4 , wherein the thermal barrier coating includes a second portion claim 4 , the second portion extending from the first portion and over the linear ridge and being characterized by a second radius of ...

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Номер записи: 63
06-01-2022 дата публикации

AIRFOIL TIP POCKET WITH AUGMENTATION FEATURES

Номер: US20220003120A1
Автор: Gautschi Steven Bruce, Propheter-Hinckley Tracy A., QUACH San
Принадлежит:

A component for a gas turbine engine includes, among other things, an airfoil that includes a pressure sidewall and a suction sidewall that meet together at both a leading edge and a trailing edge, the airfoil extending radially from a platform to a tip, a tip pocket formed in the tip and terminating prior to the trailing edge, and one or more heat transfer augmentation devices formed in the tip pocket. 1. A component for a gas turbine engine comprising:an airfoil extending in a chordwise direction between a leading edge and a trailing edge and in a thickness direction between a pressure sidewall and a suction sidewall that meet together at both the leading edge and the trailing edge, the airfoil extending in a radial direction from a platform to a tip;a tip pocket formed in the tip and terminating prior to the trailing edge;wherein the tip pocket includes a suction side lip, a pressure side lip, a leading edge lip and a trailing edge lip that each extend outwardly in the radial direction from a floor to the tip;one or more heat transfer augmentation devices formed in the tip pocket, the one or more heat transfer augmentation devices including at least one rib extending from one of the suction side lip and the pressure side lip such that a wall of the at least one rib is spaced apart from another one of the suction side lip and the pressure side lip;wherein the at least one rib extends outwardly in the radial direction from the floor towards the tip such that a length of the at least one rib is slanted at an acute angle in the chordwise direction toward either the leading edge lip or the trailing edge lip relative to the floor; anda plurality of cooling holes defined in the floor that fluidly connect the tip pocket to at least one internal cooling cavity formed inside the airfoil.2. The component as recited in claim 1 , wherein at least one of the plurality of cooling holes is angled relative to the floor.3. The component as recited in claim 1 , wherein the at least ...

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Номер записи: 64
05-01-2017 дата публикации

ROTOR BLADE OR GUIDE VANE ASSEMBLY

Номер: US20170002661A1
Автор: FERBER Joergen, HOEVEL Simone, OPDERBECKE Thomas
Принадлежит: General Electric Technology GmbH

The disclosure refers to a method of assembling or disassembling a rotor blade or guide vane assembly, wherein the rotor blade or guide vane includes an airfoil and additional structured peripheral members forming at least an inner and/or an outer platform of the rotor blade or guide vane. The airfoil includes at least one airfoil sub-structure designed for anchoring at least one superposed component for the purpose of a thermal protection. The connection between the airfoil sub-structure and the superposed component is supported on friction-locked device, wherein the airfoil sub-structure is formed by a spar and the superposed component includes at least one flow-charged outer shell. Connection between the spar or airfoil understructure and flow-charged outer shell is formed by a force-fit and/or a form-fit fixation or a shrinking joint, wherein the airfoil and additional structured peripheral members is formed by friction-locked device with a detachable, permanent or semi-permanent fixation. 1. Modular rotor blade or guide vane , at least comprising: an airfoil , a platform and a root , wherein the airfoil includes an inner core structure , designed for anchoring at least one shell , and one or more shells , encasing the inner core structure as a whole or in part , one of said shells being an outer shell , forming the outer contour of the blade or vane airfoil and being flow-charged in operating mode , wherein the connection between the inner core structure and the flow-charged outer shell is formed by a force-fit or form-fit fixation or a shrinking joint.2. Modular rotor blade or guide vane according to claim 1 , wherein at least the outer shell is designed in a closed one-piece configuration.3. Modular rotor blade or guide vane according to claim 1 , wherein at least the outer shell is designed in an envelope configuration claim 1 , to be closed after wrapping.4. Modular rotor blade or guide vane according to claim 3 , wherein the joint claim 3 , formed by the ...

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Номер записи: 65
05-01-2017 дата публикации

Gas turbine engine airfoil with bi-axial skin core

Номер: US20170002662A1
Автор: Benjamin F. Hagan, Matthew A. Devore, Ryan Alan WAITE
Принадлежит: United Technologies Corp

An airfoil for a gas turbine engine includes a body that extends in a radial direction that provides an exterior airfoil surface. The body includes pressure and suction side walls that extend from a leading edge to a trailing edge in a chord-wise direction. A core cooling passage is provided between the pressure and suction side walls and extends in the radial direction. A skin passage is arranged in one of the pressure and suction side walls between the core cooling passage and the exterior airfoil surface. The skin passage includes a first passageway that extends in the radial direction. The first passageway turns to a second passageway that extends in the chord-wise direction to terminate at an exit arranged near the trailing edge.

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Номер записи: 66
05-01-2017 дата публикации

GAS TURBINE ENGINE AIRFOIL SQUEALER POCKET COOLING HOLE CONFIGURATION

Номер: US20170002663A1
Автор: Chapdelaine David Donald, Corcoran Christopher, Gautschi Steven Bruce, King Christopher, Lewis Scott D., QUACH San
Принадлежит:

A gas turbine engine airfoil includes a body that provides an exterior airfoil surface that extends in a radial direction to a tip. The exterior surface has a leading edge in a forward direction and a trailing edge in an aft direction. The tip includes a squealer pocket that has a recess surface. A cooling passage is arranged in the body. Each of the cooling holes extends from an inlet at the cooling passage to an outlet at the recessed surface. The inlet and outlet are arranged at an angle in an angular direction relative to the recessed surface. The angular direction is toward at least one of the forward and aft directions. 1. A gas turbine engine airfoil comprising:a body that provides an exterior airfoil surface that extends in a radial direction to a tip, the exterior surface has a leading edge in a forward direction and a trailing edge in an aft direction, the tip includes a squealer pocket that has a recess surface, a cooling passage is arranged in the body, and each of the cooling holes extend from an inlet at the cooling passage to an outlet at the recessed surface, the inlet and outlet are arranged at an angle in an angular direction relative to the recessed surface, the angular direction is toward at least one of the forward and aft directions.2. The airfoil according to claim 1 , wherein the angular direction of at least one of the cooling holes is toward the forward direction.3. The airfoil according to claim 1 , wherein the angular direction of at least one of the cooling holes is toward the aft direction.4. The airfoil according to claim 1 , wherein the angular direction of at least one of the cooling holes is toward the forward direction and the angular direction of at least another one of the holes is toward the aft direction.5. The airfoil according to claim 4 , wherein the exterior airfoil surface includes pressure and suction side joined at the leading and trailing edges claim 4 , wherein the angular directions of one set of cooling holes nearest ...

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Номер записи: 67
05-01-2017 дата публикации

TURBINE BLADE

Номер: US20170002664A1
Автор: Manning Robert Francis, McKenzie Christopher Lloyd, Reddy Bhanu Mahasamudram
Принадлежит:

A turbine blade includes an airfoil having an internal cooling circuit. The cooling circuit includes a body cooling passage with at least one turn, and a tip cooling channel that forms a cooling barrier to thermally isolate the turn from at least a portion of the exterior surface of the airfoil. 1. A turbine blade comprising:an airfoil extending between a root and a tip, and having a pressure sidewall and a suction sidewall joined together to define a leading edge and a trailing edge;a body cooling passage located within the airfoil and having at least one tip turn located proximate the tip; anda tip cooling channel extending along the tip and enveloping the at least one tip turn to form a cooling barrier between the at least one tip turn and an exterior surface of the airfoil on all sides of the at least one tip turn.2. The turbine blade of claim 1 , wherein the body cooling passage further comprises a serpentine cooling passage.3. The turbine blade of claim 1 , wherein the tip cooling channel is in fluid communication with the body cooling passage downstream of the at least one tip turn.4. The turbine blade of and further comprising a plurality of tip turns located proximate the tip claim 1 , wherein the tip cooling channel envelops the tip turns to form a cooling barrier between the tip turns and an exterior surface of the airfoil on all sides of the tip turns.5. The turbine blade of and further comprising a plurality of body cooling passages located within the airfoil claim 1 , wherein each of the plurality of body cooling passages has at least one tip turn located proximate the tip.6. The turbine blade of and further comprising a trailing edge cooling channel having at least one film hole located along the trailing edge and fluidly coupled to the tip cooling channel.7. A turbine blade comprising:an airfoil extending between a root and a tip, and having a pressure sidewall and a suction sidewall joined together to define a leading edge and a trailing edge;a body ...

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Номер записи: 68
05-01-2017 дата публикации

GAS TURBINE BLADE

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

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

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Номер записи: 69
07-01-2016 дата публикации

GAS TURBINE ENGINE COMPONENT COOLING WITH INTERLEAVED FACING TRIP STRIPS

Номер: US20160003055A1
Автор: Corcoran Christopher, Ghigliotty Jaime G., JR. Dominic J., Mongillo, Pietraszkiewicz Edward F.
Принадлежит:

A gas turbine engine component includes first and second walls spaced apart from one another to provide a cooling passage. First and second trip strips are respectively provided on the first and second walls and arranged to face one another. The first and second trip strips are arranged in an interleaved fashion with respect to one another in a direction. 1. A gas turbine engine component comprising:first and second walls spaced apart from one another to provide a cooling passage, first and second trip strips respectively provided on the first and second walls and arranged to face one another, the first and second trip strips arranged in an interleaved fashion with respect to one another in a direction.2. The gas turbine engine component according to claim 1 , wherein the gas turbine engine component is an airfoil claim 1 , and the direction is a radial direction of the airfoil.3. The gas turbine engine component according to claim 2 , wherein the cooling passage is provided near a leading edge of the airfoil.4. The gas turbine engine component according to claim 3 , wherein trip strips are chevron trip strips arranged asymmetrically claim 3 , an apex of the chevron trip strips shifted within the cooling passage toward a leading edge.5. The gas turbine engine component according to claim 2 , wherein the cooling passage is provided near a trailing edge of the airfoil.6. The gas turbine engine component according to claim 1 , wherein the trip strips are chevron trip strips.7. The gas turbine engine component according to claim 1 , wherein the trip strips extend from an inner surface a distance e claim 1 , and the first and second walls respectively include first and second inner surfaces that are spaced a distance H from one another claim 1 , an e/H ratio is provided in the range of 0.05-0.40.8. The gas turbine engine component according to claim 7 , wherein the trip strips are spaced from an opposing surface a distance in the range 0.035-0.045 inch (0.89-1.14 mm) ...

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Номер записи: 70
07-01-2016 дата публикации

GAS TURBINE ENGINE SHAPED FILM COOLING HOLE

Номер: US20160003056A1
Автор: Xu JinQuan
Принадлежит:

A component for a gas turbine engine includes a wall that adjoins an interior cooling passage and provides an exterior surface. A film cooling hole fluidly connects the interior cooling passage and the exterior surface. The film cooling passage includes inlet and outlet passages that fluidly interconnect and adjoin one another in a misaligned non-line of sight relationship. 1. A component for a gas turbine engine comprising:a wall adjoining an interior cooling passage and providing an exterior surface, and a film cooling hole fluidly connecting the interior cooling passage and the exterior surface, the film cooling passage including inlet and outlet passages fluidly interconnecting and adjoining one another in a misaligned non-line of sight relationship.2. The component according to claim 1 , wherein the inlet and outlet passages are generally linear.3. The component according to claim 2 , wherein are arranged at an angle relative to one another.4. The component according to claim 3 , wherein the angle is acute.5. The component according to claim 1 , wherein the outlet passage provides a diffuser shape.6. The component according to claim 1 , wherein the inlet passage provides a metering section having a cross-sectional area that is less than a cross-sectional area of the outlet portion.7. The component according to claim 6 , wherein the inlet passage includes first and second metering portions claim 6 , the second metering portion adjoining the outlet passage and including a length L and a diameter D having an L/D ratio of greater than 1.8. The component according to claim 8 , wherein the L/D ratio is greater than 3.9. The component according to claim 1 , wherein the film cooling hole is additively manufactured.10. The component according to claim 1 , wherein the component is one of an airfoil combustor BOAS and platform.11. A method of manufacturing airfoil component for a gas turbine engine claim 1 , comprising:depositing multiple layers of a powdered metal onto ...

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Номер записи: 71
01-01-2015 дата публикации

APPARATUS FOR REDUCING A TEMPERATURE GRADIENT OF MAINSTREAM FLUID DOWNSTREAM OF AN AIRFOIL IN A GAS TURBINE ENGINE

Номер: US20150003962A1
Автор: Smith Bruce L.
Принадлежит:

An apparatus () is presented for reducing a temperature gradient () of mainstream fluid () downstream of an airfoil () in a gas turbine engine (). The apparatus includes a passage () in a trailing edge () of the airfoil having an inlet () and an outlet (). The apparatus also includes a cooling fluid source () coupled to the inlet to transmit cooling fluid into the passage. The apparatus also includes a vortex generator () within the passage effective to generate a vortex fluid () at the outlet. The outlet is positioned to inject the vortex fluid into the mainstream fluid with sufficient mixing energy to cause a reduced temperature gradient () downstream of the airfoil. 1. An airfoil of a gas turbine engine comprising:a root section;an airfoil section comprising a trailing edge connected to the root section;a cooling fluid passage in the trailing edge; anda vortex generator within the trailing edge cooling fluid passage effective to generate vorticity in cooling fluid that is ejected from the passage.2. The airfoil of claim 1 , wherein the vorticity is effective to reduce a temperature gradient downstream of the airfoil.3. The airfoil of claim 1 , wherein the vortex generator comprises a swirler inserted within an outlet of the passage.4. The airfoil of claim 1 , wherein the vortex generator comprises a plurality of swirler channels within the passage claim 1 , and wherein said swirler channels are positioned in an outer portion of the passage.5. The airfoil of claim 4 , wherein the vortex generator further comprises a solid core within a central portion of the passage claim 4 , wherein the swirler channels are disposed about the solid core such that said solid core is configured to redirect fluid from the central portion into the swirler channels.6. The airfoil of claim 4 , wherein the vortex generator further comprises a hollow core within a central portion of the passage claim 4 , and wherein the swirler channels are disposed about the hollow core.7. The airfoil ...

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Номер записи: 72
01-01-2015 дата публикации

Turbine airfoil with ambient cooling system

Номер: US20150003999A1
Автор: Christian X. Campbell, Jr., Jan H. Marsh, John J. Marra
Принадлежит: Siemens Energy Inc

A turbine airfoil usable in a turbine engine and having at least one ambient air cooling system is disclosed. At least a portion of the cooling system may include one or more cooling channels configured to receive ambient air at about atmospheric pressure. The ambient air cooling system may have a tip static pressure to ambient pressure ratio of at least 0.5, and in at least one embodiment, may include a tip static pressure to ambient pressure ratio of between about 0.5 and about 3.0. The cooling system may also be configured such that an under root slot chamber in the root is large to minimize supply air velocity. One or more cooling channels of the ambient air cooling system may terminate at an outlet at the tip such that the outlet is aligned with inner surfaces forming the at least one cooling channel in the airfoil to facilitate high mass flow.

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Номер записи: 73
01-01-2015 дата публикации

TURBINE BLADE

Номер: US20150004001A1
Автор: SCHNIEDER Martin, Shchukin Sergey
Принадлежит:

A turbine vane for a rotary turbomachine is described having a turbine blade which is delimited by a concave pressure-side wall and a convex suction-side wall which are connected in the region of a vane front edge which can be assigned to the turbine blade and enclose a cavity which extends in the longitudinal extent of the vane front edge and is delimited on the inner wall by the pressure-side wall and the suction-side wall in the region of the vane front edge and by an intermediate wall which extends in the longitudinal direction to the vane front edge and connects the suction-side wall and the pressure-side wall on the inner wall. The disclosed vane is distinguished by the fact that the intermediate wall has a perforation at least in sections in the connecting region to the suction-side wall and/or pressure-side wall, in order to increase the elasticity of the intermediate wall. 1. A turbine blade for a rotating turbomachine; the turbine blade comprising:having a blade airfoil which is bounded by a concave pressure side wall and a convex suction side wall which are connected in the region of a blade leading edge which can be assigned to the blade airfoil, and which enclose a cavity which extends in the longitudinal extent of the blade leading edge and is delimited internally by the pressure side wall and suction side wall in the region of the blade leading edge and by an intermediate wall which extends in the longitudinal direction to the blade leading edge and connects the suction side wall and pressure side wall internally, whereinthe intermediate wall has, at least in sections, a perforation in a connection region to the suction side wall and/or pressure side wall in order to increase the elasticity of the intermediate wall in the connection region.2. The turbine blade as claimed in claim 1 , wherein the perforation comprises a row of cylindrical holes.3. The turbine blade as claimed in claim 1 , wherein the perforation comprises a row of longitudinal holes or ...

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Номер записи: 74
04-01-2018 дата публикации

INTERIOR COOLING CONFIGURATIONS FOR TURBINE ROTOR BLADES

Номер: US20180003061A1
Автор: DuBose John Lesley, III John Rogers, Mason
Принадлежит: GENERAL ELECTRIC COMPANY

A turbine rotor blade that includes an interior cooling configuration having a section configuration that includes: a main channel divided into three non-overlapping segments in which an upstream segment connects to a downstream segment via a transition segment positioned therebetween; and one or more branching channels extending from the main channel via connections each makes to the transition segment. The transition segment includes a variable cross-sectional flow area that accommodates a main channel flow area reduction occurring between the upstream segment and the downstream segment. The one or more branching channels having a total branching channel flow area. The section configuration is configured according to a section channel ratio that is defined as the main channel flow area reduction divided by the total branching channel flow area, with the value of the section channel ratio being configured according a desired coolant flow characteristic. 1. A rotor blade for use in a turbine of a gas turbine that includes:an airfoil defined between a concave pressure face and a laterally opposed convex suction face, wherein the pressure face and the suction face extend axially between opposite leading and trailing edges and radially between an outboard tip and an inboard end that attaches to a root configured to couple the rotor blade to a rotor disc; a main channel divided into three non-overlapping segments in which an upstream segment connects to a downstream segment via a transition segment positioned therebetween, wherein the transition segment comprising a variable cross-sectional flow area that accommodates a main channel flow area reduction occurring between the upstream segment and the downstream segment; and', 'one or more branching channels extending from the main channel via connections each of the one or more branching channels makes to the transition segment, wherein the one or more branching channels comprise a total branching channel flow area;, 'an ...

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Номер записи: 75
04-01-2018 дата публикации

Gas turbine blade

Номер: US20180003062A1
Автор: Ik Sang LEE, Myeong Hwan Bang
Принадлежит: Doosan Heavy Industries and Construction Co Ltd

A gas turbine blade includes a plurality of guide ribs spaced apart from each other in a movement direction of a cooling fluid in order to guide movement of the cooling fluid flowing along a cooling passage formed in the turbine blade, and an opening formed in each of the guide ribs in order to guide a portion of the cooling fluid to a bottom of the cooling passage between the guide ribs or to side walls adjacent to the bottom.

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Номер записи: 76
02-01-2020 дата публикации

TURBINE COMPONENT WITH SHAPED COOLING PINS

Номер: US20200003059A1
Автор: Crites Daniel C., Kamrath Benjamin Dosland, Smoke Jason
Принадлежит: HONEYWELL INTERNATIONAL INC.

A cooling circuit to receive a cooling fluid includes at least one shaped cooling pin disposed in the cooling circuit. The at least one shaped cooling pin has a first end and a second end extending along an axis. The first end has a first curved surface defined by a minor diameter and the second end has a second curved surface defined by a major diameter. The first curved surface is to be upstream in the cooling fluid and the minor diameter is less than the major diameter. 1. A cooling circuit adapted to receive a cooling fluid , comprising:at least one shaped cooling pin disposed in the cooling circuit, the at least one shaped cooling pin having a first end and a second end extending along an axis, the first end having a first curved surface defined by a minor diameter and the second end having a second curved surface defined by a major diameter, the first curved surface is configured to be upstream in the cooling fluid and the minor diameter is less than the major diameter.2. The cooling circuit of claim 1 , wherein the first curved surface is spaced apart from the second curved surface by a length.3. The cooling circuit of claim 1 , wherein the first curved surface and the second curved surface are interconnected by a pair of surfaces defined by a pair of planes substantially tangent to a respective one of the first curved surface and the second curved surface.4. The cooling circuit of claim 2 , wherein claim 2 , in cross-section claim 2 , the second curved surface is defined by a first circle having a first center point and the first curved surface is defined by a second circle having a second center point claim 2 , and the length is defined between the first center point and the second center point.5. The cooling circuit of claim 4 , wherein the at least one shaped cooling pin comprises a plurality of shaped cooling pins that are arranged in a pattern that includes at least one row of a first sub-plurality of the plurality of shaped cooling pins and at least ...

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Номер записи: 77
07-01-2021 дата публикации

TURBINE STATOR VANE COMPRISING AN INNER COOLING WALL PRODUCED BY ADDITIVE MANUFACTURING

Номер: US20210003015A1
Автор: BREINING Jean-Luc, HALLOUIN Baptiste, LABERNY Damien
Принадлежит: SAFRAN HELICOPTER ENGINES

A stator vane of a turbine of a gas turbine engine, including an outer platform and an inner platform between which there extends an outer wall forming an outer skin, wherein it includes an inner wall, forming an inner skin, facing the outer wall so as to define an inter-skin cavity between the outer wall and the inner wall, the inner wall including a plurality of cooling orifices for impingement cooling of the outer wall, the outer wall and inner wall being produced by additive manufacturing. 1. A turbine distributor vane of a gas turbine engine , comprising an outer platform and an inner platform between which extends an outer wall , forming an outer skin , wherein said turbine distribution vane includes an inner wall , forming an inner skin , facing the outer wall so as to define an inter-skin cavity between the outer and inner walls , the inner wall including a plurality of cooling orifices for cooling the outer wall by impact , the outer and inner walls being produced by additive manufacturing.2. The vane according to claim 1 , further comprising a hook claim 1 , at the outer platform claim 1 , allowing to connect the vane to a turbine ring claim 1 , and wherein the inner wall is radially superimposed on the front end of the hook.3. The vane according to claim 1 , wherein said vane is made in one piece by additive manufacturing.4. The vane according to claim 1 , wherein the orifices for cooling the inner wall are produced by additive manufacturing.5. The vane according to claim 1 , wherein the outer and inner walls are made of the same material and wherein there is a continuity of material between them.6. The vane according to claim 1 , wherein a junction between the outer and inner walls is formed radially at the outer platform.7. A turbine of a gas turbine engine claim 1 , comprising a plurality of distributor vanes according to .8. A turbomachine claim 7 , comprising a turbine according to .9. A method for manufacturing a distributor vane according to claim ...

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Номер записи: 78
07-01-2021 дата публикации

Turbine rotor blade and gas turbine

Номер: US20210003018A1
Автор: Koichiro Iida, Ryuta Ito
Принадлежит: Mitsubishi Heavy Industries Ltd

A second outer surface (49ab) of a top plate (49) is recessed from a first outer surface (49aa) of the top plate (49) in the direction away from the inner peripheral surface (34a) of a turbine casing (34) so that a step (50) is formed between the second outer surface (49ab) and the first outer surface (49aa). At least part of the discharge opening (53B) of a cooling hole (53) is disposed in the second outer surface (49ab). The cooling hole (53) extends so as to be tilted relative to the second outer surface (49ab) so that the cooling hole (53) discharges a cooling medium to the upstream side of a combustion gas flowing between the second outer surface (49ab) and the inner peripheral surface (34a) of the turbine casing (34).

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Номер записи: 79
07-01-2021 дата публикации

ENGINE COMPONENT WITH SET OF COOLING HOLES

Номер: US20210003020A1
Автор: Brassfield Steven Robert, Gabregiorgish Helen Ogbazion, GARAY Gregory Terrence, PANG Tingfan, Webster Zachary Daniel
Принадлежит:

An apparatus and method an engine component for a turbine engine comprising an outer wall bounding an interior and defining a pressure side and an opposing suction side, with both sides extending between a leading edge and a trailing edge to define a chord-wise direction, and extending between a root and a tip to define a span-wise direction, at least one cooling passage located within the interior, a set of cooling holes having an inlet fluidly coupled to the cooling passage, an outlet located on one of the pressure side or suction side, with a connecting passage fluidly coupling the inlet to the outlet. 130-. (canceled)31. An airfoil for a turbine engine comprising:an outer wall bounding an interior and defining a pressure side and an opposing suction side, with both sides extending between a leading edge and a trailing edge to define a chord-wise direction, and extending between a root and a tip to define a span-wise direction;at least one cooling passage located within the interior;a first set of cooling holes having first inlet fluidly coupled to the cooling passage, a first outlet located on the pressure side, with a first connecting passage having a curvilinear centerline fluidly coupling the first inlet to the first outlet, and the first connecting passage having a portion extending along the suction side;a second set of cooling holes having second inlet fluidly coupled to the cooling passage, a second outlet located on the suction side, with a second connecting passage having a curvilinear centerline fluidly coupling the second inlet to the second outlet, and the second connecting passage having a portion extending along the pressure side; anda third set of cooling holes having a third outlet proximate the leading edge of the outer wall.321. The airfoil of claim wherein the at least one cooling passage is multiple cooling passages separated by an interior wall and the inlet is located along the interior wall.331. The airfoil of claim wherein at least one of ...

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Номер записи: 80
03-01-2019 дата публикации

TURBOMACHINE ROTOR BLADE

Номер: US20190003311A1
Автор: Brittingham Robert Alan
Принадлежит:

The present disclosure is directed to a rotor blade for a turbomachine. The rotor blade includes an airfoil defining a cooling passage and a tip shroud coupled to the airfoil. The tip shroud includes first and second walls that at least partially define a cooling core fluidly coupled to the cooling passage. The rotor blade also includes a plurality of ribs positioned within the cooling core and coupled to the first and second walls. The reinforcing structure includes a plurality of interconnected ribs having a first rib with a first orientation and a second rib with a second orientation. The first and second orientations are different in three spatial dimensions. 1. A rotor blade for a turbomachine , rotor blade comprising:an airfoil defining a cooling passage;a tip shroud coupled to the airfoil, the tip shroud including first and second walls that at least partially define a cooling core fluidly coupled to the cooling passage; anda reinforcing structure positioned within the cooling core and coupled to the first and second walls, the reinforcing structure comprising a plurality of interconnected ribs including a first rib having a first orientation and a second rib having a second orientation, the first and second orientations being different in three spatial dimensions.2. The rotor blade of claim 1 , wherein one of the plurality of ribs extends from the first wall to the second wall.3. The rotor blade of claim 1 , wherein the plurality of ribs are non-uniformly arranged within the cooling core.4. The rotor blade of claim 1 , wherein a pair of the plurality of ribs intersect.5. The rotor blade of claim 1 , wherein the plurality of ribs are uniformly arranged within the cooling core.6. The rotor blade of claim 5 , wherein the plurality of ribs are arranged to form a lattice structure within the cooling core.7. The rotor blade of claim 1 , wherein the reinforcing structure defines a plurality of spaces between the plurality of ribs through which a coolant flows.8. ...

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Номер записи: 81
03-01-2019 дата публикации

COMPONENT FOR A GAS TURBINE ENGINE WITH A FILM HOLE

Номер: US20190003314A1
Автор: Dyson Thomas Earl, Smith Aaron Ezekiel, Zlatinov Metodi Blagoev
Принадлежит:

A component is provided and comprises at least one wall comprising a first and a second surface. At least one film cooling hole extends through the wall between the first and second surfaces and has an outlet region at the second surface. The film cooling hole includes a first expansion section being a side diffusion portion and a second expansion section being a layback diffusion portion, wherein the side diffusion portion is upstream and spaced from the layback diffusion portion. 1. A component for a gas turbine engine comprising:a hot side exposed to a hot air flow;a cool side exposed to a cooling air flow;a film hole passage extending between the cool side and the hot side with an inlet on the cool side and an outlet on the hot side, the film hole passage defining a diameter, the film hole passage further defining a side diffusion portion defining a side diffusion length between a start of the side diffusion portion and the outlet, and a layback diffusion portion defining a layback length between a start of the layback diffusion portion and the outlet, wherein the side diffusion length is greater than the layback diffusion length.2. The component of wherein the side diffusion portion is upstream and spaced from the layback diffusion portion.3. The component of wherein the side diffusion portion defines a side diffusion angle claim 1 , α claim 1 , relative to a centerline for the film hole passage claim 1 , and the side diffusion angle is less than 12.5 degrees.4. The component of wherein the layback diffusion portion defines a layback diffusion angle claim 3 , ß claim 3 , relative to a centerline for the film hole passage claim 3 , and the layback diffusion angle is less than 12 degrees.5. The component of wherein the layback diffusion length is less than 4 times the diameter.6. The component of wherein the layback diffusion length is equal to or less than zero.7. The component of wherein the layback diffusion length is less than four times the diameter and the ...

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Номер записи: 82
03-01-2019 дата публикации

FLUID COOLING SYSTEMS FOR A GAS TURBINE ENGINE

Номер: US20190003315A1
Автор: Erno Daniel Jason, Gerstler William Dwight, Joshi Narendra Digamber
Принадлежит:

A heat exchanger includes an airfoil configured to be positioned in a coolant stream. The airfoil includes a pressure sidewall and a suction sidewall coupled to the pressure sidewall. The suction sidewall and the pressure sidewall define a leading edge and a trailing edge opposite the leading edge. The leading edge defines an impingement zone wherein the coolant stream is configured to impinge the airfoil. The heat exchanger also includes at least one channel defined within the airfoil between the pressure sidewall and the suction sidewall. The at least one channel is at least partially defined within the impingement zone proximate the leading edge. 1. A heat exchanger comprising: a pressure sidewall; and', 'a suction sidewall coupled to said pressure sidewall, said suction sidewall and said pressure sidewall define a leading edge and a trailing edge opposite said leading edge, said leading edge defines an impingement zone wherein the coolant stream is configured to impinge said airfoil; and, 'an airfoil configured to be positioned in a coolant stream, said airfoil comprisingat least one channel defined within said airfoil between said pressure sidewall and said suction sidewall, said at least one channel at least partially defined within the impingement zone proximate said leading edge.2. The heat exchanger in accordance with claim 1 , wherein said at least one channel is configured to receive a fluid stream such that heat is removed from the fluid stream at least in part through the coolant stream impinging on said leading edge.3. The heat exchanger in accordance with claim 1 , wherein said suction sidewall and said pressure sidewall further define a root portion and a tip portion opposite said root portion claim 1 , said at least one channel comprises:an inlet section extending from said root portion to adjacent said tip portion proximate said leading edge; andan outlet section extending from adjacent said tip portion to said root portion such that said at least ...

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Номер записи: 83
03-01-2019 дата публикации

Helical skin cooling passages for turbine airfoils

Номер: US20190003316A1
Автор: Jeffrey T. Morton, San Quach
Принадлежит: United Technologies Corp

An airfoil for a gas turbine engine may comprise an airfoil body having an outer diameter (OD) end extending between a leading edge and a trailing edge and having an ID end located opposite the airfoil body from the OD end. The airfoil body defines a helical skin cooling passage extending between the ID end of the airfoil and the OD end of the airfoil. The airfoil body may further define a main body core. The helical skin cooling passage may thermally shield the main body core.

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Номер записи: 84
03-01-2019 дата публикации

TURBOMACHINE ROTOR BLADE

Номер: US20190003317A1
Автор: Brittingham Robert Alan
Принадлежит:

The present disclosure is directed to a rotor blade that includes an airfoil defining a cooling passage and a tip shroud coupled to the airfoil. The tip shroud and the airfoil define a cooling core in fluid communication with the cooling passage. The cooling core includes a first cooling channel and a second cooling channel. The first cooling channel is radially spaced apart from the second cooling channel. Coolant flows in a first direction through the first cooling channel and in a second direction through the second cooling channel. The first direction is different than the second direction. 1. A rotor blade for a turbomachine , the rotor blade comprising:an airfoil defining a cooling passage; anda tip shroud coupled to the airfoil, the tip shroud and the airfoil defining a cooling core in fluid communication with the cooling passage, the cooling core including a first cooling channel and a second cooling channel, the first cooling channel being radially spaced apart from the second cooling channel,wherein coolant flows in a first direction through the first cooling channel and in a second direction through the second cooling channel, the first direction being different than the second direction.2. The rotor blade of claim 1 , wherein the first direction is opposite of the second direction.3. The rotor blade of claim 1 , wherein the tip shroud includes an interior wall positioned within the cooling core claim 1 , the interior wall at least partially defining the first cooling channel and the second cooling channel.4. The rotor blade of claim 2 , wherein the tip shroud further comprises a fillet wall that partially defines the first cooling channel.5. The rotor blade of claim 2 , wherein the tip shroud further comprises a radially outer wall that partially defines the second cooling channel.6. The rotor blade of claim 5 , wherein the interior wall is radially spaced apart from the radially outer wall.7. The rotor blade of claim 1 , wherein the first cooling ...

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Номер записи: 85
03-01-2019 дата публикации

TURBOMACHINE ROTOR BLADE

Номер: US20190003318A1
Автор: Brittingham Robert Alan, Jones Mark Andrew
Принадлежит:

The present disclosure is directed to a rotor blade for a turbomachine. The rotor blade includes an airfoil having a trailing edge surface and defining a cooling passage. The rotor blade also includes a tip shroud coupled to the airfoil. The tip shroud includes a radially inner surface. The tip shroud defines a cooling core fluidly coupled to the cooling passage. The cooling core includes at least one of a first outlet aperture having a first opening defined by the radially inner surface or a second outlet aperture having a second opening defined by the trailing edge surface of the airfoil. The first or second outlet apertures eject coolant from the cooling core in a direction of a local flow of combustion gases external to the tip shroud. 1. A rotor blade for a turbomachine , the rotor blade comprising:an airfoil including a trailing edge surface, the airfoil defining a cooling passage; anda tip shroud coupled to the airfoil, the tip shroud comprising a radially inner surface, the tip shroud defining a cooling core fluidly coupled to the cooling passage, the cooling core comprising at least one of a first outlet aperture having a first opening defined by the radially inner surface or a second outlet aperture having a second opening defined by the trailing edge surface of the airfoil,wherein the first or second outlet apertures are configured to eject coolant from the cooling core in a direction of a local flow of combustion gases external to the tip shroud.2. The rotor blade of claim 1 , wherein the first outlet aperture is configured to eject coolant from the cooling core through the first opening substantially parallel to a camber line at the radially inner surface.3. The rotor blade of claim 1 , wherein the second outlet aperture is configured to eject coolant from the cooling core through the second opening substantially parallel to a camber line at the trailing edge surface of the airfoil.4. The rotor blade of claim 1 , wherein the tip shroud comprises a ...

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Номер записи: 86
03-01-2019 дата публикации

COOLING CONFIGURATION FOR A GAS TURBINE ENGINE AIRFOIL

Номер: US20190003319A1
Автор: Heneveld Benjamin, Lee Ching-Pang
Принадлежит:

A gas turbine engine airfoil includes an outer wall including a suction side, a pressure side, a leading edge, and a trailing edge, the outer wall defining an interior chamber of the airfoil. The airfoil further includes cooling structure provided in the interior chamber. The cooling structure defines an interior cooling cavity and includes a plurality of cooling fluid outlet holes, at least one of which is in communication with a pressure side cooling circuit and at least one of which is in communication with a suction side cooling circuit. At least one of the pressure and suction side cooling circuits includes: a plurality of rows of airfoils, wherein radially adjacent airfoils within a row define segments of cooling channels. Outlets of the segments in one row align aerodynamically with inlets of segments in an adjacent downstream row such that the cooling channels have a serpentine shape. 1. A gas turbine engine airfoil comprising:an outer wall including a radially inner end, a radially outer end, a suction side, a pressure side, a leading edge, and a trailing edge, the outer wall defining an interior chamber of the airfoil;cooling structure provided in the interior chamber, the cooling structure:located closer to the leading edge of the outer wall than to the trailing edge of the outer wall;defining an interior cooling cavity; andincluding a plurality of cooling fluid outlet holes, at least one of the outlet holes in communication with a pressure side cooling circuit and discharging cooling fluid from the interior cooling cavity of the cooling structure at least partially in a direction toward the leading edge of the outer wall, and at least one of the outlet holes in communication with a suction side cooling circuit and discharging cooling fluid from the interior cooling cavity of the cooling structure at least partially in a direction toward the leading edge of the outer wall;each of the pressure side cooling circuit and the suction side cooling circuit ...

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Номер записи: 87
03-01-2019 дата публикации

TURBOMACHINE ROTOR BLADE

Номер: US20190003320A1
Автор: Balkcum, Brittingham Robert Alan, III James Tyson
Принадлежит:

The present disclosure is directed to a rotor blade for a turbomachine. The rotor blade includes an airfoil defining a cooling passage and a tip shroud coupled to the airfoil. The tip shroud and the airfoil define a cooling core in fluid communication with the cooling passage. The tip shroud including a forward exterior wall, an aft exterior wall spaced apart from the forward exterior wall along an axial direction, a radially inner exterior wall, a radially outer exterior wall spaced apart from the radially inner wall along a radial direction, a pressure side wall, and a suction side wall spaced apart from the pressure side wall along a circumferential direction. The tip shroud further includes first and second interior walls positioned within the cooling core. The first interior wall is non-coplanar with the second interior wall in the axial, radial, and circumferential directions. 1. A rotor blade for a turbomachine , the rotor blade defining an axial direction , a radial direction , and a circumferential direction , the rotor blade comprising:an airfoil defining a cooling passage; anda tip shroud coupled to the airfoil, the tip shroud and the airfoil defining a cooling core in fluid communication with the cooling passage, the tip shroud comprising a forward exterior wall, an aft exterior wall spaced apart from the forward exterior wall along the axial direction, a radially inner exterior wall, a radially outer exterior wall spaced apart from the radially inner wall along the radial direction, a pressure side wall, and a suction side wall spaced apart from the pressure side wall along the circumferential direction, the tip shroud further comprising first and second interior walls positioned within the cooling core, the first interior wall being non-coplanar with the second interior wall in the axial, radial, and circumferential directions.2. The rotor blade of claim 1 , wherein the first or second interior walls are curved.3. The rotor blade of claim 2 , wherein ...

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Номер записи: 88
12-01-2017 дата публикации

Orifice element for turbine stator and/or rotor vanes

Номер: US20170009590A1
Автор: Ulf Nilsson
Принадлежит: SIEMENS AG

An orifice element is adapted to be inserted into a recess formed at an external opening of a channel in a turbine stator or rotor vane, the channel being adapted for leading a cooling fluid through the vane. The orifice element has a mounting part formed of a solid material, and an opening part leaving an opening between a first side of the orifice element and a second side of the orifice element, the second side being opposite to the first side.

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Номер записи: 89
27-01-2022 дата публикации

TURBINE ENGINE BLADE WITH IMPROVED COOLING

Номер: US20220025771A1
Автор: AUZILLON Pierre Guillaume, CARIOU Romain Pierre, DE ROCQUIGNY Thomas Olivier Michel Pierre, ENEAU Patrice, OSTINO Léandre, ROLLINGER Adrien Bernard Vincent, SIMON Vianney, SLUSARZ Michel, Tang Ba-Phuc
Принадлежит:

A turbine blade including a root carrying an impeller terminated by a tip in the form of a squealer tip. This impeller also includes a serpentine median circuit, including a first radial pipe collecting air at the root and that is connected by a first bend to a second radial pipe that is connected by a second bend to a third radial pipe, a cavity under the squealer tip running along the pressure side wall, extending from a central region of the tip to the trailing edge, and a radial central pipe collecting air at the root extending between at least two of the three pipes of the median circuit and directly supplying the cavity under the squealer tip. 114-. (canceled)15. A turbine vane of a turbomachine , for being mounted about an axis of rotation on a rotor disc rotating about an axis of rotation , comprising a root for mounting thereof in a cell of the disc , and a hollow blade extending from the root in a radial spanwise direction and terminating in a top forming a bathtub , the blade comprising a lower surface wall and an upper surface wall , as well as a leading edge , a trailing edge and a top wall delimiting a bottom of the bathtub , and with which the lower surface wall is connected to the upper surface wall , said blade also comprising:a paper clip-type median circuit, including a first radial duct collecting air at the root and which is connected through a first bend to a second radial duct which is connected through a second bend to a third radial duct;an under-bathtub cavity located on the side of the lower surface wall and the top wall and which extends from a central region of the top to the trailing edge;a central radial duct located on the side of the lower surface wall and which collects air at the root and extends between at least two of the three ducts of the median circuit and directly feeds the under-bathtub cavity.16. The vane of claim 15 , wherein one end of the third duct and at least part of the first bend are located between the under- ...

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Номер записи: 90
14-01-2016 дата публикации

GAS TURBINE ENGINE HIGH LIFT AIRFOIL COOLING IN STAGNATION ZONE

Номер: US20160010463A1
Автор: Devore Matthew A., QUACH San
Принадлежит:

An airfoil for a gas turbine engine includes pressure and suction side walls joined to one another at leading and trailing edges. A stagnation line is located on the pressure side wall aft of the leading edge. A cooling passage is provided between the pressure and suction side walls. Forward-facing cooling holes are provided adjacent to the stagnation line on the pressure side wall and oriented toward the leading edge. 1. An airfoil for a gas turbine engine , comprising:pressure and suction side walls joined to one another at leading and trailing edges, a stagnation line located on the pressure side wall aft of the leading edge, a cooling passage provided between the pressure and suction side walls, and forward-facing cooling holes provided adjacent to the stagnation line on the pressure side wall and oriented toward the leading edge.2. The airfoil according to claim 1 , comprising shower head cooling holes clustered about the leading edge claim 1 , the forward-facing cooling holes spaced aft of the shower head cooling holes.3. The airfoil according to claim 2 , wherein the shower head cooling holes include a cluster of three rows of holes extending in a radial direction claim 2 , the three rows including a first row extending along the leading edge and second and third rows respectively arranged adjacent to and on opposing sides of the first row.4. The airfoil according to claim 1 , wherein the forward-facing cooling holes extend from a midspan of the airfoil to a tip.5. The airfoil according to claim 1 , comprising aft-facing cooling holes provided adjacent to the stagnation line on the suction side wall and oriented toward the trailing edge.6. The airfoil according to claim 5 , wherein the aft-facing cooling holes are aft of the stagnation line and oriented toward a tip of the airfoil.7. The airfoil according to claim 1 , wherein the stagnation line overlaps the leading edge.8. The airfoil according to claim 1 , wherein the forward-facing cooling holes are ...

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Номер записи: 91
14-01-2016 дата публикации

Gas turbine engine airfoil leading edge cooling

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

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

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Номер записи: 92
14-01-2016 дата публикации

GAS TURBINE ENGINE COMPONENT WITH TWISTED INTERNAL CHANNEL

Номер: US20160010466A1
Автор: Lamson Scott
Принадлежит:

A gas turbine engine component includes a component body that defines an internal micro-channel that extends in a lengthwise direction along a reference line. The internal micro-channel extends between a first reference position along the reference line and a second reference position along the reference line. The internal micro-channel twists at least 180 with respect to the reference line between the first reference position and the second reference position. 1. A gas turbine engine component comprising:a component body defining an internal micro-channel extending in a lengthwise direction along a reference line, the internal micro-channel extending between a first reference position along the reference line and a second reference position along the reference line, the internal micro-channel twisting at least 180° with respect to the reference line between the first reference position and the second reference position.2. The gas turbine engine component as recited in claim 1 , wherein the internal micro-channel twists at least 360° with respect to the reference line between the first reference position and the second reference position.3. The gas turbine engine component as recited in claim 1 , wherein the internal micro-channel twists multiple full revolutions with respect to the reference line.4. The gas turbine engine component as recited in claim 1 , further including at least one additional internal micro-channel also twisting at least 180° with respect to the reference line between the first reference position and the second reference position.5. The gas turbine engine component as recited in claim 1 , further including at least one additional internal micro-channel that is symmetrically arranged to the internal micro-channel with respect to the reference line.6. The gas turbine engine component as recited in claim 1 , wherein the internal micro-channel is helical.7. The gas turbine engine component as recited in claim 1 , further including a plurality of ...

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Номер записи: 93
14-01-2016 дата публикации

COOLING HOLE FOR A GAS TURBINE ENGINE COMPONENT

Номер: US20160010473A1
Автор: Mongillo Dominic, Spangler Brandon W.
Принадлежит:

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a wall having an internal surface and an outer skin, a cooling hole having an inlet extending from the internal surface and merging into a metering section, and a diffusion section downstream of the metering section that extends to an outlet located at the outer skin. The diffusion section of the cooling hole includes a first side diffusion angle, a second side diffusion angle and a downstream diffusion angle at a downstream surface of the diffusion section, the downstream diffusion angle being less than the first side diffusion angle and the second side diffusion angle. 1. A component for a gas turbine engine , comprising:a wall having an internal surface and an outer skin;a cooling hole having an inlet extending from said internal surface and merging into a metering section, and a diffusion section downstream of said metering section that extends to an outlet located at said outer skin; andwherein said diffusion section of said cooling hole includes a first side diffusion angle, a second side diffusion angle and a downstream diffusion angle at a downstream surface of said diffusion section, said downstream diffusion angle being less than said first side diffusion angle and said second side diffusion angle.2. The component as recited in claim 1 , wherein said wall is part of a vane.3. The component as recited in claim 1 , wherein said wall is part of a blade.4. The component as recited in claim 1 , wherein said wall is part of a blade outer air seal (BOAS).5. The component as recited in claim 1 , wherein said diffusion section includes a first side surface that diverges in a first axial direction from an axis of said metering section and a second side surface that diverges in a second axial direction from said axis.6. The component as recited in claim 5 , wherein said first side surface and said second side surface diverge at said first and ...

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Номер записи: 94
11-01-2018 дата публикации

COOLING SYSTEM FOR GAS TURBINE, GAS TURBINE EQUIPMENT PROVIDED WITH SAME, AND PARTS COOLING METHOD FOR GAS TURBINE

Номер: US20180010520A1
Автор: AI Toshishige, FUKUNAGA Yuya, Iwasaki Tatsuya, Iwasaki Yoshifumi
Принадлежит:

A cooling system includes: a high pressure bleed line configured to bleed high pressure compressed air from a first bleed position of a compressor and to send the air to a first hot part; a low pressure bleed line configured to bleed low pressure compressed air from a second bleed position of the compressor and to send the air to a second hot part; an orifice provided in the low pressure bleed line; a connecting line configured to connect the high pressure bleed line and the low pressure bleed line; a first valve provided in the connecting line; a bypass line configured to connect the connecting line and the low pressure bleed line; and a second valve provided in the bypass line. 120-. (canceled)21. A cooling system for a gas turbine which includes a compressor configured to compress air , a combustor configured to burn a fuel in the air compressed by the compressor to generate a combustion gas , and a turbine driven using the combustion gas , the cooling system for a gas turbine comprising:a high pressure bleed line configured to bleed air from a first bleed position of the compressor and to send the air bled from the first bleed position to a first hot part coming into contact with the combustion gas among parts constituting the gas turbine;a cooler configured to cool air passing through the high pressure bleed line;a low pressure bleed line configured to bleed air at a pressure lower than that of the air which is bled from the first bleed position from a second bleed position of the compressor, to send the air bled from the second bleed position to a second hot part coming into contact with the combustion gas and disposed under a lower pressure environment than the first hot part among the parts constituting the gas turbine, and is not provided with a cooler;a minimum flow rate securing device configured to secure a minimum flow rate of air flowing through the low pressure bleed line while limiting a flow rate of the air flowing through the low pressure bleed ...

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Номер записи: 95
10-01-2019 дата публикации

Mechanical component

Номер: US20190010808A1
Автор: Alexey Mozharov, Andrey SEDLOV, Peter Vitalievich LALETINE, Vladimir Vassiliev
Принадлежит: General Electric Technology GmbH

A mechanical component comprises an internal hollow space and a wall, the wall limiting the hollow space. The mechanical component further comprises a first channel extending inside the wall along a first direction and a second channel extending inside the wall in fluid communication with the internal hollow space and the first channel, serving as a feed channel. A cross-sectional dimension of the first channel is larger than a cross-sectional dimension of the feed channel, and the feed channel tangentially joins into the first channel. A third channel extends inside the wall in fluid communication with the first channel. The third channel extends inside the wall at least essentially parallel to a surface of the wall along at least a part of the extent of the wall in a second direction, and is a near wall cooling channel.

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Номер записи: 96
09-01-2020 дата публикации

Gas turbine engine with vane having a cooling inlet

Номер: US20200011347A1
Автор: Josiah Warren JOINER, Mohamed Musthafa THOPPIL, Monty Lee Shelton, Robert Proctor, Sesha SUBRAMANIAN, Venkateswararao GURRAM, Wojciech Sak
Принадлежит: General Electric Co

An apparatus and method of cooling a hot portion of a gas turbine engine, such as a multi-stage compressor of a gas turbine engine, by reducing an operating air temperature in a space between a seal and a blade post of adjacent stages by routing cooling air through an inlet in the vane, passing the routed cooling air through the vane, and emitting the routed cooling air into the space.

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Номер записи: 97
15-01-2015 дата публикации

COOLED GAS TURBINE ENGINE COMPONENT

Номер: US20150016944A1
Автор: Heffernan Tab M., Kwon Okey, Rhodes Jeffrey F., Vandervaart Peter L.
Принадлежит:

A gas turbine component having a cooling passage is disclosed. In one form, the passage is oriented as a turned passage capable of reversing direction of flow, such as a turned cooling hole. The gas turbine engine component can include a layered structure having cooling flow throughout a region of the component. The cooling hole can be in communication with a space in the layered structure. The gas turbine engine component can be a cast article where a mold can be constructed to produce the cooling hole having a turn. 1. An apparatus comprising:a cooled gas turbine engine component having a wall forming a boundary of an internal passage used for conveyance of a cooling fluid; anda cooling hole extending between a hot-side and a cold-side of the cooled gas turbine engine component having a first end oriented to receive cooling fluid from the internal passage and a second end having an outlet capable of discharging the cooling fluid from the gas turbine engine component, the cooling hole having opposing sides routed along a curvilinear path.2. The apparatus of claim 1 , wherein the cooled gas turbine engine component is a multi-walled cooled component claim 1 , and wherein the internal passage is situated between a hot-side wall and a cold-side wall of an inter-wall passage.3. The apparatus of claim 2 , wherein the curvilinear path of the cooling hole is near a leading edge of the multi-wall cooled component.4. The apparatus of claim 3 , wherein the inter-wall passage includes a plurality of pedestals claim 3 , and wherein the cooling hole is substantially free of pedestals.5. The apparatus of claim 1 , wherein the cooled gas turbine engine component includes a construction to permit transpiration cooling claim 1 , and wherein the cooling hole includes a plurality of cooling holes in flow communication with a transpiration cooling passage.6. The apparatus of claim 5 , wherein the plurality of cooling holes include outlets in a leading edge region of the cooled gas ...

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Номер записи: 98
15-01-2015 дата публикации

COOLED TURBINE GUIDE VANE OR BLADE FOR A TURBOMACHINE

Номер: US20150016961A1
Автор: Shepherd Andrew
Принадлежит: SIEMENS AKTIENGESELLSCHAFT

A turbine airfoil for a turbomachine is provided. The airfoil includes a suction side wall and a pressure side wall bordering an airfoil cavity, which receives a cooling fluid for cooling the airfoil. The suction side wall includes one or more protrusions extending inside the cavity. The number of protrusions on the suction side may be higher than the number of protrusions on the pressure side. The density of protrusions on the suction side may be higher than the density of protrusions on the pressure side and/or the surface of protrusions on the suction side may be larger than the surface of protrusions on the pressure side, so that heat transfer from the suction side to the cooling fluid is higher compared to heat transfer from the pressure side to the cooling fluid during operation of the turbomachine. 1. A turbine airfoil comprising a blade or a vane for a turbomachine , the airfoil comprisinga suction side wall and a pressure side wall bordering an airfoil cavity, which is adapted to be flowed through by a cooling fluid for cooling of the side walls and therefore of the airfoil,wherein the suction side wall comprises at least one protrusion extending therefrom inside the airfoil cavity, wherein the number of the at least one protrusion on the suction side wall is higher than the number of protrusions on the pressure side wall, the density of the at least one protrusion on the suction side wall is higher than the density of protrusions on the pressure side wall and/or the surface of the at least one protrusion on the suction side wall is larger than the surface of protrusions on the pressure side wall, so that the heat transfer from the suction side wall to the cooling fluid is higher compared to the heat transfer from the pressure side wall to the cooling fluid during the operation of the turbomachine such that an excess of the heat transfer from the suction side wall is generated.2. The turbine airfoil according to claim 1 , wherein at least one of the ...

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Номер записи: 99
15-01-2015 дата публикации

IMPINGEMENT COOLING OF TURBINE BLADES OR VANES

Номер: US20150016973A1
Автор: Mugglestone Jonathan
Принадлежит: SIEMENS AKTIENGESELLSCHAFT

A turbine assembly having a hollow aerofoil and impingement device, the aerofoil having a first side wall from leading to trailing edge and a cavity arranged a distance to an inner surface of the cavity for impingement cooling and a flow channel for cooling medium from the leading to trailing edge, the impingement device has first and second pieces arranged side by side, the second piece downstream of the first forming a first flow passage providing passage from one side of the aerofoil towards an opposite side. A blocking element is arranged in the flow channel between the second piece and first side wall at a suction side for blocking flow of cooling medium from leading to trailing edge denying access to a section of the flow channel downstream of the blocking element while directing cooling medium in the first flow passage away from the suction side towards pressure side. 1. A turbine assembly comprising: wherein the hollow aerofoil has at least a first side wall extending from a leading edge towards a trailing edge of the hollow aerofoil and at least a cavity in which in an assembled state of the at least one impingement device in the hollow aerofoil the at least one impingement device is arranged with a predetermined distance in respect to an inner surface of the cavity for impingement cooling of the at least one inner surface and to form a flow channel for a cooling medium extending from the leading edge towards the trailing edge and', 'wherein the at least one impingement device comprises a first piece and a second piece being arranged side by side in an axial direction with the second piece being located viewed in the axial direction downstream of the first piece and with an axial distance in respect to each other forming a first flow passage providing a passage from one side of the aerofoil towards an opposite side of the aerofoil, and, 'a basically hollow aerofoil and at least an impingement device,'}at least a first blocking element, which is arranged in ...

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Номер записи: 100