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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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Применить Всего найдено 14529. Отображено 100.

02-02-2012 дата публикации

Blade outer air seal and repair method

Номер: US20120027574A1

Автор: Philip R. Belanger, Susan M. Tholen

Принадлежит: United Technologies Corp

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

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Номер записи: 1

02-02-2012 дата публикации

Turbine nozzle segment and method of repairing same

Номер: US20120027617A1

Автор: Jose Abiel Garza, Marcin Trajer, Marcin Zak, Pawel Mankowski, Wayne Ray Grady

Принадлежит: General Electric Co

A method is provided for repairing a metallic turbine component which includes at least two airfoils interconnected by a mid-span shroud. The method includes: (a) applying a reinforcement plate to the mid-span shroud; (b) applying braze material to at least a portion of a perimeter of the reinforcement plate; (c) heating the component to melt and flow the braze material between the reinforcement plate and the mid-span shroud; and (d) allowing the braze material to cool and solidify so as to bond the reinforcement plate to the mid-span shroud.

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Номер записи: 2

01-03-2012 дата публикации

Casing body through which hot gases can flow and comprising an inner heat shield

Номер: US20120047905A1

Автор: Marc Rauch, Peter Oskar Baldischweiler

Принадлежит: Alstom Technology AG

A casing body for a hot gas flow includes an outer casing body having a hot gas side with a precisely prepared locating surface. A pin-type retainer is disposed on the locating surface, and an inner heat shield is disposed at a distance from the hot gas side of the outer casing body and fastened to the retainer.

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Номер записи: 3

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: 4

12-04-2012 дата публикации

Turbofan jet engine

Номер: US20120087787A1

Автор: Dewain Ray Brown

Принадлежит: Individual

There is a turbofan jet engine including an engine core. The engine core includes a fan and a compressor. The engine core includes a combustion chamber and a turbine functionally coupled to the compressor. The engine core includes a nozzle in fluid communication with the turbine. The turbofan jet engine includes a nacelle. The nacelle includes a forward extension proximate the fan and extending forward therefrom. The forward extension is funnel shaped to impart radial momentum to intake air during operation. The nacelle includes a vortex device disposed inside the forward extension and shaped to impart angular momentum to intake air. The vortex device includes a fixed blade extending from the interior of the forward extension and set at a rotational angle. The vortex device is shaped and positioned to direct intake air substantially perpendicular to the blades of the fan.

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Номер записи: 5

31-05-2012 дата публикации

Gas turbine of the axial flow type

Номер: US20120134779A1

Автор: Alexander Anatolievich Khanin, Valery Kostege

Принадлежит: Individual

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

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Номер записи: 6

26-07-2012 дата публикации

Axial flow turbine

Номер: US20120189441A1

Автор: Brian Robert Haller, Gursharanjit Singh

Принадлежит: Alstom Technology AG

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

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Номер записи: 7

09-08-2012 дата публикации

Core runout ceiling for turbine components

Номер: US20120201662A1

Автор: Kevin J. Ryan, Paul C. McClay, Paul M. Lutjen, Paul S. Bean

Принадлежит: Individual

A method of closing off a mold plug opening in a turbine component includes the steps of inserting a weld member into an opening to be closed, and to abut a necked portion within a passage leading from the opening. Heat is applied to the weld member, such that a surface of the weld member in contact with the necked portion liquefies, and such that the weld member adheres to the necked portion, closing off the opening. The weld member and application of heat are selected such that the entirety of the weld member does not liquefy, but remains in the opening, without ever having liquefied. A turbine component formed by the method is also disclosed.

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Номер записи: 8

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: 9

06-09-2012 дата публикации

Airfoil core shape for a turbomachine component

Номер: US20120224954A1

Автор: Julianne Edna Baer, Umesh Garg

Принадлежит: General Electric Co

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

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Номер записи: 10

20-09-2012 дата публикации

Method for producing a metal reinforcement for a turbine engine blade

Номер: US20120233859A1

Автор: Bernard Jose Michel Cattiez, Jerome Guinois, Philippe Marolle, Stephane Andre Leveque, Thierry Jean Emile Flesch

Принадлежит: SNECMA SAS

A method for making a metal reinforcement for the leading edge or trailing edge of a turbine engine blade, including: positioning a preform using an equipment positioning the preform in a position such that the preform, at one end thereof, has an area which is capable of receiving a filler metal; and, after the positioning, constructing a base for the metal reinforcement by hard-surfacing with filler metal in the area, in the form of metal beads.

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Номер записи: 11

04-10-2012 дата публикации

Plasma Actuated Vortex Generators

Номер: US20120248072A1

Автор: Dennis B. Finley, Paul D. McClure, Sergey Macheret

Принадлежит: Lockheed Martin Corp

A plasma-actuated vortex generator arrangement includes a plurality of spaced-apart vortex generators, and a plasma actuator distributed amongst the plurality of vortex generators.

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Номер записи: 12

01-11-2012 дата публикации

High area ratio turbine vane

Номер: US20120275922A1

Автор: Eunice Allen-Bradley, Norbert Huebner, Thomas J. Praisner

Принадлежит: Individual

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

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Номер записи: 13

20-12-2012 дата публикации

Method for making a metal reinforcement for the blade of a turbine engine

Номер: US20120317810A1

Автор: Dominique Magnaudeix, Gilles Charles Casimir Klein, Philippe Marolle, Stephane Andre Leveque

Принадлежит: SNECMA SAS

A method for making a metal reinforcement for the leading edge or trailing edge of the blade of a turbine engine that includes making a metal insert defining the base of the metal reinforcement; positioning the metal insert at the end of a blank of a shaping tool, the blank repeating the shape of the turbine-engine blade; shaping a planar metal sheet on the metal insert and the blank of the shaping tool using a superplastic hot-shaping method.

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Номер записи: 14

20-12-2012 дата публикации

Plug assembly for blade outer air seal

Номер: US20120319360A1

Автор: Ken F. Blaney, Paul M. Lutjen

Принадлежит: Individual

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

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Номер записи: 15

10-01-2013 дата публикации

Gas turbine shroud arrangement

Номер: US20130008176A1

Автор: Jun Shi, Kevin E. Green, Paul F. Croteau

Принадлежит: United Technologies Corp

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

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Номер записи: 16

07-02-2013 дата публикации

Resistance weld repairing of casing flange holes

Номер: US20130032578A1

Автор: Gregory Christopher Firestone, Manish Deepak Dighe, Timothy Joseph Trapp

Принадлежит: General Electric Co

A method for repairing a degraded bolt hole in a casing flange by reaming and removing at least some corrosion on an inside and around the hole to form a reamed hole, mounting the flange to float relative to upper and lower electrodes of a welding machine, radially and axially clamping an area of the flange surrounding the reamed hole, placing upper and lower filler slugs in the reamed hole, placing the electrodes against upper and lower filler slugs and applying a welding current through the electrodes while applying pressure to the filler slugs with the electrodes and resistively heating and melting the filler slugs to form a weldment, and pulsing the welding current on and off. Pulsing may be performed with progressively increasing amounts of current. In situ tempering under the pressure of the electrodes may be performed on a substantially liquid pool formed by the welding current.

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Номер записи: 17

14-02-2013 дата публикации

Vortex generators

Номер: US20130037657A1

Автор: Robert E. Breidenthal

Принадлежит: Ramgen Power Systems LLC

A vortex generator, or an array of vortex generators, for attenuating flow separation during flow of fluid over a surface. Vortex generators include a base with a forward end and a leading edge extending outward and rearward from the forward end to an outward end. The leading edge includes a first angular discontinuity at a height H 1 above the base, and a second angular discontinuity at a height H 2 above the base. The vortex generator(s) are configured for generating, adjacent a surface, at least two (2) vortices V 1 and V 2 in a fluid, and turning the outermost generated vortice toward the surface over which the fluid is passing.

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Номер записи: 18

14-02-2013 дата публикации

Stator for supersonic compressor

Номер: US20130039748A1

Автор: Shawn P. Lawlor

Принадлежит: Ramgen Power Systems LLC

A stator. The stator may be used in a supersonic compressor that utilizes a rotor to deliver a gas at supersonic conditions to the stator. The stator includes a plurality of aerodynamic ducts that have converging and diverging portions, for deceleration of gas to subsonic conditions and then for expansion of subsonic gas, to change kinetic energy of the gas to static pressure. The aerodynamic ducts include structures for changing the effective contraction ratio to enable starting even when the aerodynamic ducts are designed for high pressure ratios, and structures for boundary layer control.

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Номер записи: 19

14-02-2013 дата публикации

Turbomachine component having an airfoil core shape

Номер: US20130039771A1

Автор: Benjamin Michael Check, Craig Allen Bielek, Hal Feaster Pollard, Sivaraman Vedhagiri

Принадлежит: General Electric Co

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

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Номер записи: 20

28-02-2013 дата публикации

Turbine shroud segment

Номер: US20130052007A1

Автор: Eric Durocher, Guy Lefebvre

Принадлежит: Pratt and Whitney Canada Corp

A turbine shroud segment is metal injection molded (MIM) about a core to provide a composite structure. In one aspect, the core is held in position in an injection mold and then the MIM material is injected in the mold to form the body of the shroud segment about the core. Any suitable combination of materials can be used for the core and the MIM shroud body, each material selected for its own characteristics. The core may be imbedded in the shroud platform to provide a multilayered reinforced platform, which may offer resistance against crack propagation.

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Номер записи: 21

21-03-2013 дата публикации

Vibration damping blade for fluid

Номер: US20130071251A1

Автор: Hiroaki Hattori, Hisayuki Motoi, Jose Javier Bayod Relancio

Принадлежит: IHI Corp

The vibration damping blade for fluid of the present invention has an integrally formed wedge damper, in which a thickness h(x) at a distance x from an imaginary line outside of an outer edge is h(x)=εx n (where ε is a positive constant, and n is a real number of 1 or more). As a result, it is possible to offer a vibration damping blade for fluid which can be easily manufactured, and which obtains damping effects across a wide range of frequency regions without disturbing the flow of fluid.

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Номер записи: 22

28-03-2013 дата публикации

Turbocharger variable-nozzle assembly with vane sealing arrangement

Номер: US20130078082A1

Автор: Chris Groves, Denis Jeckel, Dominique Petitjean, Emmanuel Severin, Karl-Heinz Bauer, Nicolas Morand, Philippe Arnold, Pierre Barthelet

Принадлежит: Honeywell International Inc

A variable-nozzle assembly for a turbocharger includes a generally annular nozzle ring and an array of vanes rotatably mounted to the nozzle ring such that the vanes can be pivoted about their axes for regulating exhaust gas flow to the turbine wheel. A unison ring engages vane arms that are affixed to axles of the vanes, such that rotation of the unison ring causes the vanes to pivot between a closed position and an open position. The vanes have proximal ends that are adjacent a face of the nozzle ring. A vane sealing member is supported on the nozzle ring and has a portion disposed between the proximal ends of the vanes and the face of the nozzle ring. The unison ring includes cams that engage cam followers. Rotational movement of the unison ring causes the cam followers to be moved axially and thereby urge the vane sealing member against the proximal ends of the vanes.

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Номер записи: 23

02-05-2013 дата публикации

Turbine of a turbomachine

Номер: US20130104550A1

Автор: Craig Allen Bielek, Gunnar Leif Siden, Paul Kendall Smith, Thomas William Vandeputte

Принадлежит: General Electric Co

A turbine of a turbomachine is provided and includes opposing endwalls defining a pathway for a fluid flow and a plurality of interleaved blade stages and nozzle stages arranged axially along the pathway. The plurality of the blade stages includes a last blade stage at a downstream end of the pathway and a next-to-last blade stage upstream from the last blade stage. The plurality of the nozzle stages includes a last nozzle stage between the last blade stage and the next-to-last blade stage and a next-to-last nozzle stage upstream from the next-to-last blade stage. At least one of the next-to-last blade stage and the next-to-last nozzle stage includes aerodynamic elements configured to interact with the fluid flow and to define a throat distribution producing a tip strong pressure profile in the fluid flow.

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Номер записи: 24

09-05-2013 дата публикации

Fan rotor blade and fan

Номер: US20130111908A1

Автор: Hiroshi Kuroki, Kenro Obuchi, Rintaro Kajiwara, Takeshi Murooka

Принадлежит: IHI Corp

A fan rotor blade includes a blade body constructed of a composite material, a blade root constructed of the composite material, and a sheath attached to a leading edge of the blade body. The sheath includes a sheath main body and a pair of joint flanges, and is segmented into a sheath base segment and a sheath top segment. The sheath top segment has a longer length than a length of the sheath base segment along a span direction. A sheath length of the sheath main body at an assumed impact position with an obstacle is not shorter than 10% chord and not longer than 60% chord. A sheath length of the sheath along an end edge of the fan rotor blade is not shorter than 40% chord. The fan rotor blade possesses sufficient impact resistance and can be simplified and light-weighted.

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Номер записи: 25

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: 26

13-06-2013 дата публикации

Gas turbine engine with fan variable area nozzle for low fan pressure ratio

Номер: US20130145745A1

Автор: Frederick M. Schwarz, Gregory A. Kohlenberg, Sean P. Zamora

Принадлежит: Individual

A gas turbine engine includes a fan section with twenty (20) or less fan blades and a fan pressure ratio less than about 1.45.

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Номер записи: 27

20-06-2013 дата публикации

Low-ductility turbine shroud

Номер: US20130156556A1

Автор: Jason David Shapiro, Joshua Brian JAMISON, Michael John Franks, Roger Lee Doughty, Samuel Ross Rulli

Принадлежит: General Electric Co

A shroud segment for a gas turbine engine, the shroud segment constructed from a composite material including reinforcing fibers embedded in a matrix, and having a cross-sectional shape defined by opposed forward and aft walls, and opposed inner and outer walls, the walls extending between opposed first and second end faces, wherein the inner wall defines an arcuate inner flowpath surface; and wherein a compound fillet is disposed at a junction between first and second ones of the walls, the compound fillet including first and second portions, the second portion having a concave curvature extending into the first one of the walls.

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Номер записи: 28

20-06-2013 дата публикации

Annular gas turbine engine case and method of manufacturing

Номер: US20130156558A1

Автор: Barry Barnett, Czeslaw Wojtyczka, Steven Hunt

Принадлежит: Pratt and Whitney Canada Corp

The method is used for making an annular gas turbine engine case from a preform. The method comprises comprising flowforming at least one section of the preform, and then outwardly bending at least one portion of the perform.

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Номер записи: 29

27-06-2013 дата публикации

Method for manufacturing a hollow vane

Номер: US20130164145A1

Автор: Thierry Georges Paul Papin

Принадлежит: SNECMA SAS

A method for manufacturing a hollow structural turbomachine vane, the method including forming a first cavity in a first face of a first block; assembling by diffusion bonding the first block and a second block, the first face of the first block being positioned facing a second face of the second block, the first cavity thus forming a closed cavity; machining the block resulting from the assembly of the first block and the second block so as to obtain a vane including the closed cavity.

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Номер записи: 30

01-08-2013 дата публикации

Aluminum airfoil

Номер: US20130195674A1

Автор: Daniel A. Bales, Thomas J. Watson

Принадлежит: Individual

A method of making an aluminum airfoil includes brazing a first airfoil piece and a second airfoil piece together using a braze material that includes an element selected from magnesium and zinc, to form a braze joint between the first airfoil piece and the second airfoil piece. At least one of the first airfoil piece or the second airfoil piece has an aluminum alloy composition that includes greater than 0.8% by weight of zinc.

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Номер записи: 31

15-08-2013 дата публикации

Anti-Rotation Stator Segments

Номер: US20130209248A1

Автор: Jonathan J. Earl, Mark David RING

Принадлежит: Pratt and Whitney Co Inc

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

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Номер записи: 32

29-08-2013 дата публикации

Blade body and rotary machine

Номер: US20130224034A1

Автор: Eisaku Ito, Hitoshi Kitagawa, Susumu Wakazono, Yasuro Sakamoto

Принадлежит: Mitsubishi Heavy Industries Ltd

The blade body of the present invention is provided with a main body which has a dorsal face and a ventral face and also provided with a trailing edge portion which connects the dorsal face to the ventral face with a continuous curved face. The curved face of the trailing edge portion is gradually decreased in curvature radius from one of the dorsal face and the ventral face toward the rear end portion which is positioned most downstream in a direction at which a fluid flows, decreased to the greatest extent in curvature radius at the rear end portion, thereafter, gradually increased in curvature radius from the rear end portion toward the other of the dorsal face and the ventral face and arrives at the other of the dorsal face and the ventral face.

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Номер записи: 33

24-10-2013 дата публикации

Turbine engine stator and method of assembly of the same

Номер: US20130280054A1

Автор: Lewis J. HOLMES

Принадлежит: Rolls Royce PLC

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

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Номер записи: 34

07-11-2013 дата публикации

Door for thrust reverser of an aircraft nacelle

Номер: US20130292490A1

Автор: Denis Guillois, Loic Chapelain

Принадлежит: Aircelle SA

A door for a thrust reverser of a nacelle of an aircraft being pivotally amounted on a fixed structure of the nacelle, in particular, the door being fitted with deflectors deflecting air flow is disclosed. The deflectors are arranged at an upstream end of the door and mounted such that they can move in a deflection plane perpendicular to the plane of the door. Each deflector is associated at its ends with an articulation arm capable of rotating about a pivot axis perpendicular to the deflection plane, allowing the deflectors to move in a straight line in the deflection plane. The present disclosure also relates to a thrust reverser system including the door and a fixed structure on which the door is pivotally mounted between a closing position and an open position.

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Номер записи: 35

07-11-2013 дата публикации

Shaped rim cavity wing surface

Номер: US20130294897A1

Автор: Eric A. Grover

Принадлежит: United Technologies Corp

A shaped rim cavity wing includes an upper surface and a lower surface. The lower surface has a geometric shape to control the separation of airflow as it passes around the lower surface to the top surface. A point of maximum extent defines the boundary between the upper surface and the lower surface, wherein the point of maximum extent defines a corner that that separates airflow from the shaped rim cavity rim and creates a flow re-circulation adjacent to the top surface of the shaped rim cavity wing.

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Номер записи: 36

05-12-2013 дата публикации

Protecting operating margin of a gas turbine engine

Номер: US20130319009A1

Автор: Wayne P. PARENTE

Принадлежит: Individual

A method of protecting operating margin of the gas turbine engine includes calculating an aerodynamic distortion of air entering an inlet of a gas turbine engine that has a compressor section with variable vanes that are movable subject to a control parameter. The control parameter is selectively modified in response to the aerodynamic distortion.

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Номер записи: 37

09-01-2014 дата публикации

Turbomachine with variable-pitch vortex generator

Номер: US20140010638A1

Автор: Agnes Claire Marie Pesteil, Vincent Paul Gabriel Perrot

Принадлежит: SNECMA SAS

The present invention relates to a turbomachine comprising at least one bladed disk, be it mobile or static, and vortex generators ( 17 ) positioned upstream of the blading ( 1 ) of said disk, wherein the vortex generators ( 17 ) are of variable pitch.

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Номер записи: 38

30-01-2014 дата публикации

Anti-fire seal assembly and nacelle comprising such a seal

Номер: US20140026582A1

Автор: Alexandre Phi, Stephane Beilliard

Принадлежит: Aircelle SA

A seal assembly for a turbojet nacelle includes a first end fastened to a first structure, and a second end contacting against a bearing zone of a second structure. The seal assembly further includes a plurality of adjacent blades arranged along the first end and extending longitudinally and perpendicularly thereto. In particular, a portion of the blades has an accordion-shaped structure.

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Номер записи: 39

06-02-2014 дата публикации

Flow discharge device

Номер: US20140033733A1

Автор: Kevin Mark Britchford, Pamela Joan PHILLIPSON, Vincent Benoit SAVARIN

Принадлежит: Rolls Royce PLC

A bleed flow discharge device ( 136 ) adapted to discharge a bleed fluid flow into a main fluid flow, wherein the bleed flow discharge device comprises an outer wall ( 135 ) defining a passage ( 137 ) for the bleed fluid flow, the outer wall comprising a wave-shaped edge ( 139 ) where the bleed fluid flow meets the main fluid flow.

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Номер записи: 40

06-02-2014 дата публикации

Anti-rotation lug for a gas turbine engine stator assembly

Номер: US20140037442A1

Автор: Neil L. Tatman, Richard K. Hayford

Принадлежит: Individual

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

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Номер записи: 41

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: 42

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: 43

06-03-2014 дата публикации

System and method to improve performance of a compressor device comprising variable diffuser vanes

Номер: US20140064920A1

Автор: Dale Eugene HUSTED, Timothy Daniel HILGART

Принадлежит: Dresser LLC

Embodiments of a system and method can modify the position of diffuser vanes to improve performance of a compressor device, e.g., a centrifugal compressor. These embodiments form a feedback loop to manage the position of the diffuser vanes relative to one or more operating parameters on the compressor device. In one embodiment, the system and method measure input power with the diffuser vanes at a first position and a second position. Changes in the input power will identify other positions for the diffuser vanes that optimize performance of the compressor device, e.g., to reduce power consumption and to achieve and maintain peak compressor efficiency within the entire operating envelope of the compressor device.

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Номер записи: 44

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: 45

06-01-2022 дата публикации

SEAL ASSEMBLY WITH REDUCED PRESSURE LOAD ARRANGEMENT

Номер: US20220003126A1

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

Принадлежит:

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

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Номер записи: 46

06-01-2022 дата публикации

SEALING BETWEEN A ROTOR DISC AND A STATOR OF A TURBOMACHINE

Номер: US20220003127A1

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

Принадлежит:

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

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Номер записи: 47

07-01-2016 дата публикации

GAS TURBINE ENGINE COMPONENT MANUFACTURING METHOD AND CORE FOR MAKING SAME

Номер: US20160001354A1

Автор: Page Richard H., Pinero Hector M.

Принадлежит:

A method of manufacturing a gas turbine engine component includes providing a core having a brittle feature, supporting the feature with a first meltable material, arranging the core with the first meltable material in a first mold, and surrounding the core and the first meltable material with a second meltable material to provide a component shape. The method also includes coating the second meltable material with a refractory material to produce a second mold, removing the first and second meltable material, and casting a component in the second mold. 1. A method of manufacturing a gas turbine engine component comprising:providing a core having a brittle feature;supporting the feature with a first meltable material;arranging the core with the first meltable material in a first mold;surrounding the core and the first meltable material with a second meltable material to provide a component shape;coating the second meltable material with a refractory material to produce a second mold; removing the first and second meltable material; andcasting a component in the second mold.2. The method according to claim 1 , wherein the core and feature are constructed from ceramic.3. The method according to claim 2 , wherein the feature has a thickness of less than 0.013 inch and a width of greater than 0.100 inch.4. The method according to claim 1 , wherein the core is an airfoil trailing edge core.5. The method according to claim 4 , wherein the trailing edge core has a thickness of less than 0.013 inch and a width of greater than 0.100 inch claim 4 , and the core includes an integral adjacent core structure that has a thickness of greater than 0.013 inch.6. The method according to claim 5 , wherein the airfoil trailing edge core has multiple holes claim 5 , and the supporting step includes having the first meltable material extend through the holes.7. The method according to claim 5 , wherein the supporting step includes having the first meltable material adjoin the adjacent ...

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Номер записи: 48

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: 49

05-01-2017 дата публикации

BULGED NOZZLE FOR CONTROL OF SECONDARY FLOW AND OPTIMAL DIFFUSER PERFORMANCE

Номер: US20170002670A1

Автор: Bhaumik Soumyik Kumar, Chouhan Rohit

Принадлежит:

A turbine nozzle disposed in a turbine includes a suction side extending between a leading edge of the nozzle and a trailing edge of the turbine nozzle in an axial direction and transverse to a longitudinal axis of the turbine nozzle, and extending a height of the nozzle in a radial direction along the longitudinal axis, a pressure side disposed opposite the suction side and extending between the leading edge of the turbine nozzle and the trailing edge of the turbine nozzle in the axial direction, and extending the height of the nozzle in the radial direction, and a bulge disposed on the suction side of the nozzle protruding relative to the other portion of the suction side in a direction transverse to a both the radial and axial directions. 1. A turbine nozzle configured to be disposed in a turbine comprising:a suction side extending between a leading edge of the turbine nozzle and a trailing edge of the turbine nozzle in an axial direction and transverse to a longitudinal axis of the turbine nozzle, and extending a height of the turbine nozzle in a radial direction along the longitudinal axis;a pressure side disposed opposite the suction side and extending between the leading edge of the turbine nozzle and the trailing edge of the turbine nozzle in the axial direction, and extending the height of the nozzle in the radial direction; anda bulge disposed on the suction side of the turbine nozzle protruding relative to the other portion of the suction side in a direction transverse to both the radial and axial directions.2. The turbine nozzle of claim 1 , wherein the bulge begins to protrude at a starting height at a first percentage of the height of the nozzle claim 1 , reaches a maximum protrusion at a second percentage of the height of the nozzle claim 1 , and ceases to protrude at an ending height at a third percentage of the height of the nozzle.3. The turbine nozzle of claim 2 , wherein the first percentage of the height of the nozzle is between about 0% and ...

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Номер записи: 50

05-01-2017 дата публикации

AXIAL TRANSFER TUBE

Номер: US20170002671A1

Автор: Hagan Benjamin F., Jacques Jeffrey Michael, Waite Ryan Alan

Принадлежит:

What is described is a transfer tube for use with an airfoil of a gas turbine engine coupled to a platform. The transfer tube includes a main body having a first end defining an inlet configured to receive a flow of fluid and a second end defining an outlet for the flow of fluid, the main body having a curved section. The transfer tube also includes a first mating face coupled to the first end of the main body. The transfer tube also includes a second mating face coupled to the second end of the main body. At least one of the first mating face or the second mating face is configured to be coupled to a platform body of the outer diameter platform. 1. A transfer tube for use with an airfoil of a gas turbine engine coupled to a platform , comprising:a main body having a first end defining an inlet configured to receive a flow of fluid and a second end defining an outlet for the flow of fluid, the main body having a curved section;a first mating face coupled to the first end of the main body; anda second mating face coupled to the second end of the main body,wherein at least one of the first mating face or the second mating face is configured to be coupled to a platform body of the platform.2. The transfer tube of claim 1 , wherein the curved section has an angle of at least 20 degrees.3. The transfer tube of claim 2 , wherein the angle is between 80 degrees and 100 degrees.4. The transfer tube of claim 1 , further comprising a first flange coupled to the first end and a second flange coupled to the second end claim 1 , the first flange defining the first mating face and the second flange defining the second mating face.5. The transfer tube of claim 1 , wherein the transfer tube is configured to allow fluid to flow between a hook of the platform and the platform body of the platform.6. The transfer tube of claim 5 , wherein the platform is an outer diameter platform claim 5 , the first mating face is configured to be coupled to the hook such that the inlet receives the ...

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Номер записи: 51

05-01-2017 дата публикации

Turbine shroud with clamped flange attachment

Номер: US20170002674A1

Автор: Daniel K. Vetters, Robert J. Shinavski, Todd Engel

Принадлежит: Rolls Royce High Temperature Composites Inc, Rolls Royce North American Technologies Inc

A turbine engine including a turbine shroud for positioning radially outside of blades of the turbine rotor. The turbine shroud includes a carrier, a retention assembly, and a blade track. The blade track is clamped by the retention assembly, and the retention assembly is supported by the carrier.

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Номер записи: 52

05-01-2017 дата публикации

ELECTRIC ACTUATOR FOR ENGINE CONTROL

Номер: US20170002679A1

Автор: Balsiger Derick S., Bloxham Keith

Принадлежит:

An electric actuator for control of an engine includes an electric motor coupled to a drive shaft that extends to align a gear interface of the electric actuator with a variable geometry adjustment interface of the engine. A position feedback shaft extends coaxially with respect to the drive shaft. The position feedback shaft is coupled to an output shaft of the gear interface at a gear interface end of the position feedback shaft. A rotational position sensor is coupled to a motor end of the position feedback shaft proximate the electric motor. The drive shaft and the position feedback shaft are sized to position an output ring gear of the output shaft in contact with the variable geometry adjustment interface within a casing of the engine and to further position the electric motor and the rotational position sensor external to the casing of the engine. 1. An electric actuator for control of an engine , the electric actuator comprising:an electric motor coupled to a drive shaft that extends to align a gear interface of the electric actuator with a variable geometry adjustment interface of the engine;a position feedback shaft that extends coaxially with respect to the drive shaft, wherein the position feedback shaft is coupled to an output shaft of the gear interface at a gear interface end of the position feedback shaft; anda rotational position sensor coupled to a motor end of the position feedback shaft proximate the electric motor, wherein the drive shaft and the position feedback shaft are sized to position an output ring gear of the output shaft in contact with the variable geometry adjustment interface within a casing of the engine and to further position the electric motor and the rotational position sensor external to the casing of the engine.2. The electric actuator according to claim 1 , further comprising a retracting mechanism configured to selectively retract the drive shaft and a portion of the gear interface to decouple the drive shaft from the ...

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Номер записи: 53

05-01-2017 дата публикации

Nacelle compression rods

Номер: US20170002684A1

Автор: Stuart J. Byrne

Принадлежит: Rohr Inc

A compression rod may include a plunger and a spring. A proximal end and a distal end of the compression rod may contact engagement features in a core cowl of a gas turbine engine. The compression rod may transmit loads between halves of the core cowl. The spring may cause the plunger to extend and contract in response to vibrations or other relative movement between halves of the core cowl.

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Номер записи: 54

05-01-2017 дата публикации

Guide vane of a gas turbine engine, in particular of an aircraft engine

Номер: US20170002685A1

Автор: Predrag Todorovic

Принадлежит: Rolls Royce Deutschland Ltd and Co KG

A guide vane of a gas turbine engine, in particular of an aircraft engine, which has a pressure-side wall, a suction-side wall, a guide vane root, a guide vane tip, a guide vane leading edge area that is impinged by a cooling air flow of a cooling system, a guide vane trailing edge area that is facing away from the guide vane leading edge area, and at least one channel for conducting a fluid to be cooled arranged in an internal space of the guide vane. At that, during operation of the gas turbine engine, a first part of the cooling air flow flows around a pressure-side wall, and a second part of the cooling air flow flows around the suction-side wall, and a third part of the cooling air flow flows through the internal space including the channel. What is further suggested is a gas turbine engine with at least one such static guide vane.

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Номер записи: 55

05-01-2017 дата публикации

Unducted propeller turboshaft engine provided with a reinforcing shell integrating pipe segments

Номер: US20170002688A1

Автор: Bruna Ramos, Bruno Albert Beutin, Maria Natal, Nuria Llamas Castro

Принадлежит: Safran Aircraft Engines SAS

An airplane unducted propeller turboshaft engine having a gas generator and a receiver including a propulsion assembly carrying least one propeller, the engine including a first casing, a second casing, and a third casing, the third casing being provided between the first and second casings and surrounding at least a portion of the gas generator, a reinforcing shell presenting a first attachment zone mounted on the first casing second attachment zone mounted on the second casing, and a wall provided between the first and second attachment zones and surrounding the third casing, wherein the reinforcing shell further includes at least one pipe segment integrated in the wall.

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Номер записи: 56

07-01-2016 дата публикации

Parallel Twin-Impeller Compressor Having Swirl-Imparting Device For One Impeller

Номер: US20160003046A1

Автор: James Reyenga, Vai Man Lei, William Joseph Smith

Принадлежит: Honeywell International Inc

An turbocharger includes a parallel twin-impeller compressor and separate inlets into the two impellers. A first or outboard impeller receives one stream of inlet air and a second or inboard impeller receives its own separate stream of inlet air by way of a generally annular inlet volute surrounding the inlet to the second impeller. Air is fed from the inlet volute radially inwardly through a feed passage to the inlet to the second impeller. A plurality of inlet guide vanes are located in the feed passage for the second impeller, the inlet guide vanes creating a swirling air stream into the second impeller. Introduction of swirl to the second impeller alters the flow distribution between the impellers and affects the stability of the overall stage.

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Номер записи: 57

07-01-2016 дата публикации

GAS TURBINE ENGINE STATOR VANE BAFFLE ARRANGEMENT

Номер: US20160003071A1

Автор: Devore Matthew A., Ghigliotty Jaime G.

Принадлежит:

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

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Номер записи: 58

07-01-2016 дата публикации

GAS TURBINE ENGINE THIN WALL COMPOSITE VANE AIRFOIL

Номер: US20160003072A1

Автор: Chang Hoyt Y., Duelm Shelton O.

Принадлежит:

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

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Номер записи: 59

07-01-2016 дата публикации

GUIDE VANE ASSEMBLY VANE BOX OF AN AXIAL TURBINE ENGINE COMPRESSOR

Номер: US20160003073A1

Автор: Derclaye Alain

Принадлежит:

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

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Номер записи: 60

07-01-2016 дата публикации

GAS TURBINE ENGINE STATOR VANE ASSEMBLY WITH SPLIT SHROUD

Номер: US20160003075A1

Автор: Feigleson Steven J.

Принадлежит:

A method of assembling gas turbine engine front architecture includes positioning a first shroud and a first shroud portion radially relative to one another. Multiple vanes are arranged circumferentially between the first shroud and the first shroud portion. A second shroud portion is secured to the first shroud portion about the vanes. The first and second shroud portions provide a second shroud. The vanes are mechanically isolated from the first and second shrouds. 1. A method of assembling gas turbine engine front architecture comprising the steps of:positioning a first shroud and a first shroud portion radially relative to one another;arranging multiple vanes circumferentially between the first shroud and the first shroud portion;securing a second shroud portion to the first shroud portion about the vanes, the first and second shroud portions providing a second shroud; andmechanically isolating the vanes from the first and second shrouds.2. The method according to claim 1 , wherein the first and second shrouds respectively correspond to inner and outer shrouds.3. The method according to claim 2 , wherein the arranging step includes inserting the vanes into first and second slots respectively provided in the outer and inner shrouds claim 2 , and comprising the step of applying a liquid sealant around a perimeter of the vanes and at least one of the shrouds claim 2 , and bonding and supporting the ends of vanes relative to one of the shrouds with the liquid sealant.4. The method according to claim 3 , wherein each blade includes outer and inner perimeters respectively received in the first and second slots claim 3 , and the arranging step includes providing gaps between the outer and the inner perimeters and the outer and inner shrouds at their respective first and second slots claim 3 , wherein the applying step includes laying the liquid sealant about at least one of the inner and outer perimeters within their respective gaps.5. The method according to claim 4 , ...

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Номер записи: 61

07-01-2016 дата публикации

Gas turbine engine component having variable width feather seal slot

Номер: US20160003079A1

Автор: Allison Mainelli, Jeffrey J. Degray, Kevin RAJCHEL, Mark A. Boeke, Richard M. Salzillo

Принадлежит: United Technologies Corp

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

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Номер записи: 62

07-01-2016 дата публикации

GAS TURBINE ENGINE SEAL ASSEMBLY

Номер: US20160003080A1

Автор: McGARRAH CRAIG R.

Принадлежит: UNITED TECHNOLOGIES CORPORATION

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

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Номер записи: 63

07-01-2016 дата публикации

CONTOURED BLADE OUTER AIR SEAL FOR A GAS TURBINE ENGINE

Номер: US20160003082A1

Автор: Bruskotter Michael J., Lutjen Paul M.

Принадлежит:

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

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Номер записи: 64

07-01-2016 дата публикации

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

Номер: US20160003095A1

Автор: JODET Norman Bruno André, Romano Pascal Jérôme

Принадлежит: SNECMA

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

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Номер записи: 65

07-01-2016 дата публикации

Gas turbine engine attachment structure and method therefor

Номер: US20160003104A1

Автор: David J. Sander, Shelton O. Duelm

Принадлежит: United Technologies Corp

An attachment structure for a gas turbine engine includes a frame that has a first annular case. A second annular case extends around the frame. The first annular case and the second annular case include a plurality of interlocks. Each of the interlocks includes a first member mounted on one of the first annular case or the second annular case and a corresponding second member mounted on the other of the first annular case or the second case. The first member is received in the second member such that the plurality of interlocks restricts relative circumferential and axial movement between the first annular case and the second annular case.

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Номер записи: 66

07-01-2016 дата публикации

METHOD AND APPARATUS FOR HANDLING PRE-DIFFUSER AIRFLOW FOR COOLING HIGH PRESSURE TURBINE COMPONENTS

Номер: US20160003149A1

Автор: Ackermann William K., Suciu Gabriel L.

Принадлежит:

A gas turbine engine is provided comprising a compressor section, a combustor section, a diffuser case module, and a manifold. The diffuser case module includes a multiple of struts within an annular flow path from said compressor section to said combustor section, wherein at least one of said multiple of struts defines a mid-span pre-diffuser inlet in communication with said annular flow path. The manifold is in communication with said mid-span pre-diffuser inlet and said compressor section. 1. A gas turbine engine comprising:a compressor section;a combustor section;a diffuser case module with a multiple of struts within an annular flow path from said compressor section to said combustor section, at least one of said multiple of struts defines a mid-span pre-diffuser inlet in communication with said annular flow path; anda manifold in communication with said mid-span pre-diffuser inlet and said compressor section.2. The gas turbine engine as recited in claim 1 , wherein said manifold communicates a temperature tailored airflow.3. The gas turbine engine as recited in claim 1 , wherein said manifold communicates a temperature tailored airflow thru a heat exchanger.4. The gas turbine engine as recited in claim 3 , wherein said manifold communicates said temperature tailored airflow from said heat exchanger as buffer air.5. The gas turbine engine as recited in claim 4 , wherein said buffer air is communicated thru a buffer passage to one or more bearing compartments.6. The gas turbine engine as recited in claim 1 , wherein said mid-span pre-diffuser inlet supplies a temperature tailored airflow into said manifold.7. The gas turbine engine as recited in claim 1 , wherein said manifold communicates with a high pressure turbine of said compressor section.8. The gas turbine engine as recited in claim 1 , wherein said manifold is generally annular.9. The gas turbine engine as recited in claim 1 , wherein said manifold communicates with a row of rotor blade attachments in ...

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Номер записи: 67

07-01-2016 дата публикации

GAS TURBINE ENGINE MULTI-VANED STATOR COOLING CONFIGURATION

Номер: US20160003152A1

Автор: Boeke Mark A., Gregg Shawn J., Riley Sarah

Принадлежит:

A stator for a gas turbine engine has a platform supporting multiple vanes that includes first and second vanes respectively. First and second regions are arranged at the same location on the first and second vanes. The first and second regions respectively include first and second cooling hole configurations that are different than one another. 1. A stator for a gas turbine engine comprising:a platform supporting multiple vanes including first and second vanes respectively including first and second regions, the first and second regions arranged at a same location on the first and second vanes, the first and second regions respectively including first and second cooling hole configurations that are different than one another.2. The stator according to claim 1 , wherein the first and second cooling hole configurations corresponding to cooling hole size.3. The stator according to claim 1 , wherein the first and second cooling hole configurations corresponding to cooling hole shape.4. The stator according to claim 3 , wherein the first cooling hole configuration includes an oblong exit claim 3 , and the second cooling hole configuration includes a conical exit.5. The stator according to claim 1 , wherein the first and second cooling hole configurations corresponding to cooling hole density.6. The stator according to claim 6 , wherein the first and second regions are the same size claim 6 , and the first and second cooling configurations each include a different number of cooling holes.7. The stator according to claim 1 , wherein the first and second regions are provided on airfoils.8. The stator according to claim 7 , wherein the first and second regions are provided on pressure sides.9. The stator according to claim 7 , wherein the first and second regions are provided on suction sides.10. The stator according to claim 1 , wherein the first cooling hole configuration includes a cooling hole having a cooling hole axis providing a line of sight that is obstructed by ...

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Номер записи: 68

04-01-2018 дата публикации

GEARED TURBOFAN GAS TURBINE ENGINE ARCHITECTURE

Номер: US20180003112A1

Автор: Kupratis Daniel Bernard, Schwarz Frederick M.

Принадлежит:

A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine. 1. A gas turbine engine comprising:a fan including a plurality of fan blades rotatable about an axis;a compressor section;a combustor in fluid communication with the compressor section;a turbine section in fluid communication with the combustor, the turbine section including a fan drive turbine and a second turbine, wherein the second turbine is disposed forward of the fan drive turbine and the fan drive turbine includes a plurality of turbine rotors with a ratio between the number of fan blades and the number of fan drive turbine rotors is between 2.5 and 8.5; anda speed change system configured to be driven by the fan drive turbine to rotate the fan about the axis at a different speed than the fan drive turbine;wherein the fan drive turbine has a first exit area and rotates at a first speed, the second turbine section has a second exit area and rotates at a second speed, which is faster than the first speed, said first and second speeds being redline speeds, a first performance quantity is defined as the product of the first speed squared and the first area, a second performance quantity is defined as the product of the ...

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Номер записи: 69

02-01-2020 дата публикации

Turbomachine blade and method for the manufacture of same

Номер: US20200003061A1

Автор: Nicolas Pierre LANFANT, Pierre Jean FAIVRE D'ARCIER

Принадлежит: Safran SA

A blade of a turbomachine includes a blade body of composite material having a fiber reinforcement having a three-dimensional weave and densified by a matrix, the reinforcement having a first part extended by a second, end, part including two segments separated from each other; and an insert having a pi-shaped section, the insert having a platform part and two longitudinal flanges separated from each other, the platform part including a housing delimited by a bottom wall and a rim, the bottom wall including an opening communicating with the space between the two flanges, the first part of the fiber reinforcement being clamped between the two flanges of the insert, the segments of the second part of the fiber reinforcement being folded against the bottom wall of the housing.

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Номер записи: 70

02-01-2020 дата публикации

VANE SYSTEM WITH CONNECTORS OF DIFFERENT LENGTH

Номер: US20200003064A1

Автор: Kline Bradley P., Shahab Farruqh

Принадлежит:

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

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Номер записи: 71

02-01-2020 дата публикации

GUIDE VANE ARRANGEMENT FOR TURBOMACHINE

Номер: US20200003065A1

Автор: Feldmann Manfred

Принадлежит: MTU Aero Engines AG

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

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Номер записи: 72

02-01-2020 дата публикации

GAS TURBINE ENGINE COMPONENT

Номер: US20200003066A1

Автор: Clark Thomas E., Whitney Daniel J.

Принадлежит:

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

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Номер записи: 73

02-01-2020 дата публикации

VARIABLE STATOR VANE ARRANGEMENT

Номер: US20200003073A1

Автор: KARAPURATH Neelambaran, KYLE Robert J.

Принадлежит: ROLLS-ROYCE PLC

A variable stator vane arrangement is provided in which the variable stator vanes extend from a first end at a radially inner flow boundary to a second end at a radially outer flow boundary. At least one of the radially inner flow boundary and the radially outer flow boundary is faceted, such that the surface of the faceted flow boundary comprises flat portions at the interfaces with the respective first or second end of each stator vane. The flat portions mean that the tips of the variable stator vanes can be made substantially flush with the flat casing portions. This may improve aerodynamic efficiency and/or increase the design flexibility on where to position the pivot axis of the variable stator vanes. 1. A compressor for a gas turbine engine comprising:a radially inner flow boundary;a radially outer flow boundary;an annular array of variable stator vanes, each stator vane extending from a first end at the radially inner flow boundary to a second end at the radially outer flow boundary, wherein:at least one of the radially inner flow boundary and the radially outer flow so boundary is faceted, such that the surface of the faceted flow boundary comprises flat portions at the interfaces with the respective first or second end of each stator vane.2. The compressor according to claim 1 , wherein:each stator vane is pivotable about a pivot axis; andthe flat portions of the faceted flow boundary are perpendicular to the pivot axis of the respective stator vane at each interface.3. The compressor according to claim 1 , wherein each stator vane comprises:an aerofoil portion; anda boundary interface portion, whereinthe boundary interface portion is a flat surface lying in the same plane as the adjacent flat portion of the faceted flow boundary.4. The compressor according to claim 3 , wherein the boundary interface portion is circular.5. The compressor according to claim 3 , wherein there is substantially no gap between the boundary interface portion and the surrounding ...

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Номер записи: 74

02-01-2020 дата публикации

GAS TURBINE ENGINE COMPONENT

Номер: US20200003077A1

Автор: Blaney Ken F., Clark Thomas E.

Принадлежит:

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

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Номер записи: 75

02-01-2020 дата публикации

SYSTEM INCLUDING TELESCOPING HIDDEN DRAG LINK ASSEMBLY FOR ACTUATING BLOCKER DOOR OF THRUST REVERSER

Номер: US20200003151A1

Автор: Carr Alexander Jon

Принадлежит: SPIRIT AEROSYSTEMS, INC.

A system for actuating a blocker door of a thrust reverser, in which a drag link assembly is removed from the airflow through the engine during flight. The assembly couples the door to a sleeve so that translation of the sleeve between deployed and stowed positions moves the door to open and closed positions, respectively. The assembly includes a telescoping drag link having one end rotatably coupled with a drag link anchor and another end coupled with the door. During deployment and stowage, translation of the sleeve causes the door to move and the link to extend or collapse and rotate until it contacts a stop element of the anchor, and then rotate in the opposite direction as the door opens or closes, respectively. A channel may be provided in the door to accommodate the link. The anchor may include a spring which exerts a rotational force on the link. 1. A system for actuating a blocker door of a thrust reverser , the thrust reverser including a sleeve translatable between a stowed position in which an engine airflow is directed rearwardly and a deployed position in which the engine airflow is redirected forwardly , the system comprising:a blocker door moveable between a closed position associated with the stowed position of the sleeve and an open position associated with the deployed position of the sleeve and in which the engine airflow is redirected laterally by the blocker door; and a drag link including a base section and a subsequent section collapsibly and extendably connected to the base section, with the base section including a first link end, and the subsequent section including a second link end rotatably coupled with the blocker door, and', 'a drag link anchor, with the first link end of the drag link rotatably coupled with the drag link anchor, the drag link anchor including a stop element which limits the rotation of the drag link in a first direction about the drag link anchor,, 'a drag link assembly located between the sleeve in the stowed ...

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Номер записи: 76

07-01-2021 дата публикации

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

Номер: US20210003025A1

Автор: Kime Kent L., Petersen Blake, Smoke Jason, Zurmehly Luke

Принадлежит: HONEYWELL INTERNATIONAL INC.

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

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Номер записи: 77

07-01-2021 дата публикации

DOUBLE BOX BOAS AND CARRIER SYSTEM

Номер: US20210003026A1

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

Принадлежит:

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

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Номер записи: 78

07-01-2021 дата публикации

VANE ANGLE SYSTEM ACCURACY IMPROVEMENT

Номер: US20210003029A1

Автор: Ward Thomas W.

Принадлежит:

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

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Номер записи: 79

02-01-2020 дата публикации

Combustor Heat Shield Sealing

Номер: US20200003418A1

Автор: Corsmeier Donald Michael, Dziech Aaron Michael, Proano Cadena Valeria, Radwanski Michael Todd, Stieg Michael Alan

Принадлежит:

Combustor assemblies for gas turbine engines are provided. For example, a combustor assembly comprises a combustor dome, a first heat shield having an edge, a second heat shield having an edge, and a seal extending from the edge of the first heat shield to the edge of the second heat shield such that the seal spans a gap between the first heat shield and the second heat shield. In another embodiment, the seal has a first contact portion contacting the edge of the first heat shield, a second contact portion contacting the edge of the second heat shield edge, and a connecting portion connecting the first portion and the second portion. The first contact portion and the second contact portion project away from the connecting portion. Methods for sealing between adjacent heat shields of a combustor assembly also are provided. 1. A combustor assembly for a gas turbine engine , comprising:a combustor dome;a first heat shield having an edge;a second heat shield having an edge; anda seal extending from the edge of the first heat shield to the edge of the second heat shield such that the seal spans a gap between the first heat shield and the second heat shield.2. The combustor assembly of claim 1 , wherein the seal defines a plurality of openings to allow a flow of purge air therethrough.3. The combustor assembly of claim 1 , wherein the first heat shield defines a pocket along its edge and the second heat shield defines a pocket along its edge claim 1 , and wherein the seal is received within the pocket of the first heat shield and the pocket of the second heat shield.4. The combustor assembly of claim 1 , wherein the first heat shield defines an interface surface along its edge and the second heat shield defines an interface surface along its edge claim 1 , and wherein a first portion of the seal is positioned against the interface surface of the first heat shield and a second portion of the seal is positioned against the interface surface of the second heat shield.5. ( ...

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Номер записи: 80

03-01-2019 дата публикации

Blade removal device and method

Номер: US20190003310A1

Автор: Ryota CHAKI, Tadakazu Sakaguchi

Принадлежит: Mitsubishi Hitachi Power Systems Ltd

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

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Номер записи: 81

03-01-2019 дата публикации

AIRFOIL ASSEMBLY WITH A SCALLOPED FLOW SURFACE

Номер: US20190003323A1

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

Принадлежит:

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

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Номер записи: 82

03-01-2019 дата публикации

Turbine engine component with an insert

Номер: US20190003324A1

Автор: Victor Hugo Silva Correia

Принадлежит: General Electric Co

A component for a turbine engine comprises a wall with a surface along which a hot airflow passes, a second surface along which a cooling airflow passes, and an insert mounted to the wall wherein the material used for the insert can have a higher temperature capability than that of the wall.

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Номер записи: 83

03-01-2019 дата публикации

SEALING DEVICE

Номер: US20190003328A1

Автор: HAMADA Katsuhisa, KUWAMURA Yoshihiro, NISHIKAWA Toyoharu

Принадлежит:

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

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Номер записи: 84

03-01-2019 дата публикации

THERMALLY DRIVEN SPRING VALVE FOR TURBINE GAS PATH PARTS

Номер: US20190003333A1

Автор: Propheter-Hinckley Tracy A., Thornton Lane M.

Принадлежит: UNITED TECHNOLOGIES CORPORATION

A thermally driven spring valve for turbine gas path parts is disclosed herein. A thermally driven spring valve includes a bimetallic sheet comprising a base, a first finger portion extending from the base and a second finger portion extending from the base, the first finger portion having a first curvature vector and the second finger portion have a second curvature vector, wherein an exterior surface extends from the base through the first finger portion and the second finger portion and an interior surface extends from the base through the first finger portion and the second finger portion, wherein the exterior surface of the first finger portion is disposed proximate the interior surface of the base wherein the exterior surface of the second finger portion is disposed proximate the interior surface of the base. A thermally driven spring valve may include perforations through a finger portion. 1. A thermally driven spring valve comprising:a metallic sheet comprising a base mount portion and a floating portion having a curvature vector, wherein the base mount portion is coupled to a wall of a chamber, wherein the floating portion is disposed proximate an aperture in the wall.2. The thermally driven spring valve of , wherein the metallic sheet is coupled to the wall of the chamber by at least one of brazing or welding. The thermally driven spring valve of , wherein the metallic sheet is a bimetallic sheet.4. The thermally driven spring valve of claim 1 , wherein the metallic sheet comprises at least one of steel claim 1 , titanium claim 1 , titanium alloy claim 1 , cobalt claim 1 , cobalt alloy claim 1 , platinum claim 1 , or platinum alloy.5. The thermally driven spring valve of claim 1 , wherein the metallic sheet has a coefficient of thermal expansion of between about 0.6×10/K to about 25×10/K.6. The thermally driven spring valve of claim 1 , wherein the chamber is coupled to a baffle. This application is a divisional of, and claims priority to, and the benefit ...

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Номер записи: 85

03-01-2019 дата публикации

STEAM TURBINE COOLING UNIT

Номер: US20190003334A1

Автор: HAMADA Katsuhisa, MATSUMURA Takumi, NISHIKAWA Toyoharu

Принадлежит:

A steam turbine cooling unit for a steam turbine includes a coolant steam path provided to penetrate a casing (an outer casing and an inner casing) along a superheated steam supply tube to reach a gap; and a coolant steam supplying unit configured to supply coolant steam flowing through the coolant steam path along the superheated steam supply tube to reach the gap, and having a pressure higher than and a temperature lower than those of superheated steam to be supplied by the superheated steam supply tube. This configuration provides improved cooling efficiency. 1. A steam turbine cooling unit for a steam turbine that includes a rotor which is a rotating body extending along an axial center of rotations of the rotor , a casing configured to house the rotor , a steam path provided between the rotor and the casing in an extending direction of the rotor , a steam nozzle unit attached to the casing with a gap formed between an outer surface of the steam nozzle unit and an outer circumferential surface of the rotor , the gap having an annular shape surrounding the outer circumference of the rotor and communicating with the steam path , the steam nozzle unit including a steam nozzle chamber having an annular shape formed along internal of the steam nozzle unit and an opening facing the extending direction of the rotor from the steam nozzle chamber to communicate with the steam path , and a superheated steam supply tube to which superheated steam is supplied , the superheated steam supply tube being provided to penetrate the casing from external of the casing to communicate with the steam nozzle chamber in the steam nozzle unit , a coolant steam path provided to penetrate the casing along the superheated steam supply tube to reach the gap; and', 'a coolant steam supplying unit configured to supply coolant steam flowing through the coolant steam path along the superheated steam supply tube to reach the gap, the coolant steam having a pressure higher than and a temperature ...

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Номер записи: 86

03-01-2019 дата публикации

TURBOJET ENGINE COMPRISING A NACELLE EQUIPPED WITH REVERSER FLAPS

Номер: US20190003421A1

Автор: CZAPLA Lionel, Pautis Olivier, PIARD Frédéric, RIDRAY Frédéric, TIZAC Laurent

Принадлежит:

A turbofan comprising a fan casing and a nacelle comprising a cowl translatable between an advanced position and a pushed-back position in which the mobile cowl and the fan casing define a window therebetween. The nacelle also comprises reverser flaps, each reverser flap being mounted in a linked manner on the mobile assembly between a closed position in which it obstructs the window and an open position in which it does not obstruct the window. The nacelle also has an impelling mechanism comprising a lever arm rotatable on the mobile assembly and having a roller, a connecting rod mounted rotatably between the reverser flap and the lever arm, and a channel receiving the roller, the channel having a front part parallel with the translation direction and a rear part extending after the front part and oriented inward as it progresses from the front toward the rear. 1. A turbofan comprising an engine and a nacelle surrounding the engine which comprises a fan casing and a core arranged inside the fan casing , in which a duct for a bypass flow is defined between the core and the fan casing , said nacelle comprising:a fixed structure,a fan cowl fixedly mounted on the fixed structure and a mobile assembly comprising a mobile cowl and being translatable with respect to the fixed structure in a translation direction between an advanced position in which the mobile cowl is brought closer to the fan cowl and a pushed-back position in which the mobile cowl is moved away from the fan cowl toward the rear,a window defined upstream by the fan cowl and downstream by the mobile cowl, said window being open, in the pushed-back position, between the duct and the outside of the nacelle,a reverser flap mounted on the mobile assembly tilting between a closed position in which the reverser flap obstructs the window and an open position in which the reverser flap does not obstruct the window, and a rearward translation of the mobile assembly in the translation direction in order to move the ...

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Номер записи: 87

03-01-2019 дата публикации

DUAL-EXPANDER SHORT-LENGTH AEROSPIKE ENGINE

Номер: US20190003423A1

Автор: Pelfrey Philip C.

Принадлежит:

A dual-expander, truncated aerospike rocket engine includes an aerospike nozzle with an oxidizer-cooled nozzle section and a fuel-cooled nozzle section. The engine further includes a fuel pump, an oxidizer pump, and a thrust source. The thrust source combusts fuel and oxidizer in a combustion chamber(s) of a thrust source arranged around the aerospike nozzle. Fuel and oxidizer pumps that pump the fuel and the oxidizer through the nozzle sections are driven by respective turbines. 1. A rocket engine , comprising:a truncated aerospike nozzle having a first end and a second end, a fuel-cooled nozzle section between the first end and the second end, and an oxidizer-cooled nozzle section between the first end and the second end;a fuel inlet that is in fluid communication with the fuel-cooled nozzle section to permit introduction of fuel into the fuel-cooled nozzle section;a fuel outlet formed in the truncated aerospike nozzle that allows fuel to leave the fuel-cooled nozzle section;an oxidizer inlet that is in fluid communication with the oxidizer-cooled nozzle section to permit introduction of oxidizer into the oxidizer-cooled nozzle section;an oxidizer outlet formed in the truncated aerospike nozzle that allows oxidizer to leave the oxidizer-cooled nozzle section;a fuel pump in fluid communication with the fuel inlet to pump fuel into the fuel-cooled nozzle section;an oxidizer pump in fluid communication with the oxidizer inlet to pump oxidizer into the oxidizer-cooled nozzle section; anda thrust source disposed on the truncated aerospike nozzle adjacent to the first end and that surrounds the truncated aerospike nozzle, the thrust source includes a first inlet that is in fluid communication with the fuel outlet and a second inlet that is in fluid communication with the oxidizer outlet; the thrust source has a combustion chamber that receives fuel from the first inlet and oxidizer from the second inlet, and a combustion gas outlet in fluid communication with the ...

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Номер записи: 88

03-01-2019 дата публикации

FAN MODULE HAVING VARIABLE-PITCH BLADES FOR A TURBINE ENGINE

Номер: US20190003484A1

Автор: BENSLAMA Yanis, DIEVART Jérémy, Igel Dominik, Morreale Serge René

Принадлежит:

The invention relates to a fan module having variable-pitch blades for a turbine engine, including a rotor () having blades (), a stationary casing (), and a system for adjusting and controlling the pitch of the blades (), the rotor () including a central shaft () and a ring () for supporting the blades surrounding the shaft, a front end of the ring being connected to a front end of the shaft so as to define an annular space between the ring and the shaft which is open towards the rear, said annular space of the rotor () housing said system, and the shaft () being guided by a first bearing () mounted in the stationary casing (), to the rear of the ring (), characterised in that the ring () is guided by at least one complementary bearing () located upstream of the first bearing (). 19-. (canceled)10. Fan module turning about an axis (X) having variable-pitch blades for a turbine engine , said fan module comprising a rotor carrying blades , a fixed casing and a system for adjusting and controlling the pitch of the blades , the rotor comprising a central shaft and a ring supporting the blades surrounding the shaft , a front end of the ring being connected to a front end of the shaft so as to define , between the ring and the shaft , an annular space open towards the rear , said annular space of the rotor housing said system and the shaft being guided by a first bearing mounted in the fixed casing , behind the ring , characterised in that the ring is guided by at least one complementary bearing situated upstream of the first bearing with regards to rotation axis (X) and in that the system for adjusting and controlling the pitch of the blades comprises an actuator mounted on the fixed casing , a housing of which supports , on its external radial wall , an inner track of said complementary bearing connecting the external radial wall of the housing.11. The fan module according to claim 10 , wherein the housing is mounted on a same part of the fixed casing as the first ...

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Номер записи: 89

03-01-2019 дата публикации

Gas turbine engine

Номер: US20190003487A1

Автор: Daisuke Uemura, Hiroshi Kubo, Koji Terauchi, Naoto Sakai, Toshimasa Miura, Yuki Ishitobi

Принадлежит: Kawasaki Jukogyo KK

A gas turbine engine, in which a compressed gas from a compressor is burned in a combustor and obtained combustion gas drives a turbine, includes: a compressed gas supply portion configured to supply the compressed gas obtained from the compressor to the combustor; an annular dividing guide body disposed in a diffuser that forms an upstream-side portion of the compressed gas supply portion, the dividing guide body being configured to divide the compressed gas in a radial direction; and a guide support body that supports the dividing guide body on an inner diameter side wall of the compressed gas supply portion.

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Номер записи: 90

20-01-2022 дата публикации

DEVICES AND METHODS FOR GUIDING BLEED AIR IN A TURBOFAN ENGINE

Номер: US20220018292A1

Автор: ANAND Karan, CAPRON Alexandre, YAM Tammy

Принадлежит:

Device and methods for guiding bleed air in a turbofan gas turbine engine are disclosed. The devices provided include louvers and baffles that guide bleed air toward a bypass duct of the turbofan engine. The louvers and baffles have a geometric configuration that promotes desirable flow conditions and reduced energy loss. 1. A device for guiding bleed air into a bypass duct of a turbofan engine having a central axis , the device comprising:a body defining a flow-guiding surface having opposite first and second ends defining a span of the flow-guiding surface around the central axis, the flow-guiding surface extending between a radially-inner edge of the body and a radially-outer edge of the body relative to the central axis; anda side wall adjacent the first end of the flow-guiding surface of the body, the side wall extending at least partially axially relative to the central axis, the side wall extending from a first position radially inwardly of the radially-inner edge of the body to a second position radially outwardly of the radially-inner edge of the body relative to the central axis.2. The device as defined in claim 1 , wherein the second position is adjacent the radially-outer edge of the body.3. The device as defined in claim 1 , wherein the side wall is substantially planar.4. The device as defined in claim 3 , wherein the side wall is non-parallel to a radial direction relative to the central axis.5. The device as defined in claim 1 , wherein the side wall is curved.6. The device as defined in claim 1 , wherein the side wall has a Bellmouth profile when viewed along the central axis.7. The device as defined in claim 1 , wherein the side wall has a unitary construction with the body.8. The device as defined in claim 1 , comprising a baffle disposed axially of the body to define a bleed air passage between the baffle and the flow-guiding surface of the body claim 1 , wherein a gap is defined between the side wall and the baffle.9. The device as defined in ...

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Номер записи: 91

20-01-2022 дата публикации

Compressor apparatus with bleed slot and supplemental flange

Номер: US20220018293A1

Автор: Atanu Saha, Manish Singhal, Michael Anthony Thomas, Steven Mitchell Taylor, Vishnu Das K

Принадлежит: General Electric Co

An example compressor bleed slot apparatus includes an annular compressor casing, a blade row mounted for rotation about a centerline axis inside the compressor casing, a bleed slot passing through the forward section of the compressor casing, wherein the bleed slot is bounded by inboard and outboard walls defined within the compressor casing, the bleed slot extending along a slot axis, at least a portion of the bleed slot lying within an axial extent of the blade row, an array of struts interconnecting the inboard and outboard walls, and an annular supplemental flange extending radially outward from the forward section of the compressor casing, wherein at least a portion of the supplemental flange is axially positioned within an axial extent of the bleed slot, wherein the supplemental flange includes a necked-down portion adjacent the forward section of the compressor casing, the necked-down portion positioned axially between an inlet of the bleed slot and an outlet of the bleed slot.

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Номер записи: 92

20-01-2022 дата публикации

DIFFUSION SYSTEM CONFIGURED FOR USE WITH CENTRIFUGAL COMPRESSOR

Номер: US20220018361A1

Автор: Mazur Steven

Принадлежит:

A compressor includes an impeller and a diffuser. The impeller is mounted for rotation about an axis of the gas turbine engine. The diffuser is coupled to the impeller to receive the high velocity air from the impeller. The diffuser includes a first plate, a second plate spaced apart from the first plate axially, and a plurality of vanes located between the first and second plates. 1. A diffuser adapted for use with a centrifugal compressor , the diffuser comprisinga first plate that extends circumferentially about an axis,a second plate that extends circumferentially about the axis, the second plate spaced apart axially from the first plate relative to the axis to define a flow path between the first plate and the second plate, anda plurality of vanes that extend axially between and interconnect the first plate and the second plate, the plurality of vanes including a first vane and a second vane spaced apart circumferentially from the first vane to define a throat inlet of the diffuser located at a radial throat distance from the axis, the first vane and the second vane each including a leading edge and a trailing edge spaced apart radially from the leading edge to define a camber line that extends within the respective first and second vane and interconnects the leading edge and the trailing edge of the respective first and second vane,wherein the plurality of vanes are backswept such that the camber line of each of the first vane and the second vane is curved and at least one of the first plate and the second plate diverges axially relative to the other of the first plate and the second plate beginning at a location equal to the radial throat distance or radially outward of the radial throat distance.2. The diffuser of claim 1 , wherein both the first plate and the second plate diverge axially away from the other of the first plate and the second plate as the first plate and the second plate extend radially outward relative to the axis.3. The diffuser of claim 2 ...

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Номер записи: 93

12-01-2017 дата публикации

COMPRESSOR BLADE OR VANE HAVING AN EROSION-RESISTANT HARD MATERIAL COATING

Номер: US20170009591A1

Автор: Flores Renteria Arturo, Neddemeyer Torsten

Принадлежит:

A compressor blade for a gas turbine is provided. The compressor blade has a blade substrate that consists of a metal alloy and has an aluminum diffusion zone on a surface of the blade substrate. In addition, the compressor blade has a hard material coating provided on the surface of the blade substrate. A compressor that has a compressor blade and a method of producing such a compressor blade is also provided. 111-. (canceled)12. A compressor blade or vane for a gas turbine , the compressor blade or vane comprising:a blade or vane substrate;a metal alloy;an aluminum diffusion zone on a surface of the blade or vane substrate as a result of the diffusion of aluminum into a surface on the blade or vane substrate; anda hard material coating arranged on the surface of the blade or vane substrate .13. The compressor blade or vane of claim 12 , wherein the hard material coating comprises TiN claim 12 , TiAlN claim 12 , AlTiN claim 12 , CrN as single-layer or multi-layer ceramics or comprises TiN claim 12 , TiAlN claim 12 , AlTiN claim 12 , CrN as single-layer or multi-layer ceramics.14. The compressor blade or vane of claim 12 , wherein the aluminum diffusion zone has a thickness of 10 to 30 micrometers.15. The compressor blade or vane of claim 12 , wherein the aluminum diffusion zone has an aluminum proportion of 0.05 to 0.2% by weight.16. The compressor blade or vane of claim 12 , wherein the metal alloy is a creep-resistant steel.17. A compressor for a gas turbine and having a plurality of compressor blades or vanes claim 12 , wherein at least one compressor blade or vane of the plurality of compressor blades or vanes is designed as claimed in .18. The compressor of claim 12 , wherein the plurality of compressor blades or vanes are arranged in a plurality of rows claim 12 , wherein each row of the plurality of rows has a plurality of compressor blades or vanes arranged transversely to a main direction of flow of the compressor claim 12 , and wherein the plurality of ...

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Номер записи: 94

12-01-2017 дата публикации

COMPOSITE VANE AND METHOD FOR MANUFACTURING COMPOSITE VANE

Номер: US20170009592A1

Автор: Obuchi Kenro, Yagi Hiroyuki

Принадлежит: IHI CORPORATION

A composite vane includes a composite vane body that is formed from a composite material of a thermosetting resin or a thermoplastic resin and reinforced fibers, which is obtained by molding, and a metal sheath that is bonded to a leading edge section including a leading edge of the composite vane body and a vane surface in a vicinity of the leading edge via a film adhesive formed by impregnating a mesh with a hard adhesive to cover the leading edge section, wherein an underfill section that is formed in a step of removing excessive thicknesses parts remaining on the leading edge after the molding and does not need leading edge round finish is placed on the leading edge of the leading edge section in the composite vane body. It is possible to realize reduction of manufacturing time and manufacturing cost. 1. A composite vane , comprising:a composite vane body that is formed from a composite material of a thermosetting resin or a thermoplastic resin and reinforced fibers, which is obtained by molding; anda metal sheath that is bonded to a leading edge section including a leading edge of the composite vane body and a vicinity of the leading edge via a film adhesive formed by impregnating a mesh with a hard adhesive to cover the leading edge section,wherein an underfill section that is formed in a step of removing an excessive thickness part remaining on the leading edge after the molding and does not need leading edge round finish is placed on the leading edge of the leading edge section in the composite vane body.2. The composite vane according to claim 1 ,wherein the underfill section is placed at a plurality of positions in a vane width direction of the leading edge, and functions as adhesive gathering spots for a hard adhesive in the film adhesive.3. A method for manufacturing the composite vane according to claim 1 , whereinan underfill section that does not need leading edge round finish is formed on the leading edge, in a step of removing an excessive thickness ...

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Номер записи: 95

12-01-2017 дата публикации

TURBINE STATOR VANE OF CERAMIC MATRIX COMPOSITE

Номер: US20170009593A1

Автор: WATANABE Fumiaki

Принадлежит: IHI CORPORATION

A stator vane is comprised of: an airfoil section elongated in a radial direction relative to an axis; an outer band section continuous to an outer end of the airfoil section and bent in a circumferential direction relative to the axis; a first hook section continuous to a leading end in the axial direction of the outer band section and bent outward in the radial direction; a second hook section continuous to a trailing end in the axial direction of the outer band section and bent outward in the radial direction; an inner band section continuous to an inner end of the airfoil section and bent in the circumferential direction; a flange section continuous to an end in the axial direction of the inner band section and bent inward in the radial direction; and a reinforcement fiber fabric continuous throughout these sections and unitized with a ceramic. 1. A stator vane arranged around an axis to form a turbine nozzle , comprising:an airfoil section elongated in a radial direction relative to the axis;an outer band section continuous to an outer end of the airfoil section and bent in a circumferential direction relative to the axis;a first hook section continuous to a leading end in the axial direction of the outer band section and bent outward in the radial direction;a second hook section continuous to a trailing end in the axial direction of the outer band section and bent outward in the radial direction;an inner band section continuous to an inner end of the airfoil section and bent in the circumferential direction;a flange section continuous to an end in the axial direction of the inner band section and bent inward in the radial direction; anda reinforcement fiber fabric continuous throughout the airfoil section, the outer band section, the first hook section, the second hook section, the inner band section and the flange section and unitized with a ceramic.2. The stator vane of claim 1 , further comprising:a cutout ranging from a leading edge in the axial direction ...

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Номер записи: 96

12-01-2017 дата публикации

Manufacturing of single or multiple panels

Номер: US20170009600A1

Автор: Alexander Stankowski, Piero-Daniele Grasso, Sabrina Puidokas

Принадлежит: Ansaldo Energia IP UK Ltd

A method of manufacturing of a structured cooling panel includes cutting of desized 2D ceramic into tissues; slurry infiltration in the tissues by at least one knife blade coating method; laminating the tissues in a multi-layer panel, with slurry impregnation after each layer, wherein the tissue has combined fibres and/or pre-build cooling holes; drying; de-moulding; sintering the multi-layer panel, wherein part of the combined fibres burns out during the sintering process leaving a negative architecture forming the cooling structure and/or the pre-build cooling holes define the cooling structure; finishing, using of i) post-machine, and/or ii) surface smoothening/rework, and/or iii) coating application, and/or other procedures.

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Номер записи: 97

12-01-2017 дата публикации

Thermally compliant heatshield

Номер: US20170009620A1

Автор: Jonathan C Reuben, Kelsey Honious, Pawel Socha, Richard Joseph Bulkovitch

Принадлежит: United Technologies Corp

A heat shield comprising a sidewall portion, a top portion, and a plurality of flexible tabs attached to the sidewall portion is described herein, in accordance with various embodiments. The top portion may comprise an aperture. The sidewall portion may extend at an angle between 80 degrees and 100 degrees from the top portion. The sidewall portion may bound a hexagonal void. The flexible tab may comprise an angle between 80 degrees and 100 degrees. The flexible tab may be fixed to the sidewall portion, wherein the flexible tab is configured to be attached to a fitting.

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Номер записи: 98

12-01-2017 дата публикации

Motive Air Conditioning System for Gas Turbines

Номер: US20170009654A1

Автор: Maier William C.

Принадлежит:

A motive air conditioning system for a gas turbine assembly is provided. The motive air conditioning system may include an inlet flow channel configured to be fluidly coupled with the gas turbine assembly. The motive air conditioning system may also include a filtration assembly fluidly coupled with the inlet flow channel and configured to filter motive air. The filtration assembly may include a plurality of filter modules disposed adjacent one another and further disposed circumferentially about a longitudinal axis of the inlet flow channel. 1. A motive air conditioning system for a gas turbine assembly , comprising:an inlet flow channel configured to be fluidly coupled with the gas turbine assembly; anda filtration assembly fluidly coupled with the inlet flow channel and configured to filter motive air, the filtration assembly comprising a plurality of filter modules disposed adjacent one another and further disposed circumferentially about a longitudinal axis of the inlet flow channel.2. The motive air conditioning system of claim 1 , wherein each filter module of the plurality of filter modules comprises an upper endwall claim 1 , a lower endwall claim 1 , and a plurality of sidewall panels coupled with one another and defining an inlet of the filter module.3. The motive air conditioning system of claim 2 , wherein at least a portion of the upper endwall is arcuate.4. The motive air conditioning system of claim 2 , wherein at least a portion of the lower endwall is arcuate.5. The motive air conditioning system of claim 1 , wherein the inlet flow channel comprises:an annular duct fluidly coupled with the filtration assembly and configured to receive the motive air from the filtration assembly; andan elbow fluidly coupled with the annular duct, the elbow configured to receive the motive air from the annular duct and at least partially turn the motive air toward the gas turbine assembly, and further configured to attenuate the generation of sound to control ...

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Номер записи: 99

14-01-2016 дата публикации

COMPONENTS WITH COOLING CHANNELS AND METHODS OF MANUFACTURE

Номер: US20160010464A1

Автор: Bonini Eric Richard, BUNKER Ronald Scott

Принадлежит:

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

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Номер записи: 100