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Применить Всего найдено 3188. Отображено 100.
05-01-2012 дата публикации

Wing of an aircraft and assembly of a wing comprising a device for influencing a flow

Номер: US20120001028A1
Автор: Burkhard Goelling, Heinz Hansen, Juergen Frey, Klaus-Peter Neitzke, Veit Hildebrand
Принадлежит: AIRBUS OPERATIONS GMBH

A wing of an aircraft is described, having: a main wing, at least one high lift flap which can be moved between a retracted and an extended position, and a spoiler. The main wing has ejection openings, arranged side-by-side along the main wing spanwise direction, and in the main wing chordwise direction, and which are connected via an air conduit with the outlet device of a flow delivery driver device on the main wing or on the spoiler. The spoiler has inlet openings for the intake of air, which are connected via an air conduit with the inlet device of the flow delivery driver device. The flow delivery driver device has a receiver device for the reception of command signals for purposes of adjustment of the flow delivery driver device. An arrangement of a wing with a device for purposes of flow control with such a wing is also described.

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Номер записи: 1
20-03-2008 дата публикации

КРЫЛО ОКОЛОЗВУКОВОГО САМОЛЕТА

Номер: RU0000071626U1
Автор: Муравьев Юрий Леонидович, Трощилов Валерий Иванович
Принадлежит: ООО "Инновационный Центр "Опережение"

Крыло летательного аппарата, обшивка которого содержит профилированные бороздки отличающееся тем, что на большей части верхней поверхности крыла под углом (-30)-(+50)° к набегающему потоку установлены ряды вихреобразующих профилированных углублений, глубина которых «с» в зависимости от хорды профиля «в» составляет 0,5-50 мм, а шаг рядов «т» составляет 0,5-7 глубины «с». РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) 71 626 (13) U1 (51) МПК B64C 21/02 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ, ПАТЕНТАМ И ТОВАРНЫМ ЗНАКАМ (12) ОПИСАНИЕ ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ (21), (22) Заявка: 2007138629/22 , 18.10.2007 (24) Дата начала отсчета срока действия патента: 18.10.2007 (45) Опубликовано: 20.03.2008 (73) Патентообладатель(и): ООО "Инновационный Центр "Опережение" (RU) U 1 7 1 6 2 6 R U Ñòðàíèöà: 1 U 1 Формула полезной модели Крыло летательного аппарата, обшивка которого содержит профилированные бороздки отличающееся тем, что на большей части верхней поверхности крыла под углом (-30)-(+50)° к набегающему потоку установлены ряды вихреобразующих профилированных углублений, глубина которых «с» в зависимости от хорды профиля «в» составляет 0,5-50 мм, а шаг рядов «т» составляет 0,5-7 глубины «с». 7 1 6 2 6 (54) КРЫЛО ОКОЛОЗВУКОВОГО САМОЛЕТА R U Адрес для переписки: 123154, Москва, пр-кт Маршала Жукова, 30, оф.АТИС, ООО "Инновационный Центр "Опережение", Ю.Л. Муравьеву (72) Автор(ы): Муравьев Юрий Леонидович (RU), Трощилов Валерий Иванович (RU) U 1 U 1 7 1 6 2 6 7 1 6 2 6 R U R U Ñòðàíèöà: 2 RU 5 10 15 20 25 30 35 40 45 50 71 626 U1 Полезная модель относится к области авиационной техники, а именно к управлению пограничным слоем, главным образом крыла, а так же фюзеляжа, горизонтального и вертикального оперения. Развитие современной авиационной техники характеризуется созданием целого направления в самолетостроении, а именно созданием, так называемых, административных правительственных самолетов, наряду с магистральными самолетами, крейсерская скорость которых М=0,6-0,85. В ...

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Номер записи: 2
10-04-2008 дата публикации

КРЫЛО

Номер: RU0000072197U1
Автор: Бабенко Леонид Михайлович
Принадлежит: Бабенко Леонид Михайлович

1. Крыло, включающее собственно крыло, и закрылок или другое средство механизации задней кромки крыла, и предкрылок, отличающееся тем, что на задней кромке предкрылка выполнено щелевое сопло для истечения воздуха. 2. Крыло по п.1, отличающееся тем, что для закрывания в нерабочем состоянии щелевое сопло снабжено створкой, являющейся одной из поверхностей предкрылка, примыкающей к его задней кромке и установленной на нем с помощью осевого шарнира, с осью, направленной вдоль размаха предкрылка. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) 72 197 (13) U1 (51) МПК B64C 21/02 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ, ПАТЕНТАМ И ТОВАРНЫМ ЗНАКАМ (12) ОПИСАНИЕ ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ (21), (22) Заявка: 2007145311/22 , 06.12.2007 (24) Дата начала отсчета срока действия патента: 06.12.2007 (45) Опубликовано: 10.04.2008 (72) Автор(ы): Бабенко Леонид Михайлович (RU) (73) Патентообладатель(и): Бабенко Леонид Михайлович (RU) U 1 7 2 1 9 7 R U Ñòðàíèöà: 1 U 1 Формула полезной модели 1. Крыло, включающее собственно крыло, и закрылок или другое средство механизации задней кромки крыла, и предкрылок, отличающееся тем, что на задней кромке предкрылка выполнено щелевое сопло для истечения воздуха. 2. Крыло по п.1, отличающееся тем, что для закрывания в нерабочем состоянии щелевое сопло снабжено створкой, являющейся одной из поверхностей предкрылка, примыкающей к его задней кромке и установленной на нем с помощью осевого шарнира, с осью, направленной вдоль размаха предкрылка. 7 2 1 9 7 (54) КРЫЛО R U Адрес для переписки: 344044, г.Ростов-на-Дону, ул. Портовая, 174, кв.14, Л.М. Бабенко U 1 U 1 7 2 1 9 7 7 2 1 9 7 R U R U Ñòðàíèöà: 2 RU 5 10 15 20 25 30 35 40 45 50 72 197 U1 Полезная модель относится к авиационной технике, в частности к самолетам, и может быть использовано на самолетах укороченного взлета и посадки. Известна конструкция крыла самолета, способствующая увеличению подъемной силы (далее "ПС") крыла, путем интенсификации обтекающего его потока (см. пат. Англии № ...

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Номер записи: 3
20-04-2008 дата публикации

СОСТАВНАЯ ЗАКОНЦОВКА НЕСУЩЕЙ ПОВЕРХНОСТИ

Номер: RU0000072459U1
Автор: Муравьев Юрий Леонидович, Трощилов Валерий Иванович
Принадлежит: ООО "Инновационный Центр "Опережение"

Составная законцовка крыла, содержащая сопряженную с крылом профилированную основную часть и расположенный впереди ее обтекатель, отличающаяся тем, что обтекатель выполнен в виде удлиненного ребра с относительно острой наружной кромкой, плавно сопрягаемой своей профилированной поверхностью с основной законцовкой, содержащей, например, S-образные профили, при этом размах обтекателя находится в соотношении с размахом основной законцовки, как 0,1-0,3, а корневая хорда обтекателя составляет 0,3-1,2 от корневой хорды основной законцовки. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) 72 459 (13) U1 (51) МПК B64C 21/02 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ, ПАТЕНТАМ И ТОВАРНЫМ ЗНАКАМ (12) ОПИСАНИЕ ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ (21), (22) Заявка: 2007143594/22 , 27.11.2007 (24) Дата начала отсчета срока действия патента: 27.11.2007 (45) Опубликовано: 20.04.2008 (73) Патентообладатель(и): ООО "Инновационный Центр "Опережение" (RU) U 1 7 2 4 5 9 R U Ñòðàíèöà: 1 U 1 Формула полезной модели Составная законцовка крыла, содержащая сопряженную с крылом профилированную основную часть и расположенный впереди ее обтекатель, отличающаяся тем, что обтекатель выполнен в виде удлиненного ребра с относительно острой наружной кромкой, плавно сопрягаемой своей профилированной поверхностью с основной законцовкой, содержащей, например, S-образные профили, при этом размах обтекателя находится в соотношении с размахом основной законцовки, как 0,1-0,3, а корневая хорда обтекателя составляет 0,3-1,2 от корневой хорды основной законцовки. 7 2 4 5 9 (54) СОСТАВНАЯ ЗАКОНЦОВКА НЕСУЩЕЙ ПОВЕРХНОСТИ R U Адрес для переписки: 123154, Москва, пр-кт Маршала Жукова, 30, оф.АТИС, ООО "Инновационный Центр "Опережение", Ю.Л. Муравьеву (72) Автор(ы): Трощилов Валерий Иванович (RU), Муравьев Юрий Леонидович (RU) U 1 U 1 7 2 4 5 9 7 2 4 5 9 R U R U Ñòðàíèöà: 2 RU 5 10 15 20 25 30 35 40 45 50 72 459 U1 Полезная модель относится преимущественно к области авиационной техники, например, к созданию ...

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Номер записи: 4
10-06-2008 дата публикации

БЕЗОТРЫВНЫЙ ПРОФИЛЬ КРЫЛА

Номер: RU0000073845U1
Автор: Муравьев Юрий Леонидович, Трощилов Валерий Иванович
Принадлежит: ООО "Инновационный Центр "Опережение"

Безотрывный профиль крыла, содержащий участок профилированной верхней поверхности вдоль большей части размаха крыла с установленным на нем не менее одного, в том числе прерывистого, ряда вихреобразующих профилированных углублений, глубина которых «с» в зависимости от хорды профиля «в» составляет 1-50 мм, а шаг рядов «т» составляет 0,5-3 глубины «с», отличающийся тем, что ряды профилированных углублений на верхней поверхности установлены в зоне 30-95% хорды профиля. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) 73 845 (13) U1 (51) МПК B64C 21/02 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ, ПАТЕНТАМ И ТОВАРНЫМ ЗНАКАМ (12) ОПИСАНИЕ ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ (21), (22) Заявка: 2007134636/22 , 18.09.2007 (24) Дата начала отсчета срока действия патента: 18.09.2007 (45) Опубликовано: 10.06.2008 (73) Патентообладатель(и): ООО "Инновационный Центр "Опережение" (RU) U 1 7 3 8 4 5 R U Ñòðàíèöà: 1 U 1 Формула полезной модели Безотрывный профиль крыла, содержащий участок профилированной верхней поверхности вдоль большей части размаха крыла с установленным на нем не менее одного, в том числе прерывистого, ряда вихреобразующих профилированных углублений, глубина которых «с» в зависимости от хорды профиля «в» составляет 1-50 мм, а шаг рядов «т» составляет 0,5-3 глубины «с», отличающийся тем, что ряды профилированных углублений на верхней поверхности установлены в зоне 30-95% хорды профиля. 7 3 8 4 5 (54) БЕЗОТРЫВНЫЙ ПРОФИЛЬ КРЫЛА R U Адрес для переписки: 123154, Москва, пр-кт Маршала Жукова, 30, оф. АТИС, ООО "Инновационный Центр "Опережение", Ю.Л. Муравьеву (72) Автор(ы): Трощилов Валерий Иванович (RU), Муравьев Юрий Леонидович (RU) RU 5 10 15 20 25 30 35 40 45 50 73 845 U1 Полезная модель относится к области авиационной технике, а именно к управлению пограничным слоем, например, крыла и может быть использована также в гидродинамике при разработки обтекаемых профилей - корпуса судна, винта. Наиболее освоенными в авиации и экономически целесообразными в настоящее время ...

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Номер записи: 5
20-04-2009 дата публикации

УСТРОЙСТВО ДЛЯ ИЗМЕНЕНИЯ ПОДЪЕМНОЙ СИЛЫ

Номер: RU0000082189U1
Автор: Мрдуляш Павел Брунович
Принадлежит: Мрдуляш Павел Брунович

1. Устройство для изменения подъемной силы для тела в потоке текучей среды, содержащее один или более отстоящих друг от друга элементов, выдвигающихся из обтекаемой потоком поверхности тела, отличающееся тем, что упомянутые выдвигающиеся элементы выполнены гибкими. 2. Устройство по п.1, отличающееся тем, что поперечное сечение, по меньшей мере, одного выдвигающегося гибкого элемента является круглым. 3. Устройство по п.1, отличающееся тем, что поперечное сечение, по меньшей мере, одного выдвигающегося гибкого элемента является многоугольным. 4. Устройство по п.1, отличающееся тем, что, по меньшей мере, один гибкий элемент представляет собой гибкую ленту. 5. Устройство по п.1, отличающееся тем, что оно представляет собой устройство для изменения подъемной силы крыла самолета. 6. Устройство по п.1, отличающееся тем, что множество гибких элементов установлено на большей части обтекаемой потоком поверхности. 7. Устройство по п.1, отличающееся тем, что множество гибких элементов установлено на части обтекаемой потоком поверхности. 8. Устройство по п.1, отличающееся тем, что гибкий элемент жестко прикреплен к концу штока, установленного под обтекаемой потоком поверхностью, посредством которого гибкий элемент выдвигается на заданное расстояние над упомянутой обтекаемой потоком поверхностью. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 82 189 U1 (51) МПК B64C 21/00 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ, ПАТЕНТАМ И ТОВАРНЫМ ЗНАКАМ (12) ОПИСАНИЕ ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ (21), (22) Заявка: 2008108645/22, 07.03.2008 (24) Дата начала отсчета срока действия патента: 07.03.2008 (45) Опубликовано: 20.04.2009 (72) Автор(ы): Мрдуляш Павел Брунович (RU) (73) Патентообладатель(и): Мрдуляш Павел Брунович (RU) R U Адрес для переписки: 125475, Москва, ул. Зеленоградская, 21, корп.1, кв.61, П.Б. Мрдуляшу U 1 8 2 1 8 9 R U Ñòðàíèöà: 1 ru CL U 1 Формула полезной модели 1. Устройство для изменения подъемной силы для тела в потоке текучей среды, содержащее один или более ...

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Номер записи: 6
27-03-2011 дата публикации

УСТРОЙСТВО СОЗДАНИЯ ПОДЪЕМНОЙ СИЛЫ НАД ПОВЕРХНОСТЬЮ ВОДЫ

Номер: RU0000103093U1
Автор: Андреев Юрий Петрович
Принадлежит: Андреев Юрий Петрович

1. Устройство создания подъемной силы, содержащее крыло аэродинамического сечения, полость крыла разделена перегородками с образованием сегмента с радиально-щелевыми соплами, а в каждом сегменте снизу крыла выполнено не менее чем по одной щели, ведущей внутрь сегмента, отличающееся тем, что перед входящим радиально-щелевым соплом крыла установлен насос, выполненный с возможностью подачи посредством водозаборной трубы или шланга жидкости в полость крыла. 2. Устройство создания подъемной силы по п.1, отличающееся тем, что перед щелью снизу крыла установлен вакуумный насос, выполненный с возможностью откачки воздуха из полости крыла. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) 103 093 (13) U1 (51) МПК B64C 21/00 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ, ПАТЕНТАМ И ТОВАРНЫМ ЗНАКАМ (12) ОПИСАНИЕ ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ (21)(22) Заявка: 2010150869/11, 14.12.2010 (24) Дата начала отсчета срока действия патента: 14.12.2010 (73) Патентообладатель(и): Андреев Юрий Петрович (RU) R U Приоритет(ы): (22) Дата подачи заявки: 14.12.2010 (72) Автор(ы): Андреев Юрий Петрович (RU) (45) Опубликовано: 27.03.2011 1 0 3 0 9 3 R U Формула полезной модели 1. Устройство создания подъемной силы, содержащее крыло аэродинамического сечения, полость крыла разделена перегородками с образованием сегмента с радиально-щелевыми соплами, а в каждом сегменте снизу крыла выполнено не менее чем по одной щели, ведущей внутрь сегмента, отличающееся тем, что перед входящим радиально-щелевым соплом крыла установлен насос, выполненный с возможностью подачи посредством водозаборной трубы или шланга жидкости в полость крыла. 2. Устройство создания подъемной силы по п.1, отличающееся тем, что перед щелью снизу крыла установлен вакуумный насос, выполненный с возможностью откачки воздуха из полости крыла. Ñòðàíèöà: 1 ru CL U 1 U 1 (54) УСТРОЙСТВО СОЗДАНИЯ ПОДЪЕМНОЙ СИЛЫ НАД ПОВЕРХНОСТЬЮ ВОДЫ 1 0 3 0 9 3 Адрес для переписки: 127566, Москва, Высоковольтный пр-д, 1, корп.3, кв.192, Е.В. Мохову U 1 U 1 1 ...

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Номер записи: 7
10-08-2011 дата публикации

УСТРОЙСТВО ЗАЩИТЫ АВИАЦИОННЫХ ГАЗОТУРБИННЫХ ДВИГАТЕЛЕЙ ОТ ПОПАДАНИЯ ПОСТОРОННИХ ПРЕДМЕТОВ, ПОДНЯТЫХ ВИХРЕВЫМИ ШНУРАМИ ПРИ ОТЛАДКЕ НА ГОНОЧНОЙ ПЛОЩАДКЕ

Номер: RU0000107125U1
Автор: Ахъяров Айрат Аглиевич, Зайцев Геннадий Александрович, Кагарманов Раиль Латифович, Румянцева Галина Петровна
Принадлежит: Федеральное государственное унитарное предприятие "Летно-исследовательский институт имени М.М. Громова"

1. Устройство защиты авиационных газотурбинных двигателей от попадания посторонних предметов с применением покрытия из «искусственной травы», отличающееся тем, что оно содержит установленную на гоночной площадке для отладки газотурбинных двигателей металлическую пластину прямоугольной формы, массой не менее 100 кг, размером сторон не менее двух радиусов зоны вихреобразования, толщиной 2-5 мм, на которой жестко закреплено синтетическое покрытие из «искусственной травы» со сложной формой поверхности, например ровной, волнообразной, хаотичной, выполненное с плотностью ворса 3-5 шт. ворса на 1 дм, направленного вверх, и средней высотой ворса не менее 0,06-0,07 диаметра воздухозаборника газотурбинного двигателя. 2. Устройство защиты авиационных газотурбинных двигателей от попадания посторонних предметов в воздухозаборник с применением покрытия из «искусственной травы» по п.1, отличающееся тем, что устанавливают его в начале взлетной полосы на участке страгивания и разгона самолета на взлетных режимах работы двигателей до достижения им скорости сдува вихря. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) 107 125 (13) U1 (51) МПК B64C 21/00 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ, ПАТЕНТАМ И ТОВАРНЫМ ЗНАКАМ (12) ОПИСАНИЕ ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ (21)(22) Заявка: 2011114010/11, 12.04.2011 (24) Дата начала отсчета срока действия патента: 12.04.2011 (45) Опубликовано: 10.08.2011 1 0 7 1 2 5 R U Формула полезной модели 1. Устройство защиты авиационных газотурбинных двигателей от попадания посторонних предметов с применением покрытия из «искусственной травы», отличающееся тем, что оно содержит установленную на гоночной площадке для отладки газотурбинных двигателей металлическую пластину прямоугольной формы, массой не менее 100 кг, размером сторон не менее двух радиусов зоны вихреобразования, толщиной 2-5 мм, на которой жестко закреплено синтетическое покрытие из «искусственной травы» со сложной формой поверхности, например ровной, волнообразной, хаотичной, выполненное ...

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Номер записи: 8
20-11-2011 дата публикации

УСТРОЙСТВО УПРАВЛЕНИЯ ПОГРАНИЧНЫМ СЛОЕМ ТОЛСТОГО КРЫЛА

Номер: RU0000110362U1
Автор: Коноваленко Виктор Антонович, Марков Сергей Иванович
Принадлежит: Федеральное государственное бюджетное образовательное учреждение высшего профессионального образования "Санкт-Петербургский государственный политехнический университет" (ФГБОУ ВПО "СПбГПУ")

Устройство управления пограничным слоем толстого крыла, содержащее канал, образованный верхней поверхностью корпуса транспортного средства, имеющего форму толстого крыла, выступающими над ней вертикальными антииндукционными щитами, расположенными вдоль боковых кромок указанной поверхности, и системой соединяющих щиты надкрылков в верхней части канала, в который направлен обтекающий корпус воздушный поток, ускоряемый тяговыми движителями, размещенными в кормовой части канала для противодействия встречному положительному градиенту давления, отличающееся тем, что оси вращения тяговых движителей расположены вертикально, причем количество движетелей четное с попарно-противоположным направлением вращения. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 110 362 U1 (51) МПК B64C 21/00 (2006.01) B64C 9/00 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ, ПАТЕНТАМ И ТОВАРНЫМ ЗНАКАМ (12) ОПИСАНИЕ (21)(22) Заявка: ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ 2011121470/11, 27.05.2011 (24) Дата начала отсчета срока действия патента: 27.05.2011 (45) Опубликовано: 20.11.2011 Бюл. № 32 1 1 0 3 6 2 R U Формула полезной модели Устройство управления пограничным слоем толстого крыла, содержащее канал, образованный верхней поверхностью корпуса транспортного средства, имеющего форму толстого крыла, выступающими над ней вертикальными антииндукционными щитами, расположенными вдоль боковых кромок указанной поверхности, и системой соединяющих щиты надкрылков в верхней части канала, в который направлен обтекающий корпус воздушный поток, ускоряемый тяговыми движителями, размещенными в кормовой части канала для противодействия встречному положительному градиенту давления, отличающееся тем, что оси вращения тяговых движителей расположены вертикально, причем количество движетелей четное с попарнопротивоположным направлением вращения. Стр.: 1 U 1 U 1 (54) УСТРОЙСТВО УПРАВЛЕНИЯ ПОГРАНИЧНЫМ СЛОЕМ ТОЛСТОГО КРЫЛА 1 1 0 3 6 2 Адрес для переписки: 195251, Санкт-Петербург, ул. Политехническая, 29, ФГБОУ ВПО "Санкт- ...

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Номер записи: 9
10-01-2014 дата публикации

АЭРОМОБИЛЬ

Номер: RU0000136390U1
Автор: Коноваленко Виктор Антонович, Коротин Вячеслав Георгиевич, Макаров Борис Григорьевич, Попов Юрий Гаврилович
Принадлежит: Коноваленко Виктор Антонович, Попов Юрий Гаврилович

Аэромобиль, содержащий корпус в виде толстого крыла, у которого верхней поверхностью образован канал и над которым выступают вертикальные антииндукционные щиты, расположенные вдоль боковых кромок указанной поверхности, и систему соединяющих щиты надкрылков в верхней части канала, в который направлен обтекающий упомянутый корпус воздушный поток, отличающийся тем, что в нижней части канала, в зоне возможного срыва потока, расположена по всей ширине канала поперечная щель, соединенная с всасывающими частями тяговых движителей, включающих центробежные или осевые компрессоры, размещенных внутри кормовой части упомянутого корпуса и работающих на общий коллектор-конфузор, формирующий плоский воздушный поток в горизонтальной плоскости, равномерно распределенный по всей задней кромке упомянутого корпуса и отсекающий зону пониженного давления над крылом от зоны нормального давления под крылом. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 136 390 U1 (51) МПК B60F 5/02 (2006.01) B64C 21/00 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ (21)(22) Заявка: ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ 2012158361/11, 26.12.2012 (24) Дата начала отсчета срока действия патента: 26.12.2012 (73) Патентообладатель(и): Коноваленко Виктор Антонович (RU), Попов Юрий Гаврилович (RU) (45) Опубликовано: 10.01.2014 Бюл. № 1 1 3 6 3 9 0 R U Формула полезной модели Аэромобиль, содержащий корпус в виде толстого крыла, у которого верхней поверхностью образован канал и над которым выступают вертикальные антииндукционные щиты, расположенные вдоль боковых кромок указанной поверхности, и систему соединяющих щиты надкрылков в верхней части канала, в который направлен обтекающий упомянутый корпус воздушный поток, отличающийся тем, что в нижней части канала, в зоне возможного срыва потока, расположена по всей ширине канала поперечная щель, соединенная с всасывающими частями тяговых движителей, включающих центробежные или осевые компрессоры, размещенных внутри кормовой части упомянутого корпуса и ...

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Номер записи: 10
27-01-2015 дата публикации

КРЫЛО С УПРАВЛЕНИЕМ ПОГРАНИЧНЫМ СЛОЕМ

Номер: RU0000149950U1
Автор: Житенев Андрей Славдиевич, Житенёв Евгений Андреевич
Принадлежит: Житенев Андрей Славдиевич, Житенёв Евгений Андреевич

Крыло с управлением пограничным слоем, имеющее постоянную кривизну, содержащее внутреннюю полость в теле крыла, воздухозаборный канал, соединяющий полость с внешней средой и щель для сдува пограничного слоя, отличающееся тем, что внутренняя полость выполнена сквозной, воздухозаборный канал с площадью сечения S расположен по всему размаху крыла вдоль передней его кромки, а начиная с расстояния ≈35% хорды на верхней поверхности по всему размаху крыла дополнительно расположено несколько щелей с площадями сечения S, S, S, ... S, причём необходимо соблюдение неравенства (1): где S, S, S, ... S - площади выходных щелей, S - площадь воздухозаборного канала, b - хорда крыла, L - длина верхнего обвода профиля крыла, для создания скорости воздушного потока V на выходе из щелей большей, чем местная скорость воздушного потока, которая зависит от геометрических характеристик профиля крыла, в соответствии с формулой постоянства массового расхода воздуха (2) где ρ - плотность воздуха, S - сумма площадей выпускных щелей, S - площадь воздухозаборного канала, V - скорость набегающего на крыло потока, V - скорость воздушного потока выходящего из щели, сумма площадей выходных щелей (S+S+S+...+S) должна быть меньше, чем площадь входного канала (S) в Х, во столько же Х скорость воздушного потока на выходе из щели V будет больше, чем скорость набегающего воздушного потока V на крыло. И 1 149950 ко РОССИЙСКАЯ ФЕДЕРАЦИЯ 7 ВУ‘”’ 149 950° У1 ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ИЗВЕЩЕНИЯ К ПАТЕНТУ НА ПОЛЕЗНУЮ МОДЕЛЬ МЕЭК Восстановление действия патента Дата, с которой действие патента восстановлено: 05.04.2021 Дата внесения записи в Государственный реестр: 05.04.2021 Дата публикации и номер бюллетеня: 05.04.2021 Бюл. №10 Стр.: 1 па Очбб7 1 ЕП

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Номер записи: 11
04-02-2020 дата публикации

Крыло самолета

Номер: RU0000195734U1
Автор: Ковалев Игорь Сергеевич
Принадлежит: Ковалев Игорь Сергеевич

Полезная модель относится к авиационной технике и может быть использована при разработке новых конструкций самолетов. В конструкции крыла самолета поверхность крыла над хордой профиля крыла жестко соединена с вибратором и закреплена на каркасных элементах крыла с использованием уплотнений, выполненных с возможностью поглощения вибрации. Этим обеспечивается увеличение подъемной силы и снижение лобового сопротивления на режимах взлета, посадки и горизонтального полета. 2 з.п. ф-лы, 3 ил. Ц 195734 ко РОССИЙСКАЯ ФЕДЕРАЦИЯ Во“ 195 734 91 ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ИЗВЕЩЕНИЯ К ПАТЕНТУ НА ПОЛЕЗНУЮ МОДЕЛЬ ММ9К Досрочное прекращение действия патента из-за неуплаты в установленный срок пошлины за поддержание патента в силе Дата прекращения действия патента: 09.10.2020 Дата внесения записи в Государственный реестр: 20.09.2021 Дата публикации и номер бюллетеня: 20.09.2021 Бюл. №26 Стр.: 1 па $1961 ЕП

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Номер записи: 12
07-08-2020 дата публикации

Дозвуковой летательный аппарат, имеющий толстый крыловой профиль с устройством снижения лобового сопротивления

Номер: RU0000199016U1
Автор: Пилипенко Людмила Викторовна, Пирухин Виталий Александрович, Прокопенко Евгений Алексеевич, Хлебников Сергей Германович
Принадлежит: Пилипенко Людмила Викторовна, Пирухин Виталий Александрович, Прокопенко Евгений Алексеевич, Хлебников Сергей Германович

Полезная модель относится к техническим средствам в области аэродинамики, в частности к устройствам, предназначенным для снижения силы лобового сопротивления летательных аппаратов (ЛА), движущихся с дозвуковыми скоростями. Дозвуковой летательный аппарат, имеющий толстый крыловой профиль с устройством снижения лобового сопротивления, которое выполнено в виде внутренней полуоткрытой системы внешней циркуляции воздушного потока, поступающего из пограничного слоя вдоль поверхности ЛА в эллиптическую каверну, расположенную в области отрыва потока от поверхности ЛА. Внутри эллиптической каверны организован управляемый процесс циркуляции воздушного потока за счет размещения внутри ЛА воздушного канала, выполненного в виде трубопровода, имеющего на одном конце конфузор для отсоса воздуха из каверны, а на другом конце - конфузор для выдува воздуха в каверну. Для организации управляемого процесса отсоса-выдува воздуха из каверны в трубопроводе между конфузорами расположен осевой вентилятор, приводимый во вращение с помощью ведущего вала электродвигателя. Технический результат, обеспечиваемый полезной моделью, состоит в создании условий, позволяющих увеличить грузоподъемность ЛА и обеспечить устойчивость управления на различных этапах полета ЛА, в том числе в условиях высоких ветровых нагрузок. 1 ил. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 199 016 U1 (51) МПК B64C 21/00 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ (52) СПК B64C 21/00 (2020.05) (21)(22) Заявка: 2020112779, 31.03.2020 (24) Дата начала отсчета срока действия патента: Дата регистрации: 07.08.2020 (45) Опубликовано: 07.08.2020 Бюл. № 22 1 9 9 0 1 6 R U (56) Список документов, цитированных в отчете о поиске: RU 149950 U1, 27.01.2015. RU 85446 U1, 10.08.2009. CN 208882103 U, 21.05.2019. GB 2178131 А, 04.02.1987. (54) Дозвуковой летательный аппарат, имеющий толстый крыловой профиль с устройством снижения лобового сопротивления (57) Реферат: Полезная модель ...

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Номер записи: 13
23-02-2012 дата публикации

High-lift flap, arrangement of a high-lift flap together with a device for influencing the flow on the same and aircraft comprising said arrangement

Номер: US20120043428A1
Автор: Burkhard Goelling, Christoph Jensch, Heinz Hansen, Kai-Christoph Pfingsten, Klaus-Peter Neitzke, Rolf Radespiel
Принадлежит: AIRBUS OPERATIONS GMBH

An aerodynamic body of an aircraft with an air outlet opening and an air intake opening that communicates with the air outlet opening via an air conduit is described. A flow delivery driver device for influencing the flow within the air conduit is integrated into the air conduit. The surfaces of the aerodynamic body in the body chord direction include at least one air outlet opening in the front region of the aerodynamic body, and at least one air intake opening on the upper surface of the aerodynamic body and in the rear region of the aerodynamic body and/or on the upper surface of the aerodynamic body in the trailing edge region and/or on the lower surface of the aerodynamic body in the trailing edge region. Arrangements of a main wing and an adjustable flap, and an aircraft with such an aerodynamic body are also described.

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Номер записи: 14
22-03-2012 дата публикации

Active Aircraft Pylon Noise Control System

Номер: US20120068011A1
Автор: Alaa A. Elmiliqui, Michael J. Czech, Russell H. Thomas
Принадлежит: National Aeronautics and Space Administration NASA

An active pylon noise control system for an aircraft includes a pylon structure connecting an engine system with an airframe surface of the aircraft and having at least one aperture to supply a gas or fluid therethrough, an intake portion attached to the pylon structure to intake a gas or fluid, a regulator connected with the intake portion via a plurality of pipes, to regulate a pressure of the gas or fluid, a plenum chamber formed within the pylon structure and connected with the regulator, and configured to receive the gas or fluid as regulated by the regulator, and a plurality of injectors in communication with the plenum chamber to actively inject the gas or fluid through the plurality of apertures of the pylon structure.

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Номер записи: 15
14-06-2012 дата публикации

Pylon for fixing an aircraft engine having unducted pusher propellers

Номер: US20120145824A1
Автор: Gaetan Jean Mabboux, Jean Bertucchi, Stephane Emmanuel Daniel Bensilum
Принадлежит: SNECMA SAS

A pylon for fixing an aircraft engine having unducted pusher propellers, the pylon ensuring the fixation of a propulsive system on the boattail of the aircraft, the pylon having a trailing edge, with an upper face and a lower face, for an airflow encountered by the pylon, wherein at least one of the two faces of the upper face and the lower face of the trailing edge is inclinable, at least in part.

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Номер записи: 16
31-01-2013 дата публикации

Flow body, in particular for aircraft

Номер: US20130025727A1
Автор: Geza Schrauf, Martin Gerber
Принадлежит: AIRBUS OPERATIONS GMBH

A flow body is disclosed, particularly for aircraft. The flow body includes an outer side impinged on in a predetermined manner by a fluid in a direction of impinging flow, the flow body having on its outer side at least one flow control device including micro-perforations arranged in at least one segment of the outer side, at least one connecting passage communicated with the micro-perforations via at least one suction chamber so fluid flowing through the micro-perforations flows via the suction chamber into the connecting passage, at least one suction device having a first inlet communicated with the connecting passage, a second inlet communicated with at least one ram fluid feed line, wherein the ram fluid feed line is in a region of the flow body opposite to the direction of impinging flow of the flow body, and an outlet device for discharging the fluid.

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Номер записи: 17
31-01-2013 дата публикации

Aerodynamic surface with improved properties

Номер: US20130028744A1
Автор: Göte Strindberg, Pontus Nordin
Принадлежит: SAAB AB

An article including an outer surface that serves as an aerodynamic surface when the article is subjected for an air stream. A resin matrix made of a polymeric composite laminate of at least one ply includes the outer surface. The at least one ply includes a nano structure embedded therein such that nano filaments of the nano structure in the ply essentially have the same angular orientation relative the plane of the outer surface. The outer ply is a ply of a laminate including at least two plies. Each ply includes large fibers having an orientation different from or identical to the orientation of large fibers of an adjacent ply.

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Номер записи: 18
14-03-2013 дата публикации

Surface entity for the reduction of the air resistance of an aviation vehicle

Номер: US20130062469A1
Автор: Freerk Syassen, Metin Cengiz
Принадлежит: AIRBUS OPERATIONS GMBH

The invention concerns a surface entity for the reduction of the air resistance of an aviation vehicle, in particular an aircraft. In accordance with the invention the surface entity is formed with at least one metal wire arrangement, in particular with a metal mesh and/or a composite mesh, which can be arranged in at least some sections in the region of at least one surface of the aviation vehicle immersed in the flow, in particular can be adhesively bonded and/or clamped onto the latter, wherein the metal wire arrangement has a ribbed structure with a multiplicity of ribs running essentially parallel to one another. As a consequence of the multiplicity of ribs running parallel to the flow direction, the arrangement of which can be compared with the structure of the sharkskin of known art, there ensues a reduction of the flow resistance of the surface on or over which the air flows of between 3% and 10%, as a result of which a significant fuel saving potential ensues in flight operations. In comparison to polymer films of prior known art with a surface structure similar to that of the sharkskin a significantly higher resistance to erosion ensues. The metal wire arrangement can be formed with a metal mesh and/or with a composite mesh thermally joined at the crossing points.

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Номер записи: 19
27-06-2013 дата публикации

Aircraft with an arrangement of flow-influencing devices

Номер: US20130166110A1
Автор: Burkhard GÖLLING, Rudibert King, Ulrich Notger HEINZ, Wolfgang Nitsche
Принадлежит: AIRBUS OPERATIONS GMBH

An aircraft, which has a respective arrangement of flow-influencing devices in at least one surface segment of each wing extending in the wingspan direction in order to influence the fluid flow over the surface segment, and of flow condition sensor devices for measuring the flow condition on the respective segment, and a flight control device, wherein the flight control device has a flow-influencing target parameter setting device connected with the arrangement of flow-influencing devices for generating target parameters for the flow-influencing devices of the at least one surface segment, wherein the flow-influencing devices are designed in such a way as to use the target parameters to change the local lift coefficients or correlations between the drag and lift coefficients in the segment where respectively located.

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Номер записи: 20
11-07-2013 дата публикации

Air sucking vehicle fuselage component, method for producing an air sucking vehicle fuselage component and a vehicle, especially an aircraft, with an air sucking vehicle fuselage component

Номер: US20130175402A1
Автор: Wolfgang Voege
Принадлежит: AIRBUS OPERATIONS GMBH

An air-sucking vehicle fuselage component includes an outer delimitation part, and inner delimitation part, and connecting elements. The inner delimitation part is connected via the connecting elements to the outer delimitation part. The outer delimitation part includes a continuously curved shape. The inner delimitation part includes two or more flat delimitation part components, which are each connected to one another at one edge to form a connection edge, and the connection edge is configured to be attached to the outer delimitation part. A type of framework structure is thus provided, which causes increased stability and can split impacting objects in particular if two inner delimitation part components are used on the vehicle fuselage component.

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Номер записи: 21
01-08-2013 дата публикации

STRUCTURALLY DESIGNED AERODYNAMIC RIBLETS

Номер: US20130193270A1
Автор: Mathews Mary J., McLean James D., Rawlings Diane C.
Принадлежит: The Boeing Company

An array of aerodynamic riblets is formed with a surface layer for adhering to an aerodynamic surface and a plurality of riblet tips having a parabolic cross section extending from the surface layer. 1. A method for fabricating an array of riblets comprising:{'sup': '2', 'forming a master tool having parabolic protuberances with equation y=px+h corresponding to a desired riblet arrayforming a complementary tool from the master tool;depositing high elongation elastomeric for riblet tips and a surface layer in the complimentary tool;depositing an adhesive layer to form an appliqué;removing the high elongation elastomeric appliqué from the complementary tool; and,adhering the high elongation elastomeric appliqué to an aerodynamic surface.2. The method of further comprising determining a parabolic profile factor claim 1 , p claim 1 , including selecting the profile factor consistent with a cladding on the selected riblet material.3. The method of further comprising depositing a supporting polymer layer intermediate the adhesive layer and the elastomeric tips.4. The method of further comprising depositing a UV resistant cladding over the high elongation elastomeric layer.5. The method of wherein the complimentary tool is a web tool and further comprising sputtering a UV resistant coating on the web tool prior to depositing the high elongation elastomeric.6. A method of reducing drag on an aerodynamic surface comprising:forming an array of a plurality of riblet tips with a parabolic cross section extending from a surface layer; andadhering the array of riblets to an aerodynamic surface.7. A method of enhancing the durablity of riblets on an aerodynamic surface comprising:forming an array of a plurality of riblet tips have a parabolic cross section extending from a surface layer; andadhering the array of riblets to an aerodynamic surface.8. The method of further comprising selecting the high elongation elastomer from a set of polymers and copolymers and shape memory ...

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Номер записи: 22
22-08-2013 дата публикации

NOSE FOR SUPERSONIC FLYING OBJECT

Номер: US20130214094A1
Автор: Ishikawa Hiroaki, Kawai Taro, TOKUGAWA Naoko, Tozuka Ayako, Ueda Yoshine
Принадлежит: JAPAN AEROSPACE EXPLORATION AGENCY

Provided is a nose for a supersonic flying object, which has a natural laminar flow nose shape capable of suppressing laminar-turbulent transition and drastically reducing a frictional drag. The nose for a supersonic flying object has a low resistive body shape symmetrical about a central axis as a base shape, wherein the base shape is approximately a cone shape having a linear, simple convex curve, or simple concave curve generatrix and a deformation element having a wavy shape is added to the base shape. 1. A nose for a supersonic flying object , having a body shape symmetrical about a central axis as a base shape , characterized in that:said base shape is approximately a cone shape having a linear, simple convex curve, or simple concave curve generatrix, anda deformation element having a wavy shape is added to said base shape.2. The nose for a supersonic flying object according to claim 1 , characterized in that said deformation element is at least one of a sinusoidal deformation for deforming said base shape as a whole into a wave shape in at least one of a circumferential direction and an axial direction claim 1 , and a local wavy deformation for deforming a part of said base shape locally into a wave shape.6. The nose for as supersonic flying object according to claim 1 , characterized in that said base shape is a Sears-Haack body or a flared cone.7. The nose for a supersonic flying object according to claim 1 , characterized in being placed in an airflow in an attitude having an angle of attack greater than 0°. The present invention relates to a nose for a flying object that flies at supersonic speed, and more particularly to a nose for a supersonic flying object having a shape capable of delaying a boundary layer transition so as to reduce frictional drag.It is known that when a boundary layer is laminar, frictional drag is greatly reduced in comparison with a case where the boundary layer is turbulent. Therefore, to reduce frictional drag in a nose-shaped ...

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Номер записи: 23
19-09-2013 дата публикации

HIGH LIFT SYSTEM FOR AN AIRCRAFT

Номер: US20130240676A1
Автор: GÖLLING Burkhard
Принадлежит:

A high lift system with a main wing and control flaps, wherein the respective guiding device are at least partially provided with a fairing, having a flow control device for purposes of controlling the flow around the high lift system with at least two inlet ducts running along the main wing chordwise direction with in each case at least one inlet, which device is located on or underneath the lower surface of the high lift system, wherein at least one outlet duct for air is furthermore provided, which is connected with the inlet ducts in a fluid-communicating manner, and has at least one outlet, which is located on the upper surface of at least one regulating flap and/or with respect to the main wing chordwise direction in the rear third of the main wing of the high lift system. 1. A high lift system with a main wing and control flaps and guiding devices for the positioning of the control flaps , wherein the guiding devices are at least partially provided with a fairing , the high lift system comprising:a flow control device for purposes of influencing the flow around the main wing and control flaps, with at least one inlet duct running along the main wing chordwise direction each with at least one inlet, which is located on or underneath the lower surface of the main wing, and at least one outlet duct, which is connected with the at least one inlet duct in a fluid-communicating manner, and has at least one outlet so that fluid which is streaming into the inlet duct can be discharged through the outlet, wherein the outlet is located on the upper surface of the at least one control flap and/or on the upper side of the main wing, on the main wing in the rear third of the length of the main wing with respect to the main wing chordwise direction,wherein the flow control device is arranged at least partially within the fairing, and the inlet of the at least one inlet duct is facing in opposition to the main wing chordwise direction of the main wing, andwherein at least ...

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Номер записи: 24
17-10-2013 дата публикации

AIRCRAFT HAVING AN AIRCRAFT FUSELAGE AND AN AIR SUCKING FUSELAGE COMPONENT

Номер: US20130270390A1
Автор: Schrauf Géza, Stuke Hubert
Принадлежит: AIRBUS OPERATIONS GMBH

An aircraft having an aircraft fuselage that has an outer skin includes an air sucking fuselage component with an outer surface that is perforated at least in some regions, and a suction profile body. The suction profile body is arranged on the outer skin, forms a local bulge in the outer skin, and further includes a suction opening that is arranged at a location at which there is the lowest pressure, for example at a position furthest away from the outer skin. The suction opening is connected to a suction connection of the air sucking fuselage component. In this way laminarization of the flow at the air sucking fuselage component may take place without the use of active air conveying devices. 1. An aircraft having an aircraft fuselage that has an outer skin , comprising:an air sucking fuselage component with an outer surface that is perforated at least in some regions; anda suction profile body arranged on the outer skin, that forms a local bulge on the outer skin, and includes a suction opening that is arranged at a location of the suction profile body at which there is the lowest pressure in an airflow during the flight,wherein the suction opening is connected to a suction connection of the air sucking fuselage component.2. The aircraft of claim 1 , wherein the location of the suction profile body at which the lowest pressure is present during the flight is in a region of the largest distance from the outer skin.3. The aircraft of claim 1 , wherein the suction profile body is arranged on a vertical stabilizer.4. The aircraft of claim 3 , wherein the suction profile body is arranged on an upper delimitation surface of the vertical stabilizer.5. The aircraft of claim 1 , wherein the suction profile body is arranged on the underside of the aircraft fuselage.6. The aircraft of claim 5 , wherein the suction profile body follows on upstream of an outlet valve of an air conditioning system.7. The aircraft of claim 1 , further comprising an air conveying device that is ...

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Номер записи: 25
24-10-2013 дата публикации

FLOW BODY HAVING A LEADING EDGE, A SURFACE AND AN ACTIVE FLOW CONTROL SYSTEM AND VEHICLE COMPRISING AT LEAST ONE SUCH FLOW BODY AND AN AIR SOURCE

Номер: US20130277502A1
Автор: BAUER Matthias, Goelling Burkhard, HAUCKE Frank, NITSCHE Wolfgang
Принадлежит: AIRBUS OPERATIONS GMBH

A flow body having a surface, a leading edge has an active flow control system. The active flow control system includes a plurality of openings, at least one control pressure varying device and at least one fluidic actuator with an interaction chamber having an inlet connectable to an air source, at least two outlets and at least two control pressure ports. The openings are distributed along or parallel to the leading edge in a side-by-side relationship and extend through the surface. The control pressure varying device is connected to the at least two control pressure ports in a fluidic manner, wherein the control pressure varying device is adapted to bring about the flow of the fluid at least majoritarily into a respective one of the outlets. Each of the outlets is connected to one individual opening of the plurality of openings. 1. A flow body having a surface , a leading edge and an active flow control system ,wherein the active flow control system comprises a plurality of openings, at least one control pressure varying device and at least one fluidic actuator with an inlet connectable to an air source, at least first and second outlets and at least first and second control pressure ports, wherein the fluidic actuator is configured such that air from the inlet flows to the at least first and second outlets,wherein the openings are distributed along or parallel to the leading edge in a side-by-side relationship and extend through the surface andwherein the control pressure varying device is connected to the at least first and second control pressure ports in a fluidic manner, wherein the control pressure varying device is adapted to bring about the flow of the fluid at least majoritarily into a respective one of the outlets, and wherein each of the outlets is connected to one individual opening of the plurality of openings.2. The flow body of claim 1 , wherein the fluidic actuator comprises an interaction chamber connecting the inlet with the at least first and ...

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Номер записи: 26
31-10-2013 дата публикации

FLUID ACTUATOR FOR INFLUENCING THE FLOW ALONG A FLOW SURFACE, AS WELL AS BLOW-OUT DEVICE AND FLOW BODY COMPRISING A LIKE FLUID ACTUATOR

Номер: US20130284294A1
Автор: BAUER Matthias, GÖLLING Burkhard, NITSCHE Wolfgang
Принадлежит:

The invention relates to fluid actuator for influencing the flow along a flow surface through ejection of a fluid. By means of a like fluid actuator, a continuous flow is distributed to at least two outlet openings in order to generate fluid pulses out of these outlet openings. Control of this distribution takes place inside an interaction chamber which is supplied with fluid flow via a feed line. Into this interaction chamber there merge at least two control lines via control openings to which a respective different pressure may be applied. Depending on the pressure difference at the control openings, the flow in the interaction chamber is distributed to the individual outlet openings. 1. A fluid actuator for influencing a flow along a flow surface , particularly by pulsating ejection of a fluid flowing through the fluid actuator , comprising a multiplicity of outlet devices each having at least two outlet openings for ejection of the fluid and at least two outlet lines each merging into the outlet openings , the fluid actuator comprising:at least two interaction chambers which are in fluid-communicating connection with separate outlet openings via respective separate outlet lines and which comprise one flow dividing device each in each interaction chamber for dividing the flows into the outlet lines merging from a feed line into the interaction chambers;a control pressure varying device comprising a control flow dividing device to which the control lines are connected for mutually influencing the flow, wherein control lines for supplying fluid at respective different control pressures into the at least one interaction chamber are formed so as to alternatingly bring about the flow of the fluid at least majoritarily into a respective one of the outlet lines, and each control line comprises a feedback line merging into the control flow dividing device, so that if a fluid flow is supplied into the control flow dividing device, an alternating control flow from one ...

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Номер записи: 27
19-12-2013 дата публикации

AIRCRAFT

Номер: US20130336781A1
Автор: ROLT Andrew Martin, THOMPSON Christopher Andrew
Принадлежит: ROLLS-ROYCE PLC

An aircraft includes a propulsive fan arrangement having an intake and an exhaust. The fan arrangement is mounted adjacent a gas washed surface of the aircraft in the form of a suction surface of a wing. The intake is separated from the suction surface to define a channel therebetween. The aircraft further includes a Venturi device positioned downstream of the fan exhaust to draw boundary layer air through the channel. 1. An aircraft comprising:a propulsive fan arrangement having an intake mounted adjacent a gas washed surface of the aircraft which has boundary layer air flow against the surface in use, the intake being separated from the gas washed surface by a channel therebetween; and,a suction device positioned to draw boundary layer air flow through the channel in use.2. An aircraft according to claim 1 , wherein the suction device comprises a Venturi device.3. An aircraft according to claim 1 , wherein the suction device comprises a substantially constant cross section ejector device.4. An aircraft according to claim 1 , wherein the gas washed surface comprises one of a pressure surface and a suction surface of an aerofoil of the aircraft.5. An aircraft according to claim 4 , wherein the aerofoil comprises a wing.6. An aircraft according to . wherein at least part of the suction device is located downstream of the trailing edge of the aerofoil.7. An aircraft according to claim 6 , wherein the Venturi comprises a restriction claim 6 , the restriction being located downstream of the trailing edge of the aerofoil.8. An aircraft according to claim 1 , wherein the suction device has a first inlet in fluid communication with at least the gas washed surface.9. An aircraft according to claim 8 , wherein the first inlet is in fluid communication with both the suction and pressure surfaces.10. An aircraft according to claim 1 , wherein the suction device has a second inlet in fluid communication with a driving airflow source.11. An aircraft according to claim 10 , ...

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Номер записи: 28
23-01-2014 дата публикации

Spin Resistant Aircraft Configuration

Номер: US20140021302A1
Автор: Fox Len, Gionta Matthew, Karkow Jon, Koehler Dieter, Lednicer David, Roncz John
Принадлежит:

A configuration and system for rendering an aircraft spin resistant is disclosed. Resistance of the aircraft to spinning is accomplished by constraining a stall cell to a wing region adjacent to the fuselage and distant from the wing tip. Wing features that facilitate this constraint include but are not limited to one or more cuffs, stall strips, vortex generators, wing twists, wing sweeps and horizontal stabilizers. Alone or in combination, aircraft configuration features embodied by the present invention render the aircraft spin resistant by constraining the stall cell, which allows control surfaces of the aircraft to remain operational to control the aircraft. 1. A configuration for an aircraft so as to be spin resistant , said aircraft comprising a fuselage and a wing wherein the wing includes a first region adjacent to the fuselage and a second region adjacent to a wing tip , the first region being contiguous with the second region , and wherein at high angles of attack a stall cell is constrained to remain within the first region so as to remain apart from flight controls within the second region.2. The configuration for an aircraft as defined in claim 1 , wherein the wing includes a cuff operable to form an aerodynamic boundary between the first region and the second region.3. The configuration for an aircraft as defined in claim 2 , wherein the cuff includes a delta vortex generator.4. The configuration for an aircraft as defined in claim 1 , wherein the first region includes one or more stall strips associated with a leading edge of the wing.5. The configuration for an aircraft as defined in claim 4 , wherein at least one of the one or more stall strips is coupled to the leading edge of the wing and angled downward as it extends toward the wing tip.6. The configuration for an aircraft as defined in claim 5 , wherein at least one of the one or more stalls strips is coupled to and parallel with the leading edge of the wing.7. The configuration for an aircraft ...

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Номер записи: 29
23-01-2014 дата публикации

PROFILE PLATE PORTION FOR USE AS AN OUTER WALL OF A FLOW BODY, METHOD FOR MANUFACTURING A PROFILE PLATE PORTION AND FLOW BODY COMPONENT COMPRISING A SUCTION-EXTRACTION DEVICE FOR FLUID

Номер: US20140021304A1
Автор: Gerber Martin, Syassen Freerk
Принадлежит: AIRBUS OPERATIONS GMBH

A profile plate portion is disclosed for use as an outer wall of a flow body including a first profile plate panel that is fluid permeable, a second profile plate panel extending along the first profile plate panel, and a reinforcing device for supporting the first profile plate panel and the second profile plate panel on one another. Fluid can flow through the reinforcing device, and/or fluid of the flow present at the first profile plate panel, which flows through the first profile plate panel, can flow through the reinforcing device in the local profile plate thickness direction from the first profile plate panel to the second profile plate panel and in some regions can flow through to an inside that is situated opposite the flow side. A method is disclosed for manufacturing a profile plate portion and a flow body component with a suction-extraction device for fluid. 2. The profile plate portion according to claim 1 , characterised in that the reinforcing device is formed from an open-pore metal foam layer that extends between the first profile plate panel and the second profile plate panel.3. The profile plate portion according to claim 1 , characterised in that supporting carriers extend in a spanwise direction of the profile plate portion claim 1 , which supporting carriers are attached to the first profile plate panel and the second profile plate panel claim 1 , in that at least in some of the spaces which in each case form between two adjacent supporting carriers at least one open-pore metal foam layer body is received in such a manner that fluid flowing through the first profile plate panel flows through the open-pore metal foam layer body.4. The profile plate portion according to claim 2 , characterised in that the open-pore metal foam layer bodies rest flat against the respective supporting carriers claim 2 , between which in each case a metal foam layer body is situated.5. The profile plate portion according to claim 2 , characterised in that the open- ...

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Номер записи: 30
06-03-2014 дата публикации

SURFACE ELEMENT FOR AN AIRCRAFT, AIRCRAFT AND METHOD FOR IMPROVING HIGH-LIFT GENERATION ON A SURFACE ELEMENT

Номер: US20140061387A1
Автор: Goelling Burkhard, Lengers Matthias, Reckzeh Daniel
Принадлежит: AIRBUS OPERATIONS GMBH

A surface element, e.g. a wing, of an aircraft includes a leading edge, a high lift device arrangement positioned along the leading edge and at least one add-on body positioned in a leading edge region, wherein the high lift device arrangement is interrupted in the region of at least one add-on body for preventing collision with the add-on body and wherein the surface element includes an arrangement of openings in a region covering the add-on body, which openings are connected to an air conveying device for conveying air through the openings. Thereby an additional flap for harmonizing the flow above a pylon or other add-on body can be eliminated. 1. A surface element for an aircraft , comprising:a leading edge,a high lift device arrangement positioned along the leading edge,at least one add-on body positioned in a leading edge region,wherein the high lift device arrangement is interrupted in the region of at least one add-on body for preventing collision with the add-on body,wherein the surface element comprises an arrangement of openings in a region covering the add-on body, said openings connected to an air conveying device for conveying air through the openings.2. The surface element of claim 1 , wherein the add-on body is a pylon for an engine.3. The surface element of claim 1 , wherein the arrangement of openings is positioned in a region in the proximity of the region with the highest flow instability in terms of separation tendency on the surface element with extended high lift devices.4. The surface element of claim 1 , wherein the openings are selected from a group of openings claim 1 , the group consisting of:at least one bore hole,at least one slit introduced into the surface of the surface element in a direction parallel to the leading edge,at least one slit introduced into the surface of the surface element in a direction normal to the leading edge,at least one slit introduced into the surface of the surface element in a direction at an angle to the ...

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Номер записи: 31
03-04-2014 дата публикации

Systems and Methods for Attenuation of Noise and Wakes Produced by Aircraft

Номер: US20140091180A1
Автор: Khodadoust Abdollah, Shmilovich Arvin
Принадлежит: The Boeing Company

Systems and methods for reducing the trailing vortices and lowering the noise produced by the side edges of aircraft flight control surfaces, tips of wings and winglets, and tips of rotor blades. A noise-reducing, wake-alleviating device is disclosed which incorporates an actuator and one or more air-ejecting slot-shaped openings coupled to that actuator and located on the upper and/or lower surfaces and/or the side edges of an aircraft flight control surface or the tip of a wing, winglet or blade. The actuation mechanism produces sets of small and fast-moving air jets that traverse the openings in the general streamwise direction. The actuation destabilizes the flap vortex structure, resulting in reduced intensity of trailing vortices and lower airplane noise. 1. An aircraft comprising an aerodynamic element , a source of pressurized air , an air jet actuator and a controller , wherein said aerodynamic element comprises a side edge and an opening located on or near the side edge and generally aligned with a streamwise direction; said air jet actuator comprises a rotatable element , said rotatable element comprising an interior duct and an opening in fluid communication with said interior duct; and said controller is operable to cause said interior duct of said rotatable element to be placed in fluid communication with said pressurized air source and also cause said rotatable element to rotate , as a result of which said interior duct of said rotatable element is in fluid communication with said opening of said aerodynamic element via said opening in said rotatable element , thereby enabling pressurized air from said source to exit said opening of said aerodynamic element in the form of an air jet.2. The aircraft as recited in claim 1 , wherein said opening of said rotatable element comprises a helical slot claim 1 , said opening of said control surface comprises a first slot claim 1 , and said helical slot of said rotating rotatable element causes said air jet to ...

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Номер записи: 32
10-04-2014 дата публикации

SHAPE MEMORY RIBLETS

Номер: US20140099475A1
Автор: Rawlings Diane C., SCHNEIDER TERRY L.
Принадлежит: The Boeing Company

A multilayer construction for an array of aerodynamic riblets incorporates a first layer composed of a material with protuberances, the first layer material having shape memory and a second layer composed of a material exhibiting a second characteristic with capability for adherence to a surface. 1. A multilayer construction for an array of riblets comprising:a first layer composed of a material with aerodynamic riblets, formed from a superelastic shape memory material in an austenitic form in a non-stress state;a second layer composed of a material exhibiting a second characteristic of capability for adherence to a surface.2. The multilayer construction for an array of riblets as defined in wherein:the riblets comprise a plurality of tips each with an integral base extending therefrom, said tips formed from a superelastic shape memory material in an austenitic form in a non-stress state;and said second layer includes a polymer layer engaging said tips in predetermined spaced relation, said polymer layer adhering to a vehicle surface.3. The multilayer construction for an array of riblets as defined in wherein the tips are formed from shape memory alloy (SMA) material selected from the set of copper-zinc-aluminum-nickel claim 2 , copper-aluminum-nickel claim 2 , nickel-titanium (NiTi) and nickel-free claim 2 , pseudo-elastic beta titanium alloy.4. The multilayer construction for an array of riblets as defined in wherein the first layer includes a surface layer continuously cast with the tips.5. The multilayer construction for an array of riblets as defined in further comprising a polymer support layer deposited on the surface layer opposite the tips.6. The multilayer construction for an array of riblets as defined in further comprising an adhesive layer deposited on the polymer support layer to form a multilayer appliqué claim 5 , said adhesive layer adhering the appliqué to the vehicle surface.7. The multilayer construction for an array of riblets as defined in ...

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Номер записи: 33
05-01-2017 дата публикации

TILE ASSEMBLY

Номер: US20170001715A1
Автор: MIRZA Nisar Ahmed, WARSOP Clyde
Принадлежит:

A tile assembly () which, in use, is fitted to a base structure to form at least part of a fluid washed surface. The tile assembly comprises a housing () with at least one plenum () being provided within the housing (). A wall () of the housing () is provided with a plurality of flow passages () which extend from the plenum side of the wall () to an outer surface () of the wall. Flow passage closures () are provided which are operable to open and close at least some of the flow passages (). 1. A tile assembly which , in use , is fitted to a base structure to form at least part of a fluid washed surface; the tile assembly comprising:a housing;at least one plenum being provided within an interior of the housing;a wall of the housing provided with a plurality of flow passages extending from an interior, plenum side of the wall to an outer surface of the wall, said wall forming at least part of the fluid washed surface; anda plurality of flow passage closures operable to open and close at least some of the flow passages.2. A tile assembly as claimed in wherein the housing is provided with a mechanical fixing for cooperation with a complementary fixing on the base structure to thereby lock the tile assembly to the base structure.3. A tile assembly as claimed in wherein the mechanical fixing is configured to be dis-engagable claim 2 , such that the tile assembly may be removed from the base structure.4. A tile assembly as claimed in claim 1 , wherein the tile assembly further comprises a dedicated fluid supply configured to deliver a fluid to the plenum.5. A tile assembly as claimed in claim 1 , further comprising a duct located in a wall of the housing and configured for delivery of a fluid to the plenum. Preliminary Amendment6. A tile assembly as claimed in claim 5 , wherein the duct comprises a fluid coupling configured for engagement with a complementary fluid coupling provided on the base structure.7. A tile assembly as claimed in claim 1 , wherein the housing is ...

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Номер записи: 34
02-01-2020 дата публикации

AIRCRAFT GENERATING LARGER THRUST AND LIFT BY FLUID CONTINUITY

Номер: US20200001980A1
Автор: ZHU Xiaoyi
Принадлежит:

The invention discloses an aircraft generating a larger thrust and lift by fluid continuity. First open channels used to extend fluid paths are formed in front parts and/or middle parts of windward sides of wings of the aircraft and extend from sides, close to the fuselage, of the wings to sides, away from the fuselage, of the wings, and the first open channels are concave channels or convex channels, so that a pressure difference in a direction identical with a moving direction is generated from back to front due to different flow speeds of fluid flowing over the windward sides of the wings in a lengthwise direction and a widthwise direction to reduce fluid resistance, and a larger pressure difference and lift are generated due to different flow speeds on the windward sides and leeward sides of the wings. 1. An aircraft , comprising a fuselage and wings , wherein first open channels used to extend fluid paths are formed in windward sides of the wings and extend from roots of sides , close to the fuselage , of the wings to tails of sides , away from the fuselage , of the wings , and the first open channels are concave channels or convex channels , so that a larger lift is generated due to different flow speeds of fluid flowing over the windward sides of the wings in a lengthwise direction and flowing over leeward sides of the wings in a widthwise direction.2. The aircraft according to claim 1 , wherein the first open channels are formed in front parts of the windward sides of the wings claim 1 , or the first open channels are formed in the front parts and middle parts of the windward sides of the wings claim 1 , so that a pressure difference-based thrust in a direction identical with a moving direction is generated from back to front due to different flow speeds of the fluid flowing through the front parts or middle parts of the wings and flowing through rear parts of the windward sides of the wings.3. The aircraft according to claim 1 , wherein the first open ...

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Номер записи: 35
02-01-2020 дата публикации

JET FLOW CONTROL MECHANISM AND METHOD OF USE

Номер: US20200001981A1
Автор: Alhussan Khaled Abdullah
Принадлежит:

A transonic aircraft includes a frame body extending from a fuselage to a rear tail wing section; an engine placed between the fuselage and rear tail wing section, the engine is secured to the frame body; a wing section includes a wing body with an upper surface and a lower surface that extend from a leading edge to a trailing edge, the wing body is oriented at an angle relative to an elongated length of the frame body; and a flow separation control device secured to the wing section. The flow separation control device includes a plurality of openings on the upper surface of the wing body. 1. An aircraft , comprising:a frame body extending from a fuselage to a rear tail wing section;an engine placed between the fuselage and rear tail wing section, the engine is secured to the frame body;a wing section; anda flow separation control device secured to the wing section, the flow separation control device;wherein a portion of the airstream passing over the upper surface of the wing body is affected by the plurality of openings; andwherein the plurality of openings reduces a flow separation in the portion of the airstream passing over the upper surface of the body.2. The aircraft of claim 1 , the wing section comprising:a wing body with an upper surface and a lower surface that extend from a leading edge to a trailing edge, the wing body is oriented at an angle relative to an elongated length of the frame body.3. The aircraft of claim 2 , the flow separation device claim 2 , having:a plurality of openings on the upper surface of the wing body;4. The aircraft of claim 3 , wherein the flow separation device extends the longitudinal length of the wing body.5. The aircraft of claim 3 , wherein the plurality of openings extend from a ⅓ chord length to a ⅔ chord length of the wing section.6. The aircraft of claim 1 , wherein the plurality of openings extend from a ⅓ chord length to a ⅔ chord length of the wing section.7. The aircraft of claim 1 , wherein the plurality of ...

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Номер записи: 36
02-01-2020 дата публикации

SHORT TAKE OFF AND LANDING AIRCRAFT WITH ADJUSTABLE VORTICES DEVICE

Номер: US20200001982A1
Автор: Utt Larry
Принадлежит:

An improved aircraft system is provided. The improved aircraft system comprises an adjustable vortices device that may be attached to an aircraft to create various vortices effects, which increase take-off weight and improve low-speed handling of the aircraft. The adjustable vortices device comprises a linear actuator, a pivot mechanism, and a vortex generator. The pivot mechanism is operably connected to the linear actuator in a way such that the translational energy of the linear actuator causes the pivot mechanism to rotate about a central axis. The vortex generator is moveably attached to a surface of the aircraft and coupled to the pivot mechanism in a way such that rotating the pivot mechanism causes the vortex generator to rotate about a central axis, which alters the angle the vortex generators move through the air. 1. An aircraft system comprising:a fuselage, [ an internal structure,', 'an upper surface coupled to said internal structure, and', 'a lower surface coupled to said internal structure,', 'wherein said streamlined airfoil-contoured body is arranged such that said fluid moves at a higher average velocity over said upper surface and at a lower average velocity over said lower surface,, 'a streamlined airfoil-contoured body that generates lift when propelled through a fluid at an angle of incidence of at least zero, wherein said streamlined airfoil-contoured body is defined by,'}, a linear actuator configured to convert rotational energy into translational energy,', 'wherein said translational energy of said linear actuator causes said pivot mechanism to rotate about a central axis, and', 'a pivot mechanism operably connected to said linear actuator,'}, 'wherein rotating said pivot mechanism causes said vortex generator to rotate about said central axis,', 'a vortex generator coupled to said pivot mechanism,'}], 'an adjustable vortices device operably connected to said streamlined airfoil-contoured body, wherein said adjustable vortices device ...

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Номер записи: 37
07-01-2021 дата публикации

PROFILED STRUCTURE AND ASSOCIATED TURBOMACHINE

Номер: US20210003074A1
Автор: GEA AGUILERA Fernando, GRUBER Mathieu Simon Paul, Riou Georges Jean Xavier
Принадлежит: SAFRAN AIRCRAFT ENGINES

An airflow profiled structure having a profiled leading edge. The profiled leading edge having, along a leading edge line, a serrated profile line with a succession of teeth and depressions. The airflow profiled structure also includes a porous acoustically absorbent region located towards the bottom of the depressions. 1. A profiled air flow structure comprising:a body;porous acoustically absorbent regions;an upstream leading edge and/or a downstream trailing edge; andalong the upstream leading edge and/or the downstream trailing edge line, a serrated profile line showing a succession of teeth and depressions,wherein the porous acoustically absorbent regions locally form bottoms for the depressions where the porous acoustically absorbent regions occupy a part of the body to define, together with the body, the serrated profile line at the upstream leading edge and/or the downstream trailing edge.2. The profiled structure according to further comprising: along the chord, the serrated profile line has a maximum amplitude, h, and', {'br': None, 'i': 'd=h/', '10, within 30%.'}, 'the porous acoustically absorbent region has a geometric centre located at a distance d downstream of the upstream leading edge or upstream of the downstream trailing edge, at the bottom of the depressions such that], 'between upstream and downstream, a chord in which3. The profiled structure according to further comprising: along the upstream leading edge or the downstream trailing edge, the serrated profile line has a distance between two consecutive tooth tips,', 'along the chord, the serrated profile line has a maximum amplitude, h, and', along the upstream leading edge and/or the downstream trailing edge, two limits separated by a distance a such that a is equal to one third of said distance between two consecutive tooth tips, to within 30%,', 'along the chord, two limits separated by a distance b such that b=h/3, within 30%., 'the porous acoustically absorbent region has], 'between ...

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Номер записи: 38
13-01-2022 дата публикации

Active Lift Control Device and Method

Номер: US20220009618A1
Автор: John A. Cooney, Jr., Neal E. Fine
Принадлежит: Arctura Inc

A lift control device actively controls the lift force on a lifting surface. The device has a protuberance near a trailing edge of its lifting surface, which causes flow to separate from the lifting surface, generating regions of low pressure and high pressure which combine to increase the lift force on the lifting surface. The device further includes a means to keep the flow attached around the protuberance or to modify the position of the protuberance in response to a command from a central controller, so as to provide an active control of the lift between a maximum value and a minimum value.

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Номер записи: 39
13-01-2022 дата публикации

VERTICAL TAKE OFF AND LANDING AIRCRAFT WITH FLUIDIC PROPULSION SYSTEM

Номер: US20220009627A1
Автор: Evulet Andrei
Принадлежит:

An aircraft includes a fuselage and a primary airfoil having a first upper surface. The first upper surface has a recess disposed therein. A conduit is in fluid communication with recess. An ejector is disposed within the recess. The ejector is configured to receive compressed air via the conduit. The ejector is further configured to produce a propulsive efflux stream. A secondary airfoil is coupled to the primary airfoil and has a second upper surface. The ejector is positioned such that the efflux stream flows over the second surface. The second surface is oriented so as to entrain the efflux stream to flow in a direction substantially perpendicular to the first upper surface. 1a fuselage;at least one primary airfoil having a first upper surface, the first upper surface having at least one recess disposed therein;at least one conduit in fluid communication with the at least one recess;at least one ejector disposed within the at least one recess, the at least one ejector configured to receive compressed air via the at least one conduit, the at least one ejector configured to produce a propulsive efflux stream; andat least one secondary airfoil coupled to the at least one primary airfoil and having a second upper surface, the at least one ejector being positioned such that the efflux stream flows over the second surface, the second surface being oriented so as to entrain the efflux stream to flow in a direction substantially perpendicular to the first upper surface.. An aircraft, comprising: This application claims priority to U.S. Provisional Application No. 63/016,226, filed Apr. 27, 2020. This application is a continuation-in-part of U.S. application Ser. No. 16/748,560 filed Jan. 21, 2020, which claims the benefit of U.S. Provisional Application No. 62/794,464 filed Jan. 18, 2019.This application is a continuation-in-part of U.S. application Ser. No. 16/16/680,479 filed Nov. 11, 2019 and U.S. application Ser. No. 16/681,555 filed Nov. 12, 2019, each of which ...

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Номер записи: 40
14-01-2016 дата публикации

System and Method for Distributed Active Fluidic Bleed Control

Номер: US20160009374A1
Автор: Ari Glezer, John M. Kearney
Принадлежит: Georgia Tech Research Corp

A system and method for regulating and actuating bleed over a structure exposed in a fluid motion are disclosed. The bleed inlet and outlet are formed on the surface of the structure establishing fluidic communication across surfaces. The disclosed system and method contemplates active control and regulation of the bleed to modify crossflow properties such as, aerodynamic forces, hydrodynamic forces, vorticity, and moments.

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Номер записи: 41
14-01-2016 дата публикации

Semi-open Fluid Jet VTOL Aircraft

Номер: US20160009376A1
Автор: Bucheru Bogdan Tudor
Принадлежит:

The herein invention is presenting a lift generating method based on a semi-open fluid jet flowing in a closed circuit around a lifting airfoil. A VTOL aircraft with maximized pay load room and car-like shape is the preferred embodiment of the invention. The herein aircraft uses no wings, exposed propellers, hot gas jets or other high injury risk means for propulsion and life, and it can be driven by ordinary skilled people. Furthermore the aircraft has a small footprint and can land, take off and even cruise on water in one of the preferred embodiments. 1. A lifting method using a regenerative semi-open fluid jet with an airfoil surface that creates lift by circulating on an expose to the outer medium airfoil surface and returns on the closed side of the airfoil to the source of the jet , where the source is mounted between the closed side of the airfoil and a containing body , which containment body hermetically contains the fluid jet on the return path and allows the exposure of the fluid jet to the outer medium on the open side of the airfoil.2. A lifting surface apparatus that is based on the method of claim 1 , comprising of an airfoil that is rigidly connected to an outer body in such way that a fluid jet generated by a jet source contained between the airfoil and the outer body can circulate in a close path around the said airfoil; and the containing body hermetically guides the fluid jet on the return path around the closed side of the airfoil and allows the fluid jet to be in contact with the external medium when flows on the open side of the airfoil claim 1 , therefore creating lift on that open side of the airfoil.3. A lifting element comprising one or more lifting surfaces as described in claim 2 , which can achieve VTOL and horizontal flying when connected to a power source and control unit claim 2 , and furthermore can be the propulsion system of a car-like aircraft when connected to a car-like cabin including the power source claim 2 , control unit ...

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Номер записи: 42
14-01-2016 дата публикации

POSITIONABLE EJECTOR MEMBER FOR EJECTOR ENHANCED BOUNDARY LAYER ALLEVIATION

Номер: US20160009377A1
Автор: Khalid Syed J.
Принадлежит: Rolls-Royce Corporation

An aircraft having a moveable ejector member for assisting in alleviating a boundary layer flowing along an aircraft surface is disclosed. The moveable ejector member is capable of being placed at a variety of positions between a fully open position and a nested position to entrain the boundary layer with another fluid flow. The ejector member can take the form of an ejector shroud used with a nacelle. In some forms, a gas turbine engine is used to provide an ejector flow to entrain the boundary layer through the flow path. 1. An apparatus comprising:an aircraft having a power plant useful to discharge a relatively high velocity fluid stream to provide a force useful in affecting a movement of the aircraft, the fluid stream being useful as a primary flow stream of an ejector;a flow surface over which a working fluid passes to form a boundary layer as the aircraft is in motion;an airflow member structured to be moveable relative to the flow surface between a relatively open position and a relatively closed position, the airflow member capable of producing a gap between a portion of the airflow member and the flow surface, wherein the gap serves to capture at least a portion of the boundary layer from the flow surface and provide it as a secondary flow stream of the ejector; anda controller structured to operate upon a determined condition and actuate the airflow member to a variety of positions as a function of the determined condition to control the ejector, the variety of positions corresponding to respective gap distances appropriate to entrain the at least a portion of the boundary layer.2. The apparatus of claim 1 , wherein the flow surface comprises a nacelle of the power plant.3. The apparatus of claim 2 , wherein the power plant is a gas turbine engine and the high velocity fluid stream is an exhaust of the gas turbine engine.4. The apparatus of claim 3 , wherein the airflow member comprises an ejector shroud claim 3 , wherein the relatively closed position ...

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Номер записи: 43
14-01-2016 дата публикации

Stable Low Aspect Ratio Flying Wing

Номер: US20160009391A1
Автор: Friesel Eric Walter
Принадлежит:

A low aspect ratio flying wing provides aerodynamic stability throughout the flight envelope with improved aerodynamic efficiency. Insufficient stability and reduced aerodynamic efficiency typical of low aspect ratio flying wings is improved through wing design and proper application and placement of horizontal stabilizers and boundary layer control. Lateral asymmetric boundary layer manipulation is employed to alter flying wing orientation in flight. Lateral extension and retraction of the main structure wing optimizes efficiency. This novel flying wing is not found in literature or “prior art” and provides improvement in aerodynamic stability and efficiency over previous designs. Given the large amount of research, literature, patents and activity in the field since the 1930's and the absence of a practical design indicates the non-obvious nature of these disclosures. In addition, those skilled in the art teach away from present disclosures failing to realize the better than predicted advantages. 1. A flying wing comprising:An aspect ratio greater than 0.5 and less than 1.5; andA main structure of a wing, said main structure wing with a center of gravity located forward of 50 percent chord, said main structure wing aspect ratio greater than 0.5 and less than 1.5, said main structure wing is primary lifting airfoil; andHorizontal stabilizers mirrored about a central plane, said central plane is defined by the longitudinal and vertical axis of the main structure wing, said horizontal stabilizers extending outboard of the lateral extents of the main structure wing by a distance greater than 10 percent of the main structure wing longest chord length, said horizontal stabilizers offset vertically from the main structure wing chord line by a distance greater than 10 percent of the main structure wing longest chord length, and said horizontal stabilizers positioned longitudinally aft of the main structure wing 50 percent chord point.2. A flying wing of further comprising ...

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Номер записи: 44
10-01-2019 дата публикации

APPLICATOR

Номер: US20190009297A1
Автор: Huber Helmut, Sturm Thomas
Принадлежит:

An applicator for application onto and embossing microprofiling of a fluidic medium on a substrate, in particular in the aerospace sector, and a corresponding application device having such an applicator. The applicator has a circumferentially moving die that has an embossing profile, a press for the die and a stabilizing device, in particular a hardening device, for the applied medium. In addition, the applicator has a hollow support body, surrounded by the die at a distance forming a gap, the press being arranged in the gap. The application device has, in addition to the applicator, a handling device for a relative movement between the applicator and a workpiece. 1. An applicator for applying and embossing a microprofile on a fluidic medium on an outer skin of an aircraft , the applicator comprising:a circumferentially movable die comprising an embossing profile for embossing the microprofile in the fluidic medium on the an outer skin of the aircraft;a press for the die;a frame connected in a rotationally locked manner to the die; anda feeding device;wherein the die is configured to receive the fluidic medium from the feeding device using the embossing profile when the die moves circumferentially,wherein the applicator is configured to roll the die over the an outer skin of the aircraft to transfer the fluidic medium to the an outer skin of the aircraft, andwherein the applicator comprises a hollow support body, the support body being rotatably mounted on the frame of the applicator.2. The applicator according to claim 1 , wherein the applicator comprises its own drive for circumferential movement of the die.3. The applicator according to claim 2 , wherein the support body is coupled to the drive.4. The applicator according to claim 1 , wherein the microprofile comprises a microstructure having elevations and indentions in a form of at least one stripe.5. The applicator according to claim 1 , wherein the die is configured to emboss a plurality of parallel stripes ...

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Номер записи: 45
10-01-2019 дата публикации

High efficiency stall proof airfoil and means of control

Номер: US20190009890A1
Автор: George Paul Baker, Jr.
Принадлежит: Individual

A high-efficiency, stall-proof airfoil is an aircraft wing configuration whereby a motive force directly induces gaseous fluid flow across a lifting surface of the airfoil without requiring a movement of the wing through an air space. The airfoil is provided with means to control a pitch, a roll and a yaw motion and to control a position and stability of the aircraft. When not undergoing horizontal displacement, it provides highly efficient use of fuel resources, precluding the formation of drag and its incumbent power consumption. Air pressure at a bottom of the wing remains essentially ambient. Therefore, differential pressure between a lower surface of the wing and an upper surface of the wing maintains its maximum possible quantity. Virtually, all of the power consumed is utilized in a production of lift. Additionally, because lift is generated without regard to an angle-of-attack, forward speed, nor a configuration of a leading edge of the wing, the configuration is essentially stall proof.

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Номер записи: 46
09-01-2020 дата публикации

ACTIVE LAMINAR FLOW CONTROL SYSTEM WITH COMPOSITE PANEL

Номер: US20200010175A1
Автор: Paolini Thomas, Sanz Martinez Pablo T., Schrell Johann S.
Принадлежит:

An assembly is provided for active laminar flow control. This assembly includes a panel, which panel includes an outer skin, an inner skin and a plurality of plenums between the outer skin and the inner skin. Each of the plurality of plenums is fluidly coupled with a respective array of perforations through the outer skin. The panel is constructed from fiber-reinforced composite material. 1. An assembly for active laminar flow control , comprising a panel comprising an outer skin , an inner skin and a plurality of plenums between the outer skin and the inner skin , each of the plurality of plenums fluidly coupled with a respective array of perforations through the outer skin , wherein the panel is constructed from fiber-reinforced composite material.2. The assembly of claim 1 , further comprising a suction system fluidly coupled with one or more of the plurality of plenums.3. The assembly of claim 1 , wherein the panel is configured from the fiber-reinforced composite material as a monolithic body.4. The assembly of claim 1 , wherein the panel further comprises a plurality of corrugations arranged between the outer skin and the inner skin claim 1 , and the plurality of corrugations form sidewalls of the plurality of plenums.5. The assembly of claim 4 , whereina first of the plurality of corrugations comprises a bridge, a first sidewall, a second sidewall, a first flange and a second flange;the bridge extends between the first sidewall and the second sidewall, and is connected to the inner skin;the first sidewall and the second sidewall each extend between the inner skin and the outer skin;the first flange projects out from the first sidewall and is connected to the outer skin; andthe second flange projects out from the second sidewall and is connected to the outer skin.6. The assembly of claim 4 , wherein a first of the plurality of corrugations is bonded to at least one of the outer skin and the inner skin.7. The assembly of claim 4 , whereinthe panel further ...

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Номер записи: 47
09-01-2020 дата публикации

TRANSVERSE FAN PROPULSION SYSTEM

Номер: US20200010189A1
Автор: Silkowski Peter Daniel
Принадлежит: GENERAL ELECTRIC COMPANY

A transverse fan propulsion system includes a first transverse fan configured to rotate in a first direction, and a second transverse fan configured to rotate in a second direction. The second direction is opposite to the first direction. The second transverse fan is located coaxially with and radially inward of the first transverse fan. 1. A transverse fan propulsion system comprising:a first transverse fan configured to rotate in a first direction;a second transverse fan configured to rotate in a second direction, the second direction is opposite to the first direction; andwherein the second transverse fan is located coaxially with and radially inward of the first transverse fan.2. The transverse fan propulsion system of claim 1 , wherein the first transverse fan and the second transverse fan are located in an aircraft.3. The transverse fan propulsion system of claim 1 , wherein the first transverse fan and the second transverse fan are located in an airfoil or a wing of an aircraft.4. The transverse fan propulsion system of claim 1 , further comprising:a gas turbine core located downstream from the first transverse fan and the second transverse fan, the gas turbine core located in a central region of an exhaust corridor of the first transverse fan and the second transverse fan;a variable bypass channel modifying exhaust amounts from the first transverse fan and the second transverse fan to an input of the gas turbine core, the variable bypass channel having moveable doors that admit a greater amount of the exhaust when in a first position, and a lesser amount of exhaust when in a second position.5. The transverse fan propulsion system of claim 1 , further comprising:a gas turbine core located downstream from the first transverse fan and the second transverse fan, the gas turbine core located in a central region of an exhaust corridor of the first transverse fan and the second transverse fan;a variable bypass channel modifying exhaust amounts from the first ...

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Номер записи: 48
11-01-2018 дата публикации

NACELLE AND METHOD FOR INFLUENCING FLUID FLOWS IN A NACELLE

Номер: US20180010518A1
Автор: ORTMANNS Jens
Принадлежит:

The invention relates to an engine nacelle, including: a nacelle wall that has an inner side and an outer side; an inlet lip that is embodied at that end of the engine nacelle that is formed upstream; and an engine intake that takes in the air required for the respective engine and that is formed by the inner side of the nacelle wall. It is provided that the nacelle wall includes an air-permeable structure that extends from the outer side to the inner side of the nacelle wall, and that is configured for passing air that flows against the outer side from the outer side to the inner side. The invention further relates to a method for influencing the flows inside an engine nacelle. 2. The engine nacelle according to claim 1 , wherein the air-permeable structure comprises a plurality of tubes that respectively extend to the inner side.3. The engine nacelle according to claim 2 , wherein the tubes form a one-dimensional or a two-dimensional array.4. The engine nacelle according to claim 2 , wherein the tubes are respectively formed as a nozzle in the direction of their ends which are facing towards the inner side.5. The engine nacelle according to claim 1 , wherein the air-permeable structure is formed by a porous material or contains a porous material.6. The engine nacelle according to claim 1 , wherein the air-permeable structure is formed in such a manner that it has a defined blow-in direction.7. The engine nacelle according to claim 6 , wherein the preferred blow-in direction extends substantially transversely to the longitudinal direction of the engine nacelle.8. The engine nacelle according to claim 2 , wherein the tubes extend adjacent to the outer side substantially transversely to the longitudinal direction of the engine nacelle.9. The engine nacelle according to claim 1 , wherein the air-permeable structure is formed in such a manner that it has a defined blow-out direction.10. The engine nacelle according to claim 9 , wherein the blow-out direction has a ...

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Номер записи: 49
03-02-2022 дата публикации

DRAG RECOVERY SCHEME USING BOUNDARY LAYER INGESTION

Номер: US20220033067A1
Автор: Page Mark Allan
Принадлежит: Blended Wing Aircraft, Inc.

Technologies are described herein for a drag recovery scheme using a boundary layer bypass duct system. In some examples, boundary layer air is routed around the intake of one or more of the engines and reintroduced aft of the engine fan in the nozzle duct in a mixer-ejector scheme. Mixer-ejectors mix the boundary layer flow to increase mass flow. 1. An aircraft , comprising:a blended wing body;at least one fan housed within a nacelle;at least one bypass intake duct configured to receive boundary layer air from a top surface of the blended wing body, the at least one bypass intake duct located proximate to a fan intake of the at least one fan;at least one bypass exhaust duct located proximate to a fan exhaust of the at least one fan; anda passageway, located substantially between the blended wing body and the at least one fan, fluidically connecting the at least one bypass intake duct with the at least one bypass exhaust duct and configured to direct the boundary layer air from the at least one bypass intake duct to the at least one bypass exhaust duct.2. The aircraft of claim 1 , wherein the at least one bypass exhaust duct is configured to output the boundary layer air into the fan exhaust of the at least one fan.3. The aircraft of claim 2 , wherein the at least one bypass exhaust duct has a duct height and the at least one bypass exhaust duct is located a distance no less than the duct height from a nozzle exit within a nozzle of the at least one fan.4. The aircraft of claim 2 , wherein the at least one bypass exhaust duct is configured to mix the boundary layer air and the fan exhaust of the at least one fan using a turbulent mixing cone.5. The aircraft of claim 2 , wherein the boundary layer air exits the at least one bypass exhaust duct into a mixing region in the fan exhaust of the at least one fan.6. The aircraft of claim 3 , wherein the mixing region is within the nacelle.7. The aircraft of claim 1 , wherein the nacelle is semi-buried.8. The aircraft of ...

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Номер записи: 50
03-02-2022 дата публикации

SELF-ADHERING FILM WITH AERODYNAMIC PERFORMANCE

Номер: US20220033068A1
Автор: Mikami Haruyuki, Okura Ken, Tachihara Takayuki
Принадлежит:

Provided is the film that can reduce aerodynamic drag and enhance aerodynamic performance. The film according to an embodiment is a film () to be attached to a moving body that moves in a predetermined moving direction, extends along a second direction (D) being the moving direction, and includes recesses and protrusions (A) configured to enhance aerodynamic performance of the moving body on a surface of the film. 1. A film to be attached to a moving body that moves in a predetermined moving direction , the film comprises:recesses and protrusions configured to enhance aerodynamic performance of the moving body on a surface of the film.2. The film according to claim 1 , further comprising a hydrophilic coating layer configured to coat the recesses and protrusions.3. The film according to claim 1 , further comprising a hydrophobic coating layer configured to coat the recesses and protrusions.4. The film according to claim 1 , further comprising an adhesive agent layer configured to cause the film to adhere to the moving body.5. The film according to claim 4 , further comprising an intermediate layer that is positioned between the recesses and protrusions and the adhesive agent layer. One aspect of the present disclosure relates to a film.Patent Document 1 describes a method of reducing a resistance force and a resistance-force reduction item. As the resistance-force reduction item, a sheet material is described. The sheet material includes a pattern surface on a front surface, and a cross section of a pattern layer is a serrated cross section having a plurality of mountains and a plurality of valleys. Further, the sheet material including an adhesive layer on a surface opposite to the pattern surface is described. The sheet material reduces a resistance force of an item when the adhesive layer is attached to the surface of the item.Incidentally, a moving body, for example, a vehicle, an airplane, a blade of a wind power plant, or the like exerts a function of ...

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Номер записи: 51
18-01-2018 дата публикации

AIRCRAFT HAVING AN AIRFLOW DUCT

Номер: US20180016000A1
Автор: Pastouchenko Nikolai N., Yao Jixian
Принадлежит:

An aircraft is provided including a propulsor having a plurality of fan blades. The aircraft additionally includes a fuselage, with the propulsor attached to the fuselage at an aft end of the fuselage. A stabilizer extends away from the fuselage proximate the aft end, the stabilizer including a portion positioned upstream of the plurality of fan blades. The aircraft additionally includes an airflow duct extending between an inlet and an outlet. The inlet is positioned to receive an airflow from a location outside the fuselage of the aircraft. The outlet is positioned to exhaust the airflow to at least partially offset a wake upstream of the plurality of fan blades, the wake generated by the stabilizer during operation of the aircraft. 1. An aircraft comprising:a propulsor comprising a plurality of fan blades;a fuselage extending between a forward end and an aft end, the propulsor attached to the fuselage at the aft end of the fuselage;a stabilizer extending away from the fuselage proximate the aft end of the fuselage, the stabilizer including at least a portion positioned upstream from the plurality of fan blades; andan airflow duct extending between an inlet and an outlet, the inlet positioned to receive an airflow from a location outside the fuselage of the aircraft, the outlet positioned to exhaust the airflow to at least partially offset a wake upstream of the plurality of fan blades of the propulsor, the wake generated by the stabilizer during operation of the aircraft.2. The aircraft of claim 1 , wherein the inlet of the airflow duct is defined by the fuselage of the aircraft.3. The aircraft of claim 1 , wherein the inlet of the airflow duct is defined by the fuselage of the aircraft at a bottom half of the fuselage.4. The aircraft of claim 1 , wherein the stabilizer is a vertical stabilizer attached to the fuselage.5. The aircraft of claim 1 , wherein the stabilizer defines an aft end claim 1 , and wherein the outlet of the airflow duct is defined in the ...

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Номер записи: 52
19-01-2017 дата публикации

Swirling Jet Actuator for Control of Separated and Mixing Flows

Номер: US20170016462A1
Автор: Alvi Farrukh, Munday Phillip, Taira Kunihiko
Принадлежит:

The present invention includes a method of controlling a fluid flow using momentum and/or vorticity injections. Actively controlling an actuator allows for direct, precise, and independent control of the momentum and swirl entering into the fluid system. The perturbations are added to the flow field in a systematic mater providing tunable control input, thereby modifying behavior thereof in a predictable manner to improve the flow characteristics. 1. A method of controlling a fluid flow , comprising the steps of:inputting a momentum flow into the fluid flow;inputting a swirling flow into the fluid flow; andvarying the inputting of the swirling flow and the inputting of the momentum flow independently of one another.2. The method of claim 1 , wherein the momentum flow is inputted in an orientation that is normal to a surface of a body over which the fluid flow is passing.3. The method of claim 1 , wherein the swirling flow is inputted in an orientation such that a central axis claim 1 , about which the swirling flow rotates claim 1 , is normal to a surface of a body over which the fluid flow is passing.4. The method of claim 1 , wherein the momentum flow is inputted in an orientation that is normal to a surface of a body over which the fluid flow is passing claim 1 , and the swirling flow is inputted in an orientation such that a central axis claim 1 , about which the swirling flow rotates claim 1 , is normal to the surface of the body over which the fluid flow is passing.5. The method of claim 1 , wherein the momentum flow is actively inputted.6. The method of claim 1 , wherein the swirling flow is actively inputted.7. The method of claim 1 , wherein the momentum flow and swirling flow are actively inputted.8. The method of claim 1 , wherein the inputting occurs near the time-averaged separation point on a body over which the fluid flow is passing.9. The method of claim 1 , wherein the inputting occurs at a plurality of actuator sites such that each actuator site ...

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Номер записи: 53
17-01-2019 дата публикации

INTEGRATED SLAT CHINE APPARATUS AND METHODS

Номер: US20190016439A1
Автор: Leopold David Daniel
Принадлежит:

Integrated slat chine apparatus and methods are described. An example apparatus includes a chine and a slat. The chine is coupled to an airfoil. The chine includes a lateral surface. The slat is located adjacent the lateral surface of the chine and coupled to the airfoil. The slat is movable relative to the airfoil between a stowed position and a deployed position. The slat is to expose the lateral surface of the chine when the slat is in the deployed position and to cover the lateral surface of the chine when the slat is in the stowed position. 1. An apparatus , comprising:a chine coupled to an airfoil, the chine having a lateral surface; anda slat located adjacent the lateral surface of the chine and coupled to the airfoil, the slat being movable relative to the airfoil between a stowed position and a deployed position, the slat to expose the lateral surface of the chine when the slat is in the deployed position and to cover the lateral surface of the chine when the slat is in the stowed position.2. The apparatus of claim 1 , wherein the chine is located at a leading edge of the airfoil.3. The apparatus of claim 2 , wherein the chine extends from the leading edge of the airfoil in a first direction that is parallel to a second direction claim 2 , the slat being moveable relative to the airfoil along the second direction.4. The apparatus of claim 1 , wherein the chine is located outboard of a nacelle coupled to the airfoil.5. The apparatus of claim 1 , wherein a portion of an outer mold line of the slat is to be aligned with a portion of an outer mold line of the chine when the slat is in the stowed position.6. The apparatus of claim 1 , wherein the chine is to generate a vortex to energize a boundary layer of the airfoil in response to an airflow presented at the chine when the slat is in the deployed position.7. The apparatus of claim 1 , wherein the lateral surface of the chine is a first lateral surface of the chine located opposite a second lateral surface of ...

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Номер записи: 54
17-01-2019 дата публикации

LEADING EDGE STRUCTURE FOR A FLOW CONTROL SYSTEM OF AN AIRCRAFT

Номер: US20190016444A1
Автор: HECK Christian, Schrauf Géza
Принадлежит:

A leading edge structure () for a flow control system of an aircraft, including a double-walled leading edge panel () that surrounds a plenum (), wherein the leading edge panel () includes an inner wall element () facing the plenum () and an outer wall element () in contact with the ambient flow (), wherein between the inner and outer wall elements () the leading edge panel () includes elongate stiffeners () spaced apart from one another, so that between each pair of adjacent stiffeners () a hollow chamber () is formed between the inner and outer wall elements (), wherein the outer wall element () includes micro pores () forming a fluid connection between the hollow chambers () and an ambient flow (), and wherein the inner wall element () includes openings () forming a fluid connection between the hollow chambers () and the plenum (). 1. A leading edge structure for a flow control system of an aircraft , the leading edge structure comprisinga double-walled leading edge panel that surrounds a plenum in a curved manner, the plenum extending in a span direction,wherein the leading edge panel has a first side portion extending from a leading edge point to a first attachment end,wherein the leading edge panel has a second side portion opposite the first side portion, extending from the leading edge point to a second attachment end,wherein the leading edge panel comprises an inner wall element facing the plenum and an outer wall element in contact with the ambient flow,wherein between the inner and outer wall elements, the leading edge panel comprises a plurality of elongate stiffeners spaced apart from one another, so that between each pair of adjacent stiffeners a hollow chamber is formed between the inner and outer wall elements,wherein the outer wall element comprises a plurality of micro pores forming a fluid connection between the hollow chambers and an ambient flow, andwherein the inner wall element comprises openings forming a fluid connection between the hollow ...

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Номер записи: 55
21-01-2016 дата публикации

FIBRILLAR STRUCTURES TO REDUCE VISCOUS DRAG ON AERODYNAMIC AND HYDRODYNAMIC WALL SURFACES

Номер: US20160017902A1
Автор: Aksak Burak, Castillo Luciano, Sitti Metin
Принадлежит:

An aerodynamic or hydrodynamic wall surface has an array of fibrillar structures disposed on and extending from the wall surface, wherein each fibrillar structure comprises a stalk and a tip. The stalk has a first end and a second end, wherein the first end is attached to the wall surface, and the stalk is oriented with respect to the wall surface at a stalk angle between approximately 1 degrees and 179 degrees. The tip has a first side and a second side, wherein the first side is attached proximate to the second end of the stalk, the tip has a larger cross-sectional area than the stalk, and the second side comprises a substantially planar surface that is oriented with respect to the stalk at a tip angle between approximately 0 degrees and 90 degrees.

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Номер записи: 56
16-01-2020 дата публикации

AIRCRAFT GENERATING LARGER LIFT BY REDUCTION OF FLUID RESISTANCE

Номер: US20200017198A1
Автор: ZHU Xiaoyi
Принадлежит:

The invention discloses a lift source for an aircraft comprising a fuselage and wings, wherein first channels are formed in the wings, a plurality of first inlets are formed in upper surfaces of the wings, a plurality of first pressure ports are formed in lower surfaces of the wings and are communicated with the first inlets via the first channels; and spoiler devices are arranged in the first channels and under the effect of the spoiler devices, form high-speed fluid layers on the upper surfaces of the wings, thereby generating a pressure difference from the lower surfaces of the wings which counteracts an external fluid pressure on the upper surfaces of the wings in the opposite direction, so a lift is generated by reduction of fluid resistance when fluid flows through the upper and lower surfaces of the wings, thereby developing a high-speed aircraft with a larger lift and thrust. 1. An aircraft , comprising a fuselage and wings , wherein first channels are formed in the wings , a plurality of first inlets are formed in upper surfaces of the wings and are communicated with the first channels , the first channels extend in a lengthwise direction of the wings from roots of sides close to the fuselage to tails of sides away from the fuselage and are communicated with an outside via first outlets , and the first channels are internally provided with spoiler devices used to extend fluid paths and form high-speed fluid layers together with the upper surfaces of the wings;a plurality of first pressure ports are formed in lower surfaces of the wings and are communicated with the first inlets in the upper surfaces of the wings via the first channels, and an air intake area of the first inlets is larger than that of the first pressure ports; a pressure difference is generated between the lower surfaces of the wings with a low external fluid flow speed and the high-speed fluid layers and counteracts an external fluid pressure on the upper surfaces of the wings in an ...

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Номер записи: 57
16-01-2020 дата публикации

ACTIVE FLOW CONTROL SYSTEMS FOR AIRCRAFT AND RELATED METHODS

Номер: US20200017199A1
Автор: Beckman Mary E., Duffy Michael J., LEUNG ANDREW, Mattero John Jeffrey, Woszidlo Rene
Принадлежит:

Active fluid control systems and related methods are disclosed. A disclosed example active fluid control system includes a plurality of plenums coupled together to define a fluid flow passageway, and a plurality of fluidic actuators coupled to outer surfaces of respective ones of the plenums. The fluidic actuators define actuator inlets and actuator outlets. The fluid flow passageway defined by the plenums to fluidly couple the fluidic actuators and a pressurized fluid supply source. The plenums are configured to couple to an aircraft structure supporting an aerodynamic surface to enable the actuator outlets to be mounted to the aerodynamic surface. The fluidic actuators are configured to provide the pressurized fluid to the aerodynamic surface to modify an aerodynamic characteristic of the aerodynamic surface. 1. An active flow control system for an aircraft , the system comprising:a plurality of plenums coupled together to define a fluid flow passageway; anda plurality of fluidic actuators coupled to outer surfaces of respective ones of the plenums, the fluidic actuators defining actuator inlets and actuator outlets, the fluid flow passageway defined by the plenums to fluidly couple the fluidic actuators and a pressurized fluid supply source, the plenums configured to couple to an aircraft structure supporting an aerodynamic surface to enable the actuator outlets to be mounted to the aerodynamic surface, the fluidic actuators configured to provide the pressurized fluid to the aerodynamic surface to modify an aerodynamic characteristic of the aerodynamic surface.2. The system as define in claim 1 , further including connectors to couple the plenums to define the fluid flow passageway.3. The system as defined in claim 2 , wherein the connectors include at least one of a v-band clamp claim 2 , a coupling claim 2 , and a male-to-female connector.4. The system as defined in claim 1 , wherein the fluidic actuators are fluidic oscillators.5. The system as defined in ...

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Номер записи: 58
16-01-2020 дата публикации

LOW DRAG SURFACE

Номер: US20200018333A1
Автор: WEBB Simon P.
Принадлежит: ROLLS-ROYCE PLC

A low drag surface is provided for a fluid washed object, the low drag surface comprising an aerodynamic surface comprising a cut-out region, and a continuously translatable surface comprising a surface portion. The surface portion is positioned in the cut-out region such that the aerodynamic surface and the surface portion form a fluidwash surface, and the surface portion is translatable relative to the aerodynamic surface. 2. The low drag surface according to claim 1 , wherein the surface portion is substantially flush with the aerodynamic surface.3. The low drag surface according to claim 1 , wherein the continuously translatable surface is the radially outer surface of a sphere.4. The low drag surface according to claim 1 , whereineach axis of rotation is fixed in position relative to the aerodynamic surface.5. The low drag surface according to claim 1 , the low drag surface further comprising an actuator for actuating the continuously translatable surface.6. The low drag surface according to claim 1 , wherein the continuously translatable surface is arrangeable such that a forward gap exists between the aerodynamic surface and a forward edge of the surface portion and/or a rearward gap exists between the aerodynamic surface and a rearward edge of the surface portion claim 1 , and wherein air is forced out of the forward gap and/or air is forced in to the rearward gap.7. The low drag surface according to claim 1 , wherein the low drag surface comprises a plurality of cut-out regions and a plurality of surface portions claim 1 , wherein each surface portion corresponds to and is positioned in a cut-out region claim 1 , and wherein the aerodynamic surface and the plurality of surface portions form a fluidwash surface.8. The low drag surface according to claim 7 , wherein the plurality of cut-out regions are arranged in an array claim 7 , and wherein optionally the number of cut-out regions is greater than 4 claim 7 , 6 claim 7 , 8 claim 7 , 10 or 20.9. A gas ...

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Номер записи: 59
26-01-2017 дата публикации

SPIN RESISTANT AIRCRAFT CONFIGURATION

Номер: US20170021916A1
Автор: Gionta Matthew, Karkow Jon, Koehler Dieter, Lednicer David, Roncz John
Принадлежит: ICON AIRCRAFT, INC.

A configuration and system for rendering an aircraft spin resistant is disclosed. Resistance of the aircraft to spinning is accomplished by constraining a stall cell to a wing region adjacent to the fuselage and distant from the wing tip. Wing features that facilitate this constraint include but are not limited to one or more cuffs, stall strips, vortex generators, wing twists, wing sweeps and horizontal stabilizers. Alone or in combination, aircraft configuration features embodied by the present invention render the aircraft spin resistant by constraining the stall cell, which allows control surfaces of the aircraft to remain operational to control the aircraft. 1. (canceled)2. A configuration for an aircraft so as to be spin resistant , said aircraft comprising a fuselage including a T-tail configuration and a high wing having a leading edge coupled to an upper portion of the fuselage wherein the high wing includes a first region adjacent to the fuselage and a second region adjacent to a wing tip , the first region being contiguous with the second region , and wherein at high angles of attack a stall cell and associated separated airflow is constrained using a discontinuity in the leading edge of the wing to remain within the first region and flow so as to remain apart from flight controls located within the second region and flight controls associated with the T-tail.3. The configuration for an aircraft as defined in claim 2 , wherein the leading edge discontinuity includes a cuff operable to form an aerodynamic boundary between the first region and the second region.4. The configuration for an aircraft as defined in claim 3 , wherein the cuff includes a delta vortex generator.5. The configuration for an aircraft as defined in claim 2 , wherein the first region includes one or more stall strips associated with a leading edge of the wing.6. The configuration for an aircraft as defined in claim 5 , wherein at least one of the one or more stall strips is coupled to ...

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Номер записи: 60
10-02-2022 дата публикации

RIBLET FILM AND METHOD FOR THE PRODUCTION THEREOF

Номер: US20220040908A1
Автор: Pfingsten Kai-Christoph, Seul Andreas
Принадлежит:

A riblet film has a riblet structure on a first side of the riblet film; and a fixing surface on a second side of the riblet film. The riblet film is formed from a cured embossing lacquer, in which a planar textile is embedded, and in which the riblet structure is embossed. 1. A riblet film , the riblet film comprising:a riblet structure on a first side of the riblet film; anda fixing surface on a second side of the riblet film,wherein the riblet film is formed from a cured embossing lacquer, in which a planar textile is embedded, and in which the riblet structure is embossed.2. The ribled film according to claim 1 , wherein an adhesive layer is applied to the fixing surface.3. The riblet film according to claim 1 , wherein the planar textile is a woven fabric or a knitted fabric formed from threads.4. The riblet film according to claim 1 , wherein a distance between the second side of the riblet film and the embedded planar textile is at least 5 μm claim 1 , or a distance between a base surface of the riblet structure on the first side of the film and the embedded planar textile is at least 15 μm.5. The ribled film according to claim 1 , wherein a total thickness of the riblet film is between 30 μm and 120 μm.6. The riblet film according to claim 1 , wherein the riblet film is cut into square patches having a size of between 0.3×0.3 m and 1.0×1.0 m.7. A method for producing the riblet film according to claim 1 , the method comprising:a) surrounding the planar textile with the embossing lacquer;b) embossing the riblet structure in the embossing lacquer; andc) curing the embossing lacquer.8. The method according to claim 7 , the method comprising wherein the following steps carried out before step:producing a base layer from the embossing lacquer;curing the base layer at least in part; andapplying the planar textile to the base layer, which is not yet fully cured or is cured at least in part.9. The method according to claim 7 , wherein the embossing lacquer is ...

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Номер записи: 61
10-02-2022 дата публикации

Apparatus and method for heating an aircraft structure

Номер: US20220041289A1
Автор: Henry Edwards
Принадлежит: Airbus Operations Ltd

A method of providing ice protection on a surface of an aircraft using exhaust air from a laminar flow control compressor. An aircraft structure, for example a wing, includes a skin. The skin has an external surface, on an outer face of the skin. The skin has an internal surface, located opposite the external surface on an inner face of the skin. The aircraft structure includes a laminar flow control system including a compressor. The aircraft structure is so arranged that the exhaust air from the compressor is directed onto the internal surface of the skin of the aircraft structure, for example thus providing hot exhaust air which function as an ice protection system (whether by de-icing or anti-icing).

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Номер записи: 62
10-02-2022 дата публикации

FLUIDIC PROPULSIVE SYSTEM AND THRUST AND LIFT GENERATOR FOR AERIAL VEHICLES

Номер: US20220041297A1
Автор: Evulet Andrei
Принадлежит:

A vehicle includes a main body and a gas generator producing a gas stream. At least one fore conduit and tail conduit are fluidly coupled to the generator. First and second fore ejectors are fluidly coupled to the at least one fore conduit. At least one tail ejector is fluidly coupled to the at least one tail conduit. The fore ejectors respectively include an outlet structure out of which gas from the at least one fore conduit flows. The at least one tail ejector includes an outlet structure out of which gas from the at least one tail conduit flows. First and second primary airfoil elements have leading edges respectively located directly downstream of the first and second fore ejectors. At least one secondary airfoil element has a leading edge located directly downstream of the outlet structure of the at least one tail ejector. 1. A vehicle , comprising:a main body having a fore portion, a tail portion, a starboard side and a port side;a gas generator coupled to the main body and producing a gas stream;at least one fore conduit fluidly coupled to the generator;at least one tail conduit fluidly coupled to the generator;first and second fore ejectors fluidly coupled to the at least one fore conduit, coupled to the fore portion and respectively coupled to the starboard side and port side, the fore ejectors respectively comprising an outlet structure out of which gas from the at least one fore conduit flows at a predetermined adjustable velocity;at least one tail ejector fluidly coupled to the at least one tail conduit and coupled to the tail portion, the at least one tail ejector comprising an outlet structure out of which gas from the at least one tail conduit flows at a predetermined adjustable velocity;first and second primary airfoil elements having leading edges, the primary airfoil elements respectively coupled to the starboard side and port side, the leading edges of the first and second primary airfoil elements being respectively located directly downstream of ...

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Номер записи: 63
25-01-2018 дата публикации

PHONONIC MATERIALS USED TO CONTROL TURBULENT FLOW

Номер: US20180023599A1
Автор: Biringen Sedat, Hussein Mahmoud
Принадлежит:

A phononic material and a method of using a phononic material for use in interacting with a fluid or solid flow are provided. The phononic material includes an interface surface and a subsurface feature. The interface surface is adapted to move in response to a pressure, and/or velocity gradients, associated with complex motion of a turbulent flow exhibiting a polarity of frequencies exerted on the interface surface. The subsurface feature extends from the interface surface. The subsurface feature comprises a phononic crystal or locally resonant metamaterial adapted to receive the pressure, and/or velocity gradients, from the turbulent flow via the interface surface and alter the phase and amplitude of a polarity of frequency components of the turbulent flow in order to reduce or increase the kinetic energy of the turbulent flow. The interface surface is adapted to vibrate at a polarity of frequencies, phases and amplitudes in response to the frequency, phase and amplitude of at least one component of the turbulent flow. 1. A phononic material for use in interacting with a turbulent fluid or solid flow , the phononic material comprising:an interface surface adapted to move in response to at least one of a pressure and a velocity gradient associated with complex motion of a turbulent flow exhibiting a plurality of frequencies exerted on the interface surface; anda subsurface feature extending from the interface surface, the subsurface feature comprising a phononic crystal or locally resonant metamaterial adapted to receive the at least one of the pressure and the velocity gradient from the turbulent flow via the interface surface and alter phase and amplitude of a plurality of frequency components of the turbulent flow.2. The phononic material of wherein the complex motion comprises a plurality of frequency domain components.3. The phononic material of wherein the complex motion comprises a plurality of frequency domain components across a spectrum of frequencies.4. ...

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Номер записи: 64
24-01-2019 дата публикации

RIBLET STRUCTURE AND OBJECT

Номер: US20190023379A1
Автор: Iijima Hidetoshi, Koshioka Yasuhiro, Kurita Mitsuru, Nishizawa Akira, Okabayashi Kie
Принадлежит: JAPAN AEROSPACE EXPLORATION AGENCY

The present invention provides a riblet structure that further reduces drag, which is a sum of turbulent friction drag and pressure drag, and an object including such a riblet structure. An object such as an aircraft, plant, and pipeline includes a wavelike riblet pattern on a surface. The wavelike riblet pattern includes a large number of wavelike riblets. Each of the large number of wave riblets includes a wavelike ridge line, and a height thereof changes cyclically with respect to a fluid flow direction. With such a configuration, drag, which is a sum of turbulent friction drag and pressure drag, can be further reduced. 1. A riblet structure , comprisinga plurality of wavelike riblets,a height of each of the plurality of wavelike riblets being set to become smaller as an angle formed between a ridge line and a fluid flow direction becomes larger.2. The riblet structure according to claim 1 , whereinthe ridge line of each of the plurality of wavelike riblets has a sinusoidal shape of a first wavelength, andthe height of each of the plurality of wavelike riblets increases/decreases in a sinusoidal shape of a second wavelength as a half wavelength of the first wavelength in the fluid flow direction.3. The riblet structure according to claim 2 , wherein {'br': None, 'i': a≤', 'h., '0<0.2'}, 'when a reference height of the wavelike riblets is represented by h, an amplitude a of the second wavelength satisfies'}4. The riblet structure according to claim 1 , whereinthe plurality of wavelike riblets are arranged at constant intervals.5. An object claim 1 , comprisinga wavelike riblet pattern including wavelike ridge lines, a height of the wavelike riblet pattern changing cyclically with respect to a fluid flow direction. The present invention relates to an object including a surface on which fluid flows and a riblet structure applied to such an object.It is known that, by providing a riblet pattern on a surface of an aircraft, a wall surface of a pipeline, and the like, ...

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Номер записи: 65
23-01-2020 дата публикации

Aerodynamic surface of an aircraft

Номер: US20200023945A1
Автор: Sergei Nikolaevich Nizov
Принадлежит: Individual

An aerodynamic surface of an aircraft comprises a main part having a leading and a trailing edges and having an airfoil section. The aerodynamic surface also having at least two vortex generators in the form of teeth having edges along the length thereof. The teeth are mounted on the leading edge of the main part so as to be capable of generating two vortex cores on one tooth. The edges of a tooth adjoin the leading edge of the main part of the aerodynamic surface. The radius of an edge of each tooth along the length of the vortex generator is five times less than the radius of the leading edge of the main part. The main part of the aerodynamic surface has a cambered airfoil section, wherein the teeth are mounted with a deflection towards the smallest degree of curvature of the airfoil section of the main part. The invention is intended for reducing an aerodynamic drag at low angles of attack while maintaining an increased load hearing capacity of the aerodynamic surface by generating vortex cores adjoining one of the sides thereof.

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Номер записи: 66
23-01-2020 дата публикации

FLUIDIC PROPULSIVE SYSTEM AND THRUST AND LIFT GENERATOR FOR AERIAL VEHICLES

Номер: US20200023987A1
Автор: Evulet Andrei
Принадлежит:

A vehicle includes a main body and a gas generator producing a gas stream. At least one fore conduit and tail conduit are fluidly coupled to the generator. First and second fore ejectors are fluidly coupled to the at least one fore conduit. At least one tail ejector is fluidly coupled to the at least one tail conduit. The fore ejectors respectively include an outlet structure out of which gas from the at least one fore conduit flows. The at least one tail ejector includes an outlet structure out of which gas from the at least one tail conduit flows. First and second primary airfoil elements have leading edges respectively located directly downstream of the first and second fore ejectors. At least one secondary airfoil element has a leading edge located directly downstream of the outlet structure of the at least one tail ejector. 1. A vehicle , comprising:a main body having a fore portion, a tail portion, a starboard side and a port side;a gas generator coupled to the main body and producing a gas stream;at least one fore conduit fluidly coupled to the generator;at least one tail conduit fluidly coupled to the generator;first and second fore ejectors fluidly coupled to the at least one fore conduit, coupled to the fore portion and respectively coupled to the starboard side and port side, the fore ejectors respectively comprising an outlet structure out of which gas from the at least one fore conduit flows at a predetermined adjustable velocity;at least one tail ejector fluidly coupled to the at least one tail conduit and coupled to the tail portion, the at least one tail ejector comprising an outlet structure out of which gas from the at least one tail conduit flows at a predetermined adjustable velocity;first and second primary airfoil elements having leading edges, the primary airfoil elements respectively coupled to the starboard side and port side, the leading edges of the first and second primary airfoil elements being respectively located directly downstream of ...

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Номер записи: 67
24-01-2019 дата публикации

ACOUSTIC CAVITY TAILORED SYNTHETIC JET

Номер: US20190024608A1
Автор: Griffin Steven F.
Принадлежит:

An acoustic cavity tailored synthetic jet employs a body having a cavity with a wall including a taper from a first extent to an aperture. The cavity is configured to produce a matched acoustic resonance. A drive system has a piston engaged to the cavity at the first extent. The drive system and piston are configured for oscillatory motion inducing a synthetic jet at the aperture. 1. An acoustic cavity tailored synthetic jet comprising:a body having a cavity with a wall having a taper from a first extent to an aperture and configured to produce a matched acoustic resonance;a drive system having a piston engaged to the cavity at the first extent, said drive system and piston configured for oscillatory motion inducing a synthetic jet at the aperture.2. The acoustic cavity tailored synthetic jet as defined in wherein the taper has a varying slope from the extent to the aperture.3. The acoustic cavity tailored synthetic jet as defined in wherein the varying slope increases from the extent to the aperture.4. The acoustic cavity tailored synthetic jet as defined in wherein the drive system and piston are operable at a reduced uncoupled resonant frequency.5. The acoustic cavity tailored synthetic jet as defined in wherein the drive system is configured to operate at an uncoupled resonant frequency of one half the matched acoustic resonance.6. The acoustic cavity tailored synthetic jet as defined in wherein the drive system incorporates a spring claim 5 , said spring having stiffness reduced by a factor of 4 from a stiffness for an uncoupled resonant frequency equal to the matched acoustic resonance.7. The acoustic cavity tailored synthetic jet as defined in wherein the aperture comprises a slot.8. The acoustic cavity tailored synthetic jet as defined in wherein the aperture comprises a circular hole and the cavity is symmetrical about a normal axis to a center of the circular hole.9. The acoustic cavity tailored synthetic jet as defined in wherein the taper is ...

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Номер записи: 68
28-01-2021 дата публикации

LEADING-EDGE THERMAL ANTI-ICE SYSTEMS AND METHODS

Номер: US20210024220A1
Автор: Amorosi Stephen Roger, Burton Jeffrey Michael, Silva Douglas James
Принадлежит: The Boeing Company

An anti-ice system is disclosed, including an airfoil structure having a leading-edge portion facing a wind direction. The airfoil structure includes an outer skin and an inner skin which form a plurality of channels, each channel having an air inlet and an air outlet. An air delivery duct inside the airfoil structure extends transverse to the wind direction and has a plurality of openings directed toward the air inlets of the channels. A heat exchanger is configured to provide hot air to the air delivery duct. 1. An anti-ice system , comprising:an airfoil structure including a leading-edge portion facing a wind direction, and including an outer skin and an inner skin forming a plurality of channels, each channel having an air inlet and an air outlet,an air delivery duct inside the airfoil structure extending substantially transverse to the wind direction, the air delivery duct having a plurality of openings directed toward the air inlets of the plurality of channels, andan air supply configured to provide hot air to the air delivery duct.2. The anti-ice system of claim 1 , wherein each channel has a cross-sectional dimension and each cross-sectional dimension varies between the air inlet and the air outlet.3. The anti-ice system of claim 2 , wherein each channel has an upper portion claim 2 , a middle portion claim 2 , and a lower portion claim 2 , at least one of the upper and lower portions having an increased cross-sectional area as compared to the middle portion.4. The anti-ice system of claim 2 , wherein each channel has an upper portion claim 2 , a middle portion claim 2 , and a lower portion claim 2 , the upper and lower portions each having an increased cross-sectional area as compared to the middle portion.5. The anti-ice system of claim 1 , wherein the air inlets receive air from inside the airfoil structure claim 1 , and the air outlets discharge air outside of the airfoil structure.6. The anti-ice system of claim 1 , wherein the inner skin has a front ...

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Номер записи: 69
23-01-2020 дата публикации

SWIRLING JET ACTUATOR FOR CONTROL OF SEPARATED AND MIXING FLOWS

Номер: US20200025225A1
Автор: Alvi Farrukh, Munday Phillip, Taira Kunihiko
Принадлежит:

A method of controlling a fluid flow using momentum and/or vorticity injections. Actively controlling an actuator allows for direct, precise, and independent control of the momentum and swirl entering into the fluid system. The perturbations are added to the flow field in a systematic mater providing tunable control input, thereby modifying behavior thereof in a predictable manner to improve the flow characteristics. 1. A method of controlling a fluid flow , comprising inputting a swirling flow into the fluid flow.2. The method of claim 1 , wherein the swirling flow is inputted in an orientation such that a central axis claim 1 , about which the swirling flow rotates is normal to a surface of a body over which the fluid flow is passing.3. The method of claim 1 , further including the step of adjusting flow properties of the swirling flow.4. The method of claim 1 , wherein the swirling flow is actively controllable.5. The method of claim 1 , wherein the inputting occurs at a plurality of actuator sites such that each actuator site includes a swirling flow input and each swirling flow input has an initial direction of rotation that is opposite of the initial direction of rotation of the swirling flow input of an adjacently located actuator site.6. The method of claim 1 , wherein the inputting occurs at a plurality of actuator sites such that each actuator site includes a swirling flow input and each swirling flow input has an initial direction of rotation that is in the same initial direction of rotation of the swirling flow input of an adjacently located actuator site.7. The method of claim 1 , further including a step of inputting a momentum flow.8. The method of claim 7 , wherein the momentum flow is inputted in an orientation that is normal to a surface of a body over which the fluid flow is passing.9. The method of claim 7 , wherein the momentum flow is adjustable.10. A method of controlling a fluid flow claim 7 , comprising the step of inputting a swirling flow ...

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Номер записи: 70
02-02-2017 дата публикации

CONTROL SURFACE FOR AN AIRCRAFT

Номер: US20170029090A1
Автор: CHU JAMES, Northam Robert
Принадлежит:

The present application relates to a control surface for an aircraft. The control surface has a leading edge, a trailing edge, and a chord-line defined between the leading edge and the trailing edge. A first aerodynamic surface is between the leading and trailing edges and a second surface is between the leading and trailing edges. The leading edge is formed by a nose, the nose having a hinge axis about which the control surface is deflectable. A maximum thickness of the control surface perpendicular to the chord-line between the first aerodynamic surface and the second surface is located aft of the hinge axis. The present application also relates to a control surface for an aircraft having a maximum curvature of the first aerodynamic surface of the control surface located aft of the hinge axis. The present application also relates to an aircraft or part of an aircraft comprising a fixed section and a control surface. 1. A control surface for an aircraft comprisinga leading edge, a trailing edge, and a chord-line defined between the leading edge and the trailing edge,a first aerodynamic surface between the leading and trailing edges,a second surface between the leading and trailing edges,the leading edge being formed by a nose, the nose having a hinge axis about which the control surface is deflectable,the first aerodynamic surface having an exposed airflow surface and the nose comprising an arced nose profile section extending from the exposed airflow surface wherein the arced nose profile section is configured to be at least partially exposable to airflow flowing over the control surface, wherein a maximum thickness of the control surface, perpendicular to the chord-line between the first aerodynamic surface and the second surface, is located aft of the hinge axis, anda radius of the arced nose profile section is greater than a perpendicular distance between the chord-line and the point of maximum thickness of the control surface on the first aerodynamic surface.2 ...

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Номер записи: 71
02-02-2017 дата публикации

Methods and systems for rotary wing active flow control

Номер: US20170029102A1
Автор: Daniel J. Clingman, Randy Lee M. Mallari
Принадлежит: Boeing Co

Within examples, systems for enhanced performance blades for rotor craft are provided and methods for operation. An example system for a rotary device is provided comprising a rotor blade coupled to a rotor hub. The system also comprises an air channel disposed within the rotor blade, where the air channel is sealed proximate to a distal end of the rotor blade. The system also comprises an inlet positioned at a proximal end of the rotor blade, where the inlet is in fluid communication with the air channel. The system also comprises a plurality of outlets positioned along the rotor blade, where each of the plurality of outlets are in fluid communication with the air channel.

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Номер записи: 72
17-02-2022 дата публикации

LEADING EDGE STRUCTURE FOR A FLOW CONTROL SYSTEM OF AN AIRCRAFT

Номер: US20220048612A1
Автор: BÜSCHER Alexander
Принадлежит:

A leading edge structure for a flow control system of an aircraft is disclosed having a leading edge panel that surrounds a plenum, wherein the leading edge panel has a first side portion, a second side portion opposite the first side portion, an inner surface facing the plenum and an outer surface in contact with an ambient flow, and wherein the leading edge panel comprises a plurality of micro pores forming a fluid connection between the plenum and the ambient flow, wherein the plenum is connected to an air outlet arrangement configured for causing an underpressure in the plenum, so that air from the ambient flow is drawn through the micro pores into the plenum and from there discharged through the air outlet arrangement into the ambient flow. 1. A leading edge structure for a flow control system of an aircraft , comprising:a leading edge panel that surrounds a plenum in a curved manner, the plenum extending in a span direction,wherein the leading edge panel has a first side portion extending from a leading edge point to a first attachment end,wherein the leading edge panel has a second side portion opposite the first side portion, extending from the leading edge point to a second attachment end,wherein the leading edge panel comprises an inner surface facing the plenum and an outer surface in contact with an ambient flow, andwherein the leading edge panel comprises a plurality of micro pores forming a fluid connection between the plenum and the ambient flow,wherein the plenum is connected to an air outlet arrangement configured for causing an underpressure in the plenum, so that air from the ambient flow is drawn through the micro pores into the plenum and from there discharged through the air outlet arrangement into the ambient flow,wherein the air outlet arrangement is configured to operate in a flow control mode where a first mass flow rate of air from the ambient flow is drawn through the micro pores into the plenum, and in a cleaning mode where a second mass ...

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Номер записи: 73
31-01-2019 дата публикации

MULTILAYER RIBLET APPLIQUE AND METHODS OF PRODUCING THE SAME

Номер: US20190030764A1
Автор: Burg Alan G., Granger George Michael, Keough Bruce K., Kestner James M., McGarvey James Charles, Olli Larry K., Rawlings Diane C.
Принадлежит:

Multilayer riblet applique and methods of producing the same are described herein. One disclosed example method includes applying a first high elongation polymer material to a web tool, where the web tool is to be provided from a first roll, and heating, via a first heating process, the first high elongation polymer material. The disclosed example method also includes applying a second high elongation polymer material to the first high elongation polymer material, and heating, via a second heating process, the second high elongation polymer material. The disclosed example method also includes applying, via a laminating roller, a support layer to the second high elongation polymer material. 1. A multilayer riblet applique comprising:a fluorosilicone riblet structure including riblet ridges and a base from which the riblet ridges extend;a fluorosilicone layer adjacent the fluorosilicone riblet structure; anda support layer proximate the base, the support layer including a metal sub-layer, an adhesive sub-layer, and at least one thermoplastic sub-layer.2. The multilayer riblet applique as defined in claim 1 , further including a masking adjacent the fluorosilicone riblet structure.3. The multilayer riblet applique as defined in claim 2 , further including a polymer coating between the masking and the fluorosilicone riblet structure.4. The multilayer riblet applique as defined in claim 1 , wherein the fluorosilicone layer includes a color layer.5. The multilayer riblet applique as defined in claim 1 , wherein the fluorosilicone riblet structure includes a repeating pattern of ridges and valleys.6. An aircraft component including the multilayer riblet applique as defined in applied thereto.7. An aircraft having a plurality of surfaces with the multilayer riblet applique as defined in applied thereto.8. The multilayer riblet applique as defined in claim 1 , wherein the fluorosilicone riblet structure includes a repeating pattern of ridges and valleys.9. The multilayer ...

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Номер записи: 74
31-01-2019 дата публикации

FLUIDIC DEVICE

Номер: US20190031321A1
Автор: BACIC Marko, CHMIELARCZYK Pawel, Young Richard A.
Принадлежит: ROLLS-ROYCE PLC

A fluidic device for providing analogue output control includes a main channel, a first control channel, a second control channel, a comparator which receives respective input fluid flows from the main, the first and the second control channels. The first control channel is configured such that the input fluid flow therefrom carries an oscillating pressure wave signal, the second control channel includes a flow regulator controllable to vary the mass flow rate of the input fluid flow from the second control channel, and the main channel is configured such that the input fluid flow therefrom is at a reference mass flow rate. The comparator is configured such that the input fluid flows from the first control and the second control channels act in combination on the input fluid flow from the main channel to produce an output fluid flow from the comparator having a PWM mass flow rate characteristic. 1. A fluidic device for providing analogue output control , the device including:a main channel, a first control channel and a second control channel, anda comparator which receives respective input fluid flows from the main, the first control and the second control channels;wherein the first control channel is configured such that the input fluid flow therefrom carries an oscillating pressure wave signal, the second control channel includes a flow regulator which is controllable to vary the mass flow rate of the input fluid flow from the second control channel, and the main channel is configured such that the input fluid flow therefrom is at a reference mass flow rate; andwherein the comparator is configured such that the input fluid flows from the first control and the second control channels act in combination on the input fluid flow from the main channel to produce an output fluid flow from the comparator having a pulse width modulation mass flow rate characteristic.2. A fluidic device according to claim 1 , wherein the first control channel extends in flow series ...

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Номер записи: 75
31-01-2019 дата публикации

STEALTH DESIGN WITH MULTI-FACETED DIHEDRAL PLANFORM AND INSUFFLATION MECHANISM

Номер: US20190031322A1
Автор: MICROS IOANNIS
Принадлежит:

A stealth craft's aerodynamics and flight stability are improved with the use of a multi-faceted dihedral planform. The stealth craft includes a multi-faceted dihedral planform extending in a direction from a front to a rear of a craft (or wing) and defined by a first set of facets followed by a second set of facets. In an exemplary embodiment, the first and second sets of facets have an angle of incline that is ascending and descending, respectively, with respect to the direction of the planform. Selected ones of the first and second sets of facets are configured with insufflation slots for improving aerodynamics and stability, the insufflation slots extending spanwise in a direction transverse to the direction of the planform and provided to insufflate a fluid to form a cushion of air along the multi-faceted dihedral planform for improving aerodynamics and stability. 1. A stealth craft including a multi-faceted dihedral planform extending in a direction from a front to a rear of a craft or a wing thereof and defined by a first set of facets followed by a second set of facets ,wherein the first and second sets of facets have an angle of incline that is ascending and descending, respectively, with respect to the direction of the planform, andwherein selected ones of the first and second sets of facets are configured with insufflation slots for improving aerodynamics and stability, the insufflation slots extending spanwise in a direction transverse to the direction of the planform and provided to insufflate a fluid to form a cushion of air along the multi-faceted dihedral planform for improving aerodynamics and stability.2. The stealth craft of claim 1 , further comprising recovery slots that recover insufflated fluid.3. The stealth craft of claim 2 , wherein the recovery slots include means for recirculating the recovered insufflated fluid and re-insufflating it downstream of the multi-faceted dihedral planform.4. The stealth craft of claim 1 , wherein the multi- ...

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Номер записи: 76
31-01-2019 дата публикации

SYSTEM AND METHOD FOR OPERATING A BOUNDARY LAYER INGESTION FAN

Номер: US20190031363A1
Автор: Hoisington Zachary Charles
Принадлежит:

An aircraft includes a fuselage having a tail section and an engine core coupled via an external pylon to the fuselage. The aircraft further includes a boundary layer ingestion (BLI) fan coupled to the tail section of the fuselage and coupled via a shaft to the engine core. 1. An aircraft comprising:a fuselage having a tail section;an engine core coupled to the fuselage via an external pylon; anda boundary layer ingestion (BLI) fan coupled to the tail section of the fuselage and coupled to the engine core via a shaft.2. The aircraft of claim 1 , wherein the engine core is configured to generate power to rotate the shaft claim 1 , and wherein the BLI fan is configured to generate thrust based on rotation of the shaft.3. The aircraft of claim 1 , further comprising:a second engine core configured to drive a second fan; andthe second fan coupled to the second engine core and configured to generate thrust.4. The aircraft of claim 3 , wherein the engine core corresponds to a turboshaft engine claim 3 , and wherein the second engine core corresponds to a turbofan engine claim 3 , a propfan engine claim 3 , or a turboprop engine.5. The aircraft of claim 3 , further comprising a third fan coupled to and configured to be driven by the engine core claim 3 , wherein each of the engine core and the second engine core correspond to a turbofan engine claim 3 , a propfan engine claim 3 , or a turboprop engine.6. The aircraft of claim 5 , further comprising a transmission configured to transfer power from the engine core and the second engine core to the BLI fan claim 5 , wherein the shaft and the transmission are included in a BLI fan drive system.7. The aircraft of claim 3 , further comprising a second shaft coupled to the shaft via a gear system claim 3 , the shaft coupled to the BLI fan and configured to receive power from the second shaft to rotate the BLI fan.8. The aircraft of claim 1 , wherein the BLI fan includes an inlet configured to receive air from a boundary layer of ...

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Номер записи: 77
30-01-2020 дата публикации

Flow Body Of An Aircraft And Aircraft

Номер: US20200031455A1
Автор: Kreuzer Peter
Принадлежит: AIRBUS OPERATIONS GMBH

A flow body of an aircraft includes: a flow surface exposed to an airstream during flight of the aircraft, the flow surface generating at least one region of turbulent airflow during flight of the aircraft, at least one perforated area including a plurality of openings extending through the flow surface, a manifold positioned interior to the flow surface in fluid communication with the openings, and at least one suction duct having a first end and a second end, the first end being in fluid communication with the manifold, the second end including a suction opening and being arranged in the at least one region of turbulent airflow, wherein the suction opening is adapted for inducing a suction force in the at least one suction duct when the flow surface is exposed to an airstream during flight, thereby inducing a flow of air from through the plurality of openings. 1. A flow body of an aircraft , comprising:a flow surface exposed to an airstream during flight of the aircraft, the flow surface generating at least one region of turbulent airflow during flight of the aircraft;at least one perforated area comprising a plurality of openings extending through the flow surface;a manifold positioned interior to the flow surface in fluid communication with the openings; andat least one suction duct having a first end and a second end, the first end in fluid communication with the manifold, the second end comprising a suction opening and arranged in the at least one region of turbulent airflow, wherein the suction opening is adapted for inducing a suction force in the at least one suction duct when the flow surface is exposed to an airstream during flight, thereby inducing a flow of air from through the plurality of openings.2. The flow body according to claim 1 , wherein the flow body comprises a plurality of suction ducts.3. The flow body according to claim 2 , wherein the flow surface generates a plurality of regions of turbulent airflow during flight of the aircraft being ...

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Номер записи: 78
30-01-2020 дата публикации

SYSTEMS AND METHODS FOR ACTIVE CONTROL OF SURFACE DRAG

Номер: US20200031456A1
Автор: Holloway Brian C., Hyde Roderick A., Wine David William
Принадлежит: Deep Science, LLC

A fluid control system includes a deformable surface that covers a body in at least a first and second direction. The first direction is orthogonal to the second direction. The deformable surface includes a bottom side that faces the body and a top side that is opposite the bottom side. The fluid control system also includes at least one deformer between the deformable surface and the body. The at least one deformer is configured to modify a boundary layer of a fluid that is flowing over the deformable surface by selectively deforming the top side of the surface. 1. A fluid flow control system , comprising:a surface covering a body in at least a first and second direction, the first direction orthogonal to the second direction, the surface including a bottom side facing the body and a top side opposite the bottom side; andat least one movable section on the top side of the surface, the at least one movable section configured to modify a boundary layer of a fluid flowing over the surface by moving along or out of the top side of the surface.2. The fluid control system of claim 1 , wherein:the at least one movable sections selectively generates surface waves.3. The fluid flow control system of claim 1 , wherein:the at least one movable section includes a plurality of movable sections that move in opposing directions.4. The fluid flow control system of claim 3 , wherein:the at least one movable section includes at least one actuator that selectively rotates the plurality of movable sections in an oscillatory manner to generate surface waves.5. The fluid flow control system of claim 1 , wherein:the at least one movable section includes a plurality of movable sections that move in at least one of differing speeds or differing directions.6. The fluid flow control system of claim 1 , comprising:a plurality of springs attached to the at least one movable section; anda plurality of actuators configured to selectively compress the plurality of springs, the plurality of ...

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Номер записи: 79
09-02-2017 дата публикации

Gust Alleviator

Номер: US20170036755A1
Автор: JR. Albert S., RICHARDSON
Принадлежит:

A gust alleviating aircraft wing includes a gust alleviating wing portion on the wing. The wing portion can have a leading edge, a trailing edge, and a downwardly sloping upper surface therebetween. At least one air passageway can extend through the wing portion from the leading edge to a rear or downstream location on the downwardly sloping upper surface of the wing portion. At least one spoiler can be on the upper surface of the wing portion at the rear location for selectively movably covering and uncovering an exit location of the at least one passageway. Opening the at least one passageway is capable of diverting air flow through the at least one passageway and the wing portion for counteracting upward lift caused by a gust of wind. 1. A gust alleviating aircraft wing comprising:a gust alleviating wing portion on the wing having a leading edge, a trailing edge, and a downwardly sloping upper surface therebetween, at least one passageway extending through the wing portion from the leading edge to a rear location on the downwardly sloping upper surface of the wing portion; andat least one spoiler on the upper surface of the wing portion at the rear location for selectively movably covering and uncovering an exit location of the at least one passageway, whereby opening the at least one passageway is capable of diverting air flow through the at least one passageway and the wing portion for counteracting upward lift caused by a gust of wind.2. The wing of in which the wing portion is a segment of at least one wing.3. The wing of in which the gust alleviating wing portion is a left wing portion on a left wing claim 2 , and the gust alleviating aircraft wing further comprising a gust alleviating right wing portion on a right wing.4. The wing of in which the at least one spoiler is pivotably mounted to the upper surface of the wing portion along a hinge upstream from the exit location of the at least one passageway.5. The wing of in which the at least one spoiler is at ...

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Номер записи: 80
24-02-2022 дата публикации

METHOD AND APPARATUS FOR MITIGATING TRAILING VORTEX WAKES OF LIFTING OR THRUST GENERATING BODIES

Номер: US20220055739A1
Автор: Sullivan Steven
Принадлежит:

Disclosed are methods and apparatuses for mitigating the formation of concentrated wake vortex structures generated from lifting or thrust-generating bodies and maneuvering control surfaces wherein the use of contour surface geometries promotes vortex-mixing of high and low flow fluids. The methods and apparatuses can be combined with various drag reduction techniques, such as the use of riblets of various types and/or compliant surfaces (passive and active). Such combinations form unique structures for various fluid dynamic control applications to suppress transiently growing forms of boundary layer disturbances in a manner that significantly improves performance and has improved control dynamics. 1. A method for reducing aerodynamic or hydrodynamic drag by mitigating formation of concentrated wake vortex structures , the method comprising:applying and/or incorporating at least one three-dimensional contour shaped surface into and/or onto one or more lifting and/or thrust-generating bodies, airfoils, and/or other surfaces.2. The method of claim 1 , wherein the one or more lifting and/or thrust-generating bodies and/or other surfaces comprise one or more fan blade surfaces.3. The method of claim 2 , wherein the fan blade surfaces are located on either an aircraft or watercraft.4. The method of claim 2 , wherein the at least one three-dimensional contour shaped surface promotes vortex-mixing of high and low flow fluids on the one or more fan blade surfaces.5. The method of claim 1 , wherein the applying and/or incorporating is performed on a trailing edge claim 1 , a leading edge claim 1 , and/or across a surface of the one or more lifting and/or thrust-generating bodies.6. The method of claim 5 , wherein the at least one three-dimensional contour shaped surface comprises a scallop-shaped surface.7. The method of claim 6 , wherein the scallop-shaped surface has a shape that varies across a trailing edge and/or leading edge of a surface.8. The method of claim 7 , ...

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Номер записи: 81
15-02-2018 дата публикации

ACTIVE FLOW CONTROL SYSTEMS AND METHODS FOR AIRCRAFT

Номер: US20180043995A1
Автор: Shmilovich Arvin, Vijgen Paul, Yadlin Yoram
Принадлежит:

Example active flow control systems and methods for aircraft are described herein. An example method includes supplying pressurized air to a plurality of nozzles. The nozzles arranged in an array across a control surface of an aircraft, and the nozzles are oriented to eject the pressurized air in a substantially streamwise direction. The method further includes activating the nozzles to eject the pressurized air in sequence to create a wave of air moving in a spanwise direction across the control surface. 1. A method comprising:supplying pressurized air to a plurality of nozzles, the nozzles arranged in an array across a control surface of an aircraft, the nozzles oriented to eject the pressurized air in a substantially streamwise direction; andactivating the nozzles to eject the pressurized air in sequence to create a wave of air moving in a spanwise direction across the control surface.2. The method of claim 1 , wherein activating the nozzles in sequence includes activating and deactivating the nozzles such that only one of the nozzles is activated at a time.3. The method of claim 1 , wherein activating the nozzles in sequence includes activating and deactivating the nozzles such that multiple nozzles are activated at a time.4. The method of claim 1 , wherein the wave of air is a first wave of air claim 1 , further including activating the nozzles to eject the pressurized air in sequence to create a second wave of air claim 1 , the first and second waves occurring simultaneously.5. The method of claim 4 , wherein the second wave of air is separated from the first wave of air by at least one of the nozzles.6. The method of claim 1 , wherein the nozzles are converging-diverging nozzles.7. The method of claim 1 , wherein the nozzles are activated in sequence from outboard to inboard.8. The method of claim 1 , wherein the activating of the nozzles is repeated at a frequency that prevents full separation of airflow over the control surface.9. The method of further ...

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Номер записи: 82
15-02-2018 дата публикации

INLET ASSEMBLY FOR AN AIRCRAFT AFT FAN

Номер: US20180043996A1
Автор: Ramakrishnan Kishore
Принадлежит:

The present disclosure is directed to a low-distortion inlet assembly for reducing airflow swirl distortion entering an aft fan mounted to a fuselage of an aircraft. Further, the inlet assembly includes a plurality of structural members, such as struts or strakes, mounted at predetermined locations around a circumference of the fan shaft of the fan at the inlet. The predetermined location(s) may be determined as a function of swirl distortion entering the inlet. As such, the structural member(s) are configured to reduce swirl distortion of the airflow entering the fan. In some embodiments, the inlet assembly may also include inlet guide vanes. In alternative embodiments, the inlet assembly may be absent of inlet guide vanes. 1. A boundary layer ingestion fan assembly for mounting to an aft end of a fuselage of an aircraft , the boundary layer ingestion fan assembly comprising:a fan rotatable about a central axis of the boundary layer ingestion fan, the fan comprising a plurality of fan blades rotatable about a fan shaft;a nacelle surrounding the plurality of fan blades of the fan, the nacelle defining an inlet with the fuselage of the aircraft, the inlet extending substantially around the fuselage of the aircraft when the boundary layer ingestion fan is mounted at the aft end of the aircraft; one or more inlet guide vanes configured within the inlet; and', 'one or more structural members mounted at predetermined radial locations around a circumference of the fan shaft of the fan at the inlet, the structural members configured to reduce a swirl distortion entering the inlet of the fan., 'a low-distortion inlet assembly configured with the inlet, the inlet assembly comprising2. The boundary layer ingestion fan assembly of claim 1 , wherein the one or more structural members comprise at least one of a strut or a strake.3. The boundary layer ingestion fan assembly of claim 1 , wherein one or more of the structural members are integrated with at least one of the nacelle ...

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Номер записи: 83
15-02-2018 дата публикации

INLET ASSEMBLY FOR AN AIRCRAFT AFT FAN

Номер: US20180043997A1
Автор: OTTA Shourya Prakash, Ramakrishnan Kishore
Принадлежит:

The present disclosure is directed to a low-distortion inlet assembly for reducing airflow swirl distortion entering an aft fan mounted to a fuselage of an aircraft. Further, the inlet assembly includes a plurality of structural members mounted at one or more predetermined locations around a circumference of the fan shaft of the fan. More specifically, the predetermined location(s) has a swirl distortion exceeding a predetermined threshold. Further, the inlet assembly includes at least one airflow modifying element configured within the inlet so as to reduce swirl distortion entering the fan. 1. A boundary layer ingestion fan assembly for mounting to an aft end of a fuselage of an aircraft , the boundary layer ingestion fan assembly comprising:a fan rotatable about a central axis of the boundary layer ingestion fan, the fan comprising a plurality of fan blades rotatable about a fan shaft;a nacelle surrounding the plurality of fan blades of the fan, the nacelle defining an inlet with the fuselage of the aircraft, the inlet extending substantially around the fuselage of the aircraft when the boundary layer ingestion fan is mounted at the aft end of the aircraft; one or more structural members mounted at predetermined locations around a circumference of the fan shaft of the fan within the inlet, the predetermined locations comprising a swirl distortion exceeding a predetermined threshold; and', 'at least one airflow modifying element configured within the inlet so as to reduce swirl distortion entering the fan., 'a low-distortion inlet assembly mounted within the inlet, the inlet assembly comprising2. The boundary layer ingestion fan assembly of claim 1 , wherein the one or more structural members comprise at least one of an inlet guide vane or a strut.3. The boundary layer ingestion fan assembly of claim 2 , further comprising a plurality of inlet guide vanes placed in groups at the predetermined locations around the circumference of the fan shaft.4. The boundary ...

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Номер записи: 84
15-02-2018 дата публикации

Inlet guide vane assembly for reducing airflow swirl distortion of an aircraft aft fan

Номер: US20180045205A1
Автор: Jixian Yao
Принадлежит: General Electric Co

The present disclosure is directed to an aerodynamic inlet guide vane assembly for reducing airflow swirl distortion entering an aft fan mounted to a fuselage of an aircraft. Further, the inlet guide vane assembly is configured for mounting to fan shaft and a nacelle of the aft fan. The inlet guide vane assembly includes a plurality of inlet guide vanes grouped into a plurality of inlet guide vane groups. Each of the inlet guide vanes has a shape and an orientation corresponding to airflow conditions entering the fan. Further, the inlet guide vane groups are spaced circumferentially around the central axis as a function of the airflow conditions entering the fan.

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Номер записи: 85
03-03-2022 дата публикации

AIRCRAFT EQUIPPED WITH A DISTRIBUTED PROPULSION SYSTEM HAVING SUCTION AND PRESSURE FANS

Номер: US20220063796A1
Автор: BREEZE-STRINGFELLOW Andrew, Freely Brendan Michael, McMahon Michael James, Murrow Kurt David
Принадлежит:

An aircraft equipped with a distributed fan propulsion system and methods of operating such aircraft are provided. In one aspect, an aircraft includes a wing having a top surface and a bottom surface. The aircraft also has a distributed propulsion system that includes a suction fan array having one or more fans mounted to the wing and a pressure fan array having one or more fans mounted to the wing. The fans of the suction fan array are each positioned primarily above the top surface of the wing and the fans of the pressure fan array are each positioned primarily below the bottom surface of the wing. The fans of the suction fan array are controllable independent of the fans of the pressure fan array so that the air pressure above and/or below the wing can be locally controlled, allowing for adjustment of lift on the wing.

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Номер записи: 86
14-02-2019 дата публикации

Riblet film for reducing the air resistance of aircraft

Номер: US20190047684A1
Автор: Hegenbart Matthias, Linde Peter
Принадлежит:

A riblet film for reducing the air resistance of aircraft, comprising a suspension with magnetic particles enclosed therein. Depending on the pattern of a magnetic field acting on the riblet film, the magnetic field that is acting can be made visible, at least in certain regions, by changing the orientation of the magnetic particles. The riblet film allows an inspection of the aircraft structure located under the riblet film through the riblet film. 1. A riblet film for reducing air resistance of an aircraft , comprising:a suspension with magnetic particles enclosed therein,wherein, depending on a pattern of a magnetic field acting on the riblet film, the magnetic field that is acting is made visible, at least in certain regions, by changing an orientation of the magnetic particles.2. The riblet film according to claim 1 , wherein the riblet film comprises a layer with a multiplicity of microcapsules claim 1 , wherein the suspension is distributed among the microcapsules and respectively enclosed therein.3. The riblet film according to claim 1 , wherein the magnetic particles comprise colloidal nickel.4. The riblet film according to claim 1 , wherein at least one of the suspension or a layer comprising the microcapsules is applied to a substrate serving as a carrier film.5. The riblet film according to claim 1 , wherein the riblet film comprises a plurality of ribs which are arranged substantially parallel to one another and have tips.6. The riblet film according to claim 1 , wherein the riblet film comprises a plurality of riblets each having a substantially triangular cross section.7. The riblet film according to claim 1 , wherein the riblet film further comprises has an adhesive layer arranged to adhesively attach the riblet film onto an aerodynamic surface.8. An aircraft with a riblet film according to applied to an aerodynamic surface of the aircraft.9. A method for examining an aircraft structure to which a riblet film according to has been applied claim 1 , ...

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Номер записи: 87
14-02-2019 дата публикации

EJECTOR AND AIRFOIL CONFIGURATIONS

Номер: US20190047712A1
Автор: Evulet Andrei
Принадлежит:

A propulsion system coupled to a vehicle. The system includes an ejector having an outlet structure out of which propulsive fluid flows at a predetermined adjustable velocity. A control surface having a leading edge is located directly downstream of the outlet structure such that propulsive fluid from the ejector flows over the control surface. 1. A propulsion system for a vehicle , the system comprising:a primary airfoil coupled to the vehicle;a first augmenting airfoil coupled to the vehicle and positioned downstream of fluid flowing over the primary airfoil, the first augmenting airfoil comprising a first output structure and at least one first conduit coupled to the first output structure, the at least one first conduit configured to introduce to the first output structure a primary fluid produced by the vehicle, the first output structure comprising a first terminal end configured to provide egress for the introduced primary fluid toward the primary airfoil and out of the first augmenting airfoil; anda secondary airfoil located directly downstream of the first augmenting airfoil such that the fluid flowing over the primary airfoil and the primary fluid from the first augmenting airfoil flows over the secondary airfoil.2. The system of claim 1 , further comprising a second augmenting airfoil coupled to the vehicle and positioned downstream of fluid flowing over the primary airfoil claim 1 , each of the first and second augmenting airfoils having a leading edge disposed toward the primary airfoil claim 1 , the first augmenting airfoil opposing the second augmenting airfoil whereby:the first and second augmenting airfoils define a diffusing region; andthe leading edges define an intake region configured to receive and introduce to the diffusing region the primary fluid and the fluid flowing over the primary airfoil, the diffusing region comprising a primary terminal end configured to mix and provide the introduced primary fluid and fluid flowing over the primary ...

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Номер записи: 88
13-02-2020 дата публикации

Methods, Systems, and Apparatuses For In-Line Variably Porous Surfaces

Номер: US20200047875A1
Автор: Calkins Frederick Theodore, Kreitzman Jordan, Nicholson Douglas E.
Принадлежит:

Variably porous panels and panel assemblies incorporating shape memory alloy components along with methods for actuating the shape memory alloys are disclosed to predictably alter the porosity of a substrate surface, with the shape memory alloy maintained in an orientation relative to the panel that is in-plane with a mold-line of the panel outer surface. 1. A panel assembly comprising:a panel, said panel having a panel inner surface and a panel outer surface, said panel further comprising at least one through opening;a shape memory alloy component, said shape memory alloy component in contact with the panel inner surface, said shape memory alloy component having a first dimension in a non-activated state and said shape memory alloy component having at least a second dimension or a plurality of dimensions in an activated state, said second dimension or plurality of dimensions different from the first dimension;wherein the shape memory alloy component is maintained in an orientation relative to the panel, said orientation in-plane with a mold-line of the panel outer surface; andwherein said panel assembly comprises an initial porosity in a non-activated state, and said panel assembly comprises at least one variable porosity in an activated state, said variable porosity different from the initial porosity.2. The panel assembly of claim 1 , further comprising:an actuator in communication with the shape memory alloy component.3. The panel assembly of claim 2 , wherein the actuator comprises at least one of:a heating element, a cooling element, and a magnet.4. The panel assembly of claim 1 , further comprising a controller in communication with the actuator.5. The panel assembly of claim 1 , wherein the panel comprises at least one of: a metal claim 1 , a ceramic claim 1 , a composite material claim 1 , and combinations thereof.6. The panel assembly of claim 1 , wherein the panel comprises at least one of: aluminum claim 1 , aluminum alloy claim 1 , titanium claim 1 , ...

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Номер записи: 89
14-02-2019 дата публикации

LOW-NOISE AIRFOIL FOR AN OPEN ROTOR

Номер: US20190048724A1
Автор: Tweedt Daniel Lawrence
Принадлежит:

An airfoil section of a blade for an open rotor includes: a pressure side and a suction side, the pressure side and the suction side intersecting at a leading edge and a trailing edge, wherein a chord of the airfoil section is defined as a straight-line distance between the leading edge and the trailing edge; the airfoil section has a meanline defined midway between the pressure side and the suction side; and the meanline is shaped such that, in the presence of predetermined transonic or supersonic relative velocity conditions, maximum and minimum ideal Mach numbers on the suction side will lie within a 0.08 band, between 25% and 80% percent of the chord. 1. An airfoil section of a blade for an open rotor , comprising: a pressure side and a suction side , the pressure side and the suction side intersecting at a leading edge and a trailing edge , wherein a chord of the airfoil section is defined as a straight-line distance between the leading edge and the trailing edge;the airfoil section has a meanline defined midway between the pressure side and the suction side; andthe meanline is shaped such that, in the presence of predetermined transonic or supersonic relative velocity conditions, maximum and minimum ideal Mach numbers on the suction side will lie within a 0.08 band, between 25% and 80% percent of the chord.2. The airfoil section according to claim 1 , wherein a maximum camber rise of the meanline is located forward of 50% of the chord.3. The airfoil section according to claim 1 , wherein a thickness of the airfoil section is defined as a distance measured normal to the meanline between the pressure side and the suction side claim 1 , and wherein a maximum value of the thickness occurs at a location between about 20% to about 30% of the chord.4. The airfoil section according to claim 1 , wherein the airfoil section has a maximum thickness of about 2% to about 4% of the chord.5. The airfoil section according to claim 1 , wherein the airfoil section is configured ...

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Номер записи: 90
23-02-2017 дата публикации

AERODYNAMIC DEVICE

Номер: US20170050719A1
Автор: Akhmejanov Alibi
Принадлежит:

The invention relates to aviation equipment. An object of this invention is to develop a new aerodynamic device which can extend the range of aerodynamic devices for aviation, increase the efficiency of the air flow power use, increase the efficiency of the lifting force and improve the efficiency of controlling the wing resultant forces. For this purpose, the aerodynamic device has an aerodynamic wing () with a blower () of gaseous working fluid (such as air) mounted above the wing (), in accordance with the invention, the aerodynamic wing () has a specific shape it is designed in the form of a double-curved open surface made up by a system of longitudinal grooves () along the whole wing surface The wing () has a convergent segment () and a divergent segment (); between the convergent and the divergent segments there is a smooth transitional segment (). The wing outlines have end elements (). In the convergent and the divergent segments of the wing lower surface which is not blown by air, there is a controlled drive system () for the wing surface cambering and area changing. The divergent segment tip on the wing trailing edge has a deflectable controlled element (). The structural parts of the present invention meet special conditions. 1. (canceled)2. (canceled)3. An aerodynamic device , comprising: a convergent wing segment;', 'a divergent wing segment, wherein at least one of convergent wing segment and divergent wing segment is positioned along a trajectory of fluid flow;, 'an aerodynamic wing, the wing comprisinga fluid blower, connected to the wing; anda controlled drive mechanism, connected to an area of the wing not engaged by air flow;wherein the controlled drive mechanism is configured to allow at least one of cambering and area changing of a surface of the wing.4. The device of claim 1 , wherein the wing has a form of a double-curved open surface made up by at least one longitudinal groove along the surface of the wing claim 1 , and wherein the at least ...

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Номер записи: 91
22-02-2018 дата публикации

PROPULSION ENGINE FOR AN AIRCRAFT

Номер: US20180050810A1
Автор: Cheung Lawrence Chih-hui, Marrinan Patrick Michael, Miller Brandon Wayne, Niergarth Daniel Alan, Pastouchenko Nikolai N., Vondrell Randy M.
Принадлежит:

A propulsion system for an aircraft includes an electric propulsion engine. The electric propulsion engine includes an electric motor and a fan rotatable about a central axis of the electric propulsion engine by the electric motor. The electric propulsion engine also includes a bearing supporting rotation of the fan and a thermal management system. The thermal management system includes a lubrication oil circulation assembly and a heat exchanger thermally connected to the lubrication oil circulation assembly. The lubrication oil circulation assembly is configured for providing the bearing with lubrication oil. Such an electric propulsion engine may be a relatively self-sufficient engine. 1. A propulsion system for an aircraft having an aft end , the propulsion system comprising: an electric motor;', 'a fan rotatable about the central axis of the electric propulsion engine by the electric motor;', 'a bearing supporting rotation of the fan; and', a lubrication oil circulation assembly for providing the bearing with lubrication oil; and', 'a heat exchanger thermally connected to the lubrication oil circulation assembly., 'a thermal management system comprising'}], 'an electric propulsion engine defining a central axis, the electric propulsion engine comprising'}2. The propulsion system of claim 1 , wherein the lubrication oil circulation assembly comprises a lubrication oil supply pump and a lubrication oil scavenge pump.3. The propulsion system of claim 1 , further comprising:a sump enclosing the bearing, wherein the lubrication oil circulation assembly is fluidly connected to the sump.4. The propulsion system of claim 1 , further comprising:an accessory gear box dedicated to the electric propulsion engine.5. The propulsion system of claim 4 , wherein the accessory gear box is driven by the electric motor.6. The propulsion system of claim 4 , wherein the lubrication oil circulation assembly comprises a lubrication oil supply pump and a lubrication oil scavenge pump ...

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Номер записи: 92
22-02-2018 дата публикации

THERMAL MANAGEMENT SYSTEM FOR AN ELECTRIC PROPULSION ENGINE

Номер: US20180051716A1
Автор: Cheung Lawrence Chih-hui, Miller Brandon Wayne, Pastouchenko Nikolai N., Ramakrishnan Kishore
Принадлежит:

A propulsion system for an aircraft includes an electric propulsion engine configured to be mounted at an aft end of the aircraft. The electric propulsion engine includes an electric motor and a fan rotatable about a central axis, the fan driven by the electric motor. The electric propulsion system additionally includes a cooling system operable with an airflow over the aft end the aircraft when the electric propulsion system is mounted to the aircraft. The cooling system is configured to cool the electric motor during operation of the electric propulsion engine. 1. A propulsion system for an aircraft having an aft end , the propulsion system comprising: an electric motor;', 'a fan rotatable about the central axis and driven by the electric motor; and', 'a cooling system operable with an airflow over the aft end of the aircraft when the electric propulsion engine is mounted to the aircraft, the cooling system configured to cool the electric motor during operation of the electric propulsion engine., 'an electric propulsion engine configured to be mounted at the aft end of the aircraft, the electric propulsion engine defining a central axis and comprising'}2. The propulsion system of claim 1 , wherein the cooling system comprises a closed loop in thermal communication with the electric motor claim 1 , and wherein the closed loop is configured to flow a thermal transfer fluid therethrough.3. The propulsion system of claim 2 , wherein the cooling system further comprises a heat exchanger claim 2 , wherein the heat exchanger is in thermal communication with the thermal transfer fluid in the closed loop and the airflow over the aft end of the aircraft claim 2 , and wherein the heat exchanger is configured to remove heat from the thermal transfer fluid within the closed loop and transfer such heat to the airflow over the aft end of the aircraft.4. The propulsion system of claim 3 , wherein the electric propulsion engine further comprises a forward support member and an ...

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Номер записи: 93
10-03-2022 дата публикации

Tiltrotor Propulsion System for an Aircraft

Номер: US20220073199A1
Автор: Crabtree Glenn, Murrow Kurt David, Polakowski Matthew Ryan, Vondrell Randy M., Zatorski Darek Tomasz
Принадлежит:

A propulsion system of an aircraft has at least one unducted fan and at least one ducted fan, the at least one unducted fan and the at least one ducted fan being powered by an electric power source and rotatable between a vertical thrust position and a forward thrust position, and a controller configured to distribute electrical power between the at least one unducted fan and the at least one ducted fan. During a first mode when the at least one unducted fan and the at least one ducted fan are in the vertical thrust position, the controller is configured to distribute the electrical power between the plurality of unducted fans and the plurality of ducted fans such that the at least one unducted fan is a primary source of thrust. 1. A propulsion system of an aircraft comprising:at least one unducted fan and at least one ducted fan, the at least one unducted fan and the at least one ducted fan being powered by an electric power source and rotatable between a vertical thrust position and a forward thrust position; anda controller configured to distribute electrical power between the at least one unducted fan and the at least one ducted fan,wherein, during a first mode when the at least one unducted fan and the at least one ducted fan are in the vertical thrust position, the controller is configured to distribute the electrical power between the plurality of unducted fans and the plurality of ducted fans such that the at least one unducted fan is a primary source of thrust, andwherein, during a second mode when the at least one unducted fan and the at least one ducted fan are in the forward thrust position, the controller is configured to distribute the electrical power between the plurality of unducted fans and the plurality of ducted fans such that the at least one ducted fan is the primary source of thrust.2. The propulsion system of claim 1 , wherein claim 1 , in the first mode claim 1 , the at least one unducted fan produces more thrust than the at least one ducted ...

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Номер записи: 94
21-02-2019 дата публикации

Embedded engines in hybrid blended wing body

Номер: US20190055006A1
Автор: David J. Arend, John D. Wolter, Razvan Virgil FLOREA, Thomas G. Tillman, William T. COUSINS
Принадлежит: United Technologies Corp

A hybrid wing aircraft has an engine embedded into a body of the hybrid wing aircraft. The embedded engine has a fan that is received within a nacelle. The body of the aircraft provides a boundary layer over a circumferential portion of a fan. A system delivers additional air to correct fan stability issues raised by the boundary layer.

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Номер записи: 95
01-03-2018 дата публикации

BIOMIMETIC AIRFOIL BODIES AND METHODS OF DESIGNING AND MAKING SAME

Номер: US20180057141A1
Автор: Shormann David E.
Принадлежит:

An airfoil body may include a plurality of tubercles along a leading edge of the airfoil body and a plurality of crenulations along a trailing edge of the airfoil body, wherein at least one of a position, a size, and a shape of the plurality of tubercles and the plurality of crenulations varies in a non-periodic fashion. The non-periodic fashion may be according to a Fibonacci function and may mimic the configuration of a pectoral fin of a humpback whale. The tubercles and crenulations may be defined with respect to a pivot point. The spanwise profile, including the max chord trailing edge curvature, may closely follow divine spirals and related Fibonacci proportions. The spanwise chord thickness may vary in a nonlinear pattern. Related methods are also described. 1. An airfoil body comprising:a plurality of tubercles along a leading edge of said airfoil body; anda plurality of crenulations along a trailing edge of said airfoil body;wherein at least one of a position, a size, and a shape of said plurality of tubercles and said plurality of crenulations varies in a non-periodic fashion.2. The airfoil body of wherein said non-periodic fashion comprises a Fibonacci ratio.3113. The airfoil body of wherein said plurality of tubercles comprises 13 tubercles T-T.4146789101111. The airfoil body of wherein said plurality of tubercles comprises 8 primary tubercles T claim 3 , T claim 3 , T claim 3 , T claim 3 , T claim 3 , T claim 3 , T claim 3 , and T having peaks respectively located within about ±0.05 of the following proportions: 0.38 claim 3 , 0.62 claim 3 , 0.76 claim 3 , 0.86 claim 3 , 0.9 claim 3 , 0.95 claim 3 , 0.95 claim 3 , 1.0; said proportions being defined with respect to a maximum span segment between a pivot point located inboard from a root chord of said airfoil body and said peak of said tubercle T.7. The airfoil body of wherein at least some of said tubercle peaks are located on a divine spiral.8. The airfoil body of wherein at least some of said ...

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Номер записи: 96
01-03-2018 дата публикации

Deployable assembly for a propulsor

Номер: US20180057150A1
Автор: Kishore Ramakrishnan, Lawrence Chih-Hui Cheung, Nikolai N. PASTOUCHENKO
Принадлежит: General Electric Co

An aircraft includes a fuselage extending between a forward end and an aft end. The aircraft additionally includes a propulsor mounted to the fuselage at the aft end of the fuselage, the propulsor including an outer nacelle and the outer nacelle defining an inlet. Additionally, the aircraft includes a deployable assembly attached to at least one of the fuselage or the outer nacelle, the deployable assembly movable between a stowed position and an engaged position. The deployable assembly alters an airflow towards the propulsor or into the propulsor through the inlet defined by the outer nacelle when in the engaged position to increase an efficiency of the aft fan and/or of the aircraft.

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Номер записи: 97
01-03-2018 дата публикации

NACELLE FOR AN AIRCRAFT AFT FAN

Номер: US20180057181A1
Автор: Malcevic Ivan, Pastouchenko Nikolai N., Yao Jixian
Принадлежит:

An aircraft is provided including a fuselage extending between a forward end and an aft end. An aft engine is mounted to the aft end of the fuselage and defines a centerline. The aft engine further includes a nacelle having a forward transition duct at the forward end of the nacelle. The forward transition duct also defines a centerline and the centerline of the forward transition duct is angled downward relative to the centerline of the aft engine. 1. An aircraft defining a longitudinal direction and comprising:a fuselage extending between a forward end and an aft end along the longitudinal direction of the aircraft; andan aft engine mounted to the aft end of the fuselage and defining a centerline, the aft engine further comprising a nacelle comprising a forward transition duct at a forward end of the nacelle, the forward transition duct defining a centerline, the centerline of the forward transition duct angled downward relative to the centerline of the aft engine.2. The aircraft of claim 1 , wherein the forward transition duct of the nacelle includes a bottom section and a top section claim 1 , wherein the bottom section and the top section each define a camber line claim 1 , and wherein the centerline of the transition duct of the nacelle is a midpoint line between the camber line of the top section and the camber line of the bottom section.3. The aircraft of claim 2 , wherein the camber line of the top section of the forward transition duct and the camber line of the bottom section of the forward transition duct are within about five degrees of being parallel to one another.4. The aircraft of claim 1 , wherein the centerline of the forward transition duct defines an angle greater than or equal to about ten degrees with the centerline of the aft engine.5. The aircraft of claim 1 , wherein the forward transition duct of the nacelle includes a bottom section and a top section claim 1 , wherein the bottom section of the forward transition duct defines a forward ...

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Номер записи: 98
01-03-2018 дата публикации

AFT ENGINE NACELLE SHAPE FOR AN AIRCRAFT

Номер: US20180057182A1
Автор: Becker Thomas Lee, Marrinan Patrick Michael, Murrow Kurt David, Yao Jixian
Принадлежит:

An aircraft is provided including a fuselage and an aft engine. The fuselage defines a top side, a bottom side, and a frustum located proximate an aft end of the aircraft. The frustum defines a top reference line extending along the frustum at a top side of the fuselage, and a bottom reference line extending along the frustum at a bottom side of the fuselage. The top and bottom reference lines meet at a reference point aft of the frustum. The fuselage further defines a recessed portion located aft of the frustum and indented inwardly from the bottom reference line. The aft engine includes a nacelle extending adjacent to the recessed portion of the fuselage such that the aft engine may be included with the aircraft without interfering with, e.g., a takeoff angle of the aircraft. 1. An aircraft defining a longitudinal centerline , the aircraft comprising:a fuselage, extending between a forward end and an aft end of the aircraft, and defining a top side and a bottom side;a set of landing gear extending from the bottom side of the fuselage, wherein the bottom side of the fuselage and the landing gear together define a maximum takeoff angle with the longitudinal centerline, the fuselage further defining a back portion at the bottom side of the fuselage located proximate the aft end of the aircraft and defining an angle with the longitudinal centerline greater than the maximum takeoff angle; andan aft engine located aft of the pair of wings and including a nacelle extending adjacent to the back portion of the fuselage.2. The aircraft of claim 13 , wherein the aircraft defines a mean line extending from a forward end to an aft end claim 13 , and wherein the nacelle extends radially around at least a portion the mean line of the aircraft.3. The aircraft of claim 13 , wherein the nacelle defines an inlet with the fuselage claim 13 , and wherein the inlet extends radially around at least portion of the fuselage.4. The aircraft of claim 13 , wherein the aft engine defines a ...

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Номер записи: 99
01-03-2018 дата публикации

Thermal protection and drag reduction method and system for ultra high-speed aircraft

Номер: US20180057191A1
Автор: GUO Chunhai, Song Tao, Tong Jingyu, XIANG Shuhong, Yang Yang, ZHANG Tianrun, Zhang Wenwu
Принадлежит:

Disclosed are the thermal protection and drag reduction method and system for an ultra high-speed aircraft. A cold source is and a cold source driving device are arranged inside a cavity of the ultra high-speed aircraft. A plurality of micropores are arranged on a wall surface of the cavity. The cold source driving device comprises an air pump, a cold source reservoir and a buffer. The air pump supplies compressed air to a cold source reservoir during operation. The cold source enters the buffer and is vaporized under the action of air pressure. High-pressure gas is ejected from the micropores to form a gas film on the outer surface of the cavity. The gas film not only can perform thermal protection on the ultra high-speed aircraft, but also can effectively reduce viscous drag between the aircraft and the external gas, by virtue of which the thermal barrier phenomenon is alleviated or eliminated. Therefore, security of the ultra high-speed aircraft is improved and service life is prolonged. 1. A thermal protection and drag reduction method for ultra high-speed aircraft , comprisingproviding a cold source inside a cavity of the ultra high-speed aircraft,arranging a plurality of micropores on a wall surface of the cavity of the ultra high-speed aircraft, wherein the cold source is ejected from the micropores in the form of high pressure gas under the action of driving force, so as to form a gas film on an outer surface of the cavity.2. The thermal protection and drag reduction method for ultra high-speed aircraft according to claim 1 , wherein the micropores are provided at a nose cone portion and/or an empennage portion of the cavity of the ultra high-speed aircraft.3. The thermal protection and drag reduction method for ultra high-speed aircraft according to claim 1 , wherein the micropores are regularly distributed on the wall surface of the cavity of the ultra high-speed aircraft.4. The thermal protection and drag reduction method for ultra high-speed aircraft ...

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Номер записи: 100