СПОСОБ УПРАВЛЕНИЯ ГАЗОТУРБИННОЙ УСТАНОВКОЙ, СОДЕРЖАЩЕЙ ЭЛЕКТРИЧЕСКИЙ ДВИГАТЕЛЬ

Изобретение относится к способу управления газотурбинной установкой (Т), содержащему электрический двигатель (МЕ), образующий устройство подачи крутящего момента на вращающийся вал (22) высокого давления, при этом в рамках способа определяют заданное значение QCMD расхода топлива и заданное значение TRQCMD крутящего момента, направляемое на электрический двигатель (МЕ), при этом способ управления содержит: этап применения первой цепи регулирования топлива с целью определения заданного значения QCMD расхода топлива, этап применения второй цепи регулирования крутящего момента с целью определения заданного значения TRQCMD крутящего момента, включающий в себя: i) этап определения величины поправки ΔTRQCMD крутящего момента в зависимости от заданного значения скорости перехода NHTrajAccelCons, NHTrajDecelCons, и ii) этап определения заданного значения TRQCMD крутящего момента в зависимости от величины поправки ΔTRQCMD крутящего момента. Также представлены газотурбинная установка и электронный ...

ВОЗДУХОЗАБОРНИК ДВУХКОНТУРНОГО ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ

Воздухозаборник для капота турбореактивного двигателя с всасывающим патрубком спереди содержит верхнюю часть, нижнюю часть, внутреннюю (ближнюю к фюзеляжу самолета) боковую сторону (115int) и внешнюю боковую сторону (115ext), определяющие переднюю кромку (115) между внешней стенкой (111) капота и внутренней стенкой (113), образующей канал питания турбореактивного двигателя воздухом. Передняя кромка внешней боковой стенки (115ext) находится сзади по отношению к передней кромке внутренней боковой стороны (115int). Двухконтурный турбореактивный двигатель может оснащаться таким воздухозаборником. Изобретение позволяет уменьшить шум, воспринимаемый в кабине летательного аппарата, путем переноса конуса шума к внешней стороне. 2 н. и 5 з.п. ф-лы, 4 ил.

УЗЕЛ АВИАЦИОННОГО ДВИГАТЕЛЯ И АВИАЦИОННЫЙ ДВИГАТЕЛЬ

Узел авиационного двигателя для забора воздуха и выпуска центральной струи и струи обводного контура содержит цилиндрический центральный обтекатель, цилиндрическую гондолу, множество распорных элементов, основной и вспомогательный пилоны и множество направляющих лопаток на стороне выхода вентилятора. На внутренней периферийной поверхности стенки гондолы или на наружной периферийной поверхности стенки центрального обтекателя образована выступающая часть. Выступающая часть выступает внутрь или наружу в диаметральном направлении и проходит от каждой ориентированной в направлении вдоль окружности боковой поверхности по меньшей мере одного из элементов, включающих в себя вспомогательный пилон, распорные элементы и направляющие лопатки на стороне выхода вентилятора, к стороне выпуска. Форма выступающей части, если смотреть с внутренней или наружной стороны в диаметральном направлении, представляет собой обтекаемую форму, проходящую в направлении вала двигателя, и вершина выступающей части расположена ...

УСТРОЙСТВО РЕВЕРСИРОВАНИЯ ТЯГИ ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ СО СТВОРКАМИ, ОБРАЗУЮЩИМИ КОВШИ, СВЯЗАННЫЕ С ПЕРЕДНИМ ПО ПОТОКУ ПОДВИЖНЫМ ОБТЕКАТЕЛЕМ

Изобретение относится к авиадвигателестроению. Устройство реверсирования тяги двухконтурного турбореактивного двигателя содержит полые поворотные створки (3), интегрированные при функционировании устройства в режиме прямой тяги в наружный обтекатель гондолы двигателя и образующие при функционировании устройства в режиме реверсирования тяги препятствия отклонения газового потока, образуя ковши. Подвижная часть, образующая обтекатель (23), располагается при функционировании устройства в режиме прямой тяги между передней по потоку кромкой поворотной створки (3) и передней по потоку неподвижной конструкцией (6), перекрывая заднюю по потоку часть конструкции передней рамы (8). Этот передний по потоку обтекатель (23) шарнирно закреплен на неподвижной конструкции устройства и связан со средствами управления перемещениями. При функционировании устройства в режиме реверсирования тяги этот обтекатель открывает заднюю по потоку часть конструкции передней рамы (8), которая обладает аэродинамической ...

ТУРБОВЕНТИЛЯТОРНЫЙ ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ С РЕГУЛИРУЕМЫМИ ВЕНТИЛЯТОРНЫМИ ВЫХОДНЫМИ НАПРАВЛЯЮЩИМИ ЛОПАТКАМИ (ВАРИАНТЫ)

Турбовентиляторный газотурбинный двигатель содержит секцию переднего вентилятора, включающую в себя, по меньшей мере, один ряд отстоящих по периферии передних в продольном направлении лопаток ротора вентилятора, двигатель основного контура, расположенный сзади и ниже по потоку секции переднего вентилятора, обводной канал вентиляторов и выхлопной канал. Двигатель основного контура включает в себя последовательно расположенные ниже по потоку ведомый вентилятор, компрессор, камеру сгорания и турбину высокого давления, соединенную с возможностью привода с компрессором посредством вала двигателя основного контура. Обводной канал вентиляторов расположен ниже по потоку секции переднего вентилятора и находится радиально снаружи двигателя основного контура. Радиально внешний вход обводного канала вентиляторов расположен в продольном направлении между секцией переднего вентилятора и ведомым вентилятором основного контура. Выхлопной канал расположен по потоку ниже двигателя основного контура и сообщается ...

ДИСК РОТОРА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ (ВАРИАНТЫ)

Диск ротора газотурбинного двигателя имеет ряд кольцевых ступиц, окружающих центральную линию, и каждая из ступиц соединена с венцом диска перемычкой. Множество разнесенных по окружности пазов в форме ласточкина хвоста расположены в венце, проходя по окружности между стойками диска, проходя в осевом направлении от переднего конца до заднего конца венца и проходя в радиальном направлении внутрь от наружной поверхности венца диска. Проходящие по окружности кольцевые пазы для предотвращения разрыва проходят в радиальном направлении сквозь венец в пазы в форме ласточкина хвоста между каждыми двумя соседними перемычками. В показанном здесь типичном варианте осуществления изобретения пазы в форме ласточкина хвоста являются дугообразными пазами в форме ласточкина хвоста. Выступающий в осевом направлении вперед свес расположен радиально снаружи на каждой из стоек диска, и на радиально наружном углу каждой из стоек диска на части свеса расположен скос. Множество отверстий стоек проходит в осевом ...

СПОСОБ СОЗДАНИЯ ДВИЖУЩЕЙ СИЛЫ ДЛЯ ПЕРЕМЕЩЕНИЯ ЛЕТАТЕЛЬНОГО АППАРАТА И ТУРБОРЕАКТИВНЫЙ ДВИГАТЕЛЬ ДЛЯ ЕГО ОСУЩЕСТВЛЕНИЯ

Способ создания движущей силы для перемещения летательного аппарата включает ввод воздуха и создание азимутально и аксиально движущегося потока, его сжатие компрессором, нагрев потока, вывод струи со скоростью, большей азимутальной скорости лопастей турбины, ввод дополнительного объема воздуха. Сначала передают импульс горячего потока дополнительному объему воздуха. Затем создают вращающий момент турбины и потом переводят азимутальный момент потока в аксиальное усилие. В процессе сжатия поток направляют азимутально и радиально от оси, формируют азимутальный вихрь, выводят поток за кромкой лопаток компрессора, разворачивают поток азимутально и радиально к оси, формируют азимутальный вихрь и повторяют это более чем один раз. При этом турбореактивный двигатель включает выполненные на общем валу компрессор низкого давления, содержащий лопатки со структурой профилей с крыловыми элементами и турбину, камеру сгорания, выполненную азимутальной, венцы с лопатками до и после камеры сгорания. За камерой ...

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.

КАПОТ ТУРБИННОГО ДВИГАТЕЛЯ, СПОСОБНЫЙ НАКРЫВАТЬ КОНУС ВЕНТИЛЯТОРА

Изобретение относится к капоту (20) газотурбинного двигателя, способному накрывать конус (24) вентилятора. Упомянутый капот содержит крепежное средство (27, 32, 36), способное входить в зацепление с соединительным средством (28, 33, 39) упомянутого конуса (24), чтобы удерживать упомянутый капот (20) и упомянутый конус (24) скрепленными между собой. Соединительное средство (28, 33, 39) содержит множество пазов (28) и множество канавок (33). Множество пазов (28) расположены на основании (29) конуса (24). Упомянутые пазы (28) выполнены с возможностью размещения в них зубьев (27) капота (20). Множество канавок (33) открываются в пазы (28) и выполнены с возможностью размещения в них выступающих частей (32), выдающихся наружу из зубьев (27). Геометрические несплошности, обусловливающие аэродинамические возмущения, таким образом, устраняются. Не допускается отделения капота от конуса под воздействием центробежной силы при вращении конуса. 2 н. и 8 з.п. ф-лы, 14 ил.

КОМПОНОВКА РЕДУКТОРНОГО ТУРБОВЕНТИЛЯТОРНОГО ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Газотурбинный двигатель, как правило, содержит вентиляторную секцию, компрессорную секцию, секцию камеры сгорания и турбинную секцию. Для приведения в движение вентиляторной секции может быть использован редуктор, например, представляющий собой эпициклическую зубчатую передачу, так, чтобы обеспечить возможность вращения вентиляторной секции со скоростью, отличной от скорости вращения турбинной секции, и повысить суммарный тяговый КПД двигателя. В двигателях такой конструкции вал, приводимый в движение одной из турбинных секций, приводит в действие эпициклическую зубчатую передачу, которая вращает вентилятор со скоростью, отличной от скорости вращения турбинной секции, в результате чего скорости вращения как турбинной секции, так и вентиляторной секции могут быть приближены к оптимальным, что обеспечивает повышение рабочих характеристик и производительности за счет использования требуемых сочетаний раскрытых конструктивных особенностей различных компонентов описанного газотурбинного двигателя ...

ВИНТОВЕНТИЛЯТОРНЫЙ АВИАЦИОННЫЙ ДВИГАТЕЛЬ

Винтовентиляторный авиационный двигатель содержит турбокомпрессор с корпусом, компрессором, камерой сгорания, выход из которой соединен газовым трактом с турбиной, и двухступенчатый винтовентилятор. Одна ступень винтовентилятора соединена с компрессором через магнитную муфту, а другая ступень винтовентилятора соединена с первой ступенью через реверсивный редуктор. Магнитная муфта содержит полумуфту, установленную в компрессоре, например, на его рабочих лопатках, и ведомую полумуфту, установленную на корпусе турбокомпрессора. Ступени винтовентилятора размещены внутри обтекателя. Изобретение направлено на повышение КПД и надежности авиационного двигателя. 2 з.п. ф-лы, 3 ил.

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.

НОСОВОЙ ОБТЕКАТЕЛЬ ДЛЯ ТУРБОМАШИНЫ

Носовой обтекатель для турбомашины, такой как, в частности, турбореактивный двигатель, установлен на конце вала, несущего лопасти винта, и служит для управления углом наклона поступающего воздушного потока на основаниях лопастей винта. Носовой обтекатель установлен на упомянутом валу с помощью средств управляемого осевого смещения, тем самым обеспечивая изменение осевого положения обтекателя относительно лопастей винта. Носовой обтекатель является смещаемым по оси между передним положением для обеспечения защиты от засасывания твердых веществ и частиц и задним положением для оптимизирования углов наклона воздушного потока на основаниях лопастей винта. Изобретение позволяет оптимизировать угол наклона потока воздуха на основаниях лопастей винта на различных этапах полета, при этом обеспечивая эффективную защиту от засасывания внутрь льда и твердых объектов или частиц. 2 н. и 4 з.п. ф-лы, 1 ил.

ТРЕХКАСКАДНЫЙ ДВУХКОНТУРНЫЙ ТУРБОРЕАКТИВНЫЙ ДВИГАТЕЛЬ С ВЫСОКОЙ СТЕПЕНЬЮ ДВУХКОНТУРНОСТИ

Трехкаскадный двухконтурный турбореактивный двигатель с высокой степенью двухконтурности имеет передний и задний вентиляторы спереди промежуточного корпуса, который представляет собой внешнюю несущую решетку в потоке второго контура и внутреннюю несущую решетку в основном воздушном потоке. Вентиляторы имеют лопасти, которые тянутся в радиальном направлении наружу к корпусу вентилятора. Корпус вентилятора определяет внешнюю сторону воздушного потока второго контура. Турбореактивный двигатель также имеет компрессор низкого давления для сжатия воздуха, приходящего в канал для основного воздушного потока. Передний и задний вентиляторы вращаются в прямом направлении, и отдельно, посредством двух валов, которые являются соосными. Компрессор низкого давления размещен в осевом направлении между лопастями переднего вентилятора и лопастями заднего вентилятора, и включает в себя, по меньшей мере, одно кольцо лопастей ротора, тянущихся от периферии колеса, которое приводится в движение приводным валом ...

Двухконтурный турбореактивный двигатель

Изобретение относится к двигателестроению, в частности к выходным устройствам двухконтурного турбореактивного двигателя. Известный двухконтурный турбореактивный двигатель, содержащий компрессор низкого давления, компрессор высокого давления, камеру сгорания, турбину высокого давления и турбину низкого давления, канал наружного контура, канал внутреннего контура, смеситель и общие для обоих контуров форсажную камеру и сопло, по предложению выполнен в виде чередующихся по периметру каналов, образующих выходную полость наружного контура и выходную полость внутреннего контура, установлен за турбиной низкого давления, при этом выходная полость наружного контура сообщена с каналом наружного контура, а выходная полость внутреннего контура сообщена с каналом внутреннего контура, причем отношение их площадей в поперечной плоскости равно:где- площадь выходной полости наружного контура;- площадь выходной полости внутреннего контура. Применение данного изобретения позволяет при сохранении конструкции ...

АВИАЦИОННЫЙ ДВУХКОНТУРНЫЙ ТУРБОРЕАКТИВНЫЙ ДВИГАТЕЛЬ

Авиационный двухконтурный турбореактивный двигатель содержит вентилятор, компрессор высокого давления, камеру сгорания, турбины высокого и низкого давления, смеситель и общие для обоих контуров форсажную камеру и сопло. За первой ступенью компрессора высокого давления, обеспечивающей на взлетном режиме степень повышения полного давления не более πIст *=1,4…1,5, выполнен постоянно открытый кольцевой канал со спрямляющей решеткой, через который на всех режимах работы двигателя осуществляется перепуск части воздуха из-за ступени в спутный поток воздуха наружного контура за вентилятором. Изобретение повышает степень двухконтурности двигателя и повышает его экономичность на бесфорсажных режимах работы. 3 ил.

ВИНТОВЕНТИЛЯТОРНЫЙ АВИАЦИОННЫЙ ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ

Винтовентиляторный авиационный газотурбинный двигатель содержит турбокомпрессор с корпусом, компрессором, камерой сгорания, выход из которой соединен газовым трактом с турбиной, и винтовентилятор. Винтовентилятор соединен с турбиной через магнитную муфту. Магнитная муфта содержит ведущую полумуфту, установленную в турбине, например, на ее рабочих лопатках, и ведомую полумуфту, установленную на корпусе турбокомпрессора. Турбокомпрессор выполнен двухкаскадным с возможностью вращения каскадов в противоположные стороны. Винтовентилятор выполнен двухступенчатым и содержит переднюю и заднюю ступени, расположенные с возможностью вращения в противоположные стороны. Изобретение направлено на повышение КПД и надежности авиационного двигателя. 1 з.п. ф-лы, 3 ил.

Авиационный турбореактивный двигатель

Изобретение относится к авиационным турбореактивным двигателям. Предложен авиационный турбореактивный двигатель, содержащий воздухозаборник 1, в цилиндрическом корпусе 2 по оси которого установлен вал 3, последовательно соединенный с многоступенчатым компрессором 4 и турбиной 6, а также агрегат наддува, камеры сгорания 5, кинематическую связь с приводом запуска 9 и выходной аппарат 7. Агрегат наддува выполнен автономным, установлен в воздухозаборнике 1 по оси корпуса 2 посредством трех радиальных кронштейнов 10 через промежуток перед компрессором 4 и не соединен с валом 3, состоит из единой сборки: электродвигателя 11, редуктора и двухлопастного пропеллера 12. Электродвигатель 11 соединен с бортовой электрической сетью через орган 16 автономного ручного ступенчатого управления, совмещенный с органом 18 управления авиационным турбореактивным двигателем. Пространство между плоскостью вращения пропеллера 12, стенкой цилиндрического корпуса 2 и плоскостью первой ступени компрессора 4 является ...

АБРАЗИВНО ИЗНАШИВАЕМОЕ УСТРОЙСТВО, РАЗМЕЩАЕМОЕ НА КОЖУХЕ ВЕНТИЛЯТОРА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

... 1. Газотурбинный двигатель с осью вращения В, содержащий кожух (10) вентилятора и вентилятор (V) с подвижными лопатками (17), причем между внутренней поверхностью кожуха и свободными концами лопаток вентилятора предусмотрен зазор, подшипник вентилятора соединен с неподвижными частями газотурбинного двигателя при помощи связей, разрушаемых при воздействии определенной нагрузки на лопатки вентилятора, а ось вращения вентилятора совершает колебания вокруг оси вращения В собственно газотурбинного двигателя, отличающийся тем, что кожух вентилятора содержит слой термически формуемой пены (12), размещенный против свободных концов лопаток (17) вентилятора (V), приклеенный к внутренней поверхности кожуха и располагающийся на, по меньшей мере, части протяженности упомянутого зазора, причем указанный слой термически формуемой пены (12) частично покрыт слоем абразивно изнашиваемого материала (14), причем толщина слоя абразивно изнашиваемого материала (14) выбирается такой, что свободные концы лопаток ...

ТУРБОРЕАКТИВНЫЙ ДВИГАТЕЛЬ ДЛЯ КРЕПЛЕНИЯ НА ХВОСТОВОЙ ЧАСТИ ФЮЗЕЛЯЖА ЛЕТАТЕЛЬНОГО АППАРАТА В ВЕРХНЕМПОЛОЖЕНИИ

... 1. Турбореактивный двигатель (трд) для установки сверху хвостовой части фюзеляжа (1) летательного аппарата посредством по меньшей мере одной крепежной подвески ((28, 128), (28', 128')), содержащий вентилятор (11), переднюю часть (12) корпуса, заднюю часть (18) корпуса, вспомогательные устройства (22, 23, 24, 25), расположенные по внешней стороне передней части (12) корпуса, которая содержит точки ((31, 32, 33), (31', 32' 33')) крепления крепежной подвески (28, 28'), отличающийся тем, что точки крепления расположены так, чтобы обеспечить крепление трд (10) одинаково на любой из сторон фюзеляжа (1) летательного аппарата, причем вспомогательные устройства (22, 23, 24, 25) расположены на корпусе так, что они доступны снаружи фюзеляжа (1) независимо от стороны, на которой установлен трд, при этом точки ((31, 32, 33), (31', 32' 33')) крепления крепежной подвески (28, 28') на передней части (12) корпуса распределены на каждой стороне вертикальной плоскости (26), содержащей линию, проходящую через ...

КОМПРЕССОРНОЕ УСТРОЙСТВО ГАЗОВОЙ ТУРБИНЫ И КОРПУСНОЙ ЭЛЕМЕНТ КОМПРЕССОРА

... 1. Компрессорное устройство (1) газовой турбины, содержащее газовый канал (5), секцию (8) компрессора низкого давления и секцию (9) компрессора высокого давления, предназначенные для сжатия газа в этом канале, и корпусной элемент (14) компрессора, расположенный между секцией (8) компрессора низкого давления и секцией (9) компрессора высокого давления с возможностью пропуска газового потока через газовый канал и включающий группу радиально расположенных стоек (15, 16, 21, 24, 25), предназначенных для передачи нагрузки, по меньшей мере, одна из которых выполнена полой для размещения в ней вспомогательных компонентов, отличающееся тем, что стойки (15, 16, 21, 24, 25) имеют криволинейную форму, а корпусной элемент (14) компрессора расположен по потоку непосредственно за последним ротором (10) секции (8) компрессора низкого давления и выполнен с возможностью существенного изменения направления закрученного газового потока от этого ротора (10) с помощью группы указанных стоек (15, 16, 21, 24, ...

Одновальный турбореактивный двухконтурный вентиляторный двигатель

Двухконтурный турбореактивный вентиляторный двигатель малой тяги

ЦИЛИНДРИЧЕСКИЙ КОЖУХ И РЕАКТИВНЫЙ ДВИГАТЕЛЬ

Тело (5a) кожуха для кожуха (5) вентилятора в реактивном двигателе (3) образовано путем использования композитного материала пластика, армированного углеродным волокном. Металлические кольца (5b, 5c) соответствующим образом прикреплены к передней кромке и задней кромке тела (5a) кожуха. Металлические кольца (5b, 5c) электрически соединены между собой проводящим кабелем (5d). Как металлические кольца (5b, 5c), так и проводящий кабель (5d) составляют путь тока молнии. Достигается функционирование цилиндрического кожуха как пути тока молнии, когда в самолет ударила молния, даже когда он создан с использованием композитного материала с высоким электрическим сопротивлением, и реактивного двигателя, в котором тот используется. 2 н. и 4 з.п. ф-лы, 5 ил.

РОТОРНАЯ МАШИНА (ВАРИАНТЫ), ПРОКЛАДКА И МОДУЛЬ ДЛЯ РОТОРНОЙ МАШИНЫ

... 1. Роторная машина (10), содержащая в основном цилиндрический корпус (18), лопатку (22) вентилятора, установленную с возможностью вращения внутри упомянутого корпуса (18) вокруг продольной оси упомянутой машины, кольцевую прокладку (24), установленную в упомянутом корпусе (18) между вершиной упомянутой лопатки (22) вентилятора и упомянутым корпусом (18) с возможностью герметизации зазора между упомянутой вершиной лопатки (22) вентилятора и упомянутым корпусом (18), отличающаяся тем, что упомянутая прокладка (24) содержит щеточное уплотнение (34), расположенное по внутренней окружности упомянутого корпуса и имеющее щетинки (38), расположенные проходящими в радиальном направлении вовнутрь от упомянутого щеточного уплотнения (34), и фиксирующую мембрану (36), установленную по окружности упомянутого щеточного уплотнения (34) с препятствованием выступанию упомянутых щетинок (38) в основном в радиальном направлении от упомянутого щеточного уплотнения (34) и с возможностью освобождения упомянутых ...

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 ...

ЛОПАТКА (ВАРИАНТЫ) И РОТОР ГАЗОТУРБИННОГО ДВИГАТЕЛЯ, И ТРУБОВЕНТИЛЯТОРНЫЙ ДВИГАТЕЛЬ

... 1. Лопасть для турбомашины, имеющей осевую линию, содержащая корневую часть, расположенную под углом к указанной осевой линии, и перо, отходящее от указанной корневой части, причем указанная корневая часть непосредственно примыкает к указанному перу. 2. Лопасть по п.1, отличающаяся тем, что представляет собой лопасть вентилятора турбовентиляторного двигателя. 3. Лопасть по п.1 или 2, отличающаяся тем, что указанная корневая часть имеет плавный профиль. 4. Лопасть по любому из пп.1-3, отличающаяся тем, что указанная корневая часть имеет участок увеличенного размера. 5. Лопасть по п.4, отличающаяся тем, что указанному участку увеличенного размера придан елочный профиль или профиль "ласточкина хвоста". 6. Лопасть по любому из предыдущих пунктов, отличающаяся тем, что дополнительно содержит средство для предотвращения смещения назад лопасти, установленной на диск. 7. Лопасть для турбомашины, имеющей продольную ось, содержащая: корневую часть, ориентированную по направлению указанной оси, и ...

ТУРБОРЕАКТИВНЫЙ ДВИГАТЕЛЬ С ВЕНТИЛЯТОРОМ, ПРИКРЕПЛЕННЫМ К ПРИВОДНОМУ ВАЛУ, УДЕРЖИВАЕМОМУ ПЕРВЫМ И ВТОРЫМ ПОДШИПНИКАМИ

... 1. Турбореактивный двигатель, содержащий неподвижную конструкцию, ротор (2) вентилятора, прикрепленный к приводному валу (5), удерживаемому первым подшипником (6) и вторым подшипником (7), установленными на указанной конструкции и прикрепленными при помощи опорных частей (11, 19) для подшипников, причем первый подшипник (6) установлен на неподвижной конструкции турбореактивного двигателя при помощи устройства (13), которое допускает его отсоединение от неподвижной конструкции, отличающийся тем, что второй подшипник (7) установлен на удерживающей части (19) для подшипника при помощи соединения, действующего как шаровой шарнир (23, 24), и тем, что турбореактивный двигатель также содержит средство, допускающее смещения второго подшипника (7) в осевом направлении относительно неподвижной конструкции турбореактивного двигателя, если первый подшипник (6) должен быть отсоединен. 2. Турбореактивный двигатель по п.1, отличающийся тем, что второй подшипник (7) содержит внешнее опорное кольцо (15) ...

ТРЕХКАСКАДНЫЙ ДВУХКОНТУРНЫЙ ТУРБОРЕАКТИВНЫЙ ДВИГАТЕЛЬ С ВЫСОКОЙ СТЕПЕНЬЮ ДВУХКОНТУРНОСТИ

... 1. Трехкаскадный двухконтурный турбореактивный двигатель с высокой степенью двухконтурности, причем турбореактивный двигатель имеет передний вентилятор (3) и задний вентилятор (5) спереди промежуточного корпуса (2), который представляет собой внешнюю несущую решетку (30) в потоке (F2) второго контура и внутреннюю несущую решетку (31) в основном воздушном потоке (F1), причем вентиляторы имеют лопасти (10, 14), которые тянутся в радиальном направлении наружу к корпусу (12) вентилятора, а этот корпус вентилятора определяет внешнюю сторону воздушного потока (F2) второго контура, причем турбореактивный двигатель также имеет компрессор (7) низкого давления для сжатия воздуха, приходящего в канал (8) для основного воздушного потока (F1), причем упомянутый передний вентилятор (3) и упомянутый задний вентилятор (5) вращаются прямо, и отдельно, посредством двух валов (4, 6), которые являются соосными. 2. Турбореактивный двигатель по п.1, отличающийся тем, что внешняя решетка (48) имеет множество ...

АКТИВНАЯ СИСТЕМА ДЛЯ РАСШИРЕНИЯ ЗОНЫ ПОДАВЛЕНИЯ ВИХРЯ, СОЗДАВАЕМОГО ДВИГАТЕЛЕМ САМОЛЕТА

... 1. Активная система для расширения зоны подавления наземного вихря, создаваемого авиационным двигателем, содержащая а) пневмосистему, соединенную с источником сжатого воздуха, и б) по меньшей мере один блок с рабочим органом, в котором имеется по меньшей мере одно подвижное сопло, соединенное с пневмосистемой, через которую в него из пневмосистемы подается сжатый воздух, и связанное с имеющимся в пневмосистеме приводом, управляющим движением подвижного сопла и направляющим в участок, расположенный определенным образом относительно воздухозаборника двигателя, струю выходящего из сопла воздуха, которая разрушает структуру турбулентного течения наземного вихря и препятствует его всасыванию в двигатель. 2. Активная система по п.1, в которой источником сжатого воздуха служит компрессор авиационного двигателя. 3. Активная система по п.2, в которой в пневмосистеме имеются а) запорный клапан, соединенный с компрессором двигателя, и б) привод, который соединен с выходом запорного клапана. 4. Активная ...

"Авиационный двигатель для сверхзвуковых скоростей полета "Шарм"

Использование: теплознергетика, в частности авиационное двигателестроение. Сущность изобретения: модульное исполнение двигателя, (выполнение с еговинтовен- тиляторным контуром с приводом его непосредственно от установленного дизельного модуля, снабжение контура ожижения воздуха турбодетэндером, испарителем-конденсатором и форсунками впрыска жидкого воздуха в проточную часть компрессора. При этом у двигателя каждый модуль компрессора и турбины снабжены встроенными электродвигателями-электрогенераторами , а компрессор и дизельный модуль.также снабжены рубашками охлаждения . 2 з.п. ф-лы, 2 ил.

MEHRACHSIGES BLÄSERTRIEBWERK

Ventilatorgehäuse mit Rezirculationskanälen.

Planetengetriebe und Gasturbinentriebwerk

Es wird ein Planetengetriebe mit einem Planetenträger, mit wenigstens einem auf dem Planetenträger drehbar angeordneten Planetenrad und mit wenigstens einem mit dem Planetenrad kämmenden Zahnrad sowie ein Gasturbinentriebwerk mit einem derartigen Planetengetriebe beschrieben. Der Planetenträger ist mit einer Ölzuführeinrichtung (42) ausgebildet ist, die eine Zuführleitung für Öl (78) zu wenigstens einer Öffnung (48A, 48B) für das zugeführte Öl (78) umfasst. Das Öl (78) ist aus der Öffnung (48A, 48B) zum Kühlen und Schmieren in Richtung des Planetenrades und/oder des Zahnrades ausleitbar. Die Ölzuführeinrichtung (42) umfasst in Bezug auf eine Hauptdrehrichtung (DR28) des Planetenrades und/oder des Zahnrades vor der Öffnung (48A, 48B) wenigstens einen sich von einer Außenseite (44) der Ölzuführeinrichtung (42) vorkragenden Abschirmbereich (46A, 46B), der mit der Außenseite (44) der Ölzuführeinrichtung (42) auf einer der Hauptdrehrichtung des Planetenrades und/oder des Zahnrades zugewandten ...

Kühlsystem für eine Gasturbine

Verstell-Luftschraube

Flugzeugtriebwerk und Verfahren zur Kupplung oder Entkupplung einer Kupplungsvorrichtung in einem Gasturbinentriebwerk

Die Erfindung betrifft ein Gasturbinentriebwerk (10) für ein Luftfahrzeug, umfassend: ein Kerntriebwerk (11), das eine Turbine (19), einen Verdichter (14) und eine die Turbine mit dem Verdichter verbindende Kernwelle (26) umfasst;einen Fan (23), der stromaufwärts des Kerntriebwerks (11) positioniert ist, wobei der Fan (23) mehrere Fanschaufeln umfasst; undein Planetengetriebe (30), das von der Kernwelle (26) antreibbar ist, wobei der Fan (23) mittels des Planetengetriebes (30) mit einer niedrigeren Drehzahl als die Kernwelle (26) antreibbar ist,dadurch gekennzeichnet, dassein feststehendes Teil (38, 55) des Planetengetriebes (30) über eine Kupplungsvorrichtung (50) mit einem statischen Bauteil (51) des Gasturbinentriebwerks (10) schaltbar verbunden ist und eine Ölkammer (52) zur Versorgung einer Schaltung der Kupplungsvorrichtung (50) mit dem Ölsystem des Gasturbinentriebwerks (10) verbunden ist, wobei der Öldruck (p) in der Ölkammer (52) so einstellbar ist, dass nominelle Reaktionsdrehmomente ...

Beheizung einer internen Oberfläche eines Bläsertriebwerks zur Verhinderung von Eisbildung

Gepfeilte Turbomaschinenschaufel

Kompressor und Turbinenstrahltriebwerk

Abstandshalter für Schaufel und Rotorscheibenanordnung mit solchem Abstandshalter

Gasturbinentriebwerk für ein Luftfahrzeug

Die Erfindung betrifft ein Gasturbinentriebwerk für ein Luftfahrzeug, das einen Triebwerkskern (11), einen Fan (23) und ein Getriebemodul (300) aufweist. Das Getriebemodul (300) umfasst einen Getrieberaum (7), in dem eine Öl/Luft-Atmosphäre vorliegt, wobei der Getrieberaum (7) gegenüber der Umgebung des Getriebemoduls (300) über mindestens eine Dichtung (101, 102) abgedichtet ist, und ein im Getrieberaum (7) angeordnetes Planetengetriebe, das einen Eingang von der Turbinenwelle (26) empfängt und Antrieb für den Fan (23) zum Antreiben des Fans (23) mit einer niedrigeren Drehzahl als die Turbinenwelle (26) abgibt. Es ist vorgesehen, dass die axial hinterste Dichtung (101) des Getriebemoduls (300) derart axial positioniert ist, dass sie axial vor der oder in der durch die axial vorderste Verdichterscheibe (950) gebildeten Ebene (A2) des Verdichters (90) angeordnet ist.

Gasturbinentriebwerk mit einer Schubdüse mit veränderlichem Querschnitt mit Betätigung durch Formgedächtnislegierungen

Getriebe und Gasturbinentriebwerk

Es wird ein Getriebe (30) mit einem drehbar gelagerten Bauteil (34) beschrieben, das mit wenigstens zwei annähernd rotationssymmetrischen Rinnen (41, 141) ausgebildet ist, in die jeweils ausgehend von ihrem radial inneren Bereich (42, 142) Öl aus jeweils einer gehäusefesten Ölzuführung (44, 144) einleitbar ist. Die Rinnen (41, 141) weisen jeweils in wenigstens einem radial äußeren Bereich (45, 145) jeweils wenigstens eine Auslassöffnung (46, 146) für das Öl auf. Des Weiteren ist das Öl von den Auslassöffnungen (46, 146) jeweils zu wenigstens einem hydraulischen Verbraucher über jeweils wenigstens einen Leitungsbereich (47, 147) führbar. Darüber hinaus wird ein Gasturbinentriebwerk mit dem Getriebe (30) vorgeschlagen.

Gasturbine mit gegenläufigen Niederdruckrotoren

GASTURBINENTRIEBWERK FUER UNTERSCHALLFLUG

Improvements in or relating to jet propulsion units

... 654,411. Jet propulsion plant. SOC. RATEAU, and ANXIONNAZ, R. Feb. 3, 1948, No. 3155. Convention date, Feb. 6, 1947. [Class 110(iii)] In a jet propulsion unit wherein a gas turbine drives two compressors a, k mechanically interconnected and arranged one behind the other on the same axis, the compressor k delivering air to the combustion chamber h supplying the turbine t, and the compressor a delivering dilution air which is mixed with the exhaust gases from the turbine, the delivery channels from the compressor a cross the suction channels of the compressor k, the channels being streamlined where they cross. The air inlet of the compressor a is near the common axis while the air inlet of the compressor k is arranged outside the compressor a. The compressors may be centrifugal, axial-flow or centrifugal-helical type.

LOW NOISE DUCTED FAN ASSEMBLY

... 1386481 Ducted fan engines UNITED AIRCRAFT CORP 29 Aug 1972 [10 Feb 1972] 39933/72 Headings F1C and F1J A ducted fan assembly for a gas turbine engine comprises a fan rotor having from 6 to 13 blades and a subsonic tip speed, and a stator disposed in the fan duct downstream of the rotor and having a smaller number of vanes than the number of rotor blades, the rotor, stator and duct being constructed to generate a noise spectrum such that the fundamental blade passage frequency and its predominant multiples are less than 2,500 Hz. Preferably, the number of stator vanes is substantially half the number of rotor blades, i.e., 3 to 7; this suppresses the stator-dominated noise frequencies. The low rotor blade count and low tip speed, which necessitate a low-pressure ratio of between 1.05 and 1.30, reduce the frequencies of the critical acoustic modes of the fan to below 2,500 Hz, thereby avoiding the range of frequencies to which the human ear is most sensitive, viz., 2,500 to 5,500 Hz. Noise ...

COMPRESSOR INTAKES

... 1417154 Gas turbine ducted fan engine: compressor air intake ROLLS-ROYCE (1971) Ltd 18 Oct 1973 [3 Nov 1972] 50677/72 Headings F1C, F1G and F1J The compressor air intake 16, for a ducted fan gas turbine, includes an annular opening 38 in the compressor casing, and a splitter ring 44 adjacent the opening deflecting air into the opening but allowing debris to pass on out of a rear nozzle (22), Fig. 1 (not shown). The splitter ring is supported by a row of stator blades 50, struts 54, and stator blades 56.

An exhaust assembly for an aircraft engine.

An exhaust assembly is provided for a gas turbine engine effective for propelling an aircraft from takeoff through subsonic and supersonic velocity. The exhaust assembly is effective for receiving exhaust gases discharged from an outlet of a core engine of an aircraft gas turbine engine. The assembly includes a casing, a variable area converging-diverging nozzle attached to the casing and including a first throat and an outlet for channeling exhaust gases received from the core engine. A plurality of retractable chutes are disposed upstream of the nozzle outlet and are positionable in a deployed position forming a converging nozzle having a second throat with a flow area less than that of the first throat. Means are provided for channeling air along aft facing surfaces of the chutes into the CD nozzle for mixing with the exhaust gases when the chutes are disposed in the deployed position for reducing noise from the exhaust gases.

AIRCRAFT ENGINE BLEED SYSTEM.

A gas turbine engine auxiliary compressor 24, for supplying air for de-icing, or air conditioning 70, in an aircraft is driven from a rotor of the gas turbine engine, via a variable speed drive 36, for operating the auxiliary compressor independently of the aircraft gas turbine engine compressor. Means 21 may be provided for bleeding boundary layer air oil a nacelle 10 or another part of the aircraft outer skin and feeding it to the auxiliary compressor. An air turbine 56 may be provided on a common shaft (58) with the auxiliary compressor and a valve 83 provided to direct an unused portion of the airflow from the auxiliary compressor to the air turbine to help power the auxiliary compressor. Compressed air from a ground supply 82, an on board unit 84, or another engine 88 may be supplied via valve 75 to the air turbine for on ground and in flight starting of the gas turbine engine through the variable speed drive. Alternative arrangements of the auxiliary components and alternative compressor ...

DUCTED FAN JET ENGINE

... 1341295 Gas turbine ducted fan engines H E FAHLBUSCH 19 April 1971 [18 May 1970] 22178/71 Heading F1J A subsonic multiple-spool type ducted fan jet engine has two separate inlets, one inlet serving a multi-stage compressor and the other serving a fan and comprises, an engine case containing compressor means 9, combustion means, turbine means and a primary gas nozzle 19, a rotatable multi-blade variable pitch fan 2 located directly in front of the compressor means and connected by a hollow shaft 21 to a turbine of the turbine means in the engine case, a circular section compressor inlet 3 which forms a primary flow path and which is connected with a tubular part of said fan and rotatable therewith, a mechanism 8 for varying the fan blade pitch angle, a co-axial duct 18 surrounding said fan and engine case and mounted by means of a fixed co-axial tubular fan case 7 and fixed radial stator blades 6 and consisting of a co-axial circular fan inlet located in front of said fan and which forms ...

TURBOFAN MIXED FLOW EXHAUST SYSTEM

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.

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.

GAS TURBINE FAN ENGINE

... 1275062 Gas turbine ducted fan engines ROLLS-ROYCE Ltd 7 July 1969 [11 July 1968] 33188/68 Heading F1J A gas turbine ducted fan engine comprises in flow series a plurality of compressors, combustion equipment and a plurality of turbines which drive the compressors, at least one of the compressors producing compression of the air supplied to the combustion equipment and at least another of the compressors having a single stage radially inner portion whose blading is disposed in the main flow duct and produces substantially no compression of the said air, the radially outer portion of the said blading constituting the fan of the engine and being disposed in a fan duct which surrounds the main flow duct. The engine shown comprises LP compressor 30, IP compressor 13, HP compressor 14, combustion equipment 15, HP turbine 16, IP turbine 20 and LP turbine 22 all disposed in flow series in the main flow duct 11. The HP, IP and LP turbines drive the respective compressors through shafts 17, 21 and ...

Improved gas turbine engine

... 1,125,251. Gas turbine engines; gas turbine jet engines. WILLIAMS RESEARCH CORP. Sept. 1, 1966 [Sept. 22, 1965], No. 39156/66. Headings F1G and F1J. A gas turbine engine comprises compressor means which comprises a low-pressure axial-flow compressor and a high-pressure centrifugal compressor, the L.P. compressor being driven by an L.P. turbine, and the H.P. compressor being driven by an H.P. turbine, the turbines and the associated compressors comprising independently rotatable spools. The engine shown in Fig. 1 is a gas turbine by-pass type jet engine and comprises an axial-flow L.P. compressor 16 and a centrifugal H.P. compressor 18, annular combustion equipment 22, an H.P. turbine 24 connected to drive the H.P. compressor through shaft 56, and an L.P. turbine 26, 28 connected to drive the L.P. compressor through shaft 33. The L.P. compressor comprises four stages 34, 35, 36, 37 some of the air delivered by the first two stages 34, 35 passing on to the second two stages 36, 37, the remainder ...

MULTI-SHAFT TURBOJET ENGINE

... 1351000 Jet deflector MOTOREN-UND TURBINEN-UNION MUNCHEN GmbH 9 July 1971 [25 July 1970] 32470/71 Heading B7G A multishaft turbojet engine, suitable for V/STOL aircraft, comprises a front fan 4, a bypass duct 14 supplied with air from the front fan, and a plurality of jet deflectors S supplied with air from the by-pass duct, each deflector comprising a plurality of pivotally-connected elements 26, 27. When the vanes S are straight (full lines, Fig. 1) the air continues undeflected (arrow F), as for cruising flight, while when they are in the position shown in dotted lines the air is defected (arrow V) to provide vertical lift, and at intermediate positions deflection (arrow K) aiding short take-off is afforded. Two sets of vanes S may be located in ducts 18, 19 on either side of central engine casing 15. The vanes may be operated jointly with mutually rotatable tubular sections 291, 30 and 31 (the latter having a variable nozzle 32), which constitute an exhaust jet deflector assembly ...

GAS TURBINE POWER PLANT

TURBOFAN GAS TURBINE ENGINES

Gas turbine jet-propulsion engine of the by-pass type

... 851.153. Gas turbine jet propulsion engines. ROLLS-ROYCE Ltd. Dec. 9, 1958 [March 14, 1958], No. 8262/58. Class 110 (3). In a gas turbine jet propulsion plant containing a compressor, combustion apparatus and a turbine, having a main conduit, a by-pass duct branching from the main conduit on the upstream side of the combustion chambers so as to by-pass the combustion chamber and the turbine, the by-pass duct 19 comprises two spaced apart by-pass pipes which are arranged to receive the whole of the by-pass air, the bypass pipes extending longitudinally of the main jet pipe 16 and terminating short of the downstream end thereof. The by-pass pipes are of arcuate shape in section and are arranged to fit closely about the main jet pipe. As shown the by-pass pipes are supplied with air from a compressor 11 which is driven by a low pressure turbine 15, and which also supplies air to a high pressure compressor 12, driven by a high pressure turbine 14. The by-pass pipes are provided with means 23 ...

HALF OF A HOLLOW FAN BLADE, HOLLOW FAN BLADE FOR GAS-TURBINE ENGINE, AS WELL AS APPROPRIATE MANUFACTURING PROCESS

BOUNDARY LAYER AUGUST ARRANGEMENT

PROCEDURE FOR THE REDUCTION OF NOISE EMISSIONS ON THE BACK OF A TURBO ENGINE AND IN THIS WAY IMPROVED TURBO ENGINE

SENTENCE FOR BLOWER ENGINE NOISE STEAMING

FAN FOR GAS TURBINE UNIT

EMERGENCY DEVICE FOR RESTARTING A TURBOJET IN AUTOROTATION

Dispositif de secours au rallumage d'un turboréacteur en autorotation comportant une soufflante (2) entraînée par une turbine basse pression par l'intermédiaire d'un premier arbre (3), un compresseur entraîné par une turbine haute pression par l'intermédiaire d'un second arbre (4) disposé coaxialement par rapport au premier arbre, un différentiel (5) reliant les premier et second arbres en compensant leurs différences de vitesse de rotation lors du fonctionnement normal du turboréacteur, et un système de freinage (6) relié au différentiel afin de pouvoir ralentir ou bloquer celui-ci lorsque le turboréacteur s'éteint, de sorte que le premier arbre (3) entraîne alors le second arbre (4) pour que ce dernier atteigne un régime favorisant le rallumage du turboréacteur.

ROTATING INLET COWL FOR A TURBINE ENGINE, INCLUDING AN ECCENTRIC FRONT END

La présente invention se rapporte à un capot d'entrée tournant (30) pour tarbomachine, présentant un axe de rotation (34) et dont l'extrémité avant (44) est agencée de manière excentrée par rapport à cet axe de rotation (34). De plus, un cône avant (32) du capot est tronqué par une surface de troncature (70) définissant l'extrémité avant (44) du capot d'entrée.

TURBOFAN CASE AND METHOD OF MAKING

A casing for a gas turbine includes a fan case, an intermediate case and a gas generator case integrated with one another. In another aspect, the casing provides a construction including several aspects which improve structural efficiency, such as a semi-monocoque construction, improved strut design, etc. Improved load paths and means for transmitting loads in the engine case are also disclosed.

VARIABLE AREA BYPASS INJECTOR FOR A DOUBLE BYPASS VARIABLE CYCLE GAS TURBOFAN ENGINE

A double bypass, variable cycle gas turbofan engine is provided with a variable area bypass injector which varies the area through which the inner bypass stream is injected into the outer bypass stream. The bypass injector comprises an upstream variable area, double bypass selector valve, the position of which determines whether the engine will operate in the single or double bypass mode, and a downstream static pressure valve which defines an optimum flow path for injecting the inner bypass stream into the outer bypass stream. The ability to vary the area through which the inner bypass stream is injected into the outer bypass stream permits the static pressure valve to operate as an ejector and permits control of the two bypass air streams through creation of a static pressure balance at the exit of the variable area bypass injector. The ability to control the bypass operating mode and the static pressure balance at the confluence of the two bypass streams eliminates the necessity for ...

GAS TURBINE ENGINE FOR SUBSONIC FLIGHT

BLOWING CONDUIT FOR A TURBINE ENGINE

Conduite de soufflante pour une turbomachine, comprenant deux parois cylindriques coaxiales (60, 68), respectivement interne et externe, la paroi interne (68) étant constituée d'une ossature (70) sur laquelle sont fixés de manière amovible des panneaux (72), et la paroi externe (60) étant constituée d'une structure unitaire (62) de support comportant des ouvertures (64) fermées par des panneaux amovibles (66), les ouvertures (64) de la paroi externe ayant des dimensions permettant le passage des panneaux (72) de la paroi interne et le montage et le démontage de ces panneaux sur l'ossature (70) de la paroi interne.

TURBOFAN ENGINE WITH FAN DUCT FLOW DEFLECTOR

SYSTEM AND METHOD FOR DETECTING A SHAFT EVENT ON AN ENGINE

Methods and systems for detecting a shaft event of a gas turbine engine are described. The method comprises monitoring at least one engine parameter and comparing the at least one engine parameter to a schedule for the at least one parameter defining a first threshold and a second threshold greater than the first threshold; applying a limit to the at least one engine parameter when the at least one engine parameter is inside a parameter limiting region between the first threshold and the second threshold, the first threshold separating the parameter limiting region from a normal operating region, the second threshold separating the parameter limiting region from a hazardous operating region; and detecting the shaft event when the at least one engine parameter crosses the second threshold and issuing a signal in response to the detecting.

PROPELLER BLADE ANGLE FEEDBACK ARRANGEMENT AND METHOD

A blade angle feedback ring assembly for an aircraft engine propeller having adjustable angle blades is provided. The feedback ring assembly comprises a feedback ring coupled to rotate with the propeller, the engine configured such that a first electrically-conductive path is defined between the propeller and the engine via the feedback ring, and an electric current conduction element provided between the propeller and the engine to define a second electrically-conductive path between the propeller and the engine in parallel to the first path, the second electrically-conductive path configured with a lower electrical resistance to conduction between the propeller and the engine than the first electrically-conductive path.

AIRCRAFT GAS TURBINE ENGINE VARIABLE FAN BLADE MECHANISM

A variable pitch fan assembly includes variable pitch fan blades circumscribed about engine centerline axis coupled to a drive shaft centered about the engine centerline axis. Each blade pivotable about pitch axis perpendicular to centerline axis and having blade turning lever connected thereto. One or more linear actuators non-rotatably mounted parallel to engine centerline axis and operably linked to fan blades for pivoting fan blades and connected to spider ring through thrust bearings for transmission of axial displacement of non-rotatable actuator rods of actuators while the fan blades are rotating. Spider arms extending away from spider ring towards blade roots and each spider arm connected to one of the turning levers. Turning levers may be connected and cammed to spider arms by pin and slot joint. Each spider arm may include joint pin disposed through joint slot of turning lever. Joint slot may be angled or curved.

METHODS AND APPARATUS FOR ASSEMBLING GAS TURBINE ENGINES

A method enables a gas turbine engine (10) to be assembled. The method comprises forming at least one substantially elliptically-shaped opening (80) within a flange (72) extending from a fan disk (40), inserting a fastener (120) including a first body portion (122), a second body portion (124), and an anti-rotation stop (126) extending therebetween, at least partially through the at least one flange opening, and coupling the fastener to the flange such that the fastener stop is positioned within the at least one flange opening.

AERODYNAMIC ADAPTATION OF THE REAR FAN OF A DOUBLE FAN TURBOFAN

L'invention concerne l'adaptation aérodynamique de la soufflante arrière (3) d'un turboréacteur ayant deux soufflantes (3, 5) à l'avant du carter intermédiaire (2) entraînées par deux arbres indépendants (4, 6) et un compresseur à basse pression (7). Ce compresseur (7) est disposé entre les pales (10, 14) des deux soufflantes et comporte au moins une couronne d'aubes mobiles (40) à la périphérie d'une roue (41) entraînée en rotation par l'arbre (4) de la soufflante avant (3) et au moins deux couronnes d'aubes fixes (45, 46) disposées de part et d'autre de la couronne d'aubes mobiles (40) et à l'intérieur d'un anneau porte grille (47). Une grille extérieure fixe (48) ainsi qu'un stator à calage variable (50) relient l'anneau (47) au carter de soufflante (12).

FAN ROTATING BLADE FOR TURBOFAN ENGINE

... [PROBLEMS] To provide a moving blade of a turbofan engine in which fuel cost and noise can be lowered, and the weight of the engine can be reduced by increasing a suction air amount to increase the bypass rate without increasing the diameter of the fan and the inner diameter of the casing. [MEANS FOR SOLVING THE PROBLEMS] The leading edge part (11) of the moving blade (10) for intake of air comprises a vertical hub part (12) positioned on the hub side and generally vertical to a fan rotation axis, a rearward inclined mid-span part (13) inclined to the downstream side from the hub side toward the mid-span part, and a forward inclined tip part (14) inclined to the upstream side from the mid-span side toward a tip part.

DUAL FLOW TURBINE ENGINE EQUIPPED WITH A PRECOOLER

Selon l'invention, le prérefroidisseur (30) présente une forme à section annulaire autour de l'axe (L-L) de la nacelle et est disposé à l'intérieur de la partie arrière (10R) du carénage interne (10) en contact externe avec le flux froid (9) sortant du canal de soufflante (13) et en contact interne avec un courant d'air de refroidissement (29) prélevé sur ledit flux froid (9) .

HOLDING STRUCTURE OF FAN BLADE

A disk (10) comprises a plurality of dovetail grooves (12) spaced at angular intervals in the circumferential direction and extending in the axial direction. A fan blade (20) has a dovetail part (22) axially fitted into each of the dovetail grooves and capable of transmitting a centrifugal force produced during rotation to the disk. The fan blade (20) includes a projecting part (24) positioned rearward of the dovetail part (22) and extending from the dovetail groove (12) in the circumferential outer direction. The projecting part (24) is set to such a size that the part is not interfered with the dovetail groove (12) when the dovetail part (22) is positioned on the radial inner side of the dovetail groove (12).

FAN BLADE RETAINING STRUCTURE

A structure for retaining a fan blade of a turbine includes a disk having a plurality of dovetail grooves spaced at predetermined angular intervals in a circumferential direction with the grooves extending in an axial direction. Fan blades each include a dovetail part which is fitted in one of the dovetail grooves in an axial direction and capable of transmitting a centrifugal force produced during rotation to the disk. The disk further includes a flange part which is positioned in front of the dovetail grooves and extends outwardly in a radial direction. An integral ring-shape retainer member is fitted between a front surface of the dovetail part and a rear surface of the flange part of the disk and transmits an axial forward load acting on the fan blades to the disk via the rear surface of the flange part. This structure avoids overloading of releasable securing fasteners such as securing bolts while providing a cost effective structure to manufacture and assemble.

METHOD OF MANUFACTURING FAN BLADE AND APPARATUS FOR MANUFACTURING THE SAME FAN BLADE

A raw material plate (7) comprising a plurality of main fibers (71) oriented in parallel, a plurality of auxiliary fibers (72) oriented in parallel and intersecting the main fibers (71), and a resin that integrates these main fibers (71) and auxiliary fibers (72) is sandwiched by a frame-shaped blank holder (8), in a state in which the resin is heated to a temperature at which same becomes soft. The raw material plate (7) sandwiched by the blank holder (8) is pressed on to a fan blade mold (6b) such that the direction of the main fibers (71) matches the longitudinal direction (Y) of the fan blade mold (6b), thereby applying a suitable tension to the raw material plate (7) along the main fiber (71) direction, and suppressing creases.

ELONGATED GEARED TURBOFAN WITH HIGH BYPASS RATIO

A propulsion system includes a fan, a gear, a turbine configured to drive the gear to, in turn, drive the fan. The turbine has an exit point, and a diameter (Dt) is defined at the exit point. A nacelle surrounds a core engine housing. The fan is configured to deliver air into a bypass duct defined between the nacelle and the core engine housing. A core engine exhaust nozzle is provided downstream of the exit point. A downstream most point of the core engine exhaust nozzle is defined at a distance from the exit point. A ratio of the distance to the diameter is greater than or equal to about 0.90.

FRONT ENCLOSURE WHICH IS SEALED DURING THE MODULAR DISMANTLING OF A TURBOJET WITH REDUCTION GEAR

Turboréacteur à double flux comportant une soufflante (S) entraînée, via un arbre de soufflante (3) porté par au moins deux premiers paliers (11, 12), par un arbre de turbine (4) porté par au moins un deuxième palier (10) comportant une bague fixe (25) et une bague mobile (26), ledit arbre de turbine entraînant ledit arbre de soufflante (3) au travers d'un dispositif de réduction de la vitesse de rotation (7), ledit dispositif de réduction de la vitesse de rotation et lesdits premiers et deuxième paliers étant logés dans une enceinte de lubrification (E1) dont l'enveloppe comprend des parties fixes et des parties mobiles reliées les unes aux autres par des moyens d'étanchéité (29, 30, 31), ledit dispositif de réduction comportant une roue d'entrée (27) conformée pour recevoir le couple transmis par ledit arbre de turbine par l'intermédiaire de moyens d'entraînement (8, 9) rattachés à ladite bague mobile,caractérisé en ce que l'enceinte de lubrification forme un anneau coaxial avec l'arbre ...

AIR INTAKE FOR TURBOPROP ENGINE

An air intake for efficiently channeling a flow of ambient air toward an air inlet of a turboprop or turboshaft gas turbine engine is disclosed. The air intake comprises an intake inlet for receiving the flow of air, an intake duct for channelling the flow of air, and an intake outlet for discharging the flow of air toward the air inlet of the gas turbine engine. The intake duct may be oriented toward a flow direction of the air pushed aft by a propeller coupled to the gas turbine engine.

METHOD AND APPARATUS FOR A STRUCTURAL OUTLET GUIDE VANE

An outlet guide vane (62,100) for a gas turbine engine is provided. The outlet guide vane (62,100) includes a first flange (64), a second flange (66) positioned radially outwardly from the first flange, an airfoil (102) extending between the first and second flanges and coupled thereto, the airfoil including a leading edge portion (130) and a trailing edge portion (132), the leading edge portion and the trailing edge portion fabricated from a first material, a filler portion (150) coupled in a gap formed by the leading edge portion and the trailing edge portion, the filler portion including a first side (203) and a second side (204), the filler portion fabricated from a second material, the second material different from the first material, a first skin coupled to the first side, the first skin fabricated from a composite material, and a second skin coupled to the second side, the second skin fabricated from a composite material.

MODULAR LOUVER SYSTEM

Louver systems for gas turbine bleed air systems are disclosed. An example louver system may include a bleed system discharge opening arranged to vent bleed air from a bleed flow conduit and a plurality of pivotable louvers disposed proximate the discharge opening, the pivotable louvers being pivotable between a shut position and an open position. In the shut position, individual louvers may at least partially obstruct the discharge opening. In the open position, individual louvers may at least partially control a direction of flow of the bleed air exiting the discharge opening.

TURBOFAN ENGINE

L'invention concerne un turboréacteur à double flux, comportant une roue de soufflante (1 ) portant des aubes (2) et entourée par un carter annulaire (3), le carter (3) comprenant des moyens d'aspiration d'air (9, 10) dans le jeu annulaire (8) formé entre le carter (3) et les extrémités radialement externes des aubes (2) de la roue de soufflante (1 ), les moyens d'aspiration comprenant une embouchure formée par au moins une fente d'entrée (10) réalisée dans la paroi interne (1 1 ) du carter (3) et reliée à un canal d'aspiration (9) s'étendant vers l'aval, caractérisé en ce que la fente d'entrée (10) des moyens d'aspiration est située axialement, en regard uniquement de la partie amont de la corde des aubes (2) de la roue de soufflante à leur extrémité radialement externe.

METHOD AND APPARATUS FOR WINDMILL STARTS

AIRCRAFT ENGINE DRIVESHAFT VESSEL ASSEMBLY AND METHOD OF ASSEMBLING THE SAME

A driveshaft vessel assembly for a gas turbine engine is disclosed. The driveshaft vessel assembly comprises a driveshaft vessel and an outlet guide vane. The driveshaft vessel includes a forward face, an aft face, and opposing side faces and is configured to house at least a portion of a radial driveshaft. The outlet guide vane includes a leading edge and a trailing edge. The length of the trailing edge is substantially equal to a length of the forward face of the driveshaft vessel such that the trailing edge of the outlet guide vane is faired into the forward face of the driveshaft vessel.

ROTOR BLADE AND FAN

In the present invention a rotor blade main body (37) is equipped with multiple first composite sheet groups (51) and multiple second composite sheet groups (53) provided along the thickness direction (TD). Each first composite sheet group (51) has multiple composite sheets (49) overlapping from the blade thickness center (TC) toward the dorsal surface (39). Each second composite sheet group (53) has the multiple composite sheets (49) overlapping from the blade thickness center (TC) toward the ventral surface (41). The combined direction (CD) of the orientation directions of the reinforcing fibers of the multiple composite sheets (49) in each sheet group (51, 53) is inclined 20-45 degrees toward the rear edge with respect to the span direction (SD).

GEARED TURBOFAN GAS TURBINE ENGINE ARCHITECTURE

A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine.

LIGHTWEIGHT ENGINE TURBINE BEARING SUPPORT ASSEMBLY FOR WITHSTANDING RADIAL AND AXIAL LOADS

... 13DV-9992 A gas turbine engine incorporates a bearing support assembly which mounts an inner annular bearing arrangement in radially spaced relation inwardly from a stationary outer casing. The bearing arrangement, in turn, mounts a turbine rotor within the outer casing for rotation about a central axis relative to the outer casing. A row of stationary stator vanes are fixedly attached to and extend radially inwardly from the outer casing toward the central axis. The bearing support assembly includes an inner ring structure extending about and mounting the inner bearing arrangement, elongated tie rods each having opposite outer and inner end portions and being disposed in circumferential spaced relation to one another and extending through the stationary vanes in radial relation to the central axis and between the outer casing and inner ring structure, and fastening elements fixedly clamping the outer casing and inner ring structure respectively to the outer and inner end portions of the ...

GAS TURBINE ENGINE COOLING SYSTEM

A gas turbine engine cooling system . A first turbocompressor and a heat exchanger are fluidly interconnected and are each in fluid communication to receive air of differing pressures and temperatures. Typically, such air is received from various regions of the engine low pressure compressor and the engine high pressure compressor. The system delivers air through a duct to a portion of the engine for cooling, such as the engine high pressure turbine region, at lower temperatures and higher pressures that if cooling air were directly ducted from the engine compressor to the engine turbine.

TURBOFAN ENGINE EXHAUST MIXING AREA MODIFICATION FOR IMPROVED NOISE REDUCTION

A modified exhaust tailpipe especially for use with an aircraft turbofan engine utilizing a noise suppressor. The modified tailpipe is connected to the exhaust duct of the engine, The forward end of the tailpipe is essentially the same diameter as that of the exhaust duct of the engine. The contour of the tailpipe is such that it increases in diameter to form a bulge to increase the flow area for fan gases in the area of a mixer which is supported within the forward end of the tailpipe exhaust system. Th is improves the operating efficiency of the bet engine.

THRUST REVERSER FOR A JET TURBINE ENGINE WITH DOORS FORMING SCOOPS ASSOCIATED WITH A MOVABLE DEFLECTOR

Un inverseur de poussée de turboréacteur comporte des portes creuses pivotantes (3), intégrées en jet direct dans le capotage extérieur de la nacelle et constituant en jet inversé des obstacles de déviation de flux en formant écopes. Un déflecteur mobile (24) est associé à chaque porte (3). Lors de l'ouverture de la porte (3), le déflecteur (24) se place au-dessus de la surface externe du panneau externe (4) de la porte (3) de manière à éviter l'interférence entre le flux d'échappement de la porte (3) et le flux dévié à l'extérieur de la porte (3).

Rückstossantrieb für Fahrzeuge, insbesondere Flugzeuge.

Verbrennungsturbinen-Strahlvortriebswerk an Flugzeugen.

Двухконтурная газотурбинная установка

Полезная модель содержит входное устройство, вентилятор, внутренний контур, внешний контур. Внутри внутреннего контура расположены компрессор среднего давления, компрессор высокого давления, камера сгорания, турбины. Внутри внешнего контура расположены воздуховоздушный теплообменник, выхлопные патрубки, соединяющие внутренний контур с атмосферой, свободная турбина, выходной канал. Воздух из компрессора среднего давления поступает в воздухвоздушный теплообменник и далее - в компрессор высокого давления. Воздуховоздушный теплообменник позволяет понизить температуру воздуха на входе в компрессор среднего давления и, соответственно, на входе в камеру сгорания, что позволяет иметь высокие степени повышения давления воздуха в двигателе (более 100) и, соответственно, высокий эффективный КПД (более 0,6). РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 209 432 U1 (51) МПК F02K 3/06 (2006.01) F02C 7/00 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ (52) СПК F02K 3/06 (2022.01); F02C 7/00 (2022.01) (21)(22) Заявка: 2021124272, 12.08.2021 (24) Дата начала отсчета срока действия патента: (73) Патентообладатель(и): Письменный Владимир Леонидович (RU) Дата регистрации: 16.03.2022 Приоритет(ы): (22) Дата подачи заявки: 12.08.2021 (45) Опубликовано: 16.03.2022 Бюл. № 8 2 0 9 4 3 2 R U (54) ДВУХКОНТУРНАЯ ГАЗОТУРБИННАЯ УСТАНОВКА (57) Реферат: Полезная модель содержит входное поступает в воздухвоздушный теплообменник и устройство, вентилятор, внутренний контур, далее - в компрессор высокого давления. внешний контур. Внутри внутреннего контура Воздуховоздушный теплообменник позволяет расположены компрессор среднего давления, понизить температуру воздуха на входе в компрессор высокого давления, камера сгорания, компрессор среднего давления и, соответственно, турбины. Внутри внешнего контура расположены на входе в камеру сгорания, что позволяет иметь воздуховоздушный теплообменник, выхлопные высокие степени повышения давления воздуха в ...

Cowl assembly

An assembly for a turbofan engine includes a first cowl member comprising an aft portion and a translatable cowl member comprising a forward portion configured to be received within the aft portion. The translatable cowl member is configured to be moveable with respect to the first cowl member between a first operational position wherein the forward portion is received within the aft portion of the first cowl member, and a second operational position wherein a smaller portion of the forward portion is received within the aft portion than in the first operational position. The translatable cowl member is configured to cooperate with a core cowl of the turbofan engine to define a portion of a fan duct having an exit nozzle, and the translatable cowl member is configured to define a flow control location near the exit nozzle that is associated with a controlling fan duct area.

Turbofan jet engine

There is a turbofan jet engine including an engine core. The engine core includes a fan and a compressor. The engine core includes a combustion chamber and a turbine functionally coupled to the compressor. The engine core includes a nozzle in fluid communication with the turbine. The turbofan jet engine includes a nacelle. The nacelle includes a forward extension proximate the fan and extending forward therefrom. The forward extension is funnel shaped to impart radial momentum to intake air during operation. The nacelle includes a vortex device disposed inside the forward extension and shaped to impart angular momentum to intake air. The vortex device includes a fixed blade extending from the interior of the forward extension and set at a rotational angle. The vortex device is shaped and positioned to direct intake air substantially perpendicular to the blades of the fan.

Variable area fan nozzle fan flutter management system

A system and method of controlling a fan blade flutter characteristic of a gas turbine engine includes adjusting a variable area fan nozzle in response to a neural network.

Variable area fan nozzle fan flutter management system

A gas turbine engine includes a controller that controls a fan blade flutter characteristic through control of a variable area fan nozzle.

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.

JET ENGINE

A reaction engine is disclosed mainly intended for aviation, but which as described herein can be adapted for an industrial engine, which makes use of spherical chambers and a pressure system in the blades of the rotor-stator unit which permits a perfect adjustment between said blades and the inner face of the stator, preventing pressure losses. While edges likewise articulated in the same way as the blades, execute a labyrinth seal. 121-. (canceled)22. Reaction engine comprising:a compressor block equipped with at least a compressor intended to carry out a compression of air that enters the engine and which comprises, in turn, a rotor and a stator,a combustion block which comprises at least a combustion chamber intended to house an ignition of a fuel together with high pressured air coming from the compressor block,at least a turbine actuated by exhaust gases produced in the combustion chamber and which comprises a torque transmission system with at least a shaft which is joined to at least the compressor thus carrying out the aforementioned compression of air,wherein the stator of the compressor is eccentric with respect to the rotor which permits the alternative radial displacement of an array of radial blades disposed on the rotor to carry out the closure in the radial direction of the array formed by the rotor and the stator, and said stator of the compressor comprises shock absorbers arranged on expulsion valves, said shock absorbers being controlled by electronic distribution rods disposed in compressor channels to control said shock absorbers and possible imbalances which may arise between the pressures.23. The engine of claim 22 , wherein the radial blades comprise elastic elements disposed on said blades and the rotor to provide the latter with automatic position recovery claim 22 , thus adjusting themselves on an inner face of the stator claim 22 , said elastic elements being located inside one of the lateral profiles of the blade.24. The engine of ...

Geared turbofan engine with counter-rotating shafts

A mid-turbine frame is incorporated into a turbine section of a gas turbine engine intermediate a high pressure turbine and a low pressure turbine. The high pressure and low pressure turbines rotate in opposite directions. A plurality of vanes redirect the flow downstream of the high pressure turbine as it approaches the low pressure turbine.

Pump system for tms aoc reduction

An engine includes a duct containing a flow of cool air and a pump system having an impeller with an inlet for receiving air from the duct and an outlet for discharging air into a discharge manifold. The discharge manifold containing at least one heat exchanger which forms part of a thermal management system.

Geared turbofan engine with power density range

A gas turbine engine turbine has a high pressure turbine configured to rotate with a high pressure compressor as a high pressure spool in a first direction about a central axis and a low pressure turbine configured to rotate with a low pressure compressor as a low pressure spool in the first direction about the central axis. A power density is greater than or equal to about 1.5 and less than or equal to about 5.5 lbf/cubic inches. A fan is connected to the low pressure spool via a speed changing mechanism and rotates in the first direction.

DEVICE FOR CONTROLLING A VARIABLE SECTION NOZZLE OF AN AIRCRAFT

A control device for controlling a variable section nozzle of an aircraft power plant, the variable section nozzle including one or several movable parts capable of modifying the nozzle section and connected by a mechanical transmission chain to an actuator. The control device includes a system for regulating the power plant connected to a control member configured to control the actuator. The control device includes a single control member, an immobilization unit configured to immobilize all the movable parts which are deactivated only when the regulation system controls the positional change of the movable part or parts and a determination unit configured to determine the actual position of the movable part or parts. 1. A control device for controlling a variable section nozzle of an aircraft power plant , said variable section nozzle comprising one or several movable parts configured to modify the nozzle section and connected by a mechanical transmission chain to an actuator , said control device comprising:a system configured to regulate the power plant connected to a control member configured to control the actuator, characterized in that the control device comprises a single said control member, an immobilization unit configured to immobilize all the movable parts which are deactivated only when the regulation system controls the positional change of the movable part or parts, and a determination unit configured to determine the actual position of the movable part or parts.2. The control device according to claim 1 , characterized in that the determination unit that determines the actual position of the movable part or parts is connected to the regulation system to indicate said actual position thereto.3. The control device according to claim 1 , further comprising a connection between the immobilization unit and the control member which controls their deactivation.4. The control device according to claim 1 , characterized in that the immobilization unit ...

Low Noise Compressor Rotor for Geared Turbofan Engine

A gas turbine engine has a fan, a compressor section having a low pressure portion and a high pressure portion, a combustor section, and a turbine having a low pressure portion. The low pressure turbine portion drives the low pressure compressor portion and the fan. A gear reduction effects a reduction in the speed of the fan relative to a speed of the low pressure turbine and the low pressure compressor portion. At least one of the low pressure turbine portion and low pressure compressor portion has a number of blades in each of a plurality of rows. The blades operate at least some of the time at a rotational speed. The number of blades and the rotational speed are such that the following formula holds true for at least one of the blade rows of the at least one of the low pressure turbine portion and/or the low pressure compressor sections: (number of blades×rotational speed)/60≧5500. The rotational speed is an approach speed in revolutions per minute. 1. A gas turbine engine comprising:a fan, a compressor section having a low pressure portion and a high pressure portion, a combustor section, and a turbine having a low pressure portion, the low pressure turbine portion driving said low pressure compressor portion and the fan;a gear reduction effecting a reduction in the speed of said fan relative to a speed of the low pressure turbine and the low pressure compressor portion; {'br': None, '(number of blades×rotational speed)/60≧5500; and'}, 'at least one of said low pressure turbine portion and said low pressure compressor portion having a number of blades in each of a plurality of rows, and said blades operating at least some of the time at a rotational speed, and said number of blades and said rotational speed being such that the following formula holds true for at least one of the blade rows of said at least one of the low pressure turbine portion and/or the low pressure compressor sectionssaid rotational speed being an approach speed in revolutions per minute.2. ...

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.

Gas turbine engine systems and related methods involving multiple gas turbine cores

Gas turbine engine systems and related methods involving multiple gas turbine cores are provided. In this regard, a representative gas turbine engine includes: an inlet; a blade assembly mounted to receive intake air via the inlet; and multiple gas turbine cores located downstream of the blade assembly, each of the multiple gas turbine cores being independently operative in a first state, in which rotational energy is provided to rotate the blade assembly, and a second state, in which rotational energy is not provided to rotate the blade assembly.

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.

Devices and Methods to Optimize Aircraft Power Plant and Aircraft Operations

Several improvements to optimize aircraft power plant and aircraft operations are disclosed, as well as methods of using these improvements to reduce fuel consumption, gas emission, noise, aircraft weight, maintenance costs, operating costs, aircraft incident and accidents, and improving aircraft performance. The improvements consist of a power plant fitted with a front propulsor, core engine, and aft propulsor. The fan of each propulsor is separated mechanically from the core engine. The front fan is separated mechanically from the aft fan. The aft fan is driven by free turbine that is supplied by exhaust gas of the core engine. If the core engine fails, both propulsors operate and provide thrust and reversed thrust when needed. If one propulsor fails, the other propulsor of the same power plant operates and provides thrust and reversed thrust. 1. An aircraft comprising an aircraft power plants and an Auxiliary Power Unit (APU); the aircraft power plant comprises a front propulsor , a core engine , and an aft propulsor; the aircraft power plant can be configured in several configurations; the power plant can be configured as an advanced dual fan and comprises front propulsor , core engine , and aft-propulsor wherein:a) The front propulsor comprises a front fan, a front motor-generator, and a front air turbine; the front fan is driven by a front motor-generator and an air turbine to provide the thrust or the reversed thrust when needed; the front motor-generator can be connected to the front fan through a clutch and the air turbine is may be connected to the front fan through a clutch; if the front motor-generator fails, this motor-generator can be disconnected from the front fan through its clutch, such that the air turbine drives the front fan; if the front air turbine fails this air turbine can be disconnected from the front fan through its clutch, such that the motor-generator drives the fan;The front fan is separated mechanically and aerodynamically from the ...

FAN STAGGER ANGLE FOR GEARED GAS TURBINE ENGINE

A gas turbine engine includes a spool, a turbine coupled with the spool, a propulsor coupled to be rotated about an axis by the turbine through the spool and a gear assembly coupled between the propulsor and the spool such that rotation of the spool results in rotation of the propulsor at a different speed than the spool. The propulsor includes a hub and a row of propulsor blades that extends from the hub. Each of the propulsor blades has a span between a root at the hub and a tip, and a chord between a leading edge and a trailing edge such that the chord forms a stagger angle α with the axis. The stagger angle α is less than 62° at all positions along the span, with said hub being at 0% of the span and the tip being at 100% of the span. 1. A gas turbine engine comprising:a spool;a turbine coupled with said spool;a propulsor coupled to be rotated about an axis through said spool; anda gear assembly coupled between said propulsor and said spool such that rotation of said spool results in rotation of said propulsor at a different speed than said spool,said propulsor including a hub and a row of propulsor blades extending from said hub, each of said propulsor blades having a span between a root at said hub and a tip, and a chord between a leading edge and a trailing edge such that said chord forms a stagger angle α with said axis, and said stagger angle α is less than 62° at all positions along said span, with said hub being at 0% of said span and said tip being at 100% of said span.2. The gas turbine engine as recited in claim 1 , wherein said stagger angle α at 25% of said span is less than 23°.3. The gas turbine engine as recited in claim 1 , wherein said stagger angle α at 25% of said span is 16-21°.4. The gas turbine engine as recited in claim 1 , wherein said stagger angle α at 50% of said span is less than 35°.5. The gas turbine engine as recited in claim 1 , wherein said stagger angle α at 50% of said span is 28-33°.6. The gas turbine engine as recited in claim ...

TRANSITION DUCT FOR USE IN A TURBINE ENGINE AND METHOD OF ASSEMBLY

A transition duct for use in a turbine engine is provided. The transition duct includes a radially inner wall and a radially outer wall positioned about the radially inner wall defining a flow passage therebetween. The radially outer wall extends and is contoured from an upstream end to a downstream end of the transition duct. As such, the slope of the radially outer wall increases from the upstream end to a predetermined axial location and decreases from the predetermined axial location to the downstream end. 1. A transition duct for use in a turbine engine , the transition duct comprising:a radially inner wall; anda radially outer wall positioned about said radially inner wall defining a flow passage therebetween, said radially outer wall extends and is contoured from an upstream end to a downstream end of the transition duct such that a slope of said radially outer wall increases from said upstream end to a predetermined axial location and decreases from the predetermined axial location to said downstream end.2. The transition duct in accordance with further comprising a fairing that extends radially between said radially inner wall and said radially outer wall within said flow passage claim 1 , wherein said fairing comprises an aerodynamic cross-sectional shape.3. The transition duct in accordance with claim 2 , wherein the predetermined axial location corresponds to an axial location of a thickest cross-sectional portion of said fairing such that a maximum slope of said radially outer wall is at the predetermined axial location.4. The transition duct in accordance with claim 1 , wherein the slope of said radially outer wall increases from about 0° at said upstream end to greater than about 40° at the predetermined axial location.5. The transition duct in accordance with claim 1 , wherein said radially outer wall comprises a maximum wall slope at the predetermined axial location claim 1 , the maximum wall slope from about 40° to about 50°.6. The transition duct ...

Geared Turbofan Engine With Increased Bypass Ratio and Compressor Ratio ...

A gas turbine engine is typically comprised of a fan stage, multiple compressor stages, and multiple turbine stages. These stages are made up of alternating rotating blade rows and static vane rows. The total number of blades and vanes is the airfoil count. An overall pressure ratio is greater than 30. A bypass ratio is greater than 8. A stage ratio is the product of the bypass ratio and the overall pressure ratio divided by the number of stages. An airfoil ratio is that product divided by the airfoil count. The stage ratio is greater than or equal to 22 and/or the airfoil ratio is greater than or equal to 0.12.

Rotational annular airscrew with integrated acoustic arrester

A propulsion system and methods are presented. A substantially tubular structure comprises a central axis through a longitudinal geometric center, and a first fan rotates around the central axis, and comprises a first fan hub and first fan blades. The fan hub is rotationally coupled to the substantially tubular structure, and the first fan blades are coupled to the first fan hub and increase in chord length with increasing distance from the first fan hub. A second fan is rotationally coupled to the substantially tubular structure and rotates around the central axis and contra-rotates relative to the first fan. Second fan blades are coupled to the second fan hub, and a nacelle circumscribing the first fan and the second fan is coupled to and rotates with the first fan.

MANUFACTURING ASSEMBLY AND METHOD

The present disclosure relates to an assembly for formation of a fan blade. The assembly comprises a suction panel; a pressure panel; and a membrane having a leading edge and a trailing edge. The membrane is sandwiched between the suction panel and pressure panel. The membrane comprises a gas entry slot extending in a radial direction, the gas entry slot having a radially outer receiving portion for receiving a pipe, and a radially inner portion. The radially inner portion of the gas entry slot has a substantially uniform width in a direction between the leading and trailing edge of the membrane. 1. A membrane for inclusion in an assembly for formation of a fan blade , the membrane having a leading edge and a trailing edge ,wherein the membrane comprises a gas entry slot extending in a radial direction, the gas entry slot having a radially outer receiving portion for receiving a pipe, and a radially inner portion wherein the radially inner portion of the gas entry slot has a substantially uniform width in a direction between the leading and trailing edge of the membrane.2. A membrane according to wherein the width of the radially inner portion of the gas entry slot in the direction between the leading and trailing edges of the membrane is less than 3 mm.3. A membrane according to wherein the width of the radially inner portion of the gas entry slot in the direction between the leading and trailing edges of the membrane is between 1.1 and 0.8 mm.5. A membrane according to wherein the radially inner portion of the gas entry slot comprises a hook portion where the gas entry slot deflects through greater than 90 degrees.6. A membrane according to wherein the gas entry slot deflects through substantially 135 degrees.7. A membrane according to wherein the gas entry slot comprises a meander portion radially outwards of the hook portion.8. An assembly for formation of a fan blade claim 1 , the assembly comprising:a suction panel;a pressure panel; and{'claim-ref': {'@idref': ...

AIRCRAFT PROPULSION SYSTEM WITH A LOW-FAN-PRESSURE-RATIO ENGINE IN A FORWARD OVER-WING-FLOW INSTALLATION, AND METHOD OF INSTALLING THE SAME

There is provided a propulsion system for an aircraft, the system having a low-fan-pressure-ratio engine configured to be mounted, in a forward over-wing-flow installation, to a wing of the aircraft. The engine has a core, a variable pitch fan, and a nacelle having a nacelle trailing edge with a top-most portion positioned above a wing leading edge. The engine has an L/D ratio of the nacelle in a range of from 0.6 to 1.0, and a fan-pressure-ratio in a range of from 1.10 to 1.30. The forward over-wing-flow installation enables, during all flight phases of the aircraft, a fan flow exhaust to flow behind the nacelle, and to be bifurcated by the wing leading edge, so the fan flow exhaust flows both over the wing and under the wing. During a cruise flight phase of the aircraft, the engine minimizes scrubbing drag of the fan flow exhaust to the wing. 1. A propulsion system for an aircraft , comprising: a core having a first end and a second end;', 'a variable pitch fan coupled to the first end of the core;', 'a nacelle surrounding the variable pitch fan and a portion of the core, the nacelle having a nacelle leading edge and a nacelle trailing edge, the nacelle trailing edge having a top-most portion configured to be positioned above a wing leading edge of the wing, and the nacelle configured to be positioned, in its entirety, at a forward location in front of the wing leading edge;', 'a length to diameter (L/D) ratio of the nacelle in a range of from 0.6 to 1.0; and', 'a fan-pressure-ratio in a range of from 1.10 to 1.30,, 'a low-fan-pressure-ratio engine configured to be mounted, in a forward over-wing-flow installation, to a wing of the aircraft, the low-fan-pressure-ratio engine comprisingwherein the forward over-wing-flow installation of the low-fan-pressure-ratio engine of the propulsion system enables, during all flight phases of the aircraft, a fan flow exhaust, exhausted by the variable pitch fan, to flow behind the nacelle, and to be bifurcated by the wing ...

SYSTEM AND METHOD OF OPERATING A DUCTED FAN PROPULSION SYSTEM DURING AIRCRAFT TAXI

A thrust reverser assembly and a method of operating an aircraft during a taxi mode of operation are provided. The thrust reverser assembly includes one or more actuator assemblies configured to modulate a position of a moveable portion over a continuous range of travel between a fully stowed position and a fully deployed position, such that an air flow through said thrust reverser bleed passage is correspondingly varied. The thrust reverser assembly also includes a throttle device that includes a first, ground idle power level position and a second, forward thrust mode position. Movement into the second position may be actuated separately and differently from movement into the first position. An actuator intermediate lock may inhibit actuation of the intermediate forward thrust mode of operation until a plurality of preconditions is met. 1. A thrust reverser assembly for an aircraft comprising:a moveable portion that is moveable over a continuous range of travel between a fully stowed position and a fully deployed position, wherein movement away from said fully stowed position opens a thrust reverser bleed passage;one or more actuator assemblies coupled to said moveable portion and operable in an intermediate forward thrust mode to modulate a position of said moveable portion along said continuous range of travel, such that an air flow through said thrust reverser bleed passage is correspondingly varied;a throttle device comprising a first position associated with a ground idle power level and a second position associated with the intermediate forward thrust mode, wherein movement of said throttle device into said second position is actuated separately and differently from movement of said throttle device into said first position; andan actuator intermediate lock coupled to said one or more actuator assemblies and configured to inhibit actuation of the intermediate forward thrust mode until a plurality of preconditions are met.2. The thrust reverser assembly of ...

TIP SHROUDED HIGH ASPECT RATIO COMPRESSOR STAGE

A gas turbine engine compressor stage includes a rotor. Compressor blades are supported by the rotor. The blades include an inner flow path surface each supporting an airfoil that has a chord that extends radially along a span to a tip. A shroud is supported at the tip and provides an outer flow path surface. The shroud provides a noncontiguous ring about the compressor stage. 1. A gas turbine engine compressor stage comprising:a rotor;compressor blades supported by the rotor, the blades include an inner flow path surface each supporting an airfoil that has a chord extending radially along a span to a tip, a shroud supported at the tip and providing an outer flow path surface, wherein the shroud provides a noncontiguous ring about the compressor stage.2. The compressor stage according to claim 1 , wherein the airfoils have an aspect ratio corresponding to the airfoil span to the airfoil chord claim 1 , the shroud extends the full chord claim 1 , wherein the aspect ratio is in a range of 1 to 5.3. The compressor stage according to claim 2 , wherein the aspect ratio is in a range of 1.4 to 3.0.4. The compressor stage according to claim 3 , wherein the aspect ratio is in a range of 1.6 to 2.8.5. The compressor stage according to claim 1 , wherein the shroud includes a sealing structure on a side of the shroud opposite of the outer flow path surface.6. The compressor stage according to claim 5 , wherein the sealing structure has labyrinth seals.7. The compressor stage according to claim 1 , wherein the blades each include a root received in a slot in the rotor.8. The compressor stage according to claim 1 , wherein shroud includes ring segments circumferentially spaced apart from one another and separated by gaps claim 1 , wherein multiple blades share a common ring segment.9. The compressor stage according to claim 1 , wherein blades are integrated with the rotor.10. A gas turbine engine comprising:engine static structure;first and second turbine sections rotatable ...

Guide vane of a gas turbine engine, in particular of an aircraft engine

A guide vane of a gas turbine engine, in particular of an aircraft engine, which has a pressure-side wall, a suction-side wall, a guide vane root, a guide vane tip, a guide vane leading edge area that is impinged by a cooling air flow of a cooling system, a guide vane trailing edge area that is facing away from the guide vane leading edge area, and at least one channel for conducting a fluid to be cooled arranged in an internal space of the guide vane. At that, during operation of the gas turbine engine, a first part of the cooling air flow flows around a pressure-side wall, and a second part of the cooling air flow flows around the suction-side wall, and a third part of the cooling air flow flows through the internal space including the channel. What is further suggested is a gas turbine engine with at least one such static guide vane.

DEVICE FOR RETAINING DRAINED FLUIDS FOR A PROPULSIVE ASSEMBLY

A device for retaining drained fluids for a propulsive assembly includes a cavity for storing the drained fluids and two walls mounted at the opening of said cavity. The cavity has a fluid storage volume V when the device is in a substantially vertical position, and each wall is configured such as to define a fluid storage volume (V and V respectively) in the cavity when the device is in a substantially horizontal position, each of the volumes V and V being at least equal to the volume V 1132231. Device for retaining drained liquids from a propulsion assembly , comprising a body defining a cavity that is intended for storing configured to store the drained liquids and has a volume V when the device is in a first position , for example substantially vertical , said cavity comprising an upper opening through which the liquids are conveyed into the cavity , wherein the cavity comprises two walls in the region of said opening that are at least in part positioned one above the other and define a space therebetween , a first wall designed to define a volume V for storing liquids in the cavity when the device is in a second position that is at a positive angle from the first position about a substantially horizontal axis , and a second wall designed to define a volume V for storing liquids in the cavity when the device is in a third position that is at a negative angle from the first position about a substantially horizontal axis , each volume V and V being at least equal to the volume V.2. (canceled)3. Device according to claim 1 , wherein the two walls are an upper wall and a lower wall claim 1 , the upper wall defining an orifice for introducing the liquids into said space.4. Device according to claim 3 , wherein said orifice is offset on one side from a vertical median plane (P) of the cavity.5. Device according to either claim 3 , wherein the lower wall extends below the orifice in the upper plate and defines a passage for liquids from the space to the cavity.6. ...

AUXILIARY OIL SYSTEM FOR GEARED GAS TURBINE ENGINE

A gas turbine engine comprises a fan drive turbine, a fan rotor, and a gear reduction driven by the fan drive turbine to, in turn, drive the gear architecture. A main oil supply system supplies oil to components within the gear reduction, and an auxiliary oil supply system. The auxiliary oil system operates to ensure that the gear reduction will be adequately supplied with lubricant for at least seconds at power should the main oil supply system fail. 1. A gas turbine engine comprising:a fan drive turbine, a fan rotor, and a gear reduction driven by said fan drive turbine to, in turn, drive said gear architecture, a main oil supply system for supplying oil to components within said gear reduction, and an auxiliary oil supply system; andsaid auxiliary oil system being operable to ensure that the gear reduction will be adequately supplied with lubricant for at least 30 seconds at power should the main oil supply system fail.2. The gas turbine engine as set forth in claim 1 , wherein said gear reduction includes a sun gear being driven by said fan drive turbine to drive intermediate gears that engage a ring gear.3. The gas turbine engine as set forth in claim 2 , wherein said sun gear claim 2 , said intermediate gears and said ring gear are enclosed in a bearing compartment claim 2 , which captures oil removed via a scavenge line connected to a main oil pump.4. The gas turbine engine as set forth in claim 3 , wherein said main oil pump has a gutter that directs scavenged oil to a main oil tank.5. The gas turbine engine as set forth in claim 4 , wherein oil in said main oil tank feeds a main pump pressure stage which then delivers oil to said gear reduction.6. The gas turbine engine as set forth in claim 5 , wherein oil from said main pump pressure stage passes through a lubrication system that includes at least one filter and at least one heat exchanger to cool the oil.7. The gas turbine engine as set forth in claim 4 , wherein said gear reduction is surrounded by an ...

CMC CORE COWL AND METHOD OF FABRICATING

A CMC core cowl for an aircraft gas turbine engine. The ceramic core cowl comprises an interlaced fiber structure having fibers oriented in substantially transverse directions, and a ceramic matrix surrounding the ceramic fiber structure. The core cowl further comprises several panels. The ceramic fiber and matrix are formed into a substantially cylindrical shape extending from a fore end at the fan outlet guide vanes to an aft end at the low pressure turbine outlet guide vanes. The CMC core cowl includes a means for mechanical attachment circumferentially oriented around the fore end and the aft end with mating parts. The CMC core cowl further includes additional plies oriented in a third preselected direction, thereby providing additional strength for mechanical attachment. 1. A ceramic matrix composite (CMC) core cowl for an aircraft gas turbine engine comprising: an interlaced fiber structure having ceramic fibers oriented in substantially transverse directions;', 'a ceramic matrix surrounding the ceramic fibers of the ceramic fiber structure;', 'wherein the ceramic fibers and matrix are formed into a substantially cylindrical shape having a fore end and an aft end, and having a mechanical attachment circumferentially oriented around the fore end and along the longitudinal lap joints; and', 'wherein the fore end and further includes additional CMC material having fibers oriented in a third preselected direction, thereby providing additional strength to for mechanical attachment at the fore end and at lap joints., 'a plurality of duct panels, each duct panel joined to an adjacent duct panel along a longitudinal lap joint, each duct panel further comprising;'}2. The CMC core cowl of further comprising:{'b': '160', 'a bifurcation opening () formed by a layup of CMC plies creating a duct boundary'}wherein the duct boundary forms a passageway in at least one of the duct panels.3. The CMC core cowl of wherein each of the plurality of duct panels is longitudinally ...

TURBOMACHINE FAN CLUTCH

A gas turbine engine assembly includes, among other things, a clutch configured to move from a first position to a second position in response to rotation of a gas turbine engine fan at a speed greater than a threshold speed. Whether the clutch is in the first position or the second position, the clutch permits rotation of the gas turbine engine fan in a first direction. When the clutch is in the first position, the clutch limits rotation of the gas turbine engine fan only in an opposite, second direction. The clutch is disposed within a compartment that is accessible and removable via removal of an aft engine cover structure. The clutch is removable on-wing. 1. A gas turbine engine assembly , comprising:a clutch configured to move from a first position to a second position in response to rotation of a gas turbine engine fan at a speed greater than a threshold speed,wherein, whether the clutch is in the first position or the second position, the clutch permits rotation of the gas turbine engine fan in a first direction, and when the clutch is in the first position, the clutch limits rotation of the gas turbine engine fan only in an opposite, second direction,wherein the clutch is disposed within a compartment that is accessible and removable via removal of an aft engine cover structure, whereby the clutch is removable on-wing.2. The gas turbine engine assembly of claim 1 , wherein the aft engine cover structure includes an engine exhaust cone.3. The gas turbine engine assembly of claim 2 , wherein the clutch is disposed within an aft bearing compartment and the aft engine cover structure further includes an aft bearing compartment cover plate claim 2 , disposed axially inward of the exhaust cone.4. The gas turbine engine assembly of claim 1 , wherein the clutch is positioned within a gas turbine engine such that the clutch can be moved from an installed position within the gas turbine engine to an uninstalled position without removing any blades from the gas turbine ...

ROTATING INLET COWL FOR A TURBINE ENGINE, COMPRISING AN ECCENTRIC FORWARD END

A rotating inlet cowl for a turbine engine includes a rotation axis. The rotating inlet cowl includes a forward cone defining a forward end of the inlet cowl. The forward end is configured to be eccentric relative to the rotation axis of the inlet cowl. Furthermore, the forward cone is truncated by a truncation surface defining the forward end of the inlet cowl. 1. A rotating inlet cowl of a gas turbine engine , the inlet cowl including a rotation axis and comprising:a forward cone defining a forward end of the inlet cowl,wherein the forward end is configured to be eccentric relative to the rotation axis of the inlet cowl,wherein the forward cone is truncated by a truncation surface defining the forward end of the inlet cowl, andwherein the forward cone includes an axis parallel to and coincident with the rotation axis of the inlet cowl.2. The rotating inlet cowl according to claim 1 , wherein the forward cone includes at least one balancing bead that includes a variable thickness along a circumferential direction claim 1 , to compensate for an unbalanced mass.3. A turbine or aircraft engine claim 1 , comprising the rotating inlet cowl according to .4. A rotating inlet cowl of a gas turbine engine claim 1 , the inlet cowl including a rotation axis and comprising:a forward cone defining a forward end of the inlet cowl,wherein the forward end is configured to be eccentric relative to the rotation axis of the inlet cowl,wherein the forward cone is truncated by a truncation surface defining the forward end of the inlet cowl, andwherein the truncation surface is approximately a plane that is inclined relative to a plane orthogonal to the rotation axis of the inlet cowl.5. The rotating inlet cowl according to claim 4 , wherein the forward cone includes at least one balancing bead that includes a variable thickness along a circumferential direction claim 4 , to compensate for an unbalanced mass.6. The rotating inlet cowl according to claim 4 , wherein the truncation ...

Asymmetric Fan Nozzle in High-BPR Separate-Flow Nacelle

A fan nozzle for an aircraft gas turbine engine is comprised of a core engine cowl that is disposed within a fan cowl so that an air flow area is defined therebetween. The core engine cowl and fan cowl are disposed around a horizontal central plane. The fan cowl has a substantially circular shape and is formed of an upper substantially semi-circular portion having a first radius and a lower substantially semi-circular portion having a second radius. The core engine cowl has a substantially circular shape and is formed of an upper substantially semi-circular portion having a third radius and a lower substantially semi-circular portion having a third radius. The upper substantially semi-circular portion of the core engine cowl includes a left arcuate member and a right arcuate member. The second radius is less than the first radius and the third radius is less than the fourth radius. 1. A fan nozzle for a gas turbine engine , comprising:a fan cowl having a substantially circular shape, the fan cowl being formed of an upper substantially semi-circular portion and a lower substantially semi-circular portion, the upper substantially semi-circular portion of the fan cowl having a first radius, the lower substantially semi-circular portion of the fan cowl having a second radius, the second radius being less than the first radius;a core engine cowl disposed within the fan cowl, the fan cowl and the core engine cowl positioned around a horizontal central plane, the core engine cowl having a substantially circular shape, the core engine cowl being formed of an upper substantially semi-circular portion and a lower substantially semi-circular portion, the upper substantially semi-circular portion of the core engine cowl having a third radius, the upper substantially semi-circular portion of the core engine cowl being formed of a left arcuate member and a right arcuate member, the lower substantially semi-circular portion of the core engine cowl having a fourth radius, the third ...

TURBOCHARGED GAS TURBINE ENGINE WITH ELECTRIC POWER GENERATION FOR SMALL AIRCRAFT ELECTRIC PROPULSION

A turbocharged gas turbine engine with an electric generator to provide electrical power for an aircraft (e.g., UAV) with multiple propulsor fans each driven by an electric motor, where the engine includes a low spool that drives a main fan and a high spool that drives a high speed electric generator. The low pressure compressor supplies low pressure air to an inlet of the high pressure compressor. A row of stator vanes in the high pressure turbine is cooled using cooling air bled off from the low pressure compressor outlet that is passed through an intercooler and a boost compressor, where the spent vane cooling air is discharged into the combustor. The low pressure turbine and the two compressors each include a variable inlet guide vane to control the power level of the engine. Bypass flow from the main fan is used to cool hot parts of the engine. 1. A power plant for an aircraft propelled by at least one propulsor fan , the power plant comprising:a low spool having a low pressure compressor driven by a low pressure turbine;a high spool having a high pressure compressor driven by a high pressure turbine;a combustor positioned between the high pressure compressor and the high pressure turbine;an outlet of the low pressure compressor is connected to an inlet of the high pressure compressor;the low pressure turbine includes a variable inlet guide vane;the low pressure turbine is located adjacent to the high pressure turbine and hot exhaust from the high pressure turbine flows into the low pressure turbine;a main fan driven by the low spool;an electric generator driven by the high spool;an exhaust nozzle to receive hot exhaust from the low pressure turbine;the high pressure turbine having turbine hot parts with internal cooling air passages; andan intercooler with a boost compressor connected to the low pressure compressor and the combustor through the internal cooling air passages of the turbine hot parts.2. The power plant for an aircraft propelled by at least one ...

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.

Gas turbine

A gas turbine engine, in particular an aircraft engine, includes: a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear, the sun gear is connected to the input shaft device, the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, with a rear carrier bearing device radially between the planet carrier and a part on the input shaft on the input side of the gearbox device.

GAS TURBINE

The invention relates to a gas turbine engine, in particular an aircraft engine, comprising: a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear, the sun gear is connected to the input shaft device, the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, with an inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device. 1. A gas turbine engine , in particular an aircraft engine , comprising:a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear,the sun gear is connected to the input shaft device,the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, withan inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device, witha carrier bearing system being located radially between the input shaft device and a static structure supporting the carrier bearing system, the support connection being axially in front of the input side of the gearbox device.2. The gas turbine of claim 1 , wherein the inter-shaft bearing system is located axially within or in front of a low-pressure compressor or an intermediate compressor.3. The gas turbine of claim 1 , wherein the inter-shaft bearing system is axially adjacent to the gearbox device on the input and/or the output side claim 1 , in particular with an axial distance measured from the centreline of the gearbox between 0.001 and 4 times the inner radius of the inter-shaft bearing system.4. The gas turbine of claim 1 , wherein the inter-shaft bearing device comprises at least one ball bearing.5. The gas turbine of claim 1 , wherein a fan shaft ...

COMBUSTOR SHELL ATTACHMENT

A combustor shell is provided. The combustor shell may include a first aperture at least partially defined by an inner wall of the combustor shell and passing from a diffuser-facing side of the combustor shell to a combustor-facing side of the combustor shell. The combustor shell may include a spacer comprising a first segment coupled to a first flange, wherein the first flange is disposed on the diffuser-facing side of the combustor shell, wherein an outer wall of the spacer is coupled with at least a portion of an inner wall of the combustor shell. 1. A combustor shell comprising:a first aperture at least partially defined by an inner wall of the combustor shell and passing from a diffuser-facing side of the combustor shell to a combustor-facing side of the combustor shell; anda spacer comprising a first segment coupled to a first flange, wherein the first flange is disposed on the diffuser-facing side of the combustor shell, wherein an outer wall of the spacer is coupled with at least a portion of an inner wall of the combustor shell.2. The combustor shell of claim 1 , wherein the first aperture of the combustor shell is an oblong shape.3. The combustor shell of claim 1 , wherein the spacer comprises threads on an inner wall of the spacer.4. The combustor shell of claim 1 , wherein the spacer is press fit into the combustor shell.5. The combustor shell of claim 1 , wherein the spacer is configured to engage an attachment feature of a combustor panel.6. The combustor shell of claim 5 , wherein a contact length between the spacer and the attachment feature is greater than a distance between the diffuser facing side of the combustor shell and the combustor facing side of the combustor shell.7. The combustor shell of claim 1 , wherein the spacer further comprises a second flange coupled with the first segment and disposed on the combustor-facing side of the combustor shell.8. The combustor shell of claim 1 , wherein the spacer further comprises a spacer aperture at ...

FLUID COOLING SYSTEMS FOR A GAS TURBINE ENGINE

A heat exchanger includes an airfoil configured to be positioned in a coolant stream. The airfoil includes a pressure sidewall and a suction sidewall coupled to the pressure sidewall. The suction sidewall and the pressure sidewall define a leading edge and a trailing edge opposite the leading edge. The leading edge defines an impingement zone wherein the coolant stream is configured to impinge the airfoil. The heat exchanger also includes at least one channel defined within the airfoil between the pressure sidewall and the suction sidewall. The at least one channel is at least partially defined within the impingement zone proximate the leading edge. 1. A heat exchanger comprising: a pressure sidewall; and', 'a suction sidewall coupled to said pressure sidewall, said suction sidewall and said pressure sidewall define a leading edge and a trailing edge opposite said leading edge, said leading edge defines an impingement zone wherein the coolant stream is configured to impinge said airfoil; and, 'an airfoil configured to be positioned in a coolant stream, said airfoil comprisingat least one channel defined within said airfoil between said pressure sidewall and said suction sidewall, said at least one channel at least partially defined within the impingement zone proximate said leading edge.2. The heat exchanger in accordance with claim 1 , wherein said at least one channel is configured to receive a fluid stream such that heat is removed from the fluid stream at least in part through the coolant stream impinging on said leading edge.3. The heat exchanger in accordance with claim 1 , wherein said suction sidewall and said pressure sidewall further define a root portion and a tip portion opposite said root portion claim 1 , said at least one channel comprises:an inlet section extending from said root portion to adjacent said tip portion proximate said leading edge; andan outlet section extending from adjacent said tip portion to said root portion such that said at least ...

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 ...

TURBOJET ENGINE COMPRISING A NACELLE EQUIPPED WITH REVERSER FLAPS

A turbofan comprising a fan casing and a nacelle comprising a cowl translatable between an advanced position and a pushed-back position in which the mobile cowl and the fan casing define a window therebetween. The nacelle also comprises reverser flaps, each reverser flap being mounted in a linked manner on the mobile assembly between a closed position in which it obstructs the window and an open position in which it does not obstruct the window. The nacelle also has an impelling mechanism comprising a lever arm rotatable on the mobile assembly and having a roller, a connecting rod mounted rotatably between the reverser flap and the lever arm, and a channel receiving the roller, the channel having a front part parallel with the translation direction and a rear part extending after the front part and oriented inward as it progresses from the front toward the rear. 1. A turbofan comprising an engine and a nacelle surrounding the engine which comprises a fan casing and a core arranged inside the fan casing , in which a duct for a bypass flow is defined between the core and the fan casing , said nacelle comprising:a fixed structure,a fan cowl fixedly mounted on the fixed structure and a mobile assembly comprising a mobile cowl and being translatable with respect to the fixed structure in a translation direction between an advanced position in which the mobile cowl is brought closer to the fan cowl and a pushed-back position in which the mobile cowl is moved away from the fan cowl toward the rear,a window defined upstream by the fan cowl and downstream by the mobile cowl, said window being open, in the pushed-back position, between the duct and the outside of the nacelle,a reverser flap mounted on the mobile assembly tilting between a closed position in which the reverser flap obstructs the window and an open position in which the reverser flap does not obstruct the window, and a rearward translation of the mobile assembly in the translation direction in order to move the ...

Low pressure ratio fan engine having a dimensional relationship between inlet and fan size

A gas turbine engine assembly may include, among other things, a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane, a geared architecture, a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture, a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, a forward edge on the inlet portion in a second reference plane, and a length of the inlet portion having a dimension L between the first reference plane and the second reference plane. A dimensional relationship of L/D may be between 0.30 and 0.40.

DEVICES AND METHODS FOR GUIDING BLEED AIR IN A TURBOFAN ENGINE

Device and methods for guiding bleed air in a turbofan gas turbine engine are disclosed. The devices provided include louvers and baffles that guide bleed air toward a bypass duct of the turbofan engine. The louvers and baffles have a geometric configuration that promotes desirable flow conditions and reduced energy loss. 1. A device for guiding bleed air into a bypass duct of a turbofan engine having a central axis , the device comprising:a body defining a flow-guiding surface having opposite first and second ends defining a span of the flow-guiding surface around the central axis, the flow-guiding surface extending between a radially-inner edge of the body and a radially-outer edge of the body relative to the central axis; anda side wall adjacent the first end of the flow-guiding surface of the body, the side wall extending at least partially axially relative to the central axis, the side wall extending from a first position radially inwardly of the radially-inner edge of the body to a second position radially outwardly of the radially-inner edge of the body relative to the central axis.2. The device as defined in claim 1 , wherein the second position is adjacent the radially-outer edge of the body.3. The device as defined in claim 1 , wherein the side wall is substantially planar.4. The device as defined in claim 3 , wherein the side wall is non-parallel to a radial direction relative to the central axis.5. The device as defined in claim 1 , wherein the side wall is curved.6. The device as defined in claim 1 , wherein the side wall has a Bellmouth profile when viewed along the central axis.7. The device as defined in claim 1 , wherein the side wall has a unitary construction with the body.8. The device as defined in claim 1 , comprising a baffle disposed axially of the body to define a bleed air passage between the baffle and the flow-guiding surface of the body claim 1 , wherein a gap is defined between the side wall and the baffle.9. The device as defined in ...

GAS TURBINE ENGINE AFT BEARING ARRANGEMENT

An example gas turbine engine includes a turbine and first and second spools coaxial with one another. The first spool is arranged within the second spool and extends between forward and aft ends. The aft end extends axially beyond the second spool and supports the turbine. A housing is arranged downstream from the turbine. First and second bearings are mounted to the aft end of the first spool and supported by the housing portion. 1. A bearing hub for a gas turbine engine comprising:first and second hub walls integrally formed with one another to provide a unitary structure;a radial to axial translation flange arm extending outward from an apex of the unitary structure;a translation flange extending outward from said translation flange arm;a spring arm connected to the apex for connecting the bearing hub to a canted annular flange, the spring arm including a plurality of angled flex points.2. The bearing hub of claim 1 , wherein the translation flange arm extends axially aftward from said apex of said unitary structure.3. The bearing hub of claim 1 , wherein the plurality of angled flex points comprises at least a first flex point claim 1 , a second flex point claim 1 , and a third flex point claim 1 , and wherein a stiffness of each of said first flex point claim 1 , said second flex point and said third flex point is configured to control an amount of radial vibrations translated to axial vibrations by said bearing hub.4. The bearing hub of claim 1 , wherein the first and second hub walls are inclined radially inward from an annular apex claim 1 , and a first and second bearing are respectively supported by the first and second walls opposite the apex.5. The bearing hub of claim 4 , wherein a focal node of radial vibrations of the bearing hub is the first bearing.6. The bearing hub of claim 1 , wherein said spring arm is rigidly connected to said apex.7. A gas turbine engine comprising:a fan;a compressor section fluidly connected to the fan, the compressor ...

Fan rotor for a turbo machine such as a multiple flow turbojet engine driven by a reduction gear

A forward fan rotor is disclosed with a hub of axis of rotation (X) and a cone mounted on the hub of the fan. The cone comprises an air bleed orifice which opens into an air duct of which a forward end portion passes through the fan rotor, said forward end portion comprising mechanical air entrainment means. The air bleed orifice has an annular shape and in that the cone is divided by said orifice into a front vertex portion and a rear frustoconical portion. A turbomachine forward axial spool equipped with such a fan rotor is also disclosed.

VARIABLE EXHAUST MIXER AND COOLER FOR A THREE-STREAM GAS TURBINE ENGINE

A gas turbine engine includes an outer case structure around a central longitudinal engine axis. An intermediate case structure is included inboard of the outer case structure. An inner case structure is included inboard of the intermediate case structure. A variable area exhaust mixer is included, which is movable between a closed position adjacent to the intermediate case structure and an open position adjacent to the inner case structure. 1. A gas turbine engine , comprising:an outer case structure around a central longitudinal engine axis;an intermediate case structure inboard of the outer case structure;an inner case structure inboard of the intermediate case structure; anda variable area exhaust mixer movable between a closed position adjacent to the intermediate case structure and an open position adjacent to the inner case structure.2. The gas turbine engine as recited in claim 1 , wherein the variable area exhaust mixer is downstream of a turbine section.3. The gas turbine engine as recited in claim 1 , wherein the variable area exhaust mixer includes a hinge axis radially outboard of a maximum outer diameter of a tailcone.4. The gas turbine engine as recited in claim 3 , wherein the hinge axis is axially aft of a multiple of respective flaps.5. The gas turbine engine as recited in claim 4 , wherein the multiple of respective flaps open radially inward toward the inner case structure.6. The gas turbine engine as recited in claim 4 , wherein the multiple of respective flaps are operable to extend toward a wall in the inner case structure when between the closed position and the open position.7. The gas turbine engine as recited in claim 6 , wherein the wall separates an exhaust of a second stream flow path and an exhaust of a core flow path.8. The gas turbine engine as recited in claim 7 , wherein the multiple of respective flaps are operable to extend to the wall to segregates the second stream flow path and the core flow path.9. The gas turbine engine as ...

TWO-PART GAS TURBINE ENGINE

A gas turbine engine propulsion system in which a first propulsive unit has a core engine and a first low pressure turbine arranged to be driven by combustion products from the core engine. The first propulsive unit also has a first fan rotor and a first fan shaft drivingly connecting the first turbine and the first fan rotor. The propulsion system also has a further turbine arranged in flow series with the first turbine and a second propulsive unit spaced from the first propulsive unit. The second propulsive unit has a second fan rotor driven by the rotational output of the further turbine. The further turbine may be located in the second propulsive unit and may be in fluid communication with the first turbine via an inter-turbine duct. 1. A gas turbine engine propulsion system comprising:a first propulsive unit having a core engine and a first turbine arranged to be driven by combustion products from the core engine, the first propulsive unit further comprising a first fan rotor and a first fan shaft drivingly connecting the first turbine and the first fan rotor,wherein the propulsion system comprises a further turbine arranged in flow series with the first turbine and a second propulsive unit spaced from the first propulsive unit, the second propulsive unit having a second fan rotor driven by the rotational output of the further turbine;wherein the first and second fan rotors comprise a parallel flow fan arrangement and the second fan rotor is driven by combustion products from the core engine via the further turbine;wherein the first and second propulsion units include a respective nacelle and bypass fan duct.the further turbine is connected to the second fan rotor by a second fan shaft, the further turbine being located in the second propulsive unit.2. A propulsion system according to claim 1 , wherein the further turbine is downstream of the first turbine with respect to the flow of combustion products from the core engine claim 1 , the first and further ...

NOZZLE ARRANGEMENT FOR A GAS TURBINE ENGINE

A gas turbine engine comprising: a bypass duct having a bypass nozzle; an engine core having a core nozzle; and, a mixer duct defined by a mixer fairing and having a mixer nozzle, wherein the mixer duct is arranged to receive an airflow from the bypass duct through a mixer duct inlet and an airflow from the engine core, when in use, and the geometry of the mixer duct is selectively adjustable by moving the mixer fairing relative to the bypass duct and engine core in use. 1. A gas turbine engine comprising:a bypass duct having a bypass nozzle;an engine core having a core nozzle; and,a mixer duct having a mixer duct inlet and a mixer nozzle defined by a mixer fairing which is movably mounted to the engine and,wherein the mixer duct is arranged to receive an airflow from the bypass duct through the mixer duct inlet and an airflow from the engine core, when in use, and the geometry of the mixer duct is selectively adjustable by moving the mixer fairing relative to the bypass duct and engine core in use,wherein the mixer fairing is movable between a first position and second position which simultaneously alters one or more of: an output flow area of the bypass nozzle, an output flow area of the mixer nozzle, and, a throat area of the mixer duct inlet, such that the respective in use airflows are altered.2. A gas turbine engine as claimed in claim 1 , wherein moving the mixer fairing between the first and second position alters all of: the output flow area of the bypass nozzle claim 1 , the output flow area of the mixer nozzle claim 1 , and claim 1 , the throat area of the mixer duct inlet.3. A gas turbine engine as claimed in claim 1 , wherein a portion of radially outer wall of the mixer fairing downstream of the leading edge is substantially parallel to the axis of movement such that moving the mixer fairing between a first and second position alters the output flow area of the mixer nozzle and moving the mixer fairing between a second and third position alters the ...

GEARED GAS TURBINE ENGINE

A gas turbine engine includes a fan section that includes a fan rotatable about an engine axis. A compressor section includes a low pressure compressor rotatable about the engine axis. A turbine section includes a fan drive turbine for driving the fan and the low pressure compressor. A speed reduction device connects the fan drive turbine to the fan and the low pressure compressor. The speed reduction device includes a sun gear driven by an inner shaft. A plurality of intermediate gears surround the sun gear. A carrier supports the plurality of intermediate gears for driving the low pressure compressor. A ring gear is located radially outward from the intermediate gears and includes a forward portion for driving a fan drive shaft and an aft portion. 1. A gas turbine engine comprising:a fan section including a fan rotatable about an engine axis;a compressor section including a low pressure compressor rotatable about the engine axis;a turbine section including a fan drive turbine for driving the fan and the low pressure compressor; and a sun gear driven by an inner shaft;', 'a plurality of intermediate gears surrounding the sun gear;', 'a carrier supporting the plurality of intermediate gears for driving the low pressure compressor; and', 'a ring gear located radially outward from the intermediate gears including a forward portion for driving a fan drive shaft and an aft portion., 'a speed reduction device connecting the fan drive turbine to the fan and the low pressure compressor, the speed reduction device including2. The gas turbine engine of claim 1 , wherein the carrier includes an axially forward portion and an axially aft portion claim 1 , the plurality of intermediate gears are located axially between the axially forward portion and the axially aft portion of the carrier.3. The gas turbine engine of claim 2 , wherein axially aft portion of the carrier is connected to the low pressure compressor.4. The gas turbine engine of claim 1 , wherein the fan drive ...

FAN BLADE

A blade including at least one web and a vane having a leading edge and a trailing edge, wherein, for at least one aerofoil of the vane in the vicinity of the web, a maximum sweep angle associated with a position along a chord of the aerofoil extending from the leading edge to the trailing edge of the vane corresponding to a relative chord length of at least 50%. 1. A fan blade of a bypass turbine engine , comprising at least one shank and a vane having a leading edge and a trailing edge , wherein , for at least one aerofoil of the vane in the vicinity of the shank , a maximum camber defining a point of a skeleton of the aerofoil extending from the leading edge to the trailing edge of the vane wherein a distance with a chord of the aerofoil extending from the leading edge to the trailing edge of the vane is maximum , said maximum camber being associated with a position along said chord corresponding to a relative chord length of at least 55% , with an offset of said maximum camber toward the trailing edge.2. The blade according to claim 1 , wherein said position along the chord of the aerofoil associated with the maximum camber corresponds to a relative chord length comprised between 55% and 75%.3. The blade according to claim 2 , wherein said position along the chord of the aerofoil associated with the maximum camber corresponds to a relative chord length comprised between 55% and 65%.4. The blade according to claim 1 , being made of a woven composite material.5. The blade according to claim 1 , further comprising a straight root connected to the vane with the shank.6. A fan for a bypass turbine engine comprising at least one blade according to .7. The fan according to claim 6 , comprising a disk from which said blade extends substantially radially.8. The fan according to claim 7 , wherein the shank extends outside the disk and on the inside of platforms defining the interior of the stream.9. The fan according to claim 7 , further comprising a straight root ...

NON-NEWTONIAN MATERIALS IN AIRCRAFT ENGINE AIRFOILS

A component is provided for a turbine engine. The component can include an airfoil defining a surface, and an energy absorbing composite positioned on the surface of the airfoil or within the airfoil. The energy absorbing composite includes a shear thickening fluid distributed through a matrix. 1. A component for a turbine engine , the component comprising:an airfoil defining a surface; andan energy absorbing composite positioned on the surface of the airfoil or within the airfoil, wherein the energy absorbing composite includes a shear thickening fluid distributed through a matrix.2. The component as in claim 1 , wherein the energy absorbing composite is positioned within the construction of the airfoil.3. The component as in claim 1 , wherein the energy absorbing composite is positioned on at least a portion of the surface of the airfoil.4. The component as in claim 3 , wherein the energy absorbing composite is positioned on a leading edge of the airfoil claim 3 , a side surface of the airfoil claim 3 , or both.5. The component as in claim 1 , wherein the matrix comprises a solid foamed synthetic polymer matrix.6. The component as in claim 5 , wherein the solid foamed synthetic polymer matrix comprises a synthetic elastomer.7. The component as in claim 6 , wherein the synthetic elastomer comprises an elastomeric polyurethane.8. The component as in claim 6 , wherein the synthetic elastomer comprises a first polymer-based elastic material and a second polymer-based elastic material.9. The component as in claim 8 , wherein the first polymer-based elastic material comprises an ethylene vinyl acetate or an olefin polymer claim 8 , and wherein the second polymer-based elastic material comprises a silicone polymer having dilatant properties.10. The component as in claim 5 , wherein the energy absorbing composite further comprises a polymer-based dilatant.11. The component as in claim 10 , wherein the polymer-based dilatant comprises a silicone polymer having dilatant ...

STRUT ASSEMBLY FOR AN AIRCRAFT ENGINE

A strut assembly for a gas turbine engine includes an outer structural case. The outer structural case includes a first mounting pad for mounting a first strut and a second mounting pad for mounting a second strut. The outer structural case further includes a case ligament extending between the first mounting pad and the second mounting pad in a substantially straight direction to reduce an amount of bending stress on the outer structural case. 1. A strut assembly for a gas turbine engine , the strut assembly comprising: a first mounting pad for mounting a first strut;', 'a second mounting pad for mounting a second strut; and', 'a case ligament extending between the first mounting pad and the second mounting pad, the case ligament extending in a substantially straight direction from the first mounting pad to the second mounting pad., 'an outer structural case comprising'}2. The strut assembly of claim 1 , wherein the first mounting pad claim 1 , the second mounting pad claim 1 , and the case ligament are each formed of a composite material.3. The strut assembly of claim 1 , wherein the case ligament is formed of a composite material claim 1 , wherein the composite material forming the case ligament includes a plurality of substantially aligned fibers.4. The strut assembly of claim 3 , wherein the plurality of substantially aligned fibers extend in a direction from the first mounting pad to the second mounting pad.5. The strut assembly of claim 1 , wherein the case ligament defines an inside surface claim 1 , wherein the outer structural case further comprises a plurality of wedge members positioned along the inside surface of the case ligament adjacent to the first mounting pad and adjacent to the second mounting pad.6. The strut assembly of claim 5 , wherein the plurality of wedge members are non-structural components.7. The strut assembly of claim 6 , wherein the plurality of wedge members are formed of a composite material claim 6 , wherein the case ligament is ...

SHAFT SHEAR DETECTION THROUGH SHAFT OSCILLATION

There is described a shaft shear event detection method. The method comprises obtaining a demodulated waveform of a shaft oscillation wave superimposed on a shaft rotational speed signal, comparing the amplitude to an amplitude threshold, detecting oscillation when the amplitude threshold is exceeded for a plurality of samples, and detecting a shaft shear when oscillation continues for a predetermined time limit. 1. A method for detecting a shear of a rotating shaft positioned between a source and a load , the method comprising:obtaining a demodulated waveform of a shaft oscillation wave superimposed on a shaft rotational speed signal, the waveform having an amplitude and a frequency;comparing the amplitude of the waveform to an amplitude threshold;detecting an oscillation of the shaft when the amplitude threshold is exceeded for a plurality of samples; anddetecting a shear of the rotating shaft when oscillation continues for a predetermined time limit.2. The method of claim 1 , wherein the amplitude threshold comprises a lower amplitude threshold and an upper amplitude threshold.3. The method of claim 1 , wherein the demodulated waveform is obtained through amplitude demodulation.4. The method of claim 1 , wherein the demodulated waveform is obtained through frequency demodulation.5. The method of claim 1 , wherein obtaining the demodulated waveform comprises receiving a speed sensor signal and determining the amplitude and frequency of the oscillation wave from the speed sensor signal.6. The method of claim 5 , wherein the speed sensor signal comes from a phonic wheel sensing assembly.7. The method of claim 1 , wherein the predetermined time limit is measured as a function of a period of the waveform.8. A system for detecting a shear of a rotating shaft positioned between a source and a load claim 1 , the system comprising:a processing unit; and obtaining a demodulated waveform of a shaft oscillation wave superimposed on a shaft rotational speed signal, the ...

GAS TURBINE ENGINE WITH MICROCHANNEL COOLED ELECTRIC DEVICE

A gas turbine engine includes an electrical device and a microchannel cooling system in communication with the electrical device to remove heat. 1. A gas turbine engine for use in an aircraft , the engine comprisinga low pressure spool including a fan arranged at a forward end of the engine, a low pressure turbine rotor arranged at an aft end of the engine, a low pressure drive shaft extending along an axis and rotationally coupling the fan to receive driven rotation from the low pressure turbine rotor,a high pressure spool including a compressor rotor, a high pressure turbine rotor, and a high pressure drive shaft extending along the axis and rotationally coupling the compressor rotor to receive driven rotation from the high pressure turbine rotor, andan electric device including a stator having an annular core, a rotor rotationally coupled to the low pressure drive shaft and disposed about the stator in electromagnetic communication, and a microchannel cooling system arranged radially inward of the stator in thermal communication with the annular core to pass coolant for removing heat from the stator.2. The gas turbine engine of claim 1 , wherein the microchannel cooling system includes a housing and a network of micropassageways within the housing.3. The gas turbine engine of claim 2 , wherein the micropassageways include inlet passageways for receiving coolant and outlet passageways for discharging heated coolant.4. The gas turbine engine of claim 3 , wherein each inlet passageway is connected with at least one of the outlet passageways by at least one transfer section to pass coolant in thermal communication with the annular core.5. The gas turbine engine of claim 3 , wherein the inlet and outlet passageways are arranged in alternating sequence in the circumferential direction.6. The gas turbine engine of claim 1 , wherein the stator includes electrical windings disposed radially outward of the annular core.7. The gas turbine engine of claim 1 , wherein the ...

Geared turbofan engine with targeted modular efficiency

A turbofan engine includes a fan section including a fan blade having a leading edge and hub to tip ratio of less than about 0.34 and greater than about 0.020 measured at the leading edge and a speed change mechanism with gear ratio greater than about 2.6 to 1. A first compression section includes a last blade trailing edge radial tip length that is greater than about 67% of the radial tip length of a leading edge of a first stage of the first compression section. A second compression section includes a last blade trailing edge radial tip length that is greater than about 57% of a radial tip length of a leading edge of a first stage of the first compression section.

GEAR REDUCTION FOR LOWER THRUST GEARED TURBOFAN

A gas turbine engine comprises a fan rotor having a hub and a plurality of fan blades extending radially outwardly of the hub. A compressor is positioned downstream of the fan rotor, and has a first compressor blade row defined along a rotational axis of the fan rotor and the compressor rotor. A gear reduction is positioned axially between the first compressor blade row and the fan rotor, and includes a ring gear and a carrier. The carrier has an axial length and the ring gear has an outer diameter. A ratio of the axial length to the outer diameter may be greater than or equal to about 0.20 and less than or equal to about 0.40. The gear reduction is connected to drive the hub to rotate. A method of designing a gas turbine engine is also disclosed. 1. A gas turbine engine comprising:a fan rotor having a hub and a plurality of fan blades extending radially outwardly of said hub,a compressor positioned downstream of the fan rotor, the compressor having a first compressor blade row defined along a rotational axis of said fan rotor and said compressor rotor, anda gear reduction positioned axially between said first compressor blade row and said fan rotor, said gear reduction including a ring gear and a carrier, said carrier having an axial length and said ring gear having an outer diameter, wherein a ratio of said axial length to said outer diameter may be greater than or equal to about 0.20 and less than or equal to about 0.40, and wherein said gear reduction is connected to drive said hub to rotate.2. The gas turbine engine as set forth in claim 1 , wherein a volume is defined for said carrier and said ring gear claim 1 , and said volume being greater than or equal to about 899 inchesand less than or equal to about 1349 inches.3. The gas turbine engine as set forth in claim 1 , wherein the hub has a radius defined at an inlet point of said hub claim 1 , wherein said fan blades have a radius claim 1 , and a ratio of said hub radius to said fan blade radius is less than ...

Gas turbine having a high-speed low-pressure turbine and a turbine case

A gas turbine () having a high-speed low-pressure turbine () and a turbine case () that bounds a flow path of a working fluid of the gas turbine () and an exit region (), and extends between a rotor () of the high-speed low-pressure turbine () that is the most downstream in the through flow direction of the working fluid, and an exit opening () of the turbine case (). The exit region () is designed to be free of exit guide vane assemblies. 18-. (canceled)9. A gas turbine comprising:a high-speed low-pressure turbine; anda turbine case bounding a flow path of a working fluid of the gas turbine and an exit region extending between a rotor of the high-speed low-pressure turbine, the rotor being a most downstream rotor in a through flow direction of the working fluid, the turbine case having a exit opening;wherein the exit region is free of exit guide vane assemblies.10. The gas turbine as recited in wherein the gas turbine is a turbofan engine.11. The gas turbine as recited in wherein the turbofan engine has a bypass ratio of bypass flow to primary flow of at least 1.5:1.12. The gas turbine as recited in further comprising a rotatable exit cone in a region of the high-speed low-pressure turbine.13. The gas turbine as recited in further comprising a fan coupled via a reduction gear to the high-speed low-pressure turbine.14. The gas turbine as recited in wherein claim 9 , relative to the flow path of the working fluid claim 9 , the high-speed low-pressure turbine is configured downstream of a combustion chamber or downstream of a single- or multi-stage high-pressure or intermediate-pressure turbine.15. The gas turbine as recited in wherein the high-speed low-pressure turbine is configured as a single- or multi-stage low-pressure turbine.16. The gas turbine as recited in wherein the most downstream rotor is designed in such a way that claim 9 , during operation of the gas turbine claim 9 , an average exit swirl angle of the working fluid is at most ±15° relative to an axis ...

LOW FAN NOISE TURBOJET

A double flow turbojet includes a fan including a disk centered on an axis of the fan which is provided with fan blades on its periphery, the blades having a leading edge, and an air inlet sleeve extending upstream of the fan and configured to delimit a gas flow designed to enter into the fan the air inlet sleeve having a collecting surface, the turbojet having an aspect ratio 2. The turbojet according to claim 1 , wherein the form factor is comprised between 0.1 and 0.45.3. The turbojet according to claim 1 , wherein the turbojet has a bypass ratio greater than or equal to 10.4. The turbojet according to claim 1 , wherein the diameter of the fan is comprised between 203.2 centimeters and 279.4 centimeters.530. The turbojet according to claim 1 , wherein an upstream portion of the air inlet sleeve is not symmetrical.6. The turbojet according to claim 5 , wherein a downstream portion of the air inlet sleeve is axisymmetric claim 5 , a connection between the non-symmetrical upstream portion of the air inlet sleeve and its downstream axisymmetric portion extending at a distance comprised between one and five centimeters from a plane situated at the intersection between a radially internal wall of the air inlet sleeve and a most upstream point of the leading edge of the fan blades.7. The turbojet according to claim 1 , further comprising:a primary flow space and a concentric secondary flow space,a turbine, housed in the primary flow space and in fluid communication with the fan, anda reduction mechanism, coupling the turbine and the fan, said reduction mechanism comprising a star or planetary gear reduction mechanism having a reduction ratio comprised between 2.5 and 5.8. An aircraft comprising the turbojet according to .10. The method of dimensioning according to claim 9 , further comprising a step during which the form factor is defined so that said ratio is comprised between 0.1 and 0.45.11. The turbojet according to claim 2 , wherein the form factor is comprised ...

GAS TURBINE COMPONENT WITH COOLING APERTURE HAVING SHAPED INLET AND METHOD OF FORMING THE SAME

A method of manufacturing a cooled gas turbine component includes forming a core with an outer surface. The outer surface includes a core feature. The method also includes casting an outer wall of an airfoil about the core. The outer wall has an exterior surface and an interior surface. The interior surface includes a shaped inlet portion that corresponds to the core feature. Moreover, the method includes forming an outlet portion through the outer wall to fluidly connect the outlet portion to the shaped inlet portion. The shaped inlet portion and the outlet portion cooperatively define a cooling aperture through the outer wall. 1. A cooled gas turbine component for a gas turbine engine comprising:an airfoil;an outer wall of the airfoil, the outer wall having an exterior surface and an interior surface; and a cast inlet portion included on the interior surface; and', 'an outlet portion extending through the outer wall and fluidly connected to the inlet portion;', 'wherein the inlet portion has a width and a depth, wherein the width of the inlet portion gradually reduces along the depth of the inlet portion toward the outlet portion., 'a cooling aperture that extends through the outer wall, the cooling aperture including2. The cooled gas turbine component of claim 1 , wherein the outlet portion is a hole having a substantially constant diameter through the outer wall.3. The cooled gas turbine component of claim 1 ,wherein the inlet portion is at least partially conic.4. The cooled gas turbine component of claim 1 , wherein the outlet portion extends along an axis;and wherein the axis extends at an acute angle relative to the exterior surface of the outer wall.5. The cooled gas turbine component of claim 1 , wherein the inlet portion continuously encompasses the outlet portion.6. The cooled gas turbine component of claim 1 , further comprising a thickened area that is adjacent the inlet portion.7. The cooled gas turbine component of claim 1 , wherein the depth of the ...

AIRCRAFT ENGINE OPERABILITY

A gas turbine engine has a cycle operability parameter β in a defined range to achieve improved overall performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined range of cycle operability parameter β may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox. 2. A gas turbine engine according to claim 1 , wherein at cruise conditions claim 1 , 1.0 K≤β≤1.8 K.3. A gas turbine engine according to claim 1 , wherein at cruise conditions claim 1 , 1.1 K≤β≤1.6 K claim 1 , optionally 1.10 Kto 1.50 K.4. A gas turbine engine according to claim 1 , wherein at cruise conditions claim 1 , 0.029 KgsNK≤Q≤0.036 KgsNK.5. A gas turbine engine according to claim 1 , wherein at cruise conditions claim 1 , 70 Nkgs≤specific thrust≤110 Nkgs.6. A gas turbine engine according to claim 1 , wherein a fan tip loading is defined as dH/Utip claim 1 , where dH is the enthalpy rise across the fan and Utip is the translational velocity of the fan blades at the tip of the leading edge claim 1 , and at cruise conditions claim 1 , 0.28 JkgK/(ms) Подробнее

SYSTEM AND METHOD FOR REDUCING SPECIFIC FUEL CONSUMPTION (SFC) IN A TURBINE POWERED AIRCRAFT

A system for providing auxiliary power in an aircraft. A propulsion core comprises a compressor, a combustor, a turbine, and a shaft. An accessory unit comprises an accessory combustor, an accessory turbine, and an accessory shaft. A tank is configured to hold high pressure air and operably connected to the accessory unit. An electric generator comprises an electrical output and a mechanical input, with the mechanical input operably connected to the accessory shaft and the electrical output operably connected to an electric motor operably connected to the shaft. The electrical output is operably connected to an auxiliary power consuming device in the aircraft. 1. A method for reducing the specific fuel consumption for an aircraft mission , the method comprising:predetermining characteristics of the aircraft mission;injecting high pressure air from an onboard tank into a combustor of a power turbine; and,controlling the rate of injection of the high pressure air into the combustor of the power turbine as function at least of the mass of the high pressure air in the onboard tank and predetermined characteristics of the mission.2. The method of further comprising heating the high pressure air with exhaust from the aircraft's primary propulsion system prior to injection into the combustor and injecting exhaust from the power turbine into a core combustor of the aircraft's primary propulsion system.3. The method of claim 1 , further comprising determining an expected duration of the mission; determining the amount of high pressure air available in the onboard tank claim 1 , and determining a discharge mass flow rate of the high pressure air such that exhaustion of the high pressure air substantially corresponds to the end of the mission.4. The method of claim 3 , further comprising regulating the fuel supply rate to the power turbine at least as a function of the discharge mass flow rate. This application is a divisional of U.S. Utility patent application Ser. No. 15/452 ...

Aircraft engine fan

A gas turbine engine system has an engine core and a bypass duct. A fan drives the flow through the bypass duct. A bypass efficiency is defined as the efficiency of the fan compression of the bypass flow. The bypass efficiency is a function of the bypass flow rate at a given set of conditions. The fan bypass inlet mass flow rate at the reference operating point is appreciably higher than the mass flow rate through the bypass duct at the peak bypass efficiency at a given fan reference rotational speed and cruise conditions. This results in increased design flexibility and improved overall engine performance.

TURBINE ENGINE INCLUDING BALANCED LOW PRESSURE STAGE COUNT

A turbine engine includes at least a compressor section and a turbine section, each having at least a first and second portion. A ratio of turbine section second portion stages to compressor section second portion stages is less than or equal to 1. 1. A turbine engine comprising:a fan;a compressor section having at least a first portion and a second portion, wherein said first portion is configured to exhibit a higher pressure than said second portion, and wherein said second portion of the compressor section comprises a low pressure compressor;a combustor in fluid communication with the compressor section;a turbine section in fluid communication with the combustor, wherein said turbine section includes at least a first portion and a second portion and wherein said first portion includes at least two (2) stages and is configured to exhibit a higher pressure than said second portion, and wherein said second portion of the turbine section comprises a low pressure turbine, and wherein said turbine low pressure turbine drives said fan via an epicyclic gear train geared architecture driving the fan and wherein the epicylclic geared architecture includes a speed reduction greater than about 2.3;wherein each of said compressor section second portion and said turbine section second portion includes a plurality of stages;{'b': '1', 'wherein a ratio of turbine section second portion stages to compressor section second portion stages is less than ;'}a fan bypass ratio of the turbine engine is greater than about 6.0; and{'sub': LPT', 'LPC', 'LPC', 'LPT', 'LPT', 'LPC', 'LPC', 'LPT, 'a configuration complexity metric of the low pressure compressor and low pressure turbine is in the range of about 2.63 to about 4.27, wherein the configuration complexity metric is defined by the relationship [1+N][1+[1/Nx(S)+Nx(S)]]/[N+(S)/(S)]/[2N], where, Sis the number of turbine second portion stages, Sis the number of compressor second portion stages, S/Sis the ratio of the number of ...

COMPOSITE AIRFOIL WITH METAL STRENGTH

A laminated composite airfoil assembly includes a first lamina formed of a material including metal fibers, and at least a second lamina formed of a material including at least one of metal fibers intermixed with carbon fibers, only metal fibers, only carbon fibers, a substrate including metal fibers, a substrate including carbon fibers, and combinations thereof. 1. A laminated composite airfoil assembly comprising:a first lamina formed of a material comprising metal fibers; and wherein the airfoil assembly comprises a plurality of laminae formed from materials including the first lamina and the second lamina, and', 'wherein a subset of laminae of the plurality of laminae are formed from material comprising carbon fibers, the airfoil assembly further comprising metal threads extending into the subset of laminae of the plurality of laminae., 'at least a second lamina formed of a material comprising at least one of metal fibers intermixed with carbon fibers, only metal fibers, only carbon fibers, a substrate comprising metal fibers, a substrate comprising carbon fibers, and combinations thereof,'}2. The airfoil assembly of claim 1 , wherein the material of the first lamina is a pre-preg material claim 1 , and wherein the material of the second lamina is a pre-preg material.3. The airfoil assembly of claim 2 , wherein the pre-preg material of the first lamina comprises the metal fibers oriented in a first direction and the pre-preg material of the second lamina comprises the carbon fibers oriented in a second direction.4. The airfoil assembly of claim 1 , wherein the carbon fibers are unidirectional carbon fibers oriented in a first direction and the metal fibers crisscross the carbon fibers in at least one of the first lamina and the second lamina.5. The airfoil assembly of claim 1 , wherein the metal threads extend into the subset of the plurality of laminae in a 2.5D configuration.6. The airfoil assembly of claim 1 , wherein the metal threads extend into the subset ...

FAN EXIT GUIDE VANE PLATFORM CONTOURING

A turbofan engine includes a fan section with a plurality of fan blades rotatable about an engine axis generating an airflow, a bypass passage through which the airflow passes, and a fan exit guide vane. The fan exit guide vane assembly includes a plurality of airfoils disposed between an inner platform wall and an outer platform wall. At least one of the inner platform wall and the outer platform wall includes a contoured surface between adjacent airfoils. The contoured surface includes at least one concave region and at least one convex region. A method of reducing secondary flow structures in bypass air flow with the turbofan is also disclosed.

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.

TURBOFAN ENGINE

A turbofan engine includes: a core engine, a primary flow channel, a secondary flow channel, a machine axis and a thrust nozzle that is a separate thrust nozzle for the secondary flow channel or an integral thrust nozzle for the primary flow channel and the secondary flow channel, and that has a center line and a nozzle exit surface. The thrust nozzle is tilted with respect to the machine axis while forming an articulation angle that the center line of the thrust nozzle forms with respect to the machine axis at least at the nozzle exit, and the nozzle exit surface of the thrust nozzle is oriented obliquely with respect to the center line of the thrust nozzle while forming an angle of inclination that a normal of the nozzle exit surface forms with respect to the center line of the thrust nozzle. 1. A turbofan engine that is provided and suitable for the purpose of being arranged below a wing of an airplane , wherein the turbofan engine comprises:a core engine that comprises a compressor, a combustion chamber and a turbine,a primary flow channel that leads through the core engine,a secondary flow channel that leads past the core engine,a machine axis of the core engine, a center line,', 'a nozzle exit surface,, 'a thrust nozzle which has a separate thrust nozzle for the secondary flow channel, or is an integral thrust nozzle for the primary flow channel and the secondary flow channel, and which compriseswhereinthe thrust nozzle is tilted with respect to the machine axis while forming a articulation angle that the center line of the thrust nozzle forms with respect to the machine axis at least at the nozzle exit, and that the nozzle exit surface of the thrust nozzle is obliquely oriented with respect to the center line of the thrust nozzle while forming an angle of inclination that a normal of the nozzle exit surface forms with respect to the center line of the thrust nozzle.2. The turbofan engine according to claim 1 , wherein the articulation angle and the angle of ...

TURBOFAN ENGINE WTH A SPLITTERED ROTOR FAN

A turbofan engine includes: a turbomachinery core; and a low-pressure turbine configured to drive a fan to produce a fan flow, the fan being configured such that at least a portion of the fan flow exits the engine without passing through a turbine. The fan includes: a rotor having at least one rotor stage including axial-flow rotor airfoils, and at least one stator stage including axial-flow stator airfoils. At least one of the rotor or stator stages includes an array of airfoil-shaped splitter airfoils alternating with the rotor or stator airfoils of the corresponding stage. At least one of a chord dimension of the splitter airfoils and a span dimension of the splitter airfoils is less than the corresponding dimension of the airfoils of the at least one stage. 1. A turbofan engine , comprising:a turbomachinery core operable to produce a flow of combustion gases;a low-pressure turbine configured to extract energy from the combustion gases so as to drive a fan to produce a fan flow, the fan being configured such that at least a portion of the fan flow exits the engine without passing through a turbine; a rotor comprising at least one rotor stage including a rotatable disk defining a rotor flowpath surface and an array of axial-flow rotor airfoils extending outward from the flowpath surface;', 'at least one stator stage comprising a wall defining a stator flowpath surface, and an array of axial-flow stator airfoils extending away from the stator flowpath surface; and', 'wherein at least one of the rotor or stator stages includes an array of airfoil-shaped splitter airfoils extending from at least one of the flowpath surfaces thereof, the splitter airfoils alternating with the rotor or stator airfoils of the corresponding stage, wherein at least one of a chord dimension of the splitter airfoils and a span dimension of the splitter airfoils is less than the corresponding dimension of the airfoils of the at least one stage., 'wherein the fan includes2. The engine of ...

METHOD FOR RELEASE OF FAN BLISK AIRFOIL WITH EXTERNAL SHAPED CHARGE

According to one aspect, a method of releasing a fan blade for testing a turbofan engine includes arranging an external shaped charge about an airfoil and modifying the airfoil by extricating one or more portions of material from one or more sides of the airfoil. The method further includes detonating the external shaped charge such that the airfoil is released at a selected time. 1. A method of releasing a fan blade for testing a turbofan engine , comprising:arranging an external shaped charge about an airfoil;modifying the airfoil by extricating one or more portions of material from one or more sides of the airfoil; anddetonating the external shaped charge such that the airfoil is released at a selected time.2. The method of claim 1 , further comprising:modifying the airfoil by extricating a first portion of material from the leading edge to provide a first cut out;modifying the airfoil by extricating a second portion of material from the trailing edge to provide a second cut out; andarranging one or more detonators in the first cut out or the second cut out.3. The method of claim 2 , further comprising:modifying the airfoil by extricating additional portions of material from one or more sides of the airfoil; andaligning the external shaped charge with one or more cavities formed in the one or more sides of the airfoil by the extricating of the additional portions.4. The method of claim 3 , further comprising:providing with the external shaped charge a liner between the charge and the one or more sides of the airfoil; andaligning the external shaped charge and the liner with the one or more cavities in the one or more sides of the airfoil such that the liner accelerates through a cavity provided by the one or more cavities.5. The method of claim 4 , further comprising:selectively weakening the airfoil by extricating the additional portions of material from the one or more sides of the airfoil;wherein, by aligning the external shaped charge and the liner with the ...

NOSE CONE ATTACHMENT FOR TURBOFAN ENGINE

An assembly for attaching a nose cone to a turbofan engine includes a single center bolt connector, a nose cone, and a turbofan hub. Further, the single center bolt connector couples the nose cone to the turbofan hub and axially clamps the nose cone to the turbofan hub. 1. An assembly for attaching a nose cone to a turbofan engine , comprising:a single center bolt connector;a nose cone; and wherein the single center bolt connector couples the nose cone to the turbofan hub; and', 'wherein the single center bolt connector axially clamps the nose cone to the turbofan hub., 'a turbofan hub;'}2. The assembly of claim 1 , further comprising:a nose cone base; and wherein the nose cone base abuts the hub forward face; and', 'wherein the single center bolt connector provides axial compression to hold the nose cone base against the hub forward face., 'a hub forward face;'}3. The assembly of claim 2 , wherein the abutment of the nose cone base against the hub forward face provides an aerodynamic surface claim 2 , and wherein the axial compression of the nose cone base against the hub forward face closes one or more aerodynamic gaps between the nose cone and the hub.4. The assembly of claim 2 , further comprising:an interrupted pilot; and wherein the interrupted pilot is disposed about the nose cone base;', 'wherein the one or more scallops are disposed about the hub forward face; and', 'wherein the interrupted pilot couples the nose cone base to the hub forward face by interconnecting with the one or more scallops., 'one or more scallops;'}5. The assembly of claim 4 , further comprising:one or more trim balance features disposed about the nose cone, wherein the one or more trim balance features are separated from the interrupted pilot.6. The assembly of claim 5 , further comprising:one or more forward balance features disposed about the hub forward face, wherein the one or more forward balance features are the one or more scallops; andwherein the one or more trim balance ...

AIRFOIL WITH STRESS-REDUCING FILLET ADAPTED FOR USE IN A GAS TURBINE ENGINE

The present disclosure is directed toward airfoils used in gas turbine engines. More specifically, the present disclosure teaches airfoils with fillets for managing stresses in airfoils during use in gas turbine engines. 1. A gas turbine engine comprisingan engine core including a compressor assembly, a combustor assembly, and a turbine assembly,a fan assembly including a fan rotor coupled to the engine core to be driven by the engine core, an inner guide vane assembly arranged to interact with air discharged by the fan rotor moving into the engine core, and an outer guide vane assembly arranged to interact with air discharged by the fan rotor moving around the engine core,wherein the outer guide vane assembly includes an inner band arranged around at least a portion of a central axis and an airfoil that extends radially outward from the inner band away from the central axis, the airfoil including a sheet of material that is folded to define a leading edge of the airfoil, a pressure side of the airfoil, and a suction side of the airfoil, the sheet of material being welded to define a trailing edge of the airfoil, andwherein at least one of the pressure side and the suction side of the airfoil is shaped to form a fillet at the interface of the airfoil with the inner band, the fillet shaped to taper such that the fillet increases in size as the fillet extends from the trailing edge along a chord length of the airfoil toward the leading edge of the airfoil.2. The engine of claim 1 , wherein both the pressure side and the suction side of the airfoil are shaped to form a fillet at the interface of the airfoil with the inner band claim 1 , the fillet shaped to taper such that the fillet increases in size as the fillet extends from the trailing edge along a chord length of the airfoil toward the leading edge.3. The engine of claim 1 , wherein the fillet has a first radial height at the trailing edge and a second radial height claim 1 , greater than the first radial height ...

STATIC HUB TRANSITION DUCT

An engine includes a compressor having a first compressor, a second compressor, and a static hub transition duct. The first compressor includes a rotatable portion and a static portion. The second compressor is disposed downstream of the first compressor and includes an inlet and a rotatable portion. The static hub transition duct includes an inner wall and an outer wall. The inner wall is secured to the static portion of the first compressor, and the inner wall and outer wall define a space that fluidically couples the first compressor and the inlet of the second compressor. 1. A compressor of an engine , wherein the engine includes a shaft , comprising:a first compressor, wherein the first compressor includes a first rotatable portion and a static portion, wherein the first rotatable portion is mechanically secured to the shaft;a second compressor downstream of the first compressor, wherein the second compressor includes a second rotatable portion, wherein the second rotatable portion is mechanically secured to the shaft; anda transition duct having an inner wall and an outer wall, wherein the inner wall is secured to the static portion of the first compressor, and the inner wall and outer wall define a space that fluidically couples an exit of the first compressor and an inlet of the second compressor.2. The compressor of claim 1 , wherein the first compressor is an axial compressor and the second compressor is a centrifugal compressor.3. The compressor of claim 1 , wherein the inner wall of the transition duct is secured to a final stator of the axial compressor.4. The compressor of claim 3 , wherein the outer wall of the transition duct is secured to the final stator of the axial compressor.5. The compressor of claim 4 , wherein a casing surrounds the axial compressor.6. The compressor of claim 5 , wherein an extension portion of the casing forms at least a portion of the outer wall of the static transition duct.7. The compressor of claim 3 , further including ...

Gas turbine engine hollow fan blade rib orientation

A fan blade includes first and second portions that are secured to one another and provide a cavity. The first and second portions form an exterior airfoil surface that extends in a radial direction from a root to a tip and in a chord-wise direction from a leading edge to a trailing edge. Radial ribs extend in a radial direction from the root toward the tip and are spaced apart from one another in the chord-wise direction. First and second angled ribs intersect one another at a first apex. The radial ribs intersect at least one of the first and second angled ribs. The first and second angled ribs are at an angle relative to one of the radial ribs. The angle is in a range of 45°+/−30°.

Aircraft incorporating a thrust recovery system using cabin air

An aircraft incorporating a cabin air recovery system in which the aircraft comprises a pressurizable cabin, main turbofan engines, each turbofan engine having fan blades, a gas turbine coupled with the fan blades and a by-pass duct bypassing the gas turbine. The cabin fluidly communicates with the by-pass duct downstream of the fan blades so that, during operation, cabin outflow air is discharged into the by-pass duct downstream of the fan blades. By re-utilizing excess cabin air, engine thrust and efficiency is improved, and fuel consumption is reduced.

Turbofan Engine Bearing and Gearbox Arrangement

A turbofan engine () comprising a fan shaft () configured to rotate about an axis () of the engine. A fan drive gear system () is configured to drive the fan shaft. The fan drive gear system has a centerplane (). A first spool comprises a high pressure turbine () and a high pressure compressor (). A second spool comprises an intermediate pressure turbine (), a lower pressure compressor, and a shaft () coupling the intermediate pressure turbine to the lower pressure compressor. A third spool comprises a lower pressure turbine () coupled to the fan drive gear system to drive the fan. The engine has a plurality of main bearings. The turbofan engine has a single stage fan. Of the main bearings, at least one is a shaft-engaging bearing engaging a driving shaft () coupled to the fan drive gear system. A closest () of the shaft-engaging bearings engaging the driving shaft behind the fan drive gear system has a centerplane () and a characteristic radius (RB). The half angle (θ) of a virtual cone () intersecting the core flowpath inboard boundary at the gear system centerplane () and said closest of the shaft-engaging bearings at the characteristic radius (RB) is greater than about 32°. A hub-to-tip ratio (HR:FR) of the fan is less than about 0.38. 2. The engine of wherein:{'sub': 'D', 'b': 540', '580, 'a length Lbetween the centerplane () and a centerplane () of a forward/upstreammost compressor disk is at least one of{'sub': 'G', 'about 2.0 times or less of a gear width Lof the fan drive gear system;'}{'sub': 'T2', 'b': '580', 'about 60% or less of a core flowpath inboard radius Rat the forward/upstreammost compressor disk centerplane (); and'}{'sub': 'T', 'b': '540', 'about 50% less of a core flowpath inboard radius Rat the centerplane ().'}3. The engine of wherein:{'sub': 'D', 'b': 540', '580, 'claim-text': [{'sub': 'G', 'about 1.5 times or less of the gear width Lof the fan drive gear system;'}, {'sub': 'T2', 'b': '580', 'about 50% or less of the core flowpath inboard ...

LOW NOISE TURBINE FOR GEARED TURBOFAN ENGINE

A gas turbine engine comprises a fan, a compressor section including a compressor having a low pressure portion and a high pressure portion, a combustor section, and a turbine having a downstream portion. A gear reduction effects a reduction in the speed of the fan relative to an input speed to the fan. The downstream portion of the turbine has a number of turbine blades in each of a plurality of rows of the downstream turbine portion. The turbine blades operate at least some of the time at a rotational speed. The number of blades and the rotational speed are such that the following formula holds true for at least one of the blade rows of the downstream turbine: (number of blades×speed)/60≧5500. The rotational speed is an approach speed in revolutions per minute. A method of designing a gas turbine engine and a turbine module are also disclosed. 1. A gas turbine engine comprising:a fan, a compressor section including a compressor having a low pressure portion and a high pressure portion, a combustor section, and a turbine having a downstream portion;a gear reduction effecting a reduction in the speed of said fan relative to an input speed to said fan; {'br': None, '(number of blades×speed)/60≧5500;'}, 'said downstream portion of said turbine having a number of turbine blades in each of a plurality of rows of said downstream turbine portion, and said turbine blades operating at least some of the time at a rotational speed, and said number of blades and said rotational speed being such that the following formula holds true for at least one of the blade rows of the downstream turbine'}andsaid rotational speed being an approach speed in revolutions per minute.2. The gas turbine engine as set forth in claim 1 , wherein the formula results in a number greater than or equal to 6000.3. The gas turbine engine as set forth in claim 2 , wherein said gas turbine engine is rated to produce 15 claim 2 ,000 pounds of thrust or more.4. The gas turbine engine as set forth in claim 3 ...

THERMAL SHIELDING IN A GAS TURBINE

A turbine blade having labyrinth of internal channels for circulation of coolant, the novel labyrinth geometry includes; inlet arranged on axially upstream face of terminal portion leading to an upstream duct portion having first section adjacent the inlet and a second section, the second section having reduced cross section compared to first section; a first passage intersecting first section and extending through the blade body towards tip of the blade, a proximal end of leading edge passage arranged to capture incoming air flow; a main blade passage intersecting a downstream duct portion, the downstream duct portion arranged in axial alignment with the upstream duct portion but separate from upstream duct portion; and a restrictor passage intersecting with the mid-blade passage and extending towards a mid-blade duct portion, the mid-blade duct portion in axial alignment with the upstream and downstream duct portions and in fluid communication with the upstream duct portion. 1. A turbine blade having a body enclosing a labyrinth of internal channels for the circulation of coolant received through an inlet formed in a terminal portion of the blade root , the labyrinth comprising;an inlet arranged on an axially upstream face of the terminal portion leading to a duct;a first passage intersecting the duct at a first passage intersection and extending through the blade body towards the tip of the blade, a proximal end of the first passage being arranged, in use, to capture incoming coolant flow;a second passage intersecting the duct at a second passage intersection at a position downstream of the first passage intersection;in use, a clearance space between an external wall surface of the duct and a bucket groove of a disc hub in which the blade is carried, the clearance space creating a leakage path for air directed to the inlet;the duct and/or the passage intersections configured to create a pressure drop in the duct in the direction from the inlet to the second ...

THRUST REVERSER STRUCTURE MOUNTED TO FAN CASE

A gas turbine engine has a core engine and a fan. A pivoting D-door structure includes a pivoting inner D-door portion and an outer D-door portion, which is axially moveable relative to the inner D-door portion between a stowed and thrust reverse position. A fan case surrounds the fan. Cascades are axially moveable along with the outer D-door portion between a stowed and a thrust reverse position. Structure supporting the cascades is mounted on the fan case. 1. A gas turbine engine comprising:a core engine and a fan;a pivoting D-door structure including a pivoting inner D-door portion and an outer D-door portion, which is axially moveable relative to said inner D-door portion between a stowed and thrust reverse position;a fan case surrounding said fan; andcascades axially moveable along with said outer D-door portion between a stowed and a thrust reverse position, and structure supporting said cascades being mounted on said fan case.2. The gas turbine engine as set forth in claim 1 , wherein an actuator is provided on said fan case for driving said cascades and said outer D-door portion between said stowed and thrust reverse positions.3. The gas turbine engine as set forth in claim 2 , wherein said cascades include a plurality of openings for directing air in a direction opposing further movement of an aircraft receiving the gas turbine engine.4. The gas turbine engine as set forth in claim 3 , wherein blocker doors are moved to a blocking position when said cascades and said outer D-doors portions are moved to said thrust reverse position.5. The gas turbine engine as set forth in claim 4 , wherein said actuator for said cascades is a ball screw actuator.6. The gas turbine engine as set forth in claim 5 , wherein a guide for guiding said cascades is provided with a channel receiving a portion moveable with said cascades.7. The gas turbine engine as set forth in claim 2 , wherein said actuator for said cascades is a ball screw actuator.8. The gas turbine engine as ...

EFFICIENT, LOW PRESSURE RATIO PROPULSOR FOR GAS TURBINE ENGINES

A gas turbine engine includes a bypass flow passage that has an inlet and defines a bypass ratio in a range of approximately 8.5 to 13.5. A fan is arranged within the bypass flow passage. A first turbine is a 5-stage turbine and is coupled with a first shaft, which is coupled with the fan. A first compressor is coupled with the first shaft and is a 3-stage compressor. A second turbine is coupled with a second shaft and is a 2-stage turbine. The fan includes a row of fan blades that extend from a hub. The row includes a number (N) of the fan blades, a solidity value (R) at tips of the fab blades, and a ratio of N/R that is from 14 to 16. 1. A gas turbine engine comprising:a bypass flow passage and a core flow passage, the bypass flow passage including an inlet and defining a bypass ratio in a range of approximately 8.5 to 13.5 with regard to flow through the bypass flow passage and flow through the core flow passage;a fan arranged within the bypass flow passage;a first shaft and a second shaft;a first turbine coupled with the first shaft, the first shaft coupled with the fan, wherein the first turbine is a 5-stage turbine;a first compressor coupled with the first shaft, wherein the first compressor is a 3-stage compressor; anda second turbine coupled with the second shaft, wherein the second turbine is a 2-stage turbine;wherein the fan includes a hub and a row of fan blades that extend from the hub, and the row includes a number (N) of the fan blades, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 14 to 16 to manage the propulsive losses at the lower speed.2. The gas turbine engine as recited in claim 1 , wherein the number (N) of the fan blades is 18.3. The gas turbine engine as recited in claim 2 , wherein the bypass flow passage includes an inlet and an outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine claim 2 , the ...

POWER GENERATION SYSTEM, DRIVING METHOD FOR POWER GENERATION SYSTEM, AND COMBUSTOR

In a power generation system, exhausted fuel gas exhausted from a solid oxide fuel cell (SOFC) is used as a fuel of a first combustor or a second combustor of a gas turbine, and at the same time, a part of compressed air compressed by a compressor of the gas turbine is used to drive the SOFC. The gas turbine includes the first combustor for burning fuel gas which is different from the exhausted fuel gas, a first turbine configured to be driven by combustion gas supplied from the first combustor, the second combustor for burning at least a part of the exhausted fuel gas, and a second turbine coupled with the first turbine and configured to be driven by combustion gas supplied from the second combustor. 15-. (canceled)6. A power generation system comprising:a controller configured to control to close a first main nozzle control valve and open a second main nozzle control valve when a gas turbine is started and control to open the first main nozzle control valve and restrict the second main nozzle control valve when a fuel cell is started after the gas turbine has been driven, whereinexhausted fuel gas exhausted from the fuel cell is used as a fuel of a combustor of the gas turbine, andthe combustor includesa first main nozzle,a second main nozzle,a first main nozzle fuel line which is connected to the first main nozzle and sends the exhausted fuel gas exhausted from the fuel cell,a second main nozzle fuel line which is connected to the second main nozzle and sends a fuel gas different kind from the exhausted fuel gas,the first main nozzle control valve provided in the first main nozzle fuel line, andthe second main nozzle control valve provided in the second main nozzle fuel line.7. The power generation system according to claim 6 , whereinthe combustor includes a pilot nozzle, a pilot nozzle fuel line which is connected to the pilot nozzle and sends the fuel gas, and a pilot nozzle control valve provided in the pilot nozzle fuel line, and the controller controls to ...

Geared gas turbine engine with reduced fan noise

A fan section for a gas turbine engine according to an example of the present disclosure includes, among other things, a fan rotor having fan blades, and a plurality of fan exit guide vanes positioned downstream of the fan rotor. The fan rotor is configured to be driven through a gear reduction. A ratio of a number of fan exit guide vanes to a number of fan blades is defined. The fan exit guide vanes are provided with optimized sweep and optimized lean.

FAN INTEGRATED INERTIAL PARTICLE SEPARATOR

A gas turbine engine includes a fan, an engine core, and an airflow duct assembly. The fan is mounted for rotation about a central axis of the gas turbine engine assembly to produce thrust for the gas turbine engine. The engine core is coupled to the fan and configured to drive the fan about the central axis. The airflow duct assembly defines a core passageway configured to conduct a first portion of air pushed by the fan into the engine core and a by-pass passageway configured to conduct a second portion air pushed by the fan around the engine core. 1. A gas turbine engine comprisinga fan mounted for rotation about a central axis of the gas turbine engine,an engine core coupled to the fan and configured to drive the fan about the central axis to cause the fan to push a mixture of air and particles suspended in the air to provide thrust for the gas turbine engine, andan airflow duct assembly configured to conduct the mixture of air and particles through the gas turbine engine, the airflow duct assembly defining a core passageway configured to conduct a first portion of the mixture of air and particles pushed by the fan into the engine core and a by-pass passageway configured to conduct a second portion of the mixture of air and particles pushed by the fan around the engine core, andwherein the airflow duct assembly includes a particle-separator splitter positioned in the core passageway and configured to separate the first portion of the mixture of air and particles into a clean flow substantially free of particles and a dirty flow containing the particles and the particle-separator splitter is arranged to direct the clean flow into the engine core and the dirty flow away from the engine core.2. The gas turbine engine of claim 1 , wherein the airflow duct assembly further includes an inner wall arranged circumferentially around the central axis claim 1 , an outer wall arranged circumferentially around the inner wall and the fan claim 1 , and a by-pass flow splitter ...

FAN MODULE WITH VARIABLE-PITCH BLADES FOR A TURBINE ENGINE

A ducted fan module with variable-pitch blades for a turbo engine includes a rotor carrying a plurality of blades, a stationary casing and a pitch changing and control system for adjusting and controlling the setting of the blades. The rotor includes a central shaft and a blade support ring surrounding the central shaft, a front end of the ring that is connected to a front end of the central shaft, so as to create a rearward-opening annular space between the ring and the shaft. The annular space of the rotor housing the blade changing and control system and the central shaft are guided by a first bearing mounted in a stationary casing behind the ring. The fan has a tubular part connected to the stationary casing and extending into the annular space in front of the first bearing and carrying at least a first complementary bearing for guiding the shaft. 1. A ducted fan module with variable-pitch blades for a turbo engine , comprising:a rotor carrying a plurality of blades;a stationary casing; anda pitch changing and control system configured for adjusting and controlling a setting of the plurality of blades,wherein the rotor comprises a central shaft and a blade support ring surrounding the central shaft, a front end of the blade support ring connected to a front end of the central shaft so as to create a rearward-opening annular space between the blade support ring and the central shaft, the annular space of the rotor configured to house the blade pitch changing and control system and the central shaft being guided by a first bearing mounted in the stationary casing behind the blade support ring,wherein the blade support ring accommodates a plurality of shafts, each shaft of the plurality of shafts forming a foot of one blade of the plurality of blades, andwherein the module contains a tubular part connected to the stationary casing and extending into the annular space in front of the first bearing, the tubular part carrying at least a first complementary bearing for ...

Stiffened platform

A platform for a fan of a turbomachine that will be mounted on a hub of the fan between two adjacent blades, the platform comprising a principal wall with a longitudinal principal orientation, that has a curved shape along the longitudinal direction, and a stiffener that is mounted on a lower face of the principal wall, of which the section of the stiffener, in a plane perpendicular to the longitudinal direction is globally in the form of a U open towards the principal wall. The stiffener comprises at least two straight segments inclined relative to each other and inclined relative to the longitudinal principal direction.

Systems and methods for controlling blade tip clearances

A system for controlling blade tip clearances in a gas turbine engine may comprise an active clearance control system and a controller in operable communication with the active clearance control system. The controller may be configured to identify a cruise condition, reduce a thrust limit of the gas turbine engine to a de-rated maximum climb thrust, determine a first target tip clearance based on the de-rated maximum climb thrust, and send a command signal correlating to the first target tip clearance to the active clearance control system.

GAS TURBINE ENGINE KEEL BEAM

A gas turbine engine for an aircraft is provided. The gas turbine engine comprises an engine core comprising a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor. The gas turbine engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades, the fan generating a core airflow which enters the engine core and a bypass airflow which flows through a bypass duct surrounding the engine core. The gas turbine engine further comprises a circumferential row of outer guide vanes located in the bypass duct rearwards of the fan, the outer guide vanes extending radially outwardly from an inner ring which defines a radially inner surface of the bypass duct. 1. A gas turbine engine for an aircraft comprising:an engine core comprising a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor;a fan located upstream of the engine core, the fan comprising a plurality of fan blades, the fan generating a core airflow which enters the engine core and a bypass airflow which flows through a bypass duct surrounding the engine core;a circumferential row of outer guide vanes located in the bypass duct rearwards of the fan, the outer guide vanes extending radially outwardly from an inner ring which defines a radially inner surface of the bypass duct;an inner cowl which provides an aerodynamic fairing surrounding the engine core, the inner cowl being rearwards of the inner ring, and including two door sections located on respective and opposite sides of the engine, each door section being pivotably openable about an upper edge thereof to enable maintenance access to the engine core; anda keel beam which extends rearwardly from the inner ring at bottom dead centre thereof to provide latching formations for latching to a lower edge of each door section when it is closed.2. A gas turbine engine according to claim 1 , wherein the keel beam has seal lands which extend ...

LOW PRESSURE RATIO FAN ENGINE HAVING A DIMENSIONAL RELATIONSHIP BETWEEN INLET AND FAN SIZE

According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D that is based on a dimension of the fan blades, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane, a geared architecture, a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture, a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, a forward edge on the inlet portion in a second reference plane, and a length of the inlet portion having a dimension L measured along an engine axis between the first reference plane and the second reference plane. A dimensional relationship of L/D is between 0.20 and 0.40. 1. A gas turbine engine assembly comprising:a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D that is based on a dimension of the fan blades, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane;a compressor section including a low pressure compressor and a high pressure compressor;a geared architecture including an epicyclical gear train that drives the fan at a lower speed than an input speed in the geared architecture;a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture;a low spool and a high spool, the low spool comprising the low pressure compressor and the low pressure turbine, the high spool comprising the high pressure compressor and the high pressure turbine, the low spool and the high spool rotatable about an engine central longitudinal axis;a nacelle surrounding the fan, the nacelle including ...

GEARED TURBOFAN ARRANGEMENT WITH CORE SPLIT POWER RATIO

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section including a fan having a plurality of fan blades, and a nacelle surrounding the plurality of fan blades, a compressor section including a low pressure compressor and a high pressure compressor, the low pressure compressor including a plurality of stages, and the high pressure compressor including 6 or more stages. A turbine section includes a fan drive turbine that drives the fan section through a gear arrangement, and including a second turbine that drives the high pressure compressor. A power ratio is provided by the combination of a first power input of the low pressure compressor and a second power input of the high pressure compressor, the power ratio defined by the second power input divided by the first power input, and the power ratio is less than or equal to 1.0 1. A gas turbine engine comprising:a fan section including a fan having a plurality of fan blades, and a nacelle surrounding the plurality of fan blades;a compressor section including a low pressure compressor and a high pressure compressor, the low pressure compressor including a plurality of stages, and the high pressure compressor including 6 or more stages;wherein the fan section delivers a portion of air into the compressor section, and a portion of air into a bypass duct, and a bypass ratio, which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section, is equal to or greater than 12;a turbine section including a fan drive turbine that drives the fan section through a gear arrangement, and including a second turbine that drives the high pressure compressor, anda power ratio provided by the combination of a first power input of the low pressure compressor and a second power input of the high pressure compressor, the power ratio defined by the second power input divided by the first power input, and the power ratio ...

SUPERSONIC AIRCRAFT TURBOFAN

A turbofan engine having: an engine core having a centre axis and including in flow series a compressor, a combustor and a turbine; and a bypass duct surrounding the engine core, the bypass duct has a bypass duct exit area at its downstream end. The engine further includes an exhaust nozzle assembly including: coaxially arranged inner mixer and outer exhaust nozzles, the exhaust nozzle being axially downstream of said mixer nozzle; a core flow duct defined by the mixer nozzle, the core flow duct having a core exit area; and an exhaust duct defined at least in part by the exhaust nozzle downstream of the mixer nozzle, the exhaust duct having an exhaust throat area. 1. A turbofan engine having:an engine core having a centre axis and comprising in flow series a compressor, a combustor and a turbine;a bypass duct surrounding the engine core, the bypass duct having a bypass duct exit area at its downstream end; and coaxially arranged an inner mixer nozzle and an outer exhaust nozzle, the exhaust nozzle being axially downstream of said mixer nozzle;', 'a core flow duct defined by the mixer nozzle, the core flow duct having a core exit area; and', 'an exhaust duct defined at least in part by the exhaust nozzle downstream of the mixer nozzle, the exhaust duct having an exhaust throat area,, 'an exhaust nozzle assembly comprisingwherein the bypass duct exit area and core exit area are axially aligned at a mixing plane to form a mixing plane area;wherein the turbofan engine has a transonic thrust condition, a supersonic cruise condition and a take-off condition;wherein in the supersonic cruise condition, the exhaust throat area is increased relative to the exhaust throat area in the transonic condition; andwherein in the take-off condition, the core exit area is increased relative to the core exit area in the transonic condition, and the bypass duct exit area is decreased relative to the bypass duct exit area in the transonic condition.2. The turbofan engine according to ...

THRUST REVERSER ASSEMBLY

A thrust reverser assembly for a gas turbine engine including a core engine, a nacelle surrounding at least a portion of the core engine defining a bypass duct between the nacelle and the core engine where an outer door movable between a stowed position and a deployed position extends outwards from the nacelle and a blocker door movable between a stowed position and a deployed position extends into an airflow conduit defined by the bypass duct to deflect air outwards. 1. A gas turbine engine , comprising:a core engine;a nacelle surrounding at least a portion of the core engine;a bypass duct defined by and between the nacelle and the core engine and defining a fore-to-aft airflow conduit;an outer door movable between a stowed position and a deployed position, where the outer door extends outwards from the nacelle;a blocker door extending between a fore end and an aft end and movable between a stowed position and a deployed position, where the blocker door extends into the airflow conduit to deflect airflow outwards; anda rotary actuator assembly located aft of the fore end of the blocker door and rotatable about a hinge point having a circumferential axis with respect to the core engine, operably coupled to the outer door and transferring a torque to the outer door to rotate the outer door about the hinge point between the stowed and deployed positions and operably coupled to the blocker door and transferring a torque to the blocker door to rotate the blocker door about the hinge point between the stowed and deployed positions.2. The gas turbine engine of claim 1 , further comprising a cascade element located within the nacelle.3. The gas turbine engine of wherein the blocker door is located radially inward from the cascade element.4. The gas turbine engine of wherein the nacelle comprises an outer cowl portion aft of the blocker doors and wherein the outer cowl portion is fixed.5. The gas turbine engine of wherein the rotary actuator assembly is mounted within at ...

NACELLE-INTEGRATED AIR-DRIVEN AUGMENTOR FAN FOR INCREASING PROPULSOR BYPASS RATIO & EFFICIENCY

Systems and methods are provided for an air-driven augmentor fan equipped aircraft propulsor. The augmentor fan may increase the effective bypass ratio of the aircraft propulsor and reduce fuel consumption and carbon emissions of the aircraft. The augmentor fan may be driven by air energized by a ducted fan powered by the core engine of the aircraft propulsor. The energized air may be received by an inlet, flowed through a flow path, and exhausted out the outlet to drive the augmentor fan. The exhausted energized air may impart a torque on the augmentor fan or blades of the augmentor fan. One or more of the inlet, flow path, or outlet may be variable in size to control the volume of air flowed through the flow path. 1. An aircraft propulsor comprising:a nacelle;a turbofan engine;an augmentor fan comprising a plurality of augmentor fan blades, the augmentor fan coupled to the nacelle and configured to rotate mechanically independently of any turbofan engines;a control unit configured to vary a pitch of at least one of the augmentor fan blades; andan air flow path comprising an inlet, a duct, and an outlet, wherein the inlet is positioned to receive and pass air energized by the turbofan engine to the duct, and wherein the outlet is positioned to exhaust the energized air from the duct to drive the augmentor fan.2. The aircraft propulsor of claim 1 , wherein the augmentor fan further comprises an augmentor hub ring circumscribing the nacelle and configured to allow rotation of the augmentor fan blades relative to the nacelle claim 1 , wherein each of the augmentor fan blades is coupled to the augmentor hub ring.3. The aircraft propulsor of claim 2 , wherein the control unit comprises a gear ring coupled to at least one of the augmentor fan blades and configured to rotate to vary the pitch of the at least one augmentor fan blade.4. The aircraft propulsor of claim 3 , wherein each of the augmentor fan blades comprises an augmentor vane and an augmentor blade root claim ...

STABILITY MARGIN AND CLEARANCE CONTROL USING POWER EXTRACTION AND ASSIST OF A GAS TURBINE ENGINE

A method of maintaining rotor tip clearance and compressor operational line during a transient operation of a gas turbine engine is disclosed. In various embodiments, the method includes applying high spool auxiliary power to a high speed spool for a first time period, applying low spool auxiliary power to a low speed spool for a second time period, sensing one or more operational parameters of the gas turbine engine during the transient operation, and ceasing application of power to the high speed spool, based on the one or more operational parameters. 1. A method of maintaining rotor tip clearance during a transient operation of a gas turbine engine , comprising:applying high spool auxiliary power to a high speed spool for a first time period;applying low spool auxiliary power to a low speed spool for a second time period;sensing one or more operational parameters of the gas turbine engine during the transient operation; andceasing application of power to the high speed spool, based on the one or more operational parameters.2. The method of claim 1 , wherein the applying the high spool auxiliary power to the high speed spool comprises driving a high spool shaft connected to the high speed spool.3. The method of claim 2 , wherein the applying the low spool auxiliary power to the low speed spool comprises driving a low spool shaft connected to the low speed spool.4. The method of claim 3 , wherein the driving the high spool shaft comprises powering a high spool tower-shaft via an electric motor.5. The method of claim 3 , wherein the driving the low spool shaft comprises powering a low spool tower-shaft via an electric motor.6. The method of claim 1 , wherein the one or more operational parameters includes a high speed spool rotational speed and a low speed spool rotational speed.7. The method of claim 6 , wherein the one or more operational parameters includes a high pressure compressor pressure ratio and a low pressure compressor pressure ratio.8. The method of ...

FAN BLADE ATTACHMENT ROOT WITH IMPROVED STRAIN RESPONSE

A fan blade is provided and comprises a leading edge, an attachment root extending aft of the leading edge, and a trench formed in a surface of the attachment root. An attachment root is also provided. The attachment root comprises a leading edge, a dovetail extending aft of the leading edge, and a trench formed in a surface of the dovetail. A gas turbine engine is also provided. The gas turbine engine comprises a compressor section configured to rotate about an axis, a combustor aft of the compressor section, and a turbine section aft of the compressor section and configured to rotate about the axis. A fan may be disposed forward of the turbine section and include a blade. The blade may have a trench formed in an attachment root. 1. A fan blade , comprising:a leading edge;an attachment root extending aft of the leading edge; anda trench formed in a surface of the attachment root, wherein the trench is offset from the leading edge.2. An attachment root , comprising:a leading edge;a dovetail extending aft of the leading edge; anda trench formed in a surface of the dovetail, wherein the trench is offset from the leading edge. This application is a divisional of, claims priority to, and the benefit of U.S. Ser. No. 15/075,705 filed on Mar. 21, 2016 entitled “FAN BLADE ATTACHMENT ROOT WITH IMPROVED STRAIN RESPONSE.” The '705 application is a nonprovisional of, and claims priority to, and the benefit of U.S. Provisional Application No. 62/167,100, entitled “FAN BLADE ATTACHMENT ROOT WITH IMPROVED STRAIN RESPONSE,” filed on May 27, 2015.” The contents of both are incorporated herein by reference in their entirety.The present disclosure relates to gas turbine engines, and, more specifically, to a fan blade attachment root with an improved strain response to impact events.Aircraft may collide with birds while in flight. In some instances, birds may collide with a gas turbine engine. In gas turbine engines having a fan, the fan blades may absorb the brunt of an impact. The ...

AIRCRAFT FAN WITH LOW PART-SPAN SOLIDITY

A fan for a gas turbine engine includes: an annular casing; a disk disposed inside the casing and mounted for rotation about an axial centerline, the disk including a row of fan blades extending radially outwardly therefrom; each of the fan blades including an airfoil having circumferentially opposite pressure and suction sides extending radially in span from a root to a tip, and extending axially in chord between spaced-apart leading and trailing edges, with the airfoils defining corresponding flow passages therebetween for pressurizing air; the row including no more than 21 and no less than 13 of the fan blades; and wherein each of the fan blades has a solidity defined by a ratio of the airfoil chord over a circumferential pitch of the fan blades, measured at 60% of a radial distance from the root to the tip, of less than about 1.6. 1. A fan for powering an aircraft in flight comprising:an annular casing;a disk disposed inside the casing and mounted for rotation about an axial centerline, the disk including a row of fan blades extending radially outwardly therefrom;each of the fan blades including an airfoil having circumferentially opposite pressure and suction sides extending radially in span from a root to a tip, and extending axially in chord between spaced-apart leading and trailing edges, with the airfoils defining corresponding flow passages therebetween for pressurizing air;the row including no more than 21 and no less than 13 of the fan blades; andwherein each of the fan blades has a solidity defined by a ratio of the airfoil chord over a circumferential pitch of the fan blades, measured at 60% of a radial distance from the axial centerline to the tip, of less than about 1.6.2. The fan of wherein the solidity measured at 60% of the radial distance from the axial centerline to the tip is no greater than about 1.4.3. The fan of wherein each of the fan blades has a solidity defined by a ratio of the airfoil chord over the circumferential pitch claim 2 , ...

SELF RETAINED LINKAGE AND SYSTEM INCLUDING THE SELF RETAINED LINKAGE FOR A GAS TURBINE ENGINE

A variable vane assembly for a gas turbine engine, the variable vane assembly including: a plurality of vanes arranged into a plurality of stages, each one of the plurality of vanes being configured for rotation about an axis through movement of a vane arm secured to each one of the plurality of vanes at one end and a sync ring of each one of the plurality of stages at another end; and a plurality of bell cranks operably coupling the sync ring of each one of the plurality of stages to a self-retained linkage via a stud of each one of the plurality of bell cranks, the stud of each one of the plurality of bell cranks being rotatably received in a corresponding opening of the self-retained linkage in an alternating fashion such that only a single shear interface is provided between each one of the plurality of bell cranks and the self-retained linkage. 1. A variable vane assembly for a gas turbine engine , comprising:a plurality of vanes arranged into a plurality of stages, each one of the plurality of vanes being configured for rotation about an axis through movement of a vane arm secured to each one of the plurality of vanes at one end and a sync ring of each one of the plurality of stages at another end; anda plurality of bell cranks operably coupling the sync ring of each one of the plurality of stages to a self-retained linkage via a stud of each one of the plurality of bell cranks, the stud of each one of the plurality of bell cranks being rotatably received in a corresponding opening of the self-retained linkage in an alternating fashion such that only a single shear interface is provided between each one of the plurality of bell cranks and the self-retained linkage.2. The variable vane assembly as in claim 1 , wherein each one of the plurality of bell cranks are also pivotally secured to a bracket.3. The variable vane assembly as in claim 2 , wherein each one of the plurality of bell cranks are operably coupled to the sync ring of each one of the plurality of ...

LOW POWER HYBRID TURBOFAN DRAG REDUCTION

A method of operating a turbofan engine includes coupling an electric motor to drive a fan rotatable within a nacelle, detecting an in-flight operating condition indicative of an increase in drag, and driving rotation of the fan with the electric motor to redirect a flow of air through the nacelle to reduce drag. A turbofan engine is also disclosed. 1. A method of operating a turbofan engine comprising:coupling an electric motor to drive a fan rotatable within a nacelle;detecting an in-flight operating condition indicative of an increase in drag; anddriving rotation of the fan with the electric motor to redirect a flow of air through the nacelle to reduce drag.2. The method as recited in claim 1 , wherein the in-flight operating condition comprises one of a low engine power condition or an engine off condition.3. The method as recited in claim 2 , wherein the nacelle includes a plurality of exit guide vanes aft of the fan and redirecting a flow air includes adjusting an incident angle of airflow relative to a fan rotational axis from the fan by rotating the fan with the electric motor.4. The method as recited in claim 3 , wherein the incident angle of airflow is adjusted from a first angle corresponding with a non-powered fan to a second angle corresponding with a powered rotation of the fan.5. The method as recited in claim 4 , wherein the first angle is less than the second angle.6. The method as recited in claim 2 , wherein a portion of the increase in drag is generated by an airflow stagnation region at a leading edge of the nacelle being at a first location and driving rotation of the fan comprises adjusting the airflow stagnation region at the leading edge toward a second location corresponding with a reduced drag condition.7. The method as recited in claim 6 , wherein the first location is radially inboard of the second location at the leading edge of the nacelle.8. The method as recited in claim 1 , including a controller receiving information indicative of ...

A TWO-SPOOL DESIGN FOR A TURBOSHAFT ENGINE WITH A HIGH-PRESSURE COMPRESSOR CONNECTED TO A LOW-PRESSURE TURBINE

A turboshaft engine including a low-pressure compressor, a high-pressure compressor, a low-pressure turbine, a high-pressure turbine, and a regulator for regulating a speed of rotation of the low-pressure turbine to a speed that is substantially constant. The low-pressure turbine is coupled by a first shaft to the high-pressure compressor, while the high-pressure turbine is coupled by a second shaft to the low-pressure compressor. 12-. (canceled)3. A turboshaft engine comprising:a low-pressure compressor;a high-pressure compressor;a low-pressure turbine;a high-pressure turbine; andregulator means for regulating a speed of rotation of the low-pressure turbine to a speed that is substantially constant,wherein the low-pressure turbine is coupled by a first shaft to the high-pressure compressor, while the high-pressure turbine is coupled by a second shaft to the low-pressure compressor, the first shaft passes coaxially inside the second shaft, the first and second shafts defining an axial direction,wherein the high-pressure compressor, the low-pressure compressor, the high-pressure turbine, and the low-pressure turbine are arranged in that order along the axial direction.4. A turboshaft engine according to claim 3 , forming a helicopter engine. The invention relates to the internal structure of a turboshaft engine, and more particularly to the internal structure of a helicopter engine.It should be observed that the term “turbojet” designates a gas turbine engine delivering thrust required for propulsion by reaction to the ejection of hot gas at high speed, whereas the term “turboshaft engine” designates a gas turbine engine that drives a drive shaft in rotation. For example, turboshaft engines are used as engines in helicopters, ships, trains, or indeed as industrial engines. Turboprops (turboshaft engines driving propellers) are also engines used as aeroengines.A conventional turboshaft engine generally comprises a low-pressure compressor and a high-pressure compressor ...

Gas turbine engine airfoil

An airfoil for a turbine engine includes pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a stacking offset and a span position that includes at least one positive and negative slope. The positive slope leans aftward and the negative slope leans forward relative to an engine axis. The positive slope crosses an initial axial stacking offset corresponding to the 0% span position at a zero-crossing position. A first axial stacking offset X1 is provided from the zero-crossing position to a negative-most value on the curve. A second axial stacking offset X2 is provided from the zero-crossing position to a positive-most value on the curve. A ratio of the second to first axial stacking offset X2/X1 is less than 2.0.

TURBINE SECTION WITH TIP FLOW VANES

A turbine section of a gas turbine engine comprises a gas path having an outer boundary wall. A circumferential array of turbine blades projects radially into the gas path. Each turbine blade extends in span from a hub to a tip and in chord from a leading edge to a trailing edge. A circumferential array of tip flow vanes extends radially inward from the outer boundary wall with a span corresponding generally to a radial depth of a tip leakage flow region of the turbine blades. The tip flow vanes are disposed downstream of the circumferential array of turbine blades adjacent to the trailing edge of the turbine blades. 1. A turbine section of a gas turbine engine , the turbine section comprising: a gas path having an outer boundary wall , a turbine rotor having a circumferential array of turbine blades projecting radially into the gas path , each turbine blade extending in span from a hub to a tip and extending in chord from a leading edge to a trailing edge , a circumferential array of tip flow vanes extending radially inward from the outer boundary wall with a span corresponding generally to a radial depth (h) of a tip leakage flow region of the turbine blades , the tip flow vanes being disposed downstream of the circumferential array of turbine blades and at a short axial distance therefrom to catch a tip leakage flow before it starts mixing with a mainstream flow coming from the turbine blades.2. The turbine section defined in claim 1 , wherein the tip flow vanes extend from the outer boundary wall by a distance up to a value which is directly proportional to a tip clearance (t) of the turbine blades and the radial depth (h) of tip leakage flow region and inversely proportional to the span (H) of the upstream turbine blades.3. The turbine section defined in claim 1 , wherein each of the tip flow vanes defines a twist along the span thereof.4. The turbine section defined in claim 1 , wherein each of the tip flow vanes is provided in the form of an airfoil ...

Composite Fan Blade

A composite fan blade for a gas turbine engine is disclosed. The fan blade may include a core being made of a first material and a shell enclosing the core. The shell may be made of a second material and the second material may have less plasticity than the first material.

GAS TURBINE ENGINE WITH LOW STAGE COUNT LOW PRESSURE TURBINE

A gas turbine engine including a core nacelle defined about an engine axis. A fan nacelle is mounted at least partially around the core nacelle to define a fan bypass airflow path for a fan bypass airflow. A gear train is defined along an engine axis. The gear train defines a gear reduction ratio of greater than or equal to about 2.3. A spool along the engine axis drives the gear train. The spool includes a downstream turbine having six or fewer stages. A fan is driven through the gear train by the downstream turbine. A pressure ratio across the fan is less than about 1.45. A fan variable area nozzle is axially movable relative to the fan nacelle to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation. 1. A gas turbine engine comprising:a core nacelle defined about an engine axis;a fan nacelle mounted at least partially around said core nacelle to define a fan bypass airflow path for a fan bypass airflow;a gear train defined along an engine axis, said gear train defines a gear reduction ratio of greater than or equal to about 2.3;a spool along said engine axis which drives said gear train, said spool includes a downstream turbine having six or fewer stages;a fan driven through the gear train by the downstream turbine, wherein a pressure ratio across the fan is less than about 1.45; anda fan variable area nozzle axially movable relative to said fan nacelle to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation.2. The engine as recited in claim 1 , wherein said spool is a low spool.3. The engine as recited in claim 1 , wherein said downstream turbine defines a pressure ratio that is greater than about five (5).4. The engine as recited in claim 1 , wherein said downstream turbine defines a pressure ratio that is greater than five (5).5. The engine as recited in claim 1 , wherein said fan bypass airflow defines a bypass ratio greater than about ten (10).6. The ...

NOZZLE FOR AN AIRCRAFT TURBOPROP ENGINE WITH AN UNDUCTED FAN

The present disclosure provides a nozzle for an aircraft turboprop engine with an unducted fan, including: an inner wall, an outer wall radially spaced apart from the inner wall and concentric with the inner wall, a junction area of the inner and outer walls including an opening contained in a plane transverse to a longitudinal axis of the nozzle. In particular, the junction area of the inner and outer walls includes two connecting plates and a member to secure the two connecting plates together, or in another form, the junction area includes a pad secured to the inner wall, and a pad secured to the outer wall, facing the pad of the inner wall of the nozzle. 1. A nozzle for an aircraft turboprop engine with an unducted fan , comprising:an inner wall,an outer wall being radially spaced apart from the inner wall and concentric with the inner wall,a junction area of the inner and outer walls comprising at least one opening contained in a plane substantially transverse to a longitudinal axis of said nozzle,wherein the junction area of the inner and outer walls further comprises means selected from the group consisting of:means for connecting the inner and outer walls, said connecting means comprising at least two connecting plates and means for securing said at least two connecting plates together, andat least one pad secured to the inner wall and at least one pad secured to the outer wall of the nozzle and positioned facing said at least one pad of the inner wall of the nozzle.2. The nozzle according to claim 1 , wherein each of the inner and outer walls of the nozzle comprises at least one metallic skin made of an austenite nickel-chromium based superalloy.3. The nozzle according to claim 1 , wherein the connecting means are distributed discretely on a circumference of the nozzle claim 1 , between the inner and outer walls of the nozzle.4. The nozzle according to claim 1 , wherein the inner wall of the nozzle comprises a metallic skin made of an austenite nickel- ...