Аккумуляторная система

Изобретение относится к области электротехники, а именно к аккумуляторной системе и транспортному средству с указанной системой, преимущественно, с перезаряжаемыми аккумуляторами на литиевой основе. Предложенная аккумуляторная система содержит несколько электрически соединенных друг с другом аккумуляторных элементов, каждый из которых содержит несколько электрохимических индивидуальных ячеек, при этом каждый элемент из указанных аккумуляторных элементов термически изолирован от других аккумуляторных элементов. Каждая из указанных электрохимических индивидуальных ячеек содержит контрольное устройство, которое выполнено с возможностью электрического отсоединения аккумуляторного элемента от других аккумуляторных элементов, если по меньшей мере, одно измеренное значение, контролируемое указанным контрольным устройством, лежит вне заданного диапазона значений. Повышение безопасности работы аккумуляторной системы за счет термической изоляции каждого элемента от других аккумуляторных элементов ...

ВСПОМОГАТЕЛЬНАЯ СИСТЕМА ПРИВОДА ВАЛА ПРОПУЛЬСИВНОЙ СИСТЕМЫ ВЕРТОЛЕТА

Изобретение относится к области авиации, в частности к вспомогательным системам привода воздушного винта. Вспомогательная система механического привода соединительного вала (8) пропульсивной системы вертолета содержит гидравлический двигатель (7), механически соединенный с упомянутым соединительным валом (8), гидравлический контур (10) питания жидкостью под давлением упомянутого гидравлического двигателя (7), гидравлический вентиль (11) быстрого открывания по команде, установленный на гидравлическом контуре, пирогидравлический аккумулятор (9), управляемый блоком (12) управления и соединенный с упомянутым гидравлическим вентилем (11). Пирогидравлический аккумулятор (9) содержит камеру (20), содержащую поршень (16), ограничивающий главный отсек (17), содержащий по меньшей мере один твердотопливный патрон (19), и отсек (18) для резерва жидкости. Воспламеняющий заряд (25), соединен с главным отсеком (17) и выполнен с возможностью инициирования воспламенения по меньшей мере одного твердотопливного ...

ПАНЕЛЬ ЖЁСТКОСТИ И СПОСОБ ЕЁ ИЗГОТОВЛЕНИЯ

Изобретение относится к авиации и касается панелей жесткости. Панель жесткости содержит оболочку и удлиненный элемент жесткости. При этом элемент жесткости содержит желобок, проходящий в продольном направлении и образующий вместе с оболочкой полость. Причем панель содержит средство накопления и возвращения электрической энергии, расположенное внутри полости. Во время изготовления панели оболочку наносят на средство накопления и возвращения электрической энергии так, чтобы желобок элемента жесткости и оболочка образовали вместе полость. Достигается снижение массы и объема конструкции. 2 н. и 15 з.п. ф-лы, 3 ил.

СИСТЕМА РЕКУПЕРАЦИИ И РЕГЕНЕРАЦИИ ЭНЕРГИИ

Система рекуперации и регенерации энергии содержит пироэлектрический модуль рекуперации электроэнергии (МРЭ), который генерирует напряжение в ответ на изменение температуры, трубопровод охлаждения, гидравлически связанный с источником охлаждающего агента для получения охлаждающего агента, клапан для регулировки расхода охлаждающего агента, модуль аккумулирования энергии для накопления и хранения напряжения. Поверхность пироэлектрического МРЭ, образованная непосредственно напротив нагреваемой твердой поверхности, получает тепло. Вырабатывается дополнительная электрическая энергия. 2 н. и 6 з.п. ф-лы, 14 ил.

КОРОБКА ПРИВОДОВ АГРЕГАТОВ ДЛЯ УПРАВЛЕНИЯ РУЛЯМИ ЛЕТАТЕЛЬНОГО АППАРАТА

Коробка (140) приводов агрегатов газотурбинного двигателя для летательного аппарата содержит кожух (42), тягу (115) управления рулями летательного аппарата, выполненную с возможностью скольжения в осевом направлении внутри коробки (140), и силовой цилиндр (120) привода тяги (115), установленный на упомянутом кожухе (42). Силовой цилиндр (120) содержит полый корпус (121), поршень (123), выполненный с возможностью поступательного перемещения внутри упомянутого корпуса (121), и шток (122) поршня, соединенный с упомянутым поршнем (123) и проходящий снаружи корпуса (121) силового цилиндра (120). Шток (122) соединен с тягой (115). Корпус (121) силового цилиндра (120) расположен между соединением штока (122) с тягой (115) и кожухом (42) коробки (140). Летательный аппарат содержит газотурбинный двигатель, управляемый руль и коробку (140) приводов агрегатов, в которой тяга (115) управления соединена с рулем. Группа изобретений направлена на сокращение времени монтажа и демонтажа при техническом ...

УСТРОЙСТВО ДЛЯ КРЕПЛЕНИЯ АГРЕГАТОВ

Изобретение относится к области авиационной техники и может быть использовано при креплении выхлопной трубы вспомогательной силовой установки самолета, а также в других отраслях промышленности при креплении агрегатов с последующей регулировкой и фиксацией. Устройство содержит верхний и боковые узлы крепления, включающие в себя кронштейны. В кронштейнах шарнирно закреплена по крайней мере одна регулируемая по длине тяга, которая соединена со стыкуемым агрегатом. Кроме того, каждый кронштейн боковых узлов имеет внешнюю серьгу с фиксатором вертикального перемещения. Внешняя серьга может поворачиваться в горизонтальной плоскости, а внутри нее смонтирована внутренняя серьга с фиксатором горизонтального перемещения. Внутренняя серьга шарнирно соединена со стыкуемым агрегатом посредством сферического подшипника. Упомянутые фиксаторы могут быть выполнены в виде винтов с контргайками, на концах каждого из которых шарнирно закреплена серьга. Кронштейны боковых узлов могут быть выполнены в виде жесткой ...

Гидродинамический привод-генератор

Полезная модель относится к области энергетического машиностроения и может быть использована в системах электроснабжения летательных аппаратов, а также наземных транспортных средств и направлена на повышение надежности работы и ресурса привода-генератора за счет совмещения в едином конструктивном элементе функций чувствительного элемента и гидроусилителя, что исключает необходимость передачи силы между деталями с относительным вращением в регуляторе частоты вращения генератора. Гидродинамический привод-генератор содержит генератор трехфазного переменного тока, регулятор частоты вращения генератора с чувствительным элементом и сервопоршнем, рабочие полости которого сообщены каналами с чувствительным элементом, гидродинамический преобразователь крутящего момента с насосным колесом, связанным с валом приводного двигателя, турбинным колесом, связанным с валом генератора и поворотными лопатками, кинематически связанными с сервопоршнем, и насос подачи рабочей жидкости, установленный на валу турбины ...

ЦЕПЬ ПОДАЧИ МОЩНОСТИ ДЛЯ ПРОТИВООБЛЕДЕНИТЕЛЬНОЙ СИСТЕМЫ ЛЕТАТЕЛЬНОГО АППАРАТА

Цепь подачи электропитания летательного аппарата содержит сеть (17) распределения мощности на борту летательного аппарата для электрических устройств (5b), расположенных в авиационном двигателе или вблизи упомянутого двигателя, и генератор (27) подачи мощности, встроенный в авиационный двигатель с тем, чтобы подавать мощность переменного тока в противооблединительную или антиобледенительную систему (5а). Генератор подачи мощности соединен с электромеханическим приводом реверсора тяги (5с) через выпрямитель (1) для подачи мощности постоянного тока в упомянутый привод. Уменьшаются технические требования к выпрямителю. 6 з.п. ф-лы, 1 ил.

КОРОБКА ПРИВОДОВ АГРЕГАТОВ ДЛЯ УПРАВЛЕНИЯ РУЛЯМИ ЛЕТАТЕЛЬНОГО АППАРАТА

... 1. Коробка приводов агрегатов газотурбинного двигателя для летательного аппарата, при этом упомянутая коробка (140) содержит кожух (42), тягу (115) управления рулями летательного аппарата, выполненную с возможностью скольжения в осевом направлении внутри упомянутой коробки (140), и силовой цилиндр (120) привода упомянутой тяги (115), установленный на упомянутом кожухе (42), при этом упомянутый силовой цилиндр (120) содержит полый корпус (121), поршень (123), выполненный с возможностью поступательного перемещения внутри упомянутого корпуса (121), и шток (122) поршня, соединенный с упомянутым поршнем (123) и проходящий, по меньшей мере, частично снаружи корпуса (121) силового цилиндра (120), при этом упомянутый шток (122) соединен с тягой (115), при этом коробка (140) отличается тем, что корпус (121) силового цилиндра (120) расположен между соединением штока (122) с тягой (115) и кожухом (42) коробки (140).2. Коробка по п. 1, в которой шток (122) поршня является полым и проходит через весь ...

ДРЕНАЖНАЯ МАЧТА ОТСЕКА ВСПОМОГАТЕЛЬНОЙ СИЛОВОЙ УСТАНОВКИ ЛЕТАТЕЛЬНОГО АППАРАТА

... 1. Дренажная мачта (13) для слива жидкостей из отсека (11) ВСУ летательного аппарата с первым концом (15), присоединенным к отсеку (11) ВСУ, и вторым концом (17) для выпускания упомянутых жидкостей в атмосферу, отличающаяся тем, что дренажная мачта (13) сконфигурирована по меньшей мере с сектором, имеющим площади поперечного сечения, убывающие по направлению ко второму концу (17), и тем, что площадь начального поперечного сечения упомянутого сектора является меньшей, чем площадь любого поперечного сечения дренажной мачты (13), более близкого к первому концу (15).2. Дренажная мачта (13) по п.1, при этом вся дренажная мачта (13) имеет площадь поперечного сечения, убывающую по направлению ко второму концу (17).3. Дренажная мачта (13) по п.1, при этом дренажная мачта (13) сконфигурирована коленчатой формы в направлении потока воздуха, и при этом упомянутый сектор является концевым патрубком (51).4. Дренажная мачта (13) по п.1, в которой упомянутый сектор является выпускным сектором (21) на ...

СИСТЕМА ОБЕСПЕЧЕНИЯ ЛЕТАТЕЛЬНОГО АППАРАТА РЕЗЕРВНОЙ ЭЛЕКТИЧЕСКОЙ МОЩНОСТЬЮ

Способ ускорения запуска двигатель-генераторного электромашинного преобразователя постоянного напряжения в переменное и устройство для его реализации

Группа изобретений относится к электромашинным преобразователям. Способ ускорения запуска двигатель-генераторного электромашинного преобразователя постоянного напряжения в переменное заключается в следующем. От источника постоянного тока заряжают емкостный накопитель, затем накопитель подключают к обмоткам приводного электродвигателя электромашинного преобразователя, а после разгона общего вала двигателя и генератора преобразователя обмотки двигателя переключают на питание от источника постоянного тока. К источнику постоянного тока также подключают обмотку возбуждения генератора, якорную обмотку которого подключают к выходным внешним выводам преобразователя. Причем на время разгона вала преобразователя обмотку возбуждения и якорную обмотку его генератора подключают на питание от емкостного накопителя. Причем якорную обмотку подключают через управляемый электронный инвертор, обеспечивая стартерный режим генератора с регулированием частоты и амплитуды его напряжения. Также заявлено устройство ...

УСТРОЙСТВО И СПОСОБ ПОДАЧИ АВАРИЙНОГО ЭЛЕКТРОПИТАНИЯ НА БОРТ ЛЕТАТЕЛЬНОГО АППАРАТА

... 1. Устройство подачи аварийного электропитания на борт летательного аппарата, обеспечивающее подачу питания на «существенную» часть электрической силовой цепи этого летательного аппарата, содержащее первую синхронную электрическую машину с независимым возбуждением, связанную с маховиком, и вспомогательное устройство для приведения маховика в состояние вращения и поддержания его вращения. ! 2. Устройство по п.1, также содержащее вторую синхронную электрическую машину с независимым возбуждением, связанную с турбиной с приводом от набегающего потока воздуха. ! 3. Устройство по п.2, в котором первая синхронная электрическая машина соединена: ! с вспомогательным устройством, которое соединено с «основной» шиной через первый трехфазный контактор, ! с этой шиной через второй трехфазный контактор, ! со второй синхронной электрической машиной через этот второй трехфазный контактор и третий трехфазный контактор. ! 4. Устройство по п.3, в котором вспомогательное устройство содержит два статических ...

КОНФИГУРАЦИЯ ХВОСТОВОГО ОБТЕКАТЕЛЯ ФЮЗЕЛЯЖА ЛЕТАТЕЛЬНОГО АППАРАТА СО ВСПОМОГАТЕЛЬНОЙ СИЛОВОЙ УСТАНОВКОЙ И СПОСОБ МОНТАЖА ВСПОМОГАТЕЛЬНОЙ СИЛОВОЙ УСТАНОВКИ

... 1. Хвостовой обтекатель (4) фюзеляжа летательного аппарата со вспомогательной силовой установкой (1), имеющий пожарный отсек (7), который вмещает вспомогательную силовую установку (1), причем вспомогательная силовая установка (1) расположена асимметрично со смещением в сторону относительно продольной оси (X) хвостового обтекателя (4) фюзеляжа и прикреплена непосредственно к стенке (10) несущей обшивки, образующей пожарный отсек (7). ! 2. Хвостовой обтекатель по п.1, в котором вспомогательная силовая установка (1) расположена асимметрично относительно горизонтальной плоскости летательного аппарата. ! 3. Хвостовой обтекатель по п.2, пожарный отсек (7) которого, вмещающий вспомогательную силовую установку (1), имеет по меньшей мере один технологический люк (3). ! 4. Хвостовой обтекатель по п.2, в котором вспомогательная силовая установка (1), расположенная внутри пожарного отсека (7), окружена рабочим пространством (2), определяемым объемом, окружающим вспомогательную силовую установку (1) ...

СИСТЕМЫ И СПОСОБЫ УМЕНЬШЕНИЯ ВЛИЯНИЯ ЭФФЕКТА КРУТИЛЬНЫХ КОЛЕБАНИЙ НА ВЫРАБОТКУ ЭЛЕКТРИЧЕСКОЙ ЭНЕРГИИ

ТЕРМОЭЛЕКТРИЧЕСКИЙ ГЕНЕРАТОР ГАЗОВОЙ ТУРБИНЫ

... 1. Устройство для выработки электрической энергии в газовой турбине или т.п. воздушного судна, содержащее множество термоэлектрических элементов, имеющих одну поверхность, окружающую источник тепла, причем источник холода образован холодной текучей средой, протекающей над другой поверхностью термоэлектрических элементов, при этом источник холода образован по меньшей мере одной кольцевой зоной, в которую вводится холодная текучая среда, причем стенка указанной кольцевой зоны образует тепловое соединение со вторыми поверхностями термоэлектрических элементов, при этом вторая стенка кольцевой зоны является внешней оболочкой устройства, отличающееся тем, что источник тепла образован по меньшей мере одной кольцевой зоной, в которой проходит часть отработавшего газа (образующего источник тепла) по спиральной траектории к внешнему радиусу, причем источник холода образован по меньшей мере одной кольцевой зоной, по которой проходит холодная текучая среда по спиральной траектории к внутреннему радиусу ...

УСТРОЙСТВО И СПОСОБ ЗАЩИТЫ ВЫЧИСЛИТЕЛЯ ТУРБОМАШИНЫ ЛЕТАТЕЛЬНОГО АППАРАТА ОТ ОШИБОК ИЗМЕРЕНИЯ СКОРОСТИ

Вспомогательное устройство для высотного самолета

УПРАВЛЕНИЕ ЭНЕРГОРАСПРЕДЕЛЕНИЕМ В МОДУЛЬНОЙ СИСТЕМЕ ПРЕОБРАЗОВАТЕЛЕЙ С ИСПОЛЬЗОВАНИЕМ РЕЗУЛЬТАТОВ РАСЧЕТА КОЭФФИЦИЕНТА ПОЛЕЗНОГО ДЕЙСТВИЯ

УСТРОЙСТВО И СПОСОБ РЕГУЛИРОВАНИЯ ВСПОМОГАТЕЛЬНОГО ДВИГАТЕЛЯ, ВЫПОЛНЕННОГО С ВОЗМОЖНОСТЬЮ ПОДАЧИ ТЯГОВОЙ МОЩНОСТИ НА НЕСУЩИЙ ВИНТ ВЕРТОЛЕТА

Power supply unit of electrical system for air plane, has rectangular membrane that is stretched by gas during changing of pressure and temperature, such that electric power is generated by piezoelectric film

The power supply unit (10) has a substrate (12) and rectangular membrane (14) that is provided with a piezoelectric film (18) made of polymeric polyvinylidene fluoride. A gas (22) is supplied between substrate and rectangular membrane. The rectangular membrane is stretched by gas at the time of changing the pressure and temperature, such that electric power is generated by piezoelectric film. The substrate and rectangular membrane are connected with each other with space for receiving the gas. Independent claims are included for the following: (1) an electrical system; and (2) a method for providing electrical system of air plane.

INTEGRIERTE STROMVERSORGUNG UND KLIMAREGELUNGSSYSTEM

Batterie

Die Erfindung betrifft eine Batterie (100; 100'), umfassend einen Energiespeicher (101), zwei elektrische Batteriekontakte, die elektrisch mit dem Energiespeicher (101) verbunden sind, wobei die Batterie (100; 100') Lagerungsmittel (102) umfasst, wobei die Lagerungsmittel (102) dazu ausgebildet sind, in Transportschienen (203; 204) und einer Lagerungsaufnahme (202') angeordnet zu werden.

Brennstoffzellensystemmodul, Set von Brennstoffzellensystemmodulen und Verwendung eines Brennstoffzellensystemmoduls in einem Luftfahrzeug

Brennstoffzellensystemmodul (10) zum Einsatz in einem Luftfahrzeug (32), das mindestens eine Brennstoffzellensystemkomponente umfasst, dadurch gekennzeichnet, dass das Brennstoffzellensystemmodul (10) mittels einer Befestigungseinrichtung, die dazu eingerichtet ist, das Brennstoffzellensystemmodul (10) lösbar an einem Spant einer Luftfahrzeugstruktur zu befestigen, derart modular mit einem Rumpfabschnitt (30) des Luftfahrzeugs (32) verbindbar ist, dass ein Gehäuseelement (16, 26) des Brennstoffzellensystemmoduls (10) einen Abschnitt einer Außenhülle (34) des Luftfahrzeugs (32) bildet, wenn das Brennstoffzellensystemmodul (10) mit dem Rumpfabschnitt (30) des Luftfahrzeugs (32) verbunden ist.

Energieversorgungssystem für Flugzeug

Luftkreislauf- Klimaregelungssystem

Method for water-supply and air-conditioning of aircraft cockpit and cabin, requires use of reaction product such as pure water for washing, and for drinking purposes after adding electrolytes

A method for air-conditioning and water supply in which the current, water, and heat are produced simultaneously the electrochemical way via a fuel-cell according to demand. A reaction product i.e. pure water, which can be directly used as rinsing/washing water or as drinking water after addition of electrolytes/minerals and/or for humidifying the cabin- and/or cockpit-air. Fuels based on hydrogen or hydrocarbon (e.g. methane, methanol, kerosene etc) can be used.

Versorgungssystem zur Energieversorgung in einem Luftfahrzeug, Luftfahrzeug und Verfahren zum Versorgen eines Luftfahrzeugs mit Energie

Versorgungssystem (100) zur Energieversorgung in einem Luftfahrzeug, umfassend: mindestens ein Triebwerk (101) zum Antreiben des Luftfahrzeuges; eine Brennstoffzelle (102) zum Versorgen des Luftfahrzeuges mit einer elektrischen Energie, und eine Flügelenteisungs-Einrichtung (305), die derart mit der Brennstoffzelle (102) gekoppelt ist, dass sich mittels eines beim Betrieb der Brennstoffzelle (102) anfallenden Wasserdampfes ein Flügel des Luftfahrzeuges mittels der Flügelenteisungs-Einrichtung (305) enteisen lässt, wobei der Wasserdampf, der Brennstoffprodukt der Brennstoffzelle (102) ist, vor dem Flügelenteisen zum Gewinnen von Wasser auskondensiert wird, als verbleibendes Brennstoffprodukt mittels einer Wärmepumpe erwärmt und danach der Flügelenteisungs-Einrichtung (305) zugeführt wird.

Vorrichtung zur Auflösung von Verwirbelungen in einem Eduktor

Improvements in or relating to electric power supply arrangements

... 975,311. Distribution network arrangements. GENERAL ELECTRIC CO. Ltd. Nov. 8, 1961 [Nov. 11, 1960; March 28, 1961], Nos. 38902/60 and 11409/61. Heading H2H. An electric power arrangement, particularly for aircraft has two bus-bars 3, 5, the first 3, for non-voltage sensitive loads, being connected through a low-valued resistor 6 to the second 5, for voltage-sensitive loads, which has a voltage-stabilizing means such as a battery 4 connected to it. The resistor 6 may be shunted, for engine starting purposes by a contact R1 opened by a generator voltage-sensitive relay R. Other arrangements are described in the Specification utilizing a second battery in the event of failure of the first and arrangements including a rectifier fed battery or zener diodes for use with single and three phase A.C. systems.

Thermal electrical power generation for aircraft

Stow abort mechanism for a ram air turbine

Power distribution switch for a power distribution system

Multiple source electrical power distribution in aircraft

An auxillary power system

POWER SUPPLIES FOR AIRCRAFT SERVICES

... 1316275 Ventilating aircraft BRITISH AIRCRAFT CORP Ltd 23 Dec 1970 [2 Oct 1969] 48491/69 Heading F4V [Also in Division F1] The invention relates to a power supply system for essential services of an aircraft, the aircraft having a cabin 7 which is pressurized from a source of air under pressure under the control of a pressurization and air conditioning system. Air under pressure is supplied from a source such as a tapping from a compressor of a gas turbine propulsion engine through line 1 and passes through control valve 2 to a first air cooler 3 then through a second air cooler 6 and water separator 8 to the cabin 7. The power supply system comprises two air turbines 16a, 16b which are connected through gearing 16c to drive an electric generator 17. The air turbine 16a is supplied with air under pressure from lines 1 and 18 under control of valve 19, the valve being controlled in dependence on temperature in the cabin as sensed by device 22, and the air turbine 16b is supplied either with ...

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.

Gas turbine engine auxiliary component mount

A mount system for an auxiliary component includes two side brackets (42, 44) and a top bracket (46) for rigidly attaching an auxiliary component (32) to an engine casing (22). Each side bracket (42, 44) defines mount segment (50, 52), a deformable member (SO) and two retainer members (62, 64) between the mount segments. The deformable member (60) plastically deforms during a fan-blade out event, thereby absorbing a majority of the high shock load experienced on the auxiliary component (32). The retainer members (62, 64) maintain the attachment between the auxiliary component (32) and the engine casing (22) subsequent to the fan-blade out event.

Remote control of an aircraft parking brakes and/or APU

An apparatus 102 for remotely controlling an aircraft 108 having a self-propelled nosewheel 106, on the ground, comprises parking brakes; an APU on said aircraft; control means for applying and releasing said parking brakes 118; control means for turning on and off said APU 116; transmitting means for transmitting information to said parking brakes and said APU; receiving means for receiving information at said parking brakes; and receiving means for receiving information at said APU. The remote control may further comprise means for control the speed of the aircraft, the direction of motion of the aircraft, an emergency stop button and a fire suppression feature.

Inerting arrangement, aircraft comprising the same and a method of providing inert gas

An aircraft battery 2 arranged to provide power to an aircraft electrical system 21, the battery being located in a first fire-resistant container 6 with an inlet 9 for the ingress of inert gas. A fire suppression sub-system 100 is either operable, or will activate in response to a sensor 23 indicative of battery fire or fire risk, in order to provide inert gas to the container. More than one battery/container 3, 7 may be provided, in different parts of the aircraft and powering different systems. A second suppression system 200 may be present, to deliver inert gas to the first or further batteries. The suppression system may deliver oxygen-depleted air or an alkyl halide. The container may include fire-resistant metal. The provision of inert gas to the fire-resistant container may be controlled via a squib valve 17.

Method and apparatus for providing power in an aircraft to one or more aircraft systems

An aircraft comprising a management system for transferring energy

Fuel cell-based auxiliary power unit

An auxiliary power unit (APU) (10), e.g. for an aircraft, providing electrical power to at least one load (30, 31, 38; 124) includes: an array (14; 214) of fuel cells comprising a plurality of sets (210, 212, 226) of series-connected fuel cells (216, 218, 220); at least one switch (222, 224) having a first state condition corresponding to an electrical coupling of at least two of the sets in parallel and a second state condition corresponding to an electrical uncoupling of the at least two of the sets in parallel; and a controller (12) operably coupled to the at least one switch and controlling the change of state of the switch between the first state condition and the second state condition. At least two of the sets are interconnected in parallel when the switch is in the first state condition to increase the current provided by the APU.

A generator having a disconnect mechanism

A generator 1 for an aircraft engine 2, comprising: a rotor 10; a first shaft 100 to which the rotor is mounted; a second shaft 200, at least partially disposed inside the first shaft, coupled to the first shaft by means of a translatable drive connection 40, wherein the translatable drive connection is configured to transfer torque between the first and second shafts, and enable movement of the second shaft along its axis of rotation relative to the first shaft, from an extended position to a retracted position; an actuation means 300 to retract the second shaft; such that an input portion of the second shaft can be at least partially retracted towards the rotor upon activation. A biasing member 50 maybe arranged between the first and second shaft to bias the second shaft towards the extended position. The translatable drive connection may comprise one or more splines 104, 106, on each of the first and second shafts. The first part of the disconnect mechanism may comprises a tapered protrusion ...

Power generation system for an aircraft using a fuel cell

Remote shaft driven open rotor propulsion system with electrical power generation

Method and apparatus for operating a power system architecture

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 ...

Power system for more electric aircraft

A power system (20, fig 2) for an aircraft (10, fig 1) includes a first fuel cell 202 configured to provide base power to one or more loads on the aircraft. The power system further includes a second fuel cell 214 configured to provide peak power to the one or more loads on the aircraft. The peak power can be a power exceeding the base power. The power system further includes an energy storage device 218 configured to provide transient power to the one or more loads on the aircraft. The transient power can be a power exceeding the peak power. The power system further includes a controller (290, fig 2) configured to control delivery of power from the first fuel cell 202, the second fuel cell 214, and the energy storage device 218 to the one or more loads on the aircraft. Methods (400, fig 4) for providing power to one or more loads on an aircraft are also provided.

Improvements in or relating to auxiliary power systems in or for aircraft

... 1,048,130. Refrigerating. PLESSEY-U.K. Ltd. Sept. 27, 1963 [Oct. 22, 1962], No. 39867/62. Heading F4H. Auxiliary power for an aircraft is provided by a turbine or other fluid-pressure motor which is driven by refrigerant vapour evolved during cooling of the load of a refrigeration system, auxiliary heating means being provided for heating the refrigerant when the vapour produced by the load is insufficient to supply the power required. As shown, air from the compressor of an aircraft engine is supplied through a duct 6 to a turbo-compressor 7, cooled in an inter-cooler 14 and expanded in a turbine 16 before being forwarded to an aircraft cabin through duct 12. The cooler 14 comprises the evaporator of a closed condensing-evaporating system, the refrigerant vapour from cooler 14 passing via lines 17 and 18 to drive a turbine 20 which drives an hydraulic power unit 22 and an A.C. generator 21. The speed of turbine 20 is maintained constant by an automatic valve 19. Exhausted refrigerant from ...

A vehicle comprising an engine restart system

Systems, arrangements, structures and methods for aircraft

Direct current voltage regulation of permanent magnet generator

Control system for aircraft

Aircraft power generation system

Aircraft electrical power supply system

... 1,046,030. Engine turning gear. ENGLISH ELECTRIC CO. Ltd. July 16, 1963 [July 20, 1962], No. 27965/62. Heading F1K. [Also in Divisions B7 and H2] An aircraft electric power supply system comprises an electric, e.g. A.C. generator 13 drivably coupled through a constant speed drive 12 selectively to either an air turbine 14 or an associated engine 10 of the aircraft. When a selectively operable disconnect device 11 is engaged the engine 10 can be started through a free wheel 15 by the turbine 14 which is disconnected by the free wheel when the engine is operating and drives the generator. If the engine is inoperative the device 11 is disconnected whereupon the generator is driven by the turbine 14 which is supplied with air under pressure when the aircraft is grounded from either an external supply or an auxiliary source on the aircraft, or when in flight from either the latter source or another main propulsion engine. Fig. 2 shows the gearing interconnecting the system components.

Method and apparatus for modular power distribution

System and method for operating a power distribution system

Electrical Machine and Power Electronics Converter

An integrated electrical machine and power electronics converter arrangement 10, comprising: an electrical machine comprising one or more windings 1221, a power electronics converter 131 arranged to supply or receive current from the windings, a magnetocaloric effect MCE material 16 in thermal contact with the converter, and a heat sink Q for removing heat from the MCE material. The MCE material is arranged in proximity to the windings so when in use, stray magnetic flux from the windings through the MCE material activates the material in order to create magnetic refrigeration which removes heat from the converter. The machine and converter may be in a common housing or connected together in a machine housing 12 and converter housing 13. The heat sink may be arranged to transfer heat removed from the MCE material into a cooling system 17, removing heat from both the machine and the material. The cooling system may comprise conduit in a cooling jacket for cooling fluid in thermal contact ...

High voltage converter for use as electric power supply

An electric power supply 100 has high-voltage circuitry 125 converting a high-voltage, direct-current (HVDC) input 107 to multiphase high-voltage, alternating-current (HVAC) output 108, for example to power an electric appliance (e.g. an electric motor), and low-voltage circuitry 165 to power the auxiliary and control electronics controlling and managing the high-voltage circuitry as well as handling the system monitoring. The high voltage circuitry may comprise one or more DC pre-charge capacitors 116 and one or more power transistor switches 124. The HVDC circuitry is configured to receive HVDC input power of about 320 volts and/or greater and convert the HVDC input power to multi-phase HVAC output power of about 320 volts and/or greater. The low-voltage circuitry is configured to operate on low-voltage, direct-current (LVDC), and is configured to control and monitor the multi-phase HVAC output power. DC to DC converter circuitry 130 is configured to receive and convert the HVDC input ...

INLET DOOR

CURRENT SUPPLY CIRCUIT IN an AIRCRAFT FOR the ELECTRICAL DEVICES including SNOW AND ICE REMOVAL CIRCUIT

Aircraft electric motor system

AIRCRAFT ELECTRIC MOTOR SYSTEM A method and apparatus for controlling an electric aircraft is presented. An apparatus comprises a controller (400). The controller is configured to identify a state for an electric aircraft (200). The controller is further configured to identify a group of recharging parameters for a group of electric motors (412-418) in an electric propulsion system. The electric propulsion system is configured to move the electric aircraft based on the state for the group of electric motors for the electric aircraft. The controller is still further configured to recharge a power source (402) for the electric aircraft using the group of recharging parameters (420-425) to control recharging of the power source with the group of electric motors when a recharge state (410) is present for the electric aircraft. 'P. CN ~ -. \/ "/ V- ~ ...

Method for Monitoring Performance of the Lubricant Cooler in Aircraft Auxiliary Power Unit

Abstract The present invention relates to a method for detecting performance of the lubricant cooler in aircraft auxiliary power unit APU, comprising: acquiring APU-related messages within a time period; obtaining the s operation parameters of the APU lubricant cooler, the operation parameters comprise: lubricant temperature OT and load compressor inlet temperature LCIT; the revised lubricant temperature OT is obtained by the following formula: the revised lubricant temperature = lubricant temperature OTA - the load compressor inlet temperature LCIT; determining the performance of the o APU lubricant cooler is in stable phase, decline phase or malfunction phase according to the change trend of the revised lubricant temperature OT with respect to time. Figure 2 ...

ELECTRICAL POWER DISTRIBUTION SYSTEM AND ALLOCATION METHOD FOR DISTRIBUTING ELECTRICAL POWER

ELECTRICAL POWER DISTRIBUTION SYSTEM AND ALLOCATION METHOD FOR DISTRIBUTING ELECTRICAL POWER An electrical power distribution system has at least one converter module having a converter, which is designed to convert AC voltage from one or more AC voltage sources to DC voltage and provide DC voltage power with adjustable maximum power values at a plurality of electrical output interfaces of the converter module 10 up to a maximum module power value. The power distribution system also comprises a power profile management device, which is coupled to the converter and which is designed to negotiate individual power profiles with electrical consumers that can be connected to 15 respective ones of the plurality of electrical output interfaces, according to which individual power profiles electrical power up to a negotiated maximum power value is provided to the consumers by the converter via the electrical output interface and which are stored in a power profile memory of the power profile management ...

EMERGENCY HYDRAULIC POWER SYSTEM (START BOTTLE)

AUXILIARY POWER UNIT INTAKE DUCT WITH AERO-ACOUSTIC GUIDE VANES

An auxiliary power unit intake duct for an aircraft, comprising an internal straight portion 1a, an internal first bend portion 1b and an internal second bend portion 1c, extending between an air inlet 2 and a plenum entry 3 of the auxiliary power unit, a plurality of acoustic guides vanes 4,5 located in the straight portion la of the intake duct 1, extending horizontally between opposite side portions of the straight portion 1a; and a plurality of curved aerodynamic guide vanes 6,7 located in the second bend portion 1c proximate to said plenum inlet 3, extending vertically between a top portion and a bottom portion of the second bend portion 1c, each of the curved aerodynamic guide vanes 6,7 being concentric with said bend curvature of the second bend portion 1c.

REMOTE SHAFT DRIVEN OPEN ROTOR PROPULSION SYSTEM WITH ELECTRICAL POWER GENERATION

A system for aircraft propulsion is disclosed herein. The system includes a power plant. The system also includes an open rotor module operable to rotate. The open rotor module has a plurality of variable-pitch blades. The system also includes a first linkage extending between the power plant and the open rotor module. The first linkage is operable to transmit rotational power to the open rotor module for rotating the plurality of variable-pitch blades. The system also includes an actuator operable to change a pitch of the plurality of variable-pitch blades. The system also includes a generator operable to generate electric power. The system also includes a second linkage extending between the power plant and the generator. The second linkage is operable to transmit rotational power to the generator. The generator is operable to convert the rotational power to electrical power. The system also includes a controller operably coupled to the actuator to vary a pitch of the plurality of variable-pitch ...

POWERPLANT AND RELATED CONTROL SYSTEM AND METHOD

A hydrogen fueled powerplant including an internal combustion engine that drives a motor-generator, and has a two-stage turbocharger, for an aircraft. A control system controls the operation of the motor-generator to maintain the engine at a speed selected based on controlling the engine equivalence ratio. The control system controls an afterburner, an intercooler and an aftercooler to maximize powerplant efficiency. The afterburner also adds power to the turbochargers during high-altitude restarts. The turbochargers also include motor-generators that extract excess power from the exhaust.

AIRCRAFT COMPRISING A REVERSIBLE ROTARY ELECTRICAL MACHINE

The invention concerns an aircraft comprising a reversible rotary electrical machine 30), a wheel (35) for imparting movement to a stream of air, linked to said machine (30), a first ventilation duct (10, 15) configured to supply said wheel (35) with air, an electricity distribution network (70), a reversible power converter (40) connected to said machine (30) and to said network (70), and a control circuit (50) configured to control said converter (40) for, in a motor operating mode of said machine (30), taking a supply current from said network (70) to supply said machine (30) and drive said wheel (35), and, in an alternator operating mode of said machine (30), in which a stream of air flows in said first ventilation duct (15) and imparts movement to said wheel (35), to provide supply power to said network (70) which is generated by said machine (30) driven by said wheel (35).

GYROGLIDER POWER-GENERATION, CONTROL APPARATUS AND METHOD

A power generation apparatus and method comprises at least one gyroglider rotary wing flying at an altitude above the nap of the earth. A strong and flexible tether, connected to the gyroglider frame is pulled with a force generated by the rotary wing. The force is transmit-ted to a ground station that converts the comparatively lin-ear motion of the tether being pulled upward with a lifting force. The linear motion is transferred to a rotary motion at the ground station to rotate an electrical generator. The tether is retrieved and re-coiled about a drum by control-ling the gyroglider to fly down at a speed and lift force that permit recovery of the gyroglider at a substantially re-duced amount of retrieval force compared to the lifting force during payout of the tether. Thus, the net difference in force results in a net gain of energy.

MULTIFUNCTION INTEGRATED POWER UNIT

MULTIFUNCTION INTEGRATED POWER UNIT A multifunction integrated power unit (MIPU) for use aboard aircraft provides the functions both of an auxiliary power unit and of an emergency power unit while being smaller, lighter, and less expensive than the two units it replaces. Additionally, the integration of both functions in a single unit allows elimination of duplicated service devices and a further reduction in size, cost, and weight in the aircraft itself. The MIPU includes two turbine engines operable under differing conditions, an integrating gear box receiving power from the two engines selectively, and a plurality of accessory devices receiving shaft power from either of the two engines via the gear box.

SYSTEMS AND METHODS FOR CONTROLLING AIRCRAFT ELECTRICAL POWER

Systems and methods for controlling aircraft electrical power are disclosed. A system in accordance with one embodiment includes an electric motor, an aircraft load coupled to the electric motor and powered by the electric motor, and a motor controller coupled to the electric motor to vary an output of the electric motor. The motor controller is changeable among a fixed number of preset controller modes, with individual controller modes corresponding to an operation mode of the aircraft and a non-zero output level of the electric motor. In further embodiments, motor controllers can be substituted for each other, e.g., in the event one motor controller becomes unoperational.

FUEL CELL EMERGENCY SYSTEM

... ²²²In contemporary aircraft, ram air turbines are used to provide emergency power ²given a failure of the engines or hydraulic systems. The ram air turbine ²system is mechanically complex in light of the out-folding mechanism and ²rotating components, and requires comprehensive maintenance. In an exemplary ²embodiment of the present invention, a fuel cell system is indicated for ²supplying power to an aircraft, comprising a fuel cell, a hydrogen tank and an ²oxygen tank. This advantageously provides for an emergency power supply for ²aircraft that is independent of the outside air, low maintenance and reliable.² ...

SYSTEMS FOR HARVESTING ROTATIONAL WHEEL ENERGY FOR LANDING GEAR RETRACTION

AB STRACT A system for extracting energy for landing gear retraction may comprise a wheel pump rotationally coupled to a wheel via a pinion gear. A landing gear control valve assembly may be fluidly coupled to an output of the wheel pump. A secondary pump may be fluidly coupled to the landing gear control valve assembly, and an electric motor may be operationally coupled to the secondary pump. Date Recue/Date Received 2020-1 2-2 1 ...

AIRCRAFT ENGINE WITH CLUTCH AND MECHANICAL LOCK

There is disclosed an aircraft engine assembly including an engine having an engine shaft; an output shaft; a clutch in driving engagement between the engine shaft and the output shaft. The clutch has a first component in driving engagement with the engine shaft and a second component. The clutch is operable between first and second configurations. In the first configuration, the first component is rotatable relative to the second component and the engine shaft is rotatable relative to the output shaft. In the second configuration, the first and second components are engaged with one another and the engine shaft rotates with the output shaft. A mechanical lock is operable between first and second positions. In the first position, the mechanical lock is disengaged from the first component. In the second position, the first and second components are secured for joint rotation one relative to the other.

AIRCRAFT POWER PLANT

There is disclosed a cooling system for a liquid cooled internal combustion power plant housed in an engine compartment in a tail cone of an aircraft. The cooling system has: a tail cone inlet defined through a wall of the tail cone and fluidly communicating with an environment; a wall inlet defined through a firewall of the engine compartment; a blower within the engine compartment and having a blower inlet and a blower outlet, the blower inlet fluidly communicating with the environment via the tail cone inlet, via the wall inlet, and via an interior of the engine compartment; a blower outlet defined through a wall of the aircraft and fluidly communicating with the environment; and a cooling flow path extending from the tail cone inlet to the air outlet and across the wall inlet, the cooling flow path in heat exchange relationship with the power plant.

GAS TURBINE AUXILIARY POWER UNIT

A gas turbine engine includes a supply air conduit fluidly connected to an inlet of a combustion chamber to convey supply air to the combustion chamber, an exhaust conduit fluidly connected to an exit of the combustion chamber to convey exhaust gases away from the combustion chamber, and a recuperator defining a first flow path and a second flow path therethrough, the second flow path being in heat transfer communication with the first flow path, the first flow path defining part of the supply air conduit, the second flow path defining part of the exhaust conduit. Methods of operating the engine are also provided.

AIRPLANE WITH IMPROVED SAFETY

An airplane has main propulsion engines and a first fuel supply for the main propulsion engines. The airplane further has an auxiliary propulsion engine and a second fuel supply for the auxiliary propulsion engine, this second fuel supply being separate from the first fuel supply. The auxiliary propulsion engine can be switched on independently from the main propulsion engines. Such airplane has increased safety, since it will be possible to maintain flight, particularly when at high altitude, even if all main propulsion engines have failed.

TRANSLATABLE SCOOP FOR AUXILIARY POWER UNITS AIR INTAKE

An air intake assembly for an aircraft includes an auxiliary power unit intake that is defined by a skin of the aircraft. The auxiliary power unit intake is in fluid communication with an auxiliary power unit. The air intake assembly also includes a translatable scoop that defines a stowed condition and a fully deployed condition. The translatable scoop directs air exterior to the aircraft into the auxiliary power unit intake when in the fully deployed condition and does not direct the air exterior to the aircraft into the auxiliary power unit intake when in the stowed condition.

AUXILIARY AIR SUPPLY FOR AN AIRCRAFT

A secondary power unit for an aircraft comprises: a gas turbine engine (14) having at least one drive shaft, a twin load compressors (3) including first and second load compressors both coupled with the drive shaft, a first Ram-air intake (4) in fluid communication with the twin-load compressor, and flow regulating means (6) coupled with the first and second compressors outputs to individually regulate the generated flow of compressed air. The flow regulating means are also adapted to be fluidly communicated with an Air Cycle Machine (ACM) of an aircraft, and a control system is adapted to feed the flow of compressed air as bleed air to the Air Cycle Machine (ACM) of an aircraft both when the aircraft is on ground and in flight. The secondary power unit is capable of avoiding bleed air extraction from the main engines with the aim of reducing fuel consumption.

POWER SUPPLY SYSTEM

According to the present invention: each of a plurality of power generators is configured so that a frequency relationship with power-generator effective power that is output to a corresponding AC wiring section by each of the power generators has a prescribed first drooping characteristic; a plurality of power conversion devices are configured to convert, into DC power, AC power that is input through each of the AC wiring sections and to convert, into AC power, DC power that is input through a DC wiring section; and control devices are configured to determine a target value of a control factor so that a frequency relationship with power-conversion-device effective power that is output to the corresponding AC wiring section by each of the power conversion devices has a prescribed second drooping characteristic and to generate a driving signal for each of the power conversion devices by correcting the target value of the control factor according to a DC voltage in the DC wiring section.

ELECTRIC POWER SUPPLY METHOD AND ARCHITECTURE FOR ON-BOARD DOMESTIC NETWORK

L'architecture réseau de puissance électrique embarquée d'une cabine de véhicule selon l'invention utilisant des équipements domestiques, en particulier d'une cabine d'aéronef, comporte des réseaux primaires (11, 12) connectés en parallèle à un unique réseau domestique cabine (3) d'alimentation d'équipements domestiques via des ensembles convertisseurs (41, 42; 40). Des capteurs de mesure (80) sont agencés en sortie de chaque ensemble convertisseur (41, 42; 40) et les mesures sont transmises (L1) à un moyen régulateur (9) de partage d'alimentation du réseau domestique (3) en fonction de la puissance disponible, chaque ensemble convertisseur (41, 42; 40) comporte un étage de traitement amont (E1) et un étage de traitement aval (E2), intégrant chacun un dispositif anti-perturbation (D1, D2), respectivement en liaison (B1, B2) avec le réseau primaire (11, 12) et en liaison (B3, B4) avec le réseau domestique (3). Les dispositifs anti-perturbation (D1, D2) sont couplés à un ensemble (7) d'adaptateurs ...

A COMBINATION OF A HEAT GENERATING SYSTEM WITH A FUEL CELL SYSTEM

The invention relates to a system utilizing synergetic potentials of different systems, e.g. of a fuel cell system and of an aircraft hydraulic system, through coupling them via a common cooling circuit system (4).

ELECTRIC-BASED SECONDARY POWER SYSTEM ARCHITECTURES FOR AIRCRAFT

Methods and systems for providing secondary power to aircraft systems. In one embodiment, an aircraft system architecture (300) for providing power to an environmental control system (340a, 340b) includes an electric generator (316a, 316b) operably coupled to a jet engine (310a, 310b). The jet engine (310a, 310b) can be configured to provide propulsive thrust to the aircraft, and the electric generator (316a, 316b) can be configured to receive shaft power from the jet engine (310a, 310b). The environmental control system (340a, 340b) can be configured to provide outside air to a passenger cabin (320) of the aircraft in the absence of bleed air from the jet engine.

AIRCRAFT THERMAL MANAGEMENT SYSTEM

A thermal management system for an aircraft is provided. The thermal management system may comprise a first vapor compression circuit, a second vapor compression circuit, and an intercooler. The first vapor compression circuit may define a first flowpath for fluid compression, condensation, expansion, and evaporation. The second vapor compression circuit may define a second flowpath for fluid compression, condensation, expansion, and evaporation. The intercooler may be disposed in cascading thermal communication between the first vapor compression circuit and the second vapor compression circuit. Generally, heat generated by the aircraft may be transferred to the first vapor compression circuit during aircraft operation.

SYSTEM FOR RECOVERING AND CONVERTING KINETIC ENERGY AND POTENTIAL ENERGY INTO ELECTRICAL ENERGY FOR AN AIRCRAFT

L'invention concerne un système de génération d'énergie électrique (20) pour aéronef (10) comprenant un carénage aérodynamique (21) renfermant au moins une turbine (22) logée dans la partie avant (21a) du carénage (21) et un générateur d'énergie électrique (23) relié à ladite turbine. La partie avant (21a) du carénage (21) est équipée de moyens d'admission d'air (26) qui sont mobiles entre une position d'ouverture dans laquelle la turbine (22) est exposée au flux d'air externe au carénage (21) et une position de fermeture dans laquelle la turbine (22) est masquée dans le carénage (21). L'invention permet de réduire la traînée aérodynamique habituellement créée par des tourbillons présents en bout d'aile dans le cas d'une voilure classique à bords nets durant les phases de décollage, montée et croisière et de récupérer lors des phases de descente les énergies cinétique et potentielle accumulées par l'aéronef au cours des phases de montée et de croisière.

AIRCRAFT AND SYSTEM FOR SUPPLYING ELECTRICAL POWER TO AN AIRCRAFT ELECTRICAL LOAD

An aircraft, comprising: a wing; a fuselage coupled to the wing; an engine (34) coupled to at least one of the fuselage and the wing; an electrical load (20) associated with the aircraft during flight operations; a generator (32) coupled to the engine and configured to generate electrical power for the electrical load; a conductor (38) electrically disposed between the electrical load and the generator; a conduit (46) configured to house the conductor; and a dielectric gas (44) disposed in the conduit; wherein the conduit (46) is configured to envelop the conductor (38) in the dielectric gas (44).

AUXILIARY POWER UNIT ASSEMBLY WITH REMOVABLE INLET FILTER

An auxiliary power unit assembly for an aircraft, including an engine assembly including an internal combustion engine, a first duct in fluid communication with an inlet of the internal combustion engine and configured to be in fluid communication with an environment of the aircraft, and a second duct configured to be in fluid communication with the environment. The second duct may be in fluid communication with an inlet of a compressor drivable by the internal combustion engine or with a compartment receiving the engine assembly. A filter extends across the first duct. The filter is removable from the first duct through an opening in a wall of the first duct, the opening accessible from an exterior of the engine assembly.

ELECTRIC TAXING SYSTEM OF AIRCRAFT AND METHOD FOR CONTROLLING THE SYSTEM

An electrically powered taxiing system for an aircraft (1), said system being provided with: wheels (5a, 7a) of the aircraft (1); taxiing motors (5b, 7b) that, when the aircraft (1) lands, apply a braking force to the wheels (5a, 7a) on being turned with the wheels (5a, 7a) and generate regenerative power corresponding to the braking force, and that, when the aircraft (1) is taxiing, cause the wheels (5a, 7a) to turn; and a power supply control unit (27) for feeding the regenerative power to electrically powered elements (9b, 11b, 13b, 15b, 29a, 31a) of the aircraft (1), which are driven for landing of the aircraft (1) or for performing a second takeoff in an aborted landing.

AIRCRAFT FUEL CELL SYSTEM WITH CATALYTIC BURNER SYSTEM

Disclosed are fuel cell systems used as power sources aboard aircraft and utilizing catalytic systems. Fuel cell systems can include a fuel cell assembly and a catalyst system. The fuel cell assembly can receive a hydrogen input, receive an oxygen input comprising a fluid having an initial oxygen content, and convert the hydrogen input and the oxygen input so as to yield products, such as water, thermal energy, an oxygen-depleted product comprising the fluid having a second oxygen content lower than the initial oxygen content, and electrical power. The fuel cell assembly can supply any combination of such products to aircraft operational systems. The catalyst system can receive and combust hydrogen from the fuel cell assembly and/or a hydrogen storage vessel, such as to treat exhaust from the fuel cell assembly and/or provide heat for warming water and/or for regulating operating temperatures of fuel cell system components.

HYBRID SOLAR GENERATOR

A solar generator can include a photon-enhanced thermionic emission generator with a cathode to receive solar radiation. The photon-enhanced thermionic emission generator can include an anode that in conjunction with the cathode generates a first current and waste heat from the solar radiation. A thermoelectric generator can be thermally coupled to the anode and can convert the waste heat from the anode into a second current. A circuit can connect to the photon-enhanced thermionic emission generator and to the thermoelectric generator and can combine the first and the second currents into an output current.

COMPOUND ENGINE ASSEMBLY WITH BLEED AIR

A compound engine assembly for use as an auxiliary power unit for an aircraft and including an engine core with internal combustion engine(s), a compressor having an outlet in fluid communication with an engine core inlet, a bleed conduit in fluid communication with the compressor outlet through a bleed air valve, and a turbine section having an inlet in fluid communication with the engine core outlet and configured to compound power with the engine core. The turbine section may include a first stage turbine having an inlet in fluid communication with the engine core outlet and a second stage turbine having an inlet in fluid communication the first stage turbine outlet. A method of providing compressed air and electrical power to an aircraft is also discussed.

HYBRID GAS-ELECTRIC TURBINE ENGINE

A hybrid gas-electric turbine engine for turboprop or turboshaft applications is disclosed together with associated methods. In various embodiments disclosed herein, the turbine engine comprises a turbine configured to be driven by a flow of combustion gas; a turbine shaft configured to be driven by the turbine and transfer power to a load coupled to the turbine engine and an electric motor configured to transfer power to the load coupled to the turbine engine. The rotor may have a rotor axis of rotation that is radially offset from a shaft axis of rotation of the turbine shaft. In some embodiments, the electric motor may be a multi-rotor electric motor.

AIRCRAFT LOCATION SYSTEM FOR LOCATING AIRCRAFT IN WATER ENVIRONMENTS

A method and apparatus for an aircraft location system comprising an aircraft structure and a number of acoustic reflectors associated with the aircraft structure. The number of acoustic reflectors is configured to generate first sound signals in response to receiving second sound signals.

Aircraft, propulsion system, and system for taxiing an aircraft

One embodiment of the present invention is a unique aircraft. Another embodiment is a unique aircraft propulsion system. Still another embodiment is a unique system for taxiing an aircraft without starting one or more main aircraft propulsion engines. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for aircraft taxiing and propulsion systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Systems and methods for harvesting vibrational energy from vehicles

Disclosed are methods, systems and apparatuses for harvesting vibrational energy from vehicles, as well as methods for locating vibrational energy on a vehicle. One harvesting apparatus comprises a vibrational energy harvester coupled to a substantially maximal vibrational displacement node of the structural element of a vehicle, and tuned to a frequency of vibration of the maximal vibrational displacement node. The harvester may be one of a variety of harvesting devices, such as piezoelectric devices. Methods are also provided to locate structural elements which are appropriate for vibrational energy harvesting.

Ram air turbine inlet

An inlet for a ram air turbine. The inlet is positioned forward of a turbine shaft, which is configured to rotate about a portion of an object. The inlet has a fixed geometry that chokes airflow through the inlet to prevent overspeed of the turbine shaft.

Exhaust silencer with baffles

An exhaust silencer assembly includes an exhaust duct and an exhaust silencer. The exhaust silencer is disposed about the exhaust duct and has a plurality of solid baffles and at least one perforated baffle The exhaust silencer assembly is disposed downstream of an auxiliary power unit and at least partially attenuates the downstream noise of the combustion gases that result from operation of the auxiliary power unit. In one embodiment, at least one of the plurality of solid baffles is disposed to extend axially between generally radially extending solid baffles, perforated baffle(s), or partially perforated baffles. The axially extending solid baffle allows for cavity depth variation in the silencer to optimize tuning of particular frequencies from the auxiliary power unit.

Generator Frequency Controlled Load Shifting

In a power generation system that includes a ram air turbine that drives an electric generator and at least one electric bus that couples alternating current electric power from the electric generator to at least one electric load, a method of maintaining the speed of the ram air turbine within a desired range of speeds that comprises the step of uncoupling the at least one electric bus from the electric generator when the frequency of the alternating current falls under a desired minimum frequency.

Submerged ram air turbine generating system

A submerged ram air turbine generating system adapted for use in a pod mounted to the wing of an aircraft comprises a submerged inlet extending substantially entirely around the outer surface of the pod housing, a stator in the form of adjustable or fixed inlet guide vanes which direct an air stream to a hybrid ram air turbine having alternating turbine blades and splitters, and, a number of adjustable exhaust panels which are movable with respect to one or more exhaust openings in the pod housing.

Supplementary power supply for vehicles, in particular aircraft

The invention relates to a supplementary power supply including a solar generator for ohmic consumers in a vehicle, in particular an aircraft, wherein a control circuit is electrically connected to a multiplicity of solar cells via generator terminals of the solar generator and includes at least one output terminal, and wherein the control circuit is configured to interconnect the at least one output terminal with the multiplicity of solar cells in accordance with an actual intensity of insolation of the solar cells such that a nominal voltage and/or a nominal current intensity is available at the at least one output terminal.

Fire resistant structural mount yoke and system

A mount yoke assembly comprises a first main hanger bracket and a first anchor bracket. The first main hanger bracket includes a first receiving portion for a first auxiliary power unit (APU) mount with first and second hanger arms. The first hanger arm includes a first hanger flange extending proximate a distal end of the first hanger arm. The first anchor bracket includes a first plurality of anchor holes and a first anchor flange. The first plurality of anchor holes are disposed substantially along a length of the first anchor bracket for securing the mount yoke assembly to a component of an APU assembly. The first anchor flange extends from the first anchor bracket for hingeably securing the first anchor bracket to the first main hanger bracket.

Aircraft Location System for Locating Aircraft in Water Environments

A method and apparatus for an aircraft location system comprising an aircraft structure and a number of acoustic reflectors associated with the aircraft structure. The number of acoustic reflectors is configured to generate first sound signals in response to receiving second sound signals.

Energy Control Apparatus For Controlling Hybrid Energy Sources For An Aircraft

An energy control apparatus for controlling hybrid energy sources for an aircraft is designed for determining an operating characteristic of an energy source, a demand of at least one first consumer for a first resource and a demand of a second consumer for a second resource. The energy control apparatus activates the energy source in dependence on the demand for the first resource and the second resource and, if there is no demand for a first or second resource that is generatable by the energy source and is present in excess, supplies a third consumer with the excess resource. This makes it possible for an energy system to operate different consumers independently of the type of generated resources and their dependencies during the generation. 1. An energy control apparatus for controlling hybrid energy sources for an aircraft ,wherein the energy control apparatus is configured for determining an operating characteristic of an energy source;wherein the energy control apparatus is further configured for determining a demand of at least one first consumer for a first resource and a demand of a second consumer for a second resource;wherein at least the first resource and the second resource are simultaneously generatable by at least one energy source;wherein the energy control apparatus is configured to activate the energy source in dependence on the demand for the first resource and the second resource, andwherein if there is no demand for the first or second resource generatable by the energy source and present in excess, the energy control apparatus is configured to supply the excess first or second resource to a third consumer.2. The energy control apparatus of claim 1 , wherein the third consumer is configured for being supplied with a third resource.3. The energy control apparatus of claim 1 ,wherein a resource is selected from the group consisting of electrical energy and at least one byproduct;wherein the at least one byproduct is selected from the group ...

METHOD FOR OPTIMIZING THE OVERALL ENERGY EFFICIENCY OF AN AIRCRAFT, AND MAIN POWER PACKAGE FOR IMPLEMENTING SAME

A method and system limiting specific consumption of an aircraft by matching sizing of a power supply to actual power needs of a cabin pressure control system. The method optimizes overall efficiency of energy supplied onboard an aircraft including, in an environment near the cabin, at least one main power-generating engine, sized to serve as a single pneumatic energy-generating source for the cabin and as an at most partial propulsive, hydraulic, and/or electric energy-generating source for the rest of the aircraft. The method minimizes power differential between a nominal point of the power sources when the sources are operating, and a sizing point of non-propulsive energy contributions of the sources when the main engine has failed, by equally dividing power contributions of the main engines and the main power generator under nominal operating conditions and in an event of failure of a main engine. 115-. (canceled)16. A method for optimizing overall efficiency of energy supplied aboard an aircraft , the energy being propulsive or non-propulsive , the aircraft including a passenger cabin with regulated air flow , and power sources including main engines , the method comprising:providing, in an environment located near the cabin, at least one main engine-type power-generating means sized to serve as a single other pneumatic energy-generating source for the cabin and at most partly as an other propulsive, hydraulic, and/or electric energy-generating source for a rest of the aircraft; andminimizing a power difference between a nominal point of the power sources when the power sources are functioning and a sizing point of non-propulsive energy contributions of the power sources in a situation of failure of a main engine, by equally dividing power contributions of the main engines and main power-generating means under nominal operating conditions and in event of the failure of a main engine.17. A main power unit for implementing the optimizing method according to in an ...

AIRCRAFT AIR INLET DIVERTER ASSEMBLIES WITH IMPROVED AERODYNAMIC CHARACTERISTICS

Diverter assemblies for aircraft air inlets include a diverter structure at least substantially surrounding the air inlet, and a fairing mounted to an upper edge of the diverter structure forwardly of the air inlet. 1. A diverter assembly for an aircraft air inlet comprising a diverter structure at least substantially surrounding the air inlet , and a fairing mounted to an upper edge of the diverter structure forwardly of the air inlet.2. The diverter assembly as in claim 1 , wherein the fairing is a solid structure which defines a relatively smooth ramp which slopes aftward toward the in the air inlet to thereby avoid airflow detachment forwardly of the air inlet.3. The diverter assembly as in claim 1 , wherein the air inlet includes an air scoop extending outwardly therefrom.4. The diverter assembly as in claim 1 , wherein the diverter structure has a substantially S-shaped cross-section.5. The diverter assembly as in claim 1 , comprising a fairing bracket positioned along a forward edge of the air inlet claim 1 , wherein an aft edge of the fairing is attached to the fairing bracket.6. The diverter assembly as in claim 1 , wherein the fairing comprises a forward apex region claim 1 , a concavely curved aft region defining a corresponding concavely curved aft edge and laterally divergent side regions joining the forward apex and aft regions.7. The diverter assembly as in claim 6 , comprising an S-shaped fairing bracket positioned along a forward edge of the air inlet.8. The diverter assembly as in claim 7 , wherein the fairing bracket includes a lower base flange claim 7 , an upper support flange and an intermediate support member which rigidly joins the lower base and upper support flanges to one.9. The diverter assembly as in claim 8 , wherein the upper support flange is positioned rearwardly of the lower base flange so that the aft edge of the fairing overhangs a forward edge of the air inlet.10. The diverter assembly as in claim 3 , wherein the diverter is a ...

POWER SUPPLY SYSTEM WITH HETEROGENEOUS MULTIPLE INPUT

The invention relates to a power supply system that is adapted be connected in input to N distinct power supply grids (R, . . . , R) delivering a direct or alternating input voltage (V, V) and to deliver a direct distribution voltage (V), N being an integer greater than or equal to 2, of the type comprising a distribution bus operating under the distribution voltage (V), N unidirectional converters (C, . . . , C) that can respectively be connected in input to a given power grid (R, . . . , R), and control means capable of controlling the unidirectional converters (C, . . . C). The power supply system also includes an energy storage device connected to the distribution bus and having a quantity of energy (E), and the control means are adapted to control the unidirectional converters (C, . . . C) simultaneously to regulate the electrical power (P, . . . P) delivered by each of said unidirectional converters (C, . . . , C) as a function of the quantity of energy (E). 1. A power supply system adapted to be connected in input to N distinct power supply grids (R , . . . , R) delivering a direct or alternating input voltage (V , . . . , V) , and to deliver a direct distribution voltage (V) , N being an integer greater than or equal to 2 , of the type comprising:{'sub': 'BUS', 'a distribution bus operating under the distribution voltage (V),'}{'sub': 1', 'N', '1', 'N', '1', 'N, 'N unidirectional converters (C, . . . , C) adapted to be respectively connected in input to a given power grid (R, . . . , R) and adapted to deliver an electrical power (P, . . . , P) to the distribution bus, and'}{'sub': 1', 'N, 'control means adapted to control the unidirectional converters (C, . . . , C),'}{'sub': 1', 'N', '1', 'N', '1', 'N, 'characterized in that the power supply system also includes an energy storage device connected to the distribution bus and having a quantity of energy (E), and in that the control means are adapted to control the unidirectional converters (C, . . . , C) ...

POWER SUPPLY CIRCUIT FOR AN AIRCRAFT DE-ICING SYSTEM

An aircraft electrical supply circuit including: a power distribution network, onboard the aircraft, for electrical devices located in an engine of the aircraft or near the engine; and a power supply generator built into the engine of the aircraft to supply AC voltage power to a deicing or anti-icing system. The power supply generator is connected to a thrust reverser electromechanical actuator by a rectifier for supplying DC voltage power to the actuator. 17-. (canceled)8. An aircraft comprising:an aircraft engine; andan electrical power supply circuit, the electrical power supply circuit including an electrical power distribution network, onboard the aircraft, for electrical devices located in an engine of the aircraft or in an environment of the engine, and a power supply generator integrated into the engine of the aircraft and connected to a deicing or anti-icing system to supply AC voltage power to the deicing or anti-icing system,wherein the power supply generator is connected to a thrust reverser electromechanical actuator through a rectifier for supplying DC voltage power to the actuator.9. The aircraft according to claim 8 , wherein the rectifier is a diode bridge.10. The aircraft according to claim 8 , wherein the network is connected to the actuator through a second rectifier.11. The aircraft according to claim 8 , further comprising an electronic engine control unit configured to regulate the AC voltage supplied by the power generator to control closure of a first switch located between the power supply generator and the actuator when the AC voltage reaches a predetermined level.12. The aircraft according to claim 11 , wherein the electronic engine control unit is further configured to control a second switch located between the power supply generator and the deicing or anti-icing system.13. The aircraft according to claim 11 , wherein the electronic engine control unit is further configured to control the deicing or anti-icing system to operate with ...

AUXILIARY POWER SYSTEM

This invention relates to an aircraft having an auxiliary power system, the auxiliary power system, including: a reciprocating engine having an air inlet and an exhaust system; a compressed air source for providing compressed air to the air inlet; and, an electrical generator for providing electrical power. 1. An aircraft having an auxiliary power system , the auxiliary power system , comprising:a reciprocating engine having an air inlet and an exhaust system;a compressed air source for providing compressed air to the air inlet; and,a main electrical generator which is driveably connected to the reciprocating engine,wherein the exhaust system includes an energy recovery system.2. An aircraft as claimed in wherein the compressed air source is the passenger cabin of the aircraft.3. An aircraft as claimed in claim 1 , wherein the energy recovery system is a turbine claim 1 ,4. An aircraft as claimed in further comprising a second electrical generator which is coupled to and driven by the energy recovery system.5. An aircraft as claimed in further comprising an inlet compressor downstream of the compressed air source for further compressing the compressed air.6. An aircraft as claimed in wherein the inlet compressor is driven by an electrical machine.7. An aircraft as claimed in wherein the electrical machine which drives the inlet compressor receives electrical power from the second electrical generator.8. An aircraft as claimed in wherein the compressor electrical machine and second electrical generator are mounted on a common shaft.9. An aircraft as claimed in further comprising at least one duct for providing air from the passenger cabin to the air inlet claim 1 , the duct including an acoustic system for reducing noise from the air inlet entering the cabin.10. An aircraft as claimed in claim 1 , further comprising a main compressor for pressurising the compressed air source claim 1 , wherein the exhaust recovery system is arranged to power the main compressor.11. ...

Flyweight for ram air turbine

A flyweight for use in a ram air turbine has a body with a circumferentially enlarged portion extending between flat sides spaced by an angle of between 45 degrees and 150 degrees. A pivot point is defined by a bore at a circumferentially intermediate point in the body, and in a relatively thin inner portion. The relatively thin inner portion extends beyond the pivot point to an end that will be radially inward when the flyweight is mounted in a ram air turbine. A governor, a ram air turbine and a method are also described.

HYDROGEN GENERATION BY HYDROGENATED POLYSILANES FOR OPERATING FUEL CELLS

A fuel cell supply device that generates hydrogen for fuel cells in an aircraft includes a reaction chamber which reacts hydrogenated polysilanes or mixtures thereof with water; a feed device that feeds at least one reactant into the reaction chamber; and a discharge device that leads hydrogen formed in the reaction to a fuel cell. 115.-. (canceled)16. A fuel cell supply device that generates hydrogen for fuel cells in an aircraft comprising:a reaction chamber which reacts hydrogenated polysilanes or mixtures thereof with water;a feed device that feeds at least one reactant into the reaction chamber; anda discharge device that leads hydrogen formed in the reaction to a fuel cell.17. The fuel cell supply device according to claim 16 , wherein hydrogenated polysilanes or mixtures thereof are present in a solid structure selected from the group consisting of powder claim 16 , granules claim 16 , beads claim 16 , cubes claim 16 , pellets and porous pieces.18. The fuel cell supply device according to claim 16 , wherein the reaction chamber provides a basic alkaline medium.19. The fuel cell supply device according to claim 16 , wherein the reaction chamber contains metal catalysts.20. The Fuel cell supply device according to claim 16 , further comprising a waste heat recuperator which utilizes heat released in the reaction.21. The fuel cell supply device according to claim 16 , further comprising a water storage tank that supplies water to the feed device.22. The fuel cell supply device according to claim 16 , wherein the feed device condenses and feeds in water to the water storage to and the water is formed as water of reaction in the fuel cell.23. The fuel cell supply device according to claim 16 , wherein the discharge device comprises a pressure vessel with an inlet valve and an outlet valve for intermediate storage of the hydrogen produced.24. The fuel cell supply device according to claim 16 , wherein the discharge device comprises a hydrogen measuring device that ...

Warming feature for aircraft fuel cells

A system and method for warming a fuel cell on an aircraft, the system includes at least one fuel cell. The fuel cell includes an anode and a cathode for creating thermal and electrical energy. A temperature sensor measures a first temperature of the fuel cell. A control unit is coupled to the temperature sensor. The control unit increases the first temperature to a second temperature in response to the first temperature being at least equal to a selected temperature threshold. Increasing of the first temperature is indicative of the control unit operating in a warming mode. The second temperature is higher than the selected temperature threshold.

HYBRID POWER SYSTEM ARCHITECTURE FOR AN AIRCRAFT

A hybrid power distribution system for an aircraft generates hydraulic power from one of a plurality of power sources based on which power source provides energy most efficiently. Power sources includes an electric power distribution bus that distributes electrical energy onboard the aircraft, a pneumatic distribution channel that distributes pneumatic energy onboard the aircraft, and mechanical power provided by one or more engines associated with the aircraft. 1. A hybrid power distribution system for use in an aircraft having at least one engine , the hybrid power distribution system comprising:a starter drive/generator that converts mechanical energy provided by the at least one engine to electrical energy for distribution via an electrical distribution bus;a pneumatic distribution channel for extracting low-pressure bleed air from the at least one engine; andan environmental control system (ECS) connected to receive low-pressure bleed air from the pneumatic distribution circuit, wherein the ECS includes temperature control components supplied with power from the electrical power distribution bus to provide cooling and heating to the low-pressure bleed air provided from the pneumatic distribution channel.2. The hybrid power distribution system of claim 1 , further including:an auxiliary power unit (APU) that includes a starter/generator for converting mechanical energy provided by the APU to electrical energy for supply to the electrical distribution bus, wherein the electrical energy provided by the starter/generator of the APU is supplied to the starter drive/generator for conversion to mechanical power for starting the at least one engine without pneumatic energy being supplied to aid in starting of the least one engine.3. The hybrid power distribution system of claim 1 , wherein the temperature control components include an electric driven vapor cycle system to provide cooling to the low-pressure bleed air provided from the pneumatic distribution circuit.4. ...

System and method for providing electrical power

An electrical power unit provides electrical power to an electrical component on-board an aircraft. The electrical power unit includes a hydrogen generation system configured to be positioned on-board the aircraft. The hydrogen generation system is further configured to generate hydrogen using a reaction between water and metal. The electrical power unit also includes a fuel cell configured to be positioned on-board the aircraft. The fuel cell is operatively connected to the hydrogen generation system such that the fuel cell receives hydrogen from the hydrogen generation system. The fuel cell is further configured to generate electrical power from the hydrogen received from the hydrogen generation system and to be electrically connected to the electrical component for supplying the component with electrical power.

Electromechanical actuator damping arrangement for ram air turbine

A ram air turbine actuator includes a piston rod arranged in a cylinder. The cylinder and piston rod are configured to move longitudinally relative to one another between retracted and deployed positions. An annular space provided between the cylinder and piston rod, and a fluid flow regulating feature is provided in the piston rod. The fluid flow regulating feature is progressively blocked from the retracted position to the deployed position. The RAT actuator is deployed by initiating a deploy sequence and reducing a volume of an annular space. Fluid flow is forced from the annular space through a flow regulating feature to a cavity. The size of the flow regulating feature reduces to damp the actuator during the deploy sequence.

Electric motor/generator power transfer unit

A power transfer unit includes a differential gear set, first and second pump/motors, an electric motor/generator, and first and second hydraulic circuits. The differential gear set includes a first input/output member, a second input/output member, and a third input/output member. The first pump/motor is coupled to the first input/output member. The second pump/motor is coupled to the second input/output member. The electric motor/generator is coupled to the third input/output member. The first hydraulic circuit is hydraulically coupled to the first pump/motor. The second hydraulic circuit is hydraulically coupled to the second pump/motor and hydraulically separated from the first hydraulic circuit. The power transfer unit may include a power transfer unit mode where power is transferred through the differential gear set between the first and second pump/motors and include an electric motor/pump mode where power is transferred through the differential gear set between the electric motor/generator and one of the pump/motors.

METHOD FOR GENERATING HYDRAULIC POWER IN AN AIRCRAFT, USE OF A HYBRID POWER CONTROL UNIT AND DRIVE SYSTEM

A method for generating hydraulic power in an aircraft, the aircraft having a drive system comprising at least one transmission shaft connected to a power control unit, the power control unit having an electric motor and a hydraulic displacement machine connected to a differential gear unit for driving a common output shaft. The method includes switching the hydraulic displacement machine into a pump mode, arresting the output shaft, rotating the electric motor such that the hydraulic displacement machine is driven due to the arrested output shaft and supplying the fluid flow into a hydraulic system. A drive system of an aircraft may thereby be used for either moving control surfaces or for generating hydraulic power in an aircraft. This hydraulic power may be used to cover hydraulic load peaks during aircraft operation or to power hydraulic devices without the need of additional hydraulic power generation components. 1. A method for generating hydraulic power in an aircraft , the aircraft having a drive system comprising at least one transmission shaft connected to a hybrid power control unit , the power control unit having an electric motor and a hydraulic displacement machine connected to a differential gear unit for driving a common output shaft , the method comprising:arresting the output shaft;rotating the electric motor such that the hydraulic displacement machine is driven due to the arrested output shaft; andsupplying the fluid flow into a hydraulic system.2. The method of claim 1 , wherein arresting the output shaft comprises arresting at least one first brake connected to the at least one transmission shaft.3. The method of claim 2 , wherein the at least one first brake is at least one wing tip brake.4. The method of claim 1 , wherein rotating the electric motor includes releasing a second brake at the hydraulic displacement machine and a third brake at the electric motor.5. The method of claim 4 , wherein the second brake is a pressure-off brake and ...

RAM AIR TURBINE OVERSPEED PROTECTION

A power generation system includes a ram air turbine that is connected to a generator. The power generation system may be located in a pod, for example a pod for mounting on an aircraft. The ram air turbine receives air that passes through an air path through the pod, going in through an air inlet, through the turbine to turn the turbine, and out through an air outlet. The system includes a deployable flow obstruction, such as one or more airbags, that are deployable to suddenly obstruct the flow through the air path. The obstruction may be used to cut off flow (or greatly reduce flow), when overspeed of the ram air turbine is detected. The obstruction deploys (for example, airbags deploy in an air inlet of the system) to prevent continuation of the overspeed operation of the turbine, which may damage parts of the system. 1. A power generation system comprising:a ram air turbine;a turbine air inlet that provides air to the ram air turbine;a turbine air outlet that expels air that has passed through the ram air turbine, wherein the air inlet and the air outlet together define an air path that passes air through the ram air turbine; anda deployable flow obstruction that is deployable to suddenly block at least part of the air flow through the air path.2. The power generation system of claim 1 , wherein the deployable flow obstruction blocks substantially all air flow through the air path.3. The power generation system of claim 1 , wherein the deployable flow obstruction includes one or more airbags that inflate to obstruct the air path.4. The power generation system of claim 3 , wherein the one or more airbags include at least one airbag that obstructs flow in the air inlet.5. The power generation system of claim 3 , wherein the one or more airbags include at least one airbag that obstructs flow in the air outlet.6. The power generation system of claim 3 , further comprising one or more squibs that can be fired to create pressurized gases to fill the one or more ...

Combined Hub, Counterweight and Crank Arm with Single Piece Steel Body

A combination hub, counterweight and crank arm for a governor assembly in a ram air turbine is a single piece steel body.

Flexible and fireproof APU drainage device

An aircraft auxiliary power unit (APU) flexible and fireproof drainage device including a bellows which receives leakage fluids from an APU engine; a drain pipe which transfers leakage fluids outside the aircraft from the bellows, and a support piece surrounding the drain pipe. The bellows connects the APU engine with the drain pipe. The support piece provides stability to the bellows and the drain pipe.

AIRCRAFT AND SYSTEM FOR SUPPLYING ELECTRICAL POWER TO AN AIRCRAFT ELECTRICAL LOAD

One embodiment of the present disclosure is a unique aircraft. Another embodiment is a unique system for supplying electrical power to an aircraft electrical load during flight operations. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for fluid driven actuation systems. 1. An aircraft , comprising:a wing;a fuselage coupled to the wing;an engine coupled to at least one of the fuselage and the wing;an electrical load associated with the aircraft during flight operations;a generator coupled to the engine and configured to generate electrical power for the electrical load;a conductor electrically disposed between the electrical load and the generator;a conduit configured to house the conductor; anda dielectric gas disposed in the conduit;wherein the conduit is configured to envelop the conductor in the dielectric gas.2. The aircraft of claim 1 , wherein the generator is at least partially enveloped within in the dielectric gas.3. The aircraft of claim 1 , wherein the dielectric gas is a refrigerant vapor.4. The aircraft of claim 1 , further comprising a refrigeration system and a refrigerant configured for use with the refrigeration system claim 1 , wherein the dielectric gas is a refrigerant vapor.5. The aircraft of claim 4 , wherein the refrigeration system is configured to cool the generator.6. The aircraft of claim 4 , wherein the generator is at least partially enveloped within the refrigerant vapor7. The aircraft of claim 4 , further comprising a voltage reducer electrically coupled between the conductor and the electrical load claim 4 , wherein the voltage reducer is enveloped within the refrigerant vapor; and wherein the refrigeration system is configured to cool the voltage reducer.8. The aircraft of claim 4 , wherein the refrigeration system is configured to cool the conductor.9. The aircraft of claim 1 , further comprising a voltage reducer electrically coupled between the conductor and the electrical load claim ...

PROPULSION, ELECTRICAL, AND THERMAL MANAGEMENT DEVICE FOR A SMALL UNMANNED AERIAL VEHICLE

An aircraft is provided with a gas turbine engine having a plurality of shafts. A first shaft provides power to an electrical generator and a propeller, while a second shaft provides power to a refrigeration system. The refrigeration system may be integrated to the propeller, like a ducted fan, or on the outer skin of the aircraft.

FUEL CELL DEVICES FOR FIRE PREVENTION ON-BOARD AIRCRAFT

Described are inerting systems that may be used on board an aircraft or other passenger transportation vehicle to reduce a risk of fire due to electronic components or to other elements in a compartment and to assist in preventing or extinguishing any fire or hazardous condition that may occur. The systems include a source of inert gas such as oxygen depleted air generated from a fuel cell on board the aircraft. The oxygen depleted air or other inert gas is conveyed through ducts to compartments that house the electronics, thus changing the conditions in the compartment to be less conducive to fire. 1. An inerting system for an aircraft or other passenger transportation vehicle , comprising:(a) a fuel cell system;(b) a source of inert gas comprising an oxygen depleted air output of the fuel cell system;(c) a compartment located on-board the vehicle; and(d) a conduit configured to convey inert gas from the source of inert gas to the compartment to reduce a risk of fire ignition in the compartment or to facilitate extinguishing of fire within the compartment.2. The inerting system of claim 1 , wherein the compartment contains at least one electronic component within the compartment.3. The inerting system of claim 1 , wherein the source of inert gas is used to manage a hydrogen leak in the compartment.4. The inerting system of claim 1 , further comprising a valve or other regulator configured to control a flow level of inert gas from the source to the compartment.5. The inerting system of claim 1 , further comprising:a controller;one or more processors in communication with the controller; and 'in response to information received from a sensor indicating that conditions in the compartment are indicative of fire or imminent ignition therein, cause the valve or other regulator to direct inert gas from the source to the compartment.', 'memory including instructions that, when executed by the one or more processors, cause the one or more processors to6. The inerting system ...

AIRCRAFT INSTALLATION FOR SUPPLYING PRESSURIZED AIR

An aircraft installation of a pneumatic system of an aircraft with different compressed air sources for supplying pressurized air to air consumer equipment. Either an air bleed system, electrical compressors, or a combination thereof may perform such supplying of compressed air depending on the aircraft operation condition, for instance the flight altitude or specific flight phases. Also, a turbofan engine, and a method for supplying compressed air to air consumer equipment are disclosed. 1. An aircraft installation for supplying pressurized air to an air consumer equipment of an aircraft , the aircraft installation comprising:a gas turbine engine having a single port located at a low-intermediate compressor stage of such gas turbine engine; at one end in fluid communication with the single port of the gas turbine engine, and', 'at an opposite end in fluid communication with the air consumer equipment;, 'a bleed duct,'}a cooling duct comprising an inlet by which cooling air enters into the cooling duct, and an outlet;a branch duct in fluid communication with the bleed duct and configured to divert a portion of bleed air from the bleed duct into the outlet of the cooling duct; and the intake is in fluid communication with the outlet of the cooling duct after the joining with the branch duct, and', 'the outlet is in fluid communication with an opposite end of the bleed duct which is communicated with the air consumer equipment., 'an electrical compressor with an intake and an outlet, so that'}2. The aircraft installation according to claim 1 , wherein the inlet of the cooling duct is an air scoop arranged so as to substantially face the incoming cooling air entering into the cooling duct.3. The aircraft installation according to claim 1 , wherein the inlet of the cooling duct is an air scoop arranged so as to be angled relative to the incoming cooling air.4. The aircraft installation according to claim 1 , wherein the intake of the electrical compressor is connected ...

AIR MANAGEMENT SYSTEM

An air management system with a set of compressed air sources for supplying pressurized air to air consumer equipment. In particular, either an air bleed system, electrical compressors, or a combination thereof may perform such supplying of compressed air depending on the aircraft operation condition. 1. An air management system of an aircraft for supplying pressurized air to at least one air consumer , the air management system comprising:the at least one air consumer;at least one air source;at least one gas turbine engine having a single port located at a low-intermediate compressor stage of the gas turbine engine;an air bleed system in fluid communication with the at least one gas turbine engine via the single port of the gas turbine engine, the air bleed system being configured to supply compressed air to the at least one air consumer;at least one electrical compressor in fluid communication with the air source, the at least one electrical compressor being configured to supply compressed air to the at least one air consumer; anda controller configured to receive an input relative to an aircraft operation condition and selectively operate at least one of the air bleed system and the at least one electrical compressor based on the received input.2. The air management system according to claim 1 , wherein the at least one air consumer of the air management system is at least one of the following:an environmental control system;a fuel tank inerting system;a wing anti-ice system;an engine starting system;a water and waste system; andhydraulic reservoirs pressurization.3. An air management system according to claim 2 , wherein the environmental control system comprises a vapor cycle machine configured to be operated by at least one of the air bleed system and the at least one electrical compressor.4. The air management system according to claim 2 , wherein the wing anti-ice system is electrical.5. The air management system according to claim 1 , wherein the air bleed ...

APU EXHAUST SYSTEM

An exhaust system for an APU in a gas turbine engine aircraft having a tail cone in which exhaust noise is reduced. An APU exhaust liner is attached to the APU and extending to an oval cutout at the tail cone The liner has a first constant cross section from the APU to a point proximate the tail cone, and a second, increasing cross section from the point proximate the tail cone to the liner cutout. The liner is turned toward the side of the tail. The liner cutout is further enlarged with at least one larger cutout. 1. An exhaust system for an auxiliary power unit (APU) in tail cone section of an aircraft , the exhaust system comprising:an exhaust outlet on a side of the tail cone section;an APU exhaust liner attached to the APU and extending rearward from the APU to the exhaust outlet, the liner having a first section with a constant cross section from the APU to a point proximate the tail cone; and the liner having a second, larger diameter continuously increasing cross section from the downstream end of the first section to the outlet.2. The exhaust system of claim 1 , wherein the liner is turned toward a side of the tail at the tail cone.3. The exhaust system of claim 1 , where in the first diameter is about 27.8 cm (10.95 inches) and the second diameter is about 33.0 cm (13 inches).4. The exhaust system of claim 1 , wherein the exhaust outlet is oval in shape.5. The exhaust system of claim 4 , wherein the exhaust outlet is further enlarged with a second oval cutout of larger shape.6. The exhaust system of claim 5 , wherein the exhaust outlet is still further enlarged with a third oval cutout of still larger shape.7. A method for reducing the exhaust noise in an auxiliary power unit (APU) in a gas turbine engine aircraft having a tail cone claim 5 , the method comprising:receiving exhaust gas at a liner inlet adjacent the APU;directing the exhaust gas through the liner from the liner inlet to an outlet in a side surface of a fuselage tail section; andreducing the ...

AUXILIARY POWER UNIT MOUNTING BRACKET

An auxiliary power unit bracket, according to an exemplary aspect of the present disclosure includes, among other things, a gearbox mounting face and a compartment mounting face, the gearbox mounting face configured to attach to an auxiliary power unit gearbox below a sump level line, the compartment mounting face configured to attach to a portion of an auxiliary power unit compartment above the sump level line. 1. An auxiliary power unit bracket , comprising:a gearbox mounting face and a compartment mounting face,the gearbox mounting face configured to attach to an auxiliary power unit gearbox below a sump level line,the compartment mounting face configured to attach to a portion of an auxiliary power unit compartment above the sump level line.2. The auxiliary power unit bracket of claim 1 , wherein the bracket has a “U” shape.3. The auxiliary power unit bracket of claim 1 , wherein the bracket comprises steel.4. The auxiliary power unit bracket of claim 1 , wherein the gearbox mounting face is a first gearbox mounting face and the bracket includes a second gearbox mounting face opposite the first gearbox mounting face.5. The auxiliary power unit bracket of claim 1 , wherein the compartment mounting face is a first compartment mounting face and the bracket includes a second compartment mounting face opposite the first compartment mounting face.6. The auxiliary power unit bracket of claim 1 , wherein the gearbox mounting face is configured to attach to an aluminum auxiliary power unit gearbox.7. The auxiliary power unit bracket of claim 1 , including a support strap configured to interface with a downwardly facing portion of the auxiliary power unit gearbox.8. The auxiliary power unit bracket of claim 1 , wherein the compartment mounting face is configured to attach to a strut extending from a wall of the compartment to the compartment mounting face.9. An auxiliary power unit assembly comprising:a gearbox, the gearbox including a sump level line, the gearbox ...

AIRCRAFT INCLUDING PARALLEL HYBRID GAS TURBINE ELECTRIC PROPULSION SYSTEM

A gas turbine engine includes a core having a compressor section with a first compressor and a second compressor, and a turbine section with a first turbine and a second turbine. The first compressor is connected to the first turbine via a first shaft and the second compressor is connected to the second turbine via a second shaft. An electric motor is connected to the first shaft such that rotational energy generated by the electric motor is translated to the first shaft. An electric energy storage component is electrically connected to the electric motor, and electrically connected to at least one aircraft taxiing system. The gas turbine engine is configured such that the gas turbine engine requires supplemental power from the electric motor during at least one mode of operations. 1. A gas turbine engine comprising:a core including a compressor section having a first compressor and a second compressor, a turbine section having a first turbine and a second turbine;the first compressor is connected to the first turbine via a first shaft;the second compressor is connected to the second turbine via a second shaft;an electric motor connected to the first shaft such that rotational energy generated by the electric motor is translated to the first shaft;an electric energy storage component electrically connected to said electric motor, and electrically connected to at least one aircraft taxiing system; andwherein the gas turbine engine is configured such that the gas turbine engine requires supplemental power from the electric motor during at least one mode of operations.2. The gas turbine engine of claim 1 , wherein the aircraft taxiing system is a traction drive.3. The gas turbine engine of claim 2 , wherein the traction drive is drivably connected to at least one of an aircraft landing gear wheels and transmission.4. The gas turbine engine of claim 1 , wherein the aircraft taxiing system is a fan connected to said first shaft via at least one gearing system.5. The gas ...

AIRCRAFT FUEL CELL HEAT USAGES

Embodiments of the present disclosure relate generally to the use of the fuel cell systems on board aircraft and other passenger transportation vehicles and to methods of using heat, air, and water generated by such fuel cell systems. The heat may be used to address condensation within the aircraft. The heat may be used to help evaporate excess water that would otherwise condense in the aircraft skin. The excess water collected may be used to create humidification for cabin air. In other examples, the heat, warmed air, or warmed water may be delivered to other locations or heating systems for beneficial use. 1. A fuel cell by-product use system configured for managing aircraft moisture , comprising:a fuel cell system that creates electricity as a product, and water, heat, and oxygen depleted air as by-products;at least one fluid conduit for delivering the heat generated by the fuel cell to a skin insulation blanket in order to cause evaporation of moisture in the skin insulation blanket; anda collection system for capturing the evaporated moisture and delivering the moisture to an aircraft cabin air distribution system.2. The system of claim 1 , further comprising a system for capturing heat from the fuel cell system and converting the heat to heated air for delivery to the skin insulation blanket system.3. The system of claim 1 , wherein the at least one fluid conduit routes heated air beneath a cabin floor or within cabin walls or both.4. The system of claim 1 , wherein the evaporated moisture is recycled through the aircraft cabin air distribution system via delivery to a humidification system.5. A fuel cell by-product use system configured for managing aircraft heat claim 1 , comprising:a fuel cell system that creates electricity as a product, and water, heat, and oxygen depleted air as by-products;a fluid conduit for delivering the heat generated by the fuel cell to a heating system.6. The system of claim 5 , further comprising a system for controlling the heat ...

HYDRAULICALLY ASSISTED RAT ACTUATOR LATCH MECHANISM

A ram air turbine actuator having: a housing; at least one solenoid; a latch mechanism operably connected to the at least one solenoid and located within the housing, the latch mechanism in operation moves from a first position to second position when the at least one solenoid is activated; and a hydraulic pressure system operably connected to the latch mechanism, the hydraulic pressure system in operation moves the latch mechanism from the second position to the first position using hydraulic pressure from a hydraulic fluid, when the at least one solenoid is deactivated. The hydraulic pressure system is located within the housing. The hydraulic pressure is relieved when the latch mechanism is in the first position. 1. A ram air turbine actuator comprising:a housing;at least one solenoid;a latch mechanism operably connected to the at least one solenoid and located within the housing, the latch mechanism in operation moves from a first position to second position when the at least one solenoid is activated; anda hydraulic pressure system operably connected to the latch mechanism, the hydraulic pressure system in operation moves the latch mechanism from the second position to the first position using hydraulic pressure from a hydraulic fluid, when the at least one solenoid is deactivated, wherein the hydraulic pressure system is located within the housing,wherein the hydraulic pressure is relieved when the latch mechanism is in the first position.2. The ram air turbine actuator of claim 1 , wherein the latch mechanism further comprises:a lockbolt having a first biasing mechanism, the first biasing mechanism in operation moves the lockbolt in a first direction, when the at least one solenoid is activated;a toggle roller operably connected to the lockbolt, the toggle roller prevents motion by lockbolt in the first direction when the at least one solenoid is deactivated;a toggle latch operably connected to the toggle roller and the at least one solenoid, the toggle latch ...

POWER SYSTEM FOR AIRCRAFT PARALLEL HYBRID GAS TURBINE ELECTRIC PROPULSION SYSTEM

A gas turbine engine includes a compressor section having a first compressor and a second compressor and a turbine section having a first turbine and a second turbine. The first compressor is connected to the first turbine via a first shaft and the second compressor is connected to the second turbine via a second shaft. A motor connected to the first shaft such that rotational energy generated by the motor is translated to the first shaft. A power distribution system connects the motor to a stored power system including at least one of an energy storage unit and a supplementary power unit. The power distribution system is configured to provide power from the stored power system to the motor. 1. A gas turbine engine comprising:a compressor section having a first compressor and a second compressor;a turbine section having a first turbine and a second turbine,the first compressor is connected to the first turbine via a first shaft;the second compressor is connected to the second turbine via a second shaft;a motor connected to the first shaft such that rotational energy generated by the motor is translated to the first shaft; anda power distribution system connecting the motor to a stored power system including at least one of an energy storage unit and a supplementary power unit, wherein the power distribution system is configured to provide power from the stored power system to the motor.2. The gas turbine engine of claim 1 , wherein the power distribution system is isolated within the gas turbine engine.3. The gas turbine engine of claim 1 , wherein the power distribution system is integrated with an aircraft power distribution system.4. The gas turbine engine of claim 3 , wherein the power distribution system is connected to at least one aircraft electric system claim 3 , and is configured to provide operational power to the at least one aircraft electric system.5. The gas turbine engine of claim 1 , wherein the stored power system includes an energy storage unit ...

SPRING COMPRESSION DEVICE

A spring compression device comprises a spring sleeve configured to receive a spring; a spring adjustment member provided on and engaged with the spring sleeve and configured to abut a spring mounted on the spring sleeve, wherein the spring sleeve and spring adjustment member define an operating length (L) of the spring compression device, and the spring adjustment member is configured to be moved relative to the spring sleeve to adjust the operating length of the spring compression device; and at least one blocking component for blocking movement between the spring sleeve and the spring adjustment member. 1. A spring compression device , comprising:a spring sleeve configured to receive a spring;a spring adjustment member provided on and engaged with the spring sleeve and configured to abut a spring mounted on the spring sleeve, wherein:the spring sleeve and spring adjustment member define an operating length (L) of the spring compression device, andthe spring adjustment member is configured to be moved relative to the spring sleeve to adjust the operating length of the spring compression device; andat least one blocking component for blocking movement between the spring sleeve and the spring adjustment member.2. The spring compression device according to claim 1 , wherein the spring sleeve and spring adjustment member are engaged with a threaded connection.3. The spring compression device according to claim 1 , wherein an exterior surface of the spring sleeve and an interior surface of the spring adjustment member each have at least one recess which align to form a hole for receiving the blocking component.4. The spring compression device according to claim 3 , wherein the hole extends in a direction parallel to a longitudinal axis (R) of the spring sleeve.5. The spring compression device according to claim 3 , wherein the blocking component is press-fitted into the hole.6. The spring compression device according to claim 1 , wherein the blocking component ...

RAM AIR TURBINE STRUCTURES FOR TEMPERATURE DEPENDENT DAMPING

A ram air turbine (RAT) actuator includes a housing and a piston in operable communication with the housing and at least one damping orifice in operable communication with the housing and the piston, a flow area of the at least one damping orifice being alterable to adjust damping of movement between the piston and the housing in response to changes in viscosity of a fluid related to temperature. 1. A ram air turbine (RAT) actuator , comprising:a housing;a piston in operable communication with the housing;at least one damping orifice in operable communication with the housing and the piston, a flow area of the at least one damping orifice being alterable to adjust damping of movement between the piston and the housing in response to changes in viscosity of a fluid related to temperature.2. The actuator of claim 1 , wherein the housing defines a housing cavity claim 1 ,wherein the piston is at least partially disposed within the housing cavity of the housing in a sealing relationship with the housing,wherein the piston defines a piston cavity therein and is configured to allow the housing to move axially relative to the piston between a deployed position and a stowed position, wherein a plurality of damping orifices are defined in the piston,wherein hydraulic fluid can flow between the housing cavity and the piston cavity through the plurality of damping orifices, andwherein the actuator includes a damping orifice blocking device disposed around or within the piston and configured to allow or block flow through one or more of the damping orifices as a function of temperature.3. The actuator of claim 2 , wherein the damping orifice blocking device is configured to allow flow through one or more of the damping orifices when within a first temperature range and to block the one or more damping orifices when within a second temperature range that is higher than the first temperature range.4. The actuator of claim 3 , wherein the piston is made from a first material and ...

MULTI-SLOPE DROOP CONTROL

A method of controlling two or more sources of electrical power to supply electrical power to a shared bus. Each source of electrical power has a rated power P, a maximum overload power P, and is operable along a multi-voltage droop slope, the multi-voltage droop slope including a share region and a catch region, the share region being defined by an upper voltage limit, Vand a base voltage V, and the catch region being defined by the base voltage Vand a lower voltage limit V. 1. A method of controlling two or more sources of electrical power to supply electrical power to a shared bus for provision to an electrical load;{'sub': r', 'm', 'high', 'base', 'base', 'low, 'wherein each source of electrical power has a rated power P, a maximum overload power P, and is operable along a multi-voltage droop slope, the multi-voltage droop slope including a share region and a catch region, the share region being defined by an upper voltage limit, Vand a base voltage V, and the catch region being defined by the base voltage Vand a lower voltage limit, V;'}and the method comprises:providing a scaling factor, K, for each of the sources of electrical power, the scaling factor operating on a process variable used to control an output of the respective source of electrical power in a control loop, such that the sources of electrical power operate on one of the share region and the catch region; and{'sub': r', 'r', 'r, 'when a power demand of the electrical load is less than the sum of the rated powers P, setting one or more scaling factor K such that each source of electrical power is operating in the share region of the multi-voltage droop slope in which a power split between the sources of electrical power is equal to the ratio of Pfor a given source of power to the sum of rated powers Pfor the sources of electrical power; and'}{'sub': r', 'm', 'm, 'when the power demand of the electrical load is greater than the sum of the rated powers P, setting one or more scaling factor K such ...

POWERING AIRCRAFT SENSORS USING THERMAL CAPACITORS

An electric power generation system employs a thermoelectric generator placed between an aircraft inner skin and an aircraft outer skin. The thermoelectric generator is configured to utilize a thermal differential between the inner and outer skin to generate an electric current. An electrical interface is provided for access to the electric current generated by said thermoelectric generator. 1. A power generation system comprising:a principal system having a primary function, said principal system having an associated mass providing a thermal capacitor;a thermoelectric generator placed for operational engagement between the thermal capacitor and an external environment, said thermoelectric generator configured to utilize a temperature differential between the thermal capacitor and the external environment to generate an electric current; and,an auxiliary system associated with the principal system, said auxiliary system connected to operate using the electric current generated by said thermoelectric generator.2. The power generation system as defined in claim 1 , wherein the thermoelectric generator includes a first portion for contact with the thermal capacitor and a second portion being disposed in contact with a reference element in a vehicle that is at a different temperature than the thermal capacitor claim 1 , a thermoelectric gradient being formed between the first portion and the second portion to generate the electric current.3. The power generation system as defined in claim 2 , whereinthe first portion is a cold plate in thermal contact an aircraft outer skin andthe second portion is a hot plate in thermal contact with the thermal capacitor; and further comprising,a thermoelectric stack intermediate the cold plate and hot plate for electrical power generation.4. The power generation system as defined in claim 3 , wherein the thermoelectric stack operates using at least one of the Seebeck effect claim 3 , the Peltier effect claim 3 , or the Thomson effect ...

MODULAR POWER DISTRIBUTION ASSEMBLY AND METHOD OF ASSEMBLY THEREOF

A modular power distribution apparatus and method includes a frame defined by a first set of parallel supports and a second set of parallel supports, the second set of parallel supports connecting the first set of parallel supports, the frame defining a generally planar assembly, as well as a set of components extending from a side of the frame. The set of components includes a set of power modules. 1. A modular power distribution assembly , comprising:a frame defined by a first set of parallel supports and a second set of parallel supports; a set of primary power modules, each of the set of primary power modules having a power input, a bus bar power output, and a first communication interface;', 'a set of secondary power modules, each of the set of secondary power modules having a bus bar power input, a load output, and a second communication interface; and', 'a control module having a third communication interface; and, 'a first set of components extending from a first side of the frame, the first set of components comprising a common communication backplane having backplane interfaces matching a total number of first, second, and third communication interfaces defined by the first set of components, with the common communication backplane configured to communicate instructions from the control module to the set of primary power modules and the set of secondary power modules; and', 'a bus bar electrically connecting at least one bus bar power output to at least one bus bar power input., 'a second set of components extending from a second side of the frame and comprising2. The modular power distribution assembly of wherein a number in the set of primary power modules claim 1 , a number in the set of secondary power modules claim 1 , a number of backplane interfaces claim 1 , and the bus bar are selectable based on power distribution demand.3. The modular power distribution assembly of further comprising a carrier plate carried by the first set of parallel supports. ...

UNIDIRECTIONAL MATRIX CONVERTER WITH REGENERATION SYSTEM

An AC-to-AC matrix converter includes a controlled rectifier input stage, an inverter output stage, and a regenerative unit. The controlled rectifier input stage is electrically connected to a power source and configured to convert AC input power from the power source into DC power. The controlled rectifier input stage may be unidirectional. The inverter output stage is electrically connected to a load and operates to convert the DC power into AC output power, which is then fed into the load. The regenerative unit is electrically connected to the inverter output stage and operates control energy generated by the load in a regenerative mode within the inverter output stage. The controlled rectifier input stage may be isolated from the inverter output stage in the regenerative mode. 1. A matrix converter comprising:a controlled rectifier input stage electrically connected to a power source and configured to convert AC input power from the power source into DC power, wherein the controlled rectifier input stage is configured to be unidirectional;an inverter output stage electrically connected to a load and configured to convert the DC power into AC output power, the AC output power fed into the load; anda regenerative unit electrically connected to the inverter output stage and configured to control energy generated by the load in a regenerative mode within the inverter output stage, wherein the energy does not flow into the controlled rectifier input stage in the regenerative mode.2. The matrix converter of claim 1 , wherein the regenerative unit includes an energy dissipating element configured to dissipate the energy generated by the load in the regenerative mode.3. The matrix converter of claim 1 , wherein the regenerative unit includes an energy storing element configured to store the energy generated by the load in the regenerative mode.4. The matrix converter of claim 3 , wherein the regenerative unit is configured to deliver the energy stored in the energy ...

MULTIPLE POWER TOPOLOGIES FROM SINGLE POWER GENERATOR

A power unit for producing both alternating current and direct current includes a switcher connected to a direct current source, wherein the switcher includes circuitry configured to produce both alternating current having first characteristics and direct current having second characteristics, wherein the circuitry comprises a plurality of insulated gate bipolar transistor circuits and drive circuits connected to the insulated gate bipolar transistor circuits. The switcher may receive variable voltage and frequency or constant voltage and frequency. In either case, the switcher circuitry is able to provide power of the desired characteristics. The power unit may be used to power aircraft when the aircraft is on the ground. 1. A power unit for producing both alternating current and direct current , comprising:a switcher connected to a direct current source, wherein the switcher includes circuitry configured to produce both alternating current having first characteristics and direct current having second characteristics, wherein the circuitry comprises a plurality of insulated gate bipolar transistor circuits and drive circuits connected to the insulated gate bipolar transistor circuits.2. The power unit of claim 1 , wherein the first characteristics include 3-phase claim 1 , 400 hertz alternating current.3. The power unit of claim 1 , wherein the second characteristics include 270 V direct current.4. The power unit of claim 1 , wherein the switcher is connected to a rectifier claim 1 , wherein the rectifier converts alternating current into direct current supplied to the switcher.5. The power unit of claim 4 , comprising a first claim 4 , second claim 4 , and third insulated gate bipolar transistor circuit claim 4 , and the first claim 4 , second claim 4 , and third insulated gate bipolar transistor circuits are connected in parallel to a cathode and anode terminal of the rectifier.6. The power unit of claim 1 , comprising an interface circuit configured to apply a ...

Electrical power distribution box for onboard network of an aircraft and corresponding electrical power supply system

This electrical power distribution box for an aircraft electrical power supply network, which is intended to be connected to loads for powering said loads at a first voltage level and at a second voltage level lower than the first voltage level comprises a set of distribution modules for powering loads at said first voltage level. It further comprises a solid state converter capable of converting the first voltage level into a voltage at the second voltage level for powering loads at said second voltage level.

POWER-DISTRIBUTION SYSTEM PROVIDING COMPLETE SEPARABILITY BETWEEN A CIRCUIT BREAKER PANEL AND A CHASSIS

A power-distribution panel has a circuit-breaker panel that is completely separable from an aircraft and from a back chassis connected to the aircraft. The back chassis is mechanically mounted to the aircraft and electrically connected to both a power-supply bus and a power-distribution harness of the aircraft. The circuit-breaker panel can be removably coupled to the back chassis to complete a plurality of circuits from the power supply bus to a plurality the plurality of circuits that receive power via the power-distribution harness. In some embodiments, the circuit-breaker panel can be slideably coupled to the back chassis without the use of tools. Electrical connectors may provide electrically connectivity between the circuit-breaker panel and the back chassis. The electrical connectors may have back chassis portions that align with circuit-breaker panel portions so as to provide such electrically connectivity automatically when the circuit-breaker panel is coupled to the back chassis. 1. A power-distribution panel comprising:a front circuit-breaker panel comprising a plurality of circuit breakers; and an aircraft interface configured to mechanically couple to an aircraft and configured to electrically couple to both a power-supply bus and a power-distribution harness of the aircraft; and', 'a front-panel interface configured to receive and secure to the front circuit-breaker panel,, 'a back chassis comprisingwherein the front circuit-breaker panel and the back chassis are configured to be secured together such that each of the plurality of circuit breakers is in electrical communication with the power-distribution harness in a first configuration state, andwherein the front circuit-breaker panel and the back chassis are configured to be separated from one another such that the front circuit-breaker panel is electrically isolated from the power-distribution harness and from the power-supply bus in a second configuration state.2. The power-distribution panel of ...

AIRCRAFT DRAINAGE SYSTEM

A drain for expelling fluids from an interior of an aircraft to an exterior of the aircraft, the drain including a drain tube disposed at the interior of the aircraft having a first end disposed in fluid communication with an aircraft equipment to be drained and an opposite second end, wherein the drain tube terminates at the second end at a location within the interior of the aircraft, a seal which extends between the second end of the drain tube and an outer skin of the aircraft, delimiting a drainage cavity, and a drainage pathway extending from the cavity through the outer skin to the exterior of the aircraft. 2. The drain of claim 1 , wherein the seal is composed of a fire resistant material and sealingly surrounds the second end of the drain tube and a portion of the inner side of the outer skin through which the perforation extends.3. The drain of claim 2 , wherein the outer skin which contacts the seal constitutes a portion of a movable door of the aircraft claim 2 , wherein the door is movable between an open position and a closed position claim 2 , wherein in the closed position the cavity is hermetically sealed claim 2 , and wherein in the open position the cavity is exposed to the exterior of the aircraft.4. The drain of claim 3 , wherein the seal is affixed at a first end to the interior of the aircraft claim 3 , the seal including an opposite second end which freely extends from the interior of the aircraft to contact the outer skin.5. The drain of claim 4 , wherein in the closed position the outer skin of the door contacts and seals against the free second end of the seal to delimit the cavity claim 4 , and wherein in the open position the outer skin is positioned away from the seal and the cavity is exposed to the exterior of the aircraft.6. The drain of claim 5 , wherein the equipment to be drained comprises an auxiliary power unit (APU) disposed in a tail cone of the aircraft claim 5 , the drain further comprising a bracket affixed to the first end ...

VEHICLE AND SYSTEM FOR SUPPLYING ELECTRICAL POWER TO A VEHICLE ELECTRICAL LOAD

One embodiment of the present disclosure is a unique machine. Another embodiment is a unique system for supplying electrical power to a machine, such as a vehicle, having an electrical load during operation. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for fluid driven actuation systems. 120-. (canceled)21. A machine , comprising:a fixed member;a structure coupled to the fixed member;an engine coupled to at least one of the structure and the fixed member;an electrical load of a high energy device associated with the machine during operations;a generator coupled to the engine and configured to generate electrical power at a voltage at around 270 volts for the electrical load of the high energy device during operations;first and second conductors electrically disposed between the electrical load and the generator;first and second conduits arranged along each other and configured to house the respective first and second conductors; anda dielectric gas disposed in at least one of the first and second conduits;wherein the at least one of the first and second conduits is configured to envelop the conductor in the dielectric gas.221. The machine of claim , wherein the generator is enveloped within in the dielectric gas or uses dielectric gas as a cooling loop.231. The machine of claim , wherein the dielectric gas is a refrigerant vapor.241. The machine of claim , further comprising a refrigeration system and a refrigerant configured for use with the refrigeration system , wherein the dielectric gas is a refrigerant vapor.254. The machine of claim , wherein the refrigeration system is configured to cool the generator.264. The machine of claim , wherein the generator is at least partially enveloped within the refrigerant vapor.274. The machine of claim , further comprising a voltage reducer electrically coupled between at least one of the first and second conductors and the electrical load of the high energy device and ...

STRUCTURE FOR PROVIDING AN AIRCRAFT WITH EMERGENCY ELECTRIC POWER

The invention concerns a structure () for providing emergency electric power to an aircraft () comprising a pressurized cabin (), at least one main engine and an auxiliary power unit (APU), the structure being designed to provide electric power to the aircraft in the event of failure of the main engine and the emergency auxiliary unit, the structure comprising a circuit () for recovering air from the cabin (), and an electricity-generating system () which generates electricity by depressurizing the air recovered from the cabin (). 2. The architecture according to claim 1 , further including a variable timing system at the input to the turbine configured to modulate the power of the turbine.3. An aircraft claim 1 , comprising a pressurized cabin claim 1 , at least one main engine claim 1 , and an auxiliary power unit APU claim 1 , characterized in that it further comprises an architecture for supplying emergency electrical power according to .4. A method for supplying emergency electrical power in an aircraft comprising a pressurized cabin claim 1 , at least one main engine claim 1 , an auxiliary power unit APU claim 1 , and an architecture for supplying emergency electrical power according to claim 1 , the method including:during normal operation of the aircraft, controlling the distribution valve in a first configuration such that all the air recovered from the pressurized cabin is exhausted to the outside, andin the event of detection of an engine breakdown of the aircraft, control the distribution valve to assume a second configuration wherein at least a part of the air recovered from the pressurized cabin is directed into the recovery circuit to supply the electrical power production system.5. The method according to claim 1 , wherein the implementing step in the event of detection of an engine breakdown of the aircraft further includes the simultaneous control of the disengagement of the transmission gearbox from the auxiliary power unit to connect the ...

ELECTRICITY PRODUCTION SYSTEM FOR AN AIRCRAFT, COMPRISING A FUEL CELL

An electricity production system for an aircraft comprising a fuel cell, with an anode and a cathode, supplying an electric motor, a first feed pipe between the anode and a hydrogen source, a compressor, a second feed pipe between the inlet of the compressor and an oxygen source, a first transfer pipe between the outlet of the compressor and the cathode, a valve, an upstream pipe between the outlet of the compressor and the valve, a downstream pipe between the valve and an air outlet, and a controller which controls the position of the valve and the flow rate of the compressor. Such an electricity production system thus provides better control of the air flow supplied to the fuel cell on the basis of the electrical power required for the electric motor providing propulsion. 1. An electricity production system for an aircraft comprising an electric motor , a propeller driven in rotation by the electric motor , electrical conductors electrically supplying the electric motor , a hydrogen source and an oxygen source , wherein said electricity production system comprises:a fuel cell with an anode and a cathode, and in which the fuel cell is configured to electrically supply the electrical conductors,a first feed pipe fluidically connected to the anode and configured to be fluidically connected to the hydrogen source,an additional electric motor,a compressor whose rotating part is driven in rotation by the additional electric motor,a second feed pipe fluidically connected to an inlet of the compressor and configured to be fluidically connected to the oxygen source,a first transfer pipe fluidically connected between an outlet of the compressor and the cathode,a valve configured to assume different positions between an open position and a closed position,an upstream pipe fluidically connected between the outlet of the compressor and the valve,a downstream pipe fluidically connected between the valve and an air outlet, anda controller configured to control a position of the ...

Composite Panel Power System

An embodiment of the present disclosure provides a method and apparatus for supplying an electrical current. The method comprises sending the electrical current into a device from a panel comprising a dielectric core. Further, the method comprises a first sheet with a first conductive adhesive attaching the first sheet to a first side of the dielectric core, wherein the first conductive adhesive is a first electrode for a battery. Yet further, the method comprises a second sheet with a second conductive adhesive attaching the second sheet to a second side of the dielectric core, wherein the second conductive adhesive is a second electrode for the battery. Still further, the method comprises operating the device using the electrical current from the battery in the panel. 1. A panel containing an integrated battery , comprising:a dielectric core;a first sheet;a first conductive adhesive attaching the first sheet to a first side of the dielectric core, wherein the first conductive adhesive is a first electrode for the battery;a second sheet;a second conductive adhesive attaching the second sheet to a second side of the dielectric core, wherein the second conductive adhesive is a second electrode for the battery;wherein the dielectric core, the first sheet, and the second sheet form the panel; andwherein the panel comprises a structural component of a part.2. The panel containing the integrated battery of claim 1 , wherein the part comprises a part of an aircraft.3. The panel containing the integrated battery of claim 1 , wherein the part is selected from one of an interior wall claim 1 , a ceiling claim 1 , an overhead stowage bin claim 1 , a door claim 1 , a closet claim 1 , a sidewall claim 1 , a partition claim 1 , and a floor.4. The panel containing the integrated battery of claim 1 , wherein the first sheet and the second sheet comprise a number of layers of a composite material.5. The panel containing the integrated battery of claim 1 , wherein the second sheet ...

Aircraft

An aircraft includes cruise propellers, lifting propellers, an energy battery and a power battery. The energy battery is electrically connected to the cruise propellers, and the power battery is electrically connected to the lifting propellers.

FUELL CELL SYSTEM POWERED LAVATORY

In a craft such as an aircraft, a lavatory unit may be powered by the various outputs of a fuel cell system, including by-products that typically become waste. For example, but not limited to, a combination of the water, oxygen-depleted air, thermal energy and/or electrical energy generated by the fuel cell system may be used to supply water to the faucet and the toilet of the lavatory unit, to supply the lavatory unit with its electrical needs, and to heat and/or disinfect and/or dry the lavatory unit and its surfaces. 1. A lavatory unit powered by at least one output of a fuel cell system , wherein the at least one output comprises electrical energy , thermal energy , water , or oxygen-depleted air.2. The lavatory unit of claim 1 , wherein the lavatory unit is powered by at least three of the outputs of the fuel cell system.3. The lavatory unit of claim 1 , wherein the fuel cell system is proximate the lavatory unit and the water supplied by the fuel cell system is directed by at least one conduit to appropriate areas or occupants within the lavatory unit.4. The lavatory unit of claim 1 , wherein the oxygen-depleted air is directed from the fuel cell system into the lavatory unit to heat claim 1 , dry claim 1 , or sanitize the lavatory unit or surfaces of the lavatory unit.5. The lavatory unit of claim 4 , wherein the oxygen-depleted air is directed into the lavatory unit via shrouds or ducts in the floor or ceiling.6. The lavatory unit of claim 1 , wherein the thermal energy of the fuel cell system is directed into the lavatory unit.7. The lavatory unit of claim 1 , wherein a combination of the thermal energy and the oxygen-depleted air is used to sanitize the lavatory unit.8. The lavatory unit of claim 3 , wherein the water supplied by the fuel cell system is directed to a storage tank before it is directed to the appropriate areas or occupants.9. The lavatory unit of claim 1 , wherein the lavatory unit is configured for use in an aircraft.10. The lavatory unit ...

SYSTEM AND METHOD FOR REDUCING SPECIFIC FUEL CONSUMPTION (SFC) IN A TURBINE POWERED AIRCRAFT

A system for providing auxiliary power in an aircraft. A propulsion core comprises a compressor, a combustor, a turbine, and a shaft. An accessory unit comprises an accessory combustor, an accessory turbine, and an accessory shaft. A tank is configured to hold high pressure air and operably connected to the accessory unit. An electric generator comprises an electrical output and a mechanical input, with the mechanical input operably connected to the accessory shaft and the electrical output operably connected to an electric motor operably connected to the shaft. The electrical output is operably connected to an auxiliary power consuming device in the aircraft. 1. A method for reducing the specific fuel consumption for an aircraft mission , the method comprising:predetermining characteristics of the aircraft mission;injecting high pressure air from an onboard tank into a combustor of a power turbine; and,controlling the rate of injection of the high pressure air into the combustor of the power turbine as function at least of the mass of the high pressure air in the onboard tank and predetermined characteristics of the mission.2. The method of further comprising heating the high pressure air with exhaust from the aircraft's primary propulsion system prior to injection into the combustor and injecting exhaust from the power turbine into a core combustor of the aircraft's primary propulsion system.3. The method of claim 1 , further comprising determining an expected duration of the mission; determining the amount of high pressure air available in the onboard tank claim 1 , and determining a discharge mass flow rate of the high pressure air such that exhaustion of the high pressure air substantially corresponds to the end of the mission.4. The method of claim 3 , further comprising regulating the fuel supply rate to the power turbine at least as a function of the discharge mass flow rate. This application is a divisional of U.S. Utility patent application Ser. No. 15/452 ...

INDEPENDENT POWER GENERATION IN AIRCRAFT

A method includes generating thrust at one or more engines of an aircraft. The method includes generating first electricity at a first power unit concurrently with generating second electricity at a second power unit. The first power unit and the second power unit are independent of the one or more engines of the aircraft. The first electricity has a first set of electrical characteristics and the second electricity has a second set of electrical characteristics, where the first set of electrical characteristics is different than the second set of electrical characteristics. 1. A method comprising:generating thrust at one or more engines of an aircraft;generating first electricity at a first power unit that is independent of the one or more engines of the aircraft, the first electricity having a first set of electrical characteristics;concurrently with generating the first electricity, generating second electricity at a second power unit that is independent of the one or more engines of the aircraft, the second electricity having a second set of electrical characteristics, wherein the first set of electrical characteristics is different than the second set of electrical characteristics; andreceiving the first electricity at an environmental control system (ECS) of the aircraft, wherein, during flight, the ECS is electrically and pneumatically isolated from the one or more engines.2. The method of claim 1 , further comprising claim 1 , before generating thrust at the one or more engines of the aircraft claim 1 , starting the one or more engines of the aircraft using the first electricity.3. The method of claim 1 , wherein claim 1 , during flight claim 1 , the first power unit generates the first electricity independent of the one or more engines of the aircraft.4. The method of claim 1 , wherein claim 1 , during flight claim 1 , the second power unit generates the second electricity independent of the one or more engines of the aircraft.5. The method of claim 1 , ...

SPRAYING WATER IN RAM AIR FOR FUEL CELL POWER SYSTEMS IN AIRCRAFT

A fuel cell power system including at least one fuel cell, a ram air system and a cooling circuit in which coolant is intended to circulate for regulating a temperature of the at least one fuel cell. The cooling circuit comprises a ram air heat exchanger in the ram air system and the ram air system comprises a nozzle. The fuel cell power system further comprises a water tank and the fuel cell power system is arranged to flow water from the water tank to the ram air system so as to spray water in ram air via the nozzle. Thus, dimensioning of the ram air system which includes the ram air heat exchanger is reduced. 1. A fuel cell power system comprisingat least one fuel cell,a ram air system, anda cooling circuit configured to receive a circulating coolant to regulate a temperature of the at least one fuel cell, the cooling circuit comprising a ram air heat exchanger in the ram air system,wherein the ram air system comprises a nozzle, and the fuel cell power system further comprises a water tank and the fuel cell power system is arranged to flow water from the water tank to the ram air system so as to spray water in ram air via the nozzle,wherein the fuel cell power system further comprises a controller, referred to as a cooling controller, and a water flow regulating device configured to regulate a temperature of the coolant by adjusting operation of the water flow regulating device to regulate a flow of water from the water tank to the ram air system,wherein the cooling controller is configured to operate in a first operating mode when regulating the temperature of the coolant according to a first target temperature and to operate in a second operating mode when regulating the temperature of the coolant according to a second target temperature,wherein the cooling controller is configured to switch from the first operating mode to the second operating mode when a first event is received, wherein the first event indicates that the fuel cell power system should prepare ...

MULTI-MODE GENERATOR FOR ICE PROTECTION ON AIRCRAFT

On an aircraft, a multi-mode power generator is operated in a variable voltage mode to power an electric Wng Ice Prevention System (eWIPS), and is operated in a fixed voltage mode to provide backup power. When atmospheric conditions are conducive to the formation of ice (and main generators are operative), the multi-mode power generator is operated in variable voltage mode to power the eWIPS with a first or second variable voltage. The first variable voltage, the value of which depends on atmospheric conditions, is for anti-ice operation. The second variable voltage, which can be the maximum output voltage, is for de-ice operation. Transitions between different variable voltage levels are not instantaneous which eliminates fatigue damage due to transients. If a main generator fails (or when atmospheric conditions are not conducive to the formation of ice), the multi-mode power generator is operated in fixed voltage mode to provide backup power. 1. A method of operating a first multi-mode generator to power an electric Wng Ice Protection System (eWIPS) applying heat to flying surfaces of an aircraft; comprising:monitoring atmospheric conditions;during atmospheric conditions conducive to the formation of ice and when sufficient main generators are operative and providing power to a main aircraft electrical bus, operating the first multi-mode generator in a variable voltage mode and providing a variable voltage to the eWIPS; andduring atmospheric conditions not conducive to the formation of ice, or upon a failure of one or more main generators, operating the first multi-mode generator in a fixed voltage mode and providing a predetermined fixed voltage to the main aircraft electrical bus as backup power.2. The method of further comprising:monitoring the presence of ice on the aircraft flying surfaces; and providing a first variable voltage to the eWIPS when atmospheric conditions conducive to the formation of ice but no ice is detected; and', 'providing a second ...

BATTERY POWER SOURCE

A method and apparatus for a power distribution system, includes providing, from a power source having a set of dischargeable power storage devices connected by way of selectable connections, a power supply to a voltage output, and switchably connecting the voltage output to a set of power outputs having different electrical characteristics. 1. A power supply , comprising:a set of power storage units arranged in series, each power storage unit having a respective direct current (DC) dischargeable power storage device, a first selectable connection configured to enable a voltage output of the respective power storage device, and a second selectable connection configured to enable bypassing the voltage output of the respective power storage device;a controller module communicatively connected with the set of first selectable connections and the set of second selectable connections and configured to selectively enable at least a subset of the first selectable connections or a subset of the second selectable connections; anda voltage output for the set of power storage units configured to be switchably connected with a set of power outputs, wherein the set of power outputs have different electrical characteristics.2. The power supply of wherein at least one of the set of power outputs is a DC power output.3. The power supply of wherein at least one of the set of power outputs is an alternating current (AC) power output.4. The power supply of wherein the controller module is configured to selectively enable at least a subset of the first selectable connections or a subset of the second selectable connections claim 3 , and wherein the selectively enabling emulates an AC waveform at the power output.5. The power supply of claim 3 , wherein the set of power storage units includes a first set of power storage units arranged in series and connected with the voltage output claim 3 , the first set of power storage units having a positive voltage DC dischargeable power storage ...

GIMBAL, FRAME, AND UNMANNED AERIAL VEHICLE

Embodiments of the present disclosure provides a gimbal. The gimbal includes a pitch motor, a coaxial cable, and a first support arm arranged opposite a second support arm. The pitch motor is disposed at a front end of the first support arm, a motor shaft of the pitch motor extending in a direction of the second support arm for connecting with one end of an imaging device; a limiting shaft is disposed at a front end of the second support arm, the limiting shaft extending in a direction of the first arm for connecting with another end of the imaging device; a first wiring space is formed in the limiting shaft and a second wiring space is formed in the second support arm, a first end of the coaxial cable sequentially passing through the second wiring space and the first wiring space for electrically connecting with the imaging device. 1. A gimbal , comprising:a pitch motor;a coaxial cable; anda first support arm arranged opposite a second support arm, wherein:the pitch motor is disposed at a front end of the first support arm, a motor shaft of the pitch motor extending in a direction of the second support arm for detachably connecting with one end of an imaging device;a limiting shaft is disposed at a front end of the second support arm, the limiting shaft extending in a direction of the first arm for detachably connecting with another end of the imaging device; anda first wiring space is formed in the limiting shaft and a second wiring space connected to the first wiring space is formed in the second support arm, a first end of the coaxial cable sequentially passing through the second wiring space and the first wiring space for electrically connecting with the imaging device.2. The gimbal of claim 1 , wherein:the second support arm includes a first housing and a second housing covering the first housing, and the first housing and the second housing enclose the second wiring space.3. The gimbal of claim 1 , wherein:the limiting shaft is a hollow shaft with openings at ...

RAM AIR TURBINE BLADES

A ram air turbine (RAT) is provided and includes a turbine assembly including blades and a hub to which the blades are connected, a generator or a pump and a drivetrain mechanically interposed between the turbine assembly and the generator or the pump. Each blade includes an exterior, airfoil-shaped structure defining an interior and support structures disposed within the interior which connect with an inner surface of the exterior, airfoil-shaped structure and which define hollow regions within the interior. 1. A ram air turbine (RAT) , comprising:a turbine assembly comprising blades and a hub to which the blades are connected;a generator or a pump; anda drivetrain mechanically interposed between the turbine assembly and the generator or the pump,each blade comprising:an exterior, airfoil-shaped structure defining an interior; andsupport structures disposed within the interior which connect with an inner surface of the exterior, airfoil-shaped structure and which define hollow regions within the interior.2. The RAT according to claim 1 , wherein at least one of the hub claim 1 , the generator or the pump and the drivetrain are rated for operation with each blade.3. The RAT according to claim 1 , wherein the exterior claim 1 , airfoil-shaped structure is metallic claim 1 , tapered with increasing radial distance from the hub and comprises:a leading edge;a trailing edge opposite the leading edge;a low-pressure surface extending from the leading edge to the trailing edge; anda high-pressure surface extending from the leading edge to the trailing edge and opposite the low-pressure surface.4. The RAT according to claim 1 , wherein the support structures comprise metallic materials.5. The RAT according to claim 1 , wherein the support structures are provided in a honeycomb pattern.6. The RAT according to claim 1 , wherein a pitch between proximal support structures varies with increasing radial distance from the hub.7. The RAT according to claim 1 , wherein widths of the ...

Ram-body fretting corrosion proof solution

A ram-body for use with a spherical bearing has an inner circumferential surface. The ram-body comprises at least a first concave section provided in said inner circumferential surface. In some examples, a second concave section may be provided diametrically opposite to the first concave section. In some examples, the concave sections may be lunular shaped. A method of manufacturing the ram-body is also described. A ram-body and spherical bearing assembly is also described.

Fuel cell and fuel cell system for an aircraft

A fuel cell to provide a higher power density. The fuel cell can be produced by 3D printing in ceramic and has an improved power density by virtue of its spiral shape. In order to better extract the energy generated by the fuel cell, an interconnector sheet can be fastened positively to fastening knobs of the fuel cell by holding eyes. In addition, the interconnector sheet can be fixed by glass solder.

Drainage mast of the compartment of an aircraft subjected to a negative pressure

A drainage mast able to drain effectively the liquid contained in the compartment of an aircraft when the pressure within the compartment is lower than the outside pressure without using any pressure control active device and without causing any significant perturbation to the aerodynamic behavior of the aircraft. The drainage mast is configured with a cross section area decreasing from its entry section to the compartment to its outlet section and is disposed at an acute angle α relative to the compartment in the airflow direction. The entry and outlet sections have two perpendicular symmetry axes of different length, the entry section is arranged with its longer axis of symmetry substantially parallel to the airflow direction. The drainage mast has a symmetrical configuration with respect to a plane substantially perpendicular to the fuselage envelope passing through the longer axis of symmetry of the entry section.

ELECTRIC WIRE, CABLE HARNESS AND FLYING OBJECT

There is provided an electric wire having a conductor total cross-sectional area of 2 mmor less and high reliability. The electric wire includes: a twisted wire conductor () including a plurality of strand conductors ( to ) twisted together with each other and having a total cross-sectional area of the plurality of strand conductors of 2 mmor less; and a covering member () made from a resin material having a flexural modulus of 0.6 GPa or more and covering the twisted wire conductor so that an inner wall thereof is in contact with an outer peripheral surface of the twisted wire conductor, wherein the twisted wire conductor () has, on an outer peripheral surface thereof, recesses having a maximum depth of 5% or less of a maximum diameter of the twisted wire conductor in a cross section perpendicular to a length direction of the twisted wire conductor and containing boundaries of the plurality of strand conductors. 1. An electric wire comprising:a twisted wire conductor comprising a plurality of strand conductors twisted together with each other and having a total cross-sectional area of the plurality of strand conductors of 2 mm2 or less; anda covering member made from a resin material having a flexural modulus of 0.6 GPa or more and covering the twisted wire conductor so that an inner wall of the covering member is in contact with an outer peripheral surface of the twisted wire conductor,wherein the twisted wire conductor has, on an outer peripheral surface thereof, recesses having a depth of 5% or less of a maximum diameter of the twisted wire conductor in a cross section perpendicular to a length direction of the twisted wire conductor and containing boundaries of the plurality of strand conductors.2. The electric wire according to claim 1 , wherein the depth of the recesses is 3% or less of the maximum diameter of the twisted wire conductor.3. The electric wire according to claim 1 , wherein the depth of the recesses is 0.5% or more of the maximum diameter of the ...

VEHICLE AND SYSTEM FOR SUPPLYING ELECTRICAL POWER TO A VEHICLE ELECTRICAL LOAD

A system of a vehicle may include an electrical load, a generator, and first and second conduits. The electrical load of the vehicle may include a high energy device that utilizes above 270 volts during operations of the vehicle. The generator may be coupled to an engine of the vehicle and configured to generate electrical power at a voltage above 270 volts for the electrical load of the high energy device during operations of the vehicle. The first and second conduits may be arranged along each other to house respective first and second conductors that are electrically disposed between the electrical load and the generator. 120-. (canceled)21. A system of a vehicle , the system comprising:an electrical load of the vehicle including a high energy device that utilizes above 270 volts during operations of the vehicle;a generator coupled to an engine of the vehicle and configured to generate electrical power at a voltage above 270 volts for the electrical load of the high energy device during operations of the vehicle; andfirst and second conduits arranged along each other to house respective first and second conductors that are electrically disposed between the electrical load and the generator.22. The system of claim 21 , further comprising a dielectric gas disposed in at least one of the first and second conduits.23. The system of claim 21 , wherein at least one of the first and second conduits includes multiple conduits coaxially arranged and configured to envelop at least one of the first and second conductors in coaxial layers of dielectric gas.24. The system of claim 21 , further comprising a dielectric gas enveloping the generator.25. The system of claim 21 , further comprising a dielectric gas including a refrigerant vapor.26. The system of claim 21 , wherein a high energy device includes a high power directed energy system that is above 270 volts and includes at least one of a high power laser system claim 21 , a high power microwave system claim 21 , and a ...

Simultaneous Bidirectional Power Usage of Generator Power Feeders

Aircraft power distribution systems and methods for regulating a system voltage in aircraft power distribution systems are described. An example system includes a first generator, a first power feeder, a second power feeder, a first load, a second load, and a plurality of contactors. The first power feeder and the second power feeder are coupled in parallel between the first generator and a power panel. The first load is coupled to the first power feeder, and the second load is coupled to the second power feeder. The plurality of contactors is configurable to transfer power in a first direction from the first generator to both the first load and the second load during a first mode of operation, and configurable to transfer power in a second direction from the power panel to the first load or the second load during a second mode of operation. 1. An aircraft power distribution system , comprising:a first generator;a first power feeder coupled between the first generator and a power panel;a second power feeder coupled between the first generator and the power panel, the second power feeder coupled in parallel with the first power feeder;a first load coupled to the first power feeder;a second load coupled to the second power feeder; anda plurality of contactors, transfer power in a first direction from the first generator to both the first load and the second load during a first mode of operation, and', 'transfer power in a second direction from the power panel to the first load or the second load during a second mode of operation., 'wherein the plurality of contactors is configurable to, 'wherein each contactor of the plurality of contactors is connected to at least one of the first power feeder and the second power feeder, and'}2. The aircraft power distribution system of claim 1 , wherein claim 1 , during the second mode of operation claim 1 , the plurality of contactors are configured to transfer power from a second generator to the first load or the second load by ...

CROSS ENGINE COORDINATION DURING GAS TURBINE ENGINE MOTORING

A system is provided for cross engine coordination during gas turbine engine motoring. The system includes a first gas turbine engine of a first engine system, a first air turbine starter of the first engine system, a first starter air valve of the first engine system, and a controller. The controller is operable to command the first starter air valve to control delivery of compressed air to the first air turbine starter during motoring of the first gas turbine engine, monitor cross engine data of a second gas turbine engine of a second engine system to detect a present condition or a commanded action that modifies an aspect of the compressed air received at the first starter air valve, and command an adjustment to the first engine system to compensate for the modified aspect of the compressed air based on the cross engine data. 1. A system for cross engine coordination during gas turbine engine motoring , the system comprising:a first gas turbine engine of a first engine system;a first air turbine starter of the first engine system;a first starter air valve of the first engine system; anda controller operable to command the first starter air valve to control delivery of compressed air to the first air turbine starter during motoring of the first gas turbine engine, monitor cross engine data of a second gas turbine engine of a second engine system to detect a present condition or a commanded action that modifies an aspect of the compressed air received at the first starter air valve, and command an adjustment to the first engine system to compensate for the modified aspect of the compressed air based on the cross engine data.2. The system as in claim 1 , wherein the compressed air is driven by an auxiliary power unit of the aircraft claim 1 , a ground cart claim 1 , or a cross engine bleed.3. The system as in claim 1 , wherein the controller commands the first starter air valve based on one or more of: an engine speed of a starting spool of the first gas turbine ...

SUSPENSION SYSTEM FOR AN AIRCRAFT AUXILIARY POWER UNIT

A suspension system for an aircraft auxiliary power unit located in a fuselage structure, the system including: 1. A suspension system for an aircraft auxiliary power unit located in a fuselage structure , the system comprising:a plurality of struts extending between the auxiliary power unit and the fuselage structure,a plurality of auxiliary power unit attachment brackets for connecting the struts to the auxiliary power unit,a plurality of vibration isolators for joining the struts and the auxiliary power unit attachment bracket, the vibration isolators each comprising a housing connected with at least one of the struts and enclosing an elastomeric element in a load path extending through the vibration isolator and for damping vibrations,wherein least one of the vibration isolators comprises:a cone-bolt connected at one end to the housing and, at another end, attached to the auxiliary power unit attachment brackets, wherein the cone bolt includes a longitudinal threaded hollow, wherein said longitudinal hollow is open proximate the auxiliary power unit attachment brackets,an inner bolt partially located within the hollow of the cone-bolt and threaded to the upper part of the cone-bolt,an outer bolt having a longitudinal through-hole and partially located within the hollow of the cone-bolt and comprising an external thread that engages the thread of the cone-bolt, the inner bolt extending across the hollow of said outer bolt and configured so that the outer bolt is male with respect to the to the cone-bolt and female with respect to the inner bolt.2. The suspension system according to wherein at least one of the auxiliary power unit attachment brackets is divided into at least two parts claim 1 , a first part wrapping at least a first longitudinal portion of the cone-bolt and connected to the auxiliary power unit and a second part wrapping at least a second longitudinal portion of the cone-bolt and connected to the auxiliary power unit.3. The suspension system ...

MULTI-ENGINE COORDINATION DURING GAS TURBINE ENGINE MOTORING

A system is provided for multi-engine coordination of gas turbine engine motoring in an aircraft. The system includes a controller operable to determine a motoring mode as a selection between a single engine dry motoring mode and a multi-engine dry motoring mode based on at least one temperature of a plurality of gas turbine engines and initiate dry motoring based on the motoring mode. 1. A system for multi-engine coordination of gas turbine engine motoring in an aircraft , the system comprising:a controller operable to determine a motoring mode as a selection between a single engine dry motoring mode and a multi-engine dry motoring mode based on at least one temperature of a plurality of gas turbine engines and initiate dry motoring based on the motoring mode.2. The system as in claim 1 , wherein the at least one temperature is a measured core engine temperature or an oil temperature.3. The system as in claim 1 , wherein dry motoring is inhibited when the aircraft is not on the ground.4. The system as in claim 1 , wherein the motoring mode is further determined based on a plurality of performance parameters that are based on one or more of: an ambient condition claim 1 , performance limitations of a compressed air source and each air turbine starter driven by the compressed air source claim 1 , engine drag claim 1 , and parasitic factors.5. The system as in claim 4 , wherein the performance parameters are determined based on one or more of: an ambient air temperature claim 4 , an ambient pressure claim 4 , and an oil temperature.6. The system as is claim 1 , wherein the compressed air source is an auxiliary power unit of the aircraft claim 1 , a ground cart claim 1 , or a cross engine bleed.7. The system as in claim 1 , wherein the controller is further operable to monitor a speed of each of the gas turbine engines when dry motoring is active and switch from the multi-engine dry motoring mode to the single engine dry motoring mode based on one or more of the gas ...

TWIN CENTRIFUGAL SINGLE SPOOL ENGINE

An auxiliary power unit may comprise a twin centrifugal compressor including a first blade. A turbine may be disposed aft of the twin centrifugal compressor. The turbine may include a second blade. The first blade comprises a first material and the second blade comprises a second material. The first material may the same as the second material. The twin centrifugal compressor may include forward centrifugal compressor and an aft centrifugal compressor disposed aft of the forward centrifugal compressor. 1. A twin centrifugal compressor for an auxiliary power unit , comprising:a first airflow path at an input of the twin centrifugal compressor; anda second airflow path and a third airflow path defined by the twin centrifugal compressor, wherein the second airflow path and the third airflow path are split from the first airflow path.2. The twin centrifugal compressor of claim 1 , further including a blade claim 1 , wherein the blade comprises silicon nitride.3. The twin centrifugal compressor of claim 1 , wherein the second airflow path and the third airflow path are co-axial within the twin centrifugal compressor.4. The twin centrifugal compressor of claim 1 , wherein a pressure of the second airflow path is substantially the same as a pressure of the third airflow path.5. The twin centrifugal compressor of claim 1 , wherein a pressure ratio of at least one of the second airflow path or the third airflow path to the first airflow path is 7:1 or greater.6. The twin centrifugal compressor of claim 1 , further comprising a manifold defining a fourth airflow path claim 1 , wherein the fourth airflow path is comprised of a mix of the second airflow path and the third airflow path.7. A compressor section of an auxiliary power unit claim 1 , comprising:a forward centrifugal compressor; andan aft centrifugal compressor disposed aft of the forward centrifugal compressor.8. The compressor section of claim 8 , wherein the forward centrifugal compressor further comprises a blade. ...

Aircraft Location System for Locating Aircraft in Water Environments

A method and apparatus for an aircraft location system comprising an aircraft structure and a number of acoustic reflectors associated with the aircraft structure. The number of acoustic reflectors is configured to generate first sound signals in response to receiving second sound signals. 131-. (canceled)32. An aircraft location system comprising:an aircraft structure; anda first number of signal generators associated with the aircraft structure, wherein the first number of signal generators is configured to generate first light signals in response to receiving second signals.33. The aircraft location system of claim 32 , wherein the second signals comprise sound signals having a frequency in a range selected for locating the aircraft structure.34. The aircraft location system of further comprising a second number of signal generators claim 33 , wherein the second number of signal generators comprise a number of acoustic reflectors claim 33 , the second number of signal generators configured to be activated by a particular acoustic reflector in the number of acoustic reflectors and generate signals for locating the aircraft structure when activated.35. The aircraft location system of claim 34 , wherein the number of acoustic reflectors comprises:a first acoustic reflector configured to generate a number of first sound signals having a first frequency; anda second acoustic reflector configured to generate a number of second signals having a second frequency that is different from the first frequency.36. The aircraft location system of further comprising:a number of acoustic reflectors associated with the aircraft structure, wherein the number of acoustic reflectors is configured to generate first sound signals in response to receiving the second signals.37. The aircraft location system of further comprising a power source configured to power the first number of signal generators.38. The aircraft location system of claim 33 , wherein the light signal comprises a ...

Non-propulsive utility power (npup) generation system for providing secondary power in an aircraft

An aircraft may include at least one secondary power system requiring secondary power, at least two main engines, and at least three non-propulsive utility power (NPUP) generation systems. The NPUP generation systems may each be configured to provide full-time secondary power during operation of the aircraft. The NPUP generation systems may be configured to provide at least a portion of the secondary power required by the secondary power system.

Locking and unlocking mechanism

A locking mechanism for releasably locking a lock bolt against axial movement includes a solenoid assembly arranged, in use, in proximity to an axially moveable lock bolt. The solenoid assembly includes: a solenoid, a solenoid bias spring and a solenoid plunger having a plunger tip. When the solenoid is energised, the bias spring causes the solenoid plunger to move to bring the plunger tip into locking engagement with the lock bolt to prevent axial movement thereof, and when the solenoid is not energised, the solenoid bias spring causes the solenoid plunger to move to bring the plunger tip out of locking engagement with the lock bolt thus permitting axial movement thereof.

Power management for galley with fuel cell

Described are power management systems having at least one fuel cell system, a power distribution unit, at least one galley insert, a galley network controller, a control panel, and/or at least one battery pack. The power management system compares a power output from the power distribution unit to a maximum load level and/or to a minimum load level, instructs the galley network controller to cycle the the galley insert off and/or set an operation of the galley insert to a lower power consumption level when the power output approaches and/or is above the maximum load level, and instructs the galley network controller to cycle the galley insert on and/or set the operation of the galley insert to a higher power consumption level when the power output approaches and/or is below the minimum load level.

Aircraft starting and generating system

An aircraft starting and generating system includes a starter/generator and an inverter/converter/controller ( 200 ) that is connected to the starter/generator and that generates AC power to drive the starter/generator in a start mode for starting a prime mover of the aircraft, and that converts AC power, obtained from the starter/generator after the prime mover have been started, to DC power in a generate mode of the starter/generator. A four leg inverter is coupled with a DC power output ( 452 ) of the starter/generator and has an inverter/converter/controller (ICC) ( 580 ) with a four leg MOSFET-based bridge configuration that drives the starter/generator in a start mode for starting a prime mover of the aircraft, and converts DC power to AC power in a generate mode of the starter/generator. A four leg bridge gate driver ( 560 ) is configured to drive the four leg MOSFET-based bridge ( 580 ) during start and generate mode using bi-polar pulse width modulation (PWM).

ZERO CROSSING CONTACTOR AND METHOD OF OPERATING

A contactor apparatus and method for operating the contactor apparatus can include a contactor assembly with a contactor coil operably coupled to a contactor switch. One or more sensors can be provided in the contactor assembly adapted to measure one or more aspects of the contactor assembly. Based upon the measured aspects, a controller can initiate operation of the contactor switch to effectively toggle the contactor switch at a zero-crossing point along an alternating current waveform. 1. A method of operating a contactor , the method comprising:receiving, in a controller module, at least two operational characteristics of the contactor, with each operational characteristic representative of a delay time;determining, in the controller module, a contactor time delay for at least one of disconnecting or connecting a power supply by the contactor, the contactor time delay being a summation of a set of delay timings based on the delay time of each operational characteristic; andinitiating, by the controller module, the at least one of disconnecting or connecting of the power supply by the contactor at an initiation time prior to a zero-crossing voltage of an alternating current (AC) waveform of the power supply, wherein the initiation time anticipates the zero-crossing voltage based upon the contactor time delay.2. The method of claim 1 , wherein the at least two operational characteristics of the contactor include at least two of an electrical signal representative of the AC waveform and of an electrical signal delay time claim 1 , a temperature value representative of the temperature of a contactor coil and of a coil temperature delay time claim 1 , or a contactor coil operational characteristic representative of a contactor coil operational characteristic delay time.3. The method of claim 1 , wherein an effective disconnecting or connecting of the power supply coincides with the zero-crossing voltage of the AC waveform.4. The method of wherein the determining is ...

Fuel cell system for generating energy and water

Fuel cell systems aboard means of transport can be used for generating energy and for producing water. In order to reduce the overall weight of the system, the fuel cell is controlled or regulated in dependence on a current fill level or a limit level of the water tank, as well as a predicted future water consumption. In this way, it may be possible to minimize the water quantity to be stored in the water tank.

ELECTRICAL DISTRIBUTION DEVICE COMPRISING AT LEAST ONE POWER CONTROLLER

An electrical distribution device includes at least one power controller that is connectable to at least one electrical body, at least one local controller used to interface at least one power controller with at least one external calculator, and at least one local power supply path for powering at least one power controller and each local controller. Each power controller and each local controller includes a local DC/DC-type converter for coupling to each local power supply path. The device also includes an energy reservoir coupled to each local power supply path and to the local converters of the power controllers and the local controller. 1. An electrical distribution device comprising:at least one power controller that is connectable to at least one electrical component,at least one local controller adapted to interface the at least one power controller with at least one external calculator, andlocal power supply paths adapted to supply the power controller and the local controller,wherein the at least one power controller and the at least one local controller each include a DC/DC type local converter adapted to be coupled to the local power supply paths, andthe device further comprising an energy reservoir coupled to the local power supply paths, to the local converter of the power controller and to the local converter of the local controller.2. The electrical distribution device according to claim 1 , wherein the local power supply paths are without DC/DC type converter.3. The electrical distribution device according to claim 1 , further comprising a first common mode type filter which is mutualised between the local power supply paths and/or a second differential mode type filter which is mutualised between the local power supply paths.4. The electrical distribution device according to claim 1 , wherein the energy reservoir is of direct type or commuted type.5. The electrical distribution device according to claim 1 , wherein at least two local power supply ...

CREW REST POWERED BY A FUEL CELL SYSTEM

Disclosed are crew rests that may be powered by the outputs of a fuel cell system or other suitable power source. For example, but not limited to, a combination of the water, oxygen-depleted air, thermal energy and/or electrical energy generated by the fuel cell system may be used to supply the crew rest with its various power and water needs, helping to make the crew rest autonomous from the aircraft's main power systems. 1. A crew rest powered by a power source that is integral with the crew rest such that the crew rest is configured to operate independently of an aircraft's power generation system.2. The crew rest of claim 1 , wherein the power source is a fuel cell system.3. The crew rest of claim 2 , wherein one or more outputs of the fuel cell system is supplied to the crew rest and wherein the one or more outputs comprises electrical energy claim 2 , thermal energy claim 2 , water claim 2 , or oxygen-depleted air.4. The crew rest of claim 2 , wherein all of the one or more outputs are used within the crew rest.5. The crew rest of claim 1 , wherein the crew rest is mobile.6. A crew rest powered by at least one output of a fuel cell system such that the crew rest is configured to operate independently of an aircraft's power generation system claim 1 , wherein the at least one output comprises electrical energy claim 1 , thermal energy claim 1 , water claim 1 , or oxygen-depleted air.7. The crew rest of claim 6 , wherein at least two of the outputs of the fuel cell system are supplied to the crew rest.8. The crew rest of claim 6 , wherein the fuel cell system is contained within or is proximate the crew rest and wherein the water produced by the fuel cell system is directed by at least one conduit to appropriate areas of the crew rest or to consumers within the crew rest.9. The crew rest of claim 8 , further comprising a heat exchanger that cools the water produced by the fuel cell system before or after the water is directed to the appropriate areas or ...

A SYSTEM FOR RECOVERING AND CONVERTING KINETIC ENERGY AND POTENTIAL ENERGY AS ELECTRICAL ENERGY FOR AN AIRCRAFT

The invention relates to an electrical energy generator system () for an aircraft (), the system comprising a streamlined fairing () containing at least one turbine () housed in the front portion () of the fairing (), and an electrical energy generator () connected to said turbine. The front portion () of the fairing () is fitted with air admission means () that are movable between an open position in which the turbine () is exposed to the stream of air outside the fairing () and a closed position in which the turbine () is masked inside the fairing (). The invention serves to reduce the aerodynamic drag usually caused by turbulence present at a wing tip for a conventional wing having sharp edges during stages of takeoff, climbing, and cruising; and during stages of descent it makes it possible to recover the kinetic and potential energy that has been accumulated by the aircraft during its stages of climbing and cruising. 1. An electrical energy generator system for an aircraft , wherein the system comprising a shell having the shape of a wing tip tank and containing at least one turbine housed in the front portion of the shell and an electrical energy generator connected to said turbine , the front portion of the shell being fitted with air admission means that are movable between an open position in which the turbine is exposed to the air stream outside the shell , and a closed position in which the turbine is masked inside the shell , the air admission means comprising flaps or slats that are movably fastened between the nose and the body of the shell.2. A system according to claim 1 , characterized in that the air admission means comprise flaps fastened in hinged manner to the nose of the shell claim 1 , the flaps being lowered in the open position of the air admission means and being held in alignment with the body of the shell in the closed position of the air admission means.3. A system according to claim 1 , characterized in that the air admission means ...

ELECTRICAL ARCHITECTURE WITH PAIRED SECONDARY ELECTRICAL NETWORKS FOR STARTING ENGINES OF AN AIRCRAFT

An electrical architecture for an aircraft having a primary three-phase electrical network powering a transformer-rectifier unit serving to power two secondary DC electrical networks. Each of the two secondary electrical networks includes a contactor electrically connected to a generator-starter. The contactors are linked in pairs to prevent them from being closed simultaneously. The second secondary network powers the generator-starter to start a fuel-burning engine. The first secondary network can be powered by the generator-starter to power electrical components. A secondary voltage of the second secondary electrical network is greater than the voltage of the first secondary electrical network, thus making it possible to optimize the weight of the electrical architecture. The two secondary electrical networks may also be powered from autonomous sources of electricity as a replacement for the primary network to power the generator-starter and the electrical components. 1. An electrical architecture having paired secondary networks for an aircraft , the architecture comprising:two secondary electrical networks suitable for delivering respective secondary powers at respective secondary voltages;one or more electrical components; andat least one source of electricity connected to at least one secondary electrical network;wherein a common electrical component is electrically connected to both the secondary electrical networks, each secondary electrical network includes a contactor enabling each secondary electrical network to be electrically connected with the common electrical component, and the contactors being interlinked so that one contactor cannot be closed so long as the other contactor is not open so that the common electrical component cannot be electrically connected simultaneously to both the secondary networks.2. An electrical architecture according to claim 1 , wherein a first secondary electrical network delivers secondary power at a first secondary ...

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 ...

Power system controller failsafe cycling protection

Examples provide a method and system for inhibiting aircraft electrical power system controller failsafe state cycling in air mode condition. The electrical power system controller can include a bus power control unit (BPCU) or a generator control unit (GCU). A failsafe cycling protection indicator, such as a software flag or hardware electronic latch has a first configuration to indicate failsafe cycling protection is enabled and a second configuration to indicate failsafe cycling protection is disabled. An internal low voltage power source generates an output voltage. If a monitor component monitoring the output voltage detects an undervoltage or overvoltage, the aircraft electrical power system controller goes into a failsafe state. A failsafe cycling component prevents the aircraft electrical power system controller from exiting the failsafe state if the failsafe cycling protection is enabled.

Device for supplying air to an auxiliary power unit of an aircraft and associated aircraft

A device for supplying air to an auxiliary power unit for aircraft, including the auxiliary power unit; a turbocharger, and a cabin, the auxiliary power unit including a diesel engine coupled to the turbocharger in such a way that: the turbocharger includes a first air inlet in a compressor of the turbocharger coming from the cabin for its intake into a combustion chamber of the diesel engine, the diesel engine including: a first outlet carrying the air burnt by the combustion chamber to the turbocharger, the diesel engine delivering a power making it possible to supply an electric alternator and pressurising compressor, intended to pressurise a second air inlet, wherein the first air inlet is supplied by a fraction of pressurised air from the cabin of an aircraft.

COOLING CONCEPT FOR A FUEL CELL SYSTEM FOR A VEHICLE AND AIRCRAFT HAVING SUCH A FUEL CELL SYSTEM

A fuel cell system for a vehicle comprises at least one fuel cell, at least one fuel cell heat exchanger arranged in or at the at least one fuel cell for receiving heat of the at least one fuel cell, at least one thermal dissipation unit, at least one additional heat exchanger and a cooling loop having a plurality of fluid line segments for conveying a coolant. The at least one fuel cell heat exchanger is coupled with the at least one thermal dissipation unit through the cooling loop, and the additional heat exchanger is arranged in the cooling loop and is adapted for receiving heat from an external source and for raising the temperature of a coolant flowing in the cooling loop. 1. A fuel cell system for a vehicle , the fuel cell system comprisingat least one fuel cell,at least one fuel cell heat exchanger arranged in or at the at least one fuel cell for receiving heat of the at least one fuel cell,at least one thermal dissipation unit,at least one additional heat exchanger anda cooling loop having a plurality of fluid line segments for conveying a coolant,wherein the at least one fuel cell heat exchanger is coupled with the at least one thermal dissipation unit through the cooling loop, andwherein the additional heat exchanger is arranged in the cooling loop and is adapted for receiving heat from an external source and for raising the temperature of a coolant flowing in the cooling loop.2. The fuel cell system according to claim 1 ,further comprising a power electronics unit for the control and conversion of electrical power of the at least one fuel cell,wherein the power electronics unit comprises a power electronics heat exchanger arranged in or at the power electronics unit for receiving heat of the power electronics unit, andwherein the power electronics heat exchanger is arranged in the cooling loop upstream of the at least one fuel cell heat exchanger.3. The fuel cell system according to claim 2 ,further comprising a coolant bypass parallel to the power ...

Electrical network of an aircraft

An electrical network of an aircraft comprises: several main generators, several high voltage direct current networks powered by one of the generators in normal operation, several low voltage direct current networks powered by one of the high voltage direct current networks, several converters to transfer energy from one of the high voltage direct current networks to one of the low voltage direct current networks, the different converters being independent and isolated from each other, each converter being dedicated to one of the high voltage networks and to the low voltage direct current network, a load designed to be powered in normal operation by the main generators and in backup operation by one of the low voltage direct current networks, the backup operation being put into place when at least one of the high voltage networks is no longer powered by the associated main generator in normal operation. 1. An electrical network of an aircraft comprising:several main generators,several high voltage direct current HVDC networks, each one powered by one of the generators in normal operation,several low voltage direct current LVDC networks, each one powered by one of the high voltage direct current HVDC networks,several converters to transfer energy from one of the high voltage direct current HVDC networks to one of the low voltage direct current LVDC networks, the different converters being independent and isolated from each other, each converter being dedicated to one of the high voltage HVDC networks and to the low voltage direct current LVDC network,a load designed to be powered in normal operation by the main generators and in backup operation by one of the low voltage direct current LVDC networks, the backup operation being put into place when at least one of the high voltage HVDC networks is no longer powered by the associated main generator in normal operation,in which the converters are reversible and in backup operation the load is powered in parallel by ...

POWER CONVERTER, GENERATOR AND ARCHITECTURE FOR HIGH EFFICIENCY AUXILARY POWER UNIT

A high efficiency multifunction power system for an aircraft is provided. The system includes an AC generator and a multifunction power converter-controller module including at least one multifunction power converter-controller. The at least one multifunction power converter-controller is configured to function as a power converter and a controller to perform multiple operation modes. 1. A power system for an auxiliary power unit (APU) of an aircraft , comprising:an AC generator; anda multifunction power converter-controller module electrically connected to the AC generator, the multifunction power converter-controller module including at least one multifunction power converter-controller (MPCC),wherein the AC generator is configured to use ground power to run as a motor to provide mechanical power for the APU so that the APU provides air for an environmental control system, and wherein ground power is supplied through the at least one MPCC.2. The system of claim 1 , wherein the at least one MPCC is configured to start the APU using the AC generator powered by a ground power from a ground power network.3. The system of claim 1 , wherein the APU supplies air to a bleed air system so that the bleed air system provides air to the environmental control system.4. The system of claim 1 , wherein the multifunction power converter-controller module is configured to receive an AC power input from the AC generator at a first frequency and supply AC power output to an aircraft power distribution system at a second frequency claim 1 , wherein the first frequency is different from the second frequency.5. The system of claim 4 , wherein the first frequency is less than 400 Hz and the second frequency is 400 Hz.6. The system of claim 1 , wherein the AC generator is a controllable frequency generator.7. The system of claim 6 , wherein the multifunction power converter-controller module further includes a controllable frequency generator controller to control the controllable ...

REMOTE POWER SOURCE PUMP SYSTEM

A pump system includes an actuator which can be retracted by a fluid in the pump system. Also included is a pump with a first drive coupling and a rotational power device with a second drive coupling. The second drive coupling is configured to removably mate with the first drive coupling for powering the pump. A spool valve that is spring loaded and in fluid connection with the pump when the spool valve is in a stow position is also included. In addition, the pump system includes both a relief valve and a reservoir both in fluid connection with the actuator. 1. A pump system comprising:an actuator which can be retracted by a fluid in the pump system;a pump with a first drive coupling;a rotational power device with a second drive coupling, wherein the second drive coupling is configured to removably mate with the first drive coupling for powering the pump;a spool valve that is spring loaded and in fluid connection with the pump when the spool valve is in a stow position;a relief valve in fluid connection with the actuator; anda reservoir in fluid connection with the actuator.2. The pump system of claim 1 , wherein the spool valve is spring loaded to a free flow position and is toggled to the stow position when the second drive coupling is mated with the first drive coupling.3. The pump system of claim 2 , wherein when the spool valve is in the free flow position the fluid in the pump system bypasses the pump.4. The pump system of claim 1 , wherein when the spool valve is in the stow position fluid flows in a direction from the spool valve to the pump.5. The pump system of claim 4 , wherein when the spool valve is in the stow position pressurized fluid is developed at the pump and directed from the pump to the actuator to retract the actuator.6. The pump system of claim 5 , wherein the pump develops pressurized fluid in the pump system ranging from approximately 500 psig (3447 kPa) to 1500 psig (10 claim 5 ,342 kPa).7. The pump system of claim 1 , wherein the ...

AUXILLARY POWER UNIT ASSEMBLY AND A METHOD OF USING THE SAME

A gas turbine auxiliary power unit system for an aircraft comprises an aircraft cabin provided with a pressurised air supply, a gas turbine engine comprising a compressor assembly, and a controller. 1. A gas turbine auxiliary power unit system for an aircraft , the system comprising:an aircraft cabin provided with a pressurised air supply;a gas turbine engine comprising a compressor assembly; anda controller,wherein the pressurised air supply is provided to the aircraft cabin through an aircraft cabin air inlet, and exhausted from the aircraft cabin through an aircraft cabin air exhaust,the gas turbine engine is provided with an inlet air flow from the aircraft cabin air exhaust, andthe controller is operable to control a flow rate of a fuel supplied to the gas turbine engine responsive to a power demand on the auxiliary power unit, the controller being further operable to restrict an inlet air flow into the gas turbine engine, so as to maintain an air pressure inside the aircraft cabin within a pre-determined range.2. The system as claimed in claim 1 , wherein the gas turbine engine comprises a plurality of actuatable variable inlet guide vanes claim 1 , each of the variable inlet guide vanes being controlled by the controller to restrict the inlet air flow into the gas turbine engine.3. The system as claimed in claim 2 , wherein the compressor assembly comprises the plurality of variable inlet guide vanes.4. The system as claimed in claim 1 , wherein the gas turbine engine further comprises:a coupled power turbine assembly, the coupled power turbine assembly being connected to the compressor assembly by a first shaft; anda throttle assembly in fluid communication with the compressor assembly;wherein the controller is operable to control the throttle assembly to restrict the inlet air flow into the gas turbine engine.5. The system as claimed in claim 4 , the system further comprising a first electrical generator claim 4 , the first electrical generator being ...

INFLIGHT POWER MANAGEMENT FOR AIRCRAFT

There is described herein methods and systems for selective use of an auxiliary power source inflight in order to reduce fuel consumption of an aircraft. 1. A method for power management in an aircraft having at least one main power source for providing propulsive power to the aircraft and at least one auxiliary power source for providing auxiliary power to the aircraft , the method comprising:while in flight, receiving auxiliary power source current operating conditions from at least one auxiliary power source and receiving main power source current operating conditions from at least one main power source;receiving an actual load power requirement for the aircraft;determining an allocation of power loads of the aircraft for the actual load power requirements, based on the current operating conditions, to minimize fuel consumption of the aircraft; anddistributing the power loads of the aircraft between the at least one auxiliary power source and the at least one main power source in accordance with the allocation as determined.2. The method of claim 1 , wherein determining an allocation of power loads to minimize fuel consumption comprises comparing thermal efficiencies of the at least one auxiliary power source and the at least one main power source.3. The method of claim 1 , wherein determining an allocation of power loads to minimize fuel consumption comprises minimizing a fuel flow for a required propulsion thrust claim 1 , bleed air conditions claim 1 , and shaft power extraction.4. The method of claim 1 , wherein determining an allocation and distributing the power loads is triggered by a change to the actual load requirement for the aircraft.5. The method of claim 1 , further comprising shutting down the at least one auxiliary power source when the power loads are fully allocated to the at least one main power source.6. The method of claim 1 , wherein determining an allocation comprises considering a flight regime of the aircraft.7. The method of claim 1 , ...

Aircraft heating assembly with liquid cooled internal combustion engine and heating element using waste heat

An aircraft heating assembly including an internal combustion engine having a liquid coolant system distinct from any fuel and lubricating system of the engine and including cooling passages in the internal combustion engine for circulating a liquid coolant from a coolant inlet to a coolant outlet, a coolant circulation path outside of the internal combustion engine and in fluid communication with the coolant inlet and the coolant outlet, and a heating element in heat exchange relationship with a portion of the aircraft to be heated. The coolant circulation path extends through a heat exchanger configured to remove a portion of a waste heat from the liquid coolant. The heating element is in heat exchange relationship with the coolant circulation path to receive another portion of the waste heat therefrom. A method of heating a portion of an aircraft is also discussed.

AIRCRAFT HAVING A REDUNDANT AND EFFICIENT BLEED SYSTEM

An aircraft having at least one thrust providing jet engine and a fuselage with an integrated pressurized cabin includes an environmental control system having at least one pneumatic air cycle air conditioning unit, a bleed air port in the at least one thrust providing jet engine, a non thrust providing auxiliary supply unit configured to provide electricity and compressed air on ground and during flight, and at least one compressed air line. The at least one compressed air line is coupled with the bleed air port and the auxiliary supply unit. The environmental control system is coupled with the at least one compressed air line. Further, the auxiliary supply unit is dimensioned to exclusively operate the environmental control system by delivering compressed air, and bleed air from the bleed air port is selectively providable to the environmental control system for redundancy. 1. An aircraft having at least one thrust providing jet engine and a fuselage with an integrated pressurized cabin , the aircraft comprising:an environmental control system comprising at least one pneumatic air cycle air conditioning unit;a bleed air port in the at least one thrust providing jet engine;a non thrust providing auxiliary supply unit adapted for providing electricity and compressed air on ground and during flight; andat least one compressed air line;wherein the at least one compressed air line is coupled with the bleed air port and the auxiliary supply unit;wherein the environmental control system is coupled with the at least one compressed air line;wherein the auxiliary supply unit is dimensioned to exclusively operate the environmental control system by delivering compressed air; andwherein bleed air from the bleed air port is selectively providable to the environmental control system for redundancy.2. The aircraft according to claim 1 , wherein the auxiliary supply unit comprises at least one fuel cell claim 1 , a compressor and an electric motor coupled with the compressor and ...

AUXILIARY POWER UNIT WITH VARIABLE SPEED RATIO

An auxiliary power unit for an aircraft includes a rotary intermittent internal combustion engine drivingly engaged to an engine shaft, a turbine section having an inlet in fluid communication with an outlet of the engine(s), the turbine section including at least one turbine compounded with the engine shaft, and a compressor having an inlet in fluid communication with an environment of the aircraft and an outlet in fluid communication with a bleed duct for providing bleed air to the aircraft, the compressor having a compressor rotor connected to a compressor shaft, the compressor shaft drivingly engaged to the engine shaft. The driving engagement between the compressor shaft and the engine shaft is configurable to provide at least two alternate speed ratios between the compressor shaft and the engine shaft. 1. An auxiliary power unit for an aircraft , comprising:a rotary intermittent internal combustion engine drivingly engaged to an engine shaft;a turbine section having an inlet in fluid communication with an outlet of the rotary intermittent internal combustion engine, the turbine section including at least one turbine compounded with the engine shaft; anda compressor having an inlet in fluid communication with an environment of the aircraft and an outlet in fluid communication with a bleed duct for providing bleed air to the aircraft, the compressor having a compressor rotor connected to a compressor shaft, the compressor shaft drivingly engaged to the engine shaft, the driving engagement between the compressor shaft and the engine shaft configurable to provide at least two alternate speed ratios between the compressor shaft and the engine shaft.2. The auxiliary power unit as defined in claim 1 , wherein the outlet of the compressor is also in fluid communication with an inlet of the rotary intermittent internal combustion engine.3. The auxiliary power unit as defined in claim 2 , wherein the compressor is a first compressor and the compressor shaft is a first ...

AUXILIARY POWER UNIT WITH COMBINED COOLING OF GENERATOR

An auxiliary power unit for an aircraft, including an internal combustion engine having a liquid coolant system, a generator drivingly engaged to the internal combustion engine and having a liquid coolant system distinct from the liquid coolant system of the internal combustion engine, a first heat exchanger in fluid communication with the liquid coolant system of the internal combustion engine, a second heat exchanger in fluid communication with the liquid coolant system of the generator, an exhaust duct in fluid communication with air passages of the heat exchangers, and a fan received in the exhaust duct and rotatable by the internal combustion engine for driving a cooling air flow through the air passages. The liquid coolant system of the engine may be distinct from fuel and lubricating systems of the auxiliary power unit. A method of cooling a generator and an internal combustion engine is also discussed. 1. An auxiliary power unit for an aircraft , the auxiliary power unit comprising:an internal combustion engine having a liquid coolant system distinct from any fuel and lubricating system of the auxiliary power unit;a generator drivingly engaged to the internal combustion engine, the generator having a liquid coolant system distinct from the liquid coolant system of the internal combustion engine;a first heat exchanger having first coolant passages in fluid communication with the liquid coolant system of the internal combustion engine and first air passages in heat exchange relationship with the first coolant passages;a second heat exchanger having second coolant passages in fluid communication with the liquid coolant system of the generator and second air passages in heat exchange relationship with the second coolant passages;an exhaust duct in fluid communication with the first and second air passages; anda fan received in the exhaust duct and rotatable by the internal combustion engine for driving a cooling air flow through the first and second air passages ...

SERIES BATTERIES TO REDUCE AN INTERFERING MAGNETIC FIELD

A first portion of a series battery is arranged where the first portion of the series battery produces a first magnetic field and the series battery includes a plurality of charge storage devices, a negative output terminal, a positive output terminal, and a plurality of intradevice connections connecting the plurality of charge storage devices in series. A second portion of the series battery is arranged such that a second magnetic field produced by the second portion of the series battery at least partially cancels out the first magnetic field. 121-. (canceled)22. A series battery , comprising:a plurality of charge storage devices including a first set of non-adjacent charge storage devices and a second set of non-adjacent charge storage devices;a first intradevice connection which connects the first set of non-adjacent charge storage devices;a second intradevice connection which connects the second set of non-adjacent charge storage devices,wherein the first set of non-adjacent charge storage devices and the second set of non-adjacent charge storage devices are mutually exclusive;wherein the first intradevice connection and the second intradevice connection at least partially overlap; andwherein a direction of current flow of the first intradevice connection is opposite to a direction of current flow of the second intradevice connection.23. The series battery of claim 22 , wherein:the first set of non-adjacent charge storage devices includes a first charge storage device and a second charge storage device,the first charge storage device having a positive tab on a top surface and a negative tab on an opposite bottom surface; andthe second charge storage device having a negative tab on a top surface and a positive tab on an opposite bottom surface.24. The series battery of claim 23 , wherein: the third charge storage device having a negative tab on a top surface and a positive tab on an opposite bottom surface;', 'the fourth charge storage device having a positive ...

RELEASE MECHANISM

A release mechanism comprises a locking body mounted for reciprocating movement in a first axial direction between a locked position and a released position. A force transmitting element is coupled to the locking body for transmitting a force (F) to the locking body for moving the locking body from the locked position to the released position. A biasing element acts on the locking body in a direction for moving the locking body from the released position to the locked position. The locking body comprises a slot having a first slot portion extending in said first direction, and a second slot portion extending transversely from one side of said first slot portion at an end thereof. The slot slidably receives an actuating element therein. 1. A release mechanism comprising:a locking body mounted for reciprocating movement in a first axial direction between a locked position and a released position;a force transmitting element coupled to the locking body for transmitting a force (F) to said locking body for moving the locking body from the locked position to the released position; anda biasing element acting on the locking body in a direction for moving the locking body from the released position towards the locked position;said locking body comprising a slot comprising a first slot portion extending in said first direction, and a second slot portion extending transversely from one side of said first slot portion at an end thereof; andsaid slot slidably receiving an actuating element therein, said actuating element being movable in a second direction transverse to said first axial direction and received in said second slot portion when the locking body is in the locked position and in said first slot portion when the locking body is in the released position.2. The release mechanism as claimed in claim 1 , wherein said second slot portion comprises a first axially facing wall extending transversely from a first wall of said first slot portion and an opposed second axially ...

Device for supplying electrical power to an aircraft on the ground

A device for supplying electrical power to an aircraft on the ground, including two electric generators driven by an auxiliary power unit. The first generator is connected by a connection/disconnection mechanism to an aircraft network and to an electric taxiing network, to supply either an AC voltage to the aircraft network when it is connected to the aircraft network, or an AC voltage or a power to the taxiing network when it is connected to the taxiing network. The second generator is connected by the connection/disconnection mechanism to the aircraft or taxiing networks to supply the AC voltage to one of those networks.

Cooling arrangement for a generator

A generator arranged to be driven by an aircraft engine. The generator comprising a rotor. The rotor comprises an inlet for receiving a fluid, a plurality of outlets configured to release the fluid from a radially outer region of the rotor, and a fluid distribution arrangement arranged to direct fluid from the inlet to one or more of the plurality of outlets. The fluid distribution arrangement is configured to selectively distribute fluid to one or more of the plurality of outlets in dependence on an operational parameter of the rotor.

VARIABLE CYCLE COMPENSATION IN A GAS TURBINE ENGINE

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

COMPOSITE PANEL COMPRISING AN INTEGRATED ELECTRICAL CIRCUIT AND MANUFACTURING METHOD THEREOF

Here are described composite panels comprising at least one integrated or embedded electrical circuit, their methods of manufacturing and their use in the aeronautic and aircraft industries. Also described are aircraft components comprising the composite panel as defined herein. 1. A composite panel comprising at least one integrated electrical circuit , the composite panel comprising:a matrix; and a first matrix-permeable fabric preform and a second matrix-permeable fabric preform extending in a superposed relationship;', 'at least one electrical circuit provided on a film substrate, wherein the at least one electrical circuit is placed between the first and second matrix-permeable fabric preforms, and the film substrate comprises at least one hole and/or opening; and', 'at least one electrical or electronic device electrically connected to the at least one electrical circuit., 'a matrix-permeable reinforced material assembly comprising2. The composite panel of claim 1 , wherein the first and the second matrix-permeable fabric preforms are dry fabric preforms or are pre-impregnated with the matrix.3. The composite panel of claim 1 , wherein the film substrate has a surface area of less than about 85% claim 1 , or less than about 75% claim 1 , or less than about 65% claim 1 , or less than about 55% claim 1 , or less than about 45% claim 1 , or less than about 35% claim 1 , or less than about 25% claim 1 , or less than 15% claim 1 , or less than about 5% of the surface area of the composite panel.4. The composite panel of claim 1 , wherein the first and the second matrix-permeable fabric preforms are independently selected from the group consisting of woven claim 1 , nonwoven claim 1 , knitted claim 1 , and braided fabric preforms claim 1 , and/or independently comprise high-performance multifilament fibers selected from the group consisting of glass fibers claim 1 , carbon fibers claim 1 , hybrid carbon-glass fibers claim 1 , aramid fibers claim 1 , boron fibers ...