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Небесная энциклопедия

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

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Мониторинг СМИ

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

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Поддерживает ввод нескольких поисковых фраз (по одной на строку). При поиске обеспечивает поддержку морфологии русского и английского языка
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Применить Всего найдено 440. Отображено 100.
17-01-2019 дата публикации

Gimballed Augmentation Shroud

Номер: US20190017468A1
Автор: Evulet Andrei
Принадлежит:

A system for enhancing the thrust produced by an aircraft propulsion element having a propulsion fluid outlet includes a shroud element having an inlet, an outlet and a diffusing section positioned between the shroud element inlet and shroud element outlet. The shroud element is coupled to the aircraft such that the diffusing section is positioned directly downstream of the propulsion fluid outlet. At least one actuator is operable to rotate the shroud element about a first transverse axis of the shroud element and a second transverse axis of the shroud element. 1. A system for enhancing the thrust produced by an aircraft propulsion element having a propulsion fluid outlet , the system comprising:a shroud element comprising an inlet, an outlet and a diffusing section positioned between the shroud element inlet and shroud element outlet, the shroud element being coupled to the aircraft such that the diffusing section is positioned directly downstream of the propulsion fluid outlet; andat least one actuator operable to rotate the shroud element about a first transverse axis of the shroud element and a second transverse axis of the shroud element.2. The system of claim 1 , wherein the shroud element further comprises a throat section positioned between the shroud element inlet and shroud element outlet claim 1 , the diffusing section being downstream of the throat section.3. The system of claim 1 , further comprising hinges coincident with at least a portion of the first and second transverse axes and on which the shroud element is mounted claim 1 , the hinges permitting rotation of the shroud element about the first and second transverse axes.4. The system of claim 1 , wherein the shroud element is coupled to the aircraft such that the diffusing section is positioned to simultaneously receive primary fluid from the propulsion fluid outlet and secondary fluid from the ambient.5. The system of claim 1 , wherein the shroud element outlet is serrated.6. A propulsion ...

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30-04-2015 дата публикации

METHODS AND APPARATUS FOR PASSIVE THRUST VECTORING AND PLUME DEFLECTION

Номер: US20150113946A1
Принадлежит: The Boeing Company

A flow vectoring turbofan engine employs a fixed geometry fan sleeve and core cowl forming a nozzle incorporating an asymmetric convergent/divergent (con-di) and/or curvature section which varies angularly from a midplane for reduced pressure in a first operating condition to induce flow turning and axially symmetric equal pressure in a second operating condition for substantially axial flow. 1. A flow vectoring turbofan engine , comprising:a fixed geometry fan sleeve and core cowl forming a nozzle, the nozzle incorporating asymmetric convergence/divergence (con-di) and wall curvature varying angularly from a midplane for maximum con-di in a selected portion for reduced pressure in a first operating condition to induce flow turning and axially symmetrically equal pressure in a second operating condition.2. The flow vectoring turbofan engine as defined in claim 1 , wherein the first operating condition comprises a nozzle pressure ratio below a threshold allowing unchoked flow through a throat and an exit of the nozzle.3. The flow vectoring turbofan engine as defined in claim 2 , wherein the second operating condition comprises a nozzle pressure ratio above the threshold creating a sonic wave for choked flow through the throat of the nozzle.4. The flow vectoring turbofan engine as defined in claim 1 , wherein the midplane is vertical and a selected portion having maximum con-di is a bottom portion of the nozzle for downward vectoring of flow in the first operating condition.5. The flow vectoring turbofan engine as defined in claim 4 , wherein the core cowl has a symmetrical curvature and the fan sleeve exit is not aligned with a constant nacelle station having an exit.6. The flow vectoring turbofan engine as defined in claim 4 , wherein the core cowl has a symmetrical increased curvature and the fan sleeve has decreased asymmetrical con-di in the sleeve.7. The flow vectoring turbofan engine as defined in claim 4 , wherein the core cowl has an asymmetric increasing ...

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04-05-2017 дата публикации

Convertible engine exhaust for rotocraft

Номер: US20170121033A1
Принадлежит: Sikorsky Aircraft Corp

An exhaust system for a rotary wing aircraft includes a diffuser located at an airframe of the rotary wing aircraft and operably connected to an engine of the rotary wing aircraft, and a chimney extending upwardly from a diffuser wall. A door is movably positioned at the diffuser to selectively direct an engine exhaust flow through the diffuser or through the chimney. A rotary wing aircraft includes an airframe, a rotor system positioned at the airframe and rotatable about a rotor axis, and an engine operably connected to the rotor system to drive rotation of the rotor system. The aircraft includes an engine exhaust system including a diffuser positioned at the airframe and operably connected to the engine, a chimney extending upwardly from a diffuser wall, and a door positioned at the diffuser to selectively direct an engine exhaust flow through the diffuser or through the chimney.

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25-05-2017 дата публикации

COANDA DEVICE FOR A ROUND EXHAUST NOZZLE

Номер: US20170145954A1
Принадлежит:

A gas turbine engine system is disclosed herein. The gas turbine engine system includes an engine core configured to discharge air through an exhaust nozzle along a central axis and a thrust director arranged near the exhaust nozzle and configured to redirect the discharge air by applying flow to the discharge air near the exhaust nozzle. 1. A gas turbine engine system comprisingan engine core configured to produce discharge air directed through a round exhaust nozzle along a central axis, anda thrust director arranged near the round exhaust nozzle and configured to redirect the discharge air by applying flow to the discharge air near the exhaust nozzle, the thrust director including an arcuate momentum nozzle and an arcuate coanda nozzle,wherein the arcuate momentum nozzle extends along a constant radius from the central axis and is configured to discharge flow generally perpendicular to and toward the central axis, and the arcuate coanda nozzle extends along a constant radius from the central axis and is configured to discharge flow generally parallel to and along the central axis where the coanda surface influences the discharge air to turn and exit the nozzle perpendicular to the central axis.2. The gas turbine engine system of claim 1 , wherein the thrust director includes a manifold coupled to the arcuate momentum nozzle and to the arcuate coanda nozzle.3. The gas turbine engine system of claim 2 , wherein the manifold is a round claim 2 , annular component that extends around the central axis.4. The gas turbine engine system of claim 3 , wherein the engine core includes a compressor claim 3 , a combustor claim 3 , and a turbine claim 3 , the thrust director includes at least one flow feed tube claim 3 , and the at least one flow feed tube extends from the compressor to the manifold to conduct flow from the compressor to the manifold.5. The gas turbine engine system of claim 4 , wherein the manifold is formed to include a plurality of apertures arranged to ...

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07-05-2020 дата публикации

Thrust Vector Nozzle

Номер: US20200141354A1
Автор: Wu Guanhao
Принадлежит:

A thrust vectoring exhaust nozzle is disclosed. The nozzle includes an inner nozzle for changing a first degree-of-freedom of exhaust gas, an outer nozzle for changing a second degree-of-freedom of exhaust gas, a mounting bracket, a first linear actuator, a second linear actuator, a first double universal joint, and a second double universal joint. The inner nozzle is coupled to the outer nozzle. The inner nozzle is coupled to the mounting bracket. The outer nozzle is coupled to the first and second joint. When the nozzle is mounted, the inner nozzle, the outer nozzle, and the exhaust are coaxially aligned in neutral position. Actuation of the first and second linear actuators drives the first and second double universal joints independently to each other. The independent motion of the first and second double universal joints rotates the inner and outer nozzles simultaneously about the exhaust in a horizontal direction and vertical direction enabling thrust vectoring. 1. A thrust vectoring exhaust nozzle apparatus , comprising:an inner nozzle to change a first degree-of-freedom of exhaust gas emanating from an exhaust of a gas turbine; andan outer nozzle to change a second degree-of-freedom of exhaust gas emanating from the exhaust of the gas turbine; the inner nozzle is disposed within the outer nozzle; and', 'the inner nozzle, the outer nozzle, and the exhaust nozzle of the gas turbine are coaxial aligned relative to one another in neutral position., 'wherein2. The thrust vectoring exhaust nozzle apparatus of claim 1 , further comprising:a mounting bracket removably attachable to the gas turbine;a first linear actuator coupled to the mounting bracket;a second linear actuator coupled to the mounting bracket;a first double universal joint coupled to the first linear actuator;a second double universal joint coupled to the second linear actuator; andwherein:the outer nozzle is coupled to the first double universal joint and the second double universal joint;the inner ...

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22-06-2017 дата публикации

AIRCRAFT WITH A HOT AIR EXHAUST THAT COMPRISES TWO PIVOTALLY MOUNTED EXHAUST SECTIONS

Номер: US20170175673A1
Автор: WELLHAUSEN Christian
Принадлежит: AIRBUS HELICOPTERS DEUTSCHLAND GMBH

An aircraft with at least one engine that generates a hot air flow in operation of the aircraft, wherein at least one hot air exhaust is provided for exhausting the generated hot air flow, the at least one hot air exhaust comprising at least one first exhaust section that is mounted in a rotatable manner to at least one second exhaust section via an associated off-axis swivel joint, wherein an actuating member is provided that is adapted for applying a turning moment to the at least one second exhaust section in operation of the aircraft in order to displace a longitudinal axis of the at least one second exhaust section with respect to a longitudinal axis of the at least one first exhaust section by a predetermined displacement angle. 1. An aircraft with a fuselage that defines at least one drive system accommodating region , the drive system accommodating region accommodating at least one engine that generates a hot air flow in operation of the aircraft , wherein at least one hot air exhaust is provided for exhausting the generated hot air flow , the at least one hot air exhaust comprising at least one first exhaust section and at least one second exhaust section , the at least one second exhaust section being mounted in a rotatable manner to the at least one first exhaust section via an associated off-axis swivel joint , characterized in that an actuating member is provided , the actuating member being adapted for applying a turning moment to the at least one second exhaust section in operation of the aircraft in order to displace a longitudinal axis of the at least one second exhaust section with respect to a longitudinal axis of the at least one first exhaust section by a predetermined displacement angle on the basis of at least one of: current aviation parameters of the aircraft in operation , a current temperature in a region of the fuselage to which the generated hot air flow is exhausted via the at least one hot air exhaust and/or a current operation ...

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06-06-2019 дата публикации

Thrust vector controller

Номер: US20190170087A1
Автор: Yao-Chang Lin
Принадлежит: Individual

A thrust vector controller includes an airflow guiding member, a connecting member, a first driving device, and a second driving device. The airflow guiding member is adjacent to an air exhaust opening. The airflow guiding member includes a main body, a first driving portion, a second driving portion, and a connecting portion. Airflow passes through the main body and is guided by the main body. The first driving portion, the second driving portion, and the connecting portion are connected to the main body. The connecting member is movably connected to the connecting portion and an exhaust propulsion device. The main body is movably connected to the exhaust propulsion device through the connecting portion and the connecting member. The first driving device is connected to the first driving portion and drives the first driving portion to move the airflow guiding member toward a first direction.

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27-06-2019 дата публикации

TWO-DIMENSIONAL SUPERSONIC NOZZLE THRUST VECTORING USING STAGGERED RAMPS

Номер: US20190195169A1

A system and method for vectoring the thrust of a supersonic, air-breathing engine. A thrust vectoring mechanism uses two asymmetrically staggered ramps; one placed at the throat, the other positioned at the exit lip of the nozzle of the engine to re-direct exhaust flow off-axis with the nozzle. 1. An apparatus for vectoring thrust of a air-breathing engine comprising a nozzle having an inlet , exit , and a throat , disposed between the inlet and exit , the apparatus comprising:first and second deployable ramps;wherein the first ramp is disposed on an internal wall of the nozzle at or near the throat;wherein the second ramp is and the positioned at or near the exit of the nozzle; andwherein the first and second ramps are configured to redirect exhaust flow of the nozzle.2. The apparatus of claim 1 , wherein the first and second ramps are asymmetrically staggered.3. The apparatus of claim 1 , wherein the engine comprises a supersonic engine claim 1 , and the nozzle comprises a converging section comprising the throat and a diverging section leading to the exit.4. The apparatus of claim 1 , wherein the first and second ramps each have deployed states and un-deployed states;wherein the exhaust flow is axial when the first and second ramps are in the un-deployed state; andwherein the exhaust flow is redirected to a non-axial condition when the first and second ramps are in the deployed state.5. The apparatus of :wherein the nozzle comprises a nozzle axis and exhaust flow comprises a sonic line; andwherein the sonic line is skewed off-axis with the nozzle axis to vector thrust from the engine.6. The apparatus of :wherein the first ramp is configured to separate axial flow entering from the inlet; andwherein the second ramp is configured to direct re-attaching airflow at the exit to an off-axis condition.7. The apparatus of claim 4 , wherein at least one of the first and second ramps comprise a flap pivoting about a hinge from the non-deployed state to the deployed state. ...

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09-06-1995 дата публикации

Method and device for producing thrust

Номер: RU2037066C1

FIELD: jet engine engineering. SUBSTANCE: method includes producing combustion products of high pressure and temperature, accelerating the combustion products up to a velocity which corresponds to the supercritical pressure drop, and issuing the combustion products to the ambient in the form of counter streams perpendicular to the thrust vector under the flat base. To control the thrust in magnitude, the ratio of areas of critical sections of the counter flows to the area of the flat base is changed. To control the thrust direction the flat base is turned. The device has combustion chamber with gas duct 1, turbine 2, outlet pipe 3, and central body 6 with flat base 7 perpendicular to the axis of outlet pipe 3 whose face is made up as movable squeezing casing 4. The casing and the central body define a nozzle whose outlet sections face each other. Central body 6 and squeezing casing 4 are provided with drives 18 for their movement and rotation. The central body can be made up as sections which can move one with respect to the other. EFFECT: enhanced efficiency. 3 cl, 3 dwg 9901$ 0с ПЧ ГЭ РОССИЙСКОЕ АГЕНТСТВО ПО ПАТЕНТАМ И ТОВАРНЫМ ЗНАКАМ (19) ВИ” 2 037 066. (51) МПК 13) Сл Е 02 К 9/80 12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ РОССИЙСКОЙ ФЕДЕРАЦИИ (21), (22) Заявка: 93046474/23, 01.10.1993 (46) Дата публикации: 09.06.1995 (56) Ссылки: Вопросы ракетной техники, 1963, М 3, с.62, рис.13. (71) Заявитель: Знаменский Владимир Павлович, Соколов Сергей Викторович, Чекмасов Владислав Дмитриевич (72) Изобретатель: Знаменский Владимир Павлович, Соколов Сергей Викторович, Чекмасов Владислав Дмитриевич (73) Патентообладатель: Знаменский Владимир Павлович, Соколов Сергей Викторович, Чекмасов Владислав Дмитриевич (54) СПОСОБ ПОЛУЧЕНИЯ ТЯГИ И УСТРОЙСТВО ДЛЯ ЕГО ОСУЩЕСТВЛЕНИЯ (57) Реферат: Использование: в реактивной технике, в частности в способах и устройствах для создания тяги и управления вектором тяги в двигателях. Сущность изобретения: способ заключается в выработке продуктов сгорания ...

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31-05-2018 дата публикации

Turbojet engine flat nozzle

Номер: RU2656170C1

FIELD: engines and pumps. SUBSTANCE: invention relates to the field aviation engine-building, namely to design of the turbojet engines flat nozzles. Turbojet engine flat nozzle contains body with sidewalls fixed thereto, subsonic, supersonic and external flaps, as well as longitudinal levers, subsonic flaps control levers, subsonic flaps hydraulic control cylinders and brackets. Subsonic flaps are pivotally secured to the side walls and pivotally connected to the supersonic flaps. By one end the outer flaps are pivotally connected to the side walls, and by the other one are installed in the guides made in supersonic flaps. Brackets are mounted on the body and are connected to the longitudinal levers by means of tie-rods. Brackets are provided with eyelets, to which the hydraulic cylinders bodies are pivotally attached, connected by rods to the subsonic flaps control levers. By one end the longitudinal levers are installed from the outside of the body near the nozzle horizontal symmetry plane, and by the other end are connected to the side walls near the nozzle critical cross-section. EFFECT: invention allows to increase the flat nozzle strength due to the side walls deformation reducing, to reduce the body and the flat nozzle weight, and to provide the subsonic and supersonic flaps hermetic articulation with the nozzle side walls. 1 cl, 7 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 656 170 C1 (51) МПК F02K 1/06 (2006.01) F02K 1/12 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F02K 1/006 (2018.01); F02K 1/06 (2018.01); F02K 1/1269 (2018.01) (21)(22) Заявка: 2017118864, 31.05.2017 (24) Дата начала отсчета срока действия патента: Дата регистрации: Приоритет(ы): (22) Дата подачи заявки: 31.05.2017 (45) Опубликовано: 31.05.2018 Бюл. № 16 2 6 5 6 1 7 0 R U (54) Плоское сопло турбореактивного двигателя (57) Реферат: Изобретение относится к области авиационного двигателестроения, а именно к конструкции плоских ...

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05-12-2018 дата публикации

Aircraft turbojet engine flat nozzle

Номер: RU2674232C1

FIELD: engines and pumps. SUBSTANCE: invention relates to the field aviation engine-building, namely to design of the turbojet engines flat nozzles. Flat nozzle comprises successively mounted and hingedly connected to each other housing, subsonic doors and supersonic doors, as well as external flaps connected to the hull and supersonic flaps, as well as fairings, each of which is made in a U-shaped cross-section and contacts the corresponding supersonic flap along the lateral surfaces. On the side of the cut of the flat nozzle, the fairing is in contact with the outer surface of the supersonic leaf, and from the opposite side of the fairing a cylindrical tip is made, contacting the outer surface with the response surface of the adjacent end of the corresponding outer leaf. Longitudinal axis of the cylindrical end is aligned with the axis of rotation of the corresponding supersonic leaf. Every fairing is rigidly fixed on the corresponding supersonic leaf in places of contact with the latter. Every flap is connected to the hull by means of at least one bracket, one end rigidly connected to the first and the other end pivotally connected to the latter. With the corresponding supersonic flap, any of the outer flaps is connected by means of at least one bracket rigidly connected to the latter, and at least one link connected to the first by means of a hinged connection. Axis of rotation of said articulated joint is aligned with the axis of rotation of the supersonic flap. Bracket with the linkage is articulated. EFFECT: invention makes it possible to reduce the loss of thrust of a flat nozzle by reducing the bottom resistance. 1 cl, 4 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 674 232 C1 (51) МПК F02K 1/12 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F02K 1/006 (2018.05); F02K 1/1269 (2018.05) (21)(22) Заявка: 2017135015, 05.10.2017 (24) Дата начала отсчета срока действия патента: Дата регистрации: ...

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27-07-2011 дата публикации

Jet nozzle with thrust orientation, its operating method, jet turbine engine and unpiloted aircraft equipped with such nozzle

Номер: RU2425241C2
Принадлежит: Снекма

Реактивное сопло с ориентацией тяги, сформированное таким образом, чтобы разделять основной поток создающих реактивную тягу газов, поступающих из генератора газов, на первый и второй потоки для выброса в первое и второе полусопла, включает два средства управления. Первое средство управления представляет собой средство распределения основного потока в каждое из двух полусопел посредством сужения эффективного сечения горловины одного из полусопел. Второе средство управления представляет собой средство ориентации вектора тяги, создаваемой каждым из двух полусопел, причем оба упомянутые средства управления представляют собой средства со струйной инжекцией. Еще одно изобретение группы относится к способу функционирования реактивного сопла, согласно которому средства струйной инжекции запитывают воздухом, отобранным в компрессоре генератора газа, а отбор воздуха в указанном компрессоре осуществляют непрерывно. Другие изобретения группы относятся к турбореактивному двигателю и беспилотному летательному аппарату, каждый из которых включает указанное выше реактивное сопло. Изобретение позволяет повысить надежность и снизить стоимость реактивного сопла с ориентацией тяги. 4 н. и 9 з.п. ф-лы, 6 ил. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) 2 425 241 (13) C2 (51) МПК F02K 1/28 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ, ПАТЕНТАМ И ТОВАРНЫМ ЗНАКАМ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (21)(22) Заявка: 2006123330/06, 29.06.2006 (24) Дата начала отсчета срока действия патента: 29.06.2006 (43) Дата публикации заявки: 10.01.2008 Бюл. № 1 (73) Патентообладатель(и): СНЕКМА (FR) 2 4 2 5 2 4 1 (45) Опубликовано: 27.07.2011 Бюл. № 21 2 4 2 5 2 4 1 R U Адрес для переписки: 129090, Москва, ул. Б.Спасская, 25, стр.3, ООО "Юридическая фирма Городисский и Партнеры", пат.пов. Ю.Д.Кузнецову, рег.№ 595 C 2 C 2 (56) Список документов, цитированных в отчете о поиске: US 2002/0189232 A1, 19.12.2002. US 6112513 A, 05.09.2000. GB 730573 A, 25.05.1955. Большая советская энциклопедия. Под ред ...

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06-05-2022 дата публикации

Adjustable turbojet nozzle

Номер: RU2771587C1

Изобретение относится к области авиационного двигателестроения, а именно к конструкции регулируемых сопел турбореактивных двигателей. Регулируемое сопло турбореактивного двигателя содержит последовательно установленные корпус, имеющий в выходном сечении прямоугольную форму. Каждая из дозвуковых и сверхзвуковых створок выполнена в поперечном разрезе уголковой формы, образованной двумя пластинами, соединенными по торцам под тупым углом, вершина которого направлена от продольной оси турбореактивного двигателя. Каждая из дозвуковых створок шарнирно соединена с боковыми стенками. Между верхним торцом выходного фланца и близлежащим торцом верхней дозвуковой створки, а также между нижним торцом выходного фланца и близлежащим торцом нижней дозвуковой створки установлено по уплотнительному элементу, жестко соединенному с выходным фланцем. Поверхности уплотнительных элементов, близлежащие к соответствующим им дозвуковым створкам, выполнены с возможностью контакта с последними и образованы двумя коническими поверхностями, оси конусов которых совпадают с осью вращения соответствующей дозвуковой створки. Каждая из близлежащих поверхностей дозвуковых и сверхзвуковых створок, расположенных вдоль осей их шарнирных соединений, выполнены коническими. Технический результат заключается в увеличении жесткости элементов сопла. 2 ил. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 771 587 C1 (51) МПК F02K 1/12 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F02K 1/12 (2022.02); F02K 1/006 (2022.02); F02K 1/1269 (2022.02) (21)(22) Заявка: 2021121508, 20.07.2021 (24) Дата начала отсчета срока действия патента: Дата регистрации: 06.05.2022 (45) Опубликовано: 06.05.2022 Бюл. № 13 2 7 7 1 5 8 7 R U (54) Регулируемое сопло турбореактивного двигателя (57) Реферат: Изобретение относится к области авиационного двигателестроения, а именно к конструкции регулируемых сопел турбореактивных двигателей. Регулируемое сопло турбореактивного двигателя ...

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10-06-2009 дата публикации

Nozzle shutter for aviation gas-turbine engine (versions) and method for manufacture of nozzle shutter (versions)

Номер: RU2358136C2
Принадлежит: Снекма Мотер

Створка сопла для авиационного газотурбинного двигателя выполнена в виде корпуса, имеющего форму полого усеченного конуса, приплюснутого в поперечном направлении, по прямолинейным геометрическим образующим. Корпус содержит тонкую стенку постоянной толщины, поделенную в поперечном направлении на трапециевидную горячую стенку, трапециевидную холодную стенку, параллельную горячей стенке, и две симметричные боковые стенки, соединяющие по бокам холодную стенку и горячую стенку. Тонкая стенка содержит сплошную внутреннюю поверхность, имеющую закругленные участки с радиусом кривизны, по меньшей мере, равным двум толщинам тонкой стенки. Тонкая стенка выполнена из огнеупорного композитного материала, образованного усилительными сплошными и пересекающимися друг с другом волокнами из огнеупорного материала, погруженными в матрицу из огнеупорного материала. В другом варианте трапециевидные стенки имеют толщину, увеличивающуюся от боковых стенок к центральной полосе. В этом варианте створка содержит ребра жесткости, параллельные боковым стенкам, соединяющие между собой трапециевидные стенки и сопрягающиеся с ними закругленными участками с радиусом кривизны по внутренней поверхности, по меньшей мере, равным двойной толщине боковых стенок. Способ изготовления створки сопла содержит этапы создания одной структуры со сплошными краями из пересекающихся огнеупорных волокон, натяжения структуры вокруг жестких, пористых противоположных друг другу изогнутых профилей и нанесения огнеупорной матрицы путем осаждения в парообразной фазе на структуру и через профили. Противоположные друг другу изогнутые профил� РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) 2 358 136 (13) C2 (51) МПК F02K 1/12 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ, ПАТЕНТАМ И ТОВАРНЫМ ЗНАКАМ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (21), (22) Заявка: 2004116014/06, 25.05.2004 (24) Дата начала отсчета срока действия патента: 25.05.2004 (73) Патентообладатель(и): СНЕКМА МОТЕР (FR) (43) Дата публикации заявки: 20.11.2005 ...

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30-04-2008 дата публикации

Thrust vectorable fan variable area nozzle for a gas turbine engine fan nacelle

Номер: EP1916405A2
Принадлежит: United Technologies Corp

A thrust vectorable fan variable area nozzle (FVAN) (28) includes a synchronizing ring (40), a static ring (42), and a flap assembly (44) mounted within a fan nacelle (32). An actuator assembly (48) selectively rotates synchronizing ring segments (40A ... 40D) relative the static ring (42) to adjust segments (44A ... 44D) of the flap assembly to vary the annular fan exit area and vector the thrust through asymmetrical movement of the thrust vectorable FVAN segments (28A ... 28D). In operation, adjustment of the entire periphery of the thrust vectorable FVAN (28) in which all segments (28A ... 28D) are moved simultaneously to maximize engine thrust and fuel economy during each flight regime. By separately adjusting the segments (28A ... 28D) of the thrust vectorable FVAN, engine trust is selectively vectored to provide, for example only, trim balance or thrust controlled maneuvering.

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23-06-1998 дата публикации

Aircraft thrust vectoring system

Номер: US5769317A
Принадлежит: Allison Engine Co Inc

A thrust vectoring system comprising a working fluid source mounted to an aircraft, and a vectoring nozzle connected to the working fluid source which has an extendable conduit defining a passage from an inlet to an outlet. The passage is adjustable so that the direction of the working fluid exiting the outlet is different than the working fluid entering the inlet when the conduit of the vectoring nozzle is extended. This conduit provides a way to change the direction of working fluid exiting the nozzle and correspondingly change the thrust vector of the aircraft. A directing member such as a guide vane or door is mounted to the conduit adjacent the outlet to further direct the working fluid exiting the outlet. Directing members provide a way to refine the thrust vector without changing the extension of the segmented conduit. Another aspect of the present invention is to provide an adapter to connect the vectoring nozzle to a working fluid source with an annular outlet. This adapter uses an asymmetric structure to minimize flow separation.

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31-08-2000 дата публикации

Thrust vectoring techniques

Номер: WO2000050759A2
Принадлежит: Allison Advanced Development Company

A thrust directing mechanism (430) to vector thrust and control discharge throat area is disclosed. Mechanism (430) includes a number of vanes (460a-460h) mounted across passage (418) of thrust vectoring nozzle (440). Mechanism (430) includes control linkage (486) pivotally coupled to each of vanes (460a-460h). The control linkage (486) is selectively movable to correspondingly pivot vanes (460a-460h) and has a link with at least two degrees of freedom corresponding to a two-coordinate position. A desired orientation of vanes (460a-460h) may be determined as a function of the two-coordinate position. The discharge exit area is contracted by adjusting convergence of vanes (460a-460h). During convergence, vanes (460a-460h) are pivoted to various pivot angles selected to optimize thrust efficiency when contracting the discharge throat area.

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11-01-2012 дата публикации

Thrust vectorable fan variable area nozzle for a gas turbine engine fan nacelle

Номер: EP1916405A3
Принадлежит: United Technologies Corp

A thrust vectorable fan variable area nozzle (FVAN) (28) includes a synchronizing ring (40), a static ring (42), and a flap assembly (44) mounted within a fan nacelle (32). An actuator assembly (48) selectively rotates synchronizing ring segments (40A ... 40D) relative the static ring (42) to adjust segments (44A ... 44D) of the flap assembly to vary the annular fan exit area and vector the thrust through asymmetrical movement of the thrust vectorable FVAN segments (28A ... 28D). In operation, adjustment of the entire periphery of the thrust vectorable FVAN (28) in which all segments (28A ... 28D) are moved simultaneously to maximize engine thrust and fuel economy during each flight regime. By separately adjusting the segments (28A ... 28D) of the thrust vectorable FVAN, engine trust is selectively vectored to provide, for example only, trim balance or thrust controlled maneuvering.

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12-06-2019 дата публикации

Method of varying an annular fan exit area of a gas turbine engine

Номер: EP3495650A1
Принадлежит: United Technologies Corp

A method of varying an annular fan exit area of a gas turbine engine (10) comprising the step of asymmetrically adjusting a first flap assembly segment (44A ... 44D) relative a second flap assembly segment to adjust an annular fan exit throat area.

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06-08-2019 дата публикации

Turbojet engine rotary axisymmetric nozzle

Номер: RU2696833C1

FIELD: aviation. SUBSTANCE: invention relates to the aviation equipment turbojet engines, in particular to the jet nozzles design. Rotary axisymmetric nozzle of turbojet engine includes fixed housing, movable housing, control hydraulic cylinders, as well as pneumatic cylinders. Fixed housing is made with two additional supports on its outer surface side, and movable housing is located between them and is pivotally connected to fixed housing in two diametrically opposite centres pins. Each pin is fixed in radial holes of fixed housing and additional support. Control hydraulic cylinders are hinged on one side on fixed housing, and on other side - on movable housing. Pneumatic cylinders are hinged on one side on stationary housing, and on the other - on the movable and arranged in pairs in the pins installation area so that the pneumatic cylinders of each pair are located on opposite sides from the pin of the pin. In rodless cavity of pneumatic cylinders near places of their fixation holes are made, and rod cavity of pneumatic cylinders is connected to high pressure compressor with possibility to create force P c by pneumatic cylinders, opposite in direction to axial force P af gas flow acting on moving nozzle housing. EFFECT: invention allows increasing service life of rotary axisymmetric nozzle due to increased serviceability of hinged connection of fixed housing with movable housing under conditions of large axial loads and with large number of nozzle turning cycles. 1 cl, 3 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 696 833 C1 (51) МПК F02K 1/04 (2006.01) F02K 1/78 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F02K 1/04 (2019.02); F02K 1/006 (2019.02); F02K 1/78 (2019.02) (21)(22) Заявка: 2018115740, 26.04.2018 (24) Дата начала отсчета срока действия патента: Дата регистрации: 06.08.2019 (45) Опубликовано: 06.08.2019 Бюл. № 22 2 6 9 6 8 3 3 R U (56) Список документов, цитированных в отчете о поиске: RU ...

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25-07-2007 дата публикации

Device for preventing ball and socket joint from rolling

Номер: KR100739468B1
Автор: 이상연, 이환규
Принадлежит: 국방과학연구소

본 발명은 볼 소켓 조인트의 롤 회전 방지장치에 관한 것으로서, 고정부인 소켓에 대한 운동부인 볼에 결합되는 운동체의 전방향(Omnidirectional) 경사회전 운동을 방해하지 않으면서, 볼에 결합되는 운동체의 롤 회전운동만을 효과적으로 억제하는 것을 목적으로 한다. The present invention relates to a roll rotation preventing device of a ball socket joint, wherein the roll rotation of the moving object coupled to the ball without interfering with the omnidirectional oblique rotational movement of the moving body coupled to the ball as the moving part relative to the socket which is a fixed part It aims to suppress only exercise effectively. 본 발명의 롤 회전 방지장치는, 고정부(100)와, 상기 고정부에 설치된 소켓(Socket, 122)에 임의의 방향으로 회전가능하게 삽입된 볼(Ball, 124)에 결합되는 운동체(110)를 서로 연결하여, 상기 운동체의 전방향 경사회전 운동을 방해하지 않으면서 롤 회전운동만을 억제하도록 이루어진다. 즉, 본 발명의 롤 회전 방지장치는, 상기 운동체의 외주면에 설치되는 운동부 브래킷(130)과; 상기 운동부 브래킷에 상단 일측이 피봇팅(Pivoting)되는 가이드 채널(Guide Channel, 140)과; 상기 가이드 채널에 슬라이딩 가능하게 결합되는 슬라이딩 바(Sliding Bar, 150)와; 그리고, 상기 고정부에 설치되어, 상기 슬라이딩 바가 상기 가이드 채널을 따라 슬라이딩 이동하면서 좌우로 기울어지는 거동을 하도록, 상기 슬라이딩 바의 하단과 결합되는 고정부 브래킷(160);을 포함하여 이루어진다. Roll anti-rotation device of the present invention, the moving part 110 is coupled to the fixed portion 100, the ball (124) rotatably inserted in any direction to the socket (Socket, 122) installed in the fixed portion. By connecting to each other, it is made to suppress only the roll rotation movement without disturbing the forward oblique rotation movement of the moving body. That is, the roll rotation preventing device of the present invention, the movement unit bracket 130 is installed on the outer peripheral surface of the moving body; A guide channel 140 on which one side of the upper part is pivoted on the moving part bracket; A sliding bar slidably coupled to the guide channel; And a fixing part bracket 160 installed at the fixing part and coupled to the lower end of the sliding bar so that the sliding bar slides along the guide channel and tilts left and right. 볼소켓, 조인트, 전방향, 경사, 회전, 벡터링, 롤회전, 방지  Ball Socket, Joint, Omni-directional, Inclined, Rotate, Vectoring, Roll Rotation, ...

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13-10-2010 дата публикации

Exhaust plenum for a turbine engine

Номер: EP2239428A2
Принадлежит: General Electric Co

An exhaust system for gas turbine engine (12) is provided that reduces turbulence and backflow within the exhaust system and, thus, increases the efficiency of the turbine engine (12). In various embodiments, the system includes an exhaust plenum (32) that provides a gradual expansion of the exhaust gases. The exhaust plenum (32) may also include one or more flow splitters (84, 130) that further reduce turbulence in the plenum or symmetrical arrangements and provide a more uniform gas flow in the plenum and other downstream exhaust components.

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10-06-2013 дата публикации

Jet moving body path control system

Номер: RU2484417C1
Принадлежит: Мбда Франс

Изобретение относится к реактивной технике и может быть использовано для управления траекторией реактивного движущегося тела. Реактивное движущееся тело содержит систему. Система содержит заднее основание с подвижным венцом, цилиндрическую пусковую трубу с зубчатым периферийным кольцом, неподвижное относительно заднего основания движущегося тела сопло с симметрией вращения и расширяющимся раструбом, один дефлектор потока с рычагом, расположенный снаружи сопла движущегося тела и шарнирно установленный на заднем основании, средства угловой ориентации дефлектора с приводом подъемного типа и приводом червячного типа, взаимодействующего с зубчатым периферийным кольцом и венцом заднего основания, элемент отклонения из жаростойкого материала. Длина элемента управления равна диаметру расширяющегося раструба. Изобретение позволяет уменьшить прогрессивное истирание рулевых поверхностей и массу системы. 2 н. и 7 з.п. ф-лы, 7 ил. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) 2 484 417 (13) C1 (51) МПК F42B 10/66 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (21)(22) Заявка: 2012100729/11, 08.06.2010 (24) Дата начала отсчета срока действия патента: 08.06.2010 (73) Патентообладатель(и): МБДА ФРАНС (FR) R U Приоритет(ы): (30) Конвенционный приоритет: 12.06.2009 FR 0902861 (72) Автор(ы): КАРТОН Лоран (FR) (45) Опубликовано: 10.06.2013 Бюл. № 16 2 4 8 4 4 1 7 (56) Список документов, цитированных в отчете о поиске: US 4562980 A1, 07.01.1986. US 2008/179449 A1, 31.07.2008. GB 1230760 A, 05.05.1971. RU 2276280 C1, 10.05.2006. Березиков В.В. и др. Конструкция управляемых баллистических ракет. - М.: Издательство "Воениздат", 1969, с.326-329. 2 4 8 4 4 1 7 R U (86) Заявка PCT: FR 2010/000417 (08.06.2010) C 1 C 1 (85) Дата начала рассмотрения заявки PCT на национальной фазе: 12.01.2012 (87) Публикация заявки РСТ: WO 2010/142867 (16.12.2010) Адрес для переписки: 129090, Москва, ул. Б. Спасская, 25, стр.3, ООО "Юридическая фирма Городисский и Партнеры ...

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29-09-1965 дата публикации

Improvements in or relating to gas turbine propulsion units

Номер: GB1006205A
Принадлежит: Rolls Royce PLC

1,006,205. Gas turbine jet engines. ROLLSROYCE Ltd. Nov. 1, 1963 [Oct. 5, 1963], No. 39309/63. Heading FIJ. [Also in Division B7] A gas turbine jet propulsion engine comprises in flow series a compressor, combustion equipment and a turbine driving the compressor and one or more orientable propulsion nozzles which receive the exhaust gases from the turbine and a part of the air compressed by the compressor, this air having by-passed the combustion equipment and turbine. In the two-shaft by-pass engine shown, two nozzles 28 are provided and are rotatable about an axis 30 by an air motor (Fig. 2, not shown). The cold flow passes into a manifold 20 encircling the engine through an annular duct 18 and the hot flow passes into the manifold through a chamber 25 and a duct 22, guide vanes 21, 27 imparting a circumferential velocity component to the gases. In a modification (Fig. 4, not shown) the vanes 21 are dispensed with and the vanes 27 are replaced by a series of nozzles. Alternatively (Figs. 5 and 6, not shown) the vanes 27 may be dispensed with and the vanes 21 replaced by a perforated plate. The engine may be of the single shaft type with the cold flow provided by a fan, the blades of which are mounted on the tips of the compressor blades. The engine thrust may be increased by burning fuel in the duct 18 or in the duct 22. The invention is also applicable to engines with centrifugal compressors. The engine is fitted in the fuselage of an aircraft with the axis 30 passing through the centre of gravity and two vertically disposed lift engines may also be provided, one ahead of and one behind the unit.

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15-06-1984 дата публикации

TUYERE A VARIABLE SECTION FOR A TURBOJET ENGINE AND AIRCRAFT COMPRISING SUCH AN EQUIPMENT

Номер: FR2537657A1
Принадлежит: SNECMA SAS

L'INVENTION CONCERNE UNE TUYERE A SECTION VARIABLE POUR UN TURBOREACTEUR. CETTE TUYERE COMPORTE UN CONVERGENT, FORME PAR UNE DEMI-COQUILLE FIXE 2 ET UNE DEMI-COQUILLE 3 PIVOTANTE AUTOUR D'UN AXE DIAMETRAL 4A-4B; L'ETANCHEITE DU CONVERGENT EST ASSUREE PAR UN JOINT SENSIBLEMENT SEMI-CIRCULAIRE 6 RACCORDE A DEUX JOINTS LONGITUDINAUX 7A, 7B. CETTE TUYERE EST UTILISABLE EN PARTICULIER AVEC UN REACTEUR COMPORTANT UN SYSTEME DE RECHAUFFE PARTIELLE A TAUX MODERE. THE INVENTION RELATES TO A TUBE WITH A VARIABLE SECTION FOR A TURBOREACTOR. THIS TUBE INCLUDES A CONVERGENT, SHAPED BY A FIXED HALF-SHELL 2 AND A HALF-SHELL 3 SWIVEL AROUND A DIAMETRAL AXIS 4A-4B; THE TIGHTNESS OF THE CONVERGENT IS PROVIDED BY A SENSITIVELY SEMI-CIRCULAR SEAL 6 CONNECTED TO TWO LONGITUDINAL SEALS 7A, 7B. THIS TUBE CAN BE USED IN PARTICULAR WITH A REACTOR INCLUDING A PARTIAL HEATING SYSTEM AT A MODERATE RATE.

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19-04-1991 дата публикации

ACTUATING DEVICE FOR THE PLACEMENT OF AN EXHAUST PIPE THAT CAN BE VECTORIZED.

Номер: FR2653176A1
Принадлежит: General Electric Co

On décrit un dispositif simplifié d'actionnement pour mettre en place les volets divergents (16, 18) d'une tuyère d'échappement vectorisable à deux dimensions (10). Le dispositif comprend des bras de levier (31) présentant des surfaces à came (32; 34) fixées en pivotement aux arbres de manivelle (21, 23) des volets divergents. Un vérin (42) est fixé à la paroi latérale de la tuyère et reliée fonctionnellement à une barre de glissière (36; 38) comportant deux contre-cames (40) à chaque extrémité, lesquelles sont fixes l'une par rapport à l'autre, et viennent en contact avec les surfaces à came des bras de levier. La barre de glissière coulisse sur la paroi latérale de la tuyère de manière à permettre un déplacement radial par rapport à l'axe (108) de la tuyère. Lors du mouvement des volets convergents, les bras de levier se déplacent dans le sens axial et les contre-cames, en coopération avec les surfaces à came des bras de levier, programment l'angle relatif entre les volets divergents et par conséquent la surface de sortie de la tuyère. La vectorisation est obtenue par actionnement des vérins, ce qui a pour effet de déplacer les contre-cames dans la direction radiale, vers le haut ou vers le bas, d'où la rotation des volets divergents qui se déplacent à l'unisson vers le haut ou vers le bas. Application aux moteurs à turbine à gaz. A simplified actuation device for positioning the divergent flaps (16, 18) of a vectorizable two-dimensional exhaust nozzle (10) is described. The device includes lever arms (31) having cam surfaces (32; 34) pivotally attached to the crank shafts (21, 23) of the diverging flaps. A cylinder (42) is attached to the side wall of the nozzle and operatively connected to a slide bar (36; 38) having two cams (40) at each end, which are fixed with respect to one another. other, and come into contact with the cam surfaces of the lever arms. The slide bar slides on the side wall of the nozzle so as to allow ...

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05-01-2007 дата публикации

TUYERE WITH PUSH ORIENTATION

Номер: FR2887929A1

La présent invention porte sur une tuyère à orientation de poussée conformée de manière à diviser un flux principal de gaz de propulsion issu d'au moins un générateur de gaz en un premier et un second flux pour une éjection dans une première et une seconde demi tuyères (14, 16) et comportant au moins l'un des deux moyens de pilotage suivants : moyen de pilotage de la répartition du flux principal dans chacune des deux demi tuyères et moyen d'orientation du vecteur de poussée produite par chacune des deux demi tuyères.L'invention s'applique en particulier au pilotage en lacet d'un aéronef sans empennage vertical. The present invention relates to a thrust-oriented nozzle shaped to divide a main stream of propellant gas from at least one gas generator into first and second streams for ejection in first and second half-nozzles. (14, 16) and comprising at least one of the following two control means: means for controlling the distribution of the main flow in each of the two half-nozzles and means for orienting the thrust vector produced by each of the two half-nozzles The invention applies in particular to the yaw control of an aircraft without vertical tail.

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17-05-1991 дата публикации

AXISYMMETRICAL NOZZLE TURBOREACTOR EJECTION ASSEMBLY WITH VARIABLE SECTION AND PUSH-THROUGH

Номер: FR2643947B1
Принадлежит: SNECMA SAS

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06-07-2012 дата публикации

TUYERE WITH ORIENTABLE COLLAR

Номер: FR2957385B1
Принадлежит: SNECMA Propulsion Solide SA

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03-12-2004 дата публикации

TUYERE SHUTTER WITH INCREASED LIFETIME FOR AIRPLANE TURBOMOTORS.

Номер: FR2855557A1
Принадлежит: SNECMA Moteurs SA

L'invention propose un volet de tuyère à durée de vie augmentée pour turbomoteur d'avion. Un tel volet est remarquable qu'il est constitué d'un corps (10) tronconique creux et transversalement aplati suivant des génératrices géométriques (11) rectilignes, le corps (10) formant une paroi (12) mince d'épaisseur sensiblement constante E, la paroi (12) comportant une surface intérieure continue avec un rayon de courbure au moins égal à 2xE, la paroi (12) étant en matériau composite réfractaire constitué de fibres de renfort (18) en matériau réfractaire noyées dans une matrice en matériau également réfractaire, les fibres de renfort (18) étant disposées en une pluralité de nappes de fibres (18) continues empilées les unes sur les autres, les fibres (18a) d'une nappe croisant les fibres (18b) de toute nappe qui lui est adjacente. The invention provides an extended life nozzle flap for an aircraft turbine engine. Such a shutter is remarkable that it consists of a hollow frustoconical body (10) transversely flattened along straight geometric generatrices (11), the body (10) forming a thin wall (12) of substantially constant thickness E, the wall (12) having a continuous interior surface with a radius of curvature at least equal to 2xE, the wall (12) being of refractory composite material consisting of reinforcing fibers (18) of refractory material embedded in a matrix of also refractory material , the reinforcing fibers (18) being arranged in a plurality of plies of continuous fibers (18) stacked one on the other, the fibers (18a) of a ply crossing the fibers (18b) of any ply which is adjacent to it .

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24-03-1962 дата публикации

Gas turbo engine

Номер: FR1288601A
Автор:
Принадлежит: Rolls Royce PLC

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05-05-1967 дата публикации

Jet Ejector Exhaust Nozzle

Номер: FR1479351A
Автор:
Принадлежит: United Aircraft Corp

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03-08-1990 дата публикации

ADJUSTABLE EXHAUST NOZZLE FOR A GAS TURBOMACHINE

Номер: FR2569232B1
Принадлежит: Rolls Royce PLC

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26-11-1976 дата публикации

MANEUVERABLE IN-FLIGHT EJECTION TUBE FOR GAS TURBINE ENGINE

Номер: FR2309728A1
Автор: Dudley Owen Nash
Принадлежит: General Electric Co

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07-04-1995 дата публикации

Method of obtaining thrust by a jet of fluid and thrust nozzle for its implementation.

Номер: FR2710691A1

La tuyère de poussée comprend un carter, un corps central disposé à l'intérieur du carter, une enveloppe entourant le corps central en formant avec ce dernier un passage annulaire ainsi que des lèvres de sortie de l'enveloppe délimitant avec la base aval du corps centrai dans le passage annulaire une partie à section contractée en convergent-divergent d'orientation radiale et à sortie axiale. La base aval (2) du corps central (1) est plane et est montée de manière à pouvoir être déviée par rapport à sa position perpendiculaire à l'axe longitudinal de la tuyère, l'enveloppe (5) étant dotée de lèvres de sortie profilées (6), susceptibles d'être déplacées axialement par des vérins (9) pour faire varier l'aire de la section contractée (7) de la tuyère. Application à la propulsion des aéronefs à grande maniabilité. The thrust nozzle comprises a casing, a central body arranged inside the casing, a casing surrounding the central body, forming with the latter an annular passage as well as outlet lips of the casing delimiting with the downstream base of the body. centered in the annular passage a portion with a section contracted in convergent-divergent orientation radially and axially outlet. The downstream base (2) of the central body (1) is flat and is mounted so as to be able to be deviated from its position perpendicular to the longitudinal axis of the nozzle, the casing (5) being provided with outlet lips sections (6), capable of being displaced axially by jacks (9) to vary the area of the contracted section (7) of the nozzle. Application to the propulsion of highly maneuverable aircraft.

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13-05-2014 дата публикации

Thrust vectoring nozzle

Номер: CA2551343C
Принадлежит: SNECMA SAS

La présent invention porte sur une tuyère à orientation de poussée conformée de manière à diviser un flux principal de gaz de propulsion issu d' au moins un générateur de gaz en un premier et un second flux pour une éjection dans une première et une seconde demi tuyères (14, 16) et comportant au moins l'un des deux moyens de pilotage suivants : moyen de pilotage de la répartition du flux principal dans chacune des deux demi tuyères et moyen d'orientation du vecteur de poussée produite par chacune des deux demi tuyères. L'invention s'applique en particulier au pilotage en lacet d'un aéronef sans empennage vertical.

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08-03-1985 дата публикации

Patent FR2537657B1

Номер: FR2537657B1
Автор: [UNK]
Принадлежит: SNECMA SAS

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12-11-1965 дата публикации

Jet engine nozzle

Номер: FR1417902A
Автор: William Shaw
Принадлежит: Rolls Royce PLC

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18-09-2015 дата публикации

SYSTEM FOR CONTROLLING THE TRACK OF A MU MOBILE BY REACTION.

Номер: FR2946741B1
Автор: Laurent Carton
Принадлежит: MBDA France SAS

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02-07-1993 дата публикации

EXHAUST NOZZLE WITH FULL TRANSITION AND CONVERGENT SECTION

Номер: FR2651020B1
Принадлежит: General Electric Co

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04-08-2000 дата публикации

ADJUSTABLE NOZZLE ACTIVATION SYSTEM FOR A REACTION PROPELLER USING MULTIPLE CIRCUMFERENTIALLY DISTRIBUTED ELASTIC ASSEMBLIES

Номер: FR2789122A1

The invention concerns a thrust-vectoring nozzle for a jet propulsion system comprising a fixed part (1) designed to be fixed to the booster motor, a mobile part (2) articulated on the fixed part, control means (3) for operating the mobile part, and one or several elastic assemblies (4) arranged between the fixed part and the mobile part. Said nozzle is characterised in that each elastic assembly comprises at least a fixed elastic block (40) whereof one first end is integral with the fixed part and a second end opposite the first one, is integral with a rigid part (42), and at least a mobile elastic block (41), whereof one first end is integral with the mobile part and a second end, opposite the first end, is integral with said rigid part (42). Each mobile elastic block (41) is further offset relatively to each fixed elastic block (40) in a circumferential direction of the nozzle.

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07-09-1990 дата публикации

AXISYMMETRICAL NOZZLE TURBOREACTOR EJECTION ASSEMBLY WITH VARIABLE SECTION AND PUSH-THROUGH

Номер: FR2643947A1
Принадлежит: SNECMA SAS

Un ensemble d'éjection de turboréacteur est composé d'un canal annulaire 1 et d'une tuyère axisymétrique comportant deux ensembles annulaires de volets mobiles 3, 5 reliés par des biellettes 6 et commandés par des vérins 4 et un corps central 7 associé à un dispositif de commande 15 à 19 et constitué de deux éléments creux 8, 9 de révolution coaxiaux, reliés par des bras 11, ledit corps central 7 pivotant étant fixé sur le canal 1 au moyen de supports 13 associés à des articulations à rotule 12. A turbojet ejection assembly is composed of an annular channel 1 and an axisymmetric nozzle comprising two annular assemblies of movable flaps 3, 5 connected by connecting rods 6 and controlled by jacks 4 and a central body 7 associated with a control device 15 to 19 and consisting of two hollow elements 8, 9 of coaxial revolution, connected by arms 11, said central pivoting body 7 being fixed on channel 1 by means of supports 13 associated with ball joints 12.

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08-03-1968 дата публикации

Improvements to exhaust pipes

Номер: FR1516320A
Принадлежит: Rolls Royce PLC

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20-08-1982 дата публикации

Patent FR2309727B1

Номер: FR2309727B1
Автор: [UNK]
Принадлежит: General Electric Co

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01-10-1971 дата публикации

Patent FR2073455A7

Номер: FR2073455A7
Автор: [UNK]
Принадлежит: Rolls Royce PLC

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04-12-1981 дата публикации

EJECTION TUBE FOR GAS TURBINE ENGINE

Номер: FR2483523A1
Автор: Mieczyslaw Konarski
Принадлежит: General Electric Co

TUYERE D'EJECTION ORIENTABLE, PERFECTIONNEE. ELLE COMPREND UNE STRUCTURE FIXE 20, 21; PLUSIEURS VOLETS PRIMAIRES 30, 31 AYANT CHACUN UNE PARTIE RELIEE DE FACON PIVOTANTE A LA STRUCTURE FIXE 20, 21 ET AYANT UNE SURFACE INTERIEURE CONVEXE VERS L'AVAL; PLUSIEURS VOLETS SECONDAIRES 32, 33, SITUEES EN AVAL DES VOLETS PRIMAIRES 30, 31, CHACUN DES VOLETS SECONDAIRES AYANT UNE PARTIE RELIEE DE FACON PIVOTANTE A UNE EXTREMITE AVAL D'UN DES VOLETS PRIMAIRES 30, 31; ET UN DISPOSITIF 34 QUI FAIT PIVOTER ET AMENE A UNE POSITION DETERMINEE LES VOLETS PRIMAIRES 30, 31 ET SECONDAIRES 32, 33. APPLICATION AUX MOTEURS A TURBINE A GAZ. ADJUSTABLE, PERFECTED EJECTION TUBE. IT INCLUDES A FIXED STRUCTURE 20, 21; SEVERAL PRIMARY SHUTTERS 30, 31 EACH HAVING A PART SWIVELY CONNECTED TO THE FIXED STRUCTURE 20, 21 AND HAVING AN INTERIOR SURFACE CONVEX TOWARDS DOWNSTREAM; SEVERAL SECONDARY SHUTTERS 32, 33, LOCATED DOWNSTREAM OF THE PRIMARY SHUTTERS 30, 31, EACH OF THE SECONDARY SHUTTERS HAVING A PART PIVOTALLY CONNECTED TO A DOWNSTREAM END OF ONE OF THE PRIMARY SHUTTERS 30, 31; AND A DEVICE 34 WHICH PIVOTS AND TAKES THE PRIMARY 30, 31 AND SECONDARY 32, 33 SHUTTERS TO A DETERMINED POSITION. APPLICATION TO GAS TURBINE ENGINES.

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27-02-1981 дата публикации

EJECTION TUBE WITH EJECTABLE GAS STREAM

Номер: FR2463859A1
Принадлежит: General Electric Co

TUYERE A COL A POSITION VARIABLE EN MODE DE FONCTIONNEMENT AVEC ORIENTATION DU VECTEUR DE POUSSEE. ELLE COMPREND UN PREMIER DISPOSITIF ARTICULE DE DEFLEXION DU FLUX EJECTE COMPRENANT UN VOLET 36 D'ORIENTATION DU VECTEUR DE POUSSEE, QUI PIVOTE AUTOUR D'UNE PARTIE AVAL D'UN CANAL FIXE 22 EN UN ENDROIT 37 DE CE CANAL FIXE DONT LA SECTION INTERNE 26 EST GENERALEMENT INFERIEURE; UN SECOND DISPOSITIF ARTICULE DE DEFLEXION DU FLUX EJECTE COMPRENANT UN VOLET DE REGLAGE 38 A CALAGE VARIABLE RELIE FONCTIONNELLEMENT A UNE PARTIE DU CANAL FIXE 22 ET COOPERANT AVEC LE PREMIER DISPOSITIF DE DEFLEXION 36 POUR FORMER ENTRE EUX UN COL A POSITION VARIABLE, CE COL SE TROUVANT EN AVAL DE L'ENDROIT OU LES GAZ EJECTES CHANGENT D'ORIENTATION QUAND LE VOLET 36 D'ORIENTATION DU VECTEUR DE POUSSEE EST DANS UN MODE DE DEVIATION. APPLICATION AUX MOTEURS A TURBINE A GAZ D'AVION. VARIABLE POSITION NECK TUBE IN OPERATING MODE WITH PUSH VECTOR ORIENTATION. IT INCLUDES A FIRST ARTICULATED DEVICE FOR DEFLECTING THE EJECTED FLOW INCLUDING A THREAD VECTOR ORIENTATION FLAP 36, WHICH PIVOTS AROUND A DOWNSTREAM PART OF A FIXED CHANNEL 22 IN A LOCATION 37 OF THIS FIXED CHANNEL INCLUDING THE INTERNAL SECTION 26 IS GENERALLY LOWER; A SECOND ARTICULATED DEVICE FOR DEFLECTION OF THE EJECTED FLOW INCLUDING AN ADJUSTING FLAP 38 WITH VARIABLE TIMING FUNCTIONALLY CONNECTED TO A PART OF THE FIXED CHANNEL 22 AND COOPERATING WITH THE FIRST DEFLECTION DEVICE 36 TO FORM BETWEEN THEM A VARIABLE POSITION NECK, THIS NECK LOCATED DOWNSTREAM FROM THE AREA WHERE THE EJECTED GASES CHANGE ORIENTATION WHEN THE THRUST VECTOR ORIENTATION FLAP 36 IS IN A DEVIATION MODE. APPLICATION TO AIRPLANE GAS TURBINE ENGINES.

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21-02-1986 дата публикации

ADJUSTABLE EXHAUST NOZZLE FOR A GAS TURBOMACHINE

Номер: FR2569232A1
Принадлежит: Rolls Royce PLC

UNE TUYERE ORIENTABLE COMPREND UN PREMIER CONDUIT FIXE21, UN SECOND CONDUIT ROTATIF22 TRONCONNE EN OBLIQUE A SON EXTREMITE AVAL ET UN TROISIEME CONDUIT ROTATIF23 TRONCONNE EN OBLIQUE A SON EXTREMITE AMONT. LES SECOND ET TROISIEME CONDUITS22, 23 SONT MONTES DANS DES PALIERS RESPECTIFS24, 26 ET IL EST PREVU UN TRAIN D'ENGRENAGES EPICYCLOIDAL30 POUR LES ENTRAINER EN ROTATION EN SENS OPPOSE TOUT EN MAINTENANT LA DIRECTION DES GAZ D'EJECTION PROVENANT DE LA TUYERE DANS UN PLAN UNIQUE.

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27-10-2006 дата публикации

DIRECTIONABLE EJECTION TUYERE OF AN AIRCRAFT ENGINE

Номер: FR2884874A1
Принадлежит: SNECMA Moteurs SA

Tuyère orientable pour un moteur d'aéronef, la tuyère (10) étant montée pivotante autour d'un axe transversal de lacet au moyen de deux articulations (24) monobloc élastiquement déformables montées sur deux pivots externes (22) de la tuyère et logées dans des éléments de carter du moteur, le pivotement de la tuyère étant assuré par un vérin (46) monté sur le carter du moteur et relié par une biellette (50) à un pivot (22) de la tuyère. Steerable nozzle for an aircraft engine, the nozzle (10) being pivotally mounted about a transverse yaw axis by means of two elastically deformable one-piece hinges (24) mounted on two external pivots (22) of the nozzle and housed in engine casing elements, the pivoting of the nozzle being provided by a cylinder (46) mounted on the motor housing and connected by a rod (50) to a pivot (22) of the nozzle.

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26-09-1975 дата публикации

Patent FR2262735A1

Номер: FR2262735A1
Автор: [UNK]

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15-06-2007 дата публикации

DIRECTIONABLE EJECTION TUYERE OF AN AIRCRAFT ENGINE

Номер: FR2884874B1
Принадлежит: SNECMA Moteurs SA

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02-04-1976 дата публикации

IMPROVEMENTS TO TURBOPROPELLER UNITS

Номер: FR2284042A1
Автор: [UNK]
Принадлежит: Rolls Royce PLC

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11-09-1964 дата публикации

Oscillating Discharge Cone Nozzle for Gas Turbine Jet Engine

Номер: FR1372176A
Автор: James Oswald Mortlock
Принадлежит: Rolls Royce PLC

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22-02-1991 дата публикации

EXHAUST TUBE WITH INTEGRAL TRANSITION AND CONVERGENT SECTION

Номер: FR2651020A1
Принадлежит: General Electric Co

Une tuyère d'échappement à fonctions multiples, de faible poids, comporte un carter de transition 30 de courte longueur, au contour particulier, et deux volets convergents 20 au contour correspondant. Le changement de la section transversale du trajet d'écoulement pour passer d'une section de forme circulaire à une section de forme rectangulaire se produit en partie dans le carter de transition et en partie entre les deux volets de sorte que la section de transition de la tuyère est intégrée dans la section convergente. Les volets sont montés en pivotement dans la tuyère de sorte que les gaz d'échappement qui les frappent produisent des moments d'opposition autour de l'axe de pivotement de l'arbre du volet, d'où la réduction de la charge à laquelle les actionneurs des volets sont soumis. Application aux turboréacteurs. A multi-function, low-weight exhaust nozzle has a short-length transition casing 30 with a particular contour and two converging flaps 20 with a corresponding contour. The change in the cross section of the flow path from a circular section to a rectangular section occurs partly in the transition housing and partly between the two flaps so that the transition section of the nozzle is integrated in the converging section. The flaps are pivotally mounted in the nozzle so that the exhaust gases hitting them produce opposing moments about the pivot axis of the flap shaft, hence the reduction of the load at which the shutter actuators are subject. Application to turbojets.

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24-06-1988 дата публикации

ADJUSTABLE TWO-DIMENSIONAL TUYERE, IN PARTICULAR FOR AIRCRAFT REACTOR

Номер: FR2608680A1
Принадлежит: SNECMA SAS

L'INVENTION EST RELATIVE A UNE TUYERE BIDIMENSIONNELLE REGLABLE, NOTAMMENT POUR REACTEUR D'AVION. LA TUYERE COMPREND DES PAROIS LATERALES FIXES 8 ET UN ENSEMBLE DE VOLETS MOBILES ACTIONNES PAR UN SYSTEME DE COMMANDE FORME DE QUATRE ENSEMBLES COMPORTANT CHACUN UN CADRE 17 MOBILE LONGITUDINALEMENT ET DE PREFERENCE CAPABLE DE PIVOTER AUTOUR D'UN AXE DU PLAN DE SYMETRIE DE LA TUYERE, UNE PIECE A CAME 26, COOPERANT AVEC UN GALET FIXE ET ARTICULE SUR UN VOLET AVAL 11 ET SUR LE CADRE 17, LE VOLET AMONT ETANT ARTICULE SUR LE CADRE ET SUR LE VOLET AVAL. LE CADRE, LA PIECE A CAME ET LES DEUX VOLETS FORMENT UN QUADRILATERE DEFORMABLE. THE INVENTION RELATES TO AN ADJUSTABLE TWO-DIMENSIONAL TUBE, IN PARTICULAR FOR AN AIRCRAFT REACTOR. THE TUBE CONSISTS OF FIXED SIDE WALLS 8 AND A SET OF MOBILE SHUTTERS ACTUATED BY A CONTROL SYSTEM IN THE FORM OF FOUR SETS EACH INCLUDING A LONGITUDINALLY MOBILE FRAME 17 AND PREFERABLY CAPABLE OF PIVOTING AROUND AN AXIS OF THE PLAN A CAM 26 PART, COOPERATING WITH A FIXED ROLLER AND ARTICULATED ON A DOWNSTREAM FLAP 11 AND ON FRAME 17, THE UPSTREAM FLAP BEING ARTICULATED ON THE FRAME AND ON THE DOWNSTREAM FLAP. THE FRAME, THE CAM PART AND THE TWO FLAPS FORM A DEFORMABLE QUADRILATER.

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04-03-1988 дата публикации

VARIABLE SURFACE EXHAUST NOZZLE FOR A GAS TURBINE ENGINE

Номер: FR2603342A1
Принадлежит: Rolls Royce PLC

BUSE D'ECHAPPEMENT A SURFACE VARIABLE POUR MOTEUR A TURBINE A GAZ, EN PARTICULIER UTILISE COMME GROUPE MOTO-PROPULSEUR D'HELICOPTERE COMPOSITE. ELLE COMPPRTE UN PREMIER PETALE FIXE 44 ET UN SECOND PETALE MOBILE 46, UN ACTIONNEUR 48 L'AMENANT EN UNE PREMIERE ET UNE SECONDE POSITION. EN LA PREMIERE POSITION, SON BORD AVANT 46A EST EN AVANT, ET RADIALEMENT LEGEREMENT A L'INTERIEUR DU BORD AVANT 50 DU TUBE D'ECHAPPEMENT 52, DEFINISSANT UN PETIT ESPACE D'EJECTION 54, AINSI QUE LA SURFACE D'ECHAPPEMENT DE LA BUSE 36. DANS LA SECONDE POSITION (EN TRAITS MIXTES), SON BORD AVANT 46A EST DEPLACE RADIALEMENT A L'INTERIEUR DE SON BORD ARRIERE 46B, DEFINISSANT LA SURFACE D'ECHAPPEMENT AVEC UN ESPACE D'EJECTION AUXILIAIRE 56. VARIABLE SURFACE EXHAUST NOZZLE FOR GAS TURBINE ENGINE, ESPECIALLY USED AS A COMPOSITE HELICOPTER MOTOR-PROPELLER. IT INCLUDES A FIRST FIXED PETAL 44 AND A SECOND MOBILE PETAL 46, AN ACTUATOR 48 TAKING IT INTO A FIRST AND A SECOND POSITION. IN THE FIRST POSITION, ITS FRONT EDGE 46A IS FORWARD, AND RADIALLY INTO THE FRONT EDGE 50 OF THE EXHAUST TUBE 52, DEFINING A SMALL EJECTION SPACE 54, AS WELL AS THE EXHAUST SURFACE OF THE NOZZLE 36. IN THE SECOND POSITION (IN MIXED LINES), ITS FRONT EDGE 46A IS MOVED RADIALLY WITHIN ITS REAR EDGE 46B, DEFINING THE EXHAUST SURFACE WITH AN AUXILIARY EJECTION SPACE 56.

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14-09-1962 дата публикации

Jet deflector device for the orientation and stabilization of a machine during the landing phase

Номер: FR1303806A
Автор:
Принадлежит: D AVIAT LATECOERE SOC IND

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19-01-1962 дата публикации

Improvements to jet propellants for aerial vehicles

Номер: FR1282330A
Автор:
Принадлежит: Bristol Siddeley Engines Ltd

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01-10-1965 дата публикации

Gas turbine engine

Номер: FR1412683A
Автор:
Принадлежит: Rolls Royce PLC

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17-12-2010 дата публикации

SYSTEM FOR CONTROLLING THE TRACK OF A MU MOBILE BY REACTION.

Номер: FR2946741A1
Автор: Laurent Carton
Принадлежит: MBDA France SAS

Selon l'invention, le système comporte au moins un déflecteur de flux (6), extérieur à la tuyère (3) du mobile (1) et articulé sur le fond arrière (2A) de ce dernier, et des moyens d'orientation angulaire du déflecteur (6) par rotation autour de l'axe longitudinal (X-X) de la tuyère (3). According to the invention, the system comprises at least one flow deflector (6), outside the nozzle (3) of the mobile (1) and articulated on the rear base (2A) of the latter, and angular orientation means of the deflector (6) by rotation around the longitudinal axis (XX) of the nozzle (3).

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08-04-1988 дата публикации

ADJUSTABLE EJECTION NOZZLE

Номер: FR2463859B1
Принадлежит: General Electric Co

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28-01-1994 дата публикации

VARIABLE SURFACE EXHAUST NOZZLE FOR A GAS TURBINE ENGINE

Номер: FR2603342B1
Принадлежит: Rolls Royce PLC

Подробнее
03-10-1990 дата публикации

Exhaust nozzle

Номер: GB9018245D0
Автор: [UNK]
Принадлежит: General Electric Co

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12-03-2010 дата публикации

DEVIATION SYSTEM FOR A GAS FLOW IN A TUYERE

Номер: FR2883604B1
Автор: Bruno Beutin, Jeremy Fert
Принадлежит: SNECMA Moteurs SA

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10-03-2005 дата публикации

Jet engine vectored thrust nozzle with elastic ring

Номер: RU2247851C2

FIELD: jet engines. SUBSTANCE: proposed nozzle contains stationary part designed for connection with engine, adjustable flaps and flap orientation means. The latter contain elastic ring whose first circumferential periphery is rigidly connected with end of stationary part, and second circumferential periphery is coupled with adjustable flaps, and control means making it possible to displace second circumferential periphery of elastic ring and containing first jacks. First circumferential periphery of elastic ring is formed by first ring holder used as stationary support for elastic ring and secured on stationary part. Second circumferential periphery of elastic ring is formed by second rigid holder which serves as movable support for elastic ring, being coupled with flaps, and to which free ends of first jacks are connected. Adjustable flaps are installed on end of stationary part and are hinge fastened to panels or tie-rods also secured on second rigid ring holder. EFFECT: provision of adjustable flap of simple design and low cost. 9 cl, 6 dwg ÐÎÑÑÈÉÑÊÀß ÔÅÄÅÐÀÖÈß (19) RU (51) ÌÏÊ 7 (11) (13) 2 247 851 C2 F 02 K 1/12 ÔÅÄÅÐÀËÜÍÀß ÑËÓÆÁÀ ÏÎ ÈÍÒÅËËÅÊÒÓÀËÜÍÎÉ ÑÎÁÑÒÂÅÍÍÎÑÒÈ, ÏÀÒÅÍÒÀÌ È ÒÎÂÀÐÍÛÌ ÇÍÀÊÀÌ (12) ÎÏÈÑÀÍÈÅ ÈÇÎÁÐÅÒÅÍÈß Ê ÏÀÒÅÍÒÓ (21), (22) Çà âêà: 2000125335/06, 28.01.2000 (72) Àâòîð(û): ÀÁÁÅ Ôðàíñóà (FR), ÊÀÌÈ Ïüåð (FR), ÀÁÀÐÓ Æîðæ (FR), ÒÓÀËÜ Ìèøåëü (FR) (24) Äàòà íà÷àëà äåéñòâè ïàòåíòà: 28.01.2000 (30) Ïðèîðèòåò: 29.01.1999 FR 99/01022 (73) Ïàòåíòîîáëàäàòåëü(ëè): Ñîñüåòå Íàñüîíàëü Ä`Ýòþä å äå Êîíñòðóêñüîí äå Ìîòîð Ä`Àâèàñüîí-Ñ.Í.Å.Ê.Ì.À. (FR) (45) Îïóáëèêîâàíî: 10.03.2005 Áþë. ¹ 7 2 2 4 7 8 5 1 (56) Ñïèñîê äîêóìåíòîâ, öèòèðîâàííûõ â îò÷åòå î ïîèñêå: FR 2470253 A1, 29.05.1981. US 3696999 À, 10.10.1972. US 3390899 À, 02.07.1968. US 3726480 À, 04.10.1973. US Í1381 Í, 06.12.1994. SU 1246671 A1, 10.01.1996. RU 1438330 Ñ, 30.11.1994. (85) Äàòà ïåðåâîäà çà âêè PCT íà íàöèîíàëüíóþ ôàçó: 30.10.2000 2 2 4 7 8 5 1 R U (87) Ïóáëèêàöè PCT: WO 00/45040 (03.08.2000) C 2 C 2 ...

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04-06-1957 дата публикации

Aircraft jet propulsion improvements

Номер: FR1137807A
Автор: Aubrey Douglas
Принадлежит:

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22-01-1985 дата публикации

Patent JPS602511B2

Номер: JPS602511B2
Принадлежит: General Electric Co

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11-04-1984 дата публикации

Variable geometry exhaust nozzle

Номер: GB2077360B
Автор:
Принадлежит: General Electric Co

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18-11-1997 дата публикации

Yaw and pitch thrust vectoring nozzle

Номер: US5687907A
Автор: Davis C. Holden
Принадлежит: United Technologies Corp

An exhaust nozzle (12) for providing both yaw and pitch thrust vectoring includes an upper flap assembly (18) and a lower flap assembly (20). Each flap assembly (18, 20) includes a forward stub flap (51, 53) with lateral hinges (32, 34) and a plurality of individual flap sections (52) extending longitudinally downstream and adjacently engaged for defining the upper and lower gas flow boundaries. The forward end of each flap section (52) is pivoted vertically (54) in the same plane as vertical sidewall hinges (48, 50). Yaw thrust vectoring is achieved by pivoting the sections (52) and sidewalls (22, 24), while pitch thrust vectoring is achieved by moving the flap assemblies (18, 20) about the respective forward hinges (32, 34).

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20-04-2010 дата публикации

Aircraft turbojet engine with pressure chamber

Номер: RU2386841C2
Принадлежит: Стэнли ЧАНГ

FIELD: engines and pumps. SUBSTANCE: proposed engine comprises air inlet, air compressor with supercharging, combustion chamber and rear nozzle. Pressure chamber is arranged between rear nozzle that creates forward thrust and combustion chamber. Aforesaid air compressor comprises inlet compressor, larger pressure compressor and smaller pressure compressor, both coupled with transmission front shaft. It additionally comprises front drive turbine and rear drive turbine coupled with transmission rear shaft. Gearbox is arranged between transmission front and rear shafts, gearbox front section accommodating nozzle. High-temperature gas forced into pressure chamber can effuse through rear nozzle to push aircraft forward or lower nozzle to lift or down the aircraft, or to effect whatever maneuver. EFFECT: reduced noise and higher efficiency. 6 cl, 9 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) 2 386 841 (13) C2 (51) МПК F02K 1/54 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ, ПАТЕНТАМ И ТОВАРНЫМ ЗНАКАМ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (21), (22) Заявка: 2007139807/06, 19.04.2006 (24) Дата начала отсчета срока действия патента: 19.04.2006 (72) Автор(ы): ЧАНГ Винсент (CN), ЧАНГ Стэнли (CN) (43) Дата публикации заявки: 10.06.2009 2 3 8 6 8 4 1 (45) Опубликовано: 20.04.2010 Бюл. № 11 (56) Список документов, цитированных в отчете о поиске: FR 2710109 А, 24.03.1995. SU 1099664 А, 27.12.2004. SU 1277688 A, 16.04.2004. SU 766118 A, 27.11.2005. US 3861139 A, 21.01.1975. EP 1398493 A, 17.03.2004. 2 3 8 6 8 4 1 R U (86) Заявка PCT: CN 2006/000730 (19.04.2006) C 2 C 2 (85) Дата перевода заявки PCT на национальную фазу: 30.11.2007 (87) Публикация PCT: WO 2006/116907 (09.11.2006) Адрес для переписки: 119034, Москва, Пречистенский пер., д.14, стр.1, 4-й этаж, Гоулингз Интернэшнл Инк., пат.пов. В.А.Клюкину, рег.№ 5 (54) АВИАЦИОННЫЙ ТУРБОРЕАКТИВНЫЙ ДВИГАТЕЛЬ С КАМЕРОЙ ДАВЛЕНИЯ (57) Реферат: Авиационный двигатель с камерой давления содержит впускное отверстие для воздуха, ...

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20-09-1973 дата публикации

Gas turbine hub

Номер: DE1781121B2
Принадлежит: UK Secretary of State for Defence

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27-09-2007 дата публикации

REJECTIVE SYSTEM FOR GAS FLOW IN A JET NOZZLE

Номер: RU2006109097A
Принадлежит: Снекма, Снекма (Fr)

ÐÎÑÑÈÉÑÊÀß ÔÅÄÅÐÀÖÈß RU (19) (11) 2006 109 097 (13) A (51) ÌÏÊ F02K 1/00 (2006.01) ÔÅÄÅÐÀËÜÍÀß ÑËÓÆÁÀ ÏÎ ÈÍÒÅËËÅÊÒÓÀËÜÍÎÉ ÑÎÁÑÒÂÅÍÍÎÑÒÈ, ÏÀÒÅÍÒÀÌ È ÒÎÂÀÐÍÛÌ ÇÍÀÊÀÌ (12) ÇÀßÂÊÀ ÍÀ ÈÇÎÁÐÅÒÅÍÈÅ (21), (22) Çà âêà: 2006109097/06, 22.03.2006 (71) Çà âèòåëü(è): ÑÍÅÊÌÀ (FR) (43) Äàòà ïóáëèêàöèè çà âêè: 27.09.2007 Áþë. ¹ 27 (72) Àâòîð(û): ÁÅÒÝÍ Áðþíî (FR), ÔÅÐ Æåðåìè (FR) R U Àäðåñ äë ïåðåïèñêè: 129010, Ìîñêâà, óë. Á.Ñïàññêà , 25, ñòð.3, ÎÎÎ "Þðèäè÷åñêà ôèðìà Ãîðîäèññêèé è Ïàðòíåðû", ïàò.ïîâ. Ã.Á. Åãîðîâîé, ðåã.¹ 513 2 0 0 6 1 0 9 0 9 7 R U Ñòðàíèöà: 1 RU A (57) Ôîðìóëà èçîáðåòåíè 1. Îòêëîí þùà ñèñòåìà äë ãàçîâîãî ïîòîêà â ðåàêòèâíîì ñîïëå ëåòàòåëüíîãî àïïàðàòà, ñîäåðæàùà ââîä ùåå ñðåäñòâî äë ââåäåíè ñæàòîãî ãàçà â çàäàííîì íàïðàâëåíèè â ãàçîâûé ïîòîê, òåêóùèé â ñîïëå, è óïðàâë þùåå ñðåäñòâî äë óïðàâëåíè ãàçîâûì ïîòîêîì, âûõîä ùèì èç ââîä ùåãî ñðåäñòâà; ïðè ýòîì ââîä ùåå ñðåäñòâî äë ââåäåíè ãàçà ñîäåðæèò ôèêñèðîâàííûå êîíñòðóêöèîííûå êîæóõè, êîòîðûå ïðîõîä ò âíóòðè ñîïëà ïåðïåíäèêóë ðíî ê ãàçîâîìó ïîòîêó, òåêóùåìó â ñîïëå, è â êîòîðûå ïîäàåòñ îòêëîí þùèé ãàç ÷åðåç îäèí èç èõ êîíöîâ, íàðóæíûé ïî îòíîøåíèþ ê ñîïëó; ïðè÷åì â áîêîâûõ ñòåíêàõ óïîì íóòûõ êîæóõîâ âûïîëíåíû ââîä ùèå ïðîðåçè, îðèåíòèðîâàííûå â çàäàííûõ íàïðàâëåíè õ îòíîñèòåëüíî íàïðàâëåíè òå÷åíè ãàçîâîãî ïîòîêà â ñîïëå; è ðåãóëèðóåìûå ñðåäñòâà äë ðåãóëèðóåìîãî çàêðûòè ýòèõ ââîä ùèõ ïðîðåçåé; ïðè÷åì ðåãóëèðóåìûå çàêðûâàþùèå ñðåäñòâà èìåþò ñòâîðêè, óñòàíîâëåííûå âíóòðè óïîì íóòûõ êîæóõîâ è íàïðàâë åìûå ñìåùåíèåì èõ âäîëü óïîì íóòûõ ïðîðåçåé; ïðè ýòîì óïîì íóòûå ñòâîðêè ñîåäèíåíû ñ èñïîëíèòåëüíûìè ñðåäñòâàìè, ðàñïîëîæåííûìè ñíàðóæè êîæóõîâ è ñîïëà. 2. Ñèñòåìà ïî ï.1, â êîòîðîé ñòâîðêè âûïîëíåíû ñ âîçìîæíîñòüþ ñìåùåíè ìåæäó çàêðûòûì ïîëîæåíèåì ââîä ùèõ ïðîðåçåé è äâóì êðàéíèìè ïîëîæåíè ìè äë îòêðûòè ïðîðåçåé, îðèåíòèðîâàííûõ â çàäàííîì íàïðàâëåíèè, è äë çàêðûòè ïðîðåçåé, îðèåíòèðîâàííûõ â äðóãîì íàïðàâëåíèè, è íàîáîðîò. 3. Ñèñòåìà ïî ï.1, â êîòîðîé èñïîëíèòåëüíûå ñðåäñòâà ñâ çàíû ñ îáùèì óïðàâë þùèì ñðåäñòâîì, íàïðèìåð, ...

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12-12-1967 дата публикации

Thrust directing means for aircraft

Номер: US3357645A
Принадлежит: General Electric Co

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24-09-1997 дата публикации

Flight control systems for aircraft

Номер: GB2282353B
Автор: Peter William Bishop
Принадлежит: Individual

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21-04-2010 дата публикации

Exhaust nozzle for thrust vectoring

Номер: GB2464330A
Автор: Leonard John Rodgers
Принадлежит: Rolls Royce PLC

An exhaust nozzle (10) for a gas turbine engine is described for vectoring a flow of exhaust gases issuing therefrom at an angle to a centre line (8) of the engine. The exhaust nozzle (10) comprises a plurality of radially outer (25) and radially inner flaps (26) which are circumferentially disposed around the engine. The radially outer (25) and radially inner flaps (26) alternately overlap each other to define a path through which the exhaust gases flow. The radially outer flaps (25) can be moved asymmetrically about the engine centre line (8). The radially inner flaps (26) are maintained in sealing contact with the outer flaps (25) by the flow of exhaust gases passing therethrough. The radially inner flaps (26) are divided into a number of triangular sections (29) which renders the flaps (26) torsionally flexible so that they can twist along their lengths to maintain sealing contact with adjacent radially outer flaps (25) when the outer flaps are moved asymmetrically.

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24-09-2013 дата публикации

Gas stream deflection system in a nozzle

Номер: CA2540259C
Автор: Bruno Beutin, Jeremy Fert
Принадлежит: SNECMA SAS

Système de déviation d'un flux de gaz dans une tuyère d'éjection d'un engin votant, comprenant des moyens d'injection de gaz sous pression dans une direction donnée dans le flux de gaz s'écoutant dans la tuyère et des moyens de commande du débit de gaz sortant des moyens d'injection, dans lequel les moyens d'injection de gaz comprennent des boîtiers structuraux fixes qui s'étendent à l'intérieur de la tuyère perpendiculairement au flux de gaz s'écoulant dans la tuyère et qui sont alimentés en gaz de déviation par une de leurs extrémités extérieures à la tuyère, ces boîtiers comportant dans leurs parois latérales des fentes d'injection de gaz orientées dans des directions données par rapport à la direction d'écoulement du flux de gaz dans la tuyère, et des moyens commandés d'obturation réglable de ces fentes d'injection.

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11-04-1969 дата публикации

Roller bearing, especially for engine deflector segments

Номер: FR93512E
Автор: Gerhard Kopp
Принадлежит: Entwicklungsring Sued GmbH

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16-09-2011 дата публикации

TUYERE WITH ORIENTABLE COLLAR

Номер: FR2957385A1
Принадлежит: SNECMA Propulsion Solide SA

Tuyère à section bidimensionnelle (20), ladite tuyère comprenant un bâti fixe (21) comportant un divergent (22) dans sa partie aval. La tuyère comprend en outre un col (200) disposé à l'intérieur du bâti fixe (21) en amont du divergent (22). Le col comprend au moins deux parois latérales (210, 220) montées de façon pivotante dans ledit bâti (21), lesdites parois étant reliées entre elles par deux longerons transversaux (230, 240), chaque longeron étant monté de façon coulissante sur chaque paroi (210; 220). A bidimensional section nozzle (20), said nozzle comprising a fixed frame (21) having a diverging portion (22) in its downstream portion. The nozzle further comprises a neck (200) disposed within the fixed frame (21) upstream of the diverging portion (22). The neck comprises at least two side walls (210, 220) pivotally mounted in said frame (21), said walls being interconnected by two transverse longitudinal members (230, 240), each spar being slidably mounted on each wall (210; 220).

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17-01-2001 дата публикации

System for activating an adjustable tube by means of an elastic ring for a thrust nozzle

Номер: EP1068438A1

An adjustable tube for a thrust nozzle comprising a fixed part (1) that is joined to the thrust nozzle, a plurality of adjustable flaps (21) mounted on one end of (11) of the fixed part and a means for adjusting said flaps (3, 4, 22, 31). The invention is characterized in that the means for adjusting the flaps (3) comprises an elastic ring (3), whereby the first circumference thereof is joined to the end (11) of the fixed part and a second circumference is joined to the flaps (21), in addition to control means (4) enabling the second circumference of the elastic ring to be displaced.

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08-09-1982 дата публикации

Orientable nozzle assemblies

Номер: GB2032861B
Автор:

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13-03-2020 дата публикации

Double-flow turbojet engine nozzles system

Номер: RU2716651C2

FIELD: machine building. SUBSTANCE: invention relates to aircraft engine building, particularly, to double-flow turbojet engines nozzles. System of nozzles for working gases of double-flow turbojet engine without mixing of circuits flows comprises two cold channels changing their cross-section from semi-ring behind engine fan to rectangle at outlet and one hot channel of round cross section located between cold channels. Channels terminate by nozzles, which are devices for control of value and direction of working gas jet. Invention makes it possible to fit nozzle into tail part of aircraft plane by flying wing or tailless and control thrust vector in cold channel, leaving hot channel in form of extension pipe of minimum weight. System of nozzles for working gases produced by gas generator plant of double-flow turbojet engine in aircraft includes channels and nozzles. Channels of cold circuit discharge air flow from engine fan and are formed by cross-sections having shape of semi-ring at output from cold external circuit of engine, with smooth transition of section shape to rectangle located by wide side along rear edge of aircraft wing. Upper and lower edges of such rectangle contain flaps varying the cross-section area on the nozzle edge and/or the direction of air outlet from it. Channel of hot inner circuit removes hot gases from engine gas generator and is formed in cross sections by circles. Longitudinal axis of channel of hot circuit is made in the form of smooth transition from axis of engine to axis parallel to plane of symmetry of aircraft or forming with such axis small angle to the outside in order to reduce turning moment at failure of engine. EFFECT: invention allows reducing aerodynamic resistance and weight of aircraft. 6 cl, 16 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 716 651 C2 (51) МПК F02K 1/52 (2006.01) F02K 1/54 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F02K 1/52 (2019.05); ...

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19-02-1985 дата публикации

Vtol exhaust nozzle with veer flap extension

Номер: CA1182650A
Автор: William M. Madden
Принадлежит: United Technologies Corp

VTOL Exhaust Nozzle With Veer Flap Extension Abstract A gas turbine engine exhaust nozzle assembly having vertical thrust capability comprises a two-dimensional converging/diverging nozzle and a veer flap for forward thrust operation and a rotatable hood type deflector for vertical thrust operation, wherein the forward edge of the veer flap is attached to the downstream edge of the deflector hood. During forward thrust operation, the hood and rear edge of the divergent flap portion of the converging/diverging nozzle are moved synchronously to maintain the rear edge of the divergent flap portion and the front edge of the veer flap adjacent each other such that the veer flap is aligned with and acts as an extension of the divergent flap portion of the converging/ diverging nozzle. Preferably, means are provided to independently trim the orientation of the veer flap to maximize thrust. A movable ventral flap defines the nozzle throat between itself and the deflector hood when the hood is deployed for vertical thrust.

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17-03-2020 дата публикации

Binary vector spray pipe structure and aircraft with same

Номер: CN108104974B
Принадлежит: AECC Shenyang Engine Research Institute

本发明公开了一种二元矢量喷管结构及具有其的飞机。所述二元矢量喷管结构中,上固定收敛调节片与圆转方段连接;下固定收敛调节片与圆转方段连接;上可调收敛调节片设置在上固定收敛调节片上的滑槽内;下可调收敛调节片设置在下固定收敛调节片上的滑槽内;第一作动筒与上可调收敛调节片连接;第二作动筒与下可调收敛调节片连接;上可调收敛调节片的一侧与上扩张调节片的一侧铰接;下可调收敛调节片的一侧与下扩张调节片的一侧铰接;上拉杆的一端与上固定收敛调节片铰接;上拉杆的另一端与上扩张调节片铰接;下拉杆的一端与下固定收敛调节片铰接;下拉杆的另一端与下扩张调节片铰接。本申请结构简单,易于实现运动调节。

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25-02-1977 дата публикации

ORIENTABLE THRUST EJECTION TUBE COOLING SYSTEM

Номер: FR2319771A1
Принадлежит: General Electric Co

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20-09-2012 дата публикации

TWO-SECTION NOZZLE AND AIRCRAFT CONTAINING SUCH NOZZLE

Номер: RU2011108701A

1. Сопло (20) двумерного сечения, содержащее неподвижную конструкцию (21), которая содержит диффузор (22) в своей нижней по потоку части, отличающееся тем, что дополнительно содержит горловину (200), расположенную внутри неподвижной конструкции (21) выше диффузора (22) по потоку, причем горловина содержит, по меньшей мере, две боковые стенки (210, 220), установленные с возможностью поворота внутри неподвижной конструкции (21) и соединенные друг с другом посредством поперечных удлиненных элементов (230, 240), при этом каждый удлиненный элемент установлен с возможностью скольжения на каждой стенке (210; 220). ! 2. Сопло по п.1, отличающееся тем, что каждая боковая стенка (210; 220) содержит, по меньшей мере, первые и вторые панели (211, 212; 221, 222), соединенные друг с другом посредством подвижной оси (213, 223) поворота, при этом удлиненные элементы (230, 240) установлены на первых панелях (211, 221), а каждая первая панель (211; 221) прикреплена на своем нижнем по потоку конце (2110, 2210) к оси (214, 224) поворота, соединенной с неподвижной конструкцией (21). ! 3. Сопло по п.2, отличающееся тем, что свободный конец (2121; 2221) каждой второй панели (212; 222) укреплен с возможностью скольжения вдоль боковой кромки (21а; 21b) неподвижной конструкции (21). ! 4. Сопло по п.2 или 3, отличающееся тем, что содержит, по меньшей мере, один линейный исполнительный орган (314), расположенный между концом (332) одного из удлиненных элементов (330) и соединенной с конструкцией (31) осью (325) поворота, к которой прикреплена одна из первых панелей (321). ! 5. Сопло по п.2 или 3, отличающееся тем, что каждая первая панель (211; 221) содержит, по меньшей мере, одну укосину (2115; 2215), проходящую наружу от горловины, причем каждая укосина соединен� РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) 2011 108 701 (13) A (51) МПК F02K 1/00 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ЗАЯВКА НА ИЗОБРЕТЕНИЕ (21)(22) Заявка: 2011108701/06, 10.03.2011 (71) Заявитель(и): СНЕКМА ...

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01-10-2019 дата публикации

DOUBLE-CIRCUIT TURBOREACTIVE ENGINE SYSTEM

Номер: RU2018111515A

РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2018 111 515 A (51) МПК F02K 1/00 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ЗАЯВКА НА ИЗОБРЕТЕНИЕ (21)(22) Заявка: 2018111515, 30.03.2018 (71) Заявитель(и): Арутюнов Артём Георгиевич (RU) Приоритет(ы): (22) Дата подачи заявки: 30.03.2018 (43) Дата публикации заявки: 01.10.2019 Бюл. № 28 Стр.: 1 A 2 0 1 8 1 1 1 5 1 5 R U A (57) Формула изобретения 1. Система сопел для рабочих газов в летательном аппарате, приводимом в движение горячими рабочими газами, производимыми по меньшей мере, одной двухконтурной газогенераторной установкой, содержащий холодный и горячий каналы (11, 12, 15) и сопла (13, 14), отличающийся тем, что упомянутые каналы образуют уплощенный широкой стороной в направлении задней кромки крыла газовый поток, разделенный на поток горячих газов из внутреннего контура двигателя и два потока холодного воздуха от вентиляторного контура двигателя, при этом: - горячий канал размещен между холодными, - оси всех каналов параллельны, - каждый холодный канал содержит первый цилиндрический элемент (10) канала, второй переходный элемент (11, 12) канала, причем каждый из последних выходит в одно из сопел (13, 14). 2. Система сопел по п. 1, отличающаяся тем, что сечения горячего канала, нормальные направлению потока, имеют форму окружностей. 3. Система сопел по п. 1, отличающаяся тем, что газогенераторная установка является двухконтурным турбореактивным двигателем с большой степенью двухконтурности. 4. Система сопел по п. 1, отличающаяся тем, что поперечное сечение вышеуказанного холодного канала (11, 12) имеет закон изменения площадей сечений, нормальных направлению потока, минимизирующий потери тяги. 5. Система сопел по п. 1, отличающаяся тем, что она выполнена с возможностью разделения газового потока рабочих газов на, по меньшей мере, один первый и, по меньшей мере, один второй потоки для их выпуска через, по меньшей мере, одно первое (13) и, по меньшей мере, одно второе (14) сопла и содержит, по ...

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Номер: IT1065153B
Автор:
Принадлежит: United Technologies Corp

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Автор: [UNK]
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16-11-2021 дата публикации

A throat shifting type fluidic vectoring nozzle based on translational motion for short-distance takeoff and landing

Номер: CA3087472C

A throat shifting type fluidic vectoring nozzle for short-distance takeoff and landing is provided. The nozzle comprises: a tail section of an upper wall surface of a divergent-convergent section at a front portion of a second throat being selected as a translational component. In a level flight mode, the translational component is fitted with an upper fixing portion of the divergent-convergent section at the front portion of the second throat, to form an integral upper wall surface of the divergent-convergent section and be symmetrical to a lower nozzle wall surface. When switching from the level flight mode to a short-distance takeoff and landing mode, the translational component translates toward the lower rear along a motion track, and an air curtain component for blocking air flow is arranged between the translational component and the upper fixing portion to close the flow channel, thereby realizing convenient short-distance takeoff and landing.

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