Приводная установка (варианты) и способ управления приводной установкой

Изобретение относится к приводным установкам и способу управления такими установками. Приводная установка (1) для приведения в действие нагрузки (21) содержит газотурбинный двигатель (3), выполненный с возможностью приведения в действие нагрузки (21), электрический двигатель / генератор (23), соединенный с сетью (G) распределения электроэнергии, первую нагрузочную муфту (19), соединяющую турбину (3) с нагрузкой (21), и вторую нагрузочную муфту (22), соединяющую нагрузку (21) с электродвигателем / генератором (23). Газотурбинный двигатель (3) содержит газогенератор (5) с ротором (9R, 11R) и силовую турбину (7) с ротором (7R), причем ротор силовой турбины механически отделен от ротора газогенератора или не связан с ним по крутящему моменту. Электродвигатель / генератор (23) выполнен с возможностью работы в качестве генератора и в качестве двигателя. Приводная установка (1) также содержит преобразователь (25) частоты, подсоединенный между электродвигателем / генератором (23) и сетью (G) распределения ...

РЕДУКТОР ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Редуктор газотурбинного двигателя с внутренним зацеплением содержит внешний и внутренний выходные валы, соосно установленные на его выходе, и сателлитные шестерни на стационарных опорах. Сателлитные шестерни выполнены двойными с разными по наружному диаметру внешними зубчатыми венцами. Сателлитная шестерня с большим по диаметру зубчатым венцом выполнена с возможностью зацепления с ведущим колесом и с ведомой шестерней внутреннего зацепления, установленной на внешнем выходном валу. Сателлитная шестерня с меньшим по диаметру зубчатым венцом выполнена с возможностью зацепления с ведомой шестерней внешнего зацепления, установленной на внутреннем выходном валу. Сателлитные шестерни установлены на внешних относительно этих шестерен подшипниках качения, причем подшипник качения со стороны ведущего колеса выполнен с Т-образным в поперечном сечении внутренним кольцом. Отношение внешнего диаметра наружного кольца подшипника качения со стороны ведущего колеса к диаметру сателлитной шестерни с большим ...

УСТРОЙСТВО СОЕДИНЕНИЯ ВАЛОВ ТУРБИНЫ И КОМПРЕССОРАГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Изобретение относится к газотурбинным двигателям, а именно к трансмиссии, соединяющей роторы турбины и компрессора. Устройство соединения валов турбины и компрессора газотурбинного двигателя содержит промежуточный вал, в задний хвостовик которого по шлицам через шлицевую муфту установлен вал турбины, а в передний хвостовик - вал компрессора. На наружном диаметре шлиц вала турбины в кольцевую проточку заднего хвостовика промежуточного вала установлено центрирующее кольцо. Изобретение позволяет повысить надежность работы роликоподшипника турбины и уменьшить утечки воздуха в уплотнениях за счет улучшения крепления промежуточного вала на валу турбины. 3 ил.

УСТРОЙСТВО ПЕРЕДАЧИ КРУТЯЩЕГО МОМЕНТА ОТ ВАЛА КОМПРЕССОРА К КОРОБКЕ ПРИВОДНЫХ АГРЕГАТОВ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Изобретение относится к узлам приводов авиационных газотурбинных двигателей. Устройство передачи крутящего момента от вала компрессора к коробке приводных агрегатов газотурбинного двигателя содержит прямозубое ведомое цилиндрическое колесо и коническую шестерню. Между цилиндрическим колесом и конической шестерней с помощью шлицевых соединений установлен промежуточный вал. Промежуточный вал имеет на своих концах внутренние кольцевые полости, соединенные радиальными каналами с полостями шлицевых соединений. Прямозубое колесо и коническая шестерня выполнены с внутренними буртиками, диаметр которых меньше внутреннего диаметра впадин шлицев промежуточного вала. Изобретение позволяет повысить надежность работы устройства за счет постоянной прокачки маслом шлицевых соединений, а также удаления продуктов износа шлицев, их смазки и охлаждения. 3 ил.

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ В УСТАНОВКАХ С МЕХАНИЧЕСКИМ ПРИВОДОМ И СПОСОБЫ ЕГО РАБОТЫ

Приводная система для приведения в действие по меньшей мере одного компрессора. Система содержит газотурбинный двигатель (101), выполненный и установленный с возможностью приведения в действие компрессора (103). Газотурбинный двигатель имеет горячей конец (101Н) и холодный конец (101С). Нагрузочная муфта (105) для соединения указанного газотурбинного двигателя (101) с компрессором (103) расположена на горячем конце (101Н) двигателя (101). Электрический двигатель/генератор (111) расположен на холодном конце (101С) двигателя (101). Электрический двигатель/генератор (111) электрически соединен с электроэнергетической системой (G) и выполнен с возможностью работы в качестве генератора для преобразования вырабатываемой газотурбинным двигателем (101) избыточной механической энергии в электрическую энергию, доставки этой электрической энергии в электроэнергетическую систему и в качестве электрического двигателя для добавления дополнительной приводной мощности компрессору (103). Расположение двигателя ...

Газовая турбина с двусторонним приводом

Изобретение относится к энергетике. Газотурбинная система, содержащая газовую турбину (23), первую нагрузку (71) и вторую нагрузку (72), приводимые в действие с помощью газовой турбины. Газовая турбина (23) содержит газогенератор (27), турбину (50) низкого давления и приводной вал (65), приводимый в действие турбиной (50) низкого давления. Приводной вал имеет первый конец (65Н), соединенный с возможностью передачи приводного усилия с первой нагрузкой, и второй конец (65С), соединенный с возможностью передачи приводного усилия со второй нагрузкой. Первая нагрузка и вторая нагрузки расположены на противоположных сторонах газовой турбины, при этом приводной вал (65) проходит в осевом направлении через газовую турбину от первого конца ко второму концу. Причём, первая нагрузка и вторая нагрузка содержат компрессор с вертикальным разъемом корпуса. Также представлен способ эксплуатации газотурбинной системы. Изобретение позволяет повысить КПД системы. 2 н. и 12 з.п. ф-лы, 5 ил.

КОНСТРУКЦИЯ РЕДУКТОРНОГО ГАЗОТУРБИННОГО ДВИГАТЕЛЯ, ОБЕСПЕЧИВАЮЩАЯ ПОВЫШЕННЫЙ КПД

Газотурбинный двигатель содержит редуктор, соединенный с возможностью вращения с приводным валом вентилятора, и компрессор высокого давления. Газотурбинный двигатель выполнен с возможностью поддержания температуры на выходе компрессора высокого давления в диапазоне от 621 до 732°C при взлете, а отношение скоростей истечения, определяемое как отношение скорости истечения вентиляторной струи к скорости истечения основной струи, находится в диапазоне от 0,75 до 0,90 при полете с крейсерской мощностью двигателя на высоте около 10668 метров (35000 футов) со скоростью около 0,80 числа Маха. Степень двухконтурности двигателя превышает 8,0. Обеспечивается повышение КПД двигателя и, как следствие, уменьшается расход топлива. 4 н. и 24 з.п. ф-лы, 2 ил.

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ

Изобретение относится к области двигателестроения, а именно к узлам привода авиационных газотурбинных двигателей и газотурбинных установок. Технический результат заключается в повышении надежности двигателя за счет предотвращения износа шлицев в соединении ведущего зубчатого колеса и вала ротора компрессора высокого давления путем обеспечения их смазки в процессе работы. В газотурбинном двигателе, включающем вал компрессора низкого давления, вал компрессора высокого давления, ведущее зубчатое колесо, передающее крутящий момент с вала компрессора высокого давления на центральный привод через шлицевое соединение, на валу компрессора низкого давления установлена втулка с лабиринтными гребешками, а на ведущем зубчатом колесе размещена втулка с уплотнительным покрытием, образующая с указанным зубчатым колесом кольцевые радиальные полости, которые сообщаются с входной полостью, расположенной между указанными втулками, и полостью центрального привода через отверстия, выполненные в стенках втулки ...

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ С ГИДРООБЪЕМНОЙ ПЕРЕДАЧЕЙ

Изобретение относится к газотурбинным двигателям. Газотурбинный двигатель с гидрообъемной передачей содержит турбокомпрессоры первого и второго каскадов, силовую турбину, механизм передач. Силовая турбина связана с гидромотором гидрообъемной передачи, которая, в свою очередь, имеет гидравлическую связь с гидронасосом и турбокомпрессорами через механизм передач. Изобретение позволит повысить среднюю скорость движения машины с газотурбинным двигателем на II и III передачах и уменьшить средний километровый расход топлива. 1 ил.

Реверсивная турбинная установка судового типа

Изобретение относится к судостроению, в частности к реверсивным турбинным установкам судового типа. Реверсивная турбинная установка судового типа включает установленный в корпусе силовой агрегат с противоположно вращающимися роторами, валы которых соединены с коаксиально расположенными входными валами планетарно-дифференциального механизма. Дифференциальный механизм содержит солнечное колесо, эпицикл, водило, блок сателлитов, каждый из которых выполнен в виде основной зубчатой шестерни, зубчатою передачу и тормозные устройства. Силовой агрегат выполнен в виде многоступенчатой биротативной турбины, состоящей из наружного ротора, вал которого соединен с внешним валом, и внутреннего ротора. Блок сателлитов снабжен дополнительными зубчатыми шестернями по числу основных шестерен, каждая из которых жестко соединена с соответствующей основной шестерней и введена в зацепление с эпициклом. Установка снабжена торцевым контактным уплотнением с приводом его перемещения, который установлен на корпусе ...

Механизм передачи крутящего момента агрегатам турбореактивного двигателя (ТРД), центральная коническая передача (ЦКП) ТРД, главная коническая шестерённая пара ЦКП ТРД, корпус ЦКП ТРД, ведущее зубчатое коническое колесо ЦКП, ведомое зубчатое коническое колесо ЦКП, узел ЦКП ТРД

Группа изобретений относится к области авиадвигателестроения. Единый механизм передачи крутящего момента агрегатам двухвального, двухконтурного авиационного ТРД, имеющего газодинамически связанные между собой соосные валы РВД и РНД, включает соединенные с РВД с возможностью передачи агрегатам крутящего момента от турбины высокого давления ЦКП и кинематически соединенные с ней редукторы приводов КДА и КСА. Редукторы приводов КСА сообщены по крутящему моменту с ЦКП через многоступенчатый редуктор КДА и через гибкий вал с концевыми шарнирами и сильфонами. ЦКП содержит главную шестеренную пару конических ведущего и ведомого зубчатых колес, которые имеют зубчатые венцы. Главная шестеренная пара зубчатых колес ЦКП выполнена с передаточным числом i=(1,12÷1,43) [б/р]. Ведущее колесо главной шестеренной пары размещено на валу, установленном в шарико- и роликовом подшипниках. Ведомое колесо выполнено с валом, установленным в роликовом подшипнике и в шарикоподшипнике, который установлен в крышке корпуса ...

ДВУХКОНТУРНЫЙ ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ

Изобретение относится к двухконтурным газотурбинным двигателям авиационного и наземного применения. Двухконтурный газотурбинный двигатель включает в себя валы (5) и (12) вентилятора (2) и турбины низкого давления (11), соединенные с помощью эвольвентных шлиц (13). Внутри вала (5) вентилятора установлен стяжной винт (14) на сферических кольцах (16) и (17) и ввернут в стяжную втулку (15). Втулка (15) установлена в валу (12) турбины низкого давления с помощью сферического кольца (19) и зафиксирована в окружном направлении шлицами (20) балансировочной втулки (21). Втулка (21) установлена внешними осевыми ребрами (22) во внутренней кольцевой канавке (23) вала (5) вентилятора и зафиксирована относительно осевых выступов (28) на его хвостовике (24) в осевом и в окружном направлениях радиальными выступами (25), выполненными на радиальном ребре (26), и стопорным кольцом (27) с возможностью установки в кольцевой канавке (23) вала в пазах (29) между осевыми ребрами (22) втулки (21) балансировочных ...

Узел соединения валов ротора низкого давления газотурбинного двигателя

Изобретение относится к газотурбинным двигателям (ГТД) авиационного применения, а именно к конструкции узла соединения роторов компрессора и турбины. Техническим результатом, достигаемым при использовании настоящего изобретения, является: повышение безопасности двухмоторного летательного аппарата при возникновении нештатной ситуации в работе двигателя, связанной с обрывом вала турбины низкого давления, либо при еще каких-нибудь повреждениях, требующих принудительного механического останова ротора, а также расширение области применения данного устройства. Указанный технический результат достигается тем, что известный узел соединения валов ротора низкого давления газотурбинного двигателя содержит вал компрессора низкого давления, вал турбины низкого давления, цапфы которых заведены в промежуточный вал, причем вышеупомянутые валы зафиксированы относительно друг друга в окружном и осевом направлениях, втулку, жестко соединенную со статором, поршень, установленный во втулке с возможностью осевого ...

ГАЗОВАЯ ТУРБИНА, МЕХАНИЧЕСКОЕ ТРАНСПОРТНОЕ СРЕДСТВО С ГАЗОТУРБИННЫМ ПРИВОДОМ И СПОСОБ ЕГО ТОРМОЖЕНИЯ ДВИГАТЕЛЕМ

... 1. Газовая турбина (1, 101, 201), содержащая первый компрессор (2), камеру (16) сгорания и первую турбину (11), которая через первый вал (10а, 10b) приводит во вращение компрессор, а также установленный перед первой турбиной выпускной клапан (12), через который во время торможения двигателем часть сжатого компрессором газа сбрасывается в атмосферу в обход первой турбины, отличающаяся тем, что степень открытия выпускного клапана (12), от которой зависит количество сжатого в компрессоре газа, который сбрасывается в атмосферу в обход первой турбины (11), регулируется в зависимости от давления газа на выходе из компрессора (2). 2. Газовая турбина по п.1, отличающаяся тем, что выпускной клапан (12) имеет подпружиненный открывающий его элемент (38), которым регулируется степень открытия клапана. 3. Газовая турбина по п.1 или 2, отличающаяся тем, что выпускной клапан (12) расположен перед камерой сгорания (16). 4. Газовая турбина по п.1, отличающаяся наличием расположенной за первой турбиной ( ...

УЗЕЛ СОЕДИНЕНИЯ РОТОРОВ КОМПРЕССОРА И ТУРБИНЫ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Узел соединения роторов компрессора и турбины газотурбинного двигателя, содержащий валы компрессора и турбины, соединенные между собой в осевом направлении через стяжную трубу, а в окружном направлении - через рессору, контактирующую с внутренними поверхностями валов компрессора и турбины через шлицевые соединения и уплотнительные кольца, отличающийся тем, что на внутренних поверхностях валов компрессора и турбины выполнены бурты, контактирующие с торцевыми поверхностями рессоры, концевые участки последней выполнены с диаметром, меньшим внутреннего диаметра шлицев, а уплотнительные кольца размещены на этих участках.

КОМПОНОВКА РЕДУКТОРНОГО ТУРБОВЕНТИЛЯТОРНОГО ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

... 1. Газотурбинный двигатель, содержащий:компрессорную секцию;камеру сгорания, сообщающуюся по текучей среде с компрессорной секцией;турбинную секцию, сообщающуюся по текучей среде с камерой сгорания и содержащую турбину привода вентилятора и вторую турбину, причем турбина привода вентилятора содержит множество роторов турбины;вентилятор, имеющий множество лопаток, вращаемых вокруг оси, причем отношение числа лопаток вентилятора к числу роторов составляет от приблизительно 2,5 до приблизительно 8,5; исистему изменения скорости, приводимую в действие турбиной привода вентилятора для обеспечения вращения вентилятора вокруг оси;причем турбина привода вентилятора содержит первый задний ротор, прикрепленный к первому валу, а вторая турбина содержит второй задний ротор, прикрепленный ко второму валу,причем аксиально позади первого соединения между первым задним ротором и первым валом расположен первый подшипниковый узел, а аксиально перед вторым соединением между вторым задним ротором и вторым ...

ЭНЕРГЕТИЧЕСКАЯ ГАЗОТУРБИННАЯ УСТАНОВКА

... 1. Энергетическая газотурбинная установка, содержащая газотурбинный двигатель, в котором входное устройство снабжено статорным фланцем, а ротор свободной силовой турбины с радиально-упорным подшипником, размещенным со стороны входного устройства, выполнен с валом и фланцем привода полезной нагрузки, а также с лабиринтным диском, причем статорным фланцем и лабиринтным диском с корпусом входного устройства образована разгрузочная полость свободной силовой турбины, отличающаяся тем, что лабиринтный диск установлен кольцевым радиальным фланцем на кольцевом радиальном фланце вала привода полезной нагрузки и зафиксирован в осевом направлении фланцем привода полезной нагрузки, при этом валы привода полезной нагрузки и ротора свободной силовой турбины соединены устройством для передачи крутящего момента, выполненного, например, в виде шлиц, а также соединены стяжным болтом. 2. Энергетическая газотурбинная установка по п.1, отличающаяся тем, что внешний диаметр статорного фланца в 1,2-2 раза превышает ...

Способ управления двухроторным газотурбинным двигателем самолета при останове

Изобретение относится к управлению авиационным двигателем. Способ управления двухроторным газотурбинным двигателем самолета при останове заключается в уменьшении частоты вращения вала ротора высокого давления и вала ротора низкого давления. При этом частоту вращения вала ротора высокого давления и вала ротора низкого давления уменьшают до достижения роторами одинаковой частоты вращения. Роторы зацепляют друг с другом обгонной муфтой, расположенной между валами, после чего частоту вращения роторов уменьшают до останова. Изобретение обеспечивает стабильную подачу масла к опорам двигателя на останове до полной остановки всех роторов двигателя, а также позволяет снизить эффект «прихватывания» вала ротора высокого давления при останове. 1 ил.

Теплофикационная парогазовая установка

Теплофикационная парогазовая установка с паротурбинным приводом компрессора относится к энергетике и может быть применена для тепло- и электроснабжения потребителей в новых микрорайонах городов. Теплофикационная парогазовая установка, содержащая газотурбинную установку с компрессором, камерой сгорания, газовой турбиной и электрогенератором, паровой котел-утилизатор, в котором по ходу газов размещены пароперегреватель, испаритель второй ступени, камера дожигания топлива, испаритель первой ступени, экономайзер, газоводяной подогреватель сетевой воды, котел-утилизатор вырабатывает перегретый пар средних параметров, в установке применены основная и дополнительная противодавленческие паровые турбины, сетевые подогреватели первой и второй ступеней, деаэратор. В отопительных режимах ее работы за счет сжигания дополнительного топлива в камере дожигания увеличивают паропроизводительность котла-утилизатора, перегретый пар подают в дополнительную паровую турбину, отработавший в ней пар подают в сетевой ...

Двухконтурный двухвальный турбореактивный двигатель

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ

... 1. Газотурбинный двигатель, включающий входной кок, вал двигателя и вал привода внешней нагрузки со стороны входного корпуса двигателя, отличающийся тем, что во внутренней полости кока между валами двигателя и привода внешней нагрузки размещена гибкая муфта, при этом отношение наружного диаметра гибкой муфты к расстоянию от муфты до центра подшипника вала составляет 1. . . 5, а отношение наружного диаметра гибкой муфты к диаметру входного отверстия кока составляет 1,2. . . 3. 2. Газотурбинный двигатель по п. 1, отличающийся тем, что между стенками входного кока выполнена щелевая полость, соединенная радиальными каналами с внутренней полостью кока.

УСТРОЙСТВО СИЛОВОЙ УСТАНОВКИ И ТРАНСМИССИИ ДЛЯ ПРИВОДА АГРЕГАТОВ ВИНТОКРЫЛОГО ЛЕТАТЕЛЬНОГО АППАРАТА

... 1. Устройство (100) силовой установки и трансмиссии для привода агрегатов винтокрылого летательного аппарата, имеющего вал (112) несущего винта и каркас, снабженное газотурбинным двигателем (102) для вырабатывания механической энергии, установленным в корпусе (104) двигателя и имеющим ось двигателя, расположенную в основном горизонтально при нормальном режиме полета летательного аппарата, и единой трансмиссией (106), размещенной в корпусе (104) двигателя и непосредственно соединенной с упомянутым газотурбинным двигателем (102) с возможностью передачи механической энергии от упомянутого газотурбинного двигателя (102) к упомянутым агрегатам летательного аппарата, причем упомянутая трансмиссия установлена в корпусе (108) трансмиссии, а упомянутый корпус (108) двигателя и упомянутый корпус (104) трансмиссии объединены друг с другом в виде единого узла с возможностью установки на каркасе с размещением центра тяжести упомянутого газотурбинного двигателя (102) смежно с выходом упомянутой трансмиссии ...

МУЛЬТИПЛИКАТОР К ГАЗОТУРБИННОМУ ДВИГАТЕЛЮ

Изобретение относится к мультипликатору для газотурбинного двигателя. Его турбинное колесо представляет собой механическую передачу, состоящую из ведущего корпуса (6), на внешней окружной поверхности которого размещены турбинные лопатки (8). Внутренняя рабочая поверхность корпуса (6) выполнена в виде эпитрохоидального контура (7), очерченного вершинами ведомого трехуглового ротора. В роторе (4) соосно расположен кривошип (2), который выполнен эксцентрично по отношению к единому с ним стакану (3). Ось стакана (3) совпадает с центром эпитрохоидального контура (7). Радиусы стакана (3) и кривошипа (2) соотносятся как 2:3. Эксцентриситет составляет половину радиуса стакана (3). Достигается увеличение нагрузочной способности и долговечность устройства. 1 ил.

РЕДУКТОР С ЭПИЦИКЛОИДНОЙ ПЕРЕДАЧЕЙ ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ И СПОСОБ РАЗБОРКИ ПЕРЕДНЕЙ ЧАСТИ КОНСТРУКЦИИ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

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

РЕДУКТОР СКОРОСТИ ТУРБОМАШИНЫ

ПРИВОД ГАЗОТУРБИННОЙ УСТАНОВКИ

Привод газотурбинной установки, содержащий валы турбины заднего хода и турбины переднего хода, вал отбора мощности для привода винта фиксированного шага, установленный в корпусе двухступенчатый редуктор, включающий ведущую и ведомую шестерни, разобщительную муфту, при этом вал турбины переднего хода соединен с ведущей шестерней, а вал отбора мощности соединен с валом ведомой шестерни посредством разобщительной муфты, отличающийся тем, что, с целью повышения надежности и маневренности привода, редуктор содержит дополнительную ведущую шестерню, установленную на валу турбины заднего хода, дополнительную ведомую шестерню, соединенную с валом отбора мощности, дополнительную промежуточную шестерню, установленную между дополнительными ведомой и ведущей шестернями, дополнительную разобщительную муфту и гидродинамическую муфту, при этом вал дополнительной ведомой шестерни соединен с валом обора мощности посредством дополнительной разобщительной муфты, а вал турбины переднего хода соединен с валом ...

СПОСОБ ТОРМОЖЕНИЯ СИЛОВОЙ ТУРБИНЫ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Способ торможения силовой турбины газотурбинного двигателя, включающий перевод турбины в режим компрессора путем снижения расхода газа через турбину поворотом лопаток соплового аппарата и одновременного регулирования подачи топлива в камеру сгорания, отличающийся тем, что, с целью снижения расхода топлива, в период перевода турбины в режим компрессора подачу топлива прекращают, а газ с выхода турбины подают на вход компрессора.

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 ...

KRAFTUEBERTRAGUNGSVORRICHTUNG BZW. GETRIEBE ZUM STARTEN EINES VERBRAUCHERS

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 ...

Gasturbinenanlage

GASTURBINENANTRIEBSANLAGE

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.

Vorrichtung und Verfahren zur Überwachung eines Gleitlagers

Die Erfindung betrifft eine Vorrichtung zur Überwachung eines Gleitlagers (3) für ein Planetenrad (32) eines Getriebes (30) in einem Fangetriebe-Triebwerk (1), gekennzeichnet durch ein Detektionsmittel (4) für eine Relativbewegung zwischen einem Lagerbolzen (2) des Planetenrades (32) und einem Planetenträger (34) des Planetenrades (32), wobei durch das Detektionsmittel (4) bei Eintreten der Relativbewegung ein Steuersignal (S) zur Beeinflussung des Antriebes des Getriebes (30) abgebbar ist. Die Erfindung betrifft auch ein Verfahren zur Überwachung eines Gleitlagers.

Getriebeanordnung und Verfahren zu deren Herstellung

Eine Getriebeanordnung für ein Gasturbinentriebwerk (10) umfasst ein Planetengetriebe (30) mit zumindest einem Hohlrad (38) und zumindest einem Planetenrad (32), das beim Abrollen am Hohlrad (38) darauf eine Kraft in Richtung eines Kraftvektors (K) ausübt; und eine Halteeinrichtung (40; 40') zur Befestigung des zumindest einen Hohlrades (38) an einer anderen Struktur (24), mit einem ersten Abschnitt (40a), der sich in axialer Richtung auf einer Seite des Kraftvektors (K) und/oder einer geraden Verlängerung davon erstreckt, und mit einem zweiten Abschnitt (40b), der sich in axialer Richtung auf der anderen Seite des Kraftvektors (K) und/oder der geraden Verlängerung davon erstreckt. Ferner werden ein Gasturbinentriebwerk und ein Verfahren zur Herstellung einer Getriebeanordnung bereitgestellt.

Getriebeaufhängung

Andrehvorrichtung fuer Brennkraftmaschinen

A gas turbine generator

A gas turbine generator 1 comprises a combustor (2, fig. 1), a turbine 4 driven by the combustor, a compressor (6), a drive shaft 7 and an electrical generator 22 comprising a rotor 23 and a stator. The drive shaft links the turbine and the compressor along a common axis. The rotational axis of the generator rotor is radially offset from the drive axis, and a drive means is provided which connects the generator rotor to the drive shaft such that rotation of the drive shaft causes rotation of the generator rotor. The rotor may comprise a permanent magnet 28 secured to a rotor shaft 24. The permanent magnet may be partially surrounded by a sleeve 27 configured to urge the magnet against the rotor shaft. The rotational axis of the generator rotor may be orientated parallel to the drive axis. The drive means may comprise a drive gear 30 connected to and axially aligned with the drive shaft, and a rotor gear 34 connected to the generator rotor and linked to the drive gear.

Rotor assemblies for gas turbine engines

GEARING

... 1363151 Epicyclic gearing ROLLS-ROYCE (1971) Ltd 27 April 1972 [27 April 1971] 11725/71 Heading F2Q In epicyclic gearing comprising a sun gear 10, a plurality of planet pinions 11, and an annular gear 18, the annular gear 18 is relatively flexible and is surrounded by a relatively stiff backing- ring 43 with radial clearance therebetween in which a hydro-dynamic squeeze film of oil is formed. The oil is supplied to the clearance from a tube 46. The oil film opposes deformation of the annular gear 18 and damps vibration. Preferably the stiffness of the backing ring 43 is substantially seven times the stiffness of the annular gear 18.

Vehicle gas turbines

... 723,368. Gas-turbine plant. ROVER CO., Ltd. June 12, 1953 [June 23, 1952], No. 15724/52. Class 110 (3). [Also in Group XXXI] In a vehicle gas-turbine plant having a power turbine and a compressor turbine, the compressor has combined with it mechanism whereby the compressor may be driven by motion of the vehicle, the mechanism including a unidirectional clutch and a slipable friction coupling. Gases from the combustion chamber pass first to the compressor turbine g and then to the power turbine f, the latter driving the road wheels through the shaft and gearing k, m. A shaft q carrying a pinion r is driven from the gearing k and drives a shaft t through gearing u, s. The shaft t carries a disc v forming part of the slipable coupling, the other part of the coupling comprising an axially-slidable disc w which is loaded by a spring x. The two discs are contained in a hollow rotary member y secured to a sleeve 2. The sleeve is connected to a co-axial sleeve 4 by means of a unidirectional clutch ...

Gas turbine engine

... 1,075,846. Gas turbine engines. ROLLS ROYCE Ltd. May 20, 1966 [May 27, 1965], No. 25017/65. Heading F1G. A gas turbine engine comprises in flow series an axial-flow low pressure compressor 11, an axial-flow high pressure compressor 13, combustion equipment 14, an axial-flow high pressure turbine 15 and an axial-flow low pressure turbine 16, the H. P. compressor and the H. P. turbine being mounted on the first shaft 17, and the L. P. turbine being mounted on a second shaft 20 which is mounted concentrically within the first shaft 17. The L. P. compressor 11 is mounted on a shaft 12 disposed concentrically around the shaft 20. The shaft 20 is drivingly connected to a differential gear 23, Fig. 2 (not shown), which is connected to drive both the L. P. compressor 11 and the output shaft 22. The object of the invention is to facilitate starting of the engine, the L.P. turbine initially driving only the L. P. compressor but once the engine is self-sustaining the L. P. turbine drives both the ...

Reduction gear lubrication in a propfan turbo propulsion unit

A propfan turbine engine having a reduction gearing for the propfan rotors is provided with a lubricating system of its own for the reduction gearing, the system being arranged axially between the gas turbine and the reduction gearing and formed as an annular assembly surrounding a drive shaft from a gas turbine to the reduction gearing. Cooling air is supplied to the lubricating system as a location which will not interfere with the supply of air to the compressor of the gas turbine.

POWER TURBINE AND REDUCTION GEAR ASSEMBLY

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 ...

Limiting spool speeds in a gas turbine engine

A gas turbine engine 101 for an aircraft, comprises a high-pressure (HP) spool comprising an HP compressor and a first electric machine 110 driven by an HP turbine; a low-pressure (LP) spool comprising an LP compressor and a second electric machine driven 110 by an LP turbine; an engine controller 309 configured to identify a condition to the effect that the HP spool has reached or exceeded an HP speed limit whilst the LP spool has not reached an LP speed limit, and to operate the first electric machine in a generator mode of operation to reduce the HP spool speed below the HP speed limit. The engine controller may be further configured to operate the second electric machine in a motor mode of operation and transfer power electrically thereto from the first electric machine. The engine controller may be further configured to transfer power from the first electric machine to an energy storage unit such as a batter 305 or capacitor or to a consumer such as a de-icing system 304.

Supply of air to an air-conditioning circuit of an aircraft cabin

A turbo machine comprising a compressor system

Power transmission system for a gas turbine engine

... 1,001,639. Gas turbine engines. A. C. WICKMAN. June 30, 1964 [July 16, 1963], No. 28083/63. Headings FIG and F1Q. In a gas turbine engine of the type having a first turbine arranged to drive a compressor and a second turbine having a power output shaft, the first turbine is connected to drive the compressor through an overdrive epicyclic gearing the reaction member of which is driving- ly connected to a drive shaft, and a first unidirectional clutch is provided to prevent the first turbine from rotating faster than the compressor. A second set of gearing is provided to connect the drive shaft and the power out put shaft of the second turbine to a common power output shaft, the second set of gearing being so arranged that the power output shaft of the second turbine rotor will rotate at a ratio of, but slower than, the speed of the drive shaft, a second uni-directional clutch being arranged operatively between the reaction member of the first epicyclic gearing and the second gearing whereby ...

Multi-shaft gas turbine engine

A multi-shaft gas turbine engine (10, figure 1) has plural engine spools. A first spool of the engine and a second spool of the engine are operatively connected by an electrical machine that transfers power from one of the spools to the other. The electrical machine controls the operational performance of the engine. The electrical machine may comprise a first rotor 28 rotating with the first spool and supporting a circumferential row of magnets, a second rotor 30 rotating with the second spool and supporting a circumferential row of teeth carrying a row of coils. The teeth provide paths for a magnetic flux produced by the magnets and the coils when the first rotor rotates relative to the second rotor thereby transmitting torque between the spools. The coils may be controllably switchable. The electrical machine may have a driver part associated with the first spool, and a driven part associated with the second spool each driver and driven part comprising a respective rotor which rotates ...

Generating electrical power at high thrust conditions

A gas turbine engine for an aircraft comprising a high pressure (HP) spool, HP compressor 105, a first electric machine 117 having a first maximum output power and driven by HP turbine 107, low-pressure (LP) spool, a second electric machine 119 having a second maximum output power and driven by LP turbine 108, an engine controller 123 configured to identify a condition to the effect that the engine is in a maximum take-off mode of operation or a maximum climb mode of operation, and, in response to an electrical power demand being between zero and the second maximum output power, only extract electrical power from the second electric machine to meet the electrical power demand. Maximum take-off or maximum climb modes may be identified by comparing one or more of the altitude, the Mach number and the power lever angle. Both electric machines may be motor-generators.

... 1,058,515. Bearings. UNITED AIRCRAFT CORPORATION. June 11, 1965 [June 24, 1964], No.24684/65. Heading F2A. [Also in Division F1] The invention relates to means for supporting the bearing of the turbine shaft of a gas turbine ducted fan engine. The engine shown comprises an inner engine casing 14 and an outer engine casing 16 which together define a turbine discharge passage 18, also a fan casing 22 disposed outwardly of the outer engine casing and defining therewith a fan duct 24. A downstream rotor stage of the turbine is indicated at 12, the rotor being connected to a shaft 10 which is supported by a bearing comprising an inner race 34, an outer race 36 and rollers 40. The outer race 36 of the bearing is carried by the inner flange of a cylindrical housing member 38, the member 38 being of U-section with parts of the limbs of the U cut away as shown at 48, 50 so as to afford a series of hairpin springs. The outer flange 56 of the housing member is carried by a ring-shaped structure 61 ...

Gas turbine engine

... 1,055,328. Gas turbine engines. ROLLSROYCE Ltd. July 13, 1965 [July 20, 1964], No. 29417/64. Headings F1G, F1J and F1T. A gas turbine engine comprises a first shafting on which is mounted a compressor and a turbine of the engine, a second shafting on which is mounted a further turbine, there being means for maintaining the relative speed of the first and second shaftings substantially constant at all times. The invention is described with reference to a gas turbine ducted fan engine suitable for mounting at the tip of a helicopter rotor blade and comprises a compressor 14, which comprises axial-flow and centrifugal stages, combustion equipment 20 and a turbine 21 which drives the compressor through shaft 23. Downstream of the compressor drive turbine is a separate turbine 22 which is mounted on a shaft 24, and carries fan blades 25 at its periphery, the fan blades being disposed within a duct 26 formed between the outer casing 12 and inner casing 13. The two shafts 23, 24 are designed to ...

Improvements relating to gear mechanism for power drives

... 612,709. Differential gear; securing gears on shafts. POWER JETS (RESEARCH & DEVELOPMENT), Ltd., and McLEOD, R. C. Jan. 23, 1945, No. 1854. [Class 80 (ii)] [Also in Groups XXVI and XXXIII] In a gas turbine plant of the kind having a bladed propulsion means in the nature of a propeller or fan driven from the compressor end of the shaft connecting the turbine and compressor or compressors, the propeller or fan is driven through a reverted reduction gear train consisting of a central or sun pinion 20, intermediate or planet gears 21 and an internally-toothed annulus 22 arranged within an inner wall 12 of a forward facing annular inlet duct to the compressor, the central or sun pinion 20 being connected through a flexible coupling 37 to the compressor shaft and either the intermediate gears 21 or the toothed annulus 22 being coupled directly with the fan driving shaft which is coaxial with that of the compressor. If the gear is used to drive a contrarotating propeller, the intermediate gears ...

Safety coupling for shaft drive transmission

... 1,074,419. Positive clutches. GENERAL MOTORS CORPORATION. May 12, 1966 [June 11, 1965], No. 21167/66. Heading F2C. A transmission, such as a turboprop drive, comprises a positive clutch arranged to be disengaged upon the application of a predetermined reverse torque to the driven shaft, e.g. to permit limited windmilling. As shown the driving and driven shafts 10, 12 are connected by flanged members 32, 37 and an intermediate sleeve 50, the members 32, 37 being connected to the shafts 10, 12 by straight splines 26, 38, and to sleeve 50 by pairs of oppositely inclined splines 34, 52 and 39, 54. In the position shown for forward drive, sleeve 50 is urged to the right by the inclined splines and a loading spring comprising a stack of belleville discs 64, a pair of snap rings 58 bearing against the flanged member 32. Upon reversal of torque in shaft 12, sleeve 50 is moved to the left until a second pair of rings 60 bear against the flanged member 37, the belleville discs then engaging a radial ...

GEARBOX BREATHER OUTLET

INFLATABLE CURVED TUBULAR BODIES AND METHOD OF MANUFACTURE THEREOF

... 1283200 Inflatable craft INDUSTRIE PIRELLI SpA 1 Aug 1969 [3 Aug 1968] 38594/69 Heading B7A [Also in Division F2] An inflatable, curved tubular body, e.g. for an inflatable dinghy, as shown, comprises a tube of elastic and impermeable material, e.g. rubberized parallel cord fabric, with reinforcing elements 3 arranged to produce substantial inextensibility of the tube circumferentially, and further reinforcing elements 9 arranged to produce substantial inextensibility along the axis of the body, the further reinforcing elements being provided on selected portions of the body so that, when inflated, the body will take up a curved formation predetermined by the location of the reinforcing elements. As shown, body 7 also has reinforcing elements 10 which are applied after inflation of the body to stabilize its shape. In Fig. 8 (not shown), the body is formed as a straight tube with a curved end, two bodies being joined at their curved ends to form a dinghy, Figs. 10 and 11 (not shown). Methods ...

GAS TURBINE POWER PLANT

... 1361956 Gas turbine plant GENERAL MOTORS CORP 22 Feb 1973 [20 March 1972] 8692/73 Heading F1G The invention relates to a gas turbine power plant which may either drive a power output shaft or supply air under pressure for auxiliary purposes. The plant shown is suitable for driving a road vehicle or alternatively for supplying air under pressure, for example for unloading a granular cargo. The plant comprises an air compressor 2, air pre-heater 4, combustion equipment 6 and a first turbine 7 which is connected to drive the compressor through shaft 8; gases discharging from the turbine 7 pass through line 9 to a power turbine 10, gases from which discharge through line 16 to the air pre-heater 4 and then discharge through outlet 18. The power turbine is connected through a normally-engaged clutch 12 and transmission 14 to the road wheels 15. The clutch 12 is engaged by oil pressure supplied by pump 31 through line 34 and normallyopen valve 38 to the chamber 40. The compressor-drive turbine ...

A Geared gas turbine engine

GAS TURBINE POWER PLANT

Combustion turbine power units

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

Improvements relating to aircraft gas-turbine engines

... 790,550. Gas turbine plant. ROLLS-ROYCE, Ltd. Jan. 18, 1956 [Jan. 25, 1955], No. 2259/55. Class 110 (3). [Also in Group XXIV] In a compound gas turbine engine comprising independently rotatable low-pressure and highpressure rotor systems each comprising a compressor and compressor-driving turbine in which the low-pressure compressor delivers at least in part to the high-pressure compressor and the low-pressure turbine is connected to drive an airscrew, helicopter rotor or ducted fan, the low-pressure turbine is connected to the low-pressure compressor through a freewheel or equivalent device to permit the compressor to over-run the turbine, the airscrew, helicopter rotor or ducted fan is permanently connected to the low-pressure turbine and clutching means are provided for drivingly inter-connecting the high and low-pressure compressors. A lowpressure compressor 10, Fig. 1, feeding a highpressure compressor 11 is driven by a lowpressure turbine 13 through a shaft 15 and freewheel device ...

Improvements in differentially connected twin turbine power plant

... 819,489. Gas turbine driven vehicles. HUTCHINSON, D. W. May 25, 1956 [May 25, 1955], No. 16256/56. Classes 79(1), 79(2) and 79(3) [Also in Groups XXVI and XXXIII] A gas turbine power plant for driving a road vehicle, locomotive or ships propeller comprises a pair of gas turbines each of which has a drive shaft connected to compression means the discharge from which is heated and supplied to the turbines and differential gearing connecting the two turbine drive shafts and an output shaft whereby one drive shaft provides power to drive the output shaft and assists the other drive shaft to drive its compression means. The gas turbine power plant, Fig. 1, comprises a pair of gas turbine plants which drive locomotive wheels 60 through a differential gear 42, 44, 52 and a power shaft 58. One gas turbine plant comprises a compressor 4 the discharge of which is heated in an exhaust gas heat exchanger 14 before being further heated in a combustion chamber 18 feeding a turbine 6 driving the compressor ...

A bidirectional turbine drive for a shaft

A shaft drive alternately for both directions of rotation, comprises, a shaft rotatably supported on bearings having first and second turbines connectable thereto, each including a respective first and second turbine wheel freely rotatable on the shaft in respective opposite first and second directions. The turbine wheels are driven by fluid which rotates the wheels in a selected direction and, in addition, displaces the turbine wheels so that a friction disc carried thereby is engaged with a respective first and second clutch to connect it to the shaft to impart the selected direction of rotation. When the turbine is stopped by not directing the fluid into the blades, in which case the clutch mechanism is moved out of engagement and the turbine wheel may run free of the shaft, the second turbine wheel may be connected in a similar manner by effecting engagement of the associated second clutch with the shaft when the fluid is directed to the second turbine for rotating the turbine wheel ...

Improving deceleration of a gas turbine

A gas turbine engine for an aircraft comprises a high-pressure (HP) spool, HP compressor (105, figure 1), first electric machine (117) driven by an HP turbine (107); low-pressure (LP) spool, LP compressor (104), second electric machine (119) driven by an LP turbine (108), and an engine controller (123) configured to, in response to a change of a power lever angle setting indicative of an deceleration event, reduce fuel flow to a combustion system by a fuel metering unit, and operate the first electric machine in a generator mode to reduce the HP spool rotational speed and engine core mass flow. The controller may operate the second electric machine in a generator mode to further reduce engine mass flow. Electrical power may be supplied to an anti-icing system 304, battery 305, or capacitor. A method comprising identifying a condition to the effect that current fuel-air ratio in combustor is indicative of weak extinction onset, and extracting mechanical shaft power from HP spool to prevent ...

Rotor assembly

... 595,669. Securing together machine-parts for transmitting rotary motion. POWER JETS, Ltd., BONE, G. W., McLEOD, R. C., WATSON, K., and CALLISTER, T. E. April 27, 1942, No. 5653. [Class 80(ii)] [Also in Group XXVI] A centrifugal compressor rotor 23 and a turbine disc 10 are secured together by means accessible from the further side of the compressor. The compressor rotor 23 carries a quill-shaft 20 splined internally to engage external splines on a stubshaft 11 integral with the turbine rotor. The stub-shaft is secured in position by means of a screwed end 11B and a nut 26 which bears on a shoulder within the quill-shaft. The nut is part of a spindle 26B which passes through the rotor 23, is supported in a tearing plate 27 and can be rotated by means of external splines 26C. A shaft 30 provided for driving auxiliaries through a gear box 31 has internal splines 30A which engage the splines 26C and external splines 30C which engage splines on the hollow trunnion member 28 secured to the rotor ...

TRANSMISSIONS

... 1452721 Change-speed gear; friction clutches; clutch control GENERAL MOTORS CORP 27 Feb 1975 [7 March 1974] 8218/75 Headings F2C, F2D and F2L In a vehicle transmission, for use with an engine having a high no-load idle speed such as a two-spool gas turbine 10 with power transfer clutch between the compressor and load turbines 13, 22, for providing a controlled progression from idle to maximum torque capacity drive, and wherein a fluid coupling 72 and slippable friction starting clutch 84, the torquecapacity of which latter is varied during a starting phase, are arranged in parallel between an engine-driven input shaft 56 and an intermediate shaft 78 which drives a final output shaft 211 through, e.g. a four speed and reverse planet gear, providing a definite neutral and selectively operable to establish a starting drive, with the load initially stationary and the fluid coupling stalled at maximum torque, the capacity of the fluid coupling is substantially equal only to the idle torque of ...

Gear and electric amplification of generator motor compressor and turbine drives

Double-helical device, method for manufacturing a double-helical device and a holding jig

A double-helical gear device comprising a left-hand helical gear element 101 and a right-hand helical gear element 102. The two helical gear elements are connected coaxially through a connection element 110 positioned radially inwards or outwards from teethed surfaces 106 of the helical gear elements. A method for manufacturing a double-helical gear device, and a holding jig 120 to be used for the manufacturing of the double-helical gear device, are also disclosed.

ROTOR OF A TURBOMACHINE

Drive device for a hydrostatic transmission

System and method for oxidant compression in a stoichiometric exhaust gas recirculation gas turbine system

SYSTEM AND METHOD FOR OXIDANT COMPRESSION IN A STOICHIOMETRIC EXHAUST GAS RECIRCULATION GAS TURBINE SYSTEM [00212] A system includes a gas turbine system having a turbine combustor, a turbine driven by combustion products from the turbine combustor, and an exhaust gas compressor driven by the turbine. The exhaust gas compressor is configured to compress and supply an exhaust gas to the turbine combustor. The gas turbine system also has an exhaust gas recirculation (EGR) system. The EGR system is configured to recirculate the exhaust gas along an exhaust recirculation path from the turbine to the exhaust gas compressor. The system further includes a main oxidant compression system having one or more oxidant compressors. The one or more oxidant compressors are separate from the exhaust gas compressor, and the one or more oxidant compressors are configured to supply all compressed oxidant utilized by the turbine combustor in generating the combustion products. + 5/24 I-: C'", D ( Cn L U O ...

INTEGRATED OVERBOOST PROTECTION GEAR

GAS TURBINE ENGINE WITH A SINGLE OIL CAVITY

A gas turbine engine (10) having an oil cavity architecture and bearing placement which reduce heat rejection and oil system complexity by enclosing the reduction gearbox bearings (54) and at least the shaft bearings (42, 44) supporting the high pressure shaft (38) in the same oil cavity (60).

ROTOR COUPLING HAVING INSULATED STRUCTURE

The present invention provides a rotor coupling having insulated structure which can assuredly prevent galvanic corrosion of bearing members, rotors and the like which is cause by shaft voltage, in a shaft system that requires insulation in which a generator is disposed in between a steam turbine and a gas turbine or a rotating machinery such as another steam turbine. The rotor coupling having insulated structure of the present invention is employed in power generating equipment in which a generator is disposed in between a steam turbine and a gas turbine or a rotating machinery such as another steam turbine, a generator rotor and a rotating machinery rotor are connected, a first grounding electrode is provided to the steam turbine rotor, and a second grounding electrode is provided to the generator rotor, wherein both the rotors between the generator and the steam turbine are connected in an electrically insulated state.

DEVICE FOR EXTRACTING MECHANICAL POWER BETWEEN THE HP AND LP SHAFTS OF A TWIN-SPOOL TURBINE ENGINE

La présente invention porte sur un dispositif d'entraînement de machines auxiliaires d'un turbomoteur à double corps avec un arbre BP et un arbre HP, les dites machines étant installées sur un boîtier d'accessoires (7) comportant un arbre d'entraînement commun. Le dispositif est caractérisé par le fait qu'il comprend un engrenage différentiel (5) avec un premier (5a) et un deuxième (5b) arbres d'entrée et un arbre de sortie (5c), le premier arbre étant relié par un accouplement sélectif (8f-10f) soit à l'arbre BP soit à l'arbre HP, le deuxième arbre d'entrée (5b) étant relié à l'arbre HP, et l'arbre de sortie (5c) à l'arbre d'entraînement du boîtier (7).

GEARBOX DRIVE SHAFT OF TURBOJET ACCESSORY MACHINES; MODULAR SUPPLEMENTARY MACHINE

... ²²²²La présente invention porte sur un arbre (125) d'entraînement de boîtier ²à engrenages des machines auxiliaires dans un turboréacteur avec carter ²intermédiaire (119), agencé pour être monté dans un bras radial (119c) ²du carter intermédiaire (119), ledit arbre étant relié à une première ²extrémité à un moyen de transmission mécanique (126) avec un arbre ²moteur du turboréacteur, à une deuxième extrémité à un moyen de ²transmission mécanique avec ledit boîtier et comprenant un pignon ²conique (125a1) entre les deux extrémités permettant d'assurer une ²transmission mécanique avec un équipement auxiliaire supplémentaire ²(128). Cet arbre est caractérisé par le fait qu'ensemble avec ledit pignon ²(125a1) il est contenu dans une enveloppe (131) ménageant un circuit ²d'huile étanche par rapport audit bras (119c). L'invention porte aussi sur ²un équipement auxiliaire (128) comprenant un axe d'entraînement ²(128a) coopérant avec le pignon (125a1) de l'arbre de transmission ²(125) et formant ...

REMOTE SHAFT DRIVEN OPEN ROTOR PROPULSION SYSTEM WITH ELECTRICAL POWER GENERATION

A system for aircraft propulsion is disclosed herein. The system includes a power plant. The system also includes an open rotor module operable to rotate. The open rotor module has a plurality of variable-pitch blades. The system also includes a first linkage extending between the power plant and the open rotor module. The first linkage is operable to transmit rotational power to the open rotor module for rotating the plurality of variable-pitch blades. The system also includes an actuator operable to change a pitch of the plurality of variable-pitch blades. The system also includes a generator operable to generate electric power. The system also includes a second linkage extending between the power plant and the generator. The second linkage is operable to transmit rotational power to the generator. The generator is operable to convert the rotational power to electrical power. The system also includes a controller operably coupled to the actuator to vary a pitch of the plurality of variable-pitch ...

GAS TURBINE WITH WIRED SHAFT FORMING PART OF A GENERATOR/MOTOR ASSEMBLY

ENGINE AND TRANSMISSION POWER TRAIN

SPRING COMPENSATED RADIALLY FLEXIBLE POWER TAKEOFF SHAFT

A single shaft gas turbine engine of the type having a shaft assembly including, in axial alignment, a radial turbine element, a central slinger ring element, a radial compressor element, and a geared power takeoff element. The elements of the shaft assembly are interconnected by a tie bolt. An annular combustor liner defines a combustion chamber that encircles the central slinger ring. The shaft assembly is journaled by two bearings which are respectively located outboard of the turbine and on the geared power takeoff element. Each bearing is carried by a tubular support. Each tubular support has a free end carrying the respective bearing and another end secured to the bearing. The ends are spaced and interconnected by circumferentially spaced ribs to provide a flexible support for the prospective bearing. The gear on the power takeoff element drives a larger gear to withdraw power from the engine. A coil spring has one end seated on the housing and another end in contact with the free ...

DEVICE LINKING TWO ROTARY SHAFTS, SPECIFICALLY IN A TURBOMACHINE

HIGH-SPEED HIGH-POLE COUNT GENERATORS

A system for generating supplemental electrical power from the low-pressure (LP) turbine spool (30) of a turbofan engine (10) includes a high-speed, high magnetic pole count, generator (50), a gearbox (44), a controller (52) and a power converter (54). The LP turbine spool (30)is mechanically coupled to the generator (50) portion by the gearbox (44) for driving the generator portion (50). The controller portion (52) has a speed-sensing element for sensing the LP turbine speed. The controller portion (54) disables the power converter (54) when the generator (50) exceeds a predetermined speed, and enables the power converter (54) when the generator portion (50) is less or equal to the predetermined speed. The effective load on the generator (50) is reduced to approximately zero when the LP turbine spool (30) exceeds the predetermined speed, permitting the generator (50) to be electrically bound up to the predetermined speed and mechanically bound in excess of the predetermined speed.

POWER TAKE-OFF SYSTEM AND GAS TURBINE ENGINE ASSEMBLY INCLUDING SAME

A power take-off system (100) for a gas turbine engine assembly (10) is provided. The gas turbine engine (10) includes a first spool (40) and a second spool (42). The power take off system (100) further includes a starter (102) coupled to the first spool using a first shaft (32), and a generator (104) coupled to the second spool using a second shaft, the first shaft is circumferentially offset by the second shaft by an angle .alpha..

GAS TURBINE

Gas turbine (1), in particular for model aircraft, model helicopters and other small propulsion units, comprising a drive shaft (3), extending through an annular combustion chamber (2), rotatably mounted by means of two main bearings (13,13') and connected to a compressor rotor (4) and a turbine rotor (11) and a driven shaft (14) driven by the drive shaft (3). A device with a curvic gears is provided for torque transfer from the drive shaft (3) to the driven shaft (14), between the two main bearings (13, 13').

AIRCRAFT ENGINE REDUCTION GEARBOX

An aircraft engine reduction gearbox includes a power input and a power output, and an epicyclic gear train engaged with the power input and the power output. The epicyclic gear train includes a sun gear engaged with the power input and centrally disposed to define a center axis of the epicyclic gear train. Compound planet gears are mounted to a carrier and rotatable about respective planet gear axes. Each compound planet gear has an input gear in meshed engagement with the sun gear, and output gears axially spaced from the input gear. Ring gears are axially spaced apart and rotatable about the center axis. The ring gears are engaged with the power output. Each ring gear is in meshed engagement with one of the output gears.

AUXILIARY POWER UNIT

An auxiliary power unit includes a load compressor configured to generate compressed air for an environmental control system of an aircraft; a gas turbine engine drivingly coupled to the load compressor; and a conduit establishing fluid communication between the load compressor and an injection location in a gas path of the gas turbine engine to direct at least some of the compressed air generated by the load compressor to the injection location, the injection location being upstream of a turbine of the gas turbine engine.

LOCKNUT SUN GEAR FOR GAS TURBINE ENGINE

The present disclosure is directed to a shaft assembly (95) for a turbine engine (10), wherein the turbine engine (10) defines an axial direction and a radial direction, wherein the turbine engine (10) includes a fan or propeller assembly (14) and an engine core (20), and further wherein the fan or propeller assembly (14) includes a gearbox (45). The turbine engine (10) includes a coupling shaft (100), a spacer (150), a sleeve (200), and a nut (250) in adjacent radial arrangement. The coupling shaft (100) is connected at a first end (97) to the engine core (20) and coupled at a second end (96) to the gearbox (45). The coupling shaft (100) defines an annular surface (103) extended along the axial direction and a groove (104) extended in a circumferential direction. The spacer (150) defines a first portion (151) extended inward in the radial direction and a second portion (152) extended in the axial direction. The first portion (151) is disposed in the groove (104) of the coupling shaft ( ...

METHOD AND SYSTEM FOR DETECTING AND ACCOMMODATING LOSS OF A TORQUE SIGNAL

Systems and methods for detecting and accommodating for loss of a torque signal of a gas turbine engine are described herein. An engine deterioration offset may be determined while the torque signal of the engine is available. Then, in the event that the torque signal is lost, a predicted operating offset may be determined. A synthesized torque signal may be generated when the torque signal is lost at least in part from the engine deterioration offset and the predicted operating offset.

PLANETARY GEAR SYSTEM AND AIR TURBINE STARTER

An apparatus for an air turbine starter for an engine. The air turbine starter includes a housing defining an inlet, an outlet, and a flow path extending between the inlet and the outlet for communicating a flow of gas there through. A turbine member is journaled within the housing and disposed within the flow path for rotatably extracting mechanical power from the flow of gas and having a turbine output shaft. The air turbine starter further includes a planetary gear system drivingly coupled with the turbine output shaft and including a sun gear, a ring gear mounted to the housing, and a set of planetary gears operably coupling the sun gear and the ring gear with the sun gear is coupled to the turbine output shaft.

COMPLIANT JOURNAL BEARING SHAFT ASSEMBLY

CLAIMS: 1. A carrier assembly comprising: a gear carrier having a central axis and a pair of axially spaced-apart plates, the pair of axially spaced-apart plates having inward surfaces that are axially spaced apart to define an axial gap, the pair of axially spaced-part plates having a plurality of planetary bores on a plurality of planetary axes parallel to and radially outward from the central axis, the planetary bores having a planetary bore diameter; a plurality of shaft assemblies disposed on the planetary axes and mounted within the planetary bores of the gear carrier; wherein each shaft assembly includes: a journal bearing shaft mounted in the axial gap and having a pair of compliance grooves extending axially from opposed axial ends of the shaft, an inner cylindrical surface of each compliance groove defining a shaft mounting surface; and a pair of collars, each collar having a mounting socket mating the shaft mounting surface and an external collar surface matching the planetary ...

METHOD OF REPAIRING SPLINE AND SEAL TEETH OF A MATED COMPONENT

A method of repairing spline and seal teeth (20) of a mated component (10) is disclosed. The method includes a low energy input weld to deposit a repair material (50) on the non-worn, non-pressure face (34) of the splines (20) while minimizing the HAZ grain size. The splined area (12) of the component (10) is then remachined to the original spline contour by removing original material from the worn, pressure face (40) of the splines (20) and excess repair material (SO) to produce a mated component (10) with radially re-clocked splines (20') that have original component material on the pressure face (40') of the teeth (20').

TWO-SPOOL TURBOSHAFT ENGINE WITH MEANS OF POWER TAKE-OFF ON THE LOW PRESSURE AND HIGH PRESSURE ROTORS, POWER TAKE-OFF MODULE FOR THE TURBOSHAFT ENGINE AND PROCESS FOR MOUNTING THETURBOSHAFT ENGINE

L'invention concerne un turbomoteur à double corps, comportant un rotor haute pression (2) et un rotor basse pression (1), au moins un boîtier d'accessoires, un moyen d'entraînement, entraînant des arbres coaxiaux (11, 12) de transmission de mouvement vers le boîtier d'accessoires, caractérisé par le fait que le moyen d'entraînement comporte un pignon d'entraînement haute pression (9), solidaire du rotor haute pression (2) à proximité de son extrémité amont, un pignon d'entraînement basse pression (7), solidaire du rotor basse pression (1) en amont du rotor haute pression (2), et un module (29) de prise de mouvement, directement engrené avec les pignons d'entraînement, entraînant les arbres (11, 12) de transmission de mouvement. Grâce à l'invention, les arbres (11, 12) de transmission de mouvement s'étendent coaxialement l'un à l'autre et passent donc par un seul bras. L'utilisation d'un module (29) de prise de mouvement permet de simplifier le mécanisme. Le montage du turbomoteur est simplifié ...

OVERLAY FOR REPAIRING SPLINE AND SEAL TEETH OF A MATED COMPONENT

An overlay (600) for repairing spline and seal teeth (20) of a mated component (10) is disclosed. The overlay (600) is a tube constructed of a weld repair material (50), the tube having an outer surface (620) and an inner surface (625). The inner surface (625) is dimensioned to define an aperture (630) sized to receive a plurality of radially spaced splines (20) of the component (10), the splines (20) of the component (10) arranged to matingly engage a complementary component.

TURBINE ENGINE WITH SEMI-FIXED TURBINE DRIVING A RECEIVER CONTROLLED SO AS TO PRESERVE A ROUGHLY CONSTANT ROTATION SPEED

... ²The invention relates to a turbine engine ²(1) with a semi-fixed turbine, particularly for aircraft ²driving a receiver (2) controlled so as to preserve a ²roughly constant rotation speed. The turbine engine, in ²particular via a gear system (20), drives the receiver ²and an LP compressor (6) with an LP turbine (14). ²According to the invention, the gear system has a torque ²control system (26) maintaining a constant ratio between ²the drive torque of the receiver transmitted by the gear ²system and the drive torque of the LP compressor ²transmitted by this same gear system.² ...

GEARBOX PLANET SQUEEZE FILM DAMPER

An epicyclic gearing arrangement includes a planet gear rotatable on a planet bearing that is mounted via a support pin to a carrier of the epicyclic gearing arrangement. A spring film damper is disposed between the cylindrical outer surface of the support pin and the opposing inner surface of the inner ring of the planet bearing and includes an annular gap.

Energy efficient ips blower assembly

According to the invention a method and apparatus relating to a blower that is used in an environment of particle contaminated air, as with a helicopter engine, is disclosed. The blower has a power input, a fan, a power output attaching to the fan, and a clutch disposed between the power input and the power output whereby the power input may be selectively coupled to the power output to move the fan to vent the particle contaminated air. The clutch, which may be electrically, mechanically or hydraulically engaged, is activated by a user, a particle or altitude sensor, or by a full authority digital electronic controller (“FADEC”).

Drive mechanism for a pair of contra-rotating propellers through an epicyclic gear train

A turbine driving a planet gear and an epicyclic gear train and including a planet pinion cage and a ring driving two propellers in rotation, is connected to the planet gear through a flexible sleeve surrounding a turbine support shaft rather than through a support shaft itself to achieve a flexible assembly with a limit stop position in contact with the shaft to limit parasite internal forces applied to the epicyclic gear train without tolerating a loose assembly or breakage of the sleeve due to a condition of the limit stop after a clearance has been eliminated.

Gas turbine engine with pylon mounted accessory drive

A gas turbine engine includes an accessory gearbox within an engine pylon. The accessory components may be mounted within the engine pylon to save weight and space within the core nacelle as well as provide a relatively lower temperature operating environment.

Aircraft, propulsion system, and system for taxiing an aircraft

One embodiment of the present invention is a unique aircraft. Another embodiment is a unique aircraft propulsion system. Still another embodiment is a unique system for taxiing an aircraft without starting one or more main aircraft propulsion engines. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for aircraft taxiing and propulsion systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Dual fan gas turbine engine and gear train

A gas turbine engine includes a first fan, a second fan spaced axially from the first fan, a turbine-driven fan shaft and an epicyclic gear train coupled to be driven by the turbine-driven fan shaft and coupled to drive the first fan and the second fan. The epicyclic gear train includes a carrier that supports star gears that mesh with a sun gear, and a ring gear that surrounds and meshes with the star gears. The star gears are supported on respective journal bearings.

Efficient, low pressure ratio propulsor for gas turbine engines

A gas turbine engine includes a spool, a turbine coupled to drive the spool and a propulsor that is coupled to be driven by the turbine through the spool. A gear assembly is coupled between the propulsor and the spool such that rotation of the spool drives 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. The row includes no more than 16 of the propulsor blades.

Contained shaft spring load

An input shaft assembly is movable along an axis to absorb external impact loads. A biasing member exerts an axial load in a direction counter to potential impact loads. A stop is provided to control the application of biasing loads to control application of such axial load.

Generator and accessory gearbox device with a generator

The present invention proposes a generator for arrangement on a shaft of an accessory gearbox of an engine with a stator and with a rotor which can be coupled to a shaft of the accessory gearbox of the engine and which is rotatably mounted relative to the stator, where a stator area receiving the stator can be separated from a rotor area receiving the rotor. The rotor can be supplied with cooling medium in the rotor area. It furthermore proposes an accessory gearbox of an engine with a drive shaft operatively connectable to a main shaft of the engine and with at least one generator arranged on a shaft of the accessory gearbox.

GAS TURBINES

A gas turbine engine arrangement comprising a first engine section (), the first engine section () comprising a first compressor () and a first turbine () mounted on a first shaft (), the gas turbine engine arrangement further comprising at least one further turbine () mounted on a second shaft () and arranged such that gases exiting the first engine section () are ducted to the further turbine (), wherein said first and second shafts () are not mechanically coupled to one another and have respective axes which are offset from each other. 142.-. (canceled)43. A gas turbine engine arrangement comprising a first engine section , the first engine section comprising a first compressor and a first turbine mounted on a first shaft , the gas turbine engine arrangement further comprising at least one further turbine mounted on a second shaft and arranged such that gases exiting the first engine section are ducted to the at least one further turbine , wherein said first and second shafts are not mechanically coupled to one another and have respective axes which are offset from each other , wherein the further turbine is coupled to an electrical generator , and wherein the electrical generator is arranged to start the first engine section.44. The gas turbine engine arrangement of wherein the first and second shafts are parallel to one another.45. The gas turbine engine arrangement of wherein the first turbine and the at least one further turbine are each less than 100 mm in diameter.46. The gas turbine engine arrangement of wherein the second shaft is coupled to a fan.47. The gas turbine engine arrangement of wherein a duct arrangement is provided between the fan and the first engine section.48. The gas turbine engine arrangement of wherein the duct arrangement between the fan and the first engine section comprises a valve arrangement arranged selectively to permit or prevent ducting of air between said fan and said first engine section during use of the gas turbine engine ...

Motor-generator turbomachine starter

An exemplary turbomachine starter assembly includes a motor-generator that selectively operates in a motor mode or a generator mode. The motor-generator provides a rotational input to a turbomachine when operating in the motor mode. The motor-generator generates a supply of electrical power when operating in the generator mode. The supply of electrical power is used to power accessories of the turbomachine.

NON-LUBRICATED ARCHITECTURE FOR A TURBOSHAFT ENGINE

A turbine engine for a helicopter, the helicopter including a main gearbox, a rotor, and a speed-reducing device housed entirely within the main gearbox of the helicopter while also being connected to the rotor, the turbine engine including a casing, a gas generator with a gas generator shaft, and a free turbine for being driven in rotation by a gas stream generated by the gas generator, the free turbine including a free turbine shaft. When the turbine engine is fastened to the gearbox of the helicopter, the free turbine shaft extends axially into the main gearbox of the helicopter to be connected directly to the speed-reducing device. 112-. (canceled)13. A turbine engine not including a speed reducer and for connection to a helicopter , the helicopter including a main gearbox , a rotor , and a speed-reducing device housed entirely within the main gearbox of the helicopter while also being connected to the rotor , the turbine engine comprising:a casing;a gas generator with a gas generator shaft; anda free turbine for being driven in rotation by a gas stream generated by the gas generator, the free turbine having a free turbine shaft;wherein, when the turbine engine is fastened to the main gearbox of the helicopter, the free turbine shaft extends axially into the main gearbox of the helicopter to be connected directly to the speed-reducer device.14. A turbine engine according to claim 13 , further comprising at least one non-lubricated bearing arranged radially between the casing and the free turbine shaft.15. A turbine engine according to claim 13 , further comprising at least one non-lubricated bearing arranged radially between the gas generator shaft and the engine casing.16. A turbine engine according to claim 13 , further comprising an electricity generator directly connected to the gas generator shaft.17. A turbine engine according to claim 13 , further comprising a non-lubricated axial abutment device arranged axially between the gas generator shaft and the ...

GAS TURBINE ENGINE OIL BUFFERING

A turbine engine includes a shaft, a fan, at least one bearing mounted on the shaft and rotationally supporting the fan, a fan drive gear system coupled to drive the fan, a bearing compartment around the at least one bearing and a source of pressurized air in communication with a region outside of the bearing compartment. 1. A turbine engine comprising:a shaft;a fan;at least one bearing mounted on the shaft and rotationally supporting the fan;a fan drive gear system coupled to drive the fan;a bearing compartment around the at least one bearing; anda source of pressurized air in communication with a region outside of the bearing compartment.2. The turbine engine as recited in claim 1 , wherein the fan drive gear system includes an epicyclic gear train.3. The turbine engine as recited in claim 2 , wherein the epicyclic gear train has a gear reduction ratio of greater than or equal to about 2.3.4. The turbine engine as recited in claim 2 , wherein the epicyclic gear train has a gear reduction ratio of greater than or equal to 2.3.5. The turbine engine as recited in claim 2 , wherein the epicyclic gear train has a gear reduction ratio of greater than or equal to about 2.5.6. The turbine engine as recited in claim 2 , wherein the epicyclic gear train has a gear reduction ratio of greater than or equal to 2.5.7. The turbine engine as recited in claim 1 , wherein the fan defines a bypass ratio of greater than about ten (10) with regard to a bypass airflow and a core airflow.8. The turbine engine as recited in claim 1 , wherein the fan defines a bypass ratio of greater than 10.5:1 with regard to a bypass airflow and a core airflow.9. The turbine engine as recited in claim 1 , wherein the fan defines a bypass ratio of greater than ten (10) with regard to a bypass airflow and a core airflow.10. The turbine engine as recited in claim 1 , wherein the fan defines a pressure ratio that is less than about 1.45.11. The turbine engine as recited in claim 1 , wherein the fan defines ...

Split accessory drive system

A gas turbine engine includes a spool, a first accessory gearbox, a second accessory gearbox, and a scavenge pump. The first accessory gearbox is connected to and driven by the spool. The second accessory gearbox is connected to and driven by the first accessory gearbox. The scavenge pump is connected between the first accessory gearbox and the second accessory gearbox. The first accessory gearbox drives the second accessory gearbox through the scavenge pump.

Accessory gearbox device for a jet engine

The present invention proposes an accessory gearbox ( 9 ) for an engine with a drive shaft ( 12 ) operatively connectable to a main shaft of the engine, with an extension shaft ( 29 ) being provided which can be put into a detachable operative connection to the drive shaft ( 12 ) substantially coaxially to said drive shaft ( 12 ) of the accessory gearbox ( 9 ). At least one auxiliary unit ( 39 ) can be detachably arranged on the extension shaft ( 29 ).

Power Plant Line Having a Variable-Speed Pump

The invention relates to a power plant line, comprising a steam turbine and/or gas turbine that rotates at a constant speed in order to drive an electric generator; a variable-speed pump for conveying and/or compressing a working medium in order to drive and/or supply the process of the steam turbine and/or the gas turbine or to pump and/or compress an exhaust gas produced in the process supply or in the gas turbine. The invention is characterized in that the variable-speed pump is driven by the steam turbine and/or gas turbine and a speed-controllable gear train is arranged in the driving connection, said gear train having a power split, which comprises a mechanical main branch and a hydrodynamic secondary branch, wherein driving power is branched off from the mechanical main branch via a hydrodynamic coupling or a hydrodynamic converter by means of the hydrodynamic secondary branch and fed back to the mechanical main branch at the output side of the gear train in a variable-speed manner by means of a superimposing gear train.

Gas turbine engine with fan variable area nozzle for low fan pressure ratio

A gas turbine engine includes a fan section with twenty (20) or less fan blades and a fan pressure ratio less than about 1.45.

Accessory gearbox with tower shaft removal capability

An accessory system for a gas turbine engine includes an accessory gearbox which defines an accessory gearbox axis and includes first and second sides. A first geartrain includes one or more shafts rotatable about axes perpendicular to the first side of the accessory gearbox and a second geartrain includes one or more shafts rotatable about axes perpendicular to the second side of the accessory gearbox. A driven gear set defines an input axis and drives first geartrain and the second geartrain about corresponding first and second drive axes parallel to the input axis.

Coupling system for a star gear train in a gas turbine engine

A star gear train for use in a gas turbine engine includes a sun gear, a ring gear, a plurality of star gears and a coupling system. The sun gear is rotatable by a shaft. The ring gear is secured to a ring gear shaft. Each of the plurality of star gears is rotatably mounted in a star carrier and meshes with the sun gear and the ring gear. The coupling system comprises a sun gear flexible coupling, a carrier flexible coupling and a deflection limiter. The sun gear flexible coupling connects the sun gear to the shaft. The carrier flexible coupling connects the carrier to a non-rotating mechanical ground. The deflection limiter is connected to the star carrier to limit excessive radial and circumferential displacement of the star gear train.

Turbomachine drive arrangement

An arrangement and method for driving a turbomachine having a rotor is provided. The arrangement includes an input shaft rotationally coupled to the rotor. A motor generator device has a motor mode of operation and a generator mode of operation. A differential gear device having a first portion rotationally coupled to the input shaft and a second portion rotationally coupled to the motor generator device. A hydraulic assembly is rotationally coupled between the differential gear and the input shaft. The hydraulic assembly has a plurality of pistons and a pair of wobbler plates. A controller is operably coupled to the input shaft, the motor generator device and the hydraulic assembly, wherein the controller includes a processor that is responsive to executable instructions when executed on the processor for moving the pair of wobbler plates to a first position during a motor mode of operation.

GAS TURBINE ENGINE SHAFT BEARING CONFIGURATION

A gas turbine engine includes a core housing that has an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. The shaft supports a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. The shaft includes a main shaft and a flex shaft having bellows. The flex shaft is secured to the main shaft at a first end and includes a second end opposite the first end. A geared architecture is coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture. The geared architecture includes a sun gear supported on the second end. A first bearing supports the shaft relative to the intermediate case and a second bearing supports the shaft relative to the inlet case. The second bearing is arranged radially outward from the flex shaft. 1. A gas turbine engine comprising:a core housing including an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path;a shaft supporting a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path, wherein the shaft includes a main shaft and a flex shaft having bellows, the flex shaft secured to the main shaft at a first end and including a second end opposite the first end;a geared architecture coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture, wherein the geared architecture includes a sun gear supported on the second end;a first bearing supporting the shaft relative to the intermediate case; anda second bearing supporting the shaft relative to the inlet case, and the second bearing is arranged radially outward from the flex shaft.2. The gas turbine engine according to claim 1 , wherein the shaft includes a hub secured to the main shaft claim 1 , and the compressor section includes a rotor mounted to the hub claim 1 , the hub ...

GAS EXPANDER SYSTEM

A gas expander system suitable for use in a turbomachine, the gas expander system comprising: a gas expander provided with a moveable part; a magnetic gear arrangement; and a shaft; the moveable part of the gas expander being connectable to a load via the magnetic gear arrangement and the shaft, and movement of the moveable part of the gas expander being arranged to cause movement of the shaft, wherein the magnetic gear arrangement is used in a closed loop heat recovery system, with the inner and outer rotors of the magnetic gear separated by a wall that contains the stator. 1. A turbocharger system comprising:a turbocharger, comprising a turbine and a compressor;a gas expander system comprising:a gas expander provided with a moveable part;a magnetic gear arrangement; anda shaft;the moveable part of the gas expander being connectable to a load via the magnetic gear arrangement and the shaft, and movement of the moveable part of the gas expander being arranged to cause movement of the shaft,wherein the turbocharger system comprises a source of heat arranged to heat a working fluid provided in a closed-loop system to cause expansion of that working fluid to move the moveable part of the gas expanderwherein the magnetic gear comprises a first rotor and a second rotor, the first rotor and second rotor being separated from one another by at least a part of a wall forming part of the closed loop system.2. The system of wherein the at least a part of a wall forming part of the closed loop system that separates the first rotor from the second rotor comprises a stator of the magnetic gear.3. The system of wherein the wall encloses the closed loop system.4. The system of wherein the source of heat is provided by a part of the turbocharger system claim 1 , an engine to which the turbocharger is connected claim 1 , or a fluid flowing into or out of the turbocharger or the engine.5. The system of wherein the moveable part of the gas expander is located within the closed loop ...

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.

Gas turbine engine geared architecture axial retention arrangement

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

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.

ACTUATION SYSTEM FOR A PROPULSIVE UNIT OF AN AIRPLANE

The invention relates to an actuation system () for a propulsive unit including a nacelle, a turbojet engine, and an air circulation channel defined between the nacelle and the turbojet engine, the actuation system () including: a body and a connection device for attaching the body to the nacelle, the connection device enabling the body to move relative to the nacelle along two axes of rotation; a first member for moving a flap for controlling the airflow in the channel, the first member extending from the body and having a first point for attaching the first member to the lap; a second member for moving a thrust-reversing flap, the second member extending from the body and having a second point for attaching the second member to the flap, the body being arranged between the first and second attachment points; and a transmission device () having an input (), a first output () connected to the first member (), and a second output () connected to the second member (), the transmission device () being capable of selectively transmitting input movement towards the first output, in order to move the flow-controlling flap, or towards the second output, in order to move the thrust-reversing flap. 1100100. Actuation system () for a propulsion unit including a nacelle , a turbojet engine and an air circulation conduit defined between the nacelle and the turbojet engine , the actuation system () comprising:{'b': 10', '20', '10', '20, 'a body () and a connection device () for fixing the body () to the nacelle, the connection device () for fixing the body to the nacelle, the connection device allowing a movement of the body with respect to the nacelle on two rotation axes,'}{'b': 50', '60', '50', '10', '2', '60, 'a first member () for moving a flap () regulating an air flow circulating in the conduit, the first member () extending from the body () and having a first point () for fixing the first member to the flap (),'}{'b': 70', '80', '70', '10', '1', '80', '10', '2', '1, 'a ...

METHOD FOR PULLING A BEARING BODY OFF THE ROTOR OF A GAS TURBINE AND TUBULAR SHAFT EXTENSION

A method is provided for pulling a bearing body off the rotor of a gas turbine having a casing, while the casing is closed. The method includes fixing a shaft extension on the relevant end of the rotor. In order to free the bearing body of the weight of the rotor, the method involves supporting the rotor and/or holding the rotor. The method further includes fitting sliding elements between the bearing body and the rotor, and moving the bearing body axially along the machine axis onto the shaft extension. 111-. (canceled)12. A method for pulling a bearing body off the rotor of a gas turbine having a casing , while the casing is closed , the method comprising: supporting the rotor and/or', 'holding the rotor', 'in order to free the bearing body of the weight of the rotor,, 'fixing a shaft extension on the relevant end of the rotor, and'}fitting sliding elements between the bearing body and the rotor, andmoving the bearing body axially along the machine axis onto the shaft extension.13. The method as claimed in claim 12 ,wherein the bearing body is of undivided configuration in the circumferential direction.14. The method as claimed in claim 12 , further comprising:securing the bearing body against rotation during and after axial movement with the aid of an anti-rotation device.15. The method as claimed in claim 14 , further comprising:providing at least one guide extending in the axial direction as an anti-rotation device on the shaft extension, the guide limiting the tangential movement of a pin fixed on the bearing body when there is a risk of rotary motion of the bearing body.16. The method as claimed in claim 15 ,wherein the guide is embodied as a U-profile, along the outward-facing web of which the pin can slide with a small clearance.17. The method as claimed in claim 15 ,wherein the guide is embodied as a flat profile, the flat side of which extends perpendicularly to the radius of the shaft extension and parallel to the machine axis.18. The method as claimed ...

REDUCTION GEAR WITH EPICYCLIC GEAR TRAIN HAVING ROLLER-BEARING-MOUNTED PLANET SPINDLES

A reduction gear with epicyclic gear train comprising an annulus gear (), at least one planet gear () rolling on said annulus gear and able to rotate about a planet spindle () borne by a planet carrier (), wherein said planet spindle is mounted such that it can itself rotate about its axis relative to said planet carrier. The invention is suited to a reduction gear mounted in a high bypass ratio jet engine for driving the fan thereof. 1. A reduction gear with epicyclic gear train comprising an annulus gear , at least one planet gear rolling on said annulus gear and able to rotate about a planet spindle borne by a planet carrier , said planet spindle being mounted such that it can itself rotate about its axis relative to said planet carrier ,wherein there are two modes of operation: a first mode in which said planet spindle is rigidly attached to the planet carrier and a second mode in which said planet spindle can rotate freely relative to the planet carrier.2. The reduction gear as claimed in claim 1 , in which the planet gear rotates about its planet spindle by means of a journal bearing and the planet spindle is mounted on the planet carrier via the intermediary of a rolling bearing.3. The reduction gear as claimed in claim 1 , in which switching from one mode to the other is performed as a function of the temperature of the lubrication oil of said reduction gear.4. The reduction gear as claimed in claim 3 , in which switching from one mode to the other is actuated by an actuation means positioned so as to be in contact with said lubrication oil and comprising a shape memory material.5. The reduction gear as claimed in claim 4 , in which said actuation means is a flange plate in the shape of a circular annulus claim 4 , the central bore of which is attached to said spindle and the outer end of which moves in the axial direction as a function of the temperature of the lubrication oil.6. The reduction gear as claimed in claim 5 , in which said flange plate ...

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. 1. A gas turbine engine comprising:a fan shaft driving a fan;a frame which supports said fan shaft;a plurality of gears to drive said fan shaft;a flexible support which at least partially supports said plurality of gears, said flexible support having a lesser stiffness than said frame;a first turbine section providing a drive input into said plurality of gears; anda second turbine section,wherein said first turbine section has a first exit area at a first exit point and rotates at a first speed,wherein said second turbine section has a second exit area at a second exit point and rotates at a second speed, which is faster than the first speed,wherein a first performance quantity is defined as the product of the first speed squared and the first area,wherein a second performance quantity is defined as the product of the second speed squared and the second area, andwherein a ratio of the first performance quantity to the second performance quantity is between about 0.5 and about 1.5.2. The turbine section as set forth in claim 1 , wherein said ratio is above or equal to about 0.8.3. The turbine section as set forth in claim 1 , wherein said first turbine section has at least 3 stages.4. The turbine section as set forth in claim 1 , wherein said first turbine section has up to 6 stages.5. The turbine section as set forth in claim 1 , wherein said second turbine section has 2 or fewer stages.6. The turbine section as set forth in claim 1 , wherein a pressure ratio across the first turbine section is greater than about 5:1.7. The gas turbine engine as set forth in claim 1 , including a ratio of a thrust provided by said engine claim 1 , to a volume of a turbine section including both said high pressure turbine and said low pressure turbine being greater than or equal to about 1.5 and less than or equal to about 5.5 lbf/inch.8. The ...

Generating device for aircraft

An electric power generating device ( 1 ) capable of suppressing an increase of a frontal surface area of an aircraft engine includes a transmission ( 22 ) connected with a rotary shaft ( 9 ) of the engine (E), an electric power generator ( 34 ) driven by an output of the transmission ( 22 ), an input shaft ( 27 ) having a shaft axis extending in a direction crossing the rotary shaft ( 9 ) and connected with the rotary shaft ( 9 ), and a transmitting mechanism ( 21 ) connected with the input shaft ( 27 ) to drive the transmission ( 22 ) about an axis extending in a direction perpendicular to the input shaft ( 27 ). The transmission ( 22 ) and the electric power generator ( 34 ) are disposed spaced a distance from each other in a direction circumferentially of the rotary shaft ( 9 ).

Method for optimizing the speed of a twin-spool turbojet engine fan, and architecture for implementing same

A method and system improving energy efficiency of a turbojet engine by optimizing rotating speed of a fan and operability of an engine, by freeing the fan from exclusive control of a low-pressure (LP) shaft by providing combined control with a high-pressure (HP) shaft when cruising power has been reached. The turbojet engine include at least one LP turbine and one HP turbine coupled to coaxial LP shafts and HP shafts, respectively, which can drive LP and HP compressors, respectively. The LP compressors include a fan that forms a first primary air-intake compression stage. The LP and HP shafts are mounted on one of two driving mechanisms, an inner ring gear, and a planet carrier for a planetary gear train for driving the fan, the HP shaft being mounted on a disengagement mechanism and the fan being coupled to the planetary gear train via an outer driven ring gear.

GAS TURBINE ENGINE FAN DRIVE GEAR SYSTEM DAMPER

A gas turbine engine includes a fan section. A turbine section is coupled to the fan section via a geared architecture. The geared architecture includes a torque frame and a flex support spaced apart from one another at a location. A gear train is supported by the torque frame. A viscous damper is provided between the torque frame and the flex support at the location. 1. A gas turbine engine comprising:first and second members spaced apart from one another at a location;a gear train supported by the first member; anda damper provided between the first member and the second member at the location.2. The gas turbine engine according to claim 1 , wherein the first member is a torque frame claim 1 , and the second member is a flex support having a bellow claim 1 , the flex support grounded to a static structure.3. The gas turbine engine according to claim 2 , wherein the torque frame and flex support are secured to one another by fasteners in an area spaced radially inward from the location.4. The gas turbine engine according to claim 3 , wherein multiple viscous dampers are arranged circumferentially between the torque frame and the flex support claim 3 , and the bellow is provided between the fasteners and the viscous dampers.5. The gas turbine engine according to claim 4 , wherein the torque frame supports a carrier to which star gears are mounted claim 4 , a sun gear is arranged centrally relative to and intermeshing with the star gears claim 4 , and a ring gear circumscribing and intermeshing with the star gears.6. The gas turbine engine according to claim 5 , comprising a fan coupled to the ring gear claim 5 , and a low speed spool coupled to the sun gear.7. The gas turbine engine according to claim 2 , wherein the first and second members respectively include first and second apertures aligned with one another in an axial direction claim 2 , and wherein the damper is a viscous damper extending between and received in the first and second apertures.8. The gas ...

OFF-TAKE POWER RATIO

An example method of allocating power within a gas turbine engine includes driving an off-take power delivery assembly using a first amount of power from a spool, the first amount of power corresponding to an off-take power requirement of a gas turbine engine; and driving the spool of the gas turbine engine using a second amount of power, wherein a ratio of the first amount of power to the second amount of power is greater than or equal to 0.009. 1. A method of allocating power within a gas turbine engine , comprising:driving an off-take power delivery assembly using a first amount of power from a spool, the first amount of power corresponding to an off-take power requirement of a gas turbine engine; anddriving the spool of the gas turbine engine using a second amount of power, wherein a ratio of the first amount of power to the second amount of power is greater than or equal to 0.009.2. The method of claim 1 , wherein the off-take power delivery assembly comprises an accessory gearbox.3. The method of claim 2 , including rotating a tower shaft with the spool to provide power to the accessory gearbox.4. The method of claim 2 , wherein the accessory gearbox rotatably drives a generator to provide power to non-engine accessories using the first amount of power.5. The method of claim 1 , wherein the off-take power requirement corresponds to the gas turbine engine operating at an engine operating condition claim 1 , and the spool is configured to be driven using the second amount of power when the gas turbine engine is operating at the engine operating condition.6. The method of claim 1 , wherein the spool is a high speed spool and the gas turbine engine further comprises a low speed spool claim 1 , wherein the high speed spool is configured to rotate at higher speeds relative to the low speed spool during operation.7. The method of claim 1 , including delivering the first amount of power to non-engine accessories.8. A gas turbine engine comprising:a fan including a ...

SYSTEMS AND METHODS INVOLVING MULTIPLE TORQUE PATHS FOR GAS TURBINE ENGINES

A turbofan engine includes a fan, a compressor section, a combustor in fluid communication with the compressor section, a turbine section in fluid communication with the combustor, a shaft configured to be driven by the turbine section and coupled to the compressor section through a first torque load path, and a speed reduction mechanism configured to be driven by the shaft through a second torque load path separate from the first load path for rotating the fan. 1. A turbofan engine comprising:a fan;a compressor section;a combustor in fluid communication with the compressor section;a turbine section in fluid communication with the combustor;a shaft configured to be driven by the turbine section and coupled to the compressor section through a first torque load path; anda speed reduction mechanism configured to be driven by the shaft through a second torque load path separate from the first load path for rotating the fan.2. The turbofan engine as recited in claim 1 , including an intersection between the first torque load path and the shaft and a thrust bearing located adjacent to the intersection between the first torque path and the first shaft.3. The turbofan engine as recited in claim 1 , wherein the compressor section includes a first compressor section immediately aft of the fan and the first torque load path couples the shaft to the first compressor section.4. The turbofan engine as recited in claim 1 , including a first spool segment mechanically coupling the shaft to the compressor section and defining the first torque load path.5. The turbofan engine as recited in claim 4 , including a second spool segment mechanically coupling the shaft to the speed reduction mechanism and defining the second torque load path.6. The turbofan engine as recited in claim 5 , wherein the first spool segment is operative to transfer torque from the shaft to the compressor and not to the speed reduction mechanism.7. The turbofan engine as recited in claim 6 , wherein the second ...

OIL BAFFLE FOR GAS TURBINE ENGINE FAN DRIVE GEAR SYSTEM

An exemplary turbine engine assembly includes a first shaft that is rotatably driven by a second shaft of a gas turbine engine, a compressor hub driven by the second shaft, the compressor hub within a compressor section of the gas turbine engine, an epicyclic gear train that is driven by the first shaft, and a common attachment point that secures the first shaft and the compressor hub to the second shaft. 1. A turbine engine assembly comprising:a first shaft that is rotatably driven by a second shaft of a gas turbine engine;a compressor hub driven by the second shaft, the compressor hub within a compressor section of the gas turbine engine;an epicyclic gear train that is driven by the first shaft; anda common attachment point that secures the first shaft and the compressor hub to the second shaft.2. The turbine engine assembly of claim 1 , the epicyclic gear train having:a carrier, anda sun gear and intermediate gears arranged about and intermeshing with the sun gear, the intermediate gears supported by the carrier.3. The turbine engine assembly of claim 2 , including a baffle secured to the carrier by a fastening member claim 2 , the baffle including a lubrication passage near at least one of the sun gear and intermediate gears for directing lubricant on the at least one of the sun gear and intermediate gears.4. The turbine engine assembly of claim 2 , comprising a ring gear intermeshing with the intermediate gears and a third shaft interconnected to the ring gear claim 2 , and the first shaft interconnected to the sun gear.5. The turbine engine assembly of claim 4 , wherein the carrier is fixed relative to a housing claim 4 , the third shaft drives a fan claim 4 , and the first shaft supports a compressor hub having compressor blades.6. The turbine engine assembly of claim 1 , wherein the first shaft is a compressor shaft.7. The turbine engine assembly of claim 1 , wherein the epicyclic gear train rotatably drives a third shaft.8. The turbine engine assembly of ...

MID MOUNT SLEEVE ARRANGEMENT

A shaft and sleeve assembly of a gas turbine engine includes an outer shaft extending about an engine central longitudinal axis and having a radially inner shaft surface. The radially inner shaft surface has a shaft groove define therein. A sleeve is located radially inboard of the radially inner shaft surface. The sleeve includes a sleeve groove defined in a radially outer sleeve surface. The sleeve groove is formed at an axial location between a first axial end and a second axial end of the sleeve. A retaining element is installed in the shaft groove and the sleeve groove to retain the sleeve relative to the outer shaft. The sleeve includes one or more sleeve openings formed therein, and the sleeve groove is aligned with the one or more sleeve openings. 1. A shaft and sleeve assembly of a gas turbine engine comprises:an outer shaft extending about an engine central longitudinal axis and having a radially inner shaft surface, the radially inner shaft surface having a shaft groove define therein;a sleeve disposed radially inboard of the radially inner shaft surface, the sleeve including a sleeve groove defined in a radially outer sleeve surface, the sleeve groove formed at an axial location between a first axial end and a second axial end of the sleeve; anda retaining element installed in the shaft groove and the sleeve groove to retain the sleeve relative to the outer shaft;wherein the sleeve includes one or more sleeve openings formed therein; andwherein the sleeve groove is aligned with the one or more sleeve openings.2. The shaft and sleeve assembly of claim 1 , further comprising an axial stop formed in the radially inner shaft surface at which one of the first axial end or the second axial end of the sleeve are located.3. The shaft and sleeve assembly of claim 1 , wherein the retaining element is installable and/or removable via the one or more sleeve openings.4. The shaft and sleeve assembly of claim 1 , wherein the retaining element is one of a ring claim 1 ...

FLEXIBLE COUPLING FOR GEARED TURBINE ENGINE

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

ELECTRIC ACTUATOR FOR ENGINE CONTROL

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

FAN DRIVE GEAR SYSTEM SPLINE OIL LUBRICATION SCHEME

An input coupling for a fan drive gear system includes features for maintaining lubricant within a splined interface. The fan drive gear system includes a gear rotatable about an axis that includes an inner spline. The input coupling includes an outer spline engaged to the inner spline of the gear. The input coupling includes an aft oil dam for maintaining lubricant within an interface between the outer spline and the inner spline. 1. A fan drive gear system comprising:a gear rotatable about an axis, the gear including an inner spline; andan input coupling including an outer spline engaged to the inner spline of the gear, the input coupling including an aft oil dam for maintaining lubricant within an interface between the outer spline and the inner spline.2. The fan drive gear system as recited in claim 1 , wherein the gear includes a forward tab extending radially inward from an inner surface of the gear forward of the inner spline.3. The fan drive gear system as recited in claim 1 , wherein the gear includes an aft tab extending radially inward from an inner surface of the gear aft of the inner spline and forward of the aft oil dam.4. The fan drive gear system as recited in claim 1 , wherein the aft oil dam includes a retaining ring supported within an annular channel of the input coupling.5. The fan drive gear system as recited in claim 4 , wherein the retaining ring extends radially outward into contact with an inner surface of the gear.6. The fan drive gear system as recited in claim 1 , wherein the gear comprises a sun gear.7. The fan drive gear system as recited in claim 1 , wherein the input coupling includes at least one U-shaped portion for accommodating relative movement and misalignment with the gear.8. A geared turbofan engine comprising:a fan rotatable about an engine axis;a core engine including a turbine section driving a turbine shaft;a gearbox including a sun gear driven by the turbine shaft; andan input coupling transferring power between the ...

Fan Axial Containment System

A gas turbine engine () and method for containing a fan inside an engine after a fan thrust bearing assembly failure. The engine () may comprise a fan (), a housing () including a compartment (), a fan shaft () inside the compartment () and comprising a bowl (), a support structure () inside the compartment (), a speed sensor pickup () mounted on the outer surface () of the bowl (), a speed sensor () mounted on the support structure (), and a fan thrust bearing assembly () disposed forward of the bowl (). The fan thrust bearing assembly () including a bearing (). The speed sensor () and the sensor pickup () define a defining a sensor gap (). The bearing () and the outer surface () defining a fan thrust bearing gap (), wherein the sensor gap () is less than the fan thrust bearing gap. 1. A gas turbine engine disposed about a longitudinal engine axis (A) , the engine comprising:a fan;an exterior housing including an interior compartment disposed adjacent to the fan;a fan shaft disposed inside the compartment, the fan shaft configured to drive the fan, the fan shaft comprising an elongated pole and a bowl, the bowl including an outer surface and an inner surface;a fan bearing support structure disposed inside the compartment;a speed sensor pickup mounted on the outer surface of the bowl;a speed sensor mounted on the fan bearing support structure, the speed sensor and the sensor pickup defining a sensor gap in the axial direction; anda fan thrust bearing assembly disposed forward of the bowl, the fan thrust bearing assembly including a bearing, the bearing and the outer surface of the bowl defining in the axial direction a fan thrust bearing gap, wherein the sensor gap is less than the fan thrust bearing gap.2. The gas turbine engine of claim 1 , in which the fan thrust bearing assembly includes mounting hardware claim 1 , wherein the fan thrust bearing assembly is disposed adjacent to the pole and is mounted below the speed sensor to the fan bearing support structure ...

Compartment Shielding

A gas turbine engine having an engine axis and method of manufacturing the same is disclosed. The gas turbine engine may comprise a fan configured to drive air, a low pressure compressor section having a core flow path and configured to draw in and compress air flowing along the core flow path, a spool configured to drive the fan, and geared architecture configured to adjust the fan speed. The gas turbine engine may also include a housing defining a compartment that encloses the geared architecture. The housing is disposed between the core flow path and the axis, and includes a shielded mid-section that is in thermal communication with the core flow path of the low pressure compressor section. The shielded mid-section includes an outer layer and an insulator adjacent to the outer layer. 1. A gas turbine engine having an engine axis , the engine comprising:a gas generator that includes a core flow path;a propulsor that includes a fan and geared architecture for driving the fan; anda housing defining a first compartment that separates the geared architecture and the core flow path, the housing including a shielded axial mid-section that includes an outer layer and an insulator adjacent to the outer layer, the mid-section in thermal communication with the core flow path of the gas generator.2. The gas turbine engine of claim 1 , wherein the insulator is a foam insulator.3. The gas turbine engine of claim 1 , wherein the insulator is ceramic.4. The gas turbine engine of claim 1 , in which the mid-section further includes an inner layer claim 1 , wherein the insulator is disposed between the outer layer and the inner layer.5. The gas turbine engine of claim 4 , wherein the insulator is air.6. The gas turbine engine of claim 4 , wherein the insulator is air under vacuum pressure.7. The gas turbine engine of claim 1 , in which the mid-section further includes an inner layer and a second insulator claim 1 , wherein the first and second insulators are disposed between the ...

DRIVE SHAFT ASSEMBLY

A drive shaft assembly comprising a first shaft and a second shaft is disclosed. The second shaft is axially translatable from an engaged configuration with the first shaft at one limit of its travel to a disengaged configuration from the first shaft at the opposite limit of its travel. Engagement of the first and second shafts is via cooperating helical splines provided thereon. The helical splines give rise to a disengaging axial force on the second shaft in a fault condition in which a driving torque is applied from one of the first and second shafts to the helical splines in a rotational direction tending to unscrew the helical splines. Consequently when the second shaft is in the engaged configuration the disengaging axial force translates the second shaft to the disengaged configuration. 1. A drive shaft assembly comprising a first shaft and a second shaft , the second shaft being axially translatable from an engaged configuration with the first shaft at one limit of its travel to a disengaged configuration from the first shaft at an opposite limit of its travel , engagement of the first and second shafts being via cooperating helical splines provided thereon , said helical splines giving rise to a disengaging axial force on the second shaft in a fault condition in which a driving torque is applied from one of the first and second shafts to the helical splines in a rotational direction tending to unscrew the helical splines , such that when the second shaft is in the engaged configuration the disengaging axial force translates the second shaft to the disengaged configuration , the driving torque being a torque that exceeds any torque applied from the other of the first and second shafts to the helical splines.2. A drive shaft assembly according to where the helical splines give rise to an engaging axial force on the second shaft in a reset condition in which the driving torque is applied from one of the first and second shafts to the helical splines in a ...

Combined Two Engine Cycle With At Least One Recuperated Cycle Engine For Rotor Drive

A drive architecture comprises a rotor and a gearbox for driving the rotor. A pair of input gears provides rotational drive to the gearbox. A first recuperative cycle engine drives one of the pair of gears and a second engine drives the other of the pair of gears. The first recuperative cycle engine and the second engine are both gas turbine engines. A power takeoff from a drive shaft of the second engine supplies rotational drive to drive at least one component in the first recuperative cycle drive. 1. A drive architecture comprising:a rotor and a gearbox for driving said rotor;a pair of input gears for providing rotational drive to said gearbox and a first recuperative cycle engine driving one of said pair of gears and a second engine driving the other of said pair of gears; andsaid first recuperative cycle engine and said second engine both being gas turbine engines, with a power takeoff from a drive shaft of said second engine supplying rotational drive to drive at least one component in said first recuperative cycle drive.2. The drive architecture as set for in claim 1 , wherein said power takeoff from said second engine serves to provide rotational input to drive a compressor in said first recuperative cycle engine.3. The drive architecture as set for in claim 1 , wherein air downstream of said compressor in said first recuperative cycle engine is directed through a heat exchanger downstream of a turbine section in said first recuperative cycle engine claim 1 , where the air is heated and is then returned into a combustor section claim 1 , which is intermediate said compressor and said turbine section in said first recuperative cycle engine.4. The drive architecture as set for in claim 3 , wherein air is tapped from said second engine downstream of a compressor in said second engine and passed into a second heat exchanger where it additionally provides heat to the air from said compressor in said first recuperative cycle engine before the air in said first ...

Gas turbine engine with short transition duct

A turbine section for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan drive turbine including a fan drive duct, the fan drive turbine being configured to drive a fan section through a geared architecture at a speed that is less than an input speed to the geared architecture. A fan drive turbine includes a fan drive duct, the fan drive turbine being configured to drive a fan section through a geared architecture at a speed that is less than an input speed to the geared architecture. At least one upstream turbine is configured to drive at least one compressor. The at least one upstream turbine includes a turbine duct defining a conical flow path having a conical inlet defined by a first diameter and a conical outlet defined by a second diameter greater than the first diameter. The conical outlet is in fluid communication with the fan drive duct downstream of the conical outlet. At least one row of shrouded rotor blades defines at least a portion of the conical flow path. A method of designing a gas turbine engine is also disclosed.

HIGH EFFICIENCY AIRCRAFT PARALLEL HYBRID GAS TURBINE ELECTRIC PROPULSION SYSTEM

A gas turbine engine includes a compressor section having a first compressor and a second compressor, the second compressor having a higher pressure than the first compressor, and a turbine section having a first turbine and a second turbine, the second turbine having a higher pressure than the first turbine. The first compressor is connected to the first turbine via a first shaft. The second compressor is connected to the second turbine via a second shaft. An electric motor is connected to the first shaft such that rotational energy generated by the electric motor is translated to the first shaft. A fan is connected to the first shaft via a gear system. The gas turbine engine includes at least a takeoff mode of operations, a top of climb mode of operations and a maximum cruise mode of operations. The gas turbine engine is sized to operate at peak efficiency in the maximum cruise mode of operations. 1. A gas turbine engine comprising:a compressor section having a first compressor and a second compressor, the second compressor having a higher pressure than the first compressor;a turbine section having a first turbine and a second turbine, the second turbine having a higher pressure than the first turbine,the first compressor is connected to the first turbine via a first shaft;the second compressor is connected to the second turbine via a second shaft;an electric motor connected to the first shaft such that rotational energy generated by the electric motor is translated to the first shaft;a fan connected to the first shaft via a gear system; andwherein the gas turbine engine includes at least a takeoff mode of operations, a top of climb mode of operations and a maximum cruise mode of operations, the gas turbine engine being sized to operate at peak efficiency in said maximum cruise mode of operations.2. The gas turbine engine of claim 1 , wherein operating said engine at peak efficiency comprises operating said engine at a maximum turbine inlet temperature of said ...

DESCENT OPERATION FOR AN AIRCRAFT PARALLEL HYBRID GAS TURBINE ELECTRIC PROPULSION SYSTEM

A gas turbine engine includes a core having a compressor section with a first compressor and a second compressor, a turbine section with a first turbine and a second turbine, and a primary flowpath fluidly connecting the compressor section and the turbine section. The first compressor is connected to the first turbine via a first shaft, the second compressor is connected to the second turbine via a second shaft, and a motor is connected to the first shaft such that rotational energy generated by the motor is translated to the first shaft. The gas turbine engine includes a takeoff mode of operation, a top of climb mode of operation, and at least one additional mode of operation. The gas turbine engine is undersized relative to a thrust requirement in at least one of the takeoff mode of operation and the top of climb mode of operation, and a controller is configured to control the mode of operation of the gas turbine engine. 1. A gas turbine engine comprising:a core including a compressor section having a first compressor and a second compressor, a turbine section having a first turbine and a second turbine, and a primary flowpath fluidly connecting the compressor section and the turbine section;the first compressor is connected to the first turbine via a first shaft;the second compressor is connected to the second turbine via a second shaft;a motor connected to the first shaft such that rotational energy generated by the motor is translated to the first shaft;wherein the gas turbine engine includes a takeoff mode of operation, a top of climb mode of operation, and at least one additional mode of operation, and wherein the gas turbine engine is undersized relative to a thrust requirement in at least one of the takeoff mode of operation and the top of climb mode of operation; anda controller configured to control the mode of operation of the gas turbine engine.2. The gas turbine engine of claim 1 , wherein the at least one additional mode of operation includes a ...

GAS TURBINE

A gas turbine engine, in particular an aircraft engine, including: 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.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 bearing system is radially located between a fan shaft as part of the output shaft device and a static structure 1 , in particular a static front cone structure 1 , in particular the fan shaft bearing system being axially positioned within the width of the ...

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 static structure on the input side of the gearbox device, an inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device. The input shaft device having a high rigidity or the input shaft device having a means for decreasing the rigidity, in particular a diaphragm section. 1. A gas turbine 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, witha rear carrier bearing device radially between the planet carrier and a static structure on the input side of the gearbox device,an inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device.the input shaft device having a high rigidity.2. A gas turbine 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, witha rear carrier bearing device radially between the planet carrier and a static structure on the ...

Planetary gear drive and aircraft gas turbine with a planetary gear drive

The invention relates to a planetary gearbox having a ring gear support, in particular in a geared fan engine, characterized by a bolt connection for connecting the ring gear support to a static component, wherein the bolt connections are arranged on the circumference of the ring gear support in the axial direction of the planetary gearbox, and the bolt connections are designed and arranged in such a way that there is a frictional joint with a defined frictional force between the ring gear support and the static component and that a material separation, in particular a gap, is arranged between the ring gear support and the static component.

Carrier structure for an epicyclic gear drive, epicyclic gear drive and turbo engine with an epicyclic gear drive

A carrier structure for an epicyclic gear drive is provided. The carrier structure includes carrier elements connected with at least one planet gear and the first carrier element is connected with the second carrier element through at least two struts with the at least two struts having an inclination angle of more than 20° in the direction of a rotation around a rotation axis of the carrier structure.

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

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 static structure on the input side of the gearbox device or a front carrier bearing device radially between the planet carrier and a static structure on the output side of the gearbox device. 1. A gas turbine 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, witha rear carrier bearing device radially between the planet carrier and a static structure on the input side of the gearbox device ora front carrier bearing device radially between the planet carrier and a static structure on the output side of the gearbox device.2. The gas turbine of claim 1 , wherein the rear carrier bearing device or the front carrier bearing device comprise at least one roller bearing.3. The gas turbine of claim 1 , wherein the front carrier bearing device axially adjacent to the gearbox device on the output side.4. The gas turbine of claim 1 , wherein a fan shaft bearing system is radially located between a fan shaft as part of the output shaft device and a static structure.5. The gas turbine of claim 4 , wherein the fan shaft bearing system has an outer diameter between 0.05 to 0.2 the diameter of the propulsive fan claim 4 , in particular between 0.1 and 0.15 times the diameter of the propulsive fan.6. ...

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 ...

TURBOPROP COMPRISING AN INCORPORATED ELECTRICITY GENERATOR

An electrical generator is housed in an annular cavity between the casing and the propeller shaft of a turboprop, while imposing little or no additional space requirement and with lightweight ancillary equipment. The rotor of the generator is mounted on an autonomous shaft end. A flange of the outer casing is removable in order to access the generator and to enable its easy removal and remounting. 18-. (canceled)9. Turboprop , comprising a propeller , a propeller support shaft , an electricity generator located around the propeller support shaft , a casing surround the propeller support shaft and that also surrounds the generator , carries a stator of the generator , wherein the casing surrounding the generator also carries shaft end support bearings , a rotor of the generator being supported by the shaft end , and in that the casing comprises a removable end plate that carries one of said bearings , extending perpendicular to the propeller support shaft , and exposing an opening with a larger radius that the generator in the casing when said end plate is removed.10. Turboprop according to claim 9 , wherein the casing is conical around the generator.11. Turboprop according to claim 9 , further comprising a rotation speed converter extending as far as the propeller support shaft and connected to the shaft end.12. Turboprop according to claim 9 , wherein the generator is housed between a support roller bearing of the propeller support shaft by the casing and a toothed drive wheel of said propeller support shaft by a turbine shaft of the turboprop.13. Turboprop according to claim 11 , wherein the shaft end comprises one end separated from the rotor by bearings supporting the shaft end claim 11 , and a connector of the rotation speed converter to said end.14. Turboprop according to claim 13 , wherein said end comprises a system to decouple said connector from the generator.15. Turboprop according to claim 9 , wherein the generator is housed in a chamber containing at ...

TRANSFER COUPLINGS

A transfer coupling includes a static component and a rotatable component arranged concentrically, the static component including a first number of radially extending ports and the rotatable component having a second number of radially extending ports the radially extending ports arranged in a common circumferential plane wherein the ports are configured and arranged, in use, to homogenise a flow area for a fluid being transferred through the ports and thereby create a homogenous volume flow. 1. A transfer coupling comprising a static component and a rotatable component arranged in co-axial alignment , the static component including a first number of ports and the rotatable component including a second number of ports the ports arranged in a common circumferential plane wherein the ports are configured and arranged , in use , to homogenise a flow area for a fluid being transferred through the ports and thereby create a homogenous volume flow , and wherein a first of the static component and the rotatable component comprises a circumferential array comprising a first number of ports of a first size and the second of the components comprises a circumferential array comprising a second number of ports of a second size , the first number being substantially smaller than the second number and the first size being substantially larger than the second size.2. A transfer coupling as claimed in wherein the static component and rotatable component are concentrically arranged and the first and second number of ports extend radially.3. A transfer coupling as claimed in wherein the static component and rotatable component are arranged axially adjacent to each other and the ports extend axially.4. A transfer coupling as claimed in wherein the arrays of ports on the components have similar packing factors.5. A transfer coupling as claimed in wherein the second component has an array of ports comprising a single row of equally spaced circular ports having a first diameter D claim 1 ...

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.

ELECTRIC POWER GENERATING APPARATUS FOR USE IN AIRCRAFT

An electric power generating apparatus for use in an aircraft includes: a manual transmission configured to change speed of rotational power of an aircraft engine and including a plurality of gear stages; and an electric power generator to which the rotational power which has been changed in speed by the manual transmission is transmitted. The manual transmission includes: a planetary gear mechanism; an input shaft connected to a carrier holding a planetary gear of the planetary gear mechanism; an output shaft connected to a sun gear of the planetary gear mechanism; a one-way clutch which is sandwiched between the input shaft and the output shaft and by which the rotational power of the input shaft is transmitted to the output shaft; and a brake connected to a ring gear of the planetary gear mechanism. 1. An electric power generating apparatus for use in an aircraft ,the electric power generating apparatus comprising:a manual transmission configured to change speed of rotational power of an aircraft engine and including a plurality of gear stages; andan electric power generator to which the rotational power which has been changed in speed by the manual transmission is transmitted, wherein a planetary gear mechanism,', 'an input shaft connected to a carrier holding a planetary gear of the planetary gear mechanism,', 'an output shaft connected to a sun gear of the planetary gear mechanism,', 'a one-way clutch which is sandwiched between the input shaft and the output shaft and by which the rotational power of the input shaft is transmitted to the output shaft, and', 'a brake connected to a ring gear of the planetary gear mechanism., 'the manual transmission includes'}2. The electric power generating apparatus according to claim 1 , wherein the one-way clutch is arranged at a radially inner side of the ring gear.3. The electric power generating apparatus according to claim 1 , wherein the brake is provided on an outer peripheral surface of the ring gear.4. The electric ...

GAS TURBINE ENGINE FRONT ARCHITECTURE

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

Fbo torque reducing feature in fan shaft

A fan shaft includes a first end, a second end, and an axis, the fan shaft and configured to be coupled to a gear assembly of a gas turbine engine at the first end and to a fan more proximal to the second end than the first end such that in response to being coupled, the fan shaft can transfer torque from the gear assembly to the fan. At least one axial portion of the fan shaft satisfies the relationship 0.55 ≤ T * C J * τ ≤ 0.95 , where T represents peak torque during fan blade off, C represents a distance from a centerline of the gas turbine engine to an outer fiber of the fan shaft, J represents a polar moment of inertia of the fan shaft, and τ represents yield stress in shear of the fan shaft.

DUAL-FLOW TURBOJET ENGINE ARRANGEMENT WITH EPICYCLIC OR PLANETARY REDUCTION GEAR

A dual-flow turbojet engine having a central shaft surrounded by a high-pressure body which rotate about the same longitudinal axis while being independent in rotation, and including a fan driven by the central pressure shaft; a high-pressure compressor and a high-pressure turbine mounted on the high-pressure body; an inter-turbine casing; a low-pressure turbine mounted on a low-pressure rotor surrounding the central shaft; an exhaust casing on which an output cone is mounted; a reduction gear with which the low-pressure rotor drives the central pressure shaft; two bearings mounted on the exhaust casing and respectively receiving the central shaft and the low-pressure rotor; a bearing mounted on the inter-turbine casing and receiving the low-pressure rotor. 19-. (canceled)10. A dual-flow turbojet engine comprising a central shaft surrounded by a high-pressure spool , coaxial and rotatably independent , said turbojet engine including from upstream to downstream according to the direction of circulation of the flow that passes therethrough when it is operating:a fan driven by the central shaft;a high-pressure compressor and a high-pressure turbine belonging to the high-pressure spool;an inter-turbine casing;a low-pressure turbine;an exhaust casing; 'a low-pressure rotor which surrounds the central shaft and which comprises the low-pressure turbine;', 'said turbojet engine further includinga rotor upstream journal carried by the inter-turbine casing and which rotatably guides the low-pressure rotor while being located downstream of the high-pressure compressor;a rotor downstream journal carried by the exhaust casing, and which rotatably guides the low-pressure rotor;a reduction gear through which the low-pressure rotor drives the central shaft, said reduction gear being located downstream of the rotor downstream journal;a shaft downstream journal which rotatably guides the central shaft while being located downstream of the rotor downstream journal.11. The turbojet ...

Hybrid electric fan with stall free low pressure compressor

A gas turbine engine according to an exemplary embodiment of this disclosure includes among other possible things, a fan section including a plurality of fan blades, a first electric motor assembly that provides a first drive input for driving the fan blades about an axis, a turbine section, and a geared architecture driven by the turbine section and coupled to the fan section to provide a second drive input for driving the fan blades, and second electric motor assembly is coupled to rotate the geared architecture relative to a fixed structure controls a speed of the fan blades provided by a combination of the first drive input and the second drive input.

OIL BAFFLES IN CARRIER FOR A FAN DRIVE GEAR SYSTEM

A gearbox assembly for a gas turbofan engine includes a sun gear rotatable about an axis and a plurality of intermediate gears driven by the sun gear. A baffle disposed between at least two of the plurality of intermediate gears includes a first gap distance within a first gap portion and a second gap distance within a second gap portion. The first gap portion is disposed between the baffle and one of the intermediate gears away from the meshed interface with the sun gear and the second gap portion is disposed near the interface with the sun gear. The first gap distance within the first gap portion is different than the second gap distance within the second gap portion to define a desired lubricant flow path. 1. A gearbox assembly for a gas turbofan engine comprising:a sun gear rotatable about an axis;a plurality of intermediate gears driven by the sun gear;a baffle disposed between at least two of the plurality of intermediate gears, wherein the baffle is spaced a first gap distance from the at least two intermediate gears within a first gap portion and a second gap distance different than the first gap distance from one of the at least two intermediate gears within a second gap portion including an interface with the sun gear.2. The gearbox assembly as recited in claim 1 , wherein the second gap distance is larger than the first gap distance.3. The gearbox assembly as recited in claim 1 , wherein the second gap distance is between about 1.5 and about 2.5 greater than the first gap distance.4. The gearbox assembly as recited in claim 1 , wherein the sun gear includes a cavity and the baffle includes a wedge extending into the cavity for circulating lubricant out of the cavity.5. The gearbox assembly as recited in claim 1 , including a carrier supporting the intermediate gears relative to the sun gear and a ring gear circumscribing the intermediate gears claim 1 , wherein a ring gear baffle is supported on the carrier.6. The gearbox assembly as recited in claim 5 , ...

MANIFOLD FOR GAS TURBINE

In various embodiments, a manifold assembly () for conducting one or more fluids to a gear assembly () in a gas turbine engine () is provided. The manifold assembly () may comprise a first plate () and a second plate () that rotatably couple together. The manifold assembly () may be retained and/or held together by a channel () and engagement member () arrangement. 1. A manifold assembly , comprising:a first manifold comprising an engagement member;a second manifold having a groove defined therein, wherein the engagement member is installable in the groove; andthe manifold assembly configured to conduct a fluid to a gear assembly through the first manifold and the second manifold.2. The manifold assembly of claim 1 , further comprising an anti-rotation element.3. The manifold assembly of claim 2 , wherein the anti-rotation element is at least one of a fastener claim 2 , a pin claim 2 , an adhesive claim 2 , a tensioning device and a detent assembly.4. The manifold assembly of claim 1 , wherein the groove comprises a receiving portion and a retention portion.5. The manifold assembly of claim 1 , wherein the engagement member is a tongue.6. The manifold assembly of claim 1 , wherein the first manifold is rotatably coupled to the second manifold.7. A turbine engine claim 1 , comprising;a gear assembly; a first portion having a first groove and a second groove, wherein the first groove and the second groove are defined along a diameter of the first portion; and', 'a second portion having a first engagement member installable in the first groove and a second engagement member installable in the second groove., 'a manifold operatively coupled to and in fluid communication with the gear assembly, the manifold comprising8. The turbine engine of claim 7 , wherein the first groove further comprises a first groove portion that is defined radially outward from a centerline of the turbine engine.9. The turbine engine of claim 7 , wherein the engagement member comprises a shaft ...

GEARED TURBOFAN WITH INTEGRAL FRONT SUPPORT AND CARRIER

A gas turbine engine includes a fan section including a fan hub. A speed reduction device includes a star gear system. A turbine section is connected to the fan section through the speed reduction device. A first fan bearing for supporting rotation of the fan hub is connected forward of the speed reduction device. A second fan bearing for supporting rotation of the fan hub is connected aft of the speed reduction device. A first outer race of the first fan bearing is attached to the fan hub. 1. A gas turbine engine comprising:a fan section including a fan hub;a speed reduction device including a star gear system;a turbine section connected to the fan section through the speed reduction device;a first fan bearing for supporting rotation of the fan hub connected forward of the speed reduction device;a second fan bearing for supporting rotation of the fan hub connected aft of the speed reduction device; anda first outer race of the first fan bearing attached to the fan hub.2. The gas turbine engine of including a compressor section configured to rotate with the fan section.3. The gas turbine engine of including a first inner race of the first fan bearing connected to a static structure and a second inner race of the second fan bearing connected to a static structure.4. The gas turbine engine of wherein the first bearing and the second bearing include at least one of roller bearings claim 1 , ball bearings claim 1 , or tapered bearings.5. The gas turbine engine of including a high pressure compressor with a compression ratio of at least 20:1.6. The gas turbine engine of including a low pressure compressor with a compression ratio of at least 2:1.7. The gas turbine engine of including a fan by pass ratio greater than 10.8. The gas turbine engine of claim 1 , wherein the star gear system includes a sun gear claim 1 , star gears claim 1 , a ring gear mechanically attached to the fan section claim 1 , and a carrier fixed from rotation.9. The gas turbine engine of claim 8 , ...

OIL LOSS PROTECTION FOR A FAN DRIVE GEAR SYSTEM

A fan drive gear system includes at least one intermediate gear that includes an axial gear passage for receiving and conveying a fluid suitable for cooling and/or lubricating. At least a first axial end of the intermediate gear includes a first fluid storage trap for capturing fluid entering and/or exiting the gear passage and storing the fluid therein during powered operation of the fan drive gear system. The fluid is capable of being passively supplied to the intermediate gear passage during an interrupted power event. 1. A fan drive gear system comprising:at least one intermediate gear that includes an axial gear passage for receiving and conveying a fluid suitable for cooling and/or lubricating;at least a first axial end of said intermediate gear includes a first fluid storage trap for capturing fluid entering and/or exiting the gear passage and storing the fluid therein during powered operation of the fan drive gear system; andwhereby the fluid is capable of being passively supplied to the intermediate gear passage during an interrupted power event.2. The fan drive gear system of claim 1 , further comprising;a sun gear interfaced with said intermediate gear;a ring gear interfaced with said intermediate gear; anda carrier body supporting the intermediate gear.3. The fan drive gear system of claim 2 , wherein the at least one fluid trap comprises a radially outward base portion relative to an axis defined by said carrier body and at least one radially inward base portion relative to said axis defined by said carrier body.4. The fan drive gear system of claim 3 , wherein said radially outward base portion is defined on a first axial end by a radially aligned wall segment of said carrier body relative to the axis defined by the carrier body and said radially outward base portion is defined on a second axial end by a radially aligned wall of said trap.5. The fan drive gear system of claim 4 , wherein said radially aligned wall of said fluid storage trap extends ...

GAS TURBINE ENGINE WITH FAN VARIABLE AREA NOZZLE FOR LOW FAN PRESSURE RATIO

A gas turbine engine includes a fan section with twenty (20) or less fan blades and a fan pressure ratio less than about 1.45. 1. A gas turbine engine comprising:a core nacelle defined about an engine centerline axis;a core engine at least partially disposed within the core nacelle;a fan section with twenty (20) or less fan blades;a gear system driven by the core engine to drive said fan section;a fan nacelle mounted at least partially around said fan section and said core nacelle to define a fan bypass flow path for a fan bypass airflow, said fan bypass airflow having a fan pressure ratio of the fan bypass airflow during engine operation, said fan pressure ratio less than about 1.45;a variable fan nozzle axially movable relative to the fan nacelle, the variable fan nozzle including at least two sectors; anda controller for independently adjusting each of the at least two sectors.2. (canceled)3. The engine as recited in claim 1 , wherein the controller is operable to reduce said fan nozzle exit area at a cruise flight condition.4. The engine as recited in claim 1 , wherein said controller is operable to control said fan nozzle exit area to reduce a fan instability.5. The engine as recited in claim 1 , wherein said fan variable area nozzle defines a trailing edge of said fan nacelle.6. The engine as recited in claim 1 , wherein said fan variable area nozzle is axially movable relative to said fan nacelle.7. (canceled)8. The engine as recited in claim 1 , wherein said fan section defines a corrected fan tip speed less than about 1150 ft/second.9. The engine as recited in claim 1 , wherein said core engine includes a low pressure turbine which defines a low pressure turbine pressure ratio that is greater than about five (5).10. The engine as recited in claim 7 , wherein said core engine includes a low pressure turbine which defines a low pressure turbine pressure ratio that is greater than five (5).11. The engine as recited in claim 1 , further comprising a gear system ...

PINNED MECHANICAL FUSE FOR ENGINE MOTORING SYSTEM

A motoring system for a gas turbine engine having: a reduction gear train having an input and output; a motor operably connected to the input; a clutch operably connected to the output, the clutch in operation engages and disengages the reduction gear train; and a pinned mechanical fuse operably connecting the output to the clutch, the pinned mechanical fuse having at least one shear pin. The pinned mechanical fuse having: an outer sleeve having a first section, second section, inner chamber, outer wall, and at least one through hole connecting the inner chamber to the outer wall within the first section; and an inner sleeve having a first portion, second portion, outer surface, and at least one blind hole located in the outer surface within the second portion. The second portion being located within the inner chamber and operably connected to the outer sleeve through at least one shear pin. 1. A motoring system for a gas turbine engine comprising:a reduction gear train having an input and an output;a motor operably connected to the input;a clutch operably connected to the output, the clutch in operation engages and disengages the reduction gear train; and an outer sleeve having a first section, a second section, an inner chamber, an outer wall, and at least one through hole connecting the inner chamber to the outer wall within the first section; and', 'an inner sleeve having a first portion, a second portion, an outer surface, and at least one blind hole located in the outer surface within the second portion, the second portion being located within the inner chamber and operably connected to the outer sleeve through the at least one shear pin,', 'wherein the at least one through hole is aligned with the at least one blind hole,', 'wherein the at least one shear pin is secured within the at least one through hole and the at least one blind hole, and', 'wherein the at least one shear pin in operation shears when torque on the pinned mechanical fuse is greater than or ...

GAS TURBINE ENGINE STARTER REDUCTION GEAR TRAIN WITH STACKED PLANETARY GEAR SYSTEMS

According to an aspect, a system for a gas turbine engine includes a reduction gear train operable to drive rotation of a starter gear train that interfaces to an accessory gearbox of the gas turbine engine. The reduction gear train includes a starter interface gear that engages the starter gear train, a core-turning clutch operably connected to the starter interface gear, and a plurality of stacked planetary gear systems operably connected to the core-turning clutch and a core-turning input. The system also includes a mounting pad including an interface to couple a core-turning motor to the core-turning input of the reduction gear train. 1. A system for a gas turbine engine comprising: a starter interface gear that engages the starter gear train;', 'a core-turning clutch operably connected to the starter interface gear; and', 'a plurality of stacked planetary gear systems operably connected to the core-turning clutch and a core-turning input; and, 'a reduction gear train operable to drive rotation of a starter gear train that interfaces to an accessory gearbox of the gas turbine engine, the reduction gear train comprisinga mounting pad comprising an interface to couple a core-turning motor to the core-turning input of the reduction gear train.2. The system of claim 1 , wherein the starter interface gear engages a planet gear of the starter gear train claim 1 , and the starter gear train is operably connected to the accessory gearbox through a starter clutch.3. The system of claim 1 , wherein the core-turning clutch is an overrunning clutch.4. The system of claim 1 , wherein the stacked planetary gear systems further comprise a stacked series of coaxially aligned sun gears that each drives a plurality of planet gears.5. The system of claim 4 , wherein a distal sun gear is operably connected by a drive shaft to the core-turning input claim 4 , and the distal sun gear is operably connected to a first set of the planet gears of the stacked planetary gear systems.6. The ...

GAS TURBINE ENGINE STARTER REDUCTION GEAR TRAIN WITH GEARED ROTARY ACTUATOR

According to an aspect, a system for a gas turbine engine includes a reduction gear train operable to drive rotation of a starter gear train that interfaces to an accessory gearbox of the gas turbine engine. The reduction gear train includes a starter interface gear that engages the starter gear train and a core-turning clutch operably connected to the starter interface gear. The reduction gear train also includes a geared rotary actuator including a primary planetary gear system, where the geared rotary actuator is operably connected to the core-turning clutch. The reduction gear train further includes a secondary planetary gear system operably connected to the primary planetary gear system and a core-turning input. The system also includes a mounting pad with an interface to couple a core-turning motor to the core-turning input of the reduction gear train. 1. A system for a gas turbine engine comprising: a starter interface gear that engages the starter gear train;', 'a core-turning clutch operably connected to the starter interface gear;', 'a geared rotary actuator comprising a primary planetary gear system, the geared rotary actuator operably connected to the core-turning clutch; and', 'a secondary planetary gear system operably connected to the primary planetary gear system and a core-turning input; and, 'a reduction gear train operable to drive rotation of a starter gear train that interfaces to an accessory gearbox of the gas turbine engine, the reduction gear train comprisinga mounting pad comprising an interface to couple a core-turning motor to the core-turning input of the reduction gear train.2. The system of claim 1 , wherein the starter interface gear engages a planet gear of the starter gear train claim 1 , and the starter gear train is operably connected to the accessory gearbox through a starter clutch.3. The system of claim 1 , wherein the core-turning clutch is an overrunning clutch.4. The system of claim 1 , wherein the geared rotary actuator ...

GEARED GAS TURBINE ENGINE AND A GEARBOX

A gas turbine engine comprises a gearbox comprising a sun gear, an annulus gear, a plurality of planet gears and a carrier. The carrier comprises a primary structure and at least one reinforcing structure. The primary structure comprises a first material and the at least one reinforcing structure comprises a second material. The primary structure includes a first ring, a second ring spaced axially from the first ring and a plurality of circumferentially spaced axles extending axially between the first ring and the second ring. Each planet gear is rotatably mounted on a respective one of the axles by a bearing. The reinforcing structure is secured to the primary structure and the reinforcing structure comprises a particulate reinforced material or a fibre reinforced material. The reinforcing structure increases the stiffness of the carrier and reduces the weight of the carrier. 1. A gas turbine engine comprising a gearbox , the gearbox comprising a sun gear , an annulus gear , a plurality of planet gears and a carrier , the sun gear meshing with the planet gears and the planet gears meshing with the annulus gear , the carrier comprising a primary structure and at least one reinforcing structure , the primary structure comprising a first material and the at least one reinforcing structure comprising a second material , the primary structure comprising a first ring , a second ring spaced axially from the first ring and a plurality of circumferentially spaced axles extending axially between the first ring and the second ring , each planet gear being rotatably mounted on a respective one of the axles , the at least one reinforcing structure being secured to the primary structure , and the at least one reinforcing structure comprising a reinforced material , the reinforced material being selected from the group consisting essentially of a particulate reinforced material and a fibre reinforced material.2. A gas turbine engine as claimed in wherein the at least one ...

GAS TURBINE ENGINE WITH AXIAL MOVABLE FAN VARIABLE AREA NOZZLE

A method of designing a turbofan engine according to an exemplary aspect of the present disclosure includes, among other things, providing a fan section including a plurality of fan blades, providing a low pressure turbine driving the plurality of fan blades through a gear train, providing a fan nacelle and a core nacelle, the fan nacelle at least partially surrounding the core nacelle, providing a fan bypass flow path defined between the core nacelle and the fan nacelle, and providing a fan variable area nozzle in communication with the fan bypass flow path and defining a fan nozzle exit area between the fan nacelle and the core nacelle. 1. A method of designing a turbofan engine comprising:providing a fan section including a plurality of fan blades, the plurality of fan blades having a design angle of incidence;providing a low pressure turbine driving the plurality of fan blades through a gear train, the low pressure turbine having a pressure ratio greater than 5:1, and the gear train having a gear reduction ratio of greater than 2.5:1;providing a fan nacelle and a core nacelle, the fan nacelle at least partially surrounding the core nacelle;providing a fan bypass flow path defined between the core nacelle and the fan nacelle, and a bypass ratio greater than 10:1;providing a fan variable area nozzle in communication with the fan bypass flow path and defining a fan nozzle exit area between the fan nacelle and the core nacelle;providing a controller in communication with the fan variable area nozzle; andcausing the fan variable area nozzle to vary the fan nozzle exit area to adjust fan bypass air flow in the fan bypass flow path in a plurality of flight conditions in response to the controller such that an angle of incidence on the plurality of fan blades in the plurality of flight conditions is maintained close to the design angle of incidence of the plurality of fan blades.2. The method as recited in claim 1 , further comprising causing the fan variable nozzle to ...

MECHANICAL SHEAR FUSE FOR ENGINE MOTORING SYSTEM

A motoring system for a gas turbine engine having: a reduction gear train having an input and an output; an electric motor operably connected to the input; a clutch operably connected to the output, the clutch in operation engages and disengages the reduction gear train; and a mechanical shaft fuse operably connecting the output to the clutch, the mechanical shaft fuse in operation shears when torque on the mechanical shaft fuse is greater than or equal to a selected value. The mechanical shaft fuse includes a plurality of through holes. 1. A motoring system for a gas turbine engine comprising:a reduction gear train having an input and an output;an electric motor operably connected to the input;a clutch operably connected to the output, the clutch in operation engages and disengages the reduction gear train; anda mechanical shaft fuse operably connecting the output to the clutch, the mechanical shaft fuse in operation shears when torque on the mechanical shaft fuse is greater than or equal to a selected value,wherein the mechanical shaft fuse includes a plurality of through holes.2. The motoring system of claim 1 , wherein:the plurality of through holes are oriented around an approximate axial center point of the mechanical shaft fuse.3. The motoring system of claim 1 , wherein:each of the holes has a diameter of about 0.187 inches (0.475 centimeters).4. The motoring system of claim 1 , wherein:the plurality of holes comprises six holes.5. The motoring system of claim 1 , wherein:the selected value is about 64 foot-pounds (87 newton-meters).6. The motoring system of claim 1 , wherein:the mechanical shaft fuse includes a first outer diameter of about 0.63 inches (1.6 centimeters).7. The motoring system of claim 1 , wherein:the mechanical shaft fuse is hollow and includes a thickness of about 0.09 inches (0.229 centimeters).8. The motoring system of claim 1 , wherein:the mechanical shaft fuse has a hexagonal cross-sectional shape.9. The motoring system of claim 6 , ...

Ring gear mounting arrangement with oil scavenge scheme

An epicyclic gear train for a turbine engine includes a gutter with an annular channel. A rotating structure includes a ring gear that has an aperture that is axially aligned with the annular channel. Axially spaced apart walls extend radially outward relative to the rotating structure to define a passageway. The passageway is arranged radially between and axially aligned with the aperture and the annular channel. The walls are configured to inhibit an axial flow of an oil passing from the aperture toward the annular channel.

GAS TURBINE ENGINE ELECTRICAL GENERATOR

An aircraft gas turbine engine () comprises a main engine shaft () arranged to couple a turbine () and a compressor (), the main engine shaft () defining an axial direction (). The gas turbine engine () further comprises at least one radially extending offtake shaft () coupled to the main engine shaft (), and a radially extending electric machine () coupled to the radially extending offtake shaft (). 1. An aircraft gas turbine engine comprising:a main engine shaft arranged to couple a turbine and a compressor, the main engine shaft defining an axial direction;at least one radially extending offtake shaft coupled to the main engine shaft; anda radially extending electric machine coupled to the radially extending offtake shaft.2. A gas turbine engine according to claim 1 , wherein the electric machine is directly coupled to the radially extending offtake shaft.3. A gas turbine engine according to claim 1 , wherein the electric machine is coupled to the radially extending offtake shaft by a reduction gearbox.4. A gas turbine engine according to claim 1 , wherein the electric machine comprises at least one of an electric motor configured to provide motive power to start the gas turbine engine in a starting mode claim 1 , and a generator configured to generate electrical power when in a running mode.5. A gas turbine engine according to claim 1 , wherein the electric machine comprises an axial flux electric machine in which a stator of the electric machine is axially offset relative to a rotor of the electric machine.6. A gas turbine engine according to claim 1 , wherein the electric machine comprises a radial flux electric machine claim 1 , in which a stator of the electric machine is radially inward or radially outward of a rotor of the electric machine.7. A gas turbine engine according to claim 1 , wherein the gas turbine engine comprises a plurality of radially extending offtake shafts circumferentially arrayed around the main engine shaft.8. A gas turbine engine ...

GEARED TURBOFAN WITH FOUR STAR/PLANETARY GEAR REDUCTION

A turbofan engine assembly includes a nacelle and a turbofan engine. The turbofan engine includes a fan, which includes a fan rotor having fan blades, and a nacelle enclosing the fan rotor and the blades. A turbine rotor drives the fan rotor. An epicyclic gear reduction is positioned between the fan rotor and the turbine rotor. The epicyclic gear reduction includes a ring gear, a sun gear, and four intermediate gears that engage the sun gear and the ring gear. A gear ratio of the gear reduction is greater than 3.06. The fan drive turbine drives the sun gear to, in turn, drive the fan rotor. 1. A turbofan engine comprising:a fan including a fan rotor having fan blades, and housing enclosing said fan rotor and said blades;a turbine rotor driving said fan rotor;an epicyclic gear reduction positioned between said fan rotor and said turbine rotor, said epicyclic gear reduction including a ring gear, a sun gear, and four intermediate gears that engage said sun gear and said ring gear, a gear ratio of said gear reduction is greater than 3.06;wherein said fan drive turbine drives said sun gear to, in turn, drive said fan rotor;a bypass ratio is defined as a volume of air delivered by said fan into a bypass duct inward of said housing compared to a volume of air delivered into a compressor, said bypass ratio is greater than or equal to 12.0; andwherein there is a primary oil supply supplying oil to journal bearings that support said intermediate gears.2. The turbofan engine as set forth in claim 1 , wherein said gear ratio is greater than or equal to 4.0.3. The turbofan engine as set forth in claim 2 , wherein said gear ratio is greater than or equal to 4.2.4. The turbofan engine as set forth in claim 2 , wherein said gear ratio is less than or equal to 4.4.5. The turbofan engine as set forth in claim 2 , wherein there is an auxiliary oil supply supplying oil to said journal bearings when there is windmilling of said fan rotor.6. The turbofan engine as set forth in claim 1 , ...

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 ...

MECHANICAL REDUCTION GEAR FOR AN AIRCRAFT TURBOMACHINE

A mechanical reduction gear is suitable for a turbomachine and, in particular, a turbomachine of an aircraft. The reduction gear includes a sun gear, a ring gear formed by two half-ring gears, planet gears arranged between the sun gear and the ring gear, and at least one shaft rotationally fixed to the ring gear. The reduction gear further includes an annular covering part that extends around the ring gear and is independently fixed by flanges and/or splines to each of the half-ring gears. 1. A reduction gear for a turbomachine , comprising:a central sun gear having an axis X of rotation;a ring gear extending around the axis X and the sun gear and comprising a herringbone toothing, the ring gear being formed by two coaxial half-ring gears that are spaced apart from each other by an annular space and that each comprise teeth of the toothing, the ring gear being rotatable around the axis X;planet gears disposed between the sun gear and the ring gear and supported by a planet carrier which is immobile in rotation about said axis X;at least one shaft rotationally fixed to the ring gear; andan annular covering part that extends around said annular space and at least one portion of the ring gear, the covering part being fixed to each of the half-ring gears by at least one of flanges and splines,wherein the ring gear is rotationally fixed to two shafts extending respectively on each side of the reduction gear.2. The reduction gear according to claim 1 , wherein the covering part at least one of:extends between the two shafts;covers a portion of the two shafts; andis formed integrally with one of the shafts.3. The reduction gear according to claim 1 , wherein the covering part comprises first splines for coupling to one of the half-ring gears claim 1 , and second splines for coupling to the other of the half-ring gears claim 1 , the half-ring gears being clamped axially against each other and in axial abutment against a abutment of the covering part by means of a nut ...

FLEXIBLE COUPLING FOR GEARED TURBINE ENGINE

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

METHOD FOR COUPLING TWO SUB-SHAFTS

A method for coupling a first sub-shaft, which has a first turbomachine and a generator connected to a mains supply, to a second sub-shaft, which has a second turbomachine, by means of an overrunning clutch, has the following steps: a) rotating the second sub-shaft with a starting rotational speed which is lower than the rotational speed of the first sub-shaft; b) measuring the mains frequency of the mains supply; c) measuring a differential angle between the first sub-shaft and the second sub-shaft; d) accelerating the second sub-shaft with an acceleration value which is produced using the mains frequency measured in step b), the differential angle and the starting rotational speed, and therefore the overrunning clutch couples the two sub-shafts to each other with a previously determined target coupling angle. 1. A method for coupling a first sub-shaft , which has a first fluid-flow machine and a generator connected to a mains supply , to a second sub-shaft , which has a second fluid-flow machine , by means of an overrunning clutch , comprising the steps:a) rotating the second sub-shaft with an initial rotational speed which is lower than the rotational speed of the first sub-shaft;b) measuring the mains frequency of the mains supply;c) measuring a differential angle between the first sub-shaft and the second sub-shaft;d) accelerating the second sub-shaft with an acceleration value which is produced by using the mains frequency measured in step b), the differential angle and the initial rotational speed, so that the overrunning clutch couples the two sub-shafts to each other with a previously determined target coupling angle;e) measuring a new mains frequency during the accelerations of the second sub-shaft;f) in the event that the new mains frequency is different from the mains frequency measured in step b), accelerating the second sub-shaft with a changed acceleration value, which is produced by using the new mains frequency.2. The method as claimed in claim 1 , ...

GAS TURGINE ENGINE WITH TRANSMISSION

A multi spool gas turbine engine with a differential having a selectively rotatable member which rotational speed determines a variable ratio between rotational speeds of driven and driving members of the differential. The driven member is engaged to the first spool and a rotatable shaft independent of the other spools (e.g. connected to a compressor rotor) is engaged to the driving member. First and second power transfer devices are engaged to the first spool and the selectively rotatable member, respectively. A circuit interconnects the power transfer devices and allows a power transfer therebetween, and a control unit controls the power being transferred between the power transfer devices. Power can thus be transferred between the first spool and the selectively rotatable member to change the speed ratio between the first spool and the rotatable shaft. 1. A method of adjusting a speed of a rotatable shaft of a gas turbine engine having a high pressure section including interconnected compressor and turbine rotors , the method comprising:rotating at least one rotor of a low pressure turbine with a flow of exhaust gases from the high pressure section;driving a rotation of the rotatable shaft with a power shaft through a variable transmission, the power shaft being driven by the at least one rotor of the low pressure turbine; andtransferring power between the power shaft and a rotational member of the transmission to change a ratio of rotational speeds between the rotatable shaft and the power shaft.2. The method as defined in claim 1 , wherein transferring power includes transferring power from the power shaft to the transmission to increase the rotational speed of the rotatable shaft and transferring power from the transmission to the power shaft to decrease the rotational speed of the rotatable shaft.3. The method as defined in claim 1 , wherein driving a rotation of the rotatable shaft includes driving a rotation of at least one rotor of a low pressure ...

RING GEAR MOUNTING ARRANGEMENT WITH OIL SCAVENGE SCHEME

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section including a turbo fan supported on a turbo fan shaft, a turbine section including a turbine shaft, and an epicyclic gear train interconnecting the turbo fan shaft and the turbine shaft. The epicyclic gear train includes a sun gear coupled to the turbine shaft, intermediary gears arranged circumferentially about and meshing with the sun gear, a carrier supporting the intermediary gears, and a ring gear including first and second portions each having an inner periphery with teeth, the first and second portions arranged about and intermeshing with the intermediate gears, the first and second portions abutting one another at a radial interface, the first and second portions including respective flanges extending along the radial interface radially outward from the teeth, and the teeth of the first and second portions being oppositely angled teeth. 1. A gas turbine engine comprising:a fan section including a turbo fan supported on a turbo fan shaft;a turbine section including a turbine shaft;a compressor section having compressor hubs with blades driven by the turbine shaft about an axis; and a sun gear coupled to the turbine shaft such that the sun gear is rotatable about the axis;', 'intermediary gears arranged circumferentially about and meshing with the sun gear;', 'a carrier supporting the intermediary gears; and', 'a ring gear including first and second portions each having an inner periphery with teeth, the first and second portions arranged about and intermeshing with the intermediate gears, the first and second portions abutting one another at a radial interface, the first and second portions including respective flanges extending along the radial interface radially outward from the teeth, and wherein the teeth of the first and second portions are oppositely angled teeth that force the first and second portions toward one another at the radial ...

OIL PIPE COVER AND MECHANICAL REDUCTION GEAR FOR AN AIRCRAFT TURBOMACHINE COMPRISING SUCH A COVER

An oil pipe cover for a mechanical reduction gear of a turbomachine, for example of an aircraft, is configured to be fixed to a planet carrier of the reduction gear and to be mounted on an axial end of a hydrodynamic bearing tubular support of a planet gear of the reduction gear. The oil pipe cover has an annular body extending around an axis (Y) and having a mounting central orifice on the axial end. The body has two diametrically opposed circumferential deflectors: a first deflector having a circumferential oil guiding surface located radially outwards with respect to the axis (Y), and a second deflectors having a circumferential oil guiding surface located radially inwards with respect to the axis (Y). 1. An oil pipe cover for a reduction gear of an aircraft turbomachine , the oil pipe cover being configured to be fixed to a planet carrier of the reduction gear and to be mounted on an axial end of a tubular support for a hydrodynamic bearing of a planet gear of the reduction gear , the oil pipe cover comprising an annular body extending around an axis (Y) and comprising a mounting central orifice on the axial end , wherein the annular body comprises:{'b': 1', '38, 'i': 'a', 'a first circumferential deflector having a first circumferential extent (α) around the axis (Y), the first circumferential deflector comprising a circumferential oil guiding surface () located radially outwards with respect to an axis (Y);'}a second circumferential deflector having a second circumferential extent around the axis (Y), the second circumferential deflector comprising a circumferential oil guiding surface located radially inwards with respect to an axis (Y), the first and second circumferential deflectors being diametrically opposed with respect to an axis (Y); andoil discharge pipes which pass substantially axially through the annular body and which are diametrically opposed with respect to the first circumferential deflector, the oil discharge pipes opening axially on a side of ...

ADDITIVE MANUFACTURED GEAR FOR A GEARED ARCHITECTURE GAS TURBINE ENGINE

A gear includes a multiple of gear teeth that extend from an outer portion of a rim about an axis and an inner portion of the rim about the axis, the inner portion of the rim additive manufactured. 1. A gear , comprising:a multiple of gear teeth that extend from an outer portion of a rim about an axis; andan inner portion of said rim about said axis, said inner portion of said rim additive manufactured to the outer portion at a mechanical interface.2. The gear as recited in claim 1 , wherein said inner portion of said rim forms a journal bearing surface.35-. (canceled)6. The gear as recited in claim 1 , wherein said inner portion of said rim includes a matrix.7. The gear as recited in claim 6 , wherein said matrix forms a lattice structure.8. The gear as recited in claim 1 , wherein said gear is an intermediate gear of a geared architecture for a gas turbine engine claim 1 , said intermediate gear is a double helical gear.9. The gear as recited in claim 1 , wherein said inner portion provides different characteristics along an axial length.10. A geared architecture for a gas turbine engine claim 1 , said geared architecture claim 1 , comprising:a sun gear;a ring gear that surrounds said sun gear; anda multiple of intermediate gears in meshing engagement with said sun gear and said ring gear, each of said multiple of intermediate gears including an inner portion of a rim, said inner portion of said rim being additive manufactured, and a multiple of gear teeth that extend from an outer portion of said rim of each of said multiple of intermediate gears about an axis, said inner portion of said rim additive manufactured to the outer portion at a mechanical interface, said gear teeth manufactured via subtractive manufacturing.11. (canceled)12. The geared architecture as recited in claim 10 , wherein said inner portion of said rim forms an inner periphery.13. (canceled)14. The geared architecture as recited in claim 10 , wherein said mechanical interface comprises a ...

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

APPARATUS FOR GAS TURBINE ENGINES

Apparatus for a gas turbine engine, the apparatus comprising: a core engine casing having a longitudinal axis and including: an inner wall defining at least part of a core airflow path through the gas turbine engine; an outer wall defining an external surface of the core engine casing, a first cavity being defined between the inner wall and the outer wall of the core engine casing; a plurality of guide vanes extending radially from the outer wall of the core engine casing; a torque box defined within the first cavity of the core engine casing and at least partially overlapping axially with the plurality of guide vanes, the torque box defining a second cavity; and an accessory gear box positioned within the second cavity of the torque box. 1. Apparatus for a gas turbine engine , the apparatus comprising;a core engine casing having a longitudinal axis and including: an inner wall defining at least part of a core airflow path through the gas turbine engine; an outer wall defining an external surface of the core engine casing, a first cavity being defined between the inner wall and the outer wall of the core engine casing;a plurality of guide vanes extending radially from the outer wall of the core engine casing;to a torque box defined within the first cavity of the core engine casing and at least partially overlapping axially with the plurality of guide vanes, the torque box defining a second cavity; andan accessory gear box positioned within the second cavity of he torque box.2. Apparatus as claimed in claim 1 , wherein the torque box wholly overlaps axially with the plurality of guide vanes.3. Apparatus as claimed in claim 1 , wherein each of the plurality of guide vanes includes a root portion having a leading edge at a first axial position and a trailing edge at a second axial position claim 1 , the torque box comprising a first wall located at the first axial position and a second wall located at the second axial position.4. Apparatus as claimed in claim 1 , ...

HIGH EFFICIENCY GAS TURBINE ENGINE

A gas turbine engine has a quasi-non-dimensional mass flow rate in a defined range and a specific thrust in a defined range to achieve improved over all performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined ranges of quasi-non-dimensional mass flow rate and specific thrust 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 , 0.03 KgsNK≤Q≤0.035 KgsNK.3. A gas turbine engine according to claim 1 , wherein at cruise conditions claim 1 , 031 KgsNK≤Q≤0.034 KgsNK.4. A gas turbine engine according to claim 1 , wherein at cruise conditions claim 1 , the specific thrust is less than 100 Nkgs.5. 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

GEARED TURBOFAN WITH INTEGRAL FRONT SUPPORT AND CARRIER

A gas turbine engine includes a nacelle, and a bypass flow path in a bypass duct within the nacelle of the turbofan engine. A fan section includes a fan with fan blades. The fan section drives air along the bypass flow path. A fan shaft drives a fan that has fan blades and the fan rotates about a central longitudinal axis of the turbofan engine. A speed reduction device includes an epicyclic gear system. A turbine section is connected to the fan section through the speed reduction device and the turbine section rotates about the central longitudinal axis. A first fan bearing for supporting rotation of the fan hub is located axially forward of the speed reduction device. A second fan bearing for supporting rotation of the fan hub is located axially aft of the speed reduction device. A first outer race of the first fan bearing is fixed relative to the fan hub. 1. A gas turbine engine comprising:a fan section including a fan having fan blades extending from a fan hub, wherein said fan section drives air along a bypass flow path;a fan shaft driving said fan and said fan rotates about a central longitudinal axis of said gas turbine engine;a speed reduction device including an epicyclic gear system including a plurality of intermediate gears supported on a corresponding one of a plurality of flexible posts on a static carrier of the epicyclic gear system;a turbine section connected to the fan section through the speed reduction device and said turbine section rotates about said central longitudinal axis;a first fan bearing for supporting rotation of the fan hub located axially forward of the speed reduction device;a second fan bearing for supporting rotation of the fan hub located axially aft of the speed reduction device; anda first outer race of the first fan bearing is fixed relative to the fan hub.2. The gas turbine engine of claim 1 , wherein each of the plurality of flexible posts extend from the static carrier at a proximal end to a distal free end.3. The gas ...

FLEXIBLE SUPPORT STRUCTURE FOR A GEARED ARCHITECTURE GAS TURBINE ENGINE

A gas turbine engine including a drive shaft that drives a propulsor. A frame which supports the drive shaft is a K-frame bearing support. A gear system is connected to the drive shaft. The gear system includes a gear mesh that defines a gear mesh lateral stiffness and a gear mesh transverse. A flexible support supports the gear system that defines a flexible support transverse stiffness and a flexible support lateral stiffness. The flexible support lateral stiffness is less than 8% of the gear mesh lateral stiffness. 1. A gas turbine engine , comprising:a drive shaft driving a propulsor;a frame which supports said drive shaft, wherein said frame is a K-frame bearing support;a gear system connected to said drive shaft, said gear system includes a gear mesh defining a gear mesh lateral stiffness and a gear mesh transverse; anda flexible support supporting said gear system defining a flexible support transverse stiffness and a flexible support lateral stiffness;wherein said flexible support lateral stiffness is less than 8% of said gear mesh lateral stiffness.2. The gas turbine engine of claim 1 , further comprising a bypass duct at least partially defined by a housing outward of said propulsor.3. The gas turbine engine of claim 2 , wherein a lateral stiffness refers to a perpendicular direction with respect to an axis of rotation of said gas turbine engine and a transverse stiffness refers to a pivotal bending movement with respect to said axis of rotation of said gas turbine engine.4. The gas turbine engine of claim 3 , further comprising a two stage high pressure turbine.5. The gas turbine engine of claim 3 , wherein said gear system is an epicyclic gear system.6. The gas turbine engine of claim 5 , wherein said drive shaft is mounted to a ring gear of said epicyclic gear system.7. The gas turbine engine of claim 5 , wherein said drive shaft is mounted to a planet carrier of said epicyclic gear system.8. The gas turbine engine of claim 5 , further comprising an ...

HYDRAULIC DEVICE FOR EMERGENCY STARTING A TURBINE ENGINE, PROPULSION SYSTEM OF A MULTI-ENGINE HELICOPTER PROVIDED WITH ONE SUCH DEVICE, AND CORRESPONDING HELICOPTER

Emergency start-up device for a turboshaft engine of a helicopter, comprising: a hydraulic motor which is mechanically connected to said turboshaft engine; a hydropneumatic store which is connected to said hydraulic motor by a hydraulic circuit for supplying pressurised liquid to said hydraulic motor; and a hydraulic valve which has controlled quick opening, arranged on the hydraulic circuit between said store and said hydraulic motor, and is suitable for being placed on command at least in an open position in which the liquid can supply said hydraulic motor, or in a closed position in which said hydraulic motor is no longer supplied with pressurised liquid. 1. Emergency start-up device for a turboshaft engine of a helicopter , comprising:a hydraulic motor suitable for being mechanically connected to said turboshaft engine and is suitable for setting into rotation said engine to facilitate the start-up thereof;a hydropneumatic store connected to said hydraulic motor by a hydraulic circuit for supplying pressurised liquid to said hydraulic motor;a hydraulic valve having controlled opening, the hydraulic valve being arranged on the hydraulic circuit between said store and said hydraulic motor, and is suitable for being placed on command at least in an open position in which the liquid can supply said hydraulic motor for facilitating a start-up of said turboshaft engine when the device is used with said turboshaft engine, or in a closed position in which said hydraulic motor is no longer supplied with pressurised liquid; anda reservoir for recovering liquid, the reservoir being connected to said hydraulic motor by a purge valve.2. Device according to claim 1 , wherein said hydropneumatic store is selected from the group comprising a bladder-type store claim 1 , a membrane-type store and a piston-type store.3. Device according to claim 1 , wherein said hydraulic motor comprises a propshaft suitable for being mechanically connected to a gearbox shaft of an accessory ...

GEARED ARCHITECTURE TO PROTECT CRITICAL HARDWARE DURING FAN BLADE OUT

A turbofan engine including a fan section including a plurality of fan blades rotatable about an axis, a compressor including a plurality of compressor blades, a turbine including a plurality of turbine blades and a geared architecture driven by the turbine for driving the fan section at a speed and direction different than the turbine is disclosed. A rub strip proximate at least one of the compressor blades, the turbine blades and the fan blades slows rotation when engaged. The rub strip generates a torque opposing rotation when in an engaged condition that is between 2 and 6 times a torque encountered in a non-engaged condition.

Turbofan Engine Assembly Methods

A method for assembling a turbofan engine () comprises coupling a bearing assembly () and a shaft () as a unit to a bearing support (). A transmission () and the fan shaft () are installed to a front frame assembly (). The bearing assembly and shaft are installed as a unit so that the shaft engages a central gear () of the transmission and the bearing support engages the front frame assembly.

GEARED GAS TURBINE ENGINE WITH OIL DEAERATOR

A gas turbine engine comprises a fan drive turbine for driving a gear reduction. The gear reduction drives a fan rotor. A lubrication system supplies oil to the gear reduction. The lubrication system includes a lubricant pump supplying a mixed air and oil to a deaerator inlet. The deaerator includes a separator that for separating oil, and delivering separated air to an air outlet, and for delivering separated oil back into an oil tank. The separator includes a member having lubricant flow paths on both of two opposed sides. A method of designing a gas turbine engine is also disclosed.

GEARED GAS TURBINE ENGINE AND A GEARBOX

A gas turbine engine comprises a gearbox comprises a sun gear, an annulus gear, a plurality of planet gears and a carrier. The sun gear meshes with the planet gears and the planet gears mesh with the annulus gear. The planet gear carrier comprising a first ring, a second ring spaced axially from the first ring and a plurality of circumferentially spaced axles extending axially between the first ring and the second ring. Each planet gear is rotatably mounted on a respective one of the axles and the axles are arranged at a first radius. At least one of the first ring and the second ring comprises a metal matrix composite material ring and the metal matrix composite ring comprises a ring of reinforcing fibres and the ring of reinforcing fibres having a second radius greater than the first radius. 1. A gas turbine engine comprising a gearbox , the gearbox comprising a sun gear , an annulus gear , a plurality of planet gears and a carrier , the sun gear meshing with the planet gears and the planet gears meshing with the annulus gear , the carrier comprising a first ring , a second ring spaced axially from the first ring and a plurality of circumferentially spaced axles extending axially between the first ring and the second ring , each planet gear being rotatably mounted on a respective one of the axles , the axles being arranged at a first radius , at least one of the first ring and the second ring comprising a metal matrix composite material , the metal matrix composite material comprising a ring of reinforcing fibres and the ring of reinforcing fibres having a second radius greater than the first radius.2. A gas turbine engine as claimed in wherein the first ring comprises a first metal matrix composite material and the second ring comprises a second metal matrix composite material claim 1 , the first metal matrix composite material comprises a first ring of reinforcing fibres claim 1 , the first ring of reinforcing fibres having a second radius greater than the first ...

GEARED GAS TURBINE ENGINE

A gas turbine engine comprising a gearbox, the gearbox comprising a sun gear, an annulus gear, a plurality of planet gears and a carrier, each planet gear being rotatably mounted in the carrier by at least one bearing, the sun gear meshing with the planet gears and the planet gears meshing with the annulus gear, the sun gear, the planet gears and the annulus gear comprising gear teeth, the annulus gear having a length in the axial direction and an axially forward end and an axially rearward end and wherein the annulus gear is resiliently mounted to a supporting structure at both its axially forward and axially rearward ends. 1. A gas turbine engine comprising a gearbox , the gearbox comprising a sun gear , an annulus gear , a plurality of planet gears and a carrier , each planet gear being rotatably mounted in the carrier by at least one bearing , the sun gear meshing with the planet gears and the planet gears meshing with the annulus gear , the sun gear , the planet gears and the annulus gear comprising gear teeth , the annulus gear having a length in the axial direction and an axially forward end and an axially rearward end and wherein the annulus gear is resiliently mounted to a static supporting structure at both its axially forward and axially rearward ends , wherein the static supporting structure is radially outside the annulus gear and separated from the annulus gear by a cavity configured to receive a damping volume of hydraulic fluid.2. A gas turbine engine according to claim 1 , wherein the static supporting structure has one or more conduits adapted to supply hydraulic fluid to the cavity.3. A gas turbine engine according to claim 1 , wherein the cavity comprises hydraulic seals at an axially forward end and an axially rearward end.4. A gas turbine engine according to claim 3 , wherein the static supporting structure has two conduits adapted to supply hydraulic fluid to the cavity claim 3 , wherein a first one of the two conduits is located in a first ...

AIRCRAFT ENGINE APPARATUS

An aircraft engine apparatus () includes: a rotating shaft (); a fan () driven by the rotating shaft; a fan case surrounding the fan; aircraft equipment () disposed upstream of the fan and, in a radial direction of the rotating shaft, disposed inward of a peripheral edge of the fan case; a casing () that accommodates at least part of the rotating shaft and supports the fan case; a first motive force transmitter () coupled to the rotating shaft and the fan; a second motive force transmitter () disposed inward of the first motive force transmitter in the radial direction of the rotating shaft and coupled to the rotating shaft and the aircraft equipment; and a support member () disposed between the first motive force transmitter and the second motive force transmitter, the support member coupling the casing and the aircraft equipment and supporting the aircraft equipment. 1. An aircraft engine apparatus comprising:a rotating shaft;a fan driven by the rotating shaft;a fan case surrounding the fan;aircraft equipment disposed upstream of the fan and, in a radial direction of the rotating shaft, disposed inward of a peripheral edge of the fan case;a casing that accommodates at least part of the rotating shaft and supports the fan case;a first motive force transmitter coupled to the rotating shaft and the fan;a second motive force transmitter disposed inward of the first motive force transmitter in the radial direction of the rotating shaft and coupled to the rotating shaft and the aircraft equipment;a support member disposed between the first motive force transmitter and the second motive force transmitter, the support member coupling the casing and the aircraft equipment and supporting the aircraft equipment; anda nose cone disposed upstream of the fan, whereinthe aircraft equipment is disposed inside the nose cone.2. The aircraft engine apparatus according to claim 1 , further comprising a transmission disposed between the rotating shaft and the first motive force ...

Geared Turbocharged Engine

A geared turbocharged engine, having a gearbox, a drive and output assembly, and an oil supply system. The gearbox has a housing. The oil supply system has a reservoir, a supply line, an oil pump to convey oil from the reservoir, and a return line, to return oil from the gearbox to the reservoir. The housing is mounted on a support plate of reservoir a supply-line opening and a return-line opening are introduced into the housing and the support plate such that the supply-line opening of the support plate and the supply-line opening of the housing and the return-line opening of the gearbox and the return-line opening of the support plate are flush, so that the inflow of oil to the gearbox and the return flow of oil from the gearbox take place via the gearbox housing. 1. A geared turbocharged engine , comprising: at least one drive and output assembly; a gearbox housing;', 'a central gear wheel with a gear-wheel shaft; and', 'at least one pinion, with at least one pinion shaft, meshing into the central gear wheel; and, 'a gearbox comprising a support plate;', 'an oil storage reservoir in which oil is held;', 'an oil supply line by which oil can be drawn from the oil storage reservoir and guided toward the gearbox to be lubricated;', 'an oil pump configured to convey the oil out of the oil storage reservoir, and', 'an oil return line by which oil emanating from the gearbox flows back into the oil storage reservoir;', 'wherein:', 'the gearbox housing is mounted standing on the support plate by a gearbox housing base;', 'a supply-line opening and a return-line opening are introduced into each of the gearbox housing base and the support plate such that the supply-line opening of the support plate and the supply-line opening of the gearbox housing base are flush and the return-line opening of the gearbox housing base and the return-line opening of the support plate,', 'whereby inflow of the oil to the gearbox and return flow of the oil from the gearbox take place via the ...

FAN CLUTCH FOR CONVERTIBLE ENGINE

Systems and methods include providing an aircraft with a fuselage and a convertible engine disposed within the fuselage. The convertible engine is operable as a turbofan engine in a thrust mode and a turboshaft engine in a shaft power mode. The convertible engine includes a housing, an engine core having a low pressure turbine shaft, and a bypass fan system. The bypass fan system includes a bypass fan having a fan clutch. The fan clutch selectively couples at least a portion of the bypass fan to the low pressure turbine shaft when the convertible engine is operated in the thrust mode and decouples at least a portion of the bypass fan from the low pressure turbine shaft when the convertible engine is operated in the shaft power mode. 1. A convertible engine for an aircraft , comprising:an engine core comprising a low pressure turbine shaft; and 'a bypass fan comprising a fan clutch, wherein the fan clutch is configured to selectively decouple at least a portion of the bypass fan from the low pressure turbine shaft.', 'a bypass fan system comprising2. The convertible engine of claim 1 , wherein the fan clutch is at least one of an electromechanical clutch and a piezoelectric clutch.3. The convertible engine of claim 1 , wherein the fan clutch is a magnetorheological clutch comprising a magnetorheological fluid.4. The convertible engine of claim 3 , wherein electromagnets are disposed in at least one non-rotating claim 3 , fixed reference component of the convertible engine and in close proximity to the fan clutch claim 3 , and wherein the electromagnets are configured to selectively induce a magnetic field through the magnetorheological fluid in the fan clutch to couple the at least a portion of the bypass fan to the low pressure turbine shaft.5. The convertible engine of claim 1 , wherein the bypass fan is coupled to the low pressure turbine shaft when the convertible engine is operated in a thrust mode claim 1 , and wherein the bypass fan is decoupled from the low ...

COUNTER ROTATING POWER TURBINE WITH REDUCTION GEARBOX

The present disclosure is directed to a turbine engine defining a longitudinal direction, a radial direction, and a circumferential direction. The turbine engine includes a power turbine including a first turbine rotor assembly interdigitated with a second turbine rotor assembly along the longitudinal direction; a gear assembly coupled to the first turbine rotor assembly and the second turbine rotor assembly, wherein the gear assembly includes a first input interface coupled to the first turbine rotor assembly, a second input interface coupled to the second turbine rotor assembly, and one or more third gears coupled to the first input interface and the second input interface therebetween; and a first output shaft and a second output shaft, wherein each of the first output shaft and the second output shaft are configured to couple to an electrical load device. 1. A turbine engine defining a longitudinal direction , a radial direction , and a circumferential direction , the turbine engine comprising:a power turbine comprising a first turbine rotor assembly interdigitated with a second turbine rotor assembly along the longitudinal direction;a gear assembly coupled to the first turbine rotor assembly and the second turbine rotor assembly, wherein the gear assembly comprises a first input interface coupled to the first turbine rotor assembly, a second input interface coupled to the second turbine rotor assembly, and one or more third gears coupled to the first input interface and the second input interface therebetween; anda first output shaft and a second output shaft, wherein each of the first output shaft and the second output shaft are configured to couple to an electrical load device.2. The turbine engine of claim 1 , wherein the gear assembly further comprises an output interface claim 1 , and wherein the first output shaft is coupled to the output interface of the gear assembly.3. The turbine engine of claim 1 , wherein the second turbine rotor assembly is coupled ...

GAS TURBINE ENGINE WITH AN ENGINE ROTOR ELEMENT TURNING DEVICE

A turbine engine has a core with a compressor, combustor, and turbine sections in axial flow arrangement and with corresponding rotating elements mounted to a shaft defining engine rotor elements. The turbine engine has a rotary driver operably coupled to the engine rotor elements. The turbine engine has at least one thermoelectric generator in thermal communication with the core and in electrical communication with the rotary driver to provide power to the rotary driver to turn the engine rotor elements. 1. A turbine engine comprising;a core having compressor, combustor, and turbine sections in axial flow arrangement, with corresponding rotating elements mounted to a shaft to define engine rotor elements;a rotary driver operably coupled to and turning the engine rotor elements; andat least one thermoelectric generator in thermal communication with the core and in electrical communication with the rotary driver to provide power to the rotary driver to turn the engine rotor elements.2. The turbine engine of further comprising an inner core cowl surrounding the core claim 1 , and the at least one thermoelectric generator is secured to the inner core cowl.3. The turbine engine of wherein the inner core cowl further comprises an upper bifurcation area and the at least one thermoelectric generator is secured in the upper bifurication area.4. The turbine engine of wherein the at least one thermoelectric generator is removably secured in the upper bifurication area of the inner core cowl.5. The turbine engine of further comprising an accessory gear box operably coupled to the shaft claim 1 , and the rotary driver connects to and drives the accessory gear box to rotate the shaft.6. The turbine engine of further comprising a fan wherein the rotary driver connects to and drives the fan operably connected to the engine rotor elements.7. The turbine engine of wherein the at least one thermoelectric generator is in thermal communication with one of the combustor or turbine ...

Air inlet for a gas turbine engine

A radial air inlet for a gas turbine engine. The air inlet has an inlet duct defined between two axially-spaced radially-extending annular walls and has a plurality of circumferentially-spaced axially-extending struts extending between the annular walls adjacent a radially-outer portion of the air inlet. At least one of the struts has an internal passage extending between a first opening in a forward end of the strut and a second opening in an aft end of the strut, the first and second openings being axially spaced apart. A transmission shaft extends through the internal passage of said strut.

RAPIDLY AVAILABLE ELECTRIC POWER FROM A TURBINE-GENERATOR SYSTEM HAVING AN AUXILIARY POWER SOURCE

Turbine-generator systems having an auxiliary power source and methods of operating turbine-generator systems having auxiliary power sources in a high, intermittent load environment are provided. The turbine may be sized to meet substantially all the power required by the intermittent load. The auxiliary power source may have a power rating approximately equal to the power required to rotate the unloaded generator at an operational speed. The method may include decoupling the turbine from the generator, removing the load from the generator, coupling the auxiliary power source to the unloaded generator, and maintaining the operational rotational speed of the generator with the auxiliary power source. 1. In a high intermittent load environment , wherein a gas turbine drives a generator to provide power to an intermittent load and the gas turbine is sized to meet substantially all the power required by the intermittent load , a method of maintaining an operational rotational speed of the generator in a standby mode , comprising:decoupling the gas turbine from the generator;removing the load from the generator;coupling an auxiliary power source to the unloaded generator; andmaintaining the operational rotational speed of the generator with the auxiliary power source, wherein the auxiliary power source has a continuous maximum power rating approximately equal to power required to rotate the unloaded generator at the operational rotational speed.2. The method of claim 1 , wherein the auxiliary power source is an electric power source.3. The method of claim 2 , further comprising:supplying electrical power from the electric power source to the generator; andoperating the generator as a motor.4. The method of claim 2 , further comprising:supplying electrical power from the electric power source to a secondary motor; androtating the unloaded generator with the secondary motor.5. The method of claim 2 , wherein the electric power source is selected from the group consisting ...