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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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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
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: 7
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: 8
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: 9
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: 10
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: 11
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: 12
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: 13
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: 14
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: 15
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: 16
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: 17
04-04-2013 дата публикации

COUNTER-ROTATABLE FAN GAS TURBINE ENGINE WITH AXIAL FLOW POSITIVE DISPLACEMENT WORM GAS GENERATOR

Номер: US20130081374A1
Автор: Fakunle Oladapo, Giffin Rollin George, Murrow Kurt David
Принадлежит:

A counter-rotatable fan turbine engine includes a counter-rotatable fan section, a worm gas generator, and a low pressure turbine to power the counter-rotatable fan section. The low pressure turbine maybe counter-rotatable or have a single direction of rotation in which case it powers the counter-rotatable fan section through a gearbox. The gas generator has inner and outer bodies having offset inner and outer axes extending through first, second, and third sections of a core assembly. At least one of the bodies is rotatable about its axis. The inner and outer bodies have intermeshed inner and outer helical blades wound about the inner and outer axes and extending radially outwardly and inwardly respectively. The helical blades have first, second, and third twist slopes in the first, second, and third sections respectively. A combustor section extends through at least a portion of the second section. 1. A counter-rotatable fan gas turbine engine comprising in downstream serial flow relationship a counter-rotatable fan section , a worm gas generator , and a counter-rotatable low pressure turbine operably connected to the counter-rotatable fan section.2. The engine as claimed in further comprising:the gas generator including an inlet axially spaced apart and upstream from an outlet,a core assembly including an inner body disposed within an outer body and the inner and outer bodies extending from the inlet to the outlet,the inner and outer bodies having offset inner and outer axes respectively,at least one of the inner and outer bodies being rotatable about a corresponding one of the inner and outer axes,the inner and outer bodies having intermeshed inner and outer helical blades wound about the inner and outer axes respectively,the inner and outer helical blades extending radially outwardly and inwardly respectively,the core assembly having first, second, and third sections in serial downstream flow relationship extending between the inlet and the outlet,the inner and ...

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Номер записи: 18
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: 19
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: 20
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: 21
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: 22
23-05-2013 дата публикации

EMISSIONS CONTROL SYSTEMS AND METHODS

Номер: US20130125524A1
Автор: Detar Gail Leroy, Heverley, III Lawrence, Kiran Shashi, Kulkarni Shridhar Shrikant, Plummer Jessica Lynn
Принадлежит:

Methods and systems are provided related to an emissions control system. The emissions control system has an exhaust after-treatment system defining a plurality of distinct exhaust flow passages through which at least a portion of an exhaust stream can flow, e.g., the exhaust stream is produced by an engine. The emissions control system also includes a controller for controlling injection of reductant into the exhaust stream flowing through each of the flow passages. In one example, the emissions control system is configured for use in a vehicle, such as a locomotive or other rail vehicle. 1. An emissions control system , comprising:an exhaust after-treatment system defining a plurality of distinct exhaust flow passages through which an exhaust stream can flow, the exhaust stream produced by an engine,wherein the exhaust after-treatment system includes a plurality of first exhaust after-treatment components, each in or otherwise associated with a respective one of the exhaust flow passages.2. The emissions control system of claim 1 , wherein the plurality of exhaust flow passages are positioned at least generally parallel to each other.3. The emissions control system of claim 1 , wherein each of the plurality of exhaust flow passages is defined by a respective substrate claim 1 , or a set of substrates claim 1 , though which the exhaust stream can flow.4. The emissions control system of claim 3 , wherein the substrate claim 3 , or each of the set of substrates claim 3 , is rectangular shaped or cylindrically shaped.5. The emissions control system of claim 3 , wherein at least one of the exhaust flow passages is defined at least in part by a respective set of parallel-arrayed distinct substrates that are arranged with at least one of the parallel-arrayed distinct substrates located on an upper level above at least one other of the parallel-arrayed distinct substrates located on a lower level.6. The emissions control system of claim 5 , wherein a first of the exhaust ...

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Номер записи: 23
06-06-2013 дата публикации

METHOD FOR OPERATING AN EXHAUST GAS SYSTEM OF AN INTERNAL COMBUSTION ENGINE

Номер: US20130144505A1
Автор: Handler Torsten, Neumann Joerg, Pfister Tobias, Wendt Hauke
Принадлежит: ROBERT BOSCH GMBH

A method for operating an exhaust gas system of an internal combustion engine is described, wherein NOx is reduced by means of a SCR catalyst and a NOx reduction capability of an aqueous urea solution to be introduced into the exhaust gas system is monitored, and wherein at least one first variable characterizing the ammonia content of the water is ascertained and an ageing of the aqueous urea solution is inferred from said first variable. 1. A method for operating an exhaust gas system of an internal combustion engine , wherein NOx is reduced by means of a SCR catalyst and a NOx reduction capability of an aqueous urea solution to be introduced into the exhaust gas system is monitored , the method comprising:ascertaining at least one first variable characterizing the ammonia content of the water; andmonitoring the NOx reduction capability of the aqueous urea solution.2. The method according to claim 1 , wherein at least one second variable characterizing the composition of the aqueous urea solution is ascertained and the NOx reduction capability of said aqueous urea solution is inferred from the first and the second variable.3. The method according to claim 1 , wherein the first variable is an electrical conductivity of the aqueous urea solution.4. The method according to claim 1 , wherein the second variable is a density and/or a refractive index and/or a sound velocity and/or a thermal conductivity and/or a dielectric permittivity of the aqueous urea solution.5. The method according to claim 1 , wherein the ascertained first and second variables are linked to one another by means of at least one characteristic diagram and/or a table and/or a mathematical formula in order to infer the NOx reduction capability of the aqueous urea solution.6. The method according to claim 1 , wherein a filling level of a reservoir of the reducing agent and/or a point in time of a filling of the reservoir and/or a temperature of the reducing agent are used in a complementary manner ...

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Номер записи: 24
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: 25
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: 26
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: 27
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: 28
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: 29
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: 30
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: 31
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: 32
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: 33
03-10-2013 дата публикации

Rankine Cycle System

Номер: US20130255258A1
Автор: Loveday Ronald Lee
Принадлежит:

A method and apparatus for implementing the Rankine Cycle with improved efficiency through the use of thermal exchange between the working fluid and a brine liquid. In one embodiment, the brine liquid is further processed through a multistage flash distillation unit to produce distilled water or other minerals. Geothermal or exhaust waste heat may be used to further improve the Rankine Cycle efficiency, with remaining thermal energy improving the distillation process. 1. A Rankine Cycle system comprising:a first heat exchanger configured to facilitate thermal energy transfer from a heated fluid to a working fluid,a turbine coupled to a generator, said turbine is configured to receive the working fluid from the first heat exchanger;a second heat exchanger configured to facilitate thermal energy transfer from the working fluid flowing from an output of the turbine to a brine liquid; anda multistage flash distillation unit configured to produce distilled water from the brine liquid flowing from an output of the second heat exchanger.2. The Rankine Cycle system of wherein at least a portion of the heated fluid comes from a geothermal source.3. The Rankine Cycle system of wherein at least a portion of the heated fluid from an output of the first heat exchanger provides thermal energy to the multistage flash distillation unit.4. The Rankine Cycle system of wherein the multistage flash distillation unit further comprises cooling fingers sourced at least in part from the brine fluid.5. The Rankine Cycle system of wherein at least a portion of the heated fluid from an output of the first heat exchanger is routed to an injection well.6. The Rankine Cycle system of wherein at least a portion of the brine liquid from an output of the multistage flash distillation unit is routed to an injection well.7. The Rankine Cycle system of wherein at least a portion of the brine liquid from an output of the multistage flash distillation unit is processed to retrieve at least one mineral.8 ...

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Номер записи: 34
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: 35
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: 36
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: 37
24-10-2013 дата публикации

Low Noise Compressor Rotor for Geared Turbofan Engine

Номер: US20130276424A1
Автор: Morin Burce L., Topol David A.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A gas turbine engine has a fan, a compressor section having a low pressure portion and a high pressure portion, a combustor section, and a turbine having a low pressure portion. The low pressure turbine portion drives the low pressure compressor portion and the fan. A gear reduction effects a reduction in the speed of the fan relative to a speed of the low pressure turbine and the low pressure compressor portion. At least one of the low pressure turbine portion and low pressure compressor portion has a number of blades in each of a plurality of rows. The blades operate at least some of the time at a rotational speed. The number of blades and the rotational speed are such that the following formula holds true for at least one of the blade rows of the at least one of the low pressure turbine portion and/or the low pressure compressor sections: (number of blades×rotational speed)/60≧5500. The rotational speed is an approach speed in revolutions per minute. 1. A gas turbine engine comprising:a fan, a compressor section having a low pressure portion and a high pressure portion, a combustor section, and a turbine having a low pressure portion, the low pressure turbine portion driving said low pressure compressor portion and the fan;a gear reduction effecting a reduction in the speed of said fan relative to a speed of the low pressure turbine and the low pressure compressor portion; {'br': None, '(number of blades×rotational speed)/60≧5500; and'}, 'at least one of said low pressure turbine portion and said low pressure compressor portion having a number of blades in each of a plurality of rows, and said blades operating at least some of the time at a rotational speed, and said number of blades and said rotational speed being such that the following formula holds true for at least one of the blade rows of said at least one of the low pressure turbine portion and/or the low pressure compressor sectionssaid rotational speed being an approach speed in revolutions per minute.2. ...

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Номер записи: 38
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: 39
07-11-2013 дата публикации

COMBUSTOR MIXING JOINT AND METHODS OF IMPROVING DURABILITY OF A FIRST STAGE BUCKET OF A TURBINE

Номер: US20130291548A1
Автор: Ingram Clint L., Siden Gunnar Leif
Принадлежит:

The present application and the resultant patent provide a method of improving durability of a first stage bucket of a turbine of a gas turbine engine. The method may include the steps of generating a first combustion flow in a first can combustor and a second combustion flow in a second can combustor, wherein the first can combustor and the second can combustor meet at a joint comprising a flow disruption surface; passing the first combustion flow and the second combustion flow over the flow disruption surface and to a mixing region; substantially mixing the first combustion flow and the second combustion flow in the mixing region to form a mixed combustion flow; and passing the mixed combustion flow to a first stage bucket of a turbine. 1. A method of improving durability of a first stage bucket of a turbine of a gas turbine engine , the method comprising:generating a first combustion flow in a first can combustor and a second combustion flow in a second can combustor, wherein the first can combustor and the second can combustor meet at a joint comprising a flow disruption surface;passing the first combustion flow and the second combustion flow over the flow disruption surface and to a mixing region;substantially mixing the first combustion flow and the second combustion flow in the mixing region to form a mixed combustion flow; andpassing the mixed combustion flow to a first stage bucket of a turbine.2. The method of claim 1 , wherein passing the mixed combustion flow to the first stage bucket comprises generating a substantially uniform velocity field in the first stage bucket.3. The method of claim 1 , wherein passing the mixed combustion flow to the first stage bucket comprises generating a substantially uniform temperature field in the first stage bucket.4. The method of claim 1 , wherein the first combustion flow and the second combustion flow are passed to the mixing region at a first velocity claim 1 , wherein the mixed combustion flow is passed to the ...

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Номер записи: 40
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: 41
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: 42
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: 43
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: 44
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: 45
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: 46
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: 47
09-01-2014 дата публикации

Aircraft Maintenance Apparatus

Номер: US20140007587A1
Автор: Borgess Jason, Grebe Wayne, Heydinger Peter, Spinazze Paul A., Wallace Rodney
Принадлежит: Tronair Inc.

This Mobile Driven Hydraulic Power Unit (HPU) provides a source of clean, pressurized hydraulic fluid for performing required aircraft maintenance. The unit sits on a heavy duty towable frame designed for off-road terrain. Electric start engine, fully instrumented operation panel for both the engine and hydraulic system. A gasoline engine or diesel engine may be used to power the HPU. 1. An apparatus for servicing an aircraft , comprising a mobile vehicle;a hydraulic power unit (HPU) detachably mounted to the vehicle;the HPU further comprising a computer; andthe HPU being capable of providing hydraulic fluid and outputs.2. An apparatus according to claim 1 , wherein the HPU further comprises:a hydraulic pump for producing a fluid hydraulic pressure output to the aircraft.3. An apparatus according to further comprising outputs for pressurized hydraulic fluid.4. An apparatus according to further comprising a gasoline engine for driving the HPU.5. An apparatus according to further comprising a diesel engine for driving the HPU.6. An apparatus according to capable of providing hydraulic fluid at a hydraulic pressure ranging from 250-1500 psi.7. An apparatus according to capable of providing hydraulic fluid at a hydraulic pressure ranging from 250 to 4000 psi. The present patent application is based upon and claims the benefit of provisional patent application No. 61/668,166, filed on Jul. 5, 2012.The present invention relates to the field of aviation and, more particularly, to the maintenance of aircraft.Various aircraft maintenance equipment has been developed for maintaining various portions of an aircraft. Aircraft ground servicing, specifically, provides electrical, hydraulic fluid, and gaseous inputs to aircraft at or on remote locations. An aircraft on the ground whose engine is not functioning requires a number of services to determine whether the aircraft is in a condition to fly or taxi. These services include: electrical power, hydraulic power, engine-start ...

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Номер записи: 48
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: 49
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: 50
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: 51
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: 52
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: 53
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: 54
13-03-2014 дата публикации

METHOD FOR OPERATING A THERMAL POWER PLANT

Номер: US20140069104A1
Автор: Chan Kevin, Li Hongtao, Oesterheld Joerg
Принадлежит: ALSTOM Technology Ltd

The invention relates to a method for operating a thermal power plant, which includes a gas turbine and a generator driven directly by the gas turbine by means of a shaft and being connected to an electrical grid having a grid frequency (F) via an electronic decoupling apparatus and a step-up transformer. A synthetic inertia response is achieved by said method includes the steps of: 1. A method for operating a thermal power plant , which includes a gas turbine and a generator driven directly by the gas turbine by means of a shaft and being connected to an electrical grid having a grid frequency (F) via an electronic decoupling apparatus and a step-up transformer , said method comprising the steps of:{'sub': 'G', 'sensing said grid frequency (F);'}{'sub': 'G', 'detecting if in case of an excursion of said grid frequency (F) additional inertial power (ΔP) is required or not;'}if inertial power (ΔP) is required, calculating the magnitude and duration of the additional inertial power (ΔP); andreleasing additional inertial power (ΔP) to said electrical grid in accordance with said calculations via said electronic decoupling apparatus.2. The method according to claim 1 , wherein the detecting step is based on a predefined rate of change of said grid frequency (F) and a predetermined frequency threshold of said grid frequency (F).3. The method according to claim 1 , wherein said electronic decoupling apparatus has a short-term capacity claim 1 , and within said calculating step said short-term capacity of the electronic decoupling apparatus and/or the initial operating conditions of the power plant at start of said grid frequency excursion are considered.4. The method according to claim 1 , wherein for said power releasing step set points are given to said electronic decoupling apparatus to release an active inertial power and also a reactive power to said electrical grid.5. The method according to wherein an upper and lower threshold for said grid frequency (F) or turbine ...

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Номер записи: 55
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: 56
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: 57
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: 58
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: 59
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: 60
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: 61
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: 62
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: 63
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: 64
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: 65
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: 66
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: 67
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: 68
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: 69
07-01-2016 дата публикации

TURBOMACHINE FAN CLUTCH

Номер: US20160003143A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A gas turbine engine assembly includes, among other things, a clutch configured to move from a first position to a second position in response to rotation of a gas turbine engine fan at a speed greater than a threshold speed. Whether the clutch is in the first position or the second position, the clutch permits rotation of the gas turbine engine fan in a first direction. When the clutch is in the first position, the clutch limits rotation of the gas turbine engine fan only in an opposite, second direction. The clutch is disposed within a compartment that is accessible and removable via removal of an aft engine cover structure. The clutch is removable on-wing. 1. A gas turbine engine assembly , comprising:a clutch configured to move from a first position to a second position in response to rotation of a gas turbine engine fan at a speed greater than a threshold speed,wherein, whether the clutch is in the first position or the second position, the clutch permits rotation of the gas turbine engine fan in a first direction, and when the clutch is in the first position, the clutch limits rotation of the gas turbine engine fan only in an opposite, second direction,wherein the clutch is disposed within a compartment that is accessible and removable via removal of an aft engine cover structure, whereby the clutch is removable on-wing.2. The gas turbine engine assembly of claim 1 , wherein the aft engine cover structure includes an engine exhaust cone.3. The gas turbine engine assembly of claim 2 , wherein the clutch is disposed within an aft bearing compartment and the aft engine cover structure further includes an aft bearing compartment cover plate claim 2 , disposed axially inward of the exhaust cone.4. The gas turbine engine assembly of claim 1 , wherein the clutch is positioned within a gas turbine engine such that the clutch can be moved from an installed position within the gas turbine engine to an uninstalled position without removing any blades from the gas turbine ...

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Номер записи: 70
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: 71
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: 72
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: 73
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: 74
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: 75
02-01-2020 дата публикации

TURBOCHARGED GAS TURBINE ENGINE WITH ELECTRIC POWER GENERATION FOR SMALL AIRCRAFT ELECTRIC PROPULSION

Номер: US20200003115A1
Автор: Jones Russell B., JR. Robert A., Ress
Принадлежит:

A turbocharged gas turbine engine with an electric generator to provide electrical power for an aircraft (e.g., UAV) with multiple propulsor fans each driven by an electric motor, where the engine includes a low spool that drives a main fan and a high spool that drives a high speed electric generator. The low pressure compressor supplies low pressure air to an inlet of the high pressure compressor. A row of stator vanes in the high pressure turbine is cooled using cooling air bled off from the low pressure compressor outlet that is passed through an intercooler and a boost compressor, where the spent vane cooling air is discharged into the combustor. The low pressure turbine and the two compressors each include a variable inlet guide vane to control the power level of the engine. Bypass flow from the main fan is used to cool hot parts of the engine. 1. A power plant for an aircraft propelled by at least one propulsor fan , the power plant comprising:a low spool having a low pressure compressor driven by a low pressure turbine;a high spool having a high pressure compressor driven by a high pressure turbine;a combustor positioned between the high pressure compressor and the high pressure turbine;an outlet of the low pressure compressor is connected to an inlet of the high pressure compressor;the low pressure turbine includes a variable inlet guide vane;the low pressure turbine is located adjacent to the high pressure turbine and hot exhaust from the high pressure turbine flows into the low pressure turbine;a main fan driven by the low spool;an electric generator driven by the high spool;an exhaust nozzle to receive hot exhaust from the low pressure turbine;the high pressure turbine having turbine hot parts with internal cooling air passages; andan intercooler with a boost compressor connected to the low pressure compressor and the combustor through the internal cooling air passages of the turbine hot parts.2. The power plant for an aircraft propelled by at least one ...

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Номер записи: 76
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: 77
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: 78
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: 79
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: 80
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: 81
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: 82
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: 83
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: 84
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: 85
03-01-2019 дата публикации

TURBOJET ENGINE COMPRISING A NACELLE EQUIPPED WITH REVERSER FLAPS

Номер: US20190003421A1
Автор: CZAPLA Lionel, Pautis Olivier, PIARD Frédéric, RIDRAY Frédéric, TIZAC Laurent
Принадлежит:

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

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Номер записи: 86
03-01-2019 дата публикации

FAN MODULE HAVING VARIABLE-PITCH BLADES FOR A TURBINE ENGINE

Номер: US20190003484A1
Автор: BENSLAMA Yanis, DIEVART Jérémy, Igel Dominik, Morreale Serge René
Принадлежит:

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

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Номер записи: 87
13-01-2022 дата публикации

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

Номер: US20220010689A1
Автор: Becky E. Rose, Jonathan Gilson, Robert E. Malecki, Wesley K. Lord, Yuan J. Qiu
Принадлежит: Raytheon Technologies Corp

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

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Номер записи: 88
08-01-2015 дата публикации

Rotary Turbo Rocket

Номер: US20150007549A1
Автор: Bossard John
Принадлежит:

A turbojet is combined with a co-axially integrated rotary rocket to form a propulsion system called a Rotary Turbo Rocket that can function as a turbojet, as an afterburning turbojet, as an Air Turbo Rocket, or as a rotary rocket. The Rotary Turbo Rocket can operate in any of these propulsion modes singularly, or in any combination of these propulsion modes, and can transition continuously or abruptly between operating modes. The Rotary Turbo Rocket can span the zero to orbital flight velocity speed range and/or operate continuously as it transitions from atmospheric to space flight by transitioning between operating modes. 1. A rotary turbo rocket comprising:a forward facing air inlet;a rear facing nozzle outlet;a turbojet engine having a longitudinal axis and comprising a compressor, a combustion chamber, and a turbine;a rotary rocket engine having a longitudinal axis and comprising a combustion chamber and one or more rocket outlets; anda main shaft supporting rotating components of the turbojet and rotary rocket enginewherein the turbojet and rotary rocket engines are arranged coaxially with the turbojet engine positioned behind the air inlet and in front of the rotary rocket and the nozzle outlet positioned behind the rotary rocket engine.2. The rotary turbo rocket of claim 1 , wherein said main shaft is reversibly rotationally coupled to one or both of the turbojet engine and rotary rocket engine.3. The rotary turbo rocket of claim 1 , and further comprising a mechanism for coupling or de-coupling a rotation of the rotary rocket engine to or from a rotation of the turbojet compressor.4. The rotary turbo rocket of claim 1 , and further comprising a mechanism for coupling or de-coupling a rotation of the rotary rocket engine to or from a rotation of the turbojet turbine.5. The rotary turbo rocket of claim 1 , wherein the rotary rocket claim 1 , main shaft and compressor are configured such that the compressor may be driven by power delivered from the rotary ...

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Номер записи: 89
20-01-2022 дата публикации

DEVICES AND METHODS FOR GUIDING BLEED AIR IN A TURBOFAN ENGINE

Номер: US20220018292A1
Автор: ANAND Karan, CAPRON Alexandre, YAM Tammy
Принадлежит:

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

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Номер записи: 90
12-01-2017 дата публикации

GAS TURBINE ENGINE AFT BEARING ARRANGEMENT

Номер: US20170009655A1
Автор: Savela Gregory M.
Принадлежит:

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

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Номер записи: 91
12-01-2017 дата публикации

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

Номер: US20170009656A1
Автор: Christian Vessot, Matthieu Leyko
Принадлежит: Safran Aircraft Engines SAS

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

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Номер записи: 92
12-01-2017 дата публикации

VARIABLE EXHAUST MIXER AND COOLER FOR A THREE-STREAM GAS TURBINE ENGINE

Номер: US20170009703A1
Автор: Kupratis Daniel B., Moon Francis R.
Принадлежит:

A gas turbine engine includes an outer case structure around a central longitudinal engine axis. An intermediate case structure is included inboard of the outer case structure. An inner case structure is included inboard of the intermediate case structure. A variable area exhaust mixer is included, which is movable between a closed position adjacent to the intermediate case structure and an open position adjacent to the inner case structure. 1. A gas turbine engine , comprising:an outer case structure around a central longitudinal engine axis;an intermediate case structure inboard of the outer case structure;an inner case structure inboard of the intermediate case structure; anda variable area exhaust mixer movable between a closed position adjacent to the intermediate case structure and an open position adjacent to the inner case structure.2. The gas turbine engine as recited in claim 1 , wherein the variable area exhaust mixer is downstream of a turbine section.3. The gas turbine engine as recited in claim 1 , wherein the variable area exhaust mixer includes a hinge axis radially outboard of a maximum outer diameter of a tailcone.4. The gas turbine engine as recited in claim 3 , wherein the hinge axis is axially aft of a multiple of respective flaps.5. The gas turbine engine as recited in claim 4 , wherein the multiple of respective flaps open radially inward toward the inner case structure.6. The gas turbine engine as recited in claim 4 , wherein the multiple of respective flaps are operable to extend toward a wall in the inner case structure when between the closed position and the open position.7. The gas turbine engine as recited in claim 6 , wherein the wall separates an exhaust of a second stream flow path and an exhaust of a core flow path.8. The gas turbine engine as recited in claim 7 , wherein the multiple of respective flaps are operable to extend to the wall to segregates the second stream flow path and the core flow path.9. The gas turbine engine as ...

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Номер записи: 93
14-01-2016 дата публикации

VENTILATION INLET

Номер: US20160010558A1
Автор: HUSSAIN Zahid M., IRELAND Peter T., MISHRA Nityabh
Принадлежит:

A ventilation inlet including a conduit () arranged to convey flow from a first flow zone to a second flow zone. The conduit has a mouth region () presenting to the first flow zone an entrance aperture to receive the flow therefrom. The conduit has a baffle () spanning a portion of the conduit to define a throat region (), the throat region being narrower than the entrance aperture. The throat region () is movable along the conduit () to control the flow through the ventilation inlet. 1. A ventilation inlet including a conduit arranged to convey flow from a first flow zone to a second flow zone , the conduit having a mouth region presenting to the first flow zone an entrance aperture to receive the flow therefrom; anda baffle spanning a portion of the conduit to define a throat region, the throat region being narrower than the entrance aperture; whereinthe throat region is movable along the conduit to control the flow through the ventilation inlet.2. A ventilation inlet according to wherein the baffle is movable along the conduit.3. A ventilation inlet according to wherein the baffle is movable along the general direction of flow in the conduit.4. A ventilation inlet according to wherein the baffle spans transversely across a portion of the conduit.5. A ventilation inlet according to wherein the baffle is locatable in a first position claim 1 , in which the baffle is located downstream of the mouth region so that the throat region is recessed from the entrance aperture.6. A ventilation inlet according to wherein the baffle is movable to a second position claim 1 , in which the baffle is located in the mouth region so that the throat region is presented to the first flow zone to receive the flow instead of the entrance aperture of the mouth region.7. A ventilation inlet according to wherein the conduit has a central axis extending generally in the direction of flow through the conduit.8. A ventilation inlet according to wherein the baffle is arranged to extend at ...

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Номер записи: 94
14-01-2016 дата публикации

TWO-PART GAS TURBINE ENGINE

Номер: US20160010589A1
Автор: ROLT Andrew Martin
Принадлежит:

A gas turbine engine propulsion system in which a first propulsive unit has a core engine and a first low pressure turbine arranged to be driven by combustion products from the core engine. The first propulsive unit also has a first fan rotor and a first fan shaft drivingly connecting the first turbine and the first fan rotor. The propulsion system also has a further turbine arranged in flow series with the first turbine and a second propulsive unit spaced from the first propulsive unit. The second propulsive unit has a second fan rotor driven by the rotational output of the further turbine. The further turbine may be located in the second propulsive unit and may be in fluid communication with the first turbine via an inter-turbine duct. 1. A gas turbine engine propulsion system comprising:a first propulsive unit having a core engine and a first turbine arranged to be driven by combustion products from the core engine, the first propulsive unit further comprising a first fan rotor and a first fan shaft drivingly connecting the first turbine and the first fan rotor,wherein the propulsion system comprises a further turbine arranged in flow series with the first turbine and a second propulsive unit spaced from the first propulsive unit, the second propulsive unit having a second fan rotor driven by the rotational output of the further turbine;wherein the first and second fan rotors comprise a parallel flow fan arrangement and the second fan rotor is driven by combustion products from the core engine via the further turbine;wherein the first and second propulsion units include a respective nacelle and bypass fan duct.the further turbine is connected to the second fan rotor by a second fan shaft, the further turbine being located in the second propulsive unit.2. A propulsion system according to claim 1 , wherein the further turbine is downstream of the first turbine with respect to the flow of combustion products from the core engine claim 1 , the first and further ...

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Номер записи: 95
14-01-2016 дата публикации

NOZZLE ARRANGEMENT FOR A GAS TURBINE ENGINE

Номер: US20160010590A1
Автор: ROLT Andrew Martin
Принадлежит:

A gas turbine engine comprising: a bypass duct having a bypass nozzle; an engine core having a core nozzle; and, a mixer duct defined by a mixer fairing and having a mixer nozzle, wherein the mixer duct is arranged to receive an airflow from the bypass duct through a mixer duct inlet and an airflow from the engine core, when in use, and the geometry of the mixer duct is selectively adjustable by moving the mixer fairing relative to the bypass duct and engine core in use. 1. A gas turbine engine comprising:a bypass duct having a bypass nozzle;an engine core having a core nozzle; and,a mixer duct having a mixer duct inlet and a mixer nozzle defined by a mixer fairing which is movably mounted to the engine and,wherein the mixer duct is arranged to receive an airflow from the bypass duct through the mixer duct inlet and an airflow from the engine core, when in use, and the geometry of the mixer duct is selectively adjustable by moving the mixer fairing relative to the bypass duct and engine core in use,wherein the mixer fairing is movable between a first position and second position which simultaneously alters one or more of: an output flow area of the bypass nozzle, an output flow area of the mixer nozzle, and, a throat area of the mixer duct inlet, such that the respective in use airflows are altered.2. A gas turbine engine as claimed in claim 1 , wherein moving the mixer fairing between the first and second position alters all of: the output flow area of the bypass nozzle claim 1 , the output flow area of the mixer nozzle claim 1 , and claim 1 , the throat area of the mixer duct inlet.3. A gas turbine engine as claimed in claim 1 , wherein a portion of radially outer wall of the mixer fairing downstream of the leading edge is substantially parallel to the axis of movement such that moving the mixer fairing between a first and second position alters the output flow area of the mixer nozzle and moving the mixer fairing between a second and third position alters the ...

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Номер записи: 96
11-01-2018 дата публикации

GEARED GAS TURBINE ENGINE

Номер: US20180010551A1
Автор: Sheridan William G.
Принадлежит:

A gas turbine engine includes a fan section that includes a fan rotatable about an engine axis. A compressor section includes a low pressure compressor rotatable about the engine axis. A turbine section includes a fan drive turbine for driving the fan and the low pressure compressor. A speed reduction device connects the fan drive turbine to the fan and the low pressure compressor. The speed reduction device includes a sun gear driven by an inner shaft. A plurality of intermediate gears surround the sun gear. A carrier supports the plurality of intermediate gears for driving the low pressure compressor. A ring gear is located radially outward from the intermediate gears and includes a forward portion for driving a fan drive shaft and an aft portion. 1. A gas turbine engine comprising:a fan section including a fan rotatable about an engine axis;a compressor section including a low pressure compressor rotatable about the engine axis;a turbine section including a fan drive turbine for driving the fan and the low pressure compressor; and a sun gear driven by an inner shaft;', 'a plurality of intermediate gears surrounding the sun gear;', 'a carrier supporting the plurality of intermediate gears for driving the low pressure compressor; and', 'a ring gear located radially outward from the intermediate gears including a forward portion for driving a fan drive shaft and an aft portion., 'a speed reduction device connecting the fan drive turbine to the fan and the low pressure compressor, the speed reduction device including2. The gas turbine engine of claim 1 , wherein the carrier includes an axially forward portion and an axially aft portion claim 1 , the plurality of intermediate gears are located axially between the axially forward portion and the axially aft portion of the carrier.3. The gas turbine engine of claim 2 , wherein axially aft portion of the carrier is connected to the low pressure compressor.4. The gas turbine engine of claim 1 , wherein the fan drive ...

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Номер записи: 97
11-01-2018 дата публикации

FAN BLADE

Номер: US20180010613A1
Автор: ALLAIR Frederic Alain Edouard, BOISSON Alexandre Bernard Marie, FIGEUREU Claire Marie, LANCIEN Thea Claire, MARTIN Guillaume Olivier Vartan
Принадлежит: SAFRAN AIRCRAFT ENGINES

A blade including at least one web and a vane having a leading edge and a trailing edge, wherein, for at least one aerofoil of the vane in the vicinity of the web, a maximum sweep angle associated with a position along a chord of the aerofoil extending from the leading edge to the trailing edge of the vane corresponding to a relative chord length of at least 50%. 1. A fan blade of a bypass turbine engine , comprising at least one shank and a vane having a leading edge and a trailing edge , wherein , for at least one aerofoil of the vane in the vicinity of the shank , a maximum camber defining a point of a skeleton of the aerofoil extending from the leading edge to the trailing edge of the vane wherein a distance with a chord of the aerofoil extending from the leading edge to the trailing edge of the vane is maximum , said maximum camber being associated with a position along said chord corresponding to a relative chord length of at least 55% , with an offset of said maximum camber toward the trailing edge.2. The blade according to claim 1 , wherein said position along the chord of the aerofoil associated with the maximum camber corresponds to a relative chord length comprised between 55% and 75%.3. The blade according to claim 2 , wherein said position along the chord of the aerofoil associated with the maximum camber corresponds to a relative chord length comprised between 55% and 65%.4. The blade according to claim 1 , being made of a woven composite material.5. The blade according to claim 1 , further comprising a straight root connected to the vane with the shank.6. A fan for a bypass turbine engine comprising at least one blade according to .7. The fan according to claim 6 , comprising a disk from which said blade extends substantially radially.8. The fan according to claim 7 , wherein the shank extends outside the disk and on the inside of platforms defining the interior of the stream.9. The fan according to claim 7 , further comprising a straight root ...

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Номер записи: 98
11-01-2018 дата публикации

NON-NEWTONIAN MATERIALS IN AIRCRAFT ENGINE AIRFOILS

Номер: US20180010614A1
Автор: Kray Nicholas Joseph, Sellinger Aaron Todd
Принадлежит:

A component is provided for a turbine engine. The component can include an airfoil defining a surface, and an energy absorbing composite positioned on the surface of the airfoil or within the airfoil. The energy absorbing composite includes a shear thickening fluid distributed through a matrix. 1. A component for a turbine engine , the component comprising:an airfoil defining a surface; andan energy absorbing composite positioned on the surface of the airfoil or within the airfoil, wherein the energy absorbing composite includes a shear thickening fluid distributed through a matrix.2. The component as in claim 1 , wherein the energy absorbing composite is positioned within the construction of the airfoil.3. The component as in claim 1 , wherein the energy absorbing composite is positioned on at least a portion of the surface of the airfoil.4. The component as in claim 3 , wherein the energy absorbing composite is positioned on a leading edge of the airfoil claim 3 , a side surface of the airfoil claim 3 , or both.5. The component as in claim 1 , wherein the matrix comprises a solid foamed synthetic polymer matrix.6. The component as in claim 5 , wherein the solid foamed synthetic polymer matrix comprises a synthetic elastomer.7. The component as in claim 6 , wherein the synthetic elastomer comprises an elastomeric polyurethane.8. The component as in claim 6 , wherein the synthetic elastomer comprises a first polymer-based elastic material and a second polymer-based elastic material.9. The component as in claim 8 , wherein the first polymer-based elastic material comprises an ethylene vinyl acetate or an olefin polymer claim 8 , and wherein the second polymer-based elastic material comprises a silicone polymer having dilatant properties.10. The component as in claim 5 , wherein the energy absorbing composite further comprises a polymer-based dilatant.11. The component as in claim 10 , wherein the polymer-based dilatant comprises a silicone polymer having dilatant ...

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Номер записи: 99
11-01-2018 дата публикации

STRUT ASSEMBLY FOR AN AIRCRAFT ENGINE

Номер: US20180010616A1
Автор: King Daniel Elmer, Roll Christopher Timothy, Seitzer Kenneth E.
Принадлежит:

A strut assembly for a gas turbine engine includes an outer structural case. The outer structural case includes a first mounting pad for mounting a first strut and a second mounting pad for mounting a second strut. The outer structural case further includes a case ligament extending between the first mounting pad and the second mounting pad in a substantially straight direction to reduce an amount of bending stress on the outer structural case. 1. A strut assembly for a gas turbine engine , the strut assembly comprising: a first mounting pad for mounting a first strut;', 'a second mounting pad for mounting a second strut; and', 'a case ligament extending between the first mounting pad and the second mounting pad, the case ligament extending in a substantially straight direction from the first mounting pad to the second mounting pad., 'an outer structural case comprising'}2. The strut assembly of claim 1 , wherein the first mounting pad claim 1 , the second mounting pad claim 1 , and the case ligament are each formed of a composite material.3. The strut assembly of claim 1 , wherein the case ligament is formed of a composite material claim 1 , wherein the composite material forming the case ligament includes a plurality of substantially aligned fibers.4. The strut assembly of claim 3 , wherein the plurality of substantially aligned fibers extend in a direction from the first mounting pad to the second mounting pad.5. The strut assembly of claim 1 , wherein the case ligament defines an inside surface claim 1 , wherein the outer structural case further comprises a plurality of wedge members positioned along the inside surface of the case ligament adjacent to the first mounting pad and adjacent to the second mounting pad.6. The strut assembly of claim 5 , wherein the plurality of wedge members are non-structural components.7. The strut assembly of claim 6 , wherein the plurality of wedge members are formed of a composite material claim 6 , wherein the case ligament is ...

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