ПАЯНОЕ СОЕДИНЕНИЕ МЕЖДУ МЕТАЛЛИЧЕСКОЙ ДЕТАЛЬЮ НА ОСНОВЕ ТИТАНА И ДЕТАЛЬЮ ИЗ КЕРАМИЧЕСКОГО МАТЕРИАЛА НА ОСНОВЕ КАРБИДА КРЕМНИЯ (SiC) И/ИЛИ УГЛЕРОДА

Изобретение относится к области соединения пайкой металлической детали на основе титана и детали из керамического материала на основе карбида кремния (SiC) и/или углерода. Данное соединение может быть использовано в авиации, в частности в турбомашине, для соединения деталей, входящих в состав сопла, камеры сгорания и форсажной камеры. Соединение содержит многослойную структуру, включающую в себя следующие элементы, соединенные друг с другом путем пайки: металлическая деталь (10) на основе титана, первая прокладка (11), способная деформироваться для приспосабливания к разнице в расширении металлической детали (10) и детали из керамического материала (20) на основе карбида кремния и/или углерода, вторая жесткая прокладка (12) из нитрида алюминия (AlN) или вольфрама (W) с коэффициентом расширения, близким к коэффициенту расширения упомянутой детали из керамического материала (20), и деталь из керамического материала (20). Технический результат изобретения - создание спая, позволяющего компенсировать ...

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

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

УЗЕЛ АВИАЦИОННОГО ДВИГАТЕЛЯ И АВИАЦИОННЫЙ ДВИГАТЕЛЬ

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

СПОСОБ ИЗГОТОВЛЕНИЯ КОМПОНЕНТА СТАТОРА

Способ изготовления компонента статора, предназначенного для направления потока газа и передачи усилий в процессе работы статора, осуществляется путем сборки в окружном направлении, по меньшей мере, двух отлитых в виде отдельных деталей секторов, которые устанавливают рядом друг с другом и соединяют сваркой. Каждый сектор отливают с первым и вторым элементами стенки, расположенными на расстоянии друг от друга с образованием в окружном направлении между ними канала для прохода газа. Затем соединяют друг с другом два элемента стенки, по одному от двух соседних секторов, так что они образуют в компоненте статора вытянутую в его радиальном направлении перегородку. Указанная перегородка служит в компоненте статора для направления потока газа и/или передачи возникающих при работе усилий. Изобретение позволяет обеспечить высокую производительность и снизить затраты при изготовлении компонента статора, а также повысить его прочность. 17 з.п. ф-лы, 3 ил.

ЗАГОТОВКА И МОНОБЛОЧНАЯ ЛОПАТКА ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Изобретение относится к волокнистой заготовке для лопатки газотурбинного двигателя, полученной посредством моноблочного трехмерного тканья. Согласно изобретению заготовка содержит первый продольный участок, предназначенный для формирования ножки лопатки, второй продольный участок, продолжающий первый продольный участок и предназначенный для формирования части пера, и первый поперечный участок, проходящий поперечно от соединения между первым и вторым продольными участками и предназначенный для формирования первой площадки. При этом первый поперечный участок образован по меньшей мере частями первого и второго свободных полотнищ, загнутых друг на друга. Причем первое свободное полотнище и первый продольный участок и второе свободное полотнище и второй продольный участок выполнены путем совместного тканья с разъединением. При этом на соединении между первым и вторым продольными участками имеется перекрещивание слоя так, чтобы нити первого свободного полотнища были продолжены во втором продольном ...

ВОЛОКНИСТАЯ ЗАГОТОВКА ДЛЯ ПОЛОЙ ЛОПАТКИ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Изобретение относится к волокнистой заготовке для полой лопатки газотурбинного двигателя, к такой полой лопатке и способу изготовления такой полой лопатки. Изобретение также относится к газотурбинному двигателю и летательному аппарату, содержащим такую полую лопатку. Волокнистая заготовка для полой лопатки турбины газотурбинного двигателя включает в себя основную волокнистую структуру, полученную трехмерным тканьем и содержащую по меньшей мере одну основную часть, причем основная часть проходит от первой соединительной полосы, содержит первую основную продольную часть, подходящую для формирования, по существу, стенки стороны нагнетания аэродинамического профиля, затем часть поворота на 180°, подходящую для формирования, по существу, передней кромки или задней кромки аэродинамического профиля, затем вторую основную продольную часть, обращенную напротив первой основной продольной части и подходящую для формирования, по существу, стенки стороны всасывания аэродинамического профиля, и заканчивается ...

СПОСОБ ЗАДАНИЯ ПЕРЕДАТОЧНОГО ОТНОШЕНИЯ ЗУБЧАТОЙ ПЕРЕДАЧИ ВЕНТИЛЯТОРНОГО ПРИВОДА ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Газотурбинный двигатель содержит вентиляторную секцию, содержащую вентилятор, выполненный с возможностью вращения вокруг оси, и редуктор. Редуктор соединен с вентилятором и содержит планетарную приводную зубчатую передачу с заторможенным водилом с передаточным отношением, составляющим по меньшей мере 1,5. При этом вентилятор выполнен с возможностью вращения с обеспечением величины приведенной окружной скорости концевой части лопатки вентилятора более 1150, но менее 1400 футов в секунду. При этом степень двухконтурности составляет от 11,0 до 22,0. Технический результат настоящего изобретения заключается в улучшении характеристик газотурбинных двигателей, включая повышение теплового, передаточного и тягового коэффициентов полезного действия. 9 з.п. ф-лы, 3 ил.

УЗЕЛ СТАТОРА И РОТОРА ТУРБИНЫ И ТУРБИННЫЙ ДВИГАТЕЛЬ

Группа изобретений относится к глубинному приводному буру для вращательного бурения. Узел статора и ротора турбины содержит размещенные соосно статор (1) и ротор (2). Статор (1) содержит корпус, лопатку и обод. Ротор (2) содержит корпус, лопатку и обод. Внутренняя стенка обода статора соосно размещена с внешней стенкой корпуса ротора. Линия пересечения каждой точки на внешнем контуре лопатки статора с меридиональной плоскостью, соответствующей ей, представляет собой первую линию пересечения, которая перпендикулярно пересекается с первой прямой линией проекции, продолжающейся через обод статора. Линия пересечения каждой точки на внешнем контуре лопатки ротора с меридиональной плоскостью, соответствующей ей, представляет собой вторую линию пересечения, которая перпендикулярно пересекается со второй прямой линией проекции, продолжающейся через корпус ротора. Группа изобретений направлена на обеспечение высокой гидравлической эффективности, простой конструкции, высокого крутящего момента и ...

КОРПУС ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ И ТУРБОРЕАКТИВНЫЙ ДВИГАТЕЛЬ, СОДЕРЖАЩИЙ ТАКИЕ КОРПУСА

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

СЕКТОР НАСАДКИ ДЛЯ ТУРБИННОГО ДВИГАТЕЛЯ С ДИФФЕРЕНЦИАЛЬНО ОХЛАЖДАЕМЫМИ ЛОПАТКАМИ

Изобретение относится к сектору (22) сопла для турбинного двигателя. Сектор (22) сопла для турбины (2) турбомашины (1) содержит радиально-наружную опорную полку (24) для лопаток, радиально-внутреннюю опорную полку (26) для лопаток, первую концевую лопатку (81), вторую концевую лопатку (84) и по меньшей мере одну первую центральную лопатку (82, 83) между концевыми лопатками (81, 84) вдоль окружного направления (Z-Z) полок и средства (37, 50, 44, 46, 54, 56) охлаждения для охлаждения лопаток. Лопатки (81, 82, 83, 84) проходят радиально между полками (24, 26) вдоль направления X-X размаха этих лопаток. Средства (37, 50, 44, 46, 54, 56) охлаждения для охлаждения лопаток выполнены с возможностью охлаждения каждой из лопаток (81, 82, 83, 84) посредством обеспечения прохождения через них охлаждающего воздуха и с возможностью дифференциального охлаждения центральной лопатки или каждой центральной лопатки (82, 83) по крайней мере по отношению к первой концевой лопатке (81, 84). Средства (37, 50, ...

НАПРАВЛЯЮЩАЯ ЛОПАТКА ТУРБИНЫ С ОХЛАЖДАЕМОЙ ГАЛТЕЛЬЮ

Направляющая лопатка содержит полку и перо, продолжающееся от указанной полки и соединенное с полкой посредством галтели. Инжекционная трубка вставляется в перо, ограничивая охлаждающий канал между инжекционной трубкой и боковыми стенками пера. Направляющая лопатка дополнительно содержит отклоняющую структуру, расположенную смежно галтели и которая повторяет внутренний контур галтели и ограничивает первый охлаждающий проход между галтелью и отклоняющей структурой. Первое препятствие расположено внутри пера в месте соединения галтели с боковыми стенками для отделения первого охлаждающего прохода от охлаждающего канала в пере и для того, чтобы направлять охлаждающий газ из первого охлаждающего прохода в инжекционную трубку. Изобретение направлено на повышение эффективности охлаждения. 2 н. и 13 з.п. ф-лы, 8 ил.

УСОВЕРШЕНСТВОВАННЫЙ СПОСОБ ИЗГОТОВЛЕНИЯ ПО ВЫПЛАВЛЯЕМОЙ ВОСКОВОЙ МОДЕЛИ КОЛЬЦЕВОГО ЛОПАТОЧНОГО УЗЛА ТУРБОМАШИНЫ, МЕТАЛЛИЧЕСКАЯ ФОРМА И ВОСКОВАЯ МОДЕЛЬ ДЛЯ РЕАЛИЗАЦИИ ТАКОГО СПОСОБА

Изобретение относится к литейному производству. Восковая модель (50) кольцевого лопаточного узла статора турбомашины содержит радиально внутренний (52) и радиально внешний (54) коаксиальные экраны, соединенные друг с другом лопатками (56), из которых по меньшей мере одна лопатка (58) имеет внутреннюю полость. Для получения восковой модели в металлической форме размещают стержень (40) для образования оттиска полости лопатки (18), который выполнен из металла, впрыскивают воск в форму и извлекают восковую модель, содержащую стержень (40), из формы. Стержень располагают в форме так, что его радиально внутренний конец размещается на участке (58) формы, образующем лопатку с полостью, на расстоянии от радиально внутреннего конца (52) данного участка формы. Снижается брак по восковым моделям вследствие устранения деформации стержня при инжекции воска в форму. 4 н. и 3 з.п. ф-лы, 7 ил.

ГАЗОВАЯ ТУРБИНА С КАМЕРОЙ СГОРАНИЯ, ПРИКРЕПЛЕННОЙ К СОПЛОВОМУ АППАРАТУ

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

СОПЛОВОЙ СЕГМЕНТ ДЛЯ ГАЗОВОЙ ТУРБИНЫ, ПОКРЫТЫЙ ПОКРЫТИЕМ MCrAlY И НАКЛАДКАМИ ТБП

Изобретение относится к сопловому аппарату для газовой турбины. Сопловой аппарат содержит первое перо, содержащее первую спинку и первое корыто, второе перо, содержащее вторую спинку и второе корыто, внутренний бандаж и наружный бандаж. Первое перо и второе перо расположены между внутренним бандажом и наружным бандажом, при этом первое перо и второе перо по меньшей мере частично покрыты покрытием MCrAlY, и части внутреннего и наружного бандажей покрыты покрытием MCrAlY. По меньшей мере первая спинка содержит первый участок покрытой поверхности, который покрыт термобарьерным покрытием и который представляет собой по меньшей мере часть всей поверхности первой спинки. По меньшей мере внутренний бандаж или наружный бандаж содержит дополнительный участок покрытой поверхности, который покрыт дополнительным термобарьерным покрытием. Другое изобретение группы относится к способу изготовления соплового аппарата, в котором наносят покрытие MCrAlY на части соплового аппарата и затем покрывают участки ...

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ

Газотурбинный двигатель содержит камеру сгорания и узел направляющих лопаток. Узел направляющих лопаток содержит первый и второй узлы направляющих лопаток, расположенные вдоль окружного направления турбины, а также дополнительный первый узел направляющих лопаток. Первый узел направляющих лопаток, содержащий первую платформу и первое число первых аэродинамических профилей, прикрепленных к первой платформе. Второй узел направляющих лопаток, содержащий вторую платформу и второе число вторых аэродинамических профилей, прикрепленных ко второй платформе. Первое число первых аэродинамических профилей отличается от второго, причем первый узел направляющих лопаток выполнен с более высокой теплостойкостью, чем второй узел направляющих лопаток. На первый узел направляющих лопаток нанесено первое термобарьерное покрытие, а на второй - второе термобарьерное покрытие, причем первая толщина первого термобарьерного покрытия превышает вторую толщину второго. Дополнительный первый узел направляющих лопаток ...

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ

Газотурбинный двигатель включает лопатку статора для направления горячих газов сжигания на роторные лопатки. Лопатка статора включает платформу, расположенную на радиально внутренней стороне лопатки относительно оси вращения двигателя. Платформа имеет часть задней кромки по потоку ниже относительно потока горячих газов сгорания после лопатки статора. Двигатель включает также опорную и охлаждающую систему для направления охлаждающей текучей среды на верхний по потоку конец стороны части задней кромки платформы. Сторона обращена радиально внутрь относительно оси вращения двигателя. Опорная и охлаждающая система также направляет охлаждающую текучую среду для прохождения по стороне в основном в осевом направлении к нижнему по потоку концу стороны. Охлаждающая текучая среда охлаждает часть задней кромки при прохождении по стороне. В сторону включены турболизаторы для увеличения переноса тепла с части задней кромки при прохождении потока охлаждающей текучей среды по стороне. Турболизаторы проходят ...

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

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

УЗЕЛ (ВАРИАНТЫ) И СПОСОБ УСТАНОВКИ И УПЛОТНЕНИЯ СОПЛОВОГО ЭЛЕМЕНТА ДЛЯ ГАЗОТУРБИННОЙ СИСТЕМЫ

Узел для установки и уплотнения соплового элемента для газотурбинной системы содержит сопловой элемент, стопорное кольцо, пластину уплотнения и шайбу. Сопловой элемент имеет заднюю кромку наружного бандажа и паз для штифта, предотвращающего поворот. Стопорное кольцо проходит в окружном направлении вокруг наружной поверхности наружного бандажа и содержит штифт, предотвращающий поворот, и отверстие для штифта, предотвращающего поворот. Штифт выполнен с возможностью посадки в осевой ориентации в пазу для штифта и отверстии для штифта. Пластина уплотнения расположена на наружной поверхности наружного бандажа и выполнена с обеспечением удержания штифта в пазу и отверстии. Пластина уплотнения имеет по меньшей мере одно углубление, расположенное вблизи ее первого края и вблизи края смежной пластины уплотнения. Шайба расположена в отверстии пластины уплотнения. Отверстие пластины уплотнения выровнено с отверстием в стопорном кольце, и в стопорное кольцо через отверстие пластины уплотнения проходит ...

ТУРБИННЫЙ ДВИГАТЕЛЬ ЛЕТАТЕЛЬНОГО АППАРАТА, ЕГО МОДУЛЬ, ЧАСТЬ СТАТОРА ДЛЯ ТАКОГО МОДУЛЯ, А ТАКЖЕ КОЛЬЦО ДЛЯ ТАКОГО СТАТОРА

Кольцо статора модуля турбинного двигателя летательного аппарата имеет множество сквозных отверстий, предназначенных для расположения лопатки статора. Каждое отверстие определяет среднюю линию, проходящую между первым краем, предназначенным для расположения задней кромки лопатки, и вторым краем, предназначенным для расположения передней кромки лопатки. С отверстием для расположения лопатки статора соотнесена прорезь снятия механической нагрузки, выполненная сквозной на кольце и расположенная против и на удалении от упомянутого первого края такого отверстия в направлении средней линии. Другие изобретения группы относятся к части статора, содержащей указанное выше кольцо и множество лопаток статора, к модулю турбинного двигателя летательного аппарата, содержащему указанную выше часть статора, и к турбинному двигателю, содержащему такой модуль. Группа изобретений позволяет снизить вероятность образования трещин на кольце статора в области задней кромки лопатки. 4 н. и 9 з.п. ф-лы, 7 ил.

ЛОПАТКА ТУРБОМАШИНЫ, ЛОПАТОЧНЫЙ УЗЕЛ ТУРБОМАШИНЫ, РОТОР ВЕНТИЛЯТОРА И ТУРБОМАШИНА

Изобретение относится к лопатке турбомашины. Лопатка содержит перо, хвостовик, вводимый в зацепление с канавкой диска турбомашины, и внутреннюю полку, расположенную в радиальном направлении между хвостовиком и пером. Верхний по потоку конец хвостовика соединен с радиально внутренним концом входной кромки пера посредством верхнего по потоку края соединительного участка, расположенного в радиальном направлении между хвостовиком и внутренней полкой, так, что радиально внутренний конец входной кромки пера расположен дальше вниз по потоку, чем верхний по потоку конец хвостовика, и/или нижний по потоку конец хвостовика соединен с радиально внутренним концом выходной кромки пера посредством нижнего по потоку края соединительного участка так, что радиально внутренний конец выходной кромки пера расположен дальше вверх по потоку, чем нижний по потоку конец хвостовика. Изобретение также относится к лопаточному узлу турбомашины, ротору вентилятора для турбомашины, а также турбомашине. Изобретение обеспечивает ...

ПАРОВАЯ ТУРБИНА С ПОВОРОТНЫМИ ЛОПАТКАМИ СТАТОРА

Паровая турбина (200) имеет множество ступеней (261, 262, 271) расширения и лопаток (221, 222, 231) статора перед по меньшей мере одной из ступеней (261, 262, 271) расширения; во время работы паровой турбины (200) для регулирования потока пара внутри паровой турбины (200) и максимального повышения эффективности турбины управляют угловыми положениями лопаток (221, 222, 231) статора, например, с помощью внешнего блока управления посредством, например, управляющего стержня (289, 299). Это решение обеспечивает гораздо более высокую эффективность за пределами проектных рабочих условий турбины, поскольку позволяет избежать рассеивания энергии, связанного с использованием либо дросселирования, либо парциальной дуги. В частности, эффективность паровой турбины остается близкой к проектному уровню даже за пределами проектных рабочих условий. 3 н. и 12 з.п. ф-лы, 11 ил.

ДВУХЛОПАСТНАЯ ЛОПАТКА С ПЛАСТИНКАМИ, КОЛЕСО ТУРБИНЫ И ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ, СОДЕРЖАЩИЕ ТАКИЕ ЛОПАТКИ

Лопатка газотурбинного двигателя содержит первую лопасть, вторую лопасть и, по меньшей мере, одну пластинку. Каждая из первой лопасти и второй лопасти имеет внутреннюю и внешнюю стороны, размещенные между передней и задней кромками. Первая и вторая лопасти расположены рядом таким образом, что внутренняя сторона первой лопасти размещена всей своей поверхностью напротив внутренней стороны второй лопасти. Пластинка связывает внутреннюю сторону первой лопасти и внутреннюю сторону второй лопасти и размешена до задней кромки лопатки. Задняя кромка лопатки образована задней кромкой первой лопасти и задней кромкой второй лопасти. Задняя кромка первой лопасти выровнена относительно задней кромки второй лопасти и располагается рядом с ней. Другие изобретения группы относятся к колесу турбины и газотурбинному двигателю, содержащим указанные выше лопатки. Изобретение позволяет снизить аэродинамические потери на лопатке. 3 н. и 11 з.п. ф-лы, 5 ил.

ОХЛАЖДАЕМЫЙ БЛОК ЛОПАТОК (ВАРИАНТЫ)

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

СПОСОБ ЗАДАНИЯ ПЕРЕДАТОЧНОГО ОТНОШЕНИЯ ЗУБЧАТОЙ ПЕРЕДАЧИ ВЕНТИЛЯТОРНОГО ПРИВОДА ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

... 1. Газотурбинный двигатель, содержащий:вентиляторную секцию, содержащую вентилятор, выполненный с возможностью вращения вокруг оси; иредуктор, соединенный с вентилятором и содержащий звездную приводную зубчатую передачу с передаточным отношением, составляющим по меньшей мере 1,5,при этом окружная скорость концевой части лопатки вентилятора составляет менее 1400 футов в секунду, и степень двухконтурности составляет от приблизительно 11,0 до приблизительно 22,0.2. Газотурбинный двигатель по п. 1, в котором редуктор имеет передаточное отношение звездной зубчатой передачи, составляющее по меньшей мере 2,6.3. Газотурбинный двигатель по п. 2, в котором редуктор имеет передаточное отношение зубчатой передачи, меньшее или равное 4,1.4. Газотурбинный двигатель по п. 1, в котором звездная передача содержит центральное зубчатое колесо, множество звездных зубчатых колес, кольцевое зубчатое колесо и водило.5. Газотурбинный двигатель по п. 4, в котором каждое из множества зубчатых колес содержит по меньшей ...

КОМПРЕССОРНОЕ УСТРОЙСТВО ГАЗОВОЙ ТУРБИНЫ И КОРПУСНОЙ ЭЛЕМЕНТ КОМПРЕССОРА

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

ТРУБКА ПРИНУДИТЕЛЬНОГО ОХЛАЖДЕНИЯ ДЛЯ ЛОПАТКИ ГАЗОВОЙ ТУРБИНЫ, ИМЕЮЩАЯ РАЗДЕЛИТЕЛЬНУЮ СТЕНКУ

... 1. Трубка (100) для установки в направляющей лопатке (120) турбины, содержащаястенку (101) трубки для образования канала для текучей среды иразделительную стенку (110), расположенную внутри канала для текучей среды,при этом разделительная стенка (110) имеет первый край (111) и второй край (112), отстоящий от первого края (111),причем первый край (111) прикреплен к первой секции поверхности стенки (101) трубки,при этом разделительная стенка (110) образована так, что второй край (112) упруго упирается с возможностью разъединения во вторую секцию поверхности стенки (101) трубки, так что разделительная стенка (110) разделяет канал для текучей среды на первый канал (I) и второй канал (II).2. Трубка (100) по п. 1, в которой первый край (111) приварен к первой секции поверхности стенки (101) трубки.3. Трубка (100) по п. 1 или 2, в которой разделительная стенка (110) расположена внутри канала для текучей среды так, что если давление (p1) первой текучей среды в первом канале (I) превышает давление ...

ЭЛЕМЕНТ ТУРБОМАШИНЫ С РАЗГРУЗОЧНОЙ ПОЛОСТЬЮ

СПОСОБ УЛУЧШЕНИЯ ХАРАКТЕРИСТИК ТУРБИННОГО КОМПРЕССОРА

СЕКТОР СТАТОРА ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ И СПОСОБ ЕГО ИЗГОТОВЛЕНИЯ

КОНСТРУКЦИЯ ЛОПАТКИ СТАТОРА И ТУРБОВЕНТИЛЯТОРНЫЙ РЕАКТИВНЫЙ ДВИГАТЕЛЬ С ИСПОЛЬЗОВАНИЕМ ТАКОЙ ЛОПАТКИ

РЕМОНТНЫЙ ЭЛЕМЕНТ ДЛЯ ЛОПАТОЧНОГО УЗЛА ГАЗОВОЙ ТУРБИНЫ И СПОСОБ РЕМОНТА ПОВРЕЖДЕННОЙ ЛОПАТКИ ЛОПАТОЧНОГО УЗЛА ГАЗОВОЙ ТУРБИНЫ

СПОСОБ ЭКСПЛУАТАЦИИ БЛОКА ТУРБИН, ПАРОТУРБИННАЯ ЭЛЕКТРОСТАНЦИЯ ИЛИ КОМБИНИРОВАННАЯ ГАЗОПАРОТУРБИННАЯ ЭЛЕКТРОСТАНЦИЯ (ВАРИАНТЫ), ПРИМЕНЕНИЕ ДРОССЕЛЬНОГО УСТРОЙСТВА

ОБОГАЩЕННЫЙ ПЛАЗМОЙ БЫСТРО РАСШИРЯЮЩИЙСЯ ПЕРЕХОДНЫЙ КАНАЛ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

... 1. Система (11) обогащенного плазмой быстро расширяющегося канала, содержащая межтурбинный переходный канал (114) газотурбинного двигателя, имеющий вход (64) канала и выход (66) канала позади и ниже по течению от входа канала, радиально отстоящие друг от друга конические радиально внутреннюю и внешнюю стенки (60 и 62) канала, рапположенные в осевом направлении между входом (64) канала и выходом (66) канала, и конический плазменный генератор (2) для создания конической плазмы (90) вдоль внешней стенки (62) канала. 2. Система (11) по п.1, дополнительно содержащая конический плазменный генератор (2), установленный на внешнюю стенку (62) канала. 3. Система (11) по п.2, дополнительно содержащая конический плазменный генератор (2), включающий в себя радиально внутренний и внешний электроды (3, 4), разделенные диэлектрическим материалом (5). 4. Система (11) по п.3, дополнительно содержащая источник (100) питания переменного тока для подачи потенциала высокого напряжения переменного тока на электроды ...

ОТСЕК С ЛОПАТКАМИ НАПРАВЛЯЮЩЕГО ЛОПАТОЧНОГО УСТРОЙСТВА ОСЕВОГО КОМПРЕССОРА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

УЗЕЛ ДЛЯ СПРЯМЛЕНИЯ ВОЗДУХА В ГАЗОТУРБИННОМ ДВИГАТЕЛЕ С УЛУЧШЕННЫМИ АЭРОДИНАМИЧЕСКИМИ ХАРАКТЕРИСТИКАМИ

СОПЛОВОЙ СЕГМЕНТ ДЛЯ ГАЗОВОЙ ТУРБИНЫ, ПОКРЫТЫЙ ПОКРЫТИЕМ MCRALY И НАКЛАДКАМИ ТБП

ДИФФУЗОР ДЛЯ КОЛЬЦЕВОЙ КАМЕРЫ СГОРАНИЯ, В ЧАСТНОСТИ, ДЛЯ ТУРБИННОГО ДВИГАТЕЛЯ САМОЛЕТА, А ТАКЖЕ КАМЕРА СГОРАНИЯ И АВИАЦИОННЫЙ ТУРБОВИНТОВОЙ ДВИГАТЕЛЬ, СОДЕРЖАЩИЕ ТАКОЙ ДИФФУЗОР

... 1. Диффузор для кольцевой камеры сгорания с одной зоной, в частности, для авиационного турбовинтового двигателя, содержащий разделительный элемент, разделяющий поток воздуха, выходящий из компрессора, на два кольцевых диффузионных потока, при этом разделительный элемент образован тонким листом, соединенным конструкционными рычагами с внутренней и наружной кругообразно симметричными стенками диффузора, причем угол расширения каждого диффузионного потока, образованный указанным тонким листом в диффузоре, составляет приблизительно 12-13°, при этом топливные форсунки в камере сгорания расположены на одной линии с участком нижнего по потоку конца тонкого листа, образующего разделительный элемент, и ориентированы относительно продольной оси камеры сгорания по существу так же, как и указанный участок нижнего по потоку конца. 2. Диффузор по п.1, в котором наружный диффузионный поток питает систему впрыска топлива камеры сгорания, обводной контур для обвода камеры сгорания снаружи камеры сгорания ...

СПОСОБ ИЗГОТОВЛЕНИЯ КОМПОНЕНТА СТАТОРА

... 1. Способ изготовления компонента статора реактивного двигателя, предназначенного для направления потока газа и передачи усилий, отличающийся тем, что компонент статора собирают в окружном направлении по меньшей мере из двух отлитых в виде отдельных деталей секций (1, 20), которые устанавливают рядом друг с другом и соединяют сваркой. 2. Способ по п.1, отличающийся тем, что поверхность (22, 23, 24, 25, 26), по которой каждый сектор (1, 20) сваривают с другим сектором, по меньшей мере частично проходит по внешнему краю его корпуса. 3. Способ по п.1 или 2, отличающийся тем, что поверхность (22, 23, 24, 25, 26), по которой сваривают секторы, проходит и в радиальном, и в осевом направлениях. 4. Способ по п.3, отличающийся тем, что участок (25, 26) поверхности, по которой сваривают секторы, расположенный в месте перехода радиального участка в осевой, выполнен скругленным. 5. Способ по п.1, отличающийся тем, что поверхность (22, 23, 24, 25, 26), по которой сваривают секторы, выполнена непрерывной ...

ЛОПАТКА ДВОЙНОЙ КРИВИЗНЫ ДЛЯ НАПРАВЛЯЮЩЕГО АППАРАТА ТУРБОМАШИНЫ

... 1. Лопатка направляющего аппарата для вращающегося диска турбомашины, имеющая ортогональные продольную (X), тангенциальную (Y) и радиальную (Z) оси, внутреннюю поверхность (28) и внешнюю поверхность (30), расположенные в радиальном направлении между внутренней кромкой (32) и наружной кромкой (34), а в продольном направлении - между передней кромкой (36) и задней кромкой (38), и множество сечений, центры тяжести которых расположены вдоль компоновочной оси (40), при этом у лопатки имеются внутренняя часть (42а), средняя часть (42b) и наружная часть (42с), причем внутренняя часть расположена в радиальном направлении между внутренней кромкой (32) лопатки и внутренней границей (44) средней части, а наружная часть расположена в радиальном направлении между наружной границей (46) средней части и наружной кромкой (34) лопатки, отличающаяся тем, что компоновочная ось (40) имеет во внутренней и наружной частях лопатки, по существу, радиально направленную тангенциальную составляющую (48), а в средней ...

НАПРАВЛЯЮЩИЙ АППАРАТ КОМПРЕССОРА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ С УПЛОТНЕНИЕМ

Полезная модель относится к области авиадвигателестроения и газотурбинным двигателям (ГТД), в частности к компрессорам газотурбинных двигателей. Направляющий аппарат компрессора газотурбинного двигателя с уплотнением состоит из секторов с радиальными лопатками, наружным и внутренним основаниями. Дополнительно лента плоского профиля с уплотнением устанавливается в кольцевой паз внутреннего основания сектора с возможностью дополнительной фиксации от проворота установкой П-образного фиксатора в карман на краю сектора в направлении вращения ротора. Кроме того, применено уплотнение в виде сотового уплотнения или истираемого покрытия. Применена формованная лента плоского профиля из жаропрочного материала на основе никеля. Кроме того, П-образный фиксатор изготовлен методом селективного лазерного сплавления. П-образный фиксатор установлен, например, в каждый сектор или в несколько секторов направляющего аппарата. Кроме того, сектор направляющего аппарата изготовлен, например, паяным методом или ...

Dichtvorrichtung für drehende Turbinenschaufeln

Es wird eine verbesserte Dichtung für eine Turbine beschrieben, die aufeinanderfolgende Düsendeckel enthält, die einen äußeren Träger für eine radiale Anordnung von statischen Schaufeln liefern, die abwechselnd in axialer Richtung angeordnet sind, wobei radial und axial getragene äußere Dichtungsbauteile Teil einer Dichtung sind, die den Fluidstrom um die Spitze von drehenden Schaufeln verringert, wobei der radiale Träger des äußeren Dichtungsbauteils als ein von Keilen in Stellung gehaltener Ring und eine Umfangserweiterung geformt ist, derart, dass die Keile und die Umfangserweiterung einen ausreichenden Abstand haben, um eine relative radiale Bewegung zwischen dem Ring und dem Gehäuse zu erlauben, während eine Druckdichtungsseite und ein Träger in axialer Richtung gegen das Gehäuse geliefert werden, und wobei der Ring einen ausreichenden Abstand vom Gehäuse und/oder den Düsendeckeln hat, um im Wesentlichen von einer radialen Lageveränderung im Fall einer Drehbewegung der Düsendeckel ...

Verfahren zur Beschichtung einer Turbinenschaufel

Die Erfindung nennt ein Verfahren (2) zur Beschichtung einer Turbinenschaufel (1), welche einen Flügel (4) und wenigstens eine an einem Ende des Flügels (4) angeordnete Plattform (2) umfasst, wobei die oder jede Plattform (2) eine Kontaktzone (6) und wenigstens eine an die Kontaktzone (6) angrenzende flächige Überstandszone (8a, 8b) aufweist, und an der Kontaktzone (6) den Flügel (4) abschließt, mit den Verfahrensschritten: – flügelseitiges Auftragen wenigstens einer ersten Lage (22) einer Beschichtung (24) auf die Plattform (2), und – Entfernen der wenigstens ersten Lage (22) der Beschichtung (24) von wenigstens einer Stirnfläche (10a, 10b) der Plattform (2) im Bereich der Überstandszone (8a, 8b) unter Belassen der wenigstens ersten Lage (22) der Beschichtung (24) an der Stirnfläche (10c) im Bereich der Kontaktzone (6).

DICHTUNGSTEIL, INSBESONDERE DICHTUNGSRING, FUER EIN GASTURBINENTRIEBWERK

Einstellring-Dämpfer

Offenbart wird ein Turbolader mit variabler Turbinengeometrie (100), mit einer Einstellringanordnung (45), die einen Einstellring (50) und eine Reihe von schwenkbaren Leitschaufeln (30) aufweist, die mit dem Einstellring (50) wirkverbunden sind. Ein federbelasteter Sicherungsclip (60) liegt zwischen benachbarten Schaufelhebeln (36) und ist an dem Einstellring (50) angebracht. Der Sicherungsclip (60) übt eine Kraft gegen den benachbarten Schaufelhebel (36) aus und dämpft und verringert die Bewegung der entsprechenden Komponenten.

TURBINENLEITSCHAUFEL, TURBINE UND VERFAHREN ZUM MODIFIZIEREN EINER TURBINENLEITSCHAUFEL

Eine Turbinenleitschaufel (3), die umfasst:einen Leitschaufelkörper (21), der sich im Einbauzustand in der radialen Richtung einer Turbine (T), in die die Turbinenleitschaufel (3) einzubauen ist, erstreckt,einen plattenartigen inneren Deckring oder Shroud (22), der an einem radial inneren Ende des Leitschaufelkörpers (21) vorgesehen ist, undeinen plattenartigen äußeren Deckring oder Shroud (23), der an einem radial äußeren Ende des Leitschaufelkörpers (21) vorgesehen ist, wobeider Leitschaufelkörper (21) einen Serpentinenkanal (30) aufweist, der eine Vielzahl von Hauptkanälen (31) aufweist, die miteinander kommunizieren, und welcher so ausgebildet ist, dass er sich im Inneren des Leitschaufelkörpers (21) in der radialen Richtung windet und durch welchen ein Kühlmedium (c) im Betrieb strömt, wobeiein Deckring (22,23), entweder der innere Deckring (22) oder der äußerer Deckring (23):einen Terminalkanal (31C), der mit einem am weitesten stromabwärtigen Hauptkanal (31B) der Hauptkanäle (31) ...

System und Verfahren zur Halterung von Statorkomponenten

Es ist ein System und ein Verfahren zur Halterung abnehmbarer statischer Komponenten in einer Statoranordnung eines Turbinentriebwerks beschrieben. Das Verfahren weist die Schritte auf: Verbinden einer Statoraufhängeeinrichtung (210), die an einer ersten Stelle (221) an einer ersten statischen Komponente (231) angeordnet ist, mit einem Pfosten (96), der an einer ersten statischen Struktur (91) angeordnet ist, wobei der Pfosten (96) zumindest einen Teil des Gewichts der ersten statischen Komponente (231) trägt, Verbinden eines Statoranschlags (220), der an einer zweiten Stelle (222) an der ersten statischen Komponente (231) angeordnet ist, die in Umfangsrichtung von der ersten Stelle (221) beabstandet angeordnet ist, mit der Statoraufhängeeinrichtung (210), die an einer zweiten statischen Komponente (232) angeordnet ist, und Verbinden einen der ersten statischen Komponente (231) angeordne wobei die zweite statische Struktur (92) zumindest einen Teil des Gewichts der ersten statischen Komponente ...

Reparaturverfahren für Turbinenschaufeln

Die Erfindung betrifft ein Reparaturverfahren für Leitschaufeln (10, 10‘) einer Gasturbine, insbesondere einer Fluggasturbine, umfassend folgende Schritte: Bereitstellen von wenigstens einer zu wartenden Leitschaufel (10, 10‘); Erfassen der Ist-Geometrie der zu wartenden Leitschaufel (10, 10‘) unter Anwendung von wenigstens einem Messverfahren; Vergleichen der durch das berührungslose Messverfahren erfassten Ist-Geometrie mit einer vorbestimmten Soll-Geometrie (SG) für einen entsprechenden Leitschaufeltyp; Berechnung einer Zielgeometrie (ZG) für die zu wartende Leitschaufel (10, 10‘), die soweit möglich der Soll-Geometrie (SG) entspricht, derart dass unter Verwendung von Optimierungsparametern die Soll-Geometrie (SG) der zu wartenden Leitschaufel (10, 10‘) wenigstens abschnittsweise entlang ihrer Strömungskontur angenähert wird; Aufbringen und maschinelles Abtragen von Material (22) an der zu wartenden Leitschaufel (10, 10‘), derart dass die berechnete Zielgeometrie (ZG) hergestellt wird ...

Turbinenleitapparatsegment mit bogenförmiger konkaver Vorderkante

Schaufel für ein Gasturbinentriebwerk

Verdichterstufe mit variabler Statorschaufelneigung

Die Erfindung eine Verdichterstufe (10), insbesondere für ein Nebenstromtriebwerk für Luftfahrtzeuge, mit:- einem um eine Rotationsachse (X) rotierbaren Fan (14); und- einer Mehrzahl von in Umfangsrichtung verteilten Statorschaufeln (28), die stromabwärts von dem Fan (14) angeordnet sind,wobei sich die Statorschaufeln (28) jeweils von einer radial inneren Statornabe (20) zu einem radial äußeren Außengehäuse (18) erstrecken,wobei ein Neigungswinkel (α) einer jeweiligen Statorschaufel (28) variabel ist,wobei die Statorschaufeln (28) in einem Übergangsbereich zum Außengehäuse (18) einen Neigungswinkel (α) mit einem Betrag von weniger als 30° aufweisen und in wenigstens einem radial weiter einwärts liegenden Bereich einen anderen Neigungswinkel (α).

Gasturbine mit gegenläufigen Niederdruckrotoren

VTG-Lader für Fahrzeug

Ein Variable-Turbinengeometrie-Lader (VTG-Lader) für ein Fahrzeug kann aufweisen: ein Turbinenrad; ein Turbinengehäuse, das so gestaltet ist, dass es das Turbinenrad drehbar lagert, und mit einem Kanal zum Empfangen von Abgas von einer radial äußeren Seite des Turbinenrads und zum Abgeben des Abgases in einer axialen Richtung des Turbinenrads versehen ist; einen Scheibenkörper, der in dem Kanal des Turbinengehäuses vorgesehen ist und darin mit einer Umgehungsleitung versehen ist, so dass das Abgas das Turbinenrad umgeht; und eine Vielzahl von Leitschaufeln, die zwischen dem Scheibenkörper und dem Turbinengehäuse angebracht sind, um eine variable Düse zum Steuern einer Strömung des Abgases, das radial einwärts des Turbinenrads strömt, zu steuern.

Gasturbinenleitschaufel

Methode und System zur Drehung eines Turbinenleitschaufelringes

LEITSCHAUFELKRANZ MIT ZWISCHENBODEN FUER TURBINEN

Hohlschaufel

Hohlschaufel in ungekühlter Ausführung für den heißgasbeaufschlagten Turbinenbereich einer Gasturbine, mit einem Schaufelblatt (11) und einem in dem Schaufelblatt vorhandenen Hohlraum (12), dadurch gekennzeichnet, dass in dem Hohlraum (12) ein hitzebeständiges, einen Heißgaseinritt in den Hohlraum behinderndes Füllmaterial (16) angeordnet ist, und dass das Füllmaterial (16) im Hohlraum (12) durch Verschlusselemente (17, 18) fixiert ist, die im Bereich mindestens einer radial außen liegenden Öffnung und/oder mindestens einer radial innen liegenden Öffnung des Hohlraums (12) positioniert sind, und die den Hohlraum (12) zumindest an einer der Öffnungen nicht gasdicht verschließen.

Guide vane ring for thermal fluid flow engine of aircraft, has hooks inserted into recesses of housing parts, and grooves arranged laterally near hooks, where each hook is angularly attached at radial outer guide vane base of guide vane

The ring (3) has hooks (7, 8) inserted into recesses (9, 10) of housing parts (11, 12), and grooves (21) arranged laterally near the hooks, where each hook is angularly attached at a radial outer guide vane base (6) of a guide vane (4). A ring element (24) encompasses the base, where the ring element is provided with a transverse slot, where the ring element is made of a spring elastic metallic material or heat-resistant material e.g. nickel alloy such as nimonic-90 or haynes-25. The ring element comprises circular segments that are connected in a fixed manner with each other.

Mechanical joint for a gas turbine engine

MECHANICAL FLANGE JOINT FOR A GAS TURBINE ENGINE

A joint for a gas turbine engine to reduce thermal gradients across the joints component parts includes, an annular component such as an inner nozzle support (66) having an annular, radially-extending flange 68; an annular second component such as a diffuser having an annular, radially-extending second flange 64 abutting the first flange 68; a plurality of generally radially-extending channels 80,75 passing through at least one of the flanges; a plurality of generally axially-extending channels 92 extending through the second flange 64 and communicating with one of the radial channels 80,75; and a plurality of bolts 74 clamping the flanges together. The exterior surfaces of the annular components are exposed to high temperature gases which can pass from the exterior through the central components of the flange and exit via the axial channels 90. Tabs 104 connected to the fasteners 74 may deflect the flow of hot gas as it exits the axial channels.

Improvements relating to bladed rotary fluid-flow machines

... 693,727. Centrifugal compressors and radialflow turbines. POWER JETS (RESEARCH & DEVELOPMENT), Ltd. April 25, 1951 [Jan. 25, 1950], No. 9698/51. Divided out of 693,686. Classes 110(i) and 110(iii) A diffuser system for a certrifugal compressor or a nozzle system for a radial inward-flow turbine comprises inner and outer walls 4, 5 spaced to provide between them an annular flow passage, at least one of the walls having its inner surface defined by a hyperboloid of revolution of one sheet, and radially extending stator blades 6 of which the roots lie on the straight line generators of the hyperboloid. The hyperboloid shape may be locally modified adjacent the entry or exit. The rotor 1 may be as described in Specification 693,686.

Outlet guide vane for aircraft turbine engine, presenting an improved lubricant cooling function

A heat shield

Turbine support structure

A turbine support structure for a gas turbine engine has a shroud 21 with a front end portion that engages with a rear end portion of a flange 19 of an upstream nozzle 13, and a rear end portion that engages with a front end portion of a flange of a downstream nozzle. At only one of the front end portion or the rear end portion of each flange, a superposition engagement portion 49 is formed at which a turbine casing engagement portion 45, a flange engagement portion 41 and a shroud engagement portion 43 are superimposed on each other in a radial direction, and a support pin 51 penetrates the superposition engagement portion in the radial direction. The nozzle includes a plurality of nozzle division portions (13A, fig 4) and the shroud includes shroud division portions (21A, fig 4). Each nozzle division portion has an outer circumferential flange division portion (19A, fig 4) formed with insertion grooves (63, 65, fig 4) for support pins.

STATOR VANE

Improvements in or relating to fluid pressure apparatus

... 657,366. Axial - flow compressors. WESTINGHOUSE ELECTRIC INTERNATIONAL CO. Feb. 9. 1949. No. 3541. Convention date, Feb. 25, 1948. [Class 110(i)] The diffuser in the annular delivery of an axial flow compressor is fitted with a ring of radially disposed twisted blades Adjacent blades are twisted in opposite senses. Fig. 10 is a circumferential section at the outer ends of three blades. The twist of blade 29 is opposite to that of blades 30. The blade passage on the left diverges from a narrow entry at the inner ends of the blades and converges from a broad entry at their outer ends. The blade passage on the right is similar but inverted.

Gas turbine engine

Filled static structure for axial-flow machine

A turbine comprising baffles situated between turbine blades and guide vanes

A turbine comprises an annular array of nozzle guide vanes 34 and an annular array of turbine blades (24, fig 2) mounted within a casing. An array of radially extending protrusions, eg baffles plates 60, are positioned axially upstream of the guide vanes 34 and protruding inwardly from an inner casing wall 12 so as to mix a tangential momentum component of an overtip leakage fluid flow before it reaches the array of nozzle guide vanes 34. The baffle plates 60 may be positioned in a recess (62, fig 8) which also houses tips of the turbine blades (24). Baffle plates 60 protrude the same distance as a gap between the turbine blade tips and the inner casing wall 12.

Turbine nozzle with purge cavity blend

Turbine casing comprising axially connected rings with integral stator vanes.

A gas turbine engine casing 16 is made up of a plurality of coaxial ring members 19 connected in series. Each ring member 19: is provided with an integral array of radially inwardly directed stator vanes 18; extends frusto-conically away from the stator vanes on the inside; is provided with abradable seal material 22 at the joints above the rotating blades and has a flange 20 at each of its axial extents to facilitate the interconnection of the ring members by bolts 21. The casing rings have 'isogrid' reinforcing ribs on the outer surface and may be surrounded by glass fibre fabric for blade containment.

A stator vane arrangement

The invention provides for variation in stator vane spacing as well as orientation by providing relative axial displacement of vanes 31 within an assembly forming part of a turbine engine. Slots 32 are provided in a normally tapering annular or conical surface 33 such that axial displacement alters the interstitial spacing between adjacent stator vanes 31 and also their orientation.

A stator vane assembly for a turbomachine

Impulse turbine design

An impulse turbine arrangement comprises annular offset vaneless ducts 12 and 13 which position the nozzle (guide vane) rows 8, 9 at a larger radius from the rotor axis than the rotor blades 5. The guide vanes are designed to operate in periodically reversing flows, and rotor blades may have an unconventionally high turning angle of 70 degrees to obtain peak efficiency. Performance may be further enhanced by the incorporation of a boundary layer blowing system into the guide vane design. The system may include blowing holes or slots as well as a compressor to raise the blowing pressure and/or mass flow rate.

Stator assembly including bleed ports for turbine engine compressor

A vane assembly for a gas turbine engine

REDUCING ACOUSTIC SIGNATURE USING PROFILED STATOR ENDWALLS

A stage of a turbine 201a, 201b has a stator vane 104 with an outlet adjacent a rotor blade 102 with a hub 214 and a tip 212. The stator 104 has a plurality of inner and outer end walls extending between the inlet and outlet and defining a plurality of end wall passages 202. The endwall passages having a profile 208, 210 configured to direct the working fluid to a plane substantially tangential to the inner and outer radial limits (fig.3) to reduce noise and thereby increase turbine efficiency by reducing or eliminating radial pressure gradients on rotor blades 102 and their incipient secondary flow vortices which may noisily excite downstream blade and vane rows. Instead of generating inefficient noise, the fluid energy may be properly directed into the shaft as efficient work.

Flow distribution blading comprising an improved sealing plate

The invention concerns a fixed flow distribution blading (10) in a turbomachine, comprising two coaxial annular platforms that are internal (12) and external respectively, linked to each other by a plurality of radial blades (11), in which the internal annular platform (12) comprises an annular radial partition (120) and a blade support ring (121 ) extending to each side of an outer radial end of said partition (120), the blading (10) further comprising an annular sealing plate (20) mounted on the internal annular platform (12), on the upstream side of the radial partition relative to an airflow in the blading, the blading (10) being characterised in that the sealing plate (20) is mounted on the blade support ring (121), and in that it comprises a circumferential groove. The invention also concerns a method for producing such a blading and a turbomachine comprising such a blading.

Vane support in a gas turbine engine

Article, airfoil component and method for forming article

An article 100, which may be an airfoil or airfoil component, comprises manifold 102, article wall 104, post-impingement cavity 106 and a plurality of post-impingement partitions 118. The manifold includes impingement wall 108 defining plenum 110 and a plurality of impingement apertures 112. The article wall includes a plurality of external apertures 114. The post-impingement cavity is disposed between the manifold and the article wall, and is arranged to receive fluid 116 from the plenum through the impingement apertures and exhaust through the external apertures. Post-impingement partitions divide the post-impingement cavity into a plurality of hermetically separated sub-cavities 120. The impingement wall, article wall and post-impingement partitions are integrally formed as a single, continuous article. Impingement apertures may be arranged to distribute fluid to generate higher heat transfer coefficient in a sub-cavity exposed to higher temperatures. Apertures may be arranged to distribute ...

Vane assembly

Interlocking turbine blades

Inward flow radial turbine

Inward flow radial turbine

An inward flow radial turbine 1, rotating as per arrow 2 within a housing 5, associated with an annular vessel formed by co-axial cylinders 6 and 7, and an annular piston 11 operating on a two stroke cycle. On the outward stroke of the piston, energised fluid is admitted via a valve 15 and nozzle 16 to the vessel to form a swirl. On the inward stroke of the piston, valve 15 is closed, and the fluid swirl is displaced through an annular array of vanes 9 to the inducer 3 of the rotating turbine rotor and forced inwards between the turbine blades to exit at the exducer 4, whereby fluid kinetic energy is transferred to the turbine.

Assembly for controlling variable pitch blades

An assembly for the control of variable pitch blades of a turbo-engine of an aircraft. The system comprises a control ring 10 which surrounds a casing 8 of the turbo-engine which is arranged around an axis x of the turbo-engine, the ring being connected by rods 12 to the blades. A device 14 is included for driving the control ring rotationally around the axis x and translationally parallel to the axis x. The device comprises an actuator 30, which is a cylinder having a fixed part 30a and a movable part 30b, the movable part being a grooved rod and moving parallel to the axis x and being engagement with pivot connections 33. The pivot connections rotate about a radial axis 35a and act on the radial shaft 37 which passes through a clevis 39 fixed with the control ring to rotate with is about axis x. The movement of the actuator parallel to axis x acts on the pitch of the blades by actuating pivot connections 33, causing rotation of the radial shaft, causing rotation of the connecting rods ...

Insert comprising an external surface forming part of at least one aerodynamic profile of a turbine engine test vane

Tandem guide vane

Impulse turbine

Stator vane assembly

STATOR BLADES

... 1307367 Stator blades; axial flow compressors DEFENCE SECRETARY OF STATE FOR 16 June 1970 [26 June 1969] 32369/69 Headings F1C and FIT A stator blade, e.g. for an axial flow compressor or nozzle of a gas turbine engine, having radially inner and outer shroud portions 30, 32 has a working portion curved to a shape similar to the shape which would be assumed by a straight but otherwise substantially identical blade in the predominant vibratory mode during use. In Figs.3, 4 the working portions 28, 28a are curved to the shapes of the first and second vibratory modes respectively, whilst in Fig. 5 (not shown) the shape is that of the third vibratory mode. The amplitude of the curve is greater than that assumed by the straight vibrating blade. The construction reduces blade vibration.

Turbomachine

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

Undulating stator for reducing the noise produced by interaction with a rotor

Disclosed is an assembly comprising at least one stator placed radially in a flow which passes through one or more rotors which share the same axis of rotation, the stator having a leading edge 11 and a trailing edge 12 connected by a lower face 10 and an upper face 9. At least one of the faces has radial undulations which extend axially from the leading edge to the trailing edge and have at least two bosses in the same azimuth direction, the amplitude of which is at least one centimetre on at least part of the axial length of the stator. In use, pressure fluctuations with oscillations of the temporal phase are created on the at least one undulating surface and the undulations have azimuth maximums and/or minimums in the vicinity of the zero mean dephasing regions for the pressure on the undulating face. Also disclosed is a turbine engine comprising the assembly and a method for reducing noise in an assembly.

Airfoil for a first stage nozzle guide vane

Turbine nozzles and methods of manufacturing the same

A turbine nozzle is provided and includes a first ring having a first microstructure, a vane extending from the first ring, a first porous zone between the first ring and the vane that is more porous than the first microstructure to attenuate thermo-mechanical fatigue cracking between the vane and the first ring. Methods of manufacturing the turbine nozzle are also provided.

Turbine nozzle with contoured band

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

Turbine engine compressor stator

A gas turbine engine stator segment has a shroud band and a plurality of blade sections. Each of the blade sections has a first section with a first thickness, second section with a second thickness and a fairing section transitioning between the first and second section. The second section thickness is less than the first section thickness.

Noise attenuation panel and a gas turbine component comprising a noise attenuation panel

A noise attenuation panel includes a first wall, a second wall and partition walls connected to the first and second walls and defining cells between the first and second walls. The first wall is provided with a plurality of through holes. At least two of the cells are interconnected via a communication hole. One of the through holes leads to a first of the at least two interconnected cells and a second of the interconnected cells is configured to prevent any gas flow through the second cell.

Gas Turbine Nozzle Arrangement and Gas Turbine

A sealing element is provided for sealing a leak path between a radial outer platform of a turbine nozzle and a carrier ring for carrying said radial outer platform. The carrier ring has an axially facing carrier ring surface and the radial outer platform has an axially facing platform surface. The carrier ring surface forms a first sealing surface and the platform surface forming a second sealing surface. The first and second sealing surfaces is aligned in a plane with a radial gap between them. The sealing element includes a leaf seal adapted to cover the gap between the first and second sealing surfaces, and an impingement plate for allowing impingement cooling of a radial outer surface of the radial outer platform. The impingement plate is adapted to be fixed to the turbine nozzle. The sealing element may be part of a nozzle arrangement of a gas turbine.

Aerofoil blade for an axial flow turbomachine

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

Axial flow turbine

An axial flow turbine includes in axial flow series a low pressure turbine section and a turbine exhaust system. The low pressure turbine section includes a final low pressure turbine stage having a circumferential row of static aerofoil blades followed in axial succession by a circumferential row of rotating aerofoil blades. Each aerofoil blade has a radially inner hub region and a radially outer tip region. The K value, being equal to the ratio of the throat dimension (t) to the pitch dimension (p), of each static aerofoil blade of the final low pressure turbine stage varies along the height of the static aerofoil blade, between the hub region and the tip region, according to a substantially W-shaped distribution.

compressor nozzle stage for a turbine engine

A single-piece compressor nozzle stage for a turbine engine, the stage comprising two coaxial rings, connected together by radial vanes, the inner ring including an annular cavity for housing damper means for damping vibration by friction, which damper means are secured to an annular abradable-material support.

Airfoil core shape for a turbomachine component

A turbomachine component includes a compressor stator vane having an airfoil core shape. The airfoil core shape includes a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in TABLE 1, and wherein X and Y are distances in inches which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z in inches. The profile sections at the Z distances are joined smoothly with one another to form a complete airfoil core shape.

High area ratio turbine vane

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

Turbomachine component having an airfoil core shape

A turbomachine component includes a turbine stator nozzle member having an airfoil core shape. The airfoil core shape includes a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in TABLE 1, and wherein X and Y are distances in inches which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z in inches, the profile sections at the Z distances being joined smoothly with one another to form a complete airfoil core shape.

Transition channel of a turbine unit

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

Alternate shroud width to provide mistuning on compressor stator clusters

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

STEAM TURBINE, BLADE, AND METHOD

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

STATOR VANE ARRAY

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

Stationary blade cascade, assembling method of stationary blade cascade, and steam turbine

A stationary blade cascade 29 of an embodiment includes stationary blade structures 50 and a ring-shaped support structure 40 supporting the stationary blade structures 50 . The stationary blade structures 50 each include: a stationary blade part 51 where steam passes; and an outer circumference side constituent part 52 formed on an outer circumference side of the stationary blade part 51 and having a fitting groove 56 which penetrates all along a circumferential direction and which has an opening 55 all along the circumferential direction in a downstream end surface 54 of the outer circumference side constituent part 52 . The support structure 40 includes a ring-shaped support part 42 having a fitting portion 41 fitted in the fitting grooves 56 of the outer circumference side constituent parts 52 . The plural stationary blade structures 50 are supported along the circumferential direction by the ring-shaped support part 42.

Turbomachine component including a cover plate

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

NOISE-REDUCED TURBOMACHINE

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

Stress relieving slots for turbine vane ring

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

Aluminum airfoil

A method of making an aluminum airfoil includes brazing a first airfoil piece and a second airfoil piece together using a braze material that includes an element selected from magnesium and zinc, to form a braze joint between the first airfoil piece and the second airfoil piece. At least one of the first airfoil piece or the second airfoil piece has an aluminum alloy composition that includes greater than 0.8% by weight of zinc.

Blade cascade and turbomachine

A blade cascade for a turbomachine having a plurality of blades arranged next to one another in the peripheral direction, at least two blades having a variation for generating an asymmetric outflow in the rear area, as well as a turbomachine having an asymmetric blade cascade, which is connected upstream from another blade cascade, are disclosed.

Cooled vane of a turbine and corresponding turbine

A vane is provided for use in a fluid flow of a turbine engine. The vane includes a thin-walled radially extending aerodynamic vane body having axially spaced leading and trailing edges, and a radially outer platform. The wall of the vane body includes an outer shell and an inner shell and defines an interior cavity therein for flowing a cooling medium. A radially extending load strut is arranged at the inner shell of the wall of the leading edge of the vane body.

Nozzle guide vane with cooled platform for a gas turbine

A platform for supporting a nozzle guide vane for a gas turbine is provided. The platform has a gas passage surface arranged to be in contact with a streaming operation gas, and a cooling channel for guiding a cooling fluid within the cooling channel formed in an inside of the platform. A cooling portion of an inner surface of the cooling channel is in thermal contact with the gas passage surface. The platform is an integrally formed part representing a segment in a circumferential direction of the gas turbine. The cooling channel has a first cooling channel portion and a second cooling channel portion arranged downstream of the first cooling channel portion with respect to a streaming direction of the operation gas. The first cooling channel portion and the second cooling channel portion are interconnected.

Compressor

An axial compressor is provided. The compressor has an annular flow channel, in which adjustable blades of a blade ring that extend through the flow channel are rotatably mounted. Each blade has a lug on the rotor-side of the vane for mounting an inner ring. In order to provide a wear-free, self-centering inner ring, the ring is designed as a split ring with two opposing ends and a spring element for spreading the split ring.

Vane assembly for a gas turbine engine

A vane assembly for a gas turbine engine according to an exemplary embodiment of this disclosure includes, among other possible features, a first platform, a second platform spaced from the first platform, and a first variable airfoil that extends radially across an annulus between the first platform and the second platform. One of a radial outer portion and a radial inner portion of the variable airfoil includes a rotational shaft and the other of the radial outer portion and the radial inner portion includes a ball and socket joint that rotationally connect the first variable airfoil relative to the first platform and the second platform.

TURBOJET CASING AND TURBOJET RECEIVING SUCH CASINGS

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

Turbine nozzle

A nozzle arrangement for a gas turbine engine comprising a first housing member and a second housing member. The nozzle arrangement may further include a first nozzle and a second nozzle. Each of the first nozzle and second nozzle may extend between the first housing member and the second housing member so as to form a doublet. A plurality of cooling apertures may be arranged on at least one of the first nozzle, the second nozzle, the first housing member, or the second housing member so as to provide a different degree of first order cooling across the doublet.

Turbine engine stator and method of assembly of the same

A turbine engine stator stage includes a plurality of vanes with each of the plurality of vanes having a camber angle. The plurality of vanes is arranged in a plurality of groups with each group including a pre-determined sequence of vanes. The ordering of vanes within each group is determined by the camber of the individual vanes. This results in an arrangement of vanes within the stator stage which can modify the flow characteristics of the air entering the stator stage to reduce the circumferential pressure variation in the flow region immediately downstream of the stator stage.

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.

Shaped rim cavity wing surface

A shaped rim cavity wing includes an upper surface and a lower surface. The lower surface has a geometric shape to control the separation of airflow as it passes around the lower surface to the top surface. A point of maximum extent defines the boundary between the upper surface and the lower surface, wherein the point of maximum extent defines a corner that that separates airflow from the shaped rim cavity rim and creates a flow re-circulation adjacent to the top surface of the shaped rim cavity wing.

Gas turbine engine turbine vane airfoil profile

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

Turbomachine with variable-pitch vortex generator

The present invention relates to a turbomachine comprising at least one bladed disk, be it mobile or static, and vortex generators ( 17 ) positioned upstream of the blading ( 1 ) of said disk, wherein the vortex generators ( 17 ) are of variable pitch.

STATIC VANE ASSEMBLY FOR AN AXIAL FLOW TURBINE

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

Filled static structure for axial-flow machine

A stator assembly for a rotary machine having a rotor arranged to rotate about an axis in use. The stator assembly has a circumferential support member or casing arranged about said axis and a plurality of elements extending in a substantially radial direction from the support. The elements have a platform at an end thereof for engagement within the support, wherein the elements each comprise a hollow internal cavity having an opening through the platform at the end of the element, wherein said internal cavity is filled with a vibration damping material. The elements may be filled vanes in a gas turbine engine compressor. The platforms may also be filled with the vibration damping material.

Casting of thin wall hollow airfoil sections

A casting mold assembly comprises an airfoil defining section and a casting core. The airfoil defining section includes an outer mold wall and a direct-shelled inner mold wall. The direct-shelled inner mold wall is disposed within a forward chordwise portion of the airfoil defining section. The direct-shelled inner mold wall includes an aft end having an external radius measuring more than about 0.075 in. (1.9 mm). The casting core is secured within an aft chordwise portion of the airfoil defining section, and includes an aft end with an external radius measuring less than about 0.075 in. (1.9 mm). A cast component and a method for making the casting mold assembly are also disclosed.

Method and fixture for airfoil array assembly

An example method of assembling a turbomachine airfoil array includes, among other things, securing a partial airfoil array within a fixture, the partial airfoil array having at least one existing airfoil extending radially between an inner and an outer fairing and an open area where at least one existing airfoil has been removed. The method includes mounting a positioning saddle relative to a base of the fixture, the positioning saddle aligned with the open area, holding a replacement airfoil using the positioning saddle, applying a curable material at an interface between the replacement airfoil and the inner and outer fairing, and curing the curable material while maintaining a relative position between the replacement airfoil and the inner and outer fairing.

DOUBLE FLOW TURBINE HOUSING TURBOCHARGER

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

SEALING BETWEEN A ROTOR DISC AND A STATOR OF A TURBOMACHINE

Assembly including a rotor disc, an adjacent stator and a plurality of sealing elements secured to the rotor disc, the stator including an inner platform and a root bearing at least one abradable element configured to cooperate with the sealing elements, the sealing elements being placed in an enclosure formed by the abradable element, the enclosure being open to the inside and delimited axially by an upstream abradable edge and a downstream abradable edge, the enclosure being delimited radially by an outer abradable edge, at least one of the sealing elements including a first lip configured to cooperate with the upstream abradable edge or the downstream abradable edge, and a second, separate lip configured to cooperate with the outer abradable edge. 1. An assembly for a turbomachine comprising a first mobile wheel extending around an axis and an adjacent bladed turbine stator , said bladed turbine stator being coaxial with said axis and axially offset from said first mobile wheel , said assembly comprising a plurality of sealing elements , each sealing element being secured to said first mobile wheel and projecting radially from said first mobile wheel , said bladed turbine stator comprising an inner platform intended to delimit a gas flow channel in the turbomachine and a root extending radially below the inner platform , said root bearing at a radially inner end at least one abradable element configured to cooperate with the sealing elements , characterised in that the sealing elements are placed in an enclosure formed by said at least one abradable element , said enclosure being open inwards and delimited axially by an upstream abradable edge and a downstream abradable edge , said enclosure being radially delimited by an outer abradable edge , and in that at least one of the sealing elements comprises a first lip configured to cooperate with the upstream abradable edge or the downstream abradable edge , and a second lip separate from the first lip and configured ...

GAS TURBINE ENGINE COMPONENT MANUFACTURING METHOD AND CORE FOR MAKING SAME

A method of manufacturing a gas turbine engine component includes providing a core having a brittle feature, supporting the feature with a first meltable material, arranging the core with the first meltable material in a first mold, and surrounding the core and the first meltable material with a second meltable material to provide a component shape. The method also includes coating the second meltable material with a refractory material to produce a second mold, removing the first and second meltable material, and casting a component in the second mold. 1. A method of manufacturing a gas turbine engine component comprising:providing a core having a brittle feature;supporting the feature with a first meltable material;arranging the core with the first meltable material in a first mold;surrounding the core and the first meltable material with a second meltable material to provide a component shape;coating the second meltable material with a refractory material to produce a second mold; removing the first and second meltable material; andcasting a component in the second mold.2. The method according to claim 1 , wherein the core and feature are constructed from ceramic.3. The method according to claim 2 , wherein the feature has a thickness of less than 0.013 inch and a width of greater than 0.100 inch.4. The method according to claim 1 , wherein the core is an airfoil trailing edge core.5. The method according to claim 4 , wherein the trailing edge core has a thickness of less than 0.013 inch and a width of greater than 0.100 inch claim 4 , and the core includes an integral adjacent core structure that has a thickness of greater than 0.013 inch.6. The method according to claim 5 , wherein the airfoil trailing edge core has multiple holes claim 5 , and the supporting step includes having the first meltable material extend through the holes.7. The method according to claim 5 , wherein the supporting step includes having the first meltable material adjoin the adjacent ...

CAST COMPONENT HAVING CORNER RADIUS TO REDUCE RECRYSTALLIZATION

A cast component includes a cast body that has a single crystal microstructure and an internal corner bounding an internal cavity. The single crystal microstructure defines a critical internal residual stress with respect to investment casting of the cast body using a refractory metal core beyond which the single crystal microstructure recrystallizes under a predetermined condition. The internal corner has a corner radius that is greater than a critical corner radius below which an amount of internal residual stress in the single crystal microstructure exceeds the critical internal residual stress. The internal cavity includes a cross section less than about 20 mils near the corner radius. 1. A cast component comprising:a cast body having a single crystal microstructure and an internal corner bounding an internal cavity, the single crystal microstructure defining a critical internal residual stress with respect to investment casting of the cast body using a refractory metal core beyond which the single crystal microstructure recrystallizes under a predetermined condition, the internal corner defining a corner radius (R) that is greater than a critical corner radius below which an amount of internal residual stress in the single crystal microstructure exceeds the critical internal residual stress, and wherein the internal cavity includes a cross section less than about 20 mils near the corner radius.2. The cast component as recited in claim 1 , wherein the internal cavity is a micropassage with cross section less than about 15 mils.3. The cast component as recited in claim 2 , wherein the micropassage is embedded within an exterior wall of the cast body.4. The cast component as recited in claim 3 , wherein the cast body is an airfoil claim 3 , and the exterior wall is a suction side wall or a pressure side wall of the airfoil.5. The cast component as recited in claim 2 , wherein the micropassage has a cross-section taken perpendicular to a longitudinal direction of ...

ROTOR BLADE OR GUIDE VANE ASSEMBLY

The disclosure refers to a method of assembling or disassembling a rotor blade or guide vane assembly, wherein the rotor blade or guide vane includes an airfoil and additional structured peripheral members forming at least an inner and/or an outer platform of the rotor blade or guide vane. The airfoil includes at least one airfoil sub-structure designed for anchoring at least one superposed component for the purpose of a thermal protection. The connection between the airfoil sub-structure and the superposed component is supported on friction-locked device, wherein the airfoil sub-structure is formed by a spar and the superposed component includes at least one flow-charged outer shell. Connection between the spar or airfoil understructure and flow-charged outer shell is formed by a force-fit and/or a form-fit fixation or a shrinking joint, wherein the airfoil and additional structured peripheral members is formed by friction-locked device with a detachable, permanent or semi-permanent fixation. 1. Modular rotor blade or guide vane , at least comprising: an airfoil , a platform and a root , wherein the airfoil includes an inner core structure , designed for anchoring at least one shell , and one or more shells , encasing the inner core structure as a whole or in part , one of said shells being an outer shell , forming the outer contour of the blade or vane airfoil and being flow-charged in operating mode , wherein the connection between the inner core structure and the flow-charged outer shell is formed by a force-fit or form-fit fixation or a shrinking joint.2. Modular rotor blade or guide vane according to claim 1 , wherein at least the outer shell is designed in a closed one-piece configuration.3. Modular rotor blade or guide vane according to claim 1 , wherein at least the outer shell is designed in an envelope configuration claim 1 , to be closed after wrapping.4. Modular rotor blade or guide vane according to claim 3 , wherein the joint claim 3 , formed by the ...

GAS TURBINE BLADE

A gas turbine blade includes a blade root and a blade aerofoil, a cooling fluid plenum extending inside the gas turbine blade through the blade root, the blade aerofoil and the blade tip, a blade root impingement plate in the cooling fluid plenum inside the blade root and a blade tip impingement plate in the cooling fluid plenum inside the blade tip, the blade tip impingement plate having at least one cooling fluid hole configured and arranged to enable a cooling fluid to flow from the blade tip into the blade aerofoil via the cooling fluid hole or holes, and a pipe extending in the cooling fluid plenum from the blade root impingement plate to the blade tip impingement plate. The blade root impingement plate can direct the cooling fluid from the blade root to the pipe. 1. A gas turbine blade comprising:a blade root and a blade aerofoil, the blade root being attached to a first end of the blade aerofoil;a blade tip attached to a second end of the blade aerofoil;a cooling fluid plenum extending inside the gas turbine blade through the blade root, the blade aerofoil and the blade tip;a blade root impingement plate in the cooling fluid plenum inside the blade root and a blade tip impingement plate in the cooling fluid plenum inside the blade tip, the blade tip impingement plate having at least one cooling fluid hole configured and arranged to enable a cooling fluid to flow from the blade tip into the blade aerofoil via the cooling fluid hole or holes; anda pipe extending in the cooling fluid plenum from the blade root impingement plate to the blade tip impingement plate, and the pipe being configured and arranged to transport the cooling fluid from the blade root to the blade tip; andthe blade root impingement plate being configured and arranged to direct the cooling fluid from the blade root to the pipe.2. The gas turbine blade of claim 1 , wherein the pipe is attached to the blade tip impingement plate and slidably attached to the blade root impingement plate claim 1 ...

ROTOR OFF-TAKE ASSEMBLY

According to exemplary embodiments, a rotor off-take assembly is provided by positioning an angled hole or aperture in a stator assembly. This angled hole provides improved pressure recovery and utilizes higher dynamic pressure to drive the bleed air flow into the off-take cavity. 1. A rotor off-take assembly for improved pressure recovery , comprising:a first rotor disk, including at least one first blade connected to the first rotor disk and extending radially outwardly;a second rotor disk, including at least one second blade connected to the second rotor disk and extending radially outwardly;an at least one stator assembly disposed between the first rotor disk and the second rotor disk;the stator assembly including a flow surface generally extending from adjacent the first rotor disk toward the second rotor disk;the stator assembly including an off-take aperture extending downwardly at a non-perpendicular angle through the flow surface;wherein air passes through the off-take aperture of the stator assembly reducing swirl.2. The rotor off-take assembly of further comprising a bleed air passage in rotor structure.3. The rotor off-take assembly for improved pressure recovery of further comprising an impeller tube disposed radially inward of the stator.4. The rotor off-take assembly for improved pressure recovery of claim 3 , wherein the impeller tube is of reduced weight due to reduced height.5. The rotor off-take assembly for improved pressure recovery of claim 4 , wherein the offtake aperture and the impeller tube arrangement results in decreased pressure drop between the off-take aperture and impeller tube.6. The rotor off-take assembly for improved pressure recovery of claim 1 , wherein the offtake aperture is circular in cross-sectional shape.7. The rotor off-take assembly for improved pressure recovery of claim 1 , wherein the offtake aperture is oval in cross-sectional shape.8. The rotor off-take assembly for improved pressure recovery of claim 1 , wherein ...

DETECTION METHOD OF SENSOR IN GAS TURBINE

A detection method of a sensor in a gas turbine includes adopting a pressure sensor to measure a pushing force of a push rod; measuring a first rotational angle of a guide vane where a first angle sensor is mounted; measuring a second rotational angle of the guide vane where a second angle sensor is mounted; obtaining a maximum measured rotational angle deviation from the absolute value of a difference value between the first and second rotational angles; calculating a maximum calculated deviation from the pushing force of the push rod; calculating the absolute value of a difference value between the maximum measured deviation and the maximum calculated deviation; and determining that the angle sensors and the pressure sensor have appropriate measurement accuracy; or, if the absolute value is greater than the standard value, determining that the angle and/or pressure sensors require calibration. 1. A method for sensors in a gas turbine , the gas turbine including a cylinder , a plurality of guide vanes , a first angle sensor , a second angle sensor , and a guide vane driving mechanism configured to drive the guide vanes to rotate , the guide vane driving mechanism including a driving ring , a push rod configured to push the driving ring to rotate relative to the cylinder , a pressure sensor configured to measure a thrust of the push rod , a plurality of connecting rods and adjusting rods connecting the guide vanes and the driving ring , and a plurality of elastic support bases connecting the cylinder and the driving ring , the method comprising:measuring thrust of the push rod via the pressure sensor;measuring a first rotation angle of the guide vanes in an installation position of the first angle sensor;measuring a second rotation angle of the guide vanes in an installation position of the second angle sensor;obtaining a measured maximum rotation angle offset according to an absolute value of a difference between said measured first rotation angle and said measured ...

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

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

ARTICLE WITH SECTIONS HAVING DIFFERENT MICROSTRUCTURES AND METHOD THEREFOR

An article includes a body that has a first section and a second section bonded with the first section. The first section is formed with a first material that has a first microstructure and the second section is formed of a second material that has a second, different microstructure. 1. An article comprising:a body including a first section and a second section bonded with the first section, the first section being formed of a first material and having a first microstructure and the second section being formed of a second material and having a second, different microstructure.2. The article as recited in claim 1 , wherein the first microstructure and the second microstructure differ in grain structure.3. The article as recited in claim 1 , wherein the first material and the second material are metallic alloys.4. The article as recited in claim 1 , wherein the first section includes a platform and the second section includes an airfoil claim 1 , the platform being bonded to one end of the airfoil.5. The article as recited in claim 1 , wherein the second section is metallurgically bonded with the first section.6. The article as recited in claim 1 , wherein the first microstructure is a non-single crystal microstructure and the second microstructure is a single crystal microstructure.7. The article as recited in claim 1 , wherein the first material and the second material have equivalent chemical compositions.8. The article as recited in claim 1 , wherein the first material and the second material have different chemical compositions.9. The article as recited in claim 1 , wherein the first material and the second material are superalloys.10. A method of fabricating an article claim 1 , the method comprising:forming a first section of a body of an article from a powder of a first material using additive fabrication, the first section having a first microstructure; andbonding the first section with a second section to form the body of the article, the second section ...

ARTICLES HAVING REDUCED EXPANSION AND HERMETIC ENVIRONMENTAL BARRIER COATINGS AND METHODS FOR THEIR MANUFACTURE

Articles suitable for use as high-temperature machine components include a substrate and an environmental barrier coating disposed over the substrate, where the environmental barrier coating includes at least one hermetic self-sealing layer formed from a mixture including an alkaline earth metal aluminosilicate and a rare-earth silicate, and where the at least one hermetic self-sealing layer exhibits substantially no net remnant or residual expansion when subjected to high temperature heat treatment. The environmental barrier coating can further include a bondcoat disposed between the substrate and the hermetic self-sealing layer, a topcoat disposed over the hermetic self-sealing layer, and/or an intermediate layer disposed between the hermetic self-sealing layer and the bondcoat. The intermediate layer can include a barrier material that is substantially inert with respect to silica. 1. An article comprising:a substrate; andan environmental barrier coating disposed over the substrate,wherein the environmental barrier coating comprises at least one hermetic self-sealing layer formed from a mixture comprising an alkaline earth metal aluminosilicate and a rare-earth silicate, andwherein the at least one hermetic self-sealing layer exhibits substantially no net remnant or residual expansion when subjected to high temperature heat treatment.2. The article according to claim 1 , wherein the mixture comprises the alkaline earth metal aluminosilicate in an amount of between about 10 volume percent and about 50 volume percent of the mixture.3. The article according to claim 1 , wherein the mixture comprises the rare-earth silicate in an amount of between about 50 volume percent and about 90 volume percent of the mixture.4. The article according to claim 1 , wherein the alkaline earth metal aluminosilicate comprises barium strontium aluminosilicate (BSAS).5. The article according to claim 1 , wherein the rare-earth silicate is selected from the group consisting of a rare- ...

GAS TURBINE ENGINE STATOR VANE BAFFLE ARRANGEMENT

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

GAS TURBINE ENGINE THIN WALL COMPOSITE VANE AIRFOIL

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

GUIDE VANE ASSEMBLY VANE BOX OF AN AXIAL TURBINE ENGINE COMPRESSOR

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

Gas turbine engine component having variable width feather seal slot

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a mate face and a feather seal slot axially extending along the mate face, the feather seal slot having a variable width along a portion of its axial length.

CONTOURED BLADE OUTER AIR SEAL FOR A GAS TURBINE ENGINE

A blade outer air seal (BOAS) segment according to an exemplary aspect of the present disclosure includes, among other things, a seal body having a radially inner face that circumferentially extend between a first mate face and a second mate face and axially extend between a leading edge face and a trailing edge face, wherein a radial position of the radially inner face varies at a given axial position. 1. A blade outer air seal (BOAS) segment , comprising:a seal body having a radially inner face that circumferentially extend between a first mate face and a second mate face and axially extend between a leading edge face and a trailing edge face, wherein a radial position of the radially inner face varies at a given axial position.2. The BOAS segment of claim 1 , wherein the given axial position is upstream from a rub track of the radially inner face.3. The BOAS segment of claim 2 , wherein the given axial position is a first given axial position claim 2 , and a radial position of the radially inner face varies at a second given axial position that is downstream from the rub track of the radially inner face.4. The BOAS segment of claim 1 , wherein the radial position of the radially inner face smoothly varies at the given axial position.5. The BOAS segment of claim 1 , wherein the radial position of the radially inner face undulates at the given axial position between positions that are radially closer to the a central axis and positions that are radially further from the central axis.6. The BOAS segment of claim 1 , wherein the radial position of the radially inner face is contoured.7. The BOAS segment of claim 1 , wherein the BOAS includes at least a layer of an additive manufacturing material.8. A blade outer air seal (BOAS) assembly claim 1 , comprising:a BOAS segment including a radial inner face that circumferentially extends between a first mate face and a second mate face and axially extends between a leading edge face and a trailing edge face; andat least ...

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

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

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 ...

AXIAL RETAINING RING FOR TURBINE VANES

A gas turbine engine is described which has first and second turbine vane assemblies with multiple turbine vanes within respective first and second circumferential outer shrouds. The first outer shroud has a first radially extending flange and the second outer shroud has a second radially extending flange. The radially extending first and second flanges each defining an upstream mating surface and a downstream mating surface relative to a direction of air flow through the engine in use. The downstream mating surface of the first flange mates with the upstream mating surface of the second flange. An axial retaining ring axially retains together the first and second flanges, and has an annular body extending between an upstream portion of the body abutted against the upstream mating surface of the first flange and a downstream portion of the body abutted against the downstream mating surface of the second flange. 1. A gas turbine engine having a center axis of rotation , the engine comprising:first and second turbine vane assemblies having multiple turbine vanes within respective first and second circumferential outer shrouds, the first outer shroud having a first radially extending flange and the second outer shroud having a second radially extending flange, the radially extending first and second flanges each defining an upstream mating surface and a downstream mating surface relative to a direction of air flow through the engine in use, the downstream mating surface of the first flange mating with the upstream mating surface of the second flange; andan axial retaining ring axially retaining together the first and second flanges, the axial retaining ring having an annular body extending between an upstream portion of the body abutted against the upstream mating surface of the first flange and a downstream portion of the body abutted against the downstream mating surface of the second flange.2. The turbine section of claim 1 , wherein the first and second flanges ...

Gas turbine engine attachment structure and method therefor

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

GAS TURBINE ENGINE VANE END DEVICES

A turbomachinery component of a gas turbine engine is disclosed having a number of techniques of reducing the effects of a gap flow between an airfoil member of the gas turbine engine and a wall of the gas turbine engine. The airfoil member can be a variable and in one form is a variable turbine vane. In one embodiment a brush seal is included between the vane and the wall. In another form a wear surface is disposed between the vane and the wall. In yet another form a moveable member capable of being actuated to change position can be disposed between the vane and the wall to alter the size of a gap between the two. 1. An apparatus comprising:a moveable airfoil member structured for use in a working fluid flow path of a gas turbine engine; anda brush seal disposed at an end of the moveable airfoil member, the brush seal having a plurality of extensions projecting outwardly and configured to discourage a flow of working fluid through the extensions as the working fluid traverses the working fluid flow path.2. The apparatus of claim 1 , which further includes the gas turbine engine claim 1 , the engine including a plurality of the moveable airfoil members claim 1 , wherein each of the plurality of moveable airfoil members is a rotatable vane that includes a range of travel claim 1 , and wherein the extensions contact the wall in the range of travel.3. The apparatus of claim 1 , wherein the plurality of extensions each have a first end and a second end both disposed toward a distal side of the plurality of extensions claim 1 , the first end and second end connected via a body that is looped around a central member.4. The apparatus of claim 3 , which further includes a crimp to couple the plurality of extensions to the central member claim 3 , the central member extending along a chord of the moveable airfoil member claim 3 , and wherein the moveable airfoil member is a rotatable turbine vane.5. The apparatus of claim 1 , wherein the plurality of extensions cover a ...

GAS TURBINE ENGINE WITH ATTACHED NOSECONE

A gas turbine engine includes a compressor section and a nosecone assembly. The compressor section includes an inlet guide vane assembly including an inner shroud, an outer shroud, and an inlet guide vane extending from the inner shroud to the outer shroud. The nosecone assembly is attached to the inner shroud. 1. A gas turbine engine comprising: an inner shroud;', 'an outer shroud; and', 'an inlet guide vane extending from the inner shroud to the outer shroud; and, 'an inlet guide vane assembly comprising, 'a compressor section comprisinga nosecone assembly attached to the inner shroud.2. The gas turbine engine of claim 1 , wherein the nosecone assembly comprises:a nosecone; anda nosecone support ring connected to an aft end of the nosecone.3. The gas turbine engine of claim 2 , wherein the nosecone is bolted to the nosecone support ring claim 2 , which is bolted to the inner shroud.4. The gas turbine engine of claim 2 , wherein the nosecone comprises a plurality of tabs extending axially aft of the aft end of the nosecone and wherein a plurality of bolts extend radially inward through the nosecone support ring and into the plurality of tabs.5. The gas turbine engine of claim 2 , wherein the plurality of bolts are countersunk into the nosecone support ring such that heads of the plurality of bolts are positioned flush with or below an outer surface of the nosecone support ring.6. The gas turbine engine of claim 2 , wherein a first set of bolts extend substantially radially to connect the nosecone to the nosecone support ring and wherein a second set of bolts extend substantially axially to connect the nosecone support ring to the inner shroud.7. The gas turbine engine of claim 2 , wherein the nosecone support ring comprises:a rim; anda flange ending radially inward of the rim.8. The gas turbine engine of claim 7 , wherein the flange is bolted to the inner shroud.9. The gas turbine engine of claim 7 , wherein the nosecone is bolted to the rim.10. The gas turbine ...

Steam turbine

A steam turbine 10 according to an embodiment includes rotor blades 22 implanted to a turbine rotor 21, stationary blades 26 making up a turbine stage together with the rotor blades 22, diaphragm outer rings 23 including an annular extending part 24 surrounding a periphery of the rotor blades 22, and supporting the stationary blades 26, and diaphragm inner rings 25 supporting the stationary blades 26. The steam turbine 10 further includes an annular slit 40 formed at an inner surface of the diaphragm outer ring 23 between the stationary blades 26 and the rotor blades 22 along a circumferential direction, and communication holes 50 provided in plural at an outer surface of the diaphragm outer ring 23 along the circumferential direction, communicated to the annular slit 40 from the outer surface side, and communicated to an exhaust chamber sucking water films via the annular slit 40.

METHOD AND APPARATUS FOR HANDLING PRE-DIFFUSER AIRFLOW FOR COOLING HIGH PRESSURE TURBINE COMPONENTS

A gas turbine engine is provided comprising a compressor section, a combustor section, a diffuser case module, and a manifold. The diffuser case module includes a multiple of struts within an annular flow path from said compressor section to said combustor section, wherein at least one of said multiple of struts defines a mid-span pre-diffuser inlet in communication with said annular flow path. The manifold is in communication with said mid-span pre-diffuser inlet and said compressor section. 1. A gas turbine engine comprising:a compressor section;a combustor section;a diffuser case module with a multiple of struts within an annular flow path from said compressor section to said combustor section, at least one of said multiple of struts defines a mid-span pre-diffuser inlet in communication with said annular flow path; anda manifold in communication with said mid-span pre-diffuser inlet and said compressor section.2. The gas turbine engine as recited in claim 1 , wherein said manifold communicates a temperature tailored airflow.3. The gas turbine engine as recited in claim 1 , wherein said manifold communicates a temperature tailored airflow thru a heat exchanger.4. The gas turbine engine as recited in claim 3 , wherein said manifold communicates said temperature tailored airflow from said heat exchanger as buffer air.5. The gas turbine engine as recited in claim 4 , wherein said buffer air is communicated thru a buffer passage to one or more bearing compartments.6. The gas turbine engine as recited in claim 1 , wherein said mid-span pre-diffuser inlet supplies a temperature tailored airflow into said manifold.7. The gas turbine engine as recited in claim 1 , wherein said manifold communicates with a high pressure turbine of said compressor section.8. The gas turbine engine as recited in claim 1 , wherein said manifold is generally annular.9. The gas turbine engine as recited in claim 1 , wherein said manifold communicates with a row of rotor blade attachments in ...

GAS TURBINE ENGINE MULTI-VANED STATOR COOLING CONFIGURATION

A stator for a gas turbine engine has a platform supporting multiple vanes that includes first and second vanes respectively. First and second regions are arranged at the same location on the first and second vanes. The first and second regions respectively include first and second cooling hole configurations that are different than one another. 1. A stator for a gas turbine engine comprising:a platform supporting multiple vanes including first and second vanes respectively including first and second regions, the first and second regions arranged at a same location on the first and second vanes, the first and second regions respectively including first and second cooling hole configurations that are different than one another.2. The stator according to claim 1 , wherein the first and second cooling hole configurations corresponding to cooling hole size.3. The stator according to claim 1 , wherein the first and second cooling hole configurations corresponding to cooling hole shape.4. The stator according to claim 3 , wherein the first cooling hole configuration includes an oblong exit claim 3 , and the second cooling hole configuration includes a conical exit.5. The stator according to claim 1 , wherein the first and second cooling hole configurations corresponding to cooling hole density.6. The stator according to claim 6 , wherein the first and second regions are the same size claim 6 , and the first and second cooling configurations each include a different number of cooling holes.7. The stator according to claim 1 , wherein the first and second regions are provided on airfoils.8. The stator according to claim 7 , wherein the first and second regions are provided on pressure sides.9. The stator according to claim 7 , wherein the first and second regions are provided on suction sides.10. The stator according to claim 1 , wherein the first cooling hole configuration includes a cooling hole having a cooling hole axis providing a line of sight that is obstructed by ...

TURBINE ENGINE WHEEL

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

STATOR VANE ARRANGMENT AND A METHOD OF CASTING A STATOR VANE ARRANGMENT

A stator vane arrangement for a turbomachine comprises a radially inner annular structure, a radially outer annular structure and a plurality of circumferentially spaced vanes extending radially between the inner annular structure and the outer annular structure. At least one of the vanes has a passage extending from the inner annular structure to the outer annular structure. The inner annular structure has at least one radially inwardly extending boss and each boss has a passage extending there-through. The passage in each boss is aligned with a corresponding passage in a vane. Each boss comprises a first portion having a first cross-sectional area and a second portion having a second cross-sectional area which is greater than the first cross-sectional area. The first portion of each boss is positioned between and interconnecting the second portion of the boss and the inner annular structure. 1. A stator vane arrangement comprising a radially inner annular structure , a radially outer annular structure and a plurality of circumferentially spaced vanes extending radially between the radially inner annular structure and the radially outer annular structure ,at least one of the vanes having a passage extending there-through from the radially inner annular structure to the radially outer annular structure,the radially inner annular structure having at least one boss extending radially inwardly there-from, the at least one boss having a passage extending there-through, the passage in the at least one boss being aligned with the passage in the at least one vane,the at least one boss comprising a first portion having a first cross-sectional area and a second portion having a second cross-sectional area which is greater than the first cross-sectional area, the first portion of the boss being positioned between and interconnecting the second portion of the boss and the radially inner annular structure, andthe radially inner annular structure, the radially outer annular ...

SEGMENTED FACE SEAL ASSEMBLY AND AN ASSOCIATED METHOD THEREOF

A turbomachine and a method of operating the turbomachine are disclosed. The turbomachine includes a stator, a rotor including a rotor bearing face, and a face seal assembly including a first segmented seal ring and a second segmented seal ring. The first segmented seal ring includes a plurality of joints and a first flat-contact surface and the second segmented seal ring includes a plurality of segment ends and a second flat-contact surface. One of the first and second segmented seal rings includes a seal bearing face. The second segmented seal ring is coupled to the first segmented seal ring such that the second flat-contact surface is in contact with the first flat-contact surface. The plurality of segment ends is circumferentially offset from the plurality of joints. The first segmented seal ring is slidably coupled to the stator and defines a face seal clearance between the rotor and seal bearing faces. 1. A turbomachine comprising:a stator;a rotor comprising a rotor bearing face; and a first segmented seal ring comprising a plurality of joints and a first flat-contact surface; and', 'a second segmented seal ring comprising a plurality of segment ends and a second flat-contact surface,, 'a face seal assembly comprisingwherein one of the first segmented seal ring and the second segmented seal ring comprises a seal bearing face, wherein the second segmented seal ring is coupled to the first segmented seal ring such that the second flat-contact surface is in contact with the first flat-contact surface, wherein the plurality of segment ends is circumferentially offset from the plurality of joints, and wherein the first segmented seal ring is slidably coupled to the stator and defines a face seal clearance between the rotor bearing face and the seal bearing face.2. The turbomachine of claim 1 , wherein the first segmented seal ring further comprises a circumferential slot extending inwards from a first peripheral side towards a second peripheral side of the first ...

Turbomachine blade and method for the manufacture of same

A blade of a turbomachine includes a blade body of composite material having a fiber reinforcement having a three-dimensional weave and densified by a matrix, the reinforcement having a first part extended by a second, end, part including two segments separated from each other; and an insert having a pi-shaped section, the insert having a platform part and two longitudinal flanges separated from each other, the platform part including a housing delimited by a bottom wall and a rim, the bottom wall including an opening communicating with the space between the two flanges, the first part of the fiber reinforcement being clamped between the two flanges of the insert, the segments of the second part of the fiber reinforcement being folded against the bottom wall of the housing.

AIRCRAFT TURBOFAN ENGINE HAVING VARIABLE PITCH FAN AND METHOD OF OVER-PITCHING THE VARIABLE PITCH FAN IN AN ENGINE OUT CONDITION TO REDUCE DRAG

There is provided a turbofan engine for an aircraft. The turbofan engine has a core with a fan cowl and a variable pitch fan (VPF) configured to only rotate in a first rotation direction. The VPF has a plurality of fan blades each configured to over-pitch to an over-pitch position relative to a feathered position. The turbofan engine has outer guide vanes (OGVs) axially disposed downstream of the VPF, and has a rotation control device to prevent the VPF from rotating in a second rotation direction opposite the first rotation direction, during an engine out (EO) condition of the turbofan engine. When the VPF is prevented from rotating during the EO condition, the fan blades are over-pitched to the over-pitch position relative to the feathered position, to achieve no or minimal air flow separation about the OGVs, and to reduce drag of the turbofan engine during the EO condition. 1. A turbofan engine for an aircraft , the turbofan engine comprising:a core with a fan cowl surrounding a portion of the core, the core having a first end and a second end;a variable pitch fan (VPF) coupled to the first end of the core and configured to only rotate in a first rotation direction, the variable pitch fan having a plurality of fan blades extending radially outward from the core to the fan cowl, the plurality of fan blades each configured to over-pitch to an over-pitch position relative to a feathered position of the variable pitch fan;a plurality of outer guide vanes (OGVs) axially disposed downstream of the variable pitch fan, each of the plurality of outer guide vanes (OGVs) being nonrotatable and extending radially outward from the core to the fan cowl; anda rotation control device coupled to the variable pitch fan, to prevent the variable pitch fan from rotating in a second rotation direction opposite the first rotation direction, during an engine out (EO) condition of the turbofan engine for the aircraft,wherein when the variable pitch fan is prevented from rotating in the ...

VARIABLE STATOR VANE ARRANGEMENT

A variable stator vane arrangement is provided in which the variable stator vanes extend from a first end at a radially inner flow boundary to a second end at a radially outer flow boundary. At least one of the radially inner flow boundary and the radially outer flow boundary is faceted, such that the surface of the faceted flow boundary comprises flat portions at the interfaces with the respective first or second end of each stator vane. The flat portions mean that the tips of the variable stator vanes can be made substantially flush with the flat casing portions. This may improve aerodynamic efficiency and/or increase the design flexibility on where to position the pivot axis of the variable stator vanes. 1. A compressor for a gas turbine engine comprising:a radially inner flow boundary;a radially outer flow boundary;an annular array of variable stator vanes, each stator vane extending from a first end at the radially inner flow boundary to a second end at the radially outer flow boundary, wherein:at least one of the radially inner flow boundary and the radially outer flow so boundary is faceted, such that the surface of the faceted flow boundary comprises flat portions at the interfaces with the respective first or second end of each stator vane.2. The compressor according to claim 1 , wherein:each stator vane is pivotable about a pivot axis; andthe flat portions of the faceted flow boundary are perpendicular to the pivot axis of the respective stator vane at each interface.3. The compressor according to claim 1 , wherein each stator vane comprises:an aerofoil portion; anda boundary interface portion, whereinthe boundary interface portion is a flat surface lying in the same plane as the adjacent flat portion of the faceted flow boundary.4. The compressor according to claim 3 , wherein the boundary interface portion is circular.5. The compressor according to claim 3 , wherein there is substantially no gap between the boundary interface portion and the surrounding ...

COMBUSTOR INLET MIXING SYSTEM WITH SWIRLER VANES HAVING SLOTS

A combustor inlet mixing system () formed from a plurality of circumferentially spaced swirler vanes () extending radially outward from a nozzle hub. Each of the swirler vanes () may have a length () that extends downstream along at least a portion of the combustor inlet mixing system (), and may further have a thickness () that extends along a circumference of the nozzle hub. At least one of the swirler vanes () may further have at least one slot () cut entirely through the thickness () of a portion of the swirler vane (). The slot () may separate the swirler vane () from the nozzle hub along a portion of the length () of the swirler vane (). 117-. (canceled)18. A turbine engine , comprising:at least one combustor positioned upstream from a rotor assembly, wherein the rotor assembly includes at least one row of turbine blades extending radially outward from a rotor;a compressor positioned upstream from the at least one combustor;at least one compressor exhaust plenum extending between the compressor and the at least one combustor; andat least one combustor inlet mixing system formed from a plurality of circumferentially spaced swirler vanes extending radially outward from a nozzle hub, each of the plurality of swirler vanes having a length that extends downstream along at least a portion of the at least one combustor inlet mixing system and further having a thickness that extends along a circumference of the nozzle hub, wherein at least one swirler vane of the plurality of swirler vanes further has at least one slot cut entirely through the thickness of a portion of the at least one swirler vane, the at least one slot separating the at least one swirler vane from the nozzle hub along a portion of the length of the at least one swirler vane.19. The turbine engine of claim 18 , wherein the at least one slot is configured to add a layer of at least partially non-swirling air around the nozzle hub.20. The turbine engine of claim 18 , wherein the at least one slot is ...

VANE ANGLE SYSTEM ACCURACY IMPROVEMENT

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

GAS TURBINE ENGINE WITH MORPHING VARIABLE COMPRESSOR VANES

A stator vane for a gas turbine engine section includes a stator vane having an airfoil extending between a leading edge and a trailing edge. The airfoil has a suction side and a pressure side. There is at least one piezoelectric actuator for changing a shape of at least one of the leading edge and the trailing edge. A gas turbine engine is also disclosed. 1. A stator vane for a gas turbine engine section comprising:a stator vane having an airfoil extending between a leading edge and a trailing edge, said airfoil having a suction side and a pressure side, and there being at least one piezoelectric actuator for changing a shape of at least one of said leading edge and said trailing edge.2. The stator vane as set forth in claim 1 , wherein at least one piezoelectric actuator is mounted on each of said suction and pressure sides claim 1 , with one of said piezoelectric actuators being controlled to contract and the other being controlled to expand to change the position of said leading edge relative to said trailing edge.3. The stator vane as set forth in claim 2 , wherein said piezoelectric actuators are mounted within pockets in said suction and pressure sides.4. The stator vane as set forth in claim 2 , wherein said piezoelectric actuators are operable to change a position of said leading edge about a virtual hinge axis while changing the position of the trailing edge to a lesser extent.5. The stator vane as set forth in claim 2 , wherein said piezoelectric actuators are operable to change a position of said trailing edge about a virtual hinge axis while changing the position of the leading edge to a lesser extent.6. The stator vane as set forth in claim 2 , wherein said airfoil is connected to inner and outer platforms.7. The stator vane as set forth in claim 6 , wherein there is an elastomeric material between said airfoil and said radially inner and outer platforms to accommodate movement of at least one of said leading and trailing edges.8. The stator vane as ...

Fiber-Reinforced Aircraft Component and Aircraft Comprising Same

An air inlet deflector for a structure having an air inlet. The deflector may be retractable within the structure, may be integrally formed with the structure, and may prevent the structure from ingesting foreign matter, such as birds. The deflector may include a series of ribs, spokes, or vanes that may vary in width and/or thickness from fore to aft, and/or may be curvilinear in one or more planes of view, and/or may serve double duty as inlet vanes for redirecting inlet air.

AIRFOIL ASSEMBLY WITH A SCALLOPED FLOW SURFACE

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

NON-CONTACT SEAL ASSEMBLY FOR ROTATIONAL EQUIPMENT

Assemblies are provided for rotational equipment. One of these assemblies includes a first bladed rotor assembly, a second bladed rotor assembly, a stator vane assembly, a stator structure and a seal assembly. The second bladed rotor assembly includes a rotor disk structure. The stator vane assembly is axially between the first and the second bladed rotor assemblies. The stator structure is mated with and radially within the stator vane assembly. The seal assembly is configured for sealing a gap between the stator structure and the rotor disk structure, wherein the seal assembly includes a non-contact seal. 1. An assembly for rotational equipment , the assembly comprising:a first bladed rotor assembly;a second bladed rotor assembly including a rotor disk structure;a stator vane assembly axially between the first and the second bladed rotor assemblies;a stator structure mated with and radially within the stator vane assembly; anda seal assembly configured for sealing a gap between the stator structure and a seal land of the rotor disk structure, wherein the seal assembly includes a non-contact seal, and the seal land is configured as a cantilevered tubular body.2. The assembly of claim 1 , wherein the non-contact seal is a hydrostatic non-contact seal.3. The assembly of claim 1 , wherein the non-contact seal comprises:an annular base;a plurality of shoes arranged around and radially adjacent the rotor disk structure; anda plurality of spring elements, each of the spring elements radially between and connecting a respective one of the shoes to the base.4. The assembly of claim 3 , wherein the base is configured with a monolithic full hoop body.5. The assembly of claim 1 , wherein the stator structure is floating radially within the stator vane assembly.6. The assembly of claim 1 , whereinthe first and the second bladed rotor assemblies are turbine rotor assemblies; andthe first bladed rotor assembly is upstream of the second bladed rotor assembly.7. The assembly of ...

SYSTEM AND METHOD FOR AIR INJECTION PASSAGEWAY INTEGRATION AND OPTIMIZATION IN TURBOMACHINERY

Systems and methods for air injection passageway integration and optimization in turbomachinery using surface vortex generation. An airfoil including a leading edge, a trailing edge, a pressure side, and a suction side, and is configured to influence an airflow as it passes from the leading edge to the trailing edge. The airfoil defines an aerodynamic passageway having an inlet on the pressure side and an outlet on the suction side to deliver air from the airflow through the airfoil to the suction side. The outlets are configured to inject the air at areas on either airfoil side targeted due to their propensity to generate undesirable boundary layer growth and associated flow losses. Outlet may also be included in the hub and the shroud of the turbomachine. 1. A turbomachine system comprising: the airfoil is configured to influence an airflow as it passes from the leading edge to the trailing edge,', 'the airfoil defines an aerodynamic passageway having an inlet and an outlet,', 'the aerodynamic passageway is configured to deliver air from the airflow through the airfoil to a target area, and', 'the outlet is configured to inject the air at the target area, which is targeted due to a propensity to generate undesirable flow losses., 'an airfoil including a leading edge, a trailing edge, a pressure side, and a suction side, wherein2. The system of claim 1 , comprising a shroud and a hub configured to channel airflow claim 1 , wherein the airfoil is disposed between the shroud and the hub.3. The system of claim 2 , wherein at least one of the hub and the shroud includes a part of the aerodynamic passageway.4. The system of claim 3 , wherein the passageway comprises a cavity that has a segment in each of the airfoil claim 3 , the hub and the shroud.5. The system of claim 1 , wherein the aerodynamic passageway couples the inlet with the outlet and comprises an enlarged internal cavity in the airfoil.6. The system of claim 1 , wherein the inlet comprises plural inlets ...

Steam Turbine Hollow Blade

A steam turbine hollow stationary blade is able to reduce the amount of water droplets captured on a blade surface. The steam turbine hollow stationary blade, which has a cavity therein, includes a partition wall dividing the cavity into a pressure chamber on a leading edge side and an exhaust chamber on a trailing edge side, at least one steam inlet hole connecting the pressure chamber and an outside of the stationary blade to each other, and at least one pressure conditioning hole connecting the pressure chamber and the exhaust chamber. Total opening area of the pressure conditioning hole is smaller than total opening area of the steam inlet hole. 1. A steam turbine hollow stationary blade , that has a cavity therein , comprising:a partition wall dividing the cavity into a pressure chamber on a leading edge side and an exhaust chamber on a trailing edge side;at least one steam inlet hole connecting the pressure chamber and an outside of the stationary blade to each other; andat least one pressure conditioning hole connecting the pressure chamber and the exhaust chamber, whereintotal opening area of the pressure conditioning hole is smaller than total opening area of the steam inlet hole.2. The steam turbine hollow stationary blade according to claim 1 , Whereinthe steam inlet hole is positioned on a leading edge of the stationary blade.3. The steam turbine hollow stationary blade according to claim 1 , whereinthe pressure conditioning hole is provided to the partition wall at both inner and outer circumferential sides of the partition wall.4. The steam turbine hollow stationary blade according to claim 1 , further comprising:a slit connecting the exhaust chamber to the outside of the stationary blade.5. The steam turbine hollow stationary blade according to claim 1 , further comprising:an exhaust hole connecting the exhaust chamber to a steam condenser. The present invention relates to a steam turbine hollow stationary blade.In steam turbines, during the process ...

Stator Blade Sector for an Axial Turbomachine with a Dual Means of Fixing

The present application relates to a stator blade sector configured to be fixed to a housing of an axial turbomachine, the sector having a plurality of blades with platforms juxtaposed, so as to describe an arc of a circle. At least one of the platforms comprises on its outer face a fixing screw and at least one other platform has no fixing screws, the platforms being fixed together at their adjacent edges. The application also relates to a stator or portion of stator having a housing forming a generally circular wall and blade sectors arranged along the wall. The housing includes several parts connected to each other by longitudinal flanges. Platforms with no fixing screws are located opposite the flanges. 1. A stator blade sector for attachment to a housing of an axial turbomachine , comprising:a plurality of blades with platforms juxtaposed, so as to describe the arc of a circle, and with an airfoil projecting from the inner face of each platform and directed towards the center of the circular arc described by the platforms;wherein at least one of the platforms comprises:a fixing screw on the outer face thereof and at least one other platform having no fixing screws, the platforms being fixed together at their adjacent edges.2. The stator blade sector in accordance with claim 1 , further comprising:three blades with a central blade and two lateral blades on either side of the central blade, the platform of the central blade being the platform that has no fixing screw, the two platforms of the lateral blades being the platforms having fixing screws.3. The stator blade sector in accordance with claim 2 , wherein at least one of the edges of the platforms with fixing screws forms one end of the sector and includes a shoulder configured to overlap an adjacent edge of an adjacent sector.4. The stator blade sector in accordance with claim 1 , further comprising:three blades with a central blade and two lateral blades on either side of the central blade, the platform of ...

METHOD FOR GENERATING AN AIRFOIL INCLUDING AN AERODYNAMICALLY-SHAPED FILLET AND AIRFOILS INCLUDING THE AERODYNAMICALLY-SHAPED FILLET

Methods are provided for generating an airfoil including an aerodynamically-shaped fillet. 2D fillet curve offset values are defined by obtaining a 2D airfoil image having airfoil surface line and transition region interconnecting airfoil surface line to flowpath surface line with 2D fillet curve. Flowpath offset lines are generated on image at predetermined flowpath offset values and, in the transition region, graphically represent a flowpath offset surface. Intersection point between 2D fillet curve and each flowpath offset line is determined Distance between airfoil surface line and each intersection point generates airfoil offset values. Airfoil and flowpath offset surfaces according to airfoil offset values and flowpath offset values respectively and 3D fillet streamline curves are generated on computer model to define the aerodynamically-shaped fillet. Each 3D fillet streamline curve is an intersection between the airfoil and flowpath offset surfaces. 1. A method for generating an airfoil including an aerodynamically-shaped fillet , the method comprising: obtaining a 2D image of an airfoil having an airfoil surface line and including a transition region interconnecting the airfoil surface line to a flowpath surface line with a 2D fillet curve;', 'generating a plurality of flowpath offset lines on the 2D image at pre-determined flowpath offset values, each flowpath offset line of the plurality of flowpath offset lines in the transition region graphically representing a flowpath offset surface of a plurality of flowpath offset surfaces;', 'determining an intersection point between the 2D fillet curve and each flowpath offset line;', 'computing a distance between the airfoil surface line and each intersection point to generate a plurality of airfoil offset values;, 'defining a plurality of two-dimensional (2D) fillet curve offset values bygenerating a plurality of airfoil offset surfaces according to each airfoil offset value of the plurality of airfoil offset ...

Orifice element for turbine stator and/or rotor vanes

An orifice element is adapted to be inserted into a recess formed at an external opening of a channel in a turbine stator or rotor vane, the channel being adapted for leading a cooling fluid through the vane. The orifice element has a mounting part formed of a solid material, and an opening part leaving an opening between a first side of the orifice element and a second side of the orifice element, the second side being opposite to the first side.

COMPRESSOR BLADE OR VANE HAVING AN EROSION-RESISTANT HARD MATERIAL COATING

A compressor blade for a gas turbine is provided. The compressor blade has a blade substrate that consists of a metal alloy and has an aluminum diffusion zone on a surface of the blade substrate. In addition, the compressor blade has a hard material coating provided on the surface of the blade substrate. A compressor that has a compressor blade and a method of producing such a compressor blade is also provided. 111-. (canceled)12. A compressor blade or vane for a gas turbine , the compressor blade or vane comprising:a blade or vane substrate;a metal alloy;an aluminum diffusion zone on a surface of the blade or vane substrate as a result of the diffusion of aluminum into a surface on the blade or vane substrate; anda hard material coating arranged on the surface of the blade or vane substrate .13. The compressor blade or vane of claim 12 , wherein the hard material coating comprises TiN claim 12 , TiAlN claim 12 , AlTiN claim 12 , CrN as single-layer or multi-layer ceramics or comprises TiN claim 12 , TiAlN claim 12 , AlTiN claim 12 , CrN as single-layer or multi-layer ceramics.14. The compressor blade or vane of claim 12 , wherein the aluminum diffusion zone has a thickness of 10 to 30 micrometers.15. The compressor blade or vane of claim 12 , wherein the aluminum diffusion zone has an aluminum proportion of 0.05 to 0.2% by weight.16. The compressor blade or vane of claim 12 , wherein the metal alloy is a creep-resistant steel.17. A compressor for a gas turbine and having a plurality of compressor blades or vanes claim 12 , wherein at least one compressor blade or vane of the plurality of compressor blades or vanes is designed as claimed in .18. The compressor of claim 12 , wherein the plurality of compressor blades or vanes are arranged in a plurality of rows claim 12 , wherein each row of the plurality of rows has a plurality of compressor blades or vanes arranged transversely to a main direction of flow of the compressor claim 12 , and wherein the plurality of ...

COMPOSITE VANE AND METHOD FOR MANUFACTURING COMPOSITE VANE

A composite vane includes a composite vane body that is formed from a composite material of a thermosetting resin or a thermoplastic resin and reinforced fibers, which is obtained by molding, and a metal sheath that is bonded to a leading edge section including a leading edge of the composite vane body and a vane surface in a vicinity of the leading edge via a film adhesive formed by impregnating a mesh with a hard adhesive to cover the leading edge section, wherein an underfill section that is formed in a step of removing excessive thicknesses parts remaining on the leading edge after the molding and does not need leading edge round finish is placed on the leading edge of the leading edge section in the composite vane body. It is possible to realize reduction of manufacturing time and manufacturing cost. 1. A composite vane , comprising:a composite vane body that is formed from a composite material of a thermosetting resin or a thermoplastic resin and reinforced fibers, which is obtained by molding; anda metal sheath that is bonded to a leading edge section including a leading edge of the composite vane body and a vicinity of the leading edge via a film adhesive formed by impregnating a mesh with a hard adhesive to cover the leading edge section,wherein an underfill section that is formed in a step of removing an excessive thickness part remaining on the leading edge after the molding and does not need leading edge round finish is placed on the leading edge of the leading edge section in the composite vane body.2. The composite vane according to claim 1 ,wherein the underfill section is placed at a plurality of positions in a vane width direction of the leading edge, and functions as adhesive gathering spots for a hard adhesive in the film adhesive.3. A method for manufacturing the composite vane according to claim 1 , whereinan underfill section that does not need leading edge round finish is formed on the leading edge, in a step of removing an excessive thickness ...

TURBINE STATOR VANE OF CERAMIC MATRIX COMPOSITE

A stator vane is comprised of: an airfoil section elongated in a radial direction relative to an axis; an outer band section continuous to an outer end of the airfoil section and bent in a circumferential direction relative to the axis; a first hook section continuous to a leading end in the axial direction of the outer band section and bent outward in the radial direction; a second hook section continuous to a trailing end in the axial direction of the outer band section and bent outward in the radial direction; an inner band section continuous to an inner end of the airfoil section and bent in the circumferential direction; a flange section continuous to an end in the axial direction of the inner band section and bent inward in the radial direction; and a reinforcement fiber fabric continuous throughout these sections and unitized with a ceramic. 1. A stator vane arranged around an axis to form a turbine nozzle , comprising:an airfoil section elongated in a radial direction relative to the axis;an outer band section continuous to an outer end of the airfoil section and bent in a circumferential direction relative to the axis;a first hook section continuous to a leading end in the axial direction of the outer band section and bent outward in the radial direction;a second hook section continuous to a trailing end in the axial direction of the outer band section and bent outward in the radial direction;an inner band section continuous to an inner end of the airfoil section and bent in the circumferential direction;a flange section continuous to an end in the axial direction of the inner band section and bent inward in the radial direction; anda reinforcement fiber fabric continuous throughout the airfoil section, the outer band section, the first hook section, the second hook section, the inner band section and the flange section and unitized with a ceramic.2. The stator vane of claim 1 , further comprising:a cutout ranging from a leading edge in the axial direction ...

Manufacturing of single or multiple panels

A method of manufacturing of a structured cooling panel includes cutting of desized 2D ceramic into tissues; slurry infiltration in the tissues by at least one knife blade coating method; laminating the tissues in a multi-layer panel, with slurry impregnation after each layer, wherein the tissue has combined fibres and/or pre-build cooling holes; drying; de-moulding; sintering the multi-layer panel, wherein part of the combined fibres burns out during the sintering process leaving a negative architecture forming the cooling structure and/or the pre-build cooling holes define the cooling structure; finishing, using of i) post-machine, and/or ii) surface smoothening/rework, and/or iii) coating application, and/or other procedures.

COMPONENTS WITH COOLING CHANNELS AND METHODS OF MANUFACTURE

A component is provided and includes a substrate comprising an outer and an inner surface, where the inner surface defines at least one hollow, interior space. The component defines one or more grooves, where each groove extends at least partially along the outer surface of the substrate and has a base and a top. The base is wider than the top, such that each groove comprises a re-entrant shaped groove. One or more access holes are formed through the base of a respective groove, to connect the groove in fluid communication with the respective hollow interior space. Each access hole has an exit diameter D that exceeds the opening width d of the top of the respective groove. The diameter D is an effective diameter based on the area enclosed. The component further includes at least one coating disposed over at least a portion of the surface of the substrate, wherein the groove(s) and the coating together define one or more re-entrant shaped channels for cooling the component. A method for manufacturing the component is also provided. A method for manufacturing a component is also provided, where the groove and the access hole(s) are machined as a single continuous process, such that the groove and the access hole(s) form a continuous cooling passage. 1. A component comprising:a substrate comprising an outer surface and an inner surface, wherein the inner surface defines at least one hollow, interior space, wherein the component defines one or more grooves, wherein each groove extends at least partially along the substrate and has a base and a top, wherein the base is wider than the top, such that each groove comprises a re-entrant shaped groove, wherein one or more access holes are formed through the base of a respective groove, to connect the groove in fluid communication with the respective hollow interior space, wherein each access hole has an exit diameter D that exceeds an opening width d of the top of the respective groove, wherein the diameter D is an effective ...

Gas turbine engine airfoil leading edge cooling

An example gas turbine engine component includes an airfoil having a leading edge area, a first circuit to cool a first section of the leading edge area, and a second circuit to cool a second section of the leading edge area. The first circuit separate and distinct from the second circuit within the airfoil.

GAS TURBINE ENGINE COMPONENT WITH TWISTED INTERNAL CHANNEL

A gas turbine engine component includes a component body that defines an internal micro-channel that extends in a lengthwise direction along a reference line. The internal micro-channel extends between a first reference position along the reference line and a second reference position along the reference line. The internal micro-channel twists at least 180 with respect to the reference line between the first reference position and the second reference position. 1. A gas turbine engine component comprising:a component body defining an internal micro-channel extending in a lengthwise direction along a reference line, the internal micro-channel extending between a first reference position along the reference line and a second reference position along the reference line, the internal micro-channel twisting at least 180° with respect to the reference line between the first reference position and the second reference position.2. The gas turbine engine component as recited in claim 1 , wherein the internal micro-channel twists at least 360° with respect to the reference line between the first reference position and the second reference position.3. The gas turbine engine component as recited in claim 1 , wherein the internal micro-channel twists multiple full revolutions with respect to the reference line.4. The gas turbine engine component as recited in claim 1 , further including at least one additional internal micro-channel also twisting at least 180° with respect to the reference line between the first reference position and the second reference position.5. The gas turbine engine component as recited in claim 1 , further including at least one additional internal micro-channel that is symmetrically arranged to the internal micro-channel with respect to the reference line.6. The gas turbine engine component as recited in claim 1 , wherein the internal micro-channel is helical.7. The gas turbine engine component as recited in claim 1 , further including a plurality of ...

GAS TURBINE ENGINE COMPONENT COOLING CHANNELS

A component according to an exemplary aspect of the present disclosure includes, among other things, a wall, a first channel extending at least partially through the wall to a first outlet, and a second channel adjacent to the first channel and extending to a second outlet. The first channel is configured to communicate a cooling fluid along a first swirl flow path and the second channel is configured to communicate the cooling fluid along a second swirl flow path that is opposite the first swirl flow path. 1. A component , comprising:a wall;a first channel extending at least partially through said wall to a first outlet;a second channel adjacent to said first channel and extending to a second outlet; andsaid first channel configured to communicate a cooling fluid along a first swirl flow path and said second channel configured to communicate said cooling fluid along a second swirl flow path that is opposite said first swirl flow path.2. The component as recited in claim 1 , wherein said component is one of a blade claim 1 , a vane claim 1 , a blade outer air seal (BOAS) claim 1 , a combustor liner and a turbine exhaust case liner.3. The component as recited in claim 1 , wherein at least one of said first channel and said second channel are micro-channels.4. The component as recited in claim 1 , wherein at least one of said first channel and said second channel include a maximum diameter of less than 0.635 millimeters.5. The component as recited in claim 1 , wherein each of said first channel and said second channel extend along an axis and include a plurality of twists.6. The component as recited in claim 1 , wherein at least one of said first channel and said second channel twists at least one full rotation about an axis that extends through said at least one of said first channel and said second channel.7. The component as recited in claim 1 , wherein at least one of said first channel and said second channel is helical shaped.8. The component as recited in claim ...

COMPOSITE AIRFOIL METAL LEADING EDGE ASSEMBLY

An airfoil assembly () comprises a composite airfoil () having a leading edge () and a trailing edge (), a pressure side () extending between the leading edge and the trailing edge, a suction side () extending between the leading edge and the trailing edge, opposite the leading edge, a metallic leading edge assembly () disposed over the composite air-foil, the metallic leading edge assembly including a high density base (), the metallic leading edge assembly also including a nose () disposed over the base, an adhesive bond layer disposed between the composite airfoil and the metallic leading edge assembly. 1. An airfoil assembly , comprising: a leading edge and a trailing edge;', 'a pressure side extending between said leading edge and said trailing edge;', 'a suction side extending between said leading edge and said trailing edge, opposite said leading edge;, 'a composite foil havinga metallic leading edge assembly disposed over said composite foil;said metallic leading edge assembly including a high density base;said metallic leading edge assembly also including a nose disposed one of over or under said base;an adhesive bond layer disposed between the composite foil and the metallic leading edge assembly.2. The airfoil assembly of claim 1 , wherein said high density base is formed of a uniform thickness.3. The airfoil assembly of claim 1 , wherein said high density base is formed of a varying thickness.4. The airfoil assembly of claim 1 , said base being welded to said nose.5. The airfoil assembly of claim 1 , said base being bonded to said nose.6. The airfoil of claim 1 , said base having first and second legs which are longer than side walls of said nose.7. The airfoil of claim 1 , wherein said metal leading edge assembly is formed of a single construction in a radial direction.8. The airfoil of claim 1 , wherein said metal leading edge assembly is formed of multiple segments in a radial direction.9. The airfoil of claim 1 , wherein said nose is bonded to said ...

Coating systems and methods therefor

A coating system and a method of forming the coating system capable of enabling components to survive in high temperatures environments, such as the hostile thermal environment of a gas turbine. The coating system is formed of a ceramic powder having powder particles each having an inner core formed of a first material and an outer region formed of a second material on the surface of the inner core. The inner core has a lower thermal conductivity than the outer region and the outer region has improved erosion resistance relative to the inner core.

TURBINE NOZZLE COMPONENTS HAVING REDUCED FLOW AREAS

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

CANTILEVER STATOR WITH VORTEX INITIATION FEATURE

An axial flow compressor is disclosed with a plurality of rotors. Each rotor includes a disk having an outer rim. Each outer rim is coupled to a radially outwardly extending rotor blade. The case is coupled to a plurality of radially inwardly extending stator vanes, i.e. cantilever-type stator vanes. Each stator vane is disposed between two rotor blades and extends towards one of the outer rims and terminates at a tip disposed in close proximity to one of the outer rims. At least one of the outer rims includes a serrated outer surface that faces the tip of a stator vane which results in a vortex flow causing air that would normally leak through the clearance between the stator vane and the outer rim to engage the stator vane to a greater degree thereby increasing the efficiency of the compressor. 1. An axial flow compressor , comprising:a plurality of rotors coaxially disposed within a case and coupled together, each rotor including a disk having a radially outward end that includes an outer rim, each outer rim is coupled to a radially outwardly extending rotor blade, each radially outwardly extending rotor blade terminating at a tip;the case being coupled to a plurality of radially inwardly extending stator vanes, each stator vane extending towards one of the outer rims and terminating at a tip disposed in close proximity to one of the outer rims;at least one outer rim including a serrated outer surface that faces the tip of one of the stator vanes.2. The compressor of wherein the serrated outer surface includes a plurality of adjacent grooves.3. The compressor of wherein the serrated outer surface includes a plurality of coaxial grooves separated by lands with each land disposed between two grooves claim 1 , the grooves being defined by a bottom surface disposed between two side walls claim 1 , the sidewalls being slanted in a fore direction.4. The compressor of wherein the lands are flat.5. The compressor of wherein that stator vane that faces the serrated outer ...

SHAPED RIM CAVITY WING SURFACE

A shaped rim cavity wing includes an upper surface and a lower surface. The lower surface has a geometric shape to control the separation of airflow as it passes around the lower surface to the top surface. A point of maximum extent defines the boundary between the upper surface and the lower surface, wherein the point of maximum extent defines a corner that that separates airflow from the shaped rim cavity rim and creates a flow re-circulation adjacent to the top surface of the shaped rim cavity wing. 1. A shaped rim cavity wing comprising:a body configured to extend from one of a rotating component and a stationary component of a turbomachine to inhibit airflow through a gas path between the rotating component and the stationary component;an upper surface of the body; a first concave portion;', 'a convex portion adjacent the first concave portion;', 'a first inflection point between the convex portion and the first concave portion;', 'a second concave portion adjacent the convex portion; and', 'a second inflection point between the second concave portion and the convex portion; and, 'a lower surface of the body, the lower surface having a geometric shape to control the separation of airflow as it passes around the lower surface, the geometric shape includinga point of maximum extent that defines a boundary between the upper surface and the lower surface, wherein the point of maximum extent defines a corner that separates airflow from the shaped rim cavity wing and creates flow re-circulation adjacent to the upper surface of the shaped rim cavity wing;wherein the first concave portion is located adjacent the point of maximum extent.2. The shaped rim cavity wing of claim 17 , further including:a flat portion located between the point of maximum extent and the lower surface.3. The shaped rim cavity wing of claim 18 , wherein the flat portion is vertical.4. The shaped rim cavity wing of claim 17 , wherein the body extends from the rotating component of the ...

WEAR PROTECTION ARRANGEMENT FOR A TURBOMACHINE, PROCESS AND COMPRESSOR

The present invention relates to a wear protection arrangement for a turbomachine, comprising at least one adjustable guide vane, a casing in which the guide vane is arranged in an adjustable manner, an inner ring, made from a metallic material, in or on which the guide vane is arranged in an adjustable manner, a first gap between an inner guide vane tab and the inner ring and a second gap between an outer guide vane tab and the casing, at least one wear protection coating, wherein the wear protection coating(s) is/are connected to the inner ring and/or to the inner guide vane tab and the wear protection coating(s) forms or form the first gap, at least in certain regions, and/or the wear protection coating(s) is/are connected to the casing and/or to the outer guide vane tab and the wear protection coating(s) forms/form the second gap, at least in certain regions. The invention further relates to a method for applying an abradable wear protection coating and for applying an abrasive wear protection coating and to a compressor for a turbomachine having a wear protection arrangement. 114.-. (canceled)15. A wear protection arrangement for a turbomachine , wherein the arrangement comprisesat least one adjustable guide vane,a casing in which the at least one guide vane is arranged in an adjustable manner,an inner ring made from a metallic material, in or on which ring the at least one guide vane is arranged in an adjustable manner,a first gap between an inner guide vane tab and the inner ring and a second gap between an outer guide vane tab and the casing,one or more wear protection coatings connected to the inner ring and/or to the inner guide vane tab and forming the first gap, at least in certain regions,and/or one or more wear protection coatings connected to the casing and/or to the outer guide vane tab and forming the second gap, at least in certain regions.16. The wear protection arrangement of claim 15 , wherein the one or more wear protection coatings are ...

Intercooled Compressor for a Gas Turbine Engine

A multi-stage intercooled compressor for a gas turbine engine, including multiple stages of rotating blades and cooling stator vanes, a cooling stator vane including an outer wall that defines an internal coolant fluid passage and has a length along a centerline from a leading edge to a trailing edge of the outer wall, and an internal flow divider wall disposed within the internal passage and extending along the centerline to divide the internal coolant fluid passage into an inflow pathway and an outflow pathway. 1. A multi-stage intercooled compressor for a gas turbine engine , including multiple stages of rotating blades and cooling stator vanes , the cooling stator vane including an outer wall that defines an internal coolant fluid passage and has a length along a centerline from a leading edge to a trailing edge of the outer wall , and an internal flow divider wall disposed within the internal passage and extending along the centerline to divide the internal coolant fluid passage into an inflow pathway and an outflow pathway.2. The multi-stage intercooled compressor of claim 1 , wherein the outer surface of the stator vane is substantially free of an extending cooling fin.3. The multi-stage intercooled compressor of claim 1 , wherein the surface area of an interior surface of the outer wall claim 1 , exposed to cooling fluid claim 1 , is at least about 90% of a surface area of an outer surface of the outer wall claim 1 , exposed to compression air.4. The multi-stage intercooled compressor of claim 1 , wherein a leading edge of the internal flow divider wall is connected to the leading edge of the outer wall claim 1 , and a trailing edge of the internal flow divider wall is connected to the trailing edge of the outer wall.5. The multi-stage intercooled compressor of claim 1 , wherein an outer surface of the internal flow divider is completely separated from an interior surface of the outer wall.6. The multi-stage intercooled compressor of claim 5 , wherein an ...

Low energy wake stage

The leading edge, the trailing edge, or both may be axially offset for a portion of the airfoils in a disk. By offsetting the airfoils, the downstream wake energy to the next stage of airfoils may be decreased. By staggering airfoils which are offset with airfoils that are not offset, the wake shapes from the airfoils may be out of phase and will not excite the downstream airfoils as much as conventional systems. This may decrease vibration and associated vibratory stresses in the system.

COOLING HOLE WITH SHAPED METER

A gas turbine engine component having a cooling passage includes a first wall defining an inlet of the cooling passage, a second wall generally opposite the first wall and defining an outlet of the cooling passage, a metering section extending downstream from the inlet, and a diffusing section extending from the metering section to the outlet. The metering section includes an upstream side and a downstream side generally opposite the upstream side. At least one of the upstream and downstream sides includes a first passage wall and a second passage wall where the first and second passage walls intersect to form a V-shape. 1. A gas turbine engine component having a cooling passage , the component comprising:a first wall defining an inlet of the cooling passage;a second wall generally opposite the first wall and defining an outlet of the cooling passage; an upstream side; and', 'a downstream side generally opposite the upstream side, wherein at least one of the upstream and downstream sides comprises a first passage wall and a second passage wall, and wherein the first and second passage walls intersect to form a V-shape; and, 'a metering section extending downstream from the inlet, the metering section comprisinga diffusing section extending from the metering section to the outlet.2. The component of claim 1 , wherein the first passage wall and the second passage wall are generally straight.3. The component of claim 1 , wherein the first passage wall and the second passage wall intersect to form an angle that is greater than 90 degrees.4. The component of claim 1 , wherein the first passage wall and the second passage wall are located on the upstream side.5. The component of claim 1 , wherein the first passage wall and the second passage wall are located on the downstream side.6. The component of claim 4 , wherein the downstream side comprises a third passage wall and a fourth passage wall claim 4 , and wherein the third and fourth passage walls intersect to form a ...

RING STATOR

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

Spall break for turbine component coatings

A turbine engine component can include a surface comprising at least one edge and a coating disposed upon the surface that can extend to the edge. A spall break can be disposed along a line upon the surface adjacent the edge to prevent spallation of the coating from spreading from the edge onto the surface beyond the spall break. The spall break can comprise a discontinuity of the coating. A method of coating a turbine component can include preparing a substrate to receive a coating and selecting a fail location along the substrate for a coating. One or more coating can be applied to the substrate and a spall break can be incorporated into the one or more coatings. The spall break can comprise a line of discontinuity in the one or more coatings along the fail location.

Segmented Stator Assembly

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

Attachment Faces for Clamped Turbine Stator of a Gas Turbine Engine

An airfoil fairing shell for a gas turbine engine includes an airfoil section between an outer vane endwall and an inner vane endwall, at least one of the outer vane endwall and the inner vane endwall including a radial attachment face, a suction side tangential attachment face, a pressure side tangential attachment face, and an axial attachment face. 1. An airfoil fairing shell for a gas turbine engine comprising:an airfoil section between an outer vane endwall and an inner vane endwall, at least one of said outer vane endwall and said inner vane endwall including a radial attachment face, a suction side tangential attachment face, a pressure side tangential attachment face, and an axial attachment face, said suction side tangential attachment face transverse to a resultant aerodynamic load generated by said airfoil.2. The airfoil fairing shell as recited in claim 1 , wherein said radial attachment face claim 1 , said suction side tangential attachment face claim 1 , said pressure side tangential attachment face claim 1 , and said axial attachment face are formed by a thickened region of at least one of said outer vane endwall and said inner vane endwall.3. The airfoil fairing shell as recited in claim 1 , wherein said radial attachment face claim 1 , said suction side tangential attachment face claim 1 , said pressure side tangential attachment face claim 1 , and said axial attachment face are formed by a thickened region of said inner vane endwall.4. The airfoil fairing shell as recited in claim 1 , wherein said suction side tangential attachment face is parallel to said pressure side tangential attachment face.5. The airfoil fairing shell as recited in claim 1 , wherein said suction side tangential attachment face and said pressure side tangential attachment face are non-parallel to said inner vane endwall.6. The airfoil fairing shell as recited in claim 1 , wherein said suction side tangential attachment face and said pressure side tangential attachment face ...

COOLING SYSTEM FOR GAS TURBINE, GAS TURBINE EQUIPMENT PROVIDED WITH SAME, AND PARTS COOLING METHOD FOR GAS TURBINE

A cooling system includes: a high pressure bleed line configured to bleed high pressure compressed air from a first bleed position of a compressor and to send the air to a first hot part; a low pressure bleed line configured to bleed low pressure compressed air from a second bleed position of the compressor and to send the air to a second hot part; an orifice provided in the low pressure bleed line; a connecting line configured to connect the high pressure bleed line and the low pressure bleed line; a first valve provided in the connecting line; a bypass line configured to connect the connecting line and the low pressure bleed line; and a second valve provided in the bypass line. 120-. (canceled)21. A cooling system for a gas turbine which includes a compressor configured to compress air , a combustor configured to burn a fuel in the air compressed by the compressor to generate a combustion gas , and a turbine driven using the combustion gas , the cooling system for a gas turbine comprising:a high pressure bleed line configured to bleed air from a first bleed position of the compressor and to send the air bled from the first bleed position to a first hot part coming into contact with the combustion gas among parts constituting the gas turbine;a cooler configured to cool air passing through the high pressure bleed line;a low pressure bleed line configured to bleed air at a pressure lower than that of the air which is bled from the first bleed position from a second bleed position of the compressor, to send the air bled from the second bleed position to a second hot part coming into contact with the combustion gas and disposed under a lower pressure environment than the first hot part among the parts constituting the gas turbine, and is not provided with a cooler;a minimum flow rate securing device configured to secure a minimum flow rate of air flowing through the low pressure bleed line while limiting a flow rate of the air flowing through the low pressure bleed ...