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Небесная энциклопедия

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

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Мониторинг СМИ

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

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Поддерживает ввод нескольких поисковых фраз (по одной на строку). При поиске обеспечивает поддержку морфологии русского и английского языка
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Применить Всего найдено 2110. Отображено 100.
31-05-2012 дата публикации

Fluid Turbine Having Optimized Blade Pitch Profiles

Номер: US20120134820A1
Автор: Jason Daniel Cormey, Robert Clifton Vance
Принадлежит: Broadstar Investment Co LLC, Broadstar Wind Systems Group LLC

A fluid turbine comprising a rotor, having an axis of rotation, comprising at least two rotor blades disposed at a radius from the axis of rotation, each rotor blade having a pitch axis and a variable pitch angle. The fluid turbine comprises a mechanism operable to control the pitch angle of at least one rotor blade about its pitch axis and to vary the pitch angle of the rotor blade between various pitch angles as the blade moves radially about the axis of rotation of the rotor.

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Номер записи: 1
27-12-2012 дата публикации

Method For Balancing A Propulsive System Having Non-Hull Contra-Rotating Propellers

Номер: US20120328433A1
Автор: Esteban Quiroz-Hernandez
Принадлежит: AIRBUS OPERATIONS SAS

According to the invention, at least one of the counterweights ( 40, 40.1, 40.2, 40.3, 40.4 ) is mobile mounted on a guiding slot ( 24, 34 ) coaxial to the hub envelope ( 22, 32 ) that surrounds the hub ( 21, 31 ) of the corresponding propeller ( 2, 3 ), the movement of said mobile counterweight ( 40, 40.1, 40.2, 40., 40.4 ) along said guiding slot ( 24, 34 ) being controlled on the basis of an estimation of the possible unbalance of said propulsive system ( 1 ).

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Номер записи: 2
14-02-2013 дата публикации

Variable vane actuation system and method

Номер: US20130039735A1
Автор: Andy Eifert
Принадлежит: Rolls Royce Corp

A variable vane actuation system and method is disclosed herein. The variable vane actuation system includes a first ring member disposed for pivoting movement about a centerline axis. The first ring member is operably connected with at least one vane such that the at least one vane pivots in response to the pivoting movement of the first ring member. The variable vane actuation system also includes a first pin engaged with the first ring member. The variable vane actuation system also includes a ring moving device operably engaged with the first pin to move the first ring member about the centerline axis. The ring moving device includes at least one plate having a first slot and an actuator operable to move the at least one plate. The first pin is received in the first slot and is a cam follower to a cam defined at least in part by a surface of the first slot.

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Номер записи: 3
05-09-2013 дата публикации

Device for connecting a thrust reverser front frame to a fan casing and nacelle incorporating such a device

Номер: US20130227962A1
Автор: Hazem Kioua, Jerôme LESCURE, Serge Bunel, Vincent Peyron
Принадлежит: Aircelle SA

A device for connecting a thrust reverser front frame to a fan casing includes a crenulated flange secured to the front frame, an annular component to accept this flange secured to the fan casing, and a crenulated annulus of a shape that complements the flange and pivot-mounted on the annular component. The crenulated flange, the annular component and the crenulated annulus are designed in such a way that a rotation of the annulus with respect to the annular component has the effect of locking this flange against this annular component through the collaboration between the respective crenulations of the flange and of the annulus.

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Номер записи: 4
14-11-2013 дата публикации

Inner turbine shell axial movement

Номер: US20130302147A1
Автор: Kenneth Black, Pradeep GHUTE, Rohit Pruthi, Sanjay Shankar Jadhav
Принадлежит: General Electric Co

A clearance control system for a turbine having a stator assembly and a rotor assembly includes a hydraulic or pneumatic controller that axially drives, through a shaft, one or more actuators connected to the stator assembly casing. The controller causes relative movement between the stator and rotor assemblies to adjust the clearances between portions of the stator and rotor in accordance with the varying operating conditions of the turbine. More particularly, the controller moves the stator relative to the rotor in first and second axial directions to compensate for thermal expansion and contraction during the operating conditions of the turbine.

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Номер записи: 5
26-12-2013 дата публикации

Four bar drive mechanism for bleed system

Номер: US20130341547A1
Автор: Kevin J. Cummings, Ryan Edward LeBlanc
Принадлежит: Individual

An actuation system for a bleed valve includes first and second bell cranks connected by a connecting link. The first bell crank has a first arm that is coupled to a bleed valve and a second arm that is coupled to the connecting link. The connecting link has a first end coupled to the second arm of the first bell crank and a second end coupled to a first arm of the second bell crank. A second arm of the second bell crank is coupled to an actuating element. Input is communicated through the actuating element to move the bleed valve between open and closed positions via the first and second bell cranks.

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Номер записи: 6
26-12-2013 дата публикации

Spherical-link end damper system with near constant engagement

Номер: US20130343876A1
Автор: Kevin J. Cummings, Ryan Edward LeBlanc
Принадлежит: United Technologies Corp

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

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Номер записи: 7
06-02-2014 дата публикации

Generating device for aircraft

Номер: US20140038770A1
Автор: Isao Fujii, Kenichiro Tanaka, Koji Watanabe, Tatsuhiko Goi
Принадлежит: Kawasaki Jukogyo KK

An electric power generating device ( 1 ) capable of suppressing an increase of a frontal surface area of an aircraft engine includes a transmission ( 22 ) connected with a rotary shaft ( 9 ) of the engine (E), an electric power generator ( 34 ) driven by an output of the transmission ( 22 ), an input shaft ( 27 ) having a shaft axis extending in a direction crossing the rotary shaft ( 9 ) and connected with the rotary shaft ( 9 ), and a transmitting mechanism ( 21 ) connected with the input shaft ( 27 ) to drive the transmission ( 22 ) about an axis extending in a direction perpendicular to the input shaft ( 27 ). The transmission ( 22 ) and the electric power generator ( 34 ) are disposed spaced a distance from each other in a direction circumferentially of the rotary shaft ( 9 ).

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Номер записи: 8
06-03-2014 дата публикации

Drive lever arrangement

Номер: US20140064875A1
Автор: Philip Twell
Принадлежит: SIEMENS AG

A drive lever arrangement is provided. The drive lever arrangement includes a unison ring which has a groove; a drive lever having connection means for connecting the drive lever to the unison ring; a drive lever pin having a transversal throughbore hole; and a clip, wherein the drive lever pin connects the drive lever to the unison ring, and wherein the clip is inserted in the groove through the throughbore hole of the drive lever pin, and engages with the groove of the unison ring.

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Номер записи: 9
03-01-2019 дата публикации

SYSTEM AND METHOD OF OPERATING A DUCTED FAN PROPULSION SYSTEM DURING AIRCRAFT TAXI

Номер: US20190002118A1
Автор: Blodgett Keith Edward James, Cox Melanie Zoe, Gabel Martin Kenneth, Nestico Brian Francis
Принадлежит:

A thrust reverser assembly and a method of operating an aircraft during a taxi mode of operation are provided. The thrust reverser assembly includes one or more actuator assemblies configured to modulate a position of a moveable portion over a continuous range of travel between a fully stowed position and a fully deployed position, such that an air flow through said thrust reverser bleed passage is correspondingly varied. The thrust reverser assembly also includes a throttle device that includes a first, ground idle power level position and a second, forward thrust mode position. Movement into the second position may be actuated separately and differently from movement into the first position. An actuator intermediate lock may inhibit actuation of the intermediate forward thrust mode of operation until a plurality of preconditions is met. 1. A thrust reverser assembly for an aircraft comprising:a moveable portion that is moveable over a continuous range of travel between a fully stowed position and a fully deployed position, wherein movement away from said fully stowed position opens a thrust reverser bleed passage;one or more actuator assemblies coupled to said moveable portion and operable in an intermediate forward thrust mode to modulate a position of said moveable portion along said continuous range of travel, such that an air flow through said thrust reverser bleed passage is correspondingly varied;a throttle device comprising a first position associated with a ground idle power level and a second position associated with the intermediate forward thrust mode, wherein movement of said throttle device into said second position is actuated separately and differently from movement of said throttle device into said first position; andan actuator intermediate lock coupled to said one or more actuator assemblies and configured to inhibit actuation of the intermediate forward thrust mode until a plurality of preconditions are met.2. The thrust reverser assembly of ...

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Номер записи: 10
05-01-2017 дата публикации

LINEAR MOTION MECHANISM, GOVERNING VALVE DRIVE DEVICE, AND STEAM TURBINE

Номер: US20170002680A1
Автор: HARADA Naruhito, KATAGAKE Makoto, NAGAI Naoyuki, OKAMURA Tomoaki, TONAKA Hideki
Принадлежит:

A linear motion mechanism () is provided with: a cylinder rod () into which a ball screw () can be inserted, said cylinder rod () comprising a base end section that is connected to a nut () within a piston casing () and a tip section () that is exposed on the outside of the piston casing (); a nut-side grease supply hole () that is formed in the nut () and that comprises a discharge port () that opens toward the outer circumferential surface of the ball screw (); and a cylinder rod-side grease supply hole () that is formed in the cylinder rod (), that comprises on one end thereof an inlet () that opens at a position that is exposed to the outer section of the piston casing (), and that comprises another end () that is connected to the nut-side grease supply hole (). 1. A linear motion mechanism , comprising:an electric motor;a ball screw which is rotationally driven around an axis by the electric motor;a nut which is screwed into the ball screw, and advances and retreats relative to the ball screw in an axial direction of the ball screw according to rotation of the ball screw;a casing which surrounds the ball screw and the nut;a tubular cylinder rod which includes a base end section which is connected to the nut inside the casing and a tip section which is exposed to the outside of the casing, and into which the ball screw can be inserted;a nut-side grease supply hole which is formed in the nut and includes a discharge port which opens toward an outer circumferential surface of the ball screw; anda cylinder rod-side grease supply hole which is formed in the cylinder rod, and includes an inlet which opens at a position exposed to the outside of the casing on one end of the cylinder rod-side grease supply hole, and the other end thereof which communicates with the nut-side grease supply hole.2. The linear motion mechanism according to claim 1 , further comprising:a grease supply pipe which is connected to the inlet of the cylinder rod-side grease supply hole and has ...

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Номер записи: 11
05-01-2017 дата публикации

DETECTION METHOD OF SENSOR IN GAS TURBINE

Номер: US20170002682A1
Автор: LAU Ow, Neuenhahn Thomas, REN Chao, Zheng Jie
Принадлежит: SIEMENS AKTIENGESELLSCHAFT

A detection method of a sensor in a gas turbine includes adopting a pressure sensor to measure a pushing force of a push rod; measuring a first rotational angle of a guide vane where a first angle sensor is mounted; measuring a second rotational angle of the guide vane where a second angle sensor is mounted; obtaining a maximum measured rotational angle deviation from the absolute value of a difference value between the first and second rotational angles; calculating a maximum calculated deviation from the pushing force of the push rod; calculating the absolute value of a difference value between the maximum measured deviation and the maximum calculated deviation; and determining that the angle sensors and the pressure sensor have appropriate measurement accuracy; or, if the absolute value is greater than the standard value, determining that the angle and/or pressure sensors require calibration. 1. A method for sensors in a gas turbine , the gas turbine including a cylinder , a plurality of guide vanes , a first angle sensor , a second angle sensor , and a guide vane driving mechanism configured to drive the guide vanes to rotate , the guide vane driving mechanism including a driving ring , a push rod configured to push the driving ring to rotate relative to the cylinder , a pressure sensor configured to measure a thrust of the push rod , a plurality of connecting rods and adjusting rods connecting the guide vanes and the driving ring , and a plurality of elastic support bases connecting the cylinder and the driving ring , the method comprising:measuring thrust of the push rod via the pressure sensor;measuring a first rotation angle of the guide vanes in an installation position of the first angle sensor;measuring a second rotation angle of the guide vanes in an installation position of the second angle sensor;obtaining a measured maximum rotation angle offset according to an absolute value of a difference between said measured first rotation angle and said measured ...

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Номер записи: 12
02-01-2020 дата публикации

AIR TURBINE STARTER WITH TURBINE AIR EXHAUST OUTLET VALVE

Номер: US20200003072A1
Автор: FUKUDA Takao, Simpson Scott W., Zawilinski David J.
Принадлежит:

A system includes an air turbine starter having an inlet, a turbine air exhaust outlet, an output shaft, and a turbine in fluid communication with the inlet and the turbine air exhaust outlet. The turbine is operably coupled to the output shaft. The system also includes an outlet valve assembly configured to adjust an exhaust area of the turbine air exhaust outlet. 1. A system comprising: an inlet;', 'a turbine air exhaust outlet;', 'an output shaft; and', 'a turbine in fluid communication with the inlet and the turbine air exhaust outlet, the turbine operably coupled to the output shaft; and', 'an outlet valve assembly configured to adjust an exhaust area of the turbine air exhaust outlet., 'an air turbine starter comprising2. The system as in claim 1 , further comprising an actuator in fluid communication with the outlet valve assembly claim 1 , the actuator operable to adjust the exhaust area of the turbine air exhaust outlet.3. The system as in claim 2 , wherein the outlet valve assembly further comprises:a valve housing; anda valve body arranged between the valve housing and the turbine air exhaust outlet.4. The system as in claim 3 , wherein the valve housing comprises at least one pressure port claim 3 , and a pressurized cavity is formed between the at least one pressure port and the valve body responsive a pressurized flow from the actuator.5. The system as in claim 4 , further comprising at least one spring positioned between a housing of the air turbine starter and the valve body claim 4 , the at least one spring configured to provide an opening force to slide the valve body towards an open position and increase the exhaust area of the turbine air exhaust outlet.6. The system as in claim 5 , wherein the actuator is operable to increase the pressurized flow to the pressurized cavity and provide a closing force greater than the opening force to slide the valve body towards a closed position and decrease the exhaust area of the turbine air exhaust outlet.7. ...

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Номер записи: 13
02-01-2020 дата публикации

TARGET DOOR REVERSER WITH NON-PARALLEL HINGE LINES

Номер: US20200003152A1
Автор: Gormley Timothy
Принадлежит: Rohr, Inc.

The present disclosure provides a thrust reverser comprising a stationary structure defining an annular body with a centerline, a first reverser door pivotally coupled to the stationary structure by a pair of first reverser door hinges, a first reverser door hinge axis extending through the first reverser door hinges and positioned at a first angle relative to a centerline, and a second reverser door pivotally coupled to the stationary structure by a pair of second reverser door hinges, a second hinge line axis extending through the second reverser door hinges and positioned at a second angle relative to the centerline. 1. A thrust reverser , comprising:a stationary structure defining an annular body with a centerline;a first reverser door pivotally coupled to the stationary structure by a pair of first reverser door hinges, a first reverser door hinge axis extending through the pair of first reverser door hinges and positioned at a first angle relative to the centerline; anda second reverser door pivotally coupled to the stationary structure by a pair of second reverser door hinges, a second reverser door hinge axis extending through the pair of second reverser door hinges and positioned at a second angle relative to the centerline.2. The thrust reverser of claim 1 , wherein the pair of first reverser door hinges and the pair of second reverser door hinges are offset in an opposite direction and an equal distance from the centerline.3. The thrust reverser of claim 1 , wherein the pair of first reverser door hinges and the pair of second reverser door hinges are offset from the centerline in an opposite direction claim 1 , the pair of first reverser door hinges offset from the centerline by a first distance and the pair of second reverser door hinges offset from the centerline by a second distance.4. The thrust reverser of claim 1 , wherein the first reverser door is configured to rotate about the first reverser door hinge axis and the second reverser door is ...

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Номер записи: 14
02-01-2020 дата публикации

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

Номер: US20200003154A1
Автор: Carr Alexander Jon
Принадлежит: SPIRIT AEROSYSTEMS, INC.

A system for actuating a blocker door of a thrust reverser, in which a drag link assembly is removed from the airflow through the engine during flight. The assembly couples the door to a sleeve so that translation of the sleeve between deployed and stowed positions moves the door to open and closed positions, respectively. The assembly includes a drag link track having a channel extending between first and second track ends, and a drag link having a link end which is slidable in the track channel between the first and second track ends. During deployment, translation of the sleeve causes the second link end to slide within the channel toward the second track end which causes the door to open, and during stowage, translation of the sleeve causes the second link end to slide within the channel toward the first track end which causes the door to close. 1. A system for actuating a blocker door of a thrust reverser , the system comprising:a sleeve translatable between a stowed position in which an engine airflow is directed rearwardly and a deployed position in which the engine airflow is redirected forwardly;a blocker door moveable between a closed position associated with the stowed position of the sleeve in which the engine airflow is directed rearwardly and an open position associated with the deployed position of the sleeve in which the engine airflow is redirected laterally by the blocker door; and a drag link track including a first track end, a second track end, and an elongated channel extending between the first and second track ends, and', 'a drag link including a first link end and a second link end, with the second link end being slidably received in the elongated channel of the drag link track and moveable within the elongated channel between the first track end and the second track end,, 'a drag link assembly located between the sleeve and the blocker door and configured to mechanically couple the sleeve to the blocker door so that translation of the ...

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Номер записи: 15
02-01-2020 дата публикации

Collapsible drag link

Номер: US20200003155A1
Автор: Bryce Tyler Kelford
Принадлежит: Rohr Inc

A thrust reverser system of a nacelle is provided. The thrust reverser system may include a pressure shell, a blocker door, and a drag link. The blocker door may be pivotably coupled to the pressure shell. The drag link may include a first segment pivotably coupled to a second segment and a third segment pivotably coupled to the second segment and the blocker door.

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Номер записи: 16
02-01-2020 дата публикации

THRUST REVERSER WITH DISPLACEABLE TRAILING EDGE BODY

Номер: US20200003156A1
Автор: Gormley Timothy
Принадлежит:

An assembly is provided for an aircraft propulsion system. This assembly includes a target-type thrust reverser and a tubular trailing edge body. The target-type thrust reverser includes a plurality of thrust reverser doors. Each of the thrust reverser doors is configured to pivot between a stowed position to a deployed position. The tubular trailing edge body is configured to at least partially form a gas path nozzle for the aircraft propulsion system. The tubular trailing edge body is configured to be displaced when the thrust reverser doors pivot from the stowed position to the deployed position. 1. An assembly for an aircraft propulsion system , comprising:a target-type thrust reverser comprising a plurality of thrust reverser doors, each of the thrust reverser doors configured to pivot between a stowed position to a deployed position; anda tubular trailing edge body configured to at least partially form a gas path nozzle for the aircraft propulsion system, the tubular trailing edge body configured to be displaced when the thrust reverser doors pivot from the stowed position to the deployed position.2. The assembly of claim 1 , wherein the displacement of the tubular trailing edge body comprises a radial displacement relative to an axial centerline of the aircraft propulsion system.3. The assembly of claim 1 , wherein the displacement of the tubular trailing edge body comprises an axial displacement relative to an axial centerline of the aircraft propulsion system.4. The assembly of claim 1 , wherein the displacement of the tubular trailing edge body comprises a pivotal displacement.5. The assembly of claim 1 , wherein the displacement of the tubular trailing edge body comprisesa radial displacement relative to an axial centerline of the aircraft propulsion system;an axial displacement relative to the axial centerline; anda pivotal displacement.6. The assembly of claim 1 , wherein the tubular trailing edge body is configured to completely form a trailing edge of ...

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Номер записи: 17
12-01-2017 дата публикации

HOUSING FOR DRIVING AN APPARATUS FOR A TURBINE ENGINE

Номер: US20170009660A1
Автор: BECQUERELLE Samuel Raymond Germain, BRANDT Bernard, PRUNERA-USACH Stéphane, VIEL Julien
Принадлежит:

A gearbox to be attached to a turbine engine to drive at least one apparatus annexed to the turbine engine, the gearbox including a housing; a power take-off member capable of engaging with a radial shaft of the turbine engine; at least one kinematic chain located inside the housing and capable of transmitting the rotational movement of the power take-off member to at least one rotatable shaft of an apparatus. The kinematic chain includes a first end and a second end. The power take-off member is linked to the kinematic chain by a gear having convergent axes and located between the first end and the second end of the kinematic chain. 1. A gearbox to be fixed to a turbine engine in order to drive at least one apparatus annexed to the turbine engine , the gearbox comprising:a housing;a power take-off member capable of engaging with a radial shaft of the turbine engine;{'b': '1', 'at least one kinematic chain located inside the housing and capable of transmitting a rotational movement of a power take-off to at least one rotatable shaft of an apparatus, the kinematic chain comprising a first end and a second end, p wherein the power take-off member is linked to the kinematic chain by a gear having convergent axes located within the kinematic chain.'}2. The gearbox according to a claim 1 , wherein the kinematic chain comprises at least one central shaft capable of transmitting the movement of the power take-off member to at least one rotatable shaft of an apparatus through at least one intermediate gear claim 1 , where the central shaft comprises two ends claim 1 , the power take-off member being linked to the central shaft by a gear having convergent axes claim 1 , located between the first end and the second of the central shaft.3. The gearbox according to wherein the central shaft comprises a first and second part claim 2 , with the power take-off member being linked to the first part by a first gear with convergent axes and to the second part by a second gear with ...

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Номер записи: 18
11-01-2018 дата публикации

SYSTEM FOR CONTROLLING VARIABLE PITCH BLADES FOR A TURBINE ENGINE

Номер: US20180010478A1
Автор: BENDERRADJI Kamel, BROMANN Alain Marc, COUSTILLAS Suzanne Madeleine, DUMAS Lilian Yann, Lucien
Принадлежит:

The invention relates to a system for controlling variable pitch blades for a turbine engine, comprising an annular row of variable pitch blades extending about an axis (A) and each comprising a blade connected at the radially outer end thereof to a pivot () that defines a substantially radial axis of rotation of the blade and which is connected by a lever () to control means (, ) extending about said axis. The invention is characterized in that said control means include first links () supported by said pivots and second links () extending between said first links, said first and second links extending substantially along a same circumference of said axis and being connected to one another and to actuation means (). 1. A control system for variable-pitch vanes for a turbine engine , comprising at least one annular row of variable-pitch vanes extending around an axis and each comprising a blade which is connected at the radially outer end thereof to a pivot which defines a substantially radial axis of rotation of the vane and which is connected by a lever to control means extending around said axis , wherein said control means comprise first links carried by said pivots and second links extending between said first links , said first and second links extending substantially over the same circumference around said axis and being connected to one another and to actuation means.2. The control system according to claim 1 , wherein the number of the first links is equal to the number of the second links claim 1 , which is equal to the number of levers.3. The control system according to claim 1 , wherein said first links are formed in one piece with said levers.4. The control system according to claim 1 , wherein each of the first and/or second links has an elongate shape and is connected by the longitudinal ends thereof to other links.5. The control system according to claim 1 , wherein the first links are connected by pivot and/or swivel connections to the second links. ...

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Номер записи: 19
11-01-2018 дата публикации

GEARED GAS TURBINE ENGINE AND A GEARBOX

Номер: US20180010525A1
Автор: MADGE Jason J.
Принадлежит: ROLLS-ROYCE PLC

A gas turbine engine comprises a gearbox comprising a sun gear, an annulus gear, a plurality of planet gears and a carrier. The carrier comprises a primary structure and at least one reinforcing structure. The primary structure comprises a first material and the at least one reinforcing structure comprises a second material. The primary structure includes a first ring, a second ring spaced axially from the first ring and a plurality of circumferentially spaced axles extending axially between the first ring and the second ring. Each planet gear is rotatably mounted on a respective one of the axles by a bearing. The reinforcing structure is secured to the primary structure and the reinforcing structure comprises a particulate reinforced material or a fibre reinforced material. The reinforcing structure increases the stiffness of the carrier and reduces the weight of the carrier. 1. A gas turbine engine comprising a gearbox , the gearbox comprising a sun gear , an annulus gear , a plurality of planet gears and a carrier , the sun gear meshing with the planet gears and the planet gears meshing with the annulus gear , the carrier comprising a primary structure and at least one reinforcing structure , the primary structure comprising a first material and the at least one reinforcing structure comprising a second material , the primary structure comprising a first ring , a second ring spaced axially from the first ring and a plurality of circumferentially spaced axles extending axially between the first ring and the second ring , each planet gear being rotatably mounted on a respective one of the axles , the at least one reinforcing structure being secured to the primary structure , and the at least one reinforcing structure comprising a reinforced material , the reinforced material being selected from the group consisting essentially of a particulate reinforced material and a fibre reinforced material.2. A gas turbine engine as claimed in wherein the at least one ...

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Номер записи: 20
03-02-2022 дата публикации

TELESCOPIC BALLSCREW ACTUATOR

Номер: US20220034389A1
Автор: Hawksworth Andrew, Holding James, Morgan Antony
Принадлежит:

An apparatus for a thrust reverser actuation system (“TRAS”), the apparatus comprising: an input shaft; a first component located concentrically around the input shaft; a second component located concentrically around the first component; a first ballscrew mechanism between the input shaft and the first component, and configured such that rotational movement of the input shaft causes a translational movement of the first component via the first ballscrew mechanism; and a second ballscrew mechanism between the first component and the second component, and configured such that rotational movement of the first component causes a translational movement of the second component via the second ballscrew mechanism. 1. An apparatus for use in an aircraft , the apparatus comprising:an input shaft;a first component located concentrically around the input shaft;a second component located concentrically around the first component;a first ballscrew mechanism between the input shaft and the first component, and configured such that rotational movement of the input shaft causes a translational movement of the first component via the first ballscrew mechanism; anda second ballscrew mechanism between the first component and the second component, and configured such that rotational movement of the first component causes a translational movement of the second component via the second ballscrew mechanism;wherein in a first mode the first component is configured to translate along an axis upon rotational movement of the input shaft, and in a second mode the first component is configured to rotate about the axis;wherein the first mode occurs prior to the second mode, and a transition between the first mode and the second mode is caused by a stop fixedly attached to the first component abutting a stop fixedly attached to the input shaft to limit the stroke of the first component, which abutment prevents further translation of the first component such that further rotational movement of the ...

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Номер записи: 21
21-01-2016 дата публикации

Actuator for gas turbine engine blade outer air seal

Номер: US20160017743A1
Автор: Brian Duguay
Принадлежит: United Technologies Corp

A blade outer air seal (BOAS) actuator assembly, according to an exemplary aspect of the present disclosure includes, among other things, an actuator member; and a retractor configured to move with the actuator member to move a BOAS segment from a first position to a second position that is radially outside the first position, the BOAS segment seated against a support structure when in the first position and spaced from the support structure when in the second position.

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Номер записи: 22
18-01-2018 дата публикации

ASSEMBLY FOR CONTROLLING VARIABLE PITCH VANES IN A TURBINE ENGINE

Номер: US20180016931A1
Автор: BENDERRADJI Kamel, Bergan Blaise, Bromann Alain Marc Lucien, COUSTILLAS Suzanne Madeleine, DUMAS Lilian Yann
Принадлежит: SAFRAN AIRCRAFT ENGINES

An assembly, in particular for controlling variable pitch vanes in a turbine engine, comprising an actuating ring surrounding a casing of the turbine engine and connected by rods to variable pitch vanes, in addition to a driving means for rotating the actuating ring around the casing. The assembly includes a slidingly connected passive element, one end of which is connected by a sliding pivoting link on the actuating ring and a second end is connected by a ball-joint link to the casing. 1. Assembly for controlling variable pitch vanes in a turbine engine comprising an actuating ring surrounding a casing of the turbine engine and connected by rods to variable pitch vanes in addition to a driving means for rotating the actuating ring around the casing , the assembly comprising a slidingly connected passive element , one end of which is connected by a sliding pivoting link on the actuating ring and a second end is connected by a ball-joint link to the casing.2. Assembly according to claim 1 , wherein the passive element is arranged circumferentially substantially opposite the driving means of the ring.3. Assembly according to claim 1 , wherein the passive element comprises a body bearing the first end and in which a pin bearing the second end is mounted for translational movement.4. Assembly according to claim 1 , wherein the first end of the passive element is mounted for rotation and translational movement around and according to a radial axis in a yoke of a first support element integral with the actuating ring.5. Assembly according to claim 1 , wherein the second end of the passive element is connected via a ball-joint link to a second support element integral with the casing.6. Assembly according to claim 4 , wherein the first support element is arranged axially downstream claim 4 , or upstream respectively claim 4 , from the rods and the second support element is arranged upstream claim 4 , or downstream respectively claim 4 , from the rods.7. Assembly according ...

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Номер записи: 23
18-01-2018 дата публикации

Hydraulically Driven Thrust Reverser

Номер: US20180017018A1
Автор: Morgan Antony
Принадлежит:

A drive system for a thrust reverser for a gas turbine engine comprises a plurality of mechanical actuators for driving a cowl of the thrust reverser, and a single hydraulic motor for providing power to the plurality of mechanical actuators. Power can be transmitted mechanically from the hydraulic motor to the actuators, and the hydraulic motor can be located away from fire zones in the engine. 1. A drive system for a thrust reverser for a gas turbine engine; comprising:a plurality of mechanical actuators for driving a cowl of the thrust reverser; anda hydraulic motor for providing power to the plurality of mechanical actuators.2. A drive system for a thrust reverser as claimed in claim 1 , wherein rotation drive from the hydraulic motor is transmitted to the mechanical actuators by mechanical means.3. A drive system for a thrust reverser as claimed in claim 1 , wherein the hydraulic motor is located away from fire zones in the gas turbine engine.4. A drive system for a thrust reverser as claim 1 , wherein the hydraulic motor is controlled by a hydraulic motor control unit.5. A drive system for a thrust reverser as claimed in claim 1 , wherein there is a single hydraulic motor.6. A thrust reverser for a gas turbine engine claim 1 , comprising:a plurality of cowls; and a plurality of mechanical actuators for driving one of plurality of cowls; and', 'a hydraulic motor for providing power to the plurality of mechanical actuators;, 'a plurality of drive systems, the plurality of drive systems each includingwherein each cowl ids driven by a one of the drives systems.7. A thrust reverser for a gas turbine engine claim 4 , wherein the thrust reverser contains a plurality of cowls claim 4 , each cowl being driven by a drive system as claimed in claim 4 , wherein all of the hydraulic motor control units are controlled by a single controller.8. A gas turbine engine including a thrust reverser as claimed in .9. A gas turbine engine including a thrust reverser as claimed in .10 ...

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Номер записи: 24
18-01-2018 дата публикации

Turbine engines with variable area nozzle

Номер: US20180017020A1
Автор: David Patrick Calder, Graham Frank Howarth
Принадлежит: MRA Systems LLC

A turbine engine having an engine core, an inner cowl radially surrounding the engine core, an outer cowl radially surrounding the inner cowl and spaced from the inner cowl to form an annular passage between the inner and outer cowls that defines a nozzle, at least one control surface provided on the inner cowl and movable between a retracted position, where the nozzle has a first cross-sectional area, and an extended position where the nozzle has a second cross-sectional area that is less than the first cross-sectional area and an actuator operably coupled to the control surface and configured to move the control surface to control the cross-sectional area of the nozzle.

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Номер записи: 25
18-01-2018 дата публикации

VARIABLE STATOR VANE MECHANISM

Номер: US20180017080A1
Автор: WALTERS Edward A.
Принадлежит: ROLLS-ROYCE PLC

A variable vane mechanism for adjusting the angle of stator vanes in a gas turbine engine is provided. The mechanism has a circumferentially extending drive ring that is driven by an actuator, and a guide surface that is radially inside the drive ring. The mechanism also has a centralising pin that is connected to the drive ring and also in slidable contact with the guide surface so as to be movable with the drive ring relative to the guide surface. The centralising pin allows both the drive ring to be connected to a stator vane (via a lever) in order to adjust the angle of the vane, and the radial position of the drive ring to be adjusted to ensure accurate and repeatable operation of the mechanism. 1. A variable vane mechanism for adjusting the angle of stator vanes in an axial flow gas turbine engine that defines axial , radial and circumferential directions , the variable vane mechanism comprising:a circumferentially extending drive ring arranged to be driven circumferentially by a drive mechanism;a circumferentially extending guide surface that is radially inside the drive ring;a centralising pin that is connected to the drive ring so as to move with the drive ring, a first end of the centralising pin being in slidable contact with the guide surface so as to be movable relative to the guide surface; anda lever having a first end and a second end, the first end being rotatably connected to the centralising pin so as to be moveable with the centralising pin and rotatable relative to the centralising pin, and the second end being arranged for connection to a stator vane so as to enable adjustment of the angle of the stator vane.2. A variable vane mechanism according to claim 1 , wherein the centralising pin extends in a substantially radial direction.3. A variable vane mechanism according to claim 1 , wherein the centralising pin extends through the drive ring.4. A variable vane mechanism according to claim 1 , wherein the centralising pin comprises a thread that ...

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Номер записи: 26
17-01-2019 дата публикации

ACTUATOR FOR GAS TURBINE ENGINE BLADE OUTER AIR SEAL

Номер: US20190017407A1
Автор: Duguay Brian
Принадлежит:

A method of actuating a Blade Outer Air Seal (BOAS) includes moving a retractor against a portion of a BOAS segment to move the BOAS segment from a first position to a second position that is radially outside the first position. The BOAS segment is seated against a support structure when in the first position and spaced from the support structure when in the second position.

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Номер записи: 27
21-01-2021 дата публикации

Containment Case Active Clearance Control Structure

Номер: US20210017877A1
Автор: Jain Manoj Kumar, Jain Nitesh, Kray Nicholas Joseph
Принадлежит:

A propulsion system including a casing surrounding a fan rotor assembly is provided. An example casing includes an outer layer material, an inner layer material including first openings extended partially through the inner layer material along a radial direction of a first side of the inner layer material, and second openings extended partially through the inner layer material along the radial direction of a second side of the inner layer material opposite the first side, and a spring member coupled to the outer layer material and the inner layer material. 1. A casing to surround a fan rotor assembly , the casing comprising:an outer layer material; first openings extended partially through the inner layer material along a radial direction of a first side of the inner layer material; and', 'second openings extended partially through the inner layer material along the radial direction of a second side of the inner layer material opposite the first side; and, 'an inner layer material includinga spring member coupled to the outer layer material and the inner layer material.2. The casing of claim 1 , wherein the first openings and the second openings are configured to enable expansion and contraction of the inner layer material along the radial direction.3. The casing of claim 1 , wherein the outer layer material has a first coefficient of thermal expansion (CTE) and the spring member has a second CTE greater than the first CTE.4. The casing of claim 1 , wherein the spring member is disposed between the outer layer material and the inner layer material within a flow passage defined between an inner surface of the outer layer material and the first side of the inner layer material.5. The casing of claim 4 , wherein the spring member is extended at least partially along at least one of an axial direction or a circumferential direction within the flow passage.6. The casing of claim 1 , wherein the spring member defines a geometry based on a fin claim 1 , a ligament claim 1 ...

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Номер записи: 28
28-01-2016 дата публикации

Thrust reverser unit

Номер: US20160025037A1
Автор: Luis Daniel TEIA DOS SANTOS MENDES GOMES
Принадлежит: Rolls Royce Deutschland Ltd and Co KG

A thrust reverser unit (TRU) 100 for a gas turbine engine 10 is provided that has first and second cascade elements 110, 120 . In a stowed configuration, both the first and second cascade elements, and the operating mechanism, are located inside the nacelle 40 , meaning that the TRU has no detrimental impact on the flow through the bypass duct 22 . In the deployed configuration, the first cascade element 110 extends across the bypass duct 22 . The first cascade element 110 has flow passages 112 that allow the flow to pass through, redirecting the flow towards the second cascade element 120 . The second cascade element 120 further turns the flow so as to provide decelerating reverse thrust.

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Номер записи: 29
26-01-2017 дата публикации

SYSTEM FOR LOCKING A THRUST REVERSER WITH FLAPS, COMPRISING LOCKS FOR AN INTERMEDIATE OPENING POSITION

Номер: US20170022935A1
Автор: Corentin HUE, DENIS Rodolphe, GUILLOIS Denis, TISSOT Sarah
Принадлежит: AIRCELLE

The present disclosure relates to a system for locking the position of the flaps of a thrust reverser of a turbojet nacelle, said flaps being controlled by actuators, each one swinging about a tranverse pivot in order to partially close off the air stream so as to guide it forwards, further including locks for locking the flaps in an intermediate opening position, between the closed position and the open position. 1. A thrust reverser for a nacelle of a turbojet engine comprising doors controlled by actuators , the doors tilting about a transverse pivot in order to partially close air flow so as to direct the air flow forwards , and comprising a system for locking the position of the doors , wherein the locking system comprises:locks configured to lock an intermediate opening position of the doors, between a closed position and a fully deployed position, where an actuation of said locks blocks said intermediate opening position of the doors without biasing the actuators of the doors in order to hold the intermediate opening position.2. The thrust reverser according to claim 1 , wherein the locking system includes cams rotatably linked to the transverse pivot of the doors claim 1 , on which the locks act.3. The thrust reverser according to claim 2 , wherein each cam includes a spiral-shaped external portion terminated by a step in which the lock fits so as to provide blocking of the cam.4. The thrust reverser according to claim 1 , wherein the locks are guided axially so as to slide and move in a locked position.5. The thrust reverser according to claim 4 , wherein the locking system includes two locks mounted opposite to each other claim 4 , the two locks being controlled simultaneously so as to separate the two locks claim 4 , and each of the two locks functions to lock one of the doors of the thrust reverser.6. The thrust reverser according to claim 5 , wherien the locking system includes a device which claim 5 , in the absence of effort of the actuator of the ...

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Номер записи: 30
25-01-2018 дата публикации

INFRARED SUPPRESSION SYSTEM IN A GAS TURBINE ENGINE

Номер: US20180023411A1
Автор: Froemming Kenneth William, Pesyna Kenneth M.
Принадлежит:

According to one aspect, a system for alignment of vanes to suppress infrared detection in a gas turbine engine is provided. The system includes a first vane disposed on a first component, and a second vane disposed on a second component. The second vane is configured to engage the first vane such that the second component is capable of being positioned proximal to the first component. 1. A system for alignment of vanes to suppress infrared detection in a gas turbine engine , comprising:a first vane disposed on a first component; anda second vane disposed on a second component, wherein the second vane is configured to engage the first vane such that the second component is capable of being positioned proximal to the first component.2. The system of claim 1 , wherein the first and the second vanes are S-vanes claim 1 , wherein the first S-vane comprises a first transition portion and the second S-vane comprises a second transition portion such that the second transition portion is engageable with the first transition portion.3. The system of claim 2 , further comprising:a guide track connected to the first component;a guide roller disposed in the guide track, wherein the guide roller is connected to the second component; andan actuator disposed between the first component and the second component.4. The system of claim 3 , wherein the second component is guided via the guide roller travelling within the guide track in a reciprocating movement caused by the actuator toward and away from the first component.5. The system of claim 3 , wherein the first S-vane is substantially stationary and the second S-vane is movable in relation to the first S-vane.6. The system of claim 5 , wherein the first component is a stationary blocker vane segment and the second component is a telescoping blocker vane segment.7. The system of claim 5 , wherein the second transition portion rotates relative to a common central axis of the first and second components and into the first ...

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Номер записи: 31
24-01-2019 дата публикации

GAS TURBINE ROTOR AND GAS TURBINE GENERATOR

Номер: US20190024511A1
Автор: HONMA Masaya, KIDO OSAO, NISHIYAMA Yoshitsugu, Taguchi Hidetoshi
Принадлежит:

A gas turbine rotor includes a first rotor and a second rotor. The first rotor includes a compressor impeller, a turbine wheel, and a shaft. The turbine wheel has a common rotational axis with the compressor impellor. The shaft connects the compressor impeller to the turbine wheel. The second rotor is an electric generator rotor and defines an inner hollow space. The shaft includes an insertable portion disposed in the inner hollow space of the second rotor. 1. A gas turbine rotor , comprising:a first rotor that includes a compressor impeller, a turbine wheel having a common rotational axis with the compressor impellor, and a shaft connecting the compressor impeller to the turbine wheel; anda second rotor that is an electric generator rotor and defines an inner hollow space, whereinthe shaft includes an insertable portion disposed in the inner hollow space of the second rotor.2. The gas turbine rotor according to claim 1 , whereinthe shaft has a tapered surface having a distance from the common rotational axis increasing in a first direction of axial directions, the axial directions indicating directions in which the common rotational axis extends,the second rotor has a funnel surface exposed to the inner hollow space, the funnel surface having a distance from the common rotational axis increasing in the first direction of the axial directions, andthe tapered surface fits the funnel surface.3. The gas turbine rotor according to claim 1 , further comprising:a nut; anda washer, whereinthe washer has a spring structure elastically deformable in axial directions, the axial directions indicating directions in which the common rotational axis extends,the second rotor has a support surface exposed to the inner hollow space, the support surface extending in directions intersecting the axial directions,the shaft includes a screw and a screw hole,the screw is threaded through the nut and the washer in sequence and into the screw hole in a first direction of the axial ...

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Номер записи: 32
24-01-2019 дата публикации

NON-CONTACT SEAL WITH RESILIENT BIASING ELEMENT(S)

Номер: US20190024522A1
Автор: DAmbruoso Tara L.
Принадлежит:

A seal device includes a plurality of seal shoes, a seal base, a plurality of spring elements and a resilient biasing element. The seal shoes are arranged around an axis. The seal base circumscribes the seal shoes. Each of the spring elements is radially between and connects a respective one of the seal shoes and the seal base. A first of the spring elements includes a first mount, a second mount and a spring beam. The first mount is connected to a first of the seal shoes. The second mount is connected to the seal base. The spring beam connects the first mount to the second mount. The resilient biasing element is radially between and engaged with first and second components of the seal device, where the first component is configured as or otherwise includes the first mount or the second mount. 1. An assembly for rotational equipment , comprising:a seal device comprising a plurality of seal shoes, a seal base, a plurality of spring elements and a resilient biasing element;the seal shoes arranged around an axis in an annular array;the seal base circumscribing the annular array of the seal shoes;each of the spring elements radially between and connecting a respective one of the seal shoes and the seal base, a first of the spring elements including a first mount, a second mount and a spring beam, the first mount connected to a first of the seal shoes, the second mount connected to the seal base, and the spring beam connecting the first mount to the second mount; andthe resilient biasing element radially between and engaged with first and second components of the seal device, the first component comprising the first mount or the second mount.2. The assembly of claim 1 , wherein the resilient biasing element is configured to increase a stiffness of the first of the spring elements.3. The assembly of claim 1 , wherein the resilient biasing element is configured to bias a first portion of the first of the seal shoes radially away from the seal base and a second portion of ...

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Номер записи: 33
24-01-2019 дата публикации

VARIABLE-PITCH VANE ASSEMBLY

Номер: US20190024530A1
Автор: Borja Mark, Hazzard Robert L., Propheter-Hinckley Tracy A., Surace Raymond
Принадлежит:

A variable-pitch vane assembly for a gas turbine engine includes a sync ring, a vane having a vane arm, and a pin installed through the sync ring and through the vane arm. The pin includes an anti-rotation notch located along a pin shaft. An anti-rotation spacer is engaged with the pin at the anti-rotation notch to prevent rotation of the pin. A turbine section of a gas turbine engine includes a turbine rotor and a turbine stator. The turbine stator includes one or more variable-pitch vane assemblies including a sync ring, a vane having a vane arm, and a pin installed through the sync ring and through the vane arm. The pin includes an anti-rotation notch located along a pin shaft. An anti-rotation spacer is engaged with the pin at the anti-rotation notch to prevent rotation of the pin. 1. A variable-pitch vane assembly for a gas turbine engine , comprising:a sync ring;a vane having a vane arm;a pin installed through the sync ring and through the vane arm, the pin including an anti-rotation notch disposed along a pin shaft; andan anti-rotation spacer engaged with the pin at the anti-rotation notch to prevent rotation of the pin.2. The variable-pitch vane assembly of claim 1 , further comprising a bushing disposed between the vane arm and the pin.3. The variable-pitch vane assembly of claim 2 , further comprising a threaded connection between the bushing and the pin.4. The variable-pitch vane assembly of claim 1 , further comprising a threaded connection between the sync ring and the pin.5. The variable-pitch vane assembly of claim 1 , wherein the pin has a recessed hexagonal head.6. The variable-pitch vane assembly of claim 1 , wherein the anti-rotation spacer is disposed between a pin head and the vane arm.7. The variable-pitch vane assembly of claim 1 , further comprising a locking tab washer to retain the anti-rotation spacer at the anti-rotation notch and/or engage with the anti-rotation notch.8. The variable-pitch vane assembly of claim 1 , wherein the anti- ...

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Номер записи: 34
24-01-2019 дата публикации

UNISON RING ASSEMBLY

Номер: US20190024531A1
Автор: LITTLER Graham
Принадлежит: ROLLS-ROYCE PLC

A unison ring assembly for a gas turbine engine has a unison ring and a plurality of levers extending from the unison ring. Each lever has a pin at one end that inserts through a bore of a respective bush mounted in the unison ring. Each bush is formed as separate first and second parts which are mounted to their through-hole by inserting the first part into the through-hole from one side of the unison ring and the second part into the through-hole from the opposing side of the unison ring. Each part has a respective stop which prevents that part from inserting into the through-hole by more than a predetermined amount. When both parts are inserted by their predetermined amounts, their ends join together to form the bush and prevent the parts being retracted from the through-hole. 1. A unison ring assembly for rotating a circumferential row of variable vanes of a gas turbine engine , the assembly having:{'b': '40', 'a unison ring () rotatable about a central axis;'}{'b': '43', 'a plurality of circumferentially spaced levers extending from the unison ring, each lever having a pin () at one end thereof that inserts through a bore of a respective bush mounted in a respective through-hole of the unison ring, thereby rotatably connecting the lever to the unison ring at the pin, and each lever further having an engagement formation at the other end thereof that engages the lever to a spindle projecting from an end of a respective one of the variable vanes, whereby rotation of the unison ring about its central axis causes the levers to rotate the variable vanes about their spindles;'}{'b': 41', '42, 'wherein each bush is formed as separate first () and second () parts which are mounted to their through-hole by inserting a leading end of the first part into the through-hole from one side of the unison ring and a leading end of the second part into the through-hole from the opposing side of the unison ring, each part having a respective stop which prevents that part from ...

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Номер записи: 35
23-01-2020 дата публикации

NON-CONTACT SEAL WITH RESILIENT BIASING ELEMENT(S)

Номер: US20200025006A1
Автор: DAmbruoso Tara L.
Принадлежит:

A seal device includes a plurality of seal shoes, a seal base, a plurality of spring elements and a resilient biasing element. The seal shoes are arranged around an axis. The seal base circumscribes the seal shoes. Each of the spring elements is radially between and connects a respective one of the seal shoes and the seal base. A first of the spring elements includes a first mount, a second mount and a spring beam. The first mount is connected to a first of the seal shoes. The second mount is connected to the seal base. The spring beam connects the first mount to the second mount. The resilient biasing element is radially between and engaged with first and second components of the seal device, where the first component is configured as or otherwise includes the first mount or the second mount. 1. An assembly for rotational equipment , comprising:a seal device comprising a plurality of seal shoes, a seal base, a plurality of spring elements and a coil spring;the seal shoes arranged around an axis in an annular array;the seal base circumscribing the annular array of the seal shoes;each of the spring elements radially between and connecting a respective one of the seal shoes and the seal base, a first of the spring elements including a first mount, a second mount and a spring beam, the first mount connected to a first of the seal shoes, the second mount connected to the seal base, and the spring beam connecting the first mount to the second mount; andthe coil spring radially between and engaged with first and second components of the seal device, the first component comprising the first of the spring elements.2. The assembly of claim 1 , wherein the coil spring is configured to increase a stiffness of the first of the spring elements.3. The assembly of claim 1 , wherein the coil spring is configured to bias a first portion of the first of the seal shoes radially away from the seal base and a second portion of the first of the seal shoes radially towards the seal base.4 ...

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Номер записи: 36
23-01-2020 дата публикации

SEAL ASSEMBLY FOR TURBINE ENGINE COMPONENT

Номер: US20200025007A1
Автор: McCaffrey Michael G.
Принадлежит:

A seal assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a housing mountable to an engine static structure, a seal carrier secured to the housing and configured to be selectively biased from the housing, and a wedge seal secured to the seal carrier. The wedge seal abuts against sealing surfaces of adjacent blade outer air seals in response to movement of the seal carrier relative to the housing. The wedge seal is separate and distinct from the seal carrier. A method of sealing between adjacent components of a gas turbine engine is also disclosed. 1. A seal assembly for a gas turbine engine comprising:a housing mountable to an engine static structure;a seal carrier secured to the housing and configured to be selectively biased from the housing; anda wedge seal secured to the seal carrier, wherein the wedge seal abuts against sealing surfaces of adjacent blade outer air seals in response to movement of the seal carrier relative to the housing, and wherein the wedge seal is separate and distinct from the seal carrier.2. The seal assembly as recited in claim 1 , wherein the wedge seal span across a gap defined by opposed mate faces of the adjacent blade outer air seals.3. The seal assembly as recited in claim 2 , wherein the seal carrier defines a spring cavity receiving a spring that is configured to bias the seal carrier away from the housing.4. The seal assembly as recited in claim 2 , wherein the wedge seal defines a first engagement surface and a second engagement surface joined at an apex.5. The seal assembly as recited in claim 1 , further comprising a spring that is configured to bias the seal carrier away from the housing.6. The seal assembly as recited in claim 1 , wherein the wedge seal defines a first engagement surface and a second engagement surface joined at an apex that faces away from the housing.7. The seal assembly as recited in claim 1 , comprising an insulation member arranged between ...

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Номер записи: 37
23-01-2020 дата публикации

TURBINE SHROUD SEGMENT WITH LOAD DISTRIBUTION SPRINGS

Номер: US20200025013A1
Автор: "OLeary Mark", Brandt Matthew R., Freeman Ted J., Vetters Daniel K.
Принадлежит:

A turbine shroud adapted for use in a gas turbine engine includes a plurality of metallic carrier segments and a plurality of blade track segments mounted to corresponding metallic carrier segments. Cooling air is directed onto the blade track segments to cool the blade track segments when exposed to high temperatures in a gas turbine engine. 1. A turbine shroud segment adapted for use in a gas turbine engine having a central axis , the turbine shroud segment comprisinga carrier segment comprising metallic materials, the carrier segment formed to include an attachment-receiving space,a blade track segment comprising ceramic matrix composite materials, the blade track segment formed to include a runner shaped to extend at least partway around the central axis and an attachment portion that extends radially outward from the runner into the attachment-receiving space formed by the carrier segment,an attachment assembly including a first attachment post that extends from the carrier segment through an attachment hole formed in the attachment portion of the blade track segment, a first attachment support arranged inside a cavity formed by the attachment portion of the blade track segment that is shielded by the runner of the blade track segment from the central axis and coupled to the first attachment post to block withdrawal of the attachment post through the attachment hole, and a first spring member arranged outside of the attachment-receiving space and configured to pull the first attachment support radially outward away from the central axis, wherein the first attachment support contacts the attachment portion of the blade track segment at locations spaced apart from the attachment hole formed in the attachment portion of the blade track segment.2. The turbine shroud segment of claim 1 , wherein the first spring member is arranged radially outward of the carrier segment.3. The turbine shroud segment of claim 2 , wherein the first spring member is a coil spring that ...

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Номер записи: 38
23-01-2020 дата публикации

TURBINE SHROUD SEGMENT WITH SIDE PERIMETER SEAL

Номер: US20200025015A1
Автор: "OLeary Mark", Vetters Daniel K.
Принадлежит:

A turbine shroud segment of a turbine shroud for use in a turbine of a gas turbine engine is disclosed herein. The turbine shroud segment includes a carrier segment and a blade track segment. The carrier segment includes metallic materials. The blade track segment includes ceramic matrix composite materials. A seal is arranged between the carrier segment and the blade track segment. 1. A turbine shroud segment comprisinga carrier segment comprising metallic materials,a blade track segment comprising ceramic matrix composite materials, the blade track segment formed to include a runner shaped to extend partway around a central axis and an attachment box portion that extends radially outward from the runner into an attachment channel formed by the carrier segment, anda seal assembly located radially between the carrier segment and the blade track segment to resist the flow of gas across a seal interface, the seal assembly arranged to extend around a perimeter of the attachment box portion when the blade track segment is viewed radially inward toward the central axis.2. The turbine shroud segment of claim 1 , wherein the seal assembly includes at least one unitary component that extends continuously around the perimeter of the attachment box portion when the blade track segment is viewed radially inward toward the central axis.3. The turbine shroud segment of claim 1 , wherein the seal assembly is formed from a plurality of components arranged around the perimeter of the attachment box portion.4. The turbine shroud segment of claim 3 , wherein the seal assembly includes a first seal member configured to couple to the carrier segment and the attachment box portion such that the first seal member seals between the carrier segment and the attachment box portion around a first portion of the perimeter of the attachment box portion and a second seal member configured to couple to the carrier segment and the attachment box portion such that the second seal member seals ...

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Номер записи: 39
23-01-2020 дата публикации

Flow mixer stiffener ring segmented springs

Номер: US20200025131A1
Автор: Assaf Farah, Daniel Coutu, Daniel SUMMERS-LEPINE, Philippe Boyer
Принадлежит: Pratt and Whitney Canada Corp

A gas turbine engine comprises a main gas path having an inner flow boundary wall and an outer flow boundary wall. A turbine exhaust case inner body defines a portion of the inner flow boundary wall of the main gas path. A lobed exhaust mixer defines a portion of the outer flow boundary wall of the main gas path. A stiffener ring is interconnected to at least a number lobes of the lobed exhaust mixer by a plurality of circumferentially spaced-apart struts extending through the main gas path. The stiffener ring is attached to the turbine exhaust case inner body by flexible features, such as circumferentially spaced-apart spring blades.

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Номер записи: 40
23-01-2020 дата публикации

TRANSLATING CASCADE HIDDEN BLOCKER DOOR THRUST REVERSER

Номер: US20200025134A1
Автор: Gormley Timothy
Принадлежит:

Aspects of the disclosure are directed to a thrust reverser system of an aircraft comprising: a translating cascade sleeve, a blocker door, and a kinematic mechanism configured to actuate the blocker door, the kinematic mechanism comprising: a first link coupled to the translating cascade sleeve, a second link coupled to the first link and a fixed structure, and a third link coupled to the first link and the blocker door. 1. A thrust reverser system of an aircraft comprising:a translating cascade sleeve;a blocker door; and a first link coupled to the translating cascade sleeve;', 'a second link coupled to the first link and a fixed structure; and', 'a third link coupled to the first link and the blocker door., 'a kinematic mechanism configured to actuate the blocker door, the kinematic mechanism comprising2. The thrust reverser system of claim 1 , wherein the blocker door is hinged to the translating cascade sleeve.3. The thrust reverser system of claim 1 , wherein the first link passes through a set of cascades via a slot created through the cascades irrespective of whether the thrust reverser system is operated in a stowed state or a deployed state.4. The thrust reverser system of claim 1 , wherein loads experienced by the blocker door are translated through the third link and the first link to the translating cascade sleeve.5. The thrust reverser system of claim 4 , further comprising:a plurality of ribs,wherein the loads are translated from the translating cascade sleeve to the ribs in at least one of a forward direction or an aft direction.6. The thrust reverser system of claim 1 , wherein the fixed structure includes a clevis and a pin.7. The thrust reverser system of claim 1 , wherein the second link is configured to rotate in a single direction when the thrust reverser system transitions from a stowed state to a deployed state.8. The thrust reverser system of claim 1 , further comprising:a primary sleeve.9. The thrust reverser system of claim 8 , wherein the ...

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Номер записи: 41
04-02-2016 дата публикации

VARIABLE-PITCH ROTOR WITH REMOTE COUNTERWEIGHTS

Номер: US20160032740A1
Автор: Bradley Donald Albert, Miller Brandon Wayne, Niergarth Daniel Alan
Принадлежит:

A pitch control mechanism includes: a rotor structure configured for rotation about a longitudinal axis; a row of blades carried by the rotor structure, each blade having an airfoil and a trunnion mounted for pivoting movement relative to the rotor structure, about a trunnion axis which is perpendicular to the longitudinal axis; a unison ring interconnecting the blades; an actuator connected to the unison ring and the rotor structure, operable to move the unison ring relative to the rotor structure; at least one moveable counterweight carried by the rotor structure, remote from the blades; and an interconnection between the blades and the counterweight, such that movement of the counterweight causes a change in the pitch angle of the blades. 1. A pitch control mechanism , comprising:a rotor structure configured for rotation about a longitudinal axis;a row of blades carried by the rotor structure, each blade having an airfoil and a trunnion mounted for pivoting movement relative to the rotor structure, about a trunnion axis which is perpendicular to the longitudinal axis;a unison ring interconnecting the blades;an actuator connected to the unison ring and the rotor structure, operable to move the unison ring relative to the rotor structure;at least one moveable counterweight carried by the rotor structure, remote from the blades; andan interconnection between the blades and the counterweight, such that movement of the counterweight causes a change in the pitch angle of the blades.2. The pitch control mechanism of wherein the actuator is configured to produce rotary movement between the rotor structure and the unison ring.3. The pitch control mechanism of wherein the unison ring and counterweights are interconnected by gears.4. The pitch control mechanism of wherein the rotor structure carries an array of counterweight assemblies each including: a pinion shaft claim 1 , a pinion gear claim 1 , and a counterweight with an offset mass.5. The pitch control mechanism of ...

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Номер записи: 42
04-02-2016 дата публикации

MACHINED VANE ARM OF A VARIABLE VANE ACTUATION SYSTEM

Номер: US20160032759A1
Автор: Gasmen Eugene C., Pudvah Bernard W., St. Mary Christopher, Wiecko Stanley
Принадлежит:

An exemplary variable vane actuation system includes, among other things, a vane arm with a vane stem contact surface and a radially outward facing surface. The vane stem contact surface is to contact a vane stem of a variable vane and thereby actuate the variable vane about a radially extending axis. The vane stem contact surface is angled relative to both the radially extending axis and the radially outward facing surface. 1. A variable vane actuation system , comprising:a vane arm with at least one vane stem contact surface and a radially outward facing surface, the at least one vane stem contact surface to contact a vane stem of a variable vane and thereby actuate the variable vane about a radially extending axis, the at least one vane stem contact surface angled relative to both the radially extending axis and the radially outward facing surface.2. The system of claim 1 , including an aperture extending through the radially outward facing surface to receive the vane stem claim 1 , a least a portion of the aperture having a non-circular cross-sectional profile.3. The system of claim 2 , wherein the aperture comprises a first axial section and a second axial section claim 2 , the first axial section having a generally oval-shaped cross sectional profile claim 2 , the second axial section having a generally circular-shaped cross-sectional profile.4. The system of claim 2 , wherein the at least one vane stem contact surface comprises a first vane stem contact surface and a second vane stem contact surface claim 2 , the aperture positioned between the first and second vane stem contact surfaces.5. The system of claim 1 , wherein the at least one vane stem contact surface is a machined surface.6. The system of claim 5 , wherein the at least one vane stem contact surface is a milled surface.7. The system of claim 1 , wherein the vane arm is continuous radially between the at least one vane stem contact surface and the radially outward facing surface.8. The system of ...

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Номер записи: 43
04-02-2016 дата публикации

COMPONENT ARRANGEMENT OF A GAS TURBINE

Номер: US20160032777A1
Автор: Sasse Stefan
Принадлежит:

The present invention relates to a component arrangement of a gas turbine, this arrangement having a first gas turbine component, in particular a first wall segment (10) of a gas turbine duct casing; a second gas turbine component that can be joined thereto, in particular a second wall segment (20) of a gas turbine duct casing, with a flange (21), which is particularly bent; and an eccentric clamping element (30), which is mounted rotatably about an axis of rotation (A) on the first gas turbine component and has an eccentric contour portion (31), whose radial distance (r) to the axis of rotation varies by an angle (φ) about the axis of rotation, in order to press the flange (21) of the second gas turbine component (20) against the first gas turbine component, in particular to clamp it between the first gas turbine component and the eccentric contour portion (31). 1. A component arrangement of a gas turbine , comprising:{'b': '10', 'a first wall segment () of a gas turbine duct casing;'}{'b': 20', '21, 'a second wall segment () of a gas turbine duct casing that can be joined thereto with a flange (), which is bent; and'}{'b': 30', '31', '21', '20', '10', '10', '31, 'an eccentric clamping element () that is mounted rotatably about an axis of rotation (A) on the first gas turbine component and has an eccentric contour portion (), whose radial distance (r) to the axis of rotation varies by an angle (φ) about the axis of rotation, in order to press the flange () of the second gas turbine component () against the first gas turbine component (), to clamp it between the first gas turbine component () and the eccentric contour portion ().'}23032311110. The gas turbine component arrangement according to claim 1 , wherein the eccentric clamping element () has a shaft () which is joined to the eccentric contour portion () claim 1 , ire particular and resistant to rotation claim 1 , this shaft being mounted rotatably about the axis of rotation (A) in a borehole () of the first ...

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Номер записи: 44
01-02-2018 дата публикации

DEVICE FOR THE INDIVIDUAL ADJUSTMENT OF A PLURALITY OF VARIABLE-PITCH RADIAL STATOR VANES IN A TURBOMACHINE

Номер: US20180030849A1
Автор: Bordoni Nils
Принадлежит: SAFRAN AIRCRAFT ENGINES

A device for adjusting the pitch of at least one annular row of stator vanes for a turbine engine module. The device includes a first control ring mounted to rotate freely about an axis of the turbine engine. The device also includes connecting rods for connecting the first control ring to the vanes and a second control ring mounted to rotate freely about the axis. Each vane of the at least one row is simultaneously connected to the first and second control rings by a set of at least two connecting rods. The device is suitable for use as part of a module and is suitable for use in a turbine engine. 1. Device for adjusting the pitch of at least one annular row of stator vanes for a turbine engine module , the device comprising:a first control ring mounted to rotate about an axis of the turbine engineconnecting rods configured for connecting said first control ring to the vanes; anda second control ring mounted to rotate about said axis wherein each vane of said at least one row is simultaneously connected to said first and second control rings by a set of at least two connecting rods for individually adjusting the pitch of said vanes.2. Device according to claim 1 , wherein claim 1 , with the connecting rods of said set being articulated successively pairwise about substantially radial articulation shafts claim 1 , a first connecting rod is mounted to rotate about a first and a second substantially radial pivot shaft claim 1 , the first pivot shaft being mounted on said first control ring and the second pivot shaft being configured to be positioned independent of the position of the first control ring claim 1 , and a second connecting rod is pivotally mounted on the first connecting rod about a first articulation shaft located at a first distance from said second pivot shaft that is determined for each vane.3. Device according to claim 2 , wherein said first articulation shaft is located between the first and second pivot shafts.4. Device according to claim 2 , ...

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Номер записи: 45
31-01-2019 дата публикации

NACELLE

Номер: US20190031357A1
Автор: Sidelkovskiy Dmitriy B., Zysman Steven H.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A nacelle may comprise an inlet cowling defining an inlet of the nacelle, wherein the inlet cowling comprises an inlet cowling maximum point and an inlet cowling aft edge, wherein the inlet cowling aft edge comprises an aft edge length; and a boat tail cowling disposed aft of the inlet cowling, wherein the boat tail cowling comprises a boat tail cowling forward edge having a forward edge length, and wherein the boat tail cowling forward edge is disposed adjacent to the inlet cowling aft edge. The forward edge length may be smaller than the aft edge length, forming a step, the step being defined by a portion of the inlet cowling aft edge that is radially outward of the boat tail cowling forward edge. 1. A nacelle encircling an engine about an axis of rotation , comprising:an inlet cowling defining an inlet of the nacelle, wherein the inlet cowling comprises an inlet cowling maximum point and an inlet cowling aft edge, wherein the inlet cowling aft edge comprises an aft edge length; anda boat tail cowling disposed aft of the inlet cowling, wherein the boat tail cowling comprises a boat tail cowling forward edge having a forward edge length, wherein the boat tail cowling forward edge is disposed adjacent to the inlet cowling aft edge,wherein the forward edge length is smaller than the aft edge length, forming a step, the step being defined by a portion of the inlet cowling aft edge that is radially outward of the boat tail cowling forward edge.2. The nacelle of claim 1 , wherein the step is disposed aft of the inlet cowling maximum point.3. The nacelle of claim 1 , further comprising a thrust reverser coupled to the inlet cowling and the boat tail cowling claim 1 , wherein the boat tail cowling is configured to translate in an aft direction claim 1 , wherein the boat tail cowling forward edge is disposed adjacent to the inlet cowling aft edge when the boat tail cowling is in a forward-most configuration claim 1 , and wherein the thrust reverser is configured to be in ...

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Номер записи: 46
17-02-2022 дата публикации

DISTRIBUTOR FOR A TURBOMACHINE RADIAL TURBINE, TURBOMACHINE COMPRISING SUCH A DISTRIBUTOR AND AIR CONDITIONING SYSTEM COMPRISING SUCH A TURBOMACHINE

Номер: US20220049618A1
Автор: BINDER Nicolas, CROS Sandrine, FOURAGNAN James, LAURIAU Pierre-Thomas, ROUMEAS Mathieu
Принадлежит:

The invention relates to a distributor for a turbomachine radial turbine, comprising an annular grill () extending about a central axis () and comprising a plurality of variable-pitch blades (), defining between them an air passage cross section, characterized in that each blade is rotatably mounted about a pivot shaft (), itself moveable in a translation direction, comprising at least one radial component, such that said blade may, upon actuation of control means (), be pivoted about the pivot shaft and/or moved in relation to the central axis in said translation direction so as to be able to modify the air passage cross section by respectively controlling the metal angle (α) and the radial spacing (ΔR). 1. A distributor of a turbomachine radial turbine comprising a rotor equipped with vanes adapted to be rotated about a central axis , said distributor comprising an annular grill extending about said central axis , intended to be arranged on the periphery of said rotor and comprising a plurality of variable-pitch blades arranged about said central axis delimiting between the blades a passage cross section of an air stream from said distributor to said rotor , each of the variable-pitch blades further extending in a main direction and having a leading edge and a trailing edge intended to be arranged opposite the leading edge on the rotor vanes ,wherein each variable-pitch blade is mounted to be rotatably movable about a pivot shaft extending parallel to the central axis), the pivot shaft being movable relative to the central axis in a translation direction comprising at least one radial component, so that the blade, upon actuation of control means, pivotsed about the pivot shaft and/or moves relative to the central axis in said translation direction so as to be able to modify said passage cross section of the air stream from upstream to downstream of the distributor, by controlling respectively a metal angle formed between the main direction of the blade and the ...

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Номер записи: 47
17-02-2022 дата публикации

DEFLECTION LIMITER FOR A GAS TURBINE ENGINE

Номер: US20220049629A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит:

A gas turbine engine includes a turbine section that includes a fan drive turbine. A geared architecture includes a sun gear in driving engagement with the fan drive turbine. A plurality of planet gears surrounds the sun gear. A ring gear surrounds the plurality of planet gears. A deflection limiter mechanically attaches the ring gear to an engine static structure. The deflection limiter includes a first support fixed to the ring gear that has a first interlocking feature and a second support fixed to the engine static structure that has a second interlocking feature. The first and second interlocking features define at least one of a radial clearance of between 0.005 inches (0.127 mm) and 0.080 inches (2.032 mm) or a circumferential clearance of between 0.005 inches (0.127 mm) and 0.250 inches (2.032 mm). A fan section includes a plurality of fan blades in driving engagement with the geared architecture through a fan drive shaft. 1. A gas turbine engine comprising:a turbine section including a fan drive turbine;{'claim-text': ['a sun gear in driving engagement with the fan drive turbine;', 'a plurality of planet gears surrounding the sun gear; and', 'a ring gear surrounding the plurality of planet gears;'], '#text': 'a geared architecture including:'}a deflection limiter mechanically attaching the ring gear to an engine static structure, wherein the deflection limiter includes a first support fixed to the ring gear having a first interlocking feature and a second support fixed to the engine static structure having a second interlocking feature, wherein the first and second interlocking features define at least one of a radial clearance of between 0.005 inches (0.127 mm) and 0.080 inches (2.032 mm) or a circumferential clearance of between 0.005 inches (0.127 mm) and 0.250 inches (2.032 mm); anda fan section including a plurality of fan blades in driving engagement with the geared architecture through a fan drive shaft.2. The gas turbine engine of claim 1 , wherein ...

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Номер записи: 48
30-01-2020 дата публикации

Mechanism for indicating position of a latch mechanism for an aircraft nacelle

Номер: US20200031484A1
Автор: Kushy Sachdeva, Kyle A. Thomas, Matthew Joseph Woerly, Matthew Senesac, Michael Roethle, Rhoda Rebecca Toliver
Принадлежит: Boeing Co

An aircraft nacelle is disclosed and includes a first fan cowl and a second fan cowl. An opening is defined in an exterior surface of the first fan cowl. The aircraft nacelle also includes a blade, a latch mechanism, an arm, and a linkage assembly operably connecting the arm to the blade. The blade is moveable along between a stowed position and a deployed position. The latch mechanism is moveable between a latched position an unlatched position where a portion of the latch mechanism obstructs a path of movement of the blade. The arm is unable to move from an extended position where a portion of the arm extends through the opening in the exterior surface of the aircraft nacelle into a retracted position within an interior volume of the aircraft nacelle when the blade is in the deployed position and the latch mechanism is in the unlatched position.

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Номер записи: 49
04-02-2021 дата публикации

GIMBAL ASSEMBLY AND MANUFACTURE THEREOF

Номер: US20210033044A1
Автор: Hawksworth Andrew
Принадлежит:

A gimbal assembly comprises a body, comprising at least one pivot boss projecting radially outwards along a first pivot axis (V) from an outer surface of the body; a gimbal, comprising an outer case surrounding the body and at least one hole projecting radially outwards along a second pivot axis (H) to receive a pivot pin to pivotally couple the gimbal to a fixed structure. The second pivot axis (H) is perpendicular to the first pivot axis (V) and the outer case is formed at least partially from carbon fibre-reinforced polymer matrix composite material. The outer case comprises at least one cavity on its inner surface in which the at least one pivot boss is located to pivotally couple the gimbal to the body. 1. A method of making a gimbal assembly comprising:providing a body comprising at least one pivot boss projecting radially outwards along a first pivot axis from an outer surface of the body;providing a spacer around the outer surface of the body and at least one fixing member to hold the spacer in a fixed position relative to the body, the spacer being shaped to allow the at least one pivot boss to extend therethrough;winding polymer-impregnated carbon fibres onto the spacer and around the at least one pivot boss so as to form an outer case of carbon fibre-reinforced polymer matrix composite material, the outer case surrounding the body with at least one hole projecting radially outwards along a second pivot axis, the second pivot axis being perpendicular to the first pivot axis; andremoving the spacer to leave a gimbal comprising the outer case with at least one cavity on its inner surface, wherein the at least one pivot boss is located in the at least one cavity and is thereby configured to pivotally couple the outer case to the body such that the body can pivot relative to the outer case about the first pivot axis and the gimbal assembly can pivot about the second pivot axis.2. A method of making a gimbal assembly according to claim 1 , wherein the at least ...

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Номер записи: 50
11-02-2016 дата публикации

TRANSVERSE MOUNTED ACCESSORY GEARBOX

Номер: US20160040601A1
Автор: Crestin Michael Stephen, Frost Cristopher, Hayden Mariee, Hollingshead Doug, Hwang Mingfong, Laffan James, Schmittenberg Marc
Принадлежит: HONEYWELL INTERNATIONAL INC.

An accessory gear box for a gas turbine engine having a drive shaft with a rotational axis and a tower shaft coupled to the drive shaft is provided. The accessory gear box includes a first plurality of gears arranged, which extend along a first axis substantially parallel to the rotational axis of the drive shaft. The accessory gear box includes a second plurality of gears, which extend along a second axis. The accessory gear box includes a first shaft, with one of the first plurality of gears coupled to the first shaft, and one of the second plurality of gears coupled to a second shaft. The one of the second plurality of gears coupled to the first shaft includes a first engagement surface and a second engagement surface, and the second engagement surface is coupled to another one of the second plurality of gears to drive the second shaft. 1. An accessory gear box for a gas turbine engine having a drive shaft with a rotational axis and a tower shaft coupled to the drive shaft , the accessory gear box comprising:a first plurality of gears arranged within the accessory gear box, the first plurality of gears extending along a first axis substantially parallel to the rotational axis of the drive shaft;a second plurality of gears arranged within the accessory gear box, the second plurality of gears extending along a second axis, the second axis offset from and substantially parallel to the first axis; anda first shaft, with one of the first plurality of gears coupled to the first shaft, and one of the second plurality of gears coupled to a second shaft,wherein the one of the second plurality of gears coupled to the first shaft includes a first engagement surface and a second engagement surface, the first engagement surface to engage the tower shaft and the second engagement surface is coupled to another one of the second plurality of gears to drive the second shaft.2. The accessory gear box of claim 1 , wherein the one of the first plurality of gears coupled to the first ...

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Номер записи: 51
09-02-2017 дата публикации

TURBINE SHROUD SEGMENT WITH SIDE PERIMETER SEAL

Номер: US20170037740A1
Автор: "OLeary Mark", Vetters Daniel K.
Принадлежит:

A turbine shroud segment of a turbine shroud for use in a turbine of a gas turbine engine is disclosed herein. The turbine shroud segment includes a carrier segment and a blade track segment. The carrier segment includes metallic materials. The blade track segment includes ceramic matrix composite materials. A seal is arranged between the carrier segment and the blade track segment. 1. A turbine shroud segment comprisinga carrier segment comprising metallic materials,a blade track segment comprising ceramic matrix composite materials, the blade track segment formed to include a runner shaped to extend partway around a central axis and an attachment box portion that extends radially outward from the runner into an attachment channel formed by the carrier segment, anda seal assembly configured to couple to the carrier segment and the attachment box portion of the blade track segment such that the seal assembly seals between the carrier segment and the blade track segment around a perimeter of the attachment box portion when the seal assembly is coupled to the carrier segment and the attachment box portion.2. The turbine shroud segment of claim 1 , wherein the seal assembly includes at least one unitary component that extends continuously around the perimeter of the attachment box portion.3. The turbine shroud segment of claim 1 , wherein the seal assembly is formed from a plurality of components arranged around the perimeter of the attachment box portion.4. The turbine shroud segment of claim 3 , wherein the seal assembly includes a first seal member configured to couple to the carrier segment and the attachment box portion such that the first seal member seals between the carrier segment and the attachment box portion around a first portion of the perimeter of the attachment box portion and a second seal member configured to couple to the carrier segment and the attachment box portion such that the second seal member seals between the carrier segment and the ...

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Номер записи: 52
24-02-2022 дата публикации

VARIABLE CYCLE COMPENSATION IN A GAS TURBINE ENGINE

Номер: US20220055763A1
Автор: BENGEA SORIN, Terwilliger Neil
Принадлежит:

An aspect includes a method of variable cycle compensation in a gas turbine engine. An electric component can be adjusted to compensate for a power change induced by an actuation system by operating the electric component as an electric motor to compensate for an increase in power absorption or a decrease in power production of a turbomachinery of the gas turbine engine. The actuation system is configured to adjust a variable cycle of the turbomachinery by adjusting power absorption or power production, and the electric component can be configured to add or subtract torque to a shaft of the gas turbine engine. The electric component can be operated as an electric generator to compensate for an increase in power production or a decrease in power absorption of the turbomachinery. 1. A method of variable cycle compensation in a gas turbine engine , the method comprising:adjusting, by a controller, an electric component to compensate for a power change induced by an actuation system by operating the electric component as an electric motor to compensate for an increase in power absorption or a decrease in power production of a turbomachinery of the gas turbine engine, wherein the turbomachinery comprises at least one compressor section and at least one turbine section operably coupled to a shaft of the gas turbine engine, the actuation system is configured to adjust a variable cycle of the turbomachinery by adjusting power absorption or power production, and the electric component is configured to add or subtract torque to the shaft; andoperating the electric component as an electric generator to compensate for an increase in power production or a decrease in power absorption of the turbomachinery.2. The method of claim 1 , further comprising:receiving a control input; anddetermining a plurality of current operating conditions of the gas turbine engine.3. The method of claim 2 , further comprising:calculating a plurality of commands to a plurality of power production and ...

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Номер записи: 53
07-02-2019 дата публикации

AIRCRAFT PROPULSION UNIT

Номер: US20190040816A1
Автор: ANCUTA Carmen Gina, KERBLER Olivier, LEMARCHAND Kevin Morgane, SERISET Mélody
Принадлежит:

An aircraft propulsion assembly including a bypass turbojet having a stationary inter-compressor casing positioned upstream from a space between passages, a nacelle including at its downstream end an inner wall that defines the outside of the space between passages and the inside of the bypass stream flow passage, and an outer wall arranged around the inner wall and that defines the outside of the bypass stream flow passage, at least a portion of the inner wall being suitable for moving between a maintenance position and a working position, and a thrust-reverser including a downstream element of the outer wall that is movable in translation between a retracted position and a thrust-reversal position in which it allows bypass stream deflector elements to be deployed in a radial direction, the deployment of the stream deflector elements being controlled by guide rods. 1. An aircraft propulsion assembly comprising:a bypass turbojet having a space between passages separating a core stream flow passage and a bypass stream flow passage, and a stationary inter-compressor casing positioned upstream from said space between passages;a nacelle arranged around the turbojet and including at its downstream end an inner annular wall that defines the outside of the space between passages and the inside of the bypass stream flow passage, and an outer annular wall arranged around the inner wall and that defines the outside of the bypass stream flow passage, at least a portion of the inner wall being suitable for moving between a maintenance position in which it uncovers the space between passages at least in part and a working position in which it masks the space between passages; anda thrust-reverser comprising a downstream element of the outer wall that is movable in translation between a retracted position wherein it defines the outside of the bypass stream flow passage and a thrust-reversal position wherein bypass stream deflector elements are deployed in a radial direction, the ...

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Номер записи: 54
06-02-2020 дата публикации

Variable Vane Actuation System for a Turbo Machine

Номер: US20200040759A1
Автор: Mielke Mark Joseph
Принадлежит:

A system for operating a turbo machine is generally provided, the system including a variable vane assembly including a first plurality of vanes coupled together by a first unison member and a second plurality of vanes coupled together by a second unison member. An actuation system is coupled to each of the first unison member and the second unison member and a solenoid actuator is coupled to the variable vane assembly and the actuation system. 1. A system for operating a turbo machine , the system comprising:a variable vane assembly comprising a first plurality of vanes coupled together by a first unison member and a second plurality of vanes coupled together by a second unison member;an actuation system coupled to each of the first unison member and the second unison member; anda solenoid actuator coupled to the variable vane assembly and the actuation system.2. The system of claim 1 , wherein the solenoid actuator is coupled to one or more of the unison members of the variable vane assembly.3. The system of claim 1 , wherein the actuation system comprises a coupling member claim 1 , wherein the coupling member is coupled to each of the unison members.4. The system of claim 3 , wherein the solenoid actuator is coupled to the coupling member of the actuation system.5. The system of claim 1 , wherein the solenoid actuator is coupled to the first unison member of the variable vane assembly.6. The system of claim 1 , wherein the first plurality of vanes comprises an inlet guide vane.7. The system of claim 1 , wherein the solenoid actuator defines a two-position solenoid actuator.8. The system of claim 1 , wherein the solenoid actuator defines a proportional solenoid actuator9. The system of claim 1 , the system further comprising:a controller configured to perform operations, the operations comprising:displacing the first plurality of vanes from a first rotational angle corresponding to a first position of the solenoid actuator to a second rotational angle ...

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Номер записи: 55
18-02-2021 дата публикации

HIDDEN LINK SYSTEM BLOCKER DOOR

Номер: US20210047984A1
Автор: Gormley Timothy
Принадлежит: Rohr, Inc.

A system for deploying a blocker door of a nacelle includes a master link configured to be coupled to a fixed structure of the nacelle and a master crank pivotally attached to the master link. The system further includes a first door crank and a first door link pivotally coupled to the first door crank. The system further includes a first blocker door coupled to the first door link and a first driveshaft coupled to the master crank and to the first door crank and configured to transfer motion from the master crank to the first door crank such that aft translation of a translating sleeve of the nacelle drives the master crank via the master link, which drives the first door link via the first driveshaft and the first door crank to move the first blocker door into a bypass air duct defined by the nacelle. 1. A system for deploying a blocker door of a nacelle , comprising:a gear box configured to be coupled to a translating sleeve of the nacelle and to convert linear motion of the translating sleeve to rotational motion;a first door crank;a first door link pivotally attached to the first door crank;a first blocker door coupled to the first door link; anda first driveshaft coupled to the gear box and to the first door crank and configured to transfer the rotational motion from the gear box to the first door crank, such that translation of the translating sleeve in an aft direction drives the first door link via the first door crank to move the first blocker door into a bypass air duct defined by the nacelle.2. The system of claim 1 , wherein the gear box claim 1 , the first door crank claim 1 , and the first door link are each configured to be located radially outward from the bypass air duct in response to the translating sleeve being located at its forward-most position.3. The system of claim 1 , wherein the nacelle includes at least one of a translating cascade or a fixed cascade. This application is a division of application serial number U.S. Ser. No. 16/130,759 ...

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Номер записи: 56
16-02-2017 дата публикации

SYSTEM AND METHOD FOR SUPPORTING A TURBINE SHROUD

Номер: US20170044922A1
Автор: Shapiro Jason David
Принадлежит:

In one aspect the present subject matter is directed to a system for supporting a turbine shroud. The system includes a shroud support at least partially defining a first piston sleeve and a piston assembly having a first piston head disposed within the first piston sleeve and a second piston head coupled to the first piston head. The first piston head is slideably engaged with an inner surface of the first piston sleeve. The second piston head is slideably engaged with an inner surface of a second piston sleeve. The system also includes a turbine shroud that is fixedly connected to the piston assembly and that extends radially inwardly from the shroud support. The piston assembly provides for radially inward and radially outward movement of the turbine shroud in response to a change in a radial force applied to a hot side surface of the turbine shroud. 1. A system for supporting a turbine shroud , the system comprising:a shroud support, the shroud support at least partially defining a first piston sleeve therein;a piston assembly having a first piston head disposed within the first piston sleeve and a second piston head coupled to the first piston head, wherein the first piston head is slideably engaged with an inner surface of the first piston sleeve, wherein the second piston head is slideably engaged with an inner surface of a second piston sleeve; anda turbine shroud fixedly connected to the piston assembly and extending radially inwardly from the shroud support, wherein the piston assembly provides for radially inward and radially outward movement of the turbine shroud in response to a radially outward force applied to a hot side surface of the turbine shroud.2. The system as in claim 1 , wherein a first pressure chamber is at least partially defined within the shroud support between the first piston head and the second piston head claim 1 , wherein the first pressure chamber is pressurized at a first pressure to provide a first radial force on a backside ...

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Номер записи: 57
16-02-2017 дата публикации

TURBINE SHROUD ASSEMBLY AND METHOD FOR LOADING

Номер: US20170044923A1
Автор: Hafner Matthew Troy, Morgan Victor John, ROBERTS Frederic Woodrow
Принадлежит:

A turbine shroud assembly is disclosed including an inner shroud having a surface adjacent to a hot gas path, an outer shroud, and a biasing apparatus. The biasing apparatus is arranged and disposed to bias the inner shroud in a direction away from the hot gas path, loading the inner shroud to the outer shroud. In another embodiment, the biasing apparatus is a springless biasing apparatus including at least one bellows, at least one thrust piston, or a combination of at least one bellows and at least one thrust piston. A method for loading the turbine shroud assembly is disclosed including biasing the inner shroud having a surface adjacent to a hot gas path in a direction away from the hot gas path toward the outer shroud, wherein biasing the inner shroud includes a biasing force exerted by the biasing apparatus. 1. A turbine shroud assembly , comprising:an inner shroud having a surface adjacent to a hot gas path;an outer shroud; anda biasing apparatus,wherein the biasing apparatus is arranged and disposed to bias the inner shroud in a direction away from the hot gas path, loading the inner shroud to the outer shroud.2. The turbine shroud assembly of claim 1 , wherein the biasing apparatus is a springless biasing apparatus.3. The turbine shroud assembly of claim 1 , wherein the biasing apparatus is driven by a pressurized fluid.4. The turbine shroud assembly of claim 1 , wherein the biasing apparatus includes at least one bellows.5. The turbine shroud assembly of claim 4 , wherein the at least one bellows includes a first end attached to the outer shroud claim 4 , and a second end configured to expand away from the hot gas path in response to an increased internal pressure within the at least one bellows claim 4 , the second end connecting to the inner shroud and configured to exert a biasing force on the inner shroud.6. The turbine shroud assembly of claim 5 , wherein the second end of the at least one bellows is connected to the inner shroud by an attachment ...

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Номер записи: 58
15-02-2018 дата публикации

HINGED SEAL USING WIRE MESH

Номер: US20180045064A1
Автор: Guinn Witt C., McMahon Shawn M.
Принадлежит:

A seal segment includes a flapper seal, a wire mesh acting as a hinge for the flapper seal, and a spring. The flapper seal has a first surface and a second surface. The wire mesh includes a first section connected to the first surface of the flapper seal and a second section connected to the first section. The spring includes a first end that is in contact with the first section of the wire mesh to apply pressure to the first surface of the flapper seal and a second end adjacent the second section of the wire mesh. 1. A seal segment comprising:a flapper seal having a first surface and a second surface; a first section connected to and extending across the first surface of the flapper seal; and', 'a second section connected to the first section; and, 'a wire mesh acting as a hinge for the flapper seal, the wire mesh including a first end positioned to apply pressure to the first surface of the flapper seal; and', 'a second end adjacent the second section of the wire mesh., 'a spring comprising2. The seal segment of claim 1 , further comprising a fastener connecting the wire mesh to the flapper seal.3. The seal segment of claim 1 , further comprising a backing sheet positioned between the first section of the wire mesh and the first end of the spring.4. The seal segment of claim 3 , further comprising a fastener connecting the wire mesh and the backing sheet to the flapper seal.5. The seal segment of claim 3 , wherein the backing sheet is a metal claim 3 , ceramic claim 3 , or composite material.6. The seal segment of claim 1 , further comprising a bracket including:a first end connected to the second section of the wire mesh; anda second end connected to the second end of the spring.7. The seal segment of claim 6 , further comprising:a first fastener connecting the wire mesh and the first end of the bracket; anda second fastener connecting the spring and the second end of the bracket.8. The seal segment of claim 1 , wherein the second section of the wire mesh is ...

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Номер записи: 59
15-02-2018 дата публикации

DEVICE OF A TURBOMACHINE FOR ACTUATING A SETTING DEVICE AND TURBOMACHINE WITH SUCH A DEVICE

Номер: US20180045069A1
Автор: Todorovic Predrag
Принадлежит:

A device of a turbomachine actuates a setting device to vary a flow cross-section of a flow channel of the turbomachine passable by a fluid flow. The device includes a displacement mechanism having an adjustable actuation appliance and couplable with the setting device, and a drive device for displacing the actuation appliance. The displacement mechanism has a centrifugal force appliance displaceable between a basic position and a maximally displaced working position depending on a number of revolutions of the drive device, wherein the actuation appliance is displaceable depending on the position of the centrifugal force appliance with respect to the drive device. A reset device applies a force to the centrifugal force appliance by which the centrifugal force appliance is pressed in the direction of its basic position. 1. The device of a turbomachine for actuating a setting device by means of which a flow cross-section of a flow channel of the turbomachine through which a fluid can flow may be varied , with a displacement mechanism that has an adjustable actuation appliance and can be coupled with a setting device , and with an drive device embodied for displacing the actuation appliance , wherein in that the displacement mechanism has a centrifugal force appliance that is displaceable between a basic position and a maximally displaced working position depending on a number of revolutions of the drive device , wherein the actuation appliance is displaceable depending on the position of the centrifugal force appliance with respect to the drive device , and wherein a reset device is provided that is embodied for applying a force to the centrifugal force appliance , by which the centrifugal force appliance is pressed in the direction of its basic position.2. The device according to claim 1 , wherein the centrifugal force appliance is mounted so as to be axially fixated and rotatable with respect to the actuation appliance in a first area claim 1 , and so as to be ...

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Номер записи: 60
15-02-2018 дата публикации

AIRCRAFT GAS TURBINE ENGINE NACELLE

Номер: US20180045140A1
Автор: BOND Jonathan M.
Принадлежит: ROLLS-ROYCE PLC

An aircraft gas turbine engine nacelle comprising a thrust reversal arrangement. The thrust reversal arrangement comprises at least one thrust reverser cascade box comprising a fixed structure and a plurality of thrust reverser vanes configured to direct bypass air forwardly. At least a subset of the thrust reverser vanes are pivotable and/or translatable relative to the fixed structure between a deployed position, in which the vanes define a first volume, and a stowed position, in which the pivotable vanes define a second, smaller volume. 1. An aircraft gas turbine engine nacelle comprising a thrust reversal arrangement , the thrust reversal arrangement comprising:at least one thrust reverser cascade box comprising a fixed structure and a plurality of thrust reverser vanes configured to direct bypass air forwardly;at least a subset of the thrust reverser vanes being pivotable and/or translatable relative to the fixed structure between a deployed position, in which the pivotable and/or translatable vanes define a first volume, and a stowed position, in which the pivotable and/or translatable vanes define a second volume smaller than the volume.2. A nacelle according to claim 1 , wherein the vanes are pivotable and define a first radial extent when in the deployed position claim 1 , and a second radial extent when in the stowed position claim 1 , the second radial extent being smaller than the first radial extent.3. A nacelle according to claim 1 , wherein the vanes are translatable and define a first axial extent when in the deployed position claim 1 , and a second axial extent when in the stowed position claim 1 , the second axial extent being smaller than the first axial extent.4. A nacelle according to claim 1 , wherein a further subset of the vanes comprises fixed vanes which are fixed relative to the cascade box fixed structure.5. A nacelle according to claim 2 , wherein the pivotable vanes are located downstream in fan flow relative to the fixed vanes.6. A ...

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Номер записи: 61
14-02-2019 дата публикации

BOAS SEGMENTED HEAT SHIELD

Номер: US20190048736A1
Автор: Griffin David Richard, Hagan Benjamin F., McCaffrey Michael G.
Принадлежит:

A seal assembly includes a seal arc segment defining first and second seal supports with a carriage defining first and second support members. The first support member supports the seal arc segment in a first ramped interface, and the second support member supports the seal arc segment in a second ramped interface. A spring is configured to bias the seal arc segment axially. A heat shield is radially inward of the spring. 1. A seal assembly , comprising:a seal arc segment;a carriage supporting the seal arc segment;a spring configured to bias the seal arc segment axially; anda heat shield including a portion radially inward of the spring and a radially-extending leg disposed axially between the spring and the seal arc segment.2. The seal assembly as recited in claim 1 , wherein the portion includes a first axially-extending leg radially inward of the spring.3. The seal assembly as recited in claim 2 , wherein the heat shield includes a radially outer end with a second axially-extending leg.4. The seal assembly as recited in claim 3 , wherein the axially-extending leg radially inward of the spring extends in a first axial direction and the second axially-extending leg extends in a second axial direction opposite from the first axial direction.5. The seal assembly as recited in claim 3 , wherein the second axially-extending leg is disposed radially outward of the seal arc segment.6. The seal assembly as recited in claim 1 , wherein the spring is a 360 degree disc spring.7. A gas turbine engine claim 1 , comprising:a rotor rotatable about an axis;a seal arc segment radially outward of the rotor;a carriage supporting the seal arc segment;a spring configured to bias the seal arc segment axially; anda heat shield including a portion radially inward of the spring and a radially-extending leg disposed axially between the spring and the seal arc segment.8. The gas turbine engine as recited in claim 7 , wherein the portion includes an axially-extending leg radially inward of ...

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Номер записи: 62
14-02-2019 дата публикации

System of Variable Stator Vanes For A Turbine Engine

Номер: US20190048738A1
Автор: Perez Rafael
Принадлежит:

The invention relates to a system of vanes with adjustable orientation, also called a system of variable stator vanes, for a low-pressure compressor of an axial turbine engine. The system comprises vanes, each having a vane extending radially in a flow of the turbine engine and a spindle having a cylindrical portion connected to a telescopic actuating lever. The cylindrical portion comprises radially extending slot, and the actuating lever comprises a pivot joint housed in the slot, that is configured to communicate a rotary movement to the vane about its spindle. The invention also proposes a compressor and a turbine engine. 1. A system for an axial compressor of a turbine engine , said system comprising:a telescopic actuating lever;a variable stator vane with a vane body designed to extend radially in a flow of the turbine engine; anda spindle having a cylindrical portion comprising a radially extending slot,wherein the telescopic actuating lever comprises a pivot joint housed in the slot, the telescopic actuating lever being configured to communicate a rotary movement to the variable stator vane around its spindle.2. The system according to claim 1 , wherein the slot comprises inner surfaces in contact with the pivot joint of the lever.3. The system according to claim 1 , wherein the slot passes through the cylindrical portion.4. The system according to claim 1 , wherein the vane body has an average thickness that is greater than the width of the slot.5. The system according to claim 1 , wherein the spindle comprises a radial end claim 1 , the pivot joint being positioned radially between the vane body and the radial end.6. The system according to claim 1 , wherein the spindle has a constant diameter over its height.7. The system according to claim 1 , wherein the telescopic actuating lever comprises opposed lateral surfaces that are in contact with the slot.8. A system for an axial compressor of a turbine engine claim 1 , said system comprising:a telescopic ...

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Номер записи: 63
14-02-2019 дата публикации

TURBINE ENGINE WITH BEARING ASSEMBLY

Номер: US20190048743A1
Автор: Tulej Piotr
Принадлежит:

An apparatus and method for a bearing assembly including a frame, an inner race circumscribing a shaft for a turbine engine, a bearing movable about the inner ring, an outer race circumscribing the at least one rolling element, a spring assembly comprising an inner ring circumscribing the at least one cage, and an outer ring mounted to the frame, and a set of circumferentially arranged spring fingers extending between the inner ring and the outer ring. 1. A turbine engine comprising:a frame defining a central aperture;a shaft extending in a fore to aft direction through the central aperture;a bearing assembly rotationally supporting the shaft;a spring assembly comprising a first and second set of spring fingers having differing stiffness where the second set of spring fingers has an inner ring circumscribing the bearing assembly and an outer ring mounted to the frame, and a set of circumferentially arranged spring fingers extending between the inner ring and the outer ring; anda damper circumscribing the inner ring and separating the inner ring from the frame.2. The turbine engine of wherein the bearing assembly includes at least one inner race circumscribing the shaft claim 1 , at least one rolling element movable about the inner race claim 1 , and at least one outer race circumscribing the at least one rolling element.3. The turbine engine of wherein the inner ring is a first inner ring and a second inner ring and the second inner ring is the at least one outer race of the bearing assembly.4. The turbine engine of wherein the damper comprises a damper housing integral with the first inner ring.5. The turbine engine of wherein the first and second set of spring fingers comprise a plurality of spring fingers circumferentially and alternatingly arranged with respect to each other.6. The turbine engine of wherein the first set of spring fingers has a greater stiffness than the second set of spring fingers.7. The turbine engine of wherein a gap is formed between the ...

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Номер записи: 64
23-02-2017 дата публикации

TURBINE SHROUD ASSEMBLY

Номер: US20170051627A1
Автор: Hafner Matthew Troy, Morgan Victor John, ROBERTS Frederic Woodrow
Принадлежит:

A turbine shroud assembly is disclosed including an inner shroud having a surface adjacent to a hot gas path, an outer shroud, a damper block disposed between the inner shroud and the outer shroud, a first biasing apparatus, and a second biasing apparatus. The first biasing apparatus provides a first biasing force to the inner shroud, biasing the inner shroud a first deflection distance in a direction toward the hot gas path and away from the outer shroud. The second biasing apparatus provides a second biasing force to the damper block, biasing the damper block a second deflection distance in a direction toward the hot gas path and away from the outer shroud. The second deflection distance is greater than the first deflection distance. 1. A turbine shroud assembly , comprising:an inner shroud having a surface adjacent to a hot gas path;an outer shroud;a damper block disposed between the inner shroud and the outer shroud;a first biasing apparatus providing a first biasing force to the inner shroud, biasing the inner shroud a first deflection distance in a direction toward the hot gas path and away from the outer shroud; anda second biasing apparatus providing a second biasing force to the damper block, biasing the damper block a second deflection distance in the direction toward the hot gas path and away from the outer shroud,wherein the second deflection distance is greater than the first deflection distance, loading the damper block to the inner shroud.2. The turbine shroud assembly of claim 1 , wherein the first biasing apparatus includes at least one spring claim 1 , the spring connecting to or contacting the inner shroud and configured to exert the first biasing force on the inner shroud.3. The turbine shroud assembly of claim 1 , wherein the first biasing apparatus is a springless biasing apparatus.4. The turbine shroud assembly of claim 3 , wherein the biasing apparatus is driven by a pressurized fluid.5. The turbine shroud assembly of claim 1 , wherein the ...

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Номер записи: 65
01-03-2018 дата публикации

FLOATING, NON-CONTACT SEAL WITH ANGLED BEAMS

Номер: US20180058238A1
Автор: Chuong Conway, Wong Joey
Принадлежит:

Aspects of the disclosure are directed to a floating, non-contact seal comprising: a shoe, a first beam coupled to the shoe, and a second beam coupled to the shoe, where the first beam is oriented at a first angle with respect to the shoe, the first angle having a first value that is greater than five degrees. Aspects of the disclosure are directed to an engine comprising: a first structure, a second structure configured to rotate relative to the first structure, and a floating, non-contact seal that interfaces the first structure and the second structure, where the seal includes: a shoe, a first beam coupled to the shoe, and a second beam coupled to the shoe, where the first beam is oriented at a first angle with respect to the shoe, the first angle having a first value that is greater than five degrees. 1. A floating , non-contact seal comprising:a shoe;a first beam coupled to the shoe; anda second beam coupled to the shoe,wherein the first beam is oriented at a first angle with respect to the shoe, the first angle having a first value that is greater than five degrees.2. The floating claim 1 , non-contact seal of claim 1 , wherein the second beam is oriented at a second angle with respect to the shoe claim 1 , the second angle having a second value that is greater than five degrees.3. The floating claim 2 , non-contract seal of claim 2 , wherein the second value is less than ten degrees.4. The floating claim 1 , non-contract seal of claim 1 , wherein the first value is less than ten degrees.5. The floating claim 1 , non-contact seal of claim 1 , wherein the first beam is coupled to the shoe at a first location of the shoe and the second beam is coupled to the shoe at a second location of the shoe that is different from the first location.6. The floating claim 1 , non-contact seal of claim 1 , wherein the first beam claim 1 , the second beam claim 1 , and the shoe form a four-bar linkage.7. The floating claim 6 , non-contact seal of claim 6 , wherein the four-bar ...

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Номер записи: 66
01-03-2018 дата публикации

SYSTEM AND APPARATUS FOR DIVERSIFIED GEARBOX

Номер: US20180058332A1
Автор: Duong Hung, Snape Nathan, Suciu Gabriel L.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A gas turbine engine assembly comprising, a gearbox including a first housing that includes a first auxiliary gear drive on a first portion thereof, a second housing that includes a second auxiliary gear drive on a second portion thereof, and a third housing that includes a third auxiliary gear drive on a third portion thereof, the housings being interconnected so that the first portion of the first housing, the second portion of the second housing and the third portion of the third housing form a substantially triangular polyhedron shape, with the second portion of the second housing disposed between the first portion of the first housing and the third portion of the third housing. The first auxiliary gear drive, the second auxiliary gear drive and the third auxiliary gear drive project outwardly in mutually divergent directions. 1. An intermediate housing portion of a gearbox comprising:a generally triangular polyhedron shape;a coupling with a first housing portion, wherein the first housing portion comprises a first housing portion first face, and wherein a first auxiliary gear drive is arranged within the first housing portion; anda coupling with a second housing portion, wherein the second housing portion comprises a second housing portion first face, and wherein a second auxiliary gear drive is arranged within the second housing portion,wherein the first housing portion first face and the second housing portion first face intersect along a common edge distal the intermediate housing portion.2. The intermediate housing portion of the gearbox of claim 1 , wherein at least one of the first housing portion first face or the second housing portion first face includes a removable cover.3. The intermediate housing portion of the gearbox of claim 1 , further including a third auxiliary gear drive arranged within the intermediate housing portion.4. The intermediate housing portion of the gearbox of claim 3 , wherein a first set of bevel gears interconnects the first ...

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Номер записи: 67
02-03-2017 дата публикации

SPLAYED INLET GUIDE VANES

Номер: US20170058690A1
Автор: Rice Edward C., Sumner Bradley S.
Принадлежит:

A system for directing the flow of a fluid and controlling the rate of flow of the fluid. The system comprises a channel for directing the flow of the fluid; at least a pair of articulating vanes positioned within the channel for controlling the flow rate of the fluid within the channel; and a linkage between the vanes coupling the articulation of each of the vanes to the other of the vanes, wherein each vane imparts a force on the linkage when the relative angle of attack is greater than zero, wherein the force imparted on the linkage by one of the vanes is at least partially cancelled by the force imparted on the linkage by the other of the vanes during the articulation of the vanes. 1. A system for directing the flow of a fluid and controlling the rate of flow of the fluid , said system comprising:a channel for directing the flow of the fluid;at least a pair of articulating vanes positioned within said channel for controlling the flow rate of the fluid within said channel, each of said vanes comprising a pair of lateral major surfaces forming a leading edge and a trailing edge of said vane, and an axis of articulation intersecting said vane at a point spaced from the aerodynamic center of said vane; anda linkage between said vanes coupling the articulation of each of said vanes to the other of said vanes, wherein each vane imparts a force on said linkage when the relative angle of attack is greater than zero, wherein the force imparted on said linkage by one of said vanes is at least partially cancelled by the force imparted on the linkage by the other of said vanes during the articulation of said vanes.2. The system of claim 1 , wherein the axis of articulation of one of said vanes intersects said vane between the aerodynamic center and said leading edge of said vane claim 1 , and the axis of articulation of the other of said vanes intersects said other vane between the aerodynamic center and said trailing edge of said other vane.3. The system of claim 1 , ...

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Номер записи: 68
02-03-2017 дата публикации

Morphing vane

Номер: US20170058691A1
Автор: Edward C. Rice
Принадлежит: Rolls Royce North American Technologies Inc

A system for directing the flow of a fluid which comprises a channel for containing the fluid; an articulating vane positioned within the channel for directing the flow of the fluid, the vane comprising a fixed segment rigidly connected to the channel and a first moveable segment operably connected to the fixed segment by a first hub, the first hub configured to allow relative articulation between the segments; an actuator member operably connected to the moveable segment to articulate the moveable segment about the first hub; and wherein the vane further comprises a second moveable segment operably connected to the vane by a second hub, wherein the actuator member articulates the first and second moveable segments by applying a single moment to the first hubs.

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Номер записи: 69
02-03-2017 дата публикации

Loaded turbocharger turbine wastegate control linkage joints

Номер: US20170058762A1
Автор: Manuel Gonzalez, Manuel Marques
Принадлежит: Honeywell International Inc

An assembly can include a turbine housing that includes a bore, a wastegate seat and a wastegate passage that extends to the wastegate seat; a bushing disposed at least in part in the bore; a rotatable wastegate shaft received at least in part by the bushing; a wastegate plug that extends from the wastegate shaft; a control arm operatively coupled to the wastegate shaft; a control link operatively coupled to the control arm; a pin that forms a joint between the control arm and the control link; and a biasing element coupled to the pin and to the control link.

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Номер записи: 70
10-03-2016 дата публикации

TURBOMACHINE COMPRISING A PLURALITY OF FIXED RADIAL BLADES MOUNTED UPSTREAM OF THE FAN

Номер: US20160069205A1
Автор: Bordoni Nils, Kubiak Guillaume Patrice, LECORDIX Jean-Loic Herve
Принадлежит: SNECMA

Bypass turbine engine, in particular for an aircraft, in which air flows circulate from upstream to downstream, the turbine engine extending axially and comprising: 25242530245. Turbine engine according to claim 1 , wherein the means for individually adjusting the pitch of the radial vanes () comprise a single control ring () and rods ( claim 1 , ) for connecting said control ring () to each of said radial vanes ().311121311. Turbine engine according to claim 1 , wherein the inner casing () claim 1 , the inter-duct casing () and the outer casing () are at a radial distance from one another in the turbine engine () so as to define a turbine engine () having a bypass ratio (BPR) that is greater than or equal to 15.4202. Turbine engine according to claim 1 , wherein the rotational speed of the free ends of the blades () of the movable fan () is less than 340 m/s.551. Turbine engine according to claim 1 , wherein the plurality of variable-pitch radial vanes () extend in the same plane which is transverse to the axis of the turbine engine ().6525. Turbine engine according to claim 1 , wherein the axial distance between the plurality of variable-pitch radial vanes () and the movable fan () is between 0.1 and 10 times the mean chord of a variable-pitch radial vane ().720211131. Turbine engine according to claim 1 , wherein the blades () of the movable fan () extend between the inner casing () and the outer casing () of the turbine engine ().85. Turbine engine according to claim 1 , wherein each variable-pitch radial vane () has an aerodynamic profile so as to accelerate the incident air flow in accordance with a laminar flow.95. Turbine engine according to claim 1 , wherein each variable-pitch radial vane () has a body which is movable in rotation about a radial axis.1055051. Turbine engine according to claim 1 , wherein each variable-pitch radial vane (′) has a fixed body (′) and a movable flap (′).115. Turbine engine according to claim 1 , wherein each variable-pitch ...

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Номер записи: 71
09-03-2017 дата публикации

Seal assembly for turbine engine component

Номер: US20170067354A1
Автор: Michael G. Mccaffrey
Принадлежит: United Technologies Corp

A seal assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a housing, a seal carrier secured to the housing and configured to be selectively biased from the housing, and a wedge seal secured to the seal carrier and configured to abut against at least two sealing surfaces. A method of sealing between adjacent components of a gas turbine engine is also disclosed.

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Номер записи: 72
09-03-2017 дата публикации

DEVICE FOR BOUNDING A FLOW CHANNEL OF A TURBOMACHINE

Номер: US20170067366A1
Автор: Stricker Hans
Принадлежит:

A device for bounding a flow channel of a turbomachine, comprising a wall that, viewed in the circumferential direction of the turbomachine, has a multiplicity of wall segments, and including a multiplicity of outer segments that extend radially around the outside of the wall segments; each wall segment having a uniformly curved first cross-sectional contour; each outer segment including at least one second cross sectional contour that deviates from the uniformly curved first cross-sectional contour; the second cross-sectional contour having a multiplicity of depressions that are inwardly directed in the radial direction of the gas turbine; at least a portion thereof being attached to the radially outer surface of a corresponding wall segment. 1. A device for bounding a flow channel of a turbomachine , comprising:a wall, the wall, viewed in a circumferential direction of the turbomachine, having a multiplicity of wall segments, anda multiplicity of outer segments extending radially around an outside of the wall segments;each wall segment having a uniformly curved first cross-sectional contour, and each outer segment including at least one second cross sectional contour deviating from the uniformly curved first cross-sectional contour; the second cross-sectional contour having a multiplicity of depressions inwardly directed in a radial direction of the gas turbine;at least a portion of the depressions being attached to a radially outer surface of a corresponding wall segment.2. The device as recited in wherein claim 1 , in addition to the multiplicity of depressions claim 1 , the second cross-sectional contour including a multiplicity of elevations.3. The device as recited in wherein a circumferential length of an outer segment equals that of a corresponding wall segment and claim 1 , in each case claim 1 , an outer gap between two outer segments and an inner gap between two wall segments radially oppose one another.4. The device as recited in wherein a seal covers a ...

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Номер записи: 73
15-03-2018 дата публикации

Reverse flow gas turbine engine with offset rgb

Номер: US20180073429A1
Автор: Jean Dubreuil
Принадлежит: Pratt and Whitney Canada Corp

A gas turbine engine has an engine case housing a low pressure compressor drivingly connected to a low pressure turbine by a low pressure compressor shaft extending along an engine axis. The low pressure turbine is disposed forward of the low pressure compressor. A low pressure turbine shaft is drivingly connected to the low pressure turbine and extends forwardly of the low pressure turbine. A reduction gear box (RGB) is drivingly connected to the low pressure turbine shaft. The RGB is offset from the engine axis to free an access to low pressure compressor shaft connection. The offset positioning of the RGB allows to provide an access port in an axially forwardly facing surface of the engine case to access the low pressure compressor shaft and more particularly a connection thereof to the LP turbine.

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Номер записи: 74
05-06-2014 дата публикации

Turbocharger

Номер: US20140154055A1
Автор: Dietmar Metz, Thomas Ramb
Принадлежит: BorgWarner Inc

A turbocharger ( 1 ) with variable turbine geometry (VTG) having a guide grate ( 18 ) which surrounds a turbine wheel ( 4 ) radially at the outside, which has an adjusting ring ( 5 ) operatively connected to the guide blades ( 7 ) via associated blade levers ( 20 ) which are fastened to blade shafts ( 8 ) at one of the ends thereof. Each blade lever ( 20 ) has a lever head ( 23 ) which can be placed in engagement with an associated engagement recess ( 24 ), which has a base wall ( 26 ), of the adjusting ring ( 5 ), and which has a stop ( 25 ) at least for setting the minimum throughflow through the nozzle cross sections formed by the guide blades ( 7 ). The stop is a first support point ( 25 ) on the base wall ( 26 ), wherein, in the minimum throughflow position, the lever head ( 23 ) makes contact, via a wall surface ( 27 ) facing toward the base wall ( 26 ), with said first support point.

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Номер записи: 75
14-03-2019 дата публикации

VARIABLE STATOR GUIDE VANE SYSTEM

Номер: US20190078460A1
Автор: KOSTKA Richard A., Marshall Andrew, MENHEERE David
Принадлежит:

A variable stator guide vane system for a gas turbine engine comprises a set of vanes circumferentially distributed around a central axis and rotatably mounted for rotation about respective spanwise axes. A ring gear is rotatably mounted about the central axis. Pinion gears are operatively coupled to respective ones of the vanes and in driving engagement with the ring gear. Biasing members individually bias the pinion gears in meshing engagement with the ring gear. 1. A variable stator guide vane system for a gas turbine engine , the system comprising: a set of vanes circumferentially distributed around a central axis and rotatably mounted for rotation about respective spanwise axes of the vanes; a ring gear rotatably mounted about the central axis; pinion gears operatively coupled to said vanes and in driving engagement with the ring gear; and biasing members biasing the pinion gears in meshing engagement with the ring gear.2. The system of claim 1 , wherein the pinion gears are slidably engaged on spindles projecting spanwise from the vanes claim 1 , the biasing members individually biasing the pinion gears toward the ring gear along the spanwise axes.3. The system of claim 2 , wherein the ring gear and the pinion gears are beveled.4. The system of claim 1 , wherein the spindles project through respective bushings mounted to an outer casing surrounding respective airfoil portions of the vanes.5. The system of claim 1 , wherein the biasing members include individual springs spring loading the pinion gears along the spanwise axes.6. The system of claim 1 , wherein the vanes have spindles projecting along the spanwise axes from airfoil portions claim 1 , the spindles extending through circumferentially spaced-apart apertures defined through a vane inlet casing.7. The system of claim 6 , further comprising bushings disposed in the circumferentially spaced-apart apertures for receiving the spindles of the vanes.8. The system of claim 1 , wherein the vanes are inlet ...

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Номер записи: 76
14-03-2019 дата публикации

ROTARY SHAFT SUPPORT STRUCTURE AND TURBINE AND GAS TURBINE INCLUDING THE SAME

Номер: US20190078465A1
Автор: Choi Tae Gyu
Принадлежит:

A rotary shaft support structure supports each of both ends of a rotary shaft passing through a center of a gas turbine and includes a main body arranged around a circumference of the rotary shaft and mounted to each of a compressor casing and a turbine casing of the gas turbine; a pad member disposed between the main body and the circumferential surface of the rotary shaft and biased against the circumferential surface of the rotary shaft for supporting the rotary shaft; a pivot protrusion protruding from the pad member toward the main body; and a pivot housing for receiving the pivot protrusion to rotatably support the pad member. The mechanical bias is supplied by a spring member seated in the mounting groove to bias the pivot housing against the pivot protrusion. The rotary shaft support structure exhibits performance as a bearing by absorbing some of an applied load. 1. A rotary shaft support structure comprising:a main body arranged around a circumference of a rotary shaft passing through a center of a gas turbine and mounted to each of a compressor casing and a turbine casing of the gas turbine in order to support each of both ends of the rotary shaft; anda pad member disposed between the main body and the circumferential surface of the rotary shaft and biased against the circumferential surface of the rotary shaft for supporting the rotary shaft.2. The rotary shaft support structure according to claim 1 , further comprising:a pivot protrusion protruding from the pad member toward the main body; anda pivot housing for receiving the pivot protrusion to rotatably support the pad member.3. The rotary shaft support structure according to claim 2 , wherein the pivot housing is disposed between the pad member and an inner circumferential surface of the main body.4. The rotary shaft support structure according to claim 3 , wherein the inner circumferential surface of the main body is provided with a mounting groove for receiving the pivot housing.5. The rotary shaft ...

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Номер записи: 77
22-03-2018 дата публикации

THRUST REVERSER FOR A NACELLE OF AN AIRCRAFT TURBOJET ENGINE

Номер: US20180080409A1
Автор: Moradell-Casellas Pierre
Принадлежит: Safran Nacelles

A thrust reverser for a nacelle of an aircraft turbojet engine is provided that includes a thrust reverser cowl movable along a direction parallel to a longitudinal axis of the nacelle and a variable section outlet nozzle extending from the thrust reverser cowl. The thrust reverser further includes an actuator, a first locking device to lock the thrust reverser cowl, a second locking device to lock the variable section outlet nozzle, and a reset lever. The reset lever is pivotally driven by a locking pin secured to the cowl and pivots from a rest position to a reset position. 1. A thrust reverser for a nacelle of an aircraft turbojet engine , the thrust reverser comprising:a thrust reverser cowl movable in translation along a direction substantially parallel to a longitudinal axis (A) of the nacelle between a direct jet position and a reverse jet position;a variable-section outlet nozzle arranged in a downstream extension of said thrust reverser cowl and movable between at least one reduced-section ejection position and one increased-section ejection position;an actuator comprising a body mounted on a fixed structure of the thrust reverser and an actuating rod, said actuating rod being adapted to drive the variable-section outlet nozzle and the movable thrust reverser cowl in displacement;a first device for locking the thrust reverser cowl in the direct jet position on the fixed structure of the thrust reverser, the first locking device comprising a locking hook pivotally mounted about a transverse axis (B), between a closed cowl locking position in which the locking hook cooperates with a locking pin secured to the thrust reverser cowl, and an unlocking open position of the thrust reverser cowl in which the locking hook releases said locking pin;a second device for locking the variable-section outlet nozzle on the thrust reverser cowl, the second locking device adapted to alternately occupy a position of locking the variable-section outlet nozzle on the cowl and a ...

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Номер записи: 78
23-03-2017 дата публикации

Variable nozzle unit and variable geometry turbocharger

Номер: US20170082018A1
Автор: Jose Javier BAYOD, Koutarou ITOU, Taiki YOSHIZAKI
Принадлежит: IHI Corp

A support ring is connected to a first nozzle ring by multiple connecting pins. An outer cutout is formed at a region on an outer side in a radial direction at a rim on an axially one side of each first attachment hole in the first nozzle ring. An inner cutout is formed at a region on an inner side in the radial direction at the rim on the axially one side of each first attachment hole in the first nozzle ring. An outer cutout is formed at a region on an outer side in the radial direction at a rim on an axially other side of each pin hole in the support ring. An inner cutout is formed at a region on an inner side in the radial direction at the rim on the axially other side of each pin hole in the support ring.

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Номер записи: 79
12-03-2020 дата публикации

VARIABLE CYCLE COMPENSATION IN A GAS TURBINE ENGINE

Номер: US20200079517A1
Автор: BENGEA SORIN, Terwilliger Neil
Принадлежит:

An aspect includes a variable cycle system of a gas turbine engine. The variable cycle system includes an actuation system, an electric component, and a controller. The actuation system is configured to adjust a variable cycle of turbomachinery of the gas turbine engine. The electric component is operable to provide a shaft power supply or a load corresponding respectively to an adjustment of the turbomachinery. The controller is operable to adjust an output of either or both of the actuation system and the electric component for separate control of thrust and cycle responses. 1. A variable cycle system of a gas turbine engine , the variable cycle system comprising:an actuation system configured to adjust a variable cycle of turbomachinery of the gas turbine engine;an electric component operable to provide a shaft power supply or a load corresponding respectively to an adjustment of the turbomachinery; anda controller operable to adjust an output of either or both of the actuation system and the electric component for separate control of thrust and cycle responses.2. The variable cycle system of claim 1 , wherein the controller is further operable to:receive a control input;determine a plurality of current operating conditions of the gas turbine engine;calculate a plurality of commands to a plurality of power production and absorption subsystems for adjusting the variable cycle based on the current operation condition of the gas turbine engine using a plurality of models of the subsystems that describe relationships between the commands and respective impacts on engine power production and absorption; andcommunicate the commands to the power production, power absorption, and one or more other actuation subsystems based on the control input and the current operating conditions.3. The variable cycle system of claim 2 , wherein the electric component is a motor-generator.4. The variable cycle system of claim 2 , wherein the electric component absorbs power as an ...

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Номер записи: 80
31-03-2022 дата публикации

GAS-TURBINE-ENGINE OVERSPEED PROTECTION SYSTEM

Номер: US20220098995A1
Автор: Suchezky Mark E., Thompson Jeffrey C., Williams Thomas J.
Принадлежит: WILLIAMS INTERNATIONAL CO., L.L.C.

A rotational position of a variable-vane of a variable-vane turbine nozzle upstream of a turbine of a gas-turbine engine is biased towards a corresponding rotational position that will mitigate against an overspeed condition of the turbine during operation of the gas-turbine engine. When the turbine is operating at a rotational speed that is less than an overspeed threshold, the rotational position of the variable-vane is controlled independently of the biasing using a variable-vane actuator operatively coupled to the variable-vane. Responsive to a rotational speed of the turbine in excess of the overspeed threshold, the variable-vane actuator is operatively decoupled from the variable-vane so as to provide for the variable-vane to be repositioned towards the corresponding rotational position that will mitigate against the overspeed condition. 1. A gas-turbine-engine overspeed protection system , comprising:a. a variable-vane turbine nozzle, wherein said variable-vane turbine nozzle incorporates a plurality of variable nozzle vanes;b. a nozzle-vane-angle control mechanism, wherein said nozzle-vane-angle control mechanism provides for controlling a corresponding rotational angle of each of said plurality of variable nozzle vanes;c. a variable-vane actuator, wherein in a first mode of operation, said variable-vane actuator is operatively coupled to said plurality of variable nozzle vanes via said nozzle-vane-angle control mechanism so as to provide for controlling said corresponding rotational angle of each of said plurality of variable nozzle vanes and thereby control a direction of a stream of exhaust gases exiting said variable-vane turbine nozzle and subsequently impinging on a turbine of the gas-turbine engine downstream of said variable-vane turbine nozzle, and in a second mode of operation, said variable-vane actuator is operatively decoupled from said plurality of variable nozzle vanes, and said corresponding rotational angle of each of said plurality of ...

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Номер записи: 81
31-03-2022 дата публикации

STEERING ROD FOR TURBOCHARGERS AND METHOD FOR MAKING IT

Номер: US20220099019A1
Автор: ASFALTI ANDREA, CARNEVALE DOMENICO
Принадлежит:

Steering rod for connecting an actuator to a turbocharger drive member, comprising an elongated main body (), which extends between a first end () equipped with a first through-hole (), and a second end () equipped with a second through-hole (), and two bushings (), () immovably mounted in the through-holes (), (), wherein the bushings (), () are made of a material which has a greater resistance to mechanical wear than the material delimiting the through-holes (), (), and wherein each bushing () () has an outer surface that is in contact with an inner part of the through-hole (), () and has a radial recess (), the second material tightening the bushing (), () preventing it from rotating and filling the recess () preventing the bushing (), () from being extracted from the second through-hole (), (). 1. Steering rod for connecting an actuator to a turbocharger drive member , comprising:{'b': ['2', '5', '7', '6', '8'], '#text': 'an elongated main body (), which extends between a first end () equipped with a first through-hole (), and a second end () equipped with a second through-hole ();'}{'b': ['3', '7'], '#text': 'a first bushing () fixedly mounted in the first through-hole (); and'}{'b': ['4', '8'], '#text': 'a second bushing () fixedly mounted in the second through-hole ();'}wherein, moreover:{'b': ['3', '4', '7', '8'], '#text': 'the first bushing () and the second bushing () are made with a respective first material, which has a greater resistance to mechanical wear than a second material which delimits the first through-hole () and the second through-hole () respectively;'}{'b': ['3', '7', '3', '7', '9', '10', '3', '10', '3', '3', '3', '10', '9', '3', '7'], '#text': 'the first bushing () has an outer surface that is in contact with an inner part of the first through-hole (), and wherein in the outer surface of the first bushing () that is in contact with the inner part of the first through-hole () there is a radial recess () towards a central axis () of the ...

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Номер записи: 82
25-03-2021 дата публикации

COMPACT ACCESSORY SYSTEMS FOR A GAS TURBINE ENGINE

Номер: US20210087975A1
Автор: ALSTAD Shawn, Dobosz Tomasz, Felix Rocio Chavez
Принадлежит: HONEYWELL INTERNATIONAL INC.

An accessory system for a gas turbine engine having a drive shaft with an axis of rotation is provided. Also provided is a bearing housing assembly for coupling the drive shaft of an accessory having a first gear to a gear associated with the accessory system. The bearing housing assembly includes a mount including an interface to be coupled to the accessory and defining a central bore, and a lock cylinder configured to receive the drive shaft. The lock cylinder is movable relative to the central bore and the drive shaft to adjust a contact pattern between the first gear of the drive shaft and the gear of the accessory system. 1. A bearing housing assembly for coupling a drive shaft of an accessory having a first gear to a gear associated with an accessory system of a gas turbine engine , comprising:a mount including an interface to be coupled to the accessory and defining a central bore; anda lock cylinder configured to receive the drive shaft, the lock cylinder movable relative to the central bore and the drive shaft to adjust a contact pattern between the first gear of the drive shaft and the gear of the accessory system.2. The bearing housing assembly of claim 1 , wherein the mount includes a plurality of mounting bores spaced apart about a perimeter of the interface to couple the mount to the accessory.3. The bearing housing assembly of claim 1 , wherein the central bore defines a plurality of threads claim 1 , and the lock cylinder defines a plurality of threads to couple the lock cylinder to the mount.4. The bearing housing assembly of claim 3 , wherein the plurality of threads of the lock cylinder are defined on an exterior surface of the lock cylinder claim 3 , with the lock cylinder including an interior surface opposite the exterior surface to receive the drive shaft.5. The bearing housing assembly of claim 4 , wherein the exterior surface further defines a plurality of alternating flats at an end of the lock cylinder.6. The bearing housing assembly of ...

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Номер записи: 83
29-03-2018 дата публикации

Pressure-loaded seals

Номер: US20180087390A1
Автор: Jesse Ellis Barton, Kevin Thomas Mcgovern, Richard Martin DiCintio, Victor John Morgan
Принадлежит: General Electric Co

A sealing arrangement for sealing between a stage-one nozzle and an aft frame includes a seal comprising a flexible sealing element. The flexible sealing element includes an intermediate portion, a first outer portion on one side of the intermediate portion, and a second outer portion on the other side of the intermediate portion. The intermediate portion is mechanically loaded against the first stage nozzle and the aft frame, and the first outer portion and the second outer portion are pressure-loaded against the aft frame and the stage-one nozzle.

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Номер записи: 84
29-03-2018 дата публикации

GAS TURBINE ENGINE

Номер: US20180087395A1
Автор: HARDING Adrian L.
Принадлежит: ROLLS-ROYCE PLC

A gas turbine engine comprising a rotor blade, an inner casing and an outer casing, a guide vane, located upstream of the rotor blade and between the inner and outer casing, the guide vane comprising a vane contact member, wherein the guide vane is configured to pivot upon axial displacement of the inner casing relative to either or both of the outer casing and the rotor blade; and, a seal segment located radially outwardly of the rotor blade and mounted for radial displacement relative to the rotor blade; wherein at least a portion of the seal segment axially overlaps, and slidably engages a downstream end of the guide vane so that a pivoting of the guide vane causes the vane contact member to undergo a radially outward displacement which causes a radially outward displacement of at least a portion of the seal segment. 1. A gas turbine engine for the provision of propulsive thrust , comprising a rotor blade , an inner casing and an outer casing , the gas turbine engine further comprising;a guide vane, located upstream of the rotor blade and between the inner and outer casing, the guide vane comprising a vane contact member, wherein the guide vane is configured to pivot upon axial displacement of the inner casing relative to either or both of the outer casing and the rotor blade; and,a seal segment located radially outwardly of the rotor blade and mounted for radial displacement relative to the rotor blade;wherein at least a portion of the seal segment axially overlaps, and slidably engages a downstream end of the guide vane so that in use, a pivoting of the guide vane causes the vane contact member to undergo a radially outward displacement, the radially outward displacement of the vane contact member causing a radially outward displacement of at least a portion of the seal segment.2. A gas turbine engine as claimed in claim 1 , wherein the guide vane is located radially inwardly of claim 1 , and flexibly mounted to the outer casing.3. A gas turbine engine as ...

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Номер записи: 85
29-03-2018 дата публикации

Dual turn thrust reverser cascade systems and methods

Номер: US20180087474A1
Автор: Chen Chuck, Zachariah B. Vandemark
Принадлежит: Boeing Co

Systems and methods are provided for a thrust reverser system with a straight vane thrust reverser cascade. The thrust reverser system may also include a blocker door and a turning door. The blocker door may divert air flowing within a bypass flow path of the aircraft propulsor to flow through the thrust reverser cascade. The turning door may then deflect air flowing from the thrust reverser cascade to provide reverse thrust. The straight vane thrust reverser cascade may allow for increased reverse thrust and/or a smaller, more efficient, aircraft propulsor.

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Номер записи: 86
21-03-2019 дата публикации

ROTATABLE TORQUE FRAME FOR GAS TURBINE ENGINE

Номер: US20190085698A1
Автор: Clements Jeffrey Donald, VAN DER MERWE Gert Johannes, Wesling Richard Alan, Zatorski Darek Tomasz
Принадлежит:

The present disclosure is directed to a gas turbine engine including a torque frame. The torque frame includes an inner shroud defined circumferentially around the axial centerline, an outer shroud surrounding the inner shroud and defined circumferentially around the axial centerline, and a structural member extended along the radial direction and coupled to the inner shroud and the outer shroud. The torque frame is configured to rotate around the axial centerline. 1. A gas turbine engine comprising:a torque frame comprising an inner shroud defined circumferentially around the axial centerline, an outer shroud surrounding the inner shroud and defined circumferentially around the axial centerline, a structural member extended along the radial direction and coupled to the inner shroud and the outer shroud, wherein the torque frame is configured to rotate around the axial centerline.2. The gas turbine engine of claim 1 , wherein the structural member is extended along a longitudinal direction to define a lean angle relative to the axial centerline.3. The gas turbine engine of claim 2 , wherein the lean angle is acute relative to the axial centerline claim 2 , wherein a radially outward end of the structural member is disposed upstream of a radially inward end of the structural member.4. The gas turbine engine of claim 2 , wherein the lean angle is acute relative to the axial centerline claim 2 , wherein a radially inward end of the structural member is disposed upstream of a radially outward end of the structural member.5. The gas turbine engine of claim 1 , wherein the structural member defines an airfoil defining a pressure side and a suction side.6. The gas turbine engine of claim 1 , wherein the torque frame further comprises an outer band circumferentially surrounding the outer shroud claim 1 , wherein the outer band is extended at least partially along the radial direction.7. The gas turbine engine of claim 6 , further comprising a plurality of connecting members ...

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Номер записи: 87
21-03-2019 дата публикации

TURBINE BLADE, TURBINE INCLUDING SAME TURBINE BLADE, AND GAS TURBINE INCLUDING SAME TURBINE

Номер: US20190085703A1
Автор: Lee Ik Sang
Принадлежит:

A turbine blade is provided with a rotatable leading edge to reduce pressure loss as a flow direction of combustion gas varies. The turbine blade includes an airfoil including as separate bodies a trailing-edge portion and a leading-edge portion being linked to the trailing-edge portion and disposed upstream of the trailing-edge portion, the leading-edge portion including a front surface arranged on an upstream side of the leading-edge portion; and a rotary unit connected to the leading-edge portion and configured to rotate the leading-edge portion according to an inflow angle of the combustion gas such that the front surface faces a flow of the combustion gas. A barrier wall extends from a side surface of the leading-edge portion toward an open end of the trailing edge portion, and when the leading-edge portion is rotated, the bather wall portion prevents formation of a gap between the side surface and the open end. 1. A turbine blade mounted on a turbine disk in a turbine casing of a turbine , the turbine blade rotating the turbine when combustion gas flows in the turbine casing , the turbine blade comprising:an airfoil including as separate bodies a trailing-edge portion and a leading-edge portion being linked to the trailing-edge portion and disposed upstream of the trailing-edge portion, the leading-edge portion including a front surface arranged on an upstream side of the leading-edge portion; anda rotary unit connected to the leading-edge portion and configured to rotate the leading-edge portion according to an inflow angle of the combustion gas such that the front surface faces a flow of the combustion gas.2. The turbine blade according to claim 1 , wherein the trailing-edge portion includes an open end linked to the leading-edge portion and a closed end having a streamlined shape.3. The turbine blade according to claim 2 , further comprising:a barrier wall portion extending from a side surface of the leading-edge portion toward the open end of the trailing ...

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Номер записи: 88
19-03-2020 дата публикации

TURBOCHARGER

Номер: US20200088062A1
Автор: ASANO Akihiro, INOUE Yasuko, Iwata Akitoshi, Nagai Kouji, OGASAWARA Takayasu
Принадлежит: TOYOTA JIDOSHA KABUSHIKI KAISHA

A wastegate is attached to a turbine housing of a turbocharger. The wastegate opens and closes a bypass passage. An actuator is coupled to the wastegate via a link mechanism. The link mechanism includes a link rod and a link arm. The longitudinal direction of the link rod in a state in which the bypass passage is fully closed is defined as a width direction of the link arm. The coupling center position in the link arm to which the link rod is coupled is located at a position offset in a direction in which the link rod moves to open the wastegate from a middle of the link arm in the width direction. 1. A turbocharger comprising:a turbine housing that accommodates a turbine wheel;a bypass passage defined in the turbine housing, wherein the bypass passage connects an exhaust-upstream section to an exhaust-downstream section with respect to the turbine wheel, thereby bypassing the turbine wheel;a wastegate attached to the turbine housing to selectively open and close a downstream end of the bypass passage;a link mechanism coupled to the wastegate to transmit driving force to the wastegate; andan actuator coupled to the link mechanism, whereinthe turbine housing has a valve seat portion in an inner wall surface of the turbine housing, a valve member that contacts the valve seat portion at the time the bypass passage is fully closed, and', 'a shaft that extends from the valve member and is pivotally supported by a wall portion of the turbine housing,, 'the wastegate includes'} an elongated link rod that moves from one side to the other in the longitudinal direction of the link rod by receiving the driving force of the actuator, and', 'a link arm that is coupled to the link rod, converts movement of the link rod to rotation, and transmits the rotation to the shaft, and, 'the link mechanism includes'}the longitudinal direction of the link rod in a state in which the bypass passage is fully closed is defined as a width direction of the link arm, anda coupling center position ...

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Номер записи: 89
19-03-2020 дата публикации

Hidden link system blocker door

Номер: US20200088134A1
Автор: Timothy Gormley
Принадлежит: Rohr Inc

A system for deploying a blocker door of a nacelle includes a master link configured to be coupled to a fixed structure of the nacelle and a master crank pivotally attached to the master link. The system further includes a first door crank and a first door link pivotally coupled to the first door crank. The system further includes a first blocker door coupled to the first door link and a first driveshaft coupled to the master crank and to the first door crank and configured to transfer motion from the master crank to the first door crank such that aft translation of a translating sleeve of the nacelle drives the master crank via the master link, which drives the first door link via the first driveshaft and the first door crank to move the first blocker door into a bypass air duct defined by the nacelle.

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Номер записи: 90
01-04-2021 дата публикации

LINKAGE(S) BETWEEN INNER AND OUTER COWL DOORS

Номер: US20210094697A1
Автор: Bell Martin C., Linde Erik, Pretty Sean, Ramlaoui Jihad
Принадлежит:

An assembly is provided for an aircraft propulsion system. The assembly includes a fixed structure, an inner cowl door and an outer cowl door. The inner cowl door is pivotally connected to the fixed structure. The outer cowl door is pivotally connected to the fixed structure. The outer cowl door is radially outboard of and overlaps the inner cowl door. The linkage extends between and is movably connected to the inner cowl door and the outer cowl door. 1. An assembly for an aircraft propulsion system , comprising:a fixed structure;an inner cowl door pivotally connected to the fixed structure;an outer cowl door pivotally connected to the fixed structure, the outer cowl door radially outboard of and overlapping the inner cowl door; anda linkage extending between and movably connected to the inner cowl door and the outer cowl door, the linkage arranged proximate the fixed structure.2. The assembly of claim 1 , whereinthe inner cowl door extends circumferentially between a first end and a second end with a midpoint halfway circumferentially along the inner cowl door between the first end and the second end;the inner cowl door is pivotally connected to the fixed structure at the first end; andthe linkage is movably connected to the inner cowl door at a connection point circumferentially between the midpoint and the first end.3. The assembly of claim 2 , wherein the connection point is circumferentially closer to the first end than the midpoint.4. The assembly of claim 1 , whereinthe outer cowl door extends circumferentially between a first end and a second end with a midpoint halfway circumferentially along the outer cowl door between the first end and the second end;the outer cowl door is pivotally connected to the fixed structure at the first end; andthe linkage is movably connected to the outer cowl door at a connection point circumferentially between the midpoint and the first end.5. The assembly of claim 4 , wherein the connection point is circumferentially closer to ...

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Номер записи: 91
16-04-2015 дата публикации

Gas turbine engine

Номер: US20150101331A1
Автор: Brian Richard LACE, Richard Peace, Robert GOULDS
Принадлежит: Rolls Royce PLC

A gas turbine engine having a fire wall that is configured to provide a fire resistant barrier between a first zone and a second zone in the gas turbine engine, the second zone being hotter than the first zone when the gas turbine engine is in use. The gas turbine also has an actuator that is located in the first zone and is configured to generate a mechanical force when operated, an actuatable device that is located in the second zone and is configured to be actuated by a mechanical force and a mechanical force transmitting device that extends from the actuator to the actuatable device via a hole in the fire wall. The mechanical force transmitting device is configured to, when the actuator is operated, actuate the actuatable device by transmitting a mechanical force generated by the actuator to the actuatable device.

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Номер записи: 92
28-03-2019 дата публикации

VARIABLE STATOR VANE RIGGING

Номер: US20190093502A1
Автор: Lyon Owen, SWANN Andrew B.
Принадлежит: ROLLS-ROYCE PLC

A method and apparatus for rigging a variable stator vane arrangement () for a gas turbine engine () is provided. The arrangement provides at least two sets of rigging holes to an actuator (), with the actuator extension during rigging being dependent on which of the sets of holes are aligned. This allows the actuator extension to be adjustable during rigging, thereby facilitating different relationships between actuator extension and variable vane angle in use. 1. A method of rigging a variable stator vane arrangement for a gas turbine engine , the variable stator vane arrangement comprising:an actuator comprising an actuator body and an actuator ram, the actuator ram being moveable relative to the actuator body along an actuator axis (X);an array of variable stator vanes; andan adjustable connecting arrangement that, once adjusted, fixedly connects the actuator ram to the variable stator vanes such that the angle of incidence of the variable stator vanes is dependent on the position of the actuator ram along the actuator axis, wherein:each of the actuator body and the actuator ram comprises at least one rigging hole, with at least one of the actuator body and the actuator ram comprising at least two rigging holes that are offset from each other in the direction of the actuator axis, the method comprising:aligning an actuator ram rigging hole with an actuator body rigging hole;inserting a rigging pin through the aligned actuator ram rigging hole and actuator body rigging hole so as to set a rigging position of the actuator ram;setting the angle of incidence of the variable stator vanes to a rigging angle; andadjusting the adjustable connecting arrangement so as to fixedly connect the actuator ram at the rigging position to the variable stator vanes at the rigging angle.2. The method according to claim 1 , wherein the rigging holes in the at least one of the actuator body and the actuator ram that comprises at least two rigging holes are circumferentially offset ...

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Номер записи: 93
28-03-2019 дата публикации

AXIAL RETENTION OF A FAN SHAFT IN A GAS TURBINE ENGINE

Номер: US20190093504A1
Автор: BECOULET Julien Fabien Patrick, Brault Michel Gilbert Rolant, Cuvillier Romain Guillaume, Negri Arnaud Nicolas
Принадлежит: SAFRAN AIRCRAFT ENGINES

A gas turbine engine, comprising a fan driven by a fan shaft and retractable axial retainer of the shaft mounted on a stator element and movable between a first operational position and a second non-operational position. The retainer may comprise a fluid supply configured for supplying at least one cavity with fluid. In some examples, the retainer is mounted movable or slidingly so as to generate movement of the retainer between the operation and non-operational positions. 1. A gas turbine engine , comprising a fan driven by a fan shaft and retractable axial retention means for said shaft mounted on a stator element and movable between a first operational position and a second non-operational position , wherein said retention means comprises means for supplying fluid to at least one cavity wherein said retention means is movably mounted so as to generate movement of the retention means between said first operational position and said second non-operational position.2. The engine according to claim 1 , wherein said retention means comprises at least one finger claim 1 , movable between said positions claim 1 , in a substantially radial direction with respect to an axis of rotation of said fan shaft.3. The engine according to claim 2 , wherein said finger is biased into the non-operational position.4. The engine according to claim 1 , wherein said stator element is a bearing support.5. The engine according to claim 4 , wherein said bearing support comprises a downstream truncated part and an upstream cylindrical part which surrounds a rolling bearing and comprises claim 4 , at the outer periphery thereof claim 4 , said at least one cavity.6. The engine according to claim 5 , wherein said cylindrical part is surrounded by a downstream part of a fan disc.7. The engine according to claim 5 , further comprising means for pressurizing a seal claim 5 , and means for removing and conveying pressurized air to said at least one cavity.8. The engine according to claim 5 , ...

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Номер записи: 94
06-04-2017 дата публикации

PNEUMATIC TRIP VALVE PARTIAL STROKING ARRANGEMENT

Номер: US20170096908A1
Автор: Pacelli William G., Walsh James F.
Принадлежит:

A pneumatic trip system for a turbine includes a valve member; a valve stem connected to the valve member; and an actuator assembly connected to the valve stem. The actuator assembly includes a cylinder; a piston connected to the valve stem, the piston dividing the cylinder into a first chamber and a second chamber; a biasing element disposed in the second chamber of the cylinder; and a pneumatic circuit in communication with the second chamber of the cylinder. The pneumatic circuit is configured to pressurize the second chamber of the cylinder to actuate the piston to move the valve stem and the valve member to an exercised position between the open position and the closed position while the first chamber is pressurized. 1. A pneumatic trip system for a turbine , comprising:a valve member operatively associated with a turbine flow path, the valve member being configured to be actuated to engage the turbine flow path to close the flow path and prevent fluid flow through the turbine flow path;a valve stem connected to the valve member, the valve stem being configured to move the valve member between an open position allowing fluid flow through the turbine flow path and a closed position preventing fluid flow through the turbine flow path; and a cylinder;', 'a piston operatively connected to the valve stem, the piston being movably disposed within the cylinder and dividing the cylinder into a first chamber and a second chamber;', 'a biasing element disposed in the second chamber of the cylinder, the biasing element engaging the piston to bias the valve stem and the valve member toward the closed position; and', 'a pneumatic circuit in communication with the second chamber of the cylinder,, 'an actuator assembly operatively connected to the valve stem, the actuator assembly comprisingwherein the cylinder includes a first port for placing the first chamber in communication with an exterior of the cylinder such that the first chamber of the cylinder can be pressurized to ...

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Номер записи: 95
06-04-2017 дата публикации

SANDWICH STRUCTURE HAVING HINGE ASSEMBLIES FOR ACCOMMODATING DIFFERENTIAL IN-PLANE EXPANSION OF FACE SHEETS

Номер: US20170096963A1
Автор: Brown John J., Grip Robert E.
Принадлежит:

A sandwich structure includes a first skin and a second skin spaced apart from each other and interconnected by a hinge assembly including a first hinge member and a second hinge member. The first hinge member proximal end is coupled to the first skin. The second hinge member proximal end is coupled to the second skin. The first hinge member distal end is coupled to the second member distal end at a member joint. The hinge assembly: prevents movement of the first skin relative to the second skin in an in-plane longitudinal direction of the first skin, allows movement of the first skin relative to the second skin in an in-plane transverse direction of the first skin, and allows movement of the first skin relative to the second skin in an out-of-plane direction that is normal to the in-plane longitudinal direction and the in-plane transverse direction. 1. A sandwich structure , comprising:a first skin and a second skin spaced apart from each other; a first hinge member having a proximal end coupled to the first skin at a first skin joint and allowing the first hinge member to pivot about the first skin joint; and', 'a second hinge member having a proximal end coupled to the second skin at a second skin joint and allowing the second hinge member to pivot about the second skin joint;', 'the first hinge member and the second hinge member each having a distal end, the distal end of the first hinge member coupled to the distal end of the second hinge member at a member joint located between the first skin and the second skin;, 'at least one hinge assembly joining the first skin to the second skin, the hinge assembly including preventing movement of the first skin relative to the second skin in an in-plane longitudinal direction of the first skin;', 'allowing movement of the first skin relative to the second skin in an in-plane transverse direction of the first skin that is perpendicular to the in-plane longitudinal direction of the first skin; and', 'allowing movement of ...

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Номер записи: 96
16-04-2015 дата публикации

Mounting structure of cooling-fan

Номер: US20150104304A1
Автор: Yu-Seorg Jeong
Принадлежит: Hyundai Motor Co

A mounting structure of a cooling-fan mounted in an opening formed on a Front End Module (FEM) carrier may include a locking guide fixed to the opening and along a longitudinal direction of which a guide hole is formed, a link bar along a longitudinal direction of which a sliding hole is pierced and one end of which is connected rotatively to the FEM carrier, a first pin that is slid through the respective guide hole and sliding hole and connects the link bar and the locking guide, a second pin, one end of which is connected slidingly to the sliding hole, and a motor connected to the other end of the second pin to rotate the second pin. The cooling-fan is connected to the first pin and slid through the opening by link-movement of the first pin, the second pin and the link bar.

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Номер записи: 97
12-04-2018 дата публикации

INLET COWL FOR A TURBINE ENGINE

Номер: US20180100434A1
Автор: CEDAR Richard David, Crall David William, Dindar Mustafa, DONOFRIO Joseph
Принадлежит:

The disclosure is towards an inlet cowl for a turbine engine including a surface defining an inlet with a flow path and a method towards controlling the airflow in the flow path. The inlet cowl further includes an inlet lip and inner and outer barrels. The inlet lip confronts the inner barrel at a junction defining a gap. 1. An inlet cowl for a turbine engine comprising:an annular body with a surface defining an inlet opening and an outlet opening to define a flow direction from the inlet opening to the outlet opening, with the surface having a cross-sectional profile with a protuberance in the surface of the inlet opening.2. The inlet cowl of further comprising radially spaced inner and outer barrels.3. The inlet cowl of wherein the protuberance is formed on the inner barrel.4. The inlet cowl of wherein a gap is formed between a downstream edge of the surface and an upstream edge of the inner barrel.5. The inlet cowl of wherein the protuberance faces the inlet.6. The inlet cowl of wherein the protuberance is chamfered or rounded.7. The inlet cowl of further comprising a movable element wherein the movable element selectively forms the protuberance.8. The inlet cowl of wherein the movable element is pivotally coupled within the inner barrel.9. The inlet cowl of wherein the movable element is chamfered or rounded.10. The inlet cowl of further comprising an actuator operably coupled to the movable element to move the movable element in order to adjust a height of the movable element.11. The inlet cowl of wherein the actuator is at least one of a thermal or mechanical element.12. The inlet cowl of further comprising an annular strip forming the protuberance.13. The inlet cowl of wherein the annular strip extends 180 degrees within the inner barrel.14. The inlet cowl of wherein the annular strip is disposed on a side of the inlet cowl confronting a crosswind.15. The inlet cowl of wherein the protuberance has a maximum height that is a function of a boundary layer ...

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Номер записи: 98
26-03-2020 дата публикации

Containment Case Active Clearance Control Structure

Номер: US20200095883A1
Автор: Jain Manoj Kumar, Jain Nitesh, Kray Nicholas Joseph
Принадлежит:

A propulsion system including a casing surrounding a fan rotor assembly is provided. The casing includes an outer layer material defining a first coefficient of thermal expansion (CTE) and an inner layer material. The casing further includes a spring member disposed between the outer layer material and the inner layer material coupling the outer layer material and the inner layer material. The spring member is coupled to each of the outer layer material and the inner layer material within a flow passage defined between the outer layer material and the inner layer material. The spring member defines a second CTE greater than the first CTE. 1. A propulsion system , the propulsion system comprising:a casing surrounding a fan rotor assembly, wherein the casing comprises an outer layer material defining a first coefficient of thermal expansion (CTE) and an inner layer material, and wherein the casing further comprises a spring member disposed between the outer layer material and the inner layer material coupling the outer layer material and the inner layer material, wherein the spring member is coupled to each of the outer layer material and the inner layer material within a flow passage defined between the outer layer material and the inner layer material, and further wherein the spring member defines a second CTE greater than the first CTE.2. The propulsion system of claim 1 , wherein the flow passage is defined between an inner surface of the outer layer material and an outer surface of the inner layer material.3. The propulsion system of claim 1 , wherein the spring member is disposed within the flow passage claim 1 , and further wherein the spring member is coupled directly to the inner surface of the each of the outer layer material and the inner layer material.4. The propulsion system of claim 1 , further comprising:a first bleed system configured to provide a first flow of fluid to the flow passage at the casing.5. The propulsion system of claim 4 , wherein the ...

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Номер записи: 99
26-03-2020 дата публикации

Geared Rotary Power Distribution Unit With Mechanical Differential Gearing For Multiple Actuator Systems

Номер: US20200096085A1
Автор: Mariusz Kuczaj, Paul Smith, Roy Romana, Thomas Lagger
Принадлежит: Woodward Inc

Methods and systems for nacelle door electromechanical actuation may include a power distribution unit comprising a motor and differential gears; and a plurality of electromechanical actuators, each coupled to an output of a corresponding one of the differential gears. Each of the electromechanical actuators may include a configurable brake and a mechanical output, where the power distribution unit may provide mechanical torque to one of the electromechanical actuators via the motor and the differential gears based on configuration of the configurable brakes in each of the electromechanical actuators. At least one of the configurable brakes may be an electrically configurable brake. At least one of the configurable brakes may be a mechanically configurable brake. The differential gears may include two or more differential gears for receiving an input torque and supplying an output torque to one of a plurality of outputs of the differential gears.

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Номер записи: 100