СЕТЧАТЫЙ ИОННЫЙ ДВИГАТЕЛЬ С НАХОДЯЩИМСЯ В НЕМ ТВЕРДЫМ РАБОЧИМ ТЕЛОМ

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ПУЛЬСИРУЮЩИЙ ДЕТОНАЦИОННЫЙ ДВИГАТЕЛЬ С МАГНИТОГИДРОДИНАМИЧЕСКИМ УПРАВЛЕНИЕМ ПОТОКОМ (ВАРИАНТЫ) И СПОСОБ УПРАВЛЕНИЯ ДЕТОНАЦИЕЙ

Изобретение относится к пульсирующим детонационным двигателям, в которых используется магнитогидродинамическое управление потоком. Пульсирующий детонационный двигатель включает трубу (12), имеющую открытый передний конец (16) и открытый задний конец (18), и топливно-воздушный вход (20), выполненный в трубе (12) на переднем конце (16). Зажигатель (24) расположен в трубе (12) в месте, находящемся между передним концом (16) и задним концом (18). Система магнитогидродинамического управления потоком расположена между зажигателем (24) и топливно-воздушным входом (20) для управления детонацией в трубе (12) впереди зажигателя (24). В системе магнитогидродинамического управления потоком используются магнитные и электрические поля впереди зажигателя (24) для рассеяния фронта (34) детонационного горения, распространяющегося вперед, или по меньшей мере уменьшения его потенциала зажигания. Технический результат - обеспечение работы системы управления клапанами или потоком с высокой частотой для пульсирующих ...

Двигательная установка с импульсным электрическим реактивным двигателем

Изобретение относится к электрореактивным двигателям импульсного типа и ДУ на их основе, использующим жидкофазные рабочие тела. Двигательная установка с импульсным электрическим реактивным двигателем состоит из собственно ЭРД импульсного действия с электродами и линейным разрядным промежутком на подвижной поверхности, бака хранения жидкого рабочего тела, трубопровода подачи рабочего тела с насосом подачи к капиллярному фитилю перед разрядным промежутком, после разрядного промежутка с подвижной поверхностью контактирует фитиль, сообщенный с трубопроводом отсоса рабочего тела, снабженного насосом и соединенного с баком хранения рабочего тела, зарядного устройства и накопителя электрической энергии емкостного типа. Согласно изобретению разрядные промежутки импульсного ЭРД выполнены на цилиндрической поверхности минимум двух конденсаторов с обкладками на боковых поверхностях дисков с высокой диэлектрической проницаемостью, входящих в состав емкостного накопителя, с промежуточными электродами ...

Система хранения и подачи иода

Изобретение относится к области электроракетных двигателей, в частности к системе хранения и подачи рабочего тела. В системе хранения и подачи иода, содержащей сообщенную с электроракетным двигателем трубопроводом, включающим клапан и нагреватели, цилиндрическую емкость с иодом, со стороны, противоположной трубопроводу, снабженную загрузочным фланцем и подпружиненным относительно него поршнем, контактирующим с другой стороны с кристаллическим иодом. Цилиндрическая емкость, со стороны трубопровода, содержит нагреватель и ресивер, при этом нагреватель установлен в полостях непересекающихся трубок, герметично вмонтированных в цилиндрическую поверхность емкости и размещенных по крайней мере в одной плоскости, перпендикулярной оси цилиндрической емкости, к наружным стенкам трубок, со стороны цилиндрической емкости, прикреплена металлическая сетка, при этом ресивер образован днищем цилиндрической емкости, со стороны трубопровода, и наружными стенками трубок с металлической сеткой, наружная цилиндрическая ...

ДЕТОНАЦИОННЫЙ РЕАКТИВНЫЙ ДВИГАТЕЛЬ

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КОММУТАЦИОННЫЙ УЗЕЛ ПОВОРОТНЫХ ПЕРЕКЛЮЧАТЕЛЕЙ ДЛЯ ИОННОЙ СИСТЕМЫ ОБЕСПЕЧЕНИЯ ПРОДВИЖЕНИЯ

Изобретение относится к транспорту, в частности к ионным двигателям. Система управления ионными двигателями содержит два устройства управления питанием, четыре ионных двигателя и два коммутационных узла. Один коммутационный узел соединен с двумя устройствами управления питанием и с двумя из четырех ионных двигателей. Другой коммутационный узел соединен с указанными двумя устройствами управления питанием и с другими двумя ионными двигателями. Каждый коммутационный узел имеет первое и второе коммутационные состояния, которые могут быть выбраны для обеспечения возможности подачи питания любым устройством управления питанием на любой ионный двигатель с первого по четвертый. Каждый коммутационный узел содержит полый вал, выполненный с возможностью поворота и приводимый в действие шаговым двигателем. Ионный двигатель содержит разрядный анод, разрядный катод, электрод устройства поддержания разряда, разрядный нагреватель, катод нейтрализатора, нагреватель нейтрализатора, экранную, ускорительную ...

ДВУНАПРАВЛЕННЫЙ ВОЛНОВОЙ ПЛАЗМЕННЫЙ ДВИГАТЕЛЬ ДЛЯ КОСМИЧЕСКОГО АППАРАТА

Изобретение относится к космической технике, в частности к электроракетным двигательным установкам с электрическим ракетным двигателем (ЭРД) с безэлектродным источником плазмы и ускорительной ступенью. Двунаправленный волновой плазменный двигатель для космического аппарата содержит газоразрядную камеру, определяющую ось сил тяги, антенну, модуль ВЧ-генератора, имеющий электрическую связь с антенной, магнитные системы, причем газоразрядная камера выполнена открытой во внешнюю атмосферу с двух противоположных торцов с возможностью формирования двух векторов тяги, противоположных друг другу по направлению и имеющих общую ось, являющуюся осью газоразрядной камеры, причем антенна расположена на внешней стороне газоразрядной камеры и с внешней своей стороны окружена кольцом из диэлектрического материала, при этом на каждом из противоположных концов газоразрядной камеры расположено по одной магнитной системе. При реализации изобретения обеспечивается снижение массы и габаритов двигателя, увеличение ...

Электроракетный двигатель для разгона и коррекции траектории космических аппаратов

Полезная модель относится к области ракетно-космической техники и может быть использована в космосе для разгона и коррекции траектории космических аппаратов.Электроракетный двигатель содержит корпус с накопителем и загрузочным устройством, расположенную в корпусе камеру сгорания с рабочим материалом, в качестве которого используется, преимущественно, пылеобразный материал, взятый с поверхности космических тел (луны, комет, астероидов и т.д.), высокотемпературный источник для получения твердых нано- и микрочастиц из рабочего материала, зарядную камеру и примыкающее к нему с торца разгонное устройство, где размещена система электродов, в котором первый по потоку электрод (катод) заряжен отрицательно, а последующий по потоку электрод (анод) заряжен положительно. Электроды разгонного устройства подключены к внешнему высоковольтному источнику тока, для создания разности электрических потенциалов между электродами. В качестве высокотемпературного источника нагрева используется электрический нагреватель ...

ИМПУЛЬСНЫЙ ПЛАЗМЕННЫЙ ТЕПЛОВОЙ АКТУАТОР ЭЖЕКТОРНОГО ТИПА

Изобретение относится к системам управления обтеканием летательного аппарата при дозвуковых и околозвуковых скоростях полета. Импульсный плазменный тепловой актуатор эжекторного типа содержит подводной канал с обратным клапаном, разрядную камеру со встроенными игольчатыми электродами, сопло эжектора, камеру смешения, полость разрежения со щелью, соединяющей полость разрежения с поверхностью крыла, выходной диффузор. Актуатор позволяет без перегрева рабочей области создавать истекающую из сопла высокоскоростную пульсирующую струю газа в одной области течения и одновременно осуществлять отсос пограничного слоя в другой. Изобретение направлено на расширение возможности управления обтеканием крыла летательного аппарата. 2 ил.

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Коаксиально-центробежный плазмореактивный двигатель предназначен для энергообеспечения вертикального взлета, посадки и управляемого полета дисколетов, являющихся альтернативой, опирающимся на воздух, самолетам и вертолетам. Равномерно распределенная по окружности дисколета достаточно мощная плазмореактивная тяга придает ему устойчивое равновесие, так как центр тяжести этого дисколета находится внутри, на значительном расстоянии от круговой, создающей опору плазма-тяги, а вращающаяся масса плазма-дисков образует гироскопический эффект, придающий дисколету дополнительную остойчивость при управляемом движении. К тому же, для образования этой тяги не требуется тяжелых механических узлов, снижающих ресурс двигателя и отрицательно влияющих и ограничивающих его удельную мощность. Прямое преобразование химической энергии РТ посредствам сил Лоренца и образованных ими коаксиально центробежных сил в энергию ПТ способствует высокому КПД этого преобразования. Применение плазмообразующего РТ позволяет ...

МНОГОФУНКЦИОНАЛЬНАЯ ЭЛЕКТРОРЕАКТИВНАЯ ДВИГАТЕЛЬНАЯ ПОДСИСТЕМА КОСМИЧЕСКОГО АППАРАТА

Изобретение относится к области управления движением космических аппаратов (КА) с помощью электрореактивных двигателей (ЭРД). Многофункциональная электрореактивная двигательная подсистема космического аппарата содержит блоки коррекции, силовые приборы, фильтры защиты от электростатических разрядов, разрядные фильтры, коммутаторы. В качестве источников питания применены одноканальные нерезервированные силовые приборы, один из которых находится в холодном резерве. Подключение блоков коррекции и силовых приборов осуществлено посредством одного или более коммутаторов двигателей, каждый из которых обеспечивает подключение любого силового прибора к любому блоку коррекции соответствующего коммутатора. Количество одноканальных нерезервированных силовых приборов на единицу больше, чем количество одновременно включенных двигателей. Фильтры защиты от электростатических разрядов расположены в коммутаторах двигателей для каждого блока коррекции. При использовании изобретения достигается применение минимальной ...

Волновой ионный двигатель с замкнутой газоразрядной камерой

Изобретение относится к космической технике, в частности к электроракетным двигательным установкам с электрическим ракетным двигателем (ЭРД) с безэлектродным источником плазмы и электродной ускорительной ступенью. Предложенный волновой ионный двигатель с замкнутой газоразрядной камерой содержит: газоразрядную камеру замкнутой кольцевой формы; минимум одну направляющую трубку; минимум одну антенну; минимум одну втулку (по количеству антенн); ВЧ-генератор; магнитную систему; источник питания магнитной системы; минимум одну ионно-оптическую систему (по количеству направляющих трубок); источник питания ионно-оптической системы; радиальный газоввод; систему хранения и подачи рабочего тела; модуль преобразования бортового питания; управляющий модуль. При реализации заявленного изобретения обеспечивается устранение возникновения паразитных разрядов, разрушающих элементы конструкции двигателя и малого космического аппарата; уменьшение потерь при вкладе мощности в плазму на электромагнитной линии ...

ИМПУЛЬСНЫЙ ПЛАЗМЕННЫЙ УСКОРИТЕЛЬ И СПОСОБ УСКОРЕНИЯ ПЛАЗМЫ

Изобретение относится к плазменной технике и к плазменным технологиями, а более конкретно - к плазменным ускорителям. Импульсный плазменный ускоритель содержит два электрода, установленные между электродами диэлектрические шашки, выполненные из аблирующего материала, разрядный канал с открытой торцевой частью, стенки которого образованы поверхностями электродов и диэлектрических шашек, накопитель энергии, токоподводы, соединяющие электроды с накопителем энергии, которые совместно с электродами и накопителем образуют внешнюю электрическую цепь, изолятор, установленный между электродами у торцевой части разрядного канала, противоположной открытой торцевой части, и устройство инициирования разряда. Характеристики внешней электрической цепи ускорителя выбираются из условия: 2≤C/L, где С - электрическая емкость внешней электрической цепи в мкФ, а L - индуктивность внешней электрической цепи в нГн, величина которой удовлетворяет условию: L≤100 нГн. Способ ускорения плазмы заключается в зажигании ...

Электростатический плазменный двигатель космического аппарата на заряженных частицах для работы в космическом пространстве

Изобретение относится к области ракетно-космической техники и может быть использовано в космосе для межорбитальных буксиров и длительных космических межпланетных перелетов. Двигатель содержит корпус, расположенную в корпусе камеру испарения с рабочим веществом, высокотемпературный источник для разложения рабочего вещества до атомарного уровня и формирования твердых нано- и микрочастиц, зарядную камеру и примыкающему к нему с торца разгонное устройство, где размещена система электродов, в котором первый электрод имеет отрицательный электрический потенциал и размещен на входе разгонного устройства, а второй электрод, имеющий положительный потенциал, размещен на его выходе, при этом электроды размещены друг от друга на расстоянии, исключающем электрический пробой между ними. Электроды разгонного устройства подключены к внешнему высоковольтному источнику тока, для создания разности электрических потенциалов между электродами, обеспечивающих поток заряженных нано- и микрочастиц, истекающих в ...

Способ ускоренного определения ресурса элементов двигателя с замкнутым дрейфом электронов

Использование: в космической технике при наземной отработке новых моделей двигателей с замкнутым дрейфом электронов (ДЗДЭ) и при переводе их на альтернативные рабочие вещества. Способ ускоренного определения ресурса элементов ДЗДЭ, заключающийся в последовательном выполнении циклов работы двигателя, включающих нанесение на поверхность исследуемого элемента многослойного покрытия, состоящего из чередующихся пар оптически контрастных слоев, кратковременные испытание двигателя до полного распыления нанесенного покрытия, определение профиля эрозии многослойного покрытия по картине распыления, расчетное прогнозирование профиля эрозии за заданное время, механическая обработка исследуемого элемента с целью придания ему рассчитанной формы. Отличие способа заключается в использовании многослойных покрытий для определения скорости эрозии исследуемой поверхности, что многократно снижает время эксперимента. Для определения коэффициентов пропорциональности между скоростью распыления материала элемента ...

ИМПУЛЬСНЫЙ ПЛАЗМЕННЫЙ УСКОРИТЕЛЬ И СПОСОБ УСКОРЕНИЯ ПЛАЗМЫ

... 1. Импульсный плазменный ускоритель, содержащий два электрода, установленные между электродами диэлектрические шашки, выполненные из аблирующего материала, разрядный канал с открытой торцевой частью, стенки которого образованы поверхностями электродов и диэлектрических шашек, накопитель энергии, токоподводы, соединяющие электроды с накопителем энергии, которые совместно с электродами и накопителем образуют внешнюю электрическую цепь, изолятор, установленный между электродами у торцевой части разрядного канала, противоположной открытой торцевой части, и устройство инициирования разряда, отличающийся тем, что характеристики внешней электрической цепи ускорителя выбраны из условия: 2≤ C/L, где С - электрическая емкость внешней электрической цепи в мкФ, L - индуктивность внешней электрической цепи в нГн, величина которой удовлетворяет условию: L≤ 100 нГн. 2. Ускоритель по п.1, отличающийся тем, что характеристики внешней электрической цепи ускорителя выбраны из условия: 2≤ C/L≤ 5. 3. Ускоритель ...

ДВИГАТЕЛЬ НА ЭФФЕКТЕ ХОЛЛА

... 1. Двигатель (1) на эффекте Холла, содержащий:кольцевой канал (2), ограниченный внутренней стенкой (3) и внешней стенкой (4), которые расположены концентрически вокруг центральной оси (X), причем кольцевой канал (2) имеет открытый нижний по потоку край и закрытый верхний по потоку край, а внутренняя стенка (3) выполнена подвижной в аксиальном направлении;привод (22) для перемещения указанной внутренней стенки (3) в аксиальном направлении;электрический контур (21), содержащий анод (9), расположенный на верхнем по потоку крае кольцевого канала (2), катод (19), расположенный на нижнем по потоку крае кольцевого канала (2), и источник (20) электрического напряжения между указанными анодом (9) и катодом(19);инжекционный контур (11) для инжекции потока газообразного рабочего тела в кольцевой канал (2);и магнитный контур для генерирования магнитного поля (М) на нижнем по потоку крае кольцевого канала (2),отличающийся тем, что подвижная внутренняя стенка (3) имеет диаметр, уменьшающийся в направлении ...

ПЛАЗМЕННЫЙ УСКОРИТЕЛЬ С ЗАМКНУТЫМ ДРЕЙФОМ ЭЛЕКТРОНОВ

Изобретение относится к области космической техники и может быть использовано в электроракетных двигателях, в частности в стационарных плазменных двигателях (СПД), а также в технологических плазменных ускорителях, применяемых в вакуумно-плазменной технологии. В плазменном ускорителе с замкнутым дрейфом электронов, включающем по меньшей мере один катод-компенсатор, разрядную систему и магнитную систему, содержащую тыльный магнитопровод, внутренний и наружный магнитопроводы, внутренний и наружный магнитные полюса и по меньшей мере один источник намагничивающей силы, концевые участки внутреннего и наружного магнитных полюсов выполнены дугообразной формы с радиусом кривизны таким, что в зону выхода разрядной системы выдвинуты только края внутреннего и наружного магнитных полюсов. Внутренний магнитопровод с внутренним магнитным полюсом и наружный магнитопровод с наружным магнитным полюсом сопряжены дугообразно. Тыльный магнитопровод с внутренним магнитопроводом могут быть сопряжены дугообразно ...

ЭЛЕКТРИЧЕСКИЙ ИМИТАТОР СТАЦИОНАРНОГО ПЛАЗМЕННОГО ДВИГАТЕЛЯ

Изобретение относится к средствам и методам имитации работы стационарных плазменных двигателей при проведении проверок и испытаний систем питания и управления электроракетных двигательных установок. Имитатор включает блоки имитации разрядного промежутка, поджига катода, нагревателя катода, магнитной системы, регулятора расхода рабочего тела и клапанов подачи рабочего тела. Имитатор содержит микроконтроллер с аналого-цифровым преобразователем, вход которого подключен к шине аналоговых сигналов, выход микроконтроллера подключен к шине адреса данных. Каждый блок имитации включает датчики тока и напряжения силовой цепи, выходы которых подключены к шине аналоговых сигналов, и преобразователь сигнала. Каждый преобразователь сигнала содержит гальваническую развязку, вход которой подключен к шине адреса данных, дешифратор адреса, умножающий цифро-аналоговый преобразователь, вход данных которого подключен к выходу гальванической развязки, а вход записи данных - к выходу дешифратора адреса, делитель ...

ПЛАЗМЕННЫЙ УСКОРИТЕЛЬ С ЗАМКНУТЫМ ДРЕЙФОМ ЭЛЕКТРОНОВ

Изобретение относится к области космической техники. Плазменный ускоритель с замкнутым дрейфом электронов включает по меньшей мере один катод-компенсатор, разрядную систему, содержащую разрядную камеру, образованную со стороны выхода внутренним и наружным кольцами, примыкающими соответственно к внутреннему и наружному торцам полого магнитного анода. Анод состоит из внешней и внутренней магнитопроводящей стенок, между которыми образована полость газового распределителя с каналами подвода и инжекции рабочего тела в разрядную камеру. Магнитная система содержит магнитопровод, внутренний и наружный магнитные полюса, по меньшей мере один источник намагничивающей силы, а также внутреннюю магнитопроводящую стенку анода, расположенную с немагнитными зазорами относительно внутреннего и наружного магнитных полюсов. Соответственно внешняя стенка полого анода выполнена из магнитопроводящего материала так, что по меньшей мере прианодная область в разрядной камере магнитоизолирована. Каналы инжекции рабочего ...

СИСТЕМА, УСТРОЙСТВО И СПОСОБ ГЕНЕРИРОВАНИЯ НАПРАВЛЕННЫХ СИЛ ПУТЕМ ВВОДА УПРАВЛЯЕМОЙ ПЛАЗМЕННОЙ СРЕДЫ В АСИММЕТРИЧНЫЙ КОНДЕНСАТОР

... 1. Способ создания силы с использованием двигателя на основе асимметричного конденсатора, характеризующийся тем, что: ! облучают электромагнитным излучением частицы, находящиеся в среде вблизи двигателя на основе асимметричного конденсатора, содержащего, по меньшей мере, три электрода с разными площадями поверхности, и находящиеся на расстоянии друг от друга; ! прикладывают напряжение, по меньшей мере, к одному из электродов для генерирования результирующей силы в двигателе на основе асимметричного конденсатора; и ! изменяют силу путем приложения напряжения, излучения или их комбинации к разным комбинациям электродов. ! 2. Способ по п.1, в котором двигатель на основе асимметричного конденсатора содержит, по меньшей мере, один анод и, по меньшей мере, первый катод и второй катод, при этом, по меньшей мере, первый катод расположен под другим углом относительно анода, чем второй катод, для создания комбинации из анода и первого катода и комбинации из анода и второго катода. ! 3. Способ по ...

УСТРОЙСТВО ДЛЯ ФОРМИРОВАНИЯ КВАЗИНЕЙТРАЛЬНОГО ПУЧКА ПРОТИВОПОЛОЖНО ЗАРЯЖЕННЫХ ЧАСТИЦ

Двигательная установка с импульсным электрическим реактивным двигателем

ДЕТОНАЦИОННЫЙ РЕАКТИВНЫЙ ДВИГАТЕЛЬ

ЖИДКОСТНЫЙ РАКЕТНЫЙ ДВИГАТЕЛЬ С ИОНИЗАТОРОМ В КАМЕРЕ СГОРАНИЯ И ПОДОГРЕВОМ ГАЗОВ ПЛАЗМОЙ В СОПЛЕ

Изобретение относится к ракетной технике для использования в космических аппаратах, в том числе стартующих с поверхности земли или летящих по баллистическим траекториям. Двигатель содержит малогабаритную камеру сгорания, керамическое коническое сопло, тракты их охлаждения, одну или несколько форсунок. Для увеличения тяги жидкостного ракетного двигателя предложено комбинированное техническое решение, позволяющее значительно увеличить удельный импульс при некотором оптимальном увеличении веса дополнительных деталей, узлов и приборов. Предложено использовать подогрев потока газов плазмой в керамическом, охлаждаемом, коническом сопле индуктором электромагнитного высокочастотного поля с высокой удельной мощностью, превышающей удельную тепловую мощность камеры сгорания. Количество индукторов может быть более одного, работающих на кратных частотах в разрешенных диапазонах и размещенных вдоль конического сопла определенных размеров. Основной индуктор (индукторы) обязательно размещается между критическим ...

Двигательная подсистема довыведения и коррекции орбиты геостационарного космического аппарата

Изобретение относится к области управления движением космических аппаратов (КА) с помощью электрореактивных двигателей (ЭРД). Двигательная подсистема содержит блоки коррекции на базе высоковольтных плазменных двигателей, бак высокого давления, блок подачи рабочего тела, систему преобразования и управления, фильтры защиты от электростатических разрядов, разрядные фильтры, коммутаторы. Двигательная подсистема реализует режим довыведения космического аппарата на геостационарную орбиту (ГСО) с повышенной тягой двигателя и режим коррекции орбиты с наибольшим значением удельного импульса двигателя, перечисленные режимы обеспечиваются одним и тем же набором оборудования двигательной подсистемы за счет изменения напряжения и тока разряда двигателей, формируемых системой преобразования и управления, при одинаковой потребляемой мощности. Изобретение позволяет решить задачу увеличения тяги в режиме довыведения КА на ГСО, сокращения времени довыведения и обеспечения высокой экономичности в режиме коррекции ...

Многополостной катод для плазменного двигателя

Полезная модель относится к плазменной технике, а именно к катодам, которые могут быть использованы в плазменных ракетных двигателях. Многополостной катод для плазменного двигателя содержит катодную трубку из тугоплавкого металлического материала, с одного из торцов которой расположена распределительная полость для подачи газа, и множество излучателей, выполненных в виде цилиндрических стержней из тугоплавкого металлического материала, закрепленных на противоположном от распределительной полости торце трубки, размещенных параллельно друг другу с образованием цилиндрического пучка и сопряженных между собой с образованием между ними полостей в виде продольных каналов. Катодная трубка из тугоплавкого металлического материала в области распределительной полости имеет внутренний диаметр больше внешнего диаметра окружности, описанной вокруг пучка цилиндрических стержней, расположенных внутри нее и не касающихся стенки вышеупомянутой трубки, а в противоположной от распределительной полости области ...

Устройство электропитания стационарного плазменного двигателя

Изобретение относится к электроракетной технике и может применяться при разработке систем питания и управления стационарными плазменными двигателями и двигательных установок на их основе. Устройство электропитания стационарного плазменного двигателя, включающее в себя источник питания разряда и поджига с функцией ограничения тока нагрузки, источник питания накала, источник питания магнита и поджига с функцией ограничения напряжения холостого хода, полупроводниковый диод и три ключа. Предложена конструкция устройства электропитания стационарного плазменного двигателя, которая позволяет изменять ток магнитных катушек двигателя во время его работы без изменения тока основного разряда, обеспечивает питание поджига постоянным током, использует одни и те же источники питания для питания поджига, магнита и основного разряда. При использовании изобретения повышается надежность и технологичность изделия при уменьшении массогабаритных характеристик, расширяются возможности управления двигателем и ...

Energiegewinnung mittels thermonuklearer Fusion sowie deren Nutzung zum Antrieb von Raumfahrzeugen

Verfahren und Vorrichtung zur Energiegewinnung sowie deren Nutzung zum Antrieb von Raumfahrzeugen mittels thermonuklearer Fusion durch Impact makroskopischer Massen hochbeschleunigtem Fusionsmaterials oder anderweitig erzeugtes hochdichtes Plasma wobei durch gleichzeitige Bestrahlung dieses Plasma instantan weiter aufgeheizt und so die Bedingungen für eine thermonukleare Reaktion geschaffen werden, wobei im Besonderen durch Neutronenstrahlen mit dem Fusionsmaterial entsprechend der exothermen Reaktion 6Li + n He + t +4.78 [MeV]Reaktionsgleichung 3genügend Energie freigesetzt wird um in einem Teil des Plasmas das Lawson Kriterium für die thermonukleare Fusion zu erfüllen. Von dort breitet sich die Reaktion 31 H + 21 H 10 n + 42 He H = 18,5 MeVReaktionsgleichung 1 solange weiter aus wie der Trägheitseinschluss erhalten bleibt. Die so freigesetzte Energie kann für verschiedene Anwendungen genutzt werden.

Plasma thruster

A plasma thruster comprises a plasma chamber 10 having first and second axial ends 14,16. The first end 14 is open and an anode 18 is located at the second axial end 16. A cathode 20 is located at the open end 14 of the chamber. The cathode and anode are arranged to produce an electric field having at least a component in the axial direction of the thruster. A propellant inlet is located at the closed end of the chamber and is connected to supply of propellant such as krypton (Kr), argon or xenon. A magnet system comprising a plurality of magnets 22 is spaced around the thruster axis, each magnet having its north and south poles spaced around the axis. The polarities of the magnet ends 22a, 22b are alternated. Coils 24 are arranged around the magnets and the current flowing in them is varied by a controller 26. The krypton propellant is ionised in the chamber by electrons and the ions are accelerated out of the chamber to provide thrust controlled by the controller 26.

Plasma propulsion engine

A plasma propulsion engine comprises a disc with a number of gas filled chambers located around its outer circumference. The disc is rotated at a high angular velocity by a motor, or similar device, and an electromagnetic field is used to switch the gas in the chambers from an unexcited state to a state of plasma discharge and back again. The switching of state is controlled such that gas contained in the chambers on one side of the rotating disc is always in a state of plasma discharge while the gas on the other side is always in an unexcited state. Rotating the disc and controlling plasma discharge as described above is said to produce a net thrust in a direction perpendicular to the axis of rotation. The disc, or a combination of discs, can be used to produce a stable platform with reduced weight, or to provide thrust for the propulsion of craft in the vacuum of space.

Resonant power supply for a thruster

Flying craft with magnetic field/electric arc vertical thrust producing means

A saucer-shaped craft has a central electrode on a post 8, and a ring electrode 13 whereby an arc may be struck therebetween to heat/ignite fuel introduced via conduits 7 and nozzles 11, to produce vertical thrust. Superconducting coils, forming rings 15, can generate a magnetic field to cause the arc to rotate. The magnetic field may also interact with the earth's magnetic field to cause a propulsive effect. Alternatively, a fan (18, fig. 3) may produce a downwards airflow subsequently heated by the arc struck between the electrodes.

High frequency burst pursed plasma thruster and high current thermo-controlled transistor

Thermo-electric propulsion device, method of operating a thermo-electric propulsion device and spacecraft

Improvements in or relating to apparatus for producing a jet consisting of a plasma of ions and electrons

... 897,577. Jet-propulsion plant. BRISTOL SIDDELEY ENGINES Ltd. June 27, 1960 [July 15, 1959], No. 24273/59. Class 110 (3). [Also in Groups XI and XL (a)] A "pulsed-electro ram-jet " for propulsion in space comprises a particle separator and accelerator 18 (see Group XL (a)) through which positive ions and electrons from concentrated ionization regions in space are introduced into a chamber 11 containing a cooled anode 12 and cathode 13 between which an arc is maintained by a constant potential to increase the energy of the particles. A coil 23 is supplied with a pulsed current so phased as to accelerate successive groups of particles which leave the jet nozzle at high velocity and produce useful thrust. In an alternative arrangement (Fig. 1, see Group XI), the chamber 11 is closed at the rear end and a gaseous or liquid working fluid, e.g. the products of combustion from a rocket combustion chamber, is introduced tangentially. The coolant may be liquid metal, e.g. sodium, potassium or mercury ...

KATODENANORDNUNG TO THE ATOMIZATION OF A SWIVELLING TARGET PIPE

DEVICE FOR THE DISTRIBUTION OF A GASEOUS COMPONENT IN IONIZATION CHAMBERS FOR SPACE PROPULSION

PROCEDURE AND DEVICE FOR THE CLUSTER FRAGMENTATION

Verfahren zur Ablenkung eines ionisierten Fluidumstromes aus seiner normalen Strömungsrichtung und Einrichtung zur Durchführung dieses Verfahrens.

Verfahren zur Herstellung eines Acylthio-hydroxy-steroids

Verfahren und Einrichtung zum Fördern von dielektrischen, strömungsfähigen Medien

Plasmagenerator und Verfahren zu dessen Betrieb

Plasmabeschleuniger

Process and devices allowing to increase the ionization of gases

Satellite propulsion device with plasma E-section for orientation control

Le dispositif embarqué permet de contrôler l'attitude d'un satellite à l'aide de l'énergie d'un faisceau de lumière cohérente. Le faisceau de lumière cohérente est obtenu à partir d'un générateur laser (3) excité directement par la lumière du soleil préalablement filtrée et concentrée (1)(2). Le faisceau laser est focalisé sur une matière appropriée à l'intérieure des tuyères (5) préalablement stockée (7) et répartie (4)(6) provocant par réaction les mouvements du satellite.

INDUCTION PLASMA GENERATOR WITH HIGH VELOCITY SHEATH

SPACE PROPULSION MODULE HAVING ELECTRIC AND SOLID-FUEL CHEMICAL PROPULSION

A space propulsion module in particular for equipping spacecraft such as satellites, probes, or indeed upper stages of rockets. According to the invention, this space propulsion module comprises a solid-fuel chemical thruster (10), having a main body (11), and at least one electric thruster (30), said at least one electric thruster (30) being mounted on said main body (11) of the solid-fuel chemical thruster (10).

FLUID FEED SYSTEM

GAS IONIZER DEVOID OF COAXIAL ELECTRODES

Method of obtaining mechanical energy utilizing H2O plasma generated in multiple steps

A method of obtaining mechanical energy utilizing H2O-plasma that is generated in multiple steps. The general field of art of the invention is that of producing a reactive thrust by using plasma. The mechanical energy provided by the invention is produced by explosion of electrically conductive plasma which is generated by dissociating H2O. At the first step H2O (gas) produced by a gasifier is reduced to a plasmatic state by electrical discharge. At the second step the plasmatic gas is treated by a further and stronger electrical discharge and by high-frequency induction heating, and the energy level of the plasma is raised to a point at which a plasma jet is ready to be produced. At the third step, the plasma jet is generated by periodically modulating the high voltage for the second electrical discharge, and a high-pressure thermal explosion reaction is caused by synchronizing the generation with compression of the plasma jet. The result is that energy produced by the plasmatic reaction ...

Closed drift hollow cathode

In accordance with one specific embodiment of the present invention, the closed drift hollow cathode comprises an axisymmetric discharge region into which an ionizable gas is introduced, an annular electron emitting cathode insert disposed laterally about that discharge region, a surrounding enclosure, an aperture in that enclosure disposed near the axis of symmetry and at one end of that region, and a magnetic field within that region which is both axisymmetric and generally disposed transverse to a path from the cathode insert to the aperture. An electrical discharge is established between the cathode insert and the enclosure. The electrons emitted from the cathode insert drift in closed paths around the axis, collide with molecules of ionizable gas, and sustain the discharge plasma by generating additional electron-ion pairs. Ions from the plasma bombard the cathode insert, thereby maintaining an emissive temperature. Electrons from the plasma diffuse to and escape through the aperture ...

Captive cover carton

СПОСОБ УПРАВЛЕНИЯ СТАЦИОНАРНЫМ ПЛАЗМЕННЫМ ДВИГАТЕЛЕМ

Изобретение относится к исследованию и эксплуатации электроракетных стационарных плазменных двигателей. В способе, включающем запуск двигателя, сравнение измеренных значений разрядного тока с верхним допустимым его значением, и в случае превышения предельного значения выключение двигателя с последующим его запуском. Перед запуском двигателя определяют диапазон превышения разрядным током своего допустимого значения определяют для каждого значения диапазона допустимый интервал времени пребывания двигателя под аномальной токовой нагрузкой и интервал времени защиты двигателя от аномальной токовой нагрузки, а в процессе работы двигателя, в случае превышения допустимого интервала времени пребывания двигателя под аномальной нагрузкой, производят его выключение с последующим включением через интервал времени защиты двигателя от аномальной токовой нагрузки. В случае соответствия допустимому интервалу времени пребывания двигателя под аномальной нагрузкой фиксируют частоту аномальных превышений на ...

Блок питания электроракетной двигательной установки спутника и система управления электроракетной двигательной установкой спутника

Группа изобретений относится к блоку питания и системе управления электроракетной двигательной установкой спутника. Блок питания содержит внутренний источник электроэнергии, внешний вход, первый и второй внешний выходы, выполненные с возможностью подачи в качестве выхода первого и второго электропитания, первый и второй переключательные элементы. Первый переключательный элемент снабжен первым внутренним входом, соединенным с внутренним источником, и первым внешним и внутренним выходами. Второй переключательный элемент снабжен выходом, соответствующим второму внешнему выходу, и внешним и вторым внутренним входами. Система управления содержит электронный управляющий блок, блоки питания электроракетной двигательной установки, электроракетные двигатели малой тяги. Повышается надежность электроракетных двигательных систем. 3 н. и 7 з.п. ф-лы, 13 ил.

ПЛАЗМЕННЫЙ ДВИГАТЕЛЬ И СПОСОБ ГЕНЕРИРОВАНИЯ ДВИЖУЩЕЙ ПЛАЗМЕННОЙ ТЯГИ

Изобретение относится к миниатюрному плазменному двигателю, при этом согласно изобретению: производят возбуждение плазмы микроразрядом с полым катодом вблизи выхода и внутри средства инжекции газообразного рабочего тела, при этом указанное средство инжекции является магнитным и содержит заострение на своем выходном конце, электроны намагниченной плазмы приводят в циклотронное вращение на уровне выходного конца указанного средства инжекции. Плазму поддерживают за счет электронно-циклотронного резонанса (ECR), при этом указанное средство инжекции выполняют металлическим и используют в качестве антенны электромагнитного (ЭМ) излучения, при этом объем плазмы в режиме резонанса ECR на выходе указанного средства инжекции используют в качестве резонатора электромагнитной волны, плазму ускоряют в магнитном реактивном сопле при помощи диамагнитной силы, при этом выбрасываемая плазма является электрически нейтральной. Изобретение направлено на повышение КПД двигателя при уменьшении его размеров.

Электротермический двигатель

Изобретение относится к реактивным двигателям, в частности к электротермическим двигателям. Электротермический двигатель содержит корпус, изолятор с капиллярным каналом, в котором с одной стороны установлен электрод, имеющий дросселирующий канал для подачи рабочей жидкости, а с другой стороны - сверхзвуковое сопло. В капиллярном канале создается плазменный разряд, и плазма подается в сверхзвуковое сопло. Дросселирующий канал электрода удален от торца электрода, обращенного в капиллярный канал, и при электроэрозионном износе торца электрода геометрические параметры дросселирующего канала не изменяются, поэтому не изменяются массовый расход жидкости и параметры двигателя. 2 ил.

Плазменный реактивный двигатель, использующий для создания тяги вытекающую через магнитное сопло плазму, нагретую мощным электромагнитным излучением, и способ создания реактивной тяги

Изобретение относится к способам создания реактивной тяги на основе безэлектродной плазменной магнитогидродинамики и позволяет повысить эффективность использования рабочего вещества в реактивных двигателях за счет увеличения скорости его вытекания. В предлагаемом способе используют излучение микроволнового диапазона длин волн в условиях ЭЦР в бесстолкновительном режиме, при этом энергия излучения передается в поперечную по отношению к магнитному полю энергию электронов, а в выходном магнитном сопле создают такое пространственное распределение магнитного поля, которое обеспечивает адиабатическое расширение потока плазмы, чем создают условия для перехода поперечной энергии электронов в продольную, а эффективное ускорение ионов плазмы происходит за счет амбиполярного поля разделения зарядов. В разработанном устройстве используется источник излучения микроволнового диапазона, квазиоптическая электродинамическая система, направляющая излучение в зону, где выполняются условия ЭЦР и происходит ...

Коаксиальный абляционный импульсный плазменный двигатель с векторизацией тяги

Изобретение относится к космической технике, в частности к электрическим ракетным двигателям (ЭРД), предназначенным для установки на космических аппаратах (КА). Коаксиальный абляционный импульсный плазменный двигатель с векторизацией тяги содержит втулку рабочего тела, поджигное устройство, внутренний электрод, внешний электрод, магнитное сопло, отклоняющие электромагниты, втулку внешних электромагнитов, внешние электромагниты, устройство питания электромагнитов, устройство питания емкостного накопителя энергии и генератора импульсов; емкостный накопитель энергии, генератор импульсов, согласующее устройство. Внутренний и внешний электроды образуют коаксиальный канал, имеющий закрытый и открытый концы. С закрытого конца канала размещается втулка рабочего тела. За срезом свободного конца коаксиального канала расположено магнитное сопло, на внешней стороне которого жестко присоединены отклоняющие электромагниты. Магнитное сопло и отклоняющие электромагниты находятся во внутренней полости втулки ...

ГАЗОЭЛЕКТРИЧЕСКАЯ РАЗВЯЗКА

Изобретение относится к плазменной технике. Газоэлектрическая развязка (ГЭР) входит в состав тракта подачи рабочего тела в газоразрядную камеру источника заряженных частиц. Входной и выходной патрубки (1, 2) выполнены из электропроводящего материала. В диэлектрическом корпусе (3) образован проточный канал, протяженность которого превышает его максимальный диаметр. Проточный канал (4) имеет осесимметричную форму и выполнен с сужением, образованным кольцеобразным выступом (5) на поверхности канала. Газопроницаемая вставка выполнена из электропроводящего материала и установлена на входе в выходной патрубок (2) с образованием электрического контакта с выходным патрубком. Вставка образована последовательно установленными газопроницаемыми элементами (6) с пористой структурой. Поверхности близлежащих газопроницаемых элементов (6), через которые осуществляется газообмен, контактируют между собой. Элементы (6) изготовлены из порошкового электропроводящего материала методом порошковой металлургии ...

Устройство для измерения массы рабочего тела, газообразного при нормальных условиях, в баллоне электроракетной двигательной установки и способ определения его массы

СПОСОБ УПРАВЛЕНИЯ СТАЦИОНАРНЫМ ПЛАЗМЕННЫМ ДВИГАТЕЛЕМ

ДВИГАТЕЛЬНАЯ УСТАНОВКА КОСМИЧЕСКОГО ЛЕТАТЕЛЬНОГО АППАРАТА И СПОСОБ

Система хранения и подачи иода

Электротермический узел парового двигателя для наноспутников

Электротермический узел парового двигателя для наноспутников относится к двигательным установкам. В частности, настоящее изобретение относится к узлу парового двигателя для малых космических аппаратов или наноспутников. Узел включает в себя теплообменники, каждый из которых содержит каналы, входное отверстие подачи рабочего тела в каналы первого по ходу потока теплообменника, по меньшей мере один нагревательный элемент, сверхзвуковое микросопло и множество стержней, образовывающих ферменную конструкцию. Теплообменники размещаются друг в друге и соединены таким образом, чтобы обеспечивалось последовательное прохождение по ним рабочего тела. При реализации изобретения повышается эффективность работы двигателя, за счет чего снижается энергопотребление при сохранении его небольших размеров. 8 з.п. ф-лы, 4 ил.

Einrichtung zur Erzeugung eines energiereichen Plasmas mittels eines elektrischen Lichtbogens

The unified field theory drive engine

The UFT drive engine consist basically of any material body that can allow electromagnetic waves to pass through it, which is then acted upon by linearly distributed magnets which collectively produces magnetic or @ fields only or produces simultaneously magnetic or @ fields which act on the electromagnetic waves to produce a deflecting force. This deflecting force produced acts on the material body through which the electromagnetic waves passes through to produce motion.

The unifield field theory (uft) drive engine

Optical recirculation with ablative drive

Remotely powered propulsion system

Plasma Thrusters

Abstract Plasma Thrusters A plasma thruster comprises a plasma chamber having first and second 5 axial ends, the first of which is open, an anode located at the second axial end, and a cathode. The cathode and anode are arranged to produce an electric field having at least a component in the axial direction of the thruster. A magnet system comprising a plurality of magnets is spaced around the thruster axis, each magnet having its north and south poles 10 spaced around the axis.

SOLID-STATE FLOW GENERATOR AND RELATED SYSTEMS, APPLICATIONS, AND METHODS

The invention, in various embodiments, is directed to a solid-state flow generator and related systems, methods and applications.

ROTARY SWITCH ASSEMBLY FOR ION PROPULSION SYSTEM

A gridded ion propulsion system comprising two power controllers, four ion thrusters, and two switch assemblies. One switch assembly is connected to the two power controllers and to two of the four ion thrusters. The other switch assembly is connected to the two power controllers and to the other two ion thrusters. Each switch assembly has first and second switching states which can be selected to enable either power controller to supply power to any one of the four ion thrusters. Each switch assembly comprises a respective movable body and a respective multiplicity of switches which change state in unison when the movable body changes position. For example, the movable body may be a rotatable hollow shaft driven by a stepper motor.

Engine with plasma for purposes of alternative pinchings

DEVICE FOR PRODUCING AND MANIPULATING PLASMA FROM A FLUID WITH RESONANT STRUCTURE AND ACTUATOR MICROWAVE PLASMA BY ELECTRIC FIELD

CATHODE ARRANGEMENT FOR SPUTTERING A ROTATABLE TARGET PIPE

sistema, aparelho, e método para gerar forças direcionais pela introdução de um ambiente de plasma controlado em um capacitor assimétrico

Method of obtaining mechanical energy utilizing H2 O plasma generated in multiple steps

A method of obtaining mechanical energy utilizing H2 O-plasma that is generated in multiple steps. The general field of art of the invention is that of producing a reactive thrust by using plasma. The mechanical energy provided by the invention is produced by explosion of electrically conductive plasma which is generated by dissociating H2 O. At the first step H2 O (gas) produced by a gasifier is reduced to a plasmatic state by electrical discharge. At the second step the plasmatic gas is treated by a further and stronger electrical discharge and by high-frequency induction heating, and the energy level of the plasma is raised to a point at which a plasma jet is ready to be produced. At the third step, the plasma jet is generated by periodically modulating the high voltage for the second electrical discharge, and a high-pressure thermal explosion reaction is caused by synchronizing the generation with compression of the plasma jet. The result is that energy produced by the plasmatic reaction ...

Systems and methods for generating electric current from hyperthermal chemical reaction

An electric generator is disclosed that includes a duct configured to direct hyperthermal air molecules toward a source of exothermic fuel, a fuel dispenser configured to dispense fuel into a flow of the hyperthermal air molecules to cause a hyperthermic chemical reaction between the fuel and the hyperthermal air molecules that produces highly internally excited air molecules. The excited molecules amplify a seed current. Power harvesting cells are configured to capture and convert the amplified current to electricity.

MAGNETIC GAS ENGINE AND METHOD OF EXTRACTING WORK

The present subject matter overcomes the deficiencies in the prior art by introducing or generating charged particles in an air stream and manipulating the air stream with magnetic fields operating on the charged particles. Embodiments of the present subject mater compress the air stream by accelerating charged particles with a moving magnetic field, where the magnetic field has a velocity perpendicular to its flux lines. The increased velocity of the charged particles increases the statistical mean particle velocity and thereby increases the pressure in the air stream. The compressed air stream is then heated and expanded through a second magnetic field. The expansion of the air stream substantially increases the velocity of the air stream and the charged particles therein. The interaction of the high velocity charged particles and the magnetic field imparts a force perpendicular to the flux lines, this force powers the movement of the magnetic field and can also be extracted in the form ...

Rotary Switch Assembly for Ion Propulsion System

A gridded ion propulsion system comprising two power controllers, four ion thrusters, and two switch assemblies. One switch assembly is connected to the two power controllers and to two of the four ion thrusters. The other switch assembly is connected to the two power controllers and to the other two ion thrusters. Each switch assembly has first and second switching states which can be selected to enable either power controller to supply power to any one of the four ion thrusters. Each switch assembly comprises a respective movable body and a respective multiplicity of switches which change state in unison when the movable body changes position. For example, the movable body may be a rotatable hollow shaft driven by a stepper motor. 1. An ion propulsion system comprising first and second power controllers , first and second ion thrusters and a first switch assembly having at least first and second switching states , wherein when said first and second power controllers are on , said first and second ion thrusters receive power from said first and second power controllers respectively via said first switch assembly when said first switch assembly is in said first switching state and receive power from said second and first power controllers respectively via said first switch assembly when said first switch assembly is in said second switching state , and wherein said first switch assembly comprises a first body which is movable and a first multiplicity of switches which change state in unison when said first body changes position , said first switch assembly being in said first switching state when said first body is in a first position and being in said second switching state when said first body is in a second position.2. The system as recited in claim 1 , wherein said first body comprises a center shaft that is rotatable and said first and second positions are first and second angular positions respectively of said center shaft.3. The system as recited in claim 2 , ...

Cathode assembly for sputtering a rotatable tubular target

Импульсный плазменный электрический реактивный двигатель

Изобретение относится к электрореактивным двигателям импульсного действия, использующим в качестве рабочего тела жидкофазную рабочую среду. Двигатель состоит из анода и катода с разрядным промежутком линейного типа, сформированных на цилиндрической поверхности диэлектрика, смоченного жидкофазным рабочим телом. Рабочая поверхность диэлектрика цилиндрического типа и разрядный промежуток с электродами катод и анод, установленными на изоляторе, выполнены подвижными. Возможность перемещения изолятора, содержащего катод и анод, по боковой поверхности диэлектрика цилиндрического типа, смоченного жидкофазным рабочим телом, обеспечивает изменения направления вектора тяги. 1 ил.

ПЛАЗМЕННЫЙ ДВИГАТЕЛЬ С ЗАМКНУТЫМ ДРЕЙФОМ ЭЛЕКТРОНОВ

Изобретение относится к области космической техники и может быть использовано в электроракетных двигателях, а также в технологических плазменных ускорителях, применяемых в вакуумно-плазменной технологии. В плазменном двигателе с замкнутым дрейфом электронов, содержащем разрядную камеру с наружной и внутренней кольцеобразными стенками, образующими ускорительный канал, полый газораспределитель с каналами подвода и отверстиями подачи рабочего газа, анод с по большей мере наружным и внутренним козырьками, расположенный в ускорительном канале, магнитопровод, внутренний и наружный магнитные полюса, образующие рабочий межполюсный промежуток, по меньшей мере один источник намагничивающей силы и по меньшей мере один катод-компенсатор, козырек выполнен таким, что между его нависающим краем и внешней поверхностью полого газораспределителя образована кольцеобразная щель. В козырьках могут быть выполнены распределительные отверстия рабочего тела, которые в азимутальном направлении чередуются с предшествующими ...

КАСКАДНЫЙ СВЕРХПРОВОДНИКОВЫЙ РАКЕТНЫЙ ДВИГАТЕЛЬ

Каскадный сверхпроводниковый ракетный двигатель относится к области особых способов и устройств для создания реактивной тяги, не отнесенных к другим подклассам (от использования продуктов сгорания F02K). Данная полезная модель служит для упрощения управления ракетным двигателем, повышения его экологических качеств, удельного импульса, тяги и безопасности. Каскадные сверхпроводниковые ракетные двигатели могут использоваться организациями, государственными органами для: космических кораблей и иных летательных аппаратов.Сущностью описываемого устройства являются, во-первых, возможность использования сверхпроводниковых индуктивных накопителей энергии (СПИНЭ) в качестве мощных и достаточно длительных источников питания ракетного двигателя, которые последовательно разряжаются, нагревая хладагент уже разряженных СПИНЭ, во-вторых, использование большинства сжиженных газов в качестве рабочих тел ракетного двигателя, в-третьих, возможности эргономичного сочетания хранения рабочих тел и использования ...

ДВИГАТЕЛЬ ДЛЯ КОСМИЧЕСКОГО АППАРАТА И КОСМИЧЕСКИЙ АППАРАТ, СОДЕРЖАЩИЙ ТАКОЙ ДВИГАТЕЛЬ

Двигатель (10) космического аппарата, содержащий химический маневровый двигатель, имеющий сопло (30) для испускания газа сгорания, вместе с маневровым реактивным двигателем на основе эффекта Холла. Двигатель сконфигурирован таким образом, что сопло служит в качестве канала испускания для частиц, выбрасываемых реактивный двигателем на основе эффекта Холла, когда он работает. Двигатель может обеспечить высокую тягу с низким удельным импульсом или относительно низкую тягу с большим удельным импульсом. 2 н. и 10 з.п. ф-лы, 2 ил.

Импульсный детонационный ракетный двигатель

Изобретение относится к ракетной технике и предназначено для создания импульсных ракетных двигателей систем ориентации космических аппаратов и старта с поверхности и посадки на планеты с малой гравитацией, например Луну. Импульсный детонационный ракетный двигатель, в котором система подачи и поджига выполнена в виде прозрачной диэлектрической трубки, заполненной инертным газом, на торцах которой установлены анод и катод, а рабочее тело выполнено в виде цилиндрического усеченного конуса из светопоглощающего материала, обращенного широким основанием в сторону к сверхзвуковому соплу. При этом диэлектрическая прозрачная трубка установлена по оси симметрии цилиндрического усеченного конуса. Изобретение позволяет облегчить инициирование разряда, увеличить скорость истечения рабочего тела и увеличить долю сжигаемого рабочего тела, что приводит к получению сверхзвуковых скоростей на выходе из сопла, а также к упрощению системы поджига и подачи рабочего тела. 2 з.п. ф-лы, 1 ил.

СИСТЕМЫ И СПОСОБЫ НАЗЕМНЫХ ИСПЫТАНИЙ РЕАКТИВНЫХ ДВИГАТЕЛЕЙ МАЛОЙ ТЯГИ

Импульсный плазменный электрический реактивный двигатель

Блок питания электроракетной двигательной установки спутника и система управления электроракетной двигательной установкой спутника

МАГНИТОРЕЗОНАНСНЫЙ ПЛАЗМЕННЫЙ ДВИГАТЕЛЬ

Предложен плазменный двигатель. Двигатель содержит соленоиды, расположенные во внешнем сердечнике-ферромагнетике, плазменный ускоритель и дуанты, катод-компенсатор, автономный источник низкотемпературной плазмы, корпус ускорителя, канал подачи рабочего тела в ионизатор, газовые трубки. Дополнительно содержит генератор переменного поля дуантов. В окне сброса пучка на дуантах зафиксирована юстированная площадка, в щелях которой располагается профилированный электростатический дефлектор. Соленоиды установлены в корпуса катушек постоянного электромагнита. При этом соленоиды намотаны на вкладыши с сохранением кольцевых каналов коридорного типа, предназначенных для протекания охлаждающего теплоносителя. Подводящие каналы теплоносителя расположены на крышках катушек, а отводящие каналы - на корпусе катушки. Автономный источник низкотемпературной плазмы дополнительно содержит электрод-коллектор и электрод-экстрактор. При реализации изобретения обеспечивается уменьшение тепловых потерь мощности ...

Hall thruster, cosmonautic vehicle, and propulsion method

In a Hall thruster 10, an acceleration channel 12 ionizes propellant flowing into an annular discharge space 11 to generate ions, and accelerates and discharges the generated ions. A distributor 37 supplies propellant from a plurality of holes 13 arranged azimuthally, via an anode 14 penetrating to the discharge space 11 of the acceleration channel 12, to the discharge space 11 of the acceleration channel 12, an amount of the propellant varying according to positions of the plurality of holes 13, thereby generating a plurality of regions, between adjacent ones of which the mass flow rate of the propellant is different, azimuthally in the discharge space 11 of the acceleration channel 12. During that time, the distributor 37 adjusts, with respect to the mass flow rate of the propellant in the discharge space 11 of the acceleration channel 12, a differential within a range of 5 to 15% between the mass flow rate of the propellant in a region with a large mass flow rate of the propellant and the mass flow rate of the propellant in a region with a small mass flow rate of the propellant. Thus, the width of a operation parameter region with reduced discharge current oscillation of the Hall thruster 10 is expanded.

SPACECRAFT THRUSTER

A thruster () has a main chamber () defined within a tube (). The tube has a longitudinal axis which defines an axis () of thrust; an injector () injects ionizable gas within the tube, at one end of the main chamber. An ionizer () is adapted to ionize the injected gas within the main chamber (). A first magnetic field generator () and an electromagnetic field generator () are adapted to generate a magnetized ponderomotive accelerating field downstream of said ionizer () along the direction of thrust on said axis (). The thruster () ionizes the gas, and subsequently accelerates both electrons and ions by the magnetized ponderomotive force. 1. A thruster comprising:a main chamber defining an axis of thrust;an injector adapted to inject ionizable gas within the main chamber;an ionizer adapted to ionize the injected gas within the main chamber;a first magnetic field generator and an electromagnetic field generator adapted to generate a magnetized ponderomotive accelerating field downstream of the ionizer along the direction of thrust on the axis; andat least one resonant cavity;wherein the electromagnetic field generator is adapted to control the mode of the resonant cavity.2. The thruster of wherein the electromagnetic field generator further comprises a housing adapted to generate stationary electromagnetic waves within the resonant cavity.3. The thruster of wherein the housing is adapted to contain at least partly the resonant cavity.4. The thruster of wherein the mode of the resonant cavity is controlled by selecting stationary waves of the resonant cavity to get electromagnetic energy maxima where desired.5. A thruster comprising:a main chamber defining an axis of thrust;an injector injecting ionizable gas within the main chamber;an ionizer ionizing the injected gas within the main chamber;at least one resonant cavity; anda first magnetic field generator and an electromagnetic field generator generating a magnetized ponderomotive accelerating field downstream of ...

ELECTRIC THRUSTER, A METHOD OF STOPPING AN ELECTRIC ENGINE INCLUDED IN SUCH A THRUSTER, AND A SATELLITE INCLUDING SUCH A THRUSTER

An electric thruster includes at least one electric engine, a feed system for the engine including a high-pressure tank of ionizable gas, a low-pressure buffer tank connected to the high-pressure tank by a valve, and a system of pipes for conveying the gas from the low-pressure buffer tank to an anode and to a cathode of the engine. The low-pressure buffer tank is in open connection with the engine. The thruster detects that a magnitude of the discharge current between the anode and the cathode is less than a threshold value and switches off the discharge voltage as a result of the detection. The thruster can be for use in a satellite, 17-. (canceled)8. An electric thruster comprising:at least one electric engine including an anode, a cathode, and a gas manifold; a high-pressure tank of ionizable gas;', 'a low-pressure buffer tank connected to the high-pressure tank by a means for dropping a pressure of the gas;', 'at least one valve configured to open, close, or regulate a flow rate of gas between the high-pressure tank and the low-pressure buffer tank; and', 'a system of pipes for conveying the gas from the low-pressure buffer tank to the engine; and, 'a feed system for the engine, the feed system comprisingpower electronics configured to deliver or not deliver electric power to the engine by applying or interrupting a discharge voltage between the anode and the cathode;wherein the low-pressure buffer tank is in open connection with the gas manifold; andfurther comprising means for detecting that a magnitude of the discharge current between the anode and the cathode is less than a threshold value and for interrupting the discharge voltage as a result of making the detection.9. An electric thruster according to claim 8 , wherein the threshold value is about 1 mA for an engine when a nominal value of the magnitude of the discharge current is about 1 A.10. An electric thruster according to claim 8 , further comprising a restrictor between the high-pressure tank and ...

EXTERNALLY POWERED VEHICLE PROPULSION SYSTEM

A vehicle propulsion system comprises a propellant source, a microwave energy source; an ionizer, a heater, and a propellant accelerator. The ionizer is configured for receiving propellant and for also receiving microwave energy from the microwave energy source so as to produce ionized propellant. The heater comprises a heater shell that defines a plasma heating cavity and is configured for receiving the ionized propellant from the ionizer. The heater shell is configured to transmit microwave energy received from the microwave energy source to the ionized propellant in the plasma heating cavity and to thereby facilitate absorption of microwave energy by the ionized propellant to produce heated ionized propellant. The propellant accelerator is configured for receiving the heated ionized propellant from the heater, accelerating the heated ionized propellant to produce accelerated propellant, and expelling the accelerated propellant in a desired direction to impose a reaction force (i.e., thrust) upon the vehicle. 1. A propulsion system for a vehicle , the propulsion system comprising:a propellant source;an ionizer defining an ionizing chamber, the ionizer configured for receiving propellant from the propellant source into the ionizing chamber and receiving microwave energy from a microwave energy source external to the vehicle so as to ionize the propellant to produce ionized propellant;a heater comprising a heater shell that defines a plasma heating cavity, the heater configured for receiving the ionized propellant from the ionizer into the plasma heating cavity, the heater shell configured to transmit microwave energy received from the microwave energy source to the ionized propellant in the plasma heating cavity and to thereby facilitate absorption of microwave energy by the ionized propellant in the plasma heating cavity to produce heated ionized propellant; anda propellant accelerator for receiving the heated ionized propellant from the heater, accelerating the ...

METALLIC WALL HALL THRUSTERS

A Hall thruster apparatus having walls constructed from a conductive material, such as graphite, and having magnetic shielding of the walls from the ionized plasma has been demonstrated to operate with nearly the same efficiency as a conventional non-magnetically shielded design using insulators as wall components. The new design is believed to provide the potential of higher power and uniform operation over the operating life of a thruster device. 1. A Hall thruster having a conductive wall , comprising:an annular discharge chamber having a conductive wall and having a rear surface with an aperture defined therein, said conductive wall of said annular discharge chamber having a selected one of a wall shape, a profile and a cross section fabricated in accordance with a respective calculated wall shape, calculated profile, and calculated cross section deduced to be substantially a respective one of a wall shape, a profile and a cross section that would be present in said Hall thruster at an end of life operating state;an anode/gas distributor having an anode electrical terminal, said anode/gas distributor situated in said aperture defined in said rear surface of said annular discharge chamber, said anode/gas distributor having at least one inlet configured to receive an ionizable gas and configured to distribute said ionizable gas for use as a propellant;a cathode neutralizer configured to provide electrons, said cathode neutralizer having a cathode electrical terminal that can be connected to said anode electrical terminal by way of a power supply and a switch, said cathode neutralizer and said anode/gas distributor when operating generating an axial electrical field within said annular discharge chamber; anda magnetic circuit having a magnetic yoke, an inner magnetic coil and an outer magnetic coil, said magnetic circuit configured to be switchably powered by a power supply, said magnetic circuit configured to provide a substantially radial magnetic field across ...

HALL-EFFECT THRUSTER

A Hall effect thruster includes at least one tank of gas under high pressure, a pressure regulator module, a gas flow rate control device, an ionization channel, a cathode placed in a vicinity of an outlet from the ionization channel, an anode associated with the ionization channel, an electrical power supply unit, an electric filter, coils for creating a magnetic field around the ionization channel, and an additional electrical power supply unit for applying a pulsating voltage between the anode and the cathode. 18-. (canceled)9. A Hall effect thruster comprising:at least one tank of gas under high pressure;a pressure regulator module;a gas flow rate control device;an ionization channel;at least one cathode placed in a vicinity of an outlet from the ionization channel;an anode associated with the ionization channel;an electrical power supply unit;an electric filter;coils for creating a magnetic field around the ionization channel;an additional electrical power supply unit for applying a pulsating voltage between the anode and the at least one cathode, and wherein the additional electrical power supply unit produces in alternation a first discharge voltage for a first duration in a range of 5 μs to 15 μs, and a second discharge voltage for a second duration in a range of 5 μs to 15 μs.10. A thruster according to claim 9 , wherein the additional electrical power supply unit produces in alternation the first discharge voltage in a range of 150 V to 250 V and second discharge voltage lying in a range of 300 V to 1200 V.11. A thruster according to claim 9 , wherein the first duration lies in a range of 5 μs to 10 μs claim 9 , and the second duration lies in a range of 5 μs to 10 μs.12. A thruster according to claim 9 , wherein the first discharge voltage lies in a range of 180 V to 220 V claim 9 , and the second discharge voltage lies in a range of 400 V to 1000 V.13. A thruster according to claim 9 , wherein the additional electrical power supply unit includes at least ...

HALL EFFECT THRUSTER

A Hall effect thruster including a downstream end of its annular channel presenting a cross-section that is variable to vary a thrust and a specific impulse of the thruster. 17-. (canceled)8. A Hall effect thruster comprising:an annular channel defined by an inner wall and an outer wall, which walls are coaxial about a central axis, the annular channel presenting a downstream end that is open and an upstream end that is closed, and the inner wall being movable in an axial direction and presenting a diameter that decreases in a downstream direction;an actuator for moving the movable inner wall in the axial direction;an electric circuit including an anode situated at the upstream end of the annular channel, a cathode at the downstream end of the annular channel, and an electric voltage source between the anode and cathode;an injection circuit for injecting a flow of propulsion gas into the annular channel;a magnetic circuit for generating a magnetic field at the downstream end of the annular channel; anda control unit that is connected at least to the electric circuit, to the propulsion gas injection circuit, and to the actuator, and that is configured to vary a flow and/or a voltage and to regulate a position of the movable inner wall as a function of the variable flow and/or of the variable voltage to adapt a cross-section of the downstream end of the annular channel to maintain a plasma density at the downstream end of the annular channel within a predetermined range.9. The Hall effect thruster according to claim 8 , wherein the actuator is a piezoelectric actuator.10. The Hall effect thruster according to claim 9 , wherein the piezoelectric actuator is an ultrasonic motor.11. The Hall effect thruster according to claim 8 , wherein the inner and outer walls are made of ceramic material.12. A Hall effect thruster comprising:an annular channel defined by an inner wall and an outer wall, which walls are coaxial about a central axis, the annular channel presenting a ...

MODULAR MICRO-CATHODE ARC THRUSTER

A modular micro-cathode arc thruster for use in satellites. An exemplary satellite has a plurality of stacked modular arc thrusters, each having an external anode, an internal cathode, and an insulator therebetween. The arc thrusters are situated in a housing, wherein the housing has an opening to eject exhausted thrusters. Once an arc thruster is expended, the push rod ejects that arc thruster and the next arc thruster takes its place. 1. A modular arc thruster satellite comprising:a plurality of arc thrusters, wherein each of the plurality of arc thrusters has an outer electrode, an inner electrode, and an insulator therebetween;a thruster housing having an open end leading to an exterior of the satellite, the thruster housing receiving said plurality of arc thrusters;a push rod; anda motor for selectively operating the push rod to eject one of the plurality of arc thrusters from the satellite via the open end of said thruster housing once that one of the plurality of arc thrusters has been expended.2. The satellite of claim 1 , wherein the each thruster has a triggering system lasting for over 10pulses.3. The satellite of claim 1 , further comprising a spring-loaded door at the open end of the thruster housing.4. The satellite of claim 1 , wherein the inner electrode is comprised of titanium.5. The satellite of claim 1 , wherein the insulator is comprised of ceramic.6. The satellite of claim 1 , wherein said plurality of arc thrusters are cylindrical and stacked with respect to one another within the thruster housing.7. The satellite of claim 1 , wherein the outer electrode forms an inward lip that retains the inner electrode and the insulator.8. The satellite of claim 1 , wherein each of said plurality of thrusters is a discrete device.9. The satellite of claim 1 , wherein each of said plurality of thrusters is modular.10. The satellite of claim 1 , wherein said plurality of arc thrusters are stacked with respect to each other within said thruster housing.11. ...

GENERATING ELECTROSPRAY FROM A FERROFLUID

An electrospray device for generating electrospray from a ferrofluid. The electrospray device includes an emitter, an extraction electrode, and a magnet. The emitter is configured to receive a ferrofluidic liquid. The extraction electrode includes an aperture and is positioned a first distance from the emitter. The magnet generates a magnetic field in a first direction toward the emitter. The magnetic field causes Rosensweig instability in the ferrofluidic liquid, and generates a ferrofluidic peak in the ferrofluidic liquid. The magnet is positioned a second distance from the emitter, and the emitter is positioned between the extraction electrode and the magnet. The ferrofluidic liquid is biased at a first electrical potential and the extraction electrode is biased at a second electrical potential. A difference between the first electrical potential and the second electrical potential is sufficient to generate an electric field at the ferrofluidic peak that generates electrospray from the ferrofluidic peak. 1. An electrospray device comprising:an emitter configured to receive a ferrofluidic liquid;an extraction electrode positioned a first distance from the emitter; anda magnet operable to generate a magnetic field in a first direction toward the emitter, the magnetic field sufficient to cause Rosensweig instability in the ferrofluidic liquid, the Rosensweig instability generating a ferrofluidic peak in the ferrofluidic liquid, the magnet positioned a second distance from the emitter, the emitter positioned between the extraction electrode and the magnet,wherein the ferrofluidic liquid is biased at a first electrical potential and the extraction electrode is biased at a second electrical potential, andwherein a difference between the first electrical potential and the second electrical potential is sufficient to generate an electric field at the ferrofluidic peak that generates electrospray from the ferrofluidic peak.2. The electrospray device of claim 1 , wherein ...

ARCJET PROPULSION SYSTEMS FOR SPACECRAFT

An arcjet thruster system for a spacecraft is provided. The arcjet thruster system may include a power supply that includes a radio-frequency start power supply and a continuous direct-current power supply, each selectively coupled to electrodes of an arcjet for initiation and maintenance of an arc between the electrodes. A radio-frequency/direct-current control module may be provided for selectively coupling the radio-frequency start power supply and a continuous direct-current power supply. The radio-frequency start power supply may be used to initiate an arc that is then sustained by the continuous direct-current power supply. 1. An arcjet thruster for a spacecraft , the arcjet thruster comprising: an anode,', 'a cathode, and', 'a propellant valve configured to direct a propellant between the cathode and the anode; and, 'an arcjet having a radio-frequency start power supply,', 'a direct-current continuous power supply, and', 'a radio-frequency/direct-current control module coupled to the arcjet, the radio-frequency start power supply, and the direct-current continuous power supply., 'a power module comprising2. The arcjet thruster of claim 1 , wherein the radio-frequency/direct-current control module is configured to provide a radio-frequency signal from the radio-frequency start power supply to the arcjet via a coaxial feedline coupled to the radio-frequency/direct-current control module claim 1 , the cathode claim 1 , and the anode of the arcjet to initiate an electrical discharge arc between the anode and the cathode.3. The arcjet thruster of claim 2 , wherein the radio-frequency/direct-current control module is further configured to provide a direct-current voltage difference from the direct-current continuous power supply claim 2 , via the coaxial feedline claim 2 , to the cathode and the anode of the arcjet to sustain the electrical discharge arc.4. The arcjet thruster of claim 3 , wherein the radio-frequency/direct-current control module is configured to ...

IGNITION PROCESS FOR NARROW CHANNEL HALL THRUSTER

Disclosed is a closed drift, narrow channel Hall thruster configured to operate at powers <30 W. The thruster includes a thruster body and a neutralizing cathode. The thruster body includes a magnetic circuit including a magnetic source and two magnetic poles, a metallic, annular thruster channel formed by the magnetic poles with a downstream channel width smaller than about 3 mm and an upstream channel width greater than the downstream channel width, an anode positioned at the channel's entry, and a gas distributor configured to release a propellant gas into the thruster channel. The magnetic circuit is configured to generate a magnetic field in the thruster channel for trapping electrons therein. The channel walls (the magnetic poles) are under bias potential. The anode and the cathode are configured to generate a substantially axial electric field in the thruster channel. In operation, propellant gas atoms ionized by trapped electrons in the thruster channel, accelerate axially, exiting via the channel's exit. 137-. (canceled)39. The narrow channel Hall thruster of claim 38 , wherein the anode has at least one of:a width which is greater than the second channel width, ora mounting position perpendicular to the axial direction of the annular thruster channel, such that a surface of the anode which faces the first extremity of the channel, is substantially flat.40. The narrow channel Hall thruster of wherein the generated magnetic field is substantially radial at or near the second extremity of the annular thruster channel.41. The narrow channel Hall thruster of claim 38 , wherein the magnetic circuit is further configured such that the generated magnetic field increases in strength from the first extremity to the second extremity of the annular thruster channel.42. The narrow channel Hall thruster of claim 38 , wherein the thruster body further comprises a discharge chamber claim 38 , which includes the thruster channel and which is fluidly coupled to the outside ...

PLASMA ACCELERATING APPARATUS AND PLASMA ACCELERATING METHOD

A plasma accelerating apparatus includes: a cathode () configured to supply electrons to a plasma acceleration region; a anode (); a power supply () configured to apply a voltage between the cathode and the anode; a supply port () arranged on an outer circumference side than the cathode to supply a propellant to the plasma acceleration region; and a first magnetic field generator () configured to generate a first axial direction magnetic field in the upstream side region of the plasma acceleration region to suppress that the electrons supplied from the cathode head for the anode. Thus, the plasma accelerating apparatus and the plasma accelerating method having high thrust efficiency can be provided. 112-. (canceled)13. A plasma accelerating apparatus comprising:a cathode configured to emit electrons to a direction of a predetermined center axis to supply the electrons to an upstream side region of a plasma acceleration region;an anode having a ring shape when viewing from the direction of the center axis and arranged around the center axis;a power supply configured to apply a voltage between the cathode and the anode;a supply port arranged on an outer circumference side than the cathode to supply a propellant before plasmatization or a propellant after plasmatization to the plasma acceleration region; anda first magnetic field generator arranged in a second direction from the plasma acceleration region when a motion direction of the electrons emitted from the cathode is defined as a first direction and a direction opposite to the first direction is defined as the second direction, and configured to generate a first axial direction magnetic field in the upstream side region of the plasma acceleration region to suppress that the electrons supplied from the cathode head for the anode,wherein the first axial direction magnetic field has an axial direction component which is a component parallel to the center axis and monotonously degreases as heading for the first ...

Plasma Propulsion System Feedback Control

Systems and methods can support a plasma propulsion system. The system may include a thrust head comprising a plasma generator and a thrust generator. A propellant handling assembly may be directly coupled to the thrust head. The propellant handling assembly may comprise a manifold and a plurality of valves. A propellant storage vessel may be directly coupled to the propellant handling assembly. A propulsion control module may be operable to receive inputs associated with the plasma propulsion system, generate control outputs associated with the plasma propulsion system, establish and train models relating the inputs and the control outputs, apply the inputs to the models to update the output parameters, and apply the output parameters to control the plasma propulsion system. 1. A plasma propulsion system , comprising:a thrust head comprising a plasma generator and a thrust generator;a propellant handling assembly directly coupled to the thrust head, wherein the propellant handling assembly comprises a manifold and a plurality of valves;a propellant storage vessel directly coupled to the propellant handling assembly; anda propulsion control module operable to:receive inputs associated with the plasma propulsion system,generate control outputs associated with the plasma propulsion system,establish and train one or more models relating the inputs and the control outputs,apply received inputs to the one or more models to update the generated output parameters, andapply the generated output parameters to control the plasma propulsion system.2. The plasma propulsion system of claim 1 , wherein the one or more models seek to regulate performance metrics associated with the plasma propulsion system.3. The plasma propulsion system of claim 1 , wherein the plurality of valves comprise piezo micro-valves.4. The plasma propulsion system of claim 1 , wherein the manifold comprises a test port to access to a flow path between the plurality of valves.5. The plasma propulsion ...

PLASMA THRUSTER AND METHOD FOR GENERATING A PLASMA PROPULSION THRUST

The invention, which relates to a miniaturisable plasma thruster, consists of: —igniting the plasma by microhollow cathode discharge close to the outlet and inside the means for injecting the propellant gas, said injection means being magnetic and comprising a tip at the downstream end thereof; —bringing the electrons of the magnetised plasma into gyromagnetic rotation, at the outlet end of said injection means; —sustaining the plasma by means of Electron Cyclotron Resonance (ECR), said injection means being metal and being used as an antenna for electromagnetic (EM) emission, the volume of ECR plasma at the outlet of said injection means being used as a resonant cavity of the EM wave; —accelerating the plasma in a magnetic nozzle by diamagnetic force, the ejected plasma being electrically neutral. 1. A plasma thruster comprising: a discharge chamber comprising an internal cavity and an outlet opening; at least one injection means comprising an injection nozzle capable of injecting into the discharge chamber a propellant gas along a predefined axis , said injection nozzle having an outlet end; a magnetic field generator capable of setting electrons of the propellant gas present in the discharge chamber in gyromagnetic rotation; and an electromagnetic wave generator capable of irradiating the propellant gas present in the discharge chamber by generating at least one electromagnetic wave the electric field of which has a right-hand circular polarization and a frequency equal to the frequency , f , of gyromagnetic resonance of the electrons of the propellant gas magnetized by said magnetic field generator , [ a first local maximum of intensity inside the injection nozzle and at the outlet end of the injection nozzle;', 'field lines which determine an iso-field surface, known as the “ECR surface”, with an intensity equal to that allowing a cyclotron resonance of the electrons under the effect of said electromagnetic wave, said ECR surface enveloping the outlet end of ...

FIELD EMISSION NEUTRALIZER

A field emission neutralizer is provided. The field emission neutralizer includes a bottom plate and a field emission cathode unit located on the bottom plate. The field emission cathode unit includes a substrate, a shell located on the substrate, a cathode emitter located inside the shell, a mesh grid insulated from the cathode emitter, and a shielding layer insulated from the mesh grid. The cathode emitter includes a cathode substrate and a graphitized carbon nanotube array. The graphitized carbon nanotube array is in electrical contact with the cathode substrate. The graphitized carbon nanotube array is fixed on a surface of the substrate body, and the carbon nanotubes of the graphitized carbon nanotube array are substantially perpendicular to the cathode substrate. 1. A field emission neutralizer comprising:a bottom plate; andat least one field emission cathode unit located on the bottom plate, each of the at least one field emission cathode unit comprising:a substrate;a shell located on the substrate and comprising an opening;a cathode emitter located in the shell, and comprising a cathode substrate and a graphitized carbon nanotube array, wherein the graphitized carbon nanotube array comprises a plurality of carbon nanotubes, the graphitized carbon nanotube array is electrically connected with the cathode substrate;a mesh grid comprising a plurality of gate holes and being insulated from the cathode emitter; anda shielding layer comprising a through-hole and being electrically insulated from the mesh grid,wherein the opening, the plurality of gate holes, and the through-hole communicate with each other, the graphitized carbon nanotube array is fixed on a surface of the substrate, and the plurality of carbon nanotubes of the graphitized carbon nanotube array are substantially perpendicular to the cathode substrate.2. The field emission neutralizer of claim 1 , wherein the cathode substrate comprises a substrate body and an adhesive layer claim 1 , and the ...

HALL EFFECT THRUSTER AND A SPACE VEHICLE INCLUDING SUCH A THRUSTER

A Hall effect thruster arranged inside a wall and including a magnetic circuit and an electric circuit including an anode, a first cathode, and a voltage source. The magnetic circuit and the electric circuit are arranged in such a manner as to generate magnetic and electric fields around the wall. In every meridian section, the magnetic circuit presents an upstream magnetic pole and a downstream magnetic pole arranged at the surface of the wall and spaced apart from each other; and the anode and the first cathode are situated on either side of the upstream magnetic pole. 1. A Hall effect thruster for developing thrust along a thrust axis and comprising:a magnetic circuit for generating a magnetic field; andan electric circuit comprising an anode, a first cathode, and a voltage source for emitting electrons via at least the first cathode and attracting electrons via the anode; the thruster is arranged inside a wall formed around the thrust axis;', 'the magnetic circuit and the electric circuit are arranged so as to generate magnetic and electric fields around the wall; and', the magnetic circuit presents an upstream magnetic pole and a downstream magnetic pole arranged substantially at the surface of the wall and spaced apart from each other; and', 'the anode and the first cathode are situated on either side of the upstream magnetic pole., 'in all sections parallel to the thrust axis and perpendicular to the wall], 'wherein2. The thruster according to claim 1 , wherein the magnetic circuit is arranged in such a manner that the magnetic field is oriented in a direction that is generally perpendicular to the surface of the wall claim 1 , next to the upstream magnetic pole.3. The thruster according to claim 1 , wherein the electric circuit also includes a second cathode arranged axially between the upstream magnetic pole and the downstream magnetic pole.4. The thruster according to claim 1 , wherein the electric circuit also includes an additional cathode arranged ...

IODINE PROPELLANT RF ION THRUSTER WITH RF CATHODE

A thrust producing system includes an RF ion thruster with a discharge chamber having a gas inlet and an outlet, and a coil about the discharge chamber. The system further includes an RF cathode proximate the discharge chamber outlet of the RF ion thruster for ion beam neutralization. The RF cathode includes a discharge chamber having a gas inlet and an outlet and a coil about the discharge chamber. A tank for containing iodine in solid form and a heater associated with said tank to produce iodine vapor. A feed subsystem fluidly couples the tank with the RF ion thruster discharge chamber gas inlet and with the RF cathode discharge chamber gas inlet. 1. An RF ion thruster system comprising: a discharge chamber having a gas inlet and an outlet, and', 'a coil about the discharge chamber;, 'an RF ion thruster including a discharge chamber having a gas inlet and an outlet, and', 'a coil about the discharge chamber;, 'an RF cathode proximate the discharge chamber outlet of the RF ion thruster and includinga tank for containing iodine in solid form;a heater associated with said tank to produce iodine vapor; anda feed subsystem fluidly coupling said tank with the RF ion thruster discharge chamber gas inlet and with the RF cathode discharge chamber gas inlet.2. The system of further including an igniter associated with the RF cathode discharge chamber inlet and an optional igniter associated with the RF ion thruster discharge chamber inlet.3. The system of in which the RF cathode discharge chamber inlet includes a conduit and the igniter includes spaced conductive electrodes in the conduit coupled to a voltage source and a voltage bias source.4. The system of in which the RF ion thruster includes a conductive grid subsystem proximate the RF ion thruster discharge chamber outlet.5. The system of in which the conductive grid subsystem of the RF ion thruster includes at least two conductive plates with orifices therethrough claim 4 , one said plate supplied with a positive ...

Dipole Drive for Space Propulsion

The dipole drive is a new propulsion system which uses ambient space plasma as propellant, thereby avoiding the need to carry any of its own. The dipole drive is constructed from two parallel screens, one charged positive, the other negative, creating an electric field between them with no significant field outside. Ambient solar wind protons entering the dipole drive field from the negative screen side are reflected out, with the angle of incidence equaling the angle of reflection, thereby providing lift if the screen is placed at an angle to the plasma wind. Protons entering from the positive side are accelerated out the negative screen, producing thrust. The dipole drive can achieve more than 3 mN/kWe in interplanetary space and better than 10 mN/kWe in Earth, Venus, Mars, or Jupiter orbit and offers potential as a means of achieving ultra-high velocities necessary for interstellar flight. 1. A system for space propulsion without a requirement for onboard propellant comprising:Two parallel charged screens, including a positive charged screen and a negative charged screen; and,A power supply that is used to maintain their charges.2. A system of the type described in which is used to generate drag against a plasma wind by positioning its screens perpendicular to the wind claim 1 , with its negative screen in the forward position towards the incoming wind and its positive screen behind the negative screen as seen by the wind.3. A system of the type described in which is used to generate lift against a plasma wind by positioning its screens at an angle to the wind claim 1 , with its negative screen in the forward position towards the incoming wind and its positive screen behind the negative screen as seen by the wind.4. A system of the type described in which is used to generate thrust towards an incoming plasma wind by positioning its screens perpendicular to the wind claim 1 , with its positive screen in the forward position towards the incoming wind and its ...

FIBER-FED ADVANCED PULSED PLASMA THRUSTER (FPPT)

A Fiber-fed Pulsed Plasma Thruster (FPPT) will enable enhanced low Earth orbit, cis-lunar, and deep space missions for small satellites. FPPT technology utilizes an electric motor to feed PTFE fiber to its discharge region, enabling high PPT propellant throughput and variable exposed fuel area. An innovative, parallel ceramic capacitor bank dramatically lowers system specific mass. FPPT minimizes range safety concerns by the use of non-pressurized, non-toxic, inert propellant and construction materials. Estimates are that a 1 U (10 cm×10 cm×10 cm, or 1 liter) volume FPPT thruster package may provide more than 10,000 N-s total impulse and a delta-V of 1.4 km/s delta-V for an 8 kg CubeSat. 1. A pulsed plasma thruster comprising:a spool having a fiber propellant wound thereon;a stepper motor in communication with the fiber propellant to pull the fiber propellant from the spool;an insulated tube configured to have one end in communication with the stepper motor such that the fiber propellant is fed into the insulated tube;an anode bored through and having one end in communication with the insulated tube, such that the fiber propellant travels through the anode, the anode having an exit end defined with a flange extending radially Inward configured to create a stop between the edge of the flange and the exit end of the anode, wherein the fiber propellant fed through the anode stops at the stop;a coaxial insulator positioned about the exit end of the anode;a cathode connected to the insulator, the cathode having an interior profile shaped into a nozzle region; andan igniter fitted through an opening in the cathode, wherein when the igniter is pulsed the igniter is configured to expel electrons toward the anode to ignite a primary high current, high magnetic field discharge between the anode and cathode thereby creating a plasma that vaporizes the fiber propellant at the stop, and wherein the vaporizing fiber propellant combines with the high current discharge to create a ...

PLASMA DRIVE

A plasma drive includes a plurality of plasma thrusters arrayed in each of at least one array of plasma thrusters. Plasma thrust may be generated sequentially or in a pulse from each array. Circuitry is adapted to selectively fire each thruster in each array according to a digitally controlled progression. The controlled firing progression collectively provides a cumulative thrust vector for each array. In a turbine drive embodiment the controlled progression causes sequential firing of the thrusters in each array, and the arrays in sequence. The controlled progression allows for directional control of the combined cumulative thrust vectors. 1. A plasma drive comprising:a plurality of plasma thrusters arrayed in each of at least one array of said plasma thrusters, and wherein, when said at least one array of said plasma thrusters includes a plurality of said arrays of said plasma thrusters, said arrays are adapted to provide for sequentially providing plasma thrust sequentially or in a pulse from each said array in said plurality of said arrays so that a plasma thrust associated with said each plasma thruster, when energized, in each said array has a cumulative thrust vector in a desired thrust direction,circuitry operatively associated with said array, wherein said circuitry is adapted to selectively energized and de-energize said each thruster in said array according to a controlled progression,a digital processor controlling said controlled progression,wherein said controlled progression causes energizing and de-energizing of said each plasma thruster in said each array so as to collectively provide said cumulative thrust vector for said each array.and wherein said controlled progression causes energizing and de-energizing of said plurality of said arrays.2. The plasma drive of wherein said each array is substantially planar.3. The plasma drive of where said each array is substantially parallel to a next adjacent said array in said plurality of said arrays.4. The ...

Magnetic Gas Engine and Method of Extracting Work

The present subject matter overcomes the deficiencies in the prior art by introducing or generating charged particles in an air stream and manipulating the air stream with magnetic fields operating on the charged particles. Embodiments of the present subject mater compress the air stream by accelerating charged particles with a moving magnetic field, where the magnetic field has a velocity perpendicular to its flux lines. The increased velocity of the charged particles increases the statistical mean particle velocity and thereby increases the pressure in the air stream. The compressed air stream is then heated and expanded through a second magnetic field. The expansion of the air stream substantially increases the velocity of the air stream and the charged particles therein. The interaction of the high velocity charged particles and the magnetic field imparts a force perpendicular to the flux lines, this force powers the movement of the magnetic field. 1. A method for providing thrust to a craft across subsonic to supersonic regimes receiving an air stream through an inlet, said air stream having a velocity relative to the craft prior to being received in the inlet;', 'compressing the air stream to a static pressure at or below the stagnation pressure associated with the air stream by decelerating the air stream; and', 'compressing the air stream above the stagnation pressure by adding work to the stream via a magnetic or electric field when the relative air stream velocity is subsonic or less than a predetermined Mach number greater than one;', 'heating the air stream and expanding the heated air stream through a nozzle to provide thrust., 'comprising2. The method of claim 1 , wherein the step of compressing the air stream above the stagnation pressure comprises ionizing the air stream to a net charge.3. The method of claim 2 , wherein the step of compressing the air stream above the stagnation pressure comprises the steps of: providing a gas stream having a net ...

FIBER-FED ADVANCED PULSED PLASMA THRUSTER (FPPT)

A Fiber-fed Pulsed Plasma Thruster (FPPT) utilizes a motor to feed PTFE fiber to its discharge region, enabling high PPT propellant throughput and variable exposed fuel area. A highly parallel ceramic capacitor bank lowers system specific mass. Impulse bits (I-bits) from 0.057-0.241 mN-s have been measured on a thrust stand with a specific impulse (Isp) of 900-2400 s, representing an enhancement from state-of-the-art PPT technology. A 1U (10 cm×10 cm×10 cm, or 1 liter) volume FPPT thruster package will provide 2900-7700 N-s total impulse, enabling 0.6-1.6 km/s delta-V for a 5 kg CubeSat. A 1U design variation with 590 g propellant enables as much as ˜10,000 N-s and a delta-V of 2 km/s for a 5 kg CubeSat. Increasing the form factor to 2U increases propellant mass to 1.4 kg and delta-V to 10.7 km/s for an 8 kg CubeSat. 1. A pulsed plasma thruster comprising:a spool having a fiber propellant wound thereon;a stepper motor in communication with the fiber propellant to pull the fiber propellant from the spool;an insulated tube configured to have one end in communication with the stepper motor such that the fiber propellant is fed into the insulated tube;an anode bored through and having one end in communication with the insulated tube, such that the fiber propellant travels through the anode, the anode having an exit end, wherein the fiber propellant fed through the anode exits at the exit end;a power processing unit electrically connected in parallel to a capacitor bank, the capacitor bank having a positive electrical connection to the anode and the capacitor bank having a negative electrical connection to a cathode, and wherein the capacitor bank is configured to lower an equivalent series resistance raising a pulse current and raising a {right arrow over (j)}×{right arrow over (B)} thrust generated by the pulsed plasma thruster;a coaxial insulator positioned about the exit end of the anode;the cathode further positioned about the insulator and having an interior ...

EXTERNALLY POWERED HYBRID PROPULSION SYSTEM

A hybrid propulsion system for a vehicle comprises a propellant tank for providing a supply of propellant, a propellant heater, and an exhaust nozzle. The propellant tank is in fluid communication with the propellant heater and is configured for providing a flow of propellant to the propellant heater. The propellant heater is in fluid communication with the propellant tank and the nozzle and initially comprises a supply of oxidizer that is configured for reacting chemically with the propellant to produce heat. The propellant heater is further configured for receiving a beam of microwave energy and facilitating transmission of the beam of microwave energy to the propellant. The nozzle is in fluid communication with the propellant heater and is configured for receiving a flow of propellant from the propellant heater, for accelerating the propellant, and for expelling the propellant so as to produce thrust. 1. A hybrid propulsion system for a vehicle , the hybrid propulsion system comprising:a propellant tank for providing a flow of propellant;a propellant heater; andan exhaust nozzle;the propellant tank being in fluid communication with the propellant heater and configured for providing a flow of propellant to the propellant heater;the propellant heater being in fluid communication with the propellant tank and the exhaust nozzle and initially comprising a quantity of oxidizer, the quantity of oxidizer being configured for reacting chemically with the flow of propellant to produce heat;the propellant heater being further configured for receiving a beam of microwave energy and facilitating transmission of the beam of microwave energy to the flow of propellant; andthe exhaust nozzle being in fluid communication with the propellant heater and being for receiving a flow of propellant from the propellant heater, for accelerating the flow of propellant, and for expelling the flow of propellant so as to produce thrust.2. The hybrid propulsion system of claim 1 , wherein the ...

SELF CONTAINED ION POWERED AIRCRAFT

A self-contained ion powered aircraft assembly is provided. The aircraft assembly includes a collector assembly, an emitter assembly, and a control circuit operatively connected to at least the emitter and collector assemblies and comprising a power supply configured to provide voltage to the emitter and collector assemblies. The assembly is configured, such that, when the voltage is provided from an on board power supply, the aircraft provides sufficient thrust to lift each of the collector assembly, the emitter assembly, and the entire power supply against gravity. 1. A self-contained ion powered aircraft assembly comprising:a collector assembly;an emitter assembly; anda control circuit operatively connected to at least the emitter and collector assemblies and comprising a power supply configured to provide voltage to the emitter and collector assemblies, such that, when the voltage is provided, the self contained ion powered aircraft provides sufficient thrust to lift each of the collector assembly, the emitter assembly, and the control circuit against gravity.2. The self-contained ion powered aircraft assembly of claim 1 , wherein the collector assembly comprises a plurality of substantially concentric elements claim 1 , with a central support of the device located at a common centroid of the plurality of concentric elements.3. The self-contained ion powered aircraft assembly of claim 2 , wherein each of the plurality of concentric elements are substantially hexagonal.4. The self-contained ion powered aircraft assembly of claim 2 , wherein the control circuit is implemented on or within the central support.5. The self-contained ion powered aircraft assembly of claim 4 , wherein the central support is formed from a flexible printed circuit board rolled into a tube.6. The self-contained ion powered aircraft assembly of claim 2 , further comprising a plurality of peripheral supports claim 2 , each of the plurality of peripheral supports extending perpendicularly to ...

Propulsion Method Based on Liquid Carbon Dioxide Phase Change and Propulsion Device Thereof

The present disclosure discloses a propulsion method based on liquid carbon dioxide phase change and a propulsion device. The method includes the following steps of: accommodating carbon dioxide in a thermally insulated container in a liquid phase form; transiently heating to convert the carbon dioxide from a liquid phase to a gas phase; and jetting carbon dioxide gas after the phase change in a predetermined direction by a predetermined jet-out amount so as to obtain a propulsion force. 1. A propulsion method based on liquid carbon dioxide phase change , comprising the following steps:{'b': '1', 'in a first step (S), accommodating carbon dioxide in a thermally insulated container in a liquid phase form;'}{'b': '2', 'in a second step (S), transiently heating to convert the carbon dioxide from a liquid phase to a gas phase; and'}{'b': '3', 'in a third step (S), jetting carbon dioxide gas after the phase change in a predetermined direction by a predetermined jet-out amount so as to obtain a propulsion force.'}21. The method of claim 1 , wherein in the first step (S) claim 1 , the carbon dioxide is accommodated in the thermally insulated container in the liquid phase form at a predetermined temperature claim 1 , and the predetermined temperature is room temperature lower than a liquid-gas phase change temperature of carbon dioxide.31. The method of claim 2 , wherein in the first step (S) claim 2 , the predetermined temperature is 10° C.42. The method of claim 1 , wherein in the second step (S) claim 1 , time consumption of transient heating is in a millisecond grade.52. The method of claim 1 , wherein in the second step (S) claim 1 , transient heating is implemented via heat transfer claim 1 , heat exchange or other energy conversions.62. The method of claim 1 , wherein in the second step (S) claim 1 , a temperature rise of transient heating does not exceed 21° C.73. The method of claim 1 , wherein in the third step (S) claim 1 , the carbon dioxide gas after the phase ...

ION THRUSTER WITH EXTERNAL PLASMA DISCHARGE

An ion thruster is provided allowing a plasma discharge to be generated and confined in an external confinement space created by an external magnetic field B. 2. The ion thruster as claimed in claim 1 , wherein the anode makes contact with the diffuser.3. The ion thruster as claimed in claim 1 , wherein the diffuser is a porous or cellular diffuser or a chicaned mechanical diffuser.4. The ion thruster as claimed in claim 1 , wherein the anode has at least one orifice opening into the annular space.5. The ion thruster as claimed in claim 1 , wherein the anode is made of a material chosen from metals such as aluminum claim 1 , tantalum claim 1 , steels or carbon and mixtures thereof.6. The ion thruster as claimed in claim 1 , wherein the anode and the diffuser are one and the same element.7. The ion thruster as claimed in claim 1 , wherein the cathode is:placed in the central arm of the capital E, oris annular and is placed in the cross section S in the upper or lower arms E of the body.8. The ion thruster as claimed in claim 1 , wherein a protection of the magnetic poles covers the serif of the central arm of the capital E and an annular protection of the magnetic poles covers the serif of the lower and upper arms of the capital E.9. The use of an ion thruster such as defined in claim 1 , to propel a space vehicle such as a probe claim 1 , a satellite claim 1 , a space capsule claim 1 , a space shuttle or a space station.10. A method for generating propulsive thrust for a space vehicle by means of an ion thruster such as defined in comprising the following steps:a) introducing a gas via a gas line into an internal annular space of the ion thruster,b) generating an external electric field {right arrow over (E)} via an anode and a cathode and generating an external magnetic field {right arrow over (B)} via an element and an annular element that generate a magnetic field thus creating a confinement space external to the ion thruster, the external confinement space and ...

METHOD AND SYSTEM FOR A PROGRAMMABLE AND FAULT TOLERANT PULSED PLASMA THRUSTER

A system and method provides a fault-tolerant multi-channel pulsed plasma thruster system utilizing a control unit and an embedded real time application manipulating low-level timing events with programming, with clear examples of completely flexible control techniques of a scalable micropropulsion system having many pulsed plasma thruster channels, taking into account system aging behavior and specific mission utilization requirements that may change in the mission lifetime. The system and method also covers an architecture lending itself suitable for design of a dedicated FGPA or ASIC that would tightly integrate many channels of thruster components to build a robust, resilient and versatile micropropulsion subsystem for space applications, and indirectly for advanced multi-channel spacecraft instrumentation. 1. A method for controlling trigger pulse generation in a pulsed plasma thruster system , the method comprising: generating at a processing device , independent event markers in time-units , and controlling by the event markers a Trigger Pulse activation event , Trigger Pulse deactivation event , Magnetic Coil activation event , Magnetic Coil deactivation event , End of Cycle signal event , and a spacecraft related event.2. The method of claim 1 , further comprising generating time-slices at regular intervals claim 1 , and generating the event markers for each event at a predetermined number of occurrences of the time-slices.3. The method of claim 2 , wherein each of the events is triggered at a different predetermined number of occurrences of the time-slices.4. The method of claim 1 , further comprising:generating at the processing device, a time-slice count;assigning at the processing device, a unique count value to each of the events;triggering at the processing device, an event when the time-slice count equals the unique count value for that event; and,incrementing at the processing device, the time-slice count.5. The method of claim 1 , wherein operation ...

IONIC PROPULSION SYSTEM

An ionic propulsion system for an aircraft having an airfoil includes a first conductor and a second conductor, the first conductor and the second conductor being disposed at least partially within the airfoil when not in use. The propulsion system includes an actuator for extending the first conductor and the second conductor from an end of the airfoil such that the first conductor and the second conductor are in the airstream of the aircraft, the first conductor being upstream of the second conductor in the airstream. The propulsion system includes a power supply for supplying current to the first conductor and the second conductor to ionize the air particles in the vicinity of the first conductor and the end of the airfoil to create a flow of the ionized particles from the first conductor toward the second conductor.

SYSTEM AND METHOD FOR SMALL, CLEAN, STEADY-STATE FUSION REACTORS

According to some embodiments, a system for widening and densifying a scrape-off layer (SOL) in a field reversed configuration (FRC) fusion reactor is disclosed. The system includes a gas box at one end of the reactor including a gas inlet system and walls of suitable heat bearing materials. The system further includes an exit orifice adjoining the gas box, wherein the exit orifice has a controllable radius and length to allow plasma to flow out from the gas box to populate the SOL with the plasma. The system may also include fusion products, which decrease in speed in the plasma in the SOL, allowing energy to be extracted and converted into thrust or electrical power and further allowing ash to be extracted to reduce neutron emissions and maintain high, steady-state fusion power. 1. A system for widening and densifying a scrape-off layer (SOL) in a field reversed configuration (FRC) fusion reactor , comprising:a gas box at one end of the reactor comprising a gas inlet system and walls of suitable heat bearing materials; andan exit orifice adjoining the gas box, wherein the exit orifice has a controllable radius and length to allow plasma to flow out from the gas box to populate the SOL with the plasma.2. The system of claim 1 , wherein the plasma in the gas box has a peak electron temperature in the range 1-50 eV and a peak density in the range 3×10cmto 3×10cm.3. The system of claim 1 , wherein the plasma in the SOL has a peak temperature in the range 10-200 eV and a peak density in the range 5×10cmto 3×10cm.4. The system of claim 1 , wherein the FRC reactor burns D-He claim 1 , D-D claim 1 , or a combination of D-He and D-D.5. The system of claim 1 , wherein the FRC reactor further comprises a closed field region containing core plasma.6. The system of claim 5 , wherein the SOL is contained in an open field region surrounding the closed field region.7. The system of claim 6 , wherein the FRC reactor further comprises a separatrix between the open field region and ...

ION THRUSTER AND METHOD FOR PROVIDING THRUST

An ion thruster () and a method for providing trust is disclosed. The ion thruster comprises a sputtering magnetron (), a target () arranged at the sputtering magnetron, and a second electrode (). During a first pulse, the target is at a negative potential (U) with respect to a second electrode and a plasma is sustained whereby atoms are sputtered from the target and at least a portion thereof become ionised by the plasma. During a second pulse, a reversed potential (U) is applied between the target and the second electrode. This increases the potential of a volume of the plasma adjacent to the target, which in turn accelerates ions in a direction away from the target. Thereby, thrust is provided.

MICRO-CATHODE THRUSTER AND A METHOD OF INCREASING THRUST OUTPUT FOR A MICRO-CATHODE THRUSTER

A magnetically enhanced micro-cathode thruster assembly provides long-lasting thrust. The micro-cathode thruster assembly includes a tubular housing, a tubular cathode, an insulator, an anode and a magnetic field. The tubular housing includes an open distal end. The tubular cathode is housed within the housing and includes a distal end positioned proximate the open distal end of the housing. The insulator is in contact with the cathode forming an external cathode-insulator interface. The anode is housed within the housing, proximate the open distal end of the housing. The magnetic field is positioned at or about the external cathode-insulator interface and has magnetic field lines with an incidence angle of about 20 to about 30 degrees and preferably about 30 degrees relative to the external cathode-insulator interface. 1. A micro-cathode thruster assembly comprising:a tubular housing having an open distal end;a tubular cathode housed within the housing, the cathode having a distal end positioned proximate the open distal end of the housing;an insulator in contact with the cathode forming an external cathode-insulator interface;an anode housed within the housing and proximate the open distal end of the housing; anda magnetic field having magnetic field lines positioned at the external cathode-insulator interface at an incidence angle of about 20 degrees to about 30 degrees relative to the external cathode-insulator interface.2. The micro-cathode thruster assembly of claim 1 , wherein the anode is a tubular anode or a cylindrical anode.3. The micro-cathode thruster assembly of claim 2 , wherein the insulator is in contact with a distally facing end of the cathode claim 2 , and between the cathode and the anode.4. The micro-cathode thruster assembly of claim 3 , wherein the anode is positioned distal to the cathode and the insulator.5. The micro-cathode thruster assembly of claim 1 , wherein the magnetic field has a magnetic field strength of about 0.1 tesla to about ...

Thruster

A thruster comprising: a chamber to contain a fluid; a plurality of nozzles to exhaust neutral particles derived from the fluid in the chamber, wherein each nozzle has a converging section and the converging section includes a first electrode; a second electrode located distal to the first electrode to provide a voltage differential between the first and second electrodes sufficient to create plasma ions from the fluid and the voltage differential accelerates the plasma ions on a flow path through the converging section, and wherein at least one or more of the accelerated plasma ions are neutralised to form the neutral particles by charge exchange with other neutral particles, or by recombination with electrons, on the flow path. 1. A thruster comprising:a chamber to contain a fluid;a plurality of nozzles to exhaust neutral particles derived from the fluid in the chamber, wherein each nozzle has a converging section and the converging section comprises a first electrode;a second electrode located distal to the first electrode to provide a voltage differential between the first and second electrodes sufficient to create plasma ions from the fluid and the voltage differential accelerates the plasma ions on a flow path through the converging section, andwherein at least one or more of the accelerated plasma ions are neutralised to form the neutral particles by charge exchange with other neutral particles, or by recombination with electrons, on the flow path.2. A thruster according to claim 1 , wherein the plurality of nozzles are arranged in an array.3. A thruster according to claim 1 , wherein the array comprises a two-dimensional array with regular spacing between the plurality of nozzles.4. A thruster according to claim 1 , the thruster comprising a nozzle element having the plurality of nozzles in an array.5. A thruster according to claim 4 , wherein at least a portion of the nozzle element claim 4 , having the plurality of nozzles in an array claim 4 , is ...

Relative Superluminal Propulsion Drive

A Relative Superluminal Propulsion Drive that allows for the electro-mechanical means by which a vehicle of mass can be accelerated to and maintain a relative velocity greater than the universal constant C (299,792,458 meters/second, the speed of light in a vacuum) between two fixed points in space when measured from a third fixed point in space. The propulsion drive is an array of electro-mechanical antennas positioned on the forward and aft portion of the vehicle and provides for the force of acceleration to the vehicle by lowering the pressure and density of the energy state of the area in front of the vehicle and increasing the pressure and density of the energy state behind the vehicle through the collection and re-distribution of that part of the electro-magnetic spectrum responsible for maintaining the average pressure density of the void energy of space. In addition to providing propulsion the Relative Superluminal Propulsion Drive accounts for and negates the effects of the Newton's Laws of Motion during both the acceleration and deceleration portion of the travel. 1. An relative superluminal propulsion drive comprised of:an electro-magnetic energy field generator/antenna designed to absorb (or collect) triggering boson;an electro-magnetic energy field generator/antenna designed to release (or distribute) triggering boson;and an electro-mechanical device design to redistribute triggering boson.2. The relative superluminal propulsion drive of claim 1 , wherein the electro-mechanical device design to redistribute triggering boson is attached to a body of mass.3. The relative superluminal propulsion drive of claim 2 , wherein the electro-magnetic energy field generator/antenna designed to absorb (or collect) triggering boson is affixed along the longitudinal axis in the forward direction of travel of the body of mass.4. The relative superluminal propulsion drive of claim 2 , wherein the electro-magnetic energy field generator/antenna designed to release (or ...

THRUST APPARATUSES, SYSTEMS, AND METHODS

Described herein is a thrust system for a vehicle that includes at least three electrical power controllers, at least four electrical switches each configured to receive electrical power from at least one of the at least three electrical power controllers, and at least three thrusters each configured to receive electrical power from at least one of the at least three electrical switches. The at least four electrical switches are operable to switch a supply of electrical power from any of the at least three electrical power controllers to any one of the at least three thrusters. 1. A thrust system for a vehicle , comprising:at least four electrical power controllers;at least four electrical switches each configured to receive electrical power from at least one of the at least four electrical power controllers; andat least four thrusters each configured to receive electrical power from at least one of the at least four electrical switches;wherein the at least four electrical switches are operable to switch a supply of electrical power from any of the at least four electrical power controllers to any one of the at least four thrusters.2. The system of claim 1 , wherein the at least four switches are operable to allow electrical power from the at least four electrical power controllers to be concurrently supplied to the at least four thrusters claim 1 , with each electrical power controller supplying power to a respective one of the at least four thrusters.3. The system of claim 1 , wherein one of the at least four electrical power controllers is a redundant power controller claim 1 , and wherein in a first mode a first of the at least four electrical power controllers supplies electrical power to a first of the at least four thrusters claim 1 , a second of the at least four electrical power controllers supplies electrical power to a second of the at least four thrusters claim 1 , and the redundant power controller supplies no electrical power to the first and second ...

FLYING CAPACITOR MULTILEVEL CONVERTERS FOR ANODE SUPPLIES IN HALL EFFECT THRUSTERS

A flying capacitor multilevel (FCML) converter including a gate driver circuit comprising a DC-DC flyback converter having a plurality of isolated outputs. In various examples, the FCML circuit further includes a first control circuit connected to the FCML circuit determining the load current associated with a desired power output from the load; and determining a desired output voltage associated with the load current; a second control circuit that drives an inductor current (I) through the inductor so that the output applies an output voltage comprising the desired output voltage; and a third control circuit obtaining a comparison of an average of the inductor current (I) through the inductor with a predetermined reference current (I) and setting the duty cycle so that the average does not exceed the predetermined reference current. Also described is the converter driving a load comprising a plasma and a propulsion system comprising the converter. 1. A DC-DC converter circuit , comprising:an inductor;{'sub': 1n', '2n, 'a flying capacitor multilevel (FCML) converter circuit connected to the inductor and including a plurality m of transistors including a plurality of first transistors Tand a plurality of second transistors T, where m an n are integers and the FCML further comprises{'sub': 1n', '2n, 'a plurality of n cells each including one of the first transistors T, one of the second transistors T, and a capacitor having a first terminal and a second terminal;'}a first voltage rail connected to an output of the inductor and including first transistors connected in series, each of the first transistors having a first gate;a second voltage rail comprising the second transistors connected in series, each of the second transistors having a second gate; and{'sub': 'm', 'wherein each of the cells are switched on and off to charge the capacitors in response to gate voltages Vapplied to the first gates and the second gates,'}{'sub': 1n', 'm', '2n', '1n', '1n, 'sup': th', ' ...

SEQUENTIAL IMPULSE THRUSTER

The sequential impulse thruster is a system intended to provide permanent thrust to any vehicle to which it can be applied. The thrust by the system is the result of the repulsion of a perforated disc with several holes subsequent to the expulsion of compressed air or gas through the holes of the perforated disc. The compressed air or gas is expelled in sequential impulses within a hermetic frame. In order to achieve the thrust, the system is based on the organization of a set of components: a flanged and threaded axis of rotation, two types of propellers, one to realise sequences of expulsions and the other as a potential for the flowback, a perforated disc with several holes, and finally, a tube which allows separation between the expulsion of air or compressed gas and its flowback towards its pressure source. 1. A sequential impulse thruster , wherein the sequential impulse thruster is intended to produce a thrust by expelling compressed air within a hermetic frame , and the sequential impulse thruster comprises two covers , a compression valve , a pressure-relief valve , a reinforcement plate , a flanged and threaded axis of a rotation to supply a driving force , a backflow source and at least one expulsion set ,wherein each expulsion set consists of a perforated disc with holes and a sequence propeller,wherein the sequence propeller is fixed while being carried by the flanged and threaded axis of the rotation, the perforated disc is fixed to support by a thread and mounted by a contact of collars and to a separation tube, the contact of the collars is fixed by a ring,wherein the separation tube is fixed to the hermetic frame by a fixation means having a diameter smaller than a diameter of the hermetic frame, a space between the separation tube and the hermetic frame allows a potential backflow propeller to flow back expelled air towards n expulsion source of the expelled air by an outlet of a compressed air flow,wherein the compressed air within the hermetic ...

PLASMA GENERATION APPARATUS INCLUDING MEASUREMENT DEVICE AND PLASMA THRUSTER

Provided are a plasma generation apparatus including a measurement device and capable of controlling conditions of plasma properly and stably, and a plasma thruster using the plasma generation apparatus. A plasma generation apparatus including a measurement device of the present invention includes a discharge vessel, a light-emitting monitor, a probe measuring instrument, a control device, and an optical axis driving unit. The discharge vessel ionizes gas which is introduced to an inside thereof so as to generate plasma. The light-emitting monitor measures electron density of the plasma by emission spectra of the plasma. The probe measuring instrument measures the electron density of the plasma by a probe disposed in the discharge vessel. 1. A plasma generation apparatus including a measurement device , comprising;a plasma generation unit ionizing gas introduced to an inside thereof so as to generate plasma;an emission spectra measurement device measuring electron density of the plasma by emission spectra of the plasma;a probe measurement device measuring the electron density of the plasma by a probe disposed in the plasma generation unit;a control device controlling at least one of an amount of electric power to be supplied to the plasma generation unit, magnetic field distribution, and an amount of supply gas based on measurement results of the electron density of the plasma measured by the emission spectra measurement device and probe measurement device; andan optical axis driving unit changing an optical axis of light from the plasma entering the emission spectra measurement device.2. The plasma generation apparatus including the measurement device according to claim 1 , wherein the emission spectra measurement device classifies a first emission spectrum based on a neutral particle in the plasma and a second emission spectrum based on an ion in the plasma and measures a ratio of spectral intensity between the first emission spectrum and second emission spectrum ...

Electromagnetic energy momentum thruster using tapered cavity resonator evanescent modes

An electromagnetic energy momentum thruster has a cavity resonator and an electromagnetic radiation source for emitting an electromagnetic wave in evanescence into the cavity resonator. The electromagnetic wave produces a greater electromagnetic field amplitude and a greater electromagnetic radiation pressure on a primary interior surface area of the cavity resonator than on a secondary interior surface area of the cavity resonator. The difference between the electromagnetic field amplitude on the primary interior surface area and on the secondary interior surface area of the cavity resonator forms a highly directional electromagnetic energy momentum tensor and provides a highly directional general relativistic metric tensor. As a result, a force is produced on the cavity resonator in the form of a thrust or an acceleration that propels the device in a direction substantially perpendicular to the primary interior surface area.

Apparatus and Method for Operating a Heaterless Hollow Cathode, and an Electric Space Propulsion System Employing such a Cathode

A heaterless hollow cathode provides electron emission current in an electric space propulsion system. A mechanical, thermal, and electromagnetic design of the cathode apparatus is presented, and a method of operation for rapid ignition and stabilization of the cathode is provided. The keeper of the cathode apparatus has a thickness change which reduces the flow of heat away from the cathode's emitter assembly. The method for heating the emitter assembly includes controlling applied voltages so that the current flowing from the emitter assembly to the keeper is maintained at a predetermined fixed value. By this method, damage to the electron emitting surfaces of the emitter assembly by electric arcing and/or by depletion of dopant materials is avoided. 1. A heaterless hollow cathode apparatus comprising:(a) an emitter assembly comprising an electron emitter and an emitter holder, said emitter assembly defining a gas flow path passing through an emitter orifice;(b) a keeper surrounding said emitter assembly, said keeper having a keeper orifice;(c) a gas flow regulator for supplying a regulated flow of gas through said gas flow path;(d) an electrical power supply; and (i) apply an emitter-keeper voltage between said emitter assembly and said keeper while gas is supplied to a volume between said emitter assembly and said keeper to initiate a discharge between said emitter assembly and said keeper;', '(ii) monitor the value of an emitter-keeper current, flowing between said emitter assembly and said keeper, and adjust said emitter-keeper voltage so as to maintain said emitter-keeper current at a predetermined current value;', '(iii) monitor said emitter-keeper voltage so as to detect a drop in said emitter-keeper voltage to values which remain below a predetermined voltage threshold for a predetermined minimum time duration;', '(iv) actuate a main discharge circuit in which current flows from an anode to the heaterless hollow cathode; and', '(v) set said emitter-keeper ...

Ionic Threading Apparatus

This design processes free radical flows following physical principals that explain their movement conditioned by electromagnetic fields expressed in the convergence of induced field lines, in ways apart from existing designs. It describes specific means to obtain free radicals, process, and exhaust them within uniquely designed processing chambers. 1. An ion engine comprising:A bank of dedicated ion emitters, renewable power bank that feed bank of emitters, electrically insulating processing chambers with exhaust channels, supplemental bank of varying force electromagnets, electromagnetic gates, Faraday Cages, and means of measurement and control of toroidal flows2. An ion engine comprising:A bank of dedicated ion emitters, renewable power bank that feed bank of emitters, electrically insulating processing chambers with exhaust channels, supplemental bank of varying force electromagnets, electromagnetic gates, Faraday Cages, and means of measurement and control of toroidal flows3. An ion engine comprising:A bank of dedicated primary processing chambers, power bank that feed primary processing chambers, special design that augments spins exhausted from primary processing chambers, electrically insulating secondary processing chambers with exhaust channels, supplemental bank of varying force electromagnets, electromagnetic gates, Faraday Cages, means of measurement and control of toroidal flows.4. An ion engine comprising:A bank of dedicated secondary processing chambers, power bank that feed secondary processing chambers, special design that augments spins exhausted from secondary processing chambers, electrically insulating tertiary processing chambers with exhaust channels, supplemental bank of varying force electromagnets, electromagnetic gates, Faraday Cages, means of measurement and control of toroidal flows5. An ion engine comprising:A bank of dedicated tertiary processing chambers, power bank that feed tertiary processing chambers, special design that ...

PLASMAS FOR EXTRATERRESTRIAL RESOURCES AND APPLIED TECHNOLOGIES (PERT) SPACE DEBRIS REMEDIATION, MINING, AND REFINING

A process and system for the extraction of metals and gases contained on planets and asteroids (mining and refining) and for space debris remediation may include geographically localizing a material to be extracted/remediated; performing a risk analysis on the material to determine whether the material presents a serious risk of instantaneous fracture or disaggregation; using the risk analysis to qualify or refuse the material; capturing and stabilizing the qualified material in an ablation cylinder on a plasma machine (PERT station); deploying multiple magnetic hydraulic cylinders around the qualified material; equalizing and stabilizing the PERT station and the qualified material; performing ablation and destruction of the qualified material; and transforming pure elements from the ablation cylinder. 1. A process for the extraction of metals and gases contained on planets and asteroids and for space debris remediation , the process comprising:geographically localizing a material to be extracted/remediated;performing a risk analysis on the material to determine whether the material presents a serious risk of instantaneous fracture or disaggregation;using the risk analysis to qualify or refuse the material;capturing and stabilizing the qualified material in an ablation cylinder on a plasmas for extraterrestrial resources and applied technologies (PERT) station;deploying multiple magnetic hydraulic cylinders around the qualified material;equalizing and stabilizing the PERT station and the qualified material;performing ablation and destruction of the qualified material;transforming pure elements from the ablation cylinder, an ablation cylinder designed to accept a material to be processed;', 'an ablation/destruction section operatively attached to the ablation cylinder;', 'a recycling/refining preparation section operatively attached to the ablation/destruction section; and', 'a mass spectroscopy section operatively attached to the recycling/refining preparation ...

HELICON YIELD PLASMA ELECTROMAGNETIC RAM-SCRAMJET DRIVE ROCKET ION VECTOR ENGINE

HYPERDRIVE receives continuous air breathing assistance from compressed atmospheric air through a high speed magnetically core driven turbine accelerator which resolves around a common flow path tunnel. The tunnel runs from the front to the back of the engine. It is assisted by a series of radial geometric ramjet engines that share the common flow path tunnel for hypersonic exhaust but has separate inlet air from a linear aerospike which governs mass flow of air, velocity of inlet air and pressure to the turbine and/or ramjets, as well as the positioning of the shock wave at the inlet to reduce aerodynamic drag. The ramjet is of hybrid engine design where it can also function as a scramjet, thus a ram-scramjet structure for combustion in a radial configuration about the engine (aft of an electrical compressor), where the common flow path tunnel also serves as a compression tunnel to compress air through a the constantly occurring series of compression shocks entering from and around the aerospike. 1. A turbo-ram scramjet plasma rocket engine with five engine cycles , using simultaneously , dependent on what flight phase it is operating in , both a kerosene based fuel , a hydrocarbon based feel a hydrogen ion plasma generated fuel , and a drag reducing/thrust building propulsion system for high speed ascent propulsion phase , all within the same engine architecture in the flight vehicle , thereby allowing multi-engine combustion and drag reduction systems , providing fuel and oxidizer mixtures over a wide Mach number operating range , and thus capability of single stage to orbit operation.2. The engine of further comprising a central combustion chamber aligned in parallel with the superconducting bypass compression mass flow tunnel claim 1 , wherein the combustion chamber has two distinct structural designs claim 1 , one inboard claim 1 , inside the tunnel claim 1 , and one outboard the compression tunnel.3. The engine of wherein the outboard combustion chamber ...

Propellant apparatus

A system can include a reservoir configured to hold working material, a decontamination module configured to remove contaminants from the working material, a flow control mechanism configured to regulate working material flow between the reservoir and the decontamination module, and a manifold fluidly connecting the reservoir to the decontamination module.

ELECTRIC PROPULSION POWER CIRCUIT

A circuit () comprising: a first power source () supplying first current to a load () during a first Period of Time (“PoT”); a second power source () supplying a second current to the load during a second POT; a Unidirectional Current Valve (“UCV”) in series with the first power source; a current detector () in series with the UCV (); and a switch () in parallel with a series combination of the current detector and UCV to bypass the UCV during the second PoT. The current detector determines whether the second period of time has commenced and whether the switch has closed. 1. A power supply circuit , comprising:a first power source supplying a first current to a load during a first period of time;a second power source supplying a second current to the load during a second period of time;a unidirectional current valve in series with the second power source;a current detector in series with the unidirectional current valve;a switch in parallel with a series combination of the current detector and the unidirectional current valve to bypass the unidirectional current valve during the second period of time;wherein the current detector determines whether the second period of time has commenced and whether the switch has closed.2. The power supply circuit according to claim 1 , wherein current flow through the unidirectional current valve indicates that an electrical arc has been formed between two electrodes of a reaction thruster.3. The power supply circuit according to claim 1 , further comprising a controller performing operations to close the switch responsive to a detection by the current detector of current flow through the unidirectional current valve at the end of the first period of time.4. The power supply circuit according to claim 1 , wherein closure of the switch is determined to have occurred when an absence of current flow through the unidirectional current valve is detected at the beginning of the second period of time.5. The power supply circuit according ...

HIGH-EFFICIENCY ION DISCHARGE METHOD AND APPARATUS

An ion beam generator includes a discharge chamber with a backplate and tubular sidewalk A source of propellant, for example, Xenon gas is provided to the discharge chamber. First and second annular magnets are disposed on or near the backplate, and configured with alternating polarities such that a pair of ring-cusps form on the backplate, without any magnetic ring-cusp formation on the sidewalk A cathode assembly extends into the discharge chamber to provide primary electrons to ionize the propellant. 1. An ion beam generator comprising:a discharge chamber having a first end and an outflow end, the discharge chamber comprising a backplate at the first end and a tubular sidewall;a source of propellant connected to provide propellant to the discharge chamber;a first annular magnet and a second annular magnet, wherein the first and second annular magnets each have a first pole face adjacent to or narrowly spaced from the backplate and an opposite pole face oriented away from the backplate, and wherein the first and second annular magnets are configured with opposing polarity such that they generate a magnetic field in the discharge chamber defining at least two ring-cusps at the backplate;a cathode assembly extending into the discharge chamber and configured to provide primary electrons to the discharge chamber; andan extraction grid disposed at the outflow end of the discharge chamber;wherein the ion beam generator does not include any magnet configured to form a magnetic ring-cusp at the sidewall.2. The ion beam generator of claim 1 , wherein the sidewall is insulated from the backplate.3. The ion beam generator of claim 1 , wherein the first and second annular magnets each comprise either a continuous annular magnet or a discontinuous annular magnet comprising a plurality of spaced-apart magnets.4. The ion beam generator of claim 1 , wherein the first and second annular magnets are spaced from the backplate.5. The ion beam generator of claim 1 , wherein the first ...

Micro-cathode matrix arc thrusters

A matrix thruster that may be used to reposition and/or stabilize a CubeSAT satellite. The matrix thruster includes a conductive plate with an opening, a plurality of wires within the opening, a power supply electrically connected to the conductive plate or each of the plurality of wires via an inductor, and an electrical switch. The electrical switch creates a current change that creates an electric potential spike across the inductor. The electric potential spike across the inductor initiates an arc discharge between one of the wires and the conductive plate, which forms plasma that ejects cathode particles from the matrix thruster. Using multiple wires (e.g., four titanium wires) extends the lifetime of the thruster, as each wire restores an inter-electrode film needed for the other wires to continue generating plasma.

IGNITER SYSTEM FOR USE WITH ELECTRIC PROPULSION SYSTEMS

An ignitor subsystem for use in an electric propulsion system is disclosed. The igniter subsystem includes an igniter, which includes a first electrically conducting electrode, a second electrically conducting electrode, and an electrically insulating layer sandwiched between the first and the second electrically conducting electrodes, and a voltage pulse generator electrically coupled to the first and the second electrically conducting electrodes and is adapted to generate a plurality of pulses each with sufficient voltage to cause a breakdown of the electrically insulating layer, thus causing an avalanche of electrons from one of the first and the second electrically conducting electrodes to the other, the voltage pulse generator is further adapted to limit energy transferred to the igniter in each of the plurality of pulses so as to minimize damage to the igniter. 1. An ignitor subsystem for use in an electric propulsion system , comprising: a first electrically conducting electrode,', 'a second electrically conducting electrode, and', 'an electrically insulating layer sandwiched between the first and the second electrically conducting electrodes; and, 'an igniter, comprising'}a voltage pulse generator electrically coupled to the first and the second electrically conducting electrodes and adapted to generate a plurality of pulses each with sufficient voltage to cause a breakdown of the electrically insulating layer, thus causing an avalanche of electrons from one of the first and the second electrically conducting electrodes to the other, the voltage pulse generator further adapted to limit energy transferred to the igniter in each of the plurality of pulses so as to minimize damage to the igniter.2. The igniter system of claim 1 , wherein the first and the second electrically conducting electrodes are made from one or more of copper claim 1 , gold claim 1 , silver claim 1 , titanium claim 1 , tungsten claim 1 , platinum claim 1 , cadmium claim 1 , zinc claim 1 , ...

Efficient Electric Spacecraft Propulsion

A propulsion system for spacecraft is based on an electric engine that expels propellant to achieve thrust. The propellant is first ionized to generate a plasma. Plasma particles are selectively accelerated via a pulsed laser that accelerates predominantly the electrons in the plasma. The electrons are expelled first, forming a space charge that acts as a virtual cathode to accelerate the positive ions. Interactions between the laser beam and plasma electrons are predominantly through the ponderomotive force. 1. An apparatus for imparting thrust to a spacecraft , the apparatus comprising:a plasma generator operable to generate a plasma comprising generated ions;a containment vessel operable to contain the plasma;a power supply operable to generate voltage;a first outlet within the containment vessel;a second outlet within the containment vessel;a first electrode positioned adjacent to the first outlet, coupled to the power supply, and orientated such that an electric field generated by the first electrode causes electrons to move preferentially with respect to the first outlet such that some of the electrons escape from the containment vessel through the first outlet; anda second electrode positioned adjacent to the second outlet, coupled to the power supply, and oriented such that an electric field generated by the second electrode causes the generated ions to move preferentially with respect to the second outlet such that some of the generated ions escape from the containment vessel through the second outlet.2. The apparatus of claim 1 , wherein the electrons that escape the containment vessel through the first outlet form a virtual cathode outside the containment vessel such that the virtual cathode generates an electric field causing some of the generated ions to escape the containment vessel through the second outlet.3. The apparatus of claim 1 , wherein a constant charge is maintained with respect to the first electrode.4. The apparatus of of claim 1 , wherein ...

System and method for power conversion

A polarity-selectable high voltage direct current power supply including a first drive assembly that transforms a first low voltage DC input into a first medium voltage alternating current output; a first HV output assembly that transforms the first LV AC output into a first HV DC output, wherein the first HV output assembly defines a first input stage; a polarity selector coupled between the second output junction of the first drive assembly and the first and second input stages of the first HV output assembly, the polarity selector operable between a first configuration and a second configuration; wherein in the first configuration the first HV DC output has a positive polarity; and wherein in the second configuration the first HV DC output has a negative polarity.

ENGINE FOR A SPACECRAFT, AND SPACECRAFT COMPRISING SUCH AN ENGINE

A space vehicle engine () comprising a chemical thruster having a nozzle () for ejecting combustion gas, together with a Hall effect thruster. The engine is arranged in such a manner that the nozzle serves as the ejection channel for particles ejected by the Hall effect thruster when it is in operation. The engine can deliver high thrust with low specific impulse or relatively low thrust with large specific impulse. 1. A space vehicle engine comprising a chemical thruster having a nozzle for ejecting combustion gas , wherein the engine includes a Hall effect thruster arranged in such a manner that said nozzle acts as the ejection channel for particles ejected by the Hall effect thruster when it is in operation.2. An engine according to claim 1 , wherein:the Hall effect thruster has a magnetic circuit; andin a section on a meridian half-plane, the magnetic circuit is horseshoe-shaped with an airgap open to the downstream end of the nozzle;in such a manner that the magnetic circuit is suitable for generating a magnetic field in the airgap of the magnetic circuit.3. An engine according to claim 2 , wherein the nozzle has an axial section of annular shape claim 2 , and passes through the airgap of the magnetic circuit.4. An engine according to claim 3 , wherein the magnetic circuit has at least one outer magnetic core situated around the nozzle and an inner magnetic core situated radially inside the nozzle claim 3 , and in a section on a meridian half-plane claim 3 , sections of said inner core and of said at least one outer core form branches of said horseshoe-shape.5. An engine according to claim 2 , wherein the Hall effect thruster further includes an electric circuit suitable for generating an electric field in the nozzle claim 2 , and the electric circuit includes an anode and a cathode arranged respectively upstream and downstream from said airgap.6. An engine according to claim 5 , wherein the anode comprises a portion of the nozzle.7. An engine according to ...

METHOD AND APPARATUS TO PRODUCE HIGH SPECIFIC IMPULSE AND MODERATE THRUST FROM A FUSION-POWERED ROCKET ENGINE

A system and method for producing and controlling high thrust and desirable specific impulse from a continuous fusion reaction is disclosed. The resultant relatively small rocket engine will have lower cost to develop, test, and operate that the prior art, allowing spacecraft missions throughout the planetary system and beyond. The rocket engine method and system includes a reactor chamber and a heating system produce fusion reactions the stable plasma. Magnets produce a magnetic field that confines the stable plasma. A fuel injection system and a propellant injection system are included. Cold propellant into a gas box for converting a cold propellant into a warm propellant plasma at one end of the reactor chamber. The propellant and fusion products are directed out of the reactor chamber through a magnetic nozzle and are detached from the magnetic field lines producing thrust. 1. A system comprising a rocket engine employing nuclear fusion with thrust augmentation , the system comprisinga reactor chamber for containing a stable plasma comprising a fusion fuel;a heating system for heating said plasma and increasing an ion energy of said plasma to a level sufficient for producing net power from fusion reactions in said stable plasma;a plurality of magnets coaxial to said reactor chamber, the plurality of magnets producing a magnetic sufficient to confine the stable plasma and promote rapid loss of fusion products into a scrape off layer;a fuel injection system for injecting additional quantities of said fusion fuel to sustain the power output of said fusion reaction;a gas box at one end of said scrape off layer for introducing a cold propellant to be heated in the reactor chamber into a warm propellant plasma;a propellant injection system for injecting said cold propellant into said gas box to augment the mass of said scrape off layer;a magnetic nozzle for directing said warm plasma on said scrape off layer out of the end of said reactor chamber distal from said gas ...

CHAMBER BOTTOM FOR A PLASMA THRUSTER

Chamber bottom for a plasma thruster making it possible to combine several functions in a single piece and, in particular, to fasten certain insulating parts of the plasma thruster, the chamber bottom having, in a single piece, a chamber bottom surface for closing an annular chamber formed by the chamber bottom and at least one insulating part attached to the chamber bottom, and at least a first set of tabs including fastening tabs for fastening the at least one insulating part to the chamber bottom. 1. A chamber bottom for a plasma thruster , comprising , in a single piece , a chamber bottom surface for closing an annular chamber formed by the chamber bottom and delimited by at least one insulating part attached to the chamber bottom , andat least a first set of tabs comprising fastening tabs for fastening said at least one insulating part to the chamber bottom.2. The chamber bottom according to claim 1 , made of a metallic material.3. The chamber bottom according to claim 1 , produced by additive manufacturing.4. The chamber bottom according to claim 1 , wherein the chamber bottom surface has in section a generally U-shaped profile claim 1 , having at least one point of inflection.5. The chamber bottom according to claim 1 , comprising a first inner set of tabs claim 1 , comprising fastening tabs for fastening a first insulating part to the chamber bottom claim 1 , and a second outer set of tabs claim 1 , comprising fastening tabs for fastening a second insulating part to the chamber bottom.6. The chamber bottom according to claim 1 , wherein at least one set of tabs comprises dummy tabs configured so as not to cooperate with said at least one insulating part.7. The chamber bottom according to claim 1 , further comprising at least one distribution cavity communicating with the chamber via injection orifices opening onto the chamber bottom surface.8. The chamber bottom according to claim 7 , comprising a first distribution cavity communicating with a second ...

CHEMICAL-ELECTROMAGNETIC HYBRID PROPELLER WITH VARIABLE SPECIFIC IMPULSE

A chemical-electromagnetic hybrid propeller with variable specific impulse. Fuel gas ejected out from a spraying tube of the chemical propeller through chemical propulsion enters an ionization chamber through a first magnetic mirror tube for ionization. The fuel gas after ionization is heated up by radio-frequency ion cyclotron waves in an ion cyclotron wave heating chamber so as to improve the kinetic energy. Then a second magnetic mirror tube is used, so that ions in the fuel gas after the ionization are heated up many times in a reciprocating manner between the magnetic mirror tubes, and ejected to generate forward propulsion force. By means of the propeller, the propulsion force and the specific impulse are greatly increased. 1. A chemical-electromagnetic hybrid propeller with variable specific impulse , comprising: a chemical propeller; double magnetic mirror tubes; an ionization chamber; and an ion cyclotron wave heating chamber , wherein the first magnetic mirror tube is connected on the rear end of the spraying tube of the chemical propeller in a sealing way , the other end of the first magnetic mirror tube is connected with the ion cyclotron wave heating chamber by the ionization chamber in a sealing way , and the other end of the ion cyclotron wave heating chamber is connected with a second magnetic tube in a sealing way , wherein fuel gas ejected out from the spraying tube of the chemical propeller through chemical propulsion enters the ionization chamber through the first magnetic mirror tube for ionization , the fuel gas after ionization is heated up by radio-frequency ion cyclotron waves in the ion cyclotron wave heating chamber for improving improve kinetic energy , the second magnetic mirror tube is used so that ions in the fuel gas after the ionization can be heated up many times in a reciprocating manner between the magnetic mirror tubes , and plasma jet frame is ejected to generate forward propulsion force , wherein the first magnetic mirror tube ...

CHARGE SEPARATION MECHANISM

A method of producing a charge separation in a plasma having a low particle density which comprises a plurality of electrons and a plurality of positive ions. The method includes generating a magnetic field and passing the plasma having a low particle density along a first axis through the magnetic field. The magnetic field is generated having a component which is perpendicular to the first axis and is configured so as to deflect the plurality of electrons from the first axis and allow the plurality of positive ions to travel substantially undeflected along the first axis. Also provided is a magnetohydrodynamic generator and a low earth orbit thruster making use of the charge separation mechanism. 1. A method of producing a charge separation in a plasma having a low particle density , which comprises a plurality of electrons and a plurality of positive ions , the method comprising:generating a magnetic field; andpassing the low particle density plasma along a first axis through the magnetic field,wherein the magnetic field is generated having a component which is perpendicular to the first axis and is configured so as to deflect the plurality of electrons from the first axis, and to allow the plurality of positive ions to travel substantially undeflected along the first axis.2. The method according to claim 1 , wherein the plasma having the low particle density is a substantially ideal plasma in which the plurality of positive ions and the plurality of electrons move substantially independently of one another.3. The method according to claim 1 , wherein a plasma frequency of the plasma having the low particle density is larger than a particle collision frequency of the plasma having the low particle density.4. The method according to claim 1 , wherein the particle density of the plasma is lower than 10m.5. The method according to claim 1 , wherein the magnetic field is configured so as to deflect the plurality of electrons to travel in a closed drift loop around the ...

Staging of ion propulsion thrusters

Spacecraft thruster systems are disclosed. In some instances, a spacecraft thruster system may include stacked ion thrusters and/or ion thruster layers. The ion thrusters and/or ion thruster layers may be sequentially activated and jettisoned from the thruster system after use.

ELECTROTHERMAL RADIO FREQUENCY THRUSTER AND COMPONENTS

The invention provides an electrothermal RF plasma production system and thruster design, and associated components, that may be used in terrestrial applications and/or miniaturized to the mass, volume, and power budget of Cube Satellites (CubeSats) to meet the propulsion needs of the small satellite (˜5 to ˜500 kg) constellations and larger satellite buses. 1. A plasma production device comprising:(a) a substantially cylindrical plasma production chamber having a cylinder body, a first closed end and a second open end;(b) a magnet system comprising one or more radially-disposed magnets configured to establish a magnetic field within the plasma production chamber and oriented substantially parallel to a central longitudinal axis of the plasma production chamber such that each magnet produces a magnetic field of the same polarity within the plasma production chamber;(c) a propellant tank and a flow regulator in communication with the plasma production chamber and configured to deliver a gaseous propellant into the plasma production chamber; and(d) a radio frequency (RF) antenna comprising a flat spiral region external to the plasma production chamber and disposed on an external surface of the first closed end, electrically coupled to an AC power source, and configured to deliver an RF energy to an interior region of the plasma production chamber.2. The plasma production device of claim 1 , wherein the flat spiral region has 1-10 turns.3. The plasma production device of claim 1 , wherein the flat spiral region comprises a spiral region radius and the first closed end comprises a closed end radius claim 1 , and wherein the spiral region radius is 10%-100% of the closed end radius.4. The plasma production device of claim 3 , wherein the spiral region radius is 50%-100% of the closed end radius.5. The plasma production device of claim 1 , wherein the spiral region is configured to cause a constructive interference in magnetic fields produced within the plasma production ...

ELECTROTHERMAL RADIO FREQUENCY THRUSTER AND COMPONENTS

The invention provides an electrothermal RF plasma production system and thruster design, and associated components, that may be used in terrestrial applications and/or miniaturized to the mass, volume, and power budget of Cube Satellites (CubeSats) to meet the propulsion needs of the small satellite (˜5 to ˜500 kg) constellations and larger satellite buses. 1. A plasma production device comprising:(a) a substantially cylindrical plasma production chamber having a cylinder body, a first closed end and a second open end;(b) a magnet system comprising one or more radially-disposed magnets configured to establish a magnetic field within the plasma production chamber and oriented substantially parallel to a central longitudinal axis of the plasma production chamber such that each magnet produces a magnetic field of the same polarity within the plasma production chamber;(c) a propellant tank and a flow regulator in communication with the plasma production chamber and configured to deliver a gaseous propellant into the plasma production chamber; and(d) a radio frequency (RF) antenna comprising a flat spiral region external to the plasma production chamber and disposed on an external surface of the first closed end, electrically coupled to an AC power source, and configured to deliver an RF energy to an interior region of the plasma production chamber.2. The plasma production device of claim 1 , wherein the flat spiral region has 1-10 turns.3. The plasma production device of claim 1 , wherein the flat spiral region comprises a spiral region radius and the first closed end comprises a closed end radius claim 1 , and wherein the spiral region radius is 50%-100% of the closed end radius.4. The plasma production device of claim 1 , wherein the spiral region is configured to cause a constructive interference in magnetic fields produced within the plasma production chamber.5. The plasma production device of claim 1 , wherein the antenna further comprises a coiled region disposed ...

Electrically Powered Propulsion System for Use in a Spacecraft

An electrically powered propulsion system for a spacecraft includes a first center of gravity at a first time of operation and a second center of gravity at a second time of operation, where the second center of gravity is different from the first center of gravity. The electrically powered propulsion system includes a thruster realignment mechanism and at least two thrusters coupled to the thruster realignment mechanism. Each of the at least two thrusters has an individual thrust vector. The thruster realignment mechanism is adapted such that, in a first position, the individual thrust vectors of the at least two thrusters pass through the first center of gravity and that, in a second position, the individual thrust vectors of the at least two thrusters pass through the second center of gravity. The thruster realignment mechanism holds the first position in the event all of the at least two thrusters are without any failure. In addition, the thruster realignment mechanism realigns the thrusters to the second position in the event of at least one of (i) a failure of one of the at least two thrusters, and (ii) a predetermined time criterion is fulfilled. 1. An electrically powered propulsion system fir a spacecraft that comprises a first center of gravity at a first time of operation and a second center of gravity at a second time of operation , wherein the second center of gravity is different from the first center of gravity , wherein the electrically powered propulsion system comprises a thruster realignment mechanism and at least two thrusters coupled to the thruster realignment mechanism , each of the at least two thrusters having an individual thrust vector , whereinthe thruster realignment mechanism is adapted such that, in a first position, the individual thrust vectors of the at least two thrusters pass through the first center of gravity and that, in a second position, the individual thrust vectors of the at least two thrusters pass through the second ...

APPARATUS, SYSTEMS AND METHODS FOR ESTABLISHING PLASMA AND USING PLASMA IN A ROTATING MAGNETIC FIELD

Systems and methods establish plasma in a rotating magnetic field. An exemplary embodiment is a plasma thruster that establishes a first transverse magnetic field with respect to a system axis of a plasma propulsion system; establishes a second transverse magnetic field oriented orthogonally to the first transverse magnetic field, wherein the second transverse magnetic field is out of phase with the first transverse magnetic field; and establishes a magnetic field aligned with the system axis using a plurality of magnet elements oriented along the system axis. A plasma containment portion defines an interior region, wherein an interior region of a plasma containment portion accommodates a plasma that is established by a rotating magnetic field component that is cooperatively established by the first transverse magnetic field and the second transverse magnetic field, and wherein the plasma is accelerated out of the plasma containment portion by magnetic forces to generate a propulsion force. 1. A method , comprising:establishing a first transverse magnetic field with respect to a system axis of a plasma propulsion system;establishing a second transverse magnetic field oriented orthogonally to the first transverse magnetic field, wherein the second transverse magnetic field is out of phase with the first transverse magnetic field; andestablishing a magnetic field aligned with the system axis using a plurality of magnet elements oriented along the system axis,wherein a plasma containment portion defines an interior region,wherein the interior region of the plasma containment portion accommodates a plasma that is established by a rotating magnetic field component that is cooperatively established by the first transverse magnetic field and the second transverse magnetic field, andwherein the plasma is accelerated out of the plasma containment portion by magnetic forces to generate a propulsion force.2. The method of claim 1 , further comprising:inputting a first ...

PLASMA MICRO-THRUSTER

A plasma micro-thruster, including: an elongate and substantially non-conductive tube having a first end to receive a supply of propellant gas, and an open second end to act as an exhaust; first, second, and third electrodes extending circumferentially around the tube and being mutually spaced along a longitudinal axis of the tube, the third electrode being longitudinally interposed between the first and second electrodes; wherein the tube and the first, second and third electrodes are configured to generate a plasma from propellant gas flowing though the tube from the first end of the tube when the third electrode receives radio frequency power and the first and second electrodes are electrically grounded relative to the third electrode, such that the expansion of the plasma from the open end of the tube generates a corresponding thrust. 16.-. (canceled)7. A plasma micro-thruster , including:an elongate and substantially non-conductive tube having a first end to receive a supply of propellant gas, and an open second end to act as an exhaust;first, second, and third electrodes extending circumferentially around the tube and being mutually spaced along a longitudinal axis of the tube, the third electrode being longitudinally interposed between the first and second electrodes;wherein the tube and the first, second and third electrodes are configured to generate a plasma from propellant gas flowing though the tube from the first end of the tube when the third electrode receives radio frequency power and the first and second electrodes are electrically grounded relative to the third electrode, such that the expansion of the plasma from the open end of the tube generates a corresponding thrust.8. A plasma micro-thruster , including:a tube having a length greater than its width, receiving at one end a supply of propellant gas, and having the other end open as an exhaust;a first and a second conductive electrodes in a spaced-apart arrangement surrounding the tube, each ...

DEVICE AND METHOD FOR REGULATING FLOW RATE

A flow rate regulator device is provided, including an upstream chamber, a downstream chamber, a plurality of electrically conductive capillary ducts providing parallel fluid flow connections between the upstream chamber and the downstream chamber, first and second electrical terminals configured to be connected to an electric current source, and at least one electric switch configured to connect one or more of the capillary ducts selectively between the electrical terminals. A system for feeding propellant gas to a space electric thruster is also provided, including at least one such flow rate regulator device to regulate a propellant gas flow rate. And, a flow rate regulation method is provided, using the flow rate regulator device. 110-. (canceled)11. A flow rate regulator device , comprising:an upstream chamber;a downstream chamber;a plurality of electrically conductive capillary ducts providing parallel fluid flow connections between the upstream chamber and the downstream chamber;first and second electrical terminals configured to be connected to an electric current source; andat least one electric switch configured to selectively connect one or more of the plurality of electrically conductive capillary ducts between the first and the second electrical terminals.12. The flow rate regulator device according to claim 11 , wherein at least one electric switch is configured to selectively connect one electrically conductive capillary duct of the plurality of electrically conductive capillary ducts or multiple electrically conductive capillary ducts of the plurality of said capillary ducts claim 11 , in series claim 11 , between the first and the second electrical terminals.13. The flow rate regulator device according to claim 12 ,wherein the plurality of electrically conductive capillary duct comprises at least three electrically conductive capillary ducts, andwherein the flow rate regulator device further comprises a plurality of electric switches configured to ...

MAGNETICALLY SHIELDED MINIATURE HALL THRUSTER

Magnetically shielded miniature Hall thrusters are disclosed that use a unique magnetic field topology that prevents the magnetic field lines from intersecting the discharge channel walls in the acceleration region of the thruster. Instead, the lines of force originating from both the inner and outer pole pieces curve around the downstream edges of the discharge channel and follow the channel walls towards the anode. This unique field topology results in low electron temperature at the discharge channel walls while eliminating strong electric field components that would otherwise lead to high erosion rates and power deposition from ion acceleration into the channel walls. 1. A miniature Hall thruster , comprising:a discharge component;the discharge component comprising channel walls defining a discharge channel;an anode disposed in a first end of said discharge channel;the discharge channel having an open second end opposite said first end;the discharge channel configured to create a magnetic shield disposed to at least partially encircle said channel walls of said discharge component; andwherein said magnetic shield is configured to manipulate a magnetic field associated with the discharge component such that the magnetic field extends into said discharge channel from said open second end substantially without intercepting said channel walls of said discharge component so as to prevent a plasma formed in said discharge channel from contacting said channel walls.2. A miniature Hall thruster as recited in claim 1 , wherein an output power of said Hall thruster is less than 1 kW.3. A miniature Hall thruster as recited in claim 2 , wherein the power of said Hall thruster is less than 1 kW and greater than 25 W.4. A miniature Hall thruster as recited in claim 3 , wherein a power of said Hall thruster is less than 500 kW and greater than 100 W.5. A miniature Hall thruster as recited in claim 1 , wherein said discharge channel has a diameter that is less than 10 cm.6. A ...

NTAC Augmented Nuclear Electric Propulsion and/or Nuclear Thermal Propulsion

The present disclosure is directed to a system including a nuclear thermal rocket or a nuclear reactor, at least one nuclear electric thruster coupled to the nuclear thermal rocket or the nuclear reactor, and a Nuclear Thermionic Avalanche Cell (NTAC) configured to generate electrical power. The NTAC cell may be positioned around a nuclear reactor core of the nuclear thermal rocket or the nuclear reactor, and the nuclear electric thruster may be powered by the NTAC generated electrical power. 1. A system comprising:a nuclear thermal rocket;at least one nuclear electric thruster coupled to the nuclear thermal rocket; and 'wherein the NTAC cell is positioned around a nuclear reactor core of the nuclear thermal rocket, and wherein the at least one nuclear electric thruster is powered by the generated electrical power.', 'a Nuclear Thermionic Avalanche Cell (NTAC) configured to generate electrical power;'}2. The system of claim 1 , wherein the nuclear reactor core comprises U-235.3. The system of claim 1 , wherein the nuclear reactor core comprises Pu-239.4. The system of claim 1 , wherein the nuclear reactor core comprises a ceramic encapsulated nuclear fuel.5. The system of claim 1 , wherein the nuclear reactor core comprises a tri-structural isotropic particle fuel.6. The system of claim 1 , wherein the NTAC is a radiation shield.7. The system of claim 1 , wherein the NTAC converts primary high energy photons claim 1 , energetic particles such as beta particles claim 1 , or induced high energy field of photons from the nuclear reactor core of the thermal rocket into electricity.8. The system of claim 7 , wherein the NTAC includes a plurality of layers each comprising:a collector;an insulator; and wherein the collector is positioned across a thermionic vacuum gap;', 'wherein the emitter is configured to capture gamma ray photons from the nuclear reactor core of the thermal rocket;', 'wherein the captured photons free up a large number of electrons in an avalanche ...

ELECTRIC PROPULSION SYSTEM WITH STATIONARY PLASMA THRUSTERS

An electric propulsion system includes a first stationary plasma thruster including a single first cathode, a first anode, and a first gas manifold, and a second stationary plasma thruster including a single second cathode, a second anode, and a second gas manifold. The system further includes an electrical connection device common to the first and second cathodes, first and second gas flow rate control devices with a common fluid flow device for feeding gas, and a selective control device for activating at any given instant only one of the first and second cathodes in co-operation with one or the other of the first and second anodes. 19-. (canceled)10. An electric propulsion system with stationary plasma thrusters , the system comprising:at least one device for regulated delivery of gas under high pressure;first and second power processor units;first and second external thruster switch units;first and second electrical filters; andfirst and second juxtaposed stationary plasma thrusters;the first stationary plasma thruster comprising a first ionization channel, a single first cathode arranged in a vicinity of an outlet from the first ionization channel, a first anode associated with the first ionization channel, a first gas manifold, and first devices for creating a magnetic field around the first ionization channel; andthe second stationary plasma thruster comprising a second ionization channel, a second single cathode arranged in a vicinity of an outlet from the second ionization channel, a second anode associated with the second ionization channel, a second gas manifold, and second devices for creating a magnetic field around the second ionization channel;the electric propulsion system further comprising:an electrical connection device common to the first and second cathodes;first and second gas flow rate control devices associated respectively with each of the first and second stationary plasma thrusters, with a common fluid flow device for feeding gas to the ...

SYSTEM AND METHOD FOR SMALL, CLEAN, STEADY-STATE FUSION REACTORS

According to some embodiments, a system for widening and densifying a scrape-off layer (SOL) in a field reversed configuration (FRC) fusion reactor is disclosed. The system includes a gas box at one end of the reactor including a gas inlet system and walls of suitable heat bearing materials. The system further includes an exit orifice adjoining the gas box, wherein the exit orifice has a controllable radius and length to allow plasma to flow out from the gas box to populate the SOL with the plasma. The system may also include fusion products, which decrease in speed in the plasma in the SOL, allowing energy to be extracted and converted into thrust or electrical power and further allowing ash to be extracted to reduce neutron emissions and maintain high, steady-state fusion power. 1. A method for increasing He supply for use in field reversed configuration (FRC) fusion reactors , comprising:{'sup': '3', 'burning deuterium (D) with D in a first FRC reactor to breed first helium-3 (He) and tritium (T);'}{'sup': '3', 'separating the first He from the T through a permeable membrane in the first FRC reactor;'}{'sup': '3', 'producing power in the first FRC reactor by burning the first He with the D in the first FRC reactor;'}{'sup': '3', 'storing the T for a predetermined period of time such that the T transmutes to second He; and'}{'sup': '3', 'producing power in a second FRC reactor by burning the second He with D in the second FRC reactor.'}2. The method of claim 1 , wherein breeding first He and T comprises exhausting ash that includes the first He and T through an exhaust plume in the first FRC reactor.3. The method of claim 2 , wherein the ash is exhausted before it can fuse.4. The method of claim 1 , wherein breeding first He and T comprises exhausting ash that includes the first He and T through a gas box in the first FRC reactor.5. The method of claim 1 , wherein separating the first He from the T comprises separating via superpermeation.6. The method of claim 1 ...

HALL THRUSTER WITH ANNULAR CATHODE

A Hall thruster includes an annular discharge region and an annular cathode concentric to the annular discharge region. 1. A Hall thruster comprising an annular discharge region and an annular cathode concentric to the annular discharge region.2. The Hall thruster as recited in claim 1 , wherein the annular cathode is circumscribed by the annular discharge region.3. The Hall thruster as recited in claim 1 , wherein the annular cathode circumscribes the annular discharge region.4. The Hall thruster as recited in claim 1 , further comprising an anode adjacent the annular discharge region.5. A Hall thruster comprising:inner and outer magnetic poles;an annular discharge region between the inner and outer magnetic poles, the annular discharge region defining a central axis;a propellant gas feeder operable to feed propellant gas to the annular discharge region;an annular cathode circumscribing the central axis;an anode; andat least one magnet magnetically coupled with the inner and outer magnetic poles to generate a magnetic field across the annular discharge region.6. The Hall thruster as recited in claim 5 , wherein the annular cathode is circumscribed by the annular discharge region.7. The Hall thruster as recited in claim 5 , wherein the annular cathode circumscribes the annular discharge region. Ion accelerators with closed electron drift are also known as Hall-effect thrusters or Hall thrusters. Hall thrusters can be used on space vehicles for propulsion, station-keeping, orbit changes, or counteracting drag, for example. Hall thrusters generate thrust by supplying a propellant gas to an annular channel. The annular channel has a closed end with an anode and an open end through which the gas is discharged. A cathode introduces free electrons into the area of the open end. The electrons are induced to drift circumferentially in the annular channel by a generally radially extending magnetic field in combination with a longitudinal electric field, but the electrons ...

Propellant injector system for plasma production devices and thrusters

An electrothermal plasma production device is presented. The plasma production device includes: a plasma production chamber; an RF antenna external to the plasma production chamber; a propellant tank and flow regulator external to the plasma production chamber and in communication with the plasma production chamber; and a plenum disposed between the propellant tank and the plasma production chamber. The RF antenna, in combination with an AC power source, is configured to provide an RF energy to an interior region of the plasma production chamber and to an interior region of the plenum with sufficient power to ionize at least some of the propellant in the plenum. The plasma production chamber is configured to include a propellant injector for receiving propellant at a first closed end of the plasma production chamber.

PROPELLANT TANK AND LOADING FOR ELECTROSPRAY THRUSTER

Methods and apparatus of adding propellant to a thruster assembly are described. A first end of a beaker is disposed in an opening of the tank, where the beaker contains propellant and the first end of the beaker includes a breakaway bottom. The thruster assembly and beaker are placed in a first environment, where the first environment is substantially a vacuum and/or an environment composed substantially of gases that can be absorbed by the propellant. A plunger in the beaker is depressed to cause the breakaway bottom of the beaker to break and the propellant to flow into the tank of the thruster assembly. The thruster assembly is removed from the first environment and the beaker is removed from the opening. A cap is added to complete the assembly. The assembly contains a vent to allow gases to escape the interior of the tank. 119-. (canceled)20. A method of adding a liquid propellant to a thruster assembly , the method comprising:placing the thruster assembly in an environment that is composed substantially of gas that is absorbable by the liquid propellant, wherein the thruster assembly includes a tank and a porous reservoir disposed within an interior of the tank; andfilling the tank of the thruster assembly with the liquid propellant.21. The method of claim 20 , wherein the liquid propellant comprises an ionic liquid.22. The method of claim 20 , wherein filling the tank includes filling the tank with a beaker filled with the liquid propellant claim 20 , and wherein the beaker is removably disposed in an opening of the tank.23. The method of claim 22 , further comprising breaking a bottom of the beaker to flow the liquid propellant into the tank.24. The method of claim 23 , further comprising removing the beaker from the opening.25. The method of claim 20 , further comprising displacing a plunger to flow the liquid propellant into the tank.26. The method of claim 20 , further comprising transporting the liquid propellant from the porous reservoir to a porous ...

Electron Propulsion Engine

An electron acceleration device using thermionic fission cells and an electromagnetic scoop coil for power. A power control junction and electron injector control that feeds free electrons in packets into the acceleration components that consist of a series of induction linac module units, having quadrupole magnet units in series between the induction module units. Has on-board xenon gas for a deep space electron source. At the high speed electrons exit from the device, deflector plates control the exit path of the electrons to direct the course of a craft. 1. An electron accelerator including:thermionic fission cells for generating electricity;an electron injector for converting electricity into free electron packets;an induction linear accelerator for accelerating said free electron packets converted by said electron injector;a linear electron path through said linear accelerator for acceleration of said free electron packets in a straight path.2. An electron accelerator as set forth in wherein:said linear accelerator is in the form of a series of acceleration modules;a quadrupole magnet is positioned between said electron acceleration modules to control and vector said free electron packets path.3. An electron accelerator as set forth in wherein:said electron injector is positioned at the beginning or front end of said linear accelerator for converting said electric current into free electron packets for acceleration in said linear accelerator.4. An electron accelerator as set forth in wherein:a power control junction is positioned at the front end of said linear accelerator to control and coordinate the flow of current to and from all electrical components within said electron accelerator.5. An electron accelerator as set forth in wherein:electron deflector plates are positioned at the exit from said electron accelerator for controlling the direction or vector of said electrons after they leave said electron accelerator.6. An electron accelerator as set forth in ...

BATTERY POWERED VEHICLE PROPULSION SYSTEM

A vehicle propulsion system includes an air heating chamber that receives inlet air from an air intake chamber and provides thrust through an exhaust chamber. A battery powered pulse generator generates a pulsed electrical output signal. An amplifier amplifies the pulsed electrical output signal to provide an amplified pulsed power output signal to the air heating chamber. The amplified pulsed power output signal directly heats the inlet air to generate thrust through the exhaust chamber. 1. A vehicle propulsion system , comprising:an air heating chamber that receives inlet air from an air intake chamber and provides thrust through an exhaust chamber;a battery powered pulse generator to generate a pulsed electrical output signal; andan amplifier to amplify the pulsed electrical output signal to provide an amplified pulsed power output signal to the air heating chamber, wherein amplified pulsed power output signal directly heats the inlet air to generate thrust through the exhaust chamber.2. The system of claim 1 , further comprising an electrode that is driven by the amplifier to provide the pulsed power output signal to the air heating chamber.3. The system of claim 1 , further comprising a valve disposed between the air intake chamber and the air heating chamber claim 1 , wherein the valve closes to shut off the inlet air after the amplified pulsed power output signal heats the inlet air in the air heating chamber.4. The system of claim 3 , wherein the valve opens to enable the inlet air to be received by the air inlet chamber a period of time after the thrust is generated though the exhaust chamber.5. The system of claim 4 , wherein the battery powered pulse generator generates the pulsed electrical output signal according to a frequency and duty cycle to open and close the valve according to a resonant engine cycle.6. The system of claim 1 , wherein the air intake chamber and the exhaust chamber are tuned such that at least two pressure wavefronts are generated ...

Gas Inlet for an Ion Thruster

A gas inlet (), in particular for use in an ion thruster, comprises a housing () which is made of a gas-tight ceramics material and which is provided with a first gas feed channel () and a second gas feed channel () arranged downstream of the first gas feed channel (). The gas inlet () further comprises an insert () which is arranged in the second gas feed channel () and is made of a porous ceramics material, wherein the geometry and pore structure of the insert () are such that the insert () forms a desired flow resistance for a gas stream flowing through the second gas feed channel () which is greater than a flow resistance acting on a gas stream flowing through the first gas feed channel (), and wherein a ratio of a length () of the first gas feed channel () to a length () of the insert () is at least 1:2. 1. A gas inlet for use in an ion thruster , wherein the gas inlet comprises:a housing made of a gas-tight ceramics material having a first gas feed channel and a second gas feed channel arranged downstream of the first gas feed channel; andan insert arranged in the second gas feed channel and made of a porous ceramics material,wherein a geometry and pore structure of the insert are such that the insert forms a desired flow resistance for a gas stream flowing through the second gas feed channel which is greater than a flow resistance acting on a gas stream flowing through the first gas feed channel, andwherein a ratio of a length of the first gas feed channel to a length of the insert is at least 1:2.2. The gas inlet as claimed in claim 1 ,wherein the first gas feed channel, the second gas feed channel and the insert are configured such that, at a given breakdown voltage, a product of a pressure drop in a gas stream flowing through the gas inlet and an electrode gap formed by a sum of the length of the first gas feed channel and a length of the second gas feed channel lies within a predetermined range.3. The gas inlet as claimed in claim 2 ,wherein, at the given ...

NEUTRALIZER FOR AN ION ENGINE, METHOD OF OPERATING A NEUTRALIZER AND ION ENGINE

A neutralizer suitable for use in an ion engine comprises a halogen gas source and an electrode tube comprising an inlet opening connected to the halogen gas source for supplying a halogen gas provided by the halogen gas source into the electrode tube, a discharge space for generating a plasma from the halogen gas supplied into the electrode tube, and an outlet opening for discharging the plasma generated in the discharge space and free electrons from the electrode tube. An electron emitter is arranged in the discharge space of the electrode tube, which is at least partially made of tungsten, a tungsten alloy or a tungsten composite material containing at least one of iridium, rhenium, ruthenium, rhodium and osmium. 1. A neutralizer for use in an ion engine , comprising:a halogen gas source;an electrode tube comprising an inlet opening connected to the halogen gas source for supplying a halogen gas provided by the halogen gas source into the electrode tube,a discharge space for generating a plasma from the halogen gas supplied into the electrode tube, andan outlet opening for discharging the plasma generated in the discharge space and free electrons from the electrode tube; andan electron emitter arranged in the discharge space of the electrode tube, the electron emitter being at least partially made of tungsten, a tungsten alloy or a tungsten composite material containing at least one of iridium, rhenium, ruthenium, rhodium and osmium.2. The neutralizer according to claim 1 , further comprising a heating device for heating the electron emitter.3. The neutralizer according to claim 1 , wherein the electron emitter is configured as a tungsten filament.4. The neutralizer according to claim 1 , wherein the electrode tube is configured as a hollow cathode.5. The neutralizer according to claim 1 , further comprising a keeper electrode arranged adjacent to the outlet opening of the electrode tube.6. The neutralizer according to claim 5 , wherein the electrode tube and the ...

PLASMA ACCELERATING APPARATUS AND PLASMA ACCELERATING METHOD

Plasma which is supplied from a supply passage (1) is accelerated with a Hall electric field (E) which is generated through interaction of electrons (e) emitted from a cathode (3), a radial direction magnetic field (Bd), and an electric field (Ex). 1. A plasma accelerating apparatus comprising:a magnetic field generation body;a supply passage disposed to cross a central region of the magnetic field generation body, and configured to supply plasma from an upstream side toward a downstream side;a cathode disposed on a downstream side from the magnetic field generation body;an anode disposed on an upstream side from the cathode; anda voltage applying unit configured to generate an electric field between the cathode and the anode,wherein the magnetic field generation body generates an axial direction magnetic field in a center region of the magnetic field generation body, and generates a magnetic field which contains a radial direction magnetic field on a downstream side from the magnetic field generation body,wherein the voltage applying unit generates an electric field between the cathode and the anode, andwherein the plasma supplied through the supply passage is accelerated with a Hall electric field generated through the interaction of electrons emitted from the cathode, the radial direction magnetic field, and the electric field.2. The plasma accelerating apparatus according to claim 1 , further comprising:a magnetic flux collection body disposed on a downstream side from the magnetic field generation body,wherein the radial direction magnetic field is generated by the magnetic field generation body and the magnetic flux collection body.3. The plasma accelerating apparatus according to claim 2 , wherein a region with a sparse magnetic flux density is formed by the magnetic field generation body and the magnetic flux collection body on a downstream side from the magnetic field generation body and the magnetic flux collection body claim 2 , and the plasma which ...

Spacecraft Propulsion Devices and Systems with Microwave Excitation

A multi-mode thruster system for use in a spacecraft includes a microwave source; a cavity coupled to the microwave source and including a first inlet to receive a first fluid and a second inlet to receive a second fluid; and a nozzle provided at one end of the cavity. The thruster operates in a microwave electrothermal thruster (MET) mode to (i) generate a standing wave in the cavity using the microwave source and (ii) raise a temperature of the first fluid to generate a first hot gas that exits the cavity via the nozzle to generate thrust. The thruster operates in a chemical propulsion mode to (i) produce a reduction-oxidation reaction between the first fluid and the second fluid and (ii) generate a second hot gas that exits the cavity via the nozzle to generate thrust. 1. A multi-mode thruster system for use in a spacecraft , the thruster system comprising:a microwave source;a cavity coupled to the microwave source and including a first inlet to receive a first fluid and a second inlet to receive a second fluid; anda nozzle provided at one end of the cavity;wherein:the thruster operates in a microwave electrothermal thruster (MET) mode to (i) generate a standing wave in the cavity using the microwave source and (ii) raise a temperature of the first fluid to generate a first hot gas that exits the cavity via the nozzle to generate thrust, andthe thruster operates in a chemical propulsion mode to (i) produce a reduction-oxidation reaction between the first fluid and the second fluid and (ii) generate a second hot gas that exits the cavity via the nozzle to generate thrust.2. The multi-mode thruster system of claim 1 , wherein the cavity receives at least one of (i) water claim 1 , (ii) hydrozene claim 1 , (iii) hydrogen peroxide claim 1 , or (iii) ammonia as the first fluid via the first inlet when the thruster operates in the MET mode.3. The multi-mode thruster system of claim 1 , wherein the cavity operates as a resonant cavity when the thruster operates in the ...

Ion propelled vehicle

An ion powered assembly includes a collector assembly and an emitter assembly, comprising a plurality of conductive emitter wires supported by the emitter wire support structure. A control circuit is operatively connected to at least the emitter and collector assemblies and includes a power supply configured to provide voltage to the emitter and collector assemblies.

Plasma electric propulsion device

The present disclosure provides a plasma electric propulsion device comprising a capacitive energy storage device as a power source for an engine configured to heat and/or ionize and/or accelerate a propellant due to action of an electric field and/or magnetic field. The energy storage device comprises: a first electrically conductive electrode, a second electrically conductive electrode; and at least one metadielectric layer located between the first and second conductive electrodes. The metadielectric layer comprises at least one organic compound with at least one electrically resistive substituent and at least one polarizable unit. The polarizable unit is selected from intramolecular and intermolecular polarizable units. The organic compound is selected from the list comprising compounds with rigid electro-polarizable organic units, composite organic polarizable compounds, composite electro-polarizable organic compounds, composite non-linear electro-polarizable compounds, Sharp polymers, Furuta co-polymers, para-Furuta polymers, YanLi polymers, and any combination thereof.

FLOW REGULATING SYSTEM FOR SUPPLYING PROPELLANT FLUID TO AN ELECTRIC THRUSTER OF A SPACE VEHICLE

A system for regulating the flow rate of a propellant fluid for an electrical thruster of a space vehicle, the vehicle including a tank of propellant fluid and a flow rate regulator connected to the outlet of said tank; 1. A system for regulating the flow rate of a propellant fluid for an electrical thruster of a space vehicle , the vehicle including a tank of propellant fluid and a flow rate regulator connected to the outlet of said tank;the flow rate regulator including a heater element controlled by a computer and adapted to heat the propellant fluid and to modify its physical properties so as to vary the flow rate of propellant fluid leaving the tank;wherein the computer includes a storage memory having loaded therein a plurality of empirical calibration curves that have been determined empirically for defining the flow rate of propellant fluid as a function of the magnitude of heating and as a function of environmental parameters, so that said computer also performs a function of determining the flow rate of the propellant fluid.2. A system according to claim 1 , wherein the empirical calibration curves are determined during ground testing of said regulator system under various environmental parameters.3. A system according to claim 1 , wherein said computer has a plurality of semi-emprirical calibration curves are calculated on the basis of said empirical calibration curves claim 1 , said semi-emprirical calibration curves defining the propellant fluid flow rate as a function of the magnitude of heating for environmental parameters that are different from those of the empirical calibration curves.4. A system according to claim 1 , wherein said computer is configured to use said empirical calibration curves to calculate a semi-emprirical calibration curve defining the flow rate of propellant fluid as a function of the magnitude of heating and of environmental parameters.5. A system according to claim 1 , wherein said heater element is a thermocapillary tube ...

External Discharge Hall Thruster

The invention is a Hall thruster that does not have any discharge channel, and magnetic pole piece. The Hall thruster utilizes permanent magnets to produce magnetic field with strong radial component in front of an annular anode, and expands propellant directly into vacuum through the anode acting also as a gas distributor. The invention reduces mass and complexity of conventional Hall thrusters, and offers a radical solution to discharge channel and magnetic pole piece erosion problem. 1. A Hall thruster comprising:a magnetic circuit comprising inner and outer permanent cylindrical hollow magnets concentrically aligned;an annular anode located inside the magnetic circuit, extending outside the magnetic circuit, and having plurality of holes or slots arranged azimuthally;an inner and/or outer annular insulation walls placed between the anode and the magnetic circuit;a gas distributor adjacent to the anode to inject an ionizable gas;a cathode near the anode, the anode and cathode configured to generate an electric field;2. A Hall thruster as recited in claim 1 , wherein a tube fitting delivering ionizable gas to the gas distributor is concentrically aligned with the anode;3. A Hall thruster as recited in claim 1 , wherein peak radial magnetic field strength within the anode holes or slots or cavity is larger than 600 Gauss.4. A Hall thruster as recited in claim 1 , wherein the magnetic circuit comprises permanent magnets arranged azimuthally to produce magnetic field lines similar to cylindrical hollow magnets.5. A Hall thruster as recited in claim 1 , wherein heat insulation walls are partly or fully removed.6. A Hall thruster as recited in claim 1 , wherein a planar wall claim 1 , having plurality of holes or slots arranged azimuthally is placed between the rear surface of the anode and the front surface of the magnetic circuit claim 1 , and separates the anode completely from the permanent magnets and thruster body.7. The Hall thruster of claim 1 , wherein the ...

Propulsion Boost System and Methods by Enhancing Plasma Thrust via Wake-Field Acceleration

A propulsion system involving a boost feature comprising a stationary electrical conductor, the boost feature configured to couple with a combustion engine, the stationary electrical conductor disposed in a path of a moving high-velocity plasma of exhaust from the combustion engine, and the stationary electrical conductor electrically biased, whereby the moving high-velocity plasma is accelerated, and whereby propulsion is boosted. 1. A propulsion system , comprising:a boost feature including a stationary electrical conductor, wherein the boost feature is configured to couple with a combustion engine and the stationary electrical conductor is disposed in a path of a moving high-velocity plasma of exhaust from the combustion engine, wherein the stationary electrical conductor is negatively electrically bias to repulse electrons and attract positively charged ions,whereby the moving high-velocity plasma is accelerated, andwhereby propulsion is boosted.2. The propulsion system of claim 1 , wherein the stationary electrical conductor is perpendicularly disposed in relation to the path of the moving high-velocity plasma.3. The propulsion system of claim 1 , wherein the stationary electrical conductor comprises one or more wires claim 1 , wherein the wires have a thickness ranging from 0.025 mm to 5 mm.4. (canceled)5. The propulsion system of claim 1 , wherein the stationary electrical conductor comprises tungsten.6. (canceled)7. The propulsion system of claim 1 , wherein the stationary electrical conductor is disposed at a location corresponding to a maximum exhaust velocity in the combustion engine.8. The propulsion system of claim 1 , further comprising the combustion engine claim 1 , wherein the combustion engine comprises one of a jet engine and a rocket engine.9. The propulsion system of claim 1 , wherein the stationary electrical conductor is negatively electrically biased to generate an acceleration of the moving high-velocity plasma claim 1 , whereby a wake-field ...

PLASMA ACTUATING PROPULSION SYSTEM FOR AERIAL VEHICLES

A plasma propulsion nozzle incorporates a cylinder having an inlet and an outlet. A plurality of substantially cylindrical planarly disbanded electrodes with sandwiched dielectric spacers is cascaded in an array to be concentrically expanding from the inlet through an interior chamber to the outlet for a nozzle. A voltage source applies aperiodic signal with rapidly reversing polarity to the electrodes with differential phase applied to adjacent electrodes in the array creating and expelling plasma clusters at each dielectric spacer inducing flow from the nozzle outlet to produce thrust. 1. A plasma propulsion nozzle comprising:a cylinder having an inlet and an outlet;a plurality of substantially cylindrical planarly disbanded electrodes with sandwiched dielectric spacers cascaded in an array to be concentrically expanding from the inlet through an interior chamber to the outlet for a nozzle; and,a voltage source applying a periodic signal with rapidly reversing polarity to the electrodes with differential phase applied to adjacent electrodes in the array creating and expelling plasma clusters at each dielectric spacer inducing flow from the nozzle outlet to produce thrust.2. The plasma propulsion nozzle of wherein the electrodes are between 200 microns and 1 millimeter in thickness.3. The plasma propulsion nozzle of wherein the dielectric spacers are between 20 to 200 microns in thickness.4. The plasma propulsion nozzle of wherein the periodic signal ov the voltage source is between 0.01-30 kHz.5. The plasma propulsion nozzle of wherein the voltage source supplies between 200 volts and 2 kV.6. A method for thrust generation with a plasma propulsion nozzle having an array of substantially cylindrical planarly disbanded electrodes with sandwiched dielectric spacers cascaded to be concentrically expanding from an inlet through an interior chamber to an outlet for a nozzle claim 1 , comprising:introducing air through the inlet into the interior chamber of the nozzle; ...

PROPELLANT DELIVERY SYSTEM, ELECTRIC THRUSTER, AND METHOD OF OPERATING AN ELECTRIC THRUSTER

An electric thruster comprises a propellant delivery system, wherein the propellant delivery system comprises: a pipe for carrying propellant; a valve which is adapted to adjust a volume or mass flow of the propellant in the pipe; and an expansion actuator which is adapted to actuate the valve for adjusting the volume or mass flow of the propellant. The electric thruster further comprises at least one tank which is adapted to receive propellant for the electric thruster; and a discharge chamber. The at least one tank thereby at least partially encloses an end of the discharge chamber and/or an element thermally coupled with the discharge chamber, and the valve of the propellant delivery system is arranged between the tank and the end of the discharge chamber. 114-. (canceled)15. An electric thruster of an aerospace system , comprising:a propellant delivery system;at least one tank which is adapted to receive propellant for the electric thruster; anda discharge chamber, a pipe for carrying propellant,', 'a valve which is adapted to adjust a volume or mass flow of the propellant in the pipe, and', 'an expansion actuator which is adapted to actuate the valve for adjusting the volume or mass flow of the propellant,, 'wherein the propellant delivery system compriseswherein the at least one tank at least partially encloses at least one of an end of the discharge chamber or an element thermally coupled with the discharge chamber, andwherein the valve of the propellant delivery system is arranged between the tank and the end of the discharge chamber.16. The electric thruster according to claim 15 , wherein at least one of:the element thermally coupled with the discharge chamber is a thermally conducting pipe for carrying propellant to the discharge chamber or a portion of the pipe of the propellant delivery system, or an anode; and', 'a thermally conducting insulator which is coupled with the anode,, 'the electric thruster further compriseswherein the element thermally ...

ASTEROID REDIRECTION FACILITATED BY COSMIC RAY AND MUON-CATALYZED FUSION

Asteroid redirection systems are provided that use cosmic ray and muon-catalyzed micro-fusion. These systems include a micro-fusion propulsion system providing thrust for redirecting an asteroid, as well as micro-fusion electrical generation powering an ion drive. The systems deploy deuterium-containing fuel material as a localized cloud interacting with incoming ambient cosmic rays to generate energetic fusion products. Dust or other particulate matter in the fuel material converts some cosmic rays into muons that also catalyze fusion. The fusion products provide thrusting upon the asteroid, or when produced near turbines facilitates electrical generation, which can then power an ion drive. 1. A propulsion system for asteroids , comprising:one or more engines attachable to an asteroid, each engine having a store of deuterium-containing particle fuel material and means for directing the particle fuel material outward therefrom, the material interacting with an ambient flux of cosmic rays to generate products having kinetic energy;a controller coupled to all of the engines for coordinating the directing of particle fuel material such that the kinetic-energy-containing products generate thrust for the asteroid in a desired direction.2. The propulsion system as in claim 1 , wherein the deuterium-containing particle fuel material comprises LiD.3. The propulsion system as in claim 1 , wherein the deuterium-containing particle fuel material comprises DO.4. The propulsion system as in claim 1 , wherein the deuterium-containing particle fuel material comprises D.5. The propulsion system as in claim 1 , wherein the deuterium-containing particle fuel material is in solid powder form.6. The propulsion system as in claim 1 , wherein the deuterium-containing particle fuel material is in pellet form.7. The propulsion system as in claim 1 , wherein the deuterium-containing particle fuel material is in frozen form.8. The propulsion system as in claim 1 , wherein the deuterium- ...

MAGNETIC GAS ENGINE AND METHOD OF EXTRACTING WORK

The present subject matter overcomes the deficiencies in the prior art by introducing or generating charged particles in an air stream and manipulating the air stream with magnetic fields operating on the charged particles. Embodiments of the present subject mater compress the air stream by accelerating charged particles with a moving magnetic field, where the magnetic field has a velocity perpendicular to its flux lines. The increased velocity of the charged particles increases the statistical mean particle velocity and thereby increases the pressure in the air stream. The compressed air stream is then heated and expanded through a second magnetic field. The expansion of the air stream substantially increases the velocity of the air stream and the charged particles therein. The interaction of the high velocity charged particles and the magnetic field imparts a force perpendicular to the flux lines, this force powers the movement of the magnetic field. 18.-. (canceled)9. A jet engine for providing thrust across the subsonic to supersonic regimes comprising: a duct having an inlet and an exit; a magnetic field device for providing a rotating magnetic field about an axis , said magnetic field defined by a first set of magnetic flux lines proximate to said inlet and a second set of magnetic flux lines proximate to the exit; a combustion chamber within said duct and between said inlet and said exist; wherein an ionized gas stream having a net charge enters said duct via said inlet through the first set of magnetic flux lines , passes into the combustion chamber and exits said duct through the second set of magnetic flux lines and said exit; further comprising an ionized gas generator , wherein said generator is turned off above a predetermined Mach no.10. A jet engine for providing thrust across the subsonic to supersonic regimes comprising: a duct having an inlet and an exit; a magnetic field device for providing a rotating magnetic field about an axis , said magnetic ...

ELECTRICAL/CHEMICAL THRUSTER USING THE SAME MONOPROPELLANT AND METHOD

A thruster operable in a chemical mode or in an electrospray mode using the same liquid monopropellant for operation in both modes is described having a multiplicity of a microthrusters made of a catalytic material having a bore therethrough, where, when operated in the chemical mode, the microthrusters are heated to decompose the monopropellant the monopropellant flows therethrough to generate relatively high thrust. An extractor is positioned downstream of the outlet ends of the microthrusters, such that when the system is operated in its electrospray mode the flowrate of the monopropellant through the microthrusters is substantially lower than in the chemical mode and the extractor is energized with an electric field so that ions and droplets are discharged from the microthrusters and accelerated so as to yield a relatively high specific impulse. 1. An electrical/chemical thruster for a spacecraft operable in a chemical mode and in an electrospray mode that utilizes a liquid monopropellant for operation in both of said modes , said thruster having a plurality of microthrusters , each said microthruster comprising a microtube having a bore therethrough and having an inlet end and an outlet end with said inlet ends of said microtubes being in communication with a supply of said monopropellant , each said microtube being of a catalytic material capable of being heated to a preheat temperature sufficient to substantially decompose said monopropellant as said monopropellant flows therethrough at a first flowrate so as to generate relatively high thrust when operated in said chemical mode as compared to when operated in said electrospray mode , and wherein said thruster further includes a extractor proximate the outlet ends of said microtubes , whereby when said thruster is operated in its electrospray mode with said extractor being energized with an electric field so that as said monopropellant flows through said microtubes at a second flowrate substantially less than ...

METHOD AND SYSTEM FOR A PROGRAMMABLE AND FAULT TOLERANT PULSED PLASMA THRUSTER

A system and method provides a fault-tolerant multi-channel pulsed plasma thruster system utilizing a control unit and an embedded real time application manipulating low-level timing events with programming, with clear examples of completely flexible control techniques of a scalable micropropulsion system having many pulsed plasma thruster channels, taking into account system aging behavior and specific mission utilization requirements that may change in the mission lifetime. The system and method also covers an architecture lending itself suitable for design of a dedicated FGPA or ASIC that would tightly integrate many channels of thruster components to build a robust, resilient and versatile micropropulsion subsystem for space applications, and indirectly for advanced multi-channel spacecraft instrumentation. 1. A method for controlling trigger pulse generation in a pulsed plasma thruster system , the method comprising: generating at a processing device , independent event markers in time-units , and controlling by the event markers a Trigger Pulse activation event , Trigger Pulse deactivation event , Magnetic Coil activation event , Magnetic Coil deactivation event , End of Cycle signal event , and a spacecraft related event.2. The method of claim 1 , further comprising generating time-slices at regular intervals claim 1 , and generating the event markers for each event at a predetermined number of occurrences of the time-slices.3. The method of claim 2 , wherein each of the events is triggered at a different predetermined number of occurrences of the time-slices.4. The method of claim 1 , further comprising:generating at the processing device, a time-slice count;assigning at the processing device, a unique count value to each of the events;triggering at the processing device, an event when the time-slice count equals the unique count value for that event; and,incrementing at the processing device, the time-slice count.5. The method of claim 1 , wherein operation ...

MPD THRUSTER THAT ACCELERATES ELECTRODELESS PLASMA AND ELECTRODELESS PLASMA ACCELERATING METHOD USING MPD THRUSTER

Supplying electrodeless plasma to a space (S) between a cathode () and an anode (), the resistivity in the space is downed. The electrodeless plasma is accelerated with Lorentz force induced by a radial direction magnetic field component (By) and an axial direction magnetic field component (Bx) that are generated in the space (S), and current (Iac) which flows through the space (S). 1. An MPD thruster comprising:an electrodeless plasma generating device configured to generate electrodeless plasma from propellant;an accelerating device configured to accelerate the electrodeless plasma; anda supply passage configured to supply the generated electrodeless plasma to the accelerating device,wherein the accelerating device comprises:a magnetic coil;a cathode;an anode; anda voltage applying unit configured to apply a voltage between the cathode and the anode,wherein the supply passage supplies the electrodeless plasma to a space between the cathode and the anode,wherein the magnetic coil generates an axial direction magnetic field component along a central axis direction of the magnetic coil and a radial direction magnetic field component orthogonal to the center axis in the space,wherein the voltage applying unit generates a current in the space, andwherein the electrodeless plasma supplied to the space is accelerated with Lorentz force induced by the axial direction magnetic field component, the radial direction magnetic field component, and the current.2. The MPD thruster according to claim 1 , wherein a distance between the supply passage and the center axis of the magnetic coil is larger than a distance between the cathode and the center axis and is smaller than a distance between the anode and the center axis.3. The MPD thruster according to claim 1 , wherein the cathode is arranged along the center axis of the magnetic coil.4. The MPD thruster according to claim 1 , wherein the electrodeless plasma generating device comprises an antenna arranged around the supply ...

HALL EFFECT THRUSTER

A steerable-thrust Hall effect thruster in which a final stage of a magnetic circuit includes an inner pole and a facing outer pole, the inner pole being offset axially downstream relative to the outer pole, so that a magnetic field is inclined relative to a transverse plane of the thruster. 110-. (canceled)11. A steerable-thrust Hall effect thruster comprising:an annular channel defined by two concentric walls with a central axis, the annular channel including an open end and a closed end, and including an upstream segment beside the closed end that is subdivided by radial walls into a plurality of separate compartments;an anode situated at the closed end of the annular channel;an injection circuit for injecting propulsion gas into the compartments of the annular channel, the circuit including at least one individual flow rate regulator device for each compartment;a magnetic circuit for generating a magnetic field at the open end of the annular channel and including at least a final stage with an inner pole and a facing outer pole, wherein the inner pole is offset axially downstream relative to the outer pole, so that the magnetic field is inclined relative to a transverse plane of the thruster; anda cathode downstream from the open end of the annular channel.12. The Hall effect thruster according to claim 11 , wherein the annular channel includes a downstream segment beside the open end with a meridian plane that diverges in a downstream direction.13. The Hall effect thruster according to claim 11 , wherein the annular channel is not axisymmetric.14. The Hall effect thruster according to claim 13 , wherein the annular channel presents a cross-section with a main axis of symmetry and a secondary axis of symmetry that is perpendicular to and shorter than the main axis of symmetry.15. The Hall effect thruster according to claim 11 , wherein the at least one individual flow rate regulator device is connected to a control unit.16. The Hall effect thruster according to ...

Fuel-Free Spacecraft Propelling System Based on Spatial Atomic Oxygen and Propelling Method

A fuel-free spacecraft propelling system having an open-ended outer cylinder of a propelling device and an atomic oxygen collecting device is disclosed. The latter is arranged at the forwardly-propelled front end of the outer cylinder and is hermetically connected with an RF generating device and an ion cyclotron wave heating device through a magnetic confinement device. A spiral wave discharge oxygen plasma inlet and a spiral wave discharge oxygen plasma outlet in the ion cyclotron wave heating device are respectively provided with another magnetic confinement device. The propulsion of the invention does not need to carry the propellant, which greatly reduces the launch costs, and enables a spacecraft to advantageously have an increased orbit life over existing spacecraft systems. 1. A fuel-free spacecraft propulsion system based on space atomic oxygen , said fuel-free spacecraft propulsion system comprises an outer cylinder of the propulsion device with both open ends , an atomic oxygen collecting device inside the outer cylinder of the propulsion device , an RF generating device and an ion cyclotron wave heating device , wherein the atomic oxygen collecting device is arranged at the front end of the outer cylinder of the propulsion device for propelling forwardly , and said atomic oxygen collecting device is hermetically connected with the RF generating device and the ion cyclotron wave heating device through a magnetic confinement device located in the ion cyclotron wave heating device , wherein an inlet and an outlet of a spiral wave discharge oxygen plasma in the ion cyclotron wave heating device are respectively provided with another magnetic confinement device , wherein the atomic oxygen collecting device pressurizes the space atomic oxygen entering the front end of the outer cylinder of the propulsion device while the spacecraft is moving forward , the pressurized space atomic oxygen is ionized in a spiral wave discharge mode in the RF generation device , ...

Method and device for electric satellite propulsion

An electric propulsion device for a satellite equipped with at least four active thrusters exerting a parallel thrust upon a transfer, the device comprises means for detecting a thruster failure and means for reorienting the thrusters, and comprises means for computing a reorientation angle of the thrusters remaining active upon a failure of a thruster, the value of the angle being computed to reorient at least two of the remaining thrusters in order to cancel the total torque about the center of mass of the satellite.

ELECTRODELESS PLASMA THRUSTER

A plasma propulsion system with no internal electrodes is described. Gas is flowed into an insulated axisymmetric plasma liner. A radio frequency antenna generates an inductive or helicon plasma discharge within the liner. The plasma is accelerated through a converging/diverging magnetic field out of the liner, generating thrust. 1. A plasma thruster for providing a motive force , said plasma thruster comprising:a propellant tank containing a propellant;a plasma discharge system being operably coupled to said propellant tank for receiving said propellant and outputting an accelerated plasma;a magnet system outputting magnetic field lines along at least a portion of said plasma discharge system, said magnet system operable to define a converging section, a throat section downstream of said converging section having a flow boundary smaller than a flow boundary of said converging section, and a diverging section downstream of said throat section having a flow boundary larger than said flow boundary of said throat section, said converging section, throat section, and diverging section collectively receiving and accelerating said propellant through said plasma discharge system; anda radio frequency (RF) source outputting an RF field along at least a portion of said plasma discharge system, said RF field ionizing said propellant to a plasma.2. The plasma thruster according to wherein said magnet system comprises a plurality of permanent magnets.3. The plasma thruster according to wherein said magnet system comprises at least one electromagnetic coil.4. The plasma thruster according to wherein said propellant is a gaseous propellant.5. The plasma thruster according to wherein said propellant is a liquid propellant.6. The plasma thruster according to wherein said propellant is a solid propellant.7. The plasma thruster according to wherein said RF source comprises a loop RF source.8. The plasma thruster according to wherein said RF source comprises a Boswell-type RF source.9 ...

ELECTROTHERMAL DEVICE FOR A PROPULSION SYSTEM, ESPECIALLY FOR A TURBOJET, PROPULSION SYSTEM COMPRISING SUCH AN ELECTROTHERMAL DEVICE, AND ASSOCIATED METHOD

An electrothermal device () includes a primary chamber () having an anode nozzle () provided with an inlet passage (), a cathode tip () at least partially inserted into the inlet passage (), and a primary air inlet () leading into the inlet passage (), and a voltage generator () arranged between the anode nozzle () and the cathode tip () in such a way as to generate an electric arc () on the path of the primary air flow () injected into the primary chamber (). It includes a secondary chamber () wherein a secondary air flow () circulates in a heat exchange relation with the heated primary air flow () from the primary chamber (), the secondary air flow () having a lower temperature than the heated primary air flow () leaving the primary chamber (). 11100267971071169121323314215142. An electrothermal device ( , ) for a propulsion system , including a primary chamber () including an anode nozzle () provided with an inlet passage () , a cathode spike () at least partially inserted into the inlet passage () , a primary air inlet () leading into the inlet passage () , and a voltage generator () disposed between the anode nozzle () and the cathode spike () in such a way as to generate an electric arc () on the path of the primary air flow () injected into the primary chamber () , characterized in that it includes a secondary chamber () in which a secondary air flow () circulates in a heat exchange relationship with the heated primary air flow () from the primary chamber () , the secondary air flow () having a lower temperature than the heated primary air flow () at the outlet of the primary chamber ().211001315. The device ( claim 1 , ) as claimed in claim 1 , including means for separating a compressed air flow into a primary air flow () and a secondary air flow ().3110032. The device ( claim 1 , ) as claimed in claim 1 , wherein the secondary chamber () includes the primary chamber ().4. The device as claimed in claim 1 , characterized in that the primary chamber extends ...

DEPLOYABLE GRIDDED ION THRUSTER

Various examples related to a deployable gridded ion thruster are described. A deployable gridded ion thruster can include: a thruster body including an ion generating unit; and an expandable discharge chamber configured to expand from a stored configuration to a deployed configuration. The expandable discharge chamber can include a chamber wall having a first geometric shape compressed within the thruster body when in the stored configuration and a second geometric shape expanded outward from the thruster body when in the deployed configuration. Also described herein are methods of operation for a deployable gridded ion thruster. 1. A deployable gridded ion thruster , comprising:a thruster body including an ion generating unit; andan expandable discharge chamber configured to expand from a stored configuration to a deployed configuration, where the expandable discharge chamber comprises a chamber wall having a first geometric shape compressed within the thruster body when in the stored configuration and a second geometric shape expanded outward from the thruster body when in the deployed configuration.2. The deployable gridded ion thruster of claim 1 , wherein the expandable discharge chamber is further configured to retract from the second geometric shape to the first geometric shape.3. The deployable gridded ion thruster of claim 2 , wherein the expandable discharge chamber is deployed by burning a wire.4. The deployable gridded ion thruster of claim 2 , wherein the expandable discharge chamber is deployed by releasing a latch.5. The deployable gridded ion thruster of claim 1 , wherein the second geometric shape in the deployed configuration has a large volume to surface area ratio inside the expandable discharge chamber.6. The deployable gridded ion thruster of claim 1 , wherein the expandable discharge chamber can be compressed from the second geometric shape to the first geometric shape while in orbit.7. The deployable gridded ion thruster of claim 1 , wherein ...

SPACECRAFT PROPULSION SYSTEM AND METHOD

A space propulsion system includes an electrostatic thruster with a first electrical load; a resistojet; a propellant fluid feed circuit; and an electrical power supply circuit including a first power supply line and a first switch for selecting between connecting the first power supply line to the resistojet and connecting the first power supply line to the first electrical load of the electrostatic thruster. The propulsion system thus enables a space propulsion method to be applied that includes a switching step for selecting a first propulsion mode in which the resistojet is activated, or a second propulsion mode in which the electrostatic thruster is activated. 1. A space propulsion system comprising:an electrostatic thruster with at least a first electrical load;a resistojet;a propellant fluid feed circuit; andan electrical power supply circuit comprising at least a first power supply line and a first switch for selecting between connecting said first power supply line to the resistojet and connecting said first power supply line to said first electrical load of the electrostatic thruster.2. The space propulsion system according to claim 1 , wherein said first electrical load comprises a heater element of an emitter cathode of said electrostatic thruster.3. The space propulsion system according to claim 2 , wherein said first switch serves to select between connecting said first power supply line claim 2 , without any current or voltage conversion or transformation claim 2 , to a resistor forming a heater element of the resistojet claim 2 , and connecting said first power supply line claim 2 , without any current or voltage conversion or transformation claim 2 , to a resistor forming the heater element of the emitter cathode of said electrostatic thruster.4. The space propulsion system according to claim 1 , wherein said propellant fluid feed circuit includes at least one valve for feeding the electrostatic thruster and at least one valve for feeding the ...

Structural Propellant for Ion Rockets (SPIR)

Systems, methods, and apparatus for a structural propellant for ion rockets (SPIR) are disclosed. In one or more embodiments, a method for in-space propulsion of a spacecraft involves removing, by a removal device, a portion of a structure of the spacecraft. The method further involves feeding, by the removal device, the portion into a Hall thruster system. Further, the method involves utilizing, by the Hall thruster system, the portion as propellant to produce thrust. In one or more embodiments, the structure is an upper stage of the spacecraft. In at least one embodiment, the upper stage comprises at least one structural support and/or at least one upper stage housing. In some embodiments, the structure is manufactured from magnesium, bismuth, zinc, and/or indium. 1. A method for in-space propulsion of a spacecraft , the method comprising:removing, by a removal device, a portion of a structure of the spacecraft;feeding, by the removal device, the portion into a Hall thruster system; andutilizing, by the Hall thruster system, the portion as propellant to produce thrust.2. The method of claim 1 , wherein the removal device comprises at least one of at least one mechanical cutter and at least one laser cutter.3. The method of claim 2 , wherein the at least one mechanical cutter comprises at least one of at least one cutting wheel and at least one scissors mechanism.4. The method of claim 1 , wherein the structure is an upper stage of the spacecraft.5. The method of claim 4 , wherein the upper stage comprises at least one of at least one structural support and at least one upper stage housing.6. The method of claim 1 , wherein the structure is manufactured from at least one of magnesium claim 1 , bismuth claim 1 , zinc claim 1 , and indium.7. The method of claim 1 , wherein the portion of the structure is removed by the removal device in the form of a ribbon.8. The method of claim 1 , wherein the removal device is tethered to the spacecraft via at least one cord.9. ...

12CaO-7Al2O3 ELECTRIDE HOLLOW CATHODE

The use of the electride form of 12CaO-7Al 2 O 3 , or C12A7, as a low work function electron emitter in a hollow cathode discharge apparatus is described. No heater is required to initiate operation of the present cathode, as is necessary for traditional hollow cathode devices. Because C12A7 has a fully oxidized lattice structure, exposure to oxygen does not degrade the electride. The electride was surrounded by a graphite liner since it was found that the C12A7 electride converts to it's eutectic (CA+C3A) form when heated (through natural hollow cathode operation) in a metal tube.

Hybrid electric propulsion for spacecraft

A propulsion system for spacecraft is based on an electric engine that expels propellant to achieve thrust. The propellant is first ionized to generate a plasma. Plasma particles are selectively accelerated via a pulsed laser that accelerates predominantly the electrons in the plasma. The electrons are expelled first, forming a space charge that acts as a virtual cathode to accelerate the positive ions. Interactions between the laser beam and plasma electrons are predominantly through the ponderomotive force.

THRUSTER GRID CLEAR CIRCUITS AND METHODS TO CLEAR THRUSTER GRIDS

Thruster grid clear circuits and methods to clear thruster grids are disclosed. An example apparatus includes a low voltage grid clear circuit to apply first energy to a grid at a first voltage, and a high voltage grid clear circuit to detect a failure of the applied energy to clear a short circuit condition of the grid and to apply second energy to the grid at a second voltage higher than the first voltage. 1. An apparatus , comprising:a low voltage grid clear circuit to apply first energy to a grid at a first voltage; anda high voltage grid clear circuit to apply second energy to the grid at a second voltage higher than the first voltage.2. An apparatus as defined in claim 1 , wherein the grid comprises an ion propulsion system thruster grid.3. An apparatus as defined in claim 1 , wherein the high voltage grid clear circuit is to detect a short circuit condition in the grid by:measuring a first node voltage in the low voltage grid clear circuit; andcomparing the first node voltage to a threshold.4. An apparatus as defined in claim 3 , wherein the high voltage grid clear circuit is to detect a failure of the first energy to clear the short circuit condition of the grid by:measuring a second node voltage in the low voltage grid clear circuit at a different node than the first node voltage; andcomparing a difference between the second node voltage and the first node voltage to a threshold, the high voltage grid clear circuit to apply the second energy in response to detecting the failure.5. An apparatus as defined in claim 1 , wherein the high voltage grid clear circuit is to:detect that the low voltage grid clear circuit is applying the first energy to the grid: andstop applying the second energy in response to detecting that the low voltage grid clear circuit is applying the first energy.6. An apparatus as defined in claim 5 , wherein the high voltage grid clear circuit is to detect that the low voltage grid clear circuit is applying the first energy by detecting a ...

GRIDDED ION THRUSTER WITH INTEGRATED SOLID PROPELLANT

The invention relates to an ion thruster, comprising: 3. Thruster according to claim 1 , wherein the set of means for forming the ion-electron plasma comprises:at least one coil powered by a radiofrequency AC voltage source different from the radiofrequency DC or AC voltage source connected to the means for extracting and accelerating; orat least one microwave antenna powered by a microwave AC voltage source.4. Thruster as claimed in claim 3 , wherein the voltage source connected to the means for extracting and accelerating is a radiofrequency AC voltage source claim 3 , in order to form claim 3 , at the output of the chamber claim 3 , a beam of ions and of electrons.5. Thruster according to claim 2 , wherein claim 2 , when the means for extracting and accelerating is a set of at least two grids located at one end (E) of the chamber claim 2 , the electroneutrality of the beam of ions and of electrons is obtained at least partially by adjusting the application duration of the positive and/or negative potentials coming from the radiofrequency AC voltage source connected to the means for extracting and accelerating.6. Thruster according to claim 2 , wherein claim 2 , when the means for extracting and accelerating is a set of at least two grids located at one end (E) of the chamber claim 2 , the electroneutrality of the beam of ions and of electrons is obtained at least partially by adjusting the amplitude of the positive and/or negative potentials coming from the radiofrequency AC voltage source connected to the means for extracting and accelerating.7. Thruster according to claim 3 , wherein the voltage source connected to the means for extracting and accelerating is a DC voltage source claim 3 , in order to form claim 3 , at the output of the chamber claim 3 , a beam of ions claim 3 , with the thruster further comprising means for injecting electrons into said beam of ions in order to provide electroneutrality.8. Thruster according to claim 1 , wherein the reservoir ...

HALL-EFFECT THRUSTER USABLE AT HIGH ALTITUDE

A Hall-effect thruster (), configured to be arranged inside or outside a spacecraft. 1. A Hall-effect thruster for developing thrust along a thrust axis , the thruster comprising:a channel allowing collection, acceleration and ejection of particles by the thruster when it is in operation, the channel being delimited radially by an inner wall and an outer wall;an electric circuit comprising an anode, a cathode and an electrical voltage source to emit electrons via the cathode and to attract electrons via the anode;a magnetic circuit for generating a magnetic field in the channel axially downstream of the anode, the magnetic field being directed in a substantially radial direction with respect to the thrust axis;whereinthe channel is open on an upstream side of the thruster and including a particle concentrator for collecting particles;the shape of the concentrator is defined by a continuous contour situated in a plane perpendicular to the thrust axis and surrounding it;on a major portion of the contour, each section of the concentrator perpendicular to the contour has a parabolic shape and has a focus belonging to the contour; andthe magnetic circuit is arranged so as to generate the magnetic field in the vicinity of the contour.2. The thruster according to claim 1 , wherein the magnetic circuit comprises a plurality of connection arms arranged so as to connect the inner wall and the outer wall of the channel claim 1 , the channel having an annular shape.3. A Hall-effect thruster for developing a thrust along a thrust axis claim 1 , the thruster including:a magnetic circuit for generating a magnetic field;an electrical circuit comprising an anode, a first cathode, and an electrical voltage source for emitting electrons at least via the first cathode and attracting electrons via the anode; whereinthe thruster is arranged inside a wall formed around the thrust axis;the magnetic circuit and the electric circuit are arranged so as to generate magnetic and electric fields ...