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Небесная энциклопедия

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

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Мониторинг СМИ

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

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Применить Всего найдено 9511. Отображено 100.
13-03-2018 дата публикации

Реактивный двигатель

Номер: RU0000177796U1
Автор: Борис Андреевич Шахов
Принадлежит: Борис Андреевич Шахов

Реактивный двигатель содержит обтекаемый корпус, камеры сгорания, закрепленные за корпус, размещенную в корпусе электромагнитную подушку, состоящая из пустотелых цилиндров с электромагнитами статора, обратимую электромашину, состоящую из пустотелых цилиндров с электромагнитами статора и роторов. Полезная модель позволяет расширить возможности бортового или навесного электрооборудования, позволяющим применить для создания дополнительной тяги движители с электроприводами на реактивном двигателе или как бортовые движители с электроприводом на летательных аппаратах. 2 ил. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 177 796 U1 (51) МПК F02K 3/00 (2006.01) F01D 15/10 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ (52) СПК F02K 3/00 (2018.01); F01D 15/10 (2018.01) (21)(22) Заявка: 2017121758, 20.06.2017 (24) Дата начала отсчета срока действия патента: (73) Патентообладатель(и): Шахов Борис Андреевич (RU) Дата регистрации: 13.03.2018 (56) Список документов, цитированных в отчете о поиске: RU 2014482 C2, 15.06.1994. US 2743375 A, 24.04.1956. EP 0305763 A1, 08.03.1989. GB 2288642 A, 25.10.1995. US 4253031A, 24.02.1981. US 5376827 A, 27.12.1994. (45) Опубликовано: 13.03.2018 Бюл. № 8 R U (54) РЕАКТИВНЫЙ ДВИГАТЕЛЬ (57) Реферат: Реактивный двигатель содержит обтекаемый корпус, камеры сгорания, закрепленные за корпус, размещенную в корпусе электромагнитную подушку, состоящая из пустотелых цилиндров с электромагнитами статора, обратимую электромашину, состоящую из пустотелых цилиндров с электромагнитами Стр.: 1 статора и роторов. Полезная модель позволяет расширить возможности бортового или навесного электрооборудования, позволяющим применить для создания дополнительной тяги движители с электроприводами на реактивном двигателе или как бортовые движители с электроприводом на летательных аппаратах. 2 ил. U 1 U 1 Адрес для переписки: 644023, г. Омск, ул. 25 Рабочая, 125, кв. 75, Шахову Б.А. 1 7 7 7 9 6 Приоритет(ы): (22) Дата подачи ...

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Номер записи: 1
19-06-2020 дата публикации

Двухконтурный турбопрямоточный реактивный двигатель

Номер: RU0000198144U1
Автор: Вадим Александрович Иванов
Принадлежит: Вадим Александрович Иванов

Двухконтурный турбопрямоточный реактивный двигатель состоит из вентилятора В, общего для наружного НК и внутреннего контура ВК, включающего компрессор К, первую камеру сгорания КС1, многоступенчатую турбину компрессора ТК, турбину вентилятора ТВ. На выходе из наружного НК и внутреннего ВК контуров расположен смеситель потоков контуров СМ, общая форсажная камера ФК и общее регулируемое сопло PC контуров. Между ступенями турбины компрессора ТК расположена вторая камера сгорания КС2. Это обеспечивает увеличение степени понижения давления в форсажной камере и регулируемом реактивном сопле до максимальной величины за счет перераспределения перепада давления между турбиной компрессора, турбиной вентилятора, форсажной камерой и реактивным соплом. В результате наличие второй камеры сгорания обеспечивает увеличение свободной энергии внутреннего контура при сохранении его экономичности, которое используется для уменьшения диаметра внутреннего контура и соответственного увеличения степени двухконтурности.При сохранении постоянного диаметра наружного контура и степени повышения давления в вентиляторе такое увеличение степени двухконтурности обеспечивает также постоянный полетный и общий КПД двигателя и не приводит к изменению удельной тяги и удельного расхода топлива на умеренных сверхзвуковых скоростях полета, но увеличивает площадь проходного сечения наружного контура и соответственно расход воздуха через этот контур и тягу двигателя исключительно на прямоточных режимах работы с переводом внутреннего контура на режим авторотации при больших сверхзвуковых скоростях полета. При таком увеличении степени двухконтурности уменьшается также удельный вес двигателя на всех скоростях полета. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 198 144 U1 (51) МПК F02K 3/08 (2006.01) F02K 3/12 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ (52) СПК F02K 3/08 (2020.01); F02K 3/12 (2020.01); F02C 6/003 (2020.01) (21)(22) Заявка: 2019117166, 03.06. ...

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Номер записи: 2
16-03-2022 дата публикации

Двухконтурная газотурбинная установка

Номер: RU0000209432U1
Автор: Владимир Леонидович Письменный
Принадлежит: Владимир Леонидович Письменный

Полезная модель содержит входное устройство, вентилятор, внутренний контур, внешний контур. Внутри внутреннего контура расположены компрессор среднего давления, компрессор высокого давления, камера сгорания, турбины. Внутри внешнего контура расположены воздуховоздушный теплообменник, выхлопные патрубки, соединяющие внутренний контур с атмосферой, свободная турбина, выходной канал. Воздух из компрессора среднего давления поступает в воздухвоздушный теплообменник и далее - в компрессор высокого давления. Воздуховоздушный теплообменник позволяет понизить температуру воздуха на входе в компрессор среднего давления и, соответственно, на входе в камеру сгорания, что позволяет иметь высокие степени повышения давления воздуха в двигателе (более 100) и, соответственно, высокий эффективный КПД (более 0,6). РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 209 432 U1 (51) МПК F02K 3/06 (2006.01) F02C 7/00 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ (52) СПК F02K 3/06 (2022.01); F02C 7/00 (2022.01) (21)(22) Заявка: 2021124272, 12.08.2021 (24) Дата начала отсчета срока действия патента: (73) Патентообладатель(и): Письменный Владимир Леонидович (RU) Дата регистрации: 16.03.2022 Приоритет(ы): (22) Дата подачи заявки: 12.08.2021 (45) Опубликовано: 16.03.2022 Бюл. № 8 2 0 9 4 3 2 R U (54) ДВУХКОНТУРНАЯ ГАЗОТУРБИННАЯ УСТАНОВКА (57) Реферат: Полезная модель содержит входное поступает в воздухвоздушный теплообменник и устройство, вентилятор, внутренний контур, далее - в компрессор высокого давления. внешний контур. Внутри внутреннего контура Воздуховоздушный теплообменник позволяет расположены компрессор среднего давления, понизить температуру воздуха на входе в компрессор высокого давления, камера сгорания, компрессор среднего давления и, соответственно, турбины. Внутри внешнего контура расположены на входе в камеру сгорания, что позволяет иметь воздуховоздушный теплообменник, выхлопные высокие степени повышения давления воздуха в ...

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Номер записи: 3
19-01-2012 дата публикации

Propulsion system

Номер: US20120011828A1
Автор: Hugh B. Nicholson
Принадлежит: Individual

A propulsion system may include a cylindrical support member and a tubular rotatable member rotatably mounted within the support member that may be adapted to permit fluid flow therethrough. The tubular rotatable member may extend past a down stream end of the support member. An exemplary embodiment of a propulsion system may also disclose a vane attached on an interior surface of the tubular member and may include a blade which extends in a direction toward a rotational axis of the rotatable member such that rotation of the tubular member and the vane attached thereon draws fluid into the tubular member to accelerate the fluid flow through the tubular member. Additionally, a nozzle may be attached to the down stream end of the support member and include a primary nozzle and a secondary nozzle within the primary nozzle. The secondary nozzle may be engaged with the primary nozzle by a stator.

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Номер записи: 4
02-02-2012 дата публикации

Gas turbine engine

Номер: US20120023898A1
Автор: Kenneth J. MACKIE, Mark D. Taylor
Принадлежит: Rolls Royce PLC

A gas turbine engine includes, in flow series, a generator section and a free power turbine. The generator section includes one or more generator turbine stages. One or more respective generator drive shafts extend axially forwardly from the generator turbine stages to one or more corresponding compressor stages. The free power turbine includes a first turbine stage and a contra-rotating second turbine stage. The gas turbine engine further includes a first drive shaft extending axially from the free power turbine to transmit rotational drive from the first turbine stage to a first propeller. The gas turbine engine further includes a second drive shaft extending from the free power turbine coaxially with the first drive shaft to transmit contra-rotational drive from the second turbine stage to a contra-rotating second propeller.

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Номер записи: 5
09-02-2012 дата публикации

Ventilation inlet

Номер: US20120034068A1
Автор: Zahid M. Hussain
Принадлежит: Rolls Royce PLC

A ventilation inlet comprising a ventilation pipe to receive flow from a first flow zone and to deliver the flow to a second flow zone; a divider arranged to divide a portion of the ventilation pipe into a static pressure zone and a total pressure zone; and a deflector arranged to direct flow from the total pressure zone at least partially across the static pressure zone to restrict delivery of the flow from the static pressure zone to the second flow zone dependent on the pressure of the flow in the first flow zone.

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Номер записи: 6
09-02-2012 дата публикации

Actuation mechanism for a convertible gas turbine propulsion system

Номер: US20120034080A1
Автор: Gregory E. Reinhardt, Rajendra K. Agrawal
Принадлежит: Agrawal Rajendra K, Reinhardt Gregory E

A method of powering a rotary-wing aircraft includes selectively coupling and uncoupling a first power turbine to change a power distribution between the rotor system and the secondary propulsion system.

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Номер записи: 7
05-04-2012 дата публикации

Cowl assembly

Номер: US20120079804A1
Автор: Alan Roy Stuart, Kenneth Stephan Scheffel
Принадлежит: General Electric Co

An assembly for a turbofan engine includes a first cowl member comprising an aft portion and a translatable cowl member comprising a forward portion configured to be received within the aft portion. The translatable cowl member is configured to be moveable with respect to the first cowl member between a first operational position wherein the forward portion is received within the aft portion of the first cowl member, and a second operational position wherein a smaller portion of the forward portion is received within the aft portion than in the first operational position. The translatable cowl member is configured to cooperate with a core cowl of the turbofan engine to define a portion of a fan duct having an exit nozzle, and the translatable cowl member is configured to define a flow control location near the exit nozzle that is associated with a controlling fan duct area.

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Номер записи: 8
12-04-2012 дата публикации

Turbofan jet engine

Номер: US20120087787A1
Автор: Dewain Ray Brown
Принадлежит: Individual

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

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Номер записи: 9
26-04-2012 дата публикации

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

Номер: US20120099988A1
Автор: Francois Gallet, Gilles Alain Charier
Принадлежит: SNECMA SAS

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

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Номер записи: 10
03-05-2012 дата публикации

Gas turbine engine bifurcation located fan variable area nozzle

Номер: US20120102915A1
Автор: Constantine Baltas
Принадлежит: Individual

A gas turbine engine includes a core engine defined about an axis, a gear system driven by the core engine, a fan, and a variable area flow system. The gear system defines a gear reduction ratio of greater than or equal to about 2.3. The fan is driven by the gear system about the axis to generate a bypass flow. The variable area flow system operates to effect the bypass flow.

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Номер записи: 11
10-05-2012 дата публикации

Variable area fan nozzle fan flutter management system

Номер: US20120110980A1
Автор: Robert J. Morris, Zbigniew M. Grabowski
Принадлежит: Individual

A system and method of controlling a fan blade flutter characteristic of a gas turbine engine includes adjusting a variable area fan nozzle in response to a neural network.

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Номер записи: 12
17-05-2012 дата публикации

Gas turbine engine with pylon mounted accessory drive

Номер: US20120117940A1
Автор: Michael Winter
Принадлежит: United Technologies Corp

A gas turbine engine includes an accessory gearbox within an engine pylon. The accessory components may be mounted within the engine pylon to save weight and space within the core nacelle as well as provide a relatively lower temperature operating environment.

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Номер записи: 13
24-05-2012 дата публикации

Variable area fan nozzle fan flutter management system

Номер: US20120124965A1
Автор: Ray W. Carmichael, Robert J. Morris, Zbigniew M. Grabowski
Принадлежит: Individual

A gas turbine engine includes a controller that controls a fan blade flutter characteristic through control of a variable area fan nozzle.

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Номер записи: 14
28-06-2012 дата публикации

Gas turbine engine heat exchanger

Номер: US20120159961A1
Автор: Carlos Soto-Gonzalez, Michael Stephen Krautheim
Принадлежит: Rolls Royce North American Technologies Inc

A gas turbine engine having a heat exchanger is disclosed. In one form the gas turbine engine includes a particle separator that can be used to separate particles or foreign objects and create a dirty flow and a clean flow. A blower can be used to discharge the particles or foreign objects from the separator. The heat exchanger includes a relatively warm flow path from a downstream region of a compressor and a relatively cool flow path from an upstream region of the compressor. The relatively cool flow path is merged with the dirty flow. In another embodiment, the gas turbine engine is a turbofan and the relatively cool flow path is merged with a bypass flow. In one embodiment of the engine the relatively warm flow path, after having exchanged heat with the relatively cool flow path is delivered to a working component without passing through a turbomachinery component.

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Номер записи: 15
20-09-2012 дата публикации

Cooled pusher propeller system

Номер: US20120237359A1
Автор: Alfred M. Stern
Принадлежит: Individual

A propulsion system and method includes an annular exhaust nozzle about an axis radially outboard of the annular cooling flow nozzle and ejecting an exhaust flow through an annular exhaust nozzle about an axis radially outboard of the annular cooling flow nozzle.

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Номер записи: 16
01-11-2012 дата публикации

Multiple core variable cycle gas turbine engine and method of operation

Номер: US20120272656A1
Автор: James W. Norris
Принадлежит: United Technologies Corp

A gas turbine engine system includes a fan assembly, a low pressure compressor, a low pressure turbine, a plurality of engine cores including a first engine core and a second engine core, and a control assembly. A primary flowpath is defined through the fan assembly, the low pressure compressor, the low pressure turbine, and the active engine cores. Each engine core includes a high pressure compressor, a combustor downstream from the high pressure compressor, and a high pressure turbine downstream from the combustor. The control assembly is configured to control operation of the plurality of engine cores such that in a first operational mode the first and the second engine cores are active to generate combustion products and in a second operational mode the first engine core is active to generate combustion products while the second engine core is idle.

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Номер записи: 17
15-11-2012 дата публикации

Gas turbine engine and reheat system

Номер: US20120285137A1
Автор: Carol Noëlle Nin, Jon V. EBACHER
Принадлежит: Individual

One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique reheat system for a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and reheat systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

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Номер записи: 18
06-12-2012 дата публикации

Engine for thrust or shaft output and corresponding operating method

Номер: US20120304619A1
Автор: Michael Alan Beachy Head
Принадлежит: Individual

An engine and a method of operating the engine are provided. The engine includes a gas turbine and fan that rotate together to provide an exhaust gas flow stream, which flows over a free turbine that is connected to a power take-off. The free turbine can extract energy from the exhaust gas flow stream and transfer it as shaft power to the power take-off and the amount of energy extracted by the free turbine is controlled by varying the pitch of the free turbine's blades and/or by varying the pitch or stator vanes of a stator upstream of the free turbine. The control over the amount of energy extracted by the free turbine allows the engine to be used to provide thrust from the gas turbine and fan or to provide shaft power at the power take-off, or a combination of thrust and shaft power.

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Номер записи: 19
20-12-2012 дата публикации

Mounting system

Номер: US20120321470A1
Автор: Alexander V. LAVRENKO, Kenneth F. Udall
Принадлежит: Rolls Royce PLC

A mounting system for mounting a blade to a rotor body includes a pitch control mechanism including an anchor and a pitch change rod extending radially outwardly from the anchor to join to a base of the blade. The anchor and the rod are rotatable about the longitudinal axis of the rod to vary the blade pitch. The pitch control mechanism further includes a torque-transmitting formation between the blade and the anchor such that pitch-varying torque can be transmitted to the blade through the torque-transmitting formation while allowing relative radial movement between the blade and the anchor. The system includes a primary bearing formation, which transmits blade centrifugal loads to the rotor body while accommodating variation of the blade pitch, and secondary bearing formation, which transmits pitch change mechanism centrifugal loads to the rotor body while accommodating rotation of the anchor and the rod during variation of the blade pitch.

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Номер записи: 20
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: 21
03-01-2013 дата публикации

Dual fan gas turbine engine and gear train

Номер: US20130004297A1
Автор: William G. Sheridan
Принадлежит: Individual

A gas turbine engine includes a first fan, a second fan spaced axially from the first fan, a turbine-driven fan shaft and an epicyclic gear train coupled to be driven by the turbine-driven fan shaft and coupled to drive the first fan and the second fan. The epicyclic gear train includes a carrier that supports star gears that mesh with a sun gear, and a ring gear that surrounds and meshes with the star gears. The star gears are supported on respective journal bearings.

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Номер записи: 22
10-01-2013 дата публикации

Efficient, low pressure ratio propulsor for gas turbine engines

Номер: US20130008144A1
Автор: Anthony R. Bifulco, Becky E. Rose, Edward J. Gallagher, Jason Elliott, Jun Jiang
Принадлежит: Individual

A gas turbine engine includes a spool, a turbine coupled to drive the spool and a propulsor that is coupled to be driven by the turbine through the spool. A gear assembly is coupled between the propulsor and the spool such that rotation of the spool drives the propulsor at a different speed than the spool. The propulsor includes a hub and a row of propulsor blades that extends from the hub. The row includes no more than 16 of the propulsor blades.

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Номер записи: 23
28-03-2013 дата публикации

THRUST REVERSER TRANSLATING SLEEVE ASSEMBLY

Номер: US20130075492A1
Автор: Sawyer Christopher Steven, Welch John Michael
Принадлежит: SPIRIT AEROSYSTEMS, INC.

A translating sleeve assembly for use with an aircraft engine thrust reverser broadly comprises a slider track, a lifting bolt, and a hinge. The slider track may include a tubular sidewall and a guide channel formed therein that follows a curved path along a length thereof. The lifting bolt may be slidably positioned within the slider track. The hinge may include a first hinge element with a base coupled to an outer cowl shroud, a second hinge element with a hinge rod slidably positioned within the guide channel, and a pin about which the first hinge element and the second hinge element pivot. During deployment of the thrust reverser, the lifting bolt slides aftward and the hinge rod slides within the guide channel, urging the lifting bolt to pivot and the hinge to open, thereby lifting the outer cowl shroud away from the center of the aircraft engine. 1. A translating sleeve assembly for use with an aircraft engine thrust reverser , the translating sleeve assembly comprising:a slider track having a guide channel formed therein that follows a curved path along a length thereof;a lifting bolt slidably positioned within the slider track; anda hinge coupled to an outer cowl shroud, the hinge including a hinge rod slidably positioned within the guide channel and coupled to the lifting bolt,wherein during deployment of the thrust reverser, the outer cowl shroud is pushed aftward such that the lifting bolt slides within the slider track and the hinge rod slides within the guide channel, urging the lifting bolt to pivot and the hinge to open, thereby lifting the outer cowl shroud away from the center of the aircraft engine.2. The translating sleeve assembly of claim 1 , wherein the hinge includesa first hinge element with a base coupled to the outer cowl shroud,a second hinge element coupled to the hinge rod, anda pin about which the first hinge element and the second hinge element pivot.3. The translating sleeve assembly of claim 2 , wherein the first hinge element ...

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Номер записи: 24
25-04-2013 дата публикации

Turbine engine comprising a contrarotating propeller receiver supported by a structural casing attached to the intermediate housing

Номер: US20130098066A1
Автор: Audrey Grede, Francois Gallet
Принадлежит: SNECMA SAS

The present invention relates to an open rotor type aircraft turbine engine ( 1 ), comprising a contrarotating propeller receiver ( 30 ) and a dual-body gas generator ( 14 ) comprising a low-pressure compressor ( 16 ) and a high-pressure compressor ( 18 ) separated by an intermediate housing ( 27 ), said gas generator being arranged upstream from said receiver. According to the invention, the turbine engine further comprises a structural casing ( 50 ) for supporting the receiver ( 30 ), surrounding the gas generator ( 14 ) and having a downstream end ( 50 a ) attached to said receiver and an upstream end ( 50 b ) attached to said intermediate housing ( 27 ). Furthermore, it comprises additional connection means ( 60 ) between said structural supporting casing and the gas generator, arranged between the upstream and downstream ends ( 50 b , 50 a ) of the casing.

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Номер записи: 25
02-05-2013 дата публикации

Gas turbine engine with two-spool fan and variable vane turbine

Номер: US20130104521A1
Автор: Daniel B. Kupratis
Принадлежит: Individual

A gas turbine engine and a method of operating the gas turbine engine according to an exemplary aspect of the present disclosure includes modulating a variable high pressure turbine inlet guide vane of a high pressure spool to performance match a first stage fan section of a low pressure spool and an intermediate stage fan section of an intermediate spool to maintain a generally constant engine inlet flow while varying engine thrust.

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Номер записи: 26
02-05-2013 дата публикации

Gas turbine engine with auxiliary fan

Номер: US20130104523A1
Автор: Daniel B. Kupratis
Принадлежит: Individual

A propulsion system and a method of operating the propulsion system according to an exemplary aspect of the present disclosure includes powering an auxiliary fan with a gas turbine engine, the auxiliary fan along an auxiliary axis and the gas turbine engine along an engine axis, the auxiliary axis parallel to the engine axis.

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Номер записи: 27
09-05-2013 дата публикации

Fan rotor blade and fan

Номер: US20130111908A1
Автор: Hiroshi Kuroki, Kenro Obuchi, Rintaro Kajiwara, Takeshi Murooka
Принадлежит: IHI Corp

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

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Номер записи: 28
09-05-2013 дата публикации

Transition piece aft frame

Номер: US20130111910A1
Автор: Patrick Benedict MELTON, Richard Martin DiCintio
Принадлежит: General Electric Co

A transition piece aft frame is provided and includes a manifold having an interior that is receptive of fuel and formed to define fuel injection holes configured to inject the received fuel from the manifold interior toward a main flow of products of combustion flowing through the manifold. The manifold includes a main body having an interior facing surface that faces the main flow of the products of the combustion and along which the fuel injection holes are defined.

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Номер записи: 29
13-06-2013 дата публикации

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

Номер: US20130145745A1
Автор: Frederick M. Schwarz, Gregory A. Kohlenberg, Sean P. Zamora
Принадлежит: Individual

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

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Номер записи: 30
20-06-2013 дата публикации

JET ENGINE

Номер: US20130152544A1
Автор: MARTÍNEZ CASAÑ José Ramón
Принадлежит:

A reaction engine is disclosed mainly intended for aviation, but which as described herein can be adapted for an industrial engine, which makes use of spherical chambers and a pressure system in the blades of the rotor-stator unit which permits a perfect adjustment between said blades and the inner face of the stator, preventing pressure losses. While edges likewise articulated in the same way as the blades, execute a labyrinth seal. 121-. (canceled)22. Reaction engine comprising:a compressor block equipped with at least a compressor intended to carry out a compression of air that enters the engine and which comprises, in turn, a rotor and a stator,a combustion block which comprises at least a combustion chamber intended to house an ignition of a fuel together with high pressured air coming from the compressor block,at least a turbine actuated by exhaust gases produced in the combustion chamber and which comprises a torque transmission system with at least a shaft which is joined to at least the compressor thus carrying out the aforementioned compression of air,wherein the stator of the compressor is eccentric with respect to the rotor which permits the alternative radial displacement of an array of radial blades disposed on the rotor to carry out the closure in the radial direction of the array formed by the rotor and the stator, and said stator of the compressor comprises shock absorbers arranged on expulsion valves, said shock absorbers being controlled by electronic distribution rods disposed in compressor channels to control said shock absorbers and possible imbalances which may arise between the pressures.23. The engine of claim 22 , wherein the radial blades comprise elastic elements disposed on said blades and the rotor to provide the latter with automatic position recovery claim 22 , thus adjusting themselves on an inner face of the stator claim 22 , said elastic elements being located inside one of the lateral profiles of the blade.24. The engine of ...

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Номер записи: 31
01-08-2013 дата публикации

GAS TURBINE ENGINE SHAFT BEARING CONFIGURATION

Номер: US20130192199A1
Автор: Merry Brian D., Suciu Gabriel L.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A gas turbine engine includes a core housing that has an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. The shaft supports a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. The shaft includes a main shaft and a flex shaft having bellows. The flex shaft is secured to the main shaft at a first end and includes a second end opposite the first end. A geared architecture is coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture. The geared architecture includes a sun gear supported on the second end. A first bearing supports the shaft relative to the intermediate case and a second bearing supports the shaft relative to the inlet case. The second bearing is arranged radially outward from the flex shaft. 1. A gas turbine engine comprising:a core housing including an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path;a shaft supporting a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path, wherein the shaft includes a main shaft and a flex shaft having bellows, the flex shaft secured to the main shaft at a first end and including a second end opposite the first end;a geared architecture coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture, wherein the geared architecture includes a sun gear supported on the second end;a first bearing supporting the shaft relative to the intermediate case; anda second bearing supporting the shaft relative to the inlet case, and the second bearing is arranged radially outward from the flex shaft.2. The gas turbine engine according to claim 1 , wherein the shaft includes a hub secured to the main shaft claim 1 , and the compressor section includes a rotor mounted to the hub claim 1 , the hub ...

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Номер записи: 32
01-08-2013 дата публикации

Geared turbofan engine with counter-rotating shafts

Номер: US20130195624A1
Автор: Daniel Bernard Kupratis, Frederick M. Schwarz, Gabriel L. Suclu
Принадлежит: United Technologies Corp

A mid-turbine frame is incorporated into a turbine section of a gas turbine engine intermediate a high pressure turbine and a low pressure turbine. The high pressure and low pressure turbines rotate in opposite directions. A plurality of vanes redirect the flow downstream of the high pressure turbine as it approaches the low pressure turbine.

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Номер записи: 33
08-08-2013 дата публикации

ADAPTIVE FAN SYSTEM FOR A VARIABLE CYCLE TURBOFAN ENGINE

Номер: US20130199156A1
Автор: JR. Robert A., Ress, Starr Matthew J., Stratton Steve
Принадлежит:

One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique variable cycle gas turbine engine. Another embodiment is a unique adaptive fan system for a variable cycle turbofan engine having at least one turbine. Another embodiment is a unique method for operating a variable cycle gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and related systems. 1. A gas turbine engine , comprising;a compressor structured to compress an airflow received at the compressor and to output the compressed airflow as a compressor discharge airflow;a combustor in fluid communication with said compressor, said combustor being structured to combust a mixture of a fuel and at least some of said compressor discharge airflow to generate a hot working airflow;a turbine in fluid communication with said combustor, said turbine being configured to extract a mechanical power from said hot working airflow;a shaft coupled to said turbine, said shaft being configured to receive and transmit said mechanical power from said turbine;a first rotating load powered by said shaft;a second rotating load powered by said shaft; anda transmission system coupled to said shaft, said transmission system being structured to selectively vary a speed at which power is supplied from said shaft to said second rotating load relative to a speed of at least one of said shaft and said first rotating load.2. The gas turbine engine of claim 1 , configured as a turbofan engine claim 1 , wherein said first rotating load is a fan stage.3. The gas turbine engine of claim 2 , configured as a variable cycle turbofan engine claim 2 , wherein said second rotating load is an other fan stage.4. The gas turbine engine of claim 3 , further comprising:a first bypass duct configured to bypass at least a portion of the output of said fan stage to provide a thrust component; anda second bypass duct configured to ...

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Номер записи: 34
08-08-2013 дата публикации

Fan-less Thrust Generating Component for Air-bypass Engine

Номер: US20130199186A1
Автор: Ahmed Yassin SOUEIDAN
Принадлежит: Individual

A fan-less thrust generating component for air-bypass engine includes an outer casing, a fluid moving device, and a plurality of core components. The fluid moving device and the plurality of core components are positioned within an intake opening of the outer casing. The fluid moving device is positioned in front of the plurality of core components and connected to with the plurality of core components. The fluid moving device includes a hub, a helical fin, and a fluid inlet. The helical fin is connected around the hub and the fluid inlet is positioned on the helical fin. The fluid moving device pulls air fluid into the intake opening through the fluid inlet upon rotation so that thrust can be generated within the air-bypass engine.

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Номер записи: 35
08-08-2013 дата публикации

AIRCRAFT, PROPULSION SYSTEM, AND INLET WITH SUPERSONIC COMPRESSION

Номер: US20130199621A1
Автор: Anderson Jon T., Davis Zachary S., Hajic Kenneth J., Klinge John D.
Принадлежит:

Embodiments of an inlet, aircraft, and propulsion system includes an inlet spike with a compression surface configured to decelerate air flow from supersonic speed to subsonic speed before the air flow enters a cowling throat and diffuser section of the propulsion system. 1. A propulsion system comprising:an engine;a cowling surrounding at least a portion of the engine; andan inlet including at least one compression surface upstream and external to the cowling, the at least one compression surface is configured to slow air flow from supersonic speed to subsonic speed before the air flow enters the cowling.2. The propulsion system according to further comprising:a leading edge of the at least one compression surface configured to generate an oblique shock at supersonic speed.3. The propulsion system according to further comprising:the at least one compression surface includes a first conical compression surface and isentropic compression surface, the first conical compression surface is configured between the leading edge and the isentropic compression surface.4. The propulsion system according to further comprising:a second conical compression surface configured adjacent to and downstream of the isentropic compression surface.5. The propulsion system according to further comprising:a leading edge of the cowling is configured adjacent to the second conical compression surface and is separated from the second conical compression surface to form a cowling throat through which the air flows to the engine.6. The propulsion system according to claim 1 , further comprising:a diffuser section downstream of the leading edge of the cowling, wherein the air flow does not experience a normal shock along the inlet so that a boundary layer of the air flow in the diffuser section remains undisturbed.7. The propulsion system according to claim 1 , wherein the system does not include bleed air components to smooth the air flow after a normal shock at the inlet.8. A supersonic ...

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Номер записи: 36
08-08-2013 дата публикации

Gas turbine engine nacelle

Номер: US20130199624A1
Автор: Angus Roy Smith, Dimitris Arzoglou
Принадлежит: Rolls Royce PLC

A method of designing a nacelle for a gas turbine engine includes defining an initial geometry for the nacelle, pivoting, at each of one or more positions on a highlight and in a plane of a respective longitudinal section, a front portion of a section of an intake lip lying in the respective longitudinal section about the highlight, and obtaining an altered geometry for the nacelle by adjusting surfaces of the intake neighbouring the pivoted front portion to smoothly blend the surfaces of the pivoted front portion to unadjusted surfaces of the intake further removed from the pivoted front portion.

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Номер записи: 37
08-08-2013 дата публикации

GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT

Номер: US20130202415A1
Автор: Dye Christopher M., Karl Hasel, Merry Brian D., Staubach Joseph B., Suciu Gabriel L.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A gas turbine engine includes a fan section, a gear arrangement configured to drive the fan section, a compressor section and a turbine section. The compressor section includes a low pressure compressor section and a high pressure compressor section. The turbine section is configured to drive compressor section and the gear arrangement. The engine includes a combination of quantities providing beneficial operation. 1. A gas turbine engine comprising:a fan section;a gear arrangement configured to drive said fan section;a compressor section, including both a low pressure compressor section and a high pressure compressor section;a turbine section configured to drive said compressor section and said gear arrangement; andwherein an overall pressure ratio provided by the combination of said low pressure compressor section and said high pressure compressor section being greater than or equal to about 35; wherein a pressure ratio across said low pressure compressor section being less than or equal to about 8; wherein said turbine section includes a low pressure turbine section driving said low pressure compressor section; and wherein a pressure ratio across said low pressure turbine section being greater than about 5.2. A gas turbine engine comprising:a fan section;a gear arrangement configured to drive said fan section;a compressor section, including both a low pressure compressor section and a high pressure compressor section;a turbine section configured to drive said compressor section and said gear arrangement; andwherein an overall pressure ratio provided by the combination of said low pressure compressor section and said high pressure compressor section being greater than or equal to about 35; wherein a pressure ratio across said high pressure compressor section being greater than or equal to about 7; wherein said turbine section includes a low pressure turbine section driving said low pressure compressor section; and wherein a pressure ratio across said low pressure ...

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Номер записи: 38
15-08-2013 дата публикации

Aircraft gas turbine thrust-reversing device

Номер: US20130205753A1
Автор: Todorovic Predrag
Принадлежит: ROLLS-ROYCE DEUTSCHLAND LTD & CO KG

An aircraft gas turbine thrust-reversing device has an engine cowling, the rear area of which is displaceable in the axial direction of the engine from a closed forward thrust position into a rearwardly displaced thrust reversal position, resulting in an essentially annular space to a front and stationary area of the engine cowling. The rear area of the engine cowling is coupled to deflecting elements and blocker doors, which in the forward thrust position are arranged completely inside the front area of the engine cowling. A drive element is provided between the front area and the rear area, which effects the axial displacement of the rear area. The drive element is a two-stage drive element, effecting in a first stage an axial displacement by an axial partial displacement path, and in a second stage an axial displacement by the full axial displacement path. 1. Aircraft gas turbine thrust-reversing device with an engine and with an engine cowling , the rear area of which being displaceable in the axial direction of the engine from a closed forward thrust position into a rearwardly displaced thrust reversal position , resulting in an essentially annular space to a front and stationary area of the engine cowling , with the rear area of the engine cowling being coupled to deflecting elements and blocker doors , which in the forward thrust position are arranged completely inside the front area of the engine cowling , with at least one drive element being provided between the front area and the rear area , which effects the axial displacement of the rear area , with the drive element being designed as a two-stage drive element , and effecting in a first stage an axial displacement by an axial partial displacement path , and in a second stage an axial displacement by the full axial displacement path.2. Device in accordance with claim 1 , characterized in that the drive element is designed telescopic.3. Device in accordance with claim 1 , characterized in that the drive ...

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Номер записи: 39
29-08-2013 дата публикации

Counter-rotating low pressure turbine with gear system mounted to mid turbine frame

Номер: US20130219856A1
Автор: Brian D. Merry, Gabriel L. Suciu
Принадлежит: Individual

A gas turbine engine includes a shaft defining an axis of rotation. An outer turbine rotor directly drives the shaft and includes an outer set of blades. An inner turbine rotor has an inner set of blades interspersed with the outer set of blades. The inner turbine rotor is configured to rotate in an opposite direction about the axis of rotation from the outer turbine rotor. A gear system couples the inner turbine rotor to the shaft and is configured to rotate the inner set of blades at a faster speed than the outer set of blades. The gear system is mounted to a mid-turbine frame.

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Номер записи: 40
29-08-2013 дата публикации

Counter-rotating low pressure turbine without turbine exhaust case

Номер: US20130219860A1
Автор: Brian D. Merry, Gabriel L. Suciu
Принадлежит: Individual

A gas turbine engine includes a shaft defining an axis of rotation. An inner rotor directly drives the shaft and includes an inner set of blades. An outer rotor has an outer set of blades interspersed with the inner set of blades. The outer rotor is configured to rotate in an opposite direction about the axis of rotation from the inner rotor. A gear system is engaged to the outer rotor and is positioned upstream of the inner set of blades.

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Номер записи: 41
29-08-2013 дата публикации

Geared turbofan architecture for improved thrust density

Номер: US20130219907A1
Автор: Daniel Bernard Kupratis, Frederick M. Schwarz, Gabriel L. Suciu
Принадлежит: United Technologies Corp

A turbine engine includes a fan, a compressor section having a low pressure compressor section and a high pressure compressor section, a combustor in fluid communication with the compressor section and a turbine section in fluid communication with the combustor. The turbine section includes a low pressure turbine section and a high pressure turbine section. The low pressure compressor section, the low pressure turbine section and the fan rotate in a first direction whereas the high pressure compressor section and the high pressure turbine section rotate in a second direction opposite the first direction.

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Номер записи: 42
19-09-2013 дата публикации

Pump system for tms aoc reduction

Номер: US20130239588A1
Автор: Brian D. Merry, Christopher M. Dye, Gabriel L. Suciu, Jorn A. Glahn
Принадлежит: United Technologies Corp

An engine includes a duct containing a flow of cool air and a pump system having an impeller with an inlet for receiving air from the duct and an outlet for discharging air into a discharge manifold. The discharge manifold containing at least one heat exchanger which forms part of a thermal management system.

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Номер записи: 43
03-10-2013 дата публикации

Geared turbofan engine with power density range

Номер: US20130255275A1
Автор: Daniel Bernard Kupratis, Frederick M. Schwarz
Принадлежит: United Technologies Corp

A gas turbine engine turbine has a high pressure turbine configured to rotate with a high pressure compressor as a high pressure spool in a first direction about a central axis and a low pressure turbine configured to rotate with a low pressure compressor as a low pressure spool in the first direction about the central axis. A power density is greater than or equal to about 1.5 and less than or equal to about 5.5 lbf/cubic inches. A fan is connected to the low pressure spool via a speed changing mechanism and rotates in the first direction.

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Номер записи: 44
10-10-2013 дата публикации

SELF-MOUNTED CASCADE FOR A THRUST REVERSER

Номер: US20130266423A1
Автор: VAUCHEL Guy Bernard
Принадлежит: AIRCELLE

The present disclosure relates to a self-mounted cascade for a thrust reverser of an airplane jet engine nacelle. The cascade includes an upstream portion and an opposite downstream portion, along with respective connection means so that two adjacent cascades are directly connected to one another only in the respective downstream portions by downstream connection means thereof. The downstream portion corresponds to an area extending from a downstream side edge over a length less than or equal to N times the length of the last cavity located along said downstream side edge, where N is less than 3. The present disclosure is of use in the field of airplane jet engine nacelles. 2. The cascade according to claim 1 , wherein at least one of the first and second downstream connecting means is positioned partially withdrawn from the concerned transverse edge.3. The cascade according to claim 2 , wherein one of the first and second downstream connecting means is positioned partially withdrawn relative to the concerned transverse edge toward the inside of the cascade claim 2 , and the other of the first and second downstream connecting means is positioned partially withdrawn relative to the concerned transverse edge toward the outside of the cascade.4. The cascade according to claim 1 , wherein one of the first and second downstream connecting means has at least one passage opening for a fastening member situated withdrawn relative to the concerned transverse edge toward the inside of the cascade claim 1 , while the other of the first and second downstream connecting means has at least one passage opening for a fastening member situated withdrawn relative to the concerned transverse edge toward the outside of the cascade.5. The cascade according to claim 1 , wherein the first downstream connecting means has at least one first connecting tab and the second downstream connecting means has at least one second connecting tab.6. The cascade according to claim 5 , wherein the first ...

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Номер записи: 45
17-10-2013 дата публикации

DEVICE FOR CONTROLLING A VARIABLE SECTION NOZZLE OF AN AIRCRAFT

Номер: US20130269312A1
Автор: Bulin Guillaume, Devienne Nicolas, Oberle Patrick
Принадлежит: AIRBUS OPERATIONS (S.A.S)

A control device for controlling a variable section nozzle of an aircraft power plant, the variable section nozzle including one or several movable parts capable of modifying the nozzle section and connected by a mechanical transmission chain to an actuator. The control device includes a system for regulating the power plant connected to a control member configured to control the actuator. The control device includes a single control member, an immobilization unit configured to immobilize all the movable parts which are deactivated only when the regulation system controls the positional change of the movable part or parts and a determination unit configured to determine the actual position of the movable part or parts. 1. A control device for controlling a variable section nozzle of an aircraft power plant , said variable section nozzle comprising one or several movable parts configured to modify the nozzle section and connected by a mechanical transmission chain to an actuator , said control device comprising:a system configured to regulate the power plant connected to a control member configured to control the actuator, characterized in that the control device comprises a single said control member, an immobilization unit configured to immobilize all the movable parts which are deactivated only when the regulation system controls the positional change of the movable part or parts, and a determination unit configured to determine the actual position of the movable part or parts.2. The control device according to claim 1 , characterized in that the determination unit that determines the actual position of the movable part or parts is connected to the regulation system to indicate said actual position thereto.3. The control device according to claim 1 , further comprising a connection between the immobilization unit and the control member which controls their deactivation.4. The control device according to claim 1 , characterized in that the immobilization unit ...

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Номер записи: 46
31-10-2013 дата публикации

Geared Architecture for High Speed and Small Volume Fan Drive Turbine

Номер: US20130287575A1
Автор: Daniel Bernard Kupratis, Frederick M. Schwarz, Gabriel L. Suciu, Jason Husband, Michael E. McCune, William K. Ackermann
Принадлежит: United Technologies Corp

A gas turbine engine includes a flex mount for a fan drive gear system. A very high speed fan drive turbine drives the fan drive gear system.

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Номер записи: 47
14-11-2013 дата публикации

Gas turbine engine systems and related methods involving multiple gas turbine cores

Номер: US20130298565A1
Автор: Gary D. Roberge
Принадлежит: United Technologies Corp

Gas turbine engine systems and related methods involving multiple gas turbine cores are provided. In this regard, a representative gas turbine engine includes: an inlet; a blade assembly mounted to receive intake air via the inlet; and multiple gas turbine cores located downstream of the blade assembly, each of the multiple gas turbine cores being independently operative in a first state, in which rotational energy is provided to rotate the blade assembly, and a second state, in which rotational energy is not provided to rotate the blade assembly.

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Номер записи: 48
21-11-2013 дата публикации

COMBINED TURBOJET AND RAMJET ENGINE

Номер: US20130305686A1
Автор: Conrardy Jean-Marie, Soulier Nicolas
Принадлежит: SNECMA

A combined engine includes a turbopump including a pump injecting hydrogen into a heater arranged in an outer casing downstream from a central body, and a subsonic turbine driving the pump, which turbine receives partially-expanded hydrogen collected at an outlet from the heater to apply the hydrogen to a supersonic turbine to operate the engine as a turbojet. The hydrogen from the supersonic turbine is collected in tubes inside the central body to be sent to a combustion chamber defined downstream from the central body, while the hydrogen that is partially expanded in the subsonic turbine is sent directly to the combustion chamber via injectors to operate the engine as a ramjet. 17-. (canceled)8. A combined turbojet and ramjet engine comprising:an outer casing;a central body connected to the outer casing by structural arms and co-operating with the outer casing to form an air inlet sleeve and an air flow passage;at least one first air compressor stage comprising a first turbine arranged in the central body and a first rotor comprising blades arranged in the air flow passage and capable selectively of being driven by the first turbine to operate the engine as a turbojet and being feathered to operate the engine as a ramjet; anda turbopump comprising a pump that is fed with liquid hydrogen from a hydrogen tank to inject the hydrogen into a heater arranged in the outer casing downstream from the central body and a subsonic turbine driving the pump, the turbine receiving the partially-expanded hydrogen collected at an outlet from the heater, the hydrogen that is partially expanded in the subsonic turbine being applied to the first turbine, which is a supersonic turbine, to operate the engine as a turbojet, the hydrogen from the first supersonic turbine then being collected in first tubes inside the central body to be sent to the combustion chamber defined inside the casing downstream from the central body, while the hydrogen that is partially expanded in the subsonic ...

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Номер записи: 49
05-12-2013 дата публикации

AVIATION GAS TURBINE THRUST REVERSING DEVICE

Номер: US20130318945A1
Автор: Todorovic Predrag
Принадлежит: Rolls-Royce Deutschland Ltd. & Co KG

The present invention relates to an aircraft gas turbine thrust-reversing device with an engine having an engine cowling, the rear area of which can be displaced in the axial direction of the engine from a closed forward thrust position into a rearwardly displaced thrust reversal position, resulting in an essentially annular free space towards a forward stationary area of the engine cowling, with the rear area of the engine cowling being functionally coupled to deflecting elements arranged in the forward thrust position within the front area of the engine cowling, with the deflecting elements during displacement of the rear area of the engine cowling being moveable on a partial-circular path facing the central axis of the engine and contactable by their rear end areas with a cowling of the core engine. 1. Aircraft gas turbine thrust-reversing device with an engine having an engine cowling , the rear area of which can be displaced in the axial direction of the engine from a closed forward thrust position into a rearwardly displaced thrust reversal position , resulting in an essentially annular free space towards a forward stationary area of the engine cowling , with the rear area of the engine cowling being functionally coupled to deflecting elements arranged in the forward thrust position within the front area of the engine cowling , with the deflecting elements during displacement of the rear area of the engine cowling being moveable on a partial-circular path facing the central axis of the engine and contactable by their rear end areas with a cowling of the core engine.2. Device in accordance with claim 1 , characterized in that the deflecting element includes a plurality of guiding elements.3. Device in accordance with claim 1 , characterized in that the deflecting element is designed grid-like.4. Device in accordance with claim 1 , characterized in that the deflecting element is designed cascade-like.5. Device in accordance with claim 1 , characterized in that ...

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Номер записи: 50
05-12-2013 дата публикации

Nacelle bifurcation for gas turbine engine

Номер: US20130319002A1
Автор: Dmitriy B. Sidelkovskiy, Oleg Petrenko, Robert E. Malecki
Принадлежит: Individual

A nacelle structure for a gas turbine engine includes a core engine nacelle disposed about an engine axis and an outer nacelle disposed about the core engine nacelle. A bifurcation extends between the outer nacelle and the core engine nacelle along a bifurcation axis extending between the outer nacelle and the core engine nacelle. The bifurcation includes at least one mounting surface that is disposed at a non-normal angle relative to the bifurcation axis.

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Номер записи: 51
12-12-2013 дата публикации

Devices and Methods to Optimize Aircraft Power Plant and Aircraft Operations

Номер: US20130327014A1
Автор: Moulebhar Djamal
Принадлежит:

Several improvements to optimize aircraft power plant and aircraft operations are disclosed, as well as methods of using these improvements to reduce fuel consumption, gas emission, noise, aircraft weight, maintenance costs, operating costs, aircraft incident and accidents, and improving aircraft performance. The improvements consist of a power plant fitted with a front propulsor, core engine, and aft propulsor. The fan of each propulsor is separated mechanically from the core engine. The front fan is separated mechanically from the aft fan. The aft fan is driven by free turbine that is supplied by exhaust gas of the core engine. If the core engine fails, both propulsors operate and provide thrust and reversed thrust when needed. If one propulsor fails, the other propulsor of the same power plant operates and provides thrust and reversed thrust. 1. An aircraft comprising an aircraft power plants and an Auxiliary Power Unit (APU); the aircraft power plant comprises a front propulsor , a core engine , and an aft propulsor; the aircraft power plant can be configured in several configurations; the power plant can be configured as an advanced dual fan and comprises front propulsor , core engine , and aft-propulsor wherein:a) The front propulsor comprises a front fan, a front motor-generator, and a front air turbine; the front fan is driven by a front motor-generator and an air turbine to provide the thrust or the reversed thrust when needed; the front motor-generator can be connected to the front fan through a clutch and the air turbine is may be connected to the front fan through a clutch; if the front motor-generator fails, this motor-generator can be disconnected from the front fan through its clutch, such that the air turbine drives the front fan; if the front air turbine fails this air turbine can be disconnected from the front fan through its clutch, such that the motor-generator drives the fan;The front fan is separated mechanically and aerodynamically from the ...

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Номер записи: 52
26-12-2013 дата публикации

FAN STAGGER ANGLE FOR GEARED GAS TURBINE ENGINE

Номер: US20130340406A1
Автор: Gallagher Edward J., Li Linda S., Liu Ling, Monzon Byron R., Whitlow Darryl
Принадлежит:

A gas turbine engine includes a spool, a turbine coupled with the spool, a propulsor coupled to be rotated about an axis by the turbine through the spool and a gear assembly coupled between the propulsor and the spool such that rotation of the spool results in rotation of the propulsor at a different speed than the spool. The propulsor includes a hub and a row of propulsor blades that extends from the hub. Each of the propulsor blades has a span between a root at the hub and a tip, and a chord between a leading edge and a trailing edge such that the chord forms a stagger angle α with the axis. The stagger angle α is less than 62° at all positions along the span, with said hub being at 0% of the span and the tip being at 100% of the span. 1. A gas turbine engine comprising:a spool;a turbine coupled with said spool;a propulsor coupled to be rotated about an axis through said spool; anda gear assembly coupled between said propulsor and said spool such that rotation of said spool results in rotation of said propulsor at a different speed than said spool,said propulsor including a hub and a row of propulsor blades extending from said hub, each of said propulsor blades having a span between a root at said hub and a tip, and a chord between a leading edge and a trailing edge such that said chord forms a stagger angle α with said axis, and said stagger angle α is less than 62° at all positions along said span, with said hub being at 0% of said span and said tip being at 100% of said span.2. The gas turbine engine as recited in claim 1 , wherein said stagger angle α at 25% of said span is less than 23°.3. The gas turbine engine as recited in claim 1 , wherein said stagger angle α at 25% of said span is 16-21°.4. The gas turbine engine as recited in claim 1 , wherein said stagger angle α at 50% of said span is less than 35°.5. The gas turbine engine as recited in claim 1 , wherein said stagger angle α at 50% of said span is 28-33°.6. The gas turbine engine as recited in claim ...

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Номер записи: 53
26-12-2013 дата публикации

GAS TURBINE ENGINE WITH REVERSE-FLOW CORE HAVING A BYPASS FLOW SPLITTER

Номер: US20130343867A1
Автор: Kupratis Daniel Bernard, Moon Francis R.
Принадлежит:

A gas turbine engine includes a bypass flow path. A core flow path has a reverse duct configured to reverse a direction of core flow. The core flow path includes an exhaust duct in fluid communication with the bypass flow path and is configured to introduce core exhaust flow in the core flow path back into the bypass flow path. A splitter is arranged in the bypass flow path and adjoins the exhaust duct. 1. A gas turbine engine comprising:a bypass flow path;a core flow path having a reverse duct configured to reverse a direction of core flow, the core flow path including an exhaust duct in fluid communication with the bypass flow path and configured to introduce core exhaust flow in the core flow path back into the bypass flow path; anda splitter arranged in the bypass flow path and adjoining the exhaust duct.2. The gas turbine engine according to claim 1 , comprising a low pressure compressor and a low pressure turbine mounted on a low spool near one another claim 1 , and a high pressure turbine and a high pressure compressor mounted on a high spool near one another claim 1 , the low and high pressure turbines arranged axially between the low and high pressure compressors.3. The gas turbine engine according to claim 1 , wherein the reverse duct and the exhaust duct direct the core flow in a radial direction.4. The gas turbine engine according to claim 1 , wherein the splitter is a linearly extending annular structure.5. The gas turbine engine according to claim 4 , wherein the splitter and exhaust duct provide a unitary structure.6. The gas turbine engine according to claim 5 , comprising an array of circumferentially spaced claim 5 , discrete exhaust ducts adjoining the splitter.7. The gas turbine engine according to claim 1 , wherein the splitter is hollow and includes first and second sides providing a cavity.8. The gas turbine engine according to claim 7 , wherein the splitter includes an inlet configured to communicate bypass flow to the cavity.9. The gas ...

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Номер записи: 54
30-01-2014 дата публикации

AIRCRAFT HAVING A ROTATING TURBINE ENGINE

Номер: US20140027581A1
Автор: Brothers Kamyar
Принадлежит:

The present invention describes an aircraft having a rotating engine. The aircraft comprises a means for rotating the engine 360° to allow directional control of the aircraft via rotation of the engine's thrust output. The aircraft can further comprise a plurality of flap members located on the wings of the aircraft and in communication with the rotating gas turbine engine to further control and stabilize flight of the aircraft. The gas turbine engine of the aircraft can comprise a compressor, a combustion chamber, and at least two turbines mounted oppositely to the combustion chamber, such that the gas turbine engine is capable of generating more thrust from a single engine. 1. An aircraft comprising:a main frame having a bottom domed section adapted for rotation;a gas turbine engine that is adapted for rotation when said bottom domed section rotates,whereby rotation of said bottom domed section controls directional movement of the aircraft.2. The aircraft of claim 1 , further comprising a top section that is adapted for rotation.3. The aircraft of claim 2 , wherein said gas turbine engine further comprises:an inlet positioned within said top section to admit air into said engine and at least one outlet positioned within said bottom domed section to eject products of combustion of said engine, wherein said gas turbine engine is at least partially attached with said bottom domed section;a compressor, said compressor capable of pressurizing air;a combustion chamber in communication with said compressor, said combustion chamber having a means for igniting pressurized air from said compressor to create ignited gas;a first turbine connected to a first side of said combustion chamber, wherein said ignited gas from said combustion chamber motivates said first turbine to generate force in a first direction; anda second turbine connected to a second side of said combustion chamber that is diametrical to said first side of said combustion chamber, wherein said ignited gas ...

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Номер записи: 55
30-01-2014 дата публикации

Geared fan with inner counter rotating compressor

Номер: US20140030060A1
Автор: John W. Magowan
Принадлежит: Individual

A gas turbine engine comprises a fan section including a fan hub supporting a plurality of fan blades for rotation relative to a fan case, a shaft rotatable relative to the fan case about an engine center axis, and a geared architecture driven by the shaft to provide driving output to rotate the fan hub. A compressor is positioned forward of the geared architecture and radially inward of the fan blades, with the compressor being driven by the shaft.

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Номер записи: 56
06-02-2014 дата публикации

Flow discharge device

Номер: US20140033733A1
Автор: Kevin Mark Britchford, Pamela Joan PHILLIPSON, Vincent Benoit SAVARIN
Принадлежит: Rolls Royce PLC

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

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Номер записи: 57
06-02-2014 дата публикации

Retrofitable auxiliary inlet scoop

Номер: US20140037437A1
Автор: Joshua V. WOOD, Michael Joseph Murphy
Принадлежит: Individual

A gas turbine engine includes a fan nacelle and a core nacelle spaced radially inwardly of the fan nacelle to define a bypass flowpath. The engine includes a scoop to direct air from the bypass flowpath into an engine core. The scoop is comprised of a first piece extending outwardly of an outer surface the core nacelle to define an air inlet and a second piece extending inwardly of the core nacelle to define an air outlet.

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Номер записи: 58
13-02-2014 дата публикации

FLOW CONDUCTING ASSEMBLY FOR COOLING THE LOW-PRESSURE TURBINE HOUSING OF A GAS TURBINE JET ENGINE

Номер: US20140041360A1
Автор: Eichler Christian, Feldmann Manfred, Muschkorgel Sandra, Schinko Norbert
Принадлежит:

A gas turbine jet engine having a main flow channel () and a housing structure () which radially surrounds this main flow channel, in which a housing gas flow flows in the same direction as the main flow in the main flow channel, the housing structure having a flow conducting assembly (), with the aid of which the flow of the housing gas flow to and/or along the main flow channel may be adjusted. 1. A gas turbine jet engine comprising:a main flow channel; anda housing structure radially surrounding the main flow channel and in which a housing gas flow flows in a same direction as a main flow in the main flow channel, the housing structure having an adjustable flow conducting assembly, with the aid of which the flow of the housing gas flow to and/or along the main flow channel is adjustable.2. The gas turbine jet engine as recited in wherein the flow conducting assembly including at least one flow conducting sheet dividing the housing gas flow into at least two partial flows claim 1 , one partial flow running near the main flow channel and another partial flow running at a distance from the main flow channel.3. The gas turbine jet engine as recited in wherein channels for the partial flows are closable or throttled.4. The gas turbine jet engine as recited in wherein the flow conducting sheet is adjustable or has variably closable openings for variably adjusting the partial flows.5. The gas turbine jet engine as recited in wherein the flow conducting assembly has valves for closing openings or adjusting the gas flow.6. The gas turbine as jet engine recited in wherein the valves are throttle valves.7. The gas turbine jet engine as recited in further comprising a controller or regulator controlling or regulating the housing gas flow with the aid of the flow conducting assembly.8. The gas turbine jet engine as recited in wherein the flow conducting assembly is situated in the area of a low-pressure turbine.9. The gas turbine jet engine as recited in further comprising a ...

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Номер записи: 59
06-03-2014 дата публикации

Magnetic advanced generation jet electric turbine

Номер: US20140060005A1
Автор: Richard H. Lugg
Принадлежит: Richard H. Lugg

Supersonic Magnetic Advanced Generation Jet Electric Turbine (S-MAGJET) described herein, and a subsonic derivative, MAGJET, integrate a gas power turbine, superconducting electric power and propulsion generation, and magnetic power flux field systems along with an ion plasma annular injection combustor which utilizes alternative petroleum-based fuel and combustion cycles to create a hybrid turbine turbomachine for aerospace propulsion. The propulsion unit is able to achieve a dramatic increase in horsepower, combustion and propulsion efficiency, and weight reduction. In addition, the turbomachinery structures may be disposed within an exoskeleton architecture that achieves an increase in thrust to weight ratio with a concomitant increase in fuel efficiency and power generation. The engine continuously adjusts the temperature, pressure and mass airflow requirements using an electromagnetic power management system architecture. Engine performance may be controlled across the entire desired flight envelope, whether subsonic, transonic or supersonic flight conditions.

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Номер записи: 60
06-03-2014 дата публикации

Seal design and active clearance control strategy for turbomachines

Номер: US20140064909A1
Автор: Deepak Trivedi, Murat Inalpolat, Norman Arnold Turnquist, Xiaoqing Zheng
Принадлежит: General Electric Co

A labyrinth seal design, an actuation control clearance strategy, and a method of operating a turbomachine. The labyrinth seal design including a plurality of features configured to open and close radial clearances in response to relative axial movement between a stationary component and a rotating component. The actuation control clearance strategy and method of operating a turbomachine effective to achieve relative motion between a rotating component and a stationary component of the turbomachine using active elements. Axial displacement of the rotating component relative to the stationary component provides an adjustment in a radial clearance at one or more sealing locations between the rotating component and the stationary component to suit a given operating condition of the turbomachine.

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Номер записи: 61
06-03-2014 дата публикации

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

Номер: US20140064915A1
Автор: Jean-Pierre Serey, Patrick Dunleavy, Richard Masson
Принадлежит: SNECMA SAS

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

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Номер записи: 62
27-03-2014 дата публикации

GEARED TURBOFAN PRIMARY AND SECONDARY NOZZLE INTEGRATION GEOMETRY

Номер: US20140083079A1
Автор: Kohlenberg Gregory A., Kupratis Daniel Bernard
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A disclosed example geared turbofan engine includes a fan section including a plurality of fan blades rotatable about an axis and a core engine section defined about an engine axis. The core engine section includes a primary nozzle including a primary outer diameter at a primary nozzle trailing edge and a primary maximum inner diameter forward of the primary trailing edge. A bypass passage is defined between an inner nacelle surrounding the core engine section and an outer nacelle and includes a secondary nozzle. The secondary nozzle includes an outer diameter at a secondary nozzle trailing edge and a secondary maximum inner diameter forward of the secondary trailing edge. A ratio between the maximum inner diameter of the primary nozzle and an outer diameter at the trailing edge of the primary nozzle and a ratio between the maximum inner diameter of the secondary trailing edge and the outer diameter at the trailing edge of the secondary nozzle are both less than about 0.700. 1. A gas turbine engine comprising:a core engine section defined about an engine axis;a bypass passage defined about the core engine section; anda primary nozzle defined as a part of the core engine including a primary outer diameter at a primary nozzle trailing edge and a primary maximum inner diameter forward of the primary trailing edge, wherein a ratio between the inner diameter of the primary nozzle to an outer diameter at the trailing edge of the primary nozzle is less than about 0.700.2. The gas turbine engine as recited in claim 1 , wherein the bypass passage includes a secondary nozzle including a secondary outer diameter at a secondary nozzle trailing edge and a secondary maximum inner diameter forward of the secondary trailing edge claim 1 , wherein a ratio between the inner diameter of the secondary nozzle to the outer diameter at the trailing edge of the secondary nozzle is less than about 0.700.3. The gas turbine engine as recited in claim 1 , wherein the ratio between the inner ...

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Номер записи: 63
27-03-2014 дата публикации

TRANSITION DUCT FOR USE IN A TURBINE ENGINE AND METHOD OF ASSEMBLY

Номер: US20140086739A1
Автор: Carson Scott Michael, Keith Brian David, Machnaim Joseph
Принадлежит: GENERAL ELECTRIC COMPANY

A transition duct for use in a turbine engine is provided. The transition duct includes a radially inner wall and a radially outer wall positioned about the radially inner wall defining a flow passage therebetween. The radially outer wall extends and is contoured from an upstream end to a downstream end of the transition duct. As such, the slope of the radially outer wall increases from the upstream end to a predetermined axial location and decreases from the predetermined axial location to the downstream end. 1. A transition duct for use in a turbine engine , the transition duct comprising:a radially inner wall; anda radially outer wall positioned about said radially inner wall defining a flow passage therebetween, said radially outer wall extends and is contoured from an upstream end to a downstream end of the transition duct such that a slope of said radially outer wall increases from said upstream end to a predetermined axial location and decreases from the predetermined axial location to said downstream end.2. The transition duct in accordance with further comprising a fairing that extends radially between said radially inner wall and said radially outer wall within said flow passage claim 1 , wherein said fairing comprises an aerodynamic cross-sectional shape.3. The transition duct in accordance with claim 2 , wherein the predetermined axial location corresponds to an axial location of a thickest cross-sectional portion of said fairing such that a maximum slope of said radially outer wall is at the predetermined axial location.4. The transition duct in accordance with claim 1 , wherein the slope of said radially outer wall increases from about 0° at said upstream end to greater than about 40° at the predetermined axial location.5. The transition duct in accordance with claim 1 , wherein said radially outer wall comprises a maximum wall slope at the predetermined axial location claim 1 , the maximum wall slope from about 40° to about 50°.6. The transition duct ...

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Номер записи: 64
10-04-2014 дата публикации

SYSTEMS AND METHODS INVOLVING MULTIPLE TORQUE PATHS FOR GAS TURBINE ENGINES

Номер: US20140096508A1
Автор: McCune Michael E., Merry Brian D., Suciu Gabriel L.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A turbofan engine includes a fan, a compressor section, a combustor in fluid communication with the compressor section, a turbine section in fluid communication with the combustor, a shaft configured to be driven by the turbine section and coupled to the compressor section through a first torque load path, and a speed reduction mechanism configured to be driven by the shaft through a second torque load path separate from the first load path for rotating the fan. 1. A turbofan engine comprising:a fan;a compressor section;a combustor in fluid communication with the compressor section;a turbine section in fluid communication with the combustor;a shaft configured to be driven by the turbine section and coupled to the compressor section through a first torque load path; anda speed reduction mechanism configured to be driven by the shaft through a second torque load path separate from the first load path for rotating the fan.2. The turbofan engine as recited in claim 1 , including an intersection between the first torque load path and the shaft and a thrust bearing located adjacent to the intersection between the first torque path and the first shaft.3. The turbofan engine as recited in claim 1 , wherein the compressor section includes a first compressor section immediately aft of the fan and the first torque load path couples the shaft to the first compressor section.4. The turbofan engine as recited in claim 1 , including a first spool segment mechanically coupling the shaft to the compressor section and defining the first torque load path.5. The turbofan engine as recited in claim 4 , including a second spool segment mechanically coupling the shaft to the speed reduction mechanism and defining the second torque load path.6. The turbofan engine as recited in claim 5 , wherein the first spool segment is operative to transfer torque from the shaft to the compressor and not to the speed reduction mechanism.7. The turbofan engine as recited in claim 6 , wherein the second ...

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Номер записи: 65
10-04-2014 дата публикации

Geared Turbofan Engine With Increased Bypass Ratio and Compressor Ratio ...

Номер: US20140096509A1
Автор: Karl L. Hasel
Принадлежит: United Technologies Corp

A gas turbine engine is typically comprised of a fan stage, multiple compressor stages, and multiple turbine stages. These stages are made up of alternating rotating blade rows and static vane rows. The total number of blades and vanes is the airfoil count. An overall pressure ratio is greater than 30. A bypass ratio is greater than 8. A stage ratio is the product of the bypass ratio and the overall pressure ratio divided by the number of stages. An airfoil ratio is that product divided by the airfoil count. The stage ratio is greater than or equal to 22 and/or the airfoil ratio is greater than or equal to 0.12.

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Номер записи: 66
10-04-2014 дата публикации

Duct assembly

Номер: US20140099198A1
Автор: Joseph Barnsdale Cooper
Принадлежит: Rolls Royce PLC

A duct for a gas turbine engine system is used to convey exhaust gas and bypass air away from the engine system, towards an exhaust nozzle. The duct includes an inlet face having at least two inlet portions and defines the outer extremity of a path for the gas and/or air through the duct. The duct includes a plurality of flat panel members which together define the outer extremity of the path.

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Номер записи: 67
01-01-2015 дата публикации

Rotational annular airscrew with integrated acoustic arrester

Номер: US20150000252A1
Автор: Kelly L. Boren, Matthew D. Moore
Принадлежит: Boeing Co

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

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Номер записи: 68
03-01-2019 дата публикации

MANUFACTURING ASSEMBLY AND METHOD

Номер: US20190001449A1
Автор: MASON John H.
Принадлежит: ROLLS-ROYCE PLC

The present disclosure relates to an assembly for formation of a fan blade. The assembly comprises a suction panel; a pressure panel; and a membrane having a leading edge and a trailing edge. The membrane is sandwiched between the suction panel and pressure panel. The membrane comprises a gas entry slot extending in a radial direction, the gas entry slot having a radially outer receiving portion for receiving a pipe, and a radially inner portion. The radially inner portion of the gas entry slot has a substantially uniform width in a direction between the leading and trailing edge of the membrane. 1. A membrane for inclusion in an assembly for formation of a fan blade , the membrane having a leading edge and a trailing edge ,wherein the membrane comprises a gas entry slot extending in a radial direction, the gas entry slot having a radially outer receiving portion for receiving a pipe, and a radially inner portion wherein the radially inner portion of the gas entry slot has a substantially uniform width in a direction between the leading and trailing edge of the membrane.2. A membrane according to wherein the width of the radially inner portion of the gas entry slot in the direction between the leading and trailing edges of the membrane is less than 3 mm.3. A membrane according to wherein the width of the radially inner portion of the gas entry slot in the direction between the leading and trailing edges of the membrane is between 1.1 and 0.8 mm.5. A membrane according to wherein the radially inner portion of the gas entry slot comprises a hook portion where the gas entry slot deflects through greater than 90 degrees.6. A membrane according to wherein the gas entry slot deflects through substantially 135 degrees.7. A membrane according to wherein the gas entry slot comprises a meander portion radially outwards of the hook portion.8. An assembly for formation of a fan blade claim 1 , the assembly comprising:a suction panel;a pressure panel; and{'claim-ref': {'@idref': ...

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Номер записи: 69
02-01-2020 дата публикации

AIRCRAFT PROPULSION SYSTEM WITH A LOW-FAN-PRESSURE-RATIO ENGINE IN A FORWARD OVER-WING-FLOW INSTALLATION, AND METHOD OF INSTALLING THE SAME

Номер: US20200002014A1
Автор: Dehaan Mark, Sclafani Tony J., Vassberg John C.
Принадлежит: The Boeing Company

There is provided a propulsion system for an aircraft, the system having a low-fan-pressure-ratio engine configured to be mounted, in a forward over-wing-flow installation, to a wing of the aircraft. The engine has a core, a variable pitch fan, and a nacelle having a nacelle trailing edge with a top-most portion positioned above a wing leading edge. The engine has an L/D ratio of the nacelle in a range of from 0.6 to 1.0, and a fan-pressure-ratio in a range of from 1.10 to 1.30. The forward over-wing-flow installation enables, during all flight phases of the aircraft, a fan flow exhaust to flow behind the nacelle, and to be bifurcated by the wing leading edge, so the fan flow exhaust flows both over the wing and under the wing. During a cruise flight phase of the aircraft, the engine minimizes scrubbing drag of the fan flow exhaust to the wing. 1. A propulsion system for an aircraft , comprising: a core having a first end and a second end;', 'a variable pitch fan coupled to the first end of the core;', 'a nacelle surrounding the variable pitch fan and a portion of the core, the nacelle having a nacelle leading edge and a nacelle trailing edge, the nacelle trailing edge having a top-most portion configured to be positioned above a wing leading edge of the wing, and the nacelle configured to be positioned, in its entirety, at a forward location in front of the wing leading edge;', 'a length to diameter (L/D) ratio of the nacelle in a range of from 0.6 to 1.0; and', 'a fan-pressure-ratio in a range of from 1.10 to 1.30,, 'a low-fan-pressure-ratio engine configured to be mounted, in a forward over-wing-flow installation, to a wing of the aircraft, the low-fan-pressure-ratio engine comprisingwherein the forward over-wing-flow installation of the low-fan-pressure-ratio engine of the propulsion system enables, during all flight phases of the aircraft, a fan flow exhaust, exhausted by the variable pitch fan, to flow behind the nacelle, and to be bifurcated by the wing ...

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Номер записи: 70
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: 71
05-01-2017 дата публикации

TIP SHROUDED HIGH ASPECT RATIO COMPRESSOR STAGE

Номер: US20170002659A1
Автор: Chandler Jesse M., Epstein Alan H., Suciu Gabriel L.
Принадлежит:

A gas turbine engine compressor stage includes a rotor. Compressor blades are supported by the rotor. The blades include an inner flow path surface each supporting an airfoil that has a chord that extends radially along a span to a tip. A shroud is supported at the tip and provides an outer flow path surface. The shroud provides a noncontiguous ring about the compressor stage. 1. A gas turbine engine compressor stage comprising:a rotor;compressor blades supported by the rotor, the blades include an inner flow path surface each supporting an airfoil that has a chord extending radially along a span to a tip, a shroud supported at the tip and providing an outer flow path surface, wherein the shroud provides a noncontiguous ring about the compressor stage.2. The compressor stage according to claim 1 , wherein the airfoils have an aspect ratio corresponding to the airfoil span to the airfoil chord claim 1 , the shroud extends the full chord claim 1 , wherein the aspect ratio is in a range of 1 to 5.3. The compressor stage according to claim 2 , wherein the aspect ratio is in a range of 1.4 to 3.0.4. The compressor stage according to claim 3 , wherein the aspect ratio is in a range of 1.6 to 2.8.5. The compressor stage according to claim 1 , wherein the shroud includes a sealing structure on a side of the shroud opposite of the outer flow path surface.6. The compressor stage according to claim 5 , wherein the sealing structure has labyrinth seals.7. The compressor stage according to claim 1 , wherein the blades each include a root received in a slot in the rotor.8. The compressor stage according to claim 1 , wherein shroud includes ring segments circumferentially spaced apart from one another and separated by gaps claim 1 , wherein multiple blades share a common ring segment.9. The compressor stage according to claim 1 , wherein blades are integrated with the rotor.10. A gas turbine engine comprising:engine static structure;first and second turbine sections rotatable ...

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Номер записи: 72
05-01-2017 дата публикации

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

Номер: US20170002685A1
Автор: Predrag Todorovic
Принадлежит: Rolls Royce Deutschland Ltd and Co KG

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

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Номер записи: 73
05-01-2017 дата публикации

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

Номер: US20170002688A1
Автор: Bruna Ramos, Bruno Albert Beutin, Maria Natal, Nuria Llamas Castro
Принадлежит: Safran Aircraft Engines SAS

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

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Номер записи: 74
05-01-2017 дата публикации

DEVICE FOR RETAINING DRAINED FLUIDS FOR A PROPULSIVE ASSEMBLY

Номер: US20170002689A1
Автор: Brun Gilles, LEON Alexandre
Принадлежит: SNECMA

A device for retaining drained fluids for a propulsive assembly includes a cavity for storing the drained fluids and two walls mounted at the opening of said cavity. The cavity has a fluid storage volume V when the device is in a substantially vertical position, and each wall is configured such as to define a fluid storage volume (V and V respectively) in the cavity when the device is in a substantially horizontal position, each of the volumes V and V being at least equal to the volume V 1132231. Device for retaining drained liquids from a propulsion assembly , comprising a body defining a cavity that is intended for storing configured to store the drained liquids and has a volume V when the device is in a first position , for example substantially vertical , said cavity comprising an upper opening through which the liquids are conveyed into the cavity , wherein the cavity comprises two walls in the region of said opening that are at least in part positioned one above the other and define a space therebetween , a first wall designed to define a volume V for storing liquids in the cavity when the device is in a second position that is at a positive angle from the first position about a substantially horizontal axis , and a second wall designed to define a volume V for storing liquids in the cavity when the device is in a third position that is at a negative angle from the first position about a substantially horizontal axis , each volume V and V being at least equal to the volume V.2. (canceled)3. Device according to claim 1 , wherein the two walls are an upper wall and a lower wall claim 1 , the upper wall defining an orifice for introducing the liquids into said space.4. Device according to claim 3 , wherein said orifice is offset on one side from a vertical median plane (P) of the cavity.5. Device according to either claim 3 , wherein the lower wall extends below the orifice in the upper plate and defines a passage for liquids from the space to the cavity.6. ...

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Номер записи: 75
05-01-2017 дата публикации

AUXILIARY OIL SYSTEM FOR GEARED GAS TURBINE ENGINE

Номер: US20170002738A1
Автор: Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

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

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Номер записи: 76
07-01-2016 дата публикации

CMC CORE COWL AND METHOD OF FABRICATING

Номер: US20160003094A1
Автор: KEY Caroline Elizabeth, LAUDE Anthony, RENGGLI Bernard James
Принадлежит:

A CMC core cowl for an aircraft gas turbine engine. The ceramic core cowl comprises an interlaced fiber structure having fibers oriented in substantially transverse directions, and a ceramic matrix surrounding the ceramic fiber structure. The core cowl further comprises several panels. The ceramic fiber and matrix are formed into a substantially cylindrical shape extending from a fore end at the fan outlet guide vanes to an aft end at the low pressure turbine outlet guide vanes. The CMC core cowl includes a means for mechanical attachment circumferentially oriented around the fore end and the aft end with mating parts. The CMC core cowl further includes additional plies oriented in a third preselected direction, thereby providing additional strength for mechanical attachment. 1. A ceramic matrix composite (CMC) core cowl for an aircraft gas turbine engine comprising: an interlaced fiber structure having ceramic fibers oriented in substantially transverse directions;', 'a ceramic matrix surrounding the ceramic fibers of the ceramic fiber structure;', 'wherein the ceramic fibers and matrix are formed into a substantially cylindrical shape having a fore end and an aft end, and having a mechanical attachment circumferentially oriented around the fore end and along the longitudinal lap joints; and', 'wherein the fore end and further includes additional CMC material having fibers oriented in a third preselected direction, thereby providing additional strength to for mechanical attachment at the fore end and at lap joints., 'a plurality of duct panels, each duct panel joined to an adjacent duct panel along a longitudinal lap joint, each duct panel further comprising;'}2. The CMC core cowl of further comprising:{'b': '160', 'a bifurcation opening () formed by a layup of CMC plies creating a duct boundary'}wherein the duct boundary forms a passageway in at least one of the duct panels.3. The CMC core cowl of wherein each of the plurality of duct panels is longitudinally ...

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Номер записи: 77
07-01-2016 дата публикации

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

Номер: US20160003146A1
Автор: DELAPIERRE Mickael, Gignoux Herve, Loro Gael, Wintenberger Sylvie
Принадлежит: SNECMA

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

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Номер записи: 78
07-01-2016 дата публикации

COOLING AIR SYSTEM FOR AIRCRAFT TURBINE ENGINE

Номер: US20160003153A1
Автор: Rhoden William E.
Принадлежит:

In one aspect, a turbine engine is provided. The engine includes a fan, a core engine case configured to receive a core airflow, and an inner fan case. An inner air bypass is defined between the core engine case and the inner fan case, and the engine further includes an outer fan case, an outer air bypass defined between the inner fan case and the outer fan case, a first heat exchanger arranged in the core engine case, the first heat exchanger providing heat exchange between the liquid and the core airflow to cool the core airflow, and a second heat exchanger arranged in the outer air bypass. The second heat exchanger is fluidly coupled to the first heat exchanger to receive the cooled core airflow. The second heat exchanger provides heat exchange between cooled core airflow and the second portion of airflow to further cool the cooled core airflow. 1. A turbine engine comprising:a fan;a core engine case configured to receive a core airflow;an inner fan case, wherein an inner air bypass is defined between the core engine case and the inner fan case, the inner air bypass configured to receive a first portion of an airflow from the fan;an outer fan case, wherein an outer air bypass is defined between the inner fan case and the outer fan case, the outer air bypass configured to receive a second portion of the airflow from the fan;a first heat exchanger arranged in the core engine case and configured to receive a liquid and the core airflow, the first heat exchanger providing heat exchange between the liquid and the core airflow to cool the core airflow; anda second heat exchanger arranged in the outer air bypass, the second heat exchanger fluidly coupled to the first heat exchanger to receive the cooled core airflow, the second heat exchanger providing heat exchange between the cooled core airflow and the second portion of the airflow to further cool the cooled core airflow.2. The turbine engine of claim 1 , further comprising a third heat exchanger arranged in the ...

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Номер записи: 79
07-01-2016 дата публикации

GAS TURBINE ENGINE INCLUDING A THIRD FLOWPATH EXHAUST NOZZLE

Номер: US20160003187A1
Автор: Roberge Gary D.
Принадлежит:

A turbofan engine that includes a first flowpath, a second flowpath, a third flowpath, and a third flowpath exhaust nozzle is provided. The first flowpath is radially inboard of the second flowpath at a location upstream of a core section of the turbofan engine. The third flowpath is radially outboard of the second flowpath at the location upstream of the core section. The third flowpath exhaust nozzle defines a plurality of third flowpath exhaust exit ports through which gas traveling along the third flowpath may be discharged. An area or a geometry of each of the plurality of third flowpath exhaust exit ports is independently and selectively adjustable. A method for operating the turbofan engine includes independently and selectively adjusting an area or a geometry of at least one of the plurality of third flowpath exhaust exit ports to achieve a desired engine operation. 1. A turbofan engine , comprising:a first flowpath that is radially inboard of a second flowpath at a location upstream of a core section of the turbofan engine;a third flowpath that is radially outboard of the second flowpath at the location upstream of the core section of the turbofan engine;a third flowpath exhaust nozzle that defines a plurality of third flowpath exhaust exit ports through which gas traveling along the third flowpath may be discharged; andwherein the third flowpath exhaust nozzle is configured so that an area or a geometry of each of the plurality of third flowpath exhaust exit ports is independently and selectively adjustable.2. The turbofan engine of claim 1 , wherein by independently and selectively adjusting an area or a geometry of one of the plurality of third flowpath exhaust exit ports claim 1 , a characteristic of gas passing there through may be independently and selectively adjusted.3. The turbofan engine of claim 1 , further comprising a primary exhaust nozzle that defines a primary exhaust exit port through which gas traveling along the first flowpath and the ...

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Номер записи: 80
07-01-2016 дата публикации

Asymmetric Fan Nozzle in High-BPR Separate-Flow Nacelle

Номер: US20160003194A1
Автор: Lord Wesley K., Malecki Robert E.
Принадлежит:

A fan nozzle for an aircraft gas turbine engine is comprised of a core engine cowl that is disposed within a fan cowl so that an air flow area is defined therebetween. The core engine cowl and fan cowl are disposed around a horizontal central plane. The fan cowl has a substantially circular shape and is formed of an upper substantially semi-circular portion having a first radius and a lower substantially semi-circular portion having a second radius. The core engine cowl has a substantially circular shape and is formed of an upper substantially semi-circular portion having a third radius and a lower substantially semi-circular portion having a third radius. The upper substantially semi-circular portion of the core engine cowl includes a left arcuate member and a right arcuate member. The second radius is less than the first radius and the third radius is less than the fourth radius. 1. A fan nozzle for a gas turbine engine , comprising:a fan cowl having a substantially circular shape, the fan cowl being formed of an upper substantially semi-circular portion and a lower substantially semi-circular portion, the upper substantially semi-circular portion of the fan cowl having a first radius, the lower substantially semi-circular portion of the fan cowl having a second radius, the second radius being less than the first radius;a core engine cowl disposed within the fan cowl, the fan cowl and the core engine cowl positioned around a horizontal central plane, the core engine cowl having a substantially circular shape, the core engine cowl being formed of an upper substantially semi-circular portion and a lower substantially semi-circular portion, the upper substantially semi-circular portion of the core engine cowl having a third radius, the upper substantially semi-circular portion of the core engine cowl being formed of a left arcuate member and a right arcuate member, the lower substantially semi-circular portion of the core engine cowl having a fourth radius, the third ...

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Номер записи: 81
05-01-2017 дата публикации

ANNULAR COMBUSTION CHAMBER IN A TURBINE ENGINE

Номер: US20170003028A1
Автор: Cameriano Laurent Bernard, LE PANNERER Brice, PIREYER Pierre-Francois Simon Paul
Принадлежит: SNECMA

The invention relates to a device for supporting and centring a fuel injector in a turbine engine combustion chamber, which includes means for centring a fuel injector along an axis, which are movable in a plane that is radial to the centring axis () in supporting means intended for being attached to the bottom of an annular chamber (). According to the invention, the centring means include at least two radially external tabs () each inserted respectively in a circumferential recess () of the supporting means, the device including circumferential abutment means () of the radial tabs () of the centring means in the circumferential recesses (), the circumferential abutment means being configured such as to enable a greater angular displacement of a first () one of the radial tabs in a first circumferential recess () relative to a second () one of the radial tabs in a second circumferential recess (). 1. Device for supporting and centring a fuel injector in a turbine engine combustion chamber , which includes means for centring a fuel injector along an axis , which are movable in a plane that is radial to the centring axis in supporting means intended for being attached to the bottom of an annular chamber , wherein the means of centring include at least two radially external tabs , each inserted respectively in a circumferential recess of the supporting means , wherein the device includes circumferential abutment means of the radial tabs of the means of centring in the circumferential recesses , wherein the circumferential abutment means are configured such as to enable a greater angular displacement of a first one of the radial tabs in a first circumferential recess relative to a second one of the radial tabs in a second circumferential recess.2. Device according to claim 1 , wherein the supporting means include an annular edge extending along the centring axis claim 1 , in which the first and second circumferential recesses accommodating the first and second radial ...

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Номер записи: 82
04-01-2018 дата публикации

HIGH EFFICIENCY AIRCRAFT PARALLEL HYBRID GAS TURBINE ELECTRIC PROPULSION SYSTEM

Номер: US20180003071A1
Автор: Hanlon Christopher J., Hardin Larry W., Lents Charles E., Rheaume Jonathan
Принадлежит:

A gas turbine engine includes a compressor section having a first compressor and a second compressor, the second compressor having a higher pressure than the first compressor, and a turbine section having a first turbine and a second turbine, the second turbine having a higher pressure than the first turbine. The first compressor is connected to the first turbine via a first shaft. The second compressor is connected to the second turbine via a second shaft. An electric motor is connected to the first shaft such that rotational energy generated by the electric motor is translated to the first shaft. A fan is connected to the first shaft via a gear system. The gas turbine engine includes at least a takeoff mode of operations, a top of climb mode of operations and a maximum cruise mode of operations. The gas turbine engine is sized to operate at peak efficiency in the maximum cruise mode of operations. 1. A gas turbine engine comprising:a compressor section having a first compressor and a second compressor, the second compressor having a higher pressure than the first compressor;a turbine section having a first turbine and a second turbine, the second turbine having a higher pressure than the first turbine,the first compressor is connected to the first turbine via a first shaft;the second compressor is connected to the second turbine via a second shaft;an electric motor connected to the first shaft such that rotational energy generated by the electric motor is translated to the first shaft;a fan connected to the first shaft via a gear system; andwherein the gas turbine engine includes at least a takeoff mode of operations, a top of climb mode of operations and a maximum cruise mode of operations, the gas turbine engine being sized to operate at peak efficiency in said maximum cruise mode of operations.2. The gas turbine engine of claim 1 , wherein operating said engine at peak efficiency comprises operating said engine at a maximum turbine inlet temperature of said ...

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Номер записи: 83
04-01-2018 дата публикации

GEARED TURBOFAN GAS TURBINE ENGINE ARCHITECTURE

Номер: US20180003112A1
Автор: Kupratis Daniel Bernard, Schwarz Frederick M.
Принадлежит:

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

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Номер записи: 84
02-01-2020 дата публикации

Planetary gear drive and aircraft gas turbine with a planetary gear drive

Номер: US20200003078A1
Автор: Hannes WUESTENBERG
Принадлежит: Rolls Royce Deutschland Ltd and Co KG

The invention relates to a planetary gearbox having a ring gear support, in particular in a geared fan engine, characterized by a bolt connection for connecting the ring gear support to a static component, wherein the bolt connections are arranged on the circumference of the ring gear support in the axial direction of the planetary gearbox, and the bolt connections are designed and arranged in such a way that there is a frictional joint with a defined frictional force between the ring gear support and the static component and that a material separation, in particular a gap, is arranged between the ring gear support and the static component.

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Номер записи: 85
04-01-2018 дата публикации

Carrier structure for an epicyclic gear drive, epicyclic gear drive and turbo engine with an epicyclic gear drive

Номер: US20180003289A1
Автор: Elisa ZAMBON, Luca MANTOVAN
Принадлежит: Rolls Royce Deutschland Ltd and Co KG

A carrier structure for an epicyclic gear drive is provided. The carrier structure includes carrier elements connected with at least one planet gear and the first carrier element is connected with the second carrier element through at least two struts with the at least two struts having an inclination angle of more than 20° in the direction of a rotation around a rotation axis of the carrier structure.

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Номер записи: 86
02-01-2020 дата публикации

Gas turbine engine

Номер: US20200003122A1
Автор: Johnathan H. WILSHAW
Принадлежит: Rolls Royce PLC

A gas turbine engine for an aircraft includes: a fan adjacent the engine air intake, including a plurality of fan blades; downstream of the fan, an engine core including a turbine, a compressor, and a core shaft connecting the turbine and compressor; an engine core housing at least partly encasing the core; a fan case surrounding the fan and defining at least part of a bypass duct radially outside the core; a plurality of outlet guide vanes extending between the engine core housing and an outlet guide vane support region of the case, adjacent an upstream end of the bypass duct; one or more supports extending from the case to the engine core housing, wherein: a first end of the supports fixes to the case at the outlet guide vane support region; a second end of the supports fixes to the engine core housing adjacent an engine core exhaust.

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Номер записи: 87
02-01-2020 дата публикации

Gas turbine

Номер: US20200003127A1
Автор: Mark N. BINNINGTON
Принадлежит: Rolls Royce PLC

A gas turbine engine, in particular an aircraft engine, includes: a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear, the sun gear is connected to the input shaft device, the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, with a rear carrier bearing device radially between the planet carrier and a part on the input shaft on the input side of the gearbox device.

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Номер записи: 88
02-01-2020 дата публикации

GAS TURBINE

Номер: US20200003128A1
Автор: HUGHES Geoffrey, MAGUIRE Alan R.
Принадлежит: ROLLS-ROYCE PLC

A gas turbine engine, in particular an aircraft engine, includes: a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear, the sun gear is connected to the input shaft device, the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, with a rear carrier bearing device radially between the planet carrier and a static structure on the input side of the gearbox device or a front carrier bearing device radially between the planet carrier and a static structure on the output side of the gearbox device. 1. A gas turbine engine , comprising:a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear,the sun gear is connected to the input shaft device,the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, witha rear carrier bearing device radially between the planet carrier and a static structure on the input side of the gearbox device ora front carrier bearing device radially between the planet carrier and a static structure on the output side of the gearbox device.2. The gas turbine of claim 1 , wherein the rear carrier bearing device or the front carrier bearing device comprise at least one roller bearing.3. The gas turbine of claim 1 , wherein the front carrier bearing device axially adjacent to the gearbox device on the output side.4. The gas turbine of claim 1 , wherein a fan shaft bearing system is radially located between a fan shaft as part of the output shaft device and a static structure.5. The gas turbine of claim 4 , wherein the fan shaft bearing system has an outer diameter between 0.05 to 0.2 the diameter of the propulsive fan claim 4 , in particular between 0.1 and 0.15 times the diameter of the propulsive fan.6. ...

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Номер записи: 89
02-01-2020 дата публикации

GAS TURBINE

Номер: US20200003129A1
Автор: BONIFACE Dominic, HUGHES Geoff, KNIGHT Glenn, SIMMS Paul, SWIFT Andrew, THORNTON Stewart
Принадлежит:

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

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Номер записи: 90
02-01-2020 дата публикации

Counter Rotating Turbine with Reversing Reduction Gearbox

Номер: US20200003157A1
Автор: Alan Roy Stuart, Darek Tomasz Zatorski, Jeffrey Donald Clements
Принадлежит: General Electric Co

The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction, and a circumferential direction, and an upstream end and a downstream end along the longitudinal direction. The gas turbine engine includes a turbine section, a gearbox proximate to the turbine section, and a driveshaft. The turbine section includes a first rotating component interdigitated with a second rotating component along the longitudinal direction. The first rotating component includes an outer shroud defining a plurality of outer shroud airfoils extended inward of the outer shroud along the radial direction and one or more connecting airfoils coupling the outer shroud to a radially extended rotor. The second rotating component includes an inner shroud defining a plurality of inner shroud airfoils extended outward of the inner shroud along the radial direction. The second rotating component is coupled to an input shaft connected to an input gear of the gearbox. The driveshaft is extended in the longitudinal direction and is connected to an output gear of the gearbox. The first rotating component is coupled to the driveshaft.

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Номер записи: 91
07-01-2021 дата публикации

GAS TURBINE COMPONENT WITH COOLING APERTURE HAVING SHAPED INLET AND METHOD OF FORMING THE SAME

Номер: US20210003019A1
Автор: Howe Jeff, Tapia Luis
Принадлежит: HONEYWELL INTERNATIONAL INC.

A method of manufacturing a cooled gas turbine component includes forming a core with an outer surface. The outer surface includes a core feature. The method also includes casting an outer wall of an airfoil about the core. The outer wall has an exterior surface and an interior surface. The interior surface includes a shaped inlet portion that corresponds to the core feature. Moreover, the method includes forming an outlet portion through the outer wall to fluidly connect the outlet portion to the shaped inlet portion. The shaped inlet portion and the outlet portion cooperatively define a cooling aperture through the outer wall. 1. A method of manufacturing a cooled gas component for a turbomachine , the method comprising:forming a core with an outer surface, the outer surface including a core feature that tapers to reduce in width as the core feature projects from the core;casting an outer wall about the core including casting a portion of the outer wall that covers the core feature, the portion including an exterior surface and an interior surface, the interior surface defining a shaped inlet portion that is cast to inversely correspond to the core feature; andafter casting the outer wall, forming an outlet portion through the outer wall by progressively removing material through the portion of the outer wall in a direction from the exterior surface toward the interior surface and toward the shaped inlet portion to fluidly connect the outlet portion to the shaped inlet portion, the shaped inlet portion and the outlet portion cooperatively defining a cooling aperture through the outer wall.2. The method of claim 1 , further comprising adjusting the outlet portion by changing a dimension of the outlet portion.3. The method of claim 1 , further comprising:re-casting the outer wall; andforming an adjusted outlet portion through the re-cast outer wall.4. The method of claim 3 , wherein forming the outlet portion through the outer wall includes forming the outlet ...

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Номер записи: 92
07-01-2021 дата публикации

GAS TURBINE ENGINE COMPOSITE DUCT WITH BRACKET

Номер: US20210003037A1
Автор: Banhos Jonas S., Prainito Anthony
Принадлежит:

A bracket assembly for a gas turbine engine includes a first foot that includes at least one flange fastener opening and a duct flange support surface. A second foot includes at least one body portion fastener opening and a duct body support surface. At least one leg connects the first foot relative to the second foot.

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Номер записи: 93
07-01-2021 дата публикации

DUCT ASSEMBLY FOR A GAS TURBINE ENGINE

Номер: US20210003038A1
Автор: Banhos Jonas S., Prainito Anthony
Принадлежит:

A gas turbine engine duct assembly includes a composite duct body that includes at least one rim. A flanged bracket includes a radially extending portion that defines a flange. An axially extending portion has an inner duct rim recess for accepting a portion of the at least one rim. At least one attachment plate has a projection recess for accepting a projection on the flanged bracket and an outer rim recess for accepting a portion of the at least one rim. 1. A gas turbine engine duct assembly comprising:a composite duct body including at least one rim;a flanged bracket includes a radially extending portion defining a flange and an axially extending portion having an inner duct rim recess for accepting a portion of the at least one rim; andat least one attachment plate having a projection recess for accepting a projection on the flanged bracket and an outer rim recess for accepting a portion of the at least one rim.2. The assembly of claim 1 , wherein the projection includes a first axially facing projection surface extending in a first plane perpendicular to a central longitudinal axis of the flanged bracket and a second axially facing projection surface extending in a second plane perpendicular to the central longitudinal axis of the flanged bracket.3. The assembly of claim 2 , wherein the projection recess includes a first axially facing recess surface in abutting contact with the first axially facing projection surface and a second axially facing recess surface in abutting contact with the second axially facing projection surface.4. The assembly of claim 2 , wherein the projection includes a radially outward facing projection surface having a constant radial dimension between the first axially facing projection surface and the second axially facing projection surface.5. The assembly of claim 1 , wherein an inner surface of the composite duct body at least partially defines a bypass flow path.6. The assembly of claim 1 , wherein the at least one rim includes a ...

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Номер записи: 94
02-01-2020 дата публикации

COMBUSTOR SHELL ATTACHMENT

Номер: US20200003417A1
Автор: Schlichting Kevin W.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A combustor shell is provided. The combustor shell may include a first aperture at least partially defined by an inner wall of the combustor shell and passing from a diffuser-facing side of the combustor shell to a combustor-facing side of the combustor shell. The combustor shell may include a spacer comprising a first segment coupled to a first flange, wherein the first flange is disposed on the diffuser-facing side of the combustor shell, wherein an outer wall of the spacer is coupled with at least a portion of an inner wall of the combustor shell. 1. A combustor shell comprising:a first aperture at least partially defined by an inner wall of the combustor shell and passing from a diffuser-facing side of the combustor shell to a combustor-facing side of the combustor shell; anda spacer comprising a first segment coupled to a first flange, wherein the first flange is disposed on the diffuser-facing side of the combustor shell, wherein an outer wall of the spacer is coupled with at least a portion of an inner wall of the combustor shell.2. The combustor shell of claim 1 , wherein the first aperture of the combustor shell is an oblong shape.3. The combustor shell of claim 1 , wherein the spacer comprises threads on an inner wall of the spacer.4. The combustor shell of claim 1 , wherein the spacer is press fit into the combustor shell.5. The combustor shell of claim 1 , wherein the spacer is configured to engage an attachment feature of a combustor panel.6. The combustor shell of claim 5 , wherein a contact length between the spacer and the attachment feature is greater than a distance between the diffuser facing side of the combustor shell and the combustor facing side of the combustor shell.7. The combustor shell of claim 1 , wherein the spacer further comprises a second flange coupled with the first segment and disposed on the combustor-facing side of the combustor shell.8. The combustor shell of claim 1 , wherein the spacer further comprises a spacer aperture at ...

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Номер записи: 95
07-01-2021 дата публикации

OUTER SHROUD OF AN INTERMEDIATE CASING FOR A DUAL FLOW TURBINE ENGINE FOR AN AIRCRAFT, COMPRISING IMPROVED AIR-SEALING AND FIRE-RESISTANCE DEVICES

Номер: US20210003098A1
Автор: Achbari Baghdad, BROAGE Thomas Claude, GODIN Florent Robert André, JACON Bruno Alexandre Didier, ROUDIL Joël François, SIMONOTTI Hervé
Принадлежит:

An outer shroud of an intermediate casing for a dual flow turbine engine for an aircraft, the shroud including: an annular downstream portion provided with a shroud opening passing through an annular downstream edge of the shroud; a connecting member (50) attached to the annular downstream portion, and intended to attach an arm that passes through a secondary flow path; an air-sealing and fire-resistance device including: a portion including: a pad arranged in a hollow annular area of the annular downstream edge of the shroud; a blade protruding from the pad and clamped between a circumferential end and the radially outer end of the arm, a leaf spring (pressing the pad into the hollow annular area. 11614. Outer shroud () of an intermediate casing () for a dual flow turbine engine for an aircraft , the shroud comprising:{'b': 24', '30', '32', '36', '38', '32', '30', '40', '42', '44', '22', '26, 'a downstream annular part () provided with a shroud opening () passing radially through the shroud and opening up axially downstream from the shroud, the downstream part comprising a downstream annular shroud edge (), of which a radially internal surface () delimits a hollow annular zone () open radially inwards, the downstream annular edge () being interrupted by the shroud opening () so as to have two circumferential end surfaces () facing each other and delimiting a housing space () between them for a radially outer end () of an arm () that will pass radially through a fan flow () of the turbine engine;'}{'b': 50', '50', '24', '50', '22', '54', '40', '32', '56, 'a connecting element (), formed by a perforated connection plate (), fixed on the downstream annular part () of the shroud and arranged radially inwards from the shroud, the element () also being designed to be fixed to said arm () and forming two circumferential ends () fixed to the two circumferential end surfaces () of the downstream annular shroud edge (), so as to define two junction zones () respectively; and ...

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Номер записи: 96
03-01-2019 дата публикации

FLUID COOLING SYSTEMS FOR A GAS TURBINE ENGINE

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

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

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Номер записи: 97
03-01-2019 дата публикации

TURBOMACHINE COMPRISING A SURFACE AIR-OIL HEAT EXCHANGER BUILT INTO AN INTER-FLOW COMPARTMENT

Номер: US20190003390A1
Автор: ANCUTA Carmen Gina, RAMOS Bruna Manuela
Принадлежит: SAFRAN AIRCRAFT ENGINES

A turbomachine comprises an inter-flow compartment extending radially between first and second intermediate walls, gas bypass ducts passing through the second intermediate wall and communicating with evacuation outlets opening up into a secondary gas flow; an oil reservoir arranged in the inter-flow compartment, and a surface air-oil heat exchanger communicating with the oil reservoir for fluid circulation. Circumferentially in relation to the inter-flow compartment, the surface air-oil heat exchanger extends at least partially between the evacuation outlets which pass through the first intermediate wall. 1. A turbomachine comprising:a primary gas flow, and a secondary gas flow located around the primary gas flow,a first intermediate wall to radially inwardly delimit the secondary gas flow,a second intermediate wall to radially outwardly delimit the primary gas flow,an inter-flow compartment extending radially between the first and second intermediate walls, air bypass ducts of the primary gas flow:passing through the second intermediate wall and into the inter-flow compartment, andopening up into the secondary gas flow, through the first intermediate wall, via several evacuation outlets,an oil reservoir at least partially arranged in the inter-flow compartment, anda surface air-oil heat exchanger that has fluid communication with the oil reservoir and which has a heat exchange surface facing the secondary gas flow,wherein, circumferentially in relation to the inter-flow compartment, the surface air-oil heat exchanger extends at least partially between the evacuation outlets which pass through the first intermediate wall.2. The turbomachine of claim 1 , wherein claim 1 , circumferentially claim 1 , between the evacuation outlets claim 1 , the surface air-oil heat exchanger comprises several portions to achieve a heat exchange with the gas of the secondary gas flow claim 1 , said several portions being connected one to another claim 1 , at least two by two claim 1 , ...

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Номер записи: 98
03-01-2019 дата публикации

TURBOJET ENGINE WITH THRUST TAKE-UP MEANS ON THE INTER-COMPRESSOR CASE

Номер: US20190003395A1
Автор: Bordoni Nils Edouard Romain, Brault Michel Gilbert Roland, Cuvillier Romain Guillaume, Kubiak Guillaume Patrice, Negri Arnaud Nicolas, Nowakowski Nathalie
Принадлежит: SAFRAN AIRCRAFT ENGINES

A multiflow turbojet engine generally includes an upstream fan driven by a gas generator having first and second coaxial compressors, an intake case forming a mounting for the rotors of the upstream fan and the first compressor, an inter-compressor case downstream from the intake case and forming a mounting for the rotors of the second compressor, and attachment means for thrust take-up control rods arranged in the inter-compressor case. The turbojet engine also includes a structural force shroud connecting the intake case to the inter-compressor case, of the and a floating first compressor case. 1. A turbojet engine including:an upstream ducted fan driven by a gas generator, whereby the gas generator comprises a first compressor and a second compressor that is coaxial with the first compressor;an inlet case configured to form a support for a plurality of rotors of the upstream ducted fan and the first compressor;an inter-compressor case located downstream from the inlet case, and configured to form a support for a plurality of rotors of the second compressor;attachment means for a plurality of thrust take-up rods arranged on the inter-compressor case; anda stress structural shroud configured to connect the inlet case to the inter-compressor case,wherein the first compressor comprises a floating case that forms a wall of a flow path.2. The turbojet engine according to claim 1 , wherein the floating case that forms the wall of the flow path is connected in a floating configuration to one of the inlet case and the inter-compressor case by a backlash connection.3. The turbojet engine according to claim 1 , wherein the stress structural shroud is welded to the inlet case and bolted to the inter-compressor case.4. The turbojet engine according to claim 1 , wherein the stress structural shroud is bolted on the inlet case and bolted on the inter-compressor case.5. Turbojet The turbojet engine according to claim 1 , wherein the inlet case comprises a shroud that supports a ...

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Номер записи: 99
03-01-2019 дата публикации

THRUST REVERSER SYSTEM EXHIBITING LIMITED AERODYNAMIC PERTURBATION

Номер: US20190003420A1
Автор: Pautis Olivier
Принадлежит: AIRBUS OPERATIONS S.A.S.

A thrust reverser system includes a thrust reversal cascade and a reverser door which are accommodated in a housing situated outside a secondary flow duct. The system also includes connection parts the presence of which combined with the action of the actuator gives rise simultaneously to: displacement rearwards of the cascade in the direction of a nacelle opening, cleared by a nacelle cowling entrained rearwards together with the cascade; and a combined movement of the reverser door relative to the cascade, leading to displacement of the front end of the door rearwards along the cascade, and to pivoting of this door, such as to give rise to plunging of its rear end into the secondary flow duct. 1. A thrust reverser system for a turbine engine of a double-spool aircraft , the reverser system comprising:at least one thrust reversal cascade through which the air of a secondary flow duct of the turbine engine is configured to circulate in the active configuration of the reverser system;at least one reverser door configured to close the secondary flow duct at least partly when the reverser system is in the active configuration;at least one actuator,wherein, in an inactive configuration of the reverser system, the at least one reverser door is accommodated in a housing situated outside the secondary flow duct, and in which housing there is also located the at least one thrust reverser cascade entrained by the at least one actuator, and a rear end of which cascade is integral with a nacelle outer mobile cowling, the reverser system also comprising:a first connection part, a first end of which is connected inside the nacelle outer mobile cowling by a first articulated connection, and a second end of which is connected to the at least one reverser door by a second articulated connection;a second connection part, a first end of which is connected to a rear end of the at least one reverser door by a third articulated connection, and a second end of which is connected to an ...

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Номер записи: 100