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

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

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Применить Всего найдено 2014. Отображено 100.
13-06-2017 дата публикации

Система управления основными летными функциями самолета с помощью рулевых поверхностей с электромеханическими приводами

Номер: RU0000171721U1
Автор: Валентин Евгеньевич Урсу, Владимир Валентинович Урсу
Принадлежит: Общество с ограниченной ответственностью "Научно-инженерная компания "Объектные системы автоматики" (ООО "НИК "ОСА")

Полезная модель относится к области авиации, а именно к системам управления летными функциями самолетов с помощью электромеханических приводов (ЭМП) рулевых поверхностей (РП).Система РП с ЭМП для управления основными летными функциями самолета содержит: разделенные на секции РП для управления функциями курса, тангажа и крена; как минимум один следящий ЭМП, соединенный с каждой из секций разделенной РП, мощность которого зависит от площади приводимой им секции, выходные звенья которого совершают вращательное или линейное движение; поверхности аэродинамического торможения (AT), имеющие, по меньшей мере, один следящий ЭМП, выходное звено которого совершает вращательное или линейное движение; блоки управления (БУ) следящими ЭМП РП; БУ электродвигателями следящих ЭМП, соединенные шинами с БУ следящими ЭМП РП и поверхностей AT; один центральный БУ РП каждой из функций курса, тангажа и крена и поверхностями AT, соединенный интерфейсными шинами с БУ следящими ЭМП и с бортовым компьютером; датчики углового положения каждого выходного вала следящего ЭМП, соединенные с БУ следящими ЭМП; датчики положения каждой секции РП и поверхностей AT, соединенные с БУ следящими ЭМП РП, при этом каждая РП, управляющая одной из функций курса, тангажа и крена, разделена на внешнюю и внутреннюю секции так, что произведение площади каждой секции на кратчайшее расстояние от ее геометрического центра до продольной оси самолета является постоянной величиной; каждый электродвигатель является бесколлекторным вентильным, постоянного тока, с постоянными магнитами, имеет номинальную частоту вращения ротора, выбираемую из интервала 6000…60000 мин, а соединенный с валом ротора редуктор является волновым и имеет передаточное отношение, выбираемое из интервала 500…4000. 4 з.п. ф-лы, 9 ил. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 171 721 U1 (51) МПК B64C 13/50 (2006.01) B64C 13/28 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ФОРМУЛА ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ РОССИЙСКОЙ ФЕДЕРАЦИИ (21)( ...

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Номер записи: 1
28-05-2018 дата публикации

Гидравлическая система уборки-выпуска закрылков самолета

Номер: RU0000179892U1
Автор: Виктор Иванович Сабельников, Вячеслав Александрович Трунов, Вячеслав Витальевич Писарев, Юрий Витальевич Колеватов
Принадлежит: Федеральное государственное унитарное предприятие "Сибирский научно-исследовательский институт авиации им. С.А. Чаплыгина"

Полезная модель относится к области авиации, в частности к гидравлической системе управления рулевыми поверхностями самолета.Гидравлическая система уборки-выпуска закрылков самолета содержит гидравлический блок питания, насосную станцию и гидропривод, выполненный на базе высокомоментного низкооборотного планетарного гидромотора и узлов гидросистемы, объединенных в один блок, обеспечивающий работу от двух гидросистем - основной и аварийной, и дистанционной системой управления по двум каналам.Технический результат заключается в уменьшении массы и габаритов гидропривода, улучшении технологичности и условий эксплуатации самолета. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 179 892 U1 (51) МПК B64C 13/36 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ (52) СПК B64C 13/36 (2006.01) (21)(22) Заявка: 2017129149, 15.08.2017 (24) Дата начала отсчета срока действия патента: Дата регистрации: 28.05.2018 (45) Опубликовано: 28.05.2018 Бюл. № 16 2417923 C2, 10.05.2011. RU 2251513 C1, 10.05.2005. RU 149760 U1, 20.01.2015. (54) Гидравлическая система уборки-выпуска закрылков самолета (57) Реферат: Полезная модель относится к области авиации, узлов гидросистемы, объединенных в один блок, в частности к гидравлической системе управления обеспечивающий работу от двух гидросистем рулевыми поверхностями самолета. основной и аварийной, и дистанционной системой Гидравлическая система уборки-выпуска управления по двум каналам. закрылков самолета содержит гидравлический Технический результат заключается в блок питания, насосную станцию и гидропривод, уменьшении массы и габаритов гидропривода, выполненный на базе высокомоментного улучшении технологичности и условий низкооборотного планетарного гидромотора и эксплуатации самолета. R U 1 7 9 8 9 2 (56) Список документов, цитированных в отчете о поиске: EP 2727831 A1, 07.05.2014. RU Стр.: 1 U 1 U 1 Адрес для переписки: 630051, Новосибирск, ул. Ползунова, 21, ФГУП "Сибирский научно- ...

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Номер записи: 2
08-10-2021 дата публикации

Блок питания гидравлической системы самолета

Номер: RU0000207050U1
Автор: Сергей Васильевич Семенов
Принадлежит: Общество с ограниченной ответственностью "Байкал-Инжиниринг"

Полезная модель относится к бортовым источникам гидравлической энергии. Заявляемый блок питания гидравлической системы самолета предназначен для выработки энергии, использующейся для работы самолетных подсистем.Технические результаты предлагаемой полезной модели заключаются в расширении энергетических возможностей блоков питания гидравлических систем самолета, содержащим гидробак малого объема, за счет обеспечения возможности введения дополнительных некомпенсированных потребителей, с целью использования блока питания для гидропитания потребителей гидравлических систем с суммарным некомпенсированным объемом, близким к объему гидробака при появлении в системе дополнительных некомпенсированных потребителей, а также в экономии ресурса насосной станции за счет снижения частоты включения и, таким образом, продления ее календарного срока службы.Технические результаты обеспечиваются тем, что блок питания гидравлической системы включает в себя насосную станцию, в состав которой входят электроприводной гидронасос, электрогидрокран и гидробак малого объема, питающий гидронасос, основную напорную линию между электрогидрокраном и выходом электроприводного гидронасоса, линию слива рабочей жидкости в гидробак, а также по крайней мере одну дополнительную напорную линию высокого давления, в состав которой входит, в частности, гидроаккумулятор. Гидроаккумулятор состоит из корпуса, разделенного плавающим плунжером на две полости, одна из которых заполнена газом, а вторая разделена перегородкой на две гидравлические камеры одинаковой вместимости - высокого и низкого давления, причем плавающий плунжер связан тягой, проходящей через перегородку, с поршнем, расположенным в гидравлической камере низкого давления. Дополнительная напорная линия высокого давления снабжена также обратным клапаном и клапаном зарядным, соединенным с газовой полостью гидроаккумулятора. При этом гидравлическая камера высокого давления гидроаккумулятора соединена с первым сигнализатором и обратным клапаном, ...

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Номер записи: 3
05-04-2012 дата публикации

Device for actuating a control surface of an aircraft

Номер: US20120080557A1
Автор: Fernand Rodrigues
Принадлежит: Sagem Defense Securite SA

The invention relates to a device ( 4 ) for actuating a control surface ( 2 ) of an aircraft, comprising:-a frame intended to be mounted fixed in relation to a structure of an aircraft, -a bell crank part ( 7 ) mobile in rotation in relation to the frame around an axis of rotation (X, X′) and adapted in order to be connected to a mechanical unit ( 1 ) for displacement of the control surface ( 2 ), -a slider ( 6 ) intended to be connected to a drive member ( 5 ), the slider ( 6 ) being mobile in translation in relation to the frame according to a direction of translation, parallel to the axis of rotation (X, X′) of the bell crank part ( 7 ), -first connecting means ( 61 ) between the slider ( 6 ) and the bell crank part ( 7 ) in order to convert a displacement in translation of the slider ( 6 ) generated by the drive member ( 5 ) into a displacement in rotation of the bell crank part ( 7 ), in order to actuate the mechanical unit ( 1 ) for displacement of the control surface ( 2 ),an input part ( 53, 16 ) able to be driven in rotation by the motor, -a control rod mounted fixed in relation to the frame and extending parallel to the axis of rotation (X) of the input part and at a distance from the latter, in order to block in rotation the slider in relation to the frame and authorize a translation of the slider in relation to the frame, the control rod extends through the slider ( 6 ). Application specific to the actuating of control surfaces incorporated into a thin airfoil.

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Номер записи: 4
03-01-2013 дата публикации

Horizontal stabilizer trim actuator failure detection system and method using position sensors

Номер: US20130001357A1
Автор: Luc P. Cyrot
Принадлежит: Individual

An actuator assembly having a primary load path for tightly coupling an actuated surface to a reference structure and a secondary load path having a backlash portion for coupling the actuated surface to the reference structure with backlash, wherein the secondary load path is unloaded during an operative state of the primary load path and loaded during a failure state of the primary load path. The actuator assembly includes at least one sensor configured to sense the failure state of the primary load path when a relative displacement between a portion of the primary load path and a portion of the secondary load path exceeds a predetermined value or is within a predetermined range of values.

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Номер записи: 5
21-03-2013 дата публикации

MULTIPLE-ROW EPICYCLIC GEAR

Номер: US20130072344A1
Автор: Ratmanov Eduard Vladimirovich, Volkov Gleb Yurievich
Принадлежит: STATOIL ASA

A multiple-row epicyclic gear includes a carrier with planet gears arranged in rows, as well as a composite central ring gear, having major crowns, with friction spacer rings disposed there between, are also mounted in the housing. The elements of the central ring gear are axially spring loaded. The end faces of the frictional spacer rings and counter-oriented end faces of the major crowns have tapered surfaces being in frictional contact. The friction spacer rings are spliced and arranged to bring their outer cylindrical surfaces into frictional contact with the inner cylindrical surface of the housing. The planet gears are disposed on the carrier in rows in corresponding apertures (or through openings) formed in the carrier. The optimal cone apex angle (δ) of the contacting tapered end faces of the major crowns and tapered end faces of the friction spacer rings is in the range 60-90°. 1. A multiple-row epicyclic gear comprising a housing with an externally toothed central shaft mounted therein; a carrier with planet gears mounted in rows thereon; and a central internally toothed ring gear being composite and having major crowns , and friction spacer rings disposed between said major crowns , said major crowns and said friction spacer rings being pressed in an axial direction , wherein end faces of the friction spacer rings and counter-oriented end faces of the major crowns have tapered surfaces being in frictional contact , wherein the friction spacer rings are spliced and arranged to bring their outer cylindrical surfaces into frictional contact with the inner cylindrical surface of the housing.2. The multiple-row epicyclic gear according to claim 1 , wherein the planet gears are mounted in rows in corresponding apertures or through openings formed in the carrier so that each planet gear is disposed in a separate aperture or through opening.3. The multiple-row epicyclic gear according to claim 1 , wherein the externally toothed central shaft is single-crowned and ...

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Номер записи: 6
02-05-2013 дата публикации

DEVICE FOR DETECTING BREAKAGE OF A PRIMARY PATH IN A FLIGHT CONTROL ACTUATOR

Номер: US20130105623A1
Автор: Medina Raphael, Moulon Thomas
Принадлежит:

A device () for detecting the breakage of a primary path in a flight control actuator, said actuator having a primary path () comprising a rotary hollow screw (), a secondary path () comprising a safety rod () that reacts the load passing through the screw (), said device () being characterized in that it comprises a position sensor (), connected to the screw (), to measure information representative of the angular position thereof, and a disconnection system () able to disconnect the screw () position sensor () in the event of relative movement of the rod () with respect to the screw () if there is a break in the primary path (). 1131210321315217152321. A device () for detecting breakage of a primary load path in a flight control actuator , said actuator having a primary load path () comprising a hollow rotary screw () , a secondary load path () comprising a load-assuming safety rod () passing through the screw () , said device () being characterized in that it includes a position sensor () , connected to the screw () for measuring information representing its angular position , and a disconnection system () , capable of disconnecting the position sensor () from the screw () in the event of relative displacement of the rod () with respect to the screw () upon a break of the primary load path ().2181521917. The device according to claim 1 , additionally including a calculator () configured to compare the information measured by the position sensor () claim 1 , and information representing the angular position of the screw () measured by a second position sensor () independent of the disconnection system ().3181. The device according to claim 2 , wherein the calculator () is configured to a break in the primary load path () when the comparison is greater or less than a predetermined threshold.4317152323221. The device according to one of through claims 1 , wherein the disconnection system () is capable of disconnecting the position sensor () from the screw () when ...

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Номер записи: 7
16-05-2013 дата публикации

TRAILING-EDGE FLAP SYSTEM

Номер: US20130119194A1
Автор: Dostmann Tom, Eichhorn Jochen, Schlipf Bernhard
Принадлежит: AIRBUS OPERATIONS GMBH

A trailing-edge flap system is described. The trailing-edge flap system includes a trailing-edge flap and a movement device. The movement device includes at least a translatory mover for the translation of the trailing-edge flap and at least a rotational mover for the rotation of the trailing-edge flap. 1. A trailing-edge flap system comprising a trailing-edge flap and a movement device that moveably couples the trailing-edge flap to an aerofoil , the movement device comprising at least a translatory mover that translatory moves the trailing-edge flap and at least a rotational mover that rotationally moves the trailing-edge flap ,wherein the translatory mover is configured as a telescope drive comprising several individual telescopic boxes which are arranged nested in one another and can be moved relative to one another for moving the trailing-edge flap along a line of translatory movement, wherein, when seen from the aerofoil, the trailing-edge flap is rotationally coupled to a last telescopic box whereby a pivot bearing point is defined,whereinthe rotational mover is configured as a lever kinematics comprising: a stationary pivot being fixed to the aerofoil and being provided for a single lever that is coupled to the trailing-edge flap in a second pivot at the end of a connection bar.2. The trailing-edge flap system according to claim 1 , wherein the pivot bearing point lies in the region from 20% to 60% of the flap depth of the trailing-edge flap in the direction of flight.3. The trailing-edge flap system according to claim 1 , wherein a translation angle between the line of translatory movement and the chord line of the aerofoil belonging to the trailing-edge flap lies within the range between 10 degrees and 20 degrees such that the pivot bearing point for the pivot of rotation of the trailing-edge flap can be given from the aerodynamic demands placed on the translatory movement and rotational movement claim 1 , which pivot bearing point in particular lies ...

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Номер записи: 8
23-05-2013 дата публикации

DEPLOYMENT SYSTEM

Номер: US20130126670A1
Автор: Vaghela Naresh
Принадлежит: AIRBUS OPERATIONS LIMITED

A deployment mechanism is disclosed for deploying a deployable member supported on tracks relative to a base member and comprising a brake system operable in response to relative skewing of the tracks. 1. A deployment system for deploying a deployable member relative to a base member , said deployment mechanism comprising:a base member;a deployable member for deployment relative to said base member;a plurality of tracks fixed to said deployable member, supported by said base member and operable for parallel simultaneous movement relative to said base member;actuator means operable to drive said tracks so as to deploy said deployable member;sensor means operable to detect relative skew of said tracks; andbrake means operable in response to detection of said relative skew by said sensor means to brake said track so as to substantially prevent further skewing of said tracks.2. A deployment system according to further comprising limiter means associated with said actuator means claim 1 , said limiter means being arranged to limit the force applied by said actuator means to said tracks.3. A deployment system according to in which said actuator is rotary and said limiter means comprises a torque limiter.4. A deployment system according to in which said sensor means are provided for each said track.5. A deployment system according to in which a plurality of said sensors are provided for each track.6. A deployment system according to in which said brake means is provided for each sensor.7. A deployment system according to in which said brake means is provided adjacent the or each said sensor.8. A deployment system according to in which each said sensor is mechanically operable to detect said relative skew.9. A deployment system according to in which said brake means is operated mechanically by said sensor in response to said detection of said relative skew.10. A deployment system according to in which said mechanical operation comprises a predetermined amount of float.11. A ...

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Номер записи: 9
13-06-2013 дата публикации

WING CONTROL SYSTEM

Номер: US20130146705A1
Автор: Van Zyl Louwrens Hermias
Принадлежит: CSIR

The invention provides a wing control system aimed at countering the aeroelastic effect of twisting of a length of wing of an aircraft due to dynamic air pressure acting on an aileron of the wing. The wing control system includes a shaft extending along the length of wing and actuation means responsive to aileron control inputs to induce a variable torque T in the shaft. Outboard of the length of wing, the system operatively transfers T partially to the aileron to pivot the aileron and partially to the wing at an outboard end of the length of wing to counter twisting of the length of wing due to dynamic air pressure on the aileron. Inboard of the length of wing, the system operatively transfers T to the aircraft, e.g. to its fuselage, thereby effectively balancing the sum of the moments transferred respectively to the aileron and the wing. 1. A wing control system operatively installed on an aircraft , the aircraft including a wing including an aileron and a length of wing in which dynamic air pressure on the aileron , when pivoted during forward motion of the aircraft , would tend to induce twisting , the system including:a shaft extending along the length of wing and including an outboard end and an inboard end;actuation means responsive to aileron control inputs received from an aileron control system of the aircraft to induce a variable torque T in the shaft;outboard moment transfer means for transferring the torque T from the outboard end of the shaft partially as a moment MA to the aileron and partially as a moment MW to the wing at an outboard end of the length of wing, with the moment MW being in the same direction as the moment MA to counter twisting of the length of wing due to dynamic air pressure on the aileron, when pivoted during forward motion of the aircraft, and with the moments MA and MW together balancing the torque T; andinboard moment transfer means for transferring the torque T from the inboard end of the shaft to the aircraft at a position ...

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Номер записи: 10
20-06-2013 дата публикации

Automatically locking linear actuator

Номер: US20130152717A1
Автор: Joseph Thomas Kopecek
Принадлежит: GE AVIATION SYSTEMS LLC

An automatically locking actuator includes a first end connector, a rotatable drive screw operably coupled to the first end connector, a nut assembly threadably mounted on the drive screw, a second end connector operably coupled to the nut assembly and a rotary lock having a rotor and where the actuator moves between extended and retracted positions in response to rotation of the rotor.

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Номер записи: 11
20-06-2013 дата публикации

TRANSLATING CABLE DEVICE SEALING

Номер: US20130153713A1
Автор: HADLEY Paul, TEODORU Alexandru, Wilson James
Принадлежит: AIRBUS OPERATIONS LIMITED

An aircraft wing assembly, comprising a wing having a fixed leading edge, a slat mounted for movement between a retracted position and an extended position with respect to the fixed leading edge, and a translating cable device for electrically connecting the slat to the wing and having a strut coupled at one end to the slat, the fixed leading edge having an aperture to accommodate the strut, and a seal assembly for sealing between the strut and the aperture. 1. An aircraft wing assembly , comprising a wing having a fixed leading edge , a slat mounted for movement between a retracted position and an extended position with respect to the fixed leading edge , and a translating cable device for electrically connecting the slat to the wing and having a strut coupled at one end to the slat , the fixed leading edge having an aperture to accommodate the strut , and a seal assembly for sealing between the strut and the aperture.2. An aircraft wing assembly according to claim 1 , wherein the seal assembly includes a first seal fixed adjacent the aperture claim 1 , and a second seal fixed to the strut of the translating cable device.3. An aircraft wing assembly according to claim 2 , wherein the first and second seals cooperate when the slat is moved to one or more predetermined positions.4. An aircraft wing assembly according to claim 1 , wherein the translating cable device has a proximal end mounted to the wing and a distal end coupled to the slat.5. An aircraft wing assembly according to claim 1 , wherein the seal assembly includes a flap seal mounted to the fixed leading edge and projecting into the aperture.6. An aircraft wing assembly according to claim 5 , wherein the flap seal includes a plurality of flap seal sections with a gap between adjacent sections.7. An aircraft wing assembly according to claim 5 , wherein the flap seal is mounted to a portion of the fixed leading edge substantially surrounding the aperture.8. An aircraft wing assembly according to claim 5 , ...

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Номер записи: 12
20-06-2013 дата публикации

MULTI-STAGE REDUCTION GEAR

Номер: US20130157795A1
Автор: Kobayashi Tsune, MISADA Yasuharu, OONO Yoshihiro, Yamamori Motoyasu
Принадлежит: JTEKT CORPORATION

The multi-stage reducer includes a first-stage reduction part to which the rotation from a motor is input and a second-stage reduction part to which the rotation having undergone speed reduction by the first-stage reduction part is input. The second-stage reduction part is disposed on an outer circumference of the first-stage reduction part. 1. A multi-stage reducer comprising a plurality of reduction parts each comprising an input shaft to which a rotation from a drive source is input , a cam provided on the input shaft , an external gear which is supported by an outer periphery of the cam and which revolves around a rotation center of the input shaft , a carrier which is coupled with the external gear and which is linked with a rotation movement of the external gear , and an internal gear which is meshed with the external gear and which rotates around the rotation center of the input shaft , the multi-stage reducer being configured to reduce the rotation from the drive source at a predetermined reduction ratio based on a difference between a number of gear teeth of the external gear and a number of gear teeth of the internal gear and to output the rotation having undergone speed reduction from either one of the internal gear and the carrier , whereinthe plurality of reduction parts includes:a first-stage reduction part to which the rotation from the drive source is input; anda second-stage reduction part to which the rotation having undergone speed reduction by the first-stage reduction part is input, andthe second-stage reduction part is disposed on an outer circumference of the first-stage reduction part.2. The multi-stage reducer according to claim 1 , whereinthe first-stage reduction part comprises a first input shaft, a first cam, a first external gear, a first carrier, and a first internal gear,the second-stage reduction part comprises a second input shaft, a second cam, a second external gear, a second carrier, and a second internal gear,the second cam is ...

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Номер записи: 13
18-07-2013 дата публикации

ADJUSTMENT SYSTEM OF AN AEROPLANE WITH AN ADJUSTABLE FLAP

Номер: US20130181089A1
Автор: Ellmers Wilfried, Recksiek Martin, Winkelmann Christoph
Принадлежит: AIRBUS OPERATIONS GMBH

An adjustment system of an aeroplane, having: at least one adjustable flap on each of the wing of an aeroplane, adjustable with an adjustment device, a drive device for purposes of driving the adjustment devices, and a load sensor for purposes of recording the load occurring in the load path between the actuator and the adjustable flap of the respective adjustment device. The load sensor is embodied as a sensor for purposes of measuring the longitudinal force occurring in a drive rod along its longitudinal direction. 1. An adjustment system for a fault-tolerant aeroplane adjustable flap system , comprising:at least one adjustable flap, supported by an articulation device and adjustable on each of the wings of an aeroplane,at least one or more adjustment device for adjustment of the adjustable flap, wherein each adjustment device comprises: an actuator and a kinematic adjustment mechanism for kinematic coupling of the actuator to the adjustable flap with a drive rod, which via a first articulation is coupled to the actuator and via a second articulation is coupled to the adjustable flap,at least one drive device for driving the adjustment device, anda load sensor for purposes of capturing the load occurring in the load path between the actuator and the adjustable flap of the respective adjustment device, which is suitable for purposes of measuring the longitudinal force occurring in a drive rod along its longitudinal direction,a control and monitoring device, functionally connected with the at least one drive device, for adjustment of the adjustment device, which control and monitoring device is functionally connected with the load sensor for receiving the sensor signals generated by the load sensor.2. The adjustment system in accordance with claim 1 , characterised in that the load sensor is formed from at least one strain gauge claim 1 , which is fitted to the drive rod.3. The adjustment system in accordance with claim 1 , characterised in thatthe drive rod is ...

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Номер записи: 14
18-07-2013 дата публикации

HAND-HELD TOOL DEVICE

Номер: US20130184116A1
Автор: HERR Tobias
Принадлежит:

A hand-held tool device includes: a tool spindle; a striking mechanism; and a planetary transmission having at least one first planetary transmission stage which drives the striking mechanism, a second planetary transmission stage which drives the tool spindle, and a striking mechanism shut-off clutch. 1. A hand-held tool device , comprising:a tool spindle;a striking mechanism; anda planetary transmission having (i) at least one first planetary transmission stage which drives the striking mechanism, (ii) a second planetary transmission stage which drives at least the tool spindle, and (iii) a striking mechanism shut-off clutch.2. The hand-held tool device as recited in claim 1 , wherein the striking mechanism shut-off clutch is situated between the first planetary transmission stage and the second planetary transmission stage.3. The hand-held tool device as recited in claim 2 , wherein the second planetary transmission stage drives the first planetary transmission stage in at least one operating state.4. The hand-held tool device as recited in claim 2 , wherein the striking mechanism shut-off clutch has a clutch element which is supported in an axially displaceable manner.5. The hand-held tool device as recited in claim 4 , wherein the tool spindle transfers an axial clutching force in at least one operating state.6. The hand-held tool device as recited in claim 2 , wherein the striking mechanism shut-off clutch has a clutch element which is connected torsionally fixed to a planet carrier of the first planetary transmission stage.7. The hand-held tool device as recited in claim 2 , wherein the striking mechanism shut-off clutch has a clutch element which is connected torsionally fixed to a planet carrier of the second planetary transmission stage.8. The hand-held tool device as recited in claim 7 , wherein the planet carrier of the second planetary transmission stage is configured in at least two parts.9. The hand-held tool device as recited in claim 7 , wherein the ...

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Номер записи: 15
25-07-2013 дата публикации

PLANETARY GEAR SYSTEM, PLANETARY DIFFERENTIAL AND GEAR SYSTEM WITH THE PLANETARY DIFFERENTIAL

Номер: US20130190130A1
Автор: Biermann Thorsten, Doerrfuss Florian, Smetana Tomas
Принадлежит: SCHAEFFLER TECHNOLOGIES AG & CO. KG

A planetary gear system having at least two planetary steps, each of which is made up of at least one set of planets and one sun, whereby the teeth of the gearwheels in the planetary drive mesh with each other in such a way that, in each of the meshing points, at least one first tooth on a first toothing of first teeth positively engages into a tooth gap of a second toothing of second teeth. First teeth of the first toothing have a tooth flank profile concavely arched, and second teeth of the second toothing have a tooth flank profile convexly arched, so that the tooth flanks in contact with each other are arched in the same directions, at least when in contact. 110-. (canceled)11. A planetary gear system comprising:at least two planetary steps, each planetary step including at least one set of planets and one sun, teeth of gearwheels in the planetary gear system meshing with each other in such a way that, in each of the meshing points, at least one first tooth on a first toothing of first teeth positively engages into a tooth gap of a second toothing of second teeth, a first flank of the first tooth touching at least a second flank of a second tooth of the second toothing that delimits one side of the tooth gap on the second toothing in at least one point of tooth contact, the first teeth of the first toothing having a tooth flank profile that, as seen in the cross section through the toothing when the teeth mesh, is concavely arched, and in that the second teeth of the second toothing have a tooth flank profile convexly arched in the same cross section, so that the tooth flanks of the first tooth and of the second tooth, which are in contact with each other, are arched in the same directions, at least when they are in contact with each other.12. The planetary gear system as recited in wherein the sun is a sun gearwheel of the gearwheels with the first toothing claim 11 , and in that the planets of the set are each planet gearwheels of the gearwheels with the ...

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Номер записи: 16
22-08-2013 дата публикации

AERODYNAMIC BODY WITH AN ANCILLARY FLAP

Номер: US20130214092A1
Автор: Friedel Hendrik
Принадлежит: AIRBUS OPERATIONS GMBH

An aerodynamic body with at least one ancillary flap on the aerodynamic body so it can be moved with a guide mechanism, and with a drive device for actuating the ancillary flap. The guide mechanism has a connecting lever, which at its first end is articulated on the aerodynamic body by a first pivotal articulation, and which at its second end is articulated on the ancillary flap by a second pivotal articulation. The second pivotal articulation is located at some distance from a trailing edge of the ancillary flap, and from a leading edge of the ancillary flap. The guide mechanism has an actuation element, which is coupled with the drive device. The actuation element is articulated on the ancillary flap by a third pivotal articulation. The third pivotal articulation is arranged such that it can be moved in the chordwise direction of the aerodynamic body. 2. The aerodynamic body in accordance with claim 1 , characterised in that the aerodynamic lifting body is designed as a main wing of a wing.3. The aerodynamic body in accordance with claim 1 , characterised in that the aerodynamic lifting body is designed as a flap claim 1 , which is arranged on a main wing of a wing such that it can be extended.4. The aerodynamic body in accordance with claim 1 , characterised in that the ancillary flap is arranged on the pressure surface of the aerodynamic body claim 1 , and is arranged such that it can be extended in the direction of the pressure surface.5. The aerodynamic body in accordance with claim 1 , characterised in that the ancillary flap is arranged on a suction surface of the aerodynamic body claim 1 , and is arranged such that it can be extended in the direction of the suction surface.6. The aerodynamic body in accordance with claim 1 , characterised in that the third pivotal articulation is articulated on the ancillary flap in the forward region of the ancillary flap claim 1 , wherein the forward region with the ancillary flap retracted claim 1 , is designed at the ...

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Номер записи: 17
29-08-2013 дата публикации

Device for generating return forces for sticks such as aircraft control sticks

Номер: US20130220065A1
Автор: Bernard Robic, Christian Gouard, Herve Gorecki
Принадлежит: Sagem Defense Securite SA

A device ( 1 ) for generating a return force for a control stick ( 2 ) that is movable from a neutral position along a travel path ( 3 ), the device comprising mechanical connection means ( 5 ) for connecting the stick ( 2 ) to resilient return means ( 4 ) for returning the stick ( 2 ) towards the neutral position, the device being characterized in that the mechanical connection means ( 5 ) and the resilient return means ( 4 ) are arranged to produce, over a first portion ( 3 a ) of the travel path, a first strength of return force (F) and, over a second portion ( 3 b ) of the travel path, a second strength of return force (F), and in that the mechanical connection means ( 5 ) include a transmission part ( 6 ) adapted firstly to be mechanically connected to said stick ( 2 ) in such a manner that any movement of said stick ( 2 ) along the second portion ( 3 b ) of the travel path corresponds to turning movement of the transmission part ( 6 ), and secondly to transmit the return force (F) to said stick ( 2 ) at least when the stick ( 2 ) is positioned in the second portion ( 3 b ) of the travel path.

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Номер записи: 18
12-09-2013 дата публикации

LINKAGE FOR GUIDING A FLEXIBLE CABLE

Номер: US20130233967A1
Автор: Biggadike Christopher, Dutton Graeme John, Salthouse Mark Anthony
Принадлежит: ULTRA ELECTRONICS LIMITED

A linkage guides a flexible cable between two structures. The linkage includes a proximal arm with a proximal pivot joint for coupling the proximal arm to a first one of the two structures; and a distal arm which is coupled to the proximal arm by one or more intermediate pivot joints. The distal arm is shaped to follow a three-dimensional curve along a majority of its length. Shaping the distal arm to form a three-dimensional curve along a majority of its length enables the distal arm to pass through a relatively small aperture as the linkage is adjusted between its retracted and extended positions. It also enables the proximal arm to move in a locus of movement which does not interfere with other system components. The linkage can be used to guide a flexible cable between any two structures, for instance a fixed aircraft wing and a slat. 1. A linkage for guiding and protecting a flexible cable between first and second structures , the second structure being movable relative to the first structure , the linkage comprisinga proximal arm;a proximal pivot joint for coupling the proximal arm to the first structure; anda distal arm which is coupled to the proximal arm by one or more intermediate pivot joints, wherein the distal arm is shaped as a three-dimensional curve along a majority of its length.2. The linkage of wherein the intermediate pivot joint(s) permit the distal arm to rotate relative to the proximal arm about two or more axes of rotation.3. The linkage of wherein the distal arm is coupled to the proximal arm by a ball joint.4. The linkage of wherein the ball joint permits the distal arm to rotate relative to the proximal arm about three axes of rotation.5. The linkage of further comprising a distal pivot joint for coupling the distal arm to the second structure.6. The linkage of claim 5 , wherein the distal pivot joint permits the distal arm to rotate relative to the second structure about at least two axes of rotation.7. The linkage of wherein the distal ...

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Номер записи: 19
12-09-2013 дата публикации

COMPOUND PLANETARY GEAR MECHANISM

Номер: US20130237368A1
Автор: KISHI Yasuo, Nakamura Hiroshi, ZHANG Wennong
Принадлежит: KABUSHIKI KAISHA YASKAWA DENKI

A compound planetary gear mechanism includes at least two planetary gear mechanisms and a carrier. The at least two planetary gear mechanisms include a first planetary gear mechanism and a second planetary gear mechanism. The first planetary gear mechanism includes a plurality of first planetary gears. The second planetary gear mechanism includes a plurality of second planetary gears. The carrier is coupled to the plurality of first planetary gears and the plurality of second planetary gears. The carrier includes a first support shaft and a second support shaft. The first support shaft rotatably supports a first planetary gear among the plurality of first planetary gears. The second support shaft rotatably supports a second planetary gear among the plurality of second planetary gears and is independent of the first support shaft. 1. A compound planetary gear mechanism comprising: a first planetary gear mechanism comprising a plurality of first planetary gears; and', 'a second planetary gear mechanism comprising a plurality of second planetary gears; and, 'at least two planetary gear mechanisms comprising a first support shaft rotatably supporting a first planetary gear among the plurality of first planetary gears; and', 'a second support shaft rotatably supporting a second planetary gear among the plurality of second planetary gears and being independent of the first support shaft., 'a carrier coupled to the plurality of first planetary gears and the plurality of second planetary gears, the carrier comprising2. The compound planetary gear mechanism according to claim 1 , wherein the first support shaft and the second support shaft are disposed at different positions on the carrier in at least one direction among a radial direction and a circumferential direction.3. The compound planetary gear mechanism according to claim 1 ,wherein the plurality of first planetary gears each comprise a first bearing at a center portion of each of the plurality of first planetary ...

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Номер записи: 20
24-10-2013 дата публикации

HIGH-LIFT SYSTEM FOR A WING OF AN AIRCRAFT

Номер: US20130277498A1
Автор: Winkelmann Christoph
Принадлежит: AIRBUS OPERATIONS GMBH

A high-lift system for a wing of an aircraft is provided. The high-lift system includes movably held high-lift flaps, at least one drive unit, at least one transmission shaft connected to the drive unit, and several actuator devices, distributed on the transmission shaft and connected to the high-lift flaps, for moving the high-lift flaps. The actuator devices each comprise a driven element and a torque limiting means. In one example, the driven elements of two adjacent actuator devices are interconnected in a non-rotational manner with the use of a separate torque transmitting means. If one actuator device is blocked, for example as a result of a defective tooth arrangement or some other defect, the torque to be produced by the intact actuator device increases and triggers its torque limiting means. This ensures synchronous operation and in the case of malfunction prevents damage to or detachment of a flap. 1. A high-lift system for a wing of an aircraft , comprising:movably held high-lift flaps;at least one drive unit;at least one transmission shaft connected to the drive unit; andseveral actuator devices, distributed on the transmission shaft and connected to the high-lift flaps, for moving the high-lift flaps,wherein the actuator devices each include a driven element and a torque limiting means, and the driven elements of two adjacent actuator devices are interconnected in a non-rotational manner with the use of a separate torque transmitting means.2. The high-lift system of claim 1 ,wherein the torque transmitting means is a torsion shaft.3. The high-lift system of claim 2 ,wherein the torsion shaft is a hollow shaft through which the transmission shaft is fed.4. The high-lift system of claim 2 ,wherein the torsion shaft is connected to a driven element on both ends by means of an offset gear arrangement each, wherein the torsion shaft extends at a distance from and parallel to the transmission shaft.5. The high-lift system of claim 1 ,wherein high-lift flaps ...

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Номер записи: 21
28-11-2013 дата публикации

Tie Rod Lock

Номер: US20130313358A1
Автор: Hale Brian Curtis
Принадлежит: PARKER-HANNIFIN CORPORATION

A stabilizer actuator has a first end for connecting to an aircraft support structure and a second end for connecting to a stabilizer. The actuator includes a primary load path for transmittal of loads acting on the stabilizer to the aircraft support structure, and a secondary load path for transmittal of loads acting on the stabilizer to the aircraft support structure upon failure of the primary load path. The secondary load path includes a tie rod extending along a longitudinal axis, a load path locking mechanism coupled to the tie rod, a lock housing having a central bore for receiving the locking mechanism, and at least one radially movable segment that, upon failure of the primary load path, moves radially to lock the tie rod to the lock housing against axial and/or radial movement. 1. A stabilizer actuator having a first end for connecting to an aircraft support structure and a second end for connecting to a stabilizer , the actuator including a primary load path for transmittal of loads acting on the stabilizer to the aircraft support structure , and a secondary load path for transmittal of loads acting on the stabilizer to the aircraft support structure upon failure of the primary load path , the secondary load path comprising:a tie rod extending along a longitudinal axis;a load path locking mechanism coupled to the tie rod;a lock housing having a central bore for receiving the locking mechanism; andat least one radially movable segment that, upon failure of the primary load path, moves radially to lock the tie rod to the lock housing against axial and/or radial movement.2. The stabilizer actuator of claim 1 , wherein an inner surface defining the central bore includes the at least one protrusion extending radially inward from the inner surface claim 1 ,wherein the locking mechanism includes the at least one segment, and the at least one segment is configured to move radially outward from a standby position to a failsafe position to prevent relative movement ...

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Номер записи: 22
12-12-2013 дата публикации

DEVICE FOR AN ADJUSTABLE FLAP OF A WING

Номер: US20130327887A1
Автор: Dyckrup Werner, Ellmers Wilfried
Принадлежит: AIRBUS OPERATIONS GMBH

The invention concerns a device for a adjustable flap adjustably mounted on a main wing surface of an aeroplane wing, in particular, a landing flap, with at least one adjustment unit for purposes of adjustment of the adjustable flap, which has an actuator arranged, or that can be arranged, on the main wing surface, and has a kinematic adjustment mechanism running between the actuator and the adjustable flap, wherein the adjustable flap is mechanically coupled with the actuator via the kinematic adjustment mechanism. At least one damping unit for purposes of damping a dynamic loading effected by the adjustable flap on the adjustment unit, which can occur as a result of a critical malfunction event occurring in the region of the adjustable flap, is arranged, or can be arranged, between the main wing surface and the adjustable flap. 1. A device for an adjustable flap adjustably mounted on a main wing surface of an aeroplane wing , in particular , a landing flap , with at least one adjustment unit for purposes of adjustment of the adjustable flap , which has an actuator arranged , or that can be arranged , on the main wing surface and a kinematic adjustment mechanism running between the actuator and the adjustable flap , wherein the adjustable flap is mechanically coupled with the actuator via the kinematic adjustment mechanism ,characterised in that at least one damping unit for purposes of damping a dynamic loading effected by the adjustable flap on the adjustment unit, which in particular can be brought about as a result of a critical malfunction event occurring in the region of the adjustable flap, is arranged, or can be arranged, between the main wing surface and the adjustable flap.2. The device in accordance with claim 1 , characterised in that the damping unit is designed and arranged relative to the adjustment unit such that it damps out movement of the adjustable flap in the event of a fracture in the kinematic adjustment mechanism.3. The device in accordance ...

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Номер записи: 23
19-12-2013 дата публикации

FLAP PANEL SHUTTLE SYSTEM AND METHOD THEREFOR

Номер: US20130334363A1
Автор: Lam Lawrence, Lam Michael
Принадлежит:

An aircraft control system is presented. The system includes a wing including a flap track, and a shuttle connected to the flap track and configured to slide along a length of the flap track. The system includes a flap panel pivotally attached to the shuttle at a flap pivot. The flap panel is configured to rotate about the flap pivot. When the shuttle is deployed along a length of the flap track, the shuttle is configured to prevent rotation of the flap panel about the flap pivot, and when the shuttle is withdrawn into a stowed position, the shuttle is configured to allow the flap panel to rotate about the flap pivot. 1. An aircraft , comprising:a shuttle movable between a first configuration and a second configuration; anda flap panel attached to the shuttle, wherein the flap panel can move with respect to the shuttle when the shuttle is disposed in the first configuration, and the flap panel cannot move with respect to the shuttle when the shuttle is disposed in the second configuration.2. The aircraft of claim 1 , wherein the shuttle is attached to a flap track and the shuttle is configured to move along a length of the flap track.3. The aircraft of claim 2 , wherein the shuttle is configured to be disposed in the first configuration when the shuttle is in a stowed position on the flap track and to be disposed in the second configuration when the shuttle is deployed along a length of the flap track.4. The aircraft of claim 1 , including a flap sensor configured to detect a position of the flap panel.5. A shuttle for use with an aircraft wing claim 1 , the shuttle comprising:a body;a roller connected to the body, the roller being configured to slide along a length of a flap track; anda locking arm connected to the body of the shuttle, the locking arm being configured to selectively engage a flap panel, wherein, when the locking arm is engaged to the flap panel, the flap panel cannot move with respect to the shuttle, and when the locking arm is not engaged to the ...

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Номер записи: 24
19-12-2013 дата публикации

SLAT SUPPORT ASSEMBLY

Номер: US20130334364A1
Автор: PARKER Simon John
Принадлежит:

A slat support assembly is disclosed. It comprises a slat support arm which is movable to deploy a slat from a leading edge of an aircraft wing about an axis of rotation of the arm and a slat mount on a slat which is coupled to one end of said slat support arm by a joint. The joint is configured to allow the slat mount to slide in a direction of the axis of rotation of the arm. 1. A slat support assembly comprising a slat support arm which is movable to deploy a slat from a leading edge of an aircraft wing about an axis of rotation of the arm and a slat mount on a slat which is coupled to one end of said slat support arm by a joint , wherein the joint is configured to allow the slat mount to slide in a direction of the axis of rotation of the arm.2. A slat support assembly according to claim 1 , wherein the joint comprises a bearing element mounted to the slat mount claim 1 , the bearing element being slidably received in a bearing sleeve on the slat support arm.3. A slat support assembly according to claim 2 , wherein the bearing element is configured to slide in a linear direction in the bearing sleeve.4. A slat support assembly according to claim 3 , wherein the bearing element comprises a shaft which is slidable in the bearing sleeve in a direction along a longitudinal axis of the shaft.5. A slat support assembly according to claim 4 , wherein the bearing sleeve comprises opposing bushes on the slat support arm and ends of the shaft are slidably received in the bushes claim 4 , and the shaft extends between the opposing bushes.6. A slat support assembly according to claim 5 , further comprising a yoke formed at the one end of the slat support arm with two spaced yoke members claim 5 , wherein the opposing bushes are formed in the yoke members.7. A slat support assembly according to claim 6 , wherein the slat mount is disposed between the yoke members.8. A slat support assembly according to claim 5 , wherein the slat mount is mountable to the shaft of the bearing ...

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Номер записи: 25
02-01-2014 дата публикации

DEVICE FOR MECHANICAL CONNECTION OF A CONTROL SURFACE TO A FIXED STRUCTURAL ELEMENT OF AN AIRCRAFT AND AIRCRAFT WING ELEMENT EQUIPPED WITH SAID DEVICE

Номер: US20140001309A1
Автор: Andrieu Laurent, Sagansan Alain, Tieys Pierre
Принадлежит:

A device for mechanically connecting a control surface to a fixed structural element of an aircraft, including a rotary actuator for driving an element that is securely connected to the control surface in rotation in relation to an element securely connected to the fixed structural element, around an articulation axis, as well as articulation elements for articulating this control surface around this axis. These articulation elements are able to support the control surface independently of the rotary actuator. This arrangement permits taking benefit from the advantageous properties of rotary actuators while allowing removal of the rotary actuator without having first to remove the control surface. 1. A device for mechanical connection of a control surface to a fixed structural element of an aircraft , comprising articulation means for connecting the control surface to said fixed structural element according to an axis of articulation , as well as driving means for driving the control surface in rotation relative to said fixed structural element about said axis of articulation , said driving means comprising at least one rotary actuator comprising a chassis frame and an output member displaceable in rotation about an output axis of the rotary actuator relative to said chassis frame , wherein said driving means comprise:first detachable means for rotationally securing said chassis frame of the rotary actuator to a first element of said fixed structural element and said control surface, these first detachable means aligning said output axis of the rotary actuator with said axis of articulation;second detachable means for rotationally securing said output member of the rotary actuator to a second element of said fixed structural element and said control surface distinct from said first element; andwherein said articulation means are separate from said rotary actuator and support said control surface independently of said rotary actuator.2. The device according to claim ...

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Номер записи: 26
02-01-2014 дата публикации

MECHANICAL CONTROL MIXER AND METHOD THEREFOR

Номер: US20140001313A1
Автор: Lam Lawrence, Lam Michael
Принадлежит:

A mechanical control mixer configured to couple to an aircraft is provided. An axle is mounted to a frame of the mechanical control mixer, and a barrel is configured to rotate about the axle. A central rod is disposed within the barrel. The central rod is configured to rotate with respect to the barrel. A roll control input is connected to the central rod. The roll control input is configured to cause the central rod to rotate within the barrel. Output control rods are connected to the central rod. The output control rods are connected to at least one control surface of the aircraft. An air brake input is connected to the barrel. The air brake input is configured to cause the barrel to rotate about the axle to move at least one of the output control rods 1. An aircraft , comprising: an axle mounted to a frame of the mechanical control mixer,', 'a barrel configured to rotate about the axle, and', 'a central rod disposed within the barrel, the central rod being configured to rotate with respect to the barrel;, 'a mechanical control mixer coupled to the aircraft, the mechanical control mixer includinga roll control input connected to the central rod, the roll control input being configured to cause the central rod to rotate within the barrel;output control rods connected to the central rod, the output control rods being connected to at least one control surface of the aircraft; andan air brake input connected to the barrel, the air brake input being configured to cause the barrel to rotate about the axle to move at least one of the output control rods.2. The aircraft of claim 1 , wherein the air brake input is configured to move between a number of distinct positions.3. The aircraft of claim 1 , wherein the air brake input is configured to selectively deploy at least one flap of the aircraft.4. The aircraft of claim 3 , wherein an amount of flap deployment is determined by a position of the air brake input.5. The aircraft of claim 1 , wherein the air brake input is ...

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Номер записи: 27
23-01-2014 дата публикации

MEASUREMENT OF THE INERTIAL PROPERTIES OF AN AIRCRAFT MOVABLE CONTROL SURFACE

Номер: US20140021294A1
Автор: Alonso Gago Justo, Valera Rodriguez Pedro
Принадлежит: AIRBUS OPERATIONS S.L.

A method for obtaining the inertial properties of a movable rotating around a hinge line in an aircraft control surface. The method includes the steps of removing the mechanical connections of the actuators from the movable, leaving the movable free to rotate around its hinge line, further balancing the movable and calculating in a first approach its coarse static momentum. The method includes refining the coarse static momentum and obtaining simultaneously a frictional momentum of the movable; incorporating an elastic element on the control surface and configuring a second order mechanical system; inducing forced oscillations on the movable at a certain frequency, this frequency being increased until it is sensibly close to the resonance frequency of the movable to produce a wave response; calculating, from the wave response, the momentum of inertia of the movable. 1. A method for obtaining the inertial properties of a movable rotating around a hinge line in an aircraft control surface , comprising the following steps:a) removing all mechanical connections of actuators from the movable, leaving the movable free to rotate around the hinge line, further balancing the movable and calculating in a first approach a coarse static momentum of the movable;b) refining the static momentum of the movable obtained in a), and obtaining simultaneously a frictional momentum of the movable;c) incorporating an elastic element on the control surface, configuring a second order mechanical system;d) inducing forced oscillations on the movable at a certain frequency, this frequency being increased until it is sensibly close to a resonance frequency of the movable to obtain a wave response;e) calculating, from the wave response in d), a momentum of inertia of the movable.2. The method according to claim 1 , wherein the forced oscillations induced in d) are produced at a frequency of around 2 Hz.3. The method according to claim 1 , wherein the forced oscillations induced in d) are such ...

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Номер записи: 28
13-02-2014 дата публикации

DRIVE SYSTEM FOR CONTROL SURFACES OF AN AIRCRAFT

Номер: US20140042269A1
Автор: Arriola David, Benischke Stefan, Thielecke Frank, Zantz Tanno
Принадлежит:

A drive system for driving control surfaces of an aircraft includes at least one drive unit, at least one main shaft connectable to the at least one drive unit and at least two adjusting units for each control surface to be driven. Each adjusting unit includes a differential, two rotary actuators and an adjustment lever. The differential has at least one input means and two output means and is adapted to transfer torque from the at least one input means to the two output means. The input means is connectable to the main shaft, the two rotary actuators each have a rotation input means and a motion output means. The rotation input means is connectable to one of the output means of the differential each and the adjustment lever is connected to the motion output means of both rotary actuators. 1. A drive system for driving control surfaces of an aircraft , comprising:at least one drive unit;at least one main shaft connectable to the at least one drive unit,at least one adjusting units for each control surface to be driven,wherein each adjusting unit comprises a differential, first and second rotary actuators and an adjustment lever, the differential having at least one input means and first and second output means and being adapted to transfer torque from the at least one input means to the first and second output means, the input means being couplable to the main shaft, the first and second rotary actuators each having a rotation input means and a motion output means, the respective rotation input means of each of the first and second rotary actuators being coupled to one of the first and second output means and the adjustment lever being coupled to the motion output means of the first and second rotary actuators.2. The drive system of claim 1 , wherein the input means is a carrier for holding at least one planetary gear wheel of the differential claim 1 , wherein the at least one planetary gear wheel mutually engages at least one sun gear wheel.3. The drive system of ...

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Номер записи: 29
06-03-2014 дата публикации

ACTUATION SYSTEM FOR A LIFT ASSISTING DEVICE AND LINED TRACK ROLLERS USED THEREIN

Номер: US20140061381A1
Автор: Fredericksen Arnold, Gay Arthur, Gyuricsko Frederick S., Patel Nirad, Phoenix Jay
Принадлежит: ROLLER BEARING COMPANY OF AMERICA, INC.

An actuation system for deploying and retracting a lift assisting device of a leading edge of a wing of an aircraft including a track pivotally coupled to the lift assisting device. The track has first and second outer surfaces and side surfaces. The actuation system includes a shaft rotationally coupled within the wing of the aircraft and operable, in response to flight control signals, to deploy or retract the lift assisting device. The actuation system includes an actuator for actuating the lift assisting device, coupled to the shaft, between a retracted position to a deployed position along an arcuate path. The actuation system includes a plurality of track roller bearings rotatably contacting the first and second outer surfaces of the track to guide the track along the arcuate path. The plurality of track roller bearings includes one or more lined track roller assembly. 2. The actuation system of claim 1 , wherein the plurality of track roller bearings includes at least one track roller assembly in rotational contact with an upper surface of the track and at least one track roller assembly in rotational contact with a lower surface of the track.3. The actuation system of claim 1 , wherein the all of the plurality of track roller bearings are the lined track roller assembly.4. The actuation system of claim 1 , wherein the means for actuating is comprised of:an actuator arm coupled to the track; andan actuator lever coupled to the shaft and to the actuator arm;wherein when the shaft rotates in a first direction the actuator lever drives the actuator arm to move the track and the lift assisting device from the retracted to the deployed position along the arcuate path, and when the shaft rotates in a second direction the actuator lever drives the actuator arm to move the track and the lift assisting device from the deployed position to the retracted position along the arcuate path.5. The actuation system of claim 1 , wherein the actuation system further includes a ...

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Номер записи: 30
03-04-2014 дата публикации

Apparatus and method for maintaining a tension in a cable control system

Номер: US20140091174A1
Автор: Harris BUTLER
Принадлежит: Learjet Inc

A cable support apparatus for an aircraft cable control system comprises at least one cable support unit. The cable support unit includes a pulley adapted to carry a cable of the aircraft cable control system, a support link with a first joint between the pulley and the support link and a second joint between the support link and the structure of the aircraft, and a compensation link connected with the structure of the aircraft. The thermal expansion of the compensation link causes a displacement of the support link. An aircraft and a method for maintaining a tension of a cable of a cable control system of an aircraft are also provided.

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Номер записи: 31
10-04-2014 дата публикации

AIRCRAFT

Номер: US20140097292A1
Автор: Hauber Bernhard, Muenz Tanja
Принадлежит: Liebherr-Aerospace Lindenberg GmbH

The present invention relates to an aircraft having at least one landing flap arranged at the wing of the aircraft and having at least one drive unit for actuating the landing flap, wherein the aircraft furthermore has at least one control unit which controls the aileron function of the aircraft, wherein the control unit is connected to the named drive unit or units for adjusting the landing flap(s) and is configured such that it carries out the aileron function of the aircraft in at least one flight mode only or also by the operation of the named drive unit(s) and thus by the adjustment of the landing flap(s). 1. An aircraft having at least one landing flap arranged at the wing of the aircraft and having at least one drive unit for actuating the landing flap , wherein the aircraft furthermore has at least one control unit which controls the aileron function of the aircraft , wherein the control unit is connected to the named drive unit or units for adjusting the landing flap(s) and is configured such that it carries out the aileron function of the aircraft in at least one flight mode only or also by the operation of the named drive unit(s) and thus by the adjustment of the landing flap(s).2. An aircraft in accordance with claim 1 , wherein a plurality of landing flaps are arranged in the aircraft wing; and the named landing flap is the outer or outermost landing flap.3. An aircraft in accordance with claim 1 , wherein the drive unit is a hydraulic or electrical drive unit and/or the active differential gear box.4. An aircraft in accordance with claim 1 , wherein the aircraft has one claim 1 , two or more than two ailerons which are each designed with at least one drive unit for adjusting the aileron or ailerons; or the aircraft does not have any ailerons.5. An aircraft in accordance with claim 1 , wherein the aircraft has at least one aileron; and the control unit is configured such that it carries out the aileron function of the aircraft only by the operation of ...

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Номер записи: 32
07-01-2016 дата публикации

ACTIVE STRUT APPARATUS FOR USE WITH AIRCRAFT AND RELATED METHODS

Номер: US20160001874A1
Автор: Haudrich Darin P., Pitt Dale M.
Принадлежит:

Active strut apparatus for use with aircraft and related methods are disclosed. An example apparatus includes a first strut having a first end and a second end opposite the first end, the first end of the first strut is operatively coupled to a fuselage of an aircraft and the second end of the first strut is operatively coupled to a wing of the aircraft, and a first actuator is operatively coupled to the first strut to change an effective length of the first strut. 1. An apparatus comprising:a first strut having a first end and a second end opposite the first end, the first end of the first strut operatively coupled to a fuselage of an aircraft and the second end of the first strut operatively coupled to a wing of the aircraft; anda first actuator operatively coupled to the first strut to change an effective length of the first strut.2. The apparatus of claim 1 , wherein the first actuator is operatively coupled between the first end of the first strut and the fuselage of the aircraft.3. The apparatus of further comprising a second strut having a first end and a second end opposite the first end claim 2 , the first end of the second strut operatively coupled to the fuselage of the aircraft and the second end of the second strut operatively coupled to the wing of the aircraft.4. The apparatus of further comprising a second actuator operatively coupled between the first end of the second strut and the fuselage to change an effective length of the second strut.5. The apparatus of claim 4 , wherein the second end of the first strut is operatively coupled to the wing proximate a forward spar in the wing and the second end of the second strut is operatively coupled to the wing proximate an aft spar in the wing.6. The apparatus of further comprising a controller to operate the actuator to change the effective length of the first strut based on at least one of an altitude of the aircraft claim 1 , a speed of the aircraft claim 1 , a weight of the wing of the aircraft claim ...

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Номер записи: 33
05-01-2017 дата публикации

ROTATION-BLOCKING DEVICE WITH SIMPLIFIED STRUCTURE, AND ACTUATOR COMPRISING SUCH A DEVICE

Номер: US20170001714A1
Автор: Piaton Jérôme
Принадлежит:

A rotation-blocking device including a stator, a rotor, at least one blocking element mounted between the stator and the rotor to be movable between a position of interaction with the rotor and the stator and a position retracted relative to the rotor, and actuator means including a coil wound on the stator and connected to a control unit to power the coil so as to create a magnetic field, the blocking element being made of a material sensitive to the magnetic field in such a manner that powering the coil causes the blocking element to move towards one of its positions. A rotary actuator including such a device. 1. A rotation-blocking device comprising a stator , a rotor mounted to pivot about a pivot axis , at least one blocking element mounted between the stator and the rotor to be movable between a position of interaction with the rotor and the stator and a position retracted relative to the rotor , and actuator means for actuating the blocking element between these two positions , wherein the actuator means comprises a coil that is wound on the stator around an axis coaxial with the rotor and that is connected to a control unit for powering the coil in such a manner as to create a magnetic field , and in that the blocking element is made of material that is sensitive to the magnetic field in such a manner that powering the coil causes the blocking element to move towards one of its positions.2. The device according to claim 1 , wherein the actuator means comprise a resilient return element for returning the blocking element towards the other one of its positions.3. The device according to claim 1 , wherein the coil is arranged to move the blocking element from the retracted position to the interaction position claim 1 , and the resilient return element is arranged to move the blocking element from the interaction position to the retracted position.4. The device according to claim 1 , wherein the blocking element is a roller bearing against a ramp of the stator.5 ...

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Номер записи: 34
08-01-2015 дата публикации

Active Winglet

Номер: US20150008291A1
Автор: Guida Nicholas R.
Принадлежит: TAMARACK AEROSPACE GROUP, INC.

An active winglet includes a body portion substantially parallel to a wing of an aircraft, as if it were an extension of the wing. The body portion is attachable to an aircraft wing and includes a controllable airflow modification device coupled thereto. By virtue of having a controllable airflow modification device, the winglet is capable of adjusting a control surface of the controllable airflow modification device in response to in-flight conditions, to reduce wing loads, increase range, and/or increase efficiency. 120-. (canceled)21. A wing of an aircraft comprising:an angled portion coupled to the wing outboard of an aileron; anda controllable airflow modification device coupled to the wing inboard of the angled portion, the controllable airflow modification device controllable independently of the aileron, and the controllable airflow modification device configured to reduce a load on the wing.22. The wing of claim 21 , the load comprising stress in the wing caused at least in part by an aerodynamic load exerted on the angled portion.23. The wing of claim 21 , the controllable airflow modification device configured to reduce the load on the wing below a design load.24. The wing of claim 21 , the controllable airflow modification device configured to reduce a spanwise section load to a level at or below a design value for an analogous wing without an angled portion.25. The wing of claim 21 , the controllable airflow modification device being configured to adjust a control surface of the winglet at least one of electronically claim 21 , mechanically claim 21 , hydraulically claim 21 , pneumatically claim 21 , or a combination thereof.26. The wing of claim 21 , the controllable airflow modification device coupled to a control system for controlling a control surface of the controllable airflow modification device.27. The wing of claim 26 , the control system comprising a control device with control logic claim 26 , the control device being communicatively coupled ...

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Номер записи: 35
11-01-2018 дата публикации

OPTIMIZED PITCH AND ROLL CONTROL APPARATUS FOR AN AIRCRAFT

Номер: US20180009523A1
Автор: ELLIOTT JOSEPH E., SONES ANDREW, STANDLEY JOHN
Принадлежит:

An apparatus for controlling the pitch of an aircraft. The apparatus includes a horizontal control column extending from a control wheel horizontally towards a front wall of a cockpit. A pitch output link is connected to a downstream pitch control mechanism to transfer a force applied at the pitch output link to the downstream pitch control mechanism. A transfer assembly is connected to the horizontal control column and to the pitch output link. The transfer assembly translates a horizontal force applied to the horizontal control column to the pitch output link to provide the force applied to the downstream pitch control mechanism. 1. An apparatus for control of aircraft pitch comprising:a horizontal control column extending from a control wheel horizontally towards a front wall of an aircraft cockpit;a pitch output link having a first pitch output link connection point and a second pitch output link connection point connected to a downstream pitch control mechanism, wherein the pitch output link is configured to transfer a force applied at the first pitch output link connection point to the downstream pitch control mechanism; anda transfer assembly connected to the horizontal control column and to the pitch output link at the first pitch output link connection point, wherein the transfer assembly configured to translate a horizontal force applied to the horizontal control column to the pitch output link.2. The apparatus of claim 1 , further including:a roll control shaft connected to the control wheel and extending through the horizontal control column and through a front end opening of the horizontal control column; anda self-aligning bearing connected to an aircraft structure portion and positioned to support the roll control shaft, the self-aligning bearing.3. The apparatus of claim 1 , wherein the transfer assembly includes:an idler link pivotally connected to a first aircraft structure point, wherein the idler link extends down from the first aircraft ...

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Номер записи: 36
14-01-2021 дата публикации

FLAP ACTUATION SYSTEM FOR AIRCRAFT

Номер: US20210009255A1
Автор: GRUNER Bryan J., Tsai Kevin
Принадлежит:

Disclosed herein is a system for actuating a flap coupled to a wing of an aircraft in a streamwise direction. The system comprises a geared rotary actuator comprising a drive gear that is rotatable about a first rotational axis. The system also comprises a crank shaft comprising a driven gear in gear meshing engagement with the drive gear of the geared rotary actuator to rotate the crank shaft about a second rotational axis. The second rotational axis is angled relative to the first rotational axis. The system further comprises a crank arm co-rotatably coupled to the crank shaft and configured to be coupled to the flap. Rotation of the crank shaft about the second rotational axis rotates the crank arm in a direction perpendicular to the second rotational axis. 1. A system for actuating a flap coupled to a wing of an aircraft in a streamwise direction , the system comprising:a geared rotary actuator, comprising a drive gear that is rotatable about a first rotational axis;a crank shaft, comprising a driven gear in gear meshing engagement with the drive gear of the geared rotary actuator to rotate the crank shaft about a second rotational axis, wherein the second rotational axis is angled relative to the first rotational axis; anda crank arm, co-rotatably coupled to the crank shaft and configured to be coupled to the flap, wherein rotation of the crank shaft about the second rotational axis rotates the crank arm in a direction perpendicular to the second rotational axis.2. The system according to claim 1 , wherein when the second rotational axis is perpendicular relative to the streamwise direction claim 1 , the first rotational axis is parallel to a spanwise direction of the wing.3. The system according to claim 1 , wherein:the drive gear comprises a bevel gear; andthe driven gear comprises a spool gear.4. The system according to claim 1 , wherein:the crank shaft comprises a central channel that is coaxial with the second rotational axis and extends entirely through ...

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Номер записи: 37
10-01-2019 дата публикации

HOLDING DEVICE FOR AN AIRCRAFT ACTUATOR

Номер: US20190009888A1
Автор: Martens Marko
Принадлежит:

A holding device for an aircraft actuator, comprising a body that extends between a lug configured to hingedly connect the holding device to a actuator support body and an attachment flange configured to fixedly attach the holding device to a landing provided by a main body of an actuator. The holding device is configured to prevent movement of the actuator from a normal working position relative to the actuator support body in the event of structural failure of a connection body of the actuator. 1. An aircraft actuator assembly , comprising:an actuator comprising a main body and a connection body; 'the connection body being provided with a further lug configured to hingedly connect the actuator to the actuator support body;', 'a holding device comprising a body that extends between an attachment flange and a lug, wherein the attachment flange is configured to be fixedly attached to a corresponding landing provided by the main body of the actuator and the lug is configured to be hingedly connected to an actuator support body;'}wherein the holding device prevents movement of the actuator from a normal working position relative to the actuator support body in an event of structural failure of the connection body.2. The aircraft actuator assembly according to claim 1 , wherein the actuator support body forms a clevis configured to receive the lug of the holding device and the further lug of the connection body.3. The aircraft actuator assembly according to claim 1 , comprising more than one holding device.4. The aircraft actuator assembly according to claim 3 , comprising a pair of holding devices claim 3 , each holding device positioned at opposing sides of the connection body.5. The aircraft actuator assembly according to claim 1 , wherein the attachment flange and corresponding landing form a lap joint when fixedly attached to one another by a shear type fastener.6. An aircraft actuator comprising:a main body, anda connection body,wherein the main body is provided ...

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Номер записи: 38
10-01-2019 дата публикации

Overload clutch for a drive gear Mechanism for driving components of an Aircraft wing and drive gear mechanism with an overload clutch

Номер: US20190010996A1
Автор: Anton Gaile, Boris Schweitzer, Habib Stegmann
Принадлежит: Liebherr Aerospace Lindenberg GmbH

The invention relates to an overload clutch for a drive gear mechanism for driving components of an aircraft wing, in particular for driving an outer slat flap of an aircraft wing, with at least one driving and at least one driven clutch body and with at least one torque transmission body arranged therebetween. The invention further relates to a drive gear mechanism with a corresponding overload clutch.

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Номер записи: 39
10-01-2019 дата публикации

COMPOUND HARMONIC GEAR

Номер: US20190011032A1
Автор: Balsiger Derick S., Chaudhry Zaffir A., Van De Veire Nicholas R., Zhang Huan
Принадлежит:

A compound harmonic actuator is provided and includes a motor, a flex spline disposed about the motor and a wave generator radially interposable between the motor and the flex spline. The wave generator being rotatably drivable by the motor and shaped to form the flex spline into an elliptical shape with an axis such that wave generator rotations drive rotations of the axis of the ellipse of the flex spline. 1. A compound harmonic actuator , comprising:a motor;a flex spline disposed about the motor; anda wave generator radially interposable between the motor and the flex spline and rotatably drivable by the motor and shaped to form the flex spline into an elliptical shape with an axis such that wave generator rotations drive rotations of the axis of the ellipse of the flex spline,the wave generator comprising a scalloped outer surface and roller bearings arranged circumferentially about the scalloped outer surface.2. The compound harmonic actuator according to claim 1 , further comprising:a ground arm in which the motor is supportively disposable; andan output arm which is drivable by the flex spline to pivot within a predefined range of angles relative to the ground arm.3. The compound harmonic actuator according to claim 1 , wherein the wave generator is shaped to form the flex spline into an irregular elliptical shape with multiple major axes.4. The compound harmonic actuator according to claim 1 , wherein the scalloped outer surface comprises sequential ridges and lands.5. The compound harmonic actuator according to claim 4 , wherein a distal edge of each ridge is sharp.6. The compound harmonic actuator according to claim 4 , wherein a distal edge of each ridge is radially aligned with or terminal within respective central longitudinal axes of adjacent roller bearings.7. The compound harmonic actuator according to claim 1 , wherein the flex spline is formed as only a single unitary piece.8. The compound harmonic actuator according to claim 1 , wherein the flex ...

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Номер записи: 40
15-01-2015 дата публикации

COUPLING MECHANISM BETWEEN A MANUAL FLIGHT CONTROL MEMBER AND A TRIM ACTUATOR OF AN AIRCRAFT

Номер: US20150014480A1
Автор: LE-GALL Andre, LEBERNICHEUX Brice
Принадлежит:

A coupling mechanism between a control member (′) that generates manual flight commands and a trim actuator () forming part of a mechanical transmission train for transmitting flight commands in an aircraft. The coupling mechanism comprises axial engagement means between a lever arm () mechanically connected to the control member (′) and a pivot shaft () forming part of the trim actuator (). The axial engagement means comprise co-operating interlocking members () arranged on coupling members () constrained to move in rotation respectively with the lever arm () and with the pivot shaft (). Each of the coupling members () includes axial passages () enabling the other coupling member () to pass axially therethrough. An interruption of said axial engagement causes the coupling members () to move axially one through the other under the effect of axial thrust (P) exerted by elastically deformable means () used for obtaining said axial engagement. 1. A coupling mechanism between a control member generating manual flight commands and a pivot shaft of a trim actuator of a flight command transmission drive train of an aircraft , the coupling mechanism comprising:said pivot shaft mounted stationary in translation and driven in pivoting about an axis along which it generally extends by a motor;a tiltably mounted control member mechanically connected to a lever arm coupled to the pivot shaft; anda torque limiter mechanism having interlocking members that co-operate with one another by nesting under the effect of thrust exerted by elastically deformable means;whereby a force resisting transmission of movement between the pivot shaft and the control member greater than a force threshold that is predefined by a thrust force exerted by the elastically deformable means, spontaneously leads to mutual co-operation between the interlocking members being interrupted and to the control member being released from an influence exerted by the pivot shaft on the lever arm, wherein:the ...

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Номер записи: 41
03-02-2022 дата публикации

REDUNDANCY SYSTEMS FOR SMALL FLY-BY-WIRE VEHICLES

Номер: US20220033066A1
Автор: Alarie Sylvain, Esterhuizen Deon, Manzanares David James, Rey Gonzalo Javier
Принадлежит:

A universal vehicle control router for small fly-by-wire aircraft may include multiple vehicle control computers, such as flight control computers. Each flight control computer may be part of an independent channel that provides instructions to multiple actuators to control multiple vehicle components. Each channel is a distinct pathway capable of delivering a system function, such as moving an actuator. Each flight control computer may include a fully analyzable and testable voter (FAT voter). In the event of a failure to one of the flight control computers, the FAT voters may cause the failing flight control computer to be ignored or shut off power. Each flight control computer may comprise a backup battery. In the event of a power disruption from the primary power source, such as a generator and primary battery, the backup battery may power the flight control computer and all actuators. 1. A non-transitory computer-readable storage medium configured to store instructions , the instructions when executed by a processor of a control and interface system cause the control and interface system to perform steps comprising:receiving, by the control and interface system, a flight control input for a small aircraft having a single power bus;generating, by the control and interface system, an actuator instruction for an actuator based on the flight control input;performing, by the control and interface system, a self-assessment of a first flight control computer;performing, by the control and interface system, an assessment of the first flight control computer by a second flight control computer;determining, by the control and interface system, based in part on the self-assessment and the assessment of the first flight control computer by the second flight control computer, a validity of the actuator instruction generated by the first flight control computer; andtransmitting, by the control and interface system and in response to the actuator instruction being valid, the ...

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Номер записи: 42
18-01-2018 дата публикации

CROCODILE-TYPE FLIGHT CONTROL SURFACE FOR AIRCRAFT

Номер: US20180015998A1
Автор: DOUREL Pierre, ROBILLARD Jerome
Принадлежит:

A crocodile-type flight control surface comprising an upper foil flap, a lower foil flap, an actuating mechanism which guarantees the rotational displacement of each foil flap about a joint axis, either in the same direction or in different directions, and a locking mechanism alternatively adopting a locking position in which the upper foil flap and the lower foil flap are fixed with respect to each other and an unlocking position in which the upper foil flap and the lower foil flap are free with respect to the other. A crocodile-type flight control surface of this kind is therefore stiffened by the locking mechanism that joins the two foil flaps. 1. A crocodile-type flight control surface comprising:an upper foil flap,a lower foil flap,an actuating mechanism configured to guarantee the rotational displacement of each foil flap about a joint axis, either in the same direction or in different directions, anda locking mechanism alternatively adopting a locking position in which the upper foil flap and the lower foil flap are fixed, with respect to each other, and an unlocking position in which the upper foil flap and the lower foil flap are free, with respect to each other.2. The crocodile-type flight control surface according to claim 1 , wherein the actuating mechanism comprises:a first actuator,a second actuator,an upper connecting rod fixed at one end to the upper foil flap and at another end to the first actuator anda lower connecting rod fixed at one end to the lower foil flap and at another end to the second actuator,the upper connecting rod and the lower connecting rod being mounted in a rotationally movable manner about the shared axis.3. The crocodile-type flight control surface according to claim 2 , wherein one of the connecting rods comprises a first shaft claim 2 , whereas the other connecting rod comprises a second shaft which is hollow and in which is housed said first shaft claim 2 , wherein the locking mechanism comprises:a plunger cylinder, the ...

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Номер записи: 43
16-01-2020 дата публикации

SYSTEM FOR DRIVING A FLAP ARRANGEMENT BETWEEN A RETRACTED POSITION AND AN EXTENDED POSITION

Номер: US20200017193A1
Автор: GIBBERT Markus, GRIMMINGER Andreas, Schlipf Bernhard
Принадлежит:

A flap system driving a leading-edge flap between retracted and extended positions comprises a leading-edge flap having first and second flap joints, first and second scissor links, a first connecting link, and an actuator. The actuator couples with either the first scissor link or first connecting link. The first scissor link is rotatable supported on a first fixed point by a first support joint. An end of the first scissor link opposite the first support joint couples with the first flap joint. The first connecting link is rotatably supported on a second fixed point by a second support joint. An end of the first connecting link opposite the second support joint rotatably couples with an end of the second scissor link. An opposite end of the second scissor link couples with the second flap joint. The first and second scissor links are rotatably coupled to form a scissor arrangement. 1. A flap system for driving a leading-edge flap between a retracted position and an extended position , the system comprising:a leading-edge flap having a first flap joint and a second flap joint,a first scissor link,a second scissor link,a first connecting link, andan actuator,wherein the actuator is coupled with either the first scissor link or the first connecting link,wherein the first scissor link comprises a first support joint for rotatably supporting the first scissor link on a first structurally fixed point, wherein an end of the first scissor link opposed to the first support joint is coupled with the first flap joint,wherein the first connecting link comprises a second support joint for rotatably supporting the first connecting link on a second structurally fixed point and wherein an end of the first connecting link opposed to the second support joint is rotatably coupled with an end of the second scissor link,wherein an end of the second scissor link opposed to the end coupled with the first connecting link is coupled with the second flap joint,wherein additionally the ...

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Номер записи: 44
17-01-2019 дата публикации

BALL SCREW WITH SECONDARY LEAD FOR FAILURE DETECTION

Номер: US20190017580A1
Автор: Curtis Tyler Q.
Принадлежит:

A ball screw assembly for use in an actuator is provided. The ball to screw assembly may include a ball screw supported for rotation by the actuator. A ball nut may be provided around the ball screw and may be held against rotation with the ball screw by said actuator. A primary load path may be provided between the ball screw and the ball nut for operatively coupling the ball nut with the ball screw. A secondary load path may be provided between the ball screw and the ball nut, wherein the secondary load path can be disengaged during a normal operating mode of the ball screw assembly and the secondary load path can be engaged during a second operating mode of the ball screw assembly. 1. A ball screw assembly for use in an actuator , the ball screw assembly comprising:a ball screw supported for rotation by said actuator;a ball nut provided around the ball screw and held against rotation with the ball screw by said actuator;a primary load path provided between the ball screw and the ball nut for operatively coupling the ball nut with the ball screw; anda secondary load path provided between the ball screw and the ball nut, wherein the secondary load path is disengaged during a first operating mode of the ball screw assembly and the secondary load path is engaged during a second operating mode of the ball screw assembly.2. The ball screw assembly of claim 1 , wherein the ball screw has a primary groove and a secondary groove each forming a helical path extending around an outer surface of the ball screw.3. The ball screw assembly of claim 2 , wherein the primary groove and the secondary groove of the ball screw are intertwined with one another.4. The ball screw assembly of claim 1 , wherein the ball nut has an inner peripheral surface claim 1 , and a primary groove forming a helical path and a secondary thread forming a helical thread extend around the inner peripheral surface.5. The ball screw assembly of claim 4 , wherein the primary groove and the secondary thread ...

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Номер записи: 45
25-01-2018 дата публикации

ELECTRIC ACTUATOR DRIVING AND CONTROLLING DEVICE, AND AIRCRAFT

Номер: US20180022443A1
Автор: Jinno Takayuki, NAKAGAWA Shingo, NAKAJIMA Kazushige
Принадлежит:

One object is to prevent a component having low environmental resistance from being disposed in an environmentally harsh space. An electric actuator driving and controlling device is provided with a drive unit positioned in a first environment in a piece of equipment and configured to apply power to an electric actuator and a control unit positioned in a second environment in the piece of equipment and configured to transmit, to the drive unit, a power command signal including information related to power to be applied to the electric actuator. The first environment having the drive unit positioned therein is harsh compared with the second environment having the control unit positioned therein. 2. The electric actuator driving and controlling device according to claim 1 , wherein a drive element portion configured to, based on a voltage or a current inputted, apply power to the electric actuator; and', 'an interface portion configured to receive the power command signal and, based thereon, input a voltage or a current to the drive element portion., 'the drive unit comprises3. The electric actuator driving and controlling device according to claim 2 , whereinthe electric actuator is a polyphase alternating motor or a brushless DC motor,the drive element portion of the drive unit includes a plurality of switching elements corresponding to a plurality of phases of the polyphase alternating motor or the brushless DC motor, andthe interface portion inputs a voltage or a current to each of the plurality of switching elements.4. The electric actuator driving and controlling device according to claim 3 , wherein the control unit transmits the power command signal to the drive unit by serial communication.5. The electric actuator driving and controlling device according to claim 3 , wherein the control unit transmits the power command signal to the drive unit by optical communication.6. The electric actuator driving and controlling device according to claim 2 , wherein the ...

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Номер записи: 46
25-01-2018 дата публикации

ELECTRIC ACTUATOR DRIVING AND CONTROLLING DEVICE, AND AIRCRAFT

Номер: US20180022444A1
Автор: Jinno Takayuki, NAKAGAWA Shingo, NAKAJIMA Kazushige
Принадлежит:

One object is to reduce a volume occupied by a device disposed in a first space spatially limited to a large degree. An electric actuator diving and controlling device is provided with a drive unit positioned in a first space in a piece of equipment and configured to apply power to an electric actuator and a control unit positioned in a second space in the piece of equipment and configured to transmit, to the drive unit, a power command signal including information related to power to be applied to the electric actuator. The first space having the drive unit positioned therein is limited compared with the second space having the control unit positioned therein. 2. The electric actuator driving and controlling device according to claim 1 , wherein a drive element portion configured to, based on a voltage or a current inputted, apply power to the electric actuator; and', 'an interface portion configured to receive the power command signal and, based thereon, input a voltage or a current to the drive element portion., 'the drive unit comprises3. The electric actuator driving and controlling device according to claim 2 , whereinthe electric actuator is a polyphase alternating motor or a brushless DC motor,the drive element portion of the drive unit includes a plurality of switching elements corresponding to a plurality of phases of the polyphase alternating motor or the brushless DC motor, andthe interface portion inputs a voltage or a current to each of the plurality of switching elements.4. The electric actuator driving and controlling device according to claim 3 , wherein the control unit transmits the power command signal to the drive unit by serial communication.5. The electric actuator driving and controlling device according to claim 3 , wherein the control unit transmits the power command signal to the drive unit by optical communication.6. The electric actuator driving and controlling device according to claim 2 , wherein the interface portion inputs a PWM ...

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Номер записи: 47
24-01-2019 дата публикации

AERODYNAMIC CONTROL SURFACE OPERATING SYSTEM FOR AIRCRAFT USING VARIABLE TRANSMISSION

Номер: US20190023378A1
Автор: Bishop Yuniya S., Harrington Erik, Miller Christian
Принадлежит:

Control surface operating systems for controlling aerodynamic control surfaces of aircraft are provided. The systems includes a drive system operably connected to a drive shaft, the drive system including a continuously variable transmission, at least one actuator operably connected to the drive shaft and arranged to convert rotational movement of the drive shaft to move at least a portion of the aerodynamic control surface, and at least one control surface operably connected to the at least one actuator, wherein the at least one control surface is adjusted by the at least one actuator. 1. A control surface operating system for controlling aerodynamic control surfaces of an aircraft , the system comprising:a drive system operably connected to a drive shaft, the drive system including a continuously variable transmission;at least one actuator operably connected to the drive shaft and arranged to convert rotational movement of the drive shaft to move at least a portion of the aerodynamic control surface; andat least one control surface operably connected to the at least one actuator, wherein the at least one control surface is adjusted by the at least one actuator.2. The control surface operating system of claim 1 , wherein the at least one control surface is operably connected to two actuators.3. The control surface operating system of claim 1 , wherein the drive system includes a power distribution unit arranged to drive the continuously variable transmission.4. The control surface operating system of claim 1 , further comprising a controller arranged to control operation of the continuously variable transmission.5. The control surface operating system of claim 4 , wherein the controller is part of the drive system.6. The control surface operating system of claim 4 , wherein the controller controls operation of the continuously variable transmission based on a flight condition of the aircraft.7. The control surface operating system of claim 1 , wherein the at least ...

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Номер записи: 48
02-02-2017 дата публикации

ADJUSTABLE AND ROTARY RUDDER BAR FOR A ROTARY WING AIRCRAFT

Номер: US20170029092A1
Автор: FEUGAS Florian, GEMMATI Bernard, PUJOL Christophe
Принадлежит:

An adjustable and rotary rudder bar for a rotary wing aircraft. The rudder bar comprises a pedal bar, a main body, two pedals connected to the pedal bar on either side of the main body, a slider device enabling the pedal bar to slide relative to the main body, a structure fastened to a floor of the aircraft, and a shaft having a first axis (A). The shaft is secured to the main body, and the main body is movable in turning relative to the structure about the first axis (A) thus causing the main body, the pedal bar, and the pedals to turn simultaneously. 1. An adjustable and rotary rudder bar for an aircraft , the rudder bar comprising:a pedal bar;a main body;two pedals connected to the pedal bar on either side of the main body, the two pedals being positioned at the same distance D from a plane of symmetry (PS) associated with the main body;a slider device enabling the pedal bar to slide relative to the main body;a structure suitable for being fastened to a floor of the aircraft; and{'b': 1', '1, 'a shaft having a first axis (A) lying in the plane of symmetry (PS), the shaft being secured to the main body, which is movable in turning relative to the structure about the first axis (A);'}{'b': 1', '2', '1, 'wherein the structure possesses a first plane (P) substantially perpendicular to the plane of symmetry (PS) and suitable for being substantially parallel to the floor of the aircraft, and the slider device is inclined relative to the structure so that the pedal bar slides relative to the main body in a second plane (P) forming an angle β with the first plane (P) so that the vertical position of the pedals relative to the floor is modified during the sliding of the pedals, the pedals moving vertically away from the floor when the pedal bar slides towards the front of the rudder bar going away from the seat of a pilot of the aircraft.'}2. A rudder bar according to claim 1 , wherein the slider device comprises at least one lead screw connected to the main body claim 1 ...

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Номер записи: 49
04-02-2016 дата публикации

Electronic Stopper in Actuator Control

Номер: US20160031547A1
Автор: Matsui Gen
Принадлежит:

A system and method of controlling a position of a structure. A position command indicating a desired position for the structure and a position feedback signal indicating the position of the structure are received. A position control signal is generated based on a difference between the desired position and the position indicated by the position feedback signal. A stop feedback signal relative to the position of the structure is received. A stop control signal is generated based on the stop feedback signal and a stop condition for the structure. One of the position control signal and the stop control signal is selected. The selected one of the position control signal and the stop control signal is provided to an actuator for controlling the position of the structure. 1. A method of controlling a position of a structure , comprising:receiving a position command indicating a desired position for the structure;receiving a position feedback signal indicating the position of the structure;generating a position control signal based on a difference between the desired position and the position of the structure indicated by the position feedback signal;receiving a stop feedback signal relative to the position of the structure;generating a stop control signal based on the stop feedback signal and a stop condition for the structure;selecting a selected one of the position control signal and the stop control signal; andproviding the selected one of the position control signal and the stop control signal to an actuator for controlling the position of the structure.2. The method of claim 1 , wherein:the position feedback signal indicates a position of the actuator; andthe stop feedback signal indicates the position of the structure other than by identifying the position of the actuator.3. The method of claim 1 , wherein the stop feedback signal and the stop condition are selected from:the stop feedback signal indicates an angle of the structure and the stop condition indicates a ...

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Номер записи: 50
04-02-2016 дата публикации

PLANETARY GEAR BOX

Номер: US20160033010A1
Автор: Rosko Vilem
Принадлежит:

A planetary gearbox () for a renewable energy turbine is disclosed. The gearbox comprises a planetary gearbox () and at least one service cable () extending through said planetary gearbox. For example, the service cable () may be for the supply of power to and/or data from a rotor of a turbine. The service cable () enters the gearbox () through a fixed planet stage () of the gearbox and extends through a central shaft () to the stationary portion of a slip ring () assembly for providing, in use, a connection to a rotor of a turbine. The central shaft () may for example extend only partially through the gearbox. The service cable () may, for example, enter the output side of the gearbox at a location which is radially offset from the central axis of the gearbox. 1. A planetary gearbox for a renewable energy turbine comprising:a planetary gearbox; andat least one service cable extending through said planetary gearbox;wherein the service cable enters the gearbox through a fixed planet stage of the gearbox and extends through a central shaft to the stationary portion of a slip ring assembly for providing, in use, a connection to a rotor of a turbine.2. The planetary gearbox of claim 1 , wherein the slip ring assembly is proximal to the input of the gearbox.3. The planetary gearbox of or claim 1 , further comprising a cable guide defining a fixed path for the cable between the cable entrance and central shaft.4. The planetary gearbox of any preceding claim claim 1 , wherein at least a portion of the service cable extends in a radial direction through the gearbox.5. The planetary gearbox of any preceding claim claim 1 , wherein the service cable enters the output side of the gearbox at a location which is radially offset from the central axis of the gearbox.6. The planetary gearbox of any preceding claim claim 1 , wherein the service cable is routed through the planet gear carrier of said fixed planet stage of the gearbox.7. The planetary gearbox of any of to claim 1 , ...

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Номер записи: 51
17-02-2022 дата публикации

ACTUATOR WITH DECLUTCHABLE OUTPUT LEVER

Номер: US20220048611A9
Автор: ANTRAYGUE Cedric
Принадлежит:

A flight control actuator for actuating an aircraft flight control system is provided. The flight control actuator comprises a gearbox, an output shaft attached to the gearbox and an output lever provided on the output shaft. The output lever is declutchable from the output shaft. 1. A flight control actuator for actuating an aircraft flight control system , the flight control actuator comprising:a gearbox;an output shaft attached to the gearbox; andan output lever provided on the output shaft, wherein the output lever is declutchable from the output shaft.2. The flight control actuator of claim 1 , wherein:the output lever comprises an inner diameter through which the output shaft passes, wherein there is provided at least one indentation in said inner diameter;wherein the output shaft comprises a hollow cylindrical member with at least one hole provided at the axial position of the at least one indentation of the output lever;said flight control actuator further comprising:a plunger positioned in a first axial position within the output shaft;wherein one or more balls are positioned in each of said at least one hole of the output shaft and between the plunger and the at least one indentation of the output lever, such that the output lever and output shaft are rotatably coupled and the flight control actuator is clutched.3. The flight control actuator of claim 2 , wherein:the plunger is tapered in diameter such that, when moved axially with respect to the output shaft into a second axial position, the one or more balls fall out of the at least one indentation in the output lever such that the output lever and output shaft are decoupled and the flight control actuator is declutched.4. The flight control actuator of claim 3 , further comprising:a biasing member which biases the plunger into said first axial position.5. The flight control actuator of claim 4 , wherein:the plunger can be moved axially into said second axial position to decouple the output shaft from ...

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Номер записи: 52
30-01-2020 дата публикации

REDUNDANT FLY-BY-WIRE SYSTEMS WITH FAULT RESILIENCY

Номер: US20200031454A1
Автор: Wilkens Dean
Принадлежит: HONEYWELL INTERNATIONAL INC.

Aircraft fly-by-wire systems and related vehicle electrical systems are provided. In one embodiment, an electrical system suitable for use with a control surface of a vehicle, such as an aircraft, is provided. The electrical system includes a plurality of communications buses and a plurality of control modules, wherein each of the plurality of control modules is connected to a respective subset of the plurality of communications buses that is unique among the plurality of control modules, and a plurality of actuation control modules associated with the control surface, wherein each of the plurality of actuation control modules is connected to a respective subset of the plurality of communications buses that is unique among the plurality of actuation control modules. Thus, each of the control modules is isolated from at least one of the communications buses. 1. A vehicle electrical system comprising:a bus arrangement comprising a plurality of buses;a first control module coupled to a first subset of the plurality of buses, the first subset including a first bus and a second bus;a second control module coupled to a second subset of the plurality of buses, the second subset including the first bus and a third bus;a third control module coupled to a third subset of the plurality of buses, the third subset including the second bus and the third bus;a first actuation arrangement coupled to a fourth subset of the plurality of buses, the fourth subset including the first bus; and the first, second and third subsets are different; and', 'the fourth and fifth subsets are different., 'a second actuation arrangement coupled to a fifth subset of the plurality of buses, the fifth subset including the second bus, wherein2. The vehicle electrical system of claim 1 , wherein the plurality of buses comprises a plurality of controller area network (CAN) buses.3. The vehicle electrical system of claim 2 , further comprising:a user interface device coupled to each of the first, second, ...

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Номер записи: 53
04-02-2021 дата публикации

MECHANISM WITH THREE DEGREES-OF-FREEDOM (DOF) OUTPUT TO PROVIDE INDEPENDENT CONTROL OVER ROLL, PITCH, AND YAW OF OUTPUT STRUCTURE

Номер: US20210031383A1
Автор: Shannon Connor Richard
Принадлежит:

Mechanisms or apparatus convert a number of inputs via a number of input members into a number of output movements of an output structure, providing control in three degrees-of-freedom (DOF), for example control over roll, pitch and yaw of the output structure. Inputs may be rotations about a common axis of rotation, for example via a first ring, a second ring, and one or more plates, concentrically array. Rotation of the first ring may control a first DOF, rotation of the first ring may control a second DOF, and rotation of the plate may control all three DOF. Three concentrically arrayed tubular shafts may be employed, providing a through-passage or cable fluid conduit run to accommodate wires, optical fibers, fluid carrying conduits. Such may be particularly advantageous when employed as part of a robot, or other device with a tool or sensor or transducer located at or proximate a distal end thereof. 1. An apparatus that provides three degrees-of-freedom (DOF) movement , the apparatus comprising:a first input member;a second input member;a third input member;a first tubular shaft, the first tubular shaft rotatable about a first axis;a second tubular, the second tubular shaft rotatable about the first axis in response to rotation of the second input member about the first axis;a third tubular shaft, the third tubular shaft rotatable about the first axis in response to rotation of the third input member about the first axis;a first output member having a first longitudinal axis about which the first output member is rotatable;a second output member having a pivot axis about which the second output member is pivotable, the pivot axis perpendicular to the first longitudinal axis;a third output member, having a second longitudinal axis about which the third output member is rotatable, the first output member rotatably coupled to the second output member and the second output member coupled to rotate with the third output member;a plurality of gears, at least two of ...

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Номер записи: 54
04-02-2021 дата публикации

WING FOR AN AIRCRAFT

Номер: US20210031902A1
Автор: LORENZ Christian
Принадлежит:

A wing for an aircraft having a fixed wing, a foldable wing tip portion rotatably attached to the fixed wing and an actuation unit for rotating the foldable wing tip portion relative to the fixed wing about a hinge axis is disclosed. The actuation unit includes a traction means for transmitting traction between the fixed wing and the foldable wing tip portion, a main wheel attached to the foldable wing tip portion and in contact with the traction means and a drive means for generating traction to be transmitted by the traction means. The actuation unit generates traction with the drive means in the traction means so that the foldable wing tip portion can be rotated relative to the fixed wing. 1. A wing for an aircraft , comprising:a fixed wing,a foldable wing tip portion rotatably attached to the fixed wing,an actuation unit for rotating the foldable wing tip portion relative to the fixed wing about a hinge axis,wherein the actuation unit comprises a traction means for transmitting traction between the fixed wing and the foldable wing tip portion, a main wheel in contact with the traction means and a drive means for generating traction to be transmitted by the traction means, andwherein the actuation unit is configured such that by generating traction with the drive means in the traction means the foldable wing tip portion can be rotated relative to the fixed wing between an extended position and a folded position.2. The wing according to claim 1 , wherein an axis of rotation of the main wheel corresponds to the hinge axis claim 1 , wherein the main wheel is fixed to the foldable wing tip portion such that a rotation of the main wheel relative to the fixed wing corresponds to a rotation of the foldable wing tip portion relative to the fixed wing and wherein the drive means is attached to the fixed wing.3. The wing according to claim 1 , wherein the actuation unit comprises an input section where traction is generated by the drive means and an output section where ...

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Номер записи: 55
09-02-2017 дата публикации

Lower Attachment for Trimmable Horizontal Stabiliser Actuator

Номер: US20170036754A1
Автор: Medina Raphael, Ricard Quentin
Принадлежит:

A connection for a lower attachment for a trimmable horizontal stabiliser actuator (THSA) for connecting the lower attachment to a flight control surface. The attachment includes: a surface bracket for coupling to the flight control surface, the surface bracket mounted about a gimbal of the lower attachment by a bushing disposed between the gimbal and the surface bracket; a failsafe plate fittingly engaged on a first end of the bushing; and a tightening ring mounted on a second end of the bushing opposite the first end and secured to the surface bracket by at least one contact screw, such that the contact screw is operable to urge the bushing against the failsafe plate. 1. A connection for a lower attachment for a trimmable horizontal stabiliser actuator (THSA) for connecting the lower attachment to a flight control surface; comprising:{'b': 140', '130', '110, 'a surface bracket () for coupling to the flight control surface, the surface bracket mounted about a gimbal () of the lower attachment by a bushing () disposed between the gimbal and the surface bracket;'}{'b': '120', 'a failsafe plate () fittingly engaged on a first end of the bushing; and'}{'b': 170', '160, 'a tightening ring () mounted on a second end of the bushing opposite the first end and secured to the surface bracket by at least one contact screw (), such that the contact screw is operable to urge the bushing against the failsafe plate.'}2. The connection of claim 1 , wherein the failsafe plate comprises an aperture;{'b': 114', '112', '120, 'wherein the first end of the bushing comprises an insert portion () disposed within the aperture; and wherein the bushing comprises a bearing portion () disposed adjacent the failsafe plate () so as to bear against the failsafe plate.'}3. The connection of claim 1 , further comprising:{'b': 150', '130', '140, 'a spherical bearing () disposed between the gimbal () and the surface bracket () to allow articulation therebetween.'}4174170140160. The connection of ...

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Номер записи: 56
09-02-2017 дата публикации

TORQUE TUBE ASSEMBLIES FOR USE WITH AIRCRAFT HIGH LIFT DEVICES

Номер: US20170037910A1
Автор: Meyer Mark Robert, Schwartz Douglas John
Принадлежит:

Example torque tube assemblies for use with aircraft high lift devices are described herein. An example apparatus includes a spline coupling having a first yoke, a sliding splined shaft having a second yoke and a torque tube having a first end and a second end opposite the first end. A first fitting with a third yoke is coupled to the first end of the torque tube, and a second fitting with a fourth yoke is coupled to the second end of the torque tube. The third yoke is coupled to the first yoke to form a first U-joint, and the fourth yoke is coupled to the second yoke to form a second U-joint. The spline coupling is to be coupled to a first high lift device drive shaft and the sliding splined shaft is to be coupled to a second high lift device drive shaft. 1. An apparatus comprising:a spline coupling having a first yoke, the spline coupling to be coupled to a first high lift device drive shaft of an aircraft;a sliding splined shaft having a second yoke, the sliding splined shaft to be coupled to a second high lift device drive shaft of the aircraft; anda torque tube having a first end and a second end opposite the first end, a first fitting with a third yoke coupled to the first end of the torque tube, and a second fitting with a fourth yoke coupled to the second end of the torque tube, the third yoke of the first fitting coupled to the first yoke of the spline coupling to form a first U-joint, and the fourth yoke of the second fitting coupled to the second yoke of the sliding splined shaft to form a second U-joint.2. The apparatus of claim 1 , wherein the first fitting is coupled to the first end of the torque tube via an electromagnetic forming (EMF) process and the second fitting is coupled to the second end of the torque tube via an EMF process.3. The apparatus of claim 1 , wherein the spline coupling is to receive a spline gear coupled to the first high lift device drive shaft.4. The apparatus of claim 3 , wherein the spline coupling includes an aperture to ...

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Номер записи: 57
08-02-2018 дата публикации

INVERTED COMPOUND HARMONIC DRIVE

Номер: US20180038467A1
Автор: Balsiger Derick S., Van De Veire Nicholas R.
Принадлежит:

A compound harmonic drive including: a first internal ring gear; a second internal ring gear, the second internal ring gear being coaxial to the first internal ring gear; a first external gear located within the first internal ring gear and coaxial to the first internal ring gear, the first external gear meshes with the first internal ring gear; and a second external gear located within the second internal ring gear and coaxial to the second internal ring gear, the second external gear meshes with the second internal ring gear. The first internal ring gear and the second internal ring gear are composed of a flexible material. 1. A compound harmonic drive , comprising:a first internal ring gear;a second internal ring gear, the second internal ring gear being coaxial to the first internal ring gear;a first external gear located within the first internal ring gear and coaxial to the first internal ring gear, the first external gear meshes with the first internal ring gear; anda second external gear located within the second internal ring gear and coaxial to the second internal ring gear, the second external gear meshes with the second internal ring gear,wherein the first internal ring gear and the second internal ring gear are composed of a flexible material.2. The compound harmonic drive of claim 1 , wherein the first external gear includes radially-outward-extending teeth that mesh with radially-inward-extending teeth of the first internal ring gear.3. The compound harmonic drive of claim 2 , wherein the second external gear includes radially-outward-extending teeth that mesh with radially-inward-extending teeth of the second internal ring gear.4. The compound harmonic drive of claim 3 , wherein the first internal ring gear includes a different number of radially-inward-extending teeth than the second internal ring gear.5. The compound harmonic drive of claim 3 , wherein the first external gear includes a different number of radially-outward-extending teeth than the ...

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Номер записи: 58
08-02-2018 дата публикации

JOINED FLEX SPLINE FOR COMPOUND HARMONIC DRIVE

Номер: US20180038468A1
Автор: Balsiger Derick S., Van De Veire Nicholas R.
Принадлежит:

A compound harmonic drive including: a flexible ring having an inner surface and an outer surface; a first flexible gear disposed around the outer surface of the flexible ring and coaxial to the flexible ring; a second flexible gear disposed around the outer surface of the flexible ring and coaxial to the flexible ring; a first ring gear that meshes with the first flexible gear and is coaxial to the first flexible gear; and a second ring gear that meshes with the second flexible gear and is coaxial to the second flexible gear. The first flexible gear and second flexible gear are fixedly attached to the outer surface of the flexible ring. 1. A compound harmonic drive , comprising:a flexible ring having an inner surface and an outer surface;a first flexible gear disposed around the outer surface of the flexible ring and coaxial to the flexible ring;a second flexible gear disposed around the outer surface of the flexible ring and coaxial to the flexible ring;a first ring gear that meshes with the first flexible gear and is coaxial to the first flexible gear; anda second ring gear that meshes with the second flexible gear and is coaxial to the second flexible gear, wherein the first flexible gear and second flexible gear are fixedly attached to the outer surface of the flexible ring.2. The compound harmonic drive of claim 1 , wherein the first ring gear includes radially-inward-extending teeth that mesh with radially-outward-extending teeth of the first flexible gear.3. The compound harmonic drive of claim 2 , wherein the second ring gear includes radially-inward-extending teeth that mesh with radially-outward-extending teeth of the second flexible gear.4. The compound harmonic drive of claim 3 , wherein the first flexible gear includes a different number of radially-outward-extending teeth than the second flexible gear.5. The compound harmonic drive of claim 3 , wherein the first ring gear includes a different number of radially-inward-extending teeth than the second ...

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Номер записи: 59
06-02-2020 дата публикации

Lvdt-based actuator output load limited

Номер: US20200039637A1
Автор: Eric A. Polcuch
Принадлежит: Parker Hannifin Corp

An actuator assembly includes a primary load path for tightly coupling an actuated surface to a reference structure, and a secondary load path having a backlash portion for coupling the actuated surface to the reference structure with backlash, wherein the secondary load path is unloaded during an operative state of the primary load path and loaded during a failure state of the primary load path. A first sensor is configured to sense relative displacement between a portion of the primary load path and a portion of the secondary load path. A controller is operatively coupled to the first sensor, the controller configured to determine a load on the primary load path based on relative displacement sensed by the first sensor.

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Номер записи: 60
16-02-2017 дата публикации

HYDROSTATIC AUTOMATIC FLIGHT SERVO SYSTEMS

Номер: US20170043861A1
Автор: Bandera Pablo, Bartling Jon, Wilkens Dean
Принадлежит: HONEYWELL INTERNATIONAL INC.

A hydrostatic automatic flight servo system is provided. The automatic flight servo system includes a manifold that defines a first fluid chamber, and a hydraulic fluid is received in the first fluid chamber. The first fluid chamber includes a first bellows and a second bellows. The automatic flight servo system includes a stick received at least partially within the manifold and pivotally coupled to the manifold. The stick includes a control arm fixedly coupled to the first bellows, and the stick is to receive an input. The automatic flight servo system includes a flight output system pivotally coupled to the manifold. The flight output system includes a second control arm received at least partially within the manifold and coupled to the second bellows such that the pivotal movement of the stick pivots the flight output system relative to the manifold. 1. A hydrostatic automatic flight servo system comprising:a manifold that defines a first fluid chamber, with a hydraulic fluid received in the first fluid chamber, the first fluid chamber including a first bellows and a second bellows;a stick received at least partially within the manifold and pivotally coupled to the manifold, the stick including a control arm fixedly coupled to the first bellows, the stick adapted to receive an input; anda flight output system pivotally coupled to the manifold, the flight output system including a second control arm received at least partially within the manifold and coupled to the second bellows such that the pivotal movement of the stick pivots the flight output system relative to the manifold.2. The hydrostatic automatic flight servo system of claim 1 , wherein the flight output system includes an output linkage that is pivotally coupled to the manifold and the second control arm is coupled to the output linkage claim 1 , the output linkage having a first linkage end adapted to be coupled to a flight surface.3. The hydrostatic automatic flight servo system of claim 2 , wherein ...

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Номер записи: 61
16-02-2017 дата публикации

DRIVE ASSEMBLY WITH SELECTIVE DISCONNECT

Номер: US20170045129A1
Автор: Balsiger Derick
Принадлежит:

A drive assembly with selective disconnect includes a motor with a motor drive shaft; a harmonic drive coupled to one end of the motor drive shaft; an output shaft coupled to the harmonic drive; and a retracting mechanism that selectively retracts the motor drive shaft axially to decouple the motor drive shaft from the harmonic drive. 1. A method comprising:driving an output shaft with a motor drive shaft through a harmonic drive located between the output shaft and the motor drive shaft; anddisconnecting the output shaft from the motor drive shaft by selectively decoupling the motor drive shaft from the harmonic drive.2. The method of further comprising:preventing the motor drive shaft from axial movement.3. The method of claim 1 , wherein the motor drive shaft is attached to a wave generator within the harmonic drive and the wave generator is retracted from the harmonic drive when the motor drive shaft is disconnected from the harmonic drive.4. The method of claim 3 , wherein the motor drive shaft and the wave generator are selectively retracted by a spring.5. The method of claim 4 , wherein a solenoid selectively retracts a locking mechanism to unlock the motor drive shaft and allow the spring to retract the motor drive shaft and the wave generator.6. The method of claim 4 , wherein the motor drive shaft and wave generator selectively retract in response to feedback from the electric motor claim 4 , motor drive shaft claim 4 , or output shaft. This application is a divisional application under 35 U.S.C. §121 of U.S. application Ser. No. 14/086,001 filed Nov. 21, 2013 for “DRIVE ASSEMBLY WITH SELECTIVE DISCONNECT” by Derick Balsiger.This invention was made, at least in part, with U.S. Government support under the Boeing contract number 475120, awarded by the United States Air Force. The U.S. Government may have certain rights in this invention.The present invention relates to drives used with electro-mechanical actuators and, in particular, to a system and method ...

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Номер записи: 62
15-02-2018 дата публикации

Secondary load path detection

Номер: US20180045291A1
Автор: Fredric PERSOUYRE, Ghislain TEYSSIER
Принадлежит: Ratier Figeac SAS

A nut arrangement for a screw actuator is disclosed for allowing detection of wear in a primary nut of the screw actuator. The nut arrangement comprises a primary nut for providing a primary load path and a secondary nut for providing a secondary load path. An interface ring may link the secondary nut to the primary nut. A sensor is provided to detect relative axial movement between the primary and secondary nuts. During normal operation, the interface ring is seated by a flexible coupling that allows relative axial displacement of the secondary nut to the primary nut to accommodate wear in the primary nut. The sensor can be used to monitor backlash between the primary and secondary nuts to determine wear of the primary nut.

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Номер записи: 63
25-02-2016 дата публикации

Rotorcraft autopilot control

Номер: US20160052628A1
Автор: Awais Raza, Brady G. Atkins, Carl D. Griffith, Paul M. Wilson
Принадлежит: BELL HELICOPTER TEXTRON INC

A rotorcraft autopilot system includes a series actuator connecting a cockpit control component to a swashplate of a rotorcraft, the series actuator to modify a control input from the cockpit control component to the swashplate through a downstream control component. The rotorcraft autopilot system also includes a differential friction system connected to the cockpit control component, the differential friction system to control the series actuator to automatically adjust a position of the cockpit control component during rotorcraft flight based, in part, on a flight mode of the rotorcraft.

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Номер записи: 64
08-05-2014 дата публикации

WIND TURBINE GEAR MECHANISM

Номер: US20140128213A1
Автор: Kienzle Alfred, KIRSCHNER Tino, Petersen Knud Erik, WALLISER Jochen
Принадлежит:

A wind turbine gear mechanism with a planetary gear stage arranged on the drive input side and with first and second spur gear stages arranged, in a gear mechanism housing, downstream of the planetary gear stage. The first spur gear stage rotationally and functionally connects a sun gear shaft of the planetary gear stage to an intermediate shaft, the second spur gear stage rotationally and functionally connects the intermediate shaft to a drive output shaft, and only one intermediate shaft is provided. In this case, the gear mechanism housing has a separation joint along which the gear mechanism housing can be divided into a main housing portion and a housing cover in such manner that a separation joint plane passes through the central axis of the intermediate shaft and the separation joint plane extends a distance away from the central axis of the sun gear shaft. 14-. (canceled)5123452. A wind turbine gear mechanism () with a planetary gear stage () on a drive input side and with a first spur gear stage () and a second spur gear stage () , which are arranged in a gear mechanism housing () downstream of the planetary gear stage () ,{'b': 3', '6', '2', '7, 'the first spur gear stage () connecting a sun gear shaft () of the planetary gear stage (), in a rotationally functional manner, to an intermediate shaft (),'}{'b': 4', '7', '8', '7, 'the second spur gear stage () connecting the intermediate shaft (), in a rotationally functional manner, to a drive output shaft (), and only one intermediate shaft () being provided,'}{'b': 5', '9', '5', '10', '11', '9', '12', '7, 'the gear mechanism housing () having a separation joint () along which the gear mechanism housing () can be separated into a main housing portion () and a housing cover () so that a separation joint plane () passes through a central axis () of the intermediate shaft (), and'}{'b': 9', '13', '6, 'the separation joint plane () extending a distance away from the central axis () of the sun gear shaft ().'} ...

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Номер записи: 65
14-02-2019 дата публикации

DUAL RACK AND PINION ROTATIONAL INERTER SYSTEM AND METHOD FOR DAMPING MOVEMENT OF A FLIGHT CONTROL SURFACE OF AN AIRCRAFT

Номер: US20190048961A1
Автор: Fox Michael T., Howell Eric A., Roach Jeffrey M.
Принадлежит:

There is provided a dual rack and pinion rotational inerter system for damping movement of a flight control surface of an aircraft having a support structure. The system has a flexible holding structure disposed between the flight control surface and the support structure. The system has a dual rack and pinion assembly held by the flexible holding structure. The system has a first terminal and a second terminal, coupled to the dual rack and pinion assembly. The first terminal is coupled to the flight control surface. The system has a pair of inertia wheels coupled to the flexible holding structure. The system has an axle element inserted through the inertia wheels, the flexible holding structure, and the dual rack and pinion assembly, such that when the flight control surface rotates, the dual rack and pinion rotational inerter system translates and rotates, and movement of the flight control surface is dampened. 1. A dual rack and pinion rotational inerter system for damping movement of a flight control surface of an aircraft having a support structure , the dual rack and pinion rotational inerter system comprising:a flexible holding structure disposed between the flight control surface and the support structure;a dual rack and pinion assembly held by the flexible holding structure;a first terminal and a second terminal, each coupled to the dual rack and pinion assembly, the first terminal further coupled to the flight control surface;a pair of inertia wheels coupled to the flexible holding structure; andan axle element inserted through the pair of inertia wheels, the flexible holding structure, and the dual rack and pinion assembly, such that when the flight control surface rotates, the dual rack and pinion rotational inerter system translates and rotates, such that movement of the flight control surface is dampened.2. The system of claim 1 , further comprising a plurality of rod bearings inserted into interior corners of the flexible holding structure.3. The ...

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Номер записи: 66
13-02-2020 дата публикации

PLANETARY GEAR MOTOR WITH TWO COAXIAL OUTPUT SHAFTS

Номер: US20200049232A1
Автор: SMANIA Marco
Принадлежит: Unicum Transmission de Puissance

The invention relates to a geared motor comprising: 1. Geared motor wherein the geared motor comprises:an electric motor crossed, through and through, by a rotor shaft, so as to form two motor output shafts, each output shaft has a pinion;at least two planetary gearboxes with at least one stage, each geared with the pinion of a motor output shaft;two geared motor output shafts, each being coupled directly or indirectly to a planetary gearbox, the two geared motor output shafts being coaxial with each other.2. Geared motor according to claim 1 , wherein the planetary gearboxes have at least two stages.3. Geared motor according to claim 1 , wherein the geared motor comprises two planetary gearboxes coupled to each other on each side of the electric motor.4. Geared motor according to claim 1 , wherein the geared motor output shafts are each directly coupled to a planetary gearbox claim 1 , so that the geared motor output shafts are coaxial with the rotor shaft of the electric motor.5. Geared motor according to claim 1 , wherein the geared motor output shafts are coupled indirectly to the planetary gearboxes and via a single gear train claim 1 , so that the geared motor output shafts are eccentric with respect to the rotor shaft of the electric motor.6. Geared motor according to claim 1 , wherein each planetary gearbox comprises a crown claim 1 , the crowns being arranged inside a tubular housing.7. Geared motor according to claim 6 , wherein the tubular housings inside which the crowns are arranged extend around the electric motor to directly form a body of the geared motor claim 6 , the tubular housings being closed by two flanges from which the geared output shafts project.8. Geared motor according to claim 7 , wherein the flanges are square claim 7 , with sides of a length equal to a diameter of the tubular housings.9. Geared motor according to claim 1 , wherein the geared motor comprises a tubular outer housing enclosing the components of the geared motor claim 1 , ...

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Номер записи: 67
15-05-2014 дата публикации

SEAL AND ELECTRICAL CONDUCTOR FOR LINED TRACK ROLLERS USED ON ACTUATION SYSTEM FOR AIRCRAFT LIFT ASSISTING DEVICES

Номер: US20140131512A1
Автор: Gyuricsko Frederick S.
Принадлежит:

An electrical conductor for a bearing defines an annular base and is manufactured from an electrically conductive material. The electrical conductor includes a first electrical connector positioned proximate a radially outer peripheral area of the annular base. The electrical conductor includes a second electrical connector positioned proximate a radially inner peripheral area of the annular base. The second electrical connector defines one or more contact edges extending away from the annular base. The contact edge is configured for sliding electrical contact with the bearing. 1. An electrical conductor for a bearing , the electrical conductor comprising:an annular base manufactured from an electrically conductive material;a first electrical connector positioned proximate a first portion of the annular base; anda second electrical connector positioned proximate a second portion of the annular base, the second electrical connector defines at least one contact edge extending away from the annular base, the contact edge being configured for sliding electrical contact with the bearing.2. The electrical conductor of claim 1 , wherein the at least one contact edge extends continuously and entirely around the radially inner peripheral area of the annular base.3. The electrical conductor of claim 1 , wherein the at least one contact edge comprises a plurality of tabs spaced apart from one another.4. The electrical conductor of claim 1 , further comprising a flexible seal positioned on the annular base claim 1 , the flexible seal having an end configured for sliding contact with the bearing.5. An actuation system for deploying and retracting a lift assisting device of a leading edged of a wing of an aircraft claim 1 , the actuation system comprising:a track pivotally coupled to the lift assisting device, the track having first and second outer surfaces and side surfaces;a shaft rotationally coupled within the wing of the aircraft and operable, in response to flight control ...

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Номер записи: 68
05-03-2015 дата публикации

Electromechanically Controlled Decoupling Device for Actuators

Номер: US20150060602A1
Автор: Christmann Markus, NAUBERT Alexander
Принадлежит:

An aircraft includes an electromechanical actuator and a decoupling device. A form-fit between a drive element and a linkage point outside of the actuator element can be reversibly decoupled and recoupled. The decoupling device allows the actuator element to run freely in the event of a malfunction, which effectively prevents a jam. 1. An actuator element , comprising:a drive unit, having a first drive element and a second drive element, wherein the first drive element and the second drive element are arranged to functionally interact in such a manner to effect a change in length of the actuator element;a decoupling device;two linkage points, whose span can be fixed by changing the length of the actuator element, wherein one of the drive elements is connected in an effect-direct manner to one of the linkage points;an output piston connected in an effect-direct manner to the other one of the linkage points,wherein the span of the two linkage points in a coupled state of the decoupling device is adjustable by the actuator element, andwherein the span of the two linkage points in a decoupled stated of the decoupling device is adjustable by application of an external force on one of the two linkage points, a decoupling mechanism with a drive element, a form-fitting element and a retaining element,', 'wherein the retaining element is arranged to take a first position in which the form-fitting element is in a closed state,', 'wherein the retaining element is arranged to take a second position in which the form-fitting element is in an open state,', 'wherein the drive element is arranged to move the retaining element between the first position and the second position, wherein a form-fit of the form-fitting element is reversibly releasable and fixable;', 'wherein the decoupling mechanism is arranged to functionally decouple the output piston from a drive element of the actuator element so that the length change of the actuator element is achieved independent of the drive ...

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Номер записи: 69
10-03-2022 дата публикации

Gas strut, method for producing the gas strut, drive for a flap with the gas strut

Номер: US20220074462A1
Автор: Alexander Reiser, Felix Beib, Ulrich Probst
Принадлежит: Stabilus GmbH

Provided is a gas strut, including: an outer working space arranged radially to the stroke axis between the working cylinder and the equalizing cylinder, the outer working space being connected to the inner working space in a gas-conducting manner; an equalizing piston enclosing the working cylinder radially to the stroke axis, the equalizing piston) being mounted displaceably along the stroke axis, delimiting the outer working space on one side transversely to the stroke axis and being subjected to a pressure of the working medium and a pressure of the equalizing medium so as to increase the volume of the outer working space; and a restoring medium arranged in a restoring space radially to the stroke axis between the working cylinder and the equalizing cylinder, the equalizing piston being subjected to a pressure of the restoring medium so as to decrease the volume of the outer working space.

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Номер записи: 70
21-02-2019 дата публикации

CONTROL SURFACE ATTACHMENT

Номер: US20190055002A1
Автор: Bekircan Suat
Принадлежит:

An airfoil structure comprising a main body and at least one control surface attached to the main body by a flexible attachment, the flexible attachment comprising a flexible first surface and a flexible second surface opposed to the flexible first surface, each of the flexible first surface and the flexible second surface having a waveform structure. 1. An airfoil structure comprising:a main body; andat least one control surface attached to the main body by a flexible attachment, the flexible attachment comprising a flexible first surface and a flexible second surface opposed to the flexible first surface, each of the flexible first surface and the flexible second surface having a waveform structure.2. The airfoil structure of claim 1 , wherein the flexible first surface connects a first surface of the main body to a first surface of the at least one control surface claim 1 , and the flexible second surface connects a second surface of the main body to a second surface of the at least one control surface.3. The airfoil structure of claim 1 , wherein the flexible attachment further comprises flexible third and fourth surfaces joining the flexible first surface to the flexible second surface claim 1 , the flexible third and fourth surfaces comprising a waveform structure.4. The airfoil structure of claim 3 , wherein the flexible first claim 3 , second claim 3 , third and fourth surfaces completely enclose the flexible attachment claim 3 , such that there are no gaps or discontinuities between the main body and the at least one control surface.5. The airfoil structure of claim 1 , further comprising an actuator) for moving the at least one control surface.6. The airfoil structure of claim 5 , wherein the actuator comprises an electro-mechanical actuator.7. The airfoil structure of claim 5 , wherein the actuator is configured to translate the at least one control surface in a chord-wise direction relative to the main body.8. The airfoil structure of claim 5 , wherein ...

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Номер записи: 71
21-02-2019 дата публикации

ACTUATOR ASSEMBLIES FOR CONTROL SURFACES OF AN AIRCRAFT, AIRCRAFT INCLUDING THE ACTUATOR ASSSEMBLIES, AND METHODS OF UTILIZING THE SAME

Номер: US20190055005A1
Автор: Young Stuart David
Принадлежит:

Actuator assemblies and methods of utilizing the same are disclosed herein. The actuator assemblies include a base structure and an actuated arm that are pivotally coupled to the base structure. The actuator assemblies also include a drive assembly that is operatively attached to the base structure and includes an output shaft. The actuator assemblies further include a linear actuator that includes an actuator shaft and an actuated body. The actuator shaft is coupled to and configured to rotate with the output shaft about an actuator shaft axis of rotation. The actuator assemblies also include a linkage that is pivotally coupled to the actuated arm. In addition, the linkage is operatively attached to the actuated body via a joint. The joint defines a plurality of joint axes of rotation that are spaced apart from the actuator shaft axis of rotation of the actuator shaft. 1. An actuator assembly configured to move a control surface of an aircraft through a control surface range-of-motion , the actuator assembly comprising:a base structure;an actuated arm including a base mount pivotally coupled to the base structure, a surface mount configured to be pivotally coupled to the control surface, and a linkage mount;a drive assembly operatively attached to the base structure and including an output shaft;a linear actuator including:(i) an actuator shaft coupled to, and configured to rotate with, the output shaft of the drive assembly about an actuator shaft axis of rotation; andan actuated body coupled to the actuator shaft and including a joint mount, wherein the actuator shaft and the actuated body are configured such that the actuated body operatively translates linearly along a length of the actuator shaft responsive to rotation of the actuator shaft about the actuator shaft axis of rotation;a joint operatively attached to the joint mount, wherein the joint defines a plurality of joint pivot axes that all are spaced apart from the actuator shaft axis of rotation; anda ...

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Номер записи: 72
15-05-2014 дата публикации

Planetary Gear Set with Several Gear Stages

Номер: US20140135165A1
Автор: Kruselburger Markus
Принадлежит:

Planetary gear set with several gear stages, comprising an annular ring with which the planet gears are in engagement with at least one gear stage; according to the invention it is provided that at least one circular insert part with external and internal gearing is arranged in an annular ring, whereby the external gearing is in engagement with the annular ring and the internal gearing meshes with planet gears of another gear stage. 1. A planetary gear set with several gear stages , comprising an internally geared annular ring with which the planet gears are in engagement with at least one gear stage , further comprising wherein at least one circular insert part with external and internal gearing is arranged in an annular ring , whereby the external gearing is in engagement with the annular ring and the internal gearing meshes with planet gears of another gear stage.2. The planetary gear set of claim 1 , further comprising wherein the annular ring has straight gearing claim 1 , which is in engagement with the external gearing of the insert part that is designed as straight gearing.3. The planetary gear set of claim 1 , further comprising wherein the internal gearing of the insert part is designed as helical gearing.4. The planetary gear set of claim 1 , further comprising wherein (i) the annular ring is designed with a first and second internal gearing section in axial direction claim 1 , whereby the planet gears of the at least one gear stage are in engagement with the first gearing section and the external gearing of the insert part is in engagement with the second internal gearing section and (ii) for forming a transition from the first internal gearing section to the second internal gearing section claim 1 , a shoulder is provided claim 1 , which the insert part contacts with its face side.5. The planetary gear set of claim 4 , further comprising wherein for forming the shoulder claim 4 , the crown line diameter of the gearing in the area of the first gearing ...

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Номер записи: 73
01-03-2018 дата публикации

ACTUATORS FOR HIGH LIFT DEVICES ON AIRCRAFT

Номер: US20180057149A1
Автор: Atmur Robert J., Heiberg Christopher J., McCormick Patrick J.
Принадлежит:

Actuators for high lift devices on aircraft are disclosed herein. An example apparatus includes an actuator for a high lift device of an aircraft including a motor and a transmission, where the transmission includes a first gear stage and a second gear stage, the first gear stage including a first worm gear and the second gear stage including a second worm gear, the first worm gear, the second worm gear and the motor operative to prevent backdrive of the actuator. 1. An apparatus comprising an actuator for a high lift device of an aircraft including a motor and a transmission , the transmission including a first gear stage and a second gear stage , the first gear stage including a first worm gear and the second gear stage including a second worm gear , the first worm gear , the second worm gear and the motor operative to prevent backdrive of the actuator.2. The apparatus as described in claim 1 , wherein the first worm gear of the first gear stage is coupled to the motor via an input shaft.3. The apparatus as described in further including a first helical gear to mesh with the first worm gear claim 2 , the first helical gear coupled to the second worm gear via a first stage shaft.4. The apparatus as described in further including a second helical gear to mesh with the second worm gear claim 3 , the second helical gear coupled to a nut claim 3 , the second helical gear and the nut rotatable relative to a ball screw.5. The apparatus as described in claim 4 , wherein the ball screw includes a rod end to actuate the high lift device of the aircraft.6. The apparatus as described in claim 4 , wherein the first worm gear has a first lead angle and the second worm gear has a second lead angle larger than the first lead angle.7. The apparatus as described in claim 6 , wherein the first lead angle and the second lead angle are selected from a range of lead angles sufficient to prevent backdrive of the actuator.8. The apparatus as described in claim 4 , wherein the first and ...

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Номер записи: 74
04-03-2021 дата публикации

DOUBLE OVER CENTER CRANKSHAFT FLAP MECHANISM

Номер: US20210061442A1
Автор: GRUNER Bryan J., Tsai Kevin R.
Принадлежит:

A flap actuation mechanism incorporates a flap bracket attached to a flap and coupled to an underwing structure with a pivotal coupling. A crankshaft is configured for over center rotation has aligned inboard and outboard crank arms extending from axially spaced inboard and outboard journals disposed in the underwing structure and configured to rotate about a rotation axis of the inboard and outboard journals. A crank pin connected between the inboard and outboard crank arms. An actuating rod has a first end rotatably coupled to the crank pin and a second end coupled to the flap bracket. Rotation of the crankshaft displaces the actuating rod to cause rotation of the flap bracket and the flap. 1. A flap actuation mechanism , comprising:a flap bracket attached to a flap and coupled to an underwing structure with a pivotal coupling; aligned inboard and outboard crank arms extending from axially spaced inboard and outboard journals disposed in the underwing structure and configured to rotate about a rotation axis of the inboard and outboard journals;', 'a crank pin connected between the inboard and outboard crank arms; and,, 'a crankshaft configured for over center rotation having'}an actuating rod, having a first end rotatably coupled to the crank pin and a second end coupled to the flap bracket, wherein rotation of the crankshaft displaces the actuating rod to cause rotation of the flap bracket and the flap.2. The mechanism of claim 1 , wherein rotation of the crankshaft causes forward and aft movement of the actuating rod to cause rotation of the flap bracket and the flap between a stowed position and a deployed position.3. The mechanism of claim 1 , wherein the inboard and outboard journals are spaced apart such that claim 1 , during rotation of the crankshaft claim 1 , a forward portion of the actuating rod can pass through the rotation axis and between the inboard and outboard journals.4. The mechanism of claim 1 , wherein the second end of the actuating rod is ...

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Номер записи: 75
20-02-2020 дата публикации

Motion Control System For Foot-Actuated Flight Controller

Номер: US20200055588A1
Автор: Ezran Shemuel M., Hanlon James Casey
Принадлежит: Configurable Controls, LLC

The present invention discloses improvements to mechanical and electro-mechanical force simulators for aircraft pilot controllers, and particularly foot-actuated controllers. In particular, the invention replaces conventional foot-actuated, pilot controller systems—namely, those that employ multiple, discrete, motion control subsystems to control the various force simulation and trim functions used in modern aircraft to assist pilot control of a given axis of flight—with a single motion control system. The invention accomplishes this in part by eliminating the force-feel spring used in conventional, federated, foot-actuated, motor-coupled pilot controllers. Instead, in a preferred embodiment, the motion control system employs a single actuator (), such as a BLDC motor/gearhead assembly, driven by control electronics that receives inputs from force sensors () and position sensors () mounted on the actuator to, at once, provide both the “feel forces” (FEEL, FRICTION, and DAMPING) and controls trim to the pilot's foot pedals. 1. A motion control system for a foot-actuated flight controller that control a control surface of an aircraft , the system comprising:(a) an actuator coupled to a foot pedal of the controller for generating simulated feel forces on the foot pedal; and i. generate at least two feel forces on the foot pedal at least in part in response to user inputs sensed on the foot pedal; and', 'ii. control a trim function for the controller., '(b) control electronics in communication with the actuator for generating command signals to cause the actuator to'}2. The system of claim 1 , wherein the control electronics generates command signals to cause the actuator to generate at least three feel forces on the foot pedal.3. The system of claim 2 , wherein one of the simulated feel forces generated by the actuator is a mechanical inertia force.4. The system of claim 2 , further including a set of position sensors for simultaneously sensing both pedal position and ...

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Номер записи: 76
02-03-2017 дата публикации

COMPACT PLANETARY ARRANGEMENT FOR FINAL DRIVE

Номер: US20170059003A1
Автор: Turner Jeffrey S., Wagh Rahul
Принадлежит:

A drive assembly, receiving rotational power from an input shaft rotatable about a rotation axis, includes a drive housing and first and second stage planetary gear sets. The first stage planetary gear set includes a first stage sun gear that may be coupled to the input shaft to turn a plurality of first stage planet gears rotatably mounted to a first stage planet carrier that turns with respect to a first stage ring gear, which is fixed with respect to the drive housing. The second stage planetary gear set includes a second stage sun gear turned by the first stage planet carrier to turn a plurality of second stage planet gears rotatably mounted to a second stage planet carrier, which is fixed with respect to the drive housing and about which turns a second stage ring gear. An output hub is rotated about the rotation axis with respect to the drive housing by rotation of the second stage ring gear on the second stage planet carrier. 1. A drive assembly receiving rotational power from an input shaft rotatable about a rotation axis , the drive assembly comprising:a drive housing;a first stage planetary gear set contained in the drive housing, the first stage planetary gear set including a first stage sun gear configured to couple to the input shaft to turn a plurality of first stage planet gears rotatably mounted to a first stage planet carrier that turns with respect to a first stage ring gear that is fixed with respect to the drive housing;a second stage planetary gear set contained in the drive housing, the second stage planetary gear set including a second stage sun gear turned by the first stage planet carrier to turn a plurality of second stage planet gears rotatably mounted to a second stage planet carrier that is fixed with respect to the drive housing and about which turns a second stage ring gear; andan output hub rotated about the rotation axis with respect to the drive housing by rotation of the second stage ring gear on the second stage planet carrier; ...

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Номер записи: 77
04-03-2021 дата публикации

ACTUATOR

Номер: US20210062898A1
Автор: Amaral Rui, Medina Raphael
Принадлежит:

There is provided an actuator comprising screw shaft and a nut assembly. The nut assembly comprises a primary nut for transmitting load through the actuator along a primary load path, and a secondary nut for transmitting load through the actuator along a secondary load path. The secondary nut comprises first and second portions movable relative to one another. As load is transmitted through the actuator along the primary load path the secondary nut does not transmit load through the actuator, wherein upon failure of the primary load path the first and second portions move relative to each other, such relative movement causing the first and second portions to engage the screw shaft and enable transmittal of load through the secondary nut of the actuator along the secondary load path. 1. An actuator comprising:a screw shaft; anda nut assembly,wherein the nut assembly comprises a primary nut for transmitting load through the actuator along a primary load path, and a secondary nut for transmitting load through the actuator along a secondary load path,wherein the secondary nut comprises first and second portions movable relative to one another,wherein as load is transmitted through the actuator along the primary load path the secondary nut does not transmit load through the actuator and upon failure of the primary load path the first and second portions move relative to each other, such relative movement causing the first and second portions to engage the screw shaft and enable transmittal of load through the secondary nut of the actuator along the secondary load path.2. The actuator as claimed in claim 1 , wherein the first portion is configured to move between a first position in which the first portion does not engage the screw shaft claim 1 , and a second position in which the first portion engages the screw shaft to permit load to be transferred through the secondary nut via the first portion.3. The actuator as claimed in claim 2 , wherein the second portion is ...

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Номер записи: 78
22-05-2014 дата публикации

HIGH LIFT SYSTEM FOR AN AIRCRAFT WITH TWO SEPARATE DRIVE UNITS

Номер: US20140138480A1
Автор: Richter Martin
Принадлежит: AIRBUS OPERATIONS GMBH

A high-lift system on a wing of an aircraft is provided. The wing includes a right-hand and a left-hand wing half with movably held high-lift flaps and the right-hand and left-hand wing half are attached to an aircraft fuselage, thus forming a wing root. Each wing half in a region in close proximity to the wing root, includes a drive unit. In each case this drive unit is joined to a transmission shaft mechanically connected to the respective drive unit, which transmission shaft extends from the drive unit in the direction of the end of the respective wing half and is designed to mechanically move the high-lift flaps arranged in the respective wing half. By means of such an arrangement it is possible to do without deflection gear arrangements from a central drive unit to the individual wing halves. 1. A high-lift system on a wing of an aircraft , which wing comprises a left-hand wing half and a right-hand wing half with movably held high-lift flaps and is attached to an aircraft fuselage thus forming a wing root , comprising:an independently operable drive unit arranged in each of the left-hand wing half and right-hand wing half in a region in close proximity to the wing root, which drive unit is joined to a transmission shaft, mechanically coupled to the respective drive unit, that extends from the drive unit in the direction of the end of the respective one of the left-hand wing half and right-hand wing half and is designed to mechanically move the high-lift flaps arranged in the respective one of the left-hand wing half and right-hand wing half2. The high-lift system of claim 1 , further comprising at least one actuator device in each of the left-hand wing half and right-hand wing half mechanically connected to the transmission shaft and to the high-lift flaps to be moved claim 1 , wherein the drive unit claim 1 , when viewed in the direction of the wingspan claim 1 , is arranged closer to the wing root than a first actuator device.3. The high-lift system of claim ...

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Номер записи: 79
22-05-2014 дата публикации

DEPLOYMENT MECHANISM

Номер: US20140138481A1
Автор: SAKOTA Nebojsa
Принадлежит: AIRBUS OPERATIONS LIMITED

The invention provides a deployment mechanism for deploying an auxiliary wing surface device from an aircraft wing body the deployment mechanism providing a first connector portion for connecting the deployment mechanism to the aircraft wing body, a second connector portion for connecting the deployment mechanism to the auxiliary wing surface device, and a telescopic rod linking the first and second connector portions, the telescopic rod comprising an inner rod extendable from inside of an outer rod to increase the length of the telescopic rod, such that the distance between the first and second connector portions can be increased. The invention also provides an aircraft wing an aircraft and a method of operating an aircraft. 2. A deployment mechanism as claimed in claim 1 , wherein the outer rod has an internally threaded portion corresponding to an externally threaded portion of the inner rod claim 1 , such that the inner rod is extendable from inside the outer rod by a screw action of the threaded portions.3. A deployment mechanism as claimed in claim 1 , wherein the telescopic rod comprises an innermost rod claim 1 , an outermost rod and a number of intermediate rods claim 1 , each inner rod in each pair of adjacent rods being extendable from inside of an outer rod in the pair of adjacent rods.4. A deployment mechanism as claimed in claim 1 , wherein the mechanism further comprises a ball screw actuator and ball bearings in the threaded portions of either of the inner and outer rods and wherein movement of the inner rod with respect to the outer rod of the telescopic rod is actuated by the ball screw actuator.5. A deployment mechanism as claimed in claim 4 , wherein the mechanism comprises a rotating shaft and gearing for powering the ball screw actuator.6. A deployment mechanism as claimed in claim 4 , wherein the mechanism comprises an electric motor for powering the ball screw actuator.7. A deployment mechanism as claimed in claim 1 , wherein at least one of ...

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Номер записи: 80
05-03-2015 дата публикации

COMPOUND PLANETARY GEAR UNIT

Номер: US20150065292A1
Автор: Kurth Franz
Принадлежит:

A compound planetary gear unit comprising a first planetary gear set, a second planetary gear set, a planet carrier for both planetary gear sets together, a first sun gear and a second sun gear as well as comprising a first ring gear and a second ring gear, wherein the first planetary gear set is formed by first planetary gears, which are carried by the planet carrier at a radial distance from a central axis of the compound planetary gear unit in such a way that each of said first planetary gears can be rotated about its own first axis of rotation. 11481156942411885101642418116812191110171816. A compound planetary gear unit () comprising a first planetary gear set , a second planetary gear set , a planet carrier () for both planetary gear sets together , a first sun gear () and a second sun gear () as well as comprising a first ring gear () and a second ring gear () , wherein the first planetary gear set comprises first planetary gears () , which are carried by the planet carrier () at a radial distance from a central axis () of the compound planetary gear unit () in such a way that each of said first planetary gears can be rotated about its own first axis of rotation () , and which are thereby in meshing engagement with the first sun gear () and the first ring gear () , and wherein the second planetary gear set comprises second planetary gears ( , ) , which are carried by the planet carrier () at a radial distance from the central axis () in such a way that each of said second planetary gears can be rotated about an axis of rotation () , and which are thereby in meshing engagement with the second sun gear () and the second ring gear () , wherein the first sun gear () comprises a first gear tooth system () with a first number of teeth that is different from a second number of teeth of a second gear tooth system () of the second sun gear (); and that the second planetary gear set comprises at least one planetary gear () with a third gear tooth system () with a third ...

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Номер записи: 81
28-02-2019 дата публикации

ACTUATOR WITH BACKUP COMPONENT FAILURE DETECTION

Номер: US20190063567A1
Автор: Byron Joshua M., Gavriliuc Sorin, Messina Paolo J., Smit David R.
Принадлежит:

An actuator includes a ball screw, a rod provided at least partially within the ball screw, a ball nut, a ball nut restraint, a first biasing member disposed at least partially between a proximate end of the rod and a proximate end of the ball screw, and a second biasing member disposed at least partially between a distal end of the ball nut and an inner surface of the ball nut restraint. 1. An actuator , comprising:a ball screw;a rod provided at least partially within the ball screw;a ball nut;a ball nut restraint;a first biasing member disposed at least partially between a proximate end of the rod and a proximate end of the ball screw; anda second biasing member disposed at least partially between a distal end of the ball nut and an inner surface of the ball nut restraint.2. The actuator of claim 1 , including:a cover disposed at least partially over the ball nut restraint and the ball screw; anda sensor connected to the cover and configured to sense a proximity of the ball nut restraint relative to the cover.3. The actuator of claim 1 , including a strain gauge connected at or about the proximate end of the rod.4. The actuator of claim 3 , wherein the strain gauge includes a wire spool disposed rearward of the proximate end of the rod.5. The actuator of claim 1 , including:a first attachment portion disposed at the proximate end of the rod;a first attachment restraint connected to the first attachment portion;a second attachment portion disposed at a distal end of the rod; anda second attachment restraint connected to the second attachment portion.6. The actuator of claim 5 , wherein a primary load path includes the ball screw claim 5 , the ball nut claim 5 , a torque tube claim 5 , the first attachment portion claim 5 , and the second attachment portion; and a secondary load path includes the rod claim 5 , the ball nut restraint claim 5 , a torque tube restraint claim 5 , the first attachment restraint claim 5 , and the second attachment restraint.7. The ...

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Номер записи: 82
29-05-2014 дата публикации

RUDDER BIAS GAIN CHANGER

Номер: US20140145029A1
Автор: Butler, III Harris K., Veletas Mihalis
Принадлежит: LEARJET INC.

A bias gain system for aircraft rudder comprises a pair of connector ends receiving an actuation force from a bias actuator. A rudder bar interface is positioned between the connector ends. The interface rotates about the rudder bar as a function of the actuation forces from the bias actuators. A mechanism comprises links and joints between the connector ends and the rudder bar is actuatable between a contracted configuration, in which first moment arms are defined between the connector ends and the rudder bar interface, and an expanded gain configuration, in which second moment arms have a greater dimension than the first moment arms. An actuator is connected to the mechanism to actuate the mechanism to actuate the mechanism independently from the actuation forces from the bias actuators, to move the mechanism between configurations. An aircraft and a method for controlling a torque on a rudder bar of an aircraft are also provided. 1. A bias gain system for aircraft rudder , comprising:a pair of connector ends, each said connector end adapted to receive an actuation force from a respective bias actuator;a rudder bar interface positioned between the connector ends and adapted to be connected to a rudder bar for rotation therewith about an axis of the rudder bar as a function of the actuation forces from the bias actuators;a mechanism comprising links and joints between the connector ends and the rudder bar for transmission of the actuation forces to the rudder bar, the mechanism being actuatable between a contracted configuration in which first moment arms are defined between the connector ends and the rudder bar interface, and an expanded gain configuration in which second moment arms are defined between the connector ends and the rudder bar interface, the second moment arms having a greater dimension than the first moment arms; andat least one actuator connected to the mechanism to actuate the mechanism independently from the actuation forces from the bias ...

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Номер записи: 83
09-03-2017 дата публикации

BALL-BEARING CONTROL CABLE, A MECHANICAL SYSTEM, AND AN AIRCRAFT

Номер: US20170066526A1
Автор: Bihel Jean-Romain
Принадлежит:

A ball-bearing control cable comprising an outer sheath surrounding a central blade. The outer sheath comprises a central portion extending between two endpieces. Each endpiece comprises a tubular central body that surrounds the central blade, each central body extending from a proximal abutment secured to the central portion towards a free distal abutment, each abutment of an endpiece projecting from the central body of the endpiece, each endpiece including a resilient system interposed around the central body of the endpiece between an abutment of the endpiece and a ring, the resilient system of one endpiece being arranged between the distal abutment and the ring of the endpiece, and the resilient system of the other endpiece being arranged between the proximal abutment and the ring of the other endpiece. 1. A ball-bearing control cable comprising an outer sheath surrounding a central blade that is movable longitudinally relative to the outer sheath , the central blade being arranged between two rows of bearing members , each row of bearing members being arranged between a face of the central blade and a plate that is opposite an inside face of the outer sheath , the outer sheath comprising a central portion extending between an endpiece referred to as the “first” endpiece and an endpiece referred to as the “second” endpiece , wherein each endpiece comprises a hollow central body that surrounds the central blade , each central body extending longitudinally along at least one abutment referred to as the “proximal” abutment towards a free abutment referred to as the “distal” abutment , each abutment of an endpiece projecting from the central body of the endpiece towards an external medium that is outside the ball-bearing control cable , each endpiece including a resilient system interposed around the central body of the endpiece between an abutment of the endpiece and a ring , each ring being arranged around the central body of the endpiece associated therewith , ...

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Номер записи: 84
12-03-2015 дата публикации

Planetary Gearbox Transmission using Gear Bearings

Номер: US20150072828A1
Автор: Reuter Jeffrey A
Принадлежит:

The invention uses gear bearings and planet carriers to create a planetary transmission with one or more stages. The gear bearing provides the planet gears, sun gears, and ring gear portion of each planetary stage. A planet carrier connects the planet gears of each stage to the next stage, or the output shaft. The planet carrier interfaces to the gear bearing through holes in the planetary gear elements, or posts which extend from the top of the planetary gear elements. Thus each stage can be built from as little as two parts. Each stage of the planetary transmission can be attached to another planetary transmission stage to provide multiple levels of gear reduction. 1. A multistage planetary transmission consisting of one or more stages of a gear bearing and planet carrier , comprising:a gear bearing consisting of herringbone gears configured with a sun gear, planet gears and an outer ring gearintermediate planet carrier assemblies which interface to the planetary gears and drive either an output shaft or the sun gear of the next gear stage2. A multi stage planetary transmission as recited in which is comprised of one or more pairs of gear bearings and adapter plates.3. A multistage planetary transmission as recited in which uses inter-stage planet carrier assemblies claim 1 , including a post and hole assembly claim 1 , a hole and pin assembly claim 1 , and other methods of connecting the rotary motion of the planet gears to the sun gear of the next stage4. A multistage planetary transmission as recited in where the outer carrier gear of the gear bearing has added features to connect the gear bearing to the gear bearing of the next stage5. A multistage planetary transmission as recited in where the outer carrier gear of the gear bearing has added features for mounting the multistage planetary transmission to another structure The invention relates to mechanical gearboxes and transmissions, and more particularly to planetary gear transmission mechanism. The ...

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Номер записи: 85
27-02-2020 дата публикации

Selectable system controller for multi-processor computing systems

Номер: US20200065284A1
Автор: Christopher Brian Noll, Steven A. Avritch
Принадлежит: Hamilton Sundstrand Corp

A system includes a computing system and a cable connector. The computing system includes a plurality of processors and an interconnect circuit configured to connect the plurality of processors to each other. The cable connector is configured to connect to the interconnect circuit and provide a channel identifier to the computing system, and the interconnect circuit is configured to set one of the plurality of processors as a system controller based on the channel identifier.

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Номер записи: 86
19-03-2015 дата публикации

AIRCRAFT

Номер: US20150076282A1
Автор: Hauber Bernhard
Принадлежит:

The invention relates to an aircraft having at least one high lift system which is arranged at the wing of the aircraft and which comprises a drive for converting electrical or hydraulic energy into a speed-controlled rotational movement, wherein the aircraft furthermore has at least one control unit which controls the high lift system, wherein the drive comprises in accordance with the invention a main drive and an alternative drive, wherein the alternative drive is fed by a decentralized energy source. 1. An aircraft having at least one high lift system which is arranged at the wing of the aircraft and which comprises a drive for converting electrical or hydraulic energy into a speed-controlled rotational movement , wherein the aircraft furthermore has at least one control unit which controls the high lift system , whereinthe drive comprises a main drive and an alternative drive, with the alternative drive being fed from a decentralized energy source.2. An aircraft in accordance with claim 1 , wherein the main drive is fed by a power supply system of the aircraft.3. An aircraft in accordance with wherein the decentralized energy sources are batteries or ultracaps.4. An aircraft in accordance with claim 3 , wherein a plurality of batteries are combined to form a batter pack claim 3 , with the battery pack advantageously comprising batteries having different principles of action.5. An aircraft in accordance with claim 1 , wherein the main drive and the alternative drive are connected to the high lift system via a summing transmission.6. An aircraft in accordance with claim 5 , wherein the summing transmission is a spur gear transmission or a bevel gear transmission.7. An aircraft in accordance with claim 1 , wherein the main drive can be separated from the summing transmission via a coupling claim 1 , with the coupling being fed from the decentralized energy source.8. An aircraft in accordance with claim 1 , wherein the decentralized energy source can be recharged ...

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Номер записи: 87
19-03-2015 дата публикации

Flap System for an Aircraft High Lift System or an Engine Actuation and Method for Monitoring a Flap System

Номер: US20150076283A1
Автор: Bernd Schievelbusch, Manfred Fischer
Принадлежит: Liebherr Aerospace Lindenberg GmbH

The invention relates to a flap system for an aircraft high lift system or an engine actuation with a rotary shaft system, one or more drive stations as well as elements for transmitting the drive energy from the rotary shaft system to the one or more drive stations, wherein at least one drive station includes at least two independent load paths with at least one rotational transmission each for actuating the flap kinematics, and per load path at least one mechanically coupling-free synchronization unit is provided for compensating regular load fluctuations between the load paths. The invention furthermore relates to a method for monitoring a flap system with at least two redundant load paths which each comprise at least one rotational transmission, wherein it is cyclically checked whether the difference of the output-side torques of the at least two load paths exceeds a defined threshold value and/or lies within a defined limit range.

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Номер записи: 88
24-03-2022 дата публикации

SINGLE-USE NON-JAMMING STOP MODULE FOR ROTARY DRIVE ACTUATOR

Номер: US20220090663A1
Автор: LARSON Lowell Van Lund
Принадлежит:

A stop module for non-jamming end-of-stroke stoppage of a rotary drive actuator includes timing gears to articulate a stopping pawl, and a low inertia, deformable stopping disk arranged to safely dissipate excess rotational kinetic energy of the rotary actuator. The stop module does not rely on friction to stop and dissipate the excess kinetic energy, but instead relies on predictable deformation of a metallic stopping disk which may be provided in a stopping cartridge of the stop module. Use of a deformable element to dissipate excess energy allows the disclosed stop module to be lighter and smaller than conventional end-of-stroke stopping mechanisms. 1. An actuator for transmitting rotary motion from an input element to an output element , the actuator having an end-of-stroke limit , wherein the actuator comprises:a deformable element and a deforming element, wherein the deformable element is connected to the input element such that the deformable element rotates in response to rotation of the input element;a stopping element movable between a non-stopping position and a stopping position, wherein the deforming element rotates with the deformable element when the stopping element is in the non-stopping position, and wherein relative motion occurs between the deformable element and the deforming element when the stopping element is in the stopping position; anda timing gear responsive to rotation of the input element or the output element, wherein the timing gear is configured to move the stopping element to the stopping position when the end-of-stroke limit is reached;wherein the deforming element causes deformation of the deformable element when relative motion occurs between the deformable element and the deforming element;whereby kinetic energy in the actuator is dissipated through the deformation of the deformable element when the end-of-stroke limit is reached.2. The actuator according to claim 1 , wherein the actuator comprises a stopping cartridge assembly ...

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Номер записи: 89
05-03-2020 дата публикации

DISTRIBUTED TRAILING EDGE WING FLAP SYSTEMS

Номер: US20200070959A1
Автор: Huynh Neal V.
Принадлежит:

Distributed trailing edge wing flap systems are described. An example wing flap system for an aircraft includes a flap and first and second actuators. The flap is movable between a deployed position and a retracted position relative to a fixed trailing edge of a wing of the aircraft. The first and second actuators are configured to move the flap relative to the fixed trailing edge. The first actuator is operatively coupled to the second actuator via a shaft. The first actuator is actuatable via pressurized hydraulic fluid to be supplied from a hydraulic system of the aircraft to the first actuator via a hydraulic module operatively coupled to the first actuator. The first actuator is configured to control movement of the second actuator via the shaft when the hydraulic system and the hydraulic module are functional. The second actuator is actuatable via an electric motor of the second actuator. The electric motor is selectively connectable to an electrical system of the aircraft. The electric motor is connected to the electrical system in response to detection of a failure of the hydraulic system or of the hydraulic module. The second actuator is configured to control movement of the first actuator via the shaft when the electric motor is connected to the electrical system. 1. A wing flap system for an aircraft , the wing flap system comprising:a flap movable between a deployed position and a retracted position relative to a fixed trailing edge of a wing of the aircraft; andfirst and second actuators configured to move the flap relative to the fixed trailing edge, the first actuator being operatively coupled to the second actuator via a shaft, the first actuator being actuatable via pressurized hydraulic fluid to be supplied from a hydraulic system of the aircraft to the first actuator via a hydraulic module operatively coupled to the first actuator, the first actuator configured to control movement of the second actuator via the shaft when the hydraulic system and ...

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Номер записи: 90
19-03-2015 дата публикации

DEVICE AND METHOD FOR ACTIVE CONTROL OF A FORCE FEEDBACK FOR A CONTROL DEVICE

Номер: US20150081139A1
Автор: Beggiora Geoffrey, MERCIER Guillaume, Piaton Jérôme, Vendenbaviere Yann
Принадлежит:

The invention concerns a device for the active control of a force feedback for a control device, comprising a calculator, a position sensor () configured to provide the calculator with an effective position signal (Pm) of the control device, and an actuator () ensuring the displacement of the control device at the command of the calculator, the calculator being configured to use the effective position signal and modulate a setpoint current (lc) delivered to the actuator to ensure the position feedback of the displacement of the control device, characterised in that the calculator is further configured to create at least one saturation terminal (Bsat+, Bsat−) according to a predetermined function of the value of the effective position signal of the position/force law kind, and to saturate the setpoint current using the at least one saturation terminal. 132. A device for active control of a force feedback for a steering member , comprising a calculator , a position sensor () configured to provide the calculator with an effective position signal (Pm) of the steering member , and an actuator () ensuring displacement of the steering member on command from the calculator , the calculator being configured to execute the effective position signal and modulate a setpoint current (Ic) delivered to the actuator to ensure position-controlled displacement of the steering member ,{'b': 21', '23', '32, 'characterized in that the calculator is also configured to develop at least one saturation terminal (Bsat, Bsat+, Bsat−, Bsat-cor) according to a predetermined function of the value of the effective position signal of position/force law type (, , ), and to saturate the setpoint current by means of the at least one saturation terminal.'}219. The device according to claim 1 , wherein the calculator provides a position control in cascade of the actuator claim 1 , the position control in cascade being executed by means of a main position control loop () and an internal secondary speed ...

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Номер записи: 91
18-03-2021 дата публикации

FLAP ACTUATOR MECHANISM

Номер: US20210078694A1
Автор: Tsai Kevin R.
Принадлежит:

An actuator mechanism for a flap includes a first link having a rotary-driven end and a free end, and a second link having a forward end, a mid-portion, and an aft end. The rotary-driven end is pivotally connected to a base structure, the forward end is pivotally connected to the free end, and the aft end is connected to the flap. The actuator mechanism also includes a third link that includes a fixed end, an intermediate connector, and an end connector. The fixed end is pivotally connected to the base structure, and the intermediate connector is pivotally connected to the mid-portion of the second link. The actuator mechanism further includes a flap link including a first end pivotally connected to the end connector, and a second end pivotally connected to the flap. Rotation of the first link causes the flap to transition from a stowed to a fully deployed position. 1. An actuator mechanism for a flap , the actuator mechanism comprising:a first link having a rotary-driven end and a free end, the rotary-driven end pivotally connected to a base structure at a fixed axis;a second link having a forward end, a mid-portion, and an aft end, the forward end pivotally connected to the free end of the first link, and the aft end connected to an underside of the flap, wherein the underside of the flap in a stowed position defines an initial underside profile;a third link comprising a fixed end, an intermediate connector, and an end connector, the fixed end pivotally connected to the base structure, and the intermediate connector pivotally connected to the mid-portion of the second link; anda flap link having a first end and a second end, the first end pivotally connected to the end connector of the third link and the second end pivotally connected to the underside of the flap,wherein rotation of the first link in a first direction causes the flap to transition from the stowed position to a fully deployed position, and wherein the flap link remains below the initial underside ...

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Номер записи: 92
18-03-2021 дата публикации

STOWED BLADE PNEUMATIC CLAMP

Номер: US20210078695A1
Автор: Foskey Christopher Edward, Schank Troy Cyril
Принадлежит: Bell Textron Inc.

An exemplary aircraft includes a pylon rotatable relative to a wing, a proprotor assembly carried by the pylon, the proprotor assembly have a proprotor blade that is foldable to a stowed position extending substantially parallel to the pylon, and a fluidic clamp positioned with the pylon and configured to grip the proprotor blade when in the stowed position, the fluidic clamp including a first inflatable bladder positioned on a first side of an opening for receiving a portion of a proprotor blade and a pressurized fluid source in fluid connection with the first inflatable bladder. 1. A fluidic clamp for securing a proprotor blade , the fluidic clamp comprising:a first inflatable bladder positioned on a first side of an opening for receiving a portion of a proprotor blade; anda pressurized fluid source in fluid connection with the first inflatable bladder.2. The fluidic clamp of claim 1 , wherein the pressurized fluid source comprises a gas or a liquid.3. The fluidic clamp of claim 1 , further comprising a compliant member positioned on a second side of the opening opposite from the first inflatable bladder.4. The fluidic clamp of claim 1 , further comprising a second inflatable bladder positioned on a second side of the opening opposite from the first inflatable bladder.5. The fluidic clamp of claim 1 , further comprising a controller in connection with the pressurized fluid source to selectively inflate the first inflatable bladder.6. The fluidic clamp of claim 1 , further comprising a controller in connection with the pressurized fluid source to selectively inflate the first inflatable bladder; anda sensor in communication with the controller and configured to sense a position of a proprotor blade.7. The fluidic clamp of claim 6 , further comprising a compliant member positioned on a second side of the opening opposite from the first inflatable bladder.8. The fluidic clamp of claim 1 , further comprising a second inflatable bladder positioned on a second side of ...

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Номер записи: 93
26-03-2015 дата публикации

VARIABLE CAMBER FLAP SYSTEM AND METHOD

Номер: US20150083852A1
Автор: Finn Michael R., Gardner Mark J., Good Mark S., Jones Sarah A., Moser Matthew A.
Принадлежит: The Boeing Company

A variable camber system for an aircraft may include a variable camber trim unit (VCTU) positioned between an inboard device and an outboard device. The inboard device and the outboard device may be mounted to at least one of a leading edge and a trailing edge of a wing. The VCTU may include a speed sum gearbox having an inboard shaft coupled to the inboard device and an outboard shaft coupled to the outboard device. The VCTU may additionally include a VCTU electric motor engaged to the speed sum gearbox. The VCTU electric motor may be selectively operable in conjunction with the speed sum gearbox to rotate the outboard shaft independent of the inboard shaft in a manner causing the outboard device to be actuated independent of the inboard device. 1. A variable camber system for an aircraft , comprising: a speed sum gearbox having an inboard shaft coupled to the inboard device and an outboard shaft coupled to the outboard device;', 'a VCTU electric motor engaged to the speed sum gearbox; and', 'the VCTU electric motor being selectively operable in conjunction with the speed sum gearbox to rotate the outboard shaft independent of the inboard shaft in a manner causing the outboard device to be actuated independent of the inboard device., 'a variable camber trim unit (VCTU) positioned between an inboard device and an outboard device mounted to at least one of a leading edge and a trailing edge of a wing, and including2. The variable camber system of claim 1 , further comprising:a power-off brake coupled to the VCTU electric motor; andthe speed sum gearbox being configured such that when the power-off brake is applied, the outboard shaft is driven by the inboard shaft causing the inboard device and the outboard device to be actuated in unison.3. The variable camber system of claim 2 , further including:a central motor configured to drive the inboard device.4. The variable camber system of claim 3 , wherein:the central motor is coupled to a central brake preventing ...

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Номер записи: 94
26-03-2015 дата публикации

AERODYNAMIC SURFACE DRIVE MECHANISM

Номер: US20150083854A1
Автор: Da Silva Israel, FERREIRA Clóvis Augusto Eça
Принадлежит:

There is described an aerodynamic surface drive mechanism () containing at least a drive combination (), each drive combination () comprising a fixed element () associated to a fixed aircraft structure and a first mobile component () connected pivotably by a first end to the fixed element () by way of an articulation axis (E) and associated to an actuator () by an opposite end, the aerodynamic surface drive mechanism () further comprises a second mobile component () rotationally associated to the first mobile component () by way of primary swivel joints () linearly disposed along a vertical axis (Y) and rotationally connected to the aerodynamic surface () by way of secondary swivel joints () linearly disposed along a horizontal axis (Z); the first mobile component () and the second mobile component () simultaneously moving the aerodynamic surface () linearly and rotatively by means of the actuator () and of the primary swivel joints () and secondary swivel joints (′). 120505050502122213020232224244025252223403024242525. An aerodynamic surface drive mechanism () containing at least a drive combination ( , ′) , each drive combination ( , ′) comprising a fixed element () associated to a fixed aircraft structure and a first mobile component () connected pivotably by a first end to the fixed element () by way of an articulation axis (e) and associated to an actuator () by an opposite end , the aerodynamic surface drive mechanism () being characterized by further comprising a second mobile component () rotationally associated to the first mobile component () by way of primary swivel joints ( , ′) linearly disposed along a vertical axis (Y) and rotationally connected to the aerodynamic surface () by way of secondary swivel joints ( , ′) linearly disposed along a horizontal axis (Z); the first mobile component () and the second mobile component () simultaneously moving the aerodynamic surface () linearly and rotatively by means of the actuator () and of the primary swivel ...

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Номер записи: 95
12-06-2014 дата публикации

SUPPORT ASSEMBLY

Номер: US20140158822A1
Автор: King Simon, LANGLEY Anthony
Принадлежит: AIRBUS OPERATIONS LIMITED

An assembly to support a wing leading or trailing edge device during deployment and retraction of the device includes a fixed support member attachable to the support structure of an aircraft wing, an intermediate link arm having one end pivotally mounted for rotation relative to the support member about a first axis and, a primary link arm having a first end pivotally mounted to the opposite end of the intermediate link arm for rotation relative to the intermediate link arm about a second axis. A second end of the primary link arm is for attachment to the device via a bearing element so that the device can move relative to the primary link arm in any direction when the intermediate and primary link arms rotate about the first and second axes, respectively, during deployment or retraction of the device from the aircraft wing. 1. An assembly to support a wing leading or trailing edge device during deployment and retraction of said wing leading or trailing edge device from a wing of an aircraft , the assembly comprising a fixed support member attachable to the support structure of an aircraft wing , an intermediate link arm having one end pivotally mounted to the support member for rotation relative to the support member about a first axis and , a primary link arm having a first end attached to the opposite end of the intermediate link arm such that the primary link arm and intermediate link arm are movable relative to each other , a second end of the primary link arm being configured for attachment to a wing leading or trailing edge device , the arrangement being such that the assembly allows a wing leading or trailing edge device to move in any direction during deployment or retraction of said wing leading or trailing edge device from an aircraft wing.2. An assembly according to claim 1 , wherein said first end of the primary link arm is attached to said opposite end of the intermediate link arm via a bearing element so that the primary link arm claim 1 , together ...

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Номер записи: 96
24-03-2016 дата публикации

TRACK ROLLER BEARINGS WITH ROLLING ELEMENTS OR LINERS

Номер: US20160083081A1
Автор: Cowles, Fredericksen Arnold E., JR. John H., Phoenix Jay
Принадлежит: ROLLER BEARING COMPANY OF AMERICA, INC.

A track roller assembly includes a split inner ring, a split outer ring, a one piece inner ring and/or a one piece outer ring and a liner or plurality of rolling elements engaging therewith, the track roller assembly being disposed in a structure of an Airbus A-350 aircraft, an Airbus A-320 aircraft, an Airbus A320Neo aircraft, an Airbus A330 aircraft, an Airbus A330Neo aircraft, an Airbus A321 aircraft, an Airbus A340 aircraft, and an Airbus A380 aircraft. 1. A track roller assembly comprising a split inner ring and a plurality of rolling elements in rolling engagement with the split inner ring , the track roller assembly being disposed in a structure of at least one of an Airbus A-350 aircraft , an Airbus A-320 aircraft , an Airbus A320Neo aircraft , an Airbus A330 aircraft , an Airbus A330Neo aircraft , an Airbus A321 aircraft , an Airbus A340 aircraft , and an Airbus A380 aircraft.2. An actuation system for deploying and retracting a lift assisting device of an edge of a wing of an aircraft , the actuation system comprising:a track pivotally coupled to the lift assisting device, the track having first and second outer surfaces and side surfaces;a shaft rotationally coupled within the wing of the aircraft and operable, in response to flight control signals, to deploy or retract the lift assisting device;means for actuating the lift assisting device, coupled to the shaft, between a retracted position to a deployed position along an arcuate path;a plurality of track roller bearings rotatably contacting the first and second outer surfaces of the track to guide the track along the arcuate path; andthe plurality of track roller bearings including at least one track roller assembly having a split inner ring and a plurality of rolling elements in rolling engagement with the split inner ring.3. The actuation system of claim 2 , wherein the plurality of track roller bearings includes at least one track roller assembly in rotational contact with an upper surface of the ...

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Номер записи: 97
24-03-2016 дата публикации

LINK FOR COUPLING AN AIRCRAFT LIFT DEVICE TO A TRACK

Номер: US20160083082A1
Автор: Bishop Ben
Принадлежит:

A link for coupling an aircraft lift device to a track that is deployable about an axis of rotation. The link includes a track member to be fixed relative to the track, a lift device member to be fixed relative to the lift device, and an intermediate member that couples the track member to the lift device member. The link is configured to allow for sliding movement of the lift device member relative to the track member such that the lift device is slidable along a linear path relative to the track in a direction of the axis of rotation of the track. 1. A link for coupling an aircraft lift device to a track that is deployable about an axis of rotation comprising:a track member to be fixed relative to said track,a lift device member to be fixed relative to said lift device, andan intermediate member that couples the track member to the lift device member,wherein the link is configured to allow for sliding movement of the lift device member relative to the track member such that the lift device is slidable along a linear path relative to the track in a direction of an axis of rotation of the track.2. The link according to claim 1 , wherein one of the track member and the lift device member is slidable relative to the intermediate member.3. The link according to claim 2 , wherein the one of the track member and the lift device member comprises a recess claim 2 , and the intermediate member is slidably received in the recess.4. The link according to claim 2 , wherein the other one of the track member and lift device member is rotatable relative to the intermediate member about a first axis.5. The link according to claim 4 , wherein the other one of the track member and lift device member comprises first and second arms claim 4 , and the intermediate member is rotatably received between the first and second arms such that the first axis extends therebetween.6. The link according to claim 4 , wherein the one of the track member and lift device member is rotatable relative ...

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Номер записи: 98
14-03-2019 дата публикации

TORQUE TRANSMISSION DEVICE

Номер: US20190078628A1
Автор: EGAN Jason, Kracke Jeremy
Принадлежит:

A torque transmission device comprises: an input drive shaft, a housing, an earth ring, a bearing cage, and a plurality of bearings held by the bearing cage between an inner surface of the earth ring and an outer surface of the input drive shaft. The earth ring is mounted within the housing such that an outer surface of the earth ring contacts an inner surface of the housing along a contact area, thereby providing a frictional interface at the contact area for transmitting torque from the earth ring to the housing. 1. A torque transmission device comprising:an input drive shaft;a housing;an earth ring having a thickness;a bearing cage; anda plurality of bearings held by the bearing cage between an inner surface of the earth ring and an outer surface of the input drive shaft,wherein the earth ring is mounted within the housing such that an outer surface of the earth ring contacts an inner surface of the housing along a contact area, thereby providing a frictional interface at the contact area for transmitting torque from the earth ring to the housing.2. A torque transmission device according to claim 1 , wherein the contact area takes the shape of a curved surface of a cylinder claim 1 ,wherein torque is transmitted from the earth ring to the housing only via the contact area.3. A torque transmission device according to claim 1 , wherein an output is connected to the bearing cage.4. A torque transmission device according to claim 3 , wherein the torque transmission device is arranged such that claim 3 , in a first operating condition claim 3 , the output is operable to rotate on rotation of the input drive shaft claim 3 ,wherein the torque transmission device is arranged such that, in a second operating condition, torque is prevented from being transmitted from the input drive shaft to the output.5. A torque transmission device according to claim 4 , wherein the torque transmission device is a torque limiter claim 4 , wherein the first operating condition is a ...

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Номер записи: 99
14-03-2019 дата публикации

MULTI-BAR LINKAGE MECHANISM

Номер: US20190078669A1
Автор: Liu Shijie
Принадлежит:

Systems provide and methods utilize a multi-bar linkage in conjunction with an actuator to provide mechanical advantage which increases as a control surface moves towards the extremes of its operating envelope. 1. A system configured to provide a concave curve of torque against output position , comprising:a three-way linkage joint;a control link having a control ground end and a control joint end, wherein the control ground end is rotatably coupled to a fixed portion of a controlled system, and wherein the control joint end is rotatably coupled to the three-way linkage joint;a drag link having a drag drive end and a drag joint end, wherein the drag joint end is rotatably coupled to the three-way linkage joint;a drive link having a drive drag end and a drive actuator end, wherein the drive drag end is rotatably coupled to the drag drive end, and wherein the drive actuator end is operatively coupled to an actuator;an output link having an output surface end and an output joint end, wherein the output joint end is rotatably coupled to the three-way linkage joint; anda surface link having a surface output end and a surface control end, wherein the surface output end is rotatably coupled to the output surface end, and wherein the surface control end is operatively coupled to a movable surface of the controlled system.2. The system of claim 1 , further comprising the controlled system claim 1 , wherein the controlled system is a portion of an aircraft.3. The system of claim 1 , further comprising the controlled system claim 1 , wherein the controlled system is a portion of a robot.4. The system of claim 1 , further comprising the controlled system claim 1 , wherein the controlled system is a portion of a marine vessel.5. The system of claim 1 , further comprising the controlled system claim 1 , wherein the controlled system is an autonomous vehicle.6. The system of claim 1 , further comprising the movable surface and the movable surface comprises a movable control ...

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Номер записи: 100