ТУРБИНА С О ОБЕСПЕЧИВАЮЩЕЙ УПЛОТНЕНИЕ И ЛАМИНАРНОЕ ТЕЧЕНИЕ КОНФИГУРАЦИЕЙ ТРАЕКТОРИИ ПОТОКА

Турбина содержит ротор, на котором закреплено множество рабочих лопаток. Расположенные на расстоянии друг от друга в аксиальном направлении группы сопел имеют аэродинамические поверхности и внутренние и наружные бандажи на их противоположных концах. Рабочие лопатки имеют ласточкины хвосты и площадки вдоль радиально внутренних концов рабочих лопаток. Площадки, аэродинамические поверхности, внутренние и наружные бандажи частично определяют траекторию потока для потока текучей среды через турбину. На ласточкиных хвостах образованы выступающие элементы, проходящие по существу в аксиальном направлении в сторону одной из указанных групп сопел вдоль мест, расположенных радиально внутри по отношению к площадкам. На соплах из одной группы сопел предусмотрены зубцы лабиринтных уплотнений, образующие вместе с выступающими элементами уплотнение, предназначенное для уменьшения потока утечек с траектории потока в проточную часть между одним рабочим колесом и одной группой сопел. Передние края площадок ...

КОРПУС ТУРБИНЫ, ТУРБИНА, РАБОТАЮЩАЯ НА ОТРАБОТАННЫХ ГАЗАХ (ВАРИАНТЫ), И ТУРБОНАГНЕТАТЕЛЬ, СОДЕРЖАЩИЙ УПОМЯНУТУЮ ТУРБИНУ (ВАРИАНТЫ)

Корпус турбины содержит корпус подшипника для размещения вала с возможностью вращения, корпус для впуска газа, опирающийся на корпус подшипника и концентрически охватывающий его на участке опорной поверхности, а также центрирующее кольцо для центрирования корпуса для впуска газа по отношению к расположенному в корпусе подшипника валу. Центрирующее кольцо и либо корпус подшипника, либо корпус для впуска газа содержат входящие друг в друга центрирующие элементы. Центрирующие элементы содержат выемки и либо радиально, либо аксиально равнонаправленные центрирующие кулачки для вхождения в выемки. Между корпусом, не содержащим центрирующие элементы, т.е. либо между корпусом для впуска газа, либо корпусом подшипника и центрирующим кольцом имеется соединение с силовым замыканием. В соответствии с другими изобретениями группы предложена турбина, содержащая корпус турбины, выполненный как указано выше, причем корпус для впуска газа и корпус подшипника отцентрированы по отношению к валу и имеют возможность ...

СПЛАВ НА ОСНОВЕ НИКЕЛЯ

Изобретение относится к сплавам на основе никеля, имеющим желаемое соотношение прочности и сопротивления коррозии и окислению, и может быть использовано для изготовления соплового аппарата газотурбинного двигателя. Сплав состоит по существу из, мас.%: 10-25 кобальта, 20-28 хрома, 1-3 вольфрама, 0,5-1,5 алюминия, 1,5-2,8 титана, 0,8-1,45 ниобия, 0,001-0,025 бора, вплоть до 0,4 циркония, 0,02-0,15 углерода, остальное - по существу никель и случайные примеси. Сплав содержит тантал в количестве, меньшем, чем ниобий, причем Nb+0,508Ta составляет 1,15-1,45, или не содержит тантала. Сплав обладает улучшенной пластичностью и свариваемостью, но без ухудшения литейных свойств. 2 н. и 12 з.п. ф-лы, 2 табл., 6 ил.

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ С СИСТЕМОЙ ВОЗДУШНОГО ОХЛАЖДЕНИЯ ЛОПАТОК ТУРБИНЫ И СПОСОБ ОХЛАЖДЕНИЯ ПОЛОЙ ПРОФИЛЬНОЙ ЧАСТИ ЛОПАТКИ

Газотурбинный двигатель с системой воздушного охлаждения лопаток турбины, в частности профильной части лопаток, содержит проходящий через него основной газовый тракт и средства подачи охлаждающего воздуха от компрессорной секции газотурбинного двигателя к профильной части лопатки. Полки лопаток имеют отверстия, расположенные ниже внутреннего канала по направлению потока газа, для прохождения воздуха, использованного для охлаждения полки направленным потоком, в полость профильной части ниже внутреннего канала с целью увеличения давления в задней зоне профильной части. Осуществление изобретения позволяет улучшить температурный градиент по высоте профильной части лопатки. 2 с. и 1 з.п. ф-лы, 3 ил.

ОХЛАЖДАЮЩЕЕ УСТРОЙСТВО ПРОФИЛЬНОЙ ЧАСТИ ЛОПАТКИ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Охлаждающее устройство профильной части лопатки газотурбинного двигателя содержит стенку с областью передней кромки с наружной искривленной поверхностью, имеющий центр кривизны, расположенный внутри профильной части лопатки. В стенке в области передней кромки выполнено несколько каналов охлаждающего воздуха, образующих систему каналов. Каждый из каналов содержит прямой цилиндрический калиброванный участок с отверстием и диффузорную зону, формирующую выходное отверстие на пересечении с искривленной поверхностью стенки. Диффузорная зона выполнена в виде части конуса. Ось конуса совпадает в основном с осью канала и расположена на части стенки, расположенной ниже по потоку охлаждающего воздуха у выходного отверстия канала. Изобретение позволяет повысить эффективность охлаждения. 5 з.п. ф-лы, 8 ил.

УСТРОЙСТВО ФИКСИРОВАНИЯ ТУРБИННОГО СОПЛА, ВЫПОЛНЕННОЕ НА ГОРИЗОНТАЛЬНОЙ СОЕДИНИТЕЛЬНОЙ ПОВЕРХНОСТИ ДЕРЖАТЕЛЯ (ВАРИАНТЫ)

Изобретение относится к устройствам фиксирования сопел в турбине. Сопла (12) для турбины имеют выполненные в виде соединения «ласточкин хвост» основания (20) для вмещения в них соответствующих выполненных в виде соединений «ласточкин хвост» пазов (27) в половинах держателя турбины. Сопла каждой поверхности из числа горизонтальных соединительных поверхностей каждой половины держателя имеют выемки (34), выполненные вдоль их оснований, имеющих поверхность примыкания. Шпоночные вырезы (38) выполнены в горизонтальных соединительных поверхностях и вмещают в себе шпонки (40), плотно прилегающие к поверхностям примыкания. Шпонки нагартованы на половину держателя на горизонтальных соединительных поверхностях или привинчены к ней. Штифты (32) радиального нагружения контактируют с основаниями концевого сопла, чтобы смещать концевые сопла радиально внутрь, при этом не создавая помехи для шпонок, фиксирующих сопла в пазах. Изобретение обеспечивает улучшение эксплуатационных характеристик за счет того ...

НАПРАВЛЯЮЩЕЕ УСТРОЙСТВО ДЛЯ ПОТОКА ВОЗДУХА НА ВХОДЕ В КАМЕРУ СГОРАНИЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Направляющее устройство для потока воздуха на входе в камеру сгорания газотурбинного двигателя содержит спрямляющий аппарат и расположенный за ним диффузор. Спрямляющий аппарат содержит две коаксиальные обечайки, между которыми размещены лопатки, проходящие по существу в радиальном направлении. Диффузор содержит две коаксиальные стенки, представляющие собой тела вращения и связанные друг с другом при помощи радиальных перегородок. Одна из обечаек спрямляющего аппарата сформирована в виде единой детали с одной представляющей собой тело вращения стенкой диффузора. Другая обечайка спрямляющего аппарата присоединена и закреплена на другой представляющей собой тело вращения стенке диффузора. Лопатки спрямляющего аппарата жестко связаны одним концом с одной обечайкой спрямляющего аппарата и отстоят с небольшим зазором от другой обечайки на другом конце. Изобретение направлено на упрощение технологии изготовления направляющего устройства. 3 н. и 12 з.п. ф-лы, 8 ил.

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ, ВНУТРЕННЯЯ ОБОЛОЧКА КАМЕРЫ СГОРАНИЯ ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ И РОТОРНЫЙ КОЖУХ ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Газотурбинный двигатель включает компрессор, кольцеобразную камеру сгорания и турбину. Камера сгорания в переходной зоне своей оболочкой примыкает к входу в турбину с возможностью обусловленного тепловым расширением относительного движения между камерой сгорания и входом в турбину. Оболочка камеры сгорания своими распределенными по периметру опорными элементами упирается вследствие возникающего в рабочем режиме теплового расширения в конический контур на роторном кожухе, расположенном между выходом компрессора и входной зоной турбины, а также между ротором и внутренней оболочкой камеры сгорания, и опирается на него. Конический контур образует с осью газотурбинного двигателя угол, обеспечивающий скольжение оболочки камеры сгорания опорными элементами по коническому контуру. Другое изобретение группы относится к внутренней оболочке камеры сгорания, которая на выходном конце на обращенной от горячих газов стороне имеет распределенные по периметру опорные элементы со скосом. Скос опорных элементов ...

ОПОРНАЯ СТОЙКА ДЛЯ ДИАФРАГМЫ ТУРБИНЫ, ОПОРНАЯ СТОЕЧНАЯ КОНСТРУКЦИЯ ДЛЯ ДИАФРАГМЫ ТУРБИНЫ И ПАРОВАЯ ТУРБИНА

Опорная стойка (430) для диафрагмы паровой турбины содержит основную вертикальную часть (435) с утолщением (447), которое проходит от указанной части (435) по существу перпендикулярно ей. Утолщение содержит первое утолщение, проходящее от верхнего конца основной вертикальной части (435), и второе утолщение, проходящее от места вблизи нижнего конца основной вертикальной части (435). Первое утолщение отстоит на заданное расстояние от второго утолщения. Первое и второе утолщение предназначены для сопряжения с соответствующим пазом (450), выполненным в диафрагме турбины. В утолщениях имеется отверстие (455), которое проходит сквозь них вертикально и предназначено для размещения крепежного элемента (460), проходящего через первое и второе утолщения (447) с обеспечением вертикального прикрепления основной вертикальной части (435) и указанных утолщений (447) к диафрагме турбины. Достигается уменьшение продолжительности и стоимости цикла технического обслуживания, поскольку обеспечивается возможность ...

СПОСОБ ПРОФИЛИРОВАНИЯ ЗАМЕНЯЮЩЕЙ ЛОПАТКИ В КАЧЕСТВЕ ЗАМЕНЯЮЩЕЙ ЧАСТИ ДЛЯ СТАРОЙ ЛОПАТКИ ДЛЯ ТУРБОМАШИНЫ С ОСЕВЫМ НАПРАВЛЕНИЕМ ПОТОКА

Изобретение относится к способу изготовления заменяющей лопатки для турбомашины. Согласно указанному способу определяют геометрические характеристики контура ступицы и корпуса снабженного старой лопаткой проточного канала, а также осевое положение центра тяжести пера старой лопатки, которая с одной стороны зажата в ступице или в корпусе. Рассчитывают геометрические характеристики пера заменяющей лопатки, таким образом, что перо заменяющей лопатки на своей передней кромке наклоняется в направлении зажатия выше по потоку и обладает прямой стреловидностью. Задают близкую к зажатию область пера заменяющей лопатки, составляющую от 5% до 15% высоты заменяющей лопатки. Смещают участок пера заменяющей лопатки, расположенный за пределами указанной области, выше по потоку, пока осевое положение центра тяжести пера заменяющей лопатки не совпадет с осевым положением центра тяжести пера старой лопатки. В области от расположенной со стороны зажатия линии контура пера заменяющей лопатки до смещенного ...

ДЕРЖАТЕЛЬ ТРУБЫ ДЛЯ УДАЛЕНИЯ ВОЗДУХА В ТУРБОМАШИНЕ

Изобретение относится к держателю трубы для удаления насыщенного маслом воздуха из турбомашины. Держатель (5), предназначенный для удержания трубы для удаления насыщенного маслом воздуха из турбомашины, содержащий радиально внутреннюю кольцевую часть (6), предназначенную для установки вокруг упомянутой трубы, и ребра (11), простирающиеся наружу в радиальной плоскости от кольцевой части (6), образуя угол (α) с радиальным направлением. Ребра (11) содержат зоны крепления (16) к наружной периферии, наклонные в осевом направлении (А) держателя (5) для обеспечения возможности их крепления к реактивному соплу турбомашины. Также представлены турбомашина и её узел. Изобретение позволяет увеличить жесткость держателя без утяжеления его конструкции. 3 н. и 7 з.п. ф-лы, 8 ил.

КОРПУС ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ И ТУРБОРЕАКТИВНЫЙ ДВИГАТЕЛЬ, СОДЕРЖАЩИЙ ТАКИЕ КОРПУСА

Корпус турбореактивного двигателя выполнен с возможностью установки в нем множества лопаток и содержит средства крепления конца каждой лопатки, расположенные на стороне корпуса, противоположной лопаткам. Средства крепления содержат кольцевой элемент, проходящий вокруг корпуса, а корпус содержит отверстия, через которые проходят концы лопаток для их взаимодействия со средствами крепления. Корпус выполнен из длинных волокон, связанных термопластической смолой. Кольцевой элемент получен посредством пултрузии и пропитан термопластической смолой, свариваемой с термопластической смолой корпуса, причем весь узел соединен посредством горячего прессования. Другое изобретение группы относится к турбореактивному двигателю, содержащему указанный выше корпус и множество лопаток, каждая из которых имеет конец, соединенный с корпусом. Группа изобретений позволяет упростить изготовление и сборку корпуса турбореактивного двигателя. 2 н. и 3 з.п. ф-лы, 4 ил.

ВТУЛКА ШАРНИРА ДЛЯ ЛОПАТКИ С РЕГУЛИРУЕМЫМ УГЛОМ УСТАНОВКИ ТУРБОМАШИНЫ, КОЛЬЦО ТУРБОМАШИНЫ, СОДЕРЖАЩЕЕ ТАКИЕ ВТУЛКИ, А ТАКЖЕ КОМПРЕССОР ТУРБОМАШИНЫ И ТУРБОМАШИНА, СОДЕРЖАЩИЕ ТАКОЕ КОЛЬЦО

Втулка шарнира лопатки с регулируемым углом установки, устанавливаемая в выемку в кольце турбомашины, содержит трубчатый корпус вокруг продольной оси и средства для предотвращения вращения в выемке в кольце. Трубчатый корпус втулки имеет расположенный на одном его конце первый конический кольцевой выступ и расположенный на противоположном конце второй конический кольцевой выступ. Конические поверхности, образованные первым и вторым кольцевыми выступами, обращены друг к другу. Средства для предотвращения вращения втулки в выемке в кольце содержат пластину, имеющую, по меньшей мере, одну поверхность для установления по существу плоского контакта с соответствующей поверхностью пластины для предотвращения вращения соседней втулки. Еще одно изобретение группы относится к кольцу турбомашины, содержащему множество выемок для размещения шарнира и множество втулок, выполненных как указано выше. Другие изобретения группы относятся к компрессору турбомашины и турбомашине, каждый из которых содержит ...

ЛОПАТКА ГАЗОТУРБИННОЙ УСТАНОВКИ

Изобретение относится к лопатке газотурбинной установки, содержащей перо, продолжающееся в радиальном направлении от хвостовика лопатки до венца лопатки, определяя размах, равный 0% у хвостовика лопатки и 100% у венца лопатки, и продолжающееся в радиальном направлении от входной кромки до выходной кромки, которая ограничивает хорду осевой длиной хорды, определенной осевой длиной прямой линии, соединяющей входную кромку и выходную кромку пера в зависимости от размаха. Изобретение отличается тем, что осевая длина хорды увеличивается от 80% размаха до 100% размаха. Изгиб входной (9) и выходной (10) кромок определен кривизной линии (12) складывания, которая является линией на корыте (11) пера (1), продолжающейся от 0% размаха до 100% размаха в осевом положении 50%±5% осевой длины (6) хорды, при этом линия (12) складывания изогнута в области между 50%±10% размаха и 100% размаха так, что линия (12) складывания на 100% размаха образует угол α с виртуальной плоскостью (13), ориентированной ортогонально ...

НАПРАВЛЯЮЩИЙ АППАРАТ И ДИАФРАГМА НАПРАВЛЯЮЩЕГО АППАРАТА В ГАЗОТУРБИННОМ ДВИГАТЕЛЕ

Диафрагма направляющего аппарата турбины высокого давления в газотурбинном двигателе имеет внутреннюю сторону, несущую, по меньшей мере, одну направляющую лопатку, и наружную сторону, которая противоположна внутренней стороне и от которой радиально отходит фланец, формирующий со своей передней стороны канал для охлаждающего воздуха и с другой, задней стороны полость. Выходная кромка направляющей лопатки обращена к заднему по потоку краю диафрагмы. Внутренняя сторона диафрагмы снабжена покрытием, нанесенным между выходной кромкой направляющей лопатки и задним краем диафрагмы и образующим тепловой барьер, который позволяет усилить температурный градиент, вызываемый в диафрагме круговым движением воздуха в указанной полости. Изобретение обеспечивает тепловую защиту диафрагмы в той части, где не могут быть использованы другие средства охлаждения. 2 н.и 10 з.п. ф-лы, 5 ил.

СОПЛОВАЯ КОРОБКА ПАРОВОЙ ТУРБИНЫ С РАСПОЛОЖЕНИЕМ ВЫПУСКНЫХ СОПЕЛ НА ОКРУЖНОСТИ В 360° И СПОСОБ ОБРАЗОВАНИЯ СОПЛОВОЙ КОРОБКИ

Изобретение относится к паровым турбинам. Сопловая коробка содержит первую и вторую половины сопловой коробки, которые соединены болтами. Каждая половина сопловой коробки содержит сегмент соплового кольца, который несет сопла на всей его дуге в 180° так, что при соединении половин сопловой коробки образуется сопловая коробка без каких-либо разрывов непрерывности сопел на ее окружности в 360°. Сопла, поддерживаемые на каждом сегменте соплового кольца, сообщаются с впускными отверстиями и соответствующими каналами, которые перпендикулярны выходной плоскости сопловой коробки. Предложен способ изготовления такой сопловой коробки. Изобретение обеспечивает повышение КПД паровых турбин. 2 н. и 10 з.п. ф-лы, 4 ил.

ГАЗОВАЯ ТУРБИНА С КАМЕРОЙ СГОРАНИЯ, ПРИКРЕПЛЕННОЙ К СОПЛОВОМУ АППАРАТУ

Газовая турбина содержит сопловой аппарат турбины высокого давления с неподвижными лопатками, распределенными вокруг оси, совпадающей с осью камеры сгорания, внутреннюю и внешнюю металлические оболочки, а также внутреннюю и внешнюю гибкие соединительные детали. Сопловой аппарат механически соединен с задней концевой частью камеры сгорания и образует с ней единый узел. Узел, образованный камерой сгорания и сопловым аппаратом турбины, расположен между внутренней и внешней металлическими оболочками. Внутренняя и внешняя гибкие соединительные детали соединяют узел, образованный камерой сгорания и сопловым аппаратом турбины, с внутренней и внешней металлическими оболочками для поддержания указанного узла между оболочками. Газовая турбина дополнительно содержит средства блокировки разворота соплового аппарата турбины вокруг своей оси относительно, по меньшей мере, одной из металлических оболочек во избежание передачи усилий, оказываемых на лопатки соплового аппарата турбины газовым потоком, поступающим ...

ОБЛЕГЧЕННАЯ МЕЖДУЛОПАТОЧНАЯ ПЛОЩАДКА ДЛЯ ОПОРНОГОДИСКА ЛОПАТОК ВЕНТИЛЯТОРА ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ И ОПОРНЫЙ ДИСК ЛОПАТОК

Междулопаточная площадка для опорного диска лопаток вентилятора турбореактивного двигателя содержит отражательную часть с нижней стороной, оборудованной первой и второй крепежными лапками. Первая крепежная лапка содержит первое отверстие для прохождения первой крепежной шпильки. Вторая крепежная лапка содержит второе отверстие и третье отверстие для прохождения второй и третьей крепежных шпилек. Крепежные шпильки предназначены для жесткого закрепления крепежных лапок на опорном диске между двумя смежными лопатками. Другое изобретение группы касается опорного диска лопаток, содержащего множество указанных выше междулопаточных площадок, установленных между смежными парами лопаток. Изобретение позволяет упростить механическую обработку междулопаточной площадки и снизить ее вес, а также снизить повреждения лопаток вентилятора при попадании посторонних предметов в проточную часть двигателя. 2 н. и 7 з.п. ф-лы, 3 ил.

УСТРОЙСТВО ДЛЯ ПОВОРОТА РЕГУЛИРУЕМЫХ ЛОПАТОК ТУРБОМАШИНЫ

Устройство для поворота регулируемых лопаток турбомашины, в котором лопатки снабжены цапфой, установленной с возможностью поворота в цилиндрическом радиальном патрубке корпуса турбомашины, и соединены рычагом с регулировочным кольцом, охватывающим корпус. Цапфа, по меньшей мере, некоторых лопаток выступает наружу из корпуса за пределы патрубка и содержит наружный в радиальном направлении конец, установленный с возможностью поворота на неподвижном элементе, расположенном снаружи корпуса. Неподвижный элемент поддерживается и центрируется на корпусе посредством фиксирующих перемычек, равномерно распределенных вокруг продольной оси корпуса. Рычаги поддерживают и центрируют регулировочное кольцо на расстоянии от корпуса и прикреплены к цапфам лопаток между цилиндрическим патрубком и неподвижным элементом. Изобретение позволяет повысить надежность устройства за счет уменьшения изгибающих моментов, воздействующих на цапфы лопаток. 12 з.п. ф-лы, 4 ил.

КОМПРЕССОРНЫЙ УЗЕЛ ДЛЯ ТУРБОМАШИНЫ, ТУРБОМАШИНА И СПОСОБ УПРАВЛЕНИЯ РЕШЕТКОЙ ПРЕДВАРИТЕЛЬНОЙ ЗАКРУТКИ КОМПРЕССОРНОГО УЗЛА

Компрессорный узел турбомашины включает воздухозаборный канал, ступень сжатия воздуха, содержащую подвижное колесо компрессора и решетку предварительной закрутки, расположенную выше по потоку от подвижного колеса компрессора для регулирования скорости воздуха в воздушном потоке на входе подвижного колеса и содержащую множество лопаток с регулируемым углом установки. Шаг между двумя последовательными лопатками решетки превышает хорду одной из двух лопаток на данной высоте воздухозаборного канала в его верхней части. Угол предварительной закрутки воздушного потока лопатками составляет между 80° и 90° в верхней части воздухозаборного канала, когда решетка предварительной закрутки находится в закрытом рабочем положении на низких оборотах ступени сжатия. Другое изобретение группы относится к турбомашине, содержащей указанный компрессорный узел. При управлении решеткой предварительной закрутки компрессорного узла на низких рабочих оборотах ступени сжатия лопатки решетки устанавливают соответственно ...

СПОСОБ ИЗГОТОВЛЕНИЯ КОМПОНЕНТА СТАТОРА (ВАРИАНТЫ)

Способ изготовления компонента статора, предназначенного для направления потока газа, включает изготовление по меньшей мере двух секций кольца, имеющих по меньшей мере одну стенку, и сборку секций друг с другом. Секции собирают друг с другом путем соединения между собой двух смежных стенок, по одной из каждой секции, в элемент, вытянутый в радиальном направлении компонента и предназначенный для направления потока газа и/или передачи возникающих при работе компонента усилий. В другом варианте способа изготовления компонента смежные стенки соединяют между собой накладками, закрывающими зазоры между этими стенками. Изобретение позволяет снизить стоимость и трудоемкость изготовления компонента статора. 2 н. и 17 з.п. ф-лы, 12 ил.

ОХЛАЖДАЕМЫЕ ЛОПАТКИ ГАЗОВЫХ ТУРБИН, СПОСОБ ИЗГОТОВЛЕНИЯ ЛОПАТКИ

Охлаждаемая лопатка газотурбинного двигателя содержит отлитую деталь и продольный рукав, полученный посредством формовки металлического листа. Отлитая деталь включает продольный корпус с продольной полостью с первым и вторым отверстиями на своих концах. Рукав выполнен с возможностью его установки в полости с удерживанием в ней припаиванием мягким припоем или твердым припоем к стенке первого отверстия. Концевой участок этого рукава свободен для скольжения внутрь второго отверстия с образованием направляющей. Концевой участок и направляющая находятся в скользящем контакте относительно друг друга вдоль поверхностей, образованных механической обработкой. Способ изготовления лопатки заключается в изготовлении корпуса лопатки литьем с первым отверстием на одном конце и вторым отверстием на противоположном конце. Второе отверстие содержит стенку с повышенной толщиной для повторной механической обработки. Затем формируют рукав с концевым участком и осуществляют механическую обработку стенки с повышенной ...

УСТРОЙСТВО ДЛЯ УПЛОТНЕНИЯ СТЫКОВ С ПОМОЩЬЮ УПЛОТНЯЮЩИХ ПЛАСТИН

Изобретение относится к уплотнению, применяемому в турбомашинах между аэродинамическим каналом, по которому проходят горячие газы, и объемом, в который подается охлаждающий воздух, например, для охлаждения соплового аппарата. Турбомашина содержит сопловой аппарат, установленный аксиально между двумя корпусными элементами и состоящий из нескольких секторов, содержащих сопловые лопатки, расположенные между элементами внутренних и внешних полок. Элементы внутренних и внешних полок, по меньшей мере, в двух соседних секторах содержат уплотняющие прокладки, расположенные радиально в зазоре, отделяющем указанные элементы от корпусного элемента, для предотвращения протекания текучих сред между указанными элементами. Прокладки двух соседних секторов состыкованы между собой, а их стык закрыт стыковой накладкой, расположенной между указанными прокладками, и удерживающим устройством, удерживающим прокладки в герметизирующем положении. Удерживающее устройство состоит из лапы, расположенной в одном из ...

ВТУЛКА ДЛЯ ШАРНИРА ЛОПАТКИ С РЕГУЛИРУЕМЫМ УГЛОМ УСТАНОВКИ ДЛЯ ТУРБОМАШИНЫ, КОЛЬЦО ТУРБОМАШИНЫ С ТАКОЙ ВТУЛКОЙ, А ТАКЖЕ КОМПРЕССОР ТУРБОМАШИНЫ И ТУРБОМАШИНА С ТАКИМ КОЛЬЦОМ

Втулка для шарнира лопатки с регулируемым углом установки для турбомашины, устанавливаемая в отверстии в кольце турбомашины, представляет собой трубчатое тело, имеющее продольную ось, и содержит, по меньшей мере, три ребра. Форма отверстия в кольце турбомашины соответствует форме втулки, а ребра выступают по радиусу наружу по отношению к продольной оси трубчатого тела, проходят по оси по всей его высоте и равномерно распределены по его периферийной поверхности. Каждое ребро втулки выполнено в виде трубчатого ребра, имеющего по существу овальное поперечное сечение. Другое изобретение группы относится к кольцу турбомашины, имеющему множество отверстий, каждое из которых предназначено для размещения в нем направляющего шарнира лопатки с регулируемым углом установки, и множество указанных выше втулок. Еще два изобретения относятся к компрессору турбомашины и турбомашине, содержащим, по меньшей мере, одно указанное выше кольцо. Изобретения позволяют повысить качество центрирования втулки, работающей ...

СБОРКА ТУРБИННОЙ ЛОПАТКИ ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ И ГАЗОТУРБИННЫЙ СТАТОР, ВКЛЮЧАЮЩИЙ ДАННУЮ СБОРКУ

Сборка турбинной лопатки содержит кожухи сторон низкого и высокого давления, имеющие переднюю кромку, заднюю кромку и выпуклую или вогнутую поверхность между ними, образующую радиальные внешнюю и внутреннюю кромки, а также радиальные внутреннюю и внешнюю платформы. Каждый кожух включает внешний и внутренний фланцы, проходящие от внешней и внутренней кромки соответственно. Радиальная внутренняя платформа содержит множество продольных пазов для размещения радиальных внутренних фланцев кожухов сторон низкого и высокого давления. Радиальная внешняя платформа содержит множество продольных пазов для размещения радиальных внешних фланцев кожухов сторон низкого и высокого давления. Еще одно изобретение относится к газотурбинному статору, содержащему сборку лопатки и радиальные внутреннюю и внешнюю платформы. Сборка лопатки включает лопатку, радиальные внешнюю и внутреннюю торцевые заглушки и зажимной элемент между ними. Лопатка проходит радиально между заглушками и содержит кожухи сторон высокого ...

СТУПЕНЬ КОМПРЕССОРА ОСЕВОЙ ТУРБОМАШИНЫ И ОСЕВАЯ ТУРБОМАШИНА, СОДЕРЖАЩАЯ УКАЗАННУЮ СТУПЕНЬ КОМПРЕССОРА

Изобретение относится к ступени компрессора низкого давления осевой турбомашины, такой как турбореактивный двигатель. Ступень содержит ротор (12), внешняя поверхность которого содержит два кромочных уплотнителя (32), каждый из которых образует радиальное кольцевое ребро; и статор, содержащий кольцевой ряд лопастей (26) статора, проходящих в целом в радиальном направлении; и внутреннюю оболочку (28), радиальное поперечное сечение которой содержит центральную часть (40), соединенную с внутренними краями лопастей (26), боковую часть (42), проходящую с каждой стороны от центральной части до одного из двух кромочных уплотнителей (32), соответственно, образуя, таким образом, ротор с кольцевой полостью. Оболочка и ротор сконфигурированы таким образом, чтобы радиальное сечение кольцевой полости (38) имело длину L1 и высоту Н, где длина L1 превышает высоту Н, что вызывает вращательное движение содержащегося в ней воздуха. Скорость воздуха снижает его давление, что сокращает утечки ниже и выше по ...

ДВУХСТУПЕНЧАТАЯ ВЫСОКОТЕМПЕРАТУРНАЯ ГАЗОВАЯ ТУРБИНА

Двухступенчатая высокотемпературная газовая турбина содержит охлаждаемые сопловые лопатки второй ступени, передний и задний промежуточные диски, расположенные в междисковой полости и образующие лабиринтное уплотнение с нижней полкой сопловой лопатки второй ступени. На входе сопловая лопатка второй ступени соединена с полостью высокого давления и включает сопло, которое выполнено на фланце лабиринта за одно целое с нижней полкой указанной лопатки, связано с внутренней полостью ее пера и расположено на входе в лабиринтное уплотнение. Отношение расстояния от среза сопла до наружной поверхности обода переднего промежуточного диска к диаметру сопла равно 0,5…3. Изобретение позволяет повысить надежность двухступенчатой высокотемпературной газовой турбины путем организации охлаждения ободов первого и второго промежуточных дисков. 2 ил.

ДВУХСТУПЕНЧАТАЯ ТУРБИНА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Двухступенчатая турбина газотурбинного двигателя содержит внутренние полки охлаждаемых сопловых лопаток второй ступени, которые выполнены в форме четырехстенных коробок с контактными площадкам. Междисковая полость в турбине уплотнена передним и задним по потоку воздуха промежуточными дисками, образующими с коробками лабиринтное уплотнение. Уплотнение разделено на переднее между гребешками переднего промежуточного диска и ответной поверхностью передней боковой стенки коробки с увеличенным радиальным зазором Н и на заднее уплотнение. Контактные площадки выполнены на передней по потоку боковой стенке коробки и по всей высоте этой стенки. Передняя полость двуполостной сопловой лопатки на выходе соединена через канал между контактными площадками соседних лопаток со щелевой полостью переднего лабиринтного уплотнения, при этом H/h=2…5, где Н - радиальный зазор по переднему лабиринтному уплотнению, ah- радиальный зазор по заднему лабиринтному уплотнению. Изобретение направлено на повышение надежности ...

СОПЛО ГАЗОВОЙ ТУРБИНЫ С СОПЛАМИ, СФОРМИРОВАННЫМИ КАК ЕДИНОЕ ЦЕЛОЕ

Изобретение относится к соплу газовой турбины. Сопло газовой турбины включает в себя сопла, сформированные как единое целое из торцевой стенки по внутреннему периметру и торцевой стенки по внешнему периметру. Торцевая стенка по внутреннему периметру имеет соединительный участок с верхней по потоку стороны и соединительный участок с нижней по потоку стороны. Соединительный участок с верхней по потоку стороны проходит в радиальном направлении внутрь и соединяется с диафрагмой по внутреннему периметру. Соединительный участок с нижней по потоку стороны, расположенный с нижней по потоку стороны от соединительного участка с верхней по потоку стороны, проходит в радиальном направлении внутрь и соединяется с диафрагмой по внутреннему периметру. На заднем краевом участке торцевой стенки по внутреннему периметру торцевая стенка по внутреннему периметру имеет тонкостенный участок, соответствующий участку заднего краевого участка торцевой стенки по внутреннему периметру уменьшенной толщины. Изобретение ...

Регулируемый сопловой аппарат радиально-осевой турбины с устройством измерения углов поворота лопаток соплового аппарата

Полезная модель относится к машиностроению, а именно к двигателестроению, и может быть использована для исследования влияния изменяемой геометрии проточной части турбины на эффективность работы турбокомпрессора дизельного двигателя транспортных машин. Техническим результатом предложенной полезной модели является точное определение угла поворота лопаток соплового аппарата, а также возможность замены лопаток соплового аппарата на лопатки другого профиля при проведении исследований характеристик турбокомпрессора, оборудованного таким регулируемым сопловым аппаратом. Регулируемый сопловой аппарат радиально-осевой турбины с устройством измерения углов поворота лопаток соплового аппарата, содержащий регулируемый сопловой аппарат, отличающийся тем, что установлен опорный диск с градуированной шкалой и поворотное кольцо аппарата имеет указатель, а также тем, что соединение лопаток аппарата со шпильками выполнено разъемным для возможной замены лопаток аппарата на лопатки другого профиля. По сравнению ...

УСТРОЙСТВО КРЕПЛЕНИЯ ЛОПАТОК СОПЛОВЫХ АППАРАТОВ

... 1. Устройство крепления лопаток сопловых аппаратов, включающее средство для фиксации лопатки в наружном кольце, выполненное в виде штифта, часть которого утоплена в теле лопатки, причем наружное кольцо выполнено с полками для установки лопаток и разрезных колец, отличающееся тем, что разрезные кольца снабжены упорными буртами.2. Устройство по п.1, отличающееся тем, что упорные бурты жестко закреплены сваркой на разрезных кольцах и ориентированы в радиальном направлении.3. Устройство по п.1, отличающееся тем, что разрезные кольца выполнены с упорными буртами.4. Устройство по п.1, отличающееся тем, что упорные бурты выполнены за одно целое с разрезными кольцами.

Узел соединения стенки жаровой трубы камеры сгорания и соплового аппарата газотурбинного двигателя

Полезная модель относится к области двигателестроения и преимущественно может найти применение в конструкциях жаровых труб камер сгорания газотурбинных двигателей.При работе газотурбинного двигателя в условиях вибрации и температурных нагружений в узле сопряжения жаровой трубы и соплового аппарата происходят микросмещения соединительных поверхностей. Вследствие малой амплитуды микросмещений соприкасающихся поверхностей повреждения концентрируются на небольших площадках действительного контакта. Разрушение контактирующих поверхностей проявляются в появлении мелких полостей, в которых накапливаются продукты износа. Эти продукты износа образуются вследствие разрушения зон сцепления, а также повреждения от усталости микронеровностей. Количество мелких полостей с продуктами износа возрастает и сливается друг с другом. Продукты износа, накапливающиеся в небольших полостях, создают в них повышенное давление, которое в свою очередь приводит к образованию микротрещин. Некоторые микротрещины сливаются ...

УЗЕЛ КРЕПЛЕНИЯ СОПЛОВОГО АППАРАТА В НАРУЖНОМ КОРПУСЕ ТУРБИНЫ

Полезная модель относится к области газотурбинных двигателей (ГТД) и газотурбинных установок (ГТУ), в частности к конструкциям турбин, и может найти применение в узлах крепления соплового аппарата в наружном корпусе турбины.Техническим результатом предлагаемой полезной модели является уменьшение массы и габаритных размеров ступеней соплового аппарата и, следовательно, турбины двигателя, а также уменьшение утечек рабочего тела над сопловым аппаратом, и сокращение времени проведения регламентных работ по осмотру (или ремонту), замене деталей и узлов соплового аппарата и турбины, так как все работы проводят без демонтажа массивного ротора.Технический результат достигается тем, что узел крепления соплового аппарата в наружном корпусе турбины, включающий наружный корпус с горизонтальным разъемом, сопловой аппарат, состоящий из блоков сопловых лопаток и надроторных вставок, блоки сопловых лопаток и надроторные вставки закреплены в наружном корпусе турбины с помощью устройства для крепления кольцевого ...

ОБОЙМА НАПРАВЛЯЮЩИХ ЛОПАТОК ТУРБОМАШИНЫ

Предлагаемая модель решает задачу повышения эффективности работы обоймы направляющих лопаток, обеспечиваемой засчет максимального исключения перетока рабочего тела и охлаждающего воздуха сквозь зазоры в бандажной ленте обоймы. Для достижения указанного технического результата, в обойме направляющих лопаток турбомашины, использующей рабочее тело и охлаждающий воздух, содержащей скрепляющую в единый пакет направляющие лопатки бандажную ленту с фигурными пазами, в каждом из которых расположено перо лопатки, фигурные пазы в бандажной ленте выполнены увеличенными относительно размеров профиля пера соответствующей направляющей лопатки, при этом вокруг пера каждой направляющей лопатки к нижней стороне бандажной ленты прикреплен плоский промежуточный элемент с фигурным пазом, размеры которого выполнены с минимальным зазором относительно профиля пера направляющей лопатки и который, исключая переток рабочего тела и охлаждающего воздуха сквозь зазоры по профилю пера направляющей лопатки, обеспечивает ...

МНОГОВЕНЕЧНАЯ РАДИАЛЬНАЯ ТУРБИНА

... 1. Многовенечная радиальная турбина, содержащая корпус, центробежный сопловой аппарат, вал, рабочее колесо, установленное на валу и выполненное в виде диска, на двух боковых поверхностях которого размещены рабочие лопаточные венцы, причем на боковой поверхности диска рабочего колеса, обращенной к центробежному сопловому аппарату, размещен центробежный рабочий лопаточный венец, а на второй боковой поверхности диска рабочего колеса размещен центростремительный рабочий лопаточный венец, и промежуточный направляющий аппарат, охватывающий рабочее колесо по периферии и установленный с зазором по отношению к последнему, представляющий собой кольцо с размещенными на нем направляющими лопатками, поверхности которых образуют боковые стенки межлопаточных каналов, причем входные кромки направляющих лопаток размещены над центробежным рабочим лопаточным венцом, а выходные кромки направляющих лопаток размещены над центростремительным рабочим лопаточным венцом, отличающаяся тем, что направляющие лопатки ...

ТРУБЧАТАЯ КАМЕРА СГОРАНИЯ ДЛЯ ГАЗОВОЙ ТУРБИНЫ И ГАЗОВАЯ ТУРБИНА, СОДЕРЖАЩАЯ ТАКУЮ ТРУБЧАТУЮ КАМЕРУ СГОРАНИЯ

Трубчатая камера сгорания для газовой турбины, при этом трубчатая камера сгорания содержит по меньшей мере одну горелку, по меньшей мере одну жаровую трубу, ограничивающую пространство горения, имеющее ось камеры сгорания, при этом жаровая труба содержит: внутренний трубчатый элемент, наружный трубчатый элемент, перекрывающий, по меньшей мере частично, внутренний трубчатый элемент и расположенный на расстоянии от внутреннего трубчатого элемента для образования зазора для охлаждающего воздуха, при этом наружный трубчатый элемент содержит расположенный выше по потоку конец, соединенный с промежуточной частью внутреннего трубчатого элемента. Достигается обеспечение возможности относительных перемещений оболочек без влияния на размер зазора. 2 н. и 11 з.п. ф-лы, 7 ил.

СТАТОР МНОГОСТУПЕНЧАТОЙ ГАЗОВОЙ ТУРБИНЫ

Статор многоступенчатой газовой турбины содержит наружный корпус, на радиальных ребрах которого размещены наружные полки сопловых лопаток. Наружные полки зафиксированы в корпусе в окружном направлении соединениями типа «шип-паз», а в осевом направлении - размещенными ниже по потоку секторами разрезного кольца. Сектора разрезного кольца зафиксированы относительно корпуса в осевом направлении байонетными соединениями с корпусом. Сектора с байонетными соединениями зафиксированы в окружном направлении промежуточными секторами того же разрезного кольца. Промежуточные сектора выполнены без кольцевой канавки под байонетное соединение и размещены своей передней частью в выемках корпуса. Число секторов с байонетным соединением равно числу полок сопловых лопаток. Изобретение позволяет повысить надежность и ремонтопригодность статора многоступенчатой газовой турбины. 3 ил.

ОСЕВАЯ ТУРБИНА

Использование: в турбостроении. Сущность изобретения: в осевой турбине с одним рядом регулируемых направляющих лопаток 8 и одним рядом рабочих лопаток 5 расположена нерегулируемая решетка направляющих лопаток. Регулируемые лопатки 8 имеют отношение хорда/ход (s-t), которое значительно меньше такого же отношения нерегулируемых направляющих лопаток 13. Регулируемые направляющие лопатки имеют прямую контрольную линию с симметричным профилем. 5 з. п. ф-лы, 2 ил.

СИСТЕМА УПРАВЛЕНИЯ СТУПЕНЯМИ ПОВОРОТНЫХ ЛОПАТОК СТАТОРА КОМПРЕССОРА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Система управления ступенями поворотных лопаток статора компрессора газотурбинного двигателя содержит ведущий вал, включающий ведущий рычаг, связанный тягой с силовым цилиндром механизма управления, и рычаг, связанный тягой с датчиком угла установки поворотных лопаток. Ведущий рычаг содержит наружный и внутренний рычаги, соединенные в двух местах с помощью первого и второго элементов крепления. Внутренний рычаг выполнен с возможностью поворота относительно второго элемента крепления в случае поломки первого элемента крепления. Тяги силового цилиндра и датчика выполнены с возможностью перемещения в параллельных плоскостях, а датчик непосредственно связан с системой автоматического управления газотурбинного двигателя. Изобретение позволяет повысить надежность системы управления за счет получения достоверной информации о фактическом положении поворотных лопаток статора. 1 з.п. ф-лы, 3 ил.

НАПРАВЛЯЮЩИЙ АППАРАТ ТУРБОМАШИНЫ

Использование: изобретение относится к области турбостроения, в частности к конструкциям направляющих аппаратов турбомашин. Сущность изобретения: направляющий аппарат турбомашины содержит кольцевой обод с вставленными в него направляющими лопатками, имеющими цилиндрические цапфы, и стопорные элементы, выполненные в виде пластин. В ободе имеется кольцевой паз с дном и торцевая стенка, а в хвостовике лопатки - паз. 3 ил.

ОХЛАЖДАЕМЫЙ БЛОК ЛОПАТОК (ВАРИАНТЫ)

... 1. Блок неподвижных лопаток, содержащий внутреннюю полку, у которой имеются внутренняя торцевая поверхность и внутренняя полость; наружную полку, у которой имеются наружная торцевая поверхность и наружная полость, причем наружная полка смещена в радиальном направлении относительно внутренней полки, а ее наружная торцевая поверхность обращена к указанной внутренней торцевой поверхности; по меньшей мере, две лопатки, причем перо каждой лопатки, перекрывающее в радиальном направлении зазор между внутренней и наружной полками, имеет вогнутую поверхность, выпуклую поверхность, входную кромку и выходную кромку, расположенную в осевом направлении позади входной кромки, а вогнутая поверхность и выпуклая поверхность смежных лопаток обращены одна к другой; направляющий канал, ограниченный указанными вогнутой и выпуклой поверхностями смежных лопаток и торцевыми поверхностями внутренней и наружной полок; по меньшей мере, одно отверстие, имеющее входное поперечное сечение и выходное поперечное сечение ...

ЗАКРЕПЛЯЮЩАЯ ДЕТАЛЬ КОНСТРУКЦИЯ СОПЛА ТУРБИНЫ И ТУРБИНА, ИСПОЛЬЗУЮЩАЯ ЕЕ

... 1. Закрепляющая деталь конструкция сопла турбины, содержащего множество сопловых сегментов турбины, каждый из которых содержит множество статорных лопаток, расположенных кольцеобразно вокруг оси реактивного двигателя и закрепленных между корпусом турбины и гондолой реактивного двигателя, причемв сопловом сегменте турбины предусмотрен дугообразный внешний бандаж, соединяющий друг с другом каждый из дистальных концов статорных лопаток на стороне, противоположной оси;на внешнем бандаже в сопловом сегменте турбины образован крюк, взаимодействующий с опорной канавкой сопла, образованной на корпусе турбины, углубленный участок, совмещенный с противовращательной лапкой, прикрепленной к корпусу турбины, и выступающий участок, выступающий к задней части реактивного двигателя в соответствии с углублением углубленного участка на задней стороне указанного углубленного участка; причемсопловой сегмент турбины взаимодействует с крюком внешнего бандажа с опорной канавкой сопла корпуса турбины сзади реактивного ...

УЗЕЛ УПЛОТНЕНИЯ

СБОРОЧНЫЙ УЗЕЛ ЛОПАТОК С РЕГУЛИРУЕМОЙ УСТАНОВКОЙ

... 1. Сборочный узел (1) лопаток с регулируемой установкой, содержащий оси (2) лопаток, пропущенные сквозь корпус (4), регулировочные рычаги (10), насаживаемые на выступающие из корпуса наружу соответствующие части осей, и удерживающие гайки (19), наворачиваемые на резьбовые концы (18) наружных частей, причем головки (11) рычагов (10) удерживаются между гайками и корпусом, отличающийся тем, что наружные части разделены каждая на две части (15, 16), соединяемые между собой с помощью соответствующего средства (17), обеспечивающего возможность регулирования осевого их положения для изменения длины наружных частей, причем одна (16) из частей осей, которая выполнена съемной, имеет резьбовой конец и снабжена упором (20) для рычага, а рычаги (10) зажаты между этими упорами и гайками. 2. Сборочный узел лопаток с регулируемой установкой по п.1, отличающийся тем, что средство, обеспечивающее возможность регулирования осевого положения, представляет собой резьбу. 3. Сборочный узел лопаток с регулируемой ...

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ

Газотурбинный двигатель, содержащий компрессор, камеру сгорания с патрубком отбора воздуха, подключенным в зоне подачи вторичного воздуха, газовую турбину с сопловым аппаратом и рабочими колесами, отличающийся тем, что рабочие колеса установлены с возможностью вращения в противоположных направлениях, сопловый аппарат снабжен кожухом, установленным с образованием охлаждаемой полости, подключенной к патрубку отбора воздуха из камеры сгорания, и выполнен в виде кольцевого сопла, имеющего в продольном сечении форму сопла Ловаля.

ЭЛЕМЕНТ ТУРБИНЫ С ТЕПЛОЗАЩИТНЫМ ПОКРЫТИЕМ

... 1. Элемент турбины, отличающийся тем, что он содержит подложку, образованную из керамического материала, выбранного из группы, состоящей из монолитного керамического материала и композиционного керамического материала, и теплозащитное покрытие, связанное с указанной подложкой. 2. Элемент турбины по п.1, отличающийся тем, что указанный керамический материал выбран из группы, состоящей из нитрида кремния и самоупрочненного нитрида кремния. 3. Элемент турбины по п.1, отличающийся тем, что указанный керамический материал выбран из группы, состоящей из композиционного материала на основе карбидокремниевого волокна и карбидокремниевой матрицы и композиционного материала на основе углеродного волокна и карбонизированной матрицы. 4. Элемент турбины по п.1, отличающийся тем, что указанное теплозащитное покрытие содержит по меньшей мере 15 мол.% по меньшей мере одного полуторного оксида лантанида, а остальное составляет первый оксид, выбранный из группы, состоящей из диоксида циркония, оксида церия ...

СПОСОБ РАБОТЫ ПАРОВОЙ ТУРБИНЫ

Способ работы паровой турбины, содержащей проточную часть из многоступенчатого ротора с дисками, снабженными рабочими лопатками и расположенными между ними дисками диафрагм с сопловыми лопатками на ободе, заключающийся в пропуске пара через нее для вращения ротора, при этом на вогнутых поверхностях лопаток имеются отрывные зоны при истечении пара, отличающийся тем, что, с целью повышения надежности и экономичности, отрывные зоны в вогнутых поверхностях лопаток заполняются паром извне, подводимым к ним по радиально направленным каналам в дисках и лопатках от зон протечек пара помимо лопаток.

ПОСЛЕДНЯЯ СТУПЕНЬ ВЛАЖНОПАРОВОЙ ТУРБИНЫ

Изобретение относится к области энергомашиностроения, в частности паротурбостроения, и может быть использовано при проектировании паровых турбин средней и большой мощности, а именно - при разработке конструкции последних ступеней влажнопаровых турбин, имеющих элементы влагоудаления. Последняя ступень влажнопаровой турбины содержит диафрагму, выполненную из верхней и нижней части, каждая из которых содержит тело, обод, сопловую решетку, образованную полыми направляющими лопатками и внутренними и наружными бандажными лентами. На ободе диафрагмы со стороны паровхода и паровыхода, а также в периферийной зоне тела диафрагмы со стороны паровхода установлены влагоотводящие элементы. Технический результат: повышение эффективности влагоудаления при работе влажнопаровой турбины в условиях повышенной влажности перед направляющими лопатками последних ступеней, что обеспечивает повышение КПД. Повышение эффективности влагоудаления также обеспечивает снижение опасности повышенной влажнопаровой эрозии ...

ГЕРМЕТИЧНОСТЬ МЕЖДУ КАМЕРОЙ СГОРАНИЯ И НАПРАВЛЯЮЩИМ СОПЛОВЫМ АППАРАТОМ ТУРБИНЫ В ГАЗОТУРБИННОМ ДВИГАТЕЛЕ

... 1. Газотурбинный двигатель, содержащий кольцевую камеру (10) сгорания,секторальный направляющий сопловый аппарат (12) турбины, расположенный на выходе камеры, и герметизирующие средства, аксиально размещенные между камерой и направляющим сопловым аппаратом, отличающийся тем, что герметизирующие средства содержат кольцевое уплотнение (70, 70') с аксиальной упругостью, включающее средства (78) аксиального упора на выходной край камеры и выходную кольцевую кромку (74), которая выполнена секторальной, при этом каждый сектор (80) этой выходной кромки размещен на одной линии с сектором направляющего соплового аппарата и содержит средства (82) аксиального упора на входной край сектора направляющего соплового аппарата.2. Газотурбинный двигатель по п.1, отличающийся тем, что уплотнение (70, 70') выполнено аксиально предварительно напряженным в холодном состоянии между камерой (10) и направляющим сопловым аппаратом (12).3. Газотурбинный двигатель по п.1, отличающийся тем, что средства (78) аксиального ...

КОНСТРУКЦИЯ ЛОПАТКИ СТАТОРА И ТУРБОВЕНТИЛЯТОРНЫЙ РЕАКТИВНЫЙ ДВИГАТЕЛЬ С ИСПОЛЬЗОВАНИЕМ ТАКОЙ ЛОПАТКИ

ОСЕВАЯ ПРОТОЧНАЯ ТУРБИНА

Осевая проточная турбина содержит по меньшей мере один ряд изогнутых направляющих лопаток и по меньшей мере один ряд рабочих лопаток. Изгиб направляющих лопаток по их длине расположен перпендикулярно к хорде и направлен к напорной стороне соответственно соседней в окружном направлении направляющей лопатки. Направляющие лопатки сужаются в радиальном направлении. С помощью этих мероприятий уменьшаются вторичные потери, возникающие из-за отклонения пограничных слоев на направляющих лопатках.

ЛОПАТКА СОПЛОВОГО АППАРАТА

... 1. Лопатка соплового аппарата, содержащая полое перфорированное перо, термопары, установленные в его отверстиях и снабженные цилиндрическими экранами, отличающаяся тем, что, с целью повышения точности измерений, отверстия выполнены в зоне входной кромки пера, термопары снабжены трубками, имеющими выступающие над ее поверхностью участки, а экраны размещены эксцентрично последним и закреплены на пере. 2. Лопатка по п.1, отличающаяся тем, что термопары установлены в отверстиях с зазором.

ЛОПАТКА ОСЕВОЙ ТУРБОМАШИНЫ С ПЛАТФОРМАМИ, ИМЕЮЩИМИ УГЛОВОЙ ПРОФИЛЬ

... 1. Лопатка (28) статора осевой турбомашины (2), предназначенная для установки на втулку на кольцевом ряде аналогичных лопаток, при этом лопатка (28) содержит платформу (32, 132, 232, 332, 432, 532) со средствами (34, 134, 234, 334, 434, 534) крепления к втулке, позволяющую выполнять угловую регулировку лопатки, передний край (38, 138, 238, 338, 438, 538), задний край (38, 138, 238, 338, 438, 538) и два противоположных боковых края (36, 136, 236, 336, 436, 536);отличающаяся тем, чтоформа платформы содержит на каждом из своих боковых краев угловой профиль, выполненный с возможностью сопряжения со смежным краем платформы аналогичной смежной лопатки (28), для обеспечения углового позиционирования лопатки по меньшей мере в одном направлении вращения.2. Лопатка (28) по п. 1, отличающаяся тем, что угловой профиль содержит по меньшей мере одну контактную часть (48, 148, 248, 348, 448, 548), по существу, наклоненную к оси (14) турбомашины под углом, большим либо равным 20°, предпочтительно большим ...

СОПЛОВОЙ АППАРАТ ГАЗОВОЙ ТУРБИНЫ

Изобретение относится к области транспортного машиностроения, может быть использовано в газовых турбинах авиационных двигателей и позволяет повысить надежность и эффективность охлаждения соплового аппарата. При работе турбины охлаждающий воздух поступает в кольцевой канал и, вытекая из него, охлаждает внутреннюю обечайку 3. Часть расхода воздуха из зоны входной кромки 7 под козырьком 5 проходит через щелевидные прорези 6, способствуя улучшению охлаждения входной кромки 7 лопатки 1 в зонах над и под козырьком 5. 2 ил.

СОЕДИНИТЕЛЬНОЕ УСТРОЙСТВО ДЛЯ УСТАНОВЛЕНИЯ СОЕДИНЕНИЯ МЕЖДУ СОПЛОМ ТУРБОМАШИНЫ И КАМЕРОЙ ПОДАЧИ ДЛЯПОДВЕДЕНИЯ ОХЛАЖДАЮЩЕЙ ТЕКУЧЕЙ СРЕДЫ К ФОРСУНКАМ

... 1. Соединительное устройство между соплом и камерой подачи для подведения по меньшей мере к одной форсунке охлаждающей текучей среды в турбомашине, содержащее по меньшей мере одну втулку, скользящим образом установленную через ее открытые осевые концы в отверстиях в стенке сопла и в стенке камеры подачи, а также удерживающие средства для удерживания втулки в осевом направлении, при этом удерживающие средства содержат средство опоры, выступающее от цилиндрической внешней поверхности втулки между осевыми концами втулки и на расстоянии от упомянутых концов. 2. Соединительное устройство по п. 1, в котором упомянутое средство опоры содержит внешний кольцеобразный буртик, выступающий от цилиндрической поверхности втулки. 3. Соединительное устройство по п. 1, в котором средство опоры имеет радиальный размер, который превышает радиальный размер отверстий, в которых зацепляются концы втулки. 4. Соединительное устройство по п. 1, в котором втулка получена посредством механической деформации или механической ...

СОПЛО ГАЗОВОЙ ТУРБИНЫ

СПЛАВ НА ОСНОВЕ НИКЕЛЯ

... 1. Жидкотекучий, свариваемый сплав на основе никеля, по существу содержащий, по массе, 10-25% кобальта, 20-28% хрома, 1-3% вольфрама, 0,5-1,5% алюминия, 1,5-2,8% титана, 0,8-1,45% ниобия, тантал в количестве меньшем, чем ниобий, причем Nb+0,508Та составляет 1,15-1,45%, 0,001-0,025% бора, вплоть до 0,4% циркония, 0,02-0,15% углерода, остальное по существу никель и побочные примеси. 2. Сплав по п.1, в котором содержание ниобия составляет по меньшей мере 1,25%. 3. Сплав по п.1, в котором содержание тантала составляет около 0,0%. 4. Сплав по п.1, в котором содержание кобальта составляет 18,5-19,5%, содержание хрома составляет 22,2-22,8%, содержание вольфрама составляет 1,8-2,2%, содержание алюминия составляет 1,1-1,3%, содержание титана составляет 2,2-2,4%, содержание бора составляет 0,002-0,015%, содержание циркония составляет 0, 005-0,4%, и содержание углерода составляет 0,08-0,12%. 5. Сплав по п.1, который содержит по меньшей мере 18 об.% выделившейся гамма-прим-фазы. 6. Сплав по п.1, который ...

СПОСОБ СБОРКИ ДВУХЪЯРУСНОЙ ПРОТОЧНОЙ ЧАСТИ ЦИЛИНДРА ТУРБИНЫ И УЗЕЛ СОЕДИНЕНИЯ ОБОДА ДИАФРАГМЫ С КОЗЫРЬКОМ ЗАДНЕГО МЕЖЪЯРУСНОГО УПЛОТНЕНИЯ ДЛЯ ОСУЩЕСТВЛЕНИЯ СПОСОБА

... 1. Способ сборки двухъярусной проточной части цилиндра турбины, содержащий установку на ободе диафрагмы нижней половины внешнего входного козырька с уплотнительными элементами со стороны внутренней поверхности полок двухъярусных рабочих лопаток в положение, допускающее свободное опускание ротора, после установки ротора сдвиг нижней половины козырька в осевом направлении внутрь полок лопаток и осевую фиксацию козырька в рабочем положении относительно диафрагмы, отличающийся тем, что при указанной установке нижней половины козырька его фиксируют в осевом направлении соединением по подвижной посадке кольцевых паза и выступа козырька и диафрагмы, после установки ротора перед осевым сдвигом нижней половины козырька ее снимают с обода диафрагмы, осевую фиксацию козырька относительно диафрагмы в его рабочем положении выполняют, используя другое соединение по подвижной посадке кольцевых паза и выступа обода диафрагмы и козырька. 2. Узел соединения обода диафрагмы с козырьком заднего межъярусного ...

КОЛЬЦЕВАЯ ПЛАТФОРМА НАПРАВЛЯЮЩЕГО АППАРАТА ТУРБИНЫ НИЗКОГО ДАВЛЕНИЯ ТУРБОМАШИНЫ И НАПРАВЛЯЮЩИЙ АППАРАТ (ВАРИАНТЫ)

... 1. Кольцевая платформа (16, 18) соплового аппарата (10) турбины низкого давления турбомашины, имеющего продольную ось (Х-Х) и содержащего, по меньшей мере, одну сопловую лопатку (14), расположенную за, по меньшей мере, одной подвижной лопаткой (12) турбины высокого давления, причем платформа имеет заднюю часть (16а, 18а), фиксирующую сопловую лопатку (14) и ограничивающую в радиальном направлении аэродинамический канал (24), расположенный в продольном направлении между передней кромкой (14а) сопловой лопатки и задней кромкой (12а) указанной подвижной лопатки (12), отличающаяся тем, что платформа имеет также переднюю часть (16b, 18b), выступающую в продольном направлении за переднюю кромку (14а) сопловой лопатки по направлению к задней кромке (12а) указанной подвижной лопатки (12) с целью удлинения аэродинамического канала (24). 2. Платформа по п.1, отличающаяся тем, что передняя часть (16b, 18b) содержит контур охлаждения. 3. Платформа по п.2, отличающаяся тем, что контур охлаждения содержит ...

Направляющий аппарат турбомашины

Диффузор

Механизм поворота направляющих лопаток осевой турбомашины

Безлопапточный направляющий аппарат радиально-осевой турбины

Устройство крепления соплового аппарата турбины

Направляющий аппарат осевой турбомашины

Механизм поворота направляющих лопаток

Сварная диафрагма турбомашины

Паровая турбина

СОПЛОВАЯ ЛОПАТКА ТУРБОМАШИНЫ

Направляющий аппарат турбинной ступени

MECHANISM FOR ROTATION OF GUIDE BLEADES OF AXIAL TURBINE MACHINE

Регулируемый сопловой аппарат турбомашины

Изобретение относится к холодильной технике и позволяет повысить эксплуатационную надежность - соплового аппарата (СА). Кольцевая пластина 5 жестко и герметично закреплена по периметру меньшего диаметра на корпусе 1 турбины, а по периметру большего диаметра на покрывном диске (ПД) 4 со стороны корпуса. В корпусе со стороны ПД 4 выполнена кольцевая проточка 6, сообщенная с входом СА. Расход газа через СА регулируется с помощью поворотных сапло- вых лопаток 2. Поворот лопаток осуществляется с расчетным усилием, превышающим усилие поджатия ПД 4 прижимным устройством, и от усилия давления газа в зазоре между корпусом и пластиной 5, величина которого определяется из условий надежной эксплуатации СА. 1 ил.с (Л ...

Направляющий аппарат теплофикационной турбины

Изобретение может быть использовано в паровых турбинах с большим теплофикационным отбором пара и позволяет повысить экономичность и надежность турбины в работе путем уменьшения величины переменных аэродинамических сил, действующих на рабочие лопатки (Л). Во внутренних полостях 2 направляющих Л 1 установлены поперечные изогнутые пластины (П) 5 с уменьшающимся щагом 6 от корня 7 к периферии 8 Л 1, образующие решетку 9 П 5 различной густоты. Пар, обтекая направляющие Л 1, частично попадает в их внутренюю полость 2 через щели на входных кромках 4. Равномерный вдоль высоты Л 1 от их корня 7 к периферии 8 отбор пара осуществляется благодаря специально спрофилированным полостям 2 и решеткам 9 П 5. Пар, идущий на теплофикацию по полостям 2 Л 1, попадает в кольцевую отборную камеру 15. Частицы влаги с потоком через радиальные щели во входных кромках 4 Л 1 поступают в полость 2. Из последней скопивщаяся влага через влагоудаляющие каналы 16 попадает в сборную камеру 17. 2 ил. i (Л /3. со 00 ...

Входное устройство двухпоточной газовой турбины

Направляющий аппарат осевой турбомашины

Изобретение относится к турбомашиностроению. Цель изобретения - повышение КПД турбомашины. Направляющий аппарат (НА) состоит из криволинейных закрученных лопаток (Л) 1, периферийной ограничивающей поверхности (ОП) 2 и корневой ОП 3. Образующие поверхности Л 1 наклонены в плоскости, перпендикулярной оси турбины, в направлении, противоположном направлению вращения рабочего колеса, и описываются в указанной плоскости уравнением в полярной системе координат: Ω=(CTGΑгк+TGΘк.TGδк)-1{A/X[1-(R/Rк)X]+B.Rк/X+1[1-( R/Rк)X+1]+C.Rк/X+2[1-R/Rк)X+2]}, где ω -угловая координата точки образующей, за положительное направление которой принято направление вращения рабочего колеса Rк- корневой радиус в выходном сечении НА R - длина радиуса-вектора точки образующей Л 1 A, B, C - постоянные координаты, зависящие от закона распределения меридианных поточных углов по высоте Л 1 αгк- геометрический угол профиля Л 1 в корневом сечении θк- корневой угол раскрытия проточной части в меридианной плоскости δк- угол между ...

Механизм поворота направляющих лопаток осевой турбомашины

Ступень-сепаратор влажнопаровой турбины

СТУПЕНЬ-СЕПАРАТОР ВЛАЖНОПАРОВОП ТУРБИНЫ, содержащая сопловую решетку и рабочее колесо, влагоулавливающую камеру, сообщенную с межосевым зазором влагоулавливающим каналом и с системой отвода влаги, отличающаяся тем, что, с целью повьлиения эффективности и надежности , ступень снабжена жестко закрепленными на сопловой решетке в межосевом зазоре соосно с рабочим колесом концентрическими сепарационными кольцами с влагоулавливающими наклоненными в направлении выхода пара пазами, сообщенными с влагоулавливающей камерой и с системой отвода влаги. U (Л со О) со 00 го ...

Ступень влажно-паровой турбины

Сопловая лопатка для поворотной регулирующей диафрагмы паровой турбины

Направляющий аппарат турбомашины

Ступень турбины

Изобретение может быть исполь- ,зовано в газотурбинных двигателях с регулируемым сопловым аппаратом и позволяет повысить КПЛ ступени турбины . При повороте лопаток 8 кольцо 14, установленное с возможностью осепого перемещения в выходной части 15внутреннего обвода 3 диффузора 1, смещается крипошипным механизмом 16в осевом направлении в соответствии со смещением входной кромки лопаток 8, что приводит к снижению угла раскрытия диффузора 1, а также частичному перекрытию кольцом 14 корневого радиального злзора регулируемого соплового аппарата,что уменьшает перетечки.В исходном состоянии выполнение внутреннего обвода 3 диффузора 1 без перехода к сферической поверхности 6 позволяет уменьшить угол раскрытия диффузора и тем самым снизить потери. 1 ил. i ...

Сопловой аппарат газовой турбины

I.СОПЛОВОЙ АППАРАТ ГАЗОВОЙ ТУРБИНЫ, содержащий лопатки, каддая из которых имеет верхнюю полку, соединенную с наружным корпусом, и нижг нюю полку с кольцевым буртом, размещенным в пазу внутреннего обода, и внутренний корпус с кольцевым выступом , отличающийся тем, что, с целью, повышения КГЩ турбины путем уменьшения утечек, внутренний обод снабжен к ольцевым выступом, расположенным соосно выступу внутреннего корпуса, и на наружной поверхности каждого выступа консольно закреплен пакет колец, размещенных с зазорами относительно друг друга, причем кольца одного пакета частично размещены в зазорах между кольцами другого пакета. 2.Аппарат по П.1, о т л и ч а ю щ и и с я тем, что калдый пакет снабжен втулками, расположенными в зазо- pax заподлицо с его торцевой поверх- 8 ностью со стороны места закрепления па (Л ке та на выступе.

Направляющий аппарат центростремительной микротурбины

Ступень осевой турбомашины

Изобретение относится к области турбостроения. Цель изобретения - повышение КПД турбомашины. Ступень содержит установленные с осевым зазором δанаправляющие лопатки 1 и рабочие лопатки 3. В поршневом сечении Б-Б направляющих лопаток выходная кромка 3 последних наклонена под углом δкк радиальному направлению R. Величина определяется в зависимости от δа, геометрического угла α1квыхода направляющих лопаток в корневом сечении, а также радиусов Rни Rркорневых сечений соответственно направляющих на их выходе и рабочих лопаток на их входе, что препятствует образованию отрыва потока в корневых сечениях рабочих лопаток путем отклонения потока к оси турбомашины. 3 ил.

Радиальный диффузор центробежной турбомашины

Цель изобретения - повышение КПД турбомашины. Между боковыми стенками 1 установлены концентрично друг другу два ряда лопаток 2 с образованием радиального зазора между рядами, составляющего определенную часть диаметра D расположения входных кромок 3 лопаток первого ряда. Выигрыш в КПД турбомашины достигается за счет выравнивания поля скоростей и давления выходящего из лопаток первого ряда потока. 2 ил.

Радиальная турбина для наддува двигателя внутреннего сгорания

Использование; наддув двигателей внутреннего сгорания. Сущность изобретения: площадь входного сечения диффузор- ного канала выполнена больше площади входного сечения конфузорного канала. 1 ил.

Центростремительная газовая турбина

Ступень экспериментальной турбомашины

Входной направляющий аппарат турбомашины

ВХОДНОЙ НАПРАВЛЯЮЩИЙ АППАРАТ ТУРБОМАШИНЫ, , содержащий закрепленную на корпусе обечайку, установленн концентрично ей с возможностью осевого перемещения гильзу и радиальные лопатки, отличающийс я тем, что, с целью повьшения надёжности в работе, обечайка снабжена продольными окнами, гильза отверстиями; корпус - фигурными пазами , а каждая лопатка - цапфой, расположенной в окне и отверстии и имеющей эксцентричный палец, размещённый в пазу корпуса.

Направляющий аппарат турбомашины

НАПРАВЛЯ аЦИП АППАРАТ ТУРБОМАШИНЫ , содержащий наружный и внутренifi-co iL.. 13 TFv. . StSM НИИ ободы, расположенные между ними ; лопатки и перегородки, закрепленные на наружном и/или внутреннем ободе в межлопаточных каналах и расположенные передними кромками в зоне входной части межлопаточных каналов, о тличающийся тем, что, с целью повышения КПД турбомашины Путем снижения вторичных потерь, перегородка выполнена в виде упругой пластины , консольно закрепленной передней кромкой поперек межлопаточногр канала с зазором относительно обода и имеющей боковые кромки, расположенные эквидистантно поверхностям лопатки. Г 1. оо ...

Входное устройство турбомашины

ВХОДНОЕ УСТРОЙСТВО ТУРБОМАШИНЫ , содержащее подводящий патрубок с горизонтальным разъемом и расположенный в нем конфузорный кожух для размещения опоры вала турбом шины, отличающееся тем, что, с целью упрощения монтажно-демонтажных работ, кожух выполнен с горизонтальным фланцевым разъемом, и его верхняя часть закреплена на патрубке, а в проточной части последнего расположена обечайка, охватывающая верхнюю часть кожуха и закрепленная на патрубке. (Я ...

Радиально-осевая ступень турбины

Статор газовой турбины

Регулируемый направляющий аппарат турбомашины

Диафрагма паровых и газовых турбин

Сопловый аппарат для промежуточных ступеней турбины

Двухпоточная паровая турбина

Дифуззор центробежной турбомашины

Сопловой аппарат влажнопаровойТуРбиНы

Сопловой аппарат турбины

Полезная модель относится к турбинам, которые широко применяются в теплоэнергетике, судостроении. Кроме того, турбины используют в составе системы жизнеобеспечения для аппаратов, применяемых в авиации и космонавтике, в мобильных электростанциях, т.е. в тех областях техники, где предъявляют жесткие требования к мобильности и массогабаритным показателям.Поставленная задача решается тем, что в сопловом аппарате турбины, содержащем сопло, имеющее разгонный участок и выходной участок, сообщенные между собой посредством критического сечения, причем выходное сечение сопла на плоскости косого среза выполнено обтекаемой формы и имеет средний радиус изгиба, равный среднему радиусу рабочей решетки колеса турбины, короткая и длинная стенки выходного участка соединены между собой посредством боковых стенок, ограничивающих по радиусу выходной участок, причем короткая стенка выходного участка с примыкающей стенкой разгонного участка выполнены соосными, а продольная ось длинной стенки выходного участка ориентирована под острым углом к продольной оси противолежащей короткой стенки выходного участка.Кроме того, выходное сечение сопла на плоскости косого среза выполнено в виде фигурной щели, продольные стороны которой образованы боковыми стенками и выполнены в форме дуг, ограничивающих по радиусу выходной участок, и сопряжены с торцами, которые образованы короткой и длинной стенками и выполнены в форме дуг с радиусом, равным радиусу критического сечения.Технический результат выражается в повышении коэффициента скорости соплового аппарата и его КПД. 1 з.п. ф-лы, 3 ил. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 168 510 U1 (51) МПК F01D 9/02 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ФОРМУЛА ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ РОССИЙСКОЙ ФЕДЕРАЦИИ (21)(22) Заявка: 2016103347, 03.02.2016 (24) Дата начала отсчета срока действия патента: 03.02.2016 07.02.2017 Приоритет(ы): (22) Дата подачи заявки: 03.02.2016 Адрес для переписки: 690950, Приморский кр., г. Владивосток, ул. ...

Seal assembly for controlling fluid flow

A seal assembly ( 50, 60 ) for a gas turbine engine for controlling air flow between a diffuser ( 48 ) and rotor disks comprising first and second annular flange ends ( 52, 54 ) and an annular seal mid-section ( 56 ) between and operatively connected to the flange ends ( 52, 54 ). The first and second annular flange ends ( 52, 54 ) abut respective outer frame members ( 46 ) of the diffuser, whereby a fluid flow path is formed between the seal assembly ( 50, 60 ) and the rotor disks ( 42 ). The first and second end flanges ( 52, 54 ) are composed of a material having a coefficient of thermal expansion that is substantially the same as a coefficient of thermal expansion of the material of the outer frame members ( 46 ). In addition, the material of the seal mid-section ( 56 ) has a coefficient of thermal expansion that is different than that of the materials of the annular flange ends ( 52, 54 ) and outer frame members ( 46 ).

Turbine nozzle with contoured band

A turbine nozzle includes an array of turbine vanes between inner and outer bands. Each vane includes opposed pressure and suction sides extending between opposed leading and trailing edges. The vanes define a plurality of flow passages each of which is bounded between the inner band, the outer band, and adjacent first and second vanes. A surface of the inner band in each of the passages is contoured in a non-axisymmetric shape including a peak of relatively higher radial height adjoining the pressure side of the first vane adjacent its leading edge, and a trough of relatively lower radial height is disposed parallel to and spaced-away from the suction side of the second vane aft of its leading edge. The peak and trough define cooperatively define an arcuate channel extending axially along the inner band between the first and second vanes.

Aerofoil blade for an axial flow turbomachine

An exemplary aerofoil blade for an axial flow turbomachine has a radially inner platform region, a radially outer tip region, an axially forward leading edge, and an axially rearward trailing edge. The aerofoil blade has a pressure surface which is convex in a radial direction, and a suction surface which is concave in the radial direction. The axial width (W) of the aerofoil blade can vary parabolically between maximum axial widths (W max ) at the platform and tip regions, respectively and a minimum axial width (W min ) at a position between the platform region and the tip region.

High area ratio turbine vane

A vane for a turbine engine comprises an airfoil section, an inner platform and an outer platform. The airfoil section comprises pressure and suction surfaces extending from a leading edge to a trailing edge. The inner platform is attached to the airfoil section along an inner flow boundary, where the inner flow boundary extends from an upstream inlet region of the vane to a downstream outlet region of the vane. The outer platform is attached to the airfoil section along an outer flow boundary, where the outer flow boundary extends from the inlet region to the outlet region. An area ratio of the outlet region to the inlet region is greater than 2.4.

Combustion liner and transition piece

An apparatus is disclosed. The apparatus may include a body configured to flow hot gases of combustion between a forward end and an aft end. Additionally, the body may include a plurality of raised sections spaced apart circumferentially around an outer perimeter of the body. The raised sections may generally extend lengthwise between the forward and aft ends.

Transition channel of a turbine unit

A transition channel for a turbine unit with at least two components is configured as a flow channel from one component of a first pressure to a component of a second pressure. The transition channel has support ribs, extending between envelope surfaces of the transition channel and having a profile that is configured for the deflecting of a flow from an inlet cross section to an outlet cross section of the transition channel. Flow splitter blades are arranged between the support ribs, having a smaller relative profile thickness than the support ribs and/or a shorter axial design depth or profile chord length than the support ribs. Thanks to the integration of the slim and/or short flow splitter blades (tandem blades), it is possible to largely dissipate parasite secondary flows.

Stator for Progressive Cavity Pump/Motor

A stator for a progressive cavity pump or motor includes a stator housing and a stator thread made of a material moulded within the housing. The stator housing includes an outer tube and an inner anchor element within. The anchor element has radially disposed apertures and is spaced from outer tube. To fix the anchor element in the outer tube, the axial end faces of the anchor element each have a first set of protrusions and indentations around the circumference thereof. Two end caps of the housing are fixed at the end faces of the outer tube. Each end cap is provided with an interconnecting part having a second set of protrusions and indentations. The first and second sets of protrusions and indentations interengage to mechanically fix the anchor element axially and radially within the outer tube. 1. A stator for a progressive cavity pump or motor , comprising:a stator housing; anda stator thread moulded within the housing, an outer tube having a first, end and a second end opposite the first end;', 'an inner anchor element disposed within the outer tube and including a plurality of apertures extending radially therethrough, wherein the apertures provide access to one or more gaps radially disposed between the inner anchor element and the outer tube, wherein the stator thread extends through the apertures into the one or more gaps, and wherein a first end and a second end of the inner anchor element each comprise a first plurality of circumferentially-speed protrusions and a plurality of circumferentially-speed indentations, one indentation being circumferentially disposed between each pair of circumferentially-adjacent protrusions of the inner anchor element; and', 'a first end cap fixed to the first end of the outer tube;', 'a second end cap fixed to the second end of the outer tube;', 'wherein each end cap includes an interconnecting part having a second plurality of circumferentially-spaced protrusions and a plurality of circumferentially-spaced indentations, ...

Multi-fluid turbine engine

A multi-fluid turbine engine includes a turbine, a first working fluid inlet passage, and a second working fluid inlet passage. The turbine may include a first working fluid portion and a second working fluid portion. The first working fluid passage may be configured to introduce a first working fluid to the first working fluid portion to perform work on the first working fluid portion, and the second working fluid inlet passage may be configured to introduce a second working fluid to the second working fluid portion to perform work on the second working fluid portion. The first working fluid inlet passage and the second working fluid passage may be independent of each other.

AIRFLOW-STRAIGHTENING STRUCTURE FOR THE NACELLE OF AN AIRCRAFT ENGINE

An airflow-straightening structure for an aircraft engine is provided that includes a hoop inside which there are arranged a plurality of flow-straightening vanes bearing a hub of a fan. Also includes is a clevis for connecting to a suspension strut, the clevis being fixed to said hoop. The structure has at least two elements chosen from the group of the hoop, the plurality of vanes, the hub and the clevis, which are formed in one piece, that is to say without any assembly operation. Additionally, at least one of the elements chosen from the group is at least partially formed of a composite material. 2. The structure according to claim 1 , characterized in that all of said elements are made in a single piece and from a composite material.3. The structure according to claim 1 , characterized in that said hoop incorporates the case of said fan.4. The structure according to claim 1 , characterized in that said hoop incorporates an air intake shroud.5. The structure according to claim 1 , characterized in that said hoop incorporates a thrust reverser cascade edge.6. A nacelle for an aircraft engine claim 1 , characterized in that it comprises an airflow-straightening structure according to . This application is a continuation of International Application No. PCT/FR2011/051379 filed on Jun. 16, 2011, which claims the benefit of FR 10/54852, filed on Jun. 18, 2010. The disclosures of the above applications are incorporated herein by reference.The present disclosure relates to an airflow-straightening structure for an aircraft engine.The statements in this section merely provide background information related to the present disclosure and may not constitute prior art.As is known in itself, and shown in the appended , a nacelle having axis A for a dual-flow (turbojet) engine traditionally comprises an outer structure having an upstream portion forming an air intake, an intermediate portion whereof the inner skin forms a case for the fan of the engine, and a downstream ...

Alternate shroud width to provide mistuning on compressor stator clusters

A stator for a turbo-machine having a plurality of airfoils extending radially therefrom has a base from which the airfoils depend, and slits disposed in the base, each slit disposed adjacent a pair of airfoils, wherein a first set of adjacent slits and a distance between a second set of adjacent slits varies

STEAM TURBINE, BLADE, AND METHOD

A stator blade ring comprising a plurality of stator blade modules defining an annular chamber is provided. The plurality of stator blade modules comprises an elongated blade portion comprising a first and a second blade shell portion, a longitudinal passageway, and at least one opening extending through at least one of the first and the second blade shell portion to the longitudinal passageway, an inner portion brazed to a first longitudinal end of the elongated blade portion, wherein the inner portion comprises a through hole forming a portion of the annular chamber, and an inner passageway extending from the through hole to the longitudinal passageway, and an outer portion brazed to a second longitudinal end of the elongated blade portion and engaged to a steam turbine, the outer portion comprising an outer passageway open to a surface of the steam turbine and the longitudinal passageway. 1. A stator blade ring for a steam turbine , the stator blade ring comprising: an elongated blade portion comprising a first blade shell portion, a second blade shell portion brazed to the first blade shell portion, a longitudinal passageway; and at least one opening extending through at least one of the first blade shell portion and the second blade shell portion to the longitudinal passageway;', 'an inner portion brazed to a first longitudinal end of the elongated blade portion, wherein the inner portion comprises a through hole forming a portion of the annular chamber, and an inner passageway extending from the through hole to the longitudinal passageway; and', 'an outer portion brazed to a second longitudinal end of the elongated blade portion and engaged to the steam turbine, wherein the outer portion comprises an outer passageway open to a surface of the steam turbine and the longitudinal passageway., 'a plurality of stator blade modules defining an annular chamber, wherein each of the plurality of stator blade modules comprises2. The stator blade ring of claim 1 , wherein ...

STATOR VANE ARRAY

A stator vane assembly for a gas turbine engine includes circumferentially spaced vanes about a common axis. The array of vanes further includes three or more sub-arrays, which are configured such that the vane spacing in one sub-array is different from the vane spacing in the other sub-arrays. 116-. (canceled)17. A vane assembly for a gas flow machine , the vane assembly comprising an array of vanes circumferentially spaced about a common axis , wherein the array of vanes comprises three or more sub-arrays , wherein the vane spacing within one sub-array is different from the vane spacing within the other sub-arrays; wherein each sub-array comprises a plurality of adjacent vanes , each vane of a sub-array being substantially equally spaced from an adjacent vane in that sub-array.18. A vane assembly according to claim 17 , wherein each sub-array extends through a portion of a revolution about said axis and the sub-arrays are arranged in an end to end arrangement such that the array forms a complete revolution about the axis.19. A vane assembly according to wherein a step change in vane spacing occurs upon passage from one sub-array to an adjacent sub-array.20. A vane assembly according to claim 17 , wherein the vane spacing varies in a non-cyclic manner through a single revolution of the array about the axis.21. A vane assembly according to claim 17 , wherein at least a pair of sub-arrays have a vane spacing which differs from an average vane spacing for the array by an equal magnitude.22. A vane assembly according to claim 17 , comprising a casing disposed about the axis claim 17 , wherein each vane of the array depends inwardly from the casing towards the axis from the casing.23. A vane assembly according to claim 17 , comprising a stator vane assembly for a gas turbine engine.24. A gas flow machine comprising:a plurality of rows of rotor blades arranged in serial flow arrangement; anda stator vane assembly disposed in the flow path between said rows of rotor ...

Turbomachine component including a cover plate

A turbomachine component includes a body having a first end that extends to a second end. One of the first and second ends includes a mounting element, and a mounting component. A cover plate is arranged at the one of the first and second ends to establish an interface region. The cover plate includes a mounting member configured to align with the mounting element, and a mounting portion configured to align with the mounting element. A fastener member is configured and disposed to cooperate with the mounting element and the mounting member to constrain the cover plate to the body along at least two axes with the interface region being devoid of a metallurgical bond.

Multi-piece twin scroll turbine

A turbine is provided. The turbine includes a housing radially extending around a turbine rotor including a first piece defining a portion of a first scroll passage boundary and a second piece having an interface wall contiguous with an interface wall of the first piece, the second piece coupled to the first piece and including a divider defining another portion of the first scroll passage boundary and a portion of a second scroll passage boundary.

NOISE-REDUCED TURBOMACHINE

Turbomachine with an annular main flow duct () through which passes a flow, and in which is arranged at least one stator provided with stator vanes (), characterized in that the stator vanes () each have at least one recess () issuing into a flow duct () inside the stator vanes () and that the flow duct () issues into a bypass duct () of the turbomachine via at least one outflow duct () provided with a shut-off element (). 1. A turbomachine , comprising:an annular main flow duct through which passes a flow;a stator having a plurality of stator vanes arranged in the main flow duct, the stator vanes each having a recess;a flow duct inside the stator vane, the recess issuing into the flow duct;an outflow duct, the flow duct issuing into a bypass duct of the turbomachine via the outflow duct;a shut-off element for shutting-off flow through the outflow duct.2. The turbomachine of claim 1 , wherein the stator is a guide vane.3. The turbomachie of claim 2 , wherein the recess is has a slot form.4. The turbomachine of claim 2 , wherein the recess is formed as a line-type row of holes.5. The turbomachine of claim 2 , wherein the recess s located on a suction side of the stator vane.6. The turbomachine of claim 5 , wherein the recess is arranged adjacent to a stator profile trailing edge of the stator vane.7. The turbomachine of claim 6 , wherein the recess extends substantially over an entire radial length of the stator vane.8. The turbomachine of claim 7 , and further comprising a plurality of recesses arranged at at least one chosen from equal and varying spacing to one another.9. The turbomachine of claim 8 , wherein the at least one outflow duct is connected to an annular duct extending in the circumferential direction relative to a machine axis claim 8 , and the flow ducts issue into the annular duct10. The turbomachine of claim 9 , and further comprising a diffuser claim 9 , the outflow duct issuing into the bypass duct via the diffuser.11. The turbomachine of claim 10 ...

VANE ASSEMBLY FOR AN AXIAL FLOW TURBINE

A vane assembly () for an axial flow turbine engine provides a plurality of nozzles () arranged symmetrically around a turbine axis formed by an inner hub (), an outer casing () and vanes () intermediate pairs of adjacent nozzles. The nozzles have a cross-section normal to the turbine axis that changes smoothly with the position of the cross-section along the turbine axis from a first shape having a minimum radius of curvature of a first value that is no less than a limit L, providing the cross-section with no highly curved corners, to a second shape having a minimum radius of curvature of a second value less than the first value, providing corners. This allows the secondary flow structure to be designed to improve aerodynamic loss and cooling performance. 1. A vane assembly for an axial flow turbine , the vane assembly providing a plurality of nozzles arranged symmetrically around a turbine axis formed by an inner hub , an outer casing and vanes intermediate pairs of adjacent nozzles , the vanes providing a suction surface of one of the pair of adjacent nozzles and a pressure surface of the other of the pair of adjacent nozzles ,at least some of the nozzles having a cross-section normal to the turbine axis that changes smoothly with the position of the cross-section along the turbine axis from a first shape having a minimum radius of curvature of a first value that is no less than a limit L to a second shape having a minimum radius of curvature of a second value less than the first value,wheredefining a reference plane normal to the turbine axis at a position where a circle in the reference plane centred on the turbine axis and touching the trailing edge of a vane minimises the position along the turbine axis in the direction of axial flow,defining a reference radius from the turbine axis that is the average of (a) the minimum radius from the turbine axis of the nozzle at the inner hub in the reference plane and (b) the maximum radius from the turbine axis of the ...

Stress relieving slots for turbine vane ring

A turbine vane ring has a radially outer and inner annular shrouds defining therebetween an annular gaspath. Circumferentially spaced-apart airfoil vanes extend radially across the gaspath between the outer and the inner shrouds. The radially outer shroud has a circumferentially continuous cylindrical wall extending axially from a leading edge to a trailing edge. A set of circumferentially distributed stress relieving slots is defined in the leading edge of the cylindrical wall at locations adjacent to the leading edge of at least some of said airfoil vanes. The stress relieving slots extend radially through the cylindrical wall from the radially inner surface to the opposed radially outer surface thereof.

PLATFORM SEGMENT FOR SUPPORTING A NOZZLE GUIDE VANE FOR A GAS TURBINE AND NOZZLE GUIDE VANE ARRANGEMENT FOR A GAS TURBINE

A platform segment for supporting a nozzle guide vane for a gas turbine is provided. The platform segment includes a gas passage surface arranged to be in contact with a streaming gas exhausted from a combustor, wherein the streaming gas streams along the gas passage surface in a streaming direction, a cooling surface, opposite to and thermally connected to the gas passage surface and arranged to be in contact with a cooling fluid, a wall protruding from the cooling surface and extending at least partially in the streaming direction, wherein the wall is arranged circumferentially between positions, at which adjacent guide vanes are to be provided, such that cooling fluid is channeled by the wall for cooling a downstream portion of the cooling surface, and a further wall protruding from the cooling surface and extending at least partially in the streaming direction. 113-. (canceled)14. A platform segment for supporting a nozzle guide vane for a gas turbine , the platform segment comprising:a gas passage surface arranged to be in contact with a streaming gas exhausted from a combustor, wherein the streaming gas streams along the gas passage surface in a streaming direction;a cooling surface opposite to and thermally connected to the gas passage surface and arranged to be in contact with a cooling fluid;a wall protruding from the cooling surface and extending at least partially in the streaming direction, wherein the wall is arranged circumferentially between positions, at which adjacent guide vanes are to be provided, such that cooling fluid is channelled by the wall for cooling a downstream portion of the cooling surface; anda further wall protruding from the cooling surface and extending at least partially in the streaming direction,wherein a circumferential distance between the wall and the further wall decreases along the streaming direction,wherein the nozzle guide vane having a pressure surface and a suction surface is connectable at the platform segment such ...

Transition Piece Aft-Frame Seals

A transition piece aft-frame seal assembly may include an elongate body including a first side and a second side. The seal assembly may also include at least one feed hole disposed on the first side of the body. The seal assembly may also include at least one passageway extending through the body from the first side to the second side and in communication with the at least one feed hole. Moreover, the seal assembly may include at least one cooling hole disposed at the second side of the body and in communication with the at least one passageway. A flow of cooling fluid may enter the at least one feed hole, the at least one passageway, and the at least one cooling hole, wherein the at least one cooling hole directs the flow of cooling fluid to a recirculation zone about adjacent transition piece aft-frame assemblies.

TURBOJET CASING AND TURBOJET RECEIVING SUCH CASINGS

A turbojet casing adapted to receive a plurality of vanes, the casing including attachment means () for attaching one end of each vane to the casing, the casing being characterized in that the attachment means extend on a face of the casing facing away from the vanes, the casing including orifices () for passing the ends of the vanes so that they can co-operate with the attachment means of the casing. A turbojet including such a casing. 1213026. A turbojet casing adapted to receive a plurality of vanes , the casing including attachment means ( , ) for attaching one end of each vane to the casing , the casing being characterized in that the attachment means extend on a face of the casing facing away from the vanes , the casing including orifices () for passing the ends of the vanes so that they can co-operate with the attachment means of the casing.22130. A casing according to claim 1 , wherein the attachment means comprise an annular member ( claim 1 , ) extending around the casing.3212423. A casing according to claim 2 , wherein the annular member comprises at least one peripheral rail () having the ends () of fastener elements () for fastening to the ends of the vanes inserted therein.430. A casing according to claim 2 , wherein the annular member comprises a peripheral angle bar () to which the ends of the vanes are fastened directly.52130. A casing according to claim 2 , the casing being made of long fibers associated with a thermoplastic resin claim 2 , while the annular member ( claim 2 , ) is obtained by pultrusion and impregnated with a thermoplastic resin that is heat-sealable with the thermoplastic resin of the casing claim 2 , the assembly being joined together by hot compaction.6. A turbojet including at least one casing according to claim 1 , and a plurality of vanes claim 1 , each having one end connected to the casing.7. A turbojet according to claim 6 , wherein each of the vanes comprises:{'b': '2', 'an elongate one-piece front portion () cut from a ...

TURBINE AIRFOIL TRAILING EDGE COOLING SLOTS

A turbine airfoil includes pressure and suction sidewalls extending along a span from a base to a tip. Spanwise spaced apart trailing edge cooling holes in the pressure sidewall end at corresponding spanwise spaced apart trailing edge cooling slots extending chordally substantially to the trailing edge. Each cooling hole includes in downstream serial cooling flow relationship, a curved inlet, a constant area and constant width metering section, and a spanwise diverging section leading into the trailing edge cooling slot, and a spanwise height substantially greater than a hole width through the cooling hole. A pressure sidewall surface of the pressure sidewall may be planar through the metering and diverging sections. The width may be constant through the metering and diverging sections. A raised floor may include a flat up ramp in the diverging section, a flat down ramp in the slot. 1. A gas turbine engine turbine airfoil comprising:widthwise spaced apart pressure and suction sidewalls extending outwardly along a span from an airfoil base to an airfoil tip;the pressure and suction sidewalls extending chordwise between opposite leading and trailing edges;a spanwise row of spanwise spaced apart trailing edge cooling holes encased in the pressure sidewall and ending at corresponding spanwise spaced apart trailing edge cooling slots extending chordally substantially to the trailing edge;the cooling hole including in downstream serial cooling flow relationship, a curved inlet, a metering section with a constant area and constant width flow cross section, and a spanwise diverging section leading into the trailing edge cooling slot; anda spanwise height substantially greater than a hole width through the cooling hole.2. The airfoil as claimed in further comprising pressure and suction sidewall surfaces of the pressure and suction sidewalls respectively in the hole and the pressure sidewall surface being planar through the entire metering and diverging sections.3. The ...

Flexible seal for transition duct in turbine system

A turbine system is disclosed. In one embodiment, the turbine system includes a transition duct. The transition duct includes an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of the transition duct is offset from the inlet along the longitudinal axis and the tangential axis. The transition duct further includes an interface feature for interfacing with an adjacent transition duct. The turbine system further includes a flexible seal contacting the interface feature to provide a seal between the interface feature and the adjacent transition duct. The flexible seal includes a sheet having a first surface, an opposing second surface, and a peripheral edge therebetween.

TURBINE VANE

In a turbine vane and a gas turbine, an outer shroud is fixed to one end of a vane body formed in a hollow shape, an inner shroud is fixed to the other end thereof, and a partition plate is fixed to the inner portions of the vane body, the outer shroud, and the inner shroud, so that a cavity is formed so as to be continuous between the partition plate and the group of the vane body, the outer shroud, and the inner shroud. Then, the vane body, the outer shroud, and the inner shroud are provided with a plurality of cooling holes, and the partition plate is provided with a plurality of penetration holes. Accordingly, since the vane structure or the end wall structure is evenly cooled, a deformation or damage may be suppressed. 1. A turbine vane comprising: a vane structure formed in a hollow shape; an end wall structure provided in an end of the vane structure; and a partition plate for forming a cavity continuous inside the vane structure and the end wall structure , the partition plate being provided with a plurality of penetration holes.2. The turbine vane according to claim 1 , wherein the partition plate is formed in a cylindrical shape claim 1 , and an end near the end wall structure is enlarged and is fixed to the end wall structure.3. The turbine vane according to claim 1 , wherein a protrusion is provided between the vane structure and the partition plate or between the end wall structure and the partition plate so as to suppress the gap therebetween from being narrowed.4. The turbine vane according to claim 1 , wherein the end wall structure includes an outer end wall structure connected to one end of the vane structure and an inner end wall structure connected to the other end of the vane structure claim 1 , and the partition plate includes an outer partition plate inserted from the outer end wall structure and an inner partition plate inserted from the inner end wall structure.5. The turbine vane according to claim 4 , wherein the outer partition plate and ...

AIRFOIL INCLUDING ADHESIVELY BONDED SHROUD

An airfoil includes an airfoil body that extends between a leading edge and a trailing edge, a suction side and a pressure side, and a first end and a second end. At least one fitting is located on at least one of the first end and the second end. The fitting includes at least one mounting lug. At least one shroud is adhesively bonded to the at least one fitting. 1. An airfoil comprising:an airfoil body extending between a leading edge and a trailing edge, a suction side and a pressure side, and a first end and a second end;at least one fitting located on at least one of the first end and the second end, the at least one fitting including at least one mounting lug; andat least one shroud adhesively bonded to the at least one fitting.2. The airfoil as recited in claim 1 , wherein the at least one fitting includes a flange claim 1 , and the at least one shroud is adhesively bonded to the flange.3. The airfoil as recited in claim 2 , wherein the flange includes a rabbet claim 2 , and the at least one shroud is adhesively bonded to the rabbet.4. The airfoil as recited in claim 1 , wherein the at least one shroud comprises a polymeric material.5. The airfoil as recited in claim 1 , wherein the at least one fitting is metallic and the at least one shroud is polymeric.6. The airfoil as recited in claim 1 , further comprising a seal member attached at an edge of the at least one shroud.7. The airfoil as recited in claim 6 , wherein the edge of the at least one shroud includes a slot claim 6 , and the seal member is located in the slot.8. The airfoil as recited in claim 6 , wherein the seal member is adhesively bonded to the edge of the at least one shroud.9. The airfoil as recited in claim 1 , wherein the at least one shroud includes a plurality of distinct shroud pieces that claim 1 , when assembled together claim 1 , circumscribe the at least one fitting.10. The airfoil as recited in claim 1 , wherein the at least one shroud includes a shroud body extending between first ...

Gas turbine engine turbine vane airfoil profile

A turbine vane for a gas turbine engine includes inner and outer platforms joined by a radially extending airfoil. The airfoil includes leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface. The inner and outer platforms respectively include inner and outer sets of film cooling holes, wherein one of the inner and outer sets of film cooling holes are formed in substantial conformance with platform cooling hole locations described by one of the sets of Cartesian coordinates set forth in Tables 1 and 2. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate. The cooling holes with Cartesian coordinates in Tables 1 and 2 have a diametrical surface tolerance relative to the specified coordinates of 0.200 inches (5.08 mm).

NOZZLE, A NOZZLE HANGER, AND A CERAMIC TO METAL ATTACHMENT SYSTEM

A nozzle, a nozzle hanger, and a ceramic to metal attachment system are provided. The ceramic to metal attachment system includes the nozzle, a ceramic matrix composite, and the nozzle hanger, a metal. The attachment system also includes a clamping member adjacent a second surface of the nozzle and a mounting member of the nozzle. The attachment system includes a plurality of attachment members securing the nozzle, the clamping member, and the nozzle hanger together. A sealing member of the nozzle hanger seals off an airfoil of the nozzle from adjacent airflow. 1. A nozzle comprising:a first band;a second band;an airfoil joining the first band and the second band; and a cavity, the cavity extending through the airfoil;', 'a radial outer load bearing surface surrounding the cavity;', 'a radial inner load bearing surface opposite the radial outer load bearing surface;', 'a tangential interface between the radial outer load bearing surface and the radial inner load bearing surface; and', 'a moment interface surface between the radial outer load bearing surface and the radial inner load bearing surface and opposite the tangential interface;', 'wherein the mounting member attaches the nozzle to a surrounding static surface., 'a mounting member integrally formed with the second band and the airfoil, the mounting member having a first surface and a second surface, the mounting member including2. The nozzle of claim 1 , wherein the nozzle is a ceramic matrix composite or a metal.3. The nozzle of claim 1 , wherein the nozzle is mounted in a cantilevered arrangement.4. The nozzle of claim 1 , wherein the mounting member includes an axial interface adjacent the moment interface and between the radial outer load bearing surface and the radial inner load bearing surface.5. The nozzle of claim 1 , wherein mounting member is machined to form the tangential load bearing surface and the moment interface surface.6. A nozzle hanger comprising:a nozzle receiving surface;a shroud hanger ...

TURBOMACHINE WITH FAN(S) FOR AIRCRAFT, WITH MOBILE JET NOZZLE

A turbomachine for an aircraft including at least one turbine driving rotation of a fan, and a jet nozzle coaxially extending a turbine and creating, with an outlet cone that terminates the turbine, a passage of annular cross section for combustion gases that pass through the turbine. The jet nozzle can be mounted so that it can move between two extreme positions for which under action of a controller, an annular cross section of the passage for the gases is respectively at its minimum or maximum, making it possible to vary the cross section according to phases of operation of the aircraft and distribute thrust between the fan and the jet nozzle. 16-. (canceled)7. A turbomachine including at least one propeller for an aircraft , comprising:at least one turbine for driving the propeller in rotation;a nozzle that extends coaxially beyond the turbine, creating, with an exhaust cone that terminates the turbine, a passage of annular cross section for combustion gases flowing through the turbine, the nozzle being mounted to be able to move between two end positions at which the annular cross section of the passage is respectively at a minimum or a maximum; andcontrol means attached to the nozzle to change a position thereof and to vary the cross section of the passage for the combustion gases depending on functional phases of the aircraft;wherein, to vary the cross section of the passage for the combustion gases, the nozzle moves, under action of the control means, in translation parallel to the longitudinal axis of the turbomachine toward which the exhaust cone of the turbine converges.8. The turbomachine as claimed in claim 7 , wherein the control means includes actuators connecting a portion of a peripheral nacelle claim 7 , which surrounds the turbine claim 7 , to the nozzle.9. The turbomachine as claimed in claim 8 , wherein the control actuators are arranged distributed at regular angular intervals around the nozzle claim 8 , parallel to the longitudinal axis of the ...

Fixed Blade Assembly Usable in Exhaust Pump, and Exhaust Pump Provided with same

Provided are a stator blade assembly advantageous for improving the evacuation performance by shortening the evacuation time, and an exhaust pump provided with such a stator blade assembly. Inner and outer stator blade bases of a plurality of stator blades are supported by frames. A projecting portion protruding from the frame supporting the inner stator blade base, or from the frame supporting the outer stator blade base, or from both of the frames is provided in a gap in the vicinity of the inner or outer stator blade base between one of the supported stator blades and the stator blade transversely adjacent thereto. 1. A stator blade assembly which is usable in an exhaust pump in which a plurality of rotor blades protruding from an outer circumferential surface of a rotatable rotor and a plurality of stator blades protruding toward the outer circumferential surface of the rotor are alternately disposed in multiple stages along an axis of the rotor , whereinthe plurality of stator blades are configured such that inner and outer stator blade bases are supported by frames, and a projecting portion protruding from the frame supporting the inner stator blade base, or from the frame supporting the outer stator blade base, or from both of the frames is provided in a gap in the vicinity of the outer or inner stator blade base between one of the supported stator blades and a stator blade transversely adjacent thereto.2. The stator blade assembly according to claim 1 , wherein the inner and outer stator blade bases of the one stator blade and the inner and outer stator blade bases of the stator blade transversely adjacent thereto are configured as a stator blade assembly of a single-layer structure by being supported by the same frame.3. The stator blade assembly according to claim 1 , wherein the inner and outer stator blade bases of the one stator blade and the inner and outer stator blade bases of the stator blade transversely adjacent thereto are supported by separate ...

DEVICE FOR SEPARATING WATER DROPLETS FROM A GAS OR VAPOR FLOW

A moisture separator for a steam turbine power plant for separating moisture from a flow of vapor or gas that includes a bundle of vanes having a corrugated portion and a trailing edge. The trailing edge includes means to collect moisture on the surfaces of the trailing edge. These means can include a clip-like attachment, or a U-shape of the trailing edge itself. The collection means on the trailing edge contribute to the overall moisture separation efficiency of the moisture separator and a decrease of risk of damage to the steam turbine driven by the steam flow. 1. A moisture separation apparatus for a steam turbine power plant for the generation of electricity comprising a housing with an inlet for a gas or vapor or steam flow ,a bundle of vanes arranged within the housing, where the vanes are vertically arranged in parallel and evenly spaced from one another, and each vane comprises a corrugated portion arranged closest to the inlet for the gas or vapor flow and a trailing edge portion arranged opposite from said inletwhereinthe trailing edge of one or more vanes comprises a means for collecting moisture on the trailing edge's surface and direct it to the bottom end of the vane.2. The moisture separation apparatus according to wherein the means for collecting moisture on the trailing edge's surface consists of a U-shaped bend in the trailing edge.3. The moisture separation apparatus according to wherein the U-shaped bend in the trailing edge comprises a first arm extending away from the corrugated portion of the vane at an angle to the longitudinal axis of the corrugation of the vane claim 2 , a U-turn portion and a second arm extending toward the corrugation.4. The moisture separation apparatus according to wherein the means for collecting moisture on the trailing edge's surface comprises a clip attached to the trailing edge and enveloping the end face and surfaces of the trailing edge.5. The moisture separation apparatus according to wherein the clip extends ...

STATIC VANE ASSEMBLY FOR AN AXIAL FLOW TURBINE

An axial flow turbine is described having a casing defining a flow path for a working fluid therein, a rotor co-axial to the casing, a plurality of stages, each including a stationary row of vanes circumferentially mounted on the casing a rotating row blades, circumferentially mounted on the rotor, with within a stage n vanes have an extension such that at least a part of the trailing edge of each of the n vanes reaches into the annular space defined by the trailing edges of the remaining N-n vanes and the leading edges of rotating blades of the same stage. 1. An axial flow turbine comprising:a casing defining a flow path for a working fluid therein;a rotor co-axial to the casing; a row of N stationary vanes circumferentially mounted on the casing; and', 'a row of rotating blades circumferentially mounted on the rotor,, 'a plurality of stages, each comprisingwherein within a stage, n vanes have an extension such that at least a part of the trailing edge of each of the n vanes reaches into the annular space limited by the rotor and the casing and the trailing edges of the remaining N-n vanes and the leading edges of rotating blades of the same stage,wherein the number n of extended vanes is larger than zero but less than half the total number N of vanes in the stage.2. The turbine according to wherein the stage is a last stage of a low pressure steam turbine.3. The turbine according to wherein the number n is selected to be 0 Подробнее

TURBINE EXHAUST STRUCTURE AND GAS TURBINE

In a turbine exhaust structure and a gas turbine, a turbine casing () formed in an annular shape to constitute a combustion gas passage A is provided. An exhaust diffuser () formed in an annular shape to constitute a flue gas passage (B) is connected to the turbine casing () to constitute a configuration. By providing a pressure loss body () in the exhaust diffuser (), efficient pressure recovery can be carried out, which improves turbine efficiency, thereby enabling improvement of the performance. 1. A turbine exhaust structure including a casing formed in an annular shape to constitute a combustion gas passage , in which an exhaust diffuser formed in an annular shape to constitute a flue gas passage is connected to the casing ,wherein a pressure loss body is provided in the exhaust diffuser.2. The turbine exhaust structure according to claim 1 , wherein the pressure loss body includes a porous member arranged in the flue gas passage of the exhaust diffuser.3. The turbine exhaust structure according to claim 1 , wherein the pressure loss body includes a heat transfer tube in which heat is exchanged between a flue gas and a heat exchange medium flowing inside the heat transfer tube.4. The turbine exhaust structure according to claim 1 , wherein the pressure loss body is designed so as that a pressure loss in at least one of radially outer side or inner side of the exhaust diffuser is smaller than a pressure loss in a radially middle portion of the exhaust diffuser.5. The turbine exhaust structure according to claim 1 , wherein a support structure penetrating the exhaust diffuser in a radial direction is provided claim 1 , and the pressure loss body is arranged in a downstream side of the support structure in a flow direction of a flue gas.6. A gas turbine in which a fuel is supplied to the compressed air claim 1 , compressed in a compressor claim 1 , in a combustor to create combustion claim 1 , and the produced combustion gas is supplied to a turbine to obtain ...

PROCESS GAS FLOW GUIDES FOR LARGE AREA PLASMA ENHANCED CHEMICAL VAPOR DEPOSITION SYSTEMS AND METHODS

The present invention provides methods and apparatus for a gas diffusion assembly in a deposition processing chamber. The invention includes a backing plate having an inlet for providing a process gas to a process chamber, a diffusion plate including a plurality of apertures for allowing the process gas to flow into the process chamber, a blocking plate disposed between the backing plate and the diffusion plate and including a plurality of apertures, and at least one gas flow guide disposed between the blocking plate and the backing plate and adapted to direct process gas flow laterally. Numerous additional features are disclosed. 1. A gas diffusion assembly comprising:a backing plate having an inlet for providing a process gas to a process chamber;a diffusion plate including a plurality of apertures for allowing the process gas to flow into the process chamber;a blocking plate disposed between the backing plate and the diffusion plate and including a plurality of apertures; andat least one gas flow guide disposed between the blocking plate and the backing plate and adapted to direct process gas flow laterally.2. The gas diffusion assembly of wherein the at least one gas flow guide is adapted to direct process gas flow laterally toward an area above a substrate that would otherwise receive a lower deposition rate.3. The gas diffusion assembly of wherein the at least one gas flow guide is adapted to direct process gas flow laterally away from an area above a substrate that would otherwise receive a higher deposition rate.4. The gas diffusion assembly of wherein the at least one gas flow guide includes four inner gas flow guides disposed around the inlet.5. The gas diffusion assembly of wherein the inner gas flow guides are adapted to direct SiOx process gas laterally toward an area above a long edge of a substrate.6. The gas diffusion assembly of wherein the inner gas flow guides are adapted to direct SiOx process gas laterally away from an area above a corner of a ...

ASSEMBLY OF AN AXIAL TURBOMACHINE AND METHOD FOR MANUFACTURING AN ASSEMBLY OF THIS TYPE

The present invention relates to an assembly of an axial turbomachine, comprising at least one outlet guide vane of a compressor and a diffuser arranged downstream of the outlet guide vane in the flow direction. It is provided that the outlet guide vane is connected to the compressor and that the diffuser is connected to the combustion chamber, without there being a direct mechanical connection between the diffuser and the outlet guide vane. The invention furthermore relates to a method for manufacturing an assembly of this type. 1. An assembly of an axial turbomachine , comprising at least one outlet guide vane of a compressor and a diffuser arranged downstream of the outlet guide vane in the flow direction , wherein the outlet guide vane is connected to the compressor and the diffuser is connected to the combustion chamber , without there being a direct mechanical connection between the diffuser and the outlet guide vane.2. The assembly in accordance with claim 1 , wherein there is a leakage gap between the outlet guide vane and the diffuser in the flow direction.3. The assembly in accordance with claim 1 , wherein the outlet guide vane is connected to an outer casing of the compressor.4. The assembly in accordance with claim 1 , wherein the diffuser is connected to an inner combustion chamber casing of the combustion chamber.5. The assembly in accordance with claim 1 , wherein the outlet guide vane is designed with a hub gap extending between the blade ends of the outlet guide vane and the hub of the outlet guide vane.6. The assembly in accordance with claim 5 , wherein the hub of the outlet guide vane is connected to the rotor hub or the rotor disk of a compressor rotor arranged upstream of the outlet guide vane in the flow direction.7. The assembly in accordance with claim 1 , wherein the outlet guide vane is designed without a hub gap between the blade ends of the outlet guide vane and the hub of the outlet guide vane claim 1 , with the hub of the outlet guide ...

STRUCTURAL GUIDE VANE INTERNAL TOPOLOGY

A disclosed structural guide vane for a gas turbine engine includes a strut defining a plurality of channels and a cover attached over the plurality of channels. The strut includes a first side defining a first surface of an airfoil and a second side including a plurality of channels. At least one of the plurality of channels defines a wall thickness of a portion of the first side that varies in a chord wise direction. 1. A gas turbine engine comprising:a fan section including a plurality of fan blades rotatable about an axis;a fan case disposed around the fan section;a core engine section; andat least one structural guide vane extending between the core engine and the fan case, the structural guide vane including at least one channel including a wall with a first thickness transverse to the axis, wherein the first thickness varies in a chord wise direction.2. The gas turbine engine as recited in claim 1 , wherein the at least one channel extends longitudinally in a span wise direction.3. The gas turbine engine as recited in claim 1 , wherein the structural guide vane includes a cover secured over the at least one channel and defining a side of the structural guide vane.4. The gas turbine engine as recited in claim 1 , wherein the at least one channel includes a plurality of channels with each channel including a corresponding plurality of walls and each of the plurality of walls includes a thickness different than at least one of the other walls of another of the plurality of channels.5. The gas turbine engine as recited in claim 1 , wherein the at least one channel includes a trailing edge channel including a forward portion and an aft portion and a wall with a thickness that varies from the forward portion to the aft portion.6. The gas turbine engine as recited in claim 5 , wherein the wall thickness is smallest at an intermediate point between the forward portion and the aft portion.7. The gas turbine engine as recited in claim 5 , including an intermediate ...

DOUBLE FLOW TURBINE HOUSING TURBOCHARGER

Implementations of the present disclosure are directed to turbine assemblies for turbocharger systems. In some implementations, turbine housings include a body that defines an inlet for fluid communication with a fluid source, and a wall, the wall dividing the inlet into an inner inlet and an outer inlet, and a fluid guide assembly disposed within the housing, the fluid guide assembly including a plurality of vanes that demarcate an inner volute and an outer volute within the housing, the inner volute being in fluid communication with the inner inlet and the outer volute being in fluid communication with the outer inlet, each vane of the plurality of vanes being fixed at a respective angle relative to a radial direction, the plurality of vanes guiding fluid flow from the outer volute to the inner volute. 1. A turbine housing for a turbocharger , the housing comprising:a body that defines an inlet for fluid communication with a fluid source, and a wall, the wall dividing the inlet into an inner inlet and an outer inlet; anda fluid guide assembly disposed within the housing, the fluid guide assembly comprising a plurality of vanes that demarcate an inner volute and an outer volute within the housing, the inner volute being in fluid communication with the inner inlet and the outer volute being in fluid communication with the outer inlet, each vane of the plurality of vanes being fixed at a respective angle relative to a radial direction, the plurality of vanes guiding fluid flow from the outer volute to the inner volute.2. The turbine housing of claim 1 , wherein the fluid guide assembly further comprises a guide plate that is secured to the body claim 1 , the plurality of vanes being secured to the guide plate.3. The turbine housing of claim 1 , wherein at least one of the vanes is positioned at a selected angle relative to the radial center of the turbine wheel.4. The turbine housing of claim 3 , wherein the selected angle is between approximately 30° and ...

ELECTRIC ACTUATOR FOR ENGINE CONTROL

An electric actuator for control of an engine includes an electric motor coupled to a drive shaft that extends to align a gear interface of the electric actuator with a variable geometry adjustment interface of the engine. A position feedback shaft extends coaxially with respect to the drive shaft. The position feedback shaft is coupled to an output shaft of the gear interface at a gear interface end of the position feedback shaft. A rotational position sensor is coupled to a motor end of the position feedback shaft proximate the electric motor. The drive shaft and the position feedback shaft are sized to position an output ring gear of the output shaft in contact with the variable geometry adjustment interface within a casing of the engine and to further position the electric motor and the rotational position sensor external to the casing of the engine. 1. An electric actuator for control of an engine , the electric actuator comprising:an electric motor coupled to a drive shaft that extends to align a gear interface of the electric actuator with a variable geometry adjustment interface of the engine;a position feedback shaft that extends coaxially with respect to the drive shaft, wherein the position feedback shaft is coupled to an output shaft of the gear interface at a gear interface end of the position feedback shaft; anda rotational position sensor coupled to a motor end of the position feedback shaft proximate the electric motor, wherein the drive shaft and the position feedback shaft are sized to position an output ring gear of the output shaft in contact with the variable geometry adjustment interface within a casing of the engine and to further position the electric motor and the rotational position sensor external to the casing of the engine.2. The electric actuator according to claim 1 , further comprising a retracting mechanism configured to selectively retract the drive shaft and a portion of the gear interface to decouple the drive shaft from the ...

ADJUSTABLE TRANSITION SUPPORT AND METHOD OF USING THE SAME

A gas turbine engine including an adjustable transition assembly is disclosed. The gas turbine engine includes at least a combustor section and a turbine cylinder section, wherein the combustor section includes a transition piece coupled to a vane carrier assembly. The vane carrier assembly includes a plurality of holes that correspond with one or more holes of the transition piece, wherein at least one of the plurality of holes is adapted to receive a transition adjustable means therein for shifting the transition piece toward an upstream or downstream end. Also provided is a method for shifting a transition piece in a gas turbine engine, the method including at least the step of moving a transition adjustable means from a first position to a second position, where the movement shifts the transition piece in a direction towards an upstream end or a downstream end. 1. A gas turbine with an adjustable transition piece comprising:a gas turbine housing having at least a combustor section operatively connected to a turbine cylinder section;wherein said combustor section includes:a transition duct having an upstream end and a downstream end; anda transition support coupled with said transition duct proximate to said downstream end, said transition support having a plurality of recesses for coupling said transition support to a structure in a turbine section of a gas turbine and at least partially receiving a transition adjustable means therebetween;wherein said turbine cylinder section includes:a turbine support structure having a first surface and a second surface, wherein said first surface is coupled to said transition support via one or more recesses corresponding to one or more recesses of said transition support, and wherein said second surface includes one or more recesses extending at least partially through a thickness of said turbine support structure and adapted to at least partially receiving a transition adjustable means therebetween;said gas turbine further ...

GAS TURBINE ENGINE STATOR VANE BAFFLE ARRANGEMENT

A stator vane for a gas turbine engine includes an airfoil that has an exterior wall that provides a cooling cavity. The exterior surface has an interior surface that has multiple pin fins extending therefrom. A baffle is arranged in the cooling cavity and is supported by the pin fins. 1. A stator vane for a gas turbine engine comprising:an airfoil having an exterior wall providing a cooling cavity, the exterior surface has an interior surface having multiple pin fins extending therefrom; anda baffle arranged in the cooling cavity and supported by the pin fins.2. The stator vane according to claim 1 , wherein the baffle is sheet steel.3. The stator vane according to claim 2 , wherein the exterior wall provides pressure and suction sides joined at leading and trailing edges claim 2 , and the baffle includes impingement holes configured to provide impingement cooling fluid onto the exterior wall at the leading edge.4. The stator vane according to claim 2 , wherein the baffle includes a generally smooth outer contour free of protrusions.5. The stator vane according to claim 4 , wherein the outer contour is provided by plastically deformation.6. The stator vane according to claim 4 , wherein cooling holes are provided by at least one of drilling claim 4 , laser drilling claim 4 , or electro discharge machining.7. The stator vane according to claim 1 , wherein a perimeter cavity is provided between the baffle and the exterior wall claim 1 , the pin fins arranged in the perimeter cavity.8. The stator vane according to claim 7 , wherein the perimeter cavity circumscribes the baffle.9. The stator vane according to claim 8 , wherein the pin fins provide the sole support for the baffle in the perimeter cavity.10. The stator vane according to claim 1 , wherein the pin fins are arranged in rows.11. The stator vane according to claim 1 , wherein the pin fins are radially spaced from one another.12. The stator vane according to claim 1 , wherein a rib separates the cooling cavity ...

GAS TURBINE ENGINE THIN WALL COMPOSITE VANE AIRFOIL

An airfoil for a gas turbine engine has a first layer forming a cavity having transitioning from a first thickness to a second thickness through a ply drop region. A second layer is secured to the first layer. 1. An airfoil for a gas turbine engine comprising:a first layer forming a cavity having transitioning from a first thickness to a second thickness through a ply drop region; anda second layer secured to the first layer.2. The airfoil according to claim 1 , comprising a space arranged between the first and second layers claim 1 , and a filler is provided in the space.3. The airfoil according to claim 2 , wherein the second layer terminates in ends forming a V-shape at a trailing edge of the airfoil claim 2 , and the filler is provided between the first layer and second layer.4. The airfoil according to claim 3 , wherein the second thickness is provided at a location between the first thickness and the filler.5. The airfoil according to claim 2 , wherein the filler is provided near a leading edge of the airfoil.6. The airfoil according to claim 1 , wherein each layer includes multiple plies.7. The airfoil according to claim 6 , wherein the plies are constructed from a ceramic matrix composite bonded to one another by a resin.8. The airfoil according to claim 7 , wherein the ceramic matrix composite is a silicon carbide material.9. The airfoil according to claim 1 , wherein the airfoil is a vane.10. The airfoil according to claim 9 , wherein the vane is a mid turbine frame vane.11. The airfoil according to claim 9 , comprising a component passing through the cavity of the vane claim 9 , the component adjacent to the first thickness.12. The airfoil according to claim 9 , wherein a single cavity is provided in the airfoil.13. A method of forming an airfoil comprising:wrapping a first layer about a mandrel and building a thickened area with the first layer relative to an adjacent area of the first layer;applying a filler over the thickened area; andwrapping second ...

GUIDE VANE ASSEMBLY VANE BOX OF AN AXIAL TURBINE ENGINE COMPRESSOR

The invention relates to an angular sector of a bladed stator of a low-pressure compressor of an axial turbine engine. The sector comprises an outer shroud and an inner shroud in the form of circular arcs intended to be mounted in a concentric manner on the outer casing of the turbine engine compressor. The sector likewise comprises a row of stator vanes extending radially and anchored in the shrouds in such a manner as to form a bladed box. The vanes of the box comprise anchoring lugs at their outer ends, the lugs being disposed in the thickness of the outer shroud. The inner shroud comprises stubs for anchoring vanes. 1. An angular sector of a bladed stator of an axial turbine engine , said sector comprising:an arcuate segment of an outer shroud intended to be mounted on a casing of the turbine engine;an arcuate segment of inner shroud; and 'at least one anchoring portion of a box vane comprises an anchoring lug which mainly extends in the circumferential direction, and which is disposed in the thickness of one of the shrouds in such a manner as to anchor the vane to the shroud to make the box rigid.', 'a row of stator vanes extending radially from the outer shroud to the inner shroud, each of the stator vanes comprising an inner anchoring portion anchored to the inner shroud and an outer anchoring portion anchored to the outer shroud in such a manner that the stator vanes, the inner shroud and the outer shroud form a bladed box, wherein'}2. The angular sector in accordance with claim 1 , wherein each box vane comprises an airfoil extending between the shrouds in the radial direction claim 1 , the anchoring lugs extending perpendicularly to the radial direction and generally perpendicularly in respect of a chord of the associated vane.3. The angular sector in accordance with claim 1 , wherein at least one box vane comprises two lugs disposed at a same end claim 1 , the lugs being generally curved.4. The angular sector in accordance with claim 1 , wherein at least ...

GAS TURBINE ENGINE STATOR VANE PLATFORM COOLING

An airfoil component for a gas turbine engine includes a platform joined to an airfoil. The platform includes a flow path surface that extends between spaced apart lateral surfaces. The airfoil extends from the flow path surface. A contoured surface adjoins the flow path surface and one of the lateral surfaces. 1. An airfoil component for a gas turbine engine comprising:a platform joined to an airfoil, the platform includes a flow path surface extending between spaced apart lateral surfaces, the airfoil extends from the flow path surface, and a contoured surface adjoins the flow path surface and one of the lateral surfaces.2. The airfoil component according to claim 1 , comprising inner and outer platforms joined by the airfoil claim 1 , one of the inner and outer platforms providing the platform.3. The airfoil component according to claim 2 , wherein the platform is provided by an inner platform.4. The airfoil component according to claim 1 , comprising a cooling passage claim 1 , and cooling holes extend through the contoured surface and are in fluid communication with the cooling passage.5. The airfoil component according to claim 4 , wherein the contoured surface is at first and second angles with respect to the flow path surface and the lateral surface claim 4 , respectively claim 4 , the first and second angles are in the range of greater than 0° to 65°.6. The airfoil component according to claim 5 , wherein the contoured surface is curved.7. The airfoil component according to claim 5 , wherein the first and second angles are about 45°.8. The airfoil component according to claim 4 , wherein the exit of the cooling holes are directed aftward toward a trailing edge of the airfoil.9. The airfoil component according to claim 4 , comprising a thermal barrier coating provided on the inner flow path surface and the contoured surface claim 4 , the cooling holes extend through the thermal barrier coating.10. The airfoil component according to claim 1 , wherein a slot ...

FLEXIBLE FINGER SEAL FOR SEALING A GAP BETWEEN TURBINE ENGINE COMPONENTS

An assembly for a turbine engine includes a turbine engine first component, a turbine engine second component and a flexible seal that is attached to the first component. The flexible seal at least partially seals a gap between the first component and the second component. The flexible seal includes a mount and a finger seal that sealingly engages the second component. The mount includes a boss that sealingly engages the first component. 1. An assembly for a turbine engine , comprising:a turbine engine first component;a turbine engine second component; anda flexible seal attached to the first component, the flexible seal at least partially sealing a gap between the first component and the second component;wherein the flexible seal includes a mount and a finger seal that sealingly engages the second component, and the mount includes a boss that sealingly engages the first component.2. The assembly of claim 1 , further comprising a fastener that attaches the mount to the first component claim 1 , and extends through the boss.3. The assembly of claim 2 , whereinthe boss comprises a first boss, and the mount further includes a base and a second boss;the first boss and the second boss are arranged on opposing sides of the base; andthe fastener further extends through the base and the second boss.4. The assembly of claim 2 , wherein the mount further includes a base claim 2 , and the boss comprises a washer that is bonded to the base.5. The assembly of claim 2 , further comprising:a washer sealingly engaged between a flange and a first shoulder;wherein the first component includes the flange, and the fastener includes the first shoulder and a second shoulder that sealingly engages the mount; andwherein the fastener attaches the mount to the flange, and extends through mount, the flange and the washer between the first shoulder and the second shoulder.6. The assembly of claim 2 , further comprising:a second fastener that attaches the mount to the first component;wherein ...

UNDULATING STATOR FOR REDUCING THE NOISE PRODUCED BY INTERACTION WITH A ROTOR

A stator designed to be placed radially in a flow which passes through one or more rotors which share the same axis of rotation, with a leading edge and a trailing edge. The leading edge and trailing edge are connected by a lower face and an upper face, wherein at least one of the faces of the stator has radial undulations which extend axially from the leading edge to the trailing edge. The radial undulations can have at least two bosses in the same azimuth direction, the amplitude of which is at least one centimeter on at least part of the axial length of the stator. A propulsion assembly formed by the rotor and the stator, and to a turbine engine comprising such assembly is also provided. 1. Assembly comprising one or more rotors which share the same axis of rotation , and at least one stator which is designed to be placed radially in a flow which passes through said rotor(s) upstream or downstream thereof , said stator having a leading edge and a trailing edge , said leading edge and trailing edge being connected by a lower face and an upper face , wherein at least one of the faces of said stator has radial undulations which extend axially from the leading edge to the trailing edge , said radial undulations having at least two bosses in the same azimuth direction , the amplitude of which is at least one centimeter on at least part of the axial length of the stator , and in that , with the assembly being designed such that the crossing of said flow by the stator creates on said undulating surface pressure fluctuations with oscillations of the temporal phase according to the radial position , the radial undulations of said face have azimuth maximums and/or minimums in the vicinity of the zero mean dephasing regions for the pressure on the undulating face.2. Assembly according to claim 1 , wherein the radial undulations have a wavelength which is substantially constant along the radial extension of the stator.3. Assembly according to claim 1 , wherein the amplitude ...

TURBOCHARGER INTERNAL TURBINE HEAT SHIELD HAVING AXIAL FLOW TURNING VANES

A turbocharger internal heat shield () is provided having axial flow turning vanes (). Additionally, the heat shield may have a volute divider wall extender (). 112. A heat shield () for use in a turbocharger having an axial flow turbine wheel , having turning vanes () which turn the flow of exhaust gas to the axial direction.212. A heat shield () according to having 4 to 8 turning vanes ().312. A heat shield () according to having 7 turning vanes ().415. A heat shield () according to further comprising a volute divider wall extender (). 1. Field of the InventionThis invention relates to a turbocharger for an internal combustion engine. More particularly, this invention relates to turbocharger having an axial flow turbine wheel and an internal heat shield having turning vanes. Optionally, for a twin exhaust gas volute the heat shield may have a divider wall extender.2. Description of Related ArtA turbocharger is a type of forced induction system used with internal combustion engines. Turbochargers deliver compressed air to an engine intake, allowing more fuel to be combusted, thus boosting an engine's horsepower without significantly increasing engine weight. Thus, turbochargers permit the use of smaller engines that develop the same amount of horsepower as larger, normally aspirated engines. Using a smaller engine in a vehicle has the desired effect of decreasing the mass of the vehicle, increasing performance, and enhancing fuel economy. Moreover, the use of turbochargers permits more complete combustion of the fuel delivered to the engine, which contributes to the highly desirable goal of a cleaner environment.Turbochargers typically include a turbine housing connected to the engine's exhaust manifold, a compressor housing connected to the engine's intake manifold, and a center bearing housing coupling the turbine and compressor housings together. A turbine wheel in the turbine housing is rotatably driven by an inflow of exhaust gas supplied from the exhaust ...

Sealing arrangement for a turbomachine, a guide vane arrangement, and a turbomachine with such a sealing arrangement

The invention relates to a sealing arrangement , a guide vane arrangement, and a turbomachine with such a sealing arrangement , wherein the sealing arrangement is designed for a guide vane ring of a turbomachine , wherein the sealing arrangement comprises a thin-walled annular structure that is substantially closed on all sides, and wherein the annular structure delimits an annular interior space , wherein a hollow cell structure , which is designed so as to mechanically support the annular structure , is provided in the annular interior space 1156011. A sealing arrangement () for a guide vane ring () of a turbomachine () ,{'b': 15', '80, 'wherein the sealing arrangement () comprises a thin-walled annular structure () that is substantially closed on all sides, and'}{'b': 80', '105, 'wherein the annular structure () delimits an annular interior space (), wherein'}{'b': 109', '80', '105, 'a hollow cell structure (), which is designed to mechanically support the annular structure (), is provided in the annular interior space ().'}215109110110. The sealing arrangement () according to claim 1 , wherein the hollow cell structure () comprises a foam structure () claim 1 , wherein the foam structure () preferably comprises an open-pored foam claim 1 , particularly preferably claim 1 , an open-pored metallic foam claim 1 , as material.315109190190. The sealing arrangement () according to claim 1 , wherein the hollow cell structure () comprises a support structure () claim 1 , wherein the support structure () has a honeycomb design.41510920521020521080205215. The sealing arrangement () according to claim 1 , wherein the hollow cell structure () comprises at least two webs () and at least three nodal points () claim 1 , wherein the webs () are linked at the nodal points () to the annular structure () and/or the webs () are linked to one another to form a framework ().5158011513510515115135105. The sealing arrangement () according to claim 1 , wherein the annular structure () ...

TURBOMACHINE, CIRCULATION STRUCTURE AND METHOD

A turbomachine having at least one circulation structure is disclosed. The circulation structure has an annular space with baffle elements surrounding a main flow path and is open to the main flow path. A housing of the turbomachine to receive the circulation structure is divided in an axial plane into a front housing region and a rear housing region and the circulation structure is divided into a front structure region and a rear structure region in the axial plane of separation. A circulation structure divided into two parts in the axial direction and a method for the same is also disclosed. 1. A turbomachine , comprising:a housing, wherein in a first axial plane of separation, the housing is divided into a front housing region and a rear housing region; anda circulation structure accommodated in the housing, wherein the circulation structure has an annular space with baffle elements, wherein the annular space extends around a main flow path and is open to the main flow path, and wherein the circulation structure, in a second axial plane of separation, is divided into a front structure region and a rear structure region.2. The turbomachine according to claim 1 , wherein the first axial plane of separation and the second axial plane of separation are a same axial plane of separation in an assembled state.3. The turbomachine according to claim 1 , wherein the baffle elements are formed in the front structure region and wherein the front structure region is inserted into the front housing region.4. The turbomachine according to claim 1 , wherein the baffle elements are inserted directly into the front housing region.5. The turbomachine according claim 1 , wherein the rear structure region is an integral front body section of a closing element inserted into the rear housing region.6. A circulation structure for a turbomachine comprising a front structure region and a rear structure region in an axial plane of separation.7. The circulation structure according to claim ...

TURBINE ENGINE WHEEL

The invention relates to a turbine wheel () comprising a plurality of vanes connected to an annular platform () carrying annular lips (). According to the invention, one of the upstream lip () and the downstream lip () is of a first type or of a second type, with the first type corresponding to one lip () having the upstream face () which is concave curved and the downstream face () which is convex curved and the second type corresponding to a lip () having the upstream () and downstream () faces which are substantially flat and mutually parallel. 1. A wheel of a turbine engine comprising a plurality of radially extending vanes , one radially internal or external end of which is connected to an annular platform carrying annular lips extending from said platform in a direction opposite the vane between a first radial end connected to the platform and a second opposed radial free end , in order to sealingly cooperate with a radially facing ring , wherein one of the upstream lip and of the downstream lip is of a first type or of a second type , with the first type corresponding to one lip having the upstream face which is concave curved and the downstream face which is convex curved and the second type corresponding to a lip having the upstream and downstream faces which are substantially flat and mutually parallel , with the first end of said lip being arranged downstream of the second end for both the first type and the second type.2. A wheel according to claim 1 , wherein said lip is arranged at the downstream end of said platform.3. A wheel according to claim 2 , wherein said lip is of the first type.4. A wheel according to claim 3 , further comprising another annular lip arranged at the upstream end of the platform claim 3 , with the other lip being of the second type.5. A wheel according to claim 1 , wherein the generatrix of the cone of revolution going through the first end and the second end of the lip is inclined relative to a plane perpendicular to the axis ...

VANE ANGLE SYSTEM ACCURACY IMPROVEMENT

A stator vane angle system includes an engine case, a plurality of stator vanes located at an interior of the engine case. Each stator vane is rotatable about a stator vane axis. A synchronization ring is located at an exterior of the engine case. The synchronization ring is operably connected to each stator vane of the plurality of stator vanes such that movement of the synchronization ring urges rotation of each stator vane of the plurality of stator vanes about their respective stator vane axes. A plurality of impingement openings extend through the engine case from the interior of the engine case to the exterior of the engine case. The plurality of impingement openings are configured to direct flowpath gases from the interior of the engine case to impinge on the synchronization ring, thereby reducing a thermal mismatch between the engine case and the synchronization ring. 1. A stator vane angle system , comprising:an engine case;a plurality of stator vanes disposed at an interior of the engine case, each stator vane rotatable about a stator vane axis;a synchronization ring disposed at an exterior of the engine case, the synchronization ring operably connected to each stator vane of the plurality of stator vanes such that movement of the synchronization ring urges rotation of each stator vane of the plurality of stator vanes about their respective stator vane axes; anda plurality of impingement openings extending through the engine case from the interior of the engine case to the exterior of the engine case, the plurality of impingement openings configured to direct flowpath gases from the interior of the engine case to impinge on the synchronization ring.2. The stator vane angle system of claim 1 , wherein the plurality of impingement openings each have an impingement opening outlet disposed at a same axial location as the synchronization ring.3. The stator vane angle system of claim 1 , wherein the plurality of impingement openings each extend perpendicular to ...

AIRFOIL ASSEMBLY WITH A SCALLOPED FLOW SURFACE

A stage for a compressor or a turbine in a turbine engine can include an annular row of airfoils radially extending from corresponding platforms, where each platform can include a fore edge and aft edge and each airfoil can include a leading edge and trailing edge. At least one of the platforms can have a scalloped flow surface including a bulge and a trough. 1. A stage for at least one of a compressor or a turbine , the stage comprising:an annular row of airfoils radially extending from corresponding platforms, the airfoils circumferentially spaced apart to define intervening flow passages;each platform having a fore edge and an aft edge;each airfoil having an outer wall defining a pressure side and a suction side opposite the pressure side, the outer wall extending axially between a leading edge and a trailing edge defining a chord-wise direction, and the outer wall extending radially between a root and a tip defining a span-wise direction, with the root adjacent the platform and the leading edge aft of the fore edge of the platform; and the bulge having a portion extending forward of the fore edge and a local maximum located aft of the fore edge and spaced from the pressure side to define a bulge flow channel between the bulge and the pressure side, and', 'the trough extending adjacent at least a portion of the suction side with a fore portion of the trough located in front of the leading edge., 'at least one of the platforms having a scalloped flow surface including a bulge adjacent the pressure side and a trough adjacent the suction side,'}2. The stage of further comprising a fillet extending between the pressure side and the platform and located between the pressure side and the bulge.3. The stage of wherein the fillet extends between the suction side and the platform and is located between the suction side and the trough.4. The stage of wherein the fillet extends about the periphery of the outer wall.5. The stage of wherein the fore portion of the trough is ...

TURBINE NOZZLE COMPONENTS HAVING REDUCED FLOW AREAS

Embodiments of a method for controllably reducing of the flow area of a turbine nozzle component are provided, as are embodiments of turbine nozzle components having reduced flow areas. In one embodiment, the method includes the steps of obtaining a turbine nozzle component having a plurality of turbine nozzle flow paths therethrough, positioning braze preforms in the plurality of turbine nozzle flow paths and against a surface of the turbine nozzle component, and bonding the braze preforms to the turbine nozzle component to achieve a controlled reduction in the flow area of the turbine nozzle flow paths. 1. A turbine nozzle component , comprising:an inner endwall;an outer endwall radially spaced from the inner endwall;a plurality of nozzle vanes extending between the inner and outer endwalls;a plurality of turbine nozzle flow paths extending through the turbine nozzle component and generally defined by the inner endwall, the outer endwall, and the plurality of nozzle vanes; andbraze preforms positioned in the turbine nozzle flow paths and bonded to at least one of the inner endwall and outer endwall reducing the flow area of the turbine nozzle flow paths.2. The turbine nozzle component of wherein the plurality of nozzle vanes have leading and trailing edges around which the braze preforms wrap.3. The turbine nozzle component of wherein the braze preforms are further welded to at least one of the inner endwall and outer endwall.4. The turbine nozzle component of further comprising inner inter-blade flow areas provided on the inner endwall and bounding the plurality of flow paths claim 1 , the plurality of braze preforms having planform geometries substantially conformal with the inner inter-blade flow areas.5. The turbine nozzle component of further comprising outer inter-blade flow areas provided on the outer endwall and bounding the plurality of flow paths claim 1 , the plurality of braze preforms having planform geometries substantially conformal with the outer ...

AERO-ACTUATED VANES

A turbomachinery vane includes a vane body defining a longitudinal axis, a trunnion extending from the vane body and defining a pivot point for pivoting the vane body about the longitudinal axis, and a lock system operatively connected to the trunnion and configured to lock the vane body in a plurality of locked positions. A gas turbine engine includes a turbomachinery component including a row of actuated stators, wherein the actuated stator row includes a plurality of the turbomachinery vanes. A method of actuating a vane by aerodynamic loads includes moving the vane about a pivot point from a first position to a second position by a first set of by aerodynamic loads. 1. A turbomachinery vane comprising:a vane body defining a longitudinal axis; anda trunnion extending from the vane body and defining a pivot point for pivoting the vane body about the longitudinal axis.2. A vane as recited in claim 1 , further comprising a lock system operatively connected to the trunnion and configured to lock the vane body in a plurality of locked positions.3. A vane as recited in claim 2 , wherein the vane body includes a leading edge claim 2 , an opposed trailing edge claim 2 , a high pressure side claim 2 , and an opposed low pressure side claim 2 , and the trunnion is located at a position relative to the leading edge claim 2 , the trailing edge claim 2 , the high pressure side claim 2 , and the low pressure side for aerodynamic loads to pivot the vane body.4. A vane as recited in claim 3 , wherein the trunnion is located at a position relative to the leading edge claim 3 , the trailing edge claim 3 , the high pressure side claim 3 , and the low pressure side for a first set of aerodynamic loads pivot the vane body from a first locked position to the second locked position and a second set of aerodynamic loads pivot the vane body from the second locked position to a first locked position.5. A vane as recited in claim 1 , wherein the lock system is configured to release the ...

VANE ASSEMBLY FOR AN UNDUCTED THRUST PRODUCING SYSTEM

A unshrouded vane assembly for an unducted propulsion system includes a plurality of vanes which have non-uniform characteristics configured to generate a desired vane exit swirl angle. 1. A unshrouded vane assembly for an unducted propulsion system , comprising a vane assembly having a plurality of vanes which have non-uniform characteristics with respect to one another configured to generate a desired vane exit swirl angle.2. The vane assembly of claim 1 , wherein said vanes have a non-uniform characteristic selected from the group consisting of: camber claim 1 , stagger claim 1 , circumferential spacing claim 1 , axial position claim 1 , span claim 1 , tip radius claim 1 , and combinations thereof.3. The vane assembly of claim 2 , wherein said vanes have a root claim 2 , a tip claim 2 , and a span therebetween claim 2 , and wherein said non-uniform characteristic is attributed to a portion of the span of said vanes.4. The vane assembly of claim 3 , wherein said non-uniform characteristic is attributed to substantially all of the span of said vanes.5. The vane assembly of claim 1 , wherein said vanes are variable in pitch.6. The vane assembly of claim 5 , wherein said vanes are individually variable in pitch.7. The vane assembly of claim 5 , wherein a plurality of said vanes are variable in pitch in conjunction with one another.8. An unducted propulsion system claim 5 , said propulsion system comprising a rotating element having an axis of rotation and a stationary element claim 5 , said rotating element having a plurality of blades each having a blade root proximal to said axis claim 5 , a blade tip remote from said axis claim 5 , and a blade span measured between said blade root and said blade tip claim 5 , wherein said stationary element has a plurality of vanes each having a vane root proximal to said axis claim 5 , a vane tip remote from said axis claim 5 , and a vane span measured between said vane root and said vane tip configured to impart a change in ...

RING STATOR

A stator assembly for a gas turbine engine includes an annular outer shroud, an annular inner shroud radially spaced from the outer shroud and a plurality of stator vanes extending from the outer shroud to the inner shroud. A volume of potting is located at the inner shroud and at the outer shroud to retain the plurality of stator vanes thereat. A stator and case assembly for a gas turbine engine includes a case defining a working fluid flowpath for the gas turbine engine and a stator assembly located at the case. The stator assembly includes an annular outer shroud secured to the case, an annular inner shroud secured to the case and a plurality of stator vanes extending from the outer to the inner shroud. A volume of potting is located at the inner shroud and at the outer shroud to retain the plurality of stator vanes thereat. 1. A stator assembly for a gas turbine engine , comprising:an annular outer shroud;an annular inner shroud radially spaced from the outer shroud;a plurality of stator vanes extending from the outer shroud to the inner shroud; anda volume of potting disposed at the inner shroud and at the outer shroud to retain the plurality of stator vanes thereat.2. The stator assembly of claim 1 , wherein each stator vane of the plurality of stator vanes includes:an airfoil portion;an outer leg extending radially outwardly from the airfoil portion; andan inner leg extending radially inwardly from the airfoil portion.3. The stator assembly of claim 2 , wherein:the outer leg is installed into an outer shroud opening in the outer shroud; andthe inner leg is installed into an inner shroud opening in the inner shroud.4. The stator assembly of claim 3 , wherein the potting comprises:an outer grommet disposed at each outer shroud opening; andan inner grommet disposed at each inner shroud opening to retain each stator vane thereat.5. The stator assembly of claim 2 , wherein each stator vane further includes:an outer leg opening; andan inner leg opening;wherein a ...

Segmented Stator Assembly

A stator assembly for a gas turbine engine includes an arcuate outer shroud, an arcuate inner shroud radially spaced from the outer shroud and a plurality of stator vanes extending from the outer shroud to the inner shroud. A volume of potting is located at the inner shroud and at the outer shroud to retain the plurality of stator vanes thereat. A stator and case assembly includes a case defining a working fluid flowpath and a stator assembly positioned at the case. The stator assembly includes a plurality of stator segments arranged circumferentially about an engine axis, each stator segment including an arcuate outer shroud secured to the case, an arcuate inner shroud, and a plurality of stator vanes extending from the outer to inner shroud. A volume of potting is located at the inner shroud and at the outer shroud to retain the plurality of stator vanes thereat. 1. A stator assembly for a gas turbine engine , comprising:an arcuate outer shroud;an arcuate inner shroud radially spaced from the outer shroud;a plurality of stator vanes extending from the outer shroud to the inner shroud; anda volume of potting disposed at the inner shroud and at the outer shroud to retain the plurality of stator vanes thereat.2. The stator assembly of claim 1 , wherein each stator vane of the plurality of stator vanes includes:an airfoil portion;an outer leg extending radially outwardly from the airfoil portion; andan inner leg extending radially inwardly from the airfoil portion.3. The stator assembly of claim 2 , wherein:the outer leg is installed into an outer shroud opening in the outer shroud; andthe inner leg is installed into an inner shroud opening in the inner shroud.4. The stator assembly of claim 3 , wherein the potting comprises:an outer grommet disposed at each outer shroud opening; andan inner grommet disposed at each inner shroud opening to retain each stator vane thereat.5. The stator assembly of claim 2 , wherein each stator vane further includes:an outer leg opening ...

METHOD OF DESIGNING A COMMON PREFORM FOR PROVIDING PREFORMS FOR TURBINE ENGINE OUTLET GUIDE VANES OF DIFFERENT GEOMETRICAL PROFILES AND MADE OUT OF COMPOSITE MATERIAL

A method of designing a common preform for providing a plurality of preforms for turbine engine outlet guide vanes of different geometrical profiles and made out of composite material, the method including 3D geometrical modeling of the geometrical profiles of the different outlet guide vane preforms, flattening out each of the geometrical profiles of the different outlet guide vane preforms, superposing the geometrical profiles of the different flattened outlet guide vane preforms, and converging towards a single geometrical profile for an outlet guide vane preform common to all of the outlet guide vane preforms while guaranteeing identical positioning for the top and bottom regions of non-interlinking for all of the outlet guide vane preforms. 1. A method of designing a common preform for providing a plurality of preforms for turbine engine outlet guide vanes of different geometrical profiles and made out of composite material , the common preform being made from a single-piece woven fiber structure obtained by three-dimensional weaving , said fiber structure having first and second portions , each comprising a plurality of layers of yarns and each forming a fraction of the thickness of the fiber structure , the portions of the fiber structure being interlinked while leaving top and bottom regions of non-interlinking separating the first and second portions over respective fractions of the dimension of the fiber structure from opposite edges of the fiber structure , the method being performed by graphics software of a workstation and comprising:3D geometrical modeling of the geometrical profiles of the different outlet guide vane preforms;flattening out each of the geometrical profiles of the different outlet guide vane preforms;superposing the geometrical profiles of the different flattened outlet guide vane preforms; andconverging towards a single geometrical profile for an outlet guide vane preform common to all of the outlet guide vane preforms while ...

Turbomachine, Sealing Segment, and Guide Vane Segment

A turbomachine includes a sealing segment ring that is provided between a front guide vane row and a back guide vane row for sealing a radial gap between a casing section and a rotor blade row rotating between the guide vane rows, wherein the sealing segment ring has a plurality of identical sealing segments and at least one of the guide vane rows has a plurality of identical guide vane segments, wherein the sealing segments each have a plurality of engagement sites lying adjacent to one another in the peripheral direction for interaction with securing elements of this guide vane row, wherein the engagement sites and securing elements are distributed uniformly over the periphery and the engagement sites are a multiple of the securing elements, a sealing element, and a guide vane segment. 111246108461228414285056584645056584650565846. A turbomachine () with a sealing segment ring () between a front guide vane row () and a back guide vane row () for sealing a radial gap between a casing section () and a rotor blade row () rotating between the guide vane rows ( , ) , with the sealing segment ring () having a plurality of identical sealing segments () , and at least one of the guide vane rows () having a plurality of identical guide vane segments () , wherein the sealing segments () each have a plurality of engagement sites ( , , ) lying adjacent to one another in the peripheral direction for interaction with securing elements () of the guide vane row () , with the engagement sites ( , , ) and securing elements () being distributed uniformly over the periphery and the engagement sites ( , , ) being a multiple of the securing elements ().214446. The turbomachine according to claim 1 , wherein each guide vane segment () of the guide vane row () has only one securing element ().3285056585056584628144. The turbomachine according to claim 1 , wherein each sealing segment () has an identical plurality of engagement sites ( claim 1 , claim 1 , ) and the engagement sites ( ...

METHOD OF PRODUCING AN INSULATION ELEMENT AND INSULATION ELEMENT FOR A HOUSING OF AN AERO ENGINE

The invention relates to a process for producing an insulation element (), which can be arranged radially above at least one guide vane () in a housing () of a thermal gas turbine. The insulation element () is produced from a solid body () provided with a metallic shell (), the solid body () consisting at least partially of a ceramic material. The invention also relates to an insulation element (), which can be arranged radially above at least one guide vane () in a housing () of a thermal gas turbine, and to an aero engine having a housing (), in which at least one insulation element () is arranged radially above at least one guide vane (). 114.-. (canceled)15. A process for producing an insulation element which can be arranged radially above at least one guide vane in a housing of a thermal gas turbine , wherein the process comprises producing the insulation element from a solid body provided with a metallic shell , the solid body consisting at least partially of at least one ceramic material.16. The process of claim 15 , wherein the solid body is firstly produced from the at least one ceramic material and then enclosed with the metallic shell.17. The process of claim 15 , wherein firstly part of the metallic shell is provided and filled with the at least one ceramic material and claim 15 , after the solid body has been formed from the at least one ceramic material claim 15 , the metallic shell is completed.18. The process of claim 15 , wherein the solid body is produced at least partially from at least one of ceramic foam claim 15 , ceramic bodies claim 15 , ceramic powder or ceramic sheets and/or by at least one of sintering and integral bonding of ceramic particles.19. The process of claim 15 , wherein the at least one ceramic material is provided with at least one of an organic claim 15 , organometallic claim 15 , metallic or ceramic binder.20. The process of claim 15 , wherein the metallic shell is produced at least partially by at least one of a generative ...

SUBSONIC SHOCK STRUT

A gas turbine engine strut is disclosed as having a will forebody positioned upstream of a point of maximum thickness and an aft body positioned downstream of the point of maximum thickness. The aft body includes a discontinuity in a curvature distribution which provides for a “subsonic shock.” The discontinuity in curvature distribution can include in inflection point that marks a transition from a curvature associated with an upstream portion of the aft body to a second curvature associated with a downstream portion of the aft body. In some forms, the aft body can additionally include boundary layer aspiration. The gas turbine engine strut can be symmetrical about a centerline. 1. An apparatus comprisinga gas turbine engine subsonic strut having an airfoil shape structured to be disposed in a flow path having subsonic flow and including a forebody located forward of a maximum thickness of the strut and an aft body located aft of the maximum thickness, the aft body of the gas turbine engine subsonic strut having an inflection point that produces a subsonic shock pressure recovery.2. The apparatus of claim 1 , wherein the inflection point is structured to initially encourage flow separation while the aft body is structured to provide an adequate length to reattach the flow while still decelerating the flow.3. The apparatus of claim 1 , wherein the forebody upstream of a maximum thickness is configured to exert relatively little acceleration on a working fluid passing over the gas turbine engine subsonic strut.4. The apparatus of claim 1 , wherein the forebody is axially longer than the aft body.5. The apparatus of claim 1 , wherein the subsonic strut is one of a diffuser strut and a bypass duct strut claim 1 , and wherein the subsonic strut is symmetric about a plane of symmetry.6. The apparatus of claim 1 , wherein the aft body is shaped to suppress a growth in shape factor.7. The apparatus of claim 1 , wherein the inflection point is a discontinuous change in the ...

Guide vane arragement and method for mounting a guide vane

The invention relates to a guide vane arrangement and a method for mounting a guide vane. The guide vane arrangement has at least one guide vane having a radially inner journal, which engages in a borehole of a bushing. The bushing is disposed in a borehole of an inner ring segment. The journal is connected to the bushing in a rotation-resistant manner and a slide mounting is provided between the bushing and the inner ring borehole. For mounting the guide vane in an inner ring segment, the journal of the guide vane is introduced into the borehole of the bushing. Further, the journal is fastened to the bushing in a rotation-resistant manner, and the bushing is slide-mounted with the inner ring segment. The slide mounting is conducted prior to or simultaneously with or after the fastening of the journal to the bushing in a rotation-resistant manner. 1110111321113321113. A guide vane arrangement having at least one guide vane () with a radial inner journal () , which engages in a borehole of a bushing () , which is disposed in a borehole of an inner ring segment () , wherein a rotation-resistant connection is present between the journal () and the bushing () and a slide mounting is present between the bushing () and the borehole of the inner ring segment () , wherein the journal () has an outer thread and the bushing () has an inner thread screwed therewith providing a screw connection.211131113. The guide vane arrangement according claim 1 , further comprising a threaded sleeve disposed between the journal () and the bushing () claim 1 , and the journal () and the bushing () being connected to the threaded sleeve.3111. The guide vane arrangement according to claim 1 , wherein the journal () has an outer thread claim 1 , which is screwed with an inner thread of a nut.4. The guide vane arrangement according to claim 1 , further comprising a screw locking device for securing the screw connection.53378. The guide vane arrangement according to claim 4 , wherein the screw ...

GAS TURBINE COMPRESSOR STATOR VANE ASSEMBLY

A stator vane assembly for a compressor of a gas turbine, in particular of an aircraft engine, including a plurality of stator vanes whose airfoil sections form a stagger angle with an axis of rotation of the compressor, which stagger angle varies along a duct height of the stator vane assembly. Along the duct height from the inside to the outside, the stagger angle increases to a local maximum in a second section adjoining a first, radially innermost section, and decreases to an outer local minimum in a third section adjoining this second section and, along the duct height from the inside to the outside, the stagger angle decreases from the initial value to an inner local minimum in the first, radially innermost section and/or increases from the outer local minimum to a final value in a fourth, radially outermost section adjoining the third section. 1. A stator vane assembly for a compressor of a gas turbine comprising:a plurality of stator vanes whose airfoil sections form a stagger angle with an axis of rotation of the compressor, the stagger angle varying along a duct height of the stator vane assembly,wherein along the duct height between a radially innermost initial value and a radially outermost final value, the stagger angle increases to a local maximum in a second section adjoining a first, radially innermost section, and decreases to an outer local minimum in a third section adjoining said second section, andwherein, along the duct height from the inside to the outside, the stagger angle decreases from the initial value to an inner local minimum in the first, radially innermost section or increases from the outer local minimum to the final value in a fourth, radially outermost section adjoining the third section.2. The stator vane assembly as recited in wherein the stagger angle decreases monotonically in the first or third section or increases monotonically in the second or fourth section.3. The stator vane assembly as recited in wherein the monotonical ...

APPARATUS FOR GENERATING ENERGY

An apparatus for generating energy through fluid dynamics includes a fluid reservoir, an energy extractor for extracting flow energy, a back-pressure control channel for circulating the fluid, and a pressure ejector for returning fluid to the fluid reservoir. The back-pressure control channel includes a fan-like device to generate a low-pressure region and draws fluid through the energy extractor. The energy extractor includes an energy extraction rotor to convert flow energy to rotation energy and may include a nozzle to alter flow characteristics of the fluid. The apparatus for generating energy may also include a settlement chamber to reduce flow disturbances. 1. An apparatus for generating energy , comprising:a reservoir comprising a fluid;an energy extractor connected to the reservoir, wherein the energy extractor comprises an energy extraction rotor for extracting energy from the fluid;a back-pressure control channel coupled to the energy extractor, wherein the back-pressure control channel comprises a fan-like device configured to circulate the fluid and to form a low-pressure region within the back-pressure control channel; anda pressure ejector coupled to the back-pressure control channel, wherein the pressure ejector is configured to return the fluid to the reservoir.2. The apparatus for generating energy of claim 1 , wherein the back-pressure control channel comprises a first back-pressure control channel claim 1 , wherein the fan-like device comprises a first fan-like device claim 1 , wherein the low-pressure region comprises a first low-pressure region claim 1 , wherein the apparatus for generating energy comprises a second back-pressure control channel coupled to the energy extractor claim 1 , wherein the second back-pressure control channel comprises a second fan-like device configured to circulate the fluid claim 1 , and wherein the second fan-like device is further configured to form a second low-pressure region within the second back-pressure ...

A TECHNIQUE FOR COOLING A ROOT SIDE OF A PLATFORM OF A TURBOMACHINE PART

A platform cooling device directs a cooling fluid onto a root side of a platform of a turbomachine part. The platform cooling device includes a first segment to be positioned at a root of the turbomachine part and a second segment, at an angle to the first segment, to be positioned at the root side of the platform of the turbomachine part. The second segment may include at least one impingement channel having an inlet for receiving at least a part of the cooling fluid and an outlet for releasing the received cooling fluid onto the root side of the platform. The first segment and the second segment may define a path for the cooling fluid via the impingement channel. A turbomachine component includes the platform cooling device.

METHOD AND APPARATUS FOR ATTACHING A TRANSITION DUCT TO A TURBINE SECTION IN A GAS TURBINE ENGINE

An apparatus for attaching a transition duct to a turbine section of a gas turbine engine includes an annular shaped vane carrier that engages with a transition aft frame. The transition aft frame includes an outwardly extending clamp that is at least partially inserted into a radial recess of the vane carrier structure. A spacer is inserted into the recess. A first pair of mating surfaces is formed by a first end face of the spacer and an end face of the clamp and a second pair of mating surfaces is formed by a second end face of the spacer and a wall of the recess. At least one of the pairs of mating surfaces is inclined at an angle with respect to a radial plane of the vane carrier structure. A retention device secures the spacer to the vane carrier structure.

Off-Cambered Vanes for Gas Turbine Engines

An off-cambered vane for a guide vane assembly in a gas turbine engine is described. The guide vane assembly may comprise a nominal vane having a tip portion, a mid-span portion, and a hub portion. The mid-span portion of the nominal vane may adopt a nominal geometry and the hub portion of the nominal vane may adopt a common geometry. The guide vane assembly may further comprise an off-cambered vane having a tip portion, a mid-span portion, and a hub portion. The mid-span portion of the off-cambered vane may deviate variably with respect to the nominal geometry and at least one of the hub portion and the tip portion may adopt the common geometry.

AXIAL FLOW ROTATING MACHINE AND DIFFUSER

An axial flow rotating machine having: a rotor that is provided with a plurality of rotor blades; a stator provided with a plurality of stator blades; an axial flow rotating portion formed by the rotor and the stator; and a diffuser connected to the axial flow rotating portion on the downstream side of the axial flow rotating portion. The final blade portion inner-circumferential inner wall, which is a portion of the inner-circumferential inner wall of the axial flow rotating portion, is formed so that the diameter thereof at the trailing edge position of the final blade is smaller than the diameter at the leading edge position of the final blade. In addition, the diameter of all or a portion of the diffuser inner-circumferential inner wall decreases in a direction of a first side in the axial direction, the first side being the downstream side.

VANE TIP MACHINING FIXTURE ASSEMBLY

A vane tip machining fixture assembly includes a first clamp ring mountable to a fixture, the first clamp ring includes a rigid back ring mounted to a resilient ring.

RIFFLED SEAL FOR A TURBOMACHINE, TURBOMACHINE AND METHOD OF MANUFACTURING A RIFFLED SEAL FOR A TURBOMACHINE

A seal of a turbomachine reduces a leakage flow between a first and second component of the turbomachine. The first component has a first surface and the second component has a second surface, wherein the first component is stiff with regard to a first force exerted perpendicularly thereto and the second component is stiff with regard to a second force exerted perpendicularly thereto. The first surface is opposite the second surface, together defining boundaries of a fluid passage for the leakage flow. The first surface has a first surface riffle. A turbomachine has a seal described above, wherein the turbomachine is a gas turbine engine. A method of manufacturing a first component of a turbomachine with a reduced leakage flow between the first component and a second component of the turbomachine includes fabrication of a first surface riffle, in particular by grinding and/or by electrical discharge machining.

Optimized aerodynamic profile for a turbine vane, in particular for a nozzle of the sixth stage of a turbine

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P0 situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P1. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine.

OPTIMIZED AERODYNAMIC PROFILE FOR A TURBINE VANE, IN PARTICULAR FOR A NOZZLE OF THE FOURTH STAGE OF A TURBINE

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine. 1. An aerodynamic profile for a turbine vane , the profile being , when cold and in a non-coated state , substantially identical to a nominal profile determined by the Cartesian coordinates X ,Y ,Zadim given in Table 1 , in which the coordinate Zadim is the quotient D/H , where D is the distance of the point under consideration from a reference X ,Y plane situated at the base of the nominal profile , and H is the height of said profile measured from said reference plane out to the end of the vane , the measurements D and H being taken radially relative to the axis of the turbine , while the coordinate X is measured in the axial direction of the turbine.2. The aerodynamic profile as claimed in claim 1 , wherein said profile is defined within an envelope of ±1 mm in a direction normal to the surface of the nominal profile.3. The aerodynamic profile as claimed in claim 1 , wherein the coordinates X claim 1 ,Y of said profile lie within a range of ±5% relative to the coordinates X claim 1 ,Y of the nominal profile.4. The aerodynamic profile as claimed in claim 1 , wherein the vane is a nozzle vane forming a part of a stator of a turbine.5. The aerodynamic profile as claimed in claim 4 , wherein the vane is a nozzle vane of the fourth stage of the turbine.6. The aerodynamic profile as claimed in claim 4 , wherein the vane is a vane of the fourth stage nozzle ...

OPTIMIZED AERODYNAMIC PROFILE FOR AN ARM OF A STRUCTURAL CASING OF A TURBINE, AND STRUCTURAL CASING HAVING SUCH AN ARM

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine. 1. An aerodynamic profile for an arm of a structural casing of a turbine having a central hub and a shroud , the arm connecting the central hub and the shroud , the profile being , when cold and in a non-coated state , substantially identical to a nominal profile determined by the Cartesian coordinates X ,Y ,Zadim given in Table 1 , in which the coordinate Zadim is the quotient D/H , where D is the distance of the point under consideration from a first reference X ,Y plane situated at the base of the nominal profile , and H is the height of said profile measured from said first reference plane that is the intersection of the stacking axis of the arm and the axisymmetric surface of the hub , out to a second reference plane that is the intersection of said stacking axis with the axisymmetric surface of the shroud , the measurements D and H being taken radially relative to the axis of the turbine , while the coordinate X is measured in the axial direction of the turbine.2. The aerodynamic profile as claimed in claim 1 , wherein said profile is defined within an envelope of ±1 mm in a direction normal to the surface of the nominal profile.3. The aerodynamic profile as claimed in claim 1 , wherein the coordinates X claim 1 ,Y of said profile lie within a range of ±5% relative to the coordinates X claim 1 ,Y of the nominal profile.4. The aerodynamic profile ...

Housing structure with improved seal and cooling

A housing structure for a turbo-engine, especially for a gas turbine or an aircraft engine, includes an outer housing wall and an inner housing wall, the inner and outer housing walls annularly enclosing a flow channel for the turbo-engine and being spaced apart radially from the flow channel. At least one heat shield is arranged between the inner and outer housing walls, and a bar or a fixture projects at least somewhat radially from the inner housing wall. The bar or the fixture has on at least one side a broadening element that includes a sealing face against which the heat shield is positioned in a sealing manner.

TURBOMACHINE BLADE

The present invention relates to a turbomachine blade () with a base element, which has a blade part () for flow diversion and a blade root (), a first guide (), fixed on the base element, in which a first element () is movably guided, and a second guide (), fixed on the base element, in which a second element () is movably guided, wherein a dynamic of the first element in the first guide and a dynamic of the second element in the second guide are designed differently, and the first guide () is arranged in a half, nearer to the base root, of a radial height (H) of the base element, and the second guide is arranged in a half, more remote from the blade root, of the radial height of the base element. 1100200. A turbomachine blade (; ) having:{'b': 1', '2, 'a base element, which has a blade part () for flow diversion and a blade root ();'}{'b': 110', '210', '111', '211, 'a first guide (; ), fixed on the base element, in which a first element (; ) is movably guided; and'}{'b': 120', '220', '121', '221, 'a second guide (; ), fixed on the base element, in which a second element (; ) is movably guided;'}wherein a dynamic of the first element in the first guide and a dynamic of the second element in the second guide are designed differently; {'b': 110', '210, 'the first guide (; ) is arranged in a half, nearer to the blade root, of a radial height (H) of the base element, and the second guide is arranged in a half, more remote from the blade root, of the radial height of the base element.'}, 'wherein'}23434. The turbomachine blade according to claim 1 , wherein the first and/or second guide is arranged on a shroud (; ) that is radially adjacent to the blade part and/or on a side claim 1 , facing away from the blade part claim 1 , of a shroud (; ) that is radially adjacent to the blade part.3. The turbomachine blade according to claim 1 , wherein a main movement direction of the second guide claim 1 , in which the second element exhibits maximum play (s) claim 1 , forms an ...

METHODS FOR MANUFACTURING A TURBINE NOZZLE WITH SINGLE CRYSTAL ALLOY NOZZLE SEGMENTS

Methods for manufacturing a turbine nozzle are provided. A plurality of nozzle segments is formed. Each nozzle segment comprises an endwall ring portion with at least one vane. The plurality of nozzle segments are connected to an annular endwall forming a segmented annular endwall concentric to the annular endwall with the at least one vane of each nozzle segment extending between the segmented annular endwall and the annular endwall. 1. A method for manufacturing a turbine nozzle comprising:forming a plurality of nozzle segments, each nozzle segment comprising an endwall ring portion with at least one vane; andconnecting the plurality of nozzle segments to an annular endwall forming a segmented annular endwall concentric to the annular endwall with the at least one vane of each nozzle segment extending between the segmented annular endwall and the annular endwall.2. The method of claim 1 , wherein the step of forming a plurality of nozzle segments comprises forming the plurality of nozzle segments with a single crystal material.3. The method of claim 1 , wherein the step of forming a plurality of nozzle segments comprises forming by casting.4. The method of claim 1 , wherein the step of forming a plurality of nozzle segments comprises forming a plurality of singlet nozzle segments claim 1 , doublet nozzle segments claim 1 , triplet nozzle segments claim 1 , quadruplet nozzle segments claim 1 , or combinations thereof.5. The method of claim 1 , further comprising the step of processing at least one nozzle segment of the plurality of nozzle segments prior to the connecting step claim 1 , wherein the step of processing comprises applying a protective coating to at least one nozzle segment of the plurality of nozzle segments claim 1 , the annular endwall claim 1 , or both.6. The method of claim 1 , further comprising the step of processing at least one nozzle segment of the plurality of nozzle segments prior to the connecting step claim 1 , wherein the step of ...

GAS TURBINE ENGINE WITH LOW STAGE COUNT LOW PRESSURE TURBINE

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, a first compressor section including three (3) or more stages, a second compressor section including between eight (8) and thirteen (13) stages, and a first turbine section operable for driving the first compressor section, the first turbine section including between three (3) and six (6) stages, a second turbine section operable for driving the second compressor section, and a gear train defined along an engine centerline axis. One of the first turbine section and the second turbine section is operable to drive the fan section through the gear train. 1. A gas turbine engine comprising:a fan section;a first compressor section including three (3) or more stages;a second compressor section including between eight (8) and thirteen (13) stages;a first turbine section operable for driving the first compressor section, the first turbine section including between three (3) and six (6) stages;a second turbine section operable for driving the second compressor section; anda gear train defined along an engine centerline axis, wherein one of the first turbine section and the second turbine section is operable to drive the fan section through the gear train.2. The engine as recited in claim 1 , wherein the first turbine section includes three (3) stages.3. The engine as recited in claim 2 , wherein the first turbine defines a pressure ratio that is greater than about 5.0.4. The engine as recited in claim 1 , including a fan case circumscribing the fan section that defines a fan bypass airflow claim 1 , and wherein a bypass ratio of the engine is greater than about 10.0.5. The engine as recited in claim 1 , wherein said gear train defines a gear reduction ratio of greater than about 2.3.6. The gas turbine engine as set forth in claim 1 , further comprising a fan variable area nozzle to vary a fan nozzle exit area and adjust a fan pressure ratio of fan bypass ...

VARIABLE VANE DRIVE SYSTEM

An example section of a gas turbine engine includes a plurality of variable vanes circumferentially disposed about an engine axis, a first moveable annular ring disposed on an upstream side of the variable vanes, a second movable annular ring disposed on a downstream side of the variable vanes, and a plurality of vane arms, each including a first end secured to the first annular ring and a second end secured to the second annular ring. Movement of the first and second annular rings moves the vane arms, thereby actuating the plurality of variable vanes. An example variable vane assembly includes a vane arm including a portion that engages a variable vane, a first end configured to be secured to a first movable annular ring, and a second end configured to be secured to a second movable annular ring. 1. A section of a gas turbine engine comprising:a plurality of variable vanes circumferentially disposed about an engine axis;a first moveable annular ring disposed on an upstream side of the variable vanes;a second movable annular ring disposed on a downstream side of the variable vanes,a plurality of vane arms, each including a first end secured to the first annular ring and a second end secured to the second annular ring; andwherein movement of the first and second annular rings moves the vane arms, thereby actuating the plurality of variable vanes.2. The engine section of claim 1 , wherein movement of the first and second rings causes the vane arm to pivot about a radially extending axis.3. The engine section of claim 1 , further comprising a bell crank configured to move at least one of the first and second rings.4. The engine section of claim 3 , wherein the bell crank is configured to move the first and second rings in opposite circumferential directions.5. The engine section of claim 3 , further comprising an actuator configured to actuate the first bell crank.6. The engine section of claim 5 , further comprising a second engine section including a second plurality ...

GAS TURBINE ENGINE

A blade or a vane for a gas turbine engine. The blade or vane comprising an aerofoil and a leading edge member attached to the aerofoil. The leading edge comprises an electrically conductive support member and a nano-coating formed on the support member. 1. A blade or a vane for a gas turbine engine , the blade or vane comprising:an aerofoil; anda leading edge member attached to the aerofoil;wherein the leading edge comprises an electrically conductive support member and a nano-crystalline metallic layer formed on the support member.2. The blade or vane according to claim 1 , wherein the nano-crystalline metallic layer comprises a Nickel-Cobalt alloy claim 1 , Nickel or a Nickel alloy claim 1 , and/or Cobalt or a Cobalt alloy.3. The blade or vane according to claim 1 , wherein the leading edge member includes a front portion and two wings extending from the front portion towards a trailing edge of the aerofoil.4. The blade or vane according to claim 3 , wherein the front portion of the leading edge member comprises the support member.5. The blade or vane according to claim 3 , wherein the wings are bonded to the aerofoil.6. The blade or vane according to claim 1 , comprising a plurality of support members claim 1 , and wherein each support member has a surface coated with a nano-crystalline metallic layer.7. The blade or vane according to claim 6 , wherein the support members are arranged to be stacked in a chordwise direction.8. The blade or vane according to claim 6 , wherein the support members are arranged to be stacked in a spanwise direction.9. The blade or vane according to claim 1 , wherein the support member is made from a metallic or a polymeric material treated to be conductive.10. The blade or vane according to claim 1 , wherein the support member defines a curved leading edge on which the nano-crystalline metallic layer is provided.11. The blade or vane according to claim 1 , wherein the support member or at least one of the support members defines a ...

NOZZLE GUIDE VANE WITH COMPOSITE HEAT SHIELDS

A nozzle guide vane for a gas turbine engine is disclosed herein. The nozzle guide vane includes an inner endcap, an outer endcap, and at least one airfoil that extends from the inner endcap to the outer endcap. The nozzle guide vane further includes at least one composite heat shield component adapted to shield metallic components from high temperature gasses. 1. A nozzle guide vane for a gas turbine engine , the nozzle guide vane comprisinga metallic support structure including an inner endcap formed to include an inner attachment aperture and an outer endcap formed to include an outer attachment aperture, the outer endcap spaced from the inner endcap in a radial direction,an airfoil including an aerodynamic feature shaped to redirect gasses moving through a gas path between the inner end cap and the outer endcap, an inner attachment feature that extends from the aerodynamic feature into the inner attachment aperture of the inner endcap, and an outer attachment feature that extends from the aerodynamic feature into the outer attachment aperture of the outer endcap, anda ceramic-matrix composite heat shield system adapted to shield the metallic support structure from hot gasses moving through the gas path, the ceramic-matrix composite heat shield system including an inner heat shield arranged radially between the inner endcap and the gas path and an outer heat shield comprising ceramic-matrix composite materials arranged radially between the outer endcap and the gas path.2. The nozzle guide vane of claim 1 , wherein the outer heat shield is sandwiched between the aerodynamic feature and the outer endcap and the inner heat shield is sandwiched between the aerodynamic feature and the inner endcap.3. The nozzle guide vane of claim 2 , wherein the inner heat shield is formed to include an inner locator aperture claim 2 , the outer heat shield is formed to include an outer locator aperture claim 2 , the inner attachment feature extends through the inner locator aperture ...

METHOD AND SYSTEM FOR INTERFACING A CERAMIC MATRIX COMPOSITE COMPONENT TO A METALLIC COMPONENT

An airfoil assembly for a gas turbine engine and a method of transferring load from the ceramic matrix composite (CMC) airfoil assembly to a metallic vane assembly support member are provided. The airfoil assembly includes a forward end and an aft end with respect to an axial direction of the gas turbine engine. The airfoil assembly further includes a radially outer end component including a radially outwardly-facing end surface having a non-compression load-bearing feature extending radially outwardly from the outwardly-facing end surface and formed integrally with the outer end component, the feature configured to mate with a complementary feature formed in a radially inner surface of a first airfoil assembly support structure, the feature selectively positioned orthogonally to a force imparted into the airfoil assembly. The airfoil assembly also includes a radially inner end component, and a hollow airfoil body extending therebetween, the airfoil body configured to receive a strut couplable at a first end to the first airfoil assembly support structure. 1. An airfoil assembly for a gas turbine engine , said airfoil assembly comprising a ceramic matrix composite (CMC) material , said airfoil assembly comprising a forward end and an aft end with respect to an axial direction of the gas turbine engine , said airfoil assembly comprising:a radially outer end component comprising a radially outwardly-facing end surface having a non-compression load-bearing feature extending radially outwardly from said outwardly-facing end surface and formed integrally with said outer end component, said feature configured to mate with a complementary feature formed in a radially inner surface of a first airfoil assembly support structure, said feature selectively positioned orthogonally to a force imparted into said airfoil assembly;a radially inner end component configured to engage a second airfoil assembly support structure positioned radially inward from said radially inner end ...

TURBINE STATOR VANE ASSEMBLY FOR A TURBOMACHINE

A turbine stator vane assembly comprises a plurality of circumferentially arranged struts and a plurality of circumferentially arranged stator vanes. Each stator vane has a leading edge, a trailing edge and a chord length. Each strut has a leading edge, a trailing edge, a chord length, a pressure surface and a suction surface. The chord length of the struts is greater than the chord length of the stator vanes. The stator vanes are arranged in groups between adjacent pairs of struts and each group of stator vanes comprises a plurality of stator vanes. The circumferential spacing between adjacent struts is substantially the same and each stator vane in a group of stator vanes has a different chord length to the other stator vanes in that group of stator vanes. 1. A turbine stator vane assembly comprising a plurality of circumferentially arranged struts and a plurality of circumferentially arranged stator vanes , each stator vane having a leading edge , a trailing edge and a chord length , each strut having a leading edge , a trailing edge , a chord length , a pressure surface and a suction surface , the chord length of the struts being greater than the chord length of the stator vanes , the stator vanes being arranged in groups , each group of stator vanes being arranged between an adjacent pair of struts , each group of stator vanes comprising a plurality of stator vanes , wherein in at least one group of stator vanes the chord length of the stator vane nearest to the pressure surface of one of the adjacent struts is greater than the chord length of the stator vane nearest to the suction surface of the other of the adjacent struts.2. A turbine stator vane assembly as claimed in wherein in each group of stator vanes the chord length of the stator vane nearest to the pressure surface of one of the adjacent struts is greater than the chord length of the stator vane nearest to the suction surface of the other of the adjacent struts.3. A turbine stator vane assembly as ...

GAS TURBINE ENGINE COMPONENT MATEFACE SURFACES

An array of components in a gas turbine engine includes first and second structures respectively including first and second surfaces that are arranged adjacent to one another to provide a gap. The first and second surfaces respectively have first and second rounded edges at the gap that are arranged in staggered relationship relative to one another. 1. An array of components in a gas turbine engine , comprising:first and second structures respectively include first and second surfaces that are arranged adjacent to one another to provide a gap, the first and second surfaces respectively have first and second rounded edges at the gap that are arranged in staggered relationship relative to one another.2. The array according to claim 1 , wherein the gap is provided at a constant angle along a generally axial length from a forward end of the first and second structures to an aft end of the first and second structures.3. The array according to claim 2 , wherein the axial length includes first and second lengths claim 2 , the first and second lengths each in a range of 30-70% of the axial length claim 2 , the first rounded edge arranged along the first length claim 2 , and the second rounded edge arranged along the second length.4. The array according to claim 3 , wherein the first and second structures respectively include first and second matefaces facing one another at the gap claim 3 , the first and second surfaces forming generally sharp corners respectively with the first and second matefaces adjacent to the first and second rounded edges claim 3 , respectively.5. The array according to claim 4 , wherein the first and second surface and the first and second matefaces are respectively perpendicular to one another.6. The array according to claim 1 , wherein the first and second structures are one of a blade outer air seal or a platform.7. The array according to claim 6 , wherein the first and second structures are one of a stator vane or blade claim 6 , and an airfoil ...

IMPROVED TURBOMACHINE STATOR

Turbine nozzle for a turbomachine extending about an axis, the nozzle including a plurality of sectors each having at least one vane which extends radially between an internal platform and an external platform, the nozzle including at least two sealing strips each including a main portion configured to ensure the sealing between the nozzle and an element of the turbomachine which is adjacent to the nozzle, and a first portion folded relative to the main portion, the first folded portion being in contact with an outer face of at least one platform, each strip being held to the at least one platform by a clamping device configured to clamp the first folded portion between the outer face of the platform and the clamping device. 1. A nozzle for a turbomachine turbine extending about an axis , the nozzle including a plurality of sectors each comprising at least one vane which extends radially between an internal platform and an external platform , the nozzle including at least two sealing strips each including a main portion configured to ensure the sealing between the nozzle and an element of the turbomachine which is adjacent to the nozzle , and a first portion folded relative to the main portion , the first folded portion being in contact with an outer face of at least one platform , each strip being held to the at least one platform by a clamping means configured to clamp the first folded portion between the outer face of the platform and the clamping means.2. The nozzle according to claim 1 , including two strips each forming a half-ring claim 1 , the clamping means being disposed around the nozzle so as to clamp the two strips against the outer faces of the platforms of the nozzle.3. The nozzle according to claim 1 , wherein the main portion and the first folded portion of each strip form therebetween an angle comprised between 90° and 150°.4. The nozzle according to claim 1 , wherein the strips each include a second folded portion extending from the first folded ...

Gas Turbine Nozzle

The present invention provides a gas turbine nozzle capable of reducing stress related to thermal elongation caused by a rise in gas turbine nozzle temperature and thus reducing stress produced when thermal deformation occurs in the gas turbine nozzle. The gas turbine nozzle according to the present invention includes nozzles formed integrally through an inner perimeter end wall and an outer perimeter end wall. The inner perimeter end wall has an upstream connection portion and a downstream connection portion. The upstream connection portion extends radially inward to be connected to an inner perimeter diaphragm. The downstream connection portion is located downstream from the upstream connection portion and extends radially inward to be connected to the inner perimeter diaphragm. The inner perimeter end wall has a thin-walled portion in a rear edge portion of the inner perimeter end wall, the thin-walled portion corresponding to a reduced wall thickness portion of the rear edge portion of the inner perimeter end wall. 1. A gas turbine nozzle with nozzles formed integrally through an inner perimeter end wall and an outer perimeter end wall ,wherein the inner perimeter end wall has an upstream connection portion and a downstream connection portion, the upstream connection portion extending radially inward to be connected to an inner perimeter diaphragm, the downstream connection portion being located downstream from the upstream connection portion and extending radially inward to be connected to the inner perimeter diaphragm, andthe inner perimeter end wall has a thin-walled portion in a rear edge portion of the inner perimeter end wall, the thin-walled portion corresponding to a reduced wall thickness portion of the rear edge portion of the inner perimeter end wall.2. The gas turbine nozzle according to claim 1 , wherein the gas turbine nozzle has a coupled vane structure in which two of the nozzles are formed integrally through the inner perimeter end wall and the ...

FLOW ARRANGEMENT FOR PLACING IN A HOT GAS DUCT OF A TURBOMACHINE

The invention relates to a flow arrangement for placing in the hot gas duct of a turbomachine, having a first surrounding-flow structure and a second surrounding-flow structure, the surrounding-flow structures each having, in reference to the surrounding flow in the hot gas duct, a leading edge and, downstream thereof, a trailing edge, wherein the second surrounding-flow structure is provided as a deflecting blade with a suction side and a pressure side and has a lesser profile thickness than the first surrounding-flow structure, which is arranged on the suction side of the second surrounding-flow structure, and wherein, although the second surrounding-flow structure has a partial axial overlap with the first surrounding-flow structure referred to a longitudinal axis of the turbomachine, the trailing edge of the second surrounding-flow structure is, at the same time, displaced axially downstream relative to the trailing edge of the first surrounding-flow structure. 1. A flow arrangement for placing in the hot gas duct of a turbomachine , comprising:a first surrounding-flow structure;a second surrounding-flow structure;the first surrounding-flow structure and the second surrounding-flow structure each having, in reference to the surrounding flow in the hot gas duct, a leading edge and, downstream thereof, a trailing edge;wherein the second surrounding-flow structure is configured and arranged as a deflecting blade with a suction side and a pressure side and has a lesser profile thickness than the first surrounding-flow structure, which is arranged on the suction side of the second surrounding-flow structure; andwherein, the second surrounding-flow structure has a partial axial overlap with the first surrounding-flow structure referred to a longitudinal axis of the turbomachine, the trailing edge of the second surrounding-flow structure is, at the same time, displaced axially downstream relative to the trailing edge of the first surrounding-flow structure.2. The flow ...

GAS TURBINE ENGINE COMPONENT

There is disclosed a component arranged to extend across a bypass duct () of a gas turbine engine (), comprising: an outlet guide vane (); and a structural support () arranged to structurally support a radially inner component of the gas turbine engine (). The outlet guide vane () and the structural support () are contiguous at a merged portion () which is at a radially outer region of the component. 1. A component arranged to extend across a bypass duct of a gas turbine engine , comprising:an outlet guide vane; anda structural support arranged to structurally support a radially inner component of the gas turbine engine;{'b': '2', 'wherein the outlet guide vane and the structural support are contiguous at a merged portion which is at a radially outer region of the component, and the merged portion has a radial extent (L) less than the radial extent (L) of the component.'}2. A component according to claim 1 , wherein the outlet guide vane and the structural support are axially spaced apart over a portion of their radial extent so as to form a gap between the outlet guide vane and the structural support.3. A component according to claim 2 , wherein the gap is provided at a radially inner region of the component.4. A component according to claim 2 , wherein the outlet guide vane comprises a guide vane root portion and wherein the structural support comprises a structural support root portion claim 2 , and wherein the guide vane root portion and the structural support root portion are axially spaced apart so as to form the gap.53. A component according to any claim 4 , wherein the axial spacing (L) between the guide vane root portion and the structural support root portion is less than 0.5 times the radial extent (L) of the component.6. A component according to claim 1 , wherein the outlet guide vane is arranged to be upstream of the structural support.7. A component according to claim 1 , wherein the leading edge of the structural support is concave.8. A component ...

VANE CARRIER, COMPRESSOR, AND GAS TURBINE INCLUDING THE SAME

Disclosed herein are a vane carrier that is uniformly deformed by heat and a gas turbine including the same. The vane carrier includes a pair of bodies having a plurality of vanes arranged on their inner peripheral surfaces, and forming an annular shape by coupling the bodies in a semi-annular form to each other, fastening parts protruding radially from outer peripheral surfaces of respective ends of the bodies coupled to each other so as to couple the bodies, and one or more deformation prevention members disposed on the outer peripheral surfaces of the bodies and protruding radially therefrom. Accordingly, it is possible to prevent damage to components due to heat deformation since the vane carrier is relatively uniformly deformed and to improve durability by adjusting the natural frequency of the vane carrier to avoid resonance. 1. A vane carrier for supporting a vane provided in a gas turbine , comprising:a pair of bodies to have a plurality of vanes arranged on their inner peripheral surfaces, and the pair of bodies to form an annular shape by coupling the bodies in a semi-annular form to each other;fastening parts to protrude radially from outer peripheral surfaces of respective ends of the bodies coupled to each other so as to couple the bodies; andat least one deformation prevention member disposed on the outer peripheral surfaces of the bodies and to protrude radially therefrom.2. The vane carrier according to claim 1 , wherein the at least one deformation prevention member faces each other and is symmetrically disposed with respect to a center of the pair of bodies in the annular shape.3. The vane carrier according to claim 2 , wherein:the bodies are vertically disposed so that the fastening parts are horizontally arranged in parallel with each other; andthe at least one deformation prevention member vertically faces each other.4. The vane carrier according to claim 1 , wherein the at least one deformation prevention member is of a shape corresponding to ...

REAR OUTER DISCHARGE NOZZLE

A rear outer discharge nozzle for a gas turbine engine, the rear outer discharge nozzle forming a loop around an axis and comprising: a first part; a second part wrapped around the first part; and a radial slit, having first and second side walls; wherein the first part is shaped to provide the first side wall of the radial slit, and the second part is shaped to provide the second side wall of the radial slit; and wherein the second part comprises a break, splitting the second part in the axial part. 1. A rear outer discharge nozzle for a gas turbine engine , the rear outer discharge nozzle forming a loop around an axis and comprising:a first part;a second part wrapped around the first part; anda radial slit, having first and second side walls;wherein the first part is shaped to provide the first side wall of the radial slit, and the second part is shaped to provide the second side wall of the radial slit; andwherein the second part comprises a break, splitting the second part along the axial direction.2. A rear outer discharge nozzle according to claim 1 , wherein the second part comprises fastening points either side of the break.3. A rear outer discharge nozzle according to claim 2 , wherein the fastening points are attached to the second part.4. A rear outer discharge nozzle according to claim 2 , wherein the fastening points are integrally formed with the second part.5. A rear outer discharge nozzle according to claim 1 , wherein the second part is fastened to the first part.6. A rear outer discharge nozzle according to claim 1 , wherein the first and second part engage claim 1 , to inhibit angular rotation of the second part with respect to the first part around the axis.7. A rear outer discharge nozzle according to claim 6 , wherein the first part comprises a projection claim 6 , and the second part comprises a slot that fits around the projection.8. A rear outer discharge nozzle according to claim 6 , wherein the first part comprises a slot claim 6 , and the ...

MODULE FOR A TURBOMACHINE

The present invention relates to a module for a turbomachine. The guide vane arrangement has a guide vane, an inner platform, and a first guide pin, the guide vane is arranged radially outside on the inner platform and the first guide pin is arranged radially inside on the inner platform, and wherein the seal carrier is arranged radially inside the inner platform and forms an uptake, which is open radially outward, the first guide pin of the guide vane arrangement being arranged in the uptake. In the uptake, a slide body is arranged, on which the first guide pin of the guide vane arrangement, with respect to a rotation around the longitudinal axis of the module, finds a contact and is guided in a radially movable manner in this contact. The slide body is formed in one piece with a wall of the seal carrier, which axially bounds the uptake. 1. A module for a turbomachine , comprising:a guide vane arrangement; anda seal carrier;wherein the guide vane arrangement has a guide vane, an inner platform, and a first guide pin, wherein, in each case referred to a longitudinal axis of the module, the guide vane is arranged radially outside on the inner platform, and the first guide pin is arranged radially inside on the inner platform;wherein the seal carrier is arranged radially inside the inner platform and forms an uptake, which is open radially outward, the first guide pin of the guide vane arrangement being arranged in said uptake;wherein, in the uptake, a slide body is arranged, on which the first guide pin of the guide vane arrangement, with respect to a rotation around the longitudinal axis of the module, finds a contact and is guided in a radially movable manner in this contact; andwherein the slide body is formed in one piece with a wall of the seal carrier, which axially bounds the uptake.2. The module according to claim 1 , in which the slide body is monolithically formed with the wall axially bounding the uptake of the seal carrier.3. The module according to claim ...

Apparatus for Decreasing Thrust of Radial Inflow Turbine

Disclosed is an apparatus for decreasing a thrust of a radial flow turbine. The apparatus includes a rotary shaft having an axial through-hole in the interior thereof, a rotor assembled in the rotary shaft and having a rotor hub and rotor blades formed on an outer peripheral surface thereof, a casing configured to isolate the rotary shaft and the rotor from the outside, and a shaft seal configured to maintain a seal between the rotary shaft and the casing. 1. An apparatus for decreasing a thrust of a radial flow turbine , the apparatus comprising:a rotary shaft having an axial through-hole in the interior thereof;a rotor assembled in the rotary shaft and having a rotor hub and rotor blades formed on an outer peripheral surface thereof;a casing configured to isolate the rotary shaft and the rotor from the outside; anda shaft seal configured to maintain a seal between the rotary shaft and the casing.2. The apparatus of claim 1 , wherein at least one protrusion is formed on a back face of the rotor hub claim 1 , and at least one groove claim 1 , by which the at least one protrusion is guided claim 1 , is formed on an inner face of the casing claim 1 , which faces the back face of the rotor hub.3. The apparatus of claim 2 , wherein the at least one groove is at least one ring-shaped groove claim 2 , by which the at least one protrusion is guided.4. The apparatus of claim 3 , wherein the at least one protrusion is at least one protruding ring formed about the rotary shaft claim 3 , and the at least one ring is guided by the at least one groove.5. The apparatus of claim 4 , wherein the at least one ring comprises a plurality of rings having different diameters and the at least one groove comprises a plurality of grooves guiding the plurality of rings to form a multi-ring structure.6. The apparatus of claim 1 , wherein a first protruding ring and a second protruding ring having a diameter greater than that of the first ring are formed on the back face of the rotor hub ...

STEAM TURBINE NOZZLE SEGMENT HAVING TRANSITIONAL INTERFACE, AND NOZZLE ASSEMBLY AND STEAM TURBINE INCLUDING SUCH NOZZLE SEGMENT

Various embodiments include steam turbine static nozzles having transitional interfaces. In one embodiment a steam turbine static nozzle blade includes: an airfoil; an inner sidewall integral with a first side of the airfoil; and an outer sidewall integral with a second side of the airfoil; the inner sidewall and the outer sidewall each including: a first side having one of: an arcuate concave surface extending substantially an entire length of the sidewall, or an arcuate convex surface extending substantially the entire length of the sidewall; and a second side opposing the first side and having an angled interface extending substantially the entire length of the sidewall. 1. A steam turbine static nozzle blade comprising:an airfoil;an inner sidewall integral with a first side of the airfoil; andan outer sidewall integral with a second side of the airfoil; a first side having one of: an arcuate concave surface extending substantially an entire length of the sidewall, or an arcuate convex surface extending substantially the entire length of the sidewall; and', 'a second side opposing the first side and having an angled interface extending substantially the entire length of the sidewall., 'the inner sidewall and the outer sidewall each including2. The static nozzle blade of claim 1 , wherein the first side includes a pressure side claim 1 , and wherein the first side includes the arcuate concave surface.3. The static nozzle blade of claim 1 , wherein the first side includes a suction side claim 1 , and wherein the first side includes the arcuate convex surface.4. The static nozzle blade of claim 1 , wherein the arcuate concave surface or the arcuate convex surface has an arc radius of approximately 14 centimeters.5. The static nozzle blade of claim 1 , wherein the angled interface is angled to complement an adjacent static nozzle blade in an existing static nozzle blade assembly.6. The static nozzle blade of claim 1 , wherein the arcuate concave surface or the ...

MACHINED VANE ARM OF A VARIABLE VANE ACTUATION SYSTEM

An exemplary variable vane actuation system includes, among other things, a vane arm with a vane stem contact surface and a radially outward facing surface. The vane stem contact surface is to contact a vane stem of a variable vane and thereby actuate the variable vane about a radially extending axis. The vane stem contact surface is angled relative to both the radially extending axis and the radially outward facing surface. 1. A variable vane actuation system , comprising:a vane arm with at least one vane stem contact surface and a radially outward facing surface, the at least one vane stem contact surface to contact a vane stem of a variable vane and thereby actuate the variable vane about a radially extending axis, the at least one vane stem contact surface angled relative to both the radially extending axis and the radially outward facing surface.2. The system of claim 1 , including an aperture extending through the radially outward facing surface to receive the vane stem claim 1 , a least a portion of the aperture having a non-circular cross-sectional profile.3. The system of claim 2 , wherein the aperture comprises a first axial section and a second axial section claim 2 , the first axial section having a generally oval-shaped cross sectional profile claim 2 , the second axial section having a generally circular-shaped cross-sectional profile.4. The system of claim 2 , wherein the at least one vane stem contact surface comprises a first vane stem contact surface and a second vane stem contact surface claim 2 , the aperture positioned between the first and second vane stem contact surfaces.5. The system of claim 1 , wherein the at least one vane stem contact surface is a machined surface.6. The system of claim 5 , wherein the at least one vane stem contact surface is a milled surface.7. The system of claim 1 , wherein the vane arm is continuous radially between the at least one vane stem contact surface and the radially outward facing surface.8. The system of ...

Gas turbine engine end-wall component

An end-wall component of the mainstream gas annulus of a gas turbine engine having an annular arrangement of vanes, the component including a cooling arrangement having ballistic cooling holes ( 33 ) through which, in use, dilution cooling air is jetted into the mainstream gas upstream of the vanes to reduce the mainstream gas temperature adjacent the end-wall, wherein the cooling holes are arranged in one or more circumferentially extending rows and wherein the axial position of the cooling holes in the or each row varies.

Additively Manufactured Component With Integral Fuel Lines Passage

An aerospace component includes an augmentor vane and a fuel line additively manufactured within the augmentor vane. A method of manufacture including additively manufacturing an augmentor vane wall that forms an internal volume therein; and additively manufacturing a fuel line within the augmentor vane and extends through the internal volume. 1. An aerospace component , comprising:an augmentor vane; anda fuel line additively manufactured within the augmentor vane.2. The aerospace component as recited in claim 1 , wherein the augmentor vane is additively manufactured.3. The aerospace component as recited in claim 1 , wherein the component is an augmentor vane.4. The aerospace component as recited in claim 1 , wherein the fuel line is formed through an endwall.5. The aerospace component as recited in claim 4 , wherein the fuel line is spaced from the endwall.6. The aerospace component as recited in claim 5 , wherein the fuel line includes a branch that is extends about 40 degrees from a central line.7. The aerospace component as recited in claim 5 , wherein the fuel line includes a branch that extends at greater or equal to about 40 degrees from a central line.8. The aerospace component as recited in claim 1 , wherein the fuel line terminates at an exit aperture through a wall of the augmentor vane.9. The aerospace component as recited in claim 8 , wherein the wall of the augmentor vane is airfoil shaped.10. The aerospace component as recited in claim 1 , wherein the fuel line terminates at an exit aperture that defines an internal radius.11. The aerospace component as recited in claim 1 , wherein the fuel line terminates at an exit aperture that is seal less.12. An augmentor vane claim 1 , comprising:an additively manufactured augmentor vane wall that forms an internal volume therein; anda fuel line additively manufactured within the augmentor vane and extends through the internal volume, the fuel line terminates at an exit aperture that is seal less.13. The ...

FAN BLADE WITH COMPOSITE COVER

A fan blade includes a metallic body, a first composite cover, and a second composite cover. The metallic body may have a first side, a second side, a plurality of first retention slots, and a plurality of second retention slots, in accordance with various embodiments. The first and second retention slots may extend from the first side to the second side of the metallic body. The first composite cover may be coupled to the first side of the metallic body and may include a plurality of first fingers that extend through the first retention slots and are coupled to the second side of the metallic body. The second composite cover may be coupled to the second side of the metallic body and may include a plurality of second fingers that extend through the second retention slots and are coupled to the first side of the metallic body. 1. A fan blade comprising:a metallic body comprising a first side, a second side, a plurality of first retention slots, and a plurality of second retention slots, wherein the first and second retention slots extend from the first side to the second side of the metallic body;a first composite cover coupled to the first side of the metallic body, wherein the first composite cover comprises a plurality of first fingers that extend through the first retention slots and are coupled to the second side of the metallic body; anda second composite cover coupled to the second side of the metallic body, wherein the second composite cover comprises a plurality of second fingers that extend through the second retention slots and are coupled to the first side of the metallic body.2. The fan blade of claim 1 , wherein the metallic body is made from titanium.3. The fan blade of claim 1 , wherein the first and second composite covers are made from a graphite epoxy material.4. The fan blade of claim 1 , wherein the first composite cover is made from a first composite material and the second composite cover is made from a second composite material that is ...

REGENERATIVELY COOLED TRANSITION DUCT WITH TRANSVERSELY BUFFERED IMPINGEMENT NOZZLES

A cooling arrangement () having: a duct () configured to receive hot gases () from a combustor; and a flow sleeve () surrounding the duct and defining a cooling plenum () there between, wherein the flow sleeve is configured to form impingement cooling jets () emanating from dimples () in the flow sleeve effective to predominately cool the duct in an impingement cooling zone (), and wherein the flow sleeve defines a convection cooling zone () effective to cool the duct solely via a cross-flow (), the cross-flow comprising cooling fluid () exhausting from the impingement cooling zone. In the impingement cooling zone an undimpled portion () of the flow sleeve tapers away from the duct as the undimpled portion nears the convection cooling zone. The flow sleeve is configured to effect a greater velocity of the cross-flow in the convection cooling zone than in the impingement cooling zone. 1. A cooling arrangement , comprising:a duct configured to receive hot gases from a combustor can; anda flow sleeve surrounding the duct and defining a cooling plenum there between, wherein the flow sleeve is configured to form impingement cooling jets emanating from dimples in the flow sleeve effective to cool the duct in a first zone having impingement cooling, and wherein the flow sleeve is configured to form a convection cooling zone effective to cool the duct solely via a cross-flow, the cross-flow comprising spent impingement cooling fluid from the impingement cooling jets, and the cross-flow flowing from the first zone having impingement cooling into the convection cooling zone,wherein in the first zone having impingement cooling, an undimpled portion of the flow sleeve tapers away from the duct as the undimpled portion nears the convection cooling zone,wherein a cross-sectional flow area of the cooling plenum decreases in the convection cooling zone and the decreased cross-sectional flow area is effective to accelerate the cross-flow to a greater velocity in the convection ...

PASSIVE TRANSPIRATIONAL FLOW ACOUSTICALLY LINED GUIDE VANE

A passive transpirational flow acoustic liner assembly for a gas turbine engine includes a guide vane assembly and a conduit configured to deliver airflow received from the guide vane. The guide vane assembly includes an airfoil having a transpirational flow acoustic liner. The acoustic liner includes a face sheet defining a portion of an outer surface of the airfoil and having a plurality of first apertures, a segmented member coupled to the face sheet and having a plurality of chambers in fluid communication with the outer surface via the plurality of first apertures, a backing sheet having a plurality of apertures and being coupled to the segmented member such that the segmented member is positioned between the face sheet and the backing sheet, and a plenum coupled to the backing sheet opposite the segmented member and fluidly connected to the conduit. 1. A passive transpirational flow acoustic liner assembly for a gas turbine engine , the assembly comprising: an airfoil having a suction side and a pressure side; and', a face sheet defining a portion of an outer surface of the airfoil and having a plurality of first apertures;', 'a segmented member coupled to the face sheet, the segmented member comprising a plurality of chambers in fluid communication with the outer surface via the plurality of first apertures;', 'a backing sheet coupled to the segmented member such that the segmented member is positioned between the face sheet and the backing sheet, the backing sheet having a plurality of second apertures; and', 'a plenum coupled to the backing sheet opposite the segmented member; and, 'a transpirational flow acoustic liner disposed in the airfoil, the liner comprising], 'a guide vane assembly comprisinga conduit fluidly connected to the plenum and configured to deliver airflow received from the plenum.2. The assembly of claim 1 , wherein the conduit is disposed in a core housing of the gas turbine engine and fluidly connected to the plenum by an opening at a ...

ACTIVE FLOW CONTROL TRANSPIRATIONAL FLOW ACOUSTICALLY LINED GUIDE VANE

A vane for use in a gas turbine engine includes an airfoil having a suction side and a pressure side, and a transpirational flow acoustic liner disposed in the airfoil. The liner includes a face sheet defining a portion of an outer surface of the airfoil and having a plurality of first apertures, a segmented member coupled to the face sheet, and a backing sheet coupled to the segmented member the segmented member such that the segmented member is positioned between the face sheet and the backing sheet. The segmented member includes a plurality of chambers in fluid communication with the outer surface via the plurality of first apertures. The backing sheet has a plurality of second apertures. 1. A vane for use in a gas turbine engine , the vane comprising:an airfoil having a suction side and a pressure side; and{'claim-text': ['a face sheet defining a portion of an outer surface of the airfoil and having a plurality of first apertures;', 'a segmented member coupled to the face sheet, the segmented member comprising a plurality of chambers in fluid communication with the outer surface via the plurality of first apertures; and', 'a backing sheet coupled to the segmented member such that the segmented member is positioned between the face sheet and the backing sheet, the backing sheet having a plurality of second apertures.'], '#text': 'a transpirational flow acoustic liner disposed in the airfoil, the liner comprising:'}2. The vane of claim 1 , and further comprising a suction chamber coupled to the backing sheet opposite the segmented member claim 1 , the suction chamber in fluid communication with the outer surface via the plurality of second apertures claim 1 , the plurality of chambers claim 1 , the plurality of first apertures.3. The vane of claim 2 , wherein first apertures of the plurality of first apertures have a diameter in a range of 3 to 50 thousandths of an inch.4. The vane of claim 2 , wherein the second apertures of the plurality of second apertures have ...

SYSTEM AND METHOD FOR CONVERTING TURBINE COOLING NOZZLE

A stationary component of a turbine section of a turbine engine system includes a hollow vane assembly with an interior rib separating a first interior compartment from a second interior compartment. The first and second interior compartments are configured to receive the cooling fluid. The cooling fluid is allowed to communicate between the first and second interior compartments. The stationary component also includes a diaphragm attached to the hollow vane assembly. The diaphragm is configured to receive the cooling fluid from the hollow vane assembly. The diaphragm includes a chamber and a tube extending through the chamber. The tube is configured to isolate the chamber from the cooling fluid while delivering the cooling fluid from the hollow vane assembly to the wheelspace area. 1. A stationary component extending from a casing of a turbine section of a turbine engine system to a wheelspace area , the stationary component being configured to deliver cooling fluid to the wheelspace area , the stationary component comprising:a hollow vane assembly with an interior rib separating a first interior compartment from a second interior compartment, the first and second interior compartments being configured to receive the cooling fluid, the cooling fluid being allowed to communicate between the first and second interior compartments; anda diaphragm attached to the hollow vane assembly and configured to receive the cooling fluid from the hollow vane assembly, the diaphragm comprising a chamber and a tube extending through the chamber, the tube being configured to isolate the chamber from the cooling fluid while delivering the cooling fluid from the hollow vane assembly to the wheelspace area.2. The stationary component of claim 1 , wherein the rib does not extend the entire length of the hollow vane assembly to allow the cooling fluid to flow between the first and second interior compartments by flowing around the rib.3. The stationary component of claim 1 , wherein the ...

AIR-GAP FINS FOR A TURBINE ENGINE COMPRESSOR

A vane stage with a longitudinal axis designed to be fitted in a turbine engine compressor. The van stage has an annular row of mobile vanes arranged upstream from an annular row of stator vanes. The annular row of stator vanes has a radially internal annular platform bearing radial blades, an upstream annular portion of which is arranged upstream from the blades and is surrounded radially outwards by a downstream annular portion of an annular platform of the upstream row of mobile vanes. The upstream annular portion of the annular platform of the annular row of stator vanes has a radially external annular face from which fins extend, which are distributed around the longitudinal axis and extend radially outwards towards the downstream annular portion of the platform of the annular row of mobile vanes. 1. A vane stage extending around a longitudinal axis and designed to be fitted in a turbine engine compressor , the vane stage comprising an annular row of mobile vanes arranged upstream from an annular row of stator vanes , wherein the annular row of stator vanes comprises a radially internal annular platform bearing radial blades , an upstream annular portion of which is arranged upstream from said blades and is surrounded radially outwards by a downstream annular portion of an annular platform of the upstream row of mobile vanes , wherein the upstream annular portion of the annular platform of the annular row of stator vanes comprises a radially external annular face from which fins extend , said fins are distributed around the longitudinal axis and extend radially outwards towards the downstream annular portion of the platform of the annular row of mobile vanes , characterised in that the fins each comprise an intrados face and an extrados face oriented circumferentially in an identical manner to the intrados faces and extrados faces of the blades of the stator vanes.2. The vane stage of to claim 1 , wherein an angle between the longitudinal axis and the tangent ...

Moisture Separator Unit for Steam Turbine and Steam-Turbine Stationary Blade

A steam turbine stationary blade is configured to be a stationary blade structure having a hollow structure. The steam turbine stationary blade is provided with a slit disposed on a stationary blade surface. The pressure of a stationary blade hollow is reduced to suction a liquid film through the slit for removing the liquid film formed on the stationary blade surface. An opening portion of the slit is covered with a sheet in a meshed pattern formed of a fine mesh thereby reducing an accompanied steam amount so as to effectively remove the liquid film.