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 ...

РАЗБЛОКИРУЕМОЕ УСТРОЙСТВО ДЛЯ СТОПОРЕНИЯ В ОСЕВОМ НАПРАВЛЕНИИ УПЛОТНИТЕЛЬНОГО КОЛЬЦА, С КОТОРЫМ РАБОЧЕЕ КОЛЕСО РОТОРА МОДУЛЯ ТУРБОМАШИНЫ ЛЕТАТЕЛЬНОГО АППАРАТА ОСУЩЕСТВЛЯЕТ КОНТАКТ

Изобретение относится к энергетике. Устройство для стопорения в осевом направлении уплотнительного кольца, выполненного из истираемого материала и находящегося в контакте с периферией ротора модуля турбомашины летательного аппарата. Устройство содержит опору с опорным отверстием, осевую стопорную часть, причём конструкция устройства обеспечивает возможность вращения стопорной части вокруг оси между осевым стопорным положением для уплотнительного кольца и между положением для извлечения этого кольца через проход для извлечения. Также представлены модуль турбомашины летательного аппарата и турбомашина летательного аппарата, содержащие устройство для стопорения. Изобретение позволяет обеспечить выполнение невыпадающего устройства стопорения уплотнительного кольца первой ступени модуля турбомашины. 3 н. и 7 з.п. ф-лы, 11 ил.

УСТРОЙСТВО ДЛЯ ПОДВЕСКИ ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ

Изобретение относится к области авиации, в частности, к подвеске турбореактивных двигателей. Устройство для подвески турбореактивного двигателя содержит крепления с шарнирно соединенными звеньями. Крепление содержит держатель, имеющий три ветви с проходами, через которые проходит штырь. Штырь ориентирован параллельно направлению, которое является тангенциальным корпусу, и шарнирно присоединен к центральной ветви держателя посредством шарового соединения. Достигается возможность вмещения вентилятора увеличенного диаметра. 5 з.п. ф-лы, 3 ил.

ПОДВЕСКА КАНАЛА ХОЛОДНОГО ПОТОКА ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ НА ВЫПУСКНОМ КОРПУСЕ ПРИ ПОМОЩИ ТЯГ И РАДИАЛЬНЫХ ВИЛОК

Двухконтурный турбореактивный двигатель содержит цилиндрический канал холодного потока, на продольных концах которого расположены корпус, окружающий вентилятор турбореактивного двигателя, и опорное кольцо, соединенное с выпускным корпусом. Опорное кольцо установлено при помощи тяг, прикрепленных к цилиндрической наружной обечайке выпускного корпуса при помощи точек крепления. Точки крепления выпускного корпуса представляют собой вилки, проушины которых проходят радиально от наружной обечайки и расположены по существу в осевом направлении посередине обечайки. Отверстия вилок ориентированы по направлению образующих наружной обечайки. Точки крепления тяг на опорном кольце канала холодного потока расположены в осевом направлении выше по потоку от вилок наружной обечайки выпускного корпуса. Изобретение позволяет упростить конструкцию крепления опорного кольца к выпускному корпусу и обеспечить возможность их относительных перемещений. 7 з.п. ф-лы, 5 ил.

УСИЛИВАЮЩАЯ КОНСТРУКЦИЯ ГОНДОЛЫ ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ

Изобретение относится к области авиации, в частности, к гондолам турбореактивных двигателей. Гондола турбореактивного двигателя содержит кожух вентилятора, переднюю раму, средство отклонения потока, реверсор тяги и усиливающую конструкцию. Передняя рама выполнена с возможностью установки за указанным кожухом вентилятора и возможностью взаимодействия с капотом реверсора тяги также поддерживает по меньшей мере одно средство отклонения потока. Реверсор тяги установлен с возможностью скольжения из закрытого положения с перекрытием средства отклонения потока в открытое положение с открытием этого средства отклонения, обеспечивая возможность отклонения потока. Усиливающая конструкция расположена вдоль продольной оси гондолы и выполнена с возможностью установки на ней третьей линии защиты и/или удерживающего устройства между передней рамой и капотом реверса тяги. Достигается повышение уровня надежности и безопасности эксплуатации гондолы двигателя. 14 з.п. ф-лы, 9 ил.

ВОЗДУХОЗАБОРНИК ГАЗОТУРБИННОГО ДВИГАТЕЛЯ В ГОНДОЛЕ

Система газотурбинного двигателя установлена в гондоле, содержащей обтекатель воздухозаборника, образующий воздухозаборник, содержит орган отклонения посторонних объектов, образующий вместе с обтекателем воздухозаборника впускной воздушный канал, и на выходе отклоняющего органа - вторичный отклоняющий канал и главный канал подачи воздуха в двигатель. Впускной воздушный канал выполнен с возможностью отклонения, по меньшей мере, части посторонних объектов, засасываемых через воздухозаборник, в направлении вторичного отклоняющего канала. Вторичный отклоняющий канал выполнен таким образом, чтобы скорость течения проходящего через него воздуха увеличивалась от входа к выходу. Вторичный канал содержит выход с отверстием, выходящим на наружную стенку гондолы. Изобретение направлено на повышение защиты двигателя от попадания посторонних объектов при сохранении аэродинамических характеристик гондолы. 11 з.п. ф-лы, 8 ил.

УЗЕЛ ДВУХКОНТУРНОГО ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ И ДВУХКОНТУРНЫЙ ТУРБОРЕАКТИВНЫЙ ДВИГАТЕЛЬ

Узел двухконтурного турбореактивного двигателя содержит внешнее кольцо выхлопного корпуса, структурное кольцо внешнего тракта канала вентилятора, концентричного относительно внешнего кольца выхлопного корпуса, а также первый и второй кронштейны или соединительные тяги. Кронштейны или соединительные тяги образуют статически неопределимое соединение, являясь закрепленными одним концом на внешнем кольце выхлопного корпуса, а другим концом на структурном кольце. Соединение, образованное первым кронштейном или соединительной тягой, выполнено с возможностью разрыва за пределом определенной нагрузки. Второй кронштейн или тяга выполнен с возможностью формирования пути передачи усилий между кольцами, когда соединение, образованное первым кронштейном или соединительной тягой, разрывается. Другое изобретение относится к двухконтурному турбореактивному двигателю, содержащему указанный выше узел. Группа изобретений позволяет снизить массу турбореактивного двигателя. 2 н. и 8 з.п. ф-лы, 7 ил.

ВНЕШНЯЯ ОБОЛОЧКА ВОЗДУХОВОДА ВЕНТИЛЯТОРА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Двухконтурный турбореактивный двигатель содержит цилиндрическую оболочку, установленную на выходе промежуточного кожуха и ограничивающую с внешней стороны кольцевое пространство протекания вторичного потока. Цилиндрическая оболочка образована решетчатым каркасом и съемными панелями обтекателя, закрепленными на каркасе. Решетчатый каркас содержит входной кольцевой фланец крепления к промежуточному кожуху, выходной кольцевой фланец соединения с выхлопным кожухом, и жесткие балки, соединяющие оба фланца между собой. Изобретение позволяет снизить массу и упростить обслуживание двигателя. 9 з.п. ф-лы, 2 ил.

ГОНДОЛА ДЛЯ ДВУХКОНТУРНОГО ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ

Изобретения относятся к области авиации, более конкретно к гондоле для двухконтурного турбореактивного двигателя, силовой установке и летательному аппарату. Гондола (1) содержит воздухозаборник (5), расположенный перед турбореактивным двигателем (2), среднюю секцию, внутренний кожух (6а) которой охватывает вентилятор (3) турбореактивного двигателя (2) и заднюю секцию (7), содержащую внешний элемент (7а). Элемент (7а) жестко соединен с задней частью кожуха (6а) вентилятора (3) так, что поддерживает турбореактивный двигатель (2) и имеет средства соединения с пилоном, выполненным с возможностью крепления к неподвижно закрепленному элементу (13) летательного аппарата. Технический результат заключается в оптимизации зазора между лопастями вентилятора и внутренними лопатками турбореактивного двигателя с их соответствующими кожухами. 3 н. и 25 з.п. ф-лы, 18 ил.

СИЛОВАЯ УСТАНОВКА ЛЕТАТЕЛЬНОГО АППАРАТА, СОДЕРЖАЩАЯ ТУРБОРЕАКТИВНЫЙ ДВИГАТЕЛЬ С УСИЛИВАЮЩИМИ КОНСТРУКЦИЯМИ, СОЕДИНЯЮЩИМИ КОРПУС ВЕНТИЛЯТОРА С ЦЕНТРАЛЬНЫМ КОРПУСОМ

Изобретение относится к области авиации, в частности к силовым установкам летательных аппаратов. Силовая установка содержит турбореактивный двигатель (2), включающий в себя корпус вентилятора (12), промежуточный корпус (21), расположенный радиально внутри корпуса вентилятора, центральный корпус (16), кольцевую конструкцию (60), окружающую центральный корпус (16) и механически соединенную с последним с помощью промежуточных крепежных средств (62). С первым и вторым узлами (6а, 6b) подвески двигателя связана усиливающая конструкция (64а, 64b), образующая плоскость сдвига и жестко установленная на кольцевой конструкции в первой точке (68а, 68b) крепления, на корпусе вентилятора во второй точке (70а, 70b) крепления и на конструкционной связи (17) или промежуточном корпусе (21) в третьей точке (72а, 72b) крепления. Усиливающая конструкция (64а, 64b) расположена в воображаемой плоскости (66а, 66b), параллельной оси (5) турбореактивного двигателя или проходящей через эту ось, а также проходящей ...

УЗЕЛ ДВИГАТЕЛЯ САМОЛЕТА С ПОДВИЖНОЙ ГОНДОЛОЙ ДВИГАТЕЛЯ

Изобретение относится к области авиации, более конкретно к узлу двигателя самолета с подвижной гондолой двигателя. Узел (1) двигателя самолета содержит турбореактивный двигатель (2), пилон крепления (4) и гондолу (3), установленную на пилоне крепления. Гондола содержит подвижный участок (40), образующий единый цельный кожух вокруг секции турбореактивного двигателя, при этом данный участок (40) гондолы имеет кольцевую стенку (50), обеспечивающую внутреннее разграничение канала (24) кольцевого вторичного потока, и кольцевую стенку (46), обеспечивающую внешнее разграничение канала кольцевого вторичного потока. Стенки (50) и (46) оборудованы обшивкой (80) акустической защиты. Подвижный участок (40) гондолы установлен свободно с возможностью скольжения на пилоне, а задний конец (18) корпуса (12) вентилятора турбореактивного двигателя, на котором каждое из креплений (6а, 6b, 8) зафиксировано для обеспечения крепления двигателя к пилону, образует внутреннюю радиальную поддержку для подвижного ...

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ С ВЫСОКОСКОРОСТНОЙ ТУРБИННОЙ СЕКЦИЕЙ НИЗКОГО ДАВЛЕНИЯ И ХАРАКТЕРНЫМИ ОСОБЕННОСТЯМИ ОПОРЫ ПОДШИПНИКОВ

Газотурбинный двигатель содержит очень высокоскоростную турбину привода вентилятора, при этом отношение параметра, определяемого произведением площади выходного сечения турбины низкого давления на квадрат скорости вращения турбины низкого давления, к такому же параметру турбины высокого давления составляет от 0,5 до 1,5. Турбина высокого давления установлена с помощью подшипников, расположенных на внешней периферии вала, который приводится во вращение турбиной высокого давления. Достигаются увеличенный коэффициент полезного действия и уменьшенные размеры турбинной секции. 3 н. и 17 з.п. ф-лы, 3 ил.

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ, СОДЕРЖАЩИЙ КОМПОЗИТНУЮ ДЕТАЛЬ И МЕТАЛЛИЧЕСКУЮ ДЕТАЛЬ, СВЯЗАННЫЕ УСТРОЙСТВОМ УПРУГОГО КРЕПЛЕНИЯ

Газотурбинный двигатель содержит аксиальный кожух турбины низкого давления из металлического материала, на выходе которого установлен аксиальный выхлопной кожух из композитного материала, а также устройство упругого крепления, связывающее указанные кожухи между собой, элемент гибкой связи и жесткий блокирующий элемент. Элемент гибкой связи содержит первую фиксирующую часть, связанную с аксиальным кожухом турбины низкого давления первой осевой связью, и вторую фиксирующую часть, связанную с аксиальным выхлопным кожухом второй осевой связью. Элемент гибкой связи содержит расположенный по оси свободный конец. Жесткий блокирующий элемент содержит фиксирующую часть, связанную с аксиальным кожухом турбины низкого давления первой осевой связью, и расположенный аксиально свободный конец, выровненный радиально со свободным концом элемента гибкой связи и образующий средства упора последнего в случае радиальной деформации элемента гибкой связи в процессе работы двигателя. Изобретение позволяет повысить ...

КОРПУС ПОДВЕСКИ ТУРБОВИНТОВОГО ДВИГАТЕЛЯ

Полезная модель относится к авиационным двигателям, имеющим одноплоскостную подвеску. Корпус подвески двигателя сочетает в себе функции единственного пояса подвески двигателя в одной плоскости крепления и силового элемента, воспринимающего все инерционные и маневренные нагрузки объекта.Корпус подвески турбовинтового двигателя содержит цилиндрическую и конусную часть, крепежные фланцы и внутренний фланец, а также пояс подвески, который расположен на наружном участке цилиндрической части, включает крепежные площадки демпферов объекта. Корпус подвески одновременно является корпусом турбины, причем он выполнен из жаропрочного сплава единой цельной деталью, не имеющей сварных швов по наружной поверхности. На поясе подвески выполнены фланцы фиксации трубопровода, а крепежные площадки и фланцы фиксации трубопровода соединены ребрами жесткости. 4 ил.

СИЛОВАЯ УСТАНОВКА ЛЕТАТЕЛЬНОГО АППАРАТА, СОДЕРЖАЩАЯ УЗЛЫ ПОДВЕСКИ ДВИГАТЕЛЯ, СМЕЩЕННЫЕ ВНИЗ НА КОРПУСЕ ВЕНТИЛЯТОРА

Изобретение относится к области авиации, более конкретно к силовой установке летательного аппарата. Силовая установка содержит средства крепления турбореактивного двигателя (2) на жесткой конструкции (10) стойки крепления. Средства крепления содержат первый, второй и третий передние узлы (6а, 6b, 8) подвески двигателя, воспринимающие тяговые усилия и установленные на корпусе (12) вентилятора, выполненные таким образом, что третий узел (8) подвески проходит по первой диаметральной плоскости (Р1) турбореактивного двигателя. Первый и второй узлы (6а, 6b) подвески расположены по обе стороны от этой первой диаметральной плоскости (Р1). Узлы (6а, 6b) подвески установлены на корпусе (12) вентилятора соответственно в двух точках (6'а, 6'b), находящихся за пределами второй диаметральной плоскости (Р2) турбореактивного двигателя, ортогональной к первой плоскости (Р1), по отношению к третьему узлу (8) подвески двигателя. Технический результат заключается в предотвращении деформации корпуса вентилятора ...

СИСТЕМА КОМПОНОВКИ УЗЛОВ МАШИНЫ

Два отдельных компрессора - компрессор стороны низкого давления и компрессор стороны высокого давления (11А, 11В) - расположены по обе стороны приводного узла - паровой турбины (10). С внешней стороны от компрессора стороны низкого давления и компрессора стороны высокого давления (11А, 11В) установлены два отдельных детандера - детандер стороны низкого давления и детандер стороны высокого давления (12A, 12В). Паровая турбина (10), компрессоры (11A, 11В) стороны низкого давления и стороны высокого давления и детандеры (12A, 12В) стороны низкого давления и стороны высокого давления соединены валами роторов, образующими единый вал. Оптимизируется распределение крутящего момента по валам роторов, повышается компактность, надежность и ремонтопригодность. 2 з.п. ф-лы, 2 ил.

Устройство для соединения внутреннего и наружного корпусов турбомашины

Изобретение относится к области турбо- и авиадвигателестроения, а именно к устройствам соединения наружных и внутренних корпусов турбомашины. Устройство для соединения внутреннего и наружного корпусов 1 и 2 турбомашины содержит тяги 3, концы которых шарнирно соединены с соответствующими кронштейнами 5 корпуса 2 и кронштейнами 4 корпуса 1, а также промежуточный кольцевой элемент 9. Кронштейны 4 и 5 закреплены на соответствующих корпусах 1 и 2 посредством средства силового крепления. На внутренней поверхности корпуса 2, выполненного с фланцами горизонтального разъема 6, сформирован кольцевой паз, в котором установлен элемент 9, закрепленный с возможностью разъема на кронштейнах 5. При этом средства силового крепления со стороны корпуса 2 проходят насквозь через элемент 9 и образуют зазор с последним. Изобретение направлено на повышение удобства сборки, ускорение и упрощение процесса сборки при обеспечении требуемой взаимосвязи между внутренним и наружным корпусами в работе. 2 з.п. ф-лы, 3 ...

ГАЗОВАЯ ТУРБИНА С КАМЕРОЙ СГОРАНИЯ, ВЫПОЛНЕННОЙ ИЗ КОМПОЗИТНОГО МАТЕРИАЛА

... 1. Газовая турбина, содержащая кольцевую камеру (10) сгорания со стенками (12, 13), изготовленными из композитного материала с керамической матрицей, установленную внутри металлического корпуса при помощи соединительных компонентов, прикрепленных к камере методом пайки и соединяющих камеру с внутренней металлической оболочкой (30) и с внешней металлической оболочкой (40) корпуса, отличающаяся тем, что соединительные компоненты содержат внутренние соединительные пластины (50) и внешние соединительные пластины (60), которые соединяют камеру (10) сгорания соответственно с внутренней металлической оболочкой (30) и с внешней металлической оболочкой (40) корпуса, причем каждая из соединительных пластин содержит первый участок (52, 62), прикрепленный к наружной поверхности стенки (12, 13) камеры сгорания методом пайки, при этом первые участки соединительных пластин отделены друг от друга в направлении по окружности так, что паяное соединение между камерой и соединительными компонентами выполнено ...

ВНЕШНЯЯ ОБОЛОЧКА ВОЗДУХОВОДА ВЕНТИЛЯТОРА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

... 1. Двухконтурный турбореактивный двигатель, содержащий по существу цилиндрическую оболочку, установленную на выходе промежуточного кожуха и ограничивающую с внешней стороны кольцевое пространство протекания вторичного потока вокруг турбореактивного двигателя, отличающийся тем, что цилиндрическая оболочка образована решетчатым каркасом и съемными панелями обтекателя, закрепленными на каркасе, содержащим, по меньшей мере, один входной кольцевой фланец крепления к промежуточному кожуху, один выходной кольцевой фланец соединения с выхлопным кожухом, и жесткие балки, соединяющие оба фланца между собой. ! 2. Турбореактивный двигатель по п.1, отличающийся тем, что решетчатый каркас содержит балки, параллельные оси и равномерно распределенные вокруг оси. ! 3. Турбореактивный двигатель по п.1, отличающийся тем, что решетчатый каркас содержит жесткие балки, наклонные относительно оси, при этом каждая из этих балок имеет входной конец, жестко соединенный с входным кольцевым фланцем, и выходной конец ...

СТУПИЦА ЗАДНЕЙ ОПОРЫ ТУРБИНЫ НИЗКОГО ДАВЛЕНИЯ, СОДЕРЖАЩАЯ РЕБРА ЖЕСТКОСТИ, ПРЕДНАЗНАЧЕННЫЕ ДЛЯ ПЕРЕРАСПРЕДЕЛЕНИЯ НАПРЯЖЕНИЙ

... 1. Ступица в сборе задней опоры турбины низкого давления, содержащая ступицу (6), переднюю (8) и заднюю щеки (10), расположенные по одну и другую сторону от ступицы (6), а также множество манжет (14), расположенных на ступице (6), образующих угол между касательной и радиальным направлением по отношению к указанной ступице, отличающаяся тем, что, по меньшей мере, одно ребро (20, 30, 34) жесткости сформировано на основании каждой из манжет (14) в месте под критической зоной (26) напряжений соответствующей манжеты (14), причем указанное ребро (20, 30, 34) жесткости присоединено к задней щеке (8) своим задним концом, а к передней щеке (10) - своим передним концом. ! 2. Ступица в сборе п.1, отличающаяся тем, что содержит переднее ребро (30) жесткости, присоединенное к передней щеке (8), и заднее ребро (34) жесткости, присоединенное к задней щеке (10). ! 3. Ступица в сборе по любому из пп.1 или 2, отличающаяся тем, что длина манжет ступицы варьируется так, что может доходить до самой большой ...

ГОНДОЛА ДЛЯ ДВИГАТЕЛЯ ЛЕТАТЕЛЬНОГО АППАРАТА, ИМЕЮЩАЯ СОПЛО С РЕГУЛИРУЕМЫМ СЕЧЕНИЕМ

... 1. Гондола (57) для двигателя летательного аппарата, содержащая передний обтекатель (13) и задний обтекатель (1a), причем задний обтекатель (1а) установлен с возможностью смещения между верхним по потоку положением, соответствующим малому поперечному сечению сопла (9), и нижним по потоку положением, соответствующим увеличенному поперечному сечению сопла (9), отличающаяся тем, что она содержит промежуточный элемент (25), расположенный встык с указанным передним обтекателем (13) и ограничивающий собой полость (27), принимающую в себя верхний по потоку край (11) указанного заднего обтекателя (1а), когда данный обтекатель находится в своем верхнем по потоку положении. ! 2. Гондола по п.1, отличающаяся тем, что указанный промежуточный элемент (25) имеет зону (35), используемую для размещения на ней указанного переднего обтекателя (13) и расположенную выше по потоку от указанной полости (27). ! 3. Гондола по п.2, отличающаяся тем, что указанный передний обтекатель (13) снабжен средствами (39) ...

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ

Газотурбинный двигатель содержит редуктор, расположенный вдоль продольной оси двигателя, каскад, гондолу вентилятора, внутреннюю гондолу, вентилятор, вентиляторное сопло и внутренний контур. Каскад выполнен с возможностью приведения в действие редуктора и содержит турбину низкого давления с числом ступеней от трех до шести. Гондола вентилятора установлена вокруг внутренней гондолы и определяет тракт для воздушного потока наружного контура вентилятора, причем степень двухконтурности превышает шесть. Вентиляторное сопло выполнено с изменяемой площадью сечения и с возможностью перемещения в осевом направлении относительно гондолы вентилятора с целью изменения площади выходного сечения вентиляторного сопла и регулирования воздушного потока в наружном контуре вентилятора во время работы двигателя. Вентилятор выполнен с возможностью вращения со скоростью вентилятора вокруг продольной оси и приводится в действие турбиной низкого давления с помощью редуктора, причем скорость вентилятора меньше ...

РЕДУКТОРНОЕ УСТРОЙСТВО ДЛЯ ВЫСОКОСКОРОСТНОЙ И МАЛОГАБАРИТНОЙ ТУРБИНЫ ПРИВОДА ВЕНТИЛЯТОРА

... 1. Газотурбинный двигатель, содержащий:вал вентилятора, приводящий в действие вентилятор;раму, поддерживающую указанный вал вентилятора;множество зубчатых колес для приведения в действие указанного вала вентилятора;гибкую опору, по меньшей мере, частично поддерживающую указанное множество зубчатых колес, причем жесткость указанной гибкой опоры меньше жесткости указанной рамы;первую турбинную секцию, обеспечивающую вход привода в указанное множество зубчатых колес; ивторую турбинную секцию,в котором указанная первая турбинная секция имеет первую выходную площадь в первой выходной точке и способна вращаться с первой частотой, указанная вторая турбинная секция имеет вторую выходную площадь во второй выходной точке и способна вращаться со второй частотой, превышающей первую частоту, при этом первый характеризующий параметр является произведением квадрата первой частоты и первой площади выхода, второй характеризующий параметр является произведением квадрата второй частоты и второй площади выхода ...

КОЛЬЦЕВАЯ КАМЕРА СГОРАНИЯ ТУРБОМАШИНЫ

ГИБКАЯ ПОДДЕРЖИВАЮЩАЯ КОНСТРУКЦИЯ ДЛЯ ЗУБЧАТОЙ ТРАНСМИССИИ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

... 1. Газотурбинный двигатель, содержащий вал вентилятора; раму, которая поддерживает вал вентилятора, причем для указанной рамы определена собственная боковая жесткость и собственная поперечная жесткость; зубчатую трансмиссию, которая приводит во вращение вал вентилятора; гибкую опору, которая по меньшей мере частично поддерживает зубчатую трансмиссию, причем для гибкой опоры определена собственная боковая жесткость по отношению к боковой жесткости рамы, и собственная поперечная жесткость по отношению к поперечной жесткости рамы; и входную муфту зубчатой трансмиссии, причем для входной муфты определена собственная боковая жесткость по отношению к боковой жесткости рамы, и собственная поперечная жесткость по отношению к поперечной жесткости рамы.2. Газотурбинный двигатель по п. 1, отличающийся тем, что боковая жесткость гибкой опоры меньше боковой жесткости рамы.3. Газотурбинный двигатель по п. 1, отличающийся тем, что поперечная жесткость гибкой опоры меньше поперечной жесткости рамы.4. Газотурбинный ...

ПРИВОДНАЯ КОНСТРУКЦИЯ ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ (ВАРИАНТЫ)

... 1. Приводная конструкция для газотурбинного двигателя, содержащая:вал вентилятора;раму, поддерживающую вал вентилятора и задающую репер поперечной жесткости (Kframe);зубчатую систему, приводящую во вращение вал вентилятора;гибкую несущую конструкцию, по меньшей мере частично поддерживающую зубчатую систему и имеющую поперечную жесткость (KFS), определяемую относительно поперечной жесткости (Kframe) рамы, ивходной узел зубчатой системы, имеющий поперечную жесткость (KIC), определяемую относительно поперечной жесткости (Kframe) рамы.2. Конструкция по п.1, в которой рама и гибкая несущая конструкция установлены на статической конструкции.3. Конструкция по п.1, в которой рама и гибкая несущая конструкция установлены на статической конструкции в составе газотурбинного двигателя.4. Конструкция по п.1, в которой рама и гибкая несущая конструкция установлены на переднюю центральную часть корпуса газотурбинного двигателя.5. Конструкция по п.1, в которой гибкая несущая конструкция связана с водилом ...

УСТРОЙСТВО ДЛЯ КРЕПЛЕНИЯ КОЛЬЦА ТУРБИНЫ, ТУРБИНА С ТАКИМ УСТРОЙСТВОМ И ТУРБИННЫЙ ДВИГАТЕЛЬ С ТАКОЙ ТУРБИНОЙ

... 1. Устройство для крепления кольца (5) газовой турбины (1), при этом кольцо (5) предназначено для охватывания подвижных лопаток (3) турбины (1), приводимых в движение газовым потоком (G), протекающим с входа на выход, причем устройство содержит, по меньшей мере, один входной зацеп (12), обращенный к входу, предназначенный для размещения во входной канавке (8) кольца (5), открытой к выходу, и, по меньшей мере, один выходной зацеп (13), обращенный к выходу, предназначенный для размещения в выходной канавке (10) кольца (5), открытой к входу, при этом полость (С) повышенного давления, запитываемая охлаждающим газом, сформирована между входным (12) и выходным (13) зацепами, причем устройство содержит на входе входного зацепа (12) средства (16) подачи охлаждающего газа для охлаждения входного зацепа (12) и/или содержит на выходе выходного зацепа (13) средства (17, 18) подачи охлаждающего газа для охлаждения выходного зацепа (13).2. Система, состоящая из устройства по п.1 и кольца турбины, содержащего ...

УСИЛИВАЮЩАЯ КОНСТРУКЦИЯ ГОНДОЛЫ ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ

... 1. Гондола турбореактивного двигателя, содержащая неподвижную конструкцию, содержащую кожух (4) вентилятора указанного турбореактивного двигателя и переднюю раму (7), которая выполнена с возможностью установки за указанным кожухом (4) вентилятора и поддерживает, непосредственно или опосредованно по меньшей мере одно средство отклонения потока, причем указанная передняя рама выполнена с возможностью взаимодействия с капотом (6) реверса тяги, установленным с возможностью скольжения из закрытого положения с перекрытием средства отклонения потока в открытое положение с открытием этого средства отклонения, обеспечивая возможность отклонения потока, отличающаяся тем, что она дополнительно содержит, по меньшей мере, одну армирующую конструкцию (100), предназначенную для передачи усилий между кожухом (4) вентилятора и передней рамой (7), причем указанная армирующая конструкция (100) расположена вдоль продольной оси гондолы, проходя от кожуха (4) вентилятора к передней раме (7), и выполнена с возможностью ...

УЗЕЛ, ОБЕСПЕЧИВАЮЩИЙ УДЕРЖАНИЕ СРЕДСТВА СОПРЯЖЕНИЯ НЕПОДВИЖНОЙ НАРУЖНОЙ КОНСТРУКЦИИ ГОНДОЛЫ И КАРТЕРА РЕАКТИВНОГО ДВИГАТЕЛЯ

... 1. Узел (101), обеспечивающий удержание средства сопряжения неподвижной наружной конструкции (15) гондолы (3) и картера (27) реактивного двигателя (5), содержащий:- первый выступ (105), являющийся частью переднего по потоку конца неподвижной наружной конструкции (15);- второй выступ (107), являющийся частью заднего по потоку конца картера (27);причем указанные первый (105) и второй (107) выступы выполнены с возможностью размещения в соприкосновении друг с другом;- два полукольца (109), образованные стенкой, которая ограничивает собой гнездо (111), форма которого обеспечивает возможность ввода в него указанных первого (105) и второго (107) выступов, когда картер (27) и неподвижная наружная конструкция (15) установлены встык друг с другом, и упорное средство (113), выполненное с возможностью удержания указанных первого (105) и второго (107) выступов в указанном гнезде (111).2. Узел (101) по п.1, в котором указанная стенка ограничивает собой гнездо (111) по существу U-образного или V-образного ...

СИЛОВАЯ УСТАНОВКА

... 1. Силовая установка (80), содержащая газотурбинный двигатель (82), имеющий первый участок (88), второй участок (84) и продольную ось (92); первую подвижную опору (30), контактирующую с первым участком газотурбинного двигателя, при этом первая подвижная опора обеспечивает, по меньшей мере, частичное перемещение первого участка газотурбинного двигателя во время обслуживания газотурбинного двигателя и, по меньшей мере, частично поддержку первого участка газотурбинного двигателя, при перемещении первого участка газотурбинного двигателя в направлении от продольной оси, независимо от второго участка газотурбинного двигателя. ! 2. Силовая установка по п.1, характеризующаяся тем, что первая подвижная опора обеспечивает практически полную поддержку первого участка газотурбинного двигателя при перемещении первого участка газотурбинного двигателя в направлении от продольной оси, независимо от второго участка газотурбинного двигателя. ! 3. Силовая установка по п.1, характеризующаяся тем, что дополнительно ...

ИНТЕГРИРОВАННАЯ СИЛОВАЯ УСТАНОВКА С ПОДВЕСКОЙ ДЛЯ САМОЛЕТА

... 1. Интегрированная силовая установка, содержащая авиационный двухконтурный турбореактивный двигатель и гондолу, закрепленную на промежуточном картере и ограничивающую кольцевое пространство прохождения вторичного потока вокруг турбореактивного двигателя, отличающаяся тем, что гондола содержит заднюю цилиндрическую часть, которая выполнена жесткой, и содержит на своем переднем конце кольцевой фланец крепления, по меньшей мере, на 180° наружной окружности промежуточного картера, при этом цилиндрическая часть, поддерживающая и направляющая задний картер турбореактивного двигателя на заднем конце, содержит также средства крепления элементов подвески двигателя на части (22) самолета. 2. Силовая установка по п.1, отличающаяся тем, что средства крепления элементов подвески содержат продольную балку, выполненную отдельно или заодно с задней цилиндрической частью гондолы. 3. Силовая установка по п.2, отличающаяся тем, что элементы подвески двигателя на самолете содержат соединительные тяги или штанги ...

ДВИГАТЕЛЬ В СБОРЕ САМОЛЕТА, СОДЕРЖАЩИЙ КОЛЬЦЕВУЮ НЕСУЩУЮ КОНСТРУКЦИЮ, ОКРУЖАЮЩУЮ ЦЕНТРАЛЬНЫЙ КОРПУС ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ

... 1. Двигатель (1) в сборе самолета, содержащий: ! - турбореактивный двигатель (2), включающий в себя корпус (12) вентилятора, промежуточный корпус (21), расположенный внутри корпуса вентилятора и соединенный с ним несколькими ребрами (17), а также центральный корпус (16), расположенный за вышеуказанным промежуточным корпусом в направлении к задней части; ! - кольцевую несущую конструкцию (60), окружающую центральный корпус (16), и соединенную с ним посредством промежуточных крепежных средств, выполненных в виде нескольких соединительных штанг (62), при этом вышеуказанная кольцевая несущая конструкция также соединена с несколькими плоскими конструкциями (64а, 64b, 64с), расположенными по отношению к ней снаружи и поддерживающими ее соответственно в нескольких точках (68а, 68b, 68с), расположенных в окружном направлении, ! отличающийся тем, что каждая из, по меньшей мере, одной из соединительных штанг (62) связана с соответствующей из точек (68а, 68b, 68с), при этом вышеуказанная соединительная ...

ПОДВЕСКА КАНАЛА ХОЛОДНОГО ПОТОКА ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ НА ВЫПУСКНОМ КОРПУСЕ ПРИ ПОМОЩИ ТЯГ И РАДИАЛЬНЫХ ВИЛОК

... 1. Двухконтурный турбореактивный двигатель, содержащий наружный цилиндрический канал (105) холодного потока, установленный при помощи тяг (3), прикрепленных к цилиндрической наружной обечайке (11) выпускного корпуса (1) на уровне точек (4) крепления, отличающийся тем, что точки крепления выпускного корпуса представляют собой вилки, проушины которых проходят радиально от упомянутой наружной обечайки, при этом отверстие (6) упомянутых вилок ориентировано в направлении образующих наружной обечайки (11).2. Турбореактивный двигатель по п.1, в котором тяги соединяются с упомянутым наружным каналом (105) холодного потока, являясь касательными к упомянутой наружной обечайке.3. Турбореактивный двигатель по п.1, в котором вилки выполнены в четном количестве и связаны по две в одной паре, и расположены на окружности наружной обечайки (11) таким образом, чтобы тяги, закрепленные на двух соответствующих вилках, сходились в одной точке на упомянутом наружном канале (105) холодного потока.4. Турбореактивный ...

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ И СПОСОБ РАЗБОРКИ ПЕРЕДНЕЙ ЧАСТИ КОНСТРУКЦИИ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

... 1. Газотурбинный двигатель, содержащий:опору центрального узла, образующую внутреннюю кольцевую стенку для осевого контура и содержащую первое монтажное средство;узел зубчатой передачи, связывающей вал и вентилятор, установленный с возможностью вращения вокруг оси, игибкую опору, связывающую узел зубчатой передачи с опорой центрального узла и содержащую второе монтажное средство, сопрягаемое с первым монтажным средством для передачи крутящего момента от одного монтажного средства к другому.2. Двигатель по п. 1, в котором опора центрального узла содержит пространственно разделенные направляющие лопатки, расположенные радиально между внутренней и наружной кольцевыми стенками и соединяющие их.3. Двигатель по п. 2, в котором первое монтажное средство содержит взаимно смещенные по окружности группы зубцов, разделенные участками, не имеющими зубцов.4. Двигатель по п. 3, в котором направляющие лопатки согласованы по положению в радиальном направлении с участками, не имеющими зубцов.5. Двигатель ...

СПОСОБ УСТАНОВКИ ДВИГАТЕЛЯ ЛЕТАТЕЛЬНОГО АППАРАТА НА ПИЛОНЕ И КРЕПЛЕНИЕ ДВИГАТЕЛЯ ДЛЯ ОСУЩЕСТВЛЕНИЯ ТАКОГО СПОСОБА

... 1. Способ установки двигателя (40) летательного аппарата на пилоне (43), отличающийся тем, что он включает в себя следующие этапы:- размещают, по меньшей мере, один первый срезной штифт (53) в:- или в первом отверстии (51), выполненном в переднем креплении (46) двигателя, предварительно закрепленном на двигателе (40) посредством тяг,- или в первом проеме, выполненном в передней поверхности (48) пилона (43);- предварительно располагают узел, включающий в себя двигатель (40) и переднее крепление (46) двигателя, относительно пилона (43), размещая первый срезной штифт (53) против приемной полости, причем упомянутая приемная полость является:- или первым проемом в том случае, если первый срезной штифт (53) был размещен в первом отверстии (51);- или первым отверстием (51) в том случае, если первый срезной штифт (53) был размещен в первом проеме;- вставляют первый срезной штифт (53) в приемную полость.2. Способ установки двигателя (40) летательного аппарата на пилоне (43) по п.1, отличающийся ...

КОНСТРУКЦИЯ СОЕДИНИТЕЛЬНОЙ ЧАСТИ ДЛЯ ЛОПАТКИ И РЕАКТИВНЫЙ ДВИГАТЕЛЬ, ВКЛЮЧАЮЩИЙ В СЕБЯ ТО ЖЕ

... 1. Конструкция соединительной части для лопатки, которая составляет реактивный двигатель и выполнена из композитного материала из термореактивной смолы или термопластичной смолы и усилительного волокна, причем конструкция соединительной части содержит соединительную часть лопатки, в которойсоединительная часть лопатки включает в себя помещенный в нее соединительный поддерживающий элемент, причем соединительный поддерживающий элемент выполнен из металла и включает в себя пару разделенных деталей, отделенных друг от друга, причем пара разделенных деталей присоединена к концевой части лопатки с обеих сторон в направлении толщины лопатки,по меньшей мере любая одна из поверхностей стыка концевой части лопатки с парой разделенных деталей, имеет по меньшей мере один линейный выступ или канавку, образованный(ую) на ней,по меньшей мере любая одна из соответствующих концевых поверхностей стыка пары разделенных деталей соединительного поддерживающего элемента имеет канавку или линейный выступ, образованную ...

Strömungsmaschine mit einer in einer Ausnehmung einer Wandung anordenbaren Einsetzeinrichtung

Es wird eine Strömungsmaschine (10) mit einer sich zumindest bereichsweise umfangsseitig zu einer Zentralachse (1) der Strömungsmaschine (10) erstreckenden Wandung (32) und mit wenigstens einer Einsetzeinrichtung (48), die im Bereich einer in der Wandung (32) angeordneten Ausnehmung (42) lösbar mit der Wandung (32) verbindbar ist. Die Einsetzeinrichtung (48) weist eine Verbindungseinrichtung (64) auf, mittels der die Einsetzeinrichtung (48) in Wirkverbindung mit einer Verbindungseinrichtung (65) der Wandung (32) bringbar ist. Die Verbindungseinrichtung (64) der Einsetzeinrichtung (48) weist eine sich in montiertem Zustand der Einsetzeinrichtung (48) zumindest bereichsweise in radialer Richtung (R) der Strömungsmaschine (10) erstreckende Vorkragung (72) auf, die in montiertem Zustand der Einsetzeinrichtung (48) in eine Nut (80) der Verbindungseinrichtung (65) der Wandung (32) eingreift. Die Verbindungseinrichtung (64) der Wandung (32) weist eine sich in montiertem Zustand der Einsetzeinrichtung ...

Turboladerstütze

Verpuffungskammer fuer Gasturbinen

Device for connecting inner wall and outer wall of bypass filter channel of turbojet engine unit, has two wall-like supporting elements, which are partially arranged in assembled state in direction of fluid flow in bypass filter channel

The device (11) has two wall-like supporting elements (15,16), which are partially arranged behind one another in an assembled state in a direction of a fluid flow (21) in a bypass filter channel. The sections (15A,16A) overlap the supporting elements in the areas of side surfaces (15B,16B) running in the direction of the fluid flow. The supporting elements of a support device (12) are adjustable executed to each other in mutually facing areas (22,23).

Gasturbine mit gegenläufigen Niederdruckrotoren

Gasturbinenanlage mit einer Nutzleistungsturbine

Verbesserte Lagerung fuer Gasturbinenstrahltriebwerke

A transducer mount and a transducer assembly incorporating such a mount

A transducer assembly 100 for connecting a transducer 102 to a device, the transducer assembly comprising: a transducer mount 110 adapted to receive the transducer; and a clamp 120 comprising first and second jaws 122, 124, the first and second jaws being adapted to selectively clamp to a flange portion 106 of the device. The transducer mount 110 may comprise a housing 112, a resilient member 114 and first and second end walls 115, 116. The first end wall may be operatively connected to the device. The transducer 102 may be held within the housing 112 and adjacent to the first end wall 115, the resilient member 114 being disposed between the transducer 102 and the second end wall 116.

Arrangement for connecting a duct to an air-distribution casing

The invention relates to an arrangement for connecting at least one duct (12) to a casing (10) for distributing air towards said at least one duct (12), which comprises two sidewalls (14) opposite one another, and which comprises a peripheral wall (16) connecting the edges of the two planar walls, characterised in that it comprises a connecting tube which extends through the casing (10), passing through an opening associated with each of the two sidewalls (14), and which is connected to said at least one duct (12). The invention also relates to a system for controlling the clearance of a turbine engine, and to a turbine engine comprising air-injection ducts (12) that are connected to the distribution casing (10) by such an arrangement.

An exhaust gas turbocharger coupling assembly

Aircraft engine assembly, comprising an attachment device for the engine equipped with a structural cover attached on a central box

GAS TURBINE ENGINE ANNULAR SEAL

... 1434492 Seals FORD MOTOR CO Ltd 30 May 1974 [7 June 1973] 23912/74 Heading F2B [Also in Division F1] A gas turbine engine includes a seal assembly 44 disposed between annular, axially spaced, turbine stator shroud members 32, 33 and comprising a pair of resilient, divergent, annular leaves 48, 49 clamped between a U-shaped annular base 51. The latter may be located by an annular flange 52 engaging within recess 52. Flange 52 may be omitted.

Air supply device for aircraft engine turbines

A panel attachment system and a method of using the same

Duct

PREMOUNTING RETAINER

Improvements in gas-turbo-generator installations

... 719,316. Machinery beds; universal joints. FEILDEN, G. B. R., and RUSTON & HORNSBY, Ltd. March 2, 1953 [Dec. 20, 1951], No. 29856/51. Classes 80 (2) and 80 (4) [Also in Group XXVI] In order to facilitate transport over rough country and erection in the field of a gasturbo-generator installation, three separate bedplates 8, 9, 10 are provided and are bolted together on site, each plate having supporting rollers, which may be retractable, and built-in jacks and spirit levels, the plates 8, 9, 10 having detachably secured thereto an alternator 7 with its exciter, a power-turbine 21 with its reduction gearing 35, and a compressor-turbine unit 11, 12. A duct 18 conveys air from the compressor 11 to an external combustion chamber from which hot gas is conveyed to the turbine 12. Exhaust of the turbine 12 feeds the power-turbine 21 whose rotor 73, Fig. 4b, is connected by a shaft 88 and detachable flexible coupling 89 to the input shaft of the reduction gearing 35. The output shaft 92 of the gearing ...

Vibration isolating aircraft engine mount

A method and apparatus for coupling an engine to a support frame for mounting to a fuselage of an aircraft using a three point vibration isolating mounting system in which the load reactive forces at each mounting point are statically and dynamically determined. A first vibration isolating mount pivotably couples a first end of an elongated support beam to a stator portion of an engine with the pivoting action of the vibration mount being oriented such that it is pivotable about a line parallel to a center line of the engine. An aft end of the supporting frame is coupled to the engine through an additional pair of vibration isolating mounts with the mounts being oriented such that they are pivotable about a circumference of the engine. The aft mounts are symmetrically spaced to each side of the supporting frame by 45 degrees. The relative orientation between the front mount and the pair of rear mounts is such that only the rear mounts provide load reactive forces parallel to the engine ...

Disassembly of a modular fan gas turbine engine

A gas turbine fan engine with improved maintainability features is disclosed. Forward and rearward sections of the core of the engine are joined one to the other, and each are independently separable from the pylon on which the engine is mounted without the necessity of removing the other from the airplane.

Recuperative exhaust gas heat exchanger for a gas turbine engine

A recuperative exhaust gas heat exchanger 10 for a gas turbine engine comprises a crossflow matrix 16 of tubes 17 around which hot turbine exhaust gas 23 flows, a distributing tube 12 directing air into the matrix of tubes 17, a collecting tube 14 arranged parallel to the distributing tube 12 and discharges air which has been heated bypassing through the matrix of tubes 17. Both end faces of the distributing 12 and collecting 14 tubes are closed with the closed end 26 of the collecting tube 14 having a cover plate 28 connected axial and radially to a casing 30 of the turbine. The other end 24 of the tube 14 is not connected to the casing 30 and permits axial movement. Matrix 16 of tubes 17 may be U-shaped. A plurality of spacers 18 may be provided to ensure that the shape of the tubes 17 in a bundle 20 is retained. The air to be heated may be from a compressor and when heated may be fed to a combustion chamber. Distributing tube 12 may be fastened to the casing 30 in a similar manner as ...

Plates for clamping overlapping panels and bands

Panels 42 and bands 38 are joined together by overlapping to form a flowpath assembly in a gas turbine, by means of tripod plates 50 and fasteners 46. Each tripod plate 50 has a part-spherical seat 56 for receiving a washer 52, and a pair of projections 58 and a single leg 60 for respectively engaging the panel 42 and the band 38. Circumferential movement is permitted between the panels and bands while maintaining sealing. ...

A mounting for coupling a turbofan gas turbine engine to an aircraft structure

Mounting for a gas turbine engine

A gas turbine engine, is pivotally mounted about its centre of gravity to an aircraft support structure by means of a pair of radially outwardly extending trunnions. Each trunnion 28 is connected to a central conical section 16 of engine casing by means of an axially extending support member 34. The support member 34 is secured to the downstream end of the conical casing section by a flanged joint 24 at an axial location offset from the trunnion. The support member is further connected to the casing by an arrangement of links 44, 46, pins 56, 58 a spherical bearings 52, 54 in such a way that vertical loads due to the weight of the engine are transmitted only through the flanged joint 24. During operation, distortion of the less stiff conical section is reduced by virtue of the resultant shear force distribution across the casing. Alternative arrangements (involving link elements, pins and spherical bearings) for connecting the support member 34, to the casing section 16, are disclosed ( ...

Mounting a turbine nozzle on a combustion chamber having CMC walls in a gas turbine

The gas turbine comprises an annular combustion chamber (10) having inner and outer walls (12, 13) made of ceramic matrix composite material, and a high pressure turbine nozzle (20) secured to a downstream end of the combustion chamber and comprising a plurality of stationary airfoils (21) extending between the inner and outer walls (22, 23) of an annular flow path (24) through the nozzle for the gas stream coming from the combustion chamber. The turbine nozzle (20) is made of ceramic matrix composite material and it is connected to the downstream end of the combustion chamber (10) by brazing.

Fire resistant seal assembly

Mounting arrangement

Turbofan engine

A turbofan engine having a power generating portion of the engine supporting a fan and surrounded by a cowl. The cowl is split with the two sections being arcuate and hinged so that they can pivotally open for access and removability of the power generating portion of the engine. The split cowl is supported directly from a support pylon whereby there is no fan frame required.

Gas turbine engine assembly with vibration isolation mount and method for producing the same

A gas turbine engine (GTE) assembly 40 includes a strut-based vibration isolation mount 44 which comprises at least one three parameter axial strut 46-51 having a first end attached to the gas turbine engine 42 and having a second, opposing end configured to be attached to an airframe (26, fig 1). The or each three parameter axial strut is tuned to minimise the transmission of vibrations from the gas turbine engine to the airframe during operation of the gas turbine engine. Also provided is a method for producing a GTE assembly having a number of three parameter axial struts that are tuned to have stiffness and damping profiles that vary in multiple degrees of freedom.

Mechanical system for a turbine engine, turbine engine, and method for attaching a mechanical system within a turbine engine

The invention relates to a mechanical system (100) for a turbine engine, including a turbine engine part to be attached, and a plurality of threaded attachment elements (310, 330) that are mounted onto the part, one after the other, along a line (250), wherein attachment elements are arranged so as to rotatably interlock by engaging with one another. Said self-locked elements (310, 330) engage in pairs (300), each pair (300) including two self-locking elements directly adjacent along the line (250). The distance between the thread axes (250) of the two self-locked elements of at least one of said pairs (300) is strictly less than the distance between either of said two thread axes and the axis of either of the two attachment elements that are arranged on the line (250) on either side of said pair (300).

A panel attachment system and a method of using the same

A panel attachment system 100, suitably a system for an exhaust system of a gas turbine engine, comprises a panel 110, a hanger 120, a plurality of anchors 150 and a casing 170. The hanger comprises a hanger plate 130 and a plurality of bosses 140. The hanger plate comprises a first edge 132 and an opposite second edge 134 with the plurality of bosses being spaced along the first edge of the hanger plate, and a plurality of slots 132 being spaced along the second edge of the hanger plate, each of the slots having a first T�shaped cross sectional area (138, fig. 4). Each anchor comprises an elongate post portion 152 and a planar cap portion 154 with each anchor having a second T�shaped cross sectional area. The spacing between adjacent anchors corresponding to the spacing between adjacent slots in the second edge of the hanger plate, with each slot receiving a respective anchors to secure the panel to the casing.

Access to subassemblies of a turbomachine

A turbomachine has a casing 2 in which a rotor 4 is mounted and an annular supporting structure 3 provided in the casing on which subassemblies 5, such as combustion chambers and cooling systems are mounted. The annular supporting structure 3 is rotatably mounted in the casing 2 and can be locked in at least one position. When subassemblies 5 require checking or exchanging the annular supporting structure 3 may be rotated into inspection or assembly positions where they are accessible through closable assembly openings 6 in the casing 2. More than one closable assembly opening 6 may be present.

Gas turbine engine

Disclosed is a gas turbine engine (1) that combusts, with a combustor (3), compressed air (CA) compressed by a compressor (2), and drives a turbine (5) by a high-temperature, high-pressure combustion gas (G) which has been generated, the gas turbine engine comprising: a plurality of air passages (11, 12) that connect different parts within the gas turbine engine (1); and a switching device (17) that switches the flow path of air between the plurality of air passages (11, 12). Each of the plurality of air passages (11, 12) is formed by an air pipe provided outside a compressor casing (26). The switching device (17) is attached to an outer surface of the compressor casing (26) by means of a heat-shielding member (50).

Method of producing suspension for a structure in a turbojet engine using a hyperstatic trellis with pre-stressed link elements

A method of manufacturing a suspension assembly for a structure in a turbojet engine is provided. The assembly comprises: a first structure 23, arranged to be rigidly connected to a housing of a turbojet engine; a second annular structure 21 surrounding the first structure 23; and a hyperstatic trellis of connecting rods 40a, 40b, 40c, 40d, 40e, 40f maintaining the first structure 23 relative to the second 21. The method comprises a step of mounting the connecting rods of the hyperstatic trellis between the structures 21, 23 and a step for pre-stressing at least one of the connecting rods to a predetermined level. The pre-stressing step is carried out before the mounting of the connecting rod between the structures. Reference is also made to a device which is suitable for mounting the pre-stressed connecting rods.

Device and method for regulating a motor using a thrust measurement

A device 22 for regulating the flow rate of fuel supplied to an aircraft engine M produces a fuel flow rate set value Dc according to a thrust set value P* from a gas control lever 1 and a measurement of actual thrust P of the engine from an actual thrust measuring device 21 which measures deformation of a device for absorbing thrust forces interposed between the engine M and the aircraft. A fuel flow rate set value calculator 24 may produce fuel flow rate set value Dc. A comparator 25 may supply a difference signal (ΔP figure 1) between thrust set value P* and actual thrust P and calculator 24 may produce fuel flow rate set value Dc from this signal. An engine speed set value calculator 23 may produce an engine speed set value N1* from thrust set value P* and actual thrust P and a circuit 27 may produce fuel rate set value Dc from engine speed set value N1* and a measurement of the actual speed of the engine N1. The calculator 23 may use isolated measurements of actual thrust P. A module ...

Case deflection control in aircraft gas turbine engines

A gas turbine engine and structure for mounting the gas turbine engine on an aircraft is disclosed. Techniques for controlling the magnitude of engine case deflection from the axis of the engine are developed. The transfer of a portion of the gust and thrust loads externally of the engine case to a location downstream of the rearward engine mount generates a reverse moment which counteracts deflection in the engine case.

Power gas turbine

A gas turbine power unit is disclosed in which the arrangement and configuration of parts is such as to save space and weight in order to provide a compact and self-contained assembly. An air-intake casing (2) supports the upstream end of a gas generator (3), the downstream end of which is integral with a power turbine (4). The stator casing (5) of the turbine (4) is connected to a cone (6) thermally insulated and completely inserted into any exhaust casing (7) having a vertical outlet, wherein the turbine exhaust is conveyed into the exhaust casing (7) by an annular diffusing cone (8). The turbine casing is supported on four legs (15, 16). In addition, the turbine rotor (10) and thus the turbine shaft (11) are overhangingly supported by an independent structure (12), the weight of which bears on the machine base (1) outside the exhaust casing (7) and away of the power turbine space.

Improvements relating to gas-turbine power-plant installations

... 620,820. Propelling ships. STARK, F. H. Jan. 27, 1947, Nos. 2538 and 18559. Drawings to Specification. [Class 114] [Also in Group XXVI] A gas turbine plant is carried in a rigid cradle having a three-point support on the ship's hull. It drives the propeller shaft through two reduction gears, one of which is carried by the cradle. In the second Provisional Specification it is stated that the two gears may be connected by two parallel shafts, each offset from the general axis, and the propeller blade pitch-changing gear may be located between the shafts.

In-line removable heat shield for a turbomachine suspension yoke

Device for fastening turbochargers

An Aircraft Powerplant

Turbomachine

A turbomachine 1 with axially spaced guide vane rings 35, 36, wherein at least two of the guide vane supports 21, 18 are jointly fastened to and centred on the same housing flange 22. Fastening may be realised by connecting elements 17, such as screws, extending through flanges 9, 23 of the guide vane supports 21, 18 and into the housing flange. The flanges have preferably mating surfaces 3, 4, 20, 28 in an axial direction 16. Radial 15 and circumferential (12, fig. 2) adjustment of the first guide vane support 21 may be effected via first centring elements 24 extending through flange 9 and into casing flange 22. An analogous adjustment is possible for the second guide vane support 18 via second centring elements (33, fig. 3) extending through both flanges 9, 23 and into casing flange 22. Both centring elements may have a first, cylindrical section 6 received within cylindrical bores and a second section 37 of cylindrical basic contour with opposite flats (7, fig. 2) received within elongate ...

Assembly for attaching a nozzle to a structural element of a turbine engine

The invention relates to an assembly (10) comprising: an annular nozzle (4) of a turbine of a turbine engine; a structural annular element (6) of the turbine engine; the nozzle (4) and the structural annular element (6) each comprising a radial flange (16, 18), the flanges being applied axially onto one another; and at least one member (12) for axially retaining the flange (16) of the nozzle (4) with the flange (18) of the structural annular element (6), the member being applied to one of the flanges and being capable of axially retaining the two flanges on top of one another.

Force measurement system

Aircraft engine assembly, comprising an engine attachment device equipped with structural movable cowls connected to the central box

Combustion turbine power units

... 787,739. Gas turbine plants. NAPIER & SON, Ltd., D. March 23, 1956 [April 7, 1955], No. 10320/55. Class 110 (3). In a gas-turbine power plant, the compressor A supplies air under pressure to combustion equipment- B, the gases from which drive the turbine C which drives the compressor and may also supply external shaft power through shaft L1; alternatively a separate power turbine may be utilized. The power shaft L1 drives through reduction gearing contained in casing D. The power plant is supported by struts G, G1 connected to brackets G3 secured to the duct between the compressor A and combustion equipment B, and a torque transmitting coupling is provided comprising a part rigid with the reduction gear, casing D and a part which is rigidly connected to the airframe of an aircraft in which the power plant is installed. The invention is applicable in the case of a power plant driving a helicopter rotor, in which case the engine is mounted in a vertical ...

Gas turbine engine

Gas turbine engine

Turbine engine

Flexible coupling

Improvements in or relating to means for supporting gas-turbine power-plants

... 606,526. Gas turbine plant; compressors. BRITISH AEROPLANE CO., Ltd., OWNER, F. M., and HENSTRIDGE, A. G. Oct. 9, 1945, No. 26373. [Classes 110 (i) and 110 (iii)] A gas turbine plant of the kind in which a compressor, turbine and combustion equipment form a unitary structure and certain of said parts are supported from each end of an axial flow compressor is supported by a structure surrounding the axial flow compressor which is attached to the compressor by means allowing relative radial and axial movement of the compressor and the turbine is supported by a cantilever structure secured to said structure surrounding the compressor. A gas turbine power plant is supported from a mounting ring 26, which may be attached to an aircraft, by a structure 19 which comprises a truncated conical member 21 to which a pair of end pieces 22, 23 are secured, e.g. by welding. The depth of the structure is made considerable to increase its rigidity and is reinforced by longitudinal channel shaped members ...

Mounting System and Mounting Method for Gas Turbine Aero Engine

Modular gas turbine plant with a heavy duty gas turbine

A transportable gas turbine module is described, comprising: a baseplate (25) supporting at least a gas turbine (27) and a load (29) drivingly connected to the gas turbine (27). The module further comprises a structure surrounding the gas turbine (27) and the load (29) and connected to the baseplate (25). The baseplate is designed such that it can support a heavy duty gas turbine having a rated power of not less than 80 MW.

Gas turbine with primary and secondary lubricating oil cooler

A gas turbine is described, comprising a turbine package (101) housing a compressor (105; 107), a high pressure turbine (109) and a power turbine (111). The gas turbine comprises a ventilation system for cooling the interior of the turbine package, as well as a lubricating oil circuit. The lubricating oil circuit comprises a lubricating oil pump (177), a lubricating oil tank (173), a primary lubricating oil cooler (189). A secondary lubricating oil cooler (163) is arranged in the turbine package (101), in a position lower than a rotary shaft (115) of the gas turbine. The ventilation system is arranged and designed such that at least part of a package cooling airflow contacts the secondary lubricating oil cooler (163) to remove heat from the lubricating oil circulating in the secondary lubricating oil cooler.

One piece inlet lip skin design

ONE PIECE INLET LIP SKIN DESIGN An aircraft engine inlet (18) incorporates a lip skin (24) having a skin web with a plurality of stiffeners (34) integrally extending from the skin web. An aft edge land (30) integrally extends from the skin web (25) at an outer rim and an inner edge land (32) integrally extends from the skin web (25) at an inner rim. A central land (36) integrally extends from the skin web with the plurality of stiffeners (34) extending between said aft edge land (30) and the central land (36). V/ \ll FIG. 6 FIG. 6 ...

ASSEMBLY FOR HOLDING THE INTERFACE OF STATIONARY OUTER STRUCTURE OF A NACELLE AND HOUSING OF A JET ENGINE

L'invention se rapporte à un ensemble de maintien de l'interface d'une structure externe fixe (15) d'une nacelle (3) et d'un carter (27) de turboréacteur (5), comprenant : un premier élément en relief appartenant à l'extrémité amont de la structure externe fixe (15); un deuxième élément en relief appartenant à l'extrémité aval du carter (27); lesdits premier et deuxième éléments en relief étant configurés pour être mis en contact l'un avec l'autre; deux demi-anneaux (109) formés par une paroi définissant un logement, configuré pour recevoir le premier et le deuxième éléments en relief lorsque le carter (27) et la structure fixe externe (15) sont montés bord à bord, et des moyens de butée configurés pour maintenir les premier et deuxième éléments en relief dans le logement.

VIBRATION ISOLATOR

GAS TURBINE MODULE, GAS TURBINE PLANT INCLUDING THE SAME, METHOD OF UNLOADING GAS TURBINE MODULE, AND METHOD OF EXCHANGING GAS TURBINE MODULE

ABSTRACT OF THE DISCLOSURE A gas turbine module includes a gas turbine that has a gas turbine rotor and a turbine shell; an inlet plenum that is connected to an inlet of the gas turbine; an exhaust plenum that is connected to an exhaust of the gas turbine; an enclosure that covers the gas turbine; and a common base on which the gas turbine, the inlet plenum, the exhaust plenum, and the enclosure are mounted. When moving the gas turbine, the gas turbine module is moved together. CA 3099538 2020-11-16 ...

FULL TURBOMACHINERY MODULE FOR LNG PLANTS OR THE LIKE

A modular gas turbine system (1) is disclosed. The system comprises a base plate (3) and a gas turbine engine (5) mounted on the base plate (3). The gas turbine engine (5) is drivingly coupled to a rotating load (23;25) mounted on the base plate (3). A supporting frame (31) extends above the base plate (3). A first bridge crane (35) and a second bridge crane (41) are movably supported on the supporting frame.

Catcher pin assembly

A catcher pin assembly for attachment to a catcher link of an aircraft frame, the catcher pin assembly having an engine mounting element, a catcher pin, a compressible element, and a nut, which is lockable to the catcher pin, the compressible element being compressed between the pin and the mounting element and/or between the mounting element and the nut, in the assembled condition of the assembly, such that the compressible element applies a predetermined resistance to rotation of the catcher pin relative to the mounting element. The invention also relates to a method of testing a catcher pin assembly to determine if the catcher pin is carrying load from the catcher link, the method including applying a predetermined torque to the catcher pin sufficient to overcome the resistance to rotation of the catcher pin when it is unloaded, and determining whether the catcher pin rotates relative to the mounting element.

Assembly for holding the interface of stationary outer structure of a nacelle and housing of a jet engine

The invention relates to an assembly for holding the interface of a stationary outer structure ( 15 ) of a nacelle ( 3 ) and housing ( 27 ) of a jet engine ( 5 ), said assembly including: a first raised element belonging to the upstream end of the stationary outer structure ( 15 ); a second raised element belonging to the downstream end of the housing ( 27 ), said first and second raised elements being formed so as to be placed in contact with each other; two half-rings ( 109 ) formed by a wall defining a recess that is formed so as to receive the first and second raised elements when the housing ( 27 ) and the stationary outer structure ( 15 ) are mounted edge to edge; and an abutment means formed so as to keep the first and second raised elements in the recess.

Centerline generator support system and method of elevating a centerline generator from a support surface

A generator support system for a centerline mounted generator includes a generator support member configured and disposed to support a generator upon a support surface. The generator support member includes at least one lifting element having a lifting surface that faces the support surface.

Combustor Liner and Flow Sleeve Tool

A tool includes an annular frame portion including a mount portion extending radially from the frame portion, a hook portion arranged on the mount portion, the hook portion sized and shaped to engage a member of a tubular component of a turbine combustor, and a force exertion portion arranged on the mount portion, the force exertion portion operative to engage a portion of the turbine combustor.

Strut, a gas turbine engine frame comprising the strut and a gas turbine engine comprising the frame

A strut is provided for being arranged between an annular inner structural casing and an annular outer structural casing in a gas turbine engine frame for carrying loads between the inner and outer structural casing during operation. The strut includes a first tube, which is configured to house a service line or pipe between the inner and outer structural casing. The strut includes a second tube, which is configured to house a fastening element for rigidly connecting the inner and outer structural casing. The first and second tube are arranged in a side-by-side relationship and rigidly attached to each other.

Nacelle assembly having integrated afterbody mount case

A nacelle assembly for a gas turbine engine includes an integrated afterbody mount case. The integrated afterbody mount case includes an outer ring and a plurality of spokes that extend radially inwardly from the outer ring. The outer ring includes a radially outer surface and a radially inner surface. The plurality of spokes are circumferentially disposed about the radially inner surface and extend radially inwardly from the radially inner surface.

Turbine engine comprising a contrarotating propeller receiver supported by a structural casing attached to the intermediate housing

The present invention relates to an open rotor type aircraft turbine engine ( 1 ), comprising a contrarotating propeller receiver ( 30 ) and a dual-body gas generator ( 14 ) comprising a low-pressure compressor ( 16 ) and a high-pressure compressor ( 18 ) separated by an intermediate housing ( 27 ), said gas generator being arranged upstream from said receiver. According to the invention, the turbine engine further comprises a structural casing ( 50 ) for supporting the receiver ( 30 ), surrounding the gas generator ( 14 ) and having a downstream end ( 50 a ) attached to said receiver and an upstream end ( 50 b ) attached to said intermediate housing ( 27 ). Furthermore, it comprises additional connection means ( 60 ) between said structural supporting casing and the gas generator, arranged between the upstream and downstream ends ( 50 b , 50 a ) of the casing.

Gas turbine engine comprising a tension stud

A gas turbine engine including a rotor is disclosed. The rotor includes a stud extending along an axis, rotating elements of a first section, and rotating elements of a second section. The stud includes a first and second external end, the first external end adapted to engage a first pre-load nut or a shaft and the second external end adapted to engage a second pre-load nut or a shaft such that the set of rotating elements are secured. Thus stud includes a first shank connected to the first external end and a second shank connected to the second external end. The first shank is located in the first section and has a first diameter. The second shank is located in the second section and has a second diameter which is greater than the first diameter.

LATERAL TURBOJET IMPROVED IN ORDER TO LIMIT THE DEFORMATION THEREOF

A turbojet includes an intermediate casing outer shroud connected to a front suspension and a primary structure connected to a rear suspension; the shroud and the primary structure are held in a coaxial relationship by arms with at least some of the arms being shaped and/or arranged to deform in response to thrust from the turbojet by creating a deforming torque between the shroud and the primary structure in opposition to opposing stress generated by the thrust. 19-. (canceled)10. A two-stream turbojet for attaching laterally to a fuselage of an airplane via two longitudinally spaced apart suspensions including a front suspension and a rear suspension , the turbojet comprising:an intermediate casing outer shroud attached to the front suspension; anda propulsion primary structure attached to the rear suspension;the intermediate casing outer shroud and the primary structure being held in a coaxial relationship by a set of arms, each arm including a right section that is hollow and being fastened by its ends to the shroud and to the primary structure; andwherein at least some of the arms are shaped and/or arranged so as to deform in response to thrust from the turbojet by creating a deforming torque between the shroud and the primary structure, the deforming torque having a direction opposite to stress that is generated under effect of the turbojet thrust by a lever arm between the thrust axis and the front suspension.11. A turbojet according to claim 10 , wherein the at least some of the arms are configured to provide coupling between shear and twisting so that a center of torsion of a right section of the arm is situated outside a midplane of the arm claim 10 , on a side opposite from the front suspension relative to the midplane.12. A turbojet according to claim 10 , wherein at least some of the arms include a slot extending from the shroud to the primary structure.13. A turbojet according to claim 12 , wherein the slot is closed by an elastomer seal.14. A turbojet ...

COMBUSTOR LINER SUPPORT AND SEAL ASSEMBLY

A method is disclosed herein for reducing binding between a combustor liner of a gas turbine engine and a free-standing ring disposed about the combustor liner. The method comprises the step of disposing a rolling assembly to roll between the combustor liner and the free-standing ring during relative radial displacement. 1. A method for reducing binding between a combustor liner of a gas turbine engine and a free-standing ring disposed about the combustor liner comprising the of:disposing a rolling assembly to roll between the combustor liner and the free-standing ring during relative radial displacement while substantially preventing relative circumferential movement.2. The method of further comprising the step of:limiting the extent of relative radial displacement between the combustor liner and the free-standing ring to a predetermined design amount.3. The method of further comprising the step of:disposing a seal of variable width between the combustor liner and the free-standing ring, the variable width varying in at least one of a radial direction and a circumferential direction relative to an axis along which the gas turbine engine extends.4. The method of wherein the rolling assembly rolls radially along a length of a surface defined by one of the combustor liner and the free-standing ring during the relative radial displacement.5. The method of wherein the rolling assembly substantially prevents relative circumferential movement between the combustor liner and the free-standing ring while permitting the relative radial displacement between the combustor liner and the free-standing ring in a radial direction.6. The method of wherein the rolling assembly is rotatably engaged with both the combustor liner and the free-standing ring.7. The method of wherein the rolling assembly includes a plurality of pins engaged between the combustor liner and the free-standing ring.8. The method of wherein the plurality of pins are each received in one of a plurality of slots ...

SUBSTITUTION DEVICE FOR AIRCRAFT ENGINE

A substitution device for replacing a turbojet type of aircraft engine, with a nacelle comprising protective covers, and a support mast having two engine mounts, each of said protective covers being mounted to pivot between an open position and a closed position, said substitution device comprising a body, two anchors attaching to the engine mounts of the engine support mast, and bearing portions adapted to receive support portions belonging to the protective covers. A method for installing such a substitution device in a nacelle, which makes it possible to close the engine covers and authorize the movement of the aircraft. 2. The substitution device according to claim 1 , wherein the bearing portions are arranged substantially in an arc.5. The substitution device according to claim 1 , wherein the bearing portions are at least partially retractable.6. The substitution device according to claim 1 , further comprising a hydraulic unit and cylinders adapted for connecting the substitution device to at least the pair of covers.7. The substitution device according to claim 1 , further comprising wheels and a drawbar.8. The substitution device according to claim 1 , arranged in such a way that its center of gravity is located at or near the center of gravity of the engine which it replaces in the propulsion system.9. The substitution device according to claim 1 , further comprising lifting yokes. The present invention relates to substitution devices for an aircraft engine, in particular for turbojet engines.These substitution devices can be considered as support tools to assist with aircraft construction and maintenance. Under normal circumstances, such aircraft engines are usually attached to a supporting mast and protected by a nacelle which includes cowls (protective covers) surrounding the engine.According to known prior art, one (or several) engines must sometimes be removed for maintenance on said engine or for engine replacement. However, in order to remove an ...

Rigid raft

The present invention provides a rigid raft formed of rigid composite material. The raft has an electrical system and/or a fluid system embedded therein. The raft further has a tank for containing liquid integrally formed therewith. The tank can be formed of the rigid composite material. The tank can be for a gas turbine engine.

ELECTRONIC UNIT MOUNTING

An electrical assembly comprising an electrical raft and an electronic unit is provided to a gas turbine engine . The electrical raft has electrical conductors embedded in a rigid material , which may be a rigid composite material. The electrical conductors are in electrical contact with the electronic unit . When the electronic unit is installed, at least a part of it forms a part of a gas-washed surface of the engine . The electronic unit is then easily accessible from the engine , and potentially complex and/or heavy access doors/panels may not be required. 1. A gas turbine engine having an electrical assembly that comprises:an electrical raft having a rigid material with at least one electrical conductor embedded therein; andan electronic unit mounted on the electrical raft and in electrical connection with the electrical raft, wherein:the electrical raft is mechanically fixed to another component of the gas turbine engine; andthe electronic unit forms at least a part of a gas-washed surface of the gas turbine engine.2. A gas turbine engine according to claim 1 , wherein:the electrical raft is provided with a first electrical connector in electrical contact with at least one of said embedded electrical conductors, the first electrical connector being fixed relative to the electrical raft;the electronic unit is provided with a second electrical connector, that is complimentary to the first electrical connector, the second electrical connector being fixed relative to the electronic unit; andthe electrical raft and the electronic unit are in electrical connection through the first and second electrical connectors.3. A gas turbine engine according to claim 1 , wherein the electronic unit is mechanically fixed to another component of the gas turbine engine.4. A gas turbine engine according to claim 3 , wherein the electronic unit and the electrical raft are mechanically fixed to different components of the gas turbine engine.5. A gas turbine engine according to claim ...

Gas turbine engine part

The present invention provides a gas turbine engine part which has a primary purpose in the engine which is structural and/or aerodynamic. The part is formed of rigid composite material, and has an electrical system comprising electrical conductors permanently embedded in the composite material. This provides advantages in terms of weight, complexity, and build time.

ANTI-VIBRATION MOUNT

An anti-vibration mount is provided for mounting a first component to a second component. The mount has an elastomeric body which provides a recess into which the first component is received. The mount further has pair of brackets which fit to opposing sides of the elastomeric body sandwiching the first component received in the recess therebetween. At least one of the brackets is arranged to connect the anti-vibration mount and second component together. The mount further has a clamping arrangement which applies clamping pressure across the brackets and thereby compresses the elastomeric body to secure the first component in the recess. 1. An anti-vibration mount for mounting a first component to a second component , the mount having:an elastomeric body which provides a recess into which the first component is received;a pair of brackets which fit to opposing sides of the elastomeric body sandwiching the first component received in the recess therebetween, at least one of the brackets being arranged to connect the anti-vibration mount and second component together; anda clamping arrangement which applies clamping pressure across the brackets and thereby compresses the elastomeric body to secure the first component in the recess.2. An anti-vibration mount according to claim 1 , wherein the elastomeric body is in two parts which are separable from each other when the clamping pressure is removed claim 1 , the first part providing one of the opposing sides of the elastomeric body and one side of the recess and the second part providing the other opposing side of the elastomeric body and an opposing side of the recess.3. An anti-vibration mount according to claim 2 , wherein the first component has a through-hole and the first part has a projection which extends through the through-hole and is received in a matching cavity formed in the second part.4. An anti-vibration mount according to claim 1 , wherein the first component is planar and the recess is a slot.5. An ...

GAS TURBINE ENGINE SYSTEMS

A gas turbine engine comprises at least one rigid raft assembly that has a fluid passageway at least partially embedded therein. The fluid passageway is at least a part of a fluid system. In addition to the fluid passageway the rigid raft assembly also has at least a part of another system. For example, the rigid raft assembly may also include electrical conductors which are part of an electrical system. The rigid raft assembly may be lighter, easier to assemble, more robust and more compact than conventional solutions for providing systems to gas turbine engines. 1. A rigid raft assembly for a gas turbine engine , the rigid raft assembly comprising a rigid material that carries at least a part of a first gas turbine engine system and at least a part of a second gas turbine engine system , wherein:the first gas turbine engine system is a fluid system that comprises at least one fluid passage that is at least partially embedded in the rigid raft assembly.2. A rigid raft assembly according to claim 1 , wherein:the second gas turbine engine system is an electrical system that comprises electrical conductors at least partially embedded in the rigid material.3. A rigid raft assembly according to claim 1 , wherein:the fluid passage has an axial direction (p) along which, in use, fluid flows; andthe rigid material surrounds the fluid passage over at least one axial portion of the passage.4. A rigid raft assembly according to claim 1 , wherein the fluid passage is formed by a fluid pipe that is at least partially embedded in the rigid raft assembly.5. A rigid raft assembly according to claim 1 , wherein the fluid passage is formed by the rigid material.6. A rigid raft assembly according to claim 4 , wherein:the rigid raft assembly comprises two rigid rafts formed by the rigid material; andthe fluid pipe is embedded between the two rigid rafts.7. A rigid raft assembly according to claim 6 , wherein:the rigid rafts are thin elements having an upper major surface separated by ...

HINGING CRADLE FOR FAN COWLS SUPPORTED BY SAID COWLS IN CLOSED POSITION

The invention relates to an engine assembly for aircraft in which the coupling device comprises a fore aerodynamic structure having a cradle equipped with an aerodynamic cowling, the cradle being hinge mounted on the air intake of the engine, and the fan cowls being mounted to move on the cradle so as to be able to occupy an open position as well as a closed position in which they are supported by the air intake and by a thrust reverser. According to the invention, when the cowls are in open position, the cradle adopts a first configuration in which its aft end rests on a span of the engine mounting structure, and when the cowls are in closed position, the cradle, borne by said cowls, adopts a second configuration in which it is lacking any direct mechanical link with the other elements of the engine mounting structure. 1. Engine assembly for aircraft comprising an engine , a coupling device of the engine as well as a nacelle surrounding the engine and provided with fan cowls as well as an air intake , said coupling device comprising a rigid structure as well as a fore aerodynamic structure , the latter having a cradle equipped with an aerodynamic cowling , said cradle being hinge mounted at its fore end on an entity comprising a fan housing of said engine as well as the air intake , and said fan cowls being mounted to move on said cradle so as to be able to occupy an open position and a closed position in which they are supported at the fore by said entity and at the aft by a thrust reverser system ,characterised in that the assembly is designed such that when the fan cowls are in open position, said cradle adopts a first configuration in which its aft end is retained by a span of the engine mounting structure, and such that when the fan cowls are in closed position, said cradle, borne by the fan cowls, adopts a second configuration in which it is lacking any direct mechanical link with the other elements of the engine mounting structure.2. Engine assembly ...

HYPERSTATIC TRUSS COMPRISING CONNECTING RODS

A hyperstatic truss including connecting rods, used for suspension of a first ring, forming part of an engine case, inside a second ring concentric to the first ring, the connecting rods being secured at one end to the first ring and at the other end to the second ring. The tensile stiffness of the connecting rods is greater than the compressive stiffness thereof. The truss for example can be used for suspension of a ducted-fan turbine engine with an elongate bypass duct. 110-. (canceled)11. A hyperstatic truss comprising:connecting rods applied to suspending a first ring, that forms part of an engine casing, inside a second ring concentric with the first ring, the connecting rods being secured by one end to the first ring and by an other end to the second ring,wherein the connecting rods have a tensile strength that is higher than their compressive strengths, and comprise one or more parts configured to transmit compressive and tensile loads between the ends and one or more other parts configured to transmit only the tensile loads between the ends.12. The truss as claimed in claim 11 , wherein the connecting rods comprise an internal rod and a shroud surrounding the internal rod claim 11 , the internal rod configured to work in compression and in tension claim 11 , the shroud working only in tension.13. The truss as claimed in claim 12 , wherein the internal rod and the shroud of the connecting rods include a pair of surfaces bearing against one another when tension is applied to the internal rod claim 12 , the tensile loads then being transmitted between the two ends by the internal rod and the shroud.14. The truss as claimed in claim 13 , wherein the connecting rods comprise at least two pairs of bearing surfaces distributed along the axis of the rod.15. The truss as claimed in claim 12 , wherein the connecting rods include means for bearing radially against an internal surface of the shroud claim 12 , to prevent the internal rod from buckling when subjected to ...

Mounting device and method of assembling the same

A method of assembling a mounting device for an electrical harness of a gas turbine engine is provided. The electrical harness has a wire bundle. The method includes providing a first shell and providing a second shell. The method further includes coupling the first shell to the second shell with the wire bundle disposed between the first shell and the second shell such that movement of the first shell and the second shell along the wire bundle is restricted.

Method for cooling a thermal protection floor of an aft aerodynamic fairing of a structure for mounting an aircraft propulsion system

A propulsion system for an aircraft, including a dual-flow turbojet and a mounting structure for mounting this turbojet on the wing surface or on the fuselage of an aircraft. The mounting structure includes an aft aerodynamic fairing including a thermal protection floor to protect the mounting structure from the heat of a primary airstream channelled by an exhaust nozzle of the turbojet, as well as an air inlet provided in a longitudinal aerodynamic wall washed by a secondary airstream of the turbojet and delimiting together with an other similar wall a cavity isolated from the secondary airstream for extracting a cooling airstream from the secondary airstream, and air circulation means fed by the air inlet and having at least one outlet aperture emerging in a space between the thermal protection floor and the exhaust nozzle.

TURBOMACHINE COMPONENT ALIGNMENT

A rotating machine has a barrel casing and at least one component carrier disposed therein. The component carrier is aligned and supported within the barrel casing using a plurality of carrier alignment fixtures which have alignment shafts that extend through the barrel casing to the component carrier. Each alignment shaft has an eccentric key pin extending from a distal end thereof and an alignment key rotatably-mounted to each eccentric key pin. The alignment keys mate with corresponding keyway slots defined on the component carrier such that when the carrier alignment fixtures are rotated axially, the corresponding alignment keys bias against the keyway slot and shift the position of the component carrier. 1. A rotating machine , comprising:a barrel casing having a plurality of cylindrical bores extending between outer and inner circumferential surfaces of the barrel casing;a component carrier disposed within the barrel casing and defining a plurality of keyway slots;a plurality of carrier alignment fixtures arranged about the outer circumferential surface of the barrel casing, each carrier alignment fixture having an alignment shaft extending through a corresponding one of the plurality of cylindrical bores and having an eccentric key pin extending from a distal end of each alignment shaft; andan alignment key rotatably-mounted to each eccentric key pin of each corresponding carrier alignment fixture and configured to mate with a corresponding one of the plurality of keyway slots, whereby rotation of each carrier alignment fixture about a central axis of each corresponding alignment shaft adjusts a position of the component carrier with respect to the barrel casing.2. The rotating machine of claim 1 , wherein the barrel casing is unsplit and cylindrical.3. The rotating machine of claim 1 , further comprising a fluid expander component mounted within the component carrier.4. The rotating machine of claim 1 , further comprising a shaft arranged for rotation within ...

Fastening element and de-icing device of an aircraft gas-turbine engine

The present invention relates to a fastening element, in particular to its use in a de-icing device of an aircraft gas-turbine engine, for connecting two components, with the fastening element ensuring a connection of the components with a predetermined relative movability to each other, with the fastening element including two struts arranged at an angle to each other, where two first end areas, spacedly arranged to each other, can be fastened to one of the components, and the two other second end areas can be connected to each other and fastened to the other component. 1. Fastening element for connecting two components , with the fastening element ensuring a connection of the components with a predetermined relative movability to each other , with the fastening element including two struts arranged at an angle to each other , where two first end areas , spacedly arranged to each other , can be fastened to one of the components , and the two other second end areas can be connected to each other and fastened to the other component.2. Fastening element in accordance with claim 1 , characterized in that the struts are elastically bendable and/or are provided in the form of flat sheet-metal strips.3. Fastening element in accordance with claim 1 , characterized in that the second end areas can connect the other component by means of a carrier element claim 1 , attached to the second end areas and to the other component.4. De-icing device of an aircraft gas-turbine engine with an engine cowling enclosing at least one inflow region claim 1 , with the engine cowling having a double-walled design and including at least one annular tube element extending in the circumferential direction and being provided with outlet openings for passing hot air to an inflow region claim 1 , in order to de-ice it claim 1 , with the tube element in the circumferential direction being mounted on the engine cowling by several fastening elements in accordance .5. De-icing device in accordance ...

ARRANGEMENT OF COMPONENTS OF A WIND POWER PLANT

An arrangement of components () for a wind power plant. A first component () has a flange () having a flange contact surface (). A second component () has a flange-mounting surface () for a flange () of the first component (). Alternatively, the second component () has a flange () having a flange contact surface () and the flange contact surface () of the first component () and the flange-mounting surface () of the second component () are arranged opposite each other, or the flange contact surfaces () of the components are arranged opposite each other. At least one flange contact surface () of a flange () and/or the flange-mounting surface () of the second component () has an outer coating made of a chrome-steel alloy. 1121330. An arrangement of two components ( , , ) of a wind power plant , comprising:{'b': 12', '13', '15', '130', '121', '122', '133, 'a first component (, ) that has a flange (, ) having a flange contact surface (, , ), and'}{'b': 30', '131', '132', '15', '130', '12', '13, 'a second component () that has a flange-mounting surface (, ) for the flange (, ) of the first component (, ),'}{'b': 122', '133', '12', '13', '131', '132', '30, 'wherein the flange contact surface (, ) of the first component (, ) and the flange-mounting surface (, ) of the second component () are arranged opposite from each other, and'}{'b': 121', '122', '133', '15', '130', '131', '132', '30, 'wherein at least one of the flange contact surface (, , ) of the flange (, ) and the flange-mounting surface (, ) of the second component () has an outer coating made of a chrome-steel alloy.'}2. The arrangement according to claim 1 ,wherein the coating of the chrome steel alloy is applied using an arc welding method.3121330. The arrangement according to claim 1 , wherein the first component ( claim 1 , ) and the second component () are braced together using connection.4. The arrangement according to claim 1 ,wherein the chrome-steel alloy coating on material GJS 400, as a cast material ...

TURBINE SHROUD HANGER WITH DEBRIS FILTER

A turbine shroud hanger apparatus for a gas turbine engine includes: an arcuate shroud hanger having at least one cooling hole passing therethrough, the cooling hole having an inlet and an outlet; and a filter carried by the shroud hanger positioned upstream of the inlet of the cooling hole, the filter having a plurality of openings formed therethrough which are sized to permit air flow through the cooling hole while preventing the entry of debris particles larger than a preselected size into the cooling hole. 1. A turbine shroud hanger apparatus for a gas turbine engine , comprising:an arcuate shroud hanger having at least one cooling hole passing therethrough, the cooling hole having an inlet and an outlet; anda filter carried by the shroud hanger positioned upstream of the inlet of the cooling hole, the filter having a plurality of openings formed therethrough which are sized to permit air flow through the cooling hole while preventing the entry of debris particles larger than a preselected size into the cooling hole.2. The apparatus of wherein the shroud hanger comprises:an arcuate body;forward and aft flanges extending from a radially outer surface of the body; andforward and aft hooks extending from a radially inner surface of the body.3. The apparatus of wherein:an arcuate groove is formed in a forward face of the shroud hanger, the groove communicating with the cooling hole; andthe filter is received in the groove.4. The apparatus of wherein a ledge is disposed around the perimeter of the groove claim 3 , and the filter is mounted against the ledge.5. The apparatus of wherein the filter is secured to the shroud hanger by welding claim 1 , brazing claim 1 , or a combination thereof.6. The apparatus of wherein the filter has a convex cross-sectional shape.7. A turbine shroud apparatus for a gas turbine engine claim 1 , comprising:an arcuate shroud hanger having at least one cooling hole passing therethrough, the cooling hole having an inlet and an outlet;a ...

TUNGSTEN CARBIDE INSERTS AND METHOD

Systems and methods provide for wear reduction in a combustion system of a gas turbine. A system for wear reduction includes: at least one substantially H-shaped block, the substantially H-shaped block being configured to secure a transition piece of a gas turbine combustor to a support piece; a first insert including a tungsten carbide in a metal matrix, the metal matrix being selected from a group including cobalt and nickel; and a brazing material which is used in brazing the first insert to the at least one substantially H-shaped block in at least one location on an interior wear surface of the at least one substantially H-shaped block. 1. A system for wear reduction in a combustion system of a gas turbine , said system comprising:at least one substantially H-shaped block configured to secure a transition piece of a gas turbine combustor to a support piece;a first insert including a first tungsten carbide in a first metal matrix, wherein said first metal matrix is selected from a group comprising cobalt and nickel; anda first brazing material between said at least one substantially H-shaped block and said first insert, said first brazing material being configured to braze said first insert to said at least one substantially H-shaped block in at least one location on an interior wear surface of said at least one substantially H-shaped block,wherein said interior surface of said at least one substantially H-shaped block includes a first surface substantially perpendicular to a second surface which is substantially perpendicular to a third surface, said third surface being substantially parallel to and having a substantially same surface area as said first surface.2. The system of claim I , further comprising:at least one combustor liner stop configured to attach to a combustor liner of the gas turbine;a second insert including a second tungsten carbide in a second metal matrix, wherein said second metal matrix is selected from a group comprising cobalt and nickel; ...

Turbine engine case mount and dismount

A method for mounting a gas turbine engine having a compressor section, a combustor section, a turbine section, a pylon and a rear mount bracket, includes positioning the mounting bracket between the gas turbine engine and the pylon. The mounting bracket is connected to the turbine case reacting a least a vertical load, a side load, a thrust load, and a torque load from the gas turbine engine through the mounting bracket. The mounting bracket is attached to the pylon reacting the same loads from the gas turbine engine.

Turbojet engine nacelle reinforcing structure

A turbojet engine nacelle includes a fixed structure, which has a fan casing of the turbojet engine and a front frame mounted downstream of the fan casing and directly or indirectly supporting cascade vanes. The front frame is able to collaborate with a thrust reverser cowling sliding between a closed position covering the flow-diverting means and an open position exposing this flow-diverting means. At least one reinforcing structure of the engine nacelle transmits load between the fan casing and the front frame. The reinforcing structure extends along the longitudinal axis of the nacelle and supports a third line of defense and/or an inhibiting device between the front frame and the thrust reverser cowling.

Transition Duct Mounting System

The present invention is directed to a mounting system for a gas turbine transition duct. The mounting system includes a generally C-shaped mounting bracket having a generally radially extending first portion, an arc-shaped second portion, and a generally radially extending third portion. Each of the first portion and third portion includes a spherical bearing and the second portion has a plurality of sets of mounting holes. In another embodiment of the present invention, a gas turbine transition duct is provided having a panel assembly, an aft frame secured to the panel assembly and a mounting system for securing the transition duct to a turbine inlet. The panel assembly includes two formed sheets of metal secured together along axial seams. The panel assembly is secured to an aft frame that is capable of expanding in the circumferential direction due to thermal growth. The mounting system comprises a generally C-shaped mounting bracket having a generally radially extending first portion, an arc-shaped second portion, and a generally radially extending third portion. Each of the first portion and third portion includes a spherical bearing and the second portion has a plurality of sets of mounting holes. 1. A mounting system for a gas turbine transition duct comprising:a generally C-shaped mounting bracket having a generally radially extending first portion, an arc-shaped second portion, and a generally radially extending third portion;a spherical bearing located within both the first portion and third portion; anda plurality of sets of mounting holes located along the second portion of the mounting bracket.2. The mounting system of claim 1 , wherein the spherical bearings permit pivoting of the mounting bracket relative to the transition duct.3. The mounting system of claim 1 , wherein the plurality of sets of mounting holes comprises a first set and second set claim 1 , with the mounting holes of the first set spaced a first distance apart and the mounting holes ...

Gas turbine engine geared architecture axial retention arrangement

A support assembly for a geared architecture includes an engine static structure. A flex support is secured to the engine static structure and includes a bellow. A support structure is operatively secured to the flex support. A geared architecture is mounted to the support structure. First members are removably secured to one of the engine static structure and the flex support and second members are removably secured to the support structure. The first and second members are circumferentially aligned with one another and spaced apart from one another during a normal operating condition. The first and second members are configured to be engageable with one another during an extreme event to limit axial movement of the geared architecture relative to the engine static structure.

Flow path structure and gas turbine exhaust diffuser

A flow path structure includes a wall surface in which a flow path is formed, a structure configured to extend in a direction intersecting a main stream direction of a fluid flowing through the flow path from the wall surface, and a concave section forming region formed throughout a range including the structure in the main stream direction and having a concave section formed in the wall surface. As the structure occupies a partial range of the flow path in the flow path cross section intersecting the main stream, a cross-sectional area of the flow path is varied in accordance with positional variation in the main stream direction.

Geared Architecture for High Speed and Small Volume Fan Drive Turbine

A gas turbine engine includes a flex mount for a fan drive gear system. A very high speed fan drive turbine drives the fan drive gear system.

Adjustable engine mount

A forward engine mount assembly for a gas turbine engine includes a mount beam having a main body with a fore end and an aft end and a forward shackle assembly supported by the fore end of the mount beam. The forward shackle assembly comprises a first link configured to be connected to a first engine case structure and a second link configured to be connected to a second engine case structure. The first and second links are pivotally attached to each other.

Nacelle bifurcation for gas turbine engine

A nacelle structure for a gas turbine engine includes a core engine nacelle disposed about an engine axis and an outer nacelle disposed about the core engine nacelle. A bifurcation extends between the outer nacelle and the core engine nacelle along a bifurcation axis extending between the outer nacelle and the core engine nacelle. The bifurcation includes at least one mounting surface that is disposed at a non-normal angle relative to the bifurcation axis.

DEVICE FOR SUSPENDING A TURBOJET ENGINE

A turbojet is suspended by attachments including hinged links. An attachment includes a support including three branches with passages through which a pin passes, the pin being oriented generally parallel to a direction that is tangential to a casing and being hinged to a central branch of the support by a ball joint. 16-. (canceled)7. A device for suspending an airplane turbojet , wherein the turbojet is connected to a pylon fastened to a structure of the airplane by hinged link attachments , each link attachment connecting the pylon to a casing of the turbojet ,wherein at least one of the attachments comprises a pin support for a pin fastened to the pylon, and including three spaced-apart branches including passages through which the pin passes,the pin support being fastened to the pylon such that the pin is oriented generally parallel to a direction that is tangential to the casing,wherein the pin is hinge-mounted by a ball joint to a central branch of the pin support, wherein the pin passes through the other two branches with clearance, and wherein two of the links are arranged symmetrically about a midplane of the pin support and are hinged-connected to the casing and to the pin.8. A device according to claim 7 , wherein each of the two links is connected directly to the pin by a ball joint.9. A device according to claim 8 , wherein each ball joint providing a hinge connection of the respective links to the pin is situated in a vicinity of a corresponding end of the pin and on a side of one of the two other branches that is remote from the central branch of the pin support.10. A device according to claim 7 , wherein the pin support is defined using three substantially identical parts arranged side by side claim 7 , each part defining one of the branches.11. A device according to claim 7 , wherein two passages in the two branches situated on either side of the central branch are bores claim 7 , and bushings of selected thickness are mounted in the bores to ...

DRAINING DEVICE

A draining device for draining a fluid from an aircraft engine support strut includes a suction arrangement having an inlet and extending along a longitudinal main axis to form a plenum chamber and a drain tube fluidly connected to the aircraft engine support strut and extending along the longitudinal main axis, the drain tube ending by an outlet, the outlet of the drain tube terminating within the plenum chamber of the suction arrangement. An air flow entering the inlet of the suction arrangement causes a low-pressure region to be created within the plenum chamber, substantially downstream the outlet of the drain tube. 1. A draining device for draining a fluid from an aircraft engine support strut , comprising:a suction arrangement comprising an inlet and extending along a longitudinal main axis to form a plenum chamber; anda drain tube fluidly connected to the aircraft engine support strut and extending along the longitudinal main axis, the drain tube ending by an outlet, the outlet of the drain tube terminating within the plenum chamber of the suction arrangement; endwherein an air flow entering the inlet of the suction arrangement causes a low-pressure region to be created within the plenum chamber, substantially downstream the outlet of the drain tube.2. The draining device of claim 1 , wherein the suction arrangement further comprises a converging-diverging part claim 1 , the converging-diverging part forming a constricted section of tubing causing the air flow entering the inlet of the suction arrangement to be accelerated within the plenum chamber by a Venturi effect.3. The draining device of claim 1 , wherein the suction arrangement further comprises a swirling arrangement positioned upstream of the outlet of the drain tube causing the air flow entering the inlet of the suction arrangement to swirl within the plenum chamber.4. The draining device according to claim 2 , wherein the converging/diverging part forming the constricted section of tubing is ...

Four bar drive mechanism for bleed system

An actuation system for a bleed valve includes first and second bell cranks connected by a connecting link. The first bell crank has a first arm that is coupled to a bleed valve and a second arm that is coupled to the connecting link. The connecting link has a first end coupled to the second arm of the first bell crank and a second end coupled to a first arm of the second bell crank. A second arm of the second bell crank is coupled to an actuating element. Input is communicated through the actuating element to move the bleed valve between open and closed positions via the first and second bell cranks.

MID-TURBINE FRAME WITH OIL SYSTEM MOUNTS

A gas turbine engine includes a combustor, a first turbine section in fluid communication with the combustor, a second turbine section in fluid communication with the first turbine section, and a mid-turbine frame located axially between the first turbine section and the second turbine section. An oil system includes a first oil system component that houses oil. A first connector mechanically connects the first oil system component to the mid-turbine frame. 1. A gas turbine engine comprising:a combustor;a first turbine section in fluid communication with the combustor;a second turbine section in fluid communication with the first turbine section;a mid-turbine frame located axially between the first turbine section and the second turbine section;an oil system comprising a first oil system component that houses oil; anda first connector mechanically connecting the first oil system component to the mid-turbine frame.2. The gas turbine engine of claim 1 , wherein the first oil system component comprises one of an accessory gearbox and an oil reservoir.3. The gas turbine engine of claim 2 , and further comprising:a second connector connecting the first oil system component to a diffuser case; anda third connector mechanically connecting the first oil system component to the diffuser case.4. The gas turbine engine of claim 3 , wherein the second connector comprises a first bracket on the oil system component claim 3 , a second bracket on the diffuser case claim 3 , and a link connecting the first bracket to the second bracket.5. The gas turbine engine of claim 1 , wherein the first connector comprises a first bracket on the oil system component claim 1 , a second bracket on the mid turbine frame claim 1 , and a link connecting the first bracket to the second bracket.6. The gas turbine engine of claim 1 , wherein the first connector comprises a bracket mounted on a boss of the outer case of the mid turbine frame.7. The gas turbine engine of claim 6 , and further comprising ...

MID-TURBINE FRAME WITH THREADED SPOKES

A mid-turbine frame located in a gas turbine engine axially aft of a high-pressure turbine and fore of a low-pressure turbine includes an outer frame case, an inner frame case, and at least a first spoke connecting the outer frame case to the inner frame case. The first spoke includes a tie rod having a first threaded surface and a connector having a second threaded surface. The first and second threaded surfaces overlap partially but not completely. 1. A gas turbine engine comprising:a combustor;a first turbine section in fluid communication with the combustor;a second turbine section in fluid communication with the first turbine section; and an outer frame case;', 'an inner frame case; and', a tie rod having a first threaded surface; and', 'a connector having a second threaded surface, wherein the first and second threaded surfaces overlap partially but not completely., 'at least a first spoke connecting the outer frame case to the inner frame case, the first spoke comprising], 'a mid-turbine frame located axially between the first turbine section and the second turbine section, the mid-turbine frame comprising2. The gas turbine engine of claim 1 , wherein the connector is a retaining nut.3. The gas turbine engine of claim 2 , wherein the retaining nut comprises a flange extending outward from an outer surface of the retaining nut.4. The gas turbine engine of claim 3 , wherein the retaining nut is positioned in a hole of the outer frame case claim 3 , with the flange positioned radially outward of the outer frame case and the second threaded surface positioned radially inward of the outer frame case.5. The gas turbine engine of claim 3 , wherein a bolt extends through the flange into the outer frame case.6. The gas turbine engine of claim 2 , wherein the first threaded surface is on an outer surface of the tie rod claim 2 , wherein the second threaded surface is on an inner surface of the retaining nut.7. The gas turbine engine of claim 1 , wherein the tie rod ...

Transport securing device of an aircraft gas turbine

The present invention relates to a transport securing device of an aircraft gas turbine with a securing element, which at its two end areas is each provided with a first and a second connecting element, with the first connecting element being designed to be suitable for fastening to a front suspension device and the second connecting element being designed to be suitable for fastening to a rear suspension device of an aircraft gas turbine. 1. Transport securing device of an aircraft gas turbine with a securing element , which at its two end areas is each provided with a first and a second connecting element , with the first connecting element being designed to be suitable for fastening to a front suspension device and the second connecting element being designed to be suitable for fastening to a rear suspension device of an aircraft gas turbine.2. Transport securing device in accordance with claim 1 , characterized in that the first and/or the second connecting element is designed plate-like and provided with recesses for passing bolts through them.3. Transport securing device in accordance with claim 1 , characterized in that a handling attachment is arranged in the centre of gravity area.4. Transport securing device in accordance with claim 1 , characterized in that the securing element is designed in the form of a bar-like beam.5. Transport securing device in accordance with claim 1 , characterized in that the first and the second connecting element are provided with an intermediate element made of rubber or plastic. This invention relates to a transport securing device of an aircraft gas turbine.It is known from the state of the art to suspend aircraft gas turbines, with the engines being arranged underneath the wings of an aircraft, from a pylon by means of a front suspension device and a rear suspension device. To do so, the core engine is fastened at two areas to the aircraft wing such that the front suspension device absorbs the thrust forces of the engine ...

CERAMIC-TO-METAL TURBINE VOLUTE ATTACHMENT FOR A GAS TURBINE ENGINE

A means of attachment applicable to mating parts which have substantially different coefficients of thermal expansion is disclosed. The means of attachment substantially reduces the friction between the mating surfaces while still keeping the mating parts centered with respect to one another. The approach is based on radial recessed faces wherein the radial faces slide relative to each other. There may be three or more recessed/raised faces on each mating component, which when mated, maintain the alignment between the mating parts while allowing differential growth of the mating parts. This approach also the provides a much larger bearing surface for the attachment than a radial pin/slot approach, for example, and substantially eliminates areas of high stress concentration. It is thus a more robust design for components that undergo many thousands of thermal cycles. 1. A gas turbine engine , comprising:at least one turbo-compressor spool assembly, wherein the at least one turbo-compressor spool assembly comprises a compressor in mechanical communication with a turbine, a ceramic volute directing an inlet gas towards an inlet of a ceramic rotor of the turbine and a ceramic shroud adjacent to the rotor of the turbine, the ceramic shroud directing an outlet gas towards an outlet of the at least one turbo-compressor spool assembly; anda metallic housing comprising a metallic base plate having a metallic surface to engage a ceramic surface of at least one of the ceramic shroud and volute;wherein each of the engaged metallic and ceramic surfaces comprises at least one raised face and at least one recessed face, wherein, when the ceramic and metallic faces are engaged, the at least one raised face of the metallic surface opposes the at least one recessed face of the ceramic surface and the at least one recessed face of the metallic surface opposes the at least one raised face of the ceramic surface.2. The engine of claim 1 , wherein the ceramic volute interfaces with the ...

ASSEMBLY FOR A JET ENGINE OF AN AIRCRAFT

The present invention proposes a structural unit for an aircraft engine having at least one fuel pump of a fuel circuit and at least one hydraulic fluid pump of a hydraulic fluid circuit, where the structural unit can be coupled to an accessory gearbox shaft of an accessory gearbox of the engine. 1. Structural unit for an aircraft engine having at least one fuel pump of a fuel circuit and at least one hydraulic fluid pump of a hydraulic fluid circuit , where the structural unit can be coupled to an accessory gearbox shaft of an accessory gearbox of the engine.2. Structural unit in accordance with claim 1 , characterized in that the at least one fuel pump and the at least one hydraulic fluid pump are arranged inside a common casing.3. Structural unit in accordance with claim 1 , characterized in that a low-pressure fuel pump and a high-pressure fuel pump are provided which are arranged on a common shaft of the structural unit.4. Structural unit in accordance with claim 1 , characterized in that a hydraulic fluid delivery pump and at least one hydraulic fluid return pump are provided which are arranged on a common shaft of the structural unit.5. Structural unit in accordance with claim 1 , characterized in that a hydraulic fluid delivery pump and at least one hydraulic fluid return pump are provided which are arranged on separate shafts of the structural unit which in particular run parallel to one another.6. Structural unit in accordance with claim 3 , characterized in that the shafts of the structural unit are coupled to one another by a gearbox of the structural unit in particular with at least one gear stage.7. Structural unit in accordance with claim 6 , characterized in that the gearbox has a shaft area that can be coupled to an accessory gearbox shaft and is designed in particular with external teeth.8. Structural unit in accordance with claim 1 , characterized in that a heat exchanger forms part of the structural unit and is in particular integrated into the ...

THRUST REVERSER-INTEGRATED TRACK BEAM AND INNER FIXED STRUCTURE

A panel for a nacelle structure that surrounds a jet engine core includes a longitudinally extending, generally semicircular center region, an upper bifurcation region, a lower bifurcation region, a hinge beam region extending from the upper bifurcation region and configured to receive an upper thrust reverser track guide, and a latch beam region extending from the lower bifurcation region and configured to receive a lower thrust reverser track guide. A bypass duct is formed in a space between the panel and the nacelle structure. A hinge beam structure and latch beam structure are integrally formed with the panel. An inner skin layer extends continuously across a bond panel surface generally facing the engine core and an outer skin layer extends continuously across a bond panel surface generally facing the bypass duct, the inner and outer skin layers extending across the center region, upper and lower bifurcation regions, hinge beam region, and latch beam region. The outer skin layer may have an acoustic treatment, such as being perforated in at least the center region and the upper and lower bifurcation regions. In this way, a separate hinge beam and latch beam are not used, and therefore there is no need to couple the panel to the respective beams. 1. A panel for a nacelle structure that surrounds a jet engine core , the panel comprising:a center region that extends along a longitudinal axis of the jet engine core;an upper bifurcation region extending radially from the center region;a lower bifurcation region extending radially from the center region;a hinge beam region extending from the upper bifurcation region and configured to receive an upper thrust reverser track guide;a latch beam region extending from the lower bifurcation region and configured to receive a lower thrust reverser track guide;wherein a bypass duct is formed in a space between the panel and the nacelle structure;wherein the hinge beam region includes a lower flow portion that defines a ...

ASSEMBLY COMPRISING A PROTECTION DEVICE AND A TURBINE MACHINE ELEMENT FOR PROTECTING

An assembly including a protection device and an element of a turbine engine for protecting, for example, an oil tank including a magnetic particle detector, the protection device including a flexible protection blanket placed on and closed around the element for protecting and including eyelets that are engaged on fastener studs provided on the blanket or on the element for protecting. The protection device further includes at least one cable carrying an attachment mechanism for attaching to the fastener studs to prevent the studs being withdrawn from the eyelets in the blanket. 18-. (canceled)9. An assembly comprising:a protection device; andan element, or an oil tank including a magnetic particle detector, of a turbine engine for protecting;the protection device comprising a flexible protection blanket placed on and closed around the element for protecting and including eyelets that are engaged on fastener studs provided on the blanket or on the element for protecting,the protection device further comprising at least one cable carrying attachment means for attaching to the fastener studs to prevent the fastener studs being withdrawn from the eyelets in the blanket.10. An assembly according to claim 9 , wherein the attachment means comprises pins or spiral rings including branches or portions for passing through holes in the fastener studs.11. An assembly according to claim 9 , wherein the cable includes end loops in which the attachment means is engaged.12. An assembly according to claim 11 , wherein the attachment means is movable on the cable on which the end loop of the attachment means engaged in the loop form abutments claim 11 , preventing the movable attachment means being withdrawn from the cable.13. An assembly according to claim 9 , comprising at least two cables connected together via the attachment means.14. An assembly according to claim 9 , wherein the cable includes braiding wires and is surrounded by a sheath claim 9 , or a heat-shrink sheath.159. ...

METHOD FOR MOUNTING AN AIRCRAFT ENGINE ON A PYLON, AND ENGINE FASTENER FOR IMPLEMENTING SAID METHOD

A method for mounting an aircraft engine on a pylon includes positioning at least a first shear pin either in a first opening provided in a front engine fastener which is previously attached to the engine using connecting rods, or in a first bore provided in a front surface of the pylon; prepositioning the assembly including the engine and the front engine fastener relative to the pylon by placing the first shear pin opposite a receiving cavity, the receiving cavity being either the first bore, if the first shear pin has been positioned in the first opening, or the first opening, if the first shear pin has been positioned in the first bore; and inserting the first shear pin into the receiving cavity. 1. A method for mounting an aircraft engine onto a pylon , comprising: into a first opening provided in a front engine fastener, previously attached to the engine using connecting rods,', 'or into a first bore provided in a front face of the pylon;, 'positioning at least a first shear pin either either the first bore, if the first shear pin has been positioned into the first opening;', 'or the first opening, if the first shear pin has been positioned into the first bore;, 'pre-positioning the assembly including the engine and the front engine fastener, relative to the pylon by placing the first shear pin opposite a receiving cavity, said receiving cavity beinginserting the first shear pin into the receiving cavity.2. The method for mounting an aircraft engine onto a pylon according to claim 1 , wherein the pre-positioning:vertically hoisting the assembly including the engine and the front engine fastener;positioning an upper part of the front engine fastener and the front face of the pylon substantially in parallel.3. The method for mounting an aircraft engine onto a pylon according to claim 2 , wherein the pre-positioning includes claim 2 , between the hoisting and the positioning claim 2 , of shifting forwardly the assembly including the front engine fastener and the ...

METHOD FOR ASSEMBLING A NOZZLE AND AN EXHAUST CASE OF A TURBOMACHINE

A method for assembling a nozzle and an exhaust case of a turbomachine is disclosed. The exhaust case includes a hub and an outer ferrule connected to each other by a plurality of arms. The nozzle is attached to the outer ferrule of the exhaust case substantially at the trailing edges of the arms. 18-. (canceled)9. A method for assembling a nozzle and an exhaust case of a turbomachine , the exhaust case including a hub and an outer ferrule connected to each other by a plurality of arms , the method comprising:attaching the nozzle to the outer ferrule of the exhaust case in the same radial part of the turbomachine as that containing the trailing edges of the arms of the exhaust case,wherein the attachment of the nozzle to the exhaust case is performed at least partly through radial attachment flanges carried by the nozzles and the outer ferrule of the exhaust case respectively, using radial bolted joints.10. The method according to claim 9 , wherein the attachment of the nozzle to the exhaust case is performed at least partly through axial attachment flanges carried by the nozzle and the outer ferrule of the exhaust case respectively claim 9 , using axial bolted joints.11. The method according to claim 9 , wherein the nozzle includes a mixer claim 9 , the attachment being performed through the mixer.12. The method according to claim 9 , wherein at least one of the attachment flange of the outer ferrule or the attachment flange of the nozzle is scalloped claim 9 , in particular facing yokes present on the exhaust case.13. The method according to claim 12 , wherein the attachment flanges of the outer ferrule and of the nozzle are secured together by bolted joints claim 12 , and a scalloping is performed between each bolted joint.14. An exhaust system of a turbomachine comprising a nozzle and an exhaust case claim 12 , the exhaust case including a hub and an outer ferrule connected to each other by a plurality of arms claim 12 ,wherein the nozzle is attached to the ...

LINER BRACKET FOR GAS TURBINE ENGINE

A hanger for a gas turbine exhaust system includes an exhaust duct attachment structure associated with an exhaust duct and a liner attachment structure associated with a liner spaced radially inwardly of the exhaust duct. The exhaust duct attachment structure and the liner attachment structure cooperate to suspend the liner within the exhaust duct such that the exhaust duct and liner are movable relative to each other. At least one resilient member generates a resilient biasing force between the exhaust duct attachment structure and the liner attachment structure. 1. A hanger for a gas turbine exhaust system comprising:an exhaust duct attachment structure to be associated with an exhaust duct;a liner attachment structure to be associated with a liner spaced radially inwardly of the exhaust duct, wherein the exhaust duct and liner attachment structures cooperate to suspend the liner within the exhaust duct such that the exhaust duct and liner are movable relative to each other; andat least one resilient member generating a resilient biasing force between the exhaust duct attachment structure and the liner attachment structure.2. The hanger according to wherein one of the exhaust duct attachment structure and liner attachment structure comprises a housing configured to be attached to a respective one of the exhaust duct or liner claim 1 , and wherein the other of the exhaust duct attachment structure and liner attachment structure comprises a rod that is moveable relative to the housing.3. The hanger according to wherein the exhaust duct attachment structure comprises the housing and the liner attachment structure comprises the rod claim 2 , and including a rotatable member seated within the housing claim 2 , and wherein the rod has a first end associated with the rotatable member and a second end that is configured to be attached to the liner.4. The hanger according to wherein the resilient member reacts between the first end of the rod and the rotatable member.5. ...

HOT GAS EXPANDER INLET CASING ASSEMBLY AND METHOD

A gas turbine device includes an inlet casing and a stator diaphragm provided inside the inlet casing. The stator diaphragm has an integral inner stator shroud and an outer stator shroud and a plurality of stator vanes provided in a circumferential arrangement between the inner stator shroud and the outer stator shroud. A plurality of key slots is provided in a circumferential arrangement on the stator diaphragm and a corresponding plurality of key slots is provided in a circumferential arrangement on the inlet casing. The key slots provided on the stator diaphragm are aligned with the plurality of key slots provided on the inlet casing. A key is inserted into each of the plurality of key slots provided to prevent rotation of the stator diaphragm with respect to the inlet casing. The stator diaphragm is secured in an axial direction by a stator shear ring. 1. A gas turbine device , comprising:an inlet casing provided opposite a discharge casing along a longitudinal axis of the gas turbine device;a shaft disposed between the inlet casing and the discharge casing, the shaft being rotatable about the longitudinal axis of the gas turbine device;a plurality of rotor vanes extending radially from the shaft;a stator diaphragm provided inside the inlet casing, the stator diaphragm having an integral inner stator shroud and an outer stator shroud located concentric to the flow path inside the inlet casing; anda plurality of stator vanes provided in a circumferential arrangement between the inner stator shroud and the outer stator shroud.2. The gas turbine device according to claim 1 , further comprising a plurality of key slots provided in a circumferential arrangement on the stator diaphragm and a corresponding plurality of key slots provided in a circumferential arrangement on the inlet casing claim 1 , wherein the plurality of key slots provided on the stator diaphragm are aligned with the plurality of key slots provided on the inlet casing.3. The gas turbine device ...

Turboprop engine with compressor turbine shroud

A turboprop engine including an annular outer case including a mount ring for attachment to an aircraft along a plane to transfer loads from a propeller and having an annular flange extending radially inwardly therefrom within the plane, an annular turbine support case received within the outer case and connected to the outer case only through a direct connection with the annular flange allowing a limited relative pivoting motion between the turbine support case and the mount ring, and a turbine section including a rotor closely surrounded by a shroud with an annular tip clearance being defined therebetween, the shroud being directly connected to and located by the turbine support case. A method of isolating a turbine shroud from propeller loads in a propeller engine is also disclosed.

MID-TURBINE FRAME BUFFER SYSTEM

A mid-turbine frame buffer system for a gas turbine engine includes a mid-turbine frame that supports a shaft by a bearing. An air compartment and a bearing compartment are arranged radially inward of the mid-turbine frame. The bearing compartment is arranged within the air compartment and includes first and second contact seals arranged on either side of the bearing. The air compartment includes multiple air seals. A high pressure compressor is fluidly connected to the air compartment and is configured to provide high pressure air to the air compartment. A method of providing pressurized air to a buffer system includes sealing a bearing compartment with contact seals, surrounding the bearing compartment with an air compartment, and supplying high pressure air to the air compartment. 122-. (canceled)23. A method of providing pressurized air to a buffer system comprising:sealing a bearing compartment with two or more seals;surrounding the bearing compartment with an air compartment; andsupplying pressurized air to the air compartment.24. The method according to claim 23 , wherein the two or more seals comprise contact seals that include a carbon material.25. The method according to claim 23 , wherein the two or more seals comprise contact seals and two or more air seals claim 23 , and wherein at least one of the air seals is at least one of a labyrinth seal and a brush seal.26. The method according to claim 23 , wherein the pressurized air includes high pressure compressor air.27. The method according to claim 23 , wherein the bearing compartment is provided between low and high pressure turbines.28. The method according to claim 23 , comprising the step of providing a mid-turbine frame supporting a shaft by a bearing claim 23 , and the surrounding step includes the air compartment and the bearing compartment arranged radially inward of the mid-turbine frame claim 23 , the bearing compartment arranged within the air compartment claim 23 , the bearing compartment ...

PANEL SUPPORT HANGER FOR A TURBINE ENGINE

A support hanger is provided for a turbine engine. The support hanger includes a pin with a pin head, a retainer and a flexible seal. The retainer is pivotally connected to the pin head. The retainer includes a seal bearing surface facing towards the pin head. The seal includes a retainer bearing surface sealingly engaging the seal bearing surface. 1. A support hanger for a turbine engine , comprising:a pin including a pin head;a retainer pivotally connected to the pin head, the retainer including a seal bearing surface facing towards the pin head; anda flexible seal including a retainer bearing surface sealingly engaging the seal bearing surface.2. The support hanger of claim 1 , whereinthe seal extends radially between a first seal segment and a second seal segment;the first seal segment is axially displaced relative to the second seal segment; andthe second seal segment includes the retainer bearing surface.3. The support hanger of claim 1 , whereinthe seal includes a first seal segment, a second seal segment and a transition segment that extends radially and axially between the first seal segment and the second seal segment; andthe second seal segment includes the retainer bearing surface.4. The support hanger of claim 1 , further comprising:a collar including a threaded first collar segment and a second collar segment pivotally connected to the pin head;wherein the retainer has a retainer bore, and the first collar segment is mated with a threaded portion of the retainer bore.5. The support hanger of claim 4 , whereinthe pin head includes a parti-spherical collar bearing surface;the second collar segment includes a parti-spherical pin bearing surface that engages the collar bearing surface to pivotally connect the collar to the pin head; andthe pin bearing surface at least partially defines a collar bore that extending axially into the collar.6. The support hanger of claim 5 , wherein the collar bearing surface is convex claim 5 , and the pin bearing surface is ...

Thin Metal Duct Damper

A damper for damping vibration of a structural member of a gas turbine engine is disclosed. The damper may include a first metal mesh pad which abuts an outer circumferential surface of the structural member, and a garter spring which abuts the first metal mesh pad. Both the metal mesh pad and the garter spring may completely or partially encircle the structural member. Alternatively, the damper may include a damper cover which encloses the first metal mesh pad and the garter spring and which abuts the outer surface of the structural member. A second metal mesh pad may be inserted between the damper cover and the garter spring. A gas turbine engine which comprises such a damper is also disclosed. 1. A damper for damping vibration of a structural member of a turbine engine , comprising:a first metal mesh pad including a first surface which abuts an outer circumferential surface of the structural member; anda garter spring which abuts a second surface of the first metal mesh pad.2. The damper of claim 1 , wherein the first metal mesh pad encircles the structural member around the outer circumferential surface of the structural member.3. The damper of claim 1 , wherein the garter spring encircles the structural member around the outer circumferential surface of the structural member.4. The damper of claim 1 , further comprising:a damper cover which abuts the outer circumferential surface of the structural member and which forms a cavity between the damper cover and the outer circumferential surface of the structural member, at least a portion of the first metal mesh pad and at least a portion of the garter spring being in the cavity; anda second metal mesh pad between the damper cover and the garter spring, the second metal mesh pad abutting the damper cover and the garter spring.5. The damper of claim 4 , wherein the damper cover encircles the structural member around the outer circumferential surface of the structural member.6. The damper of claim 1 , wherein the ...

FLEXIBLE COUPLING FOR GEARED TURBINE ENGINE

A gas turbine engine includes a fan, a fan shaft coupled with the fan and arranged along an engine central axis, and a frame supporting the fan shaft. The frame defines a lateral frame stiffness (LFS). A non-rotatable flexible coupling and a rotatable flexible coupling support an epicyclic gear system. The couplings are subject to a Motion II of cantilever beam free end motion with respect to the engine central axis. The non-rotatable and the rotatable flexible couplings each have a stiffness of a common stiffness type under a common type of motion. The common stiffness type is a Stiffness B and the common type of motion is the Motion II. The Stiffness B of the rotatable flexible coupling is greater than the stiffness B of the non-rotatable flexible coupling, and a ratio of LFS/Stiffness B of the non-rotatable flexible coupling is in a range of 10-40. 1. A gas turbine engine comprising:a fan;a fan shaft coupled with the fan and arranged along an engine central axis;a frame supporting the fan shaft, the frame defining a lateral frame stiffness (LFS);an epicyclic gear system coupled to the fan shaft; anda non-rotatable flexible coupling and a rotatable flexible coupling supporting the epicyclic gear system, the non-rotatable flexible coupling and the rotatable flexible coupling being subject to a Motion II of cantilever beam free end motion with respect to the engine central axis,the non-rotatable flexible coupling and the rotatable flexible coupling each having a stiffness of a common stiffness type under a common type of motion with respect to the engine central axis, the common stiffness being defined with respect to the LFS, the common stiffness type is a Stiffness B and the common type of motion is the Motion II, the Stiffness B of the rotatable flexible coupling being greater than the stiffness of the non-rotatable flexible coupling, and a ratio of LFS/Stiffness B of the non-rotatable flexible coupling is in a range of 10-40.2. The gas turbine engine as recited ...

Unducted propeller turboshaft engine provided with a reinforcing shell integrating pipe segments

An airplane unducted propeller turboshaft engine having a gas generator and a receiver including a propulsion assembly carrying least one propeller, the engine including a first casing, a second casing, and a third casing, the third casing being provided between the first and second casings and surrounding at least a portion of the gas generator, a reinforcing shell presenting a first attachment zone mounted on the first casing second attachment zone mounted on the second casing, and a wall provided between the first and second attachment zones and surrounding the third casing, wherein the reinforcing shell further includes at least one pipe segment integrated in the wall.

Seals for gas turbine engine nacelle cowlings

A gas turbine engine includes an engine core and core nacelle cowling coupled to the engine core. The engine core has a core member that extends radially outward from the engine core. The core nacelle cowling has a cowling member that extends radially inward towards the engine core. The cowling member is offset axially from the core member to form a labyrinth seal that bounds a coolant inlet, thereby fluidly coupling the fan duct flow of the gas turbine engine with a core compartment defined between the engine core and core nacelle cowling.

Gas turbine engine attachment structure and method therefor

An attachment structure for a gas turbine engine includes a frame that has a first annular case. A second annular case extends around the frame. The first annular case and the second annular case include a plurality of interlocks. Each of the interlocks includes a first member mounted on one of the first annular case or the second annular case and a corresponding second member mounted on the other of the first annular case or the second case. The first member is received in the second member such that the plurality of interlocks restricts relative circumferential and axial movement between the first annular case and the second annular case.

SUPPORT LINKS WITH LOCKABLE ADJUSTMENT FEATURE

A cam-type apparatus is included in a support link between outer and inner cases of gas turbine engines for centering the cases one to another. The cam-type apparatus is lockable to allow locking an adjusted position of the cam-type apparatus. 1. A turbofan gas turbine engine comprising:a core portion of the engine;an annular bypass duct wall coaxially surrounding and supporting the core portion, thereby to define an annular bypass air passage radially between the core portion and the bypass duct wall for directing a bypass air flow passing therethrough; anda support link having a plurality of rods extending across the annular bypass air passage to interconnect the core portion and the annular bypass duct wall, the rods being connected to the core portion by a cam-type apparatus, the cam-type apparatus having a plurality of eccentric rotating pins adjustable for centering the core portion with respect to the annular bypass duct wall and having a plurality of connecting bases circumferentially spaced apart and attached to the core portion, each pin having a connecting section received in a hole of one of the rods and a base section received in a hole substantially radially extending through one of the connecting bases, the base section being eccentric to the connecting section at an eccentric distance, the base section being rotatable about a central axis thereof to adjust a position of the connecting section with the connected one of the rods, the base section being lockable to lock the adjusted position of the connecting section and the connected one of the rods.2. The turbofan gas turbine engine as defined in wherein: the cam-type apparatus is located outside the annular bypass air passage while the rods extend across the annular bypass air passage.3. The turbofan gas turbine engine as defined in wherein the rods comprise a first group of rods extending from an outer end to an inner end thereof in substantially tangential directions to the core portion claim 1 , ...

HEAT SHIELD MOUNT CONFIGURATION

An attachment interface assembly for a gas turbine engine has a component configured for attachment to an engine static structure. A fastener has a threaded body portion and an enlarged head portion. The threaded body portion is inserted through aligned holes in the component and engine static structure. A clip has a center opening that receives the threaded body portion such that the clip is positioned between the enlarged head portion and the component. 1. An attachment interface assembly for a gas turbine engine comprising:a component configured for attachment to an engine static structure;a fastener having a threaded body portion and an enlarged head portion, the threaded body portion to be inserted through aligned holes in the component and engine static structure; anda clip having a center opening that receives the threaded body portion such that the clip is positioned between the enlarged head portion and the component.2. The assembly according to wherein the component comprises a heat shield.3. The assembly according to wherein the engine static structure includes a circumferentially extending groove claim 2 , and wherein a portion of the clip is received within the groove to axially retain the clip claim 2 , fastener claim 2 , and heat shield to the engine static structure .4. The assembly according to wherein the clip comprises a generally flat body portion defining the center opening and which includes a plurality of gripping fingers extending outwardly from the flat body portion to grip the enlarged head portion.5. The assembly according to wherein the fastener comprises a tee-head bolt.6. The assembly according to wherein the heat shield comprises a thin sheet metal plate having a radially outer edge portion defined by a circumferentially extending mount flange having a first side and a second side claim 2 , and wherein the first side is in direct abutting engagement with the engine static structure and the second side is in direct abutting engagement ...

Exhaust system having a flow path liner supported by structural duct segments

The exhaust system ( 60 ) includes an exhaust flow path liner ( 62 ) surrounded and supported by a plurality of structural duct segments ( 64, 70 ). Pluralities of links ( 84 ) are secured to and extend between the duct segments ( 64, 70 ) and the liner ( 62 ). A duct end ( 88 ) of the link includes a lock member ( 96 ) having a diameter greater than a width of the stem ( 86 ). The lock member ( 96 ) is configured to be secured within a capture nest ( 98 ) defined between and within adjacent junction flanges ( 76, 80 } of the structural duct segments { 64, 70 ) when the segments ( 64, 70 ) are secured to each other to secure the segments ( 64, 70 ) together. A catch member ( 100 ) at an opposed end of the link ( 84 ) is secured to a capture node ( 102 ) at the exhaust liner ( 62 ).

Gas turbine engine

A gas turbine engine for an aircraft includes: a fan adjacent the engine air intake, including a plurality of fan blades; downstream of the fan, an engine core including a turbine, a compressor, and a core shaft connecting the turbine and compressor; an engine core housing at least partly encasing the core; a fan case surrounding the fan and defining at least part of a bypass duct radially outside the core; a plurality of outlet guide vanes extending between the engine core housing and an outlet guide vane support region of the case, adjacent an upstream end of the bypass duct; one or more supports extending from the case to the engine core housing, wherein: a first end of the supports fixes to the case at the outlet guide vane support region; a second end of the supports fixes to the engine core housing adjacent an engine core exhaust.

Method of disassembling and assembling gas turbine and gas turbine assembled thereby

A method of assembling and disassembling a gas turbine, and a gas turbine assembled thereby, improves work efficiency and reduces time and cost by carrying out various disassembly and reassembly processes depending on circumstances. In one process, a turbine section is disassembled from a gas turbine by sequential steps of disassembling an upper turbine case; disassembling a rear diffuser assembly and a rear bearing assembly; disassembling a combustor assembly; disassembling a vane assembly; and disassembling a blade assembly. In another process, first-stage to fourth-stage blade assemblies and first-stage to fourth-stage vane assemblies in a turbine section are disassembled from the gas turbine by sequential steps of disassembling an upper turbine case; disassembling a combustor assembly; disassembling a vane assembly; and disassembling a blade assembly. The gas turbine includes a compressor section, a combustor section, and the turbine section assembled in a reverse order with respect to the disassembly method.

MICRO GAS TURBINE SYSTEMS AND USES THEREOF

The present disclosure describes a micro gas turbine flameless heater, in which the heat is generated by burning fuel in a gas turbine engine, and the heater output air mixture is generated by transferring the heat in the gas turbine exhaust to the cold air drawn from the ambient environment. The present disclosure also describes component geometries and system layout for a gas turbine power generation unit that is designed for simple assembly, disassembly, and component replacement. The present disclosure also allows for quick removal of the rotating components of the gas turbine engine in order to reduce assembly and maintenance time. Furthermore, the present disclosure describes features that help to maintain safe operating temperatures for the bearings and structures of the gas turbine engine power turbine. Lastly, the present disclosure describes features of a fuel capture system that allow the injection of wellhead gas, which typically is a mixture of gaseous and liquid fuels, into the combustion chamber, and also describes methods of incorporating afterburners in the gas turbine engine, such that the gas turbine engine system can use wellhead gas to power equipment and reduce emissions from flaring in oil and gas applications. 117.-. (canceled)18. A method of operating a gas turbine heater , wherein the gas turbine heater comprises a gas turbine comprising i) an air starter; ii) a compressor; iii) a turbine; and iv) a combustion unit configured to receive compressed air for combustion from the compressor , to receive a fuel from a source , to burn the fuel to produce a combustion gas , and to supply the combustion gas to the turbine , wherein the gas turbine is configured to heat an external environment directly or indirectly using combustion gas exhausted from the turbine , the method comprising:operating a fan to pump air through an enclosure of the gas turbine heater such that the air passes from an ambient air inlet of the enclosure to an outlet of the ...

TURBOJET ENGINE WITH THRUST TAKE-UP MEANS ON THE INTER-COMPRESSOR CASE

A multiflow turbojet engine generally includes an upstream fan driven by a gas generator having first and second coaxial compressors, an intake case forming a mounting for the rotors of the upstream fan and the first compressor, an inter-compressor case downstream from the intake case and forming a mounting for the rotors of the second compressor, and attachment means for thrust take-up control rods arranged in the inter-compressor case. The turbojet engine also includes a structural force shroud connecting the intake case to the inter-compressor case, of the and a floating first compressor case. 1. A turbojet engine including:an upstream ducted fan driven by a gas generator, whereby the gas generator comprises a first compressor and a second compressor that is coaxial with the first compressor;an inlet case configured to form a support for a plurality of rotors of the upstream ducted fan and the first compressor;an inter-compressor case located downstream from the inlet case, and configured to form a support for a plurality of rotors of the second compressor;attachment means for a plurality of thrust take-up rods arranged on the inter-compressor case; anda stress structural shroud configured to connect the inlet case to the inter-compressor case,wherein the first compressor comprises a floating case that forms a wall of a flow path.2. The turbojet engine according to claim 1 , wherein the floating case that forms the wall of the flow path is connected in a floating configuration to one of the inlet case and the inter-compressor case by a backlash connection.3. The turbojet engine according to claim 1 , wherein the stress structural shroud is welded to the inlet case and bolted to the inter-compressor case.4. The turbojet engine according to claim 1 , wherein the stress structural shroud is bolted on the inlet case and bolted on the inter-compressor case.5. Turbojet The turbojet engine according to claim 1 , wherein the inlet case comprises a shroud that supports a ...

ENGINE ASSEMBLY FOR AN AIRCRAFT HAVING AN AIR-OIL EXCHANGER SYSTEM SUPPORT WITH OPTIMIZED ATTACHMENT

An engine assembly for an aircraft including a bypass turbomachine as well as a turbomachine attachment pylon including an air-oil exchanger system arranged in an inter-ducts compartment between the flow ducts, the compartment being delimited radially on the outside by an inter-ducts cowling, the exchanger system being supplied with air from a secondary flow duct of the turbomachine delimited radially on the inside by the inter-ducts cowling, and the exchanger system being supported by a support arranged in the inter-ducts compartment, this support being mechanically connected to the attachment pylon by connecting means passing through the inter-ducts cowling. 1. An engine assembly for an aircraft comprising a bypass turbomachine as well as a hooking mast of the turbomachine configured to ensure hooking of the turbomachine on a wing element of the aircraft , the turbomachine comprising an air-oil exchanger system arranged in an inter-flow compartment delimited radially outwards by an inter-flow cowling , the exchanger system being fed with air coming from a secondary flow path of the turbomachine delimited radially inwards by the inter-flow cowling , and the exchanger system being supported by a support arranged in the inter-flow compartment ,wherein said support is mechanically linked to the hooking mast by connecting means crossing the inter-flow cowling.2. The engine assembly according to claim 1 , wherein the connecting means are configured to enable the support to be displaced between an operating position placing the exchanger system inside the inter-flow compartment claim 1 , so that it radially covers one or several piece(s) of equipment within the inter-flow compartment claim 1 , and a maintenance position wherein said support is further away from a longitudinal central axis of the turbomachine than in the operating position.3. The engine assembly according to claim 2 , wherein the engine assembly is configured so that the maintenance position is accessible ...

METHOD FOR MANUFACTURING OF A GAS TURBINE ENGINE COMPONENT

The invention concerns a method for manufacturing of a gas turbine engine component () comprising an outer ring structure (), an inner ring structure (), and a plurality of circumferentially spaced elements () extending between the inner ring structure () and the outer ring structure (), wherein a primary gas channel for axial gas flow is defined between the elements (), and wherein the component () has an inlet side for gas entrance and an outlet side for gas outflow. The invention is characterized in that the method comprises the step of machining a one-piece metal blank as to form a one-piece part () comprising: a portion () of each of said elements (), wherein said portion (b) relates to a portion of an extension length of the elements () between said ring structures (); and a ring-shaped member () that connects said element portions () and that is intended to form part of one of the ring structures. The invention also concerns a gas turbine engine () comprising a component () manufactured according to the above method. 120-. (canceled)21. A method for manufacturing of a gas turbine engine component comprising an outer ring structure , an inner ring structure , and a plurality of circumferentially spaced elements extending between the inner ring structure and the outer ring structure , wherein a primary gas channel for axial gas flow is defined between the elements , and wherein the component has an inlet side for gas entrance and an outlet side for gas outflow , the method comprising: a portion of each of said elements, wherein said portion relates to a portion of an extension length of the elements between said inner ring structure and said outer ring structure; and', 'a ring-shaped member that connects said element portions and that forms part of one of the inner ring structure and the outer ring structure., 'machining a one-piece metal blank as to form a one-piece part comprising22. A method according to claim 21 , wherein the ring-shaped member forms part ...

METHOD OF PRODUCING SUSPENSION FOR A STRUCTURE IN A TURBOJET ENGINE USING A HYPERSTATIC TRELLIS WITH PRE-STRESSED LINK ELEMENTS

The invention relates to a method of manufacturing an assembly comprising a first structure (), arranged to be rigidly connected to a housing of a turbojet engine, a second annular structure () surrounding the first structure, and a hyperstatic trellis of connecting rods () maintaining the first structure () relative to the second (), said method comprising a step of mounting said connecting rods of the hyperstatic trellis between said structures and being characterised in that it comprises a step for pre-stressing at least one of said connecting rods () to a pre-determined level, carried out before the mounting thereof between said structures. It also relates to a device which is suitable for mounting the pre-stressed connecting rods. 1. Method of manufacturing an assembly comprising a first structure , arranged to be rigidly connected to a housing of a turbojet engine , a second annular structure surrounding the first structure , and a hyperstatic trellis of connecting rods maintaining the first structure relative to the second , said method comprising a step of mounting said connecting rods of the hyperstatic trellis between said structures , wherein it comprises a step of pre-stressing at least one of said connecting rods to a pre-determined level , carried out before the mounting thereof between said structures.2. Method according to claim 1 , wherein said connecting rod comprises a spindle with a yoke at each end which is suitable for fixing to one of said structures claim 1 , said pre-stressing step comprising the following sequences:applying pre-stressing at said predetermined level to the yokes of said connecting rod by suitable stressing means;installing and locking means on board said connecting rod which are suitable for maintaining the separation of the yokes, the stressing means maintaining the pre-stressing at said predetermined level;releasing the pre-stressing applied to the yokes of said connecting rod by the stressing means.3. Method according to ...

GAS TURBINE ENGINE FRONT ARCHITECTURE

A turbine engine is disclosed that includes a fan case surrounding a fan rotatable about an axis. A core is supported relative to the fan case by support structure arranged downstream from the fan. The core includes a core housing having an inlet case arranged to receive airflow from the fan. A compressor case is arranged axially adjacent to the inlet case and surrounds a compressor stage having a rotor blade with a blade trailing edge. The support structure includes a support structure leading edge facing the fan and a support structure trailing edge on a side opposite the support structure leading edge. The support structure trailing edge is arranged axially forward of the blade trailing edge. In one example, a forward attachment extends from the support structure to the inlet case. 1. A turbine engine comprising:a fan case surrounding a fan rotatable about an axis;a core supported relative to the fan case by a support structure and arranged downstream from the fan, the core including a core housing having an inlet case arranged to receive airflow from the fan, a compressor case axially adjacent to the inlet case and surrounding a compressor stage having a rotor blade with a blade trailing edge, wherein an intermediate case is arranged between the compressor case and a high pressure compressor case; andwherein the support structure includes a support structure leading edge facing the fan and a support structure trailing edge on a side opposite the support structure leading edge, and the support structure trailing edge arranged axially forward of the blade trailing edge, wherein a forward attachment extends from the support structure to the inlet case, wherein the intermediate case supports a rear portion of the compressor case near a compressed air bleed valve, and comprising a rearward attachment extending from the support structure to the intermediate case, wherein the front and rearward attachments are generally equidistant from the support structure to their ...

METHOD AND APPARATUS FOR REDUCING HIGH TRANSIENT MOUNT LOAD IN AIRCRAFT ENGINE MOUNTING SYSTEMS

Exemplary embodiments are provided to reduce the high transient mount load in an aircraft engine mounting system under extreme loading condition. The exemplary embodiments reduce the impact of the snubbing phenomenon without adding to the weight or space claim of the engine mounting system. 1. An apparatus for reducing the high transient mount load from an aircraft engine , the apparatus of the type including a housing with a sleeve for receiving a yoke pin , the apparatus comprising:a first wall having a first ovalized opening; anda second wall opposite the first wall, having a second ovalized opening;wherein the first ovalized opening and the second ovalized opening are for receiving the yoke pin therethrough.2. The apparatus of claim 1 , wherein the first ovalized opening and second ovalized opening each have a cross section that is continuous and substantially semi-circular according to a minor axis on a first half and substantially oval shaped having a major axis on a second half.3. The apparatus of claim 2 , wherein the major axis of the first ovalized opening and the major axis of the second ovalized opening are aligned and have a predetermined length.4. The apparatus of claim 3 , wherein the predetermined length is within a range of about 1.1 to about 1.4 times the minor axis.5. The apparatus of claim 4 , wherein the predetermined length is substantially equal to 1.3 times the minor axis.6. The apparatus of wherein the apparatus further comprises flexible packing material inside the housing.7. The apparatus of wherein the apparatus further comprises elastomer packing inside the housing.8. The apparatus of claim 1 , wherein the first ovalized opening comprises an ovalized area of a first predetermined thinness surrounding a half of a through-hole and the second ovalized opening comprises an ovalized area of a second predetermined thinness surrounding a half of a through-hole.9. The apparatus of wherein the first wall has a first wall thickness claim 8 , and ...

INSULATION SUPPORT SYSTEM FOR AN EXHAUST GAS SYSTEM

An insulation support system for an exhaust gas system gas turbine includes an outer casing having an inner surface and a liner sheet having an inner surface and that defines an insulation gap is defined between the inner surface of the liner sheet and the inner surface of the outer casing. The system further includes a support member that extends from the inner surface of the outer casing to the inner surface of the liner sheet and a threaded fastener that is fixedly connected to one of inner surface of the outer casing or the support member. The threaded fastener has a fixed depth and is aligned with a fastener hole of the liner sheet. The system further includes a bolt that extends through the bolt hole of the liner sheet and threadingly engages with the threaded fastener. 1. An insulation support system , comprising:an outer casing having an inner surface;a liner sheet having an inner surface and defining a fastener hole, wherein an insulation gap is defined between the inner surface of the liner sheet and the inner surface of the outer casing;a support member that extends from the inner surface of the outer casing to the inner surface of the liner sheet;a threaded fastener fixedly connected to one of inner surface of the outer casing or the support member, the threaded fastener having a fixed depth, wherein the threaded fastener is aligned with the fastener hole of the liner sheet; anda bolt that extends through the bolt hole of the liner sheet and threadingly engages with the threaded fastener.2. The insulation support system as in claim 1 , further comprising a spacer disposed between a head portion of the bolt and an outer surface of the liner sheet.3. The insulation support system as in claim 1 , wherein the threaded fastener is fixedly connected to a side wall of the support member.4. The insulation support system as in claim 1 , wherein the support member is an axially extending support member.5. The insulation support system as in claim 1 , wherein the ...

FIXTURING FOR THERMAL SPRAY COATING OF GAS TURBINE COMPONENTS

A fixture assembly includes a standoff, a backplate and a spray coating shield attached to a support. 1. A holder assembly , comprising:a standoff;a backplate positioned relative to the standoff; anda spray coating shield positioned relative to the standoff spaced from said backplate.2. The assembly as recited in claim 1 , wherein the backplate includes a multiple of apertures.3. The assembly as recited in claim 1 , wherein the backplate includes a multiple of apertures that correspond with apertures through a workpiece.4. The assembly as recited in claim 1 , wherein the backplate includes a multiple of apertures that correspond with studs that extend from a workpiece.5. The assembly as recited in claim 1 , wherein the spray coating shield is generally parallel to the backplate.6. The assembly as recited in claim 5 , wherein the spray coating shield is spaced a predetermined distance from a workpiece to mask an shadowed surface area of the workpiece with a shadow therefrom.7. The assembly as recited in claim 1 , wherein the spray coating shield is wedge shaped.8. The assembly as recited in claim 7 , wherein the spray coating shield is spaced a predetermined distance from a workpiece to mask a shadowed surface area of the workpiece with a shadow therefrom.9. The assembly as recited in claim 1 , wherein the spray coating shield is arranged radially and axially relative to the base plate.10. The assembly as recited in claim 1 , wherein the spray coating shield is arranged radially and axially relative to the backplate.11. The assembly as recited in claim 1 , wherein the spray coating shield is removable.12. The assembly as recited in claim 1 , further comprising a support claim 1 , the standoff supported by the support.13. The assembly as recited in claim 12 , wherein the standoff is a fixed ring mounted to a baseplate.14. A method of spraying a multiple of liner panels of a gas turbine engine comprising:mounting a multiple of holder assemblies to respectively support ...

TURBOMACHINE COMPONENT HANDLING ASSEMBLY

The present disclosure is directed to a component handling assembly. The component handling assembly includes a base member, an arm pivotably coupled to the base member, and a beam rotatably coupled to the base member. The component handling assembly also includes a counterweight rotatably coupled to a first end of the beam. The component handling assembly further includes a coupling assembly coupled to a second end of the beam. The coupling assembly is configured for coupling to a component. Rotating the counterweight in a first direction rotates the beam, the coupling assembly, and the component in a second direction. 1. A component handling assembly , comprising:a base member;an arm pivotably coupled to the base member;a beam rotatably coupled to the base member;a counterweight rotatably coupled to a first end of the beam; anda coupling assembly coupled to a second end of the beam, the coupling assembly being configured for coupling to a component,wherein rotating the counterweight in a first direction rotates the beam, the coupling assembly, and the component in a second direction.2. The assembly of claim 1 , further comprising:a first actuator coupled between the arm and the beam, the first actuator being configured to pivot the arm relative to the base member and the beam.3. The assembly of claim 2 , wherein the first actuator comprises a motor-driven winch.4. The assembly of claim 2 , wherein a first end of the arm is pivotably coupled to the base member and a second end of the arm defines the aperture for receiving a crane hook.5. The assembly of claim 4 , wherein the first actuator is coupled between the aperture and a portion of the beam between the base member and the first end of the beam.6. The assembly of claim 1 , further comprising:a second actuator configured to rotate the counterweight relative to the beam.7. The assembly of claim 6 , wherein the second actuator comprises a motor-driven slewing drive.8. The assembly of claim 1 , wherein the ...

SUPPORT OF A MACHINE, ROTATING MACHINE ENGINE AND A METHOD FOR ASSEMBLING OF SUCH A ROTATING MACHINE

A support for a machine part, particularly for a rotating machine part, more particularly for a gas turbine engine part, having a first support component, an interface component with a receiving unit, and a second support component. The second support component is connectable to the machine part, the second support component has a pin arranged to be engageable with the receiving unit, wherein the pin includes at least three bearing surfaces. The receiving unit is engageable with the first support component and has at least three rotatable substantially cylindrical or barrel-shaped rollers, wherein rolling surfaces of the rollers are in bearing contact with the bearing surfaces when the pin is engaged with the receiving unit. 1. A support for a machine part , comprising:a first support component,an interface component comprising a receiving unit, anda second support component, the second support component being connectable to the machine part, the second support component comprising a pin arranged to be engageable with the receiving unit, wherein the pin comprises at least three bearing surfaces,wherein the receiving unit is engageable with the first support component and comprises at least three rotatable substantially cylindrical or barrel-shaped rollers, wherein rolling surfaces of the rollers are in bearing contact with the bearing surfaces when the pin is engaged with the receiving unit.2. The support for a machine part claim 1 , according to claim 1 ,wherein the pin comprises a tapered tip and non-tapered section.3. The support for a machine part claim 1 , according to claim 1 ,wherein the pin has a longitudinal expanse and comprises a pin section along the longitudinal expanse that has a substantially triangular cross section taken perpendicularly to the longitudinal expanse.4. The support for a machine part claim 1 , according to claim 1 ,wherein the pin has a longitudinal expanse and comprises a pin section along the longitudinal expanse that has a cross ...

FAIL-SAFE ENGINE SUPPORT SYSTEM

An engine support mount including an airframe structure having a first anchor surface and a length extending along an x-axis, with a second axis defined perpendicularly to the x-axis. A support member is fixed to the airframe structure and defines a first aperture and a second aperture. A mounting assembly includes an elongated arm and first and second primary attachment assemblies respectively attaching the arm to the support member at the first and second apertures, the arm having a second anchor surface. A moment arm reduction feature includes a lug fixed to one of the anchor surfaces and a corresponding fastener fixed to the other of the anchor surfaces. The first anchor surface is positioned on the airframe structure outside of the first and second apertures, and the second anchor surface is positioned on the arm outside of the first and second attachment assemblies. 1. An engine support mount for supporting an engine having a casing , the mount comprising:an airframe structure having a length that extends along an x-axis, with a second axis being defined perpendicularly to the x-axis, the airframe structure comprising a first anchor surface;a support member fixed to the airframe structure and comprising a first lug spaced from the first anchor surface along the second axis;a mounting assembly comprising an elongated arm and a primary attachment assembly attaching the arm to the first lug, the elongated arm comprising a second anchor surface spaced from the primary attachment assembly along the second axis; anda moment arm reduction feature comprising a second lug and a corresponding fastener, the second lug being fixed to one of the anchor surfaces and the fastener being fixed to the other of the anchor surfaces,the fastener having a diameter,the second lug defining an aperture having a diameter,the fastener extending through the aperture,the diameter of the aperture being over-sized so that a load-bearing path is not established between the fastener and the ...

FLEXIBLE COUPLING FOR GEARED TURBINE ENGINE

A gas turbine engine includes a fan, a fan shaft coupled with the fan and arranged along an engine central axis, and a frame supporting the fan shaft. The frame defines a lateral frame stiffness (LFS). An epicyclic gear system is coupled to the fan shaft, and a non-rotatable flexible coupling and a rotatable flexible coupling support the epicyclic gear system. The non-rotatable flexible coupling and the rotatable flexible coupling each have a stiffness of a common stiffness type under a common type of motion with respect to the engine central axis. The stiffness is defined with respect to the LFS. The stiffness of the rotatable flexible coupling is greater than the stiffness of the non-rotatable flexible coupling. 1. A gas turbine engine comprising:a fan;a fan shaft coupled with the fan and arranged along an engine central axis;a frame supporting the fan shaft, the frame defining a lateral frame stiffness (LFS);an epicyclic gear system coupled to the fan shaft; anda non-rotatable flexible coupling and a rotatable flexible coupling supporting the epicyclic gear system,the non-rotatable flexible coupling and the rotatable flexible coupling each having a stiffness of a common stiffness type under a common type of motion with respect to the engine central axis, the stiffness being defined with respect to the LFS, the stiffness of the rotatable flexible coupling being greater than the stiffness of the non-rotatable flexible coupling.2. The gas turbine engine as recited in claim 1 , wherein the common type of motion is selected from Motion I claim 1 , Motion II claim 1 , Motion III claim 1 , or Motion IV claim 1 , where Motion I is parallel offset guided end motion claim 1 , Motion II is cantilever beam free end motion and Motion III is angular misalignment no offset motion and Motion IV is axial motion.3. The gas turbine engine as recited in claim 2 , wherein the epicyclic gear system includes a sun gear in meshed engagement with multiple intermediate gears that are ...

Modulating fuel for a turbine engine

A fuel supply system for a turbine engine that provides a modulated thrust control malfunction accommodation (TCMA). The fuel supply system can include a fuel line that fluidly connects a fuel tank and the turbine engine. A fuel pump and a fuel metering valve can be fluidly connected to the fuel line. A bypass line can fluidly connect to the fuel line. Flow through the bypass line can be controlled using a bypass valve and a balancing pressure valve. The TCMA can then modulate the fuel flow using the valves.

REDUCED STRESS BOSS GEOMETRY FOR A GAS TURBINE ENGINE

A case for a gas turbine engine includes a case wall and a boss that extends from the case wall. The boss includes a perimeter step. 120-. (canceled)21. A method for machining a case of a gas turbine engine , comprising:machining a first and second perimeter step into a case wall, the first and the second perimeter step forming a boss with a boss thickness greater than a machined case wall thickness.22. The method as recited in claim 21 , wherein machining the first perimeter step into the case wall comprises machining a radius machined around a periphery of the first perimeter step.23. The method as recited in claim 22 , wherein the perimeter step defines a radius between 0.0625″-0.25″ (1.588-6.35 mm).24. The method as recited in claim 22 , wherein machining the second perimeter step into the case wall comprises machining a radius machined around a periphery of the second perimeter step.25. The method as recited in claim 24 , wherein the perimeter step defines a radius between 0.0625″-0.25″ (1.588-6.35 mm).26. The method as recited in claim 21 , wherein the periphery of the first perimeter step and the periphery of the second perimeter step is triangular shaped.27. The method as recited in claim 21 , wherein the periphery of the first perimeter step and the periphery of the second perimeter step is rhomboid shaped.28. The method as recited in claim 21 , wherein the periphery of the first perimeter step and the periphery of the second perimeter step are of an equivalent shape.29. The method as recited in claim 21 , wherein the first perimeter step and the second perimeter step each defines a uniform height with respect to a flat top surface of the boss.30. The method as recited in claim 21 , wherein the perimeter step defines a radius between 0.0625″-0.25″ (1.588-6.35 mm).31. The method as recited in claim 21 , wherein the case wall is cylindrical.32. The method as recited in claim 21 , wherein the boss includes a through hole.33. The method as recited in claim 21 , ...

MID-TURBINE FRAME ROD AND TURBINE CASE FLANGE

A turbine section of a gas turbine engine includes a first turbine supported for rotation about an axis, a second turbine spaced axially aft of the for first turbine section for rotation about the axis, and a mid-turbine frame disposed between the first turbine and the second turbine defining a passage between the first turbine and the second turbine. A first case surrounds the first turbine and a second case surrounding the second turbine and attached to the first case. The mid-turbine frame is disposed between the first turbine section and the second turbine section and includes at least one support structure extending through an interface between the first turbine case and the second turbine case.

BENCH AFT SUB-ASSEMBLY FOR TURBINE EXHAUST CASE FAIRING

A fairing sub-assembly for a turbine frame comprises an inner ring , an outer ring and a plurality of strut-shells. The inner ring is formed of a plurality of inner segments . The outer ring is formed of a plurality of outer segments . The plurality of strut-shells connecting the inner ring and the outer ring . In another embodiment, the fairing sub-assembly comprises an inner band joining the plurality of inner segments and the plurality of strut-shells to form outer slots, and an outer band joining the plurality of outer segments and the plurality of strut-shells to form inner slots. A method of assembling a fairing comprises inserting the aforementioned fairing sub-assembly into an aft end of a turbine frame , inserting a plurality of forward strut-shells into the outer and inner slots at a forward end of the turbine frame , and joining the forward strut-shells to the fairing sub-assembly

SEALING ARRANGEMENT ON COMBUSTOR TO TURBINE INTERFACE IN A GAS TURBINE

A gas turbine unit having a combustor having a liner, a turbine, arranged downstream of the liner along a main flow gas direction and including a plurality of first stage vanes, a rotor cover support located inwardly of the vanes, and a sealing arrangement at a combustor to turbine interface, wherein the sealing arrangement includes a first dogbone seal extending between the rotor cover support and an inner downstream end of the liner or between the rotor cover support and a bulkhead located at the inner downstream end of the liner. 1. A gas turbine unit comprising:a combustor having a liner;a turbine, arranged downstream of the liner along a main flow gas direction (M) and including a plurality of first stage vanes;a rotor cover support located inwardly of the vanes;a sealing arrangement at a combustor to turbine interface, the sealing arrangement having a first dogbone seal extending between the rotor cover support and an inner downstream end of the liner, or between the rotor cover support and a bulkhead located at an inner downstream end of the liner.2. Gas turbine unit as claimed in claim 1 , wherein the first dogbone seal comprises:a central laminar portion; anda first and a second bulged edge, the edges being straight.3. Gas turbine unit as claimed in claim 2 , wherein a first end of the flat dogbone seal is housed in a groove of the rotor cover support.4. Gas turbine unit as claimed in claim 3 , wherein a portion of the rotor cover support provided with the groove is configured as an additional part.5. Gas turbine unit as claimed in claim 1 , wherein the sealing arrangement comprises:a second dogbone seal extending between the rotor cover support and a vane inner platform.6. Gas turbine unit as claimed in claim 5 , wherein the second dogbone seal comprises:a central laminar portion and a first and a second bulged edge, at least the second edge being curved.7. Gas turbine unit as claimed in claim 1 , comprising:a honeycomb seal arranged on a bulkhead portion ...

COMBUSTOR ANTI-SURGE RETENTION SYSTEM

A combustor assembly for a gas turbine engine comprises a vane support ring; an annular combustor extending around a central axis and being located radially inwards of the vane support ring, the annular combustor including an aft end, a forward end with at least one opening through which at least one fuel nozzle is received, and an annular outer shell and an annular inner shell that define an annular combustion chamber there between, the annular outer shell including a radially-outwardly extending flange including at least one scallop cut; and a stop member integrally formed with the vane support ring. 1. A combustor assembly for a gas turbine engine , comprising:a vane support ring;an annular combustor extending around a central axis and being located radially inwards of the vane support ring, the annular combustor including an aft end, a forward end with at least one opening through which at least one fuel nozzle is received, and an annular outer shell and an annular inner shell that define an annular combustion chamber there between, the annular outer shell including a radially-outwardly extending flange including at least one scallop cut; anda stop member integrally formed with the vane support ring.2. The combustor assembly according to claim 1 , wherein the radially-outwardly extending flange is located at the aft end.3. The combustor assembly according to claim 1 , wherein the radially-outwardly extending flange is annular.4. The combustor assembly according to claim 1 , wherein said at least one scallop cut and said stop member integrally formed with the vane support ring are configured for direct axial installation of the vane support ring over the combustor annular outer shell.5. The combustor assembly according to claim 1 , wherein the scallop cut in the radially-outwardly extending flange is sized to allow said stop member to pass through said radially-outwardly extending flange.6. The combustor assembly according to claim 1 , wherein the scallop cut in ...

MID-TURBINE FRAME WITH TENSIONED SPOKES

A mid-turbine frame located in a gas turbine engine axially aft of a high-pressure turbine and fore of a low-pressure turbine includes an outer frame case, an inner frame case, and at least a first spoke connecting the outer frame case to the inner frame case. The first spoke is tightened so as to be in tension during substantially all operating conditions of the gas turbine engine. 1. A gas turbine engine comprising:a combustor;a first turbine section in fluid communication with the combustor;a second turbine section in fluid communication with the first turbine section; and an outer frame case;', 'an inner frame case; and', 'at least a first spoke connecting the outer frame case to the inner frame case, wherein the first spoke is tightened so as to be in tension during substantially all operating conditions of the gas turbine engine., 'a mid-turbine frame located axially between the first turbine section and the second turbine section, the mid-turbine frame comprising2. The gas turbine engine of claim 1 , wherein the first spoke comprises:a tie rod having a first threaded surface; anda retaining nut having a second threaded surface engaged with the first threaded surface.3. The gas turbine engine of claim 2 , wherein the retaining nut comprises a flange extending outward from an outer surface of the retaining nut.4. The gas turbine engine of claim 3 , wherein the retaining nut is positioned in a hole of the outer frame case claim 3 , with the flange positioned radially outward of the outer frame case and the second threaded surface positioned radially inward of the outer frame case.5. The gas turbine engine of claim 3 , wherein a bolt extends through the flange into the outer frame case.6. The gas turbine engine of claim 2 , wherein the first threaded surface is on an outer surface of the tie rod claim 2 , wherein the second threaded surface is on an inner surface of the retaining nut claim 2 , and wherein the first and second threaded surfaces overlap partially ...

IN-LINE REMOVABLE HEAT SHIELD FOR A TURBOMACHINE SUSPENSION YOKE

Heat shield device for a means of suspending an engine from an aircraft, the said device comprising panels able to be fixed to the said means, at least one of the said panels extending radially in the direction of the axis of the engine, characterized in that it further comprises a component configured to be fixed to the said engine and comprising a groove shaped in such a way that the lower end of the said panel can be inserted into it. 1. Heat shield device for a means of suspending an engine from an aircraft , the said device comprising panels able to be fixed to the said means , at least one of the said panels extending radially in the direction of the axis of the engine , characterized in that it further comprises a component configured to be fixed to the said engine and comprising a groove shaped in such a way that the lower end of the said panel can be inserted into it.2. Device according to claim 1 , in which the said groove is formed by a bottom and at least two extensions extending in parallel and spaced-apart directions so that the lower end of the said panel can be inserted into it.3. Device according to claim 1 , in which the said panel comprises at least two ribs extending in the direction of the axis of the engine claim 1 , the said ribs being spaced apart so that one of the extensions of the said component can be inserted between them when the said lower end of the said panel is inserted into the groove of the said component.4. Device according to claim 1 , in which the said groove has a width greater than that of the said panel so as to allow it to move laterally within the said groove.5. Device according to claim 3 , in which the space in-between the two ribs of the said panel has a width greater than that of the extension of the said component claim 3 , so as to allow the said panel to move laterally with respect to the said component.651. Device according to claim 1 , shaped to surround the said means of attachment by four parallel walls ...

TURBINE SECTION OF HIGH BYPASS TURBOFAN

A turbofan engine according to an example of the present disclosure includes, among other things, a fan including an array of fan blades rotatable about an engine axis, a compressor including a first compressor section and a second compressor section, the second compressor section including a second compressor section inlet with a compressor inlet annulus area, a fan duct including a fan duct annulus area outboard of the second compressor section inlet, and a turbine having a first turbine section driving the first compressor section, a second turbine section driving the fan through an epicyclic gearbox, the second turbine section including blades and vanes, and wherein the second turbine section defines a maximum gas path radius and the fan blades define a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.6. 1. A turbofan engine comprising:a fan including an array of fan blades rotatable about an engine axis;a compressor including a first compressor section and a second compressor section, the second compressor section including a second compressor section inlet with a compressor inlet annulus area;a fan duct including a fan duct annulus area outboard of the second compressor section inlet, wherein a ratio of the fan duct annulus area to the compressor inlet annulus area defines a bypass area ratio between 8.0 and 20.0;a turbine having a first turbine section driving the first compressor section, and a second turbine section driving the fan through an epicyclic gearbox, the second turbine section including blades and vanes, and a second turbine airfoil count defined as the numerical count of all of the blades and vanes in the second turbine section;wherein a ratio of the second turbine airfoil count to the bypass area ratio is between 100 and 150;wherein the second turbine section defines a maximum gas path radius and the fan blades define a maximum radius, and a ratio of the maximum gas path radius to ...

TURBINE SECTION OF HIGH BYPASS TURBOFAN

A turbofan engine according to an example of the present disclosure includes, among other things, a fan including a circumferential array of fan blades, a compressor in fluid communication with the fan, the compressor including a first compressor section and a second compressor, the second compressor section including a second compressor section inlet with a second compressor section inlet annulus area, a fan duct including a fan duct annulus area outboard of the second compressor section inlet, a shaft assembly having a first portion and a second portion, a turbine in fluid communication with the combustor, the turbine having a first turbine section coupled to the first portion of the shaft assembly to drive the first compressor section, and a second turbine section coupled to the second portion of the shaft assembly to drive the fan, an epicyclic transmission coupled to the fan and rotatable by the second turbine section through the second portion of the shaft assembly to allow the second turbine to turn faster than the fan, wherein the second turbine section includes a maximum gas path radius and the fan blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.6.

STRUCTURAL GUIDE VANE LEADING EDGE

A structural guide vane for use in a gas turbine engine has a leading edge section, a trailing edge, a pressure surface and a suction surface. An erosion coating such as polyurethane resin is on the pressure surface and the suction surface. The leading edge of the vane is without the erosion coating and is bare metal. The vane is formed to include a plurality of pockets and bond shelves in the pressure surface side, and an epoxy bond line on the bond shelves holding a cover plate protected by the erosion coating. 1. A structural guide vane for use in a gas turbine engine , the structural guide vane comprising:a vane having a leading edge section, a trailing edge, a pressure surface and a suction surface; anda first erosion coating on the pressure side extending from the leading edge section to the trailing edge and a second erosion coating on the suction surface extending from the leading edge section to the trailing edge with both the first and second erosion coatings starting rearward of the leading edge and the leading edge is an exposed portion of the structural guide vane.2. The guide vane of claim 1 , wherein the erosion coating is a polyurethane resin.3. The guide vane of claim 1 , wherein the pressure side has a thicker erosion coating than the suction side coating.4. The guide vane of claim 3 , wherein the suction side has an erosion coating about 0.005 inches (0.0127 cm) thick and pressure side has a thicker erosion coating of about 0.010 inches (0.0254 cm).5. The guide vane of claim 1 , wherein the vane includes a plurality of pockets and bond shelves in the pressure surface side claim 1 , and an epoxy bond line on the bond shelves holding a cover plate to cover the plurality of pockets and provide a uniform pressure surface.6. The guide vane of claim 1 , wherein the erosion coating is a polyurethane resin and wherein the pressure side has a thicker erosion coating than the suction side coating.7. The guide vane of claim 6 , wherein the suction side has ...

Leaf spring hanger for exhaust duct liner

A hanger assembly for use between a first duct and a second duct has a flexible leaf spring having a body and a leg, a locking member for attaching the leg to the first duct, and a mounting member for attaching the body to the second duct.

ARRANGEMENT FOR CONNECTING A DUCT TO AN AIR-DISTRIBUTION CASING

An arrangement for connecting at least one duct with an air-distribution casing, the at least one duct including two sidewalls opposite one another and including a peripheral wall connecting edges of the two sidewalls. The arrangement includes a connecting tube that extends through the casing, passes through an orifice associated with each of the two sidewalls, and is connected to the at least one duct. A system controlling clearance of a turbo engine and a turbo engine can include air-injection ducts connected to the distribution casing by such an arrangement. 19-. (canceled)10. An arrangement for connecting at least one duct with an air-distribution casing , the at least one duct including two sidewalls opposite one another and a peripheral wall connecting edges of the two sidewalls , the arrangement comprising:a connecting tube that extends through the casing, passes through an orifice associated with each of the two sidewalls, and is connected to the at least one duct, andwherein an edge of each orifice is bent outwardly from the casing and forms a bushing to which the connecting tube is secured.11. An arrangement according to claim 10 , wherein the connecting tube is a section of the at least one duct.12. An arrangement according to claim 10 , wherein an end of the connecting tube is connected to an associated end of the at least one duct.13. An arrangement according to claim 10 , wherein the connecting tube is open in an inside volume of the casing.14. An arrangement according to claim 10 , wherein the connecting tube is sealingly secured to each of the sidewalls.15. An arrangement according to claim 14 , wherein a connection between the connecting tube and each sidewall is carried out by soldering.16. A device for controlling clearance of a turbine of a turbo engine comprising a plurality of air-injection cooling ducts that are connected to an air-distribution casing claim 14 ,{'claim-ref': {'@idref': 'CLM-00010', 'claim 10'}, 'wherein each air-injection duct ...

PYLON SHAPE WITH GEARED TURBOFAN FOR STRUCTURAL STIFFNESS

A gas turbine includes a mounting structure for mounting the engine to a pylon. A propulsor section includes a fan having a fan diameter. A geared architecture drives the fan. A generator section includes a fan drive turbine that drives the geared architecture. A turbine to fan diameter ratio is substantially less than 45%. 1. A gas turbine engine comprising: 'a propulsor section including a fan having a fan diameter; and a geared architecture driving the fan; and', 'a mounting structure for mounting the engine to a pylon; and'}a generator section including a fan drive turbine that drives the geared architecture, wherein a turbine to fan diameter ratio is substantially less than 45%.2. An assembly including the gas turbine engine of assembled to a pylon mounted to a wing.3. The assembly of wherein the fan extends forward of the wing.4. The assembly of wherein a distance of substantially 11 inches is defined between the wing and an upper portion of the gas turbine engine5. The gas turbine engine as recited in wherein the geared architecture includes an epicyclic gearbox.6. The gas turbine engine as recited in claim 1 , mounted to a pylon claim 1 , and wherein the pylon includes a forward portion and an aft portion claim 1 , the fan is attached to the forward portion of the pylon claim 1 , and the turbine section including the fan drive turbine is attached to the aft portion of the pylon.7. The gas turbine engine and mounting system as recited in wherein the turbine section is a low pressure turbine.8. The gas turbine engine and mounting system as recited in wherein the turbine to fan diameter ratio is substantially 35% to substantially 40%.9. A gas turbine engine comprising:a mounting structure for mounting the engine to a pylon, the pylon including a forward portion and an aft portion;a propulsor section including a fan having a fan diameter, wherein the fan is attached to the forward portion of the pylon;a geared architecture driving the fan;a compressor section;a ...

System and method for gas turbine engine mount with seal

A mount for coupling an engine to a vehicle includes an engine bracket to couple to the engine. The engine bracket includes a body that has a first end opposite a second end, and the body defines an offset coupling portion between the first end and the second end. The offset coupling portion protrudes from the body between the first end and the second end to define a receptacle. The mount includes a vehicle bracket to couple to the vehicle. The vehicle bracket includes a first bracket end opposite a second bracket end. The first bracket end is offset from the second bracket end and the first bracket end is received within the receptacle to couple the engine bracket to the vehicle bracket.

System and method for gas turbine engine mount with seal

A seal for a wall of a vehicle includes a first plate that defines a first slot, and the first plate is configured to be coupled to the wall. The seal includes a second plate that defines a guide that extends outwardly from the second plate. The second plate is positioned adjacent to the first plate such that the guide is in communication with the first slot. The seal includes a third plate that defines a second slot that receives the guide, and the third plate is positioned adjacent to the second plate and configured to be coupled to the wall.

CONNECTING GAS TURBINE ENGINE ANNULAR MEMBERS

A gas turbine engine assembly includes first and second annular members having different first and second thermal expansion coefficients connected together with dual arm V brackets. Brackets include first and second arms angularly spaced apart from a bracket centerline and extending axially away from bracket bases attached to a first one of the first and second annular members. Arms are attached to a second one of the first and second annular members. A turbine frame includes struts extending between outer and inner rings. An annular mixer and centerbody substantially made from a ceramic matrix composite materials is connected to and supported by the outer and inner rings with first and second sets respectively of the dual arm V brackets. Bracket bases of the first and second sets are attached to the outer and inner rings respectively. Arms of the first and second sets are attached to mixer and centerbody respectively. 1. A gas turbine engine assembly comprising:gas turbine engine first and second annular members,the second annular member substantially made from a ceramic matrix composite material,the second annular member connected to supported by the first annular member with dual arm V brackets,each of the dual arm V brackets including first and second arms angularly spaced apart from a bracket centerline,the first and second arms extending axially away from a bracket base,the bracket bases attached to a first one of the first and second annular members, andthe first and second arms attached to a second one of the first and second annular members.2. The assembly as claimed in further comprising:the bracket bases and the first and second arms being substantially flat,first and second arm holes in distal ends of the first and second arms respectively,first screws or bolts disposed through base holes in the bracket bases attaching the bracket bases to the first one of the first and second annular members, andsecond screws or bolts disposed through first and second ...

AIRCRAFT TURBOMACHINE ASSEMBLY COMPRISING AN ARTICULATED COWL

A turbomachine assembly for an aircraft comprising a nacelle with a nacelle structure, cowls attached to the structure to create an aerodynamic surface, an articulated cowl, comprising four edges, in the upper portion of the nacelle, an articulation at a first edge of the articulated cowl attached between the articulated cowl and the nacelle structure allowing the articulated cowl to move between a closed and an open position, a locking system at a second edge of the articulated cowl, opposite the first edge. The locking system comprises a handle actuable from the exterior of the nacelle, between a first position and a second position, a locking device movable between a locking position and an unlocking position, and a transmission system that moves the locking device from the locking position to the unlocking position when the handle passes from the first position to the second position, and vice versa. 1. A turbomachine assembly for an aircraft comprising a nacelle having a longitudinal axis , said nacelle comprising:a nacelle structure,cowls attached to said structure to create an aerodynamic surface,an articulated cowl comprising four edges, and being located in an upper portion of the nacelle, with a front edge located at a front of the articulated cowl in terms of a longitudinal axis of the nacelle and a rear edge located at a rear of the articulated cowl in terms of the longitudinal axis,an articulation located at the rear edge of the articulated cowl, and attached between the articulated cowl and the nacelle structure, said articulation allowing the articulated cowl to move between a closed position, in which the articulated cowl is aerodynamically continuous with the other cowls, and an open position, in which the articulated cowl frees an opening in the nacelle, a handle actuable from an exterior of the nacelle, between a first position and a second position,', 'a locking device able to move between a locking position, in which the locking device prevents ...

GAS TURBINE ENGINE

A gas turbine engine for an aircraft, comprising: an engine core; a covering configured to cover at least part of the engine core; and an engine accessory that interacts with the engine core or a component of the engine core; wherein the engine accessory is outside the covering. 1. A gas turbine engine for an aircraft , comprising;an engine core;a covering configured to cover at least part of the engine core; andan engine accessory that interacts with the engine core or a component of the engine core; whereinthe engine accessory is disposed on the outside of the covering.2. The gas turbine engine of claim 1 , whereinthe covering further comprises a socket; andthe engine accessory is removably connected to the socket.3. The gas turbine engine of claim 1 , whereinthe gas turbine engine further comprises an interface connecting the engine accessory to the engine core or the component of the engine core.4. The gas turbine engine of claim 3 , whereinthe interface comprises a socket; andthe engine accessory is removably connected to the socket.5. The gas turbine engine of claim 3 , whereinthe interface comprises a conduit that provides a fluid connection from the engine accessory to the engine core or the component of the engine core.6. The gas turbine engine of claim 3 , whereinthe interface comprises an electrical connection from the engine accessory to the engine core or the component of the engine core.7. The gas turbine engine of claim 3 , whereinthe interface comprises an connecting member that provides an actuating mechanism between the engine accessory to the engine core or the component of the engine core.8. The gas turbine engine of claim 3 , whereinthe engine accessory is detachable from the interface.9. The gas turbine engine of claim 1 , whereinthe engine accessory is disposed on the covering such that the engine accessory and the covering form a continuous surface.10. The gas turbine engine of claim 9 , wherein the engine accessory is configured to transfer ...

Turbojet engine cold stream flow path suspended from the exhaust case by radial crevice mounts and link rods

A bypass turbojet engine including a cylindrical cold stream flow path supported by link rods which are attached to a cylindrical outer shell ring of an exhaust case at attachment points. The attachment points for the exhaust case are crevice mounts including lugs that extend radially from the outer shell ring, a bore of the crevice mounts being directed in a direction of generatrices of the outer shell ring.

SEALING ARRANGEMENT FOR GAS TURBINE TRANSITION PIECES

A sealing arrangement for sealing between adjacent first and second exit frames each associated with a transition piece of a turbine. The arrangement includes a transition side seal having bristles that extend from a rail wherein the rail is bendable in a plurality of directions to accommodate bowing of an associated exit frame. The first exit frame includes a rail slot for receiving the rail. The second exit frame includes a bristle slot for receiving the bristles to form a seal between the first and second exit frames, wherein the bristle slot includes a bottom section. The bristles are separated from the bottom section by a bristle gap that is sized to accommodate movement of at least one transition piece to maintain a seal between the first and second exit frames or avoid damage to the bristles. 1. A sealing arrangement for sealing between adjacent exit frames of a turbine , wherein the exit frames are separated by a frame gap and each exit frame is associated with a transition piece , comprising:a transition side seal having bristles that extend from a rail;a first exit frame having a rail slot for receiving the rail; anda second exit frame having a bristle slot for receiving the bristles to form a seal between the first and second exit frames, wherein the bristle slot includes a bottom section and the bristles are separated from the bottom section by a bristle gap.2. The sealing arrangement according to claim 1 , wherein the bristle gap is sized to accommodate movement of at least one transition piece to maintain a seal between the first and second exit frames or avoid damage to the bristles.3. The sealing arrangement according to claim 2 , wherein movement of the at least one transition piece causes the frame gap to either increase or decrease in size.4. The sealing arrangement according to claim 2 , wherein movement of the at least one transition piece causes a misalignment of the transition pieces.5. The sealing arrangement according to claim 1 , wherein ...

COMBUSTOR REPLACEMENT METHOD AND GAS TURBINE PLANT

A combustor replacement method and a gas turbine plant capable of efficiently replacing a combustor using an existing facility. The combustor replacement method includes a step of separating, from a plurality of fuel supply systems, a first combustor that includes a plurality of nozzle systems connected to any of the plurality of fuel supply systems and supplied with fuel from the connected fuel supply systems, and removing the first combustor from a gas turbine plant. The method includes a step of attaching a second combustor that includes fewer nozzle systems than the first combustor to the gas turbine plant, and a step of providing communication between the fuel supply systems connected to the same nozzle system of the second combustor by a coupling pipe, and coupling the fuel supply systems and the second combustor. 1. A combustor replacement method for replacing a combustor of a gas turbine in a gas turbine plant including the combustor and a plurality of fuel supply systems that supplies fuel to the combustor , the method comprising the steps of:separating, from the fuel supply systems, a first combustor that comprises a plurality of nozzle systems connected to any of the plurality of fuel supply systems and supplied with the fuel from the connected fuel supply systems, and removing the first combustor from the gas turbine plant;attaching a second combustor that comprises fewer nozzle systems than the first combustor to the gas turbine plant; andproviding communication between the fuel supply systems connected to the same nozzle system of the second combustor by a coupling pipe, and coupling the fuel supply systems and the second combustor.2. The combustor replacement method according to claim 1 , whereinthe fuel supply systems comprise manifolds and branch pipes that connect the manifolds and the nozzle systems, andthe coupling pipe connects the branch pipe and the branch pipe.3. The combustor replacement method according to claim 2 , further comprising a ...

REAR MOUNT ASSEMBLY FOR GAS TURBINE ENGINE

A gas turbine engine with a rear mount assembly including link rods interconnecting the bypass duct wall and the core portion and connecting assemblies connected to the bypass duct wall. Each connecting assembly has inner and outer surfaces bordering an opening through which an outer end of a respective link rod extends. The outer surface is accessible from outside the bypass duct wall. A first locking member is engaged to the outer end in a first locked position, and includes an abutment portion located radially inwardly of the inner surface and abutting the inner surface, and an outer portion protruding radially outwardly through the opening beyond the outer surface. A second locking member is engaged the outer end in a second locked position, and has an abutment portion located radially outwardly of the outer surface and abutting the outer surface. A method of supporting a core portion is also discussed. 1. A gas turbine engine comprising:a core portion received within an annular bypass duct wall, a bypass air passage being defined by an annular space between the core portion and bypass duct wall;a front mount assembly supporting the core portion in proximity of an inlet of the bypass air passage; a plurality of link rods interconnecting the bypass duct wall and the core portion to form a load transfer path therebetween;', an inner surface facing radially inwardly and an outer surface located radially outwardly of the inner surface and facing radially outwardly, both the inner and outer surfaces bordering an opening through which an outer end of a respective one of the link rods extends, the outer surface being accessible from outside of the bypass duct wall,', 'a first locking member engaged to the outer end in a first locked position, the first locking member including a first abutment portion located radially inwardly of the inner surface and abutting the inner surface, the first locking member having an outer portion protruding radially outwardly through the ...

GAS TURBINE ENGINE COMPONENT

The invention concerns a gas turbine engine component () comprising an outer ring (), an inner ring (), a plurality of circumferentially spaced elements () extending between the inner ring () and the outer ring (), wherein an annular load transfer structure () extends circumferentially along an inner side of the inner ring () and also inwards in a radial direction of the component (). The first portion (), at least along a part of the circumference, is inclined in the radial direction in relation to a second portion (), wherein the two inclined portions () are connected in a connection zone () between the inclined portions (), and wherein a radial and/or axial position of the connection zone () varies along the circumference such that the radial/axial position of the connection zone in a location that circumferentially corresponds to a first element (), i.e. in a first circumferential location (A), is radially and/or axially different from the radial/axial position of the connection zone () in-between the first element () and an adjacent second element (), i.e. in a second 132-. (canceled)33. A gas turbine engine component , comprising:an outer ring;an inner ring;a plurality of circumferentially spaced elements extending between the inner ring and the outer ring, wherein a primary gas channel for axial gas flow is defined between the elements, wherein the component has an inlet side for gas entrance and an outlet side for gas outflow; andan annular load transfer structure for transferring loads between said elements and a bearing structure for a turbine shaft positioned centrally in the component;wherein:the annular load transfer structure extends circumferentially along an inner side of the inner ring and also inwards in a radial direction of the component, said annular load transfer structure having a first portion and a second portion, wherein the first portion is located closer to the inner ring than the second portion;the first portion, at least along a part of ...

AIRCRAFT COMPONENT AND GAS TURBINE ENGINE FOR AIRCRAFT

An aircraft component is used in a gas turbine engine for an aircraft. The aircraft component includes an annular part having an outer peripheral surface, and a boss part protruding from the outer peripheral surface of the annular part in a radial direction. In the boss part, at least two through-holes are formed to penetrate the boss part in the radial direction at predetermined intervals. In the boss part around the two through-holes, a cut-out part where a part of the boss part is cut out is formed. 1. An aircraft component used in a gas turbine engine for an aircraft , the aircraft component comprising:an annular part having an outer peripheral surface; anda boss part protruding from the outer peripheral surface of the annular part in a radial direction, whereinin the boss part, at least two through-holes are formed to penetrate the boss part in the radial direction at predetermined intervals,in the boss part around the two through-holes, a cut-out part where a part of the boss part is cut out is formed, andthe cut-out part is a recessed part that is formed from a peripheral edge of the boss part toward a space between the two through-holes.2. (canceled)3. The aircraft component according to claim 1 , wherein the recessed part is formed into a semicircular shape in the outer peripheral surface.4. The aircraft component according to claim 1 , whereinone of the two through-holes is an aperture, and the other is a bolt-hole having an aperture area smaller than the aperture, anda radius of the recessed part is the same as a radius of the bolt-hole.5. An aircraft component used in a gas turbine engine for an aircraft claim 1 , the aircraft component comprising:an annular part having an outer peripheral surface; anda boss part protruding from the outer peripheral surface of the annular part in a radial direction, whereinin the boss part, at least two through-holes are formed to penetrate the boss part in the radial direction at predetermined intervals,in the boss part ...

FLEXIBLE COUPLING FOR GEARED TURBINE ENGINE

A gas turbine engine includes a fan, a fan shaft coupled with the fan and arranged along an engine central axis, a gear system rotatably coupled to the fan shaft, and first and second flexible couplings supporting the gear system. The first flexible coupling and the second flexible coupling are subject to, with respect to the engine central axis, torsional motion and lateral motion. The first and second flexible couplings individually having a torsional stiffness under the torsional motion and a lateral stiffness under the lateral motion, and a ratio, of each of the first and second flexible couplings, of the torsional stiffness to the lateral stiffness is greater than or equal to 2. 1. A gas turbine engine comprising:a fan;a fan shaft coupled with the fan and arranged along an engine central axis;a gear system rotatably coupled to the fan shaft; andfirst and second flexible couplings supporting the gear system, the first flexible coupling and the second flexible coupling being subject to, with respect to the engine central axis, torsional motion and lateral motion,the first and second flexible couplings individually having a torsional stiffness under the torsional motion and a lateral stiffness under the lateral motion, and a ratio, of each of the first and second flexible couplings, of the torsional stiffness to the lateral stiffness is greater than or equal to 2.2. The gas turbine engine as recited in claim 1 , wherein the gear system includes a sun gear in meshed engagement with multiple intermediate gears that are rotatably mounted on bearings in a non-rotatable carrier claim 1 , each intermediate gear is in meshed engagement with a rotatable ring gear claim 1 , the sun gear is rotatably coupled to the fan shaft claim 1 , and the first flexible coupling is coupled with the non-rotatable carrier.3. The gas turbine engine as recited in claim 2 , wherein the gear system is coupled through an input shaft to a low pressure turbine claim 2 , the low pressure turbine ...

FLEXIBLE COUPLING FOR GEARED TURBINE ENGINE

A gas turbine engine includes a non-rotatable flexible coupling having Stiffnesses A, B, C, D and E, which are defined herein, and at least ratio of a frame lateral stiffness FLS as follows: FLS/Stiffness A in a range of about 6.0 to about 25.0, FLS/Stiffness B in a range of about 10.0 to about 40.0, FLS/Stiffness C in a range of about 1.5 to about 7.0, FLS/Stiffness D in a range of about 0.25 to about 0.50, or FLS/Stiffness E in a range of about 6.0 to about 40.0. 1. A gas turbine engine comprising:a fan;a fan shaft arranged along an engine central axis;a frame supporting the fan shaft, the frame defining a frame lateral stiffness (FLS);an input shaft;a high pressure turbine and a low pressure turbine, the low pressure turbine having a pressure ratio of greater than 5;a gear system rotatably coupled to the fan shaft, the gear system having a gear reduction ratio that is greater than 2.3; anda non-rotatable flexible coupling at least partially supporting the gear system, the non-rotatable flexible coupling being subject to at least one type of motion, with respect to the engine central axis, selected from the group consisting of Motion I, Motion II, Motion III, Motion IV and combinations thereof, wherein, Motion I is parallel offset guided end motion, Motion II is cantilever beam free end motion, Motion III is angular misalignment no offset motion, and Motion IV is axial motion,the non-rotatable flexible coupling having Stiffnesses A, B, C, D and E, wherein Stiffness A is axial stiffness under Motion IV, Stiffness B is radial stiffness under Motion II, Stiffness C is radial stiffness under Motion I, Stiffness D is torsional stiffness under Motion I, and Stiffness E is angular stiffness under Motion III, and at least one ofa ratio of FLS/Stiffness A of the non-rotatable flexible coupling is in a range of about 6.0 to about 25.0,a ratio of FLS/Stiffness B of the non-rotatable flexible coupling is in a range of about 10.0 to about 40.0,a ratio of FLS/Stiffness C of the ...