СИСТЕМА РЕГУЛИРОВАНИЯ ДАВЛЕНИЯ ОТБИРАЕМОГО ВОЗДУХА ДЛЯ ПРОТИВООБЛЕДЕНИТЕЛЬНОЙ СИСТЕМЫ ВОЗДУШНОГО СУДНА, ПРОТИВООБЛЕДЕНИТЕЛЬНАЯ СИСТЕМА МОТОГОНДОЛЫ ДВИГАТЕЛЯ ВОЗДУШНОГО СУДНА И СПОСОБ РЕГУЛИРОВАНИЯ ОТБОРА ВОЗДУХА В ПРОТИВООБЛЕДЕНИТЕЛЬНОЙ СИСТЕМЕ

Изобретение относится к противообледенительным системам летательных аппаратов. Система (30) регулирования давления отбираемого воздуха для противообледенительной системы воздушного летательного аппарата содержит первый расположенный выше по потоку регулирующий давление клапан (32) и второй расположенный ниже по потоку регулирующий давление клапан (34), расположенные последовательно в пути (28) отбираемого потока воздуха. Каждый из клапанов (32, 34) имеет камеру (74) регулировки давления, сообщающуюся по текучей среде с соответствующим устанавливающим давление клапаном (82, 84), каждый из которых сообщается по текучей среде с впуском (88, 96) отбираемого воздуха выше по потоку от первого регулирующего давление клапана (32). Первый регулирующий давление клапан (32) устанавливают для регулирования давления отбираемого воздуха до первого давления, и второй расположенный ниже по потоку регулирующий давление клапан (34) устанавливают для регулирования давления отбираемого воздуха до второго давления ...

ЭЛЕМЕНТ КОНСТРУКЦИИ ЛЕТАТЕЛЬНОГО АППАРАТА

Изобретение относится к элементу конструкции, способному выдерживать повышенные температуры, в частности к заднему шпангоуту гондолы летательного аппарата. Задний шпангоут воздухозаборника гондолы летательного аппарата содержит часть, окружающую отверстие, предусмотренное для прохода системы для устранения обледенения из композитного материала на базе геополимерной смолы, усиленной волокнами, и другую металлическую часть. Достигается уменьшение веса конструкции. 3 з.п. ф-лы, 5 ил.

ПЕРЕДНЯЯ КРОМКА ЛЕТАТЕЛЬНОГО АППАРАТА

Изобретение относится к области авиации, более конкретно к передней кромке летательного аппарата. Передняя кромка летательного аппарата содержит покрытие (26) для акустической обработки. Покрытие включает в себя акустически резистивный пористый слой (28), ячеистую структуру (30) и отражающий слой (32), при этом в упомянутое покрытие (26) встроена противообледенительная система. Противообледенительная система содержит вибрационный излучатель (36), размещенный в гнезде (38), открывающемся на уровне аэродинамической поверхности. Технический результат заключается в оптимизации совместного использования покрытия акустической обработки и противообледенительной системы передней кромки летательного аппарата. 2 н. и 9 з.п. ф-лы, 7 ил.

УСОВЕРШЕНСТВОВАННАЯ КОНСТРУКЦИЯ ВХОДНОГО УСТРОЙСТВА

Конструкция компактного входного устройства, включающего в себя одну перегородку и/или акустическую панель, проходящую в область кромки гондолы для уменьшения шума. Компактное входное устройство используют с низкоэнергетической системой защиты от обледенения на основе текучей среды, выполненной с возможностью предотвращения образования льда на акустической панели в области кромки гондолы. Пористая панель объединена с обшивкой входной кромки; причем низкотемпературная текучая среда на основе гликоля для защиты от обледенения просачивается через пористую панель на внешнюю поверхность обшивки входной кромки с последующим проходом назад по потоку воздуха на акустическую панель и тем самым предотвращает образование льда на пористой панели и уменьшает или предотвращает образование льда на акустической панели. Технический результат заключается в том, что термическая защита не требуется. 8 н. и 13 з.п. ф-лы, 14 ил.

СПОСОБ ИЗГОТОВЛЕНИЯ АКУСТИЧЕСКОЙ ПАНЕЛИ ДЛЯ КРОМКИ ВОЗДУХОЗАБОРНИКА ГОНДОЛЫ

Изобретение относится к области авиастроения, более конкретно, способу изготовления акустической панели для кромки воздухозаборника самолета, а также к кромке воздухозаборника, снабженной такой панелью, и гондоле газотурбинного двигателя. Способ изготовления акустической панели (12) для кромки (2) воздухозаборника гондолы (1) включает в себя этапы создания перфорированного противообледенительного узла (14), содержащего сетку из проводящих элементов, полученную методом фотолитографии, причем указанный противообледенительный узел (14) крепят к структуре (13) с ячеистой сердцевиной. Технический результат заключается в упрощении процесса изготовления акустической панели. 3 н. и 12 з.п. ф-лы, 12 ил.

РАЗДЕЛИТЕЛЬ ПОТОКА ГАЗА С УСТРОЙСТВОМ ДЛЯ ПРЕДОТВРАЩЕНИЯ ОБЛЕДЕНЕНИЯ, СОДЕРЖАЩИМ ТЕПЛОВОЙ МОСТ

Разделитель потока газа, способный разделять поток газа на первый поток и второй поток, содержит переднюю кромку разделителя и устройство для предотвращения обледенения передней кромки. Устройство для предотвращения обледенения содержит, по меньшей мере, металлическую лопатку, которая находится в тепловом контакте с передней кромкой и проходит от передней кромки к заднему краю разделителя на некотором расстоянии от передней кромки для того, чтобы находиться в тепловом контакте с источником тепла (24), расположенным на некотором расстоянии от передней кромки. Изобретение направлено на создание простого экономичного и надежного решения проблемы обледенения переднего (входного) края разделителя газового потока. 2 н. и 13 з.п. ф-лы, 5 ил.

ЭЛЕКТРИЧЕСКОЕ ПРОТИВООБЛЕДЕНИТЕЛЬНОЕ УСТРОЙСТВО И СООТВЕТСТВУЮШАЯ СИСТЕМА КОНТРОЛЯ

Изобретение относится к области авиации, более конкретно к противообледенительному устройству для одного из элементов гондолы турбореактивного двигателя. Устройство содержит электротермический противообледенитель, соединенный источником электропитания (3) и образующий таким образом группу (1) электротермических противообледенителей. Группа электротермических противообледенителей включает в себя одну или несколько подгрупп электротермических противообледенителей, каждая из которых включает в себя, в свою очередь, один или несколько электротермических противообледенителей группы, причем отдельные подгруппы электротермических противообледенителей имеют разные значения омического сопротивления. Технический результат заключается в упрощении конструкции гондолы и снижении ее веса. 3 н. и 11 з.п. ф-лы, 5 ил.

ТЕПЛОПЕРЕДАЮЩАЯ СИСТЕМА ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Теплопередающая система для газотурбинного двигателя с обтекателем воздухозаборника содержит кольцевой кожух вентилятора, кольцевой обтекатель воздухозаборника, расположенный спереди кольцевого кожуха вентилятора, теплообменник, имеющий проходящий через него источник тепла и установленный снаружи кольцевого кожуха вентилятора; и множество тепловых трубок, каждая из которых проходит между теплообменником и внутренним пространством обтекателя воздухозаборника. Каждая тепловая трубка имеет задний участок, проходящий в осевом направлении, переходный участок, проходящий в круговом направлении и имеющий дугообразную форму. Переходный участок соединяет друг с другом задний участок с передним участком, расположенным вблизи верха обтекателя воздухозаборника. Передний участок проходит внутри обтекателя воздухозаборника, при этом передние участки каждой из множества тепловых трубок расположены внутри обтекателя воздухозаборника по окружности. Изобретение направлено на повышение экономичности за счет ...

ПОКРЫТИЕ ДЛЯ АКУСТИЧЕСКОЙ ОБРАБОТКИ, ПЕРЕДНЯЯ КРОМКА И ВОЗДУХОЗАБОРНИК ЛЕТАТЕЛЬНОГО АППАРАТА, СОДЕРЖАЩИЕ ТАКОЕ ПОКРЫТИЕ

Покрытие для акустической обработки, нанесенное на передней кромке воздухозаборника гондолы летательного аппарата, содержит акустически резистивный слой, ячеистую структуру и отражающий слой. Ячеистая структура содержит множество каналов, выходящих, с одной стороны, на уровне первой поверхности и, с другой стороны, на уровне второй поверхности. Ячеистая структура содержит вырезы или отверстия, выполненные на уровне боковых стенок некоторых каналов, позволяющие устанавливать сообщение между смежными каналами таким образом, чтобы создавать сеть сообщающихся каналов, изолирующих один или группу несообщающихся каналов. По меньшей мере, один из сообщающихся каналов соединен с вводом горячего газа. Другие изобретения группы относятся к передней кромке летательного аппарата и воздухозаборнику гондолы летательного аппарата, содержащим указанное выше покрытие для акустической обработки. Изобретения позволяют совместить покрытие для акустической обработки с антиобледенительной системой. 3 н. и 7 ...

СПОСОБ ВЫПОЛНЕНИЯ ПРОТИВООБЛЕДЕНИТЕЛЬНОЙ СИСТЕМЫ НА ПАНЕЛИ ГОНДОЛЫ

Заявленное изобретение относится к области авиации, более конкретно к способу выполнения противообледенительной системы на панели (22) гондолы. Способ включает в себя следующие этапы: A) с помощью позиционирующего средства (35) на наружной обшивке (24) позиционируют вокруг отверстия или отверстий сетку из резистивных элементов; B) с помощью средства для нанесения наносят сетку из резистивных элементов в определенное на этапе A место для формирования противообледенительной системы; C) на полученную противообледенительную систему наносят поверхностное покрытие. Технический результат заключается в повышении эффективности удаления льда противообледенительной системой. 3 н. и 9 з.п. ф-лы, 14 ил.

ПОДОГРЕВАТЕЛЬ-КАПЛЕУЛОВИТЕЛЬ

Полезная модель относится к оборудованию, применяемому в энергетике, в частности в блочно-модульных электростанциях по технологии органического цикла Ренкина, газотурбинных и парогазовых установках, газоперекачивающих агрегатах на магистральных газопроводах. Подогреватель-каплеуловитель предназначен для подогрева воздуха, подаваемого в компрессор газовой турбины, за счет вторичного тепла, а также для эффективного удаления из воздуха капельной влаги и снега.Подогреватель-каплеуловитель, характеризующийся тем, что он включает: обогреваемый жидким теплоносителем парогенератор, представляющий собой кожухотрубчатый теплообменник с кипением органического рабочего тела внутри корпуса на поверхности теплообменных труб, по которым циркулирует горячий жидкий теплоноситель, проток которого регулируется клапаном теплоносителя, приваренным к вваренному в крышку входной камеры жидкого теплоносителя патрубку входа теплоносителя в парогенератор с одной стороны и к подающему трубопроводу теплоносителя с ...

ВОЗДУХООЧИСТИТЕЛЬНОЕ УСТРОЙСТВО

Изобретение относится к области машиностроения и может быть использовано в энергетике, газовой, нефтяной и других отраслях промышленности в качестве воздухоочистительного устройства (ВОУ) накопительного типа для очистки воздуха, подаваемого в газотурбинные и компрессорные установки (ГТУ) в объеме от 84 тыс. м3/ч до 165 тыс. м3/ч. Устройство состоит из двух модулей верхнего 1 и нижнего 2. В верхнем модуле 1 между блоками воздушных фильтров 8 расположен проход 9. Блоки воздушных фильтров 8 расположены инверсно и включают в себя наборы фильтроэлементов 6, каждый из которых состоит из влагоотделителя и фильтров грубой и тонкой очистки. Количество наборов фильтроэлементов в каждом блоке воздушных фильтров определяется его габаритами. Защитный колпак снабжен воздухозаборными козырьками 5, которые выполнены со сплошными боковыми стенками и снабжены защитной сеткой 10 расположенной в нижней части воздухозаборных козырьков. В верхнем модуле 1 дополнительно установлена противооблединительная система ...

ПРОТИВООБЛЕДЕНИТЕЛЬНАЯ СИСТЕМА ВХОДНОГО КОНУСА АВИАЦИОННОГО ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

... 1. Противообледенительная система входного конуса авиационного турбинного двигателя, содержащая средства воздушного диффузора, предназначенные для установки во входном конусе турбинного двигателя и обеспечивающие подачу в него горячего воздуха, отличающаяся тем, что система содержит контур удаления нагнетаемого воздуха, по меньшей мере, из одной ограниченной полости-площадки турбинного двигателя, при этом контур соединен со средством воздушного диффузора и обеспечивает подачу в них горячего воздуха. 2. Противообледенительная система по п.1, отличающаяся тем,что контур удаления нагнетаемого воздуха, по меньшей мере, из одной ограниченной полости-площадки турбинного двигателя содержит канал первичного воздуха, расположенный, по меньшей мере, частично внутри системы ведущих валов турбинного двигателя, при этом канал первичного воздуха ориентирован параллельно продольной оси турбинного двигателя. 3. Противообледенительная система по п.2, отличающаяся тем,что канал первичного воздуха установлен ...

ПЕРЕДНЯЯ КРОМКА ЛЕТАТЕЛЬНОГО АППАРАТА

... 1. Передняя кромка летательного аппарата, такая как, например, воздухозаборник (22) гондолы (14) двигательной установки (10), содержащая покрытие (26) для акустической обработки, включающее в себя, снаружи во внутрь, акустически резистивный пористый слой (28), обладающий определенной долей открытой поверхности, по меньшей мере, одну ячеистую структуру (30) и отражающий слой (32), при этом в упомянутое покрытие (26) встроена противообледенительная система, отличающаяся тем, что противообледенительная система содержит, по меньшей мере, один вибрационный излучатель (36). ! 2. Передняя кромка по п.1, отличающаяся тем, что вибрационный излучатель или излучатели (36) находятся в контакте с обрабатываемой поверхностью. ! 3. Передняя кромка по п.1 или 2, отличающаяся тем, что, по меньшей мере, один вибрационный излучатель (36) размещен в гнезде (38), открывающемся на уровне аэродинамической поверхности. ! 4. Передняя кромка по п.3, отличающаяся тем, что гнездо (38) содержит закрывающие его средства ...

СИСТЕМА УСТРАНЕНИЯ ОБЛЕДЕНЕНИЯ ПЕРЕДНЕЙ КРОМКИ ВХОДНОГО ОТВЕРСТИЯ НОСОВОГО ОБТЕКАТЕЛЯ ТУРБИННОГО ДВИГАТЕЛЯ

... 1. Система устранения обледенения передней кромки (1) входного отверстия носового обтекателя (2) турбинного двигателя, в частности, для самолета, причем упомянутая передняя кромка (1) выполнена полой и образует кольцевую камеру (4), закрытую первой внутренней перегородкой (5), содержащая: ! трубку (6, 19) подачи горячего воздуха под давлением, выполненную с возможностью подключения на ее заднем конце, противоположном упомянутой передней кромке (1), к контуру (7) подачи горячего воздуха под давлением и, на ее переднем конце, в направлении к упомянутой передней кромке (1), к инжектору (8), нагнетающему упомянутый горячий воздух под давлением в упомянутую кольцевую камеру (4) передней кромки, причем упомянутая трубка подачи пропущена через отсек (10) упомянутого входного отверстия носового обтекателя (2), который ограничен с передней стороны упомянутой первой внутренней перегородкой (5, 16, 20) и с задней стороны второй внутренней перегородкой (5, 16, 20); и ! внутренний защитный кожух (9, ...

ВПУСКНАЯ ЗАСЛОНКА ДВИГАТЕЛЯ ДЛЯ УСТАНОВКИ НА КОРПУСЕ ВОЗДУХОЗАБОРНИКА ДВИГАТЕЛЯ САМОЛЕТА, А ТАКЖЕ ДВИГАТЕЛЬ С ТАКОЙ ВПУСКНОЙ ЗАСЛОНКОЙ И САМОЛЕТНАЯ СИСТЕМА

... 1. Впускная заслонка (К) двигателя, которая предусмотренна для установки на корпусе воздухозаборника или канала воздухозаборника двигателя самолета, имеющая первый конец (Е1) и расположенный противоположно ему и в продольном направлении (L) впускной заслонки на расстоянии от него второй конец (Е2), ! при этом продольное направление (L) во время предполагаемого применения направлено против направления (S) потока поступающего в двигатель воздуха, и при этом воздушная заслонка имеет: ! - основной корпус (1) впускной заслонки с предназначенным для шарнирного соединения соединительным устройством для шарнирного соединения основного корпуса (1) впускной заслонки с корпусом воздухозаборника или каналом воздухозаборника с простирающейся вдоль второго конца (Е2) шарнирной осью (А), ! - удлинительную деталь (2) впускной заслонки, которая конструктивно интегрирована в основной корпус (1) впускной заслонки и имеет первую и вторую боковую деталь (5, 6), которые соответственно отходят от основного корпуса ...

Звукопоглощающие сопла для систем отбора тепла на вход

КОМПЛЕКСНОЕ ВОЗДУХООЧИСТИТЕЛЬНОЕ УСТРОЙСТВО

Полезная модель относится к устройствам для очистки забираемого из атмосферы воздуха и подготовки его для подачи в компрессор газотурбинного двигателя газоперекачивающего агрегата. Техническим результатом полезной модели является создание конструкции нового комплексного воздухоочистительного устройства, обеспечивающего надежность работы газоперекачивающего агрегата в зимний период в прибрежной арктической зоне Крайнего Севера. Технический результат достигается в заявляемом комплексном воздухоочистительном устройстве, содержащем воздухоприемную камеру (1), в которой установлены блоки комбинированной системы фильтрации, включающие фильтры грубой очистки (2); осадкозащитные козырьки (3), систему подогрева (4) всасываемого атмосферного воздуха, глушители шума (5), согласно полезной модели, фильтры грубой очистки (2) выполнены в виде гидрофобных коалесцирующих фильтров карманного типа, установленных против входящего потока воздуха и соединенных в блоки с кассетными фильтрами тонкой очистки ( ...

Beheizung einer internen Oberfläche eines Bläsertriebwerks zur Verhinderung von Eisbildung

Schutzeinrichtung für den mit einem Enteisungssystem versehenen Lufteinlauf eines Strahltriebwerks

Gas turbine intake anti-icing device

A gas turbine intake anti-icing device for a gas turbine electric power generation system 1 is provided, the system having a gas turbine 2 and a power generator 20 that is coupled to the gas turbine and rotationally driven to generate electrical power. The device includes a power generator cooling mechanism 21, 22, 23, 25 which takes in air from the outside and introduces the air into the power generator 20 to cool the power generator. An exhaust air supply path 31 connects a gas turbine air intake path 9 to an exhaust path 30 for air that is discharged from the power generator cooling mechanism after the power generator is cooled. The discharged air from the power generator cooling mechanism is supplied to the intake path of the gas turbine through the exhaust air supply path. The exhaust air supply path may be connected to the intake path nearest the inlet of the gas turbine engine.

Method for detecting a fluid leak in a turbomachine and system for distributing a fluid

The invention relates to a method for detecting a fluid leak in a turbomachine (10). The turbomachine (10) comprises a high temperature fluid source, at least one fluid distribution pipe (14, 15) adapted to distribute said fluid to different parts of the turbomachine (10) and/or the aircraft (20) which is intended to be equipped with said turbomachine (10), a turbomachine compartment in which the distribution pipe (14, 15) is at least partly accommodated, said compartment having in operation a low temperature relative to the high temperature of the fluid supplied by the fluid source. The method comprises the following steps: measuring a temperature variation in the compartment between two instants to obtain a temperature gradient; and detecting a fluid leak if the temperature gradient is greater than or equal to a threshold temperature gradient. The invention further relates to a system for distributing a high temperature fluid for a turbomachine and a turbomachine (10) comprising such ...

Icing prevention of a gas turbine engine pressure sensing assembly

A combination of a gas turbine pressure sensing assembly 30 and an engine electronics unit 38. The pressure sensing assembly includes a pressure manifold 34 having an air inlet, one or more air outlets, and one or more pressure sensors 32 connected to the air outlets to sense the pressure of air entering through the air inlet. The Engine Electronics Unit (EEU) produces waste heat in operation. The pressure sensing assembly further includes a heat conduction path 40 including a heat pipe 42, thermally connecting the engine electronics unit to the manifold, wherein the manifold acts as a waste heat sink. The manifold temperature rise produced by the waste heat prevents icing of the manifold. The heat conduction path may include a housing 36 surrounding the manifold and having a higher thermal conductivity than the manifold. Preferably a power supply sub-unit 39 of the EEU produces the waste heat and the EEU is an electronic engine controller or an engine health monitoring unit.

Improvements relating to the supply of hot air from a gas turbine engine for anti-icing or other purposes

... 738,889. Axial flow compressors; gas turbine plant. ARMSTRONG SIDDELEY MOTORS, Ltd. June 1, 1954 [July 3, 1953], No. 18447/53. Classes 110 (1) and 110 (3). In a gas turbine plant comprising an axial flow compressor with an outlet diffuser connected to the combustion chamber formed by inner and outer casings interconnected by aerofoil section spokes and a channel section member at the outer ends of the spokes forming part of a hollow casing which carries the engine mountings, hot air is fed from the diffuser into the hollow casing which is arranged to support a manifold having an outlet port to which hot air is supplied from the casing through a control valve. The invention is described with reference to a compound gas turbine engine comprising high and low pressure compressors driven by high and low pressure turbines. The engine is supported by two pairs of trunnions, one pair being mounted on the intake end of the compressor casing and the other pair on the outlet diffuser of the high ...

De-icing device for gas turbine plants

... 758,709. Gas turbine plant. HOLLEY CARBURETOR CO. Dec. 10, 1954, No. 35781/54. Class 110(3) [Also in Group XXIX] A gas turbine plant comprising an air compressor feeding a combustion chamber and a turbine has a plurality of hollow bars across the inlet of the compressor with elongated slots on their downstream edges, a passage for conveying hot gases from the turbine exhaust to the interior of the bars and valve means controlling the flow in the passage which is responsive to a reduction in air flow or air pressure or both between the compressor and turbine as the result of ice formation on the compressor air inlet. A compressor 26 supplies air to a combustion chamber 28 feeding a gas turbine 18 driving the compressor. A screen 12 located in the compressor entry has a series of vertical streamlined tubes 11 arranged on its upstream side to which gases from the turbine exhaust space 16 are led by a passage 14. The tubes 11 are slotted at the back so that the gases impinge directly on the ...

A Gas Turbine Ducted Fan Engine.

... 1,161,186. Gas turbine jet propulion plant; axial flow fans. ROLLS-ROYCE Ltd. May 28, 1968, No.25484/68. Headings F1C, F1G and F1J. In a gas turbine ducted fan engine having a front fan surrounded by an annular casing at least a portion 22 of the casing is formed by two spaced apart walls 20, 24, the latter wall consisting of members 25 each having an integral flange 28 extending to the wall 20 and secured thereto, e.g. by welding. Circumferentially adjacent members 25 overlap at recesses 27 and are welded together. The flanges 28 define axially extending cavities 23. The portion 22 forms part of the radially inner wall of the annular fan casing (16) upstream of the tips of the fan blades (17, Fig. 1, not shown), the outer wall of the casing being formed by a wall 21 welded to the portion 22 via annular members 33, 34 and brackets 35. In operation, hot de-icing air supplied to the downstream end of the casing passes through apertures 36, 37 in the members 33, 34 to impinge on the leading ...

De-icing system for forward parts of a turboshaft engine

Anti-ice system for jet engine inlet housings

A system for circulating heated gases within the circular leading edge of a jet engine housing to prevent ice build-up thereon, or to remove accumulated ice thereform. Hot gases such as air from a hot, high pressure section of the jet-engine are directed through a conduit. The conduit enters the annular leading edge housing, usually from the aft side through a bulkhead, then turns about 90 DEG to a direction tangential to the leading edge annulus. The hot gases exiting the tube entrain the cooler air in the housing, causing a much larger mass of air to swirl circularly around the annular housing. The entering hot gasses heat the mass of air to an intermediate, but still relatively hot, temperature. This large mass of circularly moving hot air is quite efficient in uniformly transferring heat to the skin of the leading edge without leaving any relatively cold areas and preventing the formation of ice thereon.

An apparatus for preventing ice accretion

An apparatus for preventing ice accretion on a component subjected in use to vibration, characterized in that the heat generated within the component 22 by damping of the vibration prevents the build up of ice. The damping may be provided by a viscoelastic coating 34 on gas turbine fan blade 22. The viscous elastic coating 34 may comprise epoxy, polyurethane, and/or polyethylene. A heat conductive face-sheet may cover the viscoelastic material. If ice were to build up on the fan blade, the vibrations would increase, resulting in further vibrations and therefore generation of heat, which would cause the ice to dislodge. It is also disclosed that the damping may be provided by a viscoelastic filler within the blade (see fig.3), a shape memory alloy (SMA) in the blade (see fig.4), or by a magnetic field generated around the fan blade (see fig.7).

Improvements in compressors of gas turbine units

... 636,612. Gas turbine plant. GREATREX, F. B., and MURRAY, F. R. June 20, 1947, Nos. 16413 and 25430. [Classes 110(i) and 110(iii)] To prevent icing hot gases are delivered into the air intake of a. compressor of a gas turbine engine by a distributer which entends radially from an axis about which it rotates in a plane transverse to the air intake. Hot gases from the engine are conveyed through pipes 16, an annular chamber 18 forming the leading edge of the nacelle 11, hollow struts 15 and a chamber 19 into a rotary distributer 21. The distributer 21 comprises a series of hollow, aerofoil section blade members 22 radiating from a hub 23 carried by a shaft 24. The hot gases leave the distributer through perforations in the leading or trailing edges of the blades 22. Swirl vanes 30 may be provided to give the hot gases a swirl in the direction of rotation. In the arrangement shown, the distributer is driven by the windmilling effect of the air entering the compressor, but a positive drive from ...

Systems and methods for aircraft

A thermal management system 100 has an inlet 40 and an outlet 42 of an aircraft nacelle 20 with a gas flow path extending between them and close to components to manage their temperature; the system also has a flow control suitable for controlling a mass flow along the path, preferably controlled as a function of the components temperature. The components may be an electric motor 14 and/or power electronics 22 of an aircraft propeller propulsion system, with the nacelle located in the aircraft wing. The flow control may also control the mass flow leaving the outlet, to mitigate aerodynamic interference caused by the presence of the nacelle. The flow control may have a transducer assembly turbine in the flow path to convert gas flow energy into electrical energy. The system may be arranged to direct external airflow into the inlet. The inlet may have a filter and louvre to reduce particles entering the system.

An engine arrangement

Improvements in or relating to gas turbine engines

... 920,886. Gas turbine Jet propulsion plant. ROLLS-ROYCE Ltd. Oct. 26, 1959 [Nov. 24, 1958], No. 37822/58. Class 110 (3). A gas turbine jet propulsion plant comprising a ducted fan encircling the turbine of the engine which includes rotor blades formed as radial extensions of the turbine rotor blades, has means to deliver air to parts of the turbine for cooling those parts and to deliver the air from these parts to parts of the ducted fan which are liable to icing. The rear part of a ducted fan gas turbine jet propulsion engine, Fig. 2, comprises a combustion chamber 11 feeding gases to rotor blades 17 through nozzle guide vanes 18. The blades 17 are attached to a rotor disc 16 carried by a shaft 15 supported by bearing 14 mounted in an inner casing 13. The ducted fan and its driving turbine comprise a hollow shaft 21 supported by a bearing 22 within shaft 15 and by a main thrust bearing 23. The shaft 21 has a flange 21a to which is secured a rotor disc 24 and has secured to its end a second ...

Limiting spool speeds in a gas turbine engine

A gas turbine engine 101 for an aircraft, comprises a high-pressure (HP) spool comprising an HP compressor and a first electric machine 110 driven by an HP turbine; a low-pressure (LP) spool comprising an LP compressor and a second electric machine driven 110 by an LP turbine; an engine controller 309 configured to identify a condition to the effect that the HP spool has reached or exceeded an HP speed limit whilst the LP spool has not reached an LP speed limit, and to operate the first electric machine in a generator mode of operation to reduce the HP spool speed below the HP speed limit. The engine controller may be further configured to operate the second electric machine in a motor mode of operation and transfer power electrically thereto from the first electric machine. The engine controller may be further configured to transfer power from the first electric machine to an energy storage unit such as a batter 305 or capacitor or to a consumer such as a de-icing system 304.

Systems and methods for aircraft

A thermal management system has a rotating hub 60 of an aircraft propulsion system with an aperture 64 suitable for facilitating airflow into a nacelle 20 to manage the temperature of components. The hub may have a duct extending through it with the aperture opening into the duct. The propulsion system preferably has propellers 12 and the hub is a propeller hub; the nacelle being part of an aircraft wing. An airflow generator (fig.6,50), such as a fan (fig.6,52) driven by the propeller drive shaft (fig.6,54), may be provided to draw air into the nacelle through the aperture. A fairing or nose cone 62 may be provided over the hub, also provided with an aperture and duct. The fairing may be 3D printed and/or cast with channels 66 formed in the fairing duct; when the fairing is rotated the channels, in the form of an axial compressor, draw air into the nacelle, increasing the pressure of the airflow. The components may be an electric motor 14 and/or power electronics. The system may provide ...

Systems, arrangements, structures and methods for aircraft

An electrical power management system 508 for an aircraft, the aircraft comprising a plurality of electrical power sources 512 for powering components 514 in the aircraft, the power management system 508 comprising a plurality of power management modules 510 arranged to manage electrical power from the electrical power sources 512 to the components 514, each of the plurality of power management modules 510 being arranged to manage power from at least one of the electrical power sources 512. The electrical power sources 512 may include turbogenerators 506, a battery 14, a photovoltaic panel or an energy recovery system 30 including turbines. Each power management module 510 may be dedicated to one electrical power source 512 and may function as a backup, non-dedicated power management module 510 for another power source 512 in the event of a failure of another power management module 510. The electrical power management system 508 may manage the electrical power sources 512 dependent upon ...

Dual valve apparatus for aircraft engine ice protection and related methods

Aircraft Structure

AN AIRCRAFT ENGINE NOSE COWL

... 1,247,071. Nose cowl. ROLLS ROYCE Ltd. 24 Jan., 1969, No. 4224/69. Heading B7G. [Also in Divisions F1 and H5] A nose cowl for an aeroengine comprises two portions 16 and 14, each consisting of a support plate 21 and end diaphragms 22, which are releasably secured together by captive bolts, each portion having electrical de-icing means consisting of several layers of glass cloth 18 sandwiching an electrical heating element 19. The joint between the portions 16 and 14 consist of two plates 26 having flanges 34 of such thickness that they can be sufficiently heated by elements 19 to prevent ice formation in the region of the joint, said flanges 34 having further flanges 38 securing the de-icing mats 18, 19 to support plate 21, the surface of the cowl being covered with a layer ferrosianresistant material. The heating elements are designed so that the portions 16 and 14 fit together and form a complete circuit, or so that each heating element is connected independently to a power source. The ...

Improvements in or relating to gas-turbine engine cooling arrangements

... 790,836. Gas-turbine plant. ROLLS-ROYCE, Ltd. Jan. 7, 1955 [Jan. 25, 1954], No. 229/57. Divided out of 790,834. Drawings to Specification. Class 110 (3). A gas-turbine engine having a compressor and turbine connected by a rotor shaft has means to supply cooling or sealing air to the internal components of the engine comprising an expansion turbine, drivingly connected to the rotor shaft, which supplies air to a hollow formation in the rotor shaft which acts as a part of a distributing duct for the cooling or sealing air. The major part of the subject-matter of this Specification is identical with that described with reference to Figs. 1, 2 and 4 in Specification 790,834, but the claims differ. The remainder of the Specification relates to a modification of these arrangements in which the inlet of the expansion turbine may be connected to the delivery of a stage of the compressor of the engine. In a preferred arrangement, the inlet of the expansion turbine may be connected through ducts ...

Improvements in or relating to gas turbine power plant arrangements

... 767,177. Gas turbine plant. ROLLS-ROYCE, Ltd. Oct. 5, 1953 [Oct. 10, 1952], No. 1154/56. Divided out of 767,176. Class 110 (3). A gas turbine power plant comprising means for tapping air heated by compression, heating means to produce from the power plant hot gas at a temperature above the temperature of the air abstracted from the compressor and a conduit through which the air and hot gas are led to de-icing means has valve means to control the individual flows of air and hot gas controlled by temperature-sensitive means responsive only to the temperature of the heated gas and a relief valve located in the delivery conduit preset to a selected pressure so as to control the quantity of heat transmitted to the parts to be de-iced. Compressed air is bled from the compressor 10 supplying combustion air to the combustion chambers 12 of a gas turbine plant through a port 22 and a conduit 23 controlled by a butterfly valve 24. Combustion products are bled from the combustion chambers 12 through ...

DE-ICING OF THE AIR INTAKES OF GAS TURBINE ENGINES

Systems, arrangements, structures and methods for aircraft

SPINNER OR NOSE BULLET

SNOW AND ICE REMOVAL SYSTEM FUR THE ANSTRÍMKANTE OF AN INLET HOOD OF A TURBINE ENGINE

ENGINE POD WITH ENTRANCE OPENING TO THE SNOW AND ICE REMOVAL SYSTEM

PROCEDURE FOR MANUFACTURING A HEIZKÍRPERSTRUKTUR AND HEIZKÍRPERSTRUKTUR

SOUND PROOFING LINING FOR AN AIRPLANE WITH FROST TREATMENT SYSTEM WITH JOULE MAGNETOSTRICTION

AIR SUCKING IN STRUCTURE FOR AN AIRPLANE ENGINE POD

Methods and apparatus for operating gas turbine engines

NACELLE INLET LIP ANTI-ICING WITH ENGINE OIL

A nacelle for housing a gas turbine engine having a pressurized oil system for lubricating components thereof comprises an inlet lip defining a leading edge of the nacelle, the inlet lip having a conduit therein in fluid flow communication with the pressurized oil system of the gas turbine engine and defining an oil passage for circulation of pressurized engine oil therethrough. The conduit is in heat transfer communication with an outer surface of the inlet lip.

COATED VARIABLE AREA FAN NOZZLE

A variable area fan nozzle for use with a gas turbine engine system includes a nozzle section that is movable between a plurality of positions to change an effective area associated with a bypass airflow through a fan bypass passage of a gas turbine engine. A protective coating is disposed on the nozzle section and resists change in the effective area of the nozzle section caused by environmental conditions.

ICE DETECTION SYSTEM FOR A GAS TURBINE

TURBINE ENGINE INLET CONE DEICING SYSTEM FOR AIRCRAFT

L'invention se rapporte à un système de dégivrage (2) d'un cône d'entrée (4) de turbomoteur pour aéronef, comprenant des moyens de diffusion d'air (18) destinés à équiper le cône d'entrée du turbomoteur afin de lui délivrer de l'air chaud. Selon l'invention, il comporte également un circuit (20) d'évacuation de l'air de pressurisation d'au moins une enceinte-palier du turbomoteur, ce circuit communiquant avec les moyens de diffusion d'air pour pouvoir alimenter ces derniers en air chaud.

METHODS AND APPARATUS FOR GAS TURBINE ENGINES

A gas turbine engine component 50 that includes a first side (64), an opposite second side (66), at least one capillary (84) positioned adjacent to an external surface (88) of at least one of said first and second sides, and a composite layer (90) securing the at least one said capillary to said component.

ROTATING INLET COWL FOR A TURBINE ENGINE, INCLUDING AN ECCENTRIC FRONT END

La présente invention se rapporte à un capot d'entrée tournant (30) pour tarbomachine, présentant un axe de rotation (34) et dont l'extrémité avant (44) est agencée de manière excentrée par rapport à cet axe de rotation (34). De plus, un cône avant (32) du capot est tronqué par une surface de troncature (70) définissant l'extrémité avant (44) du capot d'entrée.

CAPOT D'ENTREE NON TOURNANT DE TURBOREACTEUR A FIXATION CENTRALE ET TURBOREACTEUR AINSI EQUIPE

CAPOT D'ENTREE NON TOURNANT DE TURBOREACTEUR A FIXATION CENTRALE ET TURBOREACTEUR AINSI EQUIPE Le capot d'entrée comprend un cône extérieur (15) fixé sur la structure fixe (3-4) du turboréacteur par des moyens coopérants de rebord (18) de l'embase (16) dudit cône et de rebord amont (13) d'une bride axiale (11) de la face avant (7) de ladite structure fixe par des découpes (18a, 13a) à baïonnette et un cône intérieur (21) dont le rebord (23) de l'embase (22) coopère par des festons (23a) avec des éléments (6) de ladite structure fixe (4). Les cônes intérieure (21) et extérieur (15) sont solidarisés à l'amont par une fixation centrale comportant une pièce de verrouillage (27). Un montage est décrit sur un boîtier (4) de palier avant de turboréacteur, en adjoignant un circuit de dégivrage (34 à 38). FIGURE 1 ...

INLET FRAME FOR A GAS TURBINE ENGINE

An inlet frame (54) and a method (200) of additively manufacturing the same are provided. The inlet frame (54) includes a forward annular body (102) spaced apart from a rear annular body (104) to define an inlet passageway (106) in fluid communication with a compressor inlet (108). The inlet frame (54) may define integral wash manifolds (130) and discharge ports (140) for directing a flow of wash fluid directly through the compressor inlet (108). In addition, inlet frame (54) may define one or more integral annular heating plenums (84), (186) in fluid communication with a hot air source (150) for heating regions of the inlet frame (54) that are prone to icing conditions.

ADVANCED INLET DESIGN

A compact inlet design including a single bulkhead and/or an acoustic panel extending into nacelle lip region for noise reduction. The compact inlet is used with a low power fluid ice protection system capable of preventing ice build-up on the acoustic panel in the nacelle lip region.

DEVICE FOR DEICING AN AERONAUTICAL TURBOMACHINE SEPARATOR

L'invention concerne un dispositif de dégivrage d'un bec de séparation de turbomachine aéronautique, comprenant un bec de séparation (20) destiné à être positionné à l'aval d'une soufflante de la turbomachine pour former une séparation entre des canaux annulaires d'écoulement d'un flux primaire (16) et d'un flux secondaire issus de la turbomachine, et un carter (28) fixé au bec de séparation dans le prolongement aval de celui- ci, le carter ayant une virole interne (30) délimitant à l'extérieur le canal d'écoulement du flux primaire et comprenant au moins un conduit d'air (36) intégré à la virole interne pour former une seule et même pièce avec celle-ci, le conduit d'air s'ouvrant en aval vers une alimentation en air (38) et débouchant à l'amont à l'intérieur du bec de séparation.

COOLING DEVICE FOR A TURBO ENGINE OF AN AIRCRAFT NACELLE

L'invention concerne un dispositif (12) de refroidissement pour un turbomoteur d'une nacelle (10) d'aéronef, comportant un échangeur (26) thermique et un conduit de sortie (30) d'air, et la nacelle (10) comportant un carénage (14) avant tubulaire qui comporte une lèvre (20) avant formant bord d'attaque creux qui délimite une chambre (22) annulaire de dégivrage, caractérisé en ce qu'il comporte un conduit d'alimentation en air sous pression (48) qui s'étend depuis une extrémité d'entrée reliée sur une source d'air sous pression, jusqu'à une extrémité de sortie formant buse d'éjection (50) d'air débouchant dans la chambre (22) de dégivrage, et en ce que le conduit de sortie (30) de l'échangeur (26) thermique présente un tronçon de sortie d'air (38) qui est agencé dans la chambre (22) de dégivrage dans une position adaptée pour que la buse d'éjection (50) d'air sous pression forme pompe d'aspiration de l'air dans le conduit de sortie (30) de l'échangeur.

DE-ICING AND CONDITIONING DEVICE FOR AN AIRCRAFT

L'invention concerne un dispositif de dégivrage d'une lèvre d'entrée d'air (111) d'une nacelle (100) d'aéronef, le dit dispositif comprenant un pré- échangeur (2), un moyen de prélèvement apte à prélever de l'air basse pression en aval de la soufflante (12),deux moyens de prélèvement d'air haute pression en aval du compresseur (10, 11) ainsi que des vannes commandées (6, 7, 8, 9, 14) et des vannes anti-retour (4, 5) installées dans un réseau (1, 13) de circulation d'air remarquable en ce que le pré-échangeur(2) comprend une sortie d'air basse pression (19) apte à déboucher dans la lèvre d'entrée d'air de la nacelle (100) de l'aéronef via une canalisation (3) du réseau de circulation d'air.

DE-ICING SPLITTER LIP FOR AXIAL TURBOMACHINE COMPRESSOR

A de-icing splitter lip for a low-pressure compressor of a turbofan aircraft engine surrounds the primary flow. It has an annular splitter wall with a circular leading edge, an outer shroud connected to the splitter wall, heating means in an electric ribbon that de-ices the splitter lip. The splitter lip further includes an elastic element made of elastomer, holding the heating means on the inside of the splitter wall. The elastic element is compressed, pre-loaded, in the splitter lip. Thus, it exerts a force F clamping the heating means against the splitter wall and the outer shroud in order to help improve thermal contact. The invention also provides a method for assembling a splitter lip.

COMPOSITE SPLITTER LIP FOR AXIAL TURBOMACHINE COMPRESSOR

A splitter lip for a low-pressure compressor of an axial turbomachine for an aeroplane has an upstream annular wall made of metal with a circular leading edge, and a downstream annular partition made of an organic-matrix, short-fibre composite material. The splitter lip also supports an outer shroud for a stator upstream of the compressor. The upstream wall includes an annular anchoring portion arranged in the thickness of the downstream partition so as to anchor the partition and the wall to one another. The anchoring portion has distributed hexagonal openings in order to increase anchoring. A method for producing a bi-material splitter lip or a mixed lip includes a moulding step.

AIRCRAFT ANTI-ICING SYSTEMS HAVING DEFLECTOR VANES

Apparatus and methods to lower peak temperatures and improve performance of aircraft anti-icing systems are described herein. One described example apparatus includes a skin on an inlet side of a nacelle of an aircraft defining an annular chamber, a gas delivery system disposed within the annular chamber to provide a first gas to mix with and entrain a second gas in the annular chamber to define a flow through the annular chamber, and a deflector vane disposed within the annular chamber to redirect the flow through the annular chamber. The second gas is at a different temperature from the first gas.

Einrichtung zur Verhinderung der Eisbildung am Lufteinlass von Kompressoren an Flugzeuggasturbinen-Triebwerken.

Gasturbinen-Antriebsanlage.

Abgasleitung an einer Gasturbinenanlage.

Installation comprenant une turbine à combustion interne.

Luftfahrzeug-Gasturbinen-Kraftanlage.

Verfahren zur Verhinderung der Eisbildung in Axialkompressoren sowie Einrichtung zur Ausübung des Verfahrens.

Installation de propulsion à turbine à combustion interne.

Installation à turbine à gaz

Installation de propulsion à turbine à gaz pour un aéronef

Flugzeug-Gasturbinenanlage

Vorrichtung an Flugzeugen zur Erzeugung von Druckluft zu Hilfszwecken.

Einrichtung zum Absondern von heissem Gas aus dem Auspuffkanal eines Gasturbinentriebwerkes.

Rotationsverdichter.

Gasturbinentriebwerk.

Heizvorrichtung für die Enteisung von Streben an Gasturbinen-Triebwerken

Hohle Leitschaufel für Turbomaschinen.

Gasturbinenanlage.

Aube chauffable à l'électricité, et procédé de fabrication de cette aube.

Gasturbinenanlage

Dispositif pour pulvériser un liquide dans le courant d'air pénétrant dans le compresseur d'une installation de turbine à gaz

Einrichtung zur Vorwärmung der Ansaugluft bei einer Gasturbinenanlage mit Abgaskamin

Enteisungseinrichtung, insbesondere für die Einlassleitschaufeln am Verdichter eines Strahltriebwerkes

Systems and methods for using gas turbine areaventilation exhaust air from a.

Systeme (100) und Verfahren zur Verwendung von Ventilationsabluft aus einem Gasturbinenraum (126). Das System (100) umfasst eine Gasturbine mit einem Verdichter (104). Das System (100) umfasst ausserdem den Gasturbinenraum (126), der um die Gasturbine herum angeordnet ist. Darüber hinaus umfasst das System (100) einen Einlassabzapfwärme(IBH)-Verteiler (118), der in Strömungsverbindung mit dem Verdichter steht. Der Gasturbinenraum (126) steht in Strömungsverbindung mit dem IBH-Verteiler (118), um dem IBH-Verteiler (118) Ventilationsabluft aus dem Gasturbinenraum (126) zuzuleiten.

Systems and method for de-icing a gas turbine and for dehumidifying air intake filtersinlet filter a.

Es wird ein System zum Enteisen einer Gasturbine (101) geschaffen. Eine Zweigleitung ist mit einem Einlasssieb (135) verbunden. Eine erste Leitung ist mit einer Stufe des Verdichters (105) und einem ersten Eingang der Mischerkomponente (185) verbunden. Eine zweite Leitung ist mit dem Abgasauslass (120) und einem zweiten Eingang der Mischerkomponente (185) verbunden. Die zweite Leitung ist dafür ausgelegt, Abgase zu extrahieren, ohne den Druck am Abgasauslass zu erhöhen. Eine dritte Leitung (190) ist mit dem Auslass der Mischerkomponente (185) und der Zweigleitung verbunden. Ein Verfahren zum Enteisen eines Gasturbinen-Einlasssiebs (135) beinhaltet das Bestimmen einer Ist-Temperatur am Einlasssieb (135) und das Bestimmen einer Soll-Temperatur am Einlasssieb. Wenn die Ist-Temperatur am Einlasssieb (135) niedriger ist als die Soll-Temperatur am Einlasssieb (135), wird eine erste Strömungsrate einer Luft-Abgas-Mischung berechnet, die nötig ist, um die Soll-Einlasssiebtemperatur zu erhalten.

Intake system for a gas turbine, power plant with such and method for the conditioning of inlet air.

Ein Einlasssystem für eine Gasturbine (10) enthält: einen Einlassluftkanal (70); einen in dem Einlassluftkanal (70) angeordneten Lärmminderer (90), wobei der Lärmminderer (90) mehrere Paneele mit Zwischenräumen zwischen den Paneelen enthält; und eine Leitung (100, 105) mit Öffnungen, die zur Injektion von Einlasszapfluftwärme in jeden von den Zwischenräumen angeordnet sind. Ein Verfahren zum Konditionieren von Einlassluft für eine Gasturbine (10) enthält ein Strömenlassen von Luft durch Zwischenräume zwischen Paneelen eines Lärmminderers (90) in einem Einlassluftkanal (70) der Gasturbine (10) und Injizieren von Einlasszapfluftwärme durch Öffnungen hindurch und in jeden von den Zwischenräumen hinein ...

ANTI-ICING DEVICE FOR GAS TURBINE.

Methods and systems for removing iron from inlet lattices and for dehumidifying inlet air cleaners for gas turbines.

Es sind hierin Verfahren und Systeme zum Enteisen eines Einlassgitters und zum Entfeuchten eines Einlassluftfilters in einer Gasturbine offenbart. Das Verfahren enthält ein Bestimmen (402) einer momentanen Einlassgittertemperatur. Das Verfahren enthält ein Bestimmen (404) einer gewünschten Einlassgittertemperatur. Falls die momentane Einlassgittertemperatur geringer ist als die gewünschte Einlassgittertemperatur, kann das Verfahren ferner ein Bestimmen (408) einer ersten Menge an Gasturbinenraumbelüftungsabluft, die zum Erreichen der gewünschten Einlassgittertemperatur erforderlich ist, Entnehmen (412) der ersten Menge der Gasturbinenraumbelüftungsabluft und Fördern (414) der ersten Menge der Gasturbinenraumbelüftungsabluft zu dem Einlassgitter enthalten.

ЛОПАТКА ВХОДНОГО НАПРАВЛЯЮЩЕГО КОЛЕСА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ, ОБОРУДОВАННАЯ ПРОТИВООБЛЕДЕНИТЕЛЕМ (ВАРИАНТЫ), И ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ ЛЕТАТЕЛЬНОГО АППАРАТА

Лопатка (10) входного направляющего колеса (4) газотурбинного двигателя (2) имеет противообледенительные средства, неподвижную часть (12), расположенную спереди, и соединенный с ней подвижный щиток (18), расположенный сзади. При этом указанная неподвижная часть (12) имеет полый корпус (13) и заднюю кромку (16), и профиль указанной неподвижной части (12) имеет форму, в целом подобную букве "U", ветви которой (162, 164) направлены в целом в направлении выхода. При этом со стороны внутреннего изгиба (І) расположена ветвь (162), а со стороны спинки (Е) расположена ветвь (164). Противообледенительные средства включают, по меньшей мере, одно выпускное окно (202), в целом ориентированное в направлении "вход-выход" и расположенное только вдоль той ветви (162) не подвижной части, в целом подобной букве "U", которая находится со стороны внутреннего изгиба (І) неподвижной части (12). Выпускное окно (202) образовано гофрированным листом (152), который примыкает к указанной ветви (162) неподвижной части ...

Method for producing an acoustic treatment panel including the function of de-icing with hot air

Nacelle for an aircraft engine

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

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

Fluid impingement arrangement

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

Method and apparatus for protecting aircraft engines against icing

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

Rigid raft

The present invention provides a rigid raft formed of rigid composite material. The raft has an electrical system and/or a fluid system embedded therein. The raft further has a tank for containing liquid integrally formed therewith. The tank can be formed of the rigid composite material. The tank can be for a gas turbine engine.

Gas turbine engine part

The present invention provides a gas turbine engine part which has a primary purpose in the engine which is structural and/or aerodynamic. The part is formed of rigid composite material, and has an electrical system comprising electrical conductors permanently embedded in the composite material. This provides advantages in terms of weight, complexity, and build time.

Buffer system that communicates buffer supply air to one or more portions of a gas turbine engine

A gas turbine engine includes a buffer system that can communicate a buffer supply air to a portion of the gas turbine engine. The buffer system includes a first bleed air supply having a first pressure, a second bleed air supply having a second pressure that is greater than the first pressure, and an ejector that selectively augments the first bleed air supply to prepare the buffer supply air for communication to the portion of the gas turbine engine.

Combustion turbine inlet anti-icing resistive heating system

A resistive heating system for a combustion turbine susceptible to inlet air filter house component and compressor icing includes a plurality of heating panels (bundles) arranged in a substantially-planar array, adapted to be located on or adjacent to the turbine's inlet air filter house. Each heating panel is provided with one or more electrically-resistive heating elements; and a controller for selectively activating the resistive heating elements on each of the plurality of heating panels.

System and method for gas turbine inlet air heating

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

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

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

GAS TURBINE INTAKE ANTI-ICING DEVICE

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

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

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

Heated Screen For Air Intake Of Aircraft Engines

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

Dual pressure deicing system

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

Heating System for Backup Generator

A heating system for a backup generator for ensuring that a generator unit does not malfunction due to low ambient temperatures. The heating system includes a heater adapted to be installed within the generator unit housing and adjacent to a generator for use in providing heat to the generator. A heating pad is further provided, wherein a battery of the backup generator is adapted to be positioned on the heating pad. The heater and heating pad are operably connected to a thermostat that determines the temperature within the generator unit. The heater and heating pad are adapted to begin producing heat if the thermostat detects that the temperature has dropped below a predetermined temperature minimum. A high limit switch is further provided to turn off the heater and heating pad when the temperature within the generator housing has reached a predetermined temperature maximum. 1. A heating system for a backup generator , comprising:a heater configured to provide heat to a generator when placed adjacent thereto;a heating pad configured to provide heat to a battery when placed adjacent thereto;a thermostat operably connected to the heater and the heating pad, wherein the thermostat is adapted to detect a temperature of the generator and a temperature of the battery;wherein the thermostat is configured to activate the heater if the temperature of the generator falls below a predetermined minimum temperature;wherein the thermostat is configured to activate the heating pad if the temperature of the battery falls below a predetermined minimum temperature.2. The heating system for a backup generator of claim 1 , whereinthe thermostat is adapted to deactivate the heater if the temperature of the generator exceeds a predetermined maximum temperature.3. The heating system for a backup generator of claim 1 , whereinthe thermostat is adapted to deactivate the heating pad if the temperature of the battery exceeds a predetermined maximum temperature.4. The heating system for a ...

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

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 ...

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

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

METHOD OF REGULATING AIR PRESSURE IN ANTI-ICING SYSTEM

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

INLET BLEED HEAT SYSTEM FOR USE IN A TURBINE ENGINE

An inlet bleed heat (IBH) system for use in a turbine engine including a silencer assembly. The inlet bleed heat (IBH) system includes a feed pipe for delivering compressor discharge air. The feed pipe includes a plurality of orifices along at least a portion of a length of the feed pipe, and each orifice of the plurality of orifices extends through a wall of the feed pipe for allowing the compressor discharge air to exit the feed pipe. The system also includes a heat shielding component that extends across the feed pipe, wherein the heat shielding component is configured to reduce heat transfer between the feed pipe and the silencer assembly of the turbine engine. 1. An inlet bleed heat (IBH) system for use in a turbine engine including a silencer assembly , the inlet bleed heat (IBH) system comprising:a feed pipe for delivering compressor discharge air, said feed pipe comprising a plurality of orifices along at least a portion of a length of said feed pipe, each orifice of said plurality of orifices extending through a wall of said feed pipe for allowing the compressor discharge air to exit said feed pipe; anda heat shielding component that extends across said feed pipe, said heat shielding component configured to reduce heat transfer between said feed pipe and the silencer assembly of the turbine engine.2. The IBH system in accordance with claim 1 , wherein said heat shielding component comprises an arcuate sheet component extending about a portion of said feed pipe claim 1 , and positioned between said feed pipe and the silencer assembly.3. The IBH system in accordance with claim 2 , wherein said arcuate sheet component comprises aluminum or stainless steel.4. The IBH system in accordance with claim 1 , wherein said heat shielding component comprises a layer of thermal insulating material extending about a portion of said feed pipe claim 1 , and positioned between said feed pipe and the silencer assembly.5. The IBH system in accordance with claim 4 , wherein ...

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

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

JET ENGINE ANTI-ICING AND NOISE-ATTENUATING AIR INLETS

An embodiment of a jet engine air inlet providing anti-icing, and optionally, engine noise reduction, includes a rigid frame defining a gridwork of contiguous cells. A cap skin has an outer surface sealingly attached to an inner surface of the frame. An outer skin has an inner surface sealingly attached to an outer surface of the frame and contains a plurality of openings therein. 1. A jet engine air inlet , comprising:a rigid frame defining a gridwork of contiguous cells;a cap skin having an outer surface sealingly attached to an inner surface of the frame;an outer skin having an inner surface sealingly attached to an outer surface of the frame and containing a plurality of openings therein, each opening being disposed in fluid communication with a corresponding one of the cells;a serpentine manifold extending adjacent to an inner end of each of the cells and containing a plurality of apertures in a sidewall thereof, each aperture being disposed in fluid communication with a corresponding one of the cells; anda plenum sealingly covering an inner surface of the cap skin.2. The inlet of claim 1 , wherein the cap skin comprises a first skin containing an open channel corresponding to the manifold and a second skin sealingly disposed on the first skin and closing the channel.3. The inlet of claim 1 , wherein the manifold comprises a length of tubing corresponding to the manifold and disposed on the inner or the outer surface of the cap skin.4. The inlet of wherein the plenum comprises a portion of a D-duct of a jet engine air inlet and the outer skin comprises a lip skin of the inlet.5. The inlet of claim 1 , wherein at least one of the cells comprises a Helmholtz resonator.6. The inlet of claim 1 , further comprising:a first source of air configured to introduce air at a first pressure and a first flow rate into the manifold; anda second source of air configured to introduce air at a second pressure and a second flow rate into the plenum,wherein the first and second ...

NACELLE ANTI-ICE SYSTEM AND METHOD WITH EQUALIZED FLOW

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

NACELLE ANTI ICE SYSTEM

An anti-icing system of a nacelle inlet of an engine of an aircraft includes first and second direct acting valves and first and second control valve assemblies fluidly connected to the nacelle inlet. The first direct acting valve includes a first inlet, outlet, valve chamber, and piston. The first piston is positioned in the first direct acting valve. The first control valve assembly is fluidly connected to the first valve. The second direct acting valve includes a second inlet, outlet, valve chamber, and piston. The second piston is positioned in the second direct acting valve. The second direct acting valve is fluidly connected to the first direct acting valve in a series configuration. The second control valve assembly is fluidly connected to the second valve chamber. 1. An anti-icing system of a nacelle inlet of an engine of an aircraft , wherein the anti-icing system comprises: [ a first inlet;', 'a first valve chamber fluidly connected to the first inlet;', 'a first internal valve body circumferentially surrounding the first valve chamber;', 'a first outlet; and', 'a first piston for adjusting a rate of flow of air through the first direct acting valve, wherein the first piston is slidably engaged with the first internal valve body;, 'a first direct acting valve comprising, 'a first control valve assembly fluidly connected to the first valve chamber of the first direct acting valve;', a second inlet;', 'a second valve chamber fluidly connected to the second inlet;', 'a second internal valve body circumferentially surrounding the second valve chamber;', 'a second outlet; and', 'a second piston for adjusting a rate of flow of air through the second direct acting valve, wherein the second piston is slidably engaged with the second internal valve body, and, 'a second direct acting valve comprising], 'a valve assembly fluidly connected to the nacelle inlet, wherein the valve assembly comprises 'a second control valve assembly fluidly connected to the second valve ...

SYSTEM OF OPERATING A GAS TURBINE ENGINE

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

GAS TURBINE IMPELLER NOSE CONE

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

ENGINE ENCLOSURE AIR INLET SECTION

An air inlet section for an enclosure for an aircraft engine is provided that includes an inner barrel panel, an outer barrel panel, a lipskin and a forward bulkhead. The lipskin extends between an inner barrel end and an outer barrel end. The inner barrel end is disposed proximate the forward end of the inner barrel panel and the outer barrel end is disposed proximate the forward end of the outer barrel panel. The forward bulkhead has a panel that extends between an outer radial end and an inner radial end. The inner barrel panel, the outer barrel panel, and the lipskin define an interior annular region, and the forward bulkhead defines a sub-portion of interior annular region. The outer radial end of the forward bulkhead panel is disposed forward of the inner radial end of the forward bulkhead panel. 1. An air inlet section for an enclosure for an aircraft engine , comprising:an inner barrel panel having a forward end;an outer barrel panel having a forward end;a lip skin having an interior surface and an exterior surface, wherein the interior surface and the exterior surface oppose one another, the lipskin extending between an inner barrel end and an outer barrel end, wherein the inner barrel end is disposed proximate the forward end of the inner barrel panel and the outer barrel end is disposed proximate the forward end of the outer barrel panel; anda forward bulkhead having a panel that extends between an outer radial end and an inner radial end;wherein the air inlet section is configured as an annular structure that extends circumferentially around an axially extending centerline, with the inner barrel panel disposed radially inside of and separated from the outer barrel panel, and the inner barrel panel, the outer barrel panel, and the lipskin define an interior annular region, and the forward bulkhead defines a sub-portion of interior annular region;wherein the outer radial end of the forward bulkhead panel is disposed forward of the inner radial end of the ...

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

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

SEGREGATED ANTI-ICE DUCT CHAMBER

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

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

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

GAS TURBINE ENGINE DE-ICING SYSTEM

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

ENGINE AIR INLET HAVING A DOUBLE-PANEL HEATED WALL

An annular air inlet duct circumscribing an axis of rotation of a rotatable member of a machine comprising a forward end and an aft end is described. The air inlet duct includes an unheated wall and a heated wall adjacent to a heat source. The heated wall includes a plurality of axially-spaced wall panels forming a cavity between each of a pair of adjacent wall panels of the plurality of axially-spaced wall panels. 1. An annular air inlet duct circumscribing an axis of rotation of a rotatable member of a machine comprising a forward end and an aft end , said air inlet duct comprising:an unheated wall; anda heated wall adjacent to a heat source, said heated wall comprising a plurality of axially-spaced wall panels forming a cavity between each of a pair of adjacent wall panels of the plurality of axially-spaced wall panels.2. The annular air inlet duct of claim 1 , wherein each said cavity includes at least one of an evacuated space claim 1 , a purged space claim 1 , and a space containing an insulative material.3. The annular air inlet duct of claim 1 , wherein said unheated wall and said heated wall form a radially outer inlet mouth and a radially inner converging throat.4. The annular air inlet duct of claim 3 , wherein said radially outer inlet mouth is configured to receive a radially inward flow of air claim 3 , and wherein said radially inner converging throat is configured to turn the radially inward flow of air in an axially forward direction.5. The annular air inlet duct of claim 1 , wherein said annular air inlet duct is formed of a sintered metal material.6. The annular air inlet duct of claim 1 , wherein the heat source includes gas turbine engine components.7. The annular air inlet duct of claim 6 , wherein the gas turbine engine components include at least one of an oil tank claim 6 , an accessory gear box claim 6 , a fuel pump claim 6 , an oil pump claim 6 , and an electrical generator.8. The annular air inlet duct of claim 1 , wherein at least one of ...

EXHAUST HEAT RECOVERY FOR A GAS TURBINE SYSTEM

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

ICE PROTECTION SYSTEM FOR A COMPONENT OF AN AERODYNAMIC SYSTEM

Disclosed is an ice protection system for an aerodynamic surface of an aircraft, a surface having a flow facing side and an inwardly facing side that opposes the flow facing side, the system having: a perforated sheet configured for disposal in the surface; a heating source connected to the perforated sheet; and a suction source disposed to draw ice melted by the heating source through the perforated sheet and heating source. 1. An ice protection system for an aerodynamic surface of an aircraft , a surface having a flow facing side and an inwardly facing side that opposes the flow facing side , the system comprising:a perforated sheet configured for disposal in the surface;a heating source connected to the perforated sheet; anda suction source disposed to draw ice melted by the heating source through the perforated sheet and heating source.2. The component of claim 1 , wherein the heating source is integral with the perforated sheet.3. The component of claim 2 , wherein the heating source includes a drain hole extending between the perforated sheet and the suction source.4. The component of claim 3 , wherein the ice protection system comprises a honeycomb support structure on an inward side of the perforated sheet.5. The component of claim 4 , wherein the ice protection system comprises a water collector on an inwardly facing side of the honeycomb support structure.6. The component of claim 5 , wherein the suction source is a pump fluidly connected to the water collector.7. The component of claim 6 , wherein the ice protection system includes a rigid shell on an inward facing side of the water collector.8. The component of claim 7 , wherein the ice protection system is coupled to a leading edge thereof.9. The component of claim 8 , wherein the component is a nacelle or an aircraft control surface on a wing and empennage.10. The component of claim 8 , wherein the component is a wing and empennage.11. A method of preventing ice formation with an ice protection system ...

Nacelle air intake and nacelle comprising such an air intake

An air intake includes a substantially cylindrical inner wall, a substantially cylindrical outer wall, a front lip connecting the inner wall and the outer wall, a front mounting flange, and a support structure. The front mounting flange is configured to cooperate with a rear flange of a wall of an aircraft engine. The support structure is configured to be secured to the wall of the aircraft engine at a location longitudinally downstream of the mounting flange. The outer wall includes a downstream end configured to be positioned in a junction area flush with a front end of a fan external cowl. A portion of the outer wall being configured to bear at least against the support structure. The support structure is configured to be secured to the wall of the aircraft engine so that a load path passes directly from the outer wall towards the fan casing.

DEVICE FOR DEICING A SPLITTER NOSE AND INLET GUIDE VANES OF AN AVIATION TURBINE ENGINE

A deicer device deices both a splitter nose and inlet guide vanes of an aviation turbine engine. The device includes a splitter nose to split a stream coming from the fan into primary and secondary stream flow channels of annular shape and an inner shroud to which inlet guide vanes are fastened and including a hook that holds the inner shroud axially at its upstream end against the inner annular wall. An outer surface facing of the hook forms an angle of less than 90° with an injection orifice such that the outer surface comes progressively closer to the outer annular wall on going away from the injection orifice. The outer surface of the hook presents a minimum amount of clearance J relative to the outer annular wall such that 0.2≤J/D≤0.6 where D is the hydraulic diameter of an injection orifice. 1. An aviation turbine engine fan module comprising:a fan,a low-pressure compressor,inlet guide vanes situated upstream from the low-pressure compressor and downstream from the fan, and a splitter nose for positioning downstream from the fan of the turbine engine in order to split a stream coming from the fan into primary and secondary stream flow channels of annular shape, said nose having an outer annular wall defining the inside of the secondary stream flow channel and an inner annular wall defining an inlet of the primary stream flow channel, said inner annular wall being provided with injection orifices positioned upstream from the inlet guide vanes and through which hot air is to be blown; and', 'an inner shroud to which the inlet guide vanes are fastened and including a hook that is axisymmetric about a longitudinal axis of the turbine engine, said inner shroud being held axially at its upstream end against the inner annular wall by said hook,, 'a deicer device for deicing both a splitter nose and also inlet guide vanes of the turbine engine, said deicer device comprising {'br': None, '0.2≤J/D≤0.6'}, 'wherein the hook has an outer surface facing the outer annular ...

GAS TURBINE ENGINE

A gas turbine engine comprising: an engine core comprising a compressor; a compressor bleed valve in communication with the compressor and configured to release bleed air from the compressor; at least one component provided at the inlet of the engine core having a de-icing conduit, configured to receive the bleed air; and a flow controller, configured to provide bleed air to the de-icing conduit of the at least one component in response to either or both of a requirement to de-ice the component and a requirement to release bleed air from the compressor to optimise operation of the core. 1. A gas turbine engine comprising:an engine core comprising a compressor;a compressor bleed valve in communication with the compressor and configured to release bleed air from the compressor;at least one component provided at the inlet of the engine core having a de-icing conduit, configured to receive the bleed air; anda flow controller, configured to provide bleed air to the de-icing conduit of the at least one component in response to either or both of a requirement to de-ice the component and a requirement to release bleed air from the compressor to optimise operation of the core;wherein the flow controller is configured to provide a first mass flow rate of bleed air when the bleed air is provided in response to a requirement to de-ice the component and a second mass flow rate of bleed air, different from the first mass flow rate, when the bleed air is provided in response to a requirement to release bleed air from the compressor to optimise operation of the core, and to provide the higher of the first and second mass flow rates of bleed air when the bleed air is required both for de-icing the component and to optimise operation of the core.2. The gas turbine engine of claim 1 , further comprising an icing detector configured to detect a level of icing on the at least one component; wherein the flow controller is configured to provide a mass flow rate of bleed air to the de- ...

Inlet bleed heat system for use in a turbine engine

An inlet bleed heat (IBH) system for use in a turbine engine including a silencer assembly. The inlet bleed heat (IBH) system includes a feed pipe for delivering compressor discharge air. The feed pipe includes a plurality of orifices along at least a portion of a length of the feed pipe, and each orifice of the plurality of orifices extends through a wall of the feed pipe for allowing the compressor discharge air to exit the feed pipe. The system also includes a heat shielding component that extends across the feed pipe, wherein the heat shielding component is configured to reduce heat transfer between the feed pipe and the silencer assembly of the turbine engine.

Nacelle for an aircraft bypass turbojet engine

A nacelle for a turbojet engine has a longitudinal axis and a rear section including an annular vein formed by a wall of a fixed internal structure and a wall of an external structure. The nacelle includes a device for modulating the cross-section of a space formed by the annular vein. The device includes an injector to inject an auxiliary flow of a gas so as to vary the orientation or speed of the auxiliary flow, a suction orifice for drawing in part of the injected auxiliary flow, and an internal auxiliary flow return area in one or more walls. In particular, the internal return area allows the circulation of part of the injected auxiliary flow and the drawn-in auxiliary flow

Nacelle for an aircraft bypass turbojet engine

The disclosure relates to a nacelle for a turbojet engine, including an upstream section forming an air intake lip and defining a space for the circulation of a main air flow. The nacelle includes a device for modulating the cross-section of the space and has means for injecting an auxiliary flow of a gas by means of an induced ejector effect; suction means for drawing in the injected auxiliary flow; and an internal auxiliary flow return area in one or more walls, the area being configured to allow the circulation of the injected auxiliary flow and the drawn-in auxiliary flow and to bring part of the injected auxiliary flow into contact with the main air flow. The internal return area includes an outer front part forming an area in which the thickness of the internal return area is reduced.

RAM AIR TURBINE COOLING INLET SCREEN HEATING SYSTEM

A heating system for a ram air turbine includes a heated screen that is disposed across an air inlet defined by a strut. The strut has a first end that is connected to a generator that is operably connected to a turbine that is connected to a second end of the strut. The heated screen has a heating element embedded therein. 1. A ram air turbine system , comprising:a strut that is pivotally connected to a mounting frame that is operatively connected to an aircraft structure, the strut defining a strut body that extends between a first end and a second end, the strut body defining an air inlet;a heated screen having a heating element embedded within the heated screen;a turbine that is disposed at the second end of the strut; anda generator that is disposed proximate the first end of the strut, the generator being drivably connected to the turbine via a drive shaft.2. The ram air turbine system of claim 1 , wherein the heating element is electrically connected to the generator.3. The ram air turbine system of claim 1 , wherein the heating element is electrically isolated from an aircraft power source.4. The ram air turbine system of claim 1 , the air inlet is configured as a cooling air inlet that provides cooling air to the generator.5. The ram air turbine system of claim 4 , wherein the generator defines a cooling air outlet that exhausts the cooling air that is provided to the generator.6. A ram air turbine system claim 4 , comprising:a strut having a first end that is connected to a generator that is drivably connected to a turbine connected to a second end, the strut provided with an air inlet that is disposed proximate the first end; anda heating system being electrically connected to the generator, the heating system having a heated screen disposed over the air inlet.7. The ram air turbine system of claim 6 , wherein the heated screen includes a heating element disposed between a first screen element and a second screen element that is spaced apart from and ...

BLEED AIR VALVE A TURBINE ENGINE WITH ANTI-ICE VALVE ASSEMBLY AND METHOD OF OPERATING

An anti-ice valve, a turbine engine including an anti-ice valve assembly and a method of operating an anti-ice valve where an anti-ice valve has a housing and a valve element configured to control a flow of hot bleed air through the housing, and a muscle air passage extending through the housing, a cooling air passage extending through the housing, and a heat exchanger located within the housing and having heat transfer surfaces in thermal communication with the muscle air passage and the cooling air passage. 1. A bleed air valve assembly , comprising:a bleed air valve having a housing and a valve element configured to control a flow of hot bleed air through the housing;a muscle air passage extending through the housing;a cooling air passage extending through the housing; anda heat exchanger located within the housing and having heat transfer surfaces in thermal communication with the muscle air passage and the cooling air passage, wherein heat is transferred from muscle air within the muscle air passage to cooling air in the cooling air passage by the heat transfer surfaces to effect a cooling of the muscle air and define a lower temperature muscle air stream.2. The bleed air valve assembly of wherein the muscle air passage includes multiple passages forming a muscle air circuit and the cooling air passage includes multiple passages forming a cooling air circuit.3. The bleed air valve assembly of wherein the housing includes a servo housing and a separate valve element housing.4. The bleed air valve assembly of wherein the servo housing defines a cavity that houses the heat exchanger.5. The bleed air valve assembly of wherein the heat exchanger is integrally formed with the servo housing.6. The bleed air valve assembly of wherein the servo housing and valve element housing are spaced apart from each other.7. The bleed air valve assembly of claim 4 , further comprising a muscle air outlet fluidly coupling the servo housing to the valve element housing such that ...

ICE PROTECTION SYSTEM FOR A COMPONENT OF AN AERODYNAMIC SYSTEM

Disclosed is an ice protection system for an aerodynamic surface of an aircraft, a surface having a flow facing side and an inwardly facing side that opposes the flow facing side, the system having: a perforated sheet configured for disposal in the surface; a heating source connected to the perforated sheet; and a suction source disposed to draw ice melted by the heating source through the perforated sheet and heating source. 1. A method of preventing ice formation with an ice protection system on a surface of a component of an aerodynamic system , the surface having a flow facing side and an inwardly facing side that opposes the flow facing side ,the method comprising:heating the surface with a heating source of the ice protection system, andproviding suction on an inwardly facing side of the heating source with a suction source to draw water melted by heating the surface through the heating source into a water collector.2. The method of claim 1 , wherein a flow facing side of the heating source comprises a perforated sheet.3. The method of claim 2 , wherein the heating source includes a drain hole extending between the perforated sheet and the suction source.4. The method of claim 3 , wherein the ice protection system comprises a honeycomb support structure on an inward side of the perforated sheet.5. The method of claim 4 , wherein the ice protection system comprises a water collector on an inwardly facing side of the honeycomb support structure.6. The method of claim 5 , wherein the suction source is a pump fluidly connected to the water collector.7. The method of claim 6 , wherein the ice protection system comprises a rigid shell on an inward facing side of the water collector.8. The method of claim 7 , wherein the ice protection system is coupled to a leading edge of the component.9. The method of claim 8 , wherein the component is a nacelle.10. The method of claim 8 , wherein the component is a wing. This application is a division of U.S. application Ser. No. ...

SENSOR WITH INTEGRAL VORTEX TUBE FOR WARMING

Sensor assemblies and methods of de-icing or preventing ice formation are provided. Compressed air may be supplied to a vortex tube. The vortex tube may separate the compressed air into a first stream and a second stream, the first stream hotter than the second stream. A sensor body may be warmed by the first stream, and the second stream may be directed away from the sensor body 1. A sensor assembly comprising:a sensor body;a sensor coupled to the sensor body; anda vortex tube enclosed at least in partly within the sensor body, the vortex tube including a first outlet and a second outlet, the vortex tube configured to separate compressed air into a first stream that exits the first outlet and a second stream that exits the second outlet, the first stream hotter than the second stream, the first stream directed to the sensor body for warming the sensor body.2. The sensor assembly of claim 1 , wherein the vortex tube is integral to the sensor body.3. The sensor assembly of claim 1 , wherein the sensor body comprises an air collection manifold configured to generate the compressed air from an airflow directed at the sensor body.4. The sensor assembly of claim 3 , wherein a plurality of holes in the air collection manifold are configured to receive air from the airflow.5. The sensor assembly of claim 1 , wherein the first outlet is positioned to direct the first stream toward an inner wall of the sensor body.6. The sensor assembly of claim 1 , wherein the sensor includes a temperature sensor and/or a pressure sensor.7. The sensor assembly of claim 1 , wherein the sensor assembly includes a thermal shield between the sensor and the first stream directed to the sensor body.8. An apparatus comprising:a sensor body;a sensor positioned within the sensor body; anda means for warming the sensor body, wherein the means for warming is coupled or integral to the sensor body, the means for warming comprises a vortex tube configured to separate compressed air into a first stream ...

ANTI-ICE SYSTEMS FOR ENGINE AIRFOILS

An anti-ice system for a gas turbine engine may comprise a power generating device and a first thermally conductive applique comprising a first heating circuit. The power generating device may comprise a first component configured to rotate about an axis and generate a current. A first conductive layer may electrically couple the first heating circuit and the first component of the power generating device. 1. An anti-ice system for a gas turbine engine , comprising:a power generating device comprising a first component configured to rotate about an axis and generate a current;a first thermally conductive applique comprising a first heating circuit; anda first conductive layer electrically coupling the first heating circuit and the first component of the power generating device.2. The anti-ice system of claim 1 , wherein the first heating circuit comprising a plurality of resistor circuits.3. The anti-ice system of claim 1 , wherein the power generating device further comprises:a first magnet located over a first side of the first component; anda second magnet located over a second side of the first component opposite the first side of the first component.4. The anti-ice system of claim 3 , wherein the first component comprises a Faraday disk.5. The anti-ice system of claim 4 , wherein a first end of the first conductive layer is coupled to a positive pole of the first component and a second end of the first conductive layer opposite the first end is coupled to a negative pole of the first component.6. The anti-ice system of claim 1 , wherein the power generating device comprises a piezoelectric device.7. The anti-ice system of claim 1 , further comprising:a second thermally conductive applique comprising a second heating circuit; anda second conductive layer electrically coupling the second heating circuit and the first component of the power generating device.8. The anti-ice system of claim 7 , wherein the first thermally conductive applique is configured to attach ...

SWEEPING JET SWIRL NOZZLE

An injector head for an anti-icing system may comprise a body having a first surface, a second surface, a face, and an inlet, a first sweeping jet nozzle having a first exit port through the face, wherein the first sweeping jet nozzle comprises a fluid oscillator defining a first sweeping plane of the first sweeping jet nozzle, and a distribution manifold within the body in fluid communication with the inlet and the first sweeping jet nozzle. 1. An injector head for an anti-icing system comprising:a body having a first surface, a second surface, a face, and an inlet;a first sweeping jet nozzle having a first exit port through the face, wherein the first sweeping jet nozzle comprises a fluid oscillator defining a first sweeping plane of the first sweeping jet nozzle; anda distribution manifold within the body in fluid communication with the inlet and the first sweeping jet nozzle.2. The injector head of claim 1 , wherein the fluid oscillator is bi-stable.3. The injector head of claim 2 , wherein fluid oscillator is one of a feedback free oscillator claim 2 , a single feedback oscillator claim 2 , or a double feedback oscillator.4. The injector head of claim 3 , further comprising a second sweeping jet nozzle and a third sweeping jet nozzle claim 3 , the second sweeping jet nozzle having a second sweeping plane and a second exit port through the face claim 3 , the third sweeping jet nozzle having a third sweeping plane and a third exit port through the face.5. The injector head of claim 4 , wherein the first sweeping plane claim 4 , the second sweeping plane claim 4 , and the third sweeping plane are co-planar.6. The injector head of claim 4 , wherein the first sweeping plane is perpendicular to the second surface and the second sweeping plane is disposed at a non-orthogonal angle to the second surface.7. The injector head of claim 6 , wherein the first exit port has a first diameter claim 6 , the second exit port has a second diameter claim 6 , and the third exit ...

TURBOMACHINE COMPRISING A HEAT MANAGEMENT SYSTEM

A dual-flow turbomachine including a nacelle, compressors, turbines, a fuel supply line, a transfer line, a de-icing circuit, and a heat management system having: a first heat exchanger providing an exchange of heat between fuel in the supply line and oil in the transfer line, a loop comprising a main line and a pump which circulates a heat transfer fluid in the main line, where the main line is connected to an outlet of the pump and enters an inlet of a third heat exchanger, where at the outlet of the third heat exchanger the main line meets an inlet of the de-icing circuit, where at the outlet of the de-icing circuit the main line meets the inlet of the pump, and where the third heat exchanger transfers heat between the heat transfer fluid of the main line and the oil of the transfer line. 1. A dual-flow turbomachine for an aircraft comprising:a nacelle forming an air inlet lip,a set of compressors,a combustion chamber,a set of turbines,a supply line configured to supply fuel to the combustion chamber,a transfer line configured to circulate oil from the set of compressors to the set of turbines,a de-icing circuit for the air inlet lip, anda heat management system comprising:a first heat exchanger providing an exchange of heat between the fuel in the supply line and the oil in the transfer line,a loop comprising a main line and a pump which is designed to circulate a heat transfer fluid in the main line, where the main line is connected to the outlet of the pump and enters an inlet of a third heat exchanger, where at the outlet of the third heat exchanger the main line meets an inlet of the de-icing circuit, where at the outlet of the de-icing circuit the main line meets the inlet of the pump, and where the third heat exchanger provides an exchange of heat between the heat transfer fluid of the main line and the oil of the transfer line leaving the first heat exchanger.2. The turbomachine according to claim 1 , wherein the heat management system comprises:a first ...

Systems and methods of optimizing cooling and providing useful heating from single phase and two phase heat management in propulsion systems

Systems and methods of heat management of turbine engines including turbofans, turboprops and turboshafts and fan driven propulsion systems. The propulsion system may comprise a fan, nacelle, an electrical or mechanical heat source and a cooling system consisting of heat exchangers in the fan duct and on the nacelle and coolant pumps. The heat source can be a motor or a generator or turbine machinery or accessories rotationally coupled to rotating shafts. The heat management system transfers heat to the air in the fan flow path to provide additional fan thrust. The heat management system also transfers heat to structural members in the gas flow path that require anti-icing.

ICE PROTECTION SYSTEM FOR GAS TURBINE ENGINES

A gas turbine engine includes devices, systems, and methods for de-icing using heat of compression and expelling the de-icing medium into a gas flow path of the engine. 1. A gas turbine engine comprisinga fan mounted to rotate about an axis to draw air into the engine,an engine core that extends along the axis, the engine core including a compressor having an compressor inlet fluidly connected to receive air from the fan and a compressor discharge, a combustor having an combustor inlet fluidly connected to receive compressed air from the compressor discharge and a combustor discharge, and a turbine having a turbine inlet fluidly connected to the combustor discharge to receive combustion products therefrom, andan ice protection system including a heat exchanger having a hot side passage and a cold side passage arranged in thermal communication with each other, wherein the hot side passage is fluidly connected with the compressor discharge to receive compressed air therefrom and is fluidly connected to a downstream target to provide cooled-compressed air thereto, the cold side passage is fluidly connected with an upstream input to receive sink fluid therefrom and is fluidly connected with a de-icing location to provide heated-sink fluid thereto to provide heat for de-icing, the de-icing location is fluidly connected by a conduit with a dump point that is located within an engine gas flow path to expel spent sink fluid received from the de-icing location.2. The gas turbine engine of claim 1 , wherein the dump point is upstream of the compressor inlet.3. The gas turbine engine of claim 1 , wherein the dump point is located such that expelled sink fluid is introduced into a flow of air that includes at least a portion of air sent to a bypass duct when the gas turbine engine is operating.4. The gas turbine engine of claim 1 , wherein the dump point is located downstream of the fan.5. The gas turbine engine of claim 4 , wherein the dump point is located such that expelled ...

AIR INTAKE AND METHOD FOR DE-ICING AN AIR INTAKE OF A NACELLE OF AN AIRCRAFT JET ENGINE

A de-icing device for an air intake of an aircraft jet engine nacelle extending along an axis in which a flow of air flows from upstream to downstream, the intake comprising an inner wall and an outer wall connected by a leading edge, the inner wall comprising a plurality of blowing lines, each blowing line comprising a plurality of through-openings configured to blow elementary streams from the hot air source in order to de-ice said inner wall, the blowing lines being parallel to one another in a cylindrical projection plane, each blowing line having a depth defined along the axis X and a length defined along the axis Y in the cylindrical projection plane, two adjacent blowing lines being spaced apart by a distance, the ratio of the distances L3/D3 being between 1 and 2.

INLET FRAME FOR A GAS TURBINE ENGINE

An inlet frame and a method of additively manufacturing the same are provided. The inlet frame includes a forward annular body spaced apart from a rear annular body to define an inlet passageway in fluid communication with a compressor inlet. The inlet frame may define integral wash manifolds and discharge ports for directing a flow of wash fluid directly through the compressor inlet. In addition, inlet frame may define one or more integral annular heating plenums in fluid communication with a hot air source for heating regions of the inlet frame that are prone to icing conditions. 1. An inlet frame for a gas turbine engine , the gas turbine engine defining an axial direction , a radial direction , and a circumferential direction , the gas turbine engine comprising a compressor defining a compressor inlet and being rotatable about the axial direction for pressurizing an airflow , the inlet frame comprising:a forward annular body;a rear annular body spaced apart from the forward annular body to define an inlet passageway in fluid communication with the compressor inlet;an annular heating plenum defined by the forward annular body; anda fluid supply conduit defining a hot fluid passageway, the hot fluid passageway providing fluid communication between the annular heating plenum and a hot air source for providing a flow of heated fluid to the annular heating plenum.2. The inlet frame of claim 1 , wherein the annular heating plenum is an inlet heating plenum defined at a radially outer portion of the forward annular body proximate an inlet screen of the gas turbine engine.3. The inlet frame of claim 2 , wherein a plurality of stiffening ribs is defined within the inlet heating plenum.4. The inlet frame of claim 2 , wherein the inlet screen comprises:a forward inlet screen that extends from the inlet heating plenum outward along the radial direction, the forward inlet screen comprising a plurality of tubes in fluid communication with the inlet heating plenum for ...

ANTI ICING METHOD AND APPARATUS

A method of reducing applied heat within an inlet duct of a gas turbine generating electricity includes applying heat to the inlet duct of the gas turbine to attain an initial temperature set point and to produce conditions sufficient for preventing formation of ice within the inlet duct, measuring a position of an inlet guide vane (IGV) of the gas turbine, an inlet duct temperature, and an inlet duct relative humidity to determine a thermodynamic state in the inlet duct, evaluating the thermodynamic state to determine if the conditions are sufficient for preventing formation of ice within the inlet duct, and in response to determining that sufficient conditions exist within the inlet duct for preventing formation of ice, adjusting the applied heat to maintain the measured inlet duct temperature. 1. A method of reducing applied heat within an inlet duct of a gas turbine generating electricity , the method comprising:applying heat to the inlet duct of the gas turbine to attain an initial temperature set point and to produce conditions sufficient for preventing formation of ice within the inlet duct;measuring a position of an inlet guide vane (IGV) of the gas turbine, an inlet duct temperature, and an inlet duct relative humidity to determine a thermodynamic state in the inlet duct;evaluating the thermodynamic state to determine if the conditions are sufficient for preventing formation of ice within the inlet duct; andin response to determining that sufficient conditions exist within the inlet duct for preventing formation of ice, adjusting the applied heat to maintain the measured inlet duct temperature.2. The method of claim 1 , further comprising comparing the measured inlet duct temperature to the initial temperature set point.3. The method of claim 2 , wherein:the measured inlet duct temperature is less than the initial temperature set point when the sufficient conditions exist within the inlet duct for preventing formation of ice.4. The method of claim 1 , ...

AIRCRAFT PROPULSION ASSEMBLY COMPRISING AIR-LIQUID HEAT EXCHANGERS

An aircraft propulsion assembly includes a turbine engine surrounded by a nacelle with an annular air-intake lip extending around the turbine engine by two annular walls, inner and outer, respectively, intended for being swept across by air flows at least when the aircraft is in flight. The inner and outer walls each includes or supports at least one network of pipes forming heat exchangers. The inner wall pipe network having liquid outlet connected with a liquid intake of the outer wall pipe network. The propulsion assembly further includes means for circulating the liquid, connected to at least one liquid intake of the network of pipes of the inner wall. 1. A propulsion assembly for aircraft , comprising a turbomachine surrounded by a nacelle comprising an annular air intake lip extending around the turbomachine by two annular walls , inner and outer , respectively , configured to be swept across by air flows at least when the aircraft is in flight , wherein:said inner and outer walls each comprises or supports at least one network of pipes configured to transport a liquid in contact with said inner wall or said outer wall to form air-liquid heat exchangers respectively inner and outer, the pipes of each inner or outer heat exchanger being connected in parallel with each other,a network of pipes of the inner wall having at least one liquid outlet connected in series with at least one liquid intake of a network of pipes of the outer wall, andthe propulsion assembly comprises means for circulating the liquid, connected to at least one liquid intake of a network of pipes of the inner wall for its liquid supply, and connected to at least one liquid outlet of the network of pipes of the outer wall for the recovery of the liquid.2. The propulsion assembly according to claim 1 , wherein the turbomachine is connected to the nacelle by at least one passage of ancillaries tubular arm claim 1 , said at least one liquid intake of the network of pipes of the inner wall and ...

SYSTEM AND METHOD FOR DE-ICING A GAS TURBINE ENGINE

The invention relates generally to gas turbine engines used for electrical power generation. More specifically, embodiments of the present invention provide ways for improving gas turbine engine performance by reducing ice build-up on the inlet filter housing through heated air injection. 1. A gas turbine engine comprising:an inlet filter housing;an inlet air system in fluid communication with the inlet filter housing;a compressor in fluid communication with the inlet air system;a compressor discharge plenum in fluid communication with the compressor;an inlet bleed heat system in fluid communication with the inlet air system and the compressor discharge plenum; and,an auxiliary source of compressed air;wherein heated air is directed to the inlet filter housing to raise an operating temperature of the inlet filter housing and reduce ice formation in the inlet filter housing.2. The gas turbine engine of claim 1 , wherein the heated air is supplied by the compressor discharge plenum of the gas turbine engine.3. The gas turbine engine of claim 1 , wherein the heated air is supplied by the auxiliary source of compressed air external to the gas turbine engine.4. The gas turbine engine of claim 1 , wherein the auxiliary source of compressed air is a storage tank of compressed air.5. The gas turbine engine of claim 1 , wherein the heated air is directed to the inlet filter housing by way of a first series of air pipes and a second series of air pipes.6. The gas turbine engine of claim 1 , wherein the auxiliary source of compressed air comprises a fueled engine coupled to a multi-stage compressor where waste heat from the fueled engine heats air from the multi-stage compressor in a recuperator.7. The gas turbine engine of claim 1 , wherein the auxiliary source of compressed air provides heated air for power augmentation via the inlet bleed heat system.8. The gas turbine engine of further comprising a plurality of valves operated by a controller for regulating the heated air. ...

ADDITIVELY MANUFACTURED BOOSTER SPLITTER WITH INTEGRAL HEATING PASSAGEWAYS

A booster splitter for a gas turbine engine and a method of additively manufacturing the booster splitter are provided. The booster splitter includes an annular inner wall defining a radially outer boundary of a compressor flow path defined through a compressor section of the gas turbine engine, an annular outer wall spaced apart from the annular inner wall along the radial direction, the annular outer wall adjacent to the annular inner wall at a forward end, the forward end defining an inlet to the compressor flow path, an annular bulkhead spanning between the annular inner wall and the annular outer wall substantially along the radial direction, the bulkhead defining an inlet port, and a passageway defined within the annular outer wall, the passageway extending from the inlet port, into the bulkhead, radially outward to the outer wall, and through the annular outer wall towards the inlet defined by the forward end. 1. A gas turbine engine defining an axial direction and a radial direction , the gas turbine engine comprising:an annular inner wall defining a radially outer boundary of a compressor flow path defined through a compressor section of the gas turbine engine;an annular outer wall spaced apart from the annular inner wall along the radial direction, the annular outer wall adjacent to the annular inner wall at a forward end, the forward end defining an inlet to the compressor flow path;an annular bulkhead spanning between the annular inner wall and the annular outer wall substantially along the radial direction, the bulkhead defining an inlet port; anda passageway defined within the annular outer wall, the passageway extending from the inlet port, into the bulkhead, radially outward to the outer wall, and through the annular outer wall towards the inlet defined by the forward end.2. The gas turbine engine of claim 1 , wherein the outer wall defines heat exchange fins within the passageway.3. The gas turbine engine of claim 1 , further including an inlet ...

Gas turbine engine anti-icing system

An anti-icing system ( 100 ) for an engine section stator ( 26 ) of a gas turbine engine ( 10 ), the system ( 100 ) comprising: an environmental control system pre-cooler heat exchange system ( 116 ) configured to exchange heat between air bled from a compressor ( 14, 16 ) of the engine ( 10 ) and bypass duct air; and a conduit ( 132 ) configured to exchange heat from the pre-cooler heat exchange system ( 116 ) to a further a heat transfer medium, the conduit ( 132 ) being configured to transfer the heat from the heat transfer medium to the engine section stator ( 26 ).

INJECTOR NOZZLE CONFIGURATION FOR SWIRL ANTI-ICING SYSTEM

An anti-icing system for annular turbofan engine housings what include a substantially closed annular housing at a leading edge of the turbofan engine housing, the annular housing containing a quantity of air and a conduit extending from a source of high pressure hot gas to the annular housing. The system also includes an injector connected to the end of the conduit and extending into the annular housing; one or more nozzles extending outwardly from the injector in a direction that the quantity of air circulates in the annular housing while the turbofan engine is operating. The nozzles have an entrance in fluid contact with the injector and an exit, wherein a cross-sectional area of the entrance is less than a cross-sectional area of the exit such that gas leaving the nozzles is travelling slower than gas entering the nozzles. 1. An anti-icing system for annular turbofan engine housings comprising:a substantially closed annular housing at a leading edge of the turbofan engine housing, the annular housing containing a quantity of air;a conduit extending from a source of high pressure hot gas to the annular housing;an injector connected to the end of the conduit and extending into the annular housing; andone or more nozzles extending outwardly from the injector in a direction that the quantity of air circulates in the annular housing while the turbofan engine is operating, the nozzles having an entrance in fluid contact with the injector and an exit, wherein a cross-sectional area of the entrance is less than a cross-sectional area of the exit such that gas leaving the nozzles is travelling slower than gas entering the nozzles.2. The system of claim 1 , further comprising a turbofan jet engine compressor claim 1 , wherein the turbofan jet engine compressor is the source of the high pressure hot gas.3. The system of claim 1 , wherein the one or more nozzles includes at least three nozzles.4. The system of claim 1 , wherein at least one of the one or more nozzles ...

ANTI-ICING EXHAUST SYSTEM

An anti-icing exhaust system for an inlet of a gas turbine engine is provided. The system includes a first housing portion that at least partially defines a chamber to receive an anti-icing fluid and includes a projection. A second housing portion includes a leading edge. The system includes a support structure that couples the first housing portion to the second housing portion such that the projection of the first housing portion is spaced apart from the leading edge of the second housing portion to define a gap. The gap is in fluid communication with a manifold. The support structure defines at least one flow passage in fluid communication with the chamber and the manifold to exhaust the fluid. The system includes an insulation strip having a first end received in the gap and coupled to the leading edge; and a second end coupled to the second housing portion. 1. An anti-icing exhaust system for an inlet of a gas turbine engine , comprising:a first housing portion associated with the inlet, the first housing portion at least partially defining a chamber to receive an anti-icing fluid from a source and including a projection;a second housing portion associated with the inlet, the second housing portion including a leading edge;a support structure that couples the first housing portion to the second housing portion such that the projection of the first housing portion is spaced apart from the leading edge of the second housing portion to define a gap between the first housing portion and the second housing portion, the gap in fluid communication with a manifold defined between the projection of the first housing portion and the support structure, and the support structure defines at least one flow passage in fluid communication with the chamber and the manifold to exhaust the anti-icing fluid from the chamber through the manifold and the gap; andan insulation strip having a first end and a second end, with the first end received in the gap and coupled to the leading ...

BLEED AIR SYSTEMS FOR USE WITH AIRCRAFT AND RELATED METHODS

Bleed air systems for use with aircraft and related methods are disclosed. An example apparatus includes a compressor having a compressor inlet, a compressor outlet, and a first drive shaft. The compressor outlet is to be fluidly coupled to a system of an aircraft that receives pressurized air, and the compressor inlet is to receive bleed air from a low-pressure compressor of an engine of the aircraft. The example apparatus includes a gearbox operatively coupled to the first drive shaft to drive the compressor. The gearbox is to be operatively coupled to and powered by a second drive shaft extending from the engine. The example apparatus also includes a clutch disposed between the first drive shaft and the gearbox to selectively disconnect the first drive shaft from the gearbox. 1. A bleed air system for an aircraft , the bleed air system comprising:a turbo-compressor including a compressor, a turbine, and a first drive shaft coupled between the compressor and the turbine, the compressor having a compressor inlet to receive bleed air from an engine of the aircraft and a compressor outlet to be fluidly coupled to a system of the aircraft that receives pressurized air;a gearbox operatively coupled to and powered by a second drive shaft extending from the engine; anda freewheel disposed between the first drive shaft and the gearbox to operatively couple the first drive shaft and the gearbox.2. The bleed air system of claim 1 , wherein claim 1 , during a first operating condition of the aircraft claim 1 , the freewheel is to operatively couple the first drive shaft to the gearbox such that the gearbox drives the first drive shaft to drive the compressor and claim 1 , during a second operating condition of the aircraft claim 1 , the freewheel is to enable the first drive shaft to rotate faster than an output of the gearbox.3. The bleed air system of claim 2 , wherein claim 2 , during the second operating condition of the aircraft claim 2 , the turbine is to drive the ...

Using aircraft electric icing protection system for electrical power system quality

A method of preserving power quality on a power grid is provided. The method is implemented by a voltage overflow device that is electrically coupled to the power grid through electrical contacts. A voltage monitoring circuit of the voltage overflow device monitors a voltage via the electrical contacts on the power grid with respect to a predetermined voltage. The voltage monitoring circuit determines whether the voltage exceeds the predetermined voltage. A switch of the voltage overflow device shunt an excess voltage over the predetermined voltage to a resistive load when the voltage exceeds the predetermined voltage to preserve the power quality on the power grid.

Method for detecting a fluid leak in a turbomachine and fluid distribution system

A method for detecting a high temperature fluid leak in a turbomachine. The turbomachine includes a source of high temperature pressurized fluid, at least one fluid distribution line suitable for distributing said high temperature fluid, and a turbomachine compartment wherein the distribution line is at least partially housed. The method includes measuring at least two pressure parameters of the turbomachine compartment, including a measured pressure and a pressure variation over time; detecting a high temperature fluid leak when at least one of the two pressure parameters of the turbomachine compartment reaches a characteristic value of a high-temperature fluid leak in the compartment. A high-temperature fluid distribution system and a turbomachine comprising such a high temperature fluid distribution system.

INJECTOR NOZZLE CONFIGURATION FOR SWIRL ANTI-ICING SYSTEM

An anti-icing system for annular gas turbine engine inlet housings includes a substantially closed annular housing at a leading edge of the gas turbine engine inlet housing, the annular housing containing a quantity of air and a conduit extending from a source of high-pressure hot bleed air to the annular housing. The system also includes an injector connected to the end of the conduit and extending into the annular housing and one or more nozzles extending outwardly from the injector in a direction that the quantity of air circulates in the annular housing. The system may include one or both of an airfoil on an upstream side of the injector and an air direction element disposed one the injector that causes the quantity of air to be directed toward an outlet of the one or more nozzles. 1. An anti-icing system for annular gas turbine engine inlet housings comprising:a substantially closed annular housing at a leading edge of the gas turbine engine inlet housing, the annular housing containing a quantity of air;a conduit extending from a source of high-pressure hot bleed air to the annular housing;an injector connected to the end of the conduit and extending into the annular housing, the injector including an airfoil on an upstream side; andone or more nozzles extending outwardly from the injector in a direction that the quantity of cooler air circulates in the annular housing.2. The system of claim 1 , further comprising a gas turbine engine compressor claim 1 , wherein the gas turbine engine compressor is the source of the high-pressure hot bleed air.3. The system of claim 1 , further comprising:a first air directional element disposed on an end of the injector that causes the quantity of air to be directed toward an outlet of the one or more nozzles.4. The system of claim 1 , further comprising:a second air direction element disposed on either a top or a bottom the injector that causes the quantity of air to be directed toward an outlet of the one or more nozzles.5 ...

Composite Components Having Piezoelectric Fibers

Composite components and methods for forming composite components are provided. For example, a composite component of a gas turbine engine comprises a composite material, a plurality of piezoelectric fibers, and an anti-icing mechanism. The anti-icing mechanism is in operative communication with the piezoelectric fibers such that the anti-icing mechanism is activated by one or more electrical signals from the piezoelectric fibers. In exemplary embodiments, the composite component is a composite airfoil and the anti-icing mechanism is one or more heating elements. Methods for forming composite components may comprise forming piezoelectric plies comprising piezoelectric fibers embedded in a matrix material; forming reinforcing plies comprising reinforcing fibers embedded in the matrix material; laying up the piezoelectric and reinforcing plies to form a ply layup; and processing the ply layup to form the composite component. Methods including forming a piece of piezoelectric material that is adhered to a composite component also are provided.

Gas Turbine Anti-Icing System

Embodiments of the present disclosure are directed towards a system that includes a recirculation system. The recirculation system includes a compressor discharge air line extending from a compressor to an air intake. The compressor discharge air line is configured to flow a compressor discharge air flow, and the air intake is configured to supply an air flow to the compressor. An ejector is disposed along the compressor discharge air line between the compressor and the air intake. The ejector is configured to receive and mix the compressor discharge air flow and a turbine exhaust flow to form a first mixture. 1. A system , comprising:a recirculation system, comprising:a compressor discharge air line configured to extend from a compressor to an air intake, wherein the compressor discharge air line is configured to flow a compressor discharge air flow, and the air intake is configured to supply an air flow to the compressor; andan ejector disposed along the compressor discharge air line between the compressor and the air intake, wherein the ejector is configured to receive and mix the compressor discharge air flow and a turbine exhaust flow to form a first mixture.2. The system of claim 1 , wherein the air intake comprises a discharge air manifold claim 1 , and the discharge air manifold is configured to receive the first mixture from the compressor discharge air line.3. The system of claim 2 , wherein the discharge air manifold comprises a conduit with apertures claim 2 , and the apertures are configured to flow the first mixture into the air intake.4. The system of claim 3 , wherein the air intake comprises an inlet configured to receive an ambient air flow claim 3 , and the air intake is configured to mix the first mixture and the ambient air flow to form a second mixture.5. The system of claim 4 , wherein the air intake is configured to supply the second mixture to the compressor.6. The system of claim 4 , wherein the air intake comprises a housing and a filter ...

COMPRESSOR STAGE

The invention relates to the field of compressors, and specifically a compressor stage () comprising at least a casing () delimiting an air passage (), a stator () comprising a plurality of guide vanes () arranged radially around a central axis (X) in the air passage (), and a rotor () suitable for rotating about the central axis (X) relative to the stator () and comprising a plurality of blades () arranged radially around the central axis (X) in the air passage () downstream from the guide vanes (). Each blade () of the rotor () extends from a blade root (a) to a blade tip (b) further away from the central axis (X) than the blade root (a) and presents radial clearance (j) between the blade tip (b) and the casing (). In order to avoid ice forming on the guide vanes, and also in order to avoid blade tip clearance vortices, at least one of said guide vanes () includes an internal cavity () with a hot air inlet () for deicing the guide vane (), and the internal cavity () presents a first outlet passage () towards a trailing edge () of the guide vane () for injecting an air jet () into a boundary layer () adjacent to the casing () upstream from the blades () of the rotor (). 1. A compressor stage comprising at least:a casing delimiting an air passage;a stator comprising a plurality of guide vanes arranged radially around a central axis in the air passage; anda rotor rotatable about the central axis relative to the stator and comprising a plurality of blades arranged radially around the central axis in the air passage downstream from the guide vanes, each blade of the plurality of blades extending from a blade root to a blade tip further away from the central axis than the blade root and presenting radial clearance between the blade tip and the casing;wherein at least one guide vane of the plurality of guide vanes includes an internal cavity with a hot air inlet for deicing the at least one guide vane, and in that the internal cavity presents a first outlet passage ...

Airfoil deicing system

A gas turbine engine includes an airfoil and a deicing system. The airfoil radially extends from a hub towards a case disposed about a central longitudinal axis of the gas turbine engine. The deicing system includes an acoustic driver assembly arranged to apply acoustic energy to the airfoil to excite a predetermined vibratory mode of the airfoil.

Sound-Absorbing Panel With A Cellular Core And A De-Icing System

A sound-absorbing panel includes: an inner skin traversed by holes and intended to be oriented towards a channel in which a fluid flows, a heating mat formed by strips fixed to the inner skin on the side opposite to the channel and oriented in a first direction, wherein two adjacent strips are distant from each other in order to define a slot between them, a base fixed to the strips on the side opposite to the inner skin, wherein the base includes, on the strips side, grooves extending in a second direction different from the first direction and wherein the base has, between two successive grooves, a rib, a cellular core fixed to the base on the side opposite to the strips, and an outer panel fixed to the cellular core on the side opposite to the base. 1. A sound-absorbing panel comprising:an inner skin traversed by holes and configured to be oriented towards a channel in which a fluid flows;a heating mat comprising a plurality of strips fixed to the inner skin on the side opposite to the channel and oriented in a first direction, where two adjacent strips are distant from each other in order to define a slot between them;a base fixed to the plurality of strips on the side opposite to the inner skin, where the base comprises, on the strips side, grooves extending in a second direction different from the first direction and where the base has, between two successive grooves, a rib;a cellular core fixed to the base on the side opposite to the strips; andan outer panel fixed to the cellular core on the side opposite to the base.2. The sound-absorbing panel according to claim 1 , wherein the plurality of strips is integral with each other.3. The sound absorbing panel according to claim 1 , wherein the first direction and the second direction are perpendicular.4. The sound absorbing panel according to claim 1 , wherein the total thickness of the base at the level of the ribs is at least 4 mm.5. The sound absorbing panel according to claim 1 , wherein each of the ...

AIR INTAKE STRUCTURE OF AN AIRCRAFT NACELLE

An air intake structure for an aircraft nacelle is disclosed. The air intake structure delimits a channel and includes a lip having a U-shaped cross section oriented towards the rear, a first sound-absorbing panel fixed behind the lip and delimiting the channel, and a second sound-absorbing panel fixed behind the first sound-absorbing panel and delimiting the channel. Each sound-absorbing panel includes a cellular core which is fixed between an inner skin pierced with holes and oriented towards the channel, and an outer skin oriented in the opposite direction, where the inner skin of the first sound-absorbing panel has a thickness greater than the thickness of the inner skin of the second sound-absorbing panel, and where each of the inner skins includes a heat source which is embedded in the mass of the inner skin. 1. Air intake structure for an aircraft nacelle , the air intake structure delimiting a channel , comprising:a lip having a U-shaped cross section oriented towards a rear,a first sound-absorbing panel fixed behind the lip and delimiting the channel, anda second sound-absorbing panel fixed behind the first sound-absorbing panel and delimiting the channel,wherein each sound-absorbing panel comprises a cellular core fixed between an inner skin oriented towards the channel and pierced with holes oriented towards the channel, and an outer skin oriented in the opposite direction,wherein the inner skin of the first sound-absorbing panel has a thickness greater than the thickness of the inner skin of the second sound-absorbing panel, andwherein each of the inner skins comprises a heat source which is embedded in the mass of the inner skin.2. The air intake structure according to claim 1 , wherein the inner skin of the first sound-absorbing panel has a thickness of 3 to 6 mm.3. The air intake structure according to claim 1 , wherein the inner skin of the second sound-absorbing panel has a thickness of 0.6 to 2.1 mm.4. The air intake structure according to claim 1 ...

Pulsed deicing system

A deicing system for an aircraft may supply heat to an aircraft component in pulses. A first series of pulses may melt ice built up on the aircraft component. A second series of pulses may prevent ice from forming on the aircraft component. The length of each of the pulses in the first series of pulses may be longer than the length of each of the pulses in the second series of pulses. The pulses may be supplied by a pneumatic deicing system or an electrical deicing system.

AIR INTAKE OF AN AIRCRAFT TURBOJET ENGINE NACELLE COMPRISING VENTILATION ORIFICES FOR A DE-ICING FLOW OF HOT AIR

The invention relates to an air intake of an aircraft turbojet engine nacelle, extending along an axis X, in which an air flow circulates from upstream to downstream, the air intake extending circumferentially around the axis X and comprising an inner wall, which faces the axis X in order to guide an inner air flow, and an outer wall, which is opposite the inner wall, for guiding an external air flow, the walls being connected by a leading edge and an inner partition so as to delimit an annular cavity. The air intake comprises means for injecting at least one hot air flow into the inner cavity and at least one ventilation orifice formed in the outer wall in order to allow the hot air flow to escape after heating the internal cavity, the air intake comprising at least one disruption member of the external air flow, positioned upstream of the ventilation orifice, which extends outwardly from the outer wall. 18-. (canceled)9. An air intake of an aircraft turbojet engine nacelle extending along an axis X in which an air stream circulates from upstream to downstream , the air intake extending circumferentially about axis X and comprising an internal wall pointing to axis X to guide an internal air stream and an external wall which is opposite to the internal wall , to guide an external air stream , the walls being connected through a leading edge and an internal partition wall so as to delimit an annular cavity , the air intake comprising means for injecting at least one hot air stream into the internal cavity and at least one ventilation opening formed in the external wall to allow exhaust of the hot air stream after heating the internal cavity , which air intake is characterized in that it comprises at least one member for disturbing the external air stream , positioned upstream of the ventilation opening , which extends projecting outwardly from the external wall , the disturbance member being a polyhedral , preferably tetrahedral or pyramidal , vortex generating ...

Heated rigid electrical system

A rigid electrical raft has electrical conductors embedded in a rigid material. The electrical conductors transmit electrical signals through the rigid electrical raft, which may form part of an electrical system of a gas turbine engine. The rigid electrical raft also has electrical heating elements embedded therein. The electrical heating elements provide heat which may be used, for example, to prevent condensation and/or ice build-up and/or to raise the temperature of electrical components to be within a desired range.

ANTI-ICING APPARATUS FOR A NOSE CONE OF A GAS TURBINE ENGINE

A fan nose cone is disclosed for impeding icing and recovering momentum in a gas turbine engine. The fan nose cone comprises: an axially symmetric shell having a convex external surface and an internal surface, the shell having an opening in a forward end of the shell for communication with a source of heated pressurized air; and an axially symmetric deflector disposed forward of the opening, the deflector being configured to direct heated pressurized air exiting from the opening radially outwardly to flow in a downstream direction over the convex external surface of the shell during operation. The shell of the fan nose cone may have a rearward circumferential vent in communication with the source of heated pressurized air for directing heated pressurized from the vent in a radially outward and downstream direction toward the fan blade platforms. 1. A fan nose cone for a gas turbine engine having an axis of rotation and a forward end relative to a primary airflow path through the engine , the fan nose cone comprising:an axially symmetric shell having a convex external surface and an internal surface, the shell having an opening in a forward end of the shell, the opening adapted to be in communication with a source of heated pressurized air when the nose cone is installed on the engine; andan axially symmetric deflector disposed forward of the opening in the shell, the deflector having a rearward surface disposed forward of and cooperating with the convex external surface of the shell to define an annular air flow channel therebetween for directing heated pressurized air exiting from the opening, the rearward surface configured to radially outwardly direct said heated pressurized air to flow in a downstream direction over the convex external surface of the shell.2. The fan nose cone according to wherein the opening comprises a single central opening.3. The fan nose cone according to wherein the opening comprises a plurality of apertures symmetrically disposed about a ...

Air intake of an aircraft turbojet engine nacelle comprising ventilation orifices for a de-icing flow of hot air

The invention relates to an air intake of an aircraft turbojet engine nacelle, extending along an axis X, in which an air flow circulates from upstream to downstream, the air intake comprising an inner wall facing the axis X and an outer wall for guiding an external air flow, the walls being connected by a leading edge and an inner partition so as to delimit an annular cavity. The air intake comprises means for injecting at least one hot air flow into the inner cavity and at least one ventilation orifice formed in the outer wall to allow the hot air flow to escape after heating the inner cavity, the ventilation orifice comprising an upstream edge, the circumferential profile of which is discontinuous in order to generate turbulences, and a downstream edge, the radial profile of which is aerodynamic in order to limit the formation of pressure fluctuations.

NOZZLE FOR AN AIRCRAFT PROPULSION SYSTEM

An assembly is provided for an aircraft propulsion system. This assembly includes an inlet lip for a nacelle inlet structure, a bulkhead and a nozzle. The bulkhead is configured with the inlet lip to form a cavity between the inlet lip and the bulkhead. The nozzle includes a distribution conduit and one or more nozzle ports. The distribution conduit projects longitudinally out from the bulkhead and into the cavity along a conduit centerline. At least a first portion of the distribution conduit tapers as the distribution conduit extends longitudinally along the conduit centerline. The first portion of the distribution conduit has an elongated geometry in a plane perpendicular to the conduit centerline. The one or more nozzle ports are arranged longitudinally along the conduit centerline and fluidly coupled with an internal passage of the distribution conduit. 1. An assembly for an aircraft propulsion system , comprising:an inlet lip for a nacelle inlet structure;a bulkhead configured with the inlet lip to form a cavity between the inlet lip and the bulkhead; anda nozzle including a distribution conduit and one or more nozzle ports;the distribution conduit projecting longitudinally out from the bulkhead and into the cavity along a conduit centerline, at least a first portion of the distribution conduit tapering as the distribution conduit extends longitudinally along the conduit centerline, and the first portion of the distribution conduit having an elongated geometry in a plane perpendicular to the conduit centerline; andthe one or more nozzle ports arranged longitudinally along the conduit centerline and fluidly coupled with an internal passage of the distribution conduit.2. The assembly of claim 1 , whereinthe inlet lip includes an inner lip skin and an outer lip skin; andthe cavity extends axially along an axis between a forward end of the inlet lip and the bulkhead, the cavity extends radially between the inner lip skin and the outer lip skin, and the cavity ...

ANTI-ICE SPLITTER NOSE

Splitter apparatus for gas turbine engines are disclosed. An example splitter apparatus may include a splitter including an annular outer wall substantially defining a convex leading edge; an annular splitter support positioned radially within the outer and including a forward end disposed substantially against a splitter inner; and an annular first bulkhead spanning between the outer wall and the splitter support. The outer wall, the splitter support, and the first bulkhead may define a generally annular splitter plenum. The forward end of the splitter support may include spaced apart, radially oriented metering slots. The outer wall may include an inner portion disposed radially inward from the splitter inner surface extending aft and including spaced-apart exit slots. The splitter plenum, the metering slots, and the exit slots may conduct airflow from the plenum, through the metering slots against the splitter inner surface, and through the exit slots. 1. A splitter apparatus for a gas turbine engine , comprising:a splitter including a generally annular outer wall substantially defining a convex leading edge at a forward end thereof;a generally annular splitter support positioned radially within the outer wall, the splitter support comprising a forward end disposed substantially against a splitter inner surface of the outer wall; anda generally annular first bulkhead spanning between the outer wall and the splitter support, wherein the outer wall, the splitter support, and the first bulkhead collectively substantially define a generally annular splitter plenum;wherein the forward end of the splitter support comprises a plurality of circumferentially spaced apart, generally radially oriented metering slots therein;wherein the outer wall comprises an inner portion disposed generally radially inward from the splitter inner surface and extending generally aft, the inner portion comprising a plurality of spaced-apart exit slots; andwherein the splitter plenum, the ...

EMBEDDED AIRCRAFT HEATER REPAIR

An ice protection system for an aircraft may comprise an electric heater embedded within a composite aircraft component. In the event of malfunction of the embedded heater, a secondary heater may be attached to an interior surface of the composite aircraft component. The wiring may be detached from the embedded heater and attached to the secondary heater. The secondary heater may be used to provide ice protection to the composite aircraft component. 1. A method comprising:detecting a malfunction in an embedded electric heater in an aircraft component;coupling a secondary heater to an interior surface of the aircraft component;detaching a wire from the embedded electric heater; andcoupling the wire to the secondary heater.2. The method of claim 1 , wherein the secondary heater comprises an aperture.3. The method of claim 2 , wherein the secondary heater is coupled to the aircraft component such that an electrical contact of the embedded electric heater protrudes through the aperture.4. The method of claim 1 , further comprising at least one of supplying more power to the secondary heater than was supplied to the embedded electric heater or increasing a dwell time cycle supplied to the secondary heater.5. The method of claim 1 , further comprising anti-icing or deicing an external surface of the aircraft component using the secondary heater.6. The method of claim 1 , wherein the secondary heater comprises an electric heater.7. The method of claim 1 , wherein the embedded electric heater is located between layers of a composite laminate.8. A method of repairing an ice protection system comprising:coupling a secondary heater to an interior surface of a composite aircraft component;detaching a wire from a first electrical contact coupled to the composite aircraft component; andcoupling the wire to a second electrical contact coupled to the secondary heater.9. The method of claim 8 , wherein the first electrical contact is coupled to an embedded heater embedded within the ...

Advanced inlet design

A compact inlet design including a single bulkhead and/or an acoustic panel extending into nacelle lip region for noise reduction. The compact inlet is used with a low power fluid ice protection system capable of preventing ice build-up on the acoustic panel in the nacelle lip region.

DE-ICING AND CONDITIONING DEVICE FOR AN AIRCRAFT

The present disclosure provides a device for de-icing an air inlet lip of an aircraft nacelle. The device includes a pre-exchanger, an intake orifice of taking in low-pressure air downstream from a fan, and two high-pressure air intake orifices downstream from a compressor in addition to controlled valves and check valves installed in an air flow network. In particular, the pre-exchanger includes a low-pressure air outlet capable of opening into the air inlet lip of the aircraft nacelle via a pipe of the air flow network. 1. A device for de-icing an air intake lip of an aircraft nacelle , said device comprising:a pre-exchanger;a bleeding means configured to bleed a low-pressure air downstream of a fan; andtwo means for bleeding a high-pressure air downstream of a compressor as well as controlled valves and check valves installed in an air circulation network,wherein the pre-exchanger comprises a low-pressure air outlet configured to open into the air intake lip of the aircraft nacelle via a piping of the air circulation network.2. The de-icing device according to claim 1 , further comprising a discharge valve configured to discharge a high-pressure air circulating through the pre-exchanger.3. The de-icing device according to claim 1 , further comprising a valve configured to mix at least a part of the high-pressure air for a cabin conditioning and for an airfoil de-icing claim 1 , with the low-pressure air for the air intake lip de-icing at the low-pressure air outlet of the pre-exchanger.4. The de-icing device according to claim 1 , further comprising a valve between the low-pressure air outlet of the pre-exchanger and the air intake lip.5. The de-icing device according to claim 1 , further comprising a valve between the low-pressure air outlet of the pre-exchanger and an outside of the aircraft nacelle.6. The de-icing device according to claim 1 , further comprising a detector of a temperature of the air intake lip claim 1 , the detector configured to disable the ...

DE-ICING NOSE FOR LOW-PRESSURE COMPRESSOR OF AN AXIAL TURBINE ENGINE

The invention proposes a de-icing splitter nose at the inlet of the low-pressure compressor of a double-flow turbine engine. The splitter nose comprises an annular separation surface with a circular leading edge, the surface being adapted to separate an upstream flow into a primary flow and a secondary flow. The nose further comprises an electric de-icing device with a heating element having a series of parallel, electric, heating strips. The strips are inserted between two strata of dielectric material and are able to prevent the formation of ice on the annular surface, as close as possible to the leading edge. 1. An axial turbine engine splitter nose , said nose comprising:an annular separation surface with a circular leading edge, the surface structured and operable to separate an upstream flow into two annular flows; andan electrical de-icing device with an electrical heating element; the electrical heating element comprises at least one electric heating strip which is structured and operable to prevent the formation of ice on the annular surface.2. The axial turbine engine splitter nose according to claim 1 , wherein each strip crosses radially the circular leading edge.3. The axial turbine engine splitter nose according to claim 1 , further comprising a heated annular zone in which each strip is inset claim 1 , the at least one strip occupying at least 10% of the heated zone.4. The axial turbine engine splitter nose according to claim 1 , wherein each strip performs U-turns.5. The axial turbine engine splitter nose according to claim 1 , wherein the electrical heating element comprises a plurality of heating strips and two connectors each connected to the strips claim 1 , wherein one connector is an internal connector arranged radially on the inside of the leading edge and one connector is an external connector arranged radially on the outside of the leading edge.6. The axial turbine engine splitter nose according to claim 1 , wherein the de-icing device ...

DE-ICING DEVICE FOR AN AIR INTAKE OF AN AIRCRAFT TURBOJET ENGINE NACELLE

The invention relates to a de-icing device for an air intake of an aircraft turbojet engine nacelle extending along an X-axis in which an air stream flows from upstream to downstream, the air intake having an inner cavity, extending annularly about the X-axis, which comprises an inner wall facing the X-axis and an outer wall which is opposite the inner wall, the walls being connected by a leading edge, the de-icing device comprising at least one injector for injecting a stream of hot air into the inner cavity, the injector comprising a nozzle extending along a nozzle axis, the nozzle being configured to inject a stream of hot air having a dissymmetry along the nozzle axis. 110-. (canceled)11. A deicing device for an air intake of a nacelle of an aircraft turbojet engine extending along an axis X in which an air stream circulates from upstream to downstream , the air intake comprising an internal cavity annularly extending about axis X , which comprises an internal wall facing axis X and an external wall which is opposite to the internal wall , the walls being connected by a leading edge , the deicing device comprising at least one injector for a hot air stream into the internal cavity , the injector comprising a mouthpiece extending along a mouthpiece axis , the mouthpiece being configured to inject a hot air stream having an asymmetry along the mouthpiece axis so as to generate turbulence in the vicinity of the external wall while heating the internal wall.12. The deicing device according to claim 11 , wherein the mouthpiece comprises at least one first channel configured to lead a first elementary stream and at least one second channel configured to lead a second elementary stream so as to form the hot air stream.13. The deicing device according to claim 12 , wherein the first channel comprises at least one air deflection member.14. The deicing device according to claim 13 , wherein the air deflection member is configured to twist the first elementary stream.15. ...

DE-ICING DEVICE FOR A SPLITTER NOSE OF AN AXIAL TURBINE ENGINE COMPRESSOR

A de-icing device for a splitter nose of a dual-flow turbine engine compressor. The device comprises a splitter nose that has an inner flange and that separates a primary flow and a secondary flow, a shroud arranged on the inside of the annular wall and bearing an annular row of vanes and an abradable seal, a de-icing annular space for circulating a de-icing fluid between the nose and the shroud, a partition annularly dividing the annular space including an outer radial flange attached to the inner flange of the splitter nose using the same bolts that are used to attach the outer shroud. The nose has means for centring the partition, providing a press fit that optimizes the seal. 1. A turbine engine splitter nose de-icing device , said device comprising:a splitter nose that includes an inner flange and is structured and operable to separate a primary flow and a secondary flow of the turbine engine;a shroud arranged within the splitter nose;a de-icing annular space between the splitter nose and the shroud; anda partition dividing the annular space annularly, the partition comprising a tubular portion and an outer flange that projects radially outwards from the tubular portion and is attached to the inner flange of the splitter nose.2. The turbine engine splitter nose de-icing device of claim 1 , wherein the splitter nose includes centring means of the partition.3. The turbine engine splitter nose de-icing device of claim 2 , wherein the centring means are arranged to enable the press fitting of the partition in the centring means.4. The turbine engine splitter nose de-icing device of claim 2 , wherein the partition includes an annular centring thickening in circular contact with the centring means claim 2 ,5. The turbine engine splitter nose de-icing device of claim 2 , wherein the centring means include an inner tubular surface with a surface roughness Ra of at most 1.60 μm.6. The turbine engine splitter nose de-icing device of claim 1 , wherein the partition ...

FAN BLADE ANTI-ICING CONCEPT

A fan blade anti-icing system comprises a fan hub and a fan blade extending radially outwardly from the fan hub. The fan blade has a base and an airfoil extending radially outwardly from the fan base. The airfoil having a leading edge, a trailing edge, a convex side surface between the leading and trailing edge and a concave side surface between the leading and trailing edge. The fan blade further has a radial passage extending from a blade air inlet in the blade base in communication with a source of heated air, and a rearwardly directed passage in communication with the radial passage and having a blade air outlet forward of the trailing edge and oriented tangentially to the convex side surface or concave side surface of the airfoil. 1. A fan blade anti-icing system for a gas turbine engine comprising:a fan hub mounted for rotation about an axis; anda fan blade extending radially outwardly from the fan hub, the fan blade having a base and an airfoil extending radially outwardly from the base, the airfoil having a leading edge, a trailing edge, a convex side surface between the leading and trailing edge and a concave side surface between the leading and trailing edge, the fan blade further having a radial passage extending from a blade air inlet in the base in communication with a source of heated air, and a rearwardly directed passage in communication with the radial passage and having a blade air outlet forward of the trailing edge and oriented tangentially to the convex side surface or concave side surface of the airfoil.2. The fan blade anti-icing system according to comprising a plurality of rearwardly directed passages in communication with the radial passage and a plurality of blade air outlets in the concave side surface of the airfoil.3. The fan blade anti-icing system according to wherein the plurality of rearwardly directed passages and the plurality of blade air outlets are oriented towards the trailing edge.4. The fan blade anti-icing system according ...

Gas turbine engine with partial inlet vane

A turbofan engine including an axially extending inlet wall surrounding an inlet flow path. A radial distance between the inlet wall and the inner wall adjacent the fan defines a downstream height of the inlet flow path. A plurality of vanes are circumferentially spaced around the inlet, each of the vanes extending radially inwardly from the inlet wall, a maximum radial distance between a tip of each of the vanes and the inlet wall defining a maximum height of the vane. The maximum height of the vane is at most 50% of the downstream height of the flow path. In another embodiment, the maximum height of the vane is at most 50% of the maximum fan blade span. A method of reducing a relative Mach number at fan blade tips is also discussed.

CIRCUIT FOR DE-ICING AN AIR INLET LIP OF AN AIRCRAFT PROPULSION ASSEMBLY

Propulsion assembly, comprising a turbine engine surrounded by a nacelle comprising an annular air inlet lip, the propulsion assembly further comprising a circuit for lubricating elements of the turbine engine and a circuit for de-icing the air inlet lip, characterised in that said de-icing circuit comprises a heat exchanger comprising a primary circuit of oil supplied by said lubrication circuit and a secondary circuit of a heat transfer fluid for supplying at least one de-icing channel extending into said air inlet lip, said de-icing circuit further comprising a pump for circulating the heat transfer fluid into said at least one channel. 1. Propulsion assembly , comprising a turbine engine surrounded by a nacelle comprising an annular air inlet lip , the propulsion assembly further comprising a circuit for lubricating elements of the turbine engine and a circuit for de-icing the air inlet lip , wherein said de-icing circuit comprises a heat exchanger comprising two superimposed heat exchange modules , including:a first heat exchange module comprising a primary circuit of oil supplied by said lubrication circuit and a secondary circuit of a heat transfer fluid for supplying at least one de-icing channel extending into said air inlet lip, said de-icing circuit further comprising a pump for circulating the heat transfer fluid into said at least one de-icing channel,and a second heat exchange module of the surface type and comprising an outer surface which is intended to be swept by a flow of cooling air.2. Propulsion assembly according to claim 1 , wherein the lip comprises two skins which are superimposed and define said at least one channel therebetween.3. Propulsion assembly according to claim 2 , wherein one of the skins defines an outer surface of the lip.4. Propulsion assembly according to claim 2 , wherein the skins define a single de-icing channel therebetween claim 2 , which channel has a relatively small thickness and is designed to ensure the circulation ...

DEICING NOSE OF AN AXIAL TURBINE ENGINE COMPRESSOR

The invention relates to a de-icing splitter nose () of an axial turbine engine, notably a de-icing splitter nose of a turbo-jet engine compressor. The splitter nose includes an annular row of vanes (), each of which has a radially extending leading edge (), and a de-icing system () based on hot-air injection. The injection is pulsed, i.e. discontinuous. The system () includes an annular row of injection orifices for injecting de-icing fluid () onto the vanes () in respective injection directions (). Each injection orifice is associated with a vane such that the injection directions thereof are substantially parallel to the leading edge () of the related vane, enabling said vane () to be de-iced. 1. A de-icing splitter nose of an axial turbine engine with an axis of rotation , the splitter nose comprising:a vane with a leading edge that extends radially, anda de-icing system with an injection orifice structurally and functionally designed to inject a de-icing fluid in an injection direction,the injection direction of the injection orifice being generally parallel to the leading edge of the vane.2. The splitter nose according to claim 1 , wherein the injection direction is generally perpendicular to the axis of rotation of the turbine engine claim 1 , the injection direction being inclined of at least 80° with respect to the axis of rotation.3. The splitter nose according to claim 1 , wherein the injection direction is substantially oriented upstream.4. The splitter nose according to claim 1 , wherein the leading edge includes a mean line claim 1 , the injection direction being generally parallel to said mean line.5. The splitter nose according to claim 1 , wherein the leading edge has a curvature in a plane perpendicular to the axis of rotation.6. The splitter nose according to claim 1 , wherein the leading edge is substantially inclined in relation to the radial direction claim 1 , in a plane perpendicular to the axis of rotation and/or in a plane containing the ...

INLET AIR HEATING SYSTEM FOR A GAS TURBINE SYSTEM

An inlet air heating system for a gas turbine system includes an inlet heat exchanger configured to be positioned upstream of a compressor of the gas turbine system. The inlet air heating system also includes a heating loop fluidly coupled to the inlet heat exchanger. The heating loop is configured to provide heating fluid to the inlet heat exchanger, and the inlet heat exchanger is configured to facilitate transfer of heat from the heating fluid to an airflow into the compressor. Furthermore, the inlet air heating system includes a heat transfer assembly configured to receive cooling tower fluid from a fluid pathway extending between a steam condenser and a cooling tower. The heat transfer assembly is configured to facilitate transfer of heat from the cooling tower fluid to the heating fluid. 1. An inlet air heating system for a gas turbine system , comprising:an inlet heat exchanger configured to be positioned upstream of a compressor of the gas turbine system;a heating loop fluidly coupled to the inlet heat exchanger, wherein the heating loop is configured to provide heating fluid to the inlet heat exchanger, and the inlet heat exchanger is configured to facilitate transfer of heat from the heating fluid to an airflow into the compressor;a first heat transfer assembly configured to receive cooling tower fluid from a fluid pathway extending between a steam condenser and a cooling tower, wherein the first heat transfer assembly is configured to facilitate transfer of heat from the cooling tower fluid to the heating fluid; anda second heat transfer assembly configured to receive heated fluid from a steam turbine system, wherein the second heat transfer assembly is configured to facilitate transfer of heat from the heated fluid to the heating fluid.2. The inlet air heating system of claim 1 , wherein the first heat transfer assembly comprises a heating loop heat exchanger configured to facilitate transfer of the heat from the cooling tower fluid to the heating fluid. ...

Air Inlet Lip Of An Aircraft Engine Comprising A De-icing System

An aircraft engine air inlet lip takes an annular form about a longitudinal axis and delimits an air inlet stream, and includes: a wall having a U-shaped profile having an outer face oriented towards outside of the air inlet lip and an inner face oriented towards interior of the air inlet lip, an inner wall extending inside the wall between two zones of the inner face, so as to close an inner chamber delimited between the wall and the inner wall and filled with a gas, the inner wall having an upstream face oriented towards and a downstream face oriented away from the inner chamber, a fan configured to move the gas contained in the inner chamber, and at least one pipeline fixed to the upstream face and extending all around the air inlet lip and configured to be fed with a heat transfer fluid heated by a heat source. 1. An air inlet lip for an engine of an aircraft , said air inlet lip taking an annular form about a longitudinal axis X and delimiting an air inlet stream , said air inlet lip comprising:a wall having a U-shaped profile having an outer face oriented towards outside of the air inlet lip and an inner face oriented towards an interior of the air inlet lip;an inner wall extending inside the wall between two zones of the inner face, so as to close an inner chamber delimited between the wall and the inner wall and filled with a gas, the inner wall having an upstream face oriented towards the inner chamber and a downstream face oriented away from the inner chamber;a fan configured to move the gas contained in the inner chamber; andat least one pipeline fixed to the upstream face and extending all around the air inlet lip and configured to be fed with a heat transfer fluid heated by a heat source of the aircraft.2. The air inlet lip according to claim 1 , at least one or each of the at least one pipeline is equipped with fins oriented towards the outside of the pipeline and dipping into the inner chamber.3. The air inlet lip according to claim 1 , wherein the ...

NACELLE INLET WITH REINFORCEMENT STRUCTURE

A nacelle inlet structure for an aircraft propulsion system. This nacelle inlet structure includes an inlet lip, a bulkhead and a reinforcement structure. The inlet lip extends circumferentially about an axial centerline. The bulkhead extends circumferentially about the axial centerline. The bulkhead is configured with the inlet lip to form a cavity axially between the inlet lip and the bulkhead. The reinforcement structure extends circumferentially about the axial centerline. The reinforcement structure is connected to and extends axially between the inlet lip and the bulkhead thereby radially dividing the cavity into an inner sub-cavity and an outer sub-cavity. 1. A nacelle inlet structure for an aircraft propulsion system , comprising:an inlet lip extending circumferentially about an axial centerline;a bulkhead extending circumferentially about the axial centerline, the bulkhead configured with the inlet lip to form a cavity axially between the inlet lip and the bulkhead; anda reinforcement structure extending circumferentially about the axial centerline, the reinforcement structure connected to and extending axially between the inlet lip and the bulkhead thereby radially dividing the cavity into an inner sub-cavity and an outer sub-cavity.2. The nacelle inlet structure of claim 1 , wherein the inlet lip and the reinforcement structure are formed together as a monolithic body.3. The nacelle inlet structure of claim 1 , wherein the reinforcement structure is configured to rigidly tie the inlet lip and the bulkhead together.4. The nacelle inlet structure of claim 1 , wherein the reinforcement structure comprises a rib connected to and projecting axially out from an interior surface of the inlet lip.5. The nacelle inlet structure of claim 4 , wherein a surface of the rib is adjacent and contiguous with the interior surface of the inlet lip.6. The nacelle inlet structure of claim 4 , wherein the reinforcement structure further comprises a flange mounted to the ...

Aircraft Anti-Icing System

Various embodiments of the present disclosure provide an aircraft anti-icing system that includes an aircraft engine inlet component, a pressurized air source, and a heat source operable to: (1) heat the leading edge of the aircraft engine inlet component via the heat source to prevent ice formation on the outer surface of the leading edge; and (2) direct pressurized air from the pressurized air source so that it forces water off of the outer surface of the inlet components (and into the external air flow) as the water travels downstream from the leading edge outer surface downstream toward the trailing edge, which prevents runback ice formation. 1. A method for melting ice or preventing ice formation on a leading edge of an aircraft engine inlet component , the method comprising:heating the leading edge of the aircraft engine inlet component so that aerodynamic forces effect the flow of water along an exterior surface of the leading edge to an area downstream of an apex of the leading edge; anddirecting air from a plenum within the interior of the aircraft engine inlet component to the area downstream of the apex to force the water away from the exterior surface of the leading edge.2. The method of claim 1 , further comprising heating the air before directing the air from the plenum to the area downstream of the apex.3. The method of claim 2 , wherein heating the leading edge of the aircraft engine inlet component comprises directing the air to impinge upon an interior surface of the leading edge after heating the air.4. The method of claim 3 , wherein the air comprises engine bleed air heated by an aircraft engine claim 3 , further comprising directing the engine bleed air from the engine to the plenum.5. The method of claim 2 , further comprising directing the air from a pressurized air source to the plenum.6. The method of claim 1 , further comprising heating the leading edge via a heating device in thermal communication with the leading edge.7. The method of ...

CONTROL METHOD FOR CONTROLLING AN AIR INTAKE SYSTEM WHICH SUPPLIES AIR TO AN ENGINE OF A VEHICLE

A control method for controlling an air intake system for an engine of a vehicle; the intake system has a main air intake coupled to an air filter provided with a heating device. The control method comprises the steps of: determining a pressure difference between upstream and downstream of the air filter; determining a variation speed of the pressure difference between upstream and downstream of the air filter by calculating the first derivative in time of the pressure difference between upstream and downstream of the air filter; and turning on and/or turning off the heating device based on the variation speed of the pressure difference between upstream and downstream of the air filter. 1821849149. A control method for controlling an air intake system () which supplies air to an engine () of a vehicle (); the intake system () comprises a main air intake () coupled to an air filter () , which is provided with a heating device (); the control method comprises the step of determining a pressure difference (ΔP) between upstream and downstream of the air filter ();the control method is characterized in that it comprises the further steps of:{'b': 9', '9, 'determining a variation speed (dΔP/dt) of the pressure difference (ΔP) between upstream and downstream of the air filter () by calculating the first derivative in time of the pressure difference (ΔP) between upstream and downstream of the air filter (); and'}{'b': 14', '9, 'turning on and/or turning off the heating device () based on the variation speed (dΔP/dt) of the pressure difference (ΔP) between upstream and downstream of the air filter ().'}299. The control method according to claim 1 , wherein the variation speed (dΔP/dt) of the pressure difference (ΔP) between upstream and downstream of the air filter () is determined by applying a low-pass filter to the first derivative in time of the pressure difference (ΔP) between upstream and downstream of the air filter ().31491. The control method according to and ...

FRONT LIP OF A TURBOFAN ENGINE NACELLE COMPRISING HOT-AIR BORES UPSTREAM FROM ACOUSTIC PANELS

A front lip of a turbofan engine nacelle is provided, defined at the rear by a partition wall, the lip comprising an annular de-icing space including a tube for supplying hot air in order to de-ice the outer walls thereof, said lip also including acoustic panels arranged on the wall thereof substantially radially facing the axis of the nacelle. The front lip includes hot-air outlet bores/piercings which are arranged between the front end of the lip and said acoustic panels, which allow a flow of hot air, from said annular space, forming a substantially regular film which covers said acoustic panels with a boundary layer in order to heat said acoustic panels. 1. A front lip of a turbojet engine nacelle , delimited at a rear portion by a partition wall , the front lip including a de-icing annular volume including a hot air supply tube for performing a de-icing of its external walls , the front lip further comprising acoustic panels , the front lip further comprising hot air outlet piercings disposed between a front end of the front lip and said acoustic panels , the hot air outlet piercings providing a hot air flow rate , from said annular volume , forming a substantially uniform film covering said acoustic panels with a boundary layer of air in order to heat the acoustic panels.2. The nacelle front lip according to claim 1 , wherein the hot air outlet piercings are disposed upstream of the acoustic panels.3. The nacelle front lip according to claim 1 , wherein the acoustic panels cover a majority of a length along the nacelle.4. The nacelle front lip according to claim 1 , wherein a surface of the front lip that does not include any acoustic panel claim 1 , is heated by direct contact with hot air circulating inside the annular volume.5. The nacelle front lip according to claim 1 , wherein the annular volume is separated by internal radial partition walls into several angular sectors.6. The nacelle front lip according to claim 5 , wherein each angular sector includes ...

REGULATED OIL COOLING SYSTEM FOR A TURBINE ENGINE WITH DEICING OF THE NACELLE

The invention relates to an oil cooling system of a turbine engine installed in an aircraft comprising a circuit suitable for circulating the oil between the engine and at least one external heat exchanger placing the oil in thermal communication with a part of the lip of the nacelle, wherein it also comprises at least one air/oil heat exchanger placing the oil in thermal communication with the air circulating in the cold zone of the turbine engine, equipped with a device suitable for varying its oil cooling capacity, and with a control means for said cooling capacity variation device. 1. An assembly comprising a bypass turbine engine , a nacelle in which said turbine engine is installed , and an oil cooling system for said turbine engine comprising a circuit suitable for circulating the oil between the engine and at least one external heat exchanger consisting of oil pipelines of small diameter in contact with a metal skin constituting the surface of at least a part of the lip of the nacelle , wherein it also comprises at least one air/oil heat exchanger placing the oil in thermal communication with the air circulating in the bypass air stream of the turbine engine , equipped with a device suitable for varying its oil cooling capacity , and with a control means for said cooling capacity variation device.2. The assembly as claimed in claim 1 , in which the cooling capacity variation device is a tapping scoop with variable geometry claim 1 , for tapping air from the cold zone flow.3. The assembly as claimed in claim 2 , in which the tapping scoop can completely close the opening supplying the air/oil exchanger.4. The assembly as claimed in claim 1 , in which the external exchanger has sufficient cooling capacity to ensure on its own the temperature regulation of the oil in cruising flight conditions.5. The assembly as claimed in claim 1 , in which the external heat exchanger is placed in series with the air/oil heat exchanger.6. The assembly as claimed in claim 1 , ...

ENGINE INTAKE ASSEMBLY WITH SELECTOR VALVE

An engine assembly including an engine core including at least one internal combustion engine each including a rotor sealingly and rotationally received within a respective internal cavity to provide rotating chambers of variable volume in the respective internal cavity, a compressor having an outlet in fluid communication with an inlet of the engine core, a first intake conduit in fluid communication with an inlet of the compressor and with a first source of air, a second intake conduit in fluid communication with the inlet of the compressor and with a second source of air warmer than the first source of air, and a selector valve configurable to selectively open and close at least the fluid communication between the inlet of the compressor and the first intake conduit. A method of supplying air to a compressor is also discussed. 1. A method of supplying air to a compressor providing compressed air to an internal combustion engine core , the method comprising:directing air through an air conduit and through at least one heat exchanger extending across the air conduit;directing part of the air from the air conduit to an inlet of the compressor through a selected one of a first and second intake conduits, the first intake conduit connected to the air conduit upstream of the at least one heat exchanger and the second intake conduit connected to the air conduit downstream of the at least one heat exchanger; andpreventing the air from flowing from the air conduit to the inlet of the compressor through the other one of the first and second conduits.2. The method as defined in claim 1 , wherein the internal combustion engine core is part of a compound engine assembly for an aircraft claim 1 , the method further comprising selecting the first intake conduit to direct the part of the air from the air conduit to the inlet of the compressor during flight claim 1 , and selecting the second intake conduit to direct the part of the air from the air conduit to the inlet of the ...

AIRCRAFT PART ANTI-ICING TREATMENT METHOD

A method for the anti-icing treatment of a surface of an aircraft part made of an organic matrix composites includes a texturing step in which the surface is irradiated with femtosecond laser pulses so as to render the surface superhydrophobic. 1. A method for treating a surface of an aircraft part made of organic matrix composites , the method comprising a texturing step in which the surface is irradiated by femtosecond laser pulses so as to make the surface superhydrophobic.2. The method according to claim 1 , wherein the pulses have a duration of less than 900 femtoseconds.3. The method according to claim 1 , wherein the pulses have a duration of less than 600 femtoseconds.4. The method according to claim 1 , wherein the texturing step is preceded by a step of protecting the surface by applying an organic paint.5. The method according to claim 4 , wherein the paint comprises an epoxy base or a polyurethane base.6. An aircraft part made of organic matrix composites claim 1 , comprising a surface treated with the method according to .7. The aircraft part made of organic matrix composites according to claim 6 , wherein the surface is provided with:micro-craters having a diameter less than 1 mm and a depth less than 10 μm, andlashes having a dimension less than 1 μm, the lashes being periodically spaced.8. The aircraft part made of organic matrix composites according to claim 7 , wherein the micro-craters have a diameter between 1 μm and 100 μm.9. The aircraft part made of organic matrix composites according to claim 7 , wherein the micro-craters have a depth between 600 nm and 1 μm.10. The aircraft part made of organic matrix composites according to claim 7 , wherein the lashes have a dimension between 1 nm and 800 nm claim 7 , the lashes being periodically spaced.11. The aircraft part made of organic matrix composites according to claim 6 , the aircraft part being an air inlet section of the aircraft claim 6 , said surface being a surface of a leading edge and/or a ...

THERMAL MANAGEMENT SYSTEM FOR A GAS TURBINE ENGINE

In one exemplary embodiment, a gas turbine engine system for cooling engine components includes an engine core, a core housing containing the engine core, an engine core driven fan forward of the core housing, a nacelle surrounding the fan and the core housing, and a bypass duct defined between an outer diameter of the core housing and an inner diameter of the nacelle. Also included is a thermal management system having a coolant circuit including at least one of a first heat exchanger disposed on the inner diameter of the nacelle and a second heat exchanger disposed on a leading edge of a BiFi spanning the bypass duct. The first heat exchanger is in thermal communication with the second heat exchanger. 1. A gas turbine engine system for cooling engine components comprising:an engine core, a core housing containing the engine core, and an engine core driven fan forward of the core housing;a nacelle surrounding the fan and the core housing;a bypass duct defined between an outer diameter of the core housing and an inner diameter of the nacelle;a thermal management system including:a coolant circuit including at least one of a first heat exchanger disposed on the inner diameter of the nacelle and a second heat exchanger disposed on a leading edge of a BiFi spanning said bypass duct, wherein the first heat exchanger is in thermal communication with the second heat exchanger.2. The gas turbine engine system for cooling engine components of claim 1 , wherein the coolant circuit includes both said first heat exchanger and said second heat exchanger.3. The gas turbine engine system for cooling engine components of claim 2 , wherein a first coolant loop within said coolant circuit comprises said first heat exchanger claim 2 , said second heat exchanger claim 2 , and said cooled engine systems claim 2 , and wherein a second coolant loop in said coolant circuit includes a nacelle anti-icer.4. The gas turbine engine system for cooling engine components of claim 2 , wherein the ...

TURBINE ENGINE FAN MODULE INCLUDING A TURBINE ENGINE INLET CONE DE-ICING SYSTEM, AND A DE-ICING METHOD

The invention relates to an aviation turbine engine fan module including a de-icing system () for de-icing an inlet cone () and comprising a sheath () placed inside an inside space defined upstream by the inlet cone, said sheath comprising a first duct () having at least one hot air admission orifice (), said first duct being configured to convey hot air from a bearing enclosure () of the engine towards a wall of the inlet cone in order to heat it from the inside, the sheath further comprising a second duct () having at least one outlet situated downstream from the admission orifice of the first duct, said second duct being configured to discharge air from the first duct towards the downstream end of the engine. The invention also provides a method of de-icing a turbine engine inlet cone. 1. An aviation turbine engine fan module including an inlet cone , a lubrication bearing enclosure , and a de-icing system for de-icing the inlet cone , said de-icing system comprising a sheath placed inside an inside space defined upstream by the inlet cone , said sheath comprising a first duct-having at least one hot air admission orifice , said first duct being configured to convey hot air from the bearing enclosure towards a wall of the inlet cone in order to heat it from the inside , wherein the sheath further comprises a second duct having at least one outlet situated downstream from the admission orifice of the first duct , said second duct being configured to discharge air from the first duct towards the downstream end of the engine.2. The module according to claim 1 , wherein the sheath comprises:an inner tube centered on a longitudinal axis of the engine; andan outer tube arranged coaxially around the inner tube, the first duct being defined radially between the inner tube and the outer tube and being closed at a downstream end by an annular plate extending radially between the inner tube and the outer tube, the second duct being defined by the inside of the inner tube.3. ...

GAS TURBINE ENGINE, NACELLE THEREOF, AND ASSOCIATED METHOD OF OPERATING A GAS TURBINE ENGINE

The nacelle can have an inlet fluidly connecting a main gas path of a gas turbine engine core, the inlet having an inlet edge connecting an external skin to an internal duct wall, and a step formed in a surface of at least one of the skin and the duct wall, the step delimiting a first portion of the surface from a second portion of the surface, the second portion of the surface being recessed relative to the first portion of the surface, the second portion of the surface extending away from both the step and the inlet edge, whereas the first portion of the surface extends between the inlet edge and the step. 1. An aircraft engine nacelle comprising an inlet fluidly connecting to a main gas path of a gas turbine engine core , the inlet having an inlet edge connecting an external skin to an internal duct wall , and a step formed in a surface of at least one of the external skin and the duct wall , the step delimiting a first portion of the surface from a second portion of the surface , the second portion of the surface being recessed relative to the first portion of the surface by a height of the step , the second portion of the surface extending away from both the step and the inlet edge , whereas the first portion of the surface extends between the inlet edge and the step.2. The aircraft engine nacelle of wherein the height of the step is of between 0.010″ and 0.200″ measured normal to the surface.3. The aircraft engine nacelle of wherein the step has a riser.4. The aircraft engine nacelle of wherein the aircraft engine is a turbofan engine claim 1 , and the duct wall is an outer bypass duct wall.5. The aircraft engine nacelle of wherein the inlet edge is a portion of a D-duct claim 4 , the D-duct connecting the skin and the duct wall.6. The aircraft engine nacelle of wherein the step is formed at a junction between the D-duct and the duct wall.7. The aircraft engine nacelle of wherein a heating air conduit is provided inside the D-duct claim 5 , the heating air ...

TURBOFAN ENGINE, NACELLE THEREOF, AND ASSOCIATED METHOD OF OPERATION

The nacelle can have an inlet portion having a duct wall and an outer skin, the duct wall being annular around an axis and having a surface forming a radially-outer delimitation to a gas path upstream of a fan area, the duct wall extending from a rounded inlet edge of the nacelle to the fan area, a cavity located inside the inlet portion, a compressed air inlet leading into the cavity, and an outlet fluidly connecting the cavity to the gas path, the outlet having a plurality of apertures disposed circumferentially around the duct wall, the apertures sloping circumferentially. 1. A turbofan engine nacelle , the nacelle comprising an inlet portion having a duct wall and an outer skin , the duct wall being annular around an axis and having a surface forming a radially-outer delimitation to a gas path upstream of a fan area , the duct wall extending from a rounded inlet edge of the nacelle to the fan area , a cavity located inside the inlet portion , a compressed air inlet leading into the cavity , and an outlet fluidly connecting the cavity to the gas path , the outlet having a plurality of apertures disposed circumferentially around the duct wall , the apertures sloping circumferentially.2. The aircraft engine nacelle of wherein the apertures slope circumferentially by at least 10 degrees.3. The aircraft engine nacelle of wherein the apertures further slope radially inwardly by between 3 and 45 degrees.4. The aircraft engine nacelle of wherein the apertures are regularly circumferentially interspaced from one another along at least ¾ of the circumference of the inner skin.5. The aircraft engine nacelle of wherein the apertures are regularly circumferentially interspaced from one another along the entire circumference of the inner skin.6. The aircraft engine nacelle of wherein the plurality of apertures are delimited by a plurality of circumferentially-interspaced claim 1 , radially-oriented vanes claim 1 , the vanes sloping circumferentially.7. The aircraft engine ...

ROTOR BLADE RETENTION SYSTEM FOR A GAS TURBINE ENGINE

A rotor blade retention system for a gas turbine engine includes a rotor blade connection component defining a slot, with the rotor blade connection component including a first set of electric leads. Furthermore, the rotor blade retention system includes a rotor blade having a root section received within the slot, with the rotor blade further including a second set of electric leads. In this respect, the first and second sets of electric leads are electrically coupled together to permit electric current to be supplied to the rotor blade. 1. A rotor blade retention system for a gas turbine engine , the rotor blade retention system comprising:a rotor blade connection component defining a slot, the rotor blade connection component including a first set of electric leads; anda rotor blade including a root section received within the slot, the rotor blade further including a second set of electric leads,wherein the first and second sets of electric leads are electrically coupled together to permit electric current to be supplied to the rotor blade.2. The rotor blade retention system of claim 1 , wherein the first set of electric leads is at least partially positioned within the slot.3. The rotor blade retention system of claim 1 , wherein:the root section extends along a longitudinal direction of the gas turbine engine between an upstream surface of the root section and a downstream surface of the root section; andthe second set of electric leads includes a set of contacts positioned on the upstream surface.4. The rotor blade retention system of claim 3 , further comprising:a longitudinal retention member configured to prevent movement of the root section within the slot in the longitudinal direction, wherein the longitudinal retention member electrically couples the first and second sets of electric leads together.5. The rotor blade retention system of claim 4 , wherein the longitudinal retention member includes a set of contacts that engages the set of contacts of the ...

Nacelle air intake provided with a mixed ice protection system

An optimized protection against ice on the inner and outer faces of an aircraft engine nacelle air intake with the air intake including an outer face and an inner face meeting at a line at the longitudinally extreme, called extremum line, an acoustic panel being installed on the inner surface of a part of the inner face. An elimination system based on vibration of the ice formed is put in place on at least a part of the outer face and an ice formation prevention system using a hot fluid is put in place on at least a part of the inner face and either an ice elimination system or an ice formation prevention system using a hot fluid is installed on the inner face and on the outer face, a marking line marking the boundary between the two systems.

THREE-STREAM ENGINE HAVING A HEAT EXCHANGER

A three-stream engine is provided. The three-stream engine includes a fan section, a core engine disposed downstream of the fan section, and a core cowl annularly encasing the core engine and at least partially defining a core duct. A fan cowl is disposed radially outward from the core cowl and annularly encasing at least a portion of the core cowl. The fan cowl at least partially defining an inlet duct and a fan duct. The fan duct and the core duct at least partially co-extending axially on opposite sides of the core cowl. A heat exchanger disposed within the fan duct. The heat exchanger provides for thermal communication between a fluid flowing through fan duct and a motive fluid flowing through the heat exchanger. 1. A three-stream engine comprising:a fan sectiona core engine disposed downstream of the fan section;a core cowl annularly encasing the core engine and at least partially defining a core duct;a fan cowl disposed radially outward from the core cowl and annularly encasing at least a portion of the core cowl, the fan cowl at least partially defining an inlet duct and a fan duct, the fan duct and the core duct at least partially co-extending axially on opposite sides of the core cowl; anda heat exchanger disposed within the fan duct, wherein the heat exchanger provides for thermal communication between a fluid flowing through fan duct and a motive fluid flowing through the heat exchanger.2. The three-stream engine of claim 1 , further comprising at least one stationary strut that couples the core cowl to the fan cowl and extends through the fan duct.3. The three-stream engine of claim 2 , wherein the heat exchanger is disposed forward of the at least one stationary strut within the fan duct.4. The three-stream engine of claim 2 , wherein the heat exchanger is disposed aft of the at least one stationary strut within fan duct.5. The three-stream engine of claim 1 , wherein between about 10% and about 100% of the fluid flowing through the fan duct passes ...

INLET AIR HEATING SYSTEMS FOR COMBINED CYCLE POWER PLANTS

Inlet air heating systems for combined cycle power plants and combined cycle power plants including inlet air heating systems are disclosed. The inlet air heating systems may include a plurality of heating coil assemblies partially positioned within an inlet housing of a gas turbine system, and a vent valve in fluid communication with each of the heating coils. The inlet air heating system may also include a supply line in fluid communication with the heating coils to provide water to the heating coils, and a hot water line in fluid communication with the supply line and a component positioned downstream of a condenser of the combined cycle power plant. The hot water line may provide hot water from the combined cycle power plant to the supply line. Additionally, the inlet air heating system may include a drain line in fluid communication with the heating coils and the condenser. 1. An inlet air heating system for a gas turbine system of a combined cycle power plant , the inlet air heating system comprising:a plurality of heating coil assemblies at least partially positioned within an inlet housing of the gas turbine system;a vent valve in fluid communication with each of the plurality of heating coil assemblies, the vent valve allowing air to flow into and out of the plurality of heating coil assemblies when in an open position;a supply line in fluid communication with the plurality of heating coil assemblies, the supply line providing water to the plurality of heating coil assemblies;a hot water line in fluid communication with the supply line and a component of the combined cycle power plant positioned downstream of a condenser of the combined cycle power plant, the hot water line providing hot water from the combined cycle power plant to the supply line; anda drain line in fluid communication with the plurality of heating coil assemblies and the condenser of the combined cycle power plant, the drain line providing the water from the plurality of heating coil ...