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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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Форма поиска

Поддерживает ввод нескольких поисковых фраз (по одной на строку). При поиске обеспечивает поддержку морфологии русского и английского языка
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Применить Всего найдено 3937. Отображено 100.
16-02-2012 дата публикации

Engine and pod assembly for an aircraft, equipped with an anti-icing device

Номер: US20120036826A1
Принадлежит: Sagem Defense Securite SA

An engine and pod assembly for an aircraft includes a pod receiving an engine having an air intake. A rotating nose cone extends on the nose cone, as well as a device for limiting the formation of ice. The device includes means for creating a circumferential heterogeneity of ice on the nose cone.

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26-04-2012 дата публикации

Rotary machine having non-uniform blade and vane spacing

Номер: US20120099961A1
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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07-06-2012 дата публикации

Fluid impingement arrangement

Номер: US20120137650A1
Принадлежит: Rolls Royce PLC

A fluid impingement arrangement comprising a supply manifold and at least one nozzle exit coupled to the supply manifold. The nozzle exit is arranged as a Coanda surface having a restriction and has at least one static pressure tapping that cross-connects two regions of the restriction to induce passive oscillation in a fluid jet passing through the nozzle exit.

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09-08-2012 дата публикации

Ring element and turbomachine having such a ring element

Номер: US20120198858A1
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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27-09-2012 дата публикации

Method and apparatus for protecting aircraft engines against icing

Номер: US20120240594A1
Автор: Pavel SHAMARA
Принадлежит: Cox and Co Inc

An aircraft engine generates engine power by burning hydrocarbon fuel such as Jet-A. A minute quantity of the fuel is burned in such a manner as to generate no engine power, and the heat generated by the burning fuel is used to protect a region of a surface of a component of an aircraft. In one application, burner assemblies are located inside the splitter of a turbofan engine and the heat generated is used to deice or anti-ice the splitter and the inlet guide vanes of the engine. In another application, burner assemblies are located in an engine nacelle to deice or anti-ice the leading edge of the nacelle.

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31-01-2013 дата публикации

Compressor rotor

Номер: US20130028750A1
Принадлежит: Alstom Technology AG

A compressor rotor is provided having a rotor blade groove thereon and also includes a device for cooling the rotor in the region of a compressor rotor exit. Efficient cooling is achieved by the compressor rotor, in a compressor rotor exit region, having a ring which is pushed concentrically, and at a distance, forming a gap, over a rotor disk of the rotor, and is fastened on the disk, by the rotor blades, in the compressor rotor exit region, being inserted into corresponding grooves on the ring and being retained there, by first means for directing an axial flow of cooling medium from the compressor rotor exit through the ring, and by second means for deflecting the cooling medium which issues from the ring such that the cooling medium flows back in the axial direction through the gap between the ring and the rotor disk, encompassed by the ring.

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21-03-2013 дата публикации

Vibration damping blade for fluid

Номер: US20130071251A1
Принадлежит: IHI Corp

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

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28-03-2013 дата публикации

TURBO MACHINE WITH A DEVICE FOR PREVENTING A SEGMENT OF NOZZLE GUIDE VANES ASSEMBLY FROM ROTATING IN A CASING; ROTATION-PROOFING PEG

Номер: US20130078086A1
Принадлежит: SNECMA

A device for preventing rotation of a segment of nozzle guide vanes assembly in a form of an annulus sector housed inside an annular casing of a turbo machine with interposing of a heat shield sheet between an internal wall of the casing and an external wall of the segment of nozzle guide vanes assembly, the device including a rotation-proofing peg fitted both into a notch formed in the segment of nozzle guide vanes assembly and in a housing formed in the casing, the heat shield sheet including a tab resting against the rotation-proofing peg. A surface portion radially between the tab and the internal wall of the casing forms an end stop in event of a possible radial movement of the heat shield sheet while the turbo machine is in operation. 19-. (canceled)10. A turbo machine including a device preventing rotation of a segment of a distributor in a form of a ring sector housed inside a ring-shaped casing of the turbo machine , with interposing of a heat shield sheet between an inside wall of the casing and an outside wall of the distributor segment , comprising:a rotation-proofing peg mounted both in a notch provided on the distributor segment and a housing arranged on the casing, the heat shield sheet including a tab bearing against the rotation-proofing peg; anda surface portion arranged radially between the tab and the inside wall of the casing, forming an abutment in event of possible radial displacement of the heat shield sheet during operation of the turbo machine.11. The turbo machine as claimed in claim 10 , wherein the rotation-proofing peg includes a head housed in the notch of the distributor segment and a rod entering the housing arranged on the casing claim 10 , the surface portion being integral with the head of the rotation-proofing peg.12. The turbo machine as claimed in claim 11 , wherein the surface portion includes a shoulder machined in the head of the rotation-proofing peg.13. The turbo machine as claimed in claim 10 , wherein the surface portion ...

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06-06-2013 дата публикации

Alternate shroud width to provide mistuning on compressor stator clusters

Номер: US20130142640A1
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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13-06-2013 дата публикации

BLADE

Номер: US20130149108A1
Автор: WEBSTER John R.
Принадлежит: ROLLS-ROYCE PLC

A blade having a root portion and an aerofoil portion, wherein the aerofoil portion has a tip remote from the root portion, and a leading edge and a trailing edge, and wherein the tip of the aerofoil portion has a set-back portion extending from the leading edge or the trailing edge of the aerofoil portion part way towards the respective other edge and set back from the remainder of the tip of the aerofoil portion towards the root portion. 1. A blade comprising a root portion and an aerofoil portion , wherein the aerofoil portion has a tip remote from the root portion , and a leading and a trailing edge , and wherein the tip of the aerofoil portion has a set-back portion extending from the leading edge or the trailing edge of the aerofoil portion part way towards the respective other edge and set back from the remainder of the tip of the aerofoil portion towards the root portion.2. A blade as claimed in claim 1 , wherein the set-back portion in the tip is serrated.3. A blade as claimed in claim 2 , wherein the serrated set-back portion has shaped serration slots which extend not aligned with the circumferential direction of motion of the tip when the blade is rotating in use in a fan.4. A blade as claimed in claim 3 , wherein the serration slots are approximately perpendicular to the surface of the tip which is flow-washed when the blade is rotating in use in a fan.5. A blade as claimed in claim 3 , wherein the serration slots are 2 mm deep.6. A fan having a plurality of blades as claimed in and a fan casing around the tips of the blades claim 1 , wherein as a result of the set-back portions in the tips the tip clearance area between tips and fan casing is changed by at least 1% of fan area as compared with a case in which the set-back portions in the tips were omitted.7. (canceled)8. A blade as claimed in claim 4 , wherein the serration slots are 2 mm deep.9. A fan having a plurality of blades as claimed in and a fan casing around the tips of the blades claim 2 , ...

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13-06-2013 дата публикации

GAS TURBINE ENGINE WITH MULTIPLE COMPONENT EXHAUST DIFFUSER OPERATING IN CONJUNCTION WITH AN OUTER CASE AMBIENT EXTERNAL COOLING SYSTEM

Номер: US20130149121A1
Принадлежит:

An attachment system for attaching at least one exhaust diffuser downstream from a turbine assembly in a gas turbine engine is disclosed. The attachment system may include at least one attachment flange extending from a downstream edge and attached to a spring plate diffuser support structure and at least one attachment flange extending from side edges of the exhaust diffuser to couple sections of the exhaust diffuser together. The diffuser may also include a thermal barrier/cooling system for controlling a temperature of an outer case of the gas turbine engine. The thermal barrier/cooling system may form a flow path for an ambient air flow cooling. 1. A gas turbine engine comprising:at least one exhaust diffuser positioned downstream from a turbine assembly, extending circumferentially around a central longitudinal axis of the gas turbine engine and having an increasing cross-sectional area from an upstream edge to a downstream edge;wherein the downstream edge of the at least one exhaust diffuser includes at least one downstream attachment flange having at least one downstream attachment orifice extending therethrough.2. The gas turbine engine of claim 1 , wherein the at least one downstream attachment flange comprises a plurality of downstream attachment flanges extending generally radially outward from the at least one exhaust diffuser.3. The gas turbine engine of claim 2 , wherein adjacent downstream attachment flanges are separated by a void that is defined by a first outer surface extending between a first downstream attachment flange and the exhaust diffuser and having a radius claim 2 , a second outer surface extending between a second downstream attachment flange and the exhaust diffuser and having a radius claim 2 , and a third outer surface extending between first and second outer surfaces claim 2 , wherein the third outer surface has a radius larger than the radii of the first and second outer surfaces.4. The gas turbine engine of claim 3 , wherein the ...

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04-07-2013 дата публикации

ROTOR BEHAVIOUR DETERMINATION

Номер: US20130170947A1
Принадлежит: ROLLS-ROYCE PLC

A method of resolving vibration behaviour of a rotor assembly having a plurality of rotor blades. The method including generating a computational model of the rotor by discretization of the rotor geometry and assigning parameter values representative of a plurality of physical characteristics of the rotor. An artificial parameter feature is applied to each rotor blade such that the parameter values for the artificial parameter feature substantially depart from the physical characteristics of the rotor. A vibration response is calculated for the rotor or blade thereof for the applied artificial parameter features and compared to a predetermined vibration response so as to determine a value of the artificial parameter feature which results in a calculated vibration response that substantially matches the predetermined vibration response. 1. A method of resolving vibration behaviour of a rotor assembly having a plurality of components , the method comprising:obtaining a measured vibration response for the rotor;generating a computational model of the rotor by discretization of the rotor geometry and assigning parameter values representative of a plurality of physical characteristics of the rotor;applying a localised artificial parameter feature to each component of the rotor such that the parameter value for said artificial parameter for each component departs fromthe physical characteristics of said rotor;calculating a vibration response for the rotor or component thereof for said applied artificial parameter features;comparing the calculated vibration response against the measured vibration response so as to determine a value of the artificial parameter feature which results in a calculated vibration response that substantially matches the measured vibration response; and,modifying one or more components of the rotor in dependence on the determined artificial parameter values.2. A method according to claim 1 , wherein the artificial parameter feature comprises an ...

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11-07-2013 дата публикации

Impingement Cooling System for Use with Contoured Surfaces

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

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

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11-07-2013 дата публикации

BLADE ARRANGEMENT AND ASSOCIATED GAS TURBINE

Номер: US20130177427A1
Автор: Kayser Andreas
Принадлежит:

A blade arrangement with a rotor and a plurality of blades which are distributed in a ring along the circumference of the rotor is provided. Two immediately adjacent blades of the ring form a blade pair, between the blades of which a damping element is arranged, and wherein the respective damping element comes into contact with the two blades of the blade pair assigned to them during a rotation of the rotor about a rotor axis as a result of a centrifugal force which acts in the radial direction. In order to bring about frequency detuning of the oscillation properties of blades, as a result of which machining of the turbine blade becomes unnecessary, it is proposed that the blade ring has at least two blade pairs with different damping elements. 14-. (canceled)5. A blade arrangement , comprising:a rotor; anda plurality of blades which are distributed in a ring along the circumference of the rotor and comprise respectively in succession a blade root, a platform and a blade airfoil, wherein two immediately adjacent blades of the ring form a blade pair and are assigned at least one damping element,wherein each respective damping element comes into contact with the platforms of the two blades of the blade pair assigned to it during a rotation of the rotor about a rotor axis as a result of a centrifugal force acting in the radial direction,wherein, for adjusting the natural frequencies of the blades, the blade ring includes at least two blade pairs with different damping elements and each blade of the ring is assigned to two blade pairs and wherein two or more groups of blade pairs are provided, within each group the damping elements are in each case identical and the damping elements differ from group to group,wherein a majority of the blade pairs or each blade pair of the first group has an adjacent blade pair of the first group and an adjacent blade pair of the second group, orwherein a first group, a second group and a third group of blade pairs are provided wherein a ...

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18-07-2013 дата публикации

Air moving devices

Номер: US20130183141A1
Автор: Jianliang Tan

Disclosed herein is an air moving device which includes a housing and a motor. The air moving device may include at least one shock absorber which is configured to reduce vibrations and/or noise. The air moving device may also be configured to cancel out noises generated when operating the device.

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01-08-2013 дата публикации

Unknown

Номер: US20130195611A1
Автор: Retze Ulrich
Принадлежит: MTU AERO ENGINES GMBH

A method for vibration damping of at least one blade of a turbomachine, wherein initially at least one damping element is arranged on the blade such that it can move in the axial direction, employs a damping element having a larger permeability constant than the blade (μ>μ), and then a magnetic field acting in the radial direction is generated at least temporarily during rotation of a rotor hub of the turbomachine in order to adjust the mass of the damping element in real time. A damping device includes, for example, a ferromagnetic damping element as well as a magnetic field source, and a turbomachine. 11. A method for vibration damping of at least one blade () of a turbomachine comprising the steps:{'b': 18', '1', '1, 'sub': 'rD', 'arranging at least one damping element () at the blade () such that it can move in the axial direction, said damping element having a larger permeability constant (μ) than the blade (); and'}{'b': '26', 'generating a magnetic field () acting in the radial direction at least temporarily during rotation of a rotor hub of a turbomachine.'}226. The method according to claim 1 , wherein the magnetic field () is switched on and off during the rotation.3. The method according to claim 1 , wherein a magnetic field strength is varied as a function of the speed of rotation of the rotor hub.42618. The method according to claim 1 , wherein the magnetic field () acts radially inward on the damping element ().52618. The method according to claim 1 , wherein the magnetic field () acts radially outward on the damping element ().62618. The method according to claim 1 , wherein the magnetic field () rotates together with the damping element ().726. The method according to claim 1 , wherein the magnetic field () is fixed in position.8181. The method according to claim 1 , wherein the damping element () is arranged close to the blade mount when the blade () is braced in a blade mount subject to tolerances.9211812018. A damping device () for vibration ...

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01-08-2013 дата публикации

DRIVE SHAFT SYSTEM

Номер: US20130195615A1
Принадлежит: I.D.E. TECHNOLOGIES LTD.

A drive shaft system () comprising a shaft () connected to a motor via a coupling () and arranged to transmit a rotary movement from the motor to the rotor, at least two supports () arranged to position the shaft within a shaft housing (), each support comprising at least one bearing (), and at least one forcing element () arranged to force at least one of the supports against either an inner side of the shaft housing or an outer side of the shaft (); such as to prevent turning and radial movements of the supports within the shaft housing while enabling axial movement of the shaft within the shaft housing. The bearings may be forced at either their inner or outer rings, and may be enclosed in a bearing housing () for protection and stiffness. 1100808050100. A drive shaft system () comprising a shaft () , the shaft () connected to a motor via a coupling () and arranged to transmit a rotary movement from the motor to a rotor , the drive shaft system () characterized in that:{'b': 80', '65', '75', '90', '95', '98, 'the shaft () is positioned by at least two supports (, ) within a shaft housing (), each support comprising at least one bearing (, ),'}{'b': 95', '98', '120', '90, 'at least one of the bearings (, ) is enclosed within a bearing housing () installed within and axially moveable along the shaft housing (),'}{'b': 100', '135', '145', '120', '90', '80', '134', '120', '80', '90, 'the drive shaft system () further comprises at least one forcing element (, ) arranged to force the bearing housing () against at least one of: an inner side of the shaft housing (), and an outer side of the shaft () by applying a radial force (A) on the bearing housing (), while enabling axial movement of the shaft () within the shaft housing (), and'}{'b': 135', '145', '140', '120', '120', '90, 'the at least one forcing element (, ) further comprises a stabilizer () arranged to prevent turning and radial movements of the bearing housing () while allowing axial movements of the bearing ...

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01-08-2013 дата публикации

AIRCRAFT ICE PROTECTION SYSTEM AND AIRCRAFT PROVIDED WITH THE SAME

Номер: US20130195658A1
Принадлежит:

An aircraft ice protection system is provided for preventing ice accretion on a wing of an aircraft or removing the accreted ice. Bleed air extracted from a main engine of the aircraft and air introduced from an air intake installed on an airframe and heated by a heat source of the airframe of the aircraft are selectively supplied to a hot air chamber formed inside the wing, thereby carrying out ice protection. 1. An aircraft ice protection system for preventing ice accretion on a wing of an aircraft or removing the accreted ice ,the ice protection system comprising:a hot air chamber which is formed inside the wing of the aircraft;a bleed air supply line which supplies bleed air extracted from a main engine of the aircraft to the hot air chamber;a heated air supply line which supplies air introduced from an air intake of the aircraft to the hot air chamber via a heat source of the aircraft; andswitching device which selectively supplies the bleed air and the air heated by the heat source to the hot air chamber,wherein the wing is heated by the air supplied to the hot air chamber, and ice protection of the wing is carried out.2. The aircraft ice protection system according to claim 1 , whereinthe heat source is at least one of an oil cooler, an oil tank, and a main engine of the aircraft, the oil cooler and the oil tank being mounted on a hydraulic circuit provided on the aircraft.3. The aircraft ice protection system according to claim 2 , whereinthe oil cooler is provided with a double-pipe structured heat transfer pipe which is composed of an inner pipe through which hydraulic oil flows and an outer pipe through which the air introduced from the air intake flows between the outer pipe and the inner pipe, andthe air is heated by exchanging heat with the hydraulic oil flowing through the inner pipe.4. The aircraft ice protection system according to claim 3 , whereinprojected portions are provided on an outer face of the inner pipe.5. The aircraft ice protection ...

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08-08-2013 дата публикации

System and method for gas turbine inlet air heating

Номер: US20130199202A1
Принадлежит: General Electric Co

In one embodiment of the present disclosure, a gas turbine system for part load efficiency improvement and anti-icing within the inlet and at the compressor inlet is described. The system includes a gas turbine having a compressor which receives inlet-air. A direct-contact heat exchanger heats the inlet-air before the inlet-air flows through the inlet and to the compressor. Heating the inlet-air reduces an output of the gas turbine and extends the turndown range, and avoids ice-forming conditions within the inlet and at the compressor inlet bellmouth. The direct-contact heat exchanger may also be configured to act as an evaporative cooler, air chiller, or use liquid dessicant.

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15-08-2013 дата публикации

ROTARY DAMPING MECHANISM WITH PIVOTAL VANES

Номер: US20130209221A1
Принадлежит: C&D ZODIAC, INC.

A damping mechanism that includes a housing having fluid disposed therein and an inner circumferential surface, an axle shaft that is rotatable with respect to the housing, and a first vane having a distal end and being pivotally associated with the axle shaft. When the axle shaft rotates in a first direction, the first vane pivots to a deployed position, and when the axle shaft rotates in a second direction, the first vane pivots to a stowed position. A first clearance is defined between the distal end of the first vane and the inner circumferential surface of the housing when the first vane is in the deployed position, and a second clearance is defined between the distal end of the first vane and the inner circumferential surface of the housing when the first vane is in the stowed position. The second clearance is greater than the first clearance. 1. A damping mechanism comprising:a housing that defines a housing interior that includes a volume of fluid disposed therein, wherein the housing includes an inner circumferential surface,an axle shaft that is rotatable with respect to the housing, andat least a first vane having a distal end and being pivotally associated with the axle shaft and positioned in the housing interior and within the volume of fluid,wherein when the axle shaft and first vane rotate in a first direction, the first vane pivots to a deployed position, and wherein when the axle shaft and first vane rotate in a second direction, the first vane pivots to a stowed position,wherein a first clearance is defined between the distal end of the first vane and the inner circumferential surface of the housing when the first vane is in the deployed position, wherein a second clearance is defined between the distal end of the first vane and the inner circumferential surface of the housing when the first vane is in the stowed position, and wherein the second clearance is greater than the first clearance.2. The damping mechanism of wherein the housing has an ...

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15-08-2013 дата публикации

Turbine assembly

Номер: US20130209253A1
Принадлежит: General Electric Co

According to one aspect of the invention, a turbine assembly includes an airfoil extending from a blade and a dovetail located on a lower portion of the blade, wherein the dovetail has a dovetail contact surface. The turbine assembly also includes a member with a slot configured to couple to the airfoil via the dovetail, the slot having a slot contact surface to contact the dovetail contact surface, wherein the dovetail contact surface is reduced by a relief to alter a fundamental frequency of an assembly of the blade and member.

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22-08-2013 дата публикации

MAGNETICALLY-COUPLED DAMPER FOR TURBOMACHINERY

Номер: US20130216351A1
Автор: Griffin Timothy R.
Принадлежит: DRESSER-RAND COMPANY

A system, method, and apparatus for damping vibration in a rotor supported by primary bearings are provided. The system includes a magnetic coupling configured to magnetically engage a rotor supported by one or more primary bearings, and a piston coupled to the magnetic coupling. The system also includes a damper engaging the piston and configured to damp the rotor, wherein the damper is substantially non-load bearing. 1. A damper system for a rotor , comprising:a magnetic coupling configured to magnetically engage a rotor supported by one or more primary bearings;a piston coupled to the magnetic coupling; anda damper engaging the piston and configured to damp the rotor, wherein the damper is substantially non-load bearing.2. The damper system of claim 1 , wherein the damper includes an eddy current damper.3. The damper system of claim 2 , wherein the eddy current damper includes a housing claim 2 , and one of the housing and the piston includes a magnet and the other includes a conductive material claim 2 , such that movement of the piston in the housing induces eddy currents in the conductive material to resist relative movement between the piston and the housing.4. The damper system of claim 1 , wherein the damper includes a dashpot damper.5. The damper system of claim 1 , wherein the damper includes a sealed housing disposed around the rotor claim 1 , wherein the piston and at least a part of the magnetic coupling are disposed in the sealed housing.6. The damper system of claim 5 , wherein the piston defines orifices extending radially therethrough so as to communicate a first space at least partially defined between a radial-outside of the piston and the sealed housing with a second space at least partially defined between a radial inside of the piston and the sealed housing claim 5 , the sealed housing being substantially filled with a damping fluid claim 5 , such that radial movement of the piston forces the damping fluid through the orifices.7. The damper ...

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29-08-2013 дата публикации

Apparatus and method for conditioning air received by a power generation system

Номер: US20130219916A1
Принадлежит: General Electric Co

According to one aspect of the invention, a method for conditioning air received by a power generation system includes flowing ventilation air through a turbine system to control a temperature of the turbine system and receiving the ventilation air from the turbine system and mixing the ventilation with an ambient air to form an intake air to be directed to a compressor, wherein a temperature of the ventilation air is greater than the ambient air.

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29-08-2013 дата публикации

ANGULAR SECTOR OF A STATOR FOR A TURBINE ENGINE COMPRESSOR, A TURBINE ENGINE STATOR, AND A TURBINE ENGINE INCLUDING SUCH A SECTOR

Номер: US20130223990A1
Автор: Cloarec Yvon
Принадлежит: SNECMA

A stator angular sector for a turbine engine compressor including: an outer shroud and an inner shroud, and at least one vane extending radially between the shrouds. The outer shroud includes first and second mounting mechanisms for mounting the stator angular sector on a casing of the engine, which mechanisms are oriented parallel to the axis in opposite directions and connected together by an intermediate portion. The outer shroud includes at least one axial end portion extending from the intermediate portion with which at least one damper-forming insert is configured to come into contact, such that beyond a given value for amplitude of vibration of the end portion, the damper insert and the end portion are configured to move relative to each other to vary total moving mass moving with the end portion, thereby modifying vibratory behavior of the end portion. 117-. (canceled)18. A stator angular sector for a turbine engine compressor , the sector extending around an axis of radial symmetry , and comprising:an outer shroud and an inner shroud arranged coaxially one inside the other; andat least one vane extending radially between the shrouds and connected thereto at its radial ends;wherein the outer shroud includes first and second mounting means for mounting the stator angular sector on a casing of the engine, the first and second mounting means being oriented parallel to the axis in opposite directions and being connected together by an intermediate portion;wherein the outer shroud includes at least one axial end portion extending from the intermediate portion, provided with a free end, and connected to the radially-outer end of the vane;wherein at least one damper-forming insert is configured to come into contact with the end portion; andwherein beyond a given value for amplitude of vibration of the end portion, the damper insert and the end portion are configured to move relative to each other so as to vary total moving mass moving with the end portion, thereby ...

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29-08-2013 дата публикации

COVER HUB WITH SEALING RING

Номер: US20130224002A1
Принадлежит: SCHAEFFLER TECHNOLOGIES AG & CO. KG

A cover hub for a torque converter includes a radial wall arranged for axial alignment with a cover of the torque converter, a first circumferential surface including an opening for receiving a seal, and a second circumferential surface disposed radially inside of the first circumferential surface and arranged for sealing to a transmission input shaft. The hub has a first fluid passage exiting the hub at a first axial side of the opening and a second fluid passage, circumferentially offset from the first fluid passage and exiting the hub on a second axial side of the opening, opposite the first axial side, and a first circumferential protrusion for sealing the hub to the cover. 1. A cover hub for a torque converter comprising:a radial wall arranged for axial alignment with a cover of the torque converter;a first circumferential surface including an opening for receiving a seal;a second circumferential surface disposed radially inside of the first circumferential surface and arranged for sealing to a transmission input shaft;a first fluid passage exiting the hub at a first axial side of the opening;a second fluid passage, circumferentially offset from the first fluid passage and exiting the hub on a second axial side of the opening, opposite the first axial side; and,a first circumferential protrusion for sealing the hub to the cover.2. The cover hub of wherein at least a portion of the first fluid passage intersects the radial wall or the second circumferential surface.3. The cover hub of wherein at least a portion of the first or second fluid passage intersects the first circumferential surface.4. The cover hub of wherein the hub includes a second circumferential protrusion claim 1 , extending in a direction axially opposite of the first protrusion claim 1 , for engaging a clutch seal plate.5. A cover assembly for a torque converter comprising:a cover including a first radial wall with a circumferential recess; and, a second radial wall with a first circumferential ...

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05-09-2013 дата публикации

LEADING EDGE STRUCTURE, IN PARTICULAR FOR AN AIR INLET OF AN AIRCRAFT ENGINE NACELLE

Номер: US20130230390A1
Принадлежит: AIRCELLE

A leading edge structure for an air inlet of an aircraft nacelle includes a leading edge and an inner partition defining a longitudinal compartment that is located inside the leading edge and accommodates a de-icing and/or anti-icing mat. The leading edge structure includes a multi-axial composite structure placed on top of a heating element for anti-icing and/or de-icing. 1. A leading edge structure for an aircraft nacelle air inlet comprising: a leading edge and an inner partition defining a longitudinal compartment inside said leading edge accommodating at least one of deicing and anti-icing means , wherein said leading edge is formed from at least one multiaxial composite structure placed on top of a heating element designed for deicing and/or anti-icing.2. The structure according to claim 1 , wherein the multiaxial composite structure comprises reinforcing fibers made from carbon claim 1 , copper or aluminum.3. The structure according to claim 1 , wherein it comprises a multiaxial composite structure made using a sewing method.4. The structure according to claim 1 , wherein it comprises a multiaxial composite structure made using a needling method.5. The structure according to claim 1 , wherein the multiaxial composite structure comprises a weaving frame of the angle interlock type.6. The structure according to claim 2 , wherein the multiaxial composite structure comprises reinforcing fibers whereof the orientation is inclined relative to the normal to the plane of the structure.7. The structure according to claim 2 , wherein the multiaxial composite structure comprises reinforcing fibers arranged parallel to the normal of the plane of the structure.8. The structure according to claim 2 , wherein the reinforcing fibers do or do not pass completely through the thickness of the composite structure.9. The structure according to claim 1 , wherein the leading edge has a variable thickness along its profile claim 1 , and in particular claim 1 , for example claim 1 , ...

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12-09-2013 дата публикации

BALANCE CORRECTION WEIGHT PROVIDING CONSTANT MASS

Номер: US20130236292A1
Автор: Milner Glynn
Принадлежит:

A system is provided for balancing a movable turbine part of a turbine or of a compressor. The system includes a balancing weight element with a first hole and a second hole and a fixing element. The first hole and the second hole are formed in such a manner that the fixing element is detachably insertable in either the first hole or the second hole. The first hole is formed in such a manner that the inserted fixing element in the first hole detachably couples the balancing weight element to the movable turbine part in a spatially fixed position. The second hole is formed in such a manner that the fixing element is receivable in the second hole. 111-. (canceled)12. A system for balancing a movable turbine part of a turbine or of a compressor , the system comprisinga balancing weight element with a first hole and a second hole, anda fixing element,wherein the first hole and the second hole are formed in such a manner that the fixing element is detachably insertable in either the first hole or the second hole,wherein the first hole is formed in such a manner that the inserted fixing element in the first hole detachably couples the balancing weight element to the movable turbine part in a spatially fixed position, andwherein the second hole is formed in such a manner that the fixing element is receivable in the second hole.13. The system of claim 12 , further comprisinga further fixing element,wherein the balancing weight element is fixable in a non-detachably manner in the spatially fixed position by the further fixing element.14. The system of claim 13 , wherein the further fixing element is insertable in the first hole.15. The system of claim 14 , wherein the further fixing element is selected from the group consisting of bolts claim 14 , welding points and nails.16. The system of claim 12 , wherein the first hole is a through hole and the second hole is a blind hole.17. The system of claim 12 ,wherein the fixing element comprises a screw, in particular a grub screw ...

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26-09-2013 дата публикации

GAS TURBINE INTAKE ANTI-ICING DEVICE

Номер: US20130247541A1

The gas turbine intake anti-icing device is used for a gas turbine electric power generation system () having a gas turbine () and a power generator () coupled to the gas turbine () and rotationally driven to generate electrical power. The gas turbine intake anti-icing device includes a power generator cooling mechanism (), which takes air from the outside and introduces air into the power generator () to cool the power generator (), and exhaust air supply path () that connects intake path () of the gas turbine () to exhaust path () for air that is discharged from power generator cooling mechanism () after the power generator () is cooled. The air discharged from the power generator cooling mechanism () is supplied to the intake path () of the gas turbine () through the exhaust air supply path (). 112202. A gas turbine intake anti-icing device used for a gas turbine electric power generation system () having a gas turbine () and a power generator () that is coupled to the gas turbine () and rotationally driven to generate electrical power , the gas turbine intake anti-icing device comprising:{'b': 21', '22', '23', '25', '20', '20, 'a power generator cooling mechanism (, , , ) that takes in air from the outside and introduces the air into the power generator () to cool the power generator (); and'}{'b': 31', '61', '9', '2', '30', '21', '22', '23', '25', '20, 'an exhaust air supply path (, ) that connects an intake path () of the gas turbine () to an exhaust path () for air that is discharged from the power generator cooling mechanism (, , , ) after the power generator () is cooled;'}{'b': 21', '22', '23', '25', '9', '2', '31', '61, 'wherein the air discharged from the power generator cooling mechanism (, , , ) is supplied to the intake path () of the gas turbine () through the exhaust air supply path (, ).'}231972. The gas turbine intake anti-icing device according to claim 1 , wherein the exhaust air supply path () is connected to the intake path () nearest the ...

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26-09-2013 дата публикации

Blade Wedge Attachment

Номер: US20130247586A1
Автор: Blake J. Luczak
Принадлежит: Individual

A rotor includes a disk that has slots circumferentially arranged around its periphery. Blades include respective roots that are mounted in respective ones of the slots. The roots are smaller than the slots such that there are circumferential gaps between the roots and circumferential sides of the slots. Wedges are respectively located within the circumferential gaps. The wedges are free floating with regard to the blades and the disk.

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24-10-2013 дата публикации

Airfoil with break-way, free-floating damper member

Номер: US20130276455A1
Принадлежит: Individual

An airfoil includes an airfoil body that has a leading edge and a trailing edge and a first sidewall and a second sidewall that is spaced apart from the first sidewall. The first sidewall and the second sidewall join the leading edge and the trailing edge and at least partially define a cavity in the airfoil body. A damper member is enclosed in the cavity and is free-floating within the cavity.

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24-10-2013 дата публикации

METHOD FOR DAMPING A GAS-TURBINE BLADE, AND VIBRATION DAMPER FOR IMPLEMENTING SAME

Номер: US20130280068A1
Принадлежит: TURBOMECA

A turbine wheel with optimal-mass dampers to dampen a predetermined resonance in context of vibration of a turbine, particularly a low-speed turbine, while assisting in flexibility of adapting to bearing surfaces of recesses of the dampers, by separating mass and flexibility functions by a flexible portion for clamping against the platform, and a mass portion for controlling frictional forces. A damper includes a plate and a counterweight. The plate is stamped from a metal sheet that is substantially thinner than that of the counterweight. The plate includes a wall configured to flexibly contact a platform of the blade of the wheel, while at least partially surrounding a surface of the counterweight. The damper can be used in particular for a wheel of a turbine of a turbine engine, of a fan, or of a BP compressor having mounted blades. 17-. (canceled)8. A damping method for blades mounted on low speed wheel disks of a gas turbine , the turbine including housings recesses under a platform of a blade , configured to receive vibration dampers , the method comprising:carrying out in an independent way a flexible portion clamped against the platform and a mass portion for concentrating efforts so as to direct friction forces against the platform via the clamping action, coupling both parts together in a reversible way, the coupling of both parts being made by surrounding at least partially the mass portion through at least one clamping area of the flexible portion against the platform, the flexible portion being sufficiently flexible to be configured to a required contact level; andinserting the dampers in two parts within the housing recesses being dedicated.9. A vibration damper for implementing the blade damping method according to claim 8 , comprising:a plate and at least one counterweight,the plate being stamped from a metal sheet that is substantially thinner than the counterweight one, andthe plate including a wall configured to flexibly contact at least one ...

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31-10-2013 дата публикации

Damping means for damping a blade movement of a turbomachine

Номер: US20130287583A1
Принадлежит: MTU AERO ENGINES GMBH

A damper () for damping a blade movement of a turbomachine (), and to a method for producing the damper (). The damper () has at least one side surface (′) which can be brought into frictional contact with a friction surface of the turbomachine () in order to damp a blade movement. The side surfaces (′) are asymmetrically convex in shape. 111-. (canceled)12. A damper for damping a blade movement of a turbomachine , the damper comprising:at least one side surface intended to damp the blade movement by frictional contact with a friction surface of the turbomachine, the side surface being asymmetrically convex in shape.13. The damper as recited in wherein the side surface has at least two zones of different radii of curvature.14. The damper as recited in wherein a zone of the side surface radially farther away from a rotor axis of the turbomachine has a smaller radius of curvature than a zone of the side surface radially closer to the rotor axis.15. The damper as recited in wherein the damper has a triangular or polygonal shape in a cross section normal to a rotor axis of the turbomachine.16. The damper as recited in further comprising an anti-rotation device.17. The damper as recited in further comprising a fastener for limiting movement of the damper.18. The damper as recited in wherein the fastener limits movement in a direction of a rotor axis of the turbomachine.19. A turbomachine comprising:a rotor;at least one blade; and{'claim-ref': {'@idref': 'CLM-00012', 'claim 12'}, 'the damper as recited in .'}20. The turbomachine as recited in wherein the blade is a rotor blade coupled to the rotor.21. The turbomachine as recited in wherein the blade has an airfoil and a shroud segment at the end of the airfoil distal from the rotor claim 19 , the shroud segment having a pocket at least partially defining a cavity claim 19 , the damper being disposed in the cavity.22. The turbomachine as recited in wherein the cavity is a closed cavity.23. The turbomachine as recited in ...

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07-11-2013 дата публикации

METHOD FOR THE GENERATIVE PRODUCTION OF A COMPONENT WITH AN INTEGRATED DAMPING ELEMENT FOR A TURBOMACHINE, AND A COMPONENT PRODUCED IN A GENERATIVE MANNER WITH AN INTEGRATED DAMPING ELEMENT FOR A TURBOMACHINE

Номер: US20130294891A1
Принадлежит:

The invention provides a method for the generative manufacture of a component () with integrated damping for a turbomachine, in particular a gas turbine, having the following method steps: building up the component () in a generative manner, and introducing a damping material () into the component () during the method step of the generative building up of the component (). The invention further provides a component () with integrated damping for a turbomachine, in particular a gas turbine, wherein the component () is built up in a generative manner, and the component () has a damping material () which is introduced into the component () during the generative building up of the component (). The invention further provides a turbomachine, in particular a gas turbine, comprising such a component (). 111.-. (canceled)12. A method for the generative manufacture of a component with integrated damping for a turbomachine , wherein the method comprises building up the component in a generative manner , and introducing a damping material into the component during the generative buildup of the component.13. The method of claim 12 , wherein an unsolidified base material of the component is introduced as damping material.14. The method of claim 12 , wherein the component is built up at least in portions thereof with a cavity.15. The method of claim 14 , wherein the damping material is introduced into the cavity.16. The method of claim 14 , wherein during the generative buildup of the component an integrated supporting and/or cooling structure is built up in the cavity.17. The method of claim 15 , wherein during the generative buildup of the component an integrated supporting and/or cooling structure is built up in the cavity.18. The method of claim 17 , wherein the supporting structure is built up at least in portions thereof with a cavity into which the damping material is introduced.19. The method of claim 14 , wherein during the generative buildup of the component a cooling ...

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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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26-12-2013 дата публикации

Systems and Methods for De-Icing a Gas Turbine Engine Inlet Screen and Dehumidifying Inlet Air Filters

Номер: US20130340439A1
Принадлежит: GENERAL ELECTRIC COMPANY

A system for de-icing a gas turbine engine is provided. A manifold is coupled to an inlet screen. A first conduit is coupled to a stage of the compressor and a first input of a mixing component. A second conduit is coupled to the exhaust and a second input of the mixing component. The second conduit being adapted to extract exhaust gases without increasing the pressure at the exhaust. A third conduit is coupled to the output of the mixing component and the manifold. A method for de-icing a gas turbine engine inlet screen includes determining a current temperature at the inlet screen, and determining a desired temperature at the inlet screen. If the current temperature at the inlet screen is less than the desired temperature at the inlet screen first flow rate of an air-exhaust mixture necessary to achieve the desired inlet screen temperature is calculated. The method also includes extracting an amount of exhaust gas from a turbine exhaust subsystem without increasing a pressure at the turbine exhaust subsystem, extracting an amount of air from a compressor stage, and mixing the amount of exhaust gas with the amount of air to generate an air-exhaust mixture that is conveyed to the inlet screen. 1. A method for heating an inlet screen and an inlet air filter in a gas turbine , the method comprising:determining a current inlet screen temperature;determining a desired inlet screen temperature;if the current inlet screen temperature is less than the desired inlet screen temperature there is further included:calculating a first flow rate of an air-exhaust mixture necessary to achieve the desired inlet screen temperature;extracting an amount of exhaust gas from a turbine exhaust subsystem without increasing a pressure at the turbine exhaust subsystem;extracting an amount of air from a compressor stage;mixing the amount of exhaust gas with the amount of air to generate an air-exhaust mixture; andconveying the air-exhaust mixture to the inlet screen, the air-exhaust mixture ...

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26-12-2013 дата публикации

PROCESS FOR PRODUCING A THERMAL BARRIER IN A MULTILAYER SYSTEM FOR PROTECTING A METAL PART AND PART EQUIPPED WITH SUCH A PROTECTIVE SYSTEM

Номер: US20130344349A1
Принадлежит:

A method for producing a thermal barrier in a multilayered system for protecting a metal part made of superalloy, by producing a thermal treatment by flash sintering protection materials in layers superposed on the metal part in an SPS machine enclosure. The layers contain, on a superalloy substrate, at least two layers of zirconium-based refractory ceramics. A metal part is produced according to a SPS flash sintering method and contains a superalloy substrate, a metal sub-layer, a TGO oxide layer and the thermal barrier formed by the method. A first ceramic is an inner ceramic designed to have a substantially higher expansion coefficient. An outer ceramic is designed to have at least lower thermal conductivity, and at least one of a sintering temperature or maximum operating temperature that is substantially higher. The thermal barrier has a composition and porosity gradient from the metal sub-layer to the outer ceramic. 1. A method for producing a thermal barrier in a multilayered system for protecting a metal part made of superalloy , the method comprising producing a thermal treatment by flash sintering protection materials in layers superposed on the metal part in an SPS machine enclosure , the layers comprising , on a superalloy substrate , at least two layers of zirconium-based refractory ceramics ,whereina first ceramic layer, is an inner layer, that is chemically and thermally compatible with the substrate, and a final ceramic layer, is an outer layer, disposed over the other layers, and has a higher physicochemical resistance property, in relation to pollutants of the CMAS type, higher thermal resistance property, or both, than the inner layer.2. The method according to claim 1 , wherein the inner layer has a thermal expansion coefficient that is substantially higher than the final ceramic layer.3. The method according to claim 1 , wherein the physicochemical resistance property of the outer layer is at least one selected from the group consisting of ...

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02-01-2014 дата публикации

WIRELESS COMPONENT AND METHODS OF FABRICATING A COATED COMPONENT USING MULTIPLE TYPES OF FILLERS

Номер: US20140004310A1
Принадлежит: GENERAL ELECTRIC COMPANY

Methods of fabricating coated components using multiple types of fillers are provided. One method comprises forming one or more grooves in an outer surface of a substrate. Each groove has a base and extends at least partially along the outer surface of the substrate. The method further includes disposing a sacrificial filler within the groove(s), disposing a permanent filler over the sacrificial filler, disposing a coating over at least a portion of the substrate and over the permanent filler, and removing the first sacrificial filler from the groove(s), to define one or more channels for cooling the component. A component with a permanent filler is also provided. 1. A component comprising:a substrate comprising an outer surface and an inner surface, wherein the outer surface defines one or more grooves, wherein each of the one or more grooves extends at least partially along the outer surface of the substrate and has a base;a permanent filler disposed within and extending across a top of each of the one or more grooves; anda coating disposed over at least a portion of the substrate and over the permanent filler, wherein the one or more grooves and the permanent filler or coating together define one or more channels for cooling the component.2. The component of claim 1 , wherein the permanent filler comprises at least one of tungsten claim 1 , nickel claim 1 , cobalt claim 1 , molybdenum claim 1 , chromium claim 1 , aluminum claim 1 , and alloys thereof.3. The component of claim 1 , wherein the inner surface defines at least one hollow claim 1 , interior space claim 1 , and wherein one or more access holes extend through the base of a respective one of the one or more grooves to place the groove in fluid communication with respective ones of the at least one hollow interior space.4. The component of claim 1 , wherein each of the one or more grooves has a top claim 1 , and wherein the base of the groove is wider than the top claim 1 , such that each of the one or ...

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30-01-2014 дата публикации

METHOD OF APPLICATION FOR LAYERED THERMAL BARRIER COATING WITH BLENDED TRANSITION

Номер: US20140030446A1
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A method includes generating a plasma plume with a plasma gun, delivering a plurality of coating materials to the plasma plume with a powder feeder assembly to vaporize the coating materials. The delivery includes delivering a first (bond coat) material from a first powder feeder to the plasma gun, ceasing delivery of the first material, increasing a rate of delivery of a second (rare earth stabilized zirconia) material from a second powder feeder to the plasma plume, increasing a rate of delivery of a third material (a rare earth stabilized zirconia material different from the second material) from a third powder feeder to the plasma plume, decreasing a rate of delivery of the second material, and decreasing a rate of delivery of the third material, and depositing the plurality of coating materials on a work piece to produce a layered coating with blended transitions between coating layers. 1. A method of coating a work piece , the method comprising:positioning a work piece in a process chamber;generating a plasma plume with a plasma gun; delivering a first material from a first powder feeder to the plasma gun, wherein the first material comprises a bond coat material;', 'ceasing delivery of the first material from the first powder feeder to the plasma plume;', 'increasing a rate of delivery of a second material from a second powder feeder to the plasma plume, wherein the second material comprises a rare earth stabilized zirconia material;', 'increasing a rate of delivery of a third material from a third powder feeder to the plasma plume, wherein the third material comprises a rare earth stabilized zirconia material different from the second material;', 'decreasing a rate of delivery of the second material from the second powder feeder to the plasma plume; and', 'decreasing a rate of delivery of the third material from the third powder feeder to the plasma plume; and, 'delivering a plurality of coating materials to the plasma plume generated by the plasma gun with ...

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27-02-2014 дата публикации

ANNULAR TURBOMACHINE SEAL AND HEAT SHIELD

Номер: US20140056685A1
Принадлежит: UNITED TECHNOLOGIES CORPORATION

An example annular turbomachine seal includes an attachment portion, a first leg radially inward of the attachment portion, and a second leg radially inward of the attachment portion and configured to block flow from directly contacting an attachment associated with a bearing compartment of a turbomachine. 1. An annular turbomachine seal , comprising:an attachment portion;a first leg radially inward of the attachment portion; anda second leg radially inward of the attachment portion and configured to block flow from directly contacting an attachment associated with a bearing compartment of a turbomachine.2. The annular turbomachine seal of claim 1 , wherein the second leg is configured to seal against the bearing compartment.3. The annular turbomachine seal of claim 1 , wherein second leg extends exclusively in a radial direction.4. The annular turbomachine seal of claim 1 , wherein the second leg extends from the attachment portion radially past the attachment.5. (canceled)6. The annular turbomachine seal of claim 1 , wherein the first leg extends axially forward of the second leg.7. The annular turbomachine seal of claim 1 , wherein at least a portion of the first leg and the second leg are radially aligned.8. The annular turbomachine seal of claim 1 , wherein at least a portion of the first leg and at least a portion of the second leg are radially aligned with the attachment.9. The annular turbomachine seal of claim 1 , wherein the attachment portion claim 1 , the first leg claim 1 , and the second leg are formed as a single monolithic structure.10. The annular turbomachine seal of claim 1 , including a third leg radially outward of the attachment portion claim 1 , the third leg configured to provide a seal with a turbine section of the turbomachine.11. The annular turbomachine seal of claim 1 , wherein the attachment portion provides a plurality of apertures each configured to receive a fastener that secures the attachment portion directly to a low pressure ...

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27-02-2014 дата публикации

GERMANATE-CONTAINING THERMAL BARRIER COATING

Номер: US20140057129A1
Автор: CASSUTO James
Принадлежит: Thermatin Industries, LLC

A thermal barrier coating having a reduced high temperature thermal conductivity includes group II germanate constructs. This thermal barrier coating may be applied directly to a substrate, applied to a bond-coated substrate, and/or incorporated into a protective coating including one or more other thermal barrier coating layers. The thermal barrier coating provides improved thermal protection properties over current industry standards and materials considered for thermal protection applications. 1. A thermal barrier coating , comprising:{'sub': 2', '4, 'a group II germanate having the formula YGeO, wherein Y is chosen from Be, Mg, Ca, Sr, Ba, Ra, or a combination of two or more thereof.'}2. The thermal barrier coating of claim 1 , wherein Y comprises Mg.3. The thermal barrier coating of claim 1 , wherein the thermal barrier coating is doped with one or more metals.4. The thermal barrier coating of claim 3 , wherein the thermal barrier coating is doped with Cr claim 3 , Fe claim 3 , or a combination of Cr and Fe.5. The thermal barrier coating of claim 1 , wherein the thermal barrier coating comprises an olivine crystalline structure.6. The thermal barrier coating of claim 1 , wherein the thermal barrier coating has a thermal conductivity of less than 2.0 W.mKat about 700° C.7. The thermal barrier coating of claim 1 , wherein the thermal barrier coating has a minimum linear thermal expansion coefficient of 9.0×10° C.at about 0° C. to about 1000° C.8. The thermal barrier coating of claim 1 , wherein the thermal barrier coating comprises a maximum use temperature of greater than about 1600° C.9. The thermal barrier coating of claim 1 , wherein the thermal barrier coating has a density of less than 7.5 g/cm.10. A method of applying the thermal barrier coating of claim 1 , the method comprising applying the group II germanate by thermal spray coating claim 1 , electron beam physical vapor deposition claim 1 , or enameling.11. A coated substrate claim 1 , comprising:a ...

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06-03-2014 дата публикации

System and Method for Vibration Isolation

Номер: US20140064922A1
Принадлежит: Bell Helicopter Textron Inc.

In accordance with one embodiment of the present disclosure, a system includes a first housing, a second housing, a seal, and a spring system. The first housing includes a first volume of fluid. The first housing is capable of connecting to a first element and to a second element, and is also capable of reducing an amount of movement transferred from the first element to the second element. The second housing is connected to the first housing. The second housing includes a second volume of fluid and a volume of gas. The first volume of fluid is in fluid communication with the second volume of fluid. The seal is capable of separating the second volume of fluid from the volume of gas. The spring system is capable of applying pressure to the first volume of fluid and the second volume of fluid. 1. An aircraft , comprising:a rotor comprising a plurality of aircraft blades operable to revolve around an axis; a first portion operable to couple to a fuselage of the aircraft;', 'a second portion operable to couple to the fuselage of the aircraft;', 'a moveable portion coupled to the first portion and the second portion, the moveable portion operable to couple to the rotor; and', 'a first volume of fluid, wherein the first housing is operable to reduce an amount of movement transferred from the rotor to the fuselage of the aircraft by transferring a portion of the first volume of fluid from the second portion of the first housing to the first portion of the first housing through the moveable portion;, 'a first housing comprisinga second housing coupled to the first housing, the second housing comprising a second volume of fluid and a volume of gas, the first volume of fluid being in fluid communication with the second volume of fluid;a rubber rolling seal positioned within the second housing, the rubber rolling seal operable to separate the second volume of fluid from the volume of gas; anda mechanical spring positioned within the second housing, the mechanical spring ...

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06-03-2014 дата публикации

IMPELLER AND METHOD FOR DRIVING FLUIDS USING THE SAME

Номер: US20140064973A1
Принадлежит: JOHNSON ELECTRIC S.A.

Disclosed are mechanisms for an impeller () and method of reducing noise levels while driving fluids with impellers. Exemplary implementations include a hub () and multiple blades () separably attached to or inseparably formed on the hub. The hub may be used to effectively reduce noise levels during operations of the impeller by having a first cylindrical feature (), a first number of first ribs (), and a second number of second ribs (), while maintaining substantially similar mechanical properties or operational characteristics. The hub may further optionally include a second substantially cylindrical feature () that is separably attached to or is inseparably formed on the hub to further enhance one or more properties or characteristics of the hub while serving to reduce the noise level of the impeller. 1. An impeller , comprising:a hub that includes a top structural member and a first cylindrical feature;a number of blades attached to the first cylindrical feature;a first rib attached to the top structural member; anda second rib attached to at least a bottom portion of the first cylindrical feature and at least a portion of the top structural member.2. The impeller of claim 1 , wherein:the hub further comprises a second cylindrical feature substantially concentric with the first cylindrical feature, andboth the first cylindrical feature and the second cylindrical feature are substantially cylindrical.3. The impeller of claim 2 , wherein the first rib is separately attached to or is inseparably formed as an integral part of the second cylindrical feature claim 2 , and the number of blades are separably or inseparably attached to the first cylindrical feature.4. The impeller of claim 2 , whereinthe first cylindrical feature is separably attached to or is inseparably formed as a part of the top structural member, andthe first rib extends outward from a center of the hub and protrudes beyond an outer diameter of the second cylindrical feature and is separably or ...

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13-03-2014 дата публикации

DEVICE FOR DE-ICING A TURBOMACHINE SEPARATOR

Номер: US20140072405A1
Принадлежит: SNECMA

A turbomachine separator including a device for de-icing the turbomachine separator, and a distribution element, wherein the separator is formed by an inner ferrule and an outer ferrule, wherein the inner ferrule is fitted with a first mounting flange and a second mounting flange, the de-icing device including an internal air supply duct, able to inject air into the separator, wherein the internal supply duct is connected to an air inlet, wherein the air inlet forms a projection external to the de-icing device, allowing a flexible connection with a tube of the distribution element for conveying hot air, a first fastener constructed and arranged to be attached to the first mounting flange; and a second fastener constructed and arranged to be attached to the second mounting flange. 1. A turbomachine separator comprising a device for de-icing the turbomachine separator , and a distribution element , wherein the separator is formed by an inner ferrule and an outer ferrule , wherein the inner ferrule is fitted with a first mounting flange and a second mounting flange , the de-icing device comprising:an internal air supply duct, able to inject air into the separator, wherein said internal supply duct is connected to an air inlet, wherein the air inlet forms a projection external to the de-icing device, allowing a flexible connection with a tube of the distribution element for conveying hot air,a first fastener constructed and arranged to be attached to the first mounting flange;a second fastener constructed and arranged to be attached to the second mounting flange.2. The separator according to claim 1 , wherein the flexible connection is produced by fitting an air conveyance tube into the inlet of the de-icing device.3. The separator according to claim 1 , wherein the flexible connection is produced by a seal positioned in the air inlet such that the tube is fitted and held directly in the seal.4. The separator according to claim 1 , wherein the device forms a bridge ...

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13-03-2014 дата публикации

Filled static structure for axial-flow machine

Номер: US20140072407A1
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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20-03-2014 дата публикации

Heated Screen For Air Intake Of Aircraft Engines

Номер: US20140077039A1
Автор: Scimone Michael J.
Принадлежит: AEROSPACE FILTRATION SYSTEMS, INC.

An aircraft includes a fuselage and wings mounted on opposite sides of the fuselage for sustained forward flight. An engine is mounted in the fuselage or at least one of the wings and includes an air intake. At least a portion of the air intake generally faces the forward direction for receiving intake air during forward flight. A filter assembly is mounted adjacent the air intake and disposed to impinge air and block objects from passing therethrough. A heated screen includes a heater and is mounted adjacent the air intake and upstream of the engine such that ice entering the air intake contacts the heated screen before entering the engine. A power source is provided to supply power to the heater. 1. An aircraft comprising:a fuselage;wings mounted on opposite sides of the fuselage for sustained forward flight;an engine mounted in the fuselage or at least one of the wings and including an air intake, at least a portion of the air intake generally facing forward direction for receiving intake air during forward flight;a filter assembly mounted adjacent the air intake and disposed to impinge air and block objects from passing therethrough;a heated screen including a heater embedded therein, the screen mounted adjacent the air intake and upstream of the engine such that ice entering the air intake contacts the heated screen before entering the engine; anda power source for providing power to the heater.2. The aircraft of claim 1 , wherein the air intake includes a bypass that is movable from a closed position for directing air through the filter to an open position for allowing unfiltered air to enter engine claim 1 , the bypass inhibiting unfiltered air from entering the engine during hovering or when the aircraft is near the ground.3. The aircraft of claim 1 , further comprising a plurality of engines mounted in nacelles claim 1 , and associated rotors.4. The aircraft of claim 3 , wherein the aircraft is a tiltrotor aircraft wherein the rotation axis of each rotor is ...

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20-03-2014 дата публикации

Flat Bottom Damper Pin For Turbine Blades

Номер: US20140079529A1
Принадлежит: GENERAL ELECTRIC COMPANY

A damper system for the buckets of a gas turbine engine. The damper system may include damper pins having a generally rounded top portion and a generally flat bottom portion along substantially the entire length thereof. The generally flat bottom portion may allow addition or removal of material to or from the pin in order to achieve an optimal dynamic weight ratio. 1. A damper pin for a bucket damper slot in at least one of two adjoining buckets installed in a rotor in a turbine , the damper pin comprising:a. a generally rounded top portion, the rounded top portion configured to contact said adjoining buckets prior to full-speed rotation of said rotor; andb. a generally flat bottom portion.2. The damper pin of claim 1 , wherein the damper pin is substantially symmetrical in cross section.3. The damper pin of claim 1 , wherein the damper pin includes bossed ends.4. The damper pin of claim 1 , wherein the generally flat bottom portion extends across substantially the entire length of the damper pin.5. The damper pin of claim 1 , wherein the generally rounded top portion comprises a fully round portion.6. The damper pin of claim 1 , wherein the damper pin includes a Murphy proofing tab on the bottom thereof7. The damper pin of claim 1 , wherein said generally rounded top portion and said generally flat bottom portion are joined at a pair of rounded or beveled edges.8. A turbine engine comprising:a. a rotor having a generally circular periphery;b. a plurality of buckets mounted about the periphery of the rotor, each of said buckets including an airfoil extending outward from a platform; andc. a damper for damping vibrations of said plurality of buckets, the damper including an elongate body, the elongate body being sized and shaped for receipt within a bucket damper slot formed in undercuts of said plurality of buckets so that the elongate body frictionally engages said undercuts to dampen vibrations of said buckets, said damper being generally symmetrical in cross ...

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10-04-2014 дата публикации

DUCT DAMPER

Номер: US20140099182A1
Принадлежит:

An example damper includes a damping member configured to damp a duct wall at an interface between a duct band and the duct wall 1. A damper comprising:a damping member configured to damp a duct wall at an interface between a duct band and the duct wall.2. The damper of claim 1 , wherein the damping member comprises a viscoelastic material.3. The damper of claim 1 , wherein the damping member comprises a synthetic polymer material.4. The damper of claim 1 , wherein the damping member comprises a metallic structure.5. The damper of claim 1 , wherein the damping member comprises a synthetic fluoropolymer.6. The damper of claim 1 , wherein the damping member is secured directly to the duct band such that the damping member moves with duct band as the duct band is moved relative to the duct wall.7. The damper of claim 1 , wherein the damping member is an annular damping member and provides a plurality of radially extending apertures.8. The damper of claim 1 , wherein the duct band engages the duct wall exclusively through the damping member.9. The damper of claim 1 , wherein the damping member covers an inwardly facing surface of the duct band.10. A turbomachine damping assembly claim 1 , comprising:a duct wall between a core flowpath and a bypass flowpath of a turbomachine;a duct band disposed about a radially outer surface of the duct wall; anda damping member between the duct wall and the duct band.11. The turbomachine damping assembly of claim 10 , wherein the duct wall claim 10 , the duct band claim 10 , and the damping member each provide apertures configured to communicate flow between the core flowpath and the bypass flow path.12. The turbomachine damping assembly of claim 11 , including an actuation system that moves the duct band relative to the duct wall to selectively adjust flow through the apertures.13. The turbomachine damping assembly of claim 12 , wherein the damping member is secured to the duct band such that the damping member moves with the duct ...

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06-01-2022 дата публикации

TURBINE

Номер: US20220003129A1
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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05-01-2017 дата публикации

Nacelle compression rods

Номер: US20170002684A1
Автор: Stuart J. Byrne
Принадлежит: Rohr Inc

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

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05-01-2017 дата публикации

Dual pressure deicing system

Номер: US20170002736A1
Автор: Galdemir Botura
Принадлежит: Rohr Inc

A deicing system for an aircraft may comprise a dual pressure regulating valve. The dual pressure regulating valve may comprise a low pressure setting and a high pressure setting. The low pressure setting may supply a relatively lesser supply of bleed air to an aircraft component, and the high pressure setting may supply a relatively greater supply of bleed air to the aircraft component. The dual pressure regulating valve may be switched between the low pressure setting and the high pressure setting based on aircraft or atmospheric conditions to prevent heat damage to the aircraft component.

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07-01-2016 дата публикации

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

Номер: US20160003095A1
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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07-01-2016 дата публикации

TURBOCHARGER INTERNAL TURBINE HEAT SHIELD HAVING AXIAL FLOW TURNING VANES

Номер: US20160003096A1
Автор: Fraser Brock
Принадлежит:

A turbocharger internal heat shield () is provided having axial flow turning vanes (). Additionally, the heat shield may have a volute divider wall extender (). 112. A heat shield () for use in a turbocharger having an axial flow turbine wheel , having turning vanes () which turn the flow of exhaust gas to the axial direction.212. A heat shield () according to having 4 to 8 turning vanes ().312. A heat shield () according to having 7 turning vanes ().415. A heat shield () according to further comprising a volute divider wall extender (). 1. Field of the InventionThis invention relates to a turbocharger for an internal combustion engine. More particularly, this invention relates to turbocharger having an axial flow turbine wheel and an internal heat shield having turning vanes. Optionally, for a twin exhaust gas volute the heat shield may have a divider wall extender.2. Description of Related ArtA turbocharger is a type of forced induction system used with internal combustion engines. Turbochargers deliver compressed air to an engine intake, allowing more fuel to be combusted, thus boosting an engine's horsepower without significantly increasing engine weight. Thus, turbochargers permit the use of smaller engines that develop the same amount of horsepower as larger, normally aspirated engines. Using a smaller engine in a vehicle has the desired effect of decreasing the mass of the vehicle, increasing performance, and enhancing fuel economy. Moreover, the use of turbochargers permits more complete combustion of the fuel delivered to the engine, which contributes to the highly desirable goal of a cleaner environment.Turbochargers typically include a turbine housing connected to the engine's exhaust manifold, a compressor housing connected to the engine's intake manifold, and a center bearing housing coupling the turbine and compressor housings together. A turbine wheel in the turbine housing is rotatably driven by an inflow of exhaust gas supplied from the exhaust ...

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07-01-2016 дата публикации

INTERLOCKING ROTOR ASSEMBLY WITH THERMAL SHIELD

Номер: US20160003097A1
Принадлежит:

A rotor assembly includes a first rotor, a second rotor mounted on the first rotor and co-rotatable there with, and a thermal shield interlocked with the second rotor for co-rotation there with. 1. A rotor assembly comprising:a first rotor;a second rotor mounted on the first rotor and co-rotatable there with; anda thermal shield interlocked with the second rotor for co-rotation there with.2. The rotor assembly as recited in claim 1 , wherein the first rotor has a first outer diameter and the second rotor has a second outer diameter that is smaller than the first outer diameter.3. The rotor assembly as recited in claim 1 , wherein the second rotor includes at least one radially-extending tab and the thermal shield includes at least one radially-extending tab circumferentially interlocked with the at least one radially-extending tab of the second rotor.4. The rotor assembly as recited in claim 3 , wherein the at least one radially-extending tab of the second rotor includes a step.5. The rotor assembly as recited in claim 1 , wherein the second rotor includes a plurality of radially-extending circumferentially-spaced tabs and the thermal shield includes a plurality of radially-extending circumferentially-spaced tabs circumferentially interlocked with the plurality of radially-extending circumferentially-spaced tabs of the second rotor.6. The rotor assembly as recited in claim 5 , wherein the first rotor includes a plurality of radially-extending circumferentially-spaced tabs and the plurality of radially-extending circumferentially-spaced tabs of the second rotor are circumferentially interlocked with the plurality of radially-extending circumferentially-spaced tabs of the first rotor.7. The rotor assembly as recited in claim 6 , wherein the plurality of radially-extending circumferentially-spaced tabs of the thermal shield are circumferentially aligned with the plurality of radially-extending circumferentially-spaced tabs of the first rotor.8. The rotor assembly as ...

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07-01-2016 дата публикации

Integrated Flex Support and Front Center Body

Номер: US20160003105A1
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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07-01-2016 дата публикации

ROTATING INLET COWL FOR A TURBINE ENGINE, COMPRISING AN ECCENTRIC FORWARD END

Номер: US20160003146A1
Принадлежит: SNECMA

A rotating inlet cowl for a turbine engine includes a rotation axis. The rotating inlet cowl includes a forward cone defining a forward end of the inlet cowl. The forward end is configured to be eccentric relative to the rotation axis of the inlet cowl. Furthermore, the forward cone is truncated by a truncation surface defining the forward end of the inlet cowl. 1. A rotating inlet cowl of a gas turbine engine , the inlet cowl including a rotation axis and comprising:a forward cone defining a forward end of the inlet cowl,wherein the forward end is configured to be eccentric relative to the rotation axis of the inlet cowl,wherein the forward cone is truncated by a truncation surface defining the forward end of the inlet cowl, andwherein the forward cone includes an axis parallel to and coincident with the rotation axis of the inlet cowl.2. The rotating inlet cowl according to claim 1 , wherein the forward cone includes at least one balancing bead that includes a variable thickness along a circumferential direction claim 1 , to compensate for an unbalanced mass.3. A turbine or aircraft engine claim 1 , comprising the rotating inlet cowl according to .4. A rotating inlet cowl of a gas turbine engine claim 1 , the inlet cowl including a rotation axis and comprising:a forward cone defining a forward end of the inlet cowl,wherein the forward end is configured to be eccentric relative to the rotation axis of the inlet cowl,wherein the forward cone is truncated by a truncation surface defining the forward end of the inlet cowl, andwherein the truncation surface is approximately a plane that is inclined relative to a plane orthogonal to the rotation axis of the inlet cowl.5. The rotating inlet cowl according to claim 4 , wherein the forward cone includes at least one balancing bead that includes a variable thickness along a circumferential direction claim 4 , to compensate for an unbalanced mass.6. The rotating inlet cowl according to claim 4 , wherein the truncation ...

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07-01-2016 дата публикации

System for preventing icing on an aircraft surface operationally exposed to air

Номер: US20160003147A1
Принадлежит:

A system for preventing icing on an aircraft surface includes a plasma actuator which is applied onto an aircraft surface operationally exposed to air and which is arranged for generating at least one plasma discharge (D) for inducing a flow (F) of ionized hot-air particles towards the surface. 1. Engine nacelle including a system for preventing icing on an aircraft surface operationally exposed to air; said system being positioned at a covering lip and extending over exposed surfaces of at least one radially external side wall and one radially internal side wall adjacent to the lip;said system comprising: a dielectric barrier discharge type plasma actuator to be applied onto said surface operationally exposed to air, and arranged for generating at least one plasma discharge for inducing a flow of ionized hot-air particles towards said surface operationally exposed to air; an intermediate portion of dielectric material;', 'at least one exposed electrode portion positioned on an outer side of said intermediate portion, the at least one exposed electrode portion being exposed to air; and', 'at least one covered electrode portion of said intermediate portion, the at least one covered electrode portion is operationally shielded from air on said outer side;, 'said actuator comprisingsaid electrode portions being electrically connectable to a high-voltage electric power generator and being adapted to be energized by said electric power to generate said plasma discharge between the electrode portions;said covered electrode portion being at least partially positioned under an inner side, opposite to said outer side, of said intermediate portion;an insulating portion of dielectric material positioned under said covered electrode portion, configured for application onto said exposed surface of said aircraft.23-. (canceled)4. Engine nacelle according to claim 1 , wherein said exposed electrode portion is positioned on top of said intermediate portion.5. Engine nacelle ...

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01-01-2015 дата публикации

VIBRATION-PROOF MEMBER AND FAN ASSEMBLY HAVING THE SAME

Номер: US20150003987A1
Автор: FU LI-REN
Принадлежит:

A vibration-proof member for a cooling fan includes a plate to absorb vibration, a post perpendicularly extending out from a side of the absorbing plate, and an extending portion perpendicularly extending out from a part of a lateral side of the absorbing plate. The construction and method of installing the vibration-proof member allows any required number of the vibration-proof members to be installed and to function effectively whether the fan is installed on a vertical bracket or on a horizontal bracket. 1. A vibration-proof member , comprising:a vibration-absorbing plate;a post perpendicularly extending out from a side of the absorbing plate; andan extending portion perpendicularly extending out from a part of a lateral side of the absorbing plate.2. The vibration-proof member of claim 1 , wherein the absorbing plate is round claim 1 , the extending portion perpendicularly extends out from a part of a circumference of the absorbing plate and has an arc-shaped cross-section.3. The vibration-proof member of claim 1 , wherein an annular protrusion protrudes from a circumference of the post.4. The vibration-proof member of claim 3 , wherein a surface of the protrusion away from the post is tapered toward a distal end of the post.5. A fan assembly claim 3 , comprising:a fan comprising an end board defining a plurality of through holes; and a vibration-absorbing plate contacting an outer surface of the end board;', 'a post perpendicularly extending out from a side of the absorbing plate to extend through one of the through holes; and', 'an extending portion perpendicularly extending out from a part of a lateral side of the absorbing plate and contacting a lateral side of the end board., 'a plurality of vibration-proof members; each vibration-proof member comprising6. The fan assembly of claim 5 , wherein the end board is rectangular claim 5 , the lateral sides of four corners of the end board are round corners claim 5 , the absorbing plate is round claim 5 , the ...

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04-01-2018 дата публикации

Cavity Sealing

Номер: US20180003063A1
Автор: DOORBAR Phillip J.
Принадлежит: ROLLS-ROYCE PLC

A method of sealing one or more openings provided in a wall of an aerofoil for a gas turbine engine, the aerofoil comprising at least one cavity which is at least partly filled with a vibration damping material, the method comprising steps to provide a metallic material onto the wall of the aerofoil in order to cover the opening and bond the metallic material to the wall of the aerofoil to seal the opening. 1. A method of sealing one or more openings provided in a wall of an aerofoil for a gas turbine engine , the aerofoil comprising at least one cavity which is at least partly filled with a vibration damping material , the method comprising steps to:a) provide a metallic material onto the wall of the aerofoil in order to cover the opening; and,b) bond the metallic material to the wall of the aerofoil to seal the opening.2. A method as claimed in , wherein step a) of comprises steps to overlap the metallic material and the wall such that a weld interface between the metallic material and the wall is generally co-planar with the wall of the aerofoil.3. A method as claimed in claim 1 , wherein the metallic material comprises a metal or alloy comprising either or both of substantially identical chemical composition or mechanical properties to the wall of the aerofoil.4. A method as claimed in claim 1 , wherein the method comprises steps to form a planar section on the wall of the aerofoil in an area adjacent to or surrounding the opening.5. A method as claimed in claim 4 , wherein the steps to form a planar section on the wall of the aerofoil comprises further steps to remove at least a portion of the wall from the aerofoil.6. A method as claimed in claim 1 , wherein the metallic material comprises two or more layers.7. A method as claimed in claim 1 , wherein step b) of comprises steps to heat and plasticise at least a portion of the aerofoil wall and one or more layers of the metallic material.8. A method as claimed in claim 1 , wherein the method comprises steps to ...

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04-01-2018 дата публикации

SUPRESSING VIBRATIONS OF SHAFTS USING ADJUSTABLE BEARINGS

Номер: US20180003075A1
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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02-01-2020 дата публикации

Bearing device for load reduction

Номер: US20200003075A1
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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02-01-2020 дата публикации

METHOD OF REGULATING AIR PRESSURE IN ANTI-ICING SYSTEM

Номер: US20200003117A1
Принадлежит:

An anti-icing system of a nacelle inlet of an engine of an aircraft includes first and second direct acting valves and first and second control valve assemblies fluidly connected to the nacelle inlet. The first direct acting valve includes a first inlet, outlet, valve chamber, and piston. The first piston is positioned in the first direct acting valve. The first control valve assembly is fluidly connected to the first valve. The second direct acting valve includes a second inlet, outlet, valve chamber, and piston. The second piston is positioned in the second direct acting valve. The second direct acting valve is fluidly connected to the first direct acting valve in a series configuration. The second control valve assembly is fluidly connected to the second valve chamber. 1. A method of regulating air pressure in an anti-icing system of a nacelle inlet of an engine of an aircraft , the method comprising: a first direct acting valve with a first valve chamber, a first internal valve body, and a first piston slidably engaged with the first internal valve body;', 'a first control valve assembly with a first solenoid valve and fluidly connected to the first valve chamber of the first direct acting valve;', 'a second direct acting valve with a second valve chamber, a second internal valve body, and a second piston slidably engaged with the second internal valve body, wherein the second direct acting valve is fluidly connected to the first direct acting valve in a series configuration;', 'a second control valve assembly with a second solenoid valve and fluidly connected to the second valve chamber of the second direct acting valve; and, 'flowing air into a valve assembly comprising adjusting at least one of the first control valve assembly and the second control valve assembly in response to the temperature of the air in the outlet of the second direct acting valve by controlling an amount of electric current fed into at least one of the first solenoid valve in the first ...

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07-01-2021 дата публикации

COMPRESSOR BLADE HAVING ORGANIC VIBRATION STIFFENER

Номер: US20210003017A1
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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07-01-2021 дата публикации

TURBOMACHINERY HEAT MANAGEMENT SYSTEM

Номер: US20210003033A1
Принадлежит:

A system is provided, including a heat management system. The heat management system includes a thermal delivery system configured to providing heating, cooling, or a combination thereof, to a first zone of a turbomachinery, and a controller operatively coupled to the thermal delivery system and configured to control the heating, the cooling, or the combination thereof, of the first zone, to minimize or to eliminate positional changes, structural changes, or a combination thereof, in one or more components of the turbomachinery due to thermal energy. 1. A system , comprising: a thermal delivery system configured to providing heating, cooling, or', 'a combination thereof, to a first zone of a turbomachinery; and', 'a controller operatively coupled to the thermal delivery system and configured to control the heating, the cooling, or the combination thereof, of the first zone, to minimize or to eliminate positional changes, structural changes, or a combination thereof, in one or more components of the turbomachinery due to thermal energy., 'a heat management system, comprising2. The system of claim 1 , wherein the controller is configured to independently control the heating claim 1 , the cooling claim 1 , or the combination thereof claim 1 , of the first zone from the heating claim 1 , the cooling claim 1 , or the combination thereof claim 1 , of a second zone of the turbomachinery.3. The system of claim 1 , wherein the one or more components comprise a lower shell mechanically coupled to an upper shell of a gas turbine engine claim 1 , and wherein the controller is configured to control the heating claim 1 , the cooling claim 1 , or the combination thereof claim 1 , by deriving a temperature difference between the lower shell and the upper shell.4. The system of claim 3 , wherein the controller is configured to control the heating claim 3 , the cooling claim 3 , or the combination thereof claim 3 , by comparing the temperature different to a temperature difference ...

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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
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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08-01-2015 дата публикации

Splitter Nose with a Sheet That Forms a Surface to Guide the Flow and Acts as a De-Icing Duct

Номер: US20150007895A1
Принадлежит:

The present application relates to a splitter nose of an axial turbomachine configured to separate an annular flow into the turbomachine into a primary flow and a secondary flow, and including: a generally circular leading edge, an annular wall extending from the leading edge and bounding the secondary flow, and at least one duct for a de-icing fluid for the splitter nose extending substantially axially along the wall and opening out into the primary flow. The external surface of the wall is formed by a sheet bounding the de-icing duct. 1. A splitter nose of an axial turbomachine configured to separate a flow entering the turbomachine into a primary flow and a secondary flow , the splitter nose comprising:a generally circular leading edge;an annular wall extending from the leading edge and bounding the secondary flow;at least one duct for a de-icing fluid for the splitter nose extending substantially axially along the wall and opening into the primary flow;wherein the external surface of the wall is formed by a sheet bounding the de-icing duct.2. The splitter nose in accordance with claim 1 , wherein the duct has an essentially constant thickness over the major part of the length thereof axially along the wall.3. The splitter nose in accordance with claim 1 , wherein the sheet is at least one annular sheet and has a profile with a downstream part substantially straight and a curved upstream part which forms the leading edge.4. The splitter nose in accordance with claim 1 , wherein the annular wall comprises:a support of the sheet on which the exterior surface has a step facing the downstream edge of the sheet, so that the exterior surface of the sheet is level with that of the support at the step.5. The splitter nose in accordance with claim 1 , wherein the annular wall forms an annular hook at the leading edge with an annular groove open axially downstream.6. The splitter nose in accordance with claim 1 , further comprising: 'an annular centering surface mating ...

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11-01-2018 дата публикации

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

Номер: US20180009009A1
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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12-01-2017 дата публикации

MECHANICAL COMPONENT FOR THERMAL TURBO MACHINERY

Номер: US20170009601A1
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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12-01-2017 дата публикации

TURBO MACHINE

Номер: US20170009602A1
Принадлежит:

A turbo machine of the present disclosure includes a cylindrical bearing housing, a rotation shaft, a bearing, a bearing holder, and an end elastic body. The rotation shaft is located in the bearing housing. The bearing rotatably supports the rotation shaft at least in a radial direction of the rotation shaft. The bearing holder faces one end of the bearing and is fixed to the bearing housing. The end elastic body is disposed between the one end of the bearing and the bearing holder in the axial direction of the bearing and is in contact with the bearing and the bearing holder. The end elastic body is formed of a material having a lower modulus of elasticity than a material forming the bearing holder. 1. A turbo machine comprising:a cylindrical bearing housing;a rotation shaft that is located in the bearing housing;a bearing that is disposed between an inner surface of the bearing housing and an outer surface of the rotation shaft, that rotatably supports the rotation shaft at least in a radial direction of the rotation shaft, and that has one end in an axial direction of the bearing, the one end being located in the bearing housing;a bearing holder that is fixed to one end of the bearing housing in the axial direction of the bearing and that faces the one end of the bearing; andan end elastic body that is disposed between the one end of the bearing and the bearing holder in the axial direction of the bearing, that is in contact with the bearing and the bearing holder, and that is formed of a material having a lower modulus of elasticity than a material forming the bearing holder.2. The turbo machine according to claim 1 , wherein at least one of the one end of the bearing and a surface of the holder that is in contact with the end elastic body has a front recess that houses a portion of the end elastic body.3. The turbo machine according to claim 1 , further comprising at least one side elastic body that is disposed in a ring-shaped space formed between the inner ...

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12-01-2017 дата публикации

NACELLE ANTI-ICE SYSTEM AND METHOD WITH EQUALIZED FLOW

Номер: US20170009653A1
Принадлежит:

A gas turbine engine is provided having a nacelle and a compressor section constructed and arranged to generate hot air. An anti-icing system is constructed and arranged to discharge the hot air from the compressor section to the nacelle. An anti-icing valve is positioned in the anti-icing system and constructed and arranged to control a flow of the hot air from the compressor section to the nacelle. The anti-icing valve includes a partially open position to constrict a flow of the hot air from the compressor section to the nacelle. 1. A gas turbine engine comprising:a nacelle;a compressor section constructed and arranged to generate hot air;an anti-icing system constructed and arranged to discharge the hot air from the compressor section to the nacelle; andan anti-icing valve positioned in the anti-icing system and constructed and arranged to control a flow of the hot air from the compressor section to the nacelle, the anti-icing valve including a partially open position to constrict a flow of the hot air from the compressor section to the nacelle.2. The gas turbine engine of claim 1 , wherein the anti-icing valve includes a locking mechanism operative to lock the anti-icing valve in the partially open position.3. The gas turbine engine of claim 2 , wherein the locking mechanism is operative to lock the anti-icing valve in a ¾ open position.4. The gas turbine engine of further comprising a control operatively coupled to the anti-icing valve and operative to open and close the anti-icing valve.5. The gas turbine engine of further comprising a conduit constructed and arranged to channel the hot air from the compressor section to the nacelle.6. The gas turbine engine of further comprising a nozzle positioned adjacent the nacelle and constructed and arranged to discharge the hot air onto the nacelle.7. The gas turbine engine of further comprising a bleed valve constructed and arranged to bleed excess hot air from the compressor section.8. An anti-icing system for a gas ...

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14-01-2016 дата публикации

METHOD FOR DETUNING A ROTOR-BLADE CASCADE

Номер: US20160010461A1
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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14-01-2016 дата публикации

TURBOMACHINE BLADE

Номер: US20160010462A1
Принадлежит:

A turbomachine blade having a main body () that includes a plane, first attachment surface () having a first rim contour (), a cover () that has a plane, second attachment surface () having a second rim contour () that is welded to the first attachment surface, and a tuning body configuration having at least one tuning body (-) for contacting an inner wall () of the cover by impact therewith, a gap (s) being formed between the first and the second rim contour. 1. A turbomachine blade comprising:a main body including a plane, first attachment surface having a first rim contour; anda cover including a plane, second attachment surface having a second rim contour and welded to the first attachment surface; anda tuning body configuration having at least one tuning body for contacting an inner wall of the cover by impact therewith, a gap being formed between the first and second rim contour.2. The turbomachine blade as recited in wherein the cover is disposed on the main body in a way that allows the cover to extend freely therefrom.3. The turbomachine blade as recited in wherein at least one tuning body of the tuning body configuration is accommodated in a cavity completely formed in the cover.4. The turbomachine blade as recited in wherein the cover is a multipart cover.5. The turbomachine blade as recited in wherein at least one tuning body of the tuning body configuration is accommodated in a cavity formed at least in sections thereof in the main body.6. The turbomachine blade as recited in further comprising at least two cavities having a different shape or different volume; at least one tuning body of the tuning body configuration being accommodated in at least one of these cavities.7. The turbomachine blade as recited in wherein at least one tuning body of the tuning body configuration is accommodated in a cavity; a circumferentially extending groove being formed in the first or second attachment surface between the cavity and the second rim contour.8. The ...

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14-01-2016 дата публикации

INTERMEDIATE CASING FOR A TURBOFAN ENGINE

Номер: US20160010501A1
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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14-01-2016 дата публикации

NACELLE COMPRESSION RODS

Номер: US20160010502A1
Автор: Byrne Stuart J.
Принадлежит:

A compression rod may include a plunger and a spring. A proximal end and a distal end of the compression rod may contact engagement features in a core cowl of a gas turbine engine. The compression rod may transmit loads between halves of the core cowl. The spring may cause the plunger to extend and contract in response to vibrations or other relative movement between halves of the core cowl. 1. An aircraft nacelle comprising:a first half comprising a first engagement feature, and a second half comprising a second engagement feature, wherein the first half and the second half are rotatable about a hinge between a closed position in which the first half and the second half enclose a portion of an aircraft engine, and an open position in which the first half and the second half are separated and allow access to the portion of the aircraft engine; and a proximal end and a distal end, wherein in response to the aircraft nacelle being in the closed position, the proximal end is in contact with the first engagement feature and the distal end is in contact with the second engagement feature; and', 'a spring configured to bias the proximal end apart from the distal end, wherein, in response to the aircraft nacelle being in the closed position, the spring is in a compressed position., 'a compression rod comprising2. The aircraft nacelle of claim 1 , wherein the compression rod comprises a plunger.3. The aircraft nacelle of claim 1 , wherein the compression rod comprises a compression tube.4. The aircraft nacelle of claim 3 , wherein the compression tube comprises a plunger bore.5. The aircraft nacelle of claim 4 , wherein the spring is located within the plunger bore.6. The aircraft nacelle of claim 3 , wherein the compression tube comprises a threaded bore.7. The aircraft nacelle of claim 6 , wherein the proximal end comprises a threaded shaft located within the threaded bore.8. The aircraft nacelle of claim further comprising a pylon bracket claim 6 , wherein the ...

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11-01-2018 дата публикации

Low energy wake stage

Номер: US20180010459A1
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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11-01-2018 дата публикации

TUNED MASS DAMPER FOR TUBES

Номер: US20180010482A1
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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11-01-2018 дата публикации

SYSTEM OF OPERATING A GAS TURBINE ENGINE

Номер: US20180010527A1
Автор: ROWE Arthur L.
Принадлежит: ROLLS-ROYCE PLC

A system for operating a gas turbine engine to mitigate the risk of ice formation within the engine, the system including a controller arranged to control at least one operational parameter of the engine such that the engine operates in a safe zone; and, a processor configured to function as a determining module to make a comparison between values and determine whether the engine is operating within a safe zone based on at least a core pressure parameter relating to the pressure within the engine and a core temperature parameter relating to the temperature within the engine, wherein the safe zone is defined by the product (multiplied) of the core pressure parameter and core temperature parameter being above a safe threshold. 1. A system for operating a gas turbine engine to mitigate the risk of ice formation within the gas turbine engine , the system comprising:a controller configured to control at least one operational parameter of the gas turbine engine such that the gas turbine engine operates in a safe zone; and,a processor configured to function as a determining module to make a comparison between values and determine whether the gas turbine engine is operating within the safe zone based on at least a core pressure parameter relating to the pressure within the gas turbine engine and a core temperature parameter relating to the temperature within the gas turbine engine,wherein the safe zone is defined by the product (multiplied) of the core pressure parameter and core temperature parameter being above a safe threshold.2. A system for operating a gas turbine engine according to claim 1 , wherein the core pressure parameter relates to the static pressure within the gas turbine engine.3. A system for operating a gas turbine engine according to claim 1 , wherein the core temperature parameter relates to the stagnation temperature within the gas turbine engine.4. A system for operating a gas turbine engine according to claim 1 , wherein the core pressure parameter is ...

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14-01-2021 дата публикации

MISTUNING OF TURBINE BLADES WITH ONE OR MORE INTERNAL CAVITIES

Номер: US20210010375A1
Принадлежит:

A bladed rotor system includes first and second sets of blades with respective airfoils each having at least one internal cavity. The airfoils of both the first and second sets of blades have identical outer shapes defined by an outer surface of an outer wall of the respective airfoils. The airfoils of the first set of blades are distinguished from the airfoils of the second set of blades by a geometry and/or position of at the least one internal cavity, which is unique to blades of a given set. The natural frequency of a blade of the first set differs from the natural frequency of a blade of the second set by a predetermined amount. The blades of the first set and the second set are alternately arranged in a periodic fashion in said circumferential row, to provide a frequency mistuning to stabilize flutter of the blades. 1. A bladed rotor system for a turbomachine , comprising:a circumferential row of blades mounted on a rotor disc, each blade comprising an airfoil having an outer wall delimiting an airfoil interior, the airfoil interior comprising one or more internal cavities, the airfoils of both the first and second sets of blades have identical outer shapes defined by an outer surface of the outer wall of the respective airfoils, and', 'the airfoils of the first set of blades are distinguished from the airfoils of the second set of blades by a geometry and/or position of at least one internal cavity, which is unique to blades of a given set,, 'the row of blades comprising a first set of blades and a second set of blades, whereinwhereby the natural frequency of a blade of the first set differs from the natural frequency of a blade of the second set by a predetermined amount, andwherein blades of the first set and the second set are alternately arranged in a periodic fashion in said circumferential row, to provide a frequency mistuning to stabilize flutter of the blades.2. The bladed rotor system according to claim 1 , wherein an outer wall thickness of the ...

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14-01-2021 дата публикации

DAMPING DEVICE

Номер: US20210010391A1
Принадлежит: SAFRAN AIRCRAFT ENGINES

The invention relates to an assembly () for a turbomachine comprising: 1. A turbomachine comprising:a first rotor module comprising a first blade, the first blade having a first length;a second rotor module connected to the first rotor module and comprising a second blade, the second blade having a second length, the second length being smaller than the first length; anda damping device extending with at least one component along a turbomachine longitudinal axis, the damping device being annular while extending circumferentially around the turbomachine longitudinal axis, the damping device comprising a first radial external surface supported with friction against the first rotor module, as well as a second radial external surface supported with friction against the second rotor module, so as to couple the first rotor module with the second rotor module in order to damp vibrational movements of the first rotor module relative to the second rotor module during operation.2. The assembly of claim 1 , wherein the damping device is an annular tab claim 1 , a cross section of the damping device being shaped like a V claim 1 , a first external surface of a first branch of the damping device forming the first radial external surface claim 1 , a second external surface of a second branch of the damping device forming the second radial external surface.3. The assembly of claim 1 , wherein:the first rotor module comprises a disk centered on the turbomachine longitudinal axis;the first blade is mounted on an external periphery of the disk, the first blade thus extending from the external periphery of the disk, the first blade further comprising an airfoil, a platform, a support and a root, the root being embedded in a housing of the disk, the first radial external surface being supported with friction on a radially internal surface of the platform; andthe second rotor module comprises a ferrule, the ferrule comprising a circumferential extension extending toward the platform of ...

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10-01-2019 дата публикации

BLADE-DISC ARRANGEMENT FOR A TURBOMACHINE

Номер: US20190010825A1
Автор: Kloetzer Alexander
Принадлежит: MTU Aero Engines AG

The present invention relates to a blade-disc arrangement for a turbomachine having a disc and a plurality of grooves that are arranged in a casing surface of the disc for taking up blades, as well as a plurality of blades with blade roots that are accommodated in the grooves, wherein at least one heat shield element is arranged between at least one blade root and a groove surface of at least one groove, so that there is no direct contact between blade root and disc in the groove region, wherein the heat shield element is formed of a ceramic material, at least in part. 1. A blade-disc arrangement for a turbomachine , comprising:a disc and a plurality of grooves that are arranged in a casing surface of the disc for taking up blades, and a plurality of blades with blade roots that are accommodated in the grooves, whereinat least one heat shield element is arranged between at least one blade root and a groove surface of at least one groove, so that there is no direct contact between blade root and disc in a groove region, wherein the heat shield element is formed, at least in part, of a ceramic material that is zirconium oxide or is based on zirconium oxide.2. The blade-disc arrangement according to claim 1 , wherein the entire heat shield element is formed of a ceramic material.3. The blade-disc arrangement according to claim 1 , wherein the heat shield element is configured and arranged as a flat surface element extending along the groove or comprises a flat surface element claim 1 , which has a length along the lengthwise direction of the groove and a width crosswise to the lengthwise direction of the groove along the blade root or the groove surface claim 1 , and a thickness crosswise to the length and width which is smaller than the length and width of the flat surface element claim 1 , and wherein the flat surface element is adjacent to a surface of the blade root and/or of the groove surface or has any desired form between the groove surface and the surface of ...

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14-01-2021 дата публикации

GAS TURBINE IMPELLER NOSE CONE

Номер: US20210010421A1
Принадлежит:

A gas turbine engine is provided, with a compressor section including a fan driven by an engine shaft about a rotation axis, the engine shaft defining a bore extending axially therethrough from a hot gas inlet to a hot gas outlet, the hot gas inlet located downstream of the compressor section, a nose cone mounted to the fan and defining a cavity therewithin, a first impeller and a second impeller mounted within the cavity, the first impeller having an inlet facing a forward direction of the gas turbine engine, the inlet of the first impeller in fluid flow communication with an air inlet defined in an outer surface of the nose cone, the second impeller having an inlet facing a rearward direction of the gas turbine engine, the inlet of the second impeller being in fluid flow communication with the hot gas outlet of the engine shaft.

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09-01-2020 дата публикации

INLET – NAI EXHAUST HOLE DEFINITION FOR REDUCED D-DUCT RESONANCE NOISE AND DILUTED EXHAUST PLUME FOR THERMAL CONTROL

Номер: US20200011244A1
Принадлежит: Rohr, Inc.

An inlet for use with a nacelle having an axis includes an outer barrel. The inlet further includes a lip skin defining a plurality of elongated exit holes including a first circumferential outer hole, a second circumferential outer hole, and a plurality of center holes located between the first circumferential outer hole and the second circumferential outer hole, the first circumferential outer hole being located at least 10 degrees of an entire circumference of the inlet away from the second circumferential outer hole. 1. An inlet for use with a nacelle having an axis , comprising:an outer barrel; anda lip skin defining a plurality of elongated exit holes including a first circumferential outer hole, a second circumferential outer hole, and a plurality of center holes located between the first circumferential outer hole and the second circumferential outer hole, the first circumferential outer hole being located at least 10 degrees of an entire circumference of the inlet away from the second circumferential outer hole.2. The inlet of claim 1 , wherein the first circumferential outer hole is located at least 15 degrees of the entire circumference of the inlet away from the second circumferential outer hole.3. The inlet of claim 1 , wherein the plurality of elongated exit holes face radially outward.4. The inlet of claim 1 , wherein each of the plurality of elongated exit holes has a first dimension measured in a direction parallel to the axis and a circumferential dimension measured in a circumferential direction of the inlet.5. The inlet of claim 4 , wherein the axial dimension is at least three times the size of the circumferential dimension.6. The inlet of claim 1 , wherein each of the plurality of elongated exit holes has a rounded claim 1 , elongated shape.7. The inlet of claim 1 , wherein the plurality of elongated exit holes are non-uniformly distributed about a portion of the lip skin.8. The inlet of claim 1 , wherein each of the plurality of elongated exit ...

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09-01-2020 дата публикации

SEGREGATED ANTI-ICE DUCT CHAMBER

Номер: US20200011245A1
Принадлежит: Rohr, Inc.

The present disclosure provides an inlet assembly for a nacelle. The inlet assembly may comprise an inlet, a lip skin coupled to the inlet, the lip skin and the inlet forming a duct, and an inner duct skin situated between the inlet and the lip skin, the inner duct skin separating the duct by defining a first chamber and a second chamber. 1. An inlet assembly for a nacelle , comprising:an inlet;a lip skin coupled to the inlet, the lip skin and the inlet forming a duct; andan inner duct skin situated between the inlet and the lip skin, the inner duct skin separating the duct by defining a first chamber and a second chamber.2. The inlet assembly of claim 1 , wherein the inlet comprises an inner barrel claim 1 , an outer barrel claim 1 , and a bulkhead surface extending radially between the inner barrel and the outer barrel.3. The inlet assembly of claim 1 , wherein the lip skin comprises a first end claim 1 , a second end claim 1 , an interior surface claim 1 , and an exterior surface claim 1 , the interior surface coupled to an inner barrel and outer barrel of the inlet.4. The inlet assembly of claim 1 , wherein the first chamber is defined between an interior surface of the lip skin and an outer surface of the inner duct skin.5. The inlet assembly of claim 1 , wherein the second chamber is defined between an exterior surface of the inlet and an inner surface of the inner duct skin.6. The inlet assembly of claim 1 , further comprising a hot bleed air source extending through the inlet and the inner duct skin claim 1 , the hot bleed air source terminating in the first chamber.7. The inlet assembly of claim 5 , wherein a hot bleed air source is configured to supply bleed air to the first chamber and prevent formation of ice on at least a portion of the exterior surface of the lip skin.8. The inlet assembly of claim 2 , wherein an interior surface of the inner duct skin is coupled to the inner barrel and wherein an exterior surface of the inner duct skin is coupled to ...

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15-01-2015 дата публикации

PLANT AND METHOD FOR DAMPING ACOUSTIC VIBRATIONS IN A CORRESPONDING PLANT

Номер: US20150016951A1
Принадлежит: SIEMENS AKTIENGESELLSCHAFT

A facility, in particular a power plant, is provided having a steam turbine and a bypass station for diverting a working medium, as required, for the steam turbine around the steam turbine, wherein at least one resonance absorber is provided for the bypass station. 1. A plant , comprisinga steam turbine and a bypass station for, if necessary, diverting a working medium for the steam turbine around the steam turbine,wherein at least one resonance absorber is provided for the bypass station, andwhereinthe resonance absorber is embodied as a Helmholtz resonator.2. The plant as claimed in claim 1 ,whereinthe bypass station comprises a pipeline, andwherein the resonance absorber is formed substantially by a chamber at least partially encircling the pipeline, said chamber being connected in a sound-conducting manner to the pipeline via a plurality of through-openings that are distributed around the circumference of the pipeline.3. The plant as claimed in claim 1 ,whereinthe bypass station comprises a pipeline, andwherein the resonance absorber is formed substantially by a chamber positioned next to the pipeline, said chamber being connected in a sound-conducting manner to the pipeline via a resonator neck.4. The plant as claimed in claim 1 ,whereinthe Helmholtz resonator is embodied as a controllable Helmholtz resonator, in the case of which the resonance frequency is settable.5. The plant as claimed in claim 1 ,whereina plurality of resonance absorbers are provided to damp in each case a narrow frequency band.6. The plant as claimed in claim 1 ,whereinthe resonance absorber is positioned between a cooling medium injection and a condenser.7. The plant s claimed in claim 1 ,whereinthe resonance absorber has a resonance body, andwherein a temperature-control plant is provided for the resonance body, said temperature-control plant being used to set a substantially uniform temperature for the entire resonance body.8. The plant as claimed in claim 7 ,whereinthe resonance body ...

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03-02-2022 дата публикации

TURBOMACHINE ROTOR

Номер: US20220034228A1
Принадлежит:

The invention concerns a turbomachine rotor (), characterised in that it comprises a threaded or tapped part () and a damping nut () screwed onto the threaded or tapped part () so as to allow the threads of the nut () and of the threaded or tapped part () to rub against each other in the event of vibration of the rotor (). 110.-. (canceled)1113683683618105381053534. A turbomachine rotor () , characterised in that it comprises a threaded or tapped part ( , ) and a damping nut () screwed onto the threaded or tapped part ( , ) so as to allow the threads of the nut () and of the threaded or tapped part ( , ) to rub against one another in the event of vibration of the rotor () , the nut () comprising a stop () capable of bearing axially on a complementary stop () of the threaded or tapped part () , the stop of the nut () being formed by a radially extending flange () , the stop () of the threaded part () being formed by an end surface () or a radial shoulder , the threaded or tapped part () comprising an end area comprising teeth ().121984. A rotor () according to claim 11 , characterized in that the threads () of the nut () are located axially opposite the external teeth ().131836. A rotor () according to claim 11 , characterized in that the number of threads of the nut () engaging the threaded or tapped part ( claim 11 , ) is at least five.141836. A rotor () according to claim 12 , characterized in that the number of threads of the nut () engaging the threaded or tapped part ( claim 12 , ) is at least five.151128361. A rotor () according to claim 11 , characterized in that an elastically deformable seal () is mounted between the nut () and the threaded or tapped part ( claim 11 , ) of the rotor ().161128361. A rotor () according to claim 12 , characterized in that an elastically deformable seal () is mounted between the nut () and the threaded or tapped part ( claim 12 , ) of the rotor ().171128361. A rotor () according to claim 13 , characterized in that an ...

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03-02-2022 дата публикации

BLADED ROTOR SYSTEM AND CORRESPONDING METHOD OF SERVICING

Номер: US20220034229A1
Автор: ZHOU Yuekun
Принадлежит:

A bladed rotor system for a turbomachine includes a circumferential row of blades mounted on a rotor disc, and includes a plurality of under-platform dampers. Each damper is located between adjacent blade platforms. The plurality of dampers includes a first set of dampers and a second set of dampers. The dampers of the first set are distinguished from the dampers of the second set by a cross-sectional material distribution in the damper that is unique to the respective set. Dampers of the first set and the second set are positioned alternately in a periodic fashion in a circumferential direction, to provide a frequency mistuning to stabilize flutter of the blades. 1. A bladed rotor system for a turbomachine , comprising: a platform;', 'a root extending radially inward from the platform for mounting the blade to the rotor disc; and', 'an airfoil extending span-wise radially outward from the platform;, 'a circumferential row of blades mounted on a rotor disc, each blade comprisingwherein platforms of adjacent blades align circumferentially to define an inner diameter boundary for a working fluid flow path; anda plurality of dampers, each damper being located between adjacent platforms;wherein the plurality of dampers comprise a first set of dampers and a second set of dampers, wherein the dampers of the first set are distinguished from the dampers of the second set by a cross-sectional material distribution in the damper that is unique to the respective set, andwherein dampers of the first set and the second set are positioned alternately in a periodic fashion in a circumferential direction, to provide a frequency mistuning to stabilize flutter of the blades.2. The bladed rotor system according to claim 1 ,wherein the dampers of the first set are solid,and the dampers of the second set are hollow, each defining an internal cavity therewithin.3. The bladed rotor system according to claim 2 , wherein the dampers of the first set and the dampers of the second set are ...

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19-01-2017 дата публикации

LATERALLY REINFORCED VARIABLE PITCH ROTOR

Номер: US20170016337A1
Принадлежит:

A propulsion unit having reduced noise employs a rotor having a plurality of variable pitch blades. A tensioning joint mechanism is carried within each blade and a tension element is engaged by the tension joint mechanism. The tension element is configured to maintain tension between the blades in the rotor and the tensioning joint mechanism adapted to allow variation of pitch of the blades without altering tension in the tension element. 1. A propulsion unit having reduced noise comprising:a rotor having a plurality of variable pitch blades;a tensioning joint mechanism carried within each blade;a tension element engaged by the tension joint mechanism in each blade, said tension element configured to maintain tension between the blades in the rotor and said tensioning joint mechanism adapted to allow variation of pitch of the blades without altering tension in the tension element.2. The propulsion unit as defined in wherein the tension element has an airfoil shape.3. The propulsion unit as defined in wherein the tension element has a plurality of tension segments and an airfoil shaped fairing covering the tension segments.4. The propulsion unit as defined in wherein each blade has a cavity formed therein and the tension joint mechanism is installed in the cavity.5. The propulsion unit as defined in wherein the cavity has an elliptical shape.6. The propulsion unit as defined in wherein the tensioning joint mechanism comprises:a support bearing mounted to allow rotation of the blade relative to an axis for pitch change;a connector extending from the bearing and fixed relative to pitch change of the blade, said tension element attached to said connector.7. The propulsion unit as defined in wherein the tension element incorporates a tension segment and the tension segment is engaged by the connector with a clamping bushing.8. The propulsion unit as defined in wherein the support bearing is a ball bearing.9. A reduced noise aerodynamic rotor comprising:a plurality of ...

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19-01-2017 дата публикации

SPLITTER NOSE OF A LOW-PRESSURE COMPRESSOR OF AN AXIAL TURBOMACHINE WITH ANNULAR DEICING CONDUIT

Номер: US20170016347A1
Принадлежит:

The invention comprises a splitter nose of an axial turbomachine, in particular a compressor, the splitter nose comprising: an annular casing an annular conduit for de-icing a separation edge of the splitter nose; the conduit is connected to the casing only in a first zone in the region of a hot air inlet and in a second zone located in a position diametrically opposite the inlet, or forming relative to the axis of the turbomachine an angle α less than 30° with respect to the position so as to allow expansion deformations of the conduit. The invention also comprises a compressor and a turbomachine comprising such a splitter nose. 1. A splitter nose of an axial turbomachine , said splitter nose comprising:an annular casing that forms an annular cavity and a circular separation edge of an air flow of the turbomachine; andan annular conduit that is arranged in the annular cavity, the annular conduit being configured to de-ice the separation edge by circulation of hot air in the cavity, and the conduit comprises an air inlet that is structured and operable to be connected to a hot air supply pipe of the turbomachine, the air inlet forming a first zone; whereinthe conduit is connected to the casing only in a second zone, diametrically opposite the air inlet, and in the region of the air inlet, the second zone forming an angular portion of the annular conduit that is less than 60° so as to allow expansion deformations of the conduit.2. The splitter nose of claim 1 , wherein the second zone forms an angular portion of the annular conduit that is at most 30°.3. The splitter nose of claim 1 , wherein the annular casing comprises an internal surface that delimits the cavity claim 1 , that is free from fixation in contact with the annular conduit over at least 120°.4. The splitter nose of further comprising a flange joining the second zone of the annular conduit to the annular casing claim 1 , the flange extending radially and axially.5. The splitter nose of claim 1 , wherein ...

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21-01-2016 дата публикации

COOLING DEVICE FOR A TURBOJET ENGINE OF AN AIRCRAFT NACELLE

Номер: US20160017751A1
Автор: Caruel Pierre
Принадлежит: AIRCELLE

The present disclosure provides a cooling device for a turbo engine of an aircraft nacelle, including: a heat exchanger and an air outlet pipe. The nacelle includes a front housing that has a front lip forming a hollow leading edge that delimits an annular de-icing chamber. In particular, the cooling device further includes a pipe for supplying pressurized air that extends from an inlet end linked to a pressurized air source, to an outlet end forming an air ejection nozzle opening into the de-icing chamber, and the outlet pipe of the heat exchanger has an air outlet section that is arranged in the de-icing chamber in a position designed such that the pressurized air ejection nozzle forms an air suction pump in the outlet pipe of the exchanger.

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17-01-2019 дата публикации

PNEUMATIC DEICER WITH SENSORS

Номер: US20190016467A1
Принадлежит:

A pneumatic deicer includes a base layer, a forming layer, a first chamber, and a first sensor. The base layer has an inlet, a first side, and a second side. The forming layer is connected to the base layer along at least two seams and has inner side and an outer side with the outer side being distant from the base layer. The first chamber is formed between the base layer and the forming layer and configured to be inflated by air passing into the first chamber through the inlet in the base layer. The first sensor is situated within the first chamber. 1. A pneumatic deicer comprising:a base layer having an inlet, a first side, and a second side;a forming layer connected to the base layer along at least two seams, the forming layer having an inner side and an outer side with the outer side being distant from the base layer;a first chamber formed between the base layer and the forming layer and configured to be inflated by air passing into the first chamber through the inlet in the base layer; anda first sensor situated within the first chamber.2. The pneumatic deicer of claim 1 , wherein the first sensor is located on the first side of the base layer.3. The pneumatic deicer of claim 1 , wherein the first sensor is embedded in the first side of the base layer.4. The pneumatic deicer of claim 1 , wherein the first sensor is located on the inner side of the forming layer.5. The pneumatic deicer of claim 1 , wherein the first sensor is embedded in the inner side of the forming layer.6. The pneumatic deicer of claim 5 , further comprising:a second sensor embedded in the first side of the base layer.7. The pneumatic deicer of claim 1 , further comprising:an air line extending from an air source to the inlet in the base layer to provide air to inflate the forming layer to increase a volume of the first chamber; anda second sensor in the air line.8. The pneumatic deicer of claim 1 , wherein the first sensor measures a displacement of the forming layer relative to the base ...

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21-01-2016 дата публикации

GAS TURBINE ENGINE DE-ICING SYSTEM

Номер: US20160017803A1
Принадлежит:

A de-icing system for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a forward assembly and a rear assembly adjacent to the forward assembly. One of the forward assembly and the rear assembly is rotatable relative to the other to generate an amount of air friction between said forward and rear assemblies.

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17-04-2014 дата публикации

EXHAUST HEAT RECOVERY FOR A GAS TURBINE SYSTEM

Номер: US20140102113A1
Принадлежит: GENERAL ELECTRIC COMPANY

A system includes a gas turbine and an anti-icing system coupled to the gas turbine. The gas turbine is configured to receive air and fuel and to combust a mixture of the air and the fuel into exhaust gases. The anti-icing system is configured to use heat from the exhaust gases to heat a heat transfer fluid (HTF) and to selectively heat the fuel and the air via the HTF. 1. A system , comprising:a gas turbine configured to receive air and fuel and to combust a mixture of the air and the fuel into exhaust gases; andan anti-icing system coupled to the gas turbine and configured to use heat from the exhaust gases to heat a heat transfer fluid (HTF) and to selectively heat the fuel and the air via the HTF.2. The system of claim 1 , wherein the anti-icing system comprises:an exhaust heat exchanger disposed downstream of the gas turbine along an exhaust gas flow path and configured to selectively heat the HTF using the exhaust gases;an air heat exchanger disposed upstream of the gas turbine along an air flow path and configured to selectively heat the air using the HTF; anda fuel heat exchanger disposed upstream of the gas turbine along a fuel flow path and configured to selectively heat the fuel using the HTF.3. The system of claim 2 , wherein the anti-icing system comprises a first loop having a first HTF flow path claim 2 , wherein the HTF along the first HTF flow path is configured to bypass the fuel and air heat exchangers and to exchange heat with the exhaust gases to increase a temperature of the HTF above an HTF temperature setpoint.4. The system of claim 3 , wherein the HTF temperature setpoint is between approximately 15 and 76 degrees Celsius.5. The system of claim 3 , wherein the first loop comprises:a pump disposed along the first HTF flow path and configured to pump the HTF between a skid and the exhaust heat exchanger;a control valve disposed downstream of the pump, wherein the control valve is configured to throttle a flow rate of the HTF;a flow meter ...

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17-01-2019 дата публикации

UNSHROUDED TURBOMACHINE IMPELLER WITH IMPROVED RIGIDITY

Номер: US20190017393A1
Принадлежит:

An unshrouded turbomachine impeller is disclosed. The impeller comprises a hub and a plurality of sequentially arranged blades. Each blade extends from a blade root at the hub to a blade tip and is comprised of a first blade edge and a second blade edge. A flow vane is formed between each pair of neighboring blades. A connection member extends across each flow vane between neighboring blades and rigidly or monolithically connects a first modal displacement region of a first one of the pair of neighboring blades to a second modal displacement region of a second one of the pair of neighboring blades. 1. An unshrouded turbomachine impeller comprising:a rotation axis;a hub;a plurality of sequentially arranged blades, each blade extending from a blade root at the hub to a blade tip and comprised of a first blade edge and a second blade edge, the first blade edge and the second blade edge extending from the hub to the blade tip; anda flow vane arranged between each pair of neighboring blades;wherein a connection member extending across each flow vane between pairs of neighboring blades connects a first modal displacement region at a certain frequency of a first one of said pair of neighboring blades to a second modal displacement region at said frequency of a second one of said pair of neighboring blades; andwherein each connection member has a first end rigidly or monolithically connected to a pressure side of a first one of said pair of neighboring blades and a second end rigidly or monolithically connected to a suction side of a second one of said pair of neighboring second blades, or vice versa.2. The turbomachine impeller of claim 1 , wherein the first blade edge is located at a first radial distance from the rotation axis and the second blade edge is located at a second radial distance from the rotation axis the first radial distance being smaller than the second radial distance.3. The turbomachine impeller of claim 1 , wherein the first modal displacement region is ...

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17-01-2019 дата публикации

FLOW METERING AND RETENTION SYSTEM

Номер: US20190017412A1
Принадлежит:

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

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16-01-2020 дата публикации

COMPRESSOR ROTOR WITH COATED BLADES

Номер: US20200018176A1
Принадлежит:

A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades. 1. A compressor rotor for a gas turbine engine , the compressor rotor comprising blades extending a span length from a central hub , the blades including circumferentially alternating first and second blades having airfoils with corresponding geometric profiles , each of the airfoils including a leading edge , a trailing edge , a root , a tip and a mid-span region between the root and the tip along the span , the airfoil of the first blades having a coating on a first portion of the first blade adjacent the root with a root coating thickness , and the coating being provided on a second portion adjacent the tip of the first blade with a tip coating thickness , the root coating thickness being greater than the tip coating thickness , the coating defining a first coating structure providing the first blade with a first natural vibration frequency different from a second natural vibration frequency of the second blade.2. The compressor rotor as defined in claim 1 , wherein the coating is provided on the mid-span region of the airfoil of the first blades.3. The compressor rotor as defined in claim 2 , wherein the coating on the mid-span region has a mid-span coating thickness claim 2 , the root coating thickness being greater than the mid-span coating thickness.4. The compressor rotor as defined in claim 1 , wherein the airfoil of the second blade is free of coating.5. The compressor rotor as defined in claim 1 , wherein the airfoil of the second blade has a coating defining a second coating structure claim 1 , the coating being provided on ...

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22-01-2015 дата публикации

STARTING DEVICE

Номер: US20150023781A1
Принадлежит: AISIN AW CO., LTD.

A starting device includes a pump impeller, a turbine runner for rotating together with the pump impeller, a damper mechanism having an input element receiving power from an internal combustion engine, an output element coupled to a speed change device, an intermediate element between the input and output elements, and a dynamic damper for damping vibration at a predetermined frequency among vibration transferred to the speed change device. The starting device includes a first dynamic damper having an elastic member and a first mass body coupled the first elastic member, and coupled to the intermediate element; and a second dynamic damper having an elastic member and a second mass body connected to the second elastic member, and coupled to the intermediate element. The first mass body of the first dynamic damper or the second mass body of the second dynamic damper includes at least the turbine runner. 1. A starting device that includes a pump impeller , a turbine runner capable of rotating together with the pump impeller , a damper mechanism that has an input element to which power is input from an internal combustion engine , an output element coupled to a speed change device , and an intermediate element disposed between the input element and the output element , and a dynamic damper that damps vibration at a predetermined frequency among vibration transferred to the speed change device , the starting device comprising:a first dynamic damper including a first elastic member and a first mass body coupled to one end of the first elastic member, and coupled to the intermediate element; anda second dynamic damper including a second elastic member and a second mass body coupled to one end of the second elastic member, and coupled to the intermediate element, characterized in thatone of the first mass body of the first dynamic damper and the second mass body of the second dynamic damper includes at least the turbine runner.2. The starting device according to claim 1 , ...

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28-01-2016 дата публикации

Fan Blade Damping Device

Номер: US20160024940A1
Автор: Wilber John E.
Принадлежит:

An airfoil for a gas turbine engine and method of manufacture of the airfoil are disclosed. The airfoil may comprise a first side extending axially from a leading edge to a trailing edge and extending radially from a base to a tip, a second side opposite to the first side, a pocket disposed in the first side, a filler disposed in the pocket, and a preloaded spring disposed within the filler. 1. An airfoil for a gas turbine engine , comprising:a first side extending axially from a leading edge to a trailing edge and extending radially from a base to a tip;a second side opposite to the first side;a pocket disposed in the first side;a filler disposed in the pocket; anda preloaded spring disposed within the filler.2. The airfoil of claim 1 , wherein the preloaded spring exerts a force within the pocket for resisting flutter of the airfoil and for damping vibratory response of the airfoil.3. The airfoil of claim 1 , wherein the filler surrounds and encapsulates the preloaded spring claim 1 , and is formed to fill the pocket.4. The airfoil of claim 3 , wherein the filler dampens a vibratory response of the airfoil claim 3 , prevents decompression of the preloaded spring claim 3 , imparts force exerted from the preloaded spring to the airfoil claim 3 , and prevents contact of the preloaded spring with the pocket.5. The airfoil of claim 1 , wherein the filler is bonded to the pocket of the first side claim 1 , and the second side is bonded to the first side and the filler.6. The airfoil of claim 1 , wherein the preloaded spring and the filler are configured to change an inherent frequency of the airfoil to a predetermined frequency outside a range of resonant frequencies of the airfoil.7. The airfoil of claim 6 , wherein at least one of a composition of the filler claim 6 , thickness of the filler claim 6 , durometer of the filler claim 6 , thickness of the preloaded spring claim 6 , preload of the preloaded spring claim 6 , stiffness of the preloaded spring claim 6 , and ...

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28-01-2016 дата публикации

GAS TURBINE AND METHOD FOR OPERATING THE GAS TURBINE

Номер: US20160024960A1
Принадлежит: SIEMENS AKTIENGESELLSCHAFT

A gas turbine includes an intake tract and a compressor having a compressor flow channel. The compressor further includes an inlet guide vane row positioned in the compressor flow channel having inlet guide vanes that can be adjusted. The gas turbine has an icing sensor unit having at least one sensor arranged between a first compressor blade row and a first compressor guide vane row. The first compressor blade row is thereby arranged in the compressor flow channel directly downstream of the inlet guide vane row, and the first compressor guide vane row is arranged directly downstream of the first compressor blade row. A method detects an imminent icing of the compressor, and the compressor is safeguarded therefrom such that at least inlet guide vanes of the inlet guide vane row are adjusted such that the acceleration of an intake air mass flow is reduced. 1. A gas turbine comprising an intake section and a compressor having a compressor flow duct , wherein the compressor has an inlet guide blade row with adjustable inlet guide blades , which is positioned in the compressor flow duct , the gas turbine comprising:an icing sensor unit comprising at least one sensor arranged between a first compressor rotor blade row and a first compressor guide blade row, wherein the first compressor rotor blade row is arranged immediately downstream of the inlet guide blade row and the first compressor guide blade row is arranged immediately downstream of the first compressor rotor blade row in the compressor flow duct,wherein the icing sensor unit has at least one air humidity sensor and, in addition, at least one pressure sensor and one temperature sensor, which are both arranged between the first compressor rotor blade row and the first compressor guide blade row.2. The gas turbine as claimed in claim 1 ,wherein the icing sensor unit has at least one air humidity sensor and at least one pressure sensor and one temperature sensor, which are arranged between the first compressor ...

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26-01-2017 дата публикации

NOZZLE GUIDE VANE WITH COMPOSITE HEAT SHIELDS

Номер: US20170022829A1
Принадлежит:

A nozzle guide vane for a gas turbine engine is disclosed herein. The nozzle guide vane includes an inner endcap, an outer endcap, and at least one airfoil that extends from the inner endcap to the outer endcap. The nozzle guide vane further includes at least one composite heat shield component adapted to shield metallic components from high temperature gasses. 1. A nozzle guide vane for a gas turbine engine , the nozzle guide vane comprisinga metallic support structure including an inner endcap formed to include an inner attachment aperture and an outer endcap formed to include an outer attachment aperture, the outer endcap spaced from the inner endcap in a radial direction,an airfoil including an aerodynamic feature shaped to redirect gasses moving through a gas path between the inner end cap and the outer endcap, an inner attachment feature that extends from the aerodynamic feature into the inner attachment aperture of the inner endcap, and an outer attachment feature that extends from the aerodynamic feature into the outer attachment aperture of the outer endcap, anda ceramic-matrix composite heat shield system adapted to shield the metallic support structure from hot gasses moving through the gas path, the ceramic-matrix composite heat shield system including an inner heat shield arranged radially between the inner endcap and the gas path and an outer heat shield comprising ceramic-matrix composite materials arranged radially between the outer endcap and the gas path.2. The nozzle guide vane of claim 1 , wherein the outer heat shield is sandwiched between the aerodynamic feature and the outer endcap and the inner heat shield is sandwiched between the aerodynamic feature and the inner endcap.3. The nozzle guide vane of claim 2 , wherein the inner heat shield is formed to include an inner locator aperture claim 2 , the outer heat shield is formed to include an outer locator aperture claim 2 , the inner attachment feature extends through the inner locator aperture ...

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25-01-2018 дата публикации

EXHAUST GAS TURBOCHARGER

Номер: US20180023459A1
Принадлежит:

An exhaust gas turbocharger () including a compressor housing (), a bearing housing (), a turbine housing () in which a turbine wheel () is arranged, which has a housing inlet (), which has a turbine spiral () connecting to the housing inlet (), which has a housing outlet (), and which has a wastegate arrangement (), which in the open state brings the housing inlet () into flow connection with the housing outlet () for guiding a wastegate mass flow. The wastegate insert part () is arranged in the turbine housing () between the turbine wheel () and the housing outlet () and the open wastegate arrangement () introduces the wastegate mass flow into the wastegate insert arrangement (). 11. An exhaust gas turbocharger () comprising:{'b': '2', 'compressor housing (),'}{'b': '3', 'a bearing housing (),'}{'b': '4', 'a turbine housing ()'}{'b': '5', 'in which a turbine wheel () is arranged,'}{'b': '6', 'which has a turbine housing inlet (),'}{'b': 7', '6, 'which has a turbine spiral () connecting to the turbine housing inlet (),'}{'b': '8', 'which has a turbine housing outlet (), and'}{'b': 9', '6', '8, 'which has a wastegate arrangement (), which in the open state brings the turbine housing inlet () into flow connection with the turbine housing outlet () for guiding a wastegate mass flow, wherein'}{'b': 10', '4', '5', '8', '9', '10, 'a wastegate insert part () is arranged in the turbine housing () between the turbine wheel () and the turbine housing outlet () and the open wastegate arrangement () when open introduces the wastegate mass flow into the wastegate insert part (), and'}{'b': 10', '11, 'the wastegate insert part () is surrounded with heat insulation ().'}214. The exhaust gas turbocharger according to claim 1 , wherein the wastegate insert part () is a separately producible component which is inserted into the turbine housing () and fixed in the same.31010. The exhaust gas turbocharger according to claim 1 , wherein the wastegate insert part () is a sheet metal ...

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10-02-2022 дата публикации

DAMPER CONTROL VALVE FOR A TURBOMACHINE

Номер: US20220042423A1
Принадлежит:

A gas turbine engine having a damping system that includes features for optimizing the damping response to vibrational loads on a rotary component for a wide range of operational conditions is provided. In one aspect, the damping system includes a damper control valve. The damper control valve receives working fluid from a working fluid supply and has a valve plunger movable between a first position and a second position. When the valve plunger is in the first position, the damper control valve permits working fluid to flow to a first damper associated with a first bearing coupled with the rotary component and to a second damper associated with a second bearing coupled with the rotary component. When the valve plunger is in the second position, the damper control valve permits working fluid to flow to the first damper but not the second damper. 1. A turbomachine , comprising:a rotary component rotatable about an axis of rotation;a bearing assembly having one or more bearings each operatively coupled with the rotary component, each of the one or more bearings having a damper associated therewith, each of the dampers defining one or more chambers;a damper control valve having a valve casing defining a valve chamber, the valve chamber being in fluid communication with a working fluid supply and the one or more chambers, the damper control valve operable to receive working fluid from the working fluid supply, andwherein the damper control valve has a valve plunger movable within the valve chamber between a first position in which working fluid flows to at least two chambers of the one or more chambers and a second position in which working fluid flows to at least one less chamber than the at least two chambers to which working fluid flows when the valve plunger is in the first position.2. The turbomachine of claim 1 , wherein the damper control valve has a biasing member operable to bias the valve plunger in the first position.3. The turbomachine of claim 2 , wherein ...

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