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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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05-01-2012 дата публикации

Turbine nozzles and methods of manufacturing the same

Номер: US20120003086A1

Автор: Daniel Ryan, Donald G. Godfrey, Jason Smoke, Mark C. Morris

Принадлежит: Honeywell International Inc

A turbine nozzle is provided and includes a first ring having a first microstructure, a vane extending from the first ring, a first porous zone between the first ring and the vane that is more porous than the first microstructure to attenuate thermo-mechanical fatigue cracking between the vane and the first ring. Methods of manufacturing the turbine nozzle are also provided.

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Номер записи: 1

08-03-2012 дата публикации

Fan array fan section in air-handling systems

Номер: US20120057962A1

Автор: Lawrence G. Hopkins

Принадлежит: Huntair Inc

A fan array fan section in an air handling system includes a plurality of fan units arranged in a fan array and positioned within an air handling compartment. One preferred embodiment may include an array controller program to operate the plurality of the fan units at peak efficiency. The plurality of fan units may be arranged in a true array configuration, a space pattern array configuration, a checker board array configuration, rows slightly offset configuration, columns slightly offset configuration, or a staggered array configuration.

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Номер записи: 2

12-04-2012 дата публикации

Rotor For Turbomachines With Shrouded Blades

Номер: US20120087794A1

Автор: Paolo Calza, Rosario Grillo, Sergio Salvano

Принадлежит: Avio SpA

A rotor for turbomachines with shrouded blades comprising a row of a plurality of blades covered by an external ring, coaxial to said row of blades; the external ring being composed of a plurality of sectors, commonly referred to as shrouds, which are adjacent to one another and transversely and circumferentially attached to the end of respective blades; a damping member being arranged so as to straddle a respective pair of adjacent sectors and being pushed, in use, against radially internal surfaces of the sectors of the pair of adjacent sectors.

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Номер записи: 3

26-04-2012 дата публикации

Rotary machine having non-uniform blade and vane spacing

Номер: US20120099961A1

Автор: Brian Denver Potter, John McConnell Delvaux

Принадлежит: General Electric Co

A system, including a rotary machine including: a stator, a rotor configured to rotate relative to the stator, wherein the rotor comprises a plurality of blades having a non-uniform spacing about a circumference of the rotor.

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Номер записи: 4

26-04-2012 дата публикации

Drive mechanism for a pair of contra-rotating propellers through an epicyclic gear train

Номер: US20120099988A1

Автор: Francois Gallet, Gilles Alain Charier

Принадлежит: SNECMA SAS

A turbine driving a planet gear and an epicyclic gear train and including a planet pinion cage and a ring driving two propellers in rotation, is connected to the planet gear through a flexible sleeve surrounding a turbine support shaft rather than through a support shaft itself to achieve a flexible assembly with a limit stop position in contact with the shaft to limit parasite internal forces applied to the epicyclic gear train without tolerating a loose assembly or breakage of the sleeve due to a condition of the limit stop after a clearance has been eliminated.

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Номер записи: 5

31-05-2012 дата публикации

method of manufacturing a component

Номер: US20120135198A1

Автор: Oliver M. Strother

Принадлежит: Rolls Royce PLC

A method of manufacturing a component having first and second layers, the first and/or second layers including one or more depressions provided on a surface of the respective layer. The method including: arranging the first and second layers so that they face one another and with the depressions on inner facing surfaces of the layers; diffusion bonding the first and second layers together about their edges; applying a first differential pressure across each of the first and second layers to evacuate an inner space defined by the layers, thereby forming one or more depressions on an outer facing surface of the first or second layer; and applying a second differential pressure across each of the first and second layers to expand the inner space defined by the layers.

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Номер записи: 6

28-06-2012 дата публикации

Structure for gas turbine casing

Номер: US20120163963A1

Автор: Shinya Hashimoto, Yukihiro Hashimoto

Принадлежит: Mitsubishi Heavy Industries Ltd

A structure for a gas turbine casing, capable of preventing gas from leaking when the structure is subjected to pressure. A structure for a gas turbine casing, divided by a horizontal plane at flange sections into two halves comprising an upper-half casing ( 10 ) and a lower-half casing ( 11 ). The structure is provided with bolts ( 12 ) which are disposed in the inner surfaces of the upper-half casing ( 10 ) and the lower-half casing ( 11 ) so as to bridge therebetween, upper nuts ( 13 ) which are disposed in the inner surface of the upper-half casing ( 10 ) and attached to the bolts ( 12 ), and lower nuts ( 14 ) which are disposed in the inner surface of the lower-half casing ( 11 ) and attached to the bolts ( 12 ).

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Номер записи: 7

05-07-2012 дата публикации

Noise attenuation panel and a gas turbine component comprising a noise attenuation panel

Номер: US20120168248A1

Автор: Lars-Erik Eriksson, Markus Olander Burak

Принадлежит: Volvo Aero Corp

A noise attenuation panel includes a first wall, a second wall and partition walls connected to the first and second walls and defining cells between the first and second walls. The first wall is provided with a plurality of through holes. At least two of the cells are interconnected via a communication hole. One of the through holes leads to a first of the at least two interconnected cells and a second of the interconnected cells is configured to prevent any gas flow through the second cell.

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Номер записи: 8

02-08-2012 дата публикации

compressor nozzle stage for a turbine engine

Номер: US20120195745A1

Автор: Patrick Edmond Kapala

Принадлежит: SNECMA SAS

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

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Номер записи: 9

09-08-2012 дата публикации

Ring element and turbomachine having such a ring element

Номер: US20120198858A1

Автор: Franz-Dieter SCHMIDT, Johannes Rabe, Wilfried Schuette

Принадлежит: MTU AERO ENGINES GMBH

A ring element for a turbomachine, in particular for an aircraft gas turbine, is disclosed. The ring element has a ring element main body that has two adjacently arranged ring ends, the ring ends being connected to one another in a form-locking manner with respect to an axial plane. Also disclosed is a turbomachine having at least one such ring element.

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Номер записи: 10

01-11-2012 дата публикации

Modular fan housing with multiple modular units having sound attenuation for a fan array for an air-handling system

Номер: US20120275902A1

Автор: Lawrence G. Hopkins

Принадлежит: Huntair Inc

A modular fan housing configured to hold an array of motors and fans is provided. The modular fan housing is configured for use in an air-handling system that delivers air to a ventilation system for at least a portion of a building. The fan housing comprises a plurality of modular units configured to be stacked adjacent to one another in at least one row or column to form an array. The modular units each include an interior surface and have a front end and a back end that define a chamber. The chambers are configured to receive the motors and fans. Sound attenuation layers extend along at least a portion of the interior surface of the corresponding chambers. The sound attenuation layers are positioned between at least some of the adjacent chambers.

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Номер записи: 11

15-11-2012 дата публикации

Rotor with asymmetric blade spacing

Номер: US20120288373A1

Автор: Anthony C. Jones, Benjamin E. Fishler, Bo Zheng, GAO Yang, James C. Napier, Jay M. Francisco, Loc Quang Duong, Nagamany Thayalakhandan, Xiaolan Hu

Принадлежит: Hamilton Sundstrand Corp

A turbine apparatus comprises a rotor with a hub section defined about a rotational axis and a plurality of blades attached to the hub section. The plurality of blades comprises a first group having a first angular spacing in a first circumferential sector of the rotor, and a second group having a second angular spacing in a second circumferential sector of the rotor. The first angular spacing is different from the second angular spacing, and the rotor blades are asymmetric about the rotational axis.

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Номер записи: 12

31-01-2013 дата публикации

Acoustic Array of Polymer Material

Номер: US20130025967A1

Автор: David C. Seib

Принадлежит: Dresser Rand Co

The invention is an acoustic liner for attenuating noise in rotating machinery. The acoustic liner may include a plurality of cells coupled together to form an annular cell matrix, the plurality of cells being made of a non-metallic material, for example, plastics, polymers, thermoplastics, or thermosets. Each cell of the acoustic liner may be hexagonally-shaped such that the annular cell matrix forms a honeycomb structure.

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Номер записи: 13

28-02-2013 дата публикации

Turbocharger

Номер: US20130047605A1

Автор: Hans H. Drangel, Lorenzo Magro, Luca Trabucchi

Принадлежит: GM GLOBAL TECHNOLOGY OPERATIONS LLC

A turbocharger for an internal combustion engine, comprising a bearing housing, a rotating shaft coupled to the bearing housing, a compressor wheel mounted at one end of the rotating shaft, a compressor housing accommodating the compressor wheel and provided with an inlet and an outlet for an air stream, a turbine wheel mounted at the opposite end of the rotating shaft, and a turbine housing accommodating the turbine wheel and provided with an inlet and an outlet for an exhaust gas stream, wherein the bearing housing comprises a fastening flange suitable to be fixed to a component of the internal combustion engine, wherein the turbine housing is provided with a connecting pipe for fluidly connecting the turbine housing inlet with an exhaust manifold of the internal combustion engine, and wherein the connecting pipe includes means for compensating axial thermal deformations thereof.

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Номер записи: 14

14-03-2013 дата публикации

Aerofoil assembly

Номер: US20130064661A1

Автор: Dale E. Evans, Ian C D Care

Принадлежит: Rolls Royce PLC

This invention relates to a stator aerofoil for a gas turbine engine, comprising: a body portion; at least one panel which forms at least part of one of either the pressure or section surface of the aerofoil; at least one internal chamber which is bounded by the body portion and panel; and, at least one elastomeric component within the at least one internal chamber, wherein the elastomeric component at least partially defines a cell within the internal chamber, wherein the cell is in fluid communication with the exterior of the aerofoil.

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Номер записи: 15

21-03-2013 дата публикации

Wall structure with noise damping insulation properties and gas turbine with such a wall structure

Номер: US20130071231A1

Автор: Antun Svilicic, Eduard Bruhwiler, Marc Rauch, Mladen Matan

Принадлежит: Alstom Technology AG

A wall structure is provided with noise damping insulation properties, for an air intake manifold of a gas turbine. The wall structure includes a first structure for mechanically supporting an outer sheet, which separates the spaces on both sides of the wall in an airtight manner, and further includes a second structure for establishing noise damping insulation between the spaces on both sides of the wall. The second structure is secured to the first structure. A gas turbine including the wall structure is also provided.

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Номер записи: 16

28-03-2013 дата публикации

Interface element between a blade root and a blade root housing of a turbine disc, and turbine rotor comprising an interface element

Номер: US20130078101A1

Автор: Alain Dominique Gendraud, Fabrice Marcel Noel Garin, Romain Nicolas Lunel, Thierry Lequitte

Принадлежит: SNECMA SAS

An interface element to be mounted between a blade root and a blade root housing provided in a turbine disc of a gas turbine engine to limit wear between the root and the housing. The interface element includes: a base, configured to be aligned with a lower part of the root; and first and second upper side walls connected to the base and configured to surround the blade root up to an upper portion thereof, the first upper side wall including at least one ventilation opening to allow a flow of cooling air flowing through the rotor disc housing to flow over the upper portion of the root via the ventilation opening.

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Номер записи: 17

25-04-2013 дата публикации

Split accessory drive system

Номер: US20130098058A1

Автор: William G. Sheridan

Принадлежит: United Technologies Corp

A gas turbine engine includes a spool, a first accessory gearbox, a second accessory gearbox, and a scavenge pump. The first accessory gearbox is connected to and driven by the spool. The second accessory gearbox is connected to and driven by the first accessory gearbox. The scavenge pump is connected between the first accessory gearbox and the second accessory gearbox. The first accessory gearbox drives the second accessory gearbox through the scavenge pump.

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Номер записи: 18

23-05-2013 дата публикации

Gas turbine engine comprising a tension stud

Номер: US20130125559A1

Автор: Andrew Shepherd

Принадлежит: SIEMENS AG

A gas turbine engine including a rotor is disclosed. The rotor includes a stud extending along an axis, rotating elements of a first section, and rotating elements of a second section. The stud includes a first and second external end, the first external end adapted to engage a first pre-load nut or a shaft and the second external end adapted to engage a second pre-load nut or a shaft such that the set of rotating elements are secured. Thus stud includes a first shank connected to the first external end and a second shank connected to the second external end. The first shank is located in the first section and has a first diameter. The second shank is located in the second section and has a second diameter which is greater than the first diameter.

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Номер записи: 19

06-06-2013 дата публикации

Alternate shroud width to provide mistuning on compressor stator clusters

Номер: US20130142640A1

Автор: David P. Houston, Richard A. Lomenzo, Robert J. Morris

Принадлежит: United Technologies Corp

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

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Номер записи: 20

13-06-2013 дата публикации

Method for optimizing the control of a free turbine power package for an aircraft, and control for implementing same

Номер: US20130151112A1

Автор: Jean-Michel Haillot

Принадлежит: Turbomeca SA

A method optimizing fuel-injection control with driving speeds of apparatuses adjusted by controlling a turbine speed according to power, and optimizing control of a free turbine power package of an aircraft, including a low-pressure body, supplying power to apparatuses and linked to a high-pressure body. The method varies the low-pressure body speed to obtain a minimum speed for the high-pressure body, so power supplied by the apparatuses remains constant. Power supplied by the apparatuses is dependent upon the apparatuses driven speed by the low-pressure body, and a speed set point of the low-pressure body is dependent upon a maximum value of minimum speeds of the apparatuses, enabling required power to be optimized, upon a positive or zero incrementation added to the speed set point of the low-pressure body to minimize speed of the high-pressure body to the apparatuses power supply.

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Номер записи: 21

20-06-2013 дата публикации

Energy absorbent fan blade spacer

Номер: US20130156591A1

Автор: Phillip Alexander

Принадлежит: United Technologies Corp

A gas turbine engine fan assembly includes a rotor disk, a fan blade and a spacer. The rotor disk includes a longitudinally extending slot with a first spacer contact surface. The fan blade includes a root with a second spacer contact surface arranged within the slot. The spacer includes a leaf spring backbone and a compliant member. The leaf spring backbone includes one or more root contact segments connected longitudinally to one or more slot contact segments. A first of the root contact segments contacts the second spacer contact surface, and a first of the slot contact segments contacts the first spacer contact surface. The compliant member is radially between the first spacer contact surface and the first root contact segment, and radially between the second spacer contact surface and the first slot contact segment.

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Номер записи: 22

04-07-2013 дата публикации

Gas turbine inlet system

Номер: US20130168180A1

Автор: Dinesh Venugopal Setty, Laxmikant Merchant

Принадлежит: General Electric Co

A gas turbine inlet system includes a main inlet portion, wherein an airflow is introduced to the gas turbine inlet system. Also included is a silencer assembly. The silencer assembly includes a first silencing panel, wherein the first silencing panel is oriented substantially perpendicularly to the airflow, and wherein the first silencing panel includes a first plurality of airflow apertures. The silencer assembly also includes a second silencing panel, wherein the second silencing panel is oriented substantially perpendicularly to the airflow, and wherein the second silencing panel includes a second plurality of airflow apertures.

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Номер записи: 23

25-07-2013 дата публикации

Anisotropic vibration isolation mounting assembly

Номер: US20130186105A1

Автор: Alan Margolis, Eric Bridges, Walter Lee Meacham

Принадлежит: Honeywell International Inc

An anisotropic vibration isolation mounting assembly includes at least two three-parameter vibration isolators each having a first end configured for attachment to a rotating member assembly or a rotating member assembly housing and each having a second, opposing end configured for attachment to the rotating member assembly housing where the first end is configured for attachment to the rotating member assembly or to a system interface member where the first member is configured for attachment to the rotating member assembly housing. The at least two three-parameter vibration isolators are tuned anisotropically to minimize the transmission of vibrations during operation of the rotating machine.

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Номер записи: 24

01-08-2013 дата публикации

Stress relieving slots for turbine vane ring

Номер: US20130195643A1

Автор: Douglas MacCaul, John Pietrobon, Keppel Nyron Bharath, Vincent Paradis

Принадлежит: Pratt and Whitney Canada Corp

A turbine vane ring has a radially outer and inner annular shrouds defining therebetween an annular gaspath. Circumferentially spaced-apart airfoil vanes extend radially across the gaspath between the outer and the inner shrouds. The radially outer shroud has a circumferentially continuous cylindrical wall extending axially from a leading edge to a trailing edge. A set of circumferentially distributed stress relieving slots is defined in the leading edge of the cylindrical wall at locations adjacent to the leading edge of at least some of said airfoil vanes. The stress relieving slots extend radially through the cylindrical wall from the radially inner surface to the opposed radially outer surface thereof.

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Номер записи: 25

08-08-2013 дата публикации

Blade cascade and turbomachine

Номер: US20130202444A1

Автор: Roland Wunderer

Принадлежит: MTU AERO ENGINES GMBH

A blade cascade for a turbomachine having a plurality of blades arranged next to one another in the peripheral direction, at least two blades having a variation for generating an asymmetric outflow in the rear area, as well as a turbomachine having an asymmetric blade cascade, which is connected upstream from another blade cascade, are disclosed.

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Номер записи: 26

05-09-2013 дата публикации

Liquid level monitoring and reporting system

Номер: US20130227960A1

Автор: Brett Colin Bonner

Принадлежит: Hamilton Sundstrand Corp

An assembly includes a reservoir, a first sensor, a second sensor, and a controller. The first and second sensors are positioned in the reservoir. The controller is connected to both the first and second sensors. The controller sends a full signal when the first sensor indicates that liquid level in the reservoir is at or above a first level. The controller sends a fill signal when the second sensor indicates that liquid level in the sump is at or below a second level. The controller sends an approximate oil level signal with a value estimated based upon elapsed operating time since the reservoir was at or above the first level.

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Номер записи: 27

10-10-2013 дата публикации

Housing structure of exhaust gas turbocharger

Номер: US20130266436A1

Автор: Hiroshi Suzuki, Seiichi Ibaraki, Takashi Shiraishi

Принадлежит: Mitsubishi Heavy Industries Ltd

A casing treatment 4 including a recirculation passage 41 and a mixing piping 6 are provided. The recirculation passage 41 has a first recirculation opening 42 and a second recirculation opening 43 that are in communication with each other, the first recirculation opening 42 being formed inside a compressor housing 11 of an exhaust gas turbocharger and opening to an air passage 15 upstream of a compressor impeller 3 , the second recirculation opening 43 being formed at the outer circumferential section of the compressor impeller 3 . The mixing pipe 6 opens to the recirculation passage 41 and has a return opening 14 for introducing EGR gas and blow-by gas to the recirculation passage 41.

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Номер записи: 28

17-10-2013 дата публикации

Metallic bondcoat or alloy with a high gamma/gamma' transition temperature and a component

Номер: US20130272917A1

Автор: Anand A. Kulkarni, Jonathan E. Shipper, JR., Werner Stamm

Принадлежит: SIEMENS AG

A metallic bondcoat with phases of γ and γ′ is provided. The metallic coating or alloy is nickel based. The metallic coating or alloy has γ and γ′ phases and optionally has β-phase. The new addition in nickel based coating stabilizes the phases γ and γ′ at high temperatures leading to a reduction of local stresses.

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Номер записи: 29

31-10-2013 дата публикации

Gas turbine engine thermal management system

Номер: US20130284398A1

Автор: Federico Papa, Kathleen R. Phillips, Thomas G. Phillips

Принадлежит: United Technologies Corp

A thermal management system according to an exemplary aspect of the present disclosure includes, among other things, a first fluid circuit that selectively communicates a first portion of a first conditioned fluid having a first temperature to a first gas turbine engine system and a second portion of the first conditioned fluid having a second temperature to a second gas turbine engine system. The second temperature is a greater temperature than the first temperature. A second fluid circuit circulates a second conditioned fluid that is different from the first conditioned fluid to a third gas turbine engine system.

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Номер записи: 30

07-11-2013 дата публикации

Method for starting a turbomachine

Номер: US20130291551A1

Автор: Hubert Pascal Verdier, Luc Reberga, Philippe Etchepare, Pierre Giralt

Принадлежит: Turbomeca SA

A method of starting a turbine engine, including a re-try performed if a main injector has not ignited when a shaft has reached a first predetermined speed value, the re-try including: a stopping during which a starter and the ignitor device are stopped; a second ignition during which fuel is injected into the combustion chamber, the ignitor device being actuated, the second ignition being performed when a speed of rotation of the shaft reaches a second predetermined speed value; and a second starting during which the starter is actuated once more to drive the shaft in rotation.

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Номер записи: 31

07-11-2013 дата публикации

Gas turbine energy storage and conversion system

Номер: US20130294892A1

Автор: David William Dewis, Frank Wegner Donnelly, James Kesseli, John D. Watson, Thomas Wolf

Принадлежит: ICR Turbine Energy Corp USA

The present invention combines the principles of a gas turbine engine with an electric transmission system. A method and apparatus are disclosed for utilizing metallic and ceramic elements to store heat energy derived from a regenerative braking system. The subject invention uses this regenerated electrical energy to provide additional energy storage over conventional electrical storage methods suitable for a gas turbine engine. The subject invention provides engine braking for a gas turbine engine as well as reducing fuel consumption.

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Номер записи: 32

21-11-2013 дата публикации

Sound absorber for a gas turbine exhaust cone, and method for the production thereof

Номер: US20130306401A1

Автор: Predrag Todorovic

Принадлежит: Rolls Royce Deutschland Ltd and Co KG

A method for manufacturing a sound absorber, the outer wall of which is provided with a plurality of recesses, with funnel-like cone elements each being assigned to the recesses inside the sound absorber, said cone elements having a larger opening facing radially outwards and a smaller opening facing radially inwards, with adjacent cone elements each being provided on a strip-shaped first carrier band, with cup elements being provided radially on the inside relative to the cone elements, and each cup element receiving one cone element, with adjacent cup elements each being provided on a strip-shaped first carrier band, with the first carrier bands being arranged adjacently in a first direction, and the second carrier bands being arranged adjacently in a second direction, with the directions crossing each other and the carrier bands being joined to one another and to the outer wall to form a rigid body.

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Номер записи: 33

21-11-2013 дата публикации

Vibration damper

Номер: US20130309097A1

Автор: David Miller

Принадлежит: Rolls Royce PLC

Vibration damping is important with regard to such components as hollow turbine blades in gas turbine engines. Traditionally damping has occurred through damping elements secured at the root or tip of such blades. Such damping is not optimised and results in potential problems with wear in operational life. By providing a tube of deformable material which can be located within a hollow cavity it is possible to provide an element which through friction engagement can absorb vibration energy and therefore damp such vibration. The tube incorporates a number of cuts and/or grooves in an appropriate pattern in order to define a deformation profile once the tube is expanded in location. The tube is secured in position internally upon an expandable element which is typically an inflatable device. Once in position the tube is retained in its expanded deformable profile and the engagement between the tube and the hollow cavity wall surface results in energy absorption through vibration episodes. It is also possible to provide a tube formed from a shape memory alloy which will expand of its own right in location to engage the hollow cavity wall surfaces for energy absorption during vibration episodes.

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Номер записи: 34

05-12-2013 дата публикации

Airfoil cover system

Номер: US20130319010A1

Автор: Kwan Hui, Michael A. Weisse

Принадлежит: Individual

An example airfoil for a gas turbine engine includes a body having a first surface extending from a first edge to a second edge and a cavity disposed in the body. A first cover is at least partially disposed within the cavity. The first cover includes a first portion cooperates with a corresponding second portion. A second cover covers the first cover and forms at least a portion of the first surface with the body. The first cover is disposed between the body and the second cover. The first cover and the second cover have a different coefficient of thermal expansion than the body.

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Номер записи: 35

26-12-2013 дата публикации

Gas turbine engine aft spool bearing arrangement and hub wall configuration

Номер: US20130340435A1

Автор: Gregory M. SAVELA

Принадлежит: Individual

An example gas turbine engine includes a turbine and first and second spools coaxial with one another. The first spool is arranged within the second spool and extends between forward and aft ends. The aft end extends axially beyond the second spool and supports the turbine. A housing is arranged downstream from the turbine. First and second bearings are mounted to the aft end of the first spool and supported by the housing portion.

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Номер записи: 36

26-12-2013 дата публикации

Spherical-link end damper system with near constant engagement

Номер: US20130343876A1

Автор: Kevin J. Cummings, Ryan Edward LeBlanc

Принадлежит: United Technologies Corp

A link includes a link body with two ends, a ring bore with a ring bore axis and a bearing, a mount bore with a mount bore axis and a bearing. The link also has an end curvature at the end having the ring bore wherein the curvature axis is substantially perpendicular to the ring bore axis.

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Номер записи: 37

06-03-2014 дата публикации

Method for optimizing the speed of a twin-spool turbojet engine fan, and architecture for implementing same

Номер: US20140064915A1

Автор: Jean-Pierre Serey, Patrick Dunleavy, Richard Masson

Принадлежит: SNECMA SAS

A method and system improving energy efficiency of a turbojet engine by optimizing rotating speed of a fan and operability of an engine, by freeing the fan from exclusive control of a low-pressure (LP) shaft by providing combined control with a high-pressure (HP) shaft when cruising power has been reached. The turbojet engine include at least one LP turbine and one HP turbine coupled to coaxial LP shafts and HP shafts, respectively, which can drive LP and HP compressors, respectively. The LP compressors include a fan that forms a first primary air-intake compression stage. The LP and HP shafts are mounted on one of two driving mechanisms, an inner ring gear, and a planet carrier for a planetary gear train for driving the fan, the HP shaft being mounted on a disengagement mechanism and the fan being coupled to the planetary gear train via an outer driven ring gear.

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Номер записи: 38

06-03-2014 дата публикации

Coatings for dissipating vibration-induced stresses in components and components provided therewith

Номер: US20140065433A1

Автор: John McConnell Delvaux, Yuk-Chiu Lau

Принадлежит: General Electric Co

A coating material suitable for use in high temperature environments and capable of providing a damping effect to a component subjected to vibration-induced stresses. The coating material defines a damping coating layer of a coating system that lies on and contacts a substrate of a component and defines an outermost surface of the component. The coating system includes at least a second coating layer contacted by the damping coating layer. The damping coating layer contains a ferroelastic ceramic composition having a tetragonality ratio, c/a, of greater than 1 to 1.02, where “c” is a c axis of a unit cell of the ferroelastic ceramic composition and “a” is either of two orthogonal axes, a and b, of the ferroelastic ceramic composition.

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Номер записи: 39

13-03-2014 дата публикации

Filled static structure for axial-flow machine

Номер: US20140072407A1

Автор: James Charles PEARCE, Roy John HARRON

Принадлежит: Rolls Royce PLC

A stator assembly for a rotary machine having a rotor arranged to rotate about an axis in use. The stator assembly has a circumferential support member or casing arranged about said axis and a plurality of elements extending in a substantially radial direction from the support. The elements have a platform at an end thereof for engagement within the support, wherein the elements each comprise a hollow internal cavity having an opening through the platform at the end of the element, wherein said internal cavity is filled with a vibration damping material. The elements may be filled vanes in a gas turbine engine compressor. The platforms may also be filled with the vibration damping material.

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Номер записи: 40

27-03-2014 дата публикации

Gas turbine engine preswirler with angled holes

Номер: US20140083101A1

Автор: Edward T. RAMIREZ, Hongzhou XU, Michael D. Fox, Scott L. STAFFORD

Принадлежит: Solar Turbines Inc

A gas turbine engine preswirler ( 470 ) includes an outer ring ( 471 ) and an inner ring ( 472 ). The inner ring ( 472 ) includes a plurality of angled holes ( 490 ). Each angled hole ( 490 ) follows a vector which is angled in at least one plane. A component of the vector is located on a plane perpendicular a radial from an axis of the preswirler ( 470 ). The component of the angle is angled relative to an axial direction of the preswirler ( 470 ).

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Номер записи: 41

10-04-2014 дата публикации

Aircraft engine with an apparatus for pulsating expiration of gas into the exhaus nozzle

Номер: US20140097271A1

Автор: Daniel Redmann, Josef Steigenberger, Ludwig Schauwecker, Michael Bauer

Принадлежит: Airbus Defence and Space GmbH, EADS DEUTSCHLAND GmbH

An aircraft jet engine includes an exhaust-gas nozzle having a device configured to blow out an exhaust gas in a pulsating manner into an exhaust-gas stream so as to reduce noise. The exhaust-gas nozzle includes openings distributed along a circumference of the exhaust-gas nozzle and disposed upstream from a nozzle outlet. The openings communicate with the device.

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Номер записи: 42

01-01-2015 дата публикации

Rotational annular airscrew with integrated acoustic arrester

Номер: US20150000252A1

Автор: Kelly L. Boren, Matthew D. Moore

Принадлежит: Boeing Co

A propulsion system and methods are presented. A substantially tubular structure comprises a central axis through a longitudinal geometric center, and a first fan rotates around the central axis, and comprises a first fan hub and first fan blades. The fan hub is rotationally coupled to the substantially tubular structure, and the first fan blades are coupled to the first fan hub and increase in chord length with increasing distance from the first fan hub. A second fan is rotationally coupled to the substantially tubular structure and rotates around the central axis and contra-rotates relative to the first fan. Second fan blades are coupled to the second fan hub, and a nacelle circumscribing the first fan and the second fan is coupled to and rotates with the first fan.

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Номер записи: 43

06-01-2022 дата публикации

TURBINE

Номер: US20220003129A1

Автор: KUWAMURA Yoshihiro, SENOO Shigeki, TAKEI Sotaro, Watanabe Toshio

Принадлежит:

A turbine includes a shaft configured to rotate about a rotor axis; a pair of rotating blade rows, the pair of rotating blade rows including a pair of disks that extend radially outward from the shaft and are disposed at an interval in a direction of the rotor axis, each one of the pair of rotating blade rows including a plurality of rotating blades arranged in a circumferential direction on an outer peripheral end of the disk; and a pair of stator vane rows disposed in a one-to-one manner on a first side of the pair of rotating blade rows in the direction of the rotor axis, each one of the pair of stator vane rows including a plurality of stator vanes arranged in the circumferential direction, wherein a number of the rotating blades on each one of the pair of rotating blade rows is the same, and a number of the stator vanes on each one of the pair of stator vane rows is the same. 1. A turbine , comprising:a shaft configured to rotate about a rotor axis;a pair of rotating blade rows, the pair of rotating blade rows including a pair of disks that extend radially outward from the shaft and are disposed at an interval in a direction of the rotor axis, each one of the pair of rotating blade rows including a plurality of rotating blades arranged in a circumferential direction on an outer peripheral end of the disk; anda pair of stator vane rows disposed in a one-to-one manner on a first side of the pair of rotating blade rows in the direction of the rotor axis, each one of the pair of stator vane rows including a plurality of stator vanes arranged in the circumferential direction, whereina number of the rotating blades on each one of the pair of rotating blade rows is the same, and a number of the stator vanes on each one of the pair of stator vane rows is the same.2. The turbine according to claim 1 , whereinthe number of the stator vanes ranges from 30% to 70% of the number of the rotating blades.3. The turbine according to claim 1 , further comprising:an attachment ...

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Номер записи: 44

06-01-2022 дата публикации

FLEXIBLE COUPLING FOR GEARED TURBINE ENGINE

Номер: US20220003172A1

Автор: Otto John R., Sheridan William G.

Принадлежит:

A gas turbine engine includes a fan, a fan shaft coupled with the fan and arranged along an engine central axis, and a frame supporting the fan shaft. The frame defines a lateral frame stiffness (LFS). A non-rotatable flexible coupling and a rotatable flexible coupling support an epicyclic gear system. The couplings are subject to a Motion II of cantilever beam free end motion with respect to the engine central axis. The non-rotatable and the rotatable flexible couplings each have a stiffness of a common stiffness type under a common type of motion. The common stiffness type is a Stiffness B and the common type of motion is the Motion II. The Stiffness B of the rotatable flexible coupling is greater than the stiffness B of the non-rotatable flexible coupling, and a ratio of LFS/Stiffness B of the non-rotatable flexible coupling is in a range of 10-40. 1. A gas turbine engine comprising:a fan;a fan shaft coupled with the fan and arranged along an engine central axis;a frame supporting the fan shaft, the frame defining a lateral frame stiffness (LFS);an epicyclic gear system coupled to the fan shaft; anda non-rotatable flexible coupling and a rotatable flexible coupling supporting the epicyclic gear system, the non-rotatable flexible coupling and the rotatable flexible coupling being subject to a Motion II of cantilever beam free end motion with respect to the engine central axis,the non-rotatable flexible coupling and the rotatable flexible coupling each having a stiffness of a common stiffness type under a common type of motion with respect to the engine central axis, the common stiffness being defined with respect to the LFS, the common stiffness type is a Stiffness B and the common type of motion is the Motion II, the Stiffness B of the rotatable flexible coupling being greater than the stiffness of the non-rotatable flexible coupling, and a ratio of LFS/Stiffness B of the non-rotatable flexible coupling is in a range of 10-40.2. The gas turbine engine as recited ...

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Номер записи: 45

07-01-2016 дата публикации

APU EXHAUST SYSTEM

Номер: US20160001890A1

Автор: Jones Anthony C., Nager Eric Andrew, Napier James C., Rehman Farooq

Принадлежит: UNITED TECHNOLOGIES CORPORATION

An exhaust system for an APU in a gas turbine engine aircraft having a tail cone in which exhaust noise is reduced. An APU exhaust liner is attached to the APU and extending to an oval cutout at the tail cone The liner has a first constant cross section from the APU to a point proximate the tail cone, and a second, increasing cross section from the point proximate the tail cone to the liner cutout. The liner is turned toward the side of the tail. The liner cutout is further enlarged with at least one larger cutout. 1. An exhaust system for an auxiliary power unit (APU) in tail cone section of an aircraft , the exhaust system comprising:an exhaust outlet on a side of the tail cone section;an APU exhaust liner attached to the APU and extending rearward from the APU to the exhaust outlet, the liner having a first section with a constant cross section from the APU to a point proximate the tail cone; and the liner having a second, larger diameter continuously increasing cross section from the downstream end of the first section to the outlet.2. The exhaust system of claim 1 , wherein the liner is turned toward a side of the tail at the tail cone.3. The exhaust system of claim 1 , where in the first diameter is about 27.8 cm (10.95 inches) and the second diameter is about 33.0 cm (13 inches).4. The exhaust system of claim 1 , wherein the exhaust outlet is oval in shape.5. The exhaust system of claim 4 , wherein the exhaust outlet is further enlarged with a second oval cutout of larger shape.6. The exhaust system of claim 5 , wherein the exhaust outlet is still further enlarged with a third oval cutout of still larger shape.7. A method for reducing the exhaust noise in an auxiliary power unit (APU) in a gas turbine engine aircraft having a tail cone claim 5 , the method comprising:receiving exhaust gas at a liner inlet adjacent the APU;directing the exhaust gas through the liner from the liner inlet to an outlet in a side surface of a fuselage tail section; andreducing the ...

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Номер записи: 46

05-01-2017 дата публикации

SEAL

Номер: US20170002686A1

Автор: SCOTHERN David P.

Принадлежит: ROLLS-ROYCE PLC

A hydraulic seal arrangement including first and second rotatable components, the first defining a first annular trough defined by radially inwardly extending first and second walls, the second component defining a radially outwardly extending web between the walls. The first component includes a second trough defined by the second wall and a third wall axially spaced from the second, the second trough defining an open radially inner end defined by a radially inner end of the second wall such that the first and second annular troughs fluidly communicate around the circumference of the first component via the inner end of the second wall. At a first circumferential position, the second wall of the first component defines a second trough oil inlet providing further fluid communication between the first and second troughs and at a second circumferential position, the third wall of the first component defines a second trough oil outlet. 1. A hydraulic seal arrangement for a rotating machine comprising:first and second relatively rotatable components, the first component defining a first annular trough defined by radially inwardly extending axially spaced first and second walls, the second component defining a radially outwardly extending web extending between the radially inwardly extending walls of the first component; the first component further comprising:a second trough defined by the radially inwardly extending second wall and a radially inwardly extending third wall axially spaced from the second radially extending wall, the second trough defining an open radially inner end defined by a radially inner end of the second radially extending wall such that the first and second annular troughs fluidly communicate around substantially the whole circumference of the first component via the radially inner end of the second radially extending wall; whereinat a first circumferential position, the second radially inwardly extending wall of the first component defines a ...

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Номер записи: 47

05-01-2017 дата публикации

AUXILIARY OIL SYSTEM FOR GEARED GAS TURBINE ENGINE

Номер: US20170002738A1

Автор: Sheridan William G.

Принадлежит: UNITED TECHNOLOGIES CORPORATION

A gas turbine engine comprises a fan drive turbine, a fan rotor, and a gear reduction driven by the fan drive turbine to, in turn, drive the gear architecture. A main oil supply system supplies oil to components within the gear reduction, and an auxiliary oil supply system. The auxiliary oil system operates to ensure that the gear reduction will be adequately supplied with lubricant for at least seconds at power should the main oil supply system fail. 1. A gas turbine engine comprising:a fan drive turbine, a fan rotor, and a gear reduction driven by said fan drive turbine to, in turn, drive said gear architecture, a main oil supply system for supplying oil to components within said gear reduction, and an auxiliary oil supply system; andsaid auxiliary oil system being operable to ensure that the gear reduction will be adequately supplied with lubricant for at least 30 seconds at power should the main oil supply system fail.2. The gas turbine engine as set forth in claim 1 , wherein said gear reduction includes a sun gear being driven by said fan drive turbine to drive intermediate gears that engage a ring gear.3. The gas turbine engine as set forth in claim 2 , wherein said sun gear claim 2 , said intermediate gears and said ring gear are enclosed in a bearing compartment claim 2 , which captures oil removed via a scavenge line connected to a main oil pump.4. The gas turbine engine as set forth in claim 3 , wherein said main oil pump has a gutter that directs scavenged oil to a main oil tank.5. The gas turbine engine as set forth in claim 4 , wherein oil in said main oil tank feeds a main pump pressure stage which then delivers oil to said gear reduction.6. The gas turbine engine as set forth in claim 5 , wherein oil from said main pump pressure stage passes through a lubrication system that includes at least one filter and at least one heat exchanger to cool the oil.7. The gas turbine engine as set forth in claim 4 , wherein said gear reduction is surrounded by an ...

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Номер записи: 48

05-01-2017 дата публикации

GAS TURBINE FUEL COMPONENTS

Номер: US20170002743A1

Автор: Gauthier Donald, Jorgensen Stephen W., Matys Pawel, McNally Joshua R., Paskin Marc, Pennell Douglas, Ruck Thomas R., Sam Bernard Tam-Yen, Stuttaford Peter John

Принадлежит:

Gas turbine combustion systems and fuel cartridge assemblies are provided. An exemplary combustion system may comprise a combustor including one or more components, such as a cylindrical combustion liner, a flow sleeve, a main mixer, a radial inflow swirler, a combustor dome, and a fuel cartridge assembly. An exemplary fuel cartridge assembly may comprise first and second fuel manifolds which are connected to respective fuel circuits which supply fuel, such as liquid fuel, through a plurality of fuel passages within the fuel cartridge assembly or to other locations within an associated combustor. The fuel cartridge assembly may further include a plurality of fuel injector tips located at a tip plate of the fuel cartridge assembly through which fuel may be supplied to an associated combustor. 1. A fuel cartridge assembly , comprising: an aft portion comprising at least a first fuel manifold and a second fuel manifold;', 'a main body extending from the aft portion and having a first passageway contained therein;', 'a plurality of fuel passages extending axially from the aft portion to a tip plate, each fuel passage of the plurality of fuel passages in communication with the first fuel manifold or the second fuel manifold; and', 'the tip plate coupled to an end of the main body opposite the aft portion, the tip plate having at least one ignition opening and a plurality of openings corresponding to the plurality of axially extending fuel passages., 'a centerbody comprising2. The assembly of claim 1 , wherein the centerbody further comprises a plurality of fuel injector tips coupled to the respective plurality of fuel passages claim 1 , each of the plurality of fuel injector tips circumscribed at least partially by one of the plurality of openings in the tip plate.3. The assembly of claim 2 , wherein the fuel injector tips are positioned in the respective plurality of openings such that the fuel injector tips are movable relative to the tip plate.4. The assembly of claim ...

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Номер записи: 49

07-01-2016 дата публикации

FAN DRIVE GEAR SYSTEM SPLINE OIL LUBRICATION SCHEME

Номер: US20160003090A1

Автор: Lin Ning

Принадлежит:

An input coupling for a fan drive gear system includes features for maintaining lubricant within a splined interface. The fan drive gear system includes a gear rotatable about an axis that includes an inner spline. The input coupling includes an outer spline engaged to the inner spline of the gear. The input coupling includes an aft oil dam for maintaining lubricant within an interface between the outer spline and the inner spline. 1. A fan drive gear system comprising:a gear rotatable about an axis, the gear including an inner spline; andan input coupling including an outer spline engaged to the inner spline of the gear, the input coupling including an aft oil dam for maintaining lubricant within an interface between the outer spline and the inner spline.2. The fan drive gear system as recited in claim 1 , wherein the gear includes a forward tab extending radially inward from an inner surface of the gear forward of the inner spline.3. The fan drive gear system as recited in claim 1 , wherein the gear includes an aft tab extending radially inward from an inner surface of the gear aft of the inner spline and forward of the aft oil dam.4. The fan drive gear system as recited in claim 1 , wherein the aft oil dam includes a retaining ring supported within an annular channel of the input coupling.5. The fan drive gear system as recited in claim 4 , wherein the retaining ring extends radially outward into contact with an inner surface of the gear.6. The fan drive gear system as recited in claim 1 , wherein the gear comprises a sun gear.7. The fan drive gear system as recited in claim 1 , wherein the input coupling includes at least one U-shaped portion for accommodating relative movement and misalignment with the gear.8. A geared turbofan engine comprising:a fan rotatable about an engine axis;a core engine including a turbine section driving a turbine shaft;a gearbox including a sun gear driven by the turbine shaft; andan input coupling transferring power between the ...

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Номер записи: 50

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: 51

07-01-2016 дата публикации

Integrated Flex Support and Front Center Body

Номер: US20160003105A1

Автор: Cigal Brian P., Davis Todd A.

Принадлежит:

A gas turbine engine is provided. The gas turbine engine may include a geared architecture, a central body support and a bearing package. The geared architecture may interconnect a spool and a fan rotatable about an axis. The central body support may provide an annular wall for a core flow path and an integral flex support inwardly extending therefrom. The integral flex support may couple the geared architecture to the central body support. The bearing package may include a bearing support removably coupled to the integral flex support. 1. A gas turbine engine , comprising:a geared architecture interconnecting a spool and a fan rotatable about an axis;a central body support providing an annular wall for a core flow path and an integral flex support inwardly extending therefrom, the integral flex support coupling the geared architecture to the central body support; anda bearing package having a bearing support removably coupled to the integral flex support.2. The gas turbine engine of claim 1 , wherein the integral flex support includes at least one flex member configured to at least partially suppress vibrations within the gas turbine engine.3. The gas turbine engine of claim 2 , wherein the at least one flex member is geometrically structured to at least partially suppress vibrations within the gas turbine engine.4. The gas turbine engine of claim 2 , wherein the at least one flex member is configured to at least partially suppress vibrations between the central body support and at least the geared architecture.5. The gas turbine engine of claim 1 , wherein the integral flex support is disposed substantially aft of the geared architecture so as to provide sufficient axial clearance to an oil manifold associated with the geared architecture.6. The gas turbine engine of claim 1 , wherein at least the integral flex support and the annular wall are formed of a unitary body.7. The gas turbine engine of claim 1 , wherein the integral flex support is configured to form an ...

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Номер записи: 52

07-01-2016 дата публикации

Method for the production of a curved ceramic sound attenuation panel

Номер: US20160003106A1

Автор: Eddy Goulliane, Eric Philippe, Sebastien Jimenez, Stephanie Fouquet

Принадлежит: Herakles SA

A method of fabricating a sound attenuation panel of curved shape, the method including impregnating a fiber structure defining a cellular structure with a ceramic precursor resin; polymerizing the ceramic precursor resin while holding the fiber structure on tooling presenting a curved shape corresponding to the final shape of the cellular structure; docking the cellular structure with first and second skins, each formed by a fiber structure impregnated with a ceramic precursor resin, each skin being docked to the cellular structure before or after polymerizing the resin of the skins; pyrolyzing the assembly constituted by the cellular structure and the first and second skins; and densifying the assembly by chemical vapor infiltration.

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Номер записи: 53

07-01-2016 дата публикации

Damping device for a gas turbine, gas turbine and method for damping thermoacoustic oscillations

Номер: US20160003162A1

Автор: Christian Beck, Jaap Van Kampen

Принадлежит: SIEMENS AG

A damping device for a gas turbine has at least one Helmholtz resonator and at least one duct, wherein the Helmholtz resonator has a resonator housing and at least one resonator neck pipe and the resonator housing encloses a resonance volume of the Helmholtz resonator, into which volume acoustic vibrations can be injected by means of the resonator neck pipe. The damping device enables a particularly effective damping of thermo-acoustic vibrations. For this purpose, the duct is formed with a duct jacket and at least one outlet opening. Acoustic vibrations of a fluid stream flowing through a burner plenum and a combustion chamber can be injected into the outlet opening. A cooling fluid can be applied to the duct and the at least one resonator neck pipe opens on the hot-gas side into such a duct upstream of the at least one outlet opening.

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Номер записи: 54

05-01-2017 дата публикации

Gas turbine control system

Номер: US20170003032A1

Автор: Bernard Tam-Yen SAM, Donald Gauthier, Douglas Pennell, Joshua R. Mcnally, Marc PASKIN, Pawel MATYS, Peter John Stuttaford, Stephen W. Jorgensen, Thomas R. RUCK

Принадлежит: Ansaldo Energia IP UK Ltd

Gas turbine combustion systems and fuel cartridge assemblies are provided. An exemplary combustion system may comprise a combustor including a cylindrical combustion liner, a flow sleeve, a main mixer, a radial inflow swirler, a combustor dome, and a fuel cartridge assembly. An exemplary combustor and/or fuel cartridge assembly may comprise first and second fuel circuits or manifolds. Methods and systems are also provided for staging and controlling a flow of fuel and/or water through different fuel circuits and pilot injectors, to allow purging and ignition using different fuel circuits, pilot injectors, and fuel sources.

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Номер записи: 55

04-01-2018 дата публикации

SUPRESSING VIBRATIONS OF SHAFTS USING ADJUSTABLE BEARINGS

Номер: US20180003075A1

Автор: Chasalevris Athanasios, Dohnal Fadi

Принадлежит:

A bearing configured to actively damp vibration of a shaft in a turbine. In one implementation, the bearing can include actuating members that move in a manner that changes properties of fluid, typically a thin film of lubricant, disposed in the bearing to facilitate rotation of the shaft. These changes effectively manipulate the stiffness and damping of the thin film according to a time periodicity that matches a parametric anti-resonance of the bearing. In turn, the resulting interaction of vibrating modes is favorable to damp vibration amplitudes at critical speeds. 1. A bearing , comprising:a casing having a center axis;a first pad and a second pad disposed in the casing, each forming an arcuate carrying surface and radially moveable in a direction perpendicular to the center axis of the casing; andan actuator coupled with the first pad and the second pad,wherein the actuator is configured to displace the first pad and the second pad harmonically relative to the center axis.2. The bearing of claim 1 , wherein the first pad and the second pad are spaced apart annularly from one another about the center axis.3. The bearing of claim 1 , wherein the first pad and the second pad oppose each other on opposite sides of the center axis.4. The bearing of claim 1 , wherein the first pad and the second pad have adjacent ends that are annularly offset from one another by less than 180°.5. The bearing of claim 1 , wherein the arcuate carrying surface has an arc length that is the same as between the first pad and the second pad.6. The bearing of claim 1 , wherein the arcuate carrying surface has an arc length that is different as between the first pad and the second pad.7. The bearing of claim 1 , further comprising a third pad also having the arcuate carrying surface and radially moveable in a direction perpendicular to the center axis claim 1 , wherein the arcuate carrying surface of the first pad claim 1 , the second pad claim 1 , and the third pad claim 1 , in aggregate ...

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Номер записи: 56

04-01-2018 дата публикации

MULTIPLE RESERVOIR LUBRICATION SYSTEM

Номер: US20180003082A1

Автор: James Denman H.

Принадлежит:

A lubrication system for use with a gas turbine engine includes a first reservoir for containing a lubricant. The first reservoir includes a first discharge passage through which the lubricant is flowable in a first direction. A second reservoir contains the lubricant. The second reservoir includes a second discharge passage through which the lubricant is flowable in a second direction. The first direction is generally opposite to the second direction. A first pump pumps the lubricant from the first reservoir. A second pump pumps the lubricant from the second reservoir. A manifold distributes the lubricant to a component. The lubricant from the first pump and the second pump flows into the manifold and exits the manifold through a manifold discharge. 1. A lubrication system for use with a gas turbine engine comprising:a first reservoir for containing a lubricant, wherein the first reservoir includes a first discharge passage through which the lubricant is flowable in a first direction;a second reservoir for containing the lubricant, wherein the second reservoir includes a second discharge passage through which the lubricant is flowable in a second direction, wherein the first direction is generally opposite to the second direction;a first pump that pumps the lubricant from the first reservoir;a second pump that pumps the lubricant from the second reservoir; anda manifold to distribute a constant and uninterrupted supply of the lubricant to a bearing, wherein the lubricant from the first pump and the second pump flows into the manifold to combine into a common flow, and the common flow exits the manifold through a common manifold discharge.2. The lubrication system as recited in wherein the bearing is a fan journal bearing of a gas turbine engine.3. The lubrication system as recited in wherein the first direction is substantially upwardly and the second direction is substantially downwardly.4. The lubrication system as recited in wherein an output of each of the ...

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Номер записи: 57

04-01-2018 дата публикации

RADIALLY CONNECTED CASCADE GRIDS

Номер: US20180003129A1

Автор: Aten Michael Ray, Stein Adam

Принадлежит: Rohr, Inc.

In various embodiments, a cascade array may comprise an actuator, a first cascade having a first integral flange, and a second cascade having a second integral flange, wherein the first cascade and the second cascade are operatively coupled to one another via the first integral flange and the second integral flange to form a cascade assembly, and the actuator is disposed between the first cascade and the second cascade. 1. A cascade array , comprising:an actuator;a first cascade having a first integral flange; and wherein the first cascade and the second cascade are operatively coupled to one another via the first integral flange and the second integral flange to form a cascade assembly,', 'the actuator is disposed between the first cascade and the second cascade, and', 'the first integral flange and the second integral flange extend towards one another and extend past at least a portion of the actuator., 'a second cascade having a second integral flange,'}2. The cascade array of claim 1 , wherein the cascade assembly is configured to at least partially surround the actuator.3. The cascade array of claim 1 , wherein the cascade assembly is configured to isolate the actuator from radial loads and hoop loads.4. The cascade array of claim 1 , further comprising a splice plate.5. The cascade array of claim 4 , wherein the splice plate is configured to operatively couple the first cascade and the second cascade together to form the cascade assembly.6. The cascade array of claim 1 , wherein the cascade assembly is housed with a translating sleeve when a thrust reverser system is in the stowed position.7. The cascade array of claim 1 , wherein the cascade assembly is operatively coupled to a torque box.8. The cascade array of claim 5 , wherein the splice plate is disposed radially outward from the first integral flange and the second integral flange.9. The cascade array of claim 8 , wherein a radially inner surface of the splice plate is coupled to a radially outer surface ...

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Номер записи: 58

04-01-2018 дата публикации

INTERNAL GEAR PUMP

Номер: US20180003173A1

Автор: CHENOUX MATHIEU, Cornet Albert

Принадлежит: SAFRAN AERO BOOSTERS S.A.

An internal gear pump includes a pinion, a ring arranged around the pinion, and a cylindrical wall arranged around the ring. A support element, on which the pinion and the ring are supported, carries high-pressure liquid towards a recess located at the junction between the ring and the cylindrical wall, and also carries low-pressure liquid towards another recess located at another point of the junction between the ring and the cylindrical wall. The recess allows the load of the ring on the cylindrical wall to be reduced. 1. An internal gear pump comprising:a pinion capable of rotating about a first axis of rotation;a ring arranged around the pinion, having a cylindrical periphery and being configured to rotate about a second axis of rotation that is different to the first axis of rotation and parallel thereto, the pinion and the ring being arranged to delimit a work space comprising a high-pressure space and a low-pressure space;a first support element arranged on a first side of the ring to limit the movement thereof in a first sense along a first orientation that is substantially parallel to the first and second axes of rotation;a second support element arranged on a second side of the ring to limit the movement thereof in a second sense that is opposite the first sense and in the first orientation;a cylindrical wall arranged around the ring to limit the movement thereof in a plane that is substantially perpendicular to the second axis of rotation;a first high-pressure recess in an arc of a circle and in fluid contact with a first portion of the cylindrical periphery of the ring;a first low-pressure recess in an arc of a circle and in fluid contact with a second portion of the cylindrical periphery of the ring, the first high-pressure recess and the first low-pressure recess being separated;a first high-pressure conveying means forming a fluid connection between the high-pressure space and the first high-pressure recess; anda first low-pressure conveying means ...

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Номер записи: 59

02-01-2020 дата публикации

GAS TURBINE

Номер: US20200003068A1

Автор: FELSCHER Patrick, FONTANA Luca, HERZOG Christoph, HOEFGEN Mathias, JAHIC Hana

Принадлежит:

A gas turbine engine, in particular an aircraft engine, including: a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear, the sun gear is connected to the input shaft device, the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, with an inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device. 1. A gas turbine engine , in particular an aircraft engine , comprising:a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear,the sun gear is connected to the input shaft device,the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, withan inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device.2. The gas turbine of claim 1 , wherein the inter-shaft bearing system is located axially within or in front of a low-pressure compressor or an intermediate compressor.3. The gas turbine of claim 1 , wherein the inter-shaft bearing system is axially adjacent to the gearbox device on the input and/or the output side claim 1 , in particular with an axial distance measured from the centreline of the gearbox between 0.001 and 4 times the inner radius of the inter-shaft bearing system.4. The gas turbine of claim 1 , wherein the inter-shaft bearing device comprises at least one ball bearing.5. The gas turbine of claim 1 , wherein a fan shaft bearing system is radially located between a fan shaft as part of the output shaft device and a static structure 1 , in particular a static front cone structure 1 , in particular the fan shaft bearing system being axially positioned within the width of the ...

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Номер записи: 60

02-01-2020 дата публикации

Bearing device for load reduction

Номер: US20200003075A1

Автор: Martin BOEHME, Timo SCHATKE

Принадлежит: Rolls Royce Deutschland Ltd and Co KG

A bearing assembly for a gas turbine engine comprises a bearing; a bearing bracket, which holds the bearing and is secured by a predetermined breaking device on a connecting element, which can be connected or is connected to a support structure of the gas turbine engine; and a clutch for transmitting a torque from a first clutch element connected in a fixed manner to the rotor of the bearing to a second clutch element supported on the bearing bracket, wherein the clutch elements are spaced apart when the predetermined breaking device is intact and can be brought into contact with one another by destruction of the predetermined breaking device. A gas turbine engine and a method are furthermore provided.

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Номер записи: 61

02-01-2020 дата публикации

Jacket ring assembly for a turbomachine

Номер: US20200003076A1

Автор: Feldmann Manfred

Принадлежит:

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

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Номер записи: 62

02-01-2020 дата публикации

Acoustic panel for thrust reversers

Номер: US20200003124A1

Автор: Alexandre D. Curaudeau, Brian L. Riedel, II Taurus L. Brackett

Принадлежит: Boeing Co

An acoustic panel includes a base, a cantilevered portion, a gap, and a support member. The base has a surface defining a plurality of cavities configured to attenuate noise from an engine. The cantilevered portion extends from the base and is configured to be removably coupled with a portion of a transcowl. The gap is defined by the base and the cantilevered portion. The support member is coupled to the cantilevered portion and the base, and the supporting member is configured to support the cantilevered portion.

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Номер записи: 63

02-01-2020 дата публикации

GAS TURBINE

Номер: US20200003129A1

Автор: BONIFACE Dominic, HUGHES Geoff, KNIGHT Glenn, SIMMS Paul, SWIFT Andrew, THORNTON Stewart

Принадлежит:

The invention relates to a gas turbine engine, in particular an aircraft engine, comprising: a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear, the sun gear is connected to the input shaft device, the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, with an inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device. 1. A gas turbine engine , in particular an aircraft engine , comprising:a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear,the sun gear is connected to the input shaft device,the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, withan inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device, witha carrier bearing system being located radially between the input shaft device and a static structure supporting the carrier bearing system, the support connection being axially in front of the input side of the gearbox device.2. The gas turbine of claim 1 , wherein the inter-shaft bearing system is located axially within or in front of a low-pressure compressor or an intermediate compressor.3. The gas turbine of claim 1 , wherein the inter-shaft bearing system is axially adjacent to the gearbox device on the input and/or the output side claim 1 , in particular with an axial distance measured from the centreline of the gearbox between 0.001 and 4 times the inner radius of the inter-shaft bearing system.4. The gas turbine of claim 1 , wherein the inter-shaft bearing device comprises at least one ball bearing.5. The gas turbine of claim 1 , wherein a fan shaft ...

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Номер записи: 64

07-01-2021 дата публикации

COMPRESSOR BLADE HAVING ORGANIC VIBRATION STIFFENER

Номер: US20210003017A1

Автор: Garner Chad, JARAMILLO ANDRES

Принадлежит:

A compressor blade of a gas turbine includes a root member; an airfoil that is disposed on the root member and includes a first interior wall and a second interior wall forming a hollow space defined between the first and second interior walls; and an organic vibration stiffener (OVS) formed on at least one of the first interior wall and the second interior wall. The OVS is formed by 3D printing performed with respect to a surface of the at least one of the first interior wall and the second interior wall and includes an uneven surface formed on at least part of the at least one of the first interior wall and the second interior wall. The OVS may include a protruded or recessed portion protruding from or recessed into at least part of the at least one of the first interior wall and the second interior wall. 1. A compressor blade of a gas turbine , comprising:a root member;an airfoil that is disposed on the root member and includes a first interior wall and a second interior wall forming a hollow space defined between the first and second interior walls; andan organic vibration stiffener (OVS) formed on at least one of the first interior wall and the second interior wall.2. The compressor blade according to claim 1 , wherein the OVS is formed by 3D printing performed with respect to a surface of the at least one of the first interior wall and the second interior wall.3. The compressor blade according to claim 1 , wherein the OVS includes an uneven surface formed on at least part of the at least one of the first interior wall and the second interior wall.4. The compressor blade according to claim 1 , wherein the OVS includes a protruded portion protruding from at least part of the at least one of the first interior wall and the second interior wall.5. The compressor blade according to claim 1 , wherein the OVS includes a recessed portion recessed into at least part of the at least one of the first interior wall and the second interior wall.6. The compressor blade ...

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Номер записи: 65

07-01-2021 дата публикации

PROFILED STRUCTURE AND ASSOCIATED TURBOMACHINE

Номер: US20210003074A1

Автор: GEA AGUILERA Fernando, GRUBER Mathieu Simon Paul, Riou Georges Jean Xavier

Принадлежит: SAFRAN AIRCRAFT ENGINES

An airflow profiled structure having a profiled leading edge. The profiled leading edge having, along a leading edge line, a serrated profile line with a succession of teeth and depressions. The airflow profiled structure also includes a porous acoustically absorbent region located towards the bottom of the depressions. 1. A profiled air flow structure comprising:a body;porous acoustically absorbent regions;an upstream leading edge and/or a downstream trailing edge; andalong the upstream leading edge and/or the downstream trailing edge line, a serrated profile line showing a succession of teeth and depressions,wherein the porous acoustically absorbent regions locally form bottoms for the depressions where the porous acoustically absorbent regions occupy a part of the body to define, together with the body, the serrated profile line at the upstream leading edge and/or the downstream trailing edge.2. The profiled structure according to further comprising: along the chord, the serrated profile line has a maximum amplitude, h, and', {'br': None, 'i': 'd=h/', '10, within 30%.'}, 'the porous acoustically absorbent region has a geometric centre located at a distance d downstream of the upstream leading edge or upstream of the downstream trailing edge, at the bottom of the depressions such that], 'between upstream and downstream, a chord in which3. The profiled structure according to further comprising: along the upstream leading edge or the downstream trailing edge, the serrated profile line has a distance between two consecutive tooth tips,', 'along the chord, the serrated profile line has a maximum amplitude, h, and', along the upstream leading edge and/or the downstream trailing edge, two limits separated by a distance a such that a is equal to one third of said distance between two consecutive tooth tips, to within 30%,', 'along the chord, two limits separated by a distance b such that b=h/3, within 30%., 'the porous acoustically absorbent region has], 'between ...

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Номер записи: 66

03-01-2019 дата публикации

METHOD FOR ALTERING THE LAW OF TWIST OF THE AERODYNAMIC SURFACE OF A GAS TURBINE ENGINE FAN BLADE

Номер: US20190003313A1

Автор: COURTIER Vivien Michael

Принадлежит: SAFRAN AIRCRAFT ENGINES

A method of altering the twisting relationship for the aerodynamic surface of a fan blade of a gas turbine engine, wherein the following steps are performed: establishing, for a portion of the aerodynamic surface of the fan blade, an alteration relationship defined by variation of a pitch angle of the blade as a function of radial height along the blade, the alteration relationship including alterations that are each defined by a height along with the radial height of the fan blade and by an amplitude; and applying the alteration relationship as established in this way to an initial twisting relationship of the fan blade so as to obtain an altered twisting relationship for the fan blade, the initial twisting relationship being defined by a polynomial for the radial height of the fan blade as a function of its pitch angle. 1. A method of altering the twisting relationship for the aerodynamic surface of a fan blade of a gas turbine engine , the method comprising:establishing, for a portion of the aerodynamic surface of the fan blade, an alteration relationship defined by variation of a pitch angle of the blade as a function of radial height along the blade, said alteration relationship comprising alterations that are each defined by a height along with the radial height of the fan blade and by an amplitude; andapplying the alteration relationship as established in this way to an initial twisting relationship of the blade so as to obtain an altered twisting relationship for the fan blade, said initial twisting relationship being defined by a polynomial for the radial height of the fan blade as a function of its pitch angle.2. The method according to claim 1 , wherein the alteration relationship is defined in such a manner as to be zero and to have a derivative of zero at least one end of the portion of the aerodynamic surface of the fan blade.3. The method according to claim 1 , wherein the alteration relationship is also defined in such a manner that the amplitude of ...

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Номер записи: 67

03-01-2019 дата публикации

SYSTEM FOR MOUNTING A VIBRATION SENSOR ONTO A MACHINE

Номер: US20190003331A1

Автор: Neuman Jena Le, Sampathkumar Varadharajan, Samuelian Michael Lawrence, Tse Kwok-Kwong Ben

Принадлежит:

A system for mounting at least one vibration sensor onto a machine includes a mounting block that is funned from a pre-determined material and of a form such that the pre-determined material and form of the mounting block render the mounting block with a natural frequency lying at least 1500 Hertz (Hz) away from a range of frequencies associated with operation of the machine. 1. A system for mounting at least one vibration sensor onto a machine , the system comprising:a mounting block formed from a pre-determined material and having a form such that the pre-determined material and form of the mounting block render the mounting block with a natural frequency lying at least 1500 Hertz (Hz) away from a range of frequencies associated with Operation of the machine.2. The system of claim 1 , wherein the pre-determined material includes at least one of: a 410 Grade Stainless Steel per ASTM A479 Martensitic Grade Condition 2 or 3 claim 1 , and a 410 Stainless Steel per ASTM A276 Condition H.3. The system of claim 1 , wherein the mounting block is configured to define a mounting surface for facilitating a mounting of the at least one vibration sensor thereon.4. The system of claim 3 , wherein the mounting block includes a hole that extends through a pair of opposing sidewalls claim 3 , at least one of the opposing sidewalls being located adjacent to the mounting surface.5. The system of further comprising a securement system that is configured to engage with the hole far securing the mounting block to the machine.6. The system of claim 5 , wherein the securement system includes:a washer located on one of the opposing sidewalk and disposed about the hole; anda fastener received within a hole defined by the washer and the hole of the mounting block, the fastener configured to engage with a threaded receptacle located on the machine.7. A system for mounting at least one vibration sensor onto a machine claim 5 , the system comprising: a first portion having a mounting surface ...

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Номер записи: 68

03-01-2019 дата публикации

TRANSFER COUPLINGS

Номер: US20190003336A1

Автор: ADAMCZEWSKI Ulrich, COLEBROOKE Jack Francis, UHKÖTTER Stephan

Принадлежит:

A transfer coupling includes a static component and a rotatable component arranged concentrically, the static component including a first number of radially extending ports and the rotatable component having a second number of radially extending ports the radially extending ports arranged in a common circumferential plane wherein the ports are configured and arranged, in use, to homogenise a flow area for a fluid being transferred through the ports and thereby create a homogenous volume flow. 1. A transfer coupling comprising a static component and a rotatable component arranged in co-axial alignment , the static component including a first number of ports and the rotatable component including a second number of ports the ports arranged in a common circumferential plane wherein the ports are configured and arranged , in use , to homogenise a flow area for a fluid being transferred through the ports and thereby create a homogenous volume flow , and wherein a first of the static component and the rotatable component comprises a circumferential array comprising a first number of ports of a first size and the second of the components comprises a circumferential array comprising a second number of ports of a second size , the first number being substantially smaller than the second number and the first size being substantially larger than the second size.2. A transfer coupling as claimed in wherein the static component and rotatable component are concentrically arranged and the first and second number of ports extend radially.3. A transfer coupling as claimed in wherein the static component and rotatable component are arranged axially adjacent to each other and the ports extend axially.4. A transfer coupling as claimed in wherein the arrays of ports on the components have similar packing factors.5. A transfer coupling as claimed in wherein the second component has an array of ports comprising a single row of equally spaced circular ports having a first diameter D claim 1 ...

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Номер записи: 69

03-01-2019 дата публикации

Exhaust Stack Assemblies with Acoustic Attenuation Features

Номер: US20190003358A1

Автор: Johnsen Matthew Alan, KIPPEL Bradly Aaron, Loud Richard Lynn

Принадлежит:

An exhaust stack assembly includes an exhaust stack having an internal surface that defines an interior of the exhaust stack. The exhaust stack is configured to receive exhaust gas from at least one turbomachine component and exhaust the exhaust gas to atmosphere. The exhaust gas assembly further includes a plurality of attenuation assemblies disposed in the interior, each of the plurality of attenuation assemblies including a base substrate generally oriented in the direction of flow of the exhaust gas through the interior, each of the plurality of attenuation assemblies further including a plurality of attenuation modules mounted to the base substrate. Each of the plurality of attenuation modules includes a fiber mesh. The fiber mesh is exposed to the exhaust gas in the interior. 1. An exhaust stack assembly , comprising:an exhaust stack having an internal surface that defines an interior of the exhaust stack, the exhaust stack configured to receive exhaust gas from at least one turbomachine component and exhaust the exhaust gas to atmosphere; anda plurality of attenuation assemblies disposed in the interior, each of the plurality of attenuation assemblies comprising a base substrate generally oriented in the direction of flow of the exhaust gas through the interior, each of the plurality of attenuation assemblies further comprising a plurality of attenuation modules mounted to the base substrate, each of the plurality of attenuation modules comprising a fiber mesh, wherein the fiber mesh is exposed to the exhaust gas in the interior.2. The exhaust stack assembly of claim 1 , wherein a diameter of the fibers in the fiber mesh of each of the plurality of attenuation modules is between 0.5 and 10 microns.3. The exhaust stack assembly of claim 1 , wherein the fibers comprise at least one of ceramic fibers claim 1 , alkaline earth silicate fibers claim 1 , or polycrystalline wool fibers.4. The exhaust stack assembly of claim 1 , wherein a density of each of the ...

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Номер записи: 70

03-01-2019 дата публикации

TWO-TURBINE GEARED ENGINE WITH LOW PRESSURE ENVIRONMENTAL CONTROL SYSTEM FOR AIRCRAFT

Номер: US20190003382A1

Автор: Schwarz Frederick M., Suciu Gabriel L.

Принадлежит:

A gas turbine engine assembly includes a fan section delivering air into a main compressor section. The main compressor section compresses air and delivers air into a combustion section. Products of combustion pass from the combustion section over a turbine section to drive the fan section and main compressor sections. A gearbox is driven by the turbine section to drive the fan section. A pylon supports the gas turbine engine. An environmental control system includes a higher pressure tap at a higher pressure location in the main compressor section, and a lower pressure tap at a lower pressure location. The lower pressure location being at a lower pressure than the higher pressure location. The lower pressure tap communicates to a first passage leading to a downstream outlet and a compressor section of a turbocompressor. The higher pressure tap leads into a turbine section of the turbocompressor such that air in the higher pressure tap drives the turbine section to in turn drive the compressor section of the turbocompressor. The pylon includes a lowermost surface and the higher pressure tap does not extend above a plane including the lowermost surface. A combined outlet of the compressor section and the turbine section of the turbocompressor intermixes and passes downstream to be delivered to an aircraft use. An environmental control system is also disclosed. 1. A gas turbine engine assembly comprising:a fan section delivering air into a main compressor section, said main compressor section compressing air and delivering air into a combustion section, products of combustion passing from said combustion section over a turbine section to drive said fan section and main compressor sections, wherein a gearbox is driven by said turbine section to drive said fan section;a pylon supporting the gas turbine engine;an environmental control system including a higher pressure tap at a higher pressure location in said main compressor section, and a lower pressure tap at a lower ...

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Номер записи: 71

03-01-2019 дата публикации

Aircraft thermal management system

Номер: US20190003391A1

Автор: Brian Merry, Gabriel L. Suciu, James D. Hill, Nathan Snape

Принадлежит: United Technologies Corp

An aircraft thermal management system includes a first fluid system containing a first fluid, a fluid loop containing a thermally neutral heat transfer fluid, a second fluid system containing a second fluid, a first heat exchanger configured to transfer heat from the first fluid to the thermally neutral heat transfer fluid, and a second heat exchanger configured to transfer heat from the thermally neutral heat transfer fluid to the second fluid. The fluid loop is configured to provide the thermally neutral heat transfer fluid to the first heat exchanger at a pressure that matches the pressure of the first fluid.

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Номер записи: 72

13-01-2022 дата публикации

Low pressure ratio fan engine having a dimensional relationship between inlet and fan size

Номер: US20220010689A1

Автор: Becky E. Rose, Jonathan Gilson, Robert E. Malecki, Wesley K. Lord, Yuan J. Qiu

Принадлежит: Raytheon Technologies Corp

A gas turbine engine assembly may include, among other things, a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane, a geared architecture, a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture, a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, a forward edge on the inlet portion in a second reference plane, and a length of the inlet portion having a dimension L between the first reference plane and the second reference plane. A dimensional relationship of L/D may be between 0.30 and 0.40.

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Номер записи: 73

13-01-2022 дата публикации

ACOUSTICALLY TREATED PANELS

Номер: US20220010731A1

Автор: Bifulco Anthony R., Lee Duncan R.C.

Принадлежит: Raytheon Technologies Corporation

A system for dampening noise generated by a gas turbine engine is disclosed. In various embodiments, the system includes a fan exit guide vane having a leading edge, a trailing edge and a pocket that extends in a chordwise direction between the leading edge and the trailing edge; and an acoustic panel configured to be received within the pocket, the acoustic panel including a back sheet, a face sheet and a core disposed between the back sheet and the face sheet, the core having a first cavity extending between a first wall and a second wall and a second cavity disposed within the first cavity. 1. A system for dampening noise generated by a gas turbine engine , comprising:a fan exit guide vane having a leading edge, a trailing edge and a pocket that extends in a chordwise direction between the leading edge and the trailing edge; andan acoustic panel configured to be received within the pocket, the acoustic panel including a back sheet, a face sheet and a core disposed between the back sheet and the face sheet, the core having a first cavity extending between a first wall and a second wall and a second cavity disposed within the first cavity.2. The system of claim 1 , wherein a first plurality of openings extends through the face sheet and into the first cavity.3. The system of claim 2 , wherein a second plurality of openings extends through the face sheet and into the second cavity.4. The system of claim 3 , wherein the first cavity is configured to form a first Helmholtz resonator configured to dampen a first resonant frequency.5. The system of claim 4 , wherein the second cavity is configured to form a second Helmholtz resonator configured to dampen a second resonant frequency.6. The system of claim 1 , wherein the first cavity is characterized by a first length in the chordwise direction of the fan exit guide vane and wherein the second cavity is characterized by a second length in the chordwise direction claim 1 , the second length being less than the first ...

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Номер записи: 74

13-01-2022 дата публикации

LASER IGNITION DEVICE, SPACE ENGINE, AND AIRCRAFT ENGINE

Номер: US20220010753A1

Автор: ITOU Mitsunori, IZAWA Jun, MATSUURA Yoshiki, NAGAO Takahisa, Sasaki Masahiro, Taira Takunori

Принадлежит:

A laser ignition device includes an excitation light source that generates excitation light, and a pulsed laser oscillator connected to the excitation light source, wherein the pulsed laser oscillator generates a plurality of pulsed light beams at a time of one ignition to produce an initial flame. 1. A laser ignition device comprising:an excitation light source that generates excitation light; anda pulsed laser oscillator connected to the excitation light source, whereinthe pulsed laser oscillator generates a plurality of pulsed light beams at a time of one ignition to produce an initial flame.2. The laser ignition device according to claim 1 , whereinthe pulsed laser oscillator generates a plurality of pulsed light beams by burst light emission.3. The laser ignition device according to claim 1 , further comprisingan optical fiber that connects the excitation light source and the pulsed laser oscillator to each other.4. The laser ignition device according to claim 1 , whereinthe pulsed laser oscillator includes a laser crystal and a Q-switch that generates pulsed light beam.5. A space engine comprising:{'claim-ref': {'@idref': 'CLM-00001', 'claim 1'}, 'the laser ignition device according to ; and'}a combustor that burns a fuel.6. An aircraft engine comprising:{'claim-ref': {'@idref': 'CLM-00001', 'claims 1'}, 'the laser ignition device according to ; and'}a combustor that burns a fuel.7. The laser ignition device according to claim 2 , further comprisingan optical fiber that connects the excitation light source and the pulsed laser oscillator to each other.8. The laser ignition device according to claim 2 , whereinthe pulsed laser oscillator includes a laser crystal and a Q-switch that generates pulsed light beam.9. The laser ignition device according to claim 3 , whereinthe pulsed laser oscillator includes a laser crystal and a Q-switch that generates pulsed light beam.10. The laser ignition device according to claim 7 , whereinthe pulsed laser oscillator includes ...

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Номер записи: 75

11-01-2018 дата публикации

Dynamic Resonance System and Method for the Anti-Icing and De-Icing of Inlet Grids

Номер: US20180009009A1

Автор: Peduchi Andrew, Schneider William, Wood Yvette

Принадлежит:

In one embodiment, a system includes an inlet grid configured to reduce distortion of an incoming airflow. The system may also include a vibration device coupled to the inlet grid and a controller communicatively coupled to the vibration device. The controller may transmit a vibration signal to the vibration device causing the vibration device to vibrate the inlet grid such that the inlet grid resonates at a natural frequency inducing a mode shape in the inlet grid. The mode shape may break up and prevent ice on the inlet grid. 1. A system , comprising:an inlet grid configured to reduce distortion of an incoming airflow;a vibration device coupled to the inlet grid; anda controller communicatively coupled to the vibration device, the controller configured to transmit a vibration signal to the vibration device; andwherein the vibration signal is operable to cause the vibration device to vibrate the inlet grid such that the inlet grid resonates at a natural frequency, thereby inducing a mode shape in the inlet grid, the mode shape configured to break up and prevent ice on the inlet grid.2. The system of claim 1 , wherein:the vibration device is a first vibration device, the vibration signal is a first vibration signal, the natural frequency is a first natural frequency, and the mode shape is a first mode shape; a second vibration device coupled to the inlet grid;', 'the controller is further configured to transmit a second vibration signal to the second vibration device; and', 'the second vibration signal is operable to cause the second vibration device to vibrate the inlet grid such that the inlet grid resonates at a second natural frequency, thereby inducing a second mode shape, the second mode shape configured to break up and prevent ice on the inlet grid., 'the system further comprises3. The system of claim 2 , wherein the first vibration device has a first excitation direction and the second vibration device has a second excitation direction.4. The system of claim ...

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Номер записи: 76

27-01-2022 дата публикации

NOISE SIGNATURE PROXIMITY WARNING SYSTEM AND FEEDBACK MECHANISM FOR AIRCRAFT

Номер: US20220024600A1

Автор: CLARK Dereck, Hawley Christopher, SURACE Lawrence

Принадлежит: HONEYWELL INTERNATIONAL INC.

A noise reduction system in a vertical takeoff and landing (VTOL) vehicle is provided. The noise reduction system is configured to: identify a noise level at which the VTOL vehicle can operate; dynamically determine a motor-specific fan RPM and a motor-specific fan pitch that will allow the vehicle to not exceed the noise level based on vehicle noise characteristics and an ambient noise level; determine whether the determined motor-specific fan RPM and motor-specific fan pitch will allow the vehicle to operate within its safety envelope; and cause a motor-specific fan RPM command and a motor-specific fan pitch command to be sent to the lifter motor controller to cause the vehicle lifter motors to operate at the determined motor-specific fan RPM and motor-specific fan pitch when it is determined that the determined motor-specific fan RPM and motor-specific fan pitch will allow the VTOL vehicle to operate within its safety envelope. 1. A noise reduction system in a vertical takeoff and landing (VTOL) vehicle comprising a plurality of vehicle lifter motors and a lifter motor controller for the vehicle lifter motors , the noise reduction system comprising a controller configured to:identify a VTOL noise level at which the VTOL vehicle can operate at a specific geographical location to not cause a total noise level at the geographical location to exceed a predetermined threshold noise level;dynamically determine, for the plurality of vehicle lifter motors, a motor-specific fan RPM and a motor-specific fan pitch that will allow the VTOL vehicle to achieve sufficient lift and not exceed the identified VTOL noise level at the specific geographical location based on VTOL vehicle noise characteristics for the VTOL vehicle and an ambient noise level at the geographical location;determine whether the determined motor-specific fan RPM and motor-specific fan pitch for the plurality of vehicle lifter motors will allow the VTOL vehicle to operate within its safety envelope; ...

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Номер записи: 77

12-01-2017 дата публикации

MECHANICAL COMPONENT FOR THERMAL TURBO MACHINERY

Номер: US20170009601A1

Автор: Mueller Wolfgang, Szwedowicz Jaroslaw Leszek

Принадлежит: ANSALDO ENERGIA SWITZERLAND AG

A mechanical component for thermal turbo machinery, such as a steam or gas turbine, includes a base part and at least one additional device being mechanically coupled to the base part in order to influence the vibration characteristic of the base part during operation of the turbo machine. High-Cycle Fatigue at part-load can be reduced by enabling the mechanical coupling between the base part and the at least one additional device to change with the temperature of the at least one additional device. 1. Mechanical component for thermal turbo machinery , comprising a base part , and at least one additional device being mechanically coupled to said part in order to influence a vibration characteristic of said part during operation of the turbo machine , wherein a mechanical coupling between said part and said at least one additional device changes with a temperature of said at least one additional device.2. Component as claimed in claim 1 , wherein said at least one additional device is a device claim 1 , which changes with temperature its form and position relative to said base part in order to establish an additional mechanical contact between said part and said at least one additional device within a predetermined temperature range.3. Component as claimed in claim 2 , wherein said at least one additional device is a bi-metallic device.4. Component as claimed in claim 2 , wherein said at least one additional device is a shape-memory-alloy device.5. Component as claimed in claim 2 , wherein said additional mechanical contact is a stiffening contact claim 2 , which mechanically stiffens said part.6. Component as claimed in claim 2 , wherein said additional mechanical contact is a friction contact claim 2 , which dampens vibrations in said part.7. Component as claimed in claim 2 , wherein said at least one additional device has the form of a longitudinal beam or curved plate claim 2 , which is fixedly connected at both ends to said part claim 2 , such that it ...

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Номер записи: 78

12-01-2017 дата публикации

GAS TURBINE ENGINE AFT BEARING ARRANGEMENT

Номер: US20170009655A1

Автор: Savela Gregory M.

Принадлежит:

An example gas turbine engine includes a turbine and first and second spools coaxial with one another. The first spool is arranged within the second spool and extends between forward and aft ends. The aft end extends axially beyond the second spool and supports the turbine. A housing is arranged downstream from the turbine. First and second bearings are mounted to the aft end of the first spool and supported by the housing portion. 1. A bearing hub for a gas turbine engine comprising:first and second hub walls integrally formed with one another to provide a unitary structure;a radial to axial translation flange arm extending outward from an apex of the unitary structure;a translation flange extending outward from said translation flange arm;a spring arm connected to the apex for connecting the bearing hub to a canted annular flange, the spring arm including a plurality of angled flex points.2. The bearing hub of claim 1 , wherein the translation flange arm extends axially aftward from said apex of said unitary structure.3. The bearing hub of claim 1 , wherein the plurality of angled flex points comprises at least a first flex point claim 1 , a second flex point claim 1 , and a third flex point claim 1 , and wherein a stiffness of each of said first flex point claim 1 , said second flex point and said third flex point is configured to control an amount of radial vibrations translated to axial vibrations by said bearing hub.4. The bearing hub of claim 1 , wherein the first and second hub walls are inclined radially inward from an annular apex claim 1 , and a first and second bearing are respectively supported by the first and second walls opposite the apex.5. The bearing hub of claim 4 , wherein a focal node of radial vibrations of the bearing hub is the first bearing.6. The bearing hub of claim 1 , wherein said spring arm is rigidly connected to said apex.7. A gas turbine engine comprising:a fan;a compressor section fluidly connected to the fan, the compressor ...

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Номер записи: 79

12-01-2017 дата публикации

Fan rotor for a turbo machine such as a multiple flow turbojet engine driven by a reduction gear

Номер: US20170009656A1

Автор: Christian Vessot, Matthieu Leyko

Принадлежит: Safran Aircraft Engines SAS

A forward fan rotor is disclosed with a hub of axis of rotation (X) and a cone mounted on the hub of the fan. The cone comprises an air bleed orifice which opens into an air duct of which a forward end portion passes through the fan rotor, said forward end portion comprising mechanical air entrainment means. The air bleed orifice has an annular shape and in that the cone is divided by said orifice into a front vertex portion and a rear frustoconical portion. A turbomachine forward axial spool equipped with such a fan rotor is also disclosed.

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Номер записи: 80

12-01-2017 дата публикации

CENTRIFUGAL COMPRESSOR, TURBOCHARGER PROVIDED WITH THE CENTRIFUGAL COMPRESSOR, AND METHOD FOR PRODUCING THE CENTRIFUGAL COMPRESSOR

Номер: US20170009780A1

Автор: HAYASHI Noriyuki, Yano Seigi

Принадлежит: MITSUBISHI HEAVY INDUSTRIES, LTD.

An object is to provide a centrifugal compressor in which an axial force applied from a shrink-fit impeller to a sleeve section is ensured even if the sleeve section is separated from a clamp surface in an axial direction of an attachment hole, as well as a turbocharger provided with the centrifugal compressor and a method of producing the centrifugal compressor. An inner peripheral surface of an attachment hole formed on a hub includes a clamp surface and a diameter-widening surface, and an outer peripheral surface of a sleeve section includes a diameter-reducing surface. The diameter-widening surface and the diameter-reducing surface respectively include an impeller-side contacting portion and a sleeve-side contacting portion which contact each other, and a relationship represented by an expression θs Подробнее

Номер записи: 81

12-01-2017 дата публикации

FLOATING BUSH BEARING DEVICE AND TURBOCHARGER PROVIDED WITH THE BEARING DEVICE

Номер: US20170009810A1

Автор: FUTAE Takaya, HAYASHI Noriyuki, NAMBU Takashi, NISHIOKA Tadasuke, Suzuki Hiroshi

Принадлежит: MITSUBISHI HEAVY INDUSTRIES, LTD.

An object is to provide a floating bush bearing device including a circumferential groove over the entire circumference of an outer peripheral surface of a floating bush while ensuring that a pressing force is applied by lubricant oil to the floating bush to reduce oscillation, as well as a turbocharger provided with the bearing device. A floating bush bearing device includes: a rotary shaft disposed rotatably inside a bearing hole of a casing; a floating bush surrounding the rotary shaft; an oil-feed hole of lubricant oil having an opening on an inner peripheral surface of the bearing hole; a plurality of communication holes formed on the floating bush, each extending between an inner peripheral surface and an outer peripheral surface of the floating bush, and disposed at intervals in a circumferential direction of the floating bush; and a circumferential groove formed on the outer peripheral surface of the floating bush or the inner peripheral surface of the bearing hole and extending over an entire circumference of the outer peripheral surface of the floating bush or the inner peripheral surface of the bearing hole, the circumferential groove passing through openings of the plurality of communication holes or facing the openings of the plurality of communication holes. The circumferential groove has a cross-sectional area which varies in accordance with a circumferential position. 1. A floating bush bearing device , comprising:a casing including a bearing hole;a rotary shaft disposed rotatably inside the bearing hole;a floating bush disposed rotatably inside the bearing hole and surrounding the rotary shaft;an oil-feed hole of lubricant oil, the oil-feed hole having an opening on an inner peripheral surface of the bearing hole;a plurality of communication holes formed on the floating bush, each of the communication holes extending between an inner peripheral surface and an outer peripheral surface of the floating bush, and disposed at intervals in a ...

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Номер записи: 82

27-01-2022 дата публикации

OIL STANDPIPE ASSEMBLY FOR SERVICING AND OIL LEVEL MAINTENANCE OF A STARTER IN A GAS TURBINE ENGINE

Номер: US20220025781A1

Автор: Jakab Thomas, Pugliese Jordan

Принадлежит:

A standpipe assembly configured to connect to a starter that includes a housing is described herein. The standpipe assembly includes: a standpipe having first and second openings at opposite ends of a hollow passageway, wherein the standpipe, upon being connected to the starter, is oriented within the housing so that the standpipe drains oil through the standpipe when oil in the housing reaches an overfill level, and is oriented in parallel to internal oil flow in the starter to inhibit interference with the internal oil flow during operation of the starter; and an attachment portion, wherein the attachment portion is structured to attach the standpipe assembly to the housing of the starter and prevent movement of the standpipe assembly with respect to the housing of the starter. 1. A standpipe assembly configured to connect to a starter that includes a housing , the standpipe assembly comprising:a standpipe having first and second openings at opposite ends of a hollow passageway, wherein the standpipe, upon being connected to the starter, is oriented within the housing so that the standpipe drains oil through the standpipe when oil in the housing reaches an overfill level, and is oriented in parallel to internal oil flow in the starter to inhibit interference with the internal oil flow during operation of the starter; andan attachment portion, wherein the attachment portion is structured to attach the standpipe assembly to the housing of the starter and prevent movement of the standpipe assembly with respect to the housing of the starter.2. The standpipe assembly of claim 1 , wherein the attachment portion attaches the standpipe assembly to the housing using an interference fit.3. The standpipe assembly of claim 1 , wherein the attachment portion comprises one or more brackets to attach the standpipe assembly to the housing with one or more fasteners.4. The standpipe assembly of claim 1 , wherein the first opening of the standpipe claim 1 , upon the standpipe being ...

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Номер записи: 83

12-01-2017 дата публикации

Jet engine, flying object, and method of operating a jet engine

Номер: US20170009992A1

Автор: Shojiro Furuya, Yoshihiko Ueno

Принадлежит: Mitsubishi Heavy Industries Ltd

A jet engine includes an inlet and a combustor. the inlet takes in air. The combustor combusts fuel with the air. The combustor ( 12 ) includes an injector ( 20 ) having a plurality of openings ( 31 a, 31 b ) from which the fuel is injected. The plurality of openings ( 31 a, 31 b ) are arranged in a direction perpendicular to a direction of a flow path of the air in the combustor ( 12 ). The plurality of openings ( 31 a, 31 b ) include two types of openings different in area.

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Номер записи: 84

27-01-2022 дата публикации

NOISE REDUCTION DEVICE FOR OUTLET SIDE OF FAN AND HEAT EXCHANGE SYSTEM INCLUDING THE SAME

Номер: US20220025906A1

Автор: Cheng Xin, Ji JunJie, Li Haijun, Shi Yuling

Принадлежит:

The utility model relates to a noise reduction device for an outlet side of a fan, and a heat exchange system including the noise reduction device. The noise reduction device includes: a connecting portion configured to be connected with at least a part of the air duct cover, and form an accommodation space communicating with an airflow on the outlet side via at least one of the through holes; and at least one first chamber and/or at least one second chamber, the first chamber being located in the accommodation space and filled with a sound-absorbing material, and the second chamber being located in the accommodation space and configured as a resonant noise-reduction cavity. The utility model is easy to manufacture, install and maintain, the noise reduction effect is obvious, and therefore the utility model has significant practicability. 1. A noise reduction device for an outlet side of a fan , the outlet side being provided with an air duct cover with one or more through holes , wherein the noise reduction device comprises:a connecting portion configured to be connected with at least a part of the air duct cover, and form an accommodation space communicating with an airflow on the outlet side via at least one of the through holes; andat least one first chamber and/or at least one second chamber, the first chamber being located in the accommodation space and filled with a sound-absorbing material, and the second chamber being located in the accommodation space and configured as a resonant noise-reduction cavity.2. The noise reduction device according to claim 1 , wherein the first chamber is configured to reduce noises in a preset frequency spectrum range claim 1 , and noise peaks in the preset frequency spectrum range are all less than a preset value.3. The noise reduction device according to claim 1 , wherein the resonant noise-reduction cavity is configured to reduce preset single-frequency noises of the fan claim 1 , and when a ratio between the energy of the ...

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Номер записи: 85

27-01-2022 дата публикации

DOUBLE JOURNAL BEARING IMPELLER FOR ACTIVE DE-AERATOR

Номер: US20220026021A1

Автор: Martin Bruno

Принадлежит:

An active de-aerator for an aircraft engine is provided, with a housing having an air-oil inlet, an oil outlet and an air outlet. An impeller is received within and rotatable relative to the housing about a central axis. The active de-aerator has a first journal bearing on a first side of the impeller for rotatably supporting the impeller relative to the housing and a second journal bearing on a second side of the impeller for rotatably supporting the impeller relative to the housing, the second side being opposite the first side. 1. An active de-aerator for an aircraft engine , comprising:a housing having an air-oil inlet, an oil outlet and an air outlet;an impeller received within and rotatable relative to the housing about a central axis;a first journal bearing on a first side of the impeller for rotatably supporting the impeller relative to the housing; anda second journal bearing on a second side of the impeller for rotatably supporting the impeller relative to the housing, the second side being opposite the first side.2. The active de-aerator as defined in claim 1 , wherein the active de-aerator is adapted to be driven by an oil pump claim 1 , the impeller defining a shaft connecting portion for connecting the impeller to an end of a pump shaft.3. The active de-aerator as defined in claim 1 , wherein the impeller defines a shaft connecting portion for connecting the impeller to an end of a pump shaft claim 1 , the first journal bearing defined by a portion of the housing and the shaft connecting portion of the impeller.4. The active de-aerator as defined in claim 1 , wherein the impeller defines a flange claim 1 , the second journal bearing defined by a portion of the housing and the flange of the impeller.5. The active de-aerator as defined in claim 1 , wherein the impeller has blades circumferentially distributed about the central axis claim 1 , the first and second sides of the impeller being on opposite sides of the blades along the central axis claim 1 , ...

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Номер записи: 86

27-01-2022 дата публикации

Fuel distribution manifold

Номер: US20220026069A1

Автор: Diana Alonso, Elizabeth Grace Hildebrand, Rajendra Paniharam

Принадлежит: General Electric Co

Fuel distribution manifolds and combustors are provided. A fuel distribution manifold includes a main body and a fuel circuit that is defined within the main body. The fuel circuit includes an inlet section extending generally axially from an inlet to a first branch section and a second branch section. The first branch section and the second branch section diverge circumferentially away from each other as they extend axially from the inlet section to a respective first outlet and a respective second outlet.

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Номер записи: 87

14-01-2016 дата публикации

METHOD FOR DETUNING A ROTOR-BLADE CASCADE

Номер: US20160010461A1

Автор: de Lazzer Armin, Gronsfelder Thomas, Walkenhorst Jan

Принадлежит: SIEMENS AKTIENGESELLSCHAFT

A method for detuning a rotor-blade cascade of a turbomachine having a plurality of rotor blades includes: a) establishing at least one target natural frequency for at least one vibration mode; b) setting up a value table having discrete mass values and radial centre-of-gravity positions, and determining respective natural frequency; c) measuring the mass and radial centre-of-gravity position of one of the rotor blades; d) determining an actual natural frequency by interpolating the measured mass and radial centre-of-gravity position in the value table; e) if actual natural frequency is outside a tolerance around target natural frequency, selecting a value pair that at least approximates target natural frequency, and removing material from the rotor blade in such a way that mass and radial centre-of-gravity position correspond to the value pair; f) repeating steps c) to e) until actual natural frequency is within the tolerance around target natural frequency. 1. A method for detuning a rotor-blade cascade , comprising a multiplicity of rotor blades , of a turbomachine , the method comprising:{'sub': 'F,S', 'a) establishing for each of the rotor blades of the rotor-blade cascade at least one setpoint natural frequency νwhich the rotor blade has for at least one predetermined oscillation mode during normal operation of the turbomachine under the effect of centrifugal force, such that the oscillation load of the rotor-blade cascade under the centrifugal force lies below a tolerance limit;'}{'sub': F', 'S', 'S', 'F', 'S, 'b) compiling a value table ν(m, r) with selected value pairs of discrete mass values m and radial center-of-mass positions r, which result from variations of the nominal geometry of the rotor blade, and determining the respective natural frequency νof the predetermined oscillation mode under the centrifugal force for each selected value pair m and r;'}{'sub': I', 'S,I, 'c) measuring the mass mand the radial center-of-mass position rof one of the rotor ...

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Номер записи: 88

14-01-2016 дата публикации

GAS TURBINE SPINDLE BOLT STRUCTURE WITH REDUCED FRETTING FATIGUE

Номер: US20160010481A1

Автор: Abdo Zafir A. M., Gurao Manish S., Light Kevin M., OZBAYSAL KAZIM

Принадлежит:

A spindle bolt structure is provided in a gas turbine engine, and includes a pilot region located within a bolt hole extending through a seal disk. The pilot region includes a circumferential pilot ridge located adjacent to a downstream axial face of the seal disk and a circumferential trough portion located between a bolt shoulder and the pilot ridge. The trough portion defines a trough diameter that is less than a diameter of the bolt shoulder and less than a diameter of the pilot ridge. The bolt shoulder and pilot ridge are formed with an applied compressive residual stress and are positioned for engagement with the seal disk. 1. In a gas turbine engine , a rotor including a plurality of turbine disks for supporting rows of blades , a torque tube located on a compressor side of the turbine disks , and a seal disk located between the torque tube and a first stage turbine disk , a spindle bolt structure comprising:a spindle bolt extending through the turbine disks and disposed offset from a rotational axis of the turbine disks;the seal disk including an upstream axial face and an opposing downstream axial face, and a bolt hole extending between the upstream and downstream axial faces;the spindle bolt extending through the bolt hole and including a bolt shoulder engaged on the seal disk within the bolt hole;a pilot region formed on the spindle bolt and located within the bolt hole for effecting a reduction in fretting fatigue of the spindle bolt, the pilot region including a circumferential pilot ridge located in the bolt hole adjacent to the downstream axial face of the seal disk and a circumferential trough portion located between the bolt shoulder and the pilot ridge, the trough portion defining a trough diameter that is less than a diameter of the bolt shoulder and that is less than a diameter of the pilot ridge.2. The spindle bolt structure of claim 1 , wherein the diameter of the pilot ridge is less than the diameter of the bolt shoulder.3. The spindle bolt ...

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Номер записи: 89

14-01-2016 дата публикации

LUBRICATION SYSTEM FOR A GAS TURBINE ENGINE

Номер: US20160010499A1

Автор: Ajango Vaino E., Benson Michael L., Gatto Stephen M.

Принадлежит:

A gas turbine engine includes a lubrication system for distributing lubricant throughout the engine. The lubrication system includes a breather assembly that receives air and lubricant from various other components of the lubrication system. The breather assembly includes a baffle that redirects air and lubricant received by the breather assembly. 1. A lubrication system for a gas turbine engine , the lubrication system comprisingan oil tank,a sump housing a bearing, anda breather assembly coupled to the oil tank, the sump, and atmosphere, wherein the breather assembly is configured to redirect oil droplets from the oil tank into the sump and vent carrier air carrying the oil droplets from the oil tank to atmosphere while also venting pressurized air from the sump to the atmosphere without allowing the pressurized air to increase static pressure in the oil tank.2. The lubrication system of claim 1 , wherein the breather assembly includes a breather housing that defines a breather cavity claim 1 , a tank port arranged to transport oil droplets and carrier air from the oil tank into the breather cavity claim 1 , a vent port arranged to transport air from the breather cavity to the atmosphere claim 1 , a sump air-bleed port arranged to transport pressurized air from the sump into the breather cavity claim 1 , and a sump oil-return port arranged to transport oil from the breather cavity to the sump.3. The lubrication system of claim 2 , wherein the breather assembly includes a baffle arranged in the breather cavity between the tank port and the sump air-bleed port to block pressurized claim 2 , relatively high-velocity air from the sump air-bleed port from directly impinging into the tank port.4. The lubrication system of claim 3 , wherein the sump air-bleed port is arranged to direct pressurized air from the sump toward the tank port and to create a venture to reduce total cavity pressure.5. The lubrication system of claim 3 , wherein the baffle includes a plate formed ...

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Номер записи: 90

14-01-2016 дата публикации

Supercharging device

Номер: US20160010500A1

Автор: Alexander Baeuerle, Christoph Butscher, Joerg Eckert, Ruediger Kleinschmidt, Thomas Striedelmeyer, Ulrich Siol

Принадлежит: Bosch Mahle Turbo Systems GmbH and Co KG

A supercharging device may include a rotor mounted in a bearing housing via at least one bearing bush. The bearing bush may include an oil bore. An oil film may be disposed at least one of between the rotor and the bearing bush and between the bearing bush and the bearing housing. The bearing housing may include at least one oil feed duct for lubricating the bearing bush. At least one of the oil feed duct and the oil bore may be configured with respect to the bearing bush such that the bearing bush during the operation of the supercharging device is positively accelerated in a direction of rotation of the rotor via an oil jet communicated from the oil feed duct.

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Номер записи: 91

14-01-2016 дата публикации

INTERMEDIATE CASING FOR A TURBOFAN ENGINE

Номер: US20160010501A1

Автор: Behaghel Laurent Donatien, Cameau Jerome Etienne Robert, Contal Kévin Michel Robert, Juigne Sébastien Nicolas

Принадлежит: SNECMA

An intermediate casing comprising an inner annular hub, an outer annular barrel and an annular part for separating flows situated between the hub and the outer barrel. A primary stream is delimited between the hub and the separation part. A secondary stream is delimited between the separation part and the outer barrel. At least one hollow arm extends radially from the hub to the outer barrel, passing through the primary and secondary streams. A transmission shaft extends radially in the hollow arm. The hollow arm comprises a hydraulic-fluid outlet situated downstream of the transmission shaft. The arm further comprises a bypass channel or pocket able to bypass the transmission shaft. 1. An intermediate casing for a turbofan comprising a radially internal annular hub , a radially external annular barrel and an annular flow-separation part situated radially between the hub and the outer barrel , a primary stream for flow of a primary flow being delimited between the hub and the separation part , a secondary stream allowing flow of a secondary flow being delimited between the separation part and the outer barrel , at least one hollow arm extending radially from the hub to the outer barrel passing through the primary and secondary streams , a transmission shaft extending radially in said hollow arm , wherein the hollow arm comprises a hydraulic-fluid outlet situated downstream of the transmission shaft with respect to the direction of circulation of the primary flow or secondary flow , said arm further comprising a bypass channel or pocket able to bypass the transmission shaft and extending from upstream to downstream of said transmission shaft.2. An intermediate casing according to claim 1 , wherein the arm comprises first and second walls externally delimiting the arm claim 1 , extending radially and joining at an upstream edge claim 1 , said bypass channel or pocket being formed by a hollow region produced in the first wall and/or the second wall of the arm and ...

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Номер записи: 92

14-01-2016 дата публикации

OIL BAFFLES IN CARRIER FOR A FAN DRIVE GEAR SYSTEM

Номер: US20160010549A1

Автор: McCune Michael E., Muldoon Marc

Принадлежит:

A gearbox assembly for a gas turbofan engine includes a sun gear rotatable about an axis and a plurality of intermediate gears driven by the sun gear. A baffle disposed between at least two of the plurality of intermediate gears includes a first gap distance within a first gap portion and a second gap distance within a second gap portion. The first gap portion is disposed between the baffle and one of the intermediate gears away from the meshed interface with the sun gear and the second gap portion is disposed near the interface with the sun gear. The first gap distance within the first gap portion is different than the second gap distance within the second gap portion to define a desired lubricant flow path. 1. A gearbox assembly for a gas turbofan engine comprising:a sun gear rotatable about an axis;a plurality of intermediate gears driven by the sun gear;a baffle disposed between at least two of the plurality of intermediate gears, wherein the baffle is spaced a first gap distance from the at least two intermediate gears within a first gap portion and a second gap distance different than the first gap distance from one of the at least two intermediate gears within a second gap portion including an interface with the sun gear.2. The gearbox assembly as recited in claim 1 , wherein the second gap distance is larger than the first gap distance.3. The gearbox assembly as recited in claim 1 , wherein the second gap distance is between about 1.5 and about 2.5 greater than the first gap distance.4. The gearbox assembly as recited in claim 1 , wherein the sun gear includes a cavity and the baffle includes a wedge extending into the cavity for circulating lubricant out of the cavity.5. The gearbox assembly as recited in claim 1 , including a carrier supporting the intermediate gears relative to the sun gear and a ring gear circumscribing the intermediate gears claim 1 , wherein a ring gear baffle is supported on the carrier.6. The gearbox assembly as recited in claim 5 , ...

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Номер записи: 93

14-01-2016 дата публикации

MANIFOLD FOR GAS TURBINE

Номер: US20160010550A1

Автор: Baker Stephanie, Otto John

Принадлежит: UNITED TECHNOLOGIES CORPORATION

In various embodiments, a manifold assembly () for conducting one or more fluids to a gear assembly () in a gas turbine engine () is provided. The manifold assembly () may comprise a first plate () and a second plate () that rotatably couple together. The manifold assembly () may be retained and/or held together by a channel () and engagement member () arrangement. 1. A manifold assembly , comprising:a first manifold comprising an engagement member;a second manifold having a groove defined therein, wherein the engagement member is installable in the groove; andthe manifold assembly configured to conduct a fluid to a gear assembly through the first manifold and the second manifold.2. The manifold assembly of claim 1 , further comprising an anti-rotation element.3. The manifold assembly of claim 2 , wherein the anti-rotation element is at least one of a fastener claim 2 , a pin claim 2 , an adhesive claim 2 , a tensioning device and a detent assembly.4. The manifold assembly of claim 1 , wherein the groove comprises a receiving portion and a retention portion.5. The manifold assembly of claim 1 , wherein the engagement member is a tongue.6. The manifold assembly of claim 1 , wherein the first manifold is rotatably coupled to the second manifold.7. A turbine engine claim 1 , comprising;a gear assembly; a first portion having a first groove and a second groove, wherein the first groove and the second groove are defined along a diameter of the first portion; and', 'a second portion having a first engagement member installable in the first groove and a second engagement member installable in the second groove., 'a manifold operatively coupled to and in fluid communication with the gear assembly, the manifold comprising8. The turbine engine of claim 7 , wherein the first groove further comprises a first groove portion that is defined radially outward from a centerline of the turbine engine.9. The turbine engine of claim 7 , wherein the engagement member comprises a shaft ...

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Номер записи: 94

14-01-2016 дата публикации

GAS TURBINE SILENCER, AND GAS TURBINE PROVIDED WITH SAME

Номер: US20160010557A1

Автор: Matsuyama Keisuke, TODA Tomonori

Принадлежит: MITSUBISHI HEAVY INDUSTRIES, LTD.

Provided are a gas turbine silencer for avoiding the formation of gaps between an upstream silencer panel and a downstream silencer panel and suppressing the occurrence of secondary noise, and a gas turbine provided with this silencer. A silencer panel () has a structure that can be divided into an upstream silencer panel () and a downstream silencer panel () in an airflow direction, a stepped part () shorter in length along the alignment direction of the silencer panel () is formed in an opening-side portion of the downstream silencer panel (), and the upstream silencer panel () and downstream silencer panel () are linked by the stepped part () fitting into substantially the entire opening () in the upstream silencer panel (). 1. A gas turbine silencer installed between an air intake port and a compressor of a gas turbine , the gas turbine silencer comprising:a plurality of plate-shaped divided silencer panels aligned at predetermined intervals in a direction orthogonal to a flow direction of a fluid from the air intake port toward the compressor;the divided silencer panels wherein a surface thereof with a greatest plate area is arranged in orientation along a flow of the fluid, the divided silencer panels comprising an upstream silencer panel arranged on an upstream side in the flow direction of the fluid, and a downstream silencer panel arranged on an downstream side of the upstream silencer panel and linked with the upstream silencer panel;one silencer panel out of the upstream silencer panel and the downstream silencer panel being formed with an opening opened on a side facing the other silencer panel;the other silencer panel being formed with a fitting section fitting into the opening; andthe upstream silencer panel and the downstream silencer panel being linked by the fitting section of the other silencer panel fitting into the opening of the one silencer panel.2. The gas turbine silencer according to claim 1 , wherein the divided silencer panels are ...

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Номер записи: 95

14-01-2016 дата публикации

OIL LOSS PROTECTION FOR A FAN DRIVE GEAR SYSTEM

Номер: US20160010563A1

Автор: Sheridan William G.

Принадлежит:

A fan drive gear system includes at least one intermediate gear that includes an axial gear passage for receiving and conveying a fluid suitable for cooling and/or lubricating. At least a first axial end of the intermediate gear includes a first fluid storage trap for capturing fluid entering and/or exiting the gear passage and storing the fluid therein during powered operation of the fan drive gear system. The fluid is capable of being passively supplied to the intermediate gear passage during an interrupted power event. 1. A fan drive gear system comprising:at least one intermediate gear that includes an axial gear passage for receiving and conveying a fluid suitable for cooling and/or lubricating;at least a first axial end of said intermediate gear includes a first fluid storage trap for capturing fluid entering and/or exiting the gear passage and storing the fluid therein during powered operation of the fan drive gear system; andwhereby the fluid is capable of being passively supplied to the intermediate gear passage during an interrupted power event.2. The fan drive gear system of claim 1 , further comprising;a sun gear interfaced with said intermediate gear;a ring gear interfaced with said intermediate gear; anda carrier body supporting the intermediate gear.3. The fan drive gear system of claim 2 , wherein the at least one fluid trap comprises a radially outward base portion relative to an axis defined by said carrier body and at least one radially inward base portion relative to said axis defined by said carrier body.4. The fan drive gear system of claim 3 , wherein said radially outward base portion is defined on a first axial end by a radially aligned wall segment of said carrier body relative to the axis defined by the carrier body and said radially outward base portion is defined on a second axial end by a radially aligned wall of said trap.5. The fan drive gear system of claim 4 , wherein said radially aligned wall of said fluid storage trap extends ...

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Номер записи: 96

14-01-2016 дата публикации

Fan Blade Lubrication

Номер: US20160010795A1

Автор: Daniel F. Palmer

Принадлежит: United Technologies Corp

In accordance with one aspect of the disclosure, a lubricant application device for providing lubricant to an airfoil of a gas turbine engine is disclosed. The device may include a container for holding a lubricant, a fluid outlet in fluid communication with the container, a discharge mechanism controlling a flow of the lubricant from the container to the fluid outlet, and an extension tube in fluid communication with the fluid outlet. When the discharge mechanism is activated the lubricant may flow out of the container and through the extension tube. The extension tube may be dimensioned to fit between an airfoil root and rotor cavity of a gas turbine engine.

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Номер записи: 97

11-01-2018 дата публикации

Low energy wake stage

Номер: US20180010459A1

Автор: Charles H. ROCHE, Charles P. Gendrich

Принадлежит: United Technologies Corp

The leading edge, the trailing edge, or both may be axially offset for a portion of the airfoils in a disk. By offsetting the airfoils, the downstream wake energy to the next stage of airfoils may be decreased. By staggering airfoils which are offset with airfoils that are not offset, the wake shapes from the airfoils may be out of phase and will not excite the downstream airfoils as much as conventional systems. This may decrease vibration and associated vibratory stresses in the system.

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Номер записи: 98

09-01-2020 дата публикации

Airfoil for a rotary machine including a propellor assembly

Номер: US20200010174A1

Автор: Kishore Ramakrishnan, Trevor Howard Wood

Принадлежит: General Electric Co

In some embodiments, an airfoil comprises a proximal end; a distal end opposite said proximal end; a distal portion extending adjacent said distal end; an edge extending between said proximal end and said distal end; and a surface extending between said proximal end and said distal end, said edge and said surface defining a sweep and a cahedral through said distal portion, wherein the distal portion extends over an acoustically active portion of the airfoil.

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Номер записи: 99

11-01-2018 дата публикации

TUNED MASS DAMPER FOR TUBES

Номер: US20180010482A1

Автор: Peters Donald W., Ramlogan Amarnath, RappoIi Nico M., Sadil Andreas

Принадлежит:

A tuned mass damper for reducing vibration on a component includes a shaft connector member configured to be coupled to the component and a cable termination member. The tuned mass damper also includes at least one cable coupled to the shaft connector member and to the cable termination member such that vibration of the component is transferred to the at least one cable via the shaft connector member and increased or decreased by the at least one cable. 1. A tuned mass damper for increasing or decreasing vibration on a component , comprising:a shaft connector member configured to be coupled to the component;a cable termination member; andat least one cable coupled to the shaft connector member and to the cable termination member such that vibration of the component is transferred to the at least one cable via the shaft connector member and increased or decreased by the at least one cable.2. The tuned mass damper of claim 1 , wherein the component is a shaft and the cable termination member is configured to be annularly positioned about the shaft claim 1 , wherein the cable termination member and the shaft define a gap.3. The tuned mass damper of claim 1 , wherein the tuned mass damper is tuned to have a damper frequency that reduces vibrations at a resonant frequency of the component.4. The tuned mass damper of claim 1 , wherein the tuned mass damper is tuned by adjusting at least one of a length of the at least one cable claim 1 , a mass of the cable termination member claim 1 , a total number of cables claim 1 , a total number of strands of the at least one cable claim 1 , a diameter of the at least one cable claim 1 , or a material of the at least one cable.5. The tuned mass damper of claim 1 , wherein the component is a shaft of an augmentor spray bar of a gas turbine engine.6. The tuned mass damper of claim 1 , wherein each of the shaft connector member claim 1 , the at least one cable claim 1 , and the cable termination member include at least one of a nickel- ...

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Номер записи: 100