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

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

ОХЛАЖДАЮЩЕЕ УСТРОЙСТВО ПРОФИЛЬНОЙ ЧАСТИ ЛОПАТКИ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

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

ЛОПАТКА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ С УЛУЧШЕННЫМ ОХЛАЖДЕНИЕМ

Настоящее изобретение относится к лопатке турбины, содержащей хвостовик, перо (12), которое содержит входную кромку, выходную кромку (17), спинку и корытце, а также содержит охлаждающие выходные отверстия (26, 27) на выходной кромке (17), причем упомянутое перо также содержит первый (T1) и второй (T2) змеевидные контуры, причем каждый змеевидный контур (T1, T2) содержит несколько каналов (CA1, CM1, CT1, CA2, CM2, CV2, CT2), продолжающихся в направлении (EV) размаха и связанных друг с другом изогнутыми участками, при этом каждый змеевидный контур (T1, T2) снабжается воздухом посредством своего канала (CA1, CA2), расположенного ближе всего к входной кромке (16), причем выходные отверстия (26, 27) снабжаются воздухом из первого и второго змеевидных контуров (T1, T2). Достигается улучшение охлаждения. 3 н. и 8 з.п. ф-лы, 3 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, ...

ОГНЕУПОРНЫЙ СЕРДЕЧНИК, СОДЕРЖАЩИЙ ОСНОВНОЙ КОРПУС И КОЖУХ

Изобретение относится к литейному производству и может быть использовано для изготовления полой лопатки турбинного двигателя. Огнеупорный сердечник (12, 112) для изготовления лопатки (10) c использованием технологии литья по выплавляемой модели содержит основной корпус (14, 114) и по меньшей мере один кожух (16, 116), соединенный с основным корпусом (14, 114) и ограничивающий полость (18, 118) между основным корпусом и кожухом, причем кожух (16, 116) скомпонован таким образом, чтобы входить в контакт с лопаткой (10) в процессе изготовления, а полость (18, 118) выполнена закрытой для того, чтобы материал отливки не проникал в полость во время отливки лопатки (10). Изобретение позволяет повысить прочность лопатки. 2 н. и 9 з.п. ф-лы, 4 ил.

ТУРБИННЫЙ УЗЕЛ, СООТВЕТСТВУЮЩАЯ ТРУБКА СОУДАРИТЕЛЬНОГО ОХЛАЖДЕНИЯ И ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ

Турбинный узел содержит полую аэродинамическую часть, имеющую по меньшей мере одну полость с по меньшей мере одной трубкой соударительного охлаждения, предназначенную для введения внутрь полости полой аэродинамической части и используемую для соударительного охлаждения, по меньшей мере, внутренней поверхности полости, и по меньшей мере одну платформу, расположенную на радиальном конце полой аэродинамической части, и по меньшей мере одну охлаждающую камеру, используемую для охлаждения по меньшей мере одной платформы, и которая расположена на противоположной полой аэродинамической части стороне платформы. Охлаждающая камера ограничена на первом радиальном конце платформой, а на противоположном радиальном втором конце с помощью по меньшей мере одной закрывающей пластины. Трубка соударительного охлаждения выполнена из переднего элемента и заднего элемента, вставленных оба в по меньшей мере одну полость. Передний элемент расположен в направлении передней кромки полой аэродинамической части.

ЛОПАТКА ГАЗОТУРБИННОЙ УСТАНОВКИ

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

ЛОПАТКА ТУРБИНЫ

Лопатка турбины содержит канал охлаждения, сформированный в лопатке и проходящий в направлении ее высоты, и множество отверстий охлаждения. Спинка пера и корыто пера лопатки покрыты теплозащитным покрытием. Расчетная точка на спинке пера задана на поверхности лопатки на спинке пера в каждом сечении лопатки, перпендикулярном направлению высоты лопатки, в пределах диапазона от положения позади положения критического сечения, включая это положение, в котором расстояние между лопатками турбины является минимальным, до положения спереди от положения выходного конца последнего канала охлаждения, за исключением этого положения, которое является ближайшим к выходной кромке лопатки. Распределение толщины теплозащитного покрытия на спинке пера каждого сечения лопатки выполнено так, что толщина теплозащитного покрытия неизменна от входной кромки лопатки до расчетной точки и постепенно уменьшается от расчетной точки назад до выходной кромки лопатки. Множество отверстий охлаждения для образования пленки ...

Система (варианты) и способ охлаждения турбинных лопаток

Изобретение относится к энергетике. Система содержит турбинную лопатку, имеющую по меньшей мере один охлаждающий паз, предназначенный для транспортировки хладагента в направлении потока от внутренней части турбинной лопатки наружу. Охлаждающий паз имеет входное отверстие, соединенное с внутренней поверхностью, и сходящуюся секцию, расположенную ниже по потоку от входного отверстия. Сходящаяся секция имеет первую площадь поперечного сечения, которая уменьшается в направлении потока. Охлаждающий паз также имеет выходное отверстие, расположенное вдоль задней кромки турбинной лопатки. Также представлены вращающаяся лопатка турбины и способ изготовления лопатки. Изобретение позволяет обеспечить лучшую передачу тепла к задней кромке лопатки. 3 н. и 17 з.п. ф-лы, 6 ил.

ЛОПАТКА ТУРБИНЫ С ОПТИМИЗИРОВАННЫМ ОХЛАЖДЕНИЕМ

Лопатка турбины турбинного двигателя, такого как турбовинтовой или турбореактивный двигатель включает в себя хвостовик, перо, поддерживаемое хвостовиком, проходящее по направлению размаха, заканчиваясь концом, и содержит переднюю кромку и заднюю кромку, расположенную ниже по потоку от передней кромки, стенку стороны нагнетания и стенку стороны всасывания, расположенные на расстоянии друг от друга и соединяющие, каждая, переднюю кромку с задней кромкой. Перо содержит, по меньшей мере, один канал, выполненный с возможностью собирать охлаждающий воздух в хвостовике лопатки и осуществлять его циркуляцию в пере для его охлаждения; отверстия и/или щели, выполненные в его стенках, чтобы выпускать охлаждающий воздух из пера; верхнюю внутреннюю полость, расположенную на конце пера, чтобы охлаждать этот конец пера. При этом только один из указанных каналов подает охлаждающий воздух, собранный в хвостовике, в верхнюю полость. Указанный канал представляет собой канал непосредственной подачи, предназначенный ...

ТУРБИННАЯ ЛОПАТКА

Турбинная лопатка содержит профилированное обтекаемое рабочим газом перо, которое имеет переднюю кромку для набегания рабочего газа, а также проходящую вдоль главной оси пера лопатки от зоны хвостовика пера лопатки к противоположной зоне головки заднюю кромку для сбегания рабочего газа, и первую систему каналов и вторую систему каналов для раздельного направления двух различных подаваемых раздельно в турбинную лопатку сред. Первая канальная система заканчивается, по меньшей мере, в одном первом расположенном в зоне задней кромки выходном отверстии для выдувания первой среды в рабочий газ. По меньшей мере, одно расположенное в зоне задней кромки второе выходное отверстие для выдувания второй среды соединено со второй канальной системой. При рассматривании вдоль главной оси пера лопатки второе выходное отверстие расположено, по меньшей мере, частично на той же высоте, что и первое выходное отверстие Изобретение направлено на создание турбинной лопатки для газовой турбины, в которой надежно ...

ЛОПАТКА ГАЗОВОЙ ТУРБИНЫ С УСОВЕРШЕНСТВОВАННЫМИ КОНТУРАМИ ОХЛАЖДЕНИЯ

Лопатка газовой турбины для авиационного двигателя снабжена, по меньшей мере, тремя охлаждающими контурами. Первый охлаждающий контур содержит, по меньшей мере, одну полость на вогнутой стороне пера лопатки, вытянутую в радиальном направлении вблизи вогнутой поверхности пера. Второй охлаждающий контур независим от первого охлаждающего контура и содержит, по меньшей мере, одну полость на выпуклой стороне пера, вытянутую в радиальном направлении вблизи выпуклой поверхности пера. Третий охлаждающий контур независим от первого и второго охлаждающих контуров и содержит, по меньшей мере, одну центральную полость, расположенную в центральной части пера лопатки между полостью на вогнутой стороне пера и полостью на выпуклой стороне пера, по меньшей мере, одну полость вблизи входной кромки пера, соединительные отверстия, связывающие центральную полость и полость вблизи входной кромки пера, а также выпускные отверстия, сообщающиеся с указанной полостью вблизи входной кромки пера и выходящие на входную ...

КОМПОНЕНТ ГАЗОВОЙ ТУРБИНЫ, ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ, СПОСОБ ИЗГОТОВЛЕНИЯ КОМПОНЕНТА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Компонент газовой турбины, имеющий теплоизолирующую внешнюю поверхность для воздействия газообразных продуктов сгорания, содержит металлическую подложку, крепящий слой на поверхности подложки, теплозащитное покрытие, структуру выступающих элементов и структуру элементов в виде канавок. Теплозащитное покрытие включает слой внешнего теплозащитного покрытия, имеющий внутреннюю поверхность, нанесенную поверх и сцепленную с крепящим слоем, и внешнюю поверхность для воздействия газообразных продуктов сгорания. Структура в виде выступающих элементов имеет высоту выступов, составляющую от 2 до 75% совокупной общей толщины слоев теплозащитного покрытия. Структура элементов в виде канавок имеет канавки, сформированные в ранее нанесенном слое внешнего теплозащитного покрытия, по его внешней поверхности, и проникающие в этот слой. Структура выступающих элементов и структура элементов в виде канавок находятся в соответственно ограниченных, отделенных, трехмерных, независимо выровненных структурах, проходящих ...

ЭКСЦЕНТРИЧЕСКАЯ ФАСКА У ВХОДА ОТВЕТВЛЕНИЙ В ПРОТОЧНОМ КАНАЛЕ

Проточный канал содержит основной канал, ответвляющийся канал, в котором направление потока перпендикулярно направлению потока основного канала, и входное отверстие ответвляющегося канала, которое расположено в стенке основного канала и задано кромкой, содержащей верхнюю по потоку кромку и нижнюю по потоку кромку. У верхней по потоку кромки входного отверстия предусмотрена фаска. Нижняя по потоку кромка входного отверстия является острой кромкой, образованной прямым углом между стенкой ответвляющегося канала и стенкой основного канала. Фаска у верхней по потоку кромки имеет форму, которая является эксцентрической относительно продольной оси ответвляющегося канала. Изобретение направлено на уменьшение потерь давления. 2 н. и 5 з.п. ф-лы, 8 ил.

ТУРБИННЫЕ ЛОПАТКИ И ГАЗОТУРБИННАЯ УСТАНОВКА С ТАКИМИ ТУРБИННЫМИ ЛОПАТКАМИ

Турбинная лопатка содержит первую и вторую стеночные поверхности, соединительный канал и выступ. Первая стеночная поверхность обращена к охлаждающему каналу, по которому течет охлаждающий воздух. Вторая стеночная поверхность обращена к каналу рабочей текучей среды, по которому течет рабочая текучая среда. Соединительный канал обеспечивает сообщение между охлаждающим каналом и каналом рабочей текучей среды. Выступ предусмотрен на нижней по потоку стороне направления течения охлаждающего воздуха в отверстии соединительного канала. Отверстие образовано в первой стеночной поверхности. Выступ выступает от первой стеночной поверхности к охлаждающему каналу. Изобретение направлено на повышение эффективности охлаждения. 2 н. и 3 з.п. ф-лы, 11 ил.

СПОСОБ ИЗГОТОВЛЕНИЯ НАПРАВЛЯЮЩЕЙ ЛОПАТКИ, А ТАКЖЕ НАПРАВЛЯЮЩАЯ ЛОПАТКА

... 1. Способ изготовления турбинной лопатки (130), имеющей перо (149) лопатки, хвостовик (145) лопатки и область с ограниченной доступностью инструмента для выполнения отверстий для охлаждающего воздуха, причем эта область в переходе между хвостовиком (145) лопатки и пером (149) лопатки имеет вогнутую кромку (153), включающий в себя шаги:a) изготовление пера (149) лопатки и хвостовика (145) лопатки в виде отдельных конструктивных элементов,b) выполнение по меньшей мере одного отверстия (151) для охлаждающего воздуха в пере (149) лопатки и/или в хвостовике (145) лопатки в указанной области, иc) соединение пера (149) лопатки и хвостовика (145) лопатки после шага b),при этом ось (155) отверстия (151) для охлаждающего воздуха на наружной стороне пера (149) лопатки направлена на хвостовик (145) лопатки, или соответственно наоборот.2. Способ по п. 1, при котором изготовление отдельных конструктивных элементов (145, 149) в соответствии с шагом а) осуществляется путем литья.3. Способ по п. 1 или 2 ...

ЛОПАТКА ТУРБИНЫ

Лопатка турбины содержит канал охлаждения, сформированный в лопатке и проходящий в направлении ее высоты, и множество отверстий охлаждения. Спинка пера и корыто пера лопатки покрыты теплозащитным покрытием. Расчетная точка на спинке пера задана на поверхности лопатки на спинке пера в каждом сечении лопатки, перпендикулярном направлению высоты лопатки, в пределах диапазона от положения позади положения критического сечения, включая это положение, в котором расстояние между лопатками турбины является минимальным, до положения спереди от положения выходного конца последнего канала охлаждения, за исключением этого положения, которое является ближайшим к выходной кромке лопатки. Распределение толщины теплозащитного покрытия на спинке пера каждого сечения лопатки выполнено так, что толщина теплозащитного покрытия неизменна от входной кромки лопатки до расчетной точки и постепенно уменьшается от расчетной точки назад до выходной кромки лопатки. Множество отверстий охлаждения для образования пленки ...

ЭЛЕМЕНТ ГАЗОВОЙ ТУРБИНЫ С ПЛЕНОЧНЫМ ОХЛАЖДЕНИЕМ

Изобретение относится к элементу газовой турбины с пленочным охлаждением, имеющему подвергаемую воздействию горячего газа поверхность, в которой выполнены отверстия для пленочного охлаждения. Каждое из отверстий для пленочного охлаждения имеет в направлении своего потока канальную часть и непосредственно примыкающую к ней диффузорную часть, включающую в себя расположенную выше по потоку кромку диффузорной части, две продольные кромки и расположенную ниже по потоку кромку диффузорной части. Каждая продольная кромка сходится с расположенной ниже по потоку кромкой диффузорной части в угловой области. Чтобы создать эффективное расположение отверстий для пленочного охлаждения, охлаждающая пленка которых образуется ближе, чем прежде, за расположенной ниже по потоку кромкой диффузорной части, оси соответствующих канальных частей должны быть наклонены к локальному направлению потока горячего газа, а их диффузорные части выполнены асимметричными таким образом, что непосредственно соседние угловые ...

ПЛЕНОЧНОЕ ОХЛАЖДЕНИЕ С ПЛАЗМЕННЫМ ЭКРАНИРОВАНИЕМ

... 1. Система (11) плазменного экранирования пограничного слоя, содержащая: ! отверстия (49) для пленочного охлаждения, проходящие через стенку (26), причем отверстия (49) для пленочного охлаждения расположены под углом в направлении (D), т.е. вперед по ходу, и продолжаются от холодной поверхности (59) стенки (26) до горячей поверхности (54) стенки (26), и ! генератор (2) плазмы, расположенный впереди по ходу относительно отверстий (49) для пленочного охлаждения и производящий плазму (90), распространяющуюся поверх отверстий (49) для пленочного охлаждения. ! 2. Система (11) по п.1, дополнительно содержащая генератор (2) плазмы, установленный на стенке (26). ! 3. Система (11) по п.2, дополнительно содержащая генератор (2) плазмы, включающий в себя внутренний и внешний электроды (3, 4), разделенные диэлектриком (5). ! 4. Система (11) по п.3, дополнительно содержащая источник (100) переменного тока, подсоединенный к электродам для подачи на электроды высокого напряжения переменного тока. ! 5.

ОГНЕУПОРНЫЙ СЕРДЕЧНИК, СОДЕРЖАЩИЙ ОСНОВНОЙ КОРПУС И КОЖУХ

ЛОПАТКА ТУРБИНЫ

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

РАБОЧАЯ ЛОПАТКА ТУРБИНЫ

... 1. Рабочая лопатка турбины для газотурбинного двигателя, содержащая:аэродинамическую часть, имеющую концевую часть на наружной радиальной кромке,причем аэродинамическая часть лопатки содержит сторону повышенного давления и сторону пониженного давления, которые соединены вместе на передней кромке и на задней кромке аэродинамической части лопатки, и сторона повышенного давления и сторона пониженного давления проходят от хвостовика лопатки до концевой части,при этом концевая часть содержит пластину и выступающую кромку, расположенную вдоль периферии пластины концевой части, причем выступающая кромка имеет микроканал, соединенный с источником охлаждающей текучей среды.2. Рабочая лопатка по п.1, в которой сторона повышенного давления содержит наружную радиальную кромку, а сторона пониженного давления содержит наружную радиальную кромку, при этом аэродинамическая часть лопатки выполнена таким образом, что пластина концевой части проходит в осевом и в окружном направлениях для соединения наружной ...

Ауксетическая структура с искаженными проецированием прорезями в конфигурациях, разработанных для обеспечения поведения с отрицательным коэффициентом Пуассона и улучшенной характеристики

ОХЛАЖДАЕМАЯ ТУРБИНА

... 1. Охлаждаемая турбина, содержащая рабочее колесо с установленными на нем рабочими лопатками с двумя контурами охлаждения, последовательно соединенные с воздушными каналами в рабочем колесе, с независимыми кольцевыми диффузорными каналами, образованными на поверхности рабочего колеса, соединенными с сопловыми аппаратами закрутки и транзитными воздуховодами на их входе, сопловые лопатки, каждая из которых выполнена в виде конструктивного элемента, ограниченного верхней и нижней полками, и пространства между ними, ограниченного вогнутой и выпуклой стенками пера сопловой лопатки, в виде расположенных вдоль ее оси раздаточного коллектора входной кромки и раздаточной полости, причем раздаточный коллектор входной кромки соединен на входе с воздушной полостью камеры сгорания, а на выходе через перфорационные отверстия во входной кромке сопловой лопатки с проточной частью турбины, теплообменник, соединенный на входе с воздушной полостью камеры сгорания, а на выходе последовательно сообщенный с ...

СПОСОБ УМЕНЬШЕНИЯ ДИАМЕТРА ОТВЕРСТИЯ

... 1. Способ уменьшения диаметра отверстия (2), отличающийся тем, что содержит этап, на котором сплющивают периметр (3) отверстия (2) при помощи инструмента, контактный конец которого имеет сферическую или по существу сферическую форму или форму усеченного конуса.2. Способ по п.1, в котором упомянутое сплющивание производят при помощи инструмента (6), центрованного по отверстию (2).3. Способ по п.2, в котором контактный конец (7) упомянутого инструмента (6) содержит шарик (4).4. Способ по п.1, в котором упомянутый периметр (3) отверстия (2) является металлическим.5. Способ по любому из предыдущих пунктов, в котором упомянутый периметр (3) отверстия (2) выполнен из жаропрочного материала.6. Способ по п.1, в котором диаметр отверстия (2) составляет 0,5-3 мм.7. Способ коррекции проницаемости детали (9, 10), содержащей множество отверстий (2) для прохождения газообразной текучей среды, отличающийся тем, что содержит следующие этапы:- идентифицируют, по меньшей мере, одно отверстие (2), диаметр ...

ТУРБИННАЯ ЛОПАТКА С ОБЕСПЫЛИВАЮЩИМ ОТВЕРСТИЕМ В ОСНОВАНИИ ЛОПАСТИ

... 1. Охлаждаемая турбинная лопатка для турбомашины, содержащая лопасть (2), установленную на платформе (6), которая расположена на ножке (5), при этом упомянутая лопасть является полой с одной или несколькими полостями для циркуляции охлаждающего воздуха, причем полость (11), размещенная вдоль задней кромки (4), питается охлаждающим воздухом от питающего канала (10), выполненного в виде колена (13) внутри упомянутой ножки и связывающего воздушный вход (12), расположенный в нижней части ножки (5), с полостью (11) задней кромки, отличающаяся тем, что канал (10) содержит на оси, по существу, радиальной относительно впускного отверстия (12), нишу (14), расположенную под платформой (6) и имеющую форму колокола, при этом упомянутая ниша (6) содержит в вершине обеспыливающее отверстие (19), пересекающее упомянутую платформу, и ограничена внутри ножки (5) закрывающими ее по бокам стенками, расположенными, по существу, радиально от платформы (6).2. Турбинная лопатка по п.1, содержащая, кроме того, ...

Minimale Leckströmung zwischen Schaufelspitze und gegenüberliegender Gehäusewand

Air-cooled gas turbine blade

A turbine blade for a gas turbine engine has internal passages which are opened at their roots to supply pressure so that there is a constant flow of cooling air inside the blade in use. The blade has an airfoil surface and a number of internal passages inside the walls of the blade, some being formed adjacent the mid-chord section and extend from the root section to the tip section to define feed channels open at the blade root section. A number of radially spaced film cooling holes in the airfoil surface communicate with each feed channel to flow a film of cooling air adjacent the airfoil surface. A number of replenishment holes are radially spaced in the wall to flow air from the mid-chord section to the feed channel(s) to replenish the cooling air in the feed channel(s) that is otherwise lost in supplying air to the film cooling holes. A device communicates cooling air from the root section via the feed channels to discharge from orifices in the airfoil surface at the tip section.

Schaufelspitze mit Prallkühlung

Kühlung in Gasturbinen

Die erfindungsgemäße Vorrichtung und die erfindungsgemäßen Verfahren dienen der effizienten und zuverlässigen Kühlung von Bauteilen 210, insbesondere von Turbomaschinen, auch im Falle einer lokalen Erhöhung des statischen Drucks 234 eines heißen Fluids, das das Bauteil überströmt. Um eine ausreichende Kühlung der Bauteile 210 zu gewährleisten, ist erfindungsgemäß der Abstand der Kühlbohrungen 240 untereinander jeweils so gewählt, daß die Kühlbohrungen 240 in dem Bereich erhöhten statischen Druckes 234 des heißen Fluids einen kleineren Abstand zueinander aufweisen als in den Bereichen niedrigeren statischen Druckes. Eine typische Ausführung der Erfindung ist in Figur 3 dargestellt.

Vorrichtung zur Kühlung einer Wandung eines Bauteils einer Gasturbine

Die Erfindung bezieht sich auf eine Vorrichtung zur Kühlung einer Wandung 25 eines Bauteils einer Gasturbine, wobei eine Strömung parallel zur Wandung 25 strömt, mit zumindest einem in der Wandung 25 ausgebildeten Einströmkanal 29, welcher zur Zuführung von Kühlluft in eine Ausnehmung 30 der Wandung 25 mündet, dadurch gekennzeichnet, dass eine Mittelachse 31 des Einströmkanals 29 auf eine doppelt konvex ausgebildete Aufprallwandung 32 der Ausnehmung 30 ausgerichtet ist.

Gekühlte Schaufeln eines Gasturbinentriebwerks

Gekühlte Schaufel eines Gasturbinentriebwerks, umfassend: einen Kühlkanal, der innerhalb der gekühlten Schaufel ausgebildet ist und eine darin strömende Kühlluft aufweist; ausgebildete Filmkühllöcher, die von einer Innenwandfläche des Kühlkanals zu einer Außenwandfläche der Kühlschaufel durchdringen und einen Kühlfilm auf einer äußeren Fläche der Schaufel bilden, und ein darauf ausgebildetes Prallkühlelement, welches eine Vielzahl von die Kühlluft ausstoßenden kleinen Löchern aufweist, und gekennzeichnet ist durch: wobei das Prallkühlelement im Kühlkanal angeordnet ist und dabei einen vorbestimmten Zwischenraum entfernt von der Innenwandfläche lässt; wobei ein durch die Innenwandfläche und das Prallkühlelement ausgebildeter Zwischenraum einen darin angebrachten Dichtungsabschnitt aufweist, der den entsprechenden Zwischenraum in einer Schaufelsehnenrichtung teilt, und wobei der Dichtungsabschnitt zwischen in einer Schaufelsehnenrichtung benachbarten Filmkühllöchern angebracht ist.

Verfahren zur Herstellung eines thermisch belasteten Gussteils

Zur Herstellung eines thermisch hoch belasteten Gussteils (1, 14, 16, 17) einer thermisch Turbomaschine, welches mit einem bekannten Gussverfahren hergestellt wird, wird die Gussform aus einem Schlicker mit einem Wachsmodell und einem daran angehefteten oder in einen Hohlraum eingeführten Polymerschaum hergestellt. Auf diese Weise dringt während des Gussverfahrens die flüssige Superlegierung auch in die offenporige Struktur der Gussform ein, so dass eine integrale Kühlstruktur (7) während des Erstarrungsvorgangs des Gussteils (1, 14, 16, 17) entsteht. Vorteilhaft wird ein einkristallines oder gerichtet erstarrtes Gussteil (1, 14, 16, 17) hergestellt. Auch eine Variation der Porengröße des Polymerschaums, eine getrennte Herstellung von Kühlstruktur (7) und Grundwerkstoff und eine Beschichtung der Kühlstruktur (7) mit einer keramischen Schutzschicht (TBC) (11) ist denkbar.

Innengekühltes Blatt einer Turbinenrotorschaufel eines Gasturbinenmotors

Innengekühltes Blatt (16) einer Turbinenrotorschaufel (10) eines Gasturbinenmotors, mit - einer Kopffläche (32), einer voreilenden Kante (20), einer nacheilenden Kante (22) und einer Druckseite (18), - einer ringförmigen Abdeckung, die konzentrisch um die Blätter angeordnet ist und mit der Kopffläche (32) einen Spalt (48) begrenzt, - einem inneren Durchlass (52) zum Leiten von Kühlluft zu Auslässen (30), die benachbart zur Kopffläche (32) des Blattes (16) im Bereich der Druckseite (18) ausgebildet sind und einen Kühlluftstrahl (42) unter einem gegebenen Winkel (C) abgeben, dadurch gekennzeichnet, dass jeder der Auslässe (30) in dem Übergang zwischen der Druckseite (18) und der Kopffläche (32) in einer Ausnehmung (36) mündet, welch letztere sich von der Druckseite (18) in Richtung der Saugseite (24) bis zur saugseitigen Wand des Auslasses und, gesehen in der Richtung von der voreilenden Kante (20) zur nacheilenden Kante (22), im Wesentlichen über die gesamte Breite des betreffenden Auslasses ...

Verfahren zum Unschädlichmachen von beim Beschichten mit einer Schutzschicht entstehenden Verengungen in den Kühllöchern von gasgekühlten Teilen

Turbinenschaufel

Eine Turbinenschaufel (1) für eine Gasturbine weist ein sich von einer Plattform (3, 4) erstreckendes hohles Schaufelblatt (8) auf. Zwischen dem Schaufelblatt (2) und der Plattform (3, 4) befindet sich sowohl auf der Druckseite (8) als auch auf der Saugseite (9) des Schaufelblatts (2) ein Übergang (13). Der Übergang (13) enthält eine entlang einem Teil der Länge des Übergangs (13) verlaufende Kühlbohrung (17, 18) mit einem mit dem Inneren der Turbinenschaufel (1) in Verbindung stehenden ersten Ende (17a, 18a) zur Aufnahme eines gasförmigen Kühlmittels und einem mit dem Äußeren der Turbinenschaufel (1) in Verbindung stehenden zweiten Ende (17b, 18b). Die Kühlbohrung (17, 18) kann durch Funkerosion (EDM - electro-discharge machining) hergestellt sein.

Strömungsmaschinenschaufel mit aparter Kühlung der Anströmkante

BAUTEIL MIT FILMKÜHLLOCH

Improvements in and relating to composite blades for gas turbines

... 531,445. Gas turbines. AKT.-GES. BROWN, BOVERI, & CIE. July 22, 1939, No. 21366. Convention date, July 27, 1938. [Class 110 (iii)] An air channel leading from a passage 7 in the blade wheel, is formed between the leading edge of the blade proper 1 and a spaced cap blade 2, The air emerges from the slits 4 and spreads over the blade surface. The cap blade may be formed of specially hard metal. The blades are welded together from 8 to 9 in the foot portion and at the tip 10 and the channel is closed by a plate 11. The cap blade may have a foot flange 12 corresponding to the. flange 5 of the main blade. The cap blade may be formed as shown in Fig. 5, or, it may co-operate with a groove 17, Fig. 7, formed in the leading edge of the. main blade.

Cooled vane of a turbine guide nozzle vane

NOZZLE GUIDE VANE FOR GAS TURBINE ENGINE

Tip seal and anti-contamination for turbine blades

Gas turbine airfoil internal air system

Gas turbine aerofoil cooling with pressure attenuation chambers

A gas turbine engine aerofoil 24 has a plurality of attenuation chambers 34 positioned between a cooling air passageway 26 and its leading edge 28. Cooling air passing from the passageway 26 to the exterior surface of the leading edge is attenuated in pressure by impingement on the opposing walls of the respective chambers 34, prior to leaving the chambers 34 via exit passageways 38 which straddle the leading edge. A plurality of input passageways 36 are provided for flow from passageway 26 into the chambers 34, and each chamber may have numerically less input passageways 36 than exit passageways 38. In alternative embodiments, passageways (36, fig 5) connect to upper and/or lower ends of each chamber 36 for pressure attenuation resulting from expansion therein. Hot spots due to blockage and effects due to pressure fluctuation are reduced.

Thin-walled,lightweight cooled turbine blade

FILM COOLED VANES & TURBINES

Turbine blade

Modulating cooling airflows

Cooling airflow is modulated by means of interacting metering apertures formed in members 42, 46 that can move in relation to each other according to changes in operating temperature. The modulation method may be applied to a gas turbine engine airfoil, having a wall (42) provided with a cooling effusion hole 40 therein for film cooling the external surface 52 of the wall 42. A member 46 attached (at one edge or at a point between two metering apertures) to the internal surface 48 of the wall 42 is provided with an aperture 44 which at least partially overlaps the cooling effusion hole 40. The airfoil wall 42 and member 46 are formed from materials having different coefficients of thermal expansion whereby over a temperature range, the aperture 44 and cooling effusion hole 40 interact to a greater or lesser extent to modulate the flow of cooling air therethrough.

Vane assembly

A turbine engine wall having at least some cooling orifices that are plugged

A turbine engine wall (16, 18, Fig 1), (44, 46, Fig 2) has a cold side and a hot side. The wall has a plurality of cooling orifices 32, 72 which permit air flowing on the cold side of the wall to penetrate to the hot side in order to form a film of cooling air along the wall. The orifices are distributed in a plurality of axially spaced circumferential rows and the axes of each of the orifices are inclined at an angle relative to a wall. At least some of the orifices are plugged by a plugging material 80 so as to define a minimum level of porosity for the wall as the turbine engine is put into service. During the lifetime of the engine, the plugged orifices are progressively unplugged such that a maximum level of porosity is defined at the turbine engine end of life. The unplugging may be achieved by an increase in the temperature of the wall, and the plugging material may be a metal alloy based on brass or aluminium, and having a melting temperature range of approximately 650 Celsius to ...

A turbine blade support assembly and a turbine assembly

Film-cooled gas turbine engine component

The body of a component, such as a turbine blade, is provided with an angled film-cooling hole having an inlet 15, an outlet 17, and a partition 19. The partition may divide the outlet into at least two openings 17a, 17b, and extend to the inlet side of the cooling hole, and the outlet may be wedge-shaped (divergent).

A component having a film cooling arrangement

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 ...

An air-cooled component

A turbine ring assembly with cold setting

Blade for a turbomachine turbine, comprising internal passages for circulating cooling air

Improvements in cooled turbine blades

... 1,120,694. Turbine blades. SOC. NATIONALE D'ETUDE ET DE CONSTRUCTION DE MOTEURS D'AVIATION. July 12, 1966 [Aug. 2, 1965; June 23, 1966], No.31196/66. Heading F1T. A gas turbine blade 1 mounted on a rotor 3 has internal sealed passages 6, 7 containing a liquid coolant, and passages 2a, 2b adjacent the leading and trailing edges, respectively, supplied with air from the engine compressor by way of the rotor interior, the air exhausting from passages 2a, 2b through orifices 8a, 8b to flow over the exterior of the blade working surfaces. The liquid coolant may be water or a sodiumpotassium eutectic which is liquid at both ambient and turbine operating temperatures. Orifices 8a, 8b may be replaced by porous wall sections. Ribs 5 on the blade root 4 assist the cooling.

IMPROVEMENTS IN OR RELATING TO BLADES FOR FLUID FLOW MACHINES

... 1285369 Turbomachine blades ROLLS ROYCE Ltd 14 Dec 1970 [16 Dec 1969] 61365/69 Heading FIT A hollow turbomachine blade, e. g. for the turbine of a gas turbine engine, has a plate 34 recessed into the concave and/or convex blade surfaces, channels 36 being provided between the plate and blade wall and communicating with apertures 32, 38 in the wall and plate to provide a flow path for cooling air from the interior to the exterior of the blade. The channels 36 are formed on the inside surface of plate 34 or the outside surface of the blade wall and each has an X, L, T, swastika or other shape. The cooling air is supplied to the interior of the blade through apertures 42 and some exhausts through holes 40 in shroud 26 and some through apertures in the leading and trailing edges of the blade.

Turbine airfoil suction aided film cooling means

An aerofoil

SHOVEL FOR A FLUID-FLOW MACHINE

TURBINE BLADE WITH IMPROVED COOLING CHARACTERISTICS AND USEFUL LIFE

... ²²²²La présente invention concerne le domaine des aubes de turbine de ²turbomachine, notamment une aube de turbine (1), comportant une paroi ²intrados (2), une paroi extrados (3), au moins une première cavité radiale ²de bord de fuite (4), au moins une seconde cavité radiale (5) en amont de ²la cavité de bord de fuite (4), une paroi interne (6) séparant les cavités ²radiales (4 et 5) et comprenant au moins un canal (7) reliant les cavités (4 ²et 5) entre elles, ledit canal (7) étant orienté selon un axe (71) coupant la ²surface interne (42) de la paroi intrados (2).² ...

TURBINE VANE WITH DUSTING HOLE AT THE BASE OF THE BLADE

Aube de turbine refroidie pour turbomachine comprenant une pale (2) montée sur une plate-forme (6) portée par un pied (5), ladite pale étant creusée d'une ou plusieurs cavités pour la circulation d'air de refroidissement, la cavité (11) s'étendant le long du bord de fuite étant alimentée en air de refroidissement par un conduit d'alimentation (10) reliant une entrée d'air (12) située en partie basse du pied (5) à la cavité (11) de bord de fuite en faisant un coude (13) au sein dudit pied, caractérisée en ce que le conduit (10) comporte, sur un axe sensiblement radial par rapport à l'entrée d'air (12), une niche (14) située sous la plate- forme (6) ayant une forme en cloche, ladite niche débouchant à son sommet par un trou de dépoussiérage (19) traversant ladite plate-forme et étant délimitée à l'intérieur du pied (5) par des parois s'étendant sensiblement radialement à partir de la plate-forme (6) pour la refermer latéralement.

COMPOSITE GAS TURBINE ENGINE COMPONENT

A method for manufacturing a composite gas turbine engine component is provided. A composite structure is formed that is operable in a gas turbine engine, the composite structure is defined by a composite material and machining a geometric shape into a surface of the composite structure through at least part of the composite material using an ultrasonic machining process. The composite material is a ceramic matrix composite (CMC), a metal matrix composite (MMC), an organic matrix composite (OMC) and/or a carbon-carbon composite. The composite structure may also be a composite air foil.

AIRFOIL LEADING EDGE END WALL VORTEX REDUCING PLASMA

A leading edge vortex reducing system (11) includes a gas turbine engine airfoil (39) extending in a spanwise direction (S) away from an end wall (88), one or more plasma generators (2) extending in the spanwise direction (S) through a fillet (34) between the airfoil (39) and the end wall (88) in a leading edge region (89) near and around a leading edge (LE) of the airfoil (39) and near the fillet (34). The plasma generators (2) being operable for producing a plasma (90) extending over a portion of the fillet (34) in the leading edge region (89). The plasma generators (2) may be mounted on an outer wall (26) of the airfoil (39) with a first portion of the plasma generators (2) on a pressure side (46) of the airfoil (39) and a second portion of the plasma generators (2) on a suction side (48) of the airfoil (39). A method for operating the system (11) includes energizing one or more plasma generators (2) to form the plasma (90) in steady state or unsteady modes.

AIR COOLED TURBINE VANES

COOLING CIRCUITS FOR GAS TURBINE BLADE

Aube (1) de turbine à gaz d'un moteur d'avion, comportant dans sa partie centrale au moins un premier circuit de refroidissement central (A) comprenant au moins une première (2) et une deuxième (4) cavités s'étendant radialement du côté intrados (1a) de l'aube (1), au moins une cavité (6) s'étendant du côté extrados (1b) de l'aube, une ouverture d'admission d'air à une extrémité radiale de la première cavité intrados (2) pour alimenter le premier circuit de refroidissement central (A) en air de refroidissement, un premier passage faisant communiquer l'autre extrémité radiale de la première cavité intrados (2) à une extrémité radiale voisine de la cavité extrados (6), un second passage faisant communiquer l'autre extrémité radiale de la cavité extrados avec une extrémité radiale voisine de la deuxième cavité intrados (4), et des orifices de sortie s'ouvrant dans la deuxième cavité intrados et débouchant sur la face intrados (1a) de l'aube.

HIGH EFFICIENCY FAN COOLING HOLES FOR TURBINE AIRFOIL

A turbine airfoil (32) includes a leading edge (36) and an axially spaced-part trailing edge (38), the leading edge (36) having an axially-extending external surface curvature. A cooling circuit (50) in the airfoil includes cooling holes (60) formed in the leading edge (36) along the span axis of the airfoil. The cooling holes (60) have a diffuser section communicating with the leading edge (36) surface. The diffuser section has four opposed walls defining a generally quadralinear exit opening on the surface of the leading edge (36). One of the diffuser walls has a convex curvature that approximates the external surface curvature of the leading edge (36) whereby fluid flow from the cooling hole exits is evenly dispersed and spread along land portions of the leading edge (36) adjacent the cooling holes (60).

PROCESS FOR DRILLING HOLES IN A METALLIC WORKPIECE HAVING A THERMAL BARRIER COATING

A method is provided for drilling a hole through a metallic workpiece (1) having a thermal barrier coating (3) with a ceramic top coat (2) by laser drilling a counterbore to a depth which extends through the ceramic top coat but not substantially into the metallic workpiece and then laser drilling the hole through the workpiece aligned with the counterbore, the counterbore having a diameter larger than the hole.

REFRACTORY CORE COMPRISING A MAIN BODY AND A SHELL

Noyau réfractaire (12) pour la fabrication d'une aube creuse (10) de turbomachine selon la technique de la fonderie à la cire perdue, comprenant un corps principal (14) et au moins une coque (16) reliée au corps principal (14) et définissant une cavité (18) entre le corps principal et la coque, la coque (16) étant configurée pour venir au contact de l'aube (10) lors de la fabrication.

METHODS AND APPARATUS FOR COOLING GAS TURBINE ENGINE COMPONENTS

A gas turbine engine component (40) includes a substrate wall (50) including a first surface (52) and an opposite second surface (54), and a plurality of pores (56) extending through the wall. The component also includes a thermal barrier coating (TBC) (74) extending over the wall first surface, wherein the TBC substantially seals the pores at the first surface. The component also includes a plurality of film cooling holes (58) extending through the wall and the TBC, wherein the plurality of film cooling holes and the plurality of pores extend substantially perpendicularly through the wall and the TBC.

COOLING CHANNEL INSIDE A WALL

Paroi (1) dans laquelle est ménagé au moins un canal de refroidissement (6), cette paroi étant refroidie par l'air frais circulant dans le canal, ce canal (6) comprenant un perçage (7) et une partie de diffusion (9), le perçage (7) débouchant, d'un côté, sur la surface intérieure (3) de la paroi (1) et, de l'autre, dans la partie de diffusion (9) en formant un orifice (11), la partie de diffusion (9) s'évasant autour de cet orifice (11) et débouchant sur la surface extérieure (5) de la paroi (1), la partie de diffusion (9) comprenant un fond avant sensiblement plan, incliné, s'étendant à l'avant de l'orifice (11), et une bordure s'étendant à l'arrière, sur les côtés et à l'avant de l'orifice (11), cette bordure rejoignant les côtés du fond avant. Procédé et électrode (20) pour réaliser un tel canal de refroidissement. Aube de turbomachine présentant une telle paroi.

PROCESS FOR DRILLING HOLES IN A METALLIC WORKPIECE HAVING A THERMAL BARRIER COATING

A method is provided for drilling a hole through a metallic workpiece (1) having a thermal barrier coating (3) with a ceramic top coat (2) by laser drilling a counterbore to a depth which extends through the ceramic top coat but not substantially into the metallic workpiece and then laser drilling the hole through the workpiece aligned with the counterbore, the counterbore having a diameter larger than the hole.

TURBINE PART

A large number of film-cooling holes (13) are formed in that portion of a turbine part body (3) which is exposed to high- temperature gas. Each film- cooling hole (13) has a straight hole section (15) formed in a portion on the inner wall surface (5a) side of the turbine part body (3) and an enlarged hole section (17) formed in a portion on the outer wall surface (5b) side of the turbine part body (3). Further, each film-cooling hole (13) is formed such that the inclination of an ejection surface (17p) of the enlarged hole section (17) increases gradually from the center in the width direction (L3) of the hole toward the opposite ends of the hole, the inclination being defined with an imaginary plane (VP) as the standard, the imaginary plane being set in parallel to two directions that are the axis direction (L2) of the straight hole section (15) and the hole width direction (L3).

AIRFOIL LEADING EDGE IMPINGEMENT COOLING

An airfoil including a spar and a cover sheet. Standoffs, a leading edge wall, and a separator wall extend away from an outer surface of the spar. The standoffs are arranged to define leading grooves disposed at the pressure side of the leading edge. The cover sheet is coupled to the leading edge wall and the standoffs over the leading grooves to define cooling passageways. The cooling passageways are in communication with one or more inlet ports formed in the spar, which are in communication with a plenum disposed within the spar. The cover sheet is arranged to define outlet ports or a slot in communication with the cooling passageway. Cooling air is delivered from the cooling air plenum through the inlet port for impingement cooling at the cover sheet, and traverses downstream through the cooling passageway to the outlet ports or slot for film cooling of the leading edge.

COOLED AIRFOIL TRAILING EDGE AND METHOD OF COOLING THE AIRFOIL TRAILING EDGE

An airfoil and method of cooling a airfoil including a leading edge, a trailing edge, a suction side, a pressure side and at least one internal cooling channel configured to convey a cooling fluid, is provided. A plurality of trailing edge bleed slots are in fluid communication with the at least one internal cooling channel, wherein a downstream edge of the pressure side of the airfoil lies upstream of a downstream edge of the suction side to expose the plurality of trailing edge bleed slots proximate to the trailing edge of the airfoil. The at least one internal cooling channel is configured to supply the cooling fluid from a source of cooling fluid towards the plurality of trailing edge bleed slots. A plurality of obstruction features are disposed within the at least one internal cooling channel and at a downstream edge of the remaining pressure side. The one or more obstruction features are configured having a predefined substantially polygon shape, to distribute a flow of the cooling ...

COOLING PASSAGES IN A TURBINE COMPONENT

A turbine component has a plurality of cooling passages each extending through a body of a structure between opposite hot and cold surfaces of the structure. According to one embodiment, at least one of the cooling passages includes a plurality of upstream paths defining respective inlet openings on the cold surface and merging into a number of downstream paths defining respective outlet openings on the hot surface.

TURBINE BLADE WITH OPTIMISED COOLING AT THE TRAILING EDGE OF SAME COMPRISING UPSTREAM AND DOWNSTREAM DUCTS AND INNER SIDE CAVITIES

The invention concerns a turbine blade (91) comprising a root (P), a vane extending in a spanwise direction (EV), ending at a tip (S) and comprising a leading edge and a trailing edge and a pressure-side wall and a suction-side wall, said vane further comprising: at least one upstream duct (93) configured to collect air at the root (P) to cool the leading edge, discharging said air through holes passing through the wall of the leading edge; at least one downstream duct (96) separate from the upstream duct (93) and configured to collect air at the root (P) to cool the trailing edge, discharging said air through holes (97) passing through the pressure wall upstream from the trailing edge; an inner side cavity (101) running along the pressure-side wall to form a heat shield insulating the downstream duct (96).

GAS TURBINE ENGINE WITH FILLET FILM HOLES

An airfoil for a gas turbine engine can have an exterior wall and an interior wall, with each wall having a thickness. The walls can intersect to define a corner at the intersection. A cooling passage can be defined by the walls at or near the corner to provide fluid communication between the interior and exterior of the airfoil. A film hole can be disposed in the walls and can have a length and diameter to define a ratio of length to diameter, L/D. An arcuate fillet can be located in the corner to define an effective radius for the fillet. The effective radius can be at least 1.5 times larger than the thicknesses of the walls to provide for an increased length to diameter ratio for the film hole.

ADDITIVE MANUFACTURING METHOD FOR MAKING HOLES BOUNDED BY THIN WALLS IN TURBINE COMPONENTS

A method of forming a passage in a turbine component that includes using an additive manufacturing process to form a first support structure on a first surface of the turbine component; and forming a passage through the first support structure and the turbine component.

HOT STREAK ALIGNMENT FOR GAS TURBINE DURABILITY

Embodiments of hot streak alignment for gas turbine durability include structures and methods to align hot streaks with the leading edges of aligned first stage nozzle vanes in order to improve mixing of the hot streaks with cooling air at a stator nozzle of a first stage turbine and reduce usage of cooling air at first stage non-aligned stator nozzle vanes disposed adjacent to the aligned stator vanes.

TURBINE NOZZLE WITH IMPINGEMENT BAFFLE

A turbine nozzle apparatus includes: a vane (16) extending between inner and outer bands (12, 14), the interior of the vane (16) being open and communicating with an aperture (30) in the outer band (14), wherein the vane (16) and the bands (12, 14) are part of a monolithic whole of low-ductility material; a metallic baffle (42) inside the vane (16), the baffle (42) having upper (44) and lower ends (46) and a peripheral wall (48) including a plurality of impingement holes (56) defining an interior space, closed off by an end wall (50) at the lower end (46); and a metallic retainer (58) having a body (60) with a shape generally matching the shape of the aperture (30), the body (60) bearing against the upper end (44) of the impingement baffle (42) and being connected to the outer band (14) by a plurality of mechanical fasteners (86, 92).

ENGINE COMPONENT AND METHODS FOR AN ENGINE COMPONENT

An engine component for a gas turbine engine includes a film-cooled substrate having a hot surface facing hot combustion gas and a cooling surface facing a cooling fluid flow. The substrate includes one or more film holes that have a multi-faceted diffusing section configured to improve the adhesion of a coating on the substrate.

ENGINE COMPONENT

An engine component for a gas turbine engine includes a film-cooled substrate having a hot surface facing hot combustion gas and a cooling surface facing a cooling fluid flow. The substrate includes one or more film holes that have a passage with an inlet and an outlet, where the passage includes portions that are angled relative to each other.

SUBSTRATE WITH SHAPED COOLING HOLES AND METHODS OF MANUFACTURE

A substrate having one or more shaped effusion cooling holes formed therein. Each shaped cooling hole has a bore angled relative to an exit surface of the combustor liner. One end of the bore is an inlet formed in an inlet surface of the combustor liner. The other end of the bore is an outlet formed in the exit surface of the combustor liner. The outlet has a shaped portion that expands in only one dimension. Also methods for making the shaped cooling holes.

COOLED VANE OF A HIGH-PRESSURE TURBINE

La présente invention porte sur une aube mobile de turbomachine comportant une pale (12) avec des cavités internes de refroidissement et un pied (11) par lequel l'aube peut être montée sur un disque de rotor, le pied comprenant au moins deux canaux (11c) communiquant avec lesdites cavités internes et débouchant sur sa base (11b), ladite base comprenant au moins deux ouvertures (11b1, 11b2) à travers lesquelles débouchent les canaux, une plaquette de calibrage (20) pourvue de perçages calibrés (21,22 ) correspondant auxdites ouvertures étant fixée sur la base (11b) du pied. L'aube est caractérisée par le fait qu'un moyen mécanique (25, 26; 11m1, 11m2) formant barrière d'étanchéité entre les deux ouvertures est ménagé entre la plaquette et la base du pied.

INTERNAL TURBINE COMPONENT ELECTROPLATING

Method and apparatus are provided for electroplating a surface area of an internal wall defining a cooling cavity present in a gas turbine engine component.

FILM COOLING OF JET ENGINE COMPONENTS

Patent 13DV-10548 A jet engine component, such as an aircraft gas turbine engine rotor blade or a scramjet engine fuel injector. The component has a wall portion including a first surface exposable to a cooler, higher static pressure fluid and a second surface exposable to a hotter, lower static pressure gas flow flowing across the second surface. The component further includes a film coolant passageway having an inlet on the first surface and an outlet on the second surface. The second surface has an open groove extending from the outlet along the gas flow for improved film cooling of the second surface.

GAS TURBINE STATIONARY BLADE

Gas turbine stationary blade is improved in shapes of blade leading edge and fillets, in supporting of inserts and in blowing of cooling air, so that blade cooling efficiency is enhanced, insert supporting structure is simplified and clogging of cooling holes is prevented, thus reliability of the stationary blade is enhanced. Passages (23,24) are provided in stationary blade (10) . Front insert (2) is provided in the passage (23) and rear insert (5) in the passage (5) to be supported at two points of insert supporting portions (3a, 3b), (6a, 6b), respectively. Projection (1) is provided at blade leading edge so that portion where thermal load is high is made smaller in size and number of rows of cooling holes (11a) in this portion is lessened. Air blowing holes (4b) on dorsal side of the front insert (2) and film cooling holes (12) of the blade have diameters larger than those of other holes, so that dusts in cooling air are caused to flow out to prevent clogging of the holes. Curved surface ...

GAS TURBINE AIRFOIL COOLING

A cooling system for airfoil vanes in a turbine of a gas turbine engine wherein the airfoil platforms have openings downstream of the insert tube for passing spent platform impingement coolant air into the airfoil cavity downstream of the insert tube to increase the pressure in the aft zone of the airfoil to improve the temperature gradient across the span of the airfoil.

Maschinenteil mit Kühleinrichtung gegen schädliche Erwärmung, insbesondere Turbinenschaufel und Verfahren zu seiner Herstellung.

Mehrteilige Gasturbinenschaufel.

High pressure turbine vane cooling hole distrubution

A turbine vane for a gas turbine engine with an airfoil portion including a perimeter wall having first, second, third and fourth sets of cooling holes defined therethrough, including the holes numbered HA-1 to HA-13, HB-1 to HB-13, PA-1 to PA-9, and SA-1 to SA-3, respectively, and located such that a central axis thereof extends through the respective point 1 and point 2 having a nominal location in accordance with the X, Y Cartesian coordinate values set forth in Table 3.

Minimizing blockage of holes in turbine engine components

An airfoil for use in a gas turbine engine is provided, the airfoil having a hole therein. A ceramic plug is inserted in the hole so that the plug extends above a depth of a thermal barrier coating, such as a ceramic, to be placed on the airfoil. The airfoil is then coated by non-line of sight vapor deposition and the plugs are then removed.

Cooled turbine blades for a gas-turbine engine

The present invention relates to a cooled turbine blade for a gas-turbine engine having at least one cooling duct ( 14 ) extending radially, relative to a rotary axis of the gas-turbine engine, inside the airfoil and air-supply ducts ( 12 ) issuing into said cooling duct, characterized in that the cooling duct ( 14 ) extends into the blade root ( 6 ) in order to generate close to the wall a cooling airflow moved at high circumferential velocity and radially in helical form and that in the area of the blade root ( 6 ) at least one nozzle-shaped air-supply duct ( 12 ) issues into the cooling duct ( 14 ) tangentially or with a tangential velocity component.

Apparatus and methods for cooling platform regions of turbine rotor blades

A platform cooling configuration in a turbine rotor blade that includes platform slot formed through at least one of the pressure side slashface and the suction side slashface; a removably-engaged impingement insert that separates the platform into two radially stacked plenums, a first plenum that resides inboard of a second plenum; a high-pressure connector that connects the first plenum to the high-pressure coolant region of the interior cooling passage; a low-pressure connector that connects the second plenum to the low-pressure coolant region of the interior cooling passage.

Cooled rotor blade

A cooled turbine rotor blade for a gas turbine engine is provided. The engine has an annular flow path for conducting working fluid though the engine. The blade has an aerofoil section for extending across the annular flow path. The blade further has a root portion radially inward of the aerofoil section for joining the blade to a rotor disc of the engine. The blade further has a platform between the aerofoil section and the root portion. The platform extends laterally relative to the radial direction of the engine to form an inner boundary of the annular flow path and to provide a rear overhang portion which projects in use towards a corresponding platform of a downstream nozzle guide vane. The platform contains at least one internal elongate plenum chamber for receiving cooling air. The longitudinal axis of the plenum chamber is substantially aligned with the circumferential direction of the engine. The plenum chamber supplies the cooling air to a plurality of exit holes formed in the external surface of the rear overhang portion to cool that portion.

Curved film cooling holes for turbine airfoil and related method

A turbine bucket includes an airfoil portion at one end thereof; a root portion at an opposite end thereof; a platform portion between the airfoil portion and the root portion; at least one internal cavity within or radially inward of the platform portion having at least one film cooling hole extending between the at least one cavity and an external surface of the platform portion. The film cooling hole is curved along a length dimension of the film cooling hole.

Airfoil with wrapped leading edge cooling passage

A turbine engine airfoil includes an airfoil structure having an exterior surface providing a leading edge. A radially extending first cooling passage is arranged near the leading edge and includes first and second portions. The first portion extends to the exterior surface and forms a radially extending trench in the leading edge. The second portion is in fluid communication with a second cooling passage. In one example, the second cooling passage extends radially, and the first cooling passage wraps around a portion of the second cooling passage from a pressure side to a suction side between the second cooling passage and the exterior surface. In the example, the first portion is arranged between the pressure and suction sides. In one example, the first cooling passage is formed by arranging a core in an airfoil mold. The trench is formed by the core in one example.

Method of fabricating a component using a two-layer structural coating

A method of fabricating a component is provided. The fabrication method includes depositing a first layer of a structural coating on an outer surface of a substrate. The substrate has at least one hollow interior space. The fabrication method further includes machining the substrate through the first layer of the structural coating, to define one or more openings in the first layer of the structural coating and to form respective one or more grooves in the outer surface of the substrate. Each groove has a respective base and extends at least partially along the surface of the substrate. The fabrication method further includes depositing a second layer of the structural coating over the first layer of the structural coating and over the groove(s), such that the groove(s) and the second layer of the structural coating together define one or more channels for cooling the component. A component is also disclosed.

Angled trench diffuser

An article is disclosed that includes a substrate having a first surface and a second surface and a coating disposed on the second surface. In addition, the article includes an angled trench at least partially defined in the coating. The angled trench may include a bottom surface, a first sidewall and a second sidewall disposed downstream of the first sidewall. The first and second sidewalls may extend from the bottom surface at an angle of less than about 60 degrees. Moreover, the article may include a plurality of holes defined between the first surface and the bottom surface.

Turbine blade and engine component

A turbine blade that is used in a turbine of a gas turbine engine and cooled by cooling air, and includes a cooling channel that is formed within the turbine blade and in which the cooling air flows, plural bottomed recesses that are formed on a blade surface of the turbine blade and of which each downstream-side inner wall is inclined, and an ejection hole that is formed on each bottom of the plural bottomed recesses and communicates with the cooling channel to eject the cooling air. The ejection hole is formed so that a central line of the ejection hole extends along the downstream-side inner wall. The above turbine blade can improve cooling without reducing efficiency.

Component aperture location using computed tomography

An exemplary component measuring method includes determining a position of an aperture of a component using a computed tomography scan of a gage and a component. The gage is inserted into the aperture of the component during the computed tomography scan.

Turbine blade incorporating trailing edge cooling design

A turbine blade ( 10 ) including an airfoil ( 12 ) having multiple interior wall portions ( 70 ) each separating at least one chamber from another one of multiple chambers ( 46, 48, 50, 58, 60 ). In one embodiment a first wall portion ( 70 - 2 ) between first and second chambers ( 60, 52 ) includes first and second pluralities of flow paths ( 86 P, 86 S) extending through the first wall portion. The first wall portion includes a first region R 1 having a first thickness, t, measurable as a distance between the chambers. One of the paths extends a first path distance, d, as measured from an associated path opening ( 78 ) in the first chamber ( 60 ), through the first region and to an exit opening ( 82 ) in the second chamber ( 52 ) which path distance is greater than the first thickness.

Cooled turbine blade for gas turbine engine

A cooled turbine blade for a gas turbine engine including a pressure surface wall, a suction surface wall and a distal wall connecting the pressure surface wall and the suction surface wall, arranged so as to create in the region of the distal end of the blade at least one external cavity forming a bathtub-shaped cavity and at least one internal cavity separated by the distal wall, the blade having at least one opening for the introduction of a flow of cooling air into the external cavity, wherein, on the one hand, at least one part of the distal wall is inclined relative to the verticals of the pressure surface wall and, on the other hand, the opening is created in the vicinity of the distal wall so that the flow of cooling air is directed towards the distal end of the pressure surface wall.

Double-jet type film cooling structure

Provided is a film cooling structure capable of suppressing a cooling medium film from being separated from a wall surface, to increase a film efficiency on the wall surface and thereby-cool the wall surface effectively. One or more pairs of injection holes are formed on a wall surface facing a passage of high-temperature gas to inject a cooling medium to the passage. A single supply passage is formed inside the wall to supply the cooling medium to the injection holes. A separating section is provided between the injection holes in a location forward relative to rear ends of the injection holes to separate the cooling medium into components flowing to the injection holes. An injection direction of the cooling medium is inclined relative to a gas flow direction so that the cooling medium forms swirl flows that push the cooling medium against the wall surface.

Impingement Cooling System for Use with Contoured Surfaces

The present application provides an impingement cooling system for use with a contoured surface. The impingement cooling system may include an impingement plenum and an impingement plate with a linear shape facing the contoured surface. The impingement surface may include a number of projected area thereon with a number of impingement holes having varying sizes and varying spacings.

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.

Surface analysis for detecting closed holes and method for reopening

By means of laser triangulation measurements of an uncoated and a coated component having holes, the exact positions of the holes can be detected for reopening after coating. A method for the surface analysis of at least partially closed holes which are to be opened is provided. A process for reopening coated holes is also provided.

Components with microchannel cooling

A component includes a substrate having outer and inner surfaces, where the inner surface defines at least one hollow, interior space. The outer surface defines pressure side and suction side walls, which are joined together at leading and trailing edges of the component. The outer surface defines one or more grooves that extend at least partially along the pressure or suction side walls in a vicinity of the trailing edge. Each groove is in fluid communication with a respective hollow, interior space. The component further includes a coating disposed over at least a portion of the outer substrate surface. The coating comprises at least a structural coating that extends over the groove(s), such that the groove(s) and the structural coating together define one or more channels for cooling the trailing edge. A method of forming cooling channels in the vicinity of the trailing edge is also provided.

COMPONENTS WITH COOLING CHANNELS AND METHODS OF MANUFACTURE

A method of manufacturing a component is provided. The method includes forming one or more grooves in an outer surface of a substrate. Each groove extends at least partially along the surface of the substrate and has a base, a top and at least one discharge point. The method further includes forming a run-out region adjacent to the discharge point for each groove and disposing a coating over at least a portion of the surface of the substrate. The groove(s) and the coating define one or more channels for cooling the component. Components with cooling channels are also provided. 1. A component comprising:a substrate comprising an outer surface and an inner surface, wherein the outer surface defines one or more grooves and one or more run-out regions, wherein each groove extends at least partially along the outer surface of the substrate and has a base and at least one discharge point, and wherein each run-out region is adjacent to the respective discharge point for a respective groove; anda coating disposed over at least a portion of the outer surface of the substrate, such that the one or more grooves and the coating together define one or more channels for cooling the component.2. The component of claim 1 , wherein the inner surface defines at least one hollow claim 1 , interior space claim 1 , and wherein one or more access holes extend through the base of a respective one of the one or more grooves to place the groove in fluid communication with respective ones of the at least one hollow interior space.3. The component of claim 1 , wherein the run-out region is wider than a top of the respective groove.4. The component of claim 3 , wherein the coating does not bridge the one or more run-out regions claim 3 , such that each run-out region forms a film hole for the respective groove.5. The component of claim 1 , wherein the base of each groove is wider than the top claim 1 , such that each groove comprises a re-entrant shaped groove.6. The component of claim 5 , ...

TURBINE BLADE

A turbine blade () includes a plurality of rows of cooling portions (A, B) that have notch portions () that discharge cooling gas that has been introduced into an internal portion () of a blade portion () onto a ventral side blade surface (), and that are formed in rows that are stacked in a direction between the blade front edge and the blade rear edge. This turbine blade () is provided with: turbulence promoting cooling portions (B) that are provided in the row located furthest to the downstream side from among the plurality of rows, and that have turbulence promoting devices () in areas exposed by the notch portions (); and film cooling portions (A) that are provided in at least one of other rows from among the plurality of rows, and that form a film cooling layer in the areas exposed by the notch portions (). 1. A turbine blade that is provided with a plurality of rows of cooling portions that have notch portions that discharge cooling gas that has been introduced into an internal portion of a blade portion onto a ventral side blade surface , and that are formed in rows that are stacked in a direction between the blade front edge and the blade rear edge , comprising:turbulence promoting cooling portions that are provided in the row located closest to the downstream side from among the plurality of rows, and that have turbulence promoting devices in areas exposed by the notch portions; andfilm cooling portions that are provided in at least one of other rows from among the plurality of rows, and that form a film cooling layer in the areas exposed by the notch portions.2. The turbine blade according to claim 1 , wherein claim 1 , in at least one of the plurality of rows claim 1 , the plurality of cooling portions are placed discretely from each other in the height direction of the blade portion.3. The turbine blade according to claim 1 , wherein claim 1 , in at least one of the plurality of rows claim 1 , the cooling portions are provided continuously in an ...

AIRFOIL COOLING ARRANGEMENT

An airfoil according to an exemplary embodiment of the present disclosure can include an airfoil body having a plurality of film cooling holes that extend through an exterior surface of the airfoil body. Each of the plurality of film cooling holes break through the exterior surface at geometric coordinates in accordance with Cartesian coordinate values of X, Y and Z as set forth in Table 1. Each of the geometric coordinates is measured from a reference point on a leading edge rail of a platform of the airfoil. 1. An airfoil , comprising:an airfoil body having a plurality of film cooling holes that extend through an exterior surface of said airfoil body, wherein each of said plurality of film cooling holes break through said exterior surface at geometric coordinates in accordance with Cartesian coordinate values of X, Y and Z as set forth in Table 1, wherein each of said geometric coordinates is measured from a reference point on a leading edge rail of a platform of the airfoil.2. The airfoil as recited in claim 1 , wherein said airfoil is a first stage turbine vane.3. The airfoil as recited in claim 1 , wherein said Cartesian coordinate values of Table 1 are expressed in inches.4. The airfoil as recited in claim 1 , wherein said reference point includes a pin hole of said platform.5. The airfoil as recited in claim 1 , wherein said plurality of film cooling holes are spaced along a span of said airfoil body in multiple collinearly aligned rows.6. The airfoil as recited in claim 1 , wherein a first portion of said plurality of film cooling holes are cone shaped and a second portion of said plurality of film cooling holes are round shaped.7. The airfoil as recited in claim 1 , wherein at least a portion of said plurality of film cooling holes are arranged in a herringbone configuration.8. The airfoil as recited in claim 1 , wherein said plurality of film cooling holes are disposed on a pressure side claim 1 , a suction side and a leading edge of said airfoil body.9. ...

AIRFOIL COOLING CIRCUITS

An airfoil includes leading and trailing edges; first and second sides extending from the leading edge to the trailing edge, each side having an exterior surface; a core passage located between the first and second sides and the leading and trailing edges; and a wall structure located between the core passage and the exterior surface of the first side. The wall structure includes a plurality of cooling fluid inlets communicating with the core passage for receiving cooling fluid from the core passage, a plurality of cooling fluid outlets on the exterior surface of the first side for expelling cooling fluid and forming a cooling film along the exterior surface of the first side, and a plurality of cooling passages communicating with the plurality of cooling fluid inlets and the plurality of cooling fluid outlets. At least a portion of one cooling passage extends between adjacent cooling fluid outlets. 1. An airfoil comprising:leading and trailing edges;a first side extending from the leading edge to the trailing edge and having an exterior surface;a second side generally opposite the first side and extending from the leading edge to the trailing edge and having an exterior surface;a core passage located between the first and second sides and the leading and trailing edges; and a plurality of cooling fluid inlets communicating with the core passage for receiving cooling fluid from the core passage;', 'a plurality of cooling fluid outlets on the exterior surface of the first side for expelling cooling fluid and forming a cooling film along the exterior surface of the first side; and', 'a plurality of cooling passages communicating with the plurality of cooling fluid inlets and the plurality of cooling fluid outlets, wherein at least a portion of one cooling passage extends between adjacent cooling fluid outlets., 'a wall structure located between the core passage and the exterior surface of the first side, the wall structure comprising2. The airfoil of claim 1 , wherein ...

Gas turbine engine component cooling circuit

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a body portion, a cooling circuit disposed within the body portion and including at least a first cavity and a microcircuit in fluid communication with the first cavity. A plunged hole intersects at least a portion of the microcircuit.

TURBINE VANE

A plurality of film cooling holes are formed, so as to communicate with a front cooling passage, in a vane surface on the front-edge side of a stator vane body of a turbine stator vane. The hole cross-section of each of the film cooling holes has a rectangular long-hole shape extending in a direction parallel to the cross-section along the span direction and having a rounded corner. The hole-center line of each of the film cooling holes is inclined with respect to the thickness direction in the cross-section along the span direction. The exit-side portion of the hole wall surface of each of the film cooling holes is inclined with respect to the thickness direction by a greater degree than that of the hole-center line. 1. A turbine vane for a turbine of a gas turbine engine , and capable of being cooled by cooling air , the turbine vane comprising: a vane surface;', 'a cooling passage allowing the cooling air to flow into the vane body, and', 'a plurality of film cooling holes formed in the vane surface on a front edge side of the vane body so as to communicate with the cooling passage to jet out the cooling air through the plurality of film cooling holes, a hole cross section of each film cooling hole having a long-hole shape extending in a direction parallel to a cross section along a span direction of the vane body, a hole-center line of each film cooling hole tilting from a thickness direction of the vane body on a cross section of the vane surface along the span direction, and an exit-side and obtuse angle-side portion of a hole wall surface of each film cooling hole tilting further from the thickness direction than the hole-center line on the cross section along the span direction., 'a vane body including2. The turbine vane according to claim 1 , whereinan aspect ratio of the hole cross section of each film cooling hole is set in a range of 1.1 to 3.0.3. The turbine vane according to claim 1 , whereina tilt angle of the hole-center line of each film cooling ...

AIRFOIL DESIGN HAVING LOCALIZED SUCTION SIDE CURVATURES

An airfoil for a gas turbine engine comprises a radially extending body having a transverse cross-section. The transverse cross-section comprises a leading edge, a trailing edge, a pressure side and a suction side. The pressure side extends between the leading edge and the trailing edge with a predominantly concave curvature. The suction side extends between the leading edge and the trailing edge with a predominantly convex curvature. The suction side includes an approximately flat portion flanked by forward and aft convex portions. In another embodiment, the suction side includes a series of local curvature changes that produce inflection points in the convex curvature of the suction side spaced from the trailing edge. 1. An airfoil for a gas turbine engine , the airfoil comprising: a leading edge;', 'a trailing edge;', 'a pressure side extending between the leading edge and the trailing edge with a predominantly concave curvature; and', 'a suction side extending between the leading edge and the trailing edge with a predominantly convex curvature that includes an approximately flat portion flanked by forward and aft convex portions., 'a radially extending body having a transverse cross-section comprising2. The airfoil of wherein the approximately flat portion is formed by a series of local curvature changes that produce inflection points in the convex curvature of the suction side.3. The airfoil of wherein the approximately flat portion is defined by a plurality of changes in sign of the second derivative of a curve defining the suction side.4. The airfoil of wherein the radially extending body defines a chord length and wherein the approximately flat portion is located within a mid-chord region of the airfoil on the suction side.5. The airfoil of wherein the approximately flat portion joins the forward convex portion at a mid-chord point of the airfoil.6. The airfoil of wherein the approximately flat portion joins the forward convex portion at a throat of the ...

Cooled blade for a gas turbine

The invention relates to a cooled blade for a gas turbine that includes a radially extending aerofoil with a leading edge, a trailing edge, a suction side and a pressure side, and wherein a lip overhang is provided on the suction side of the trailing edge The blade also includes a plurality of radial internal flow channels connected via flow bends to form a multi-pass serpentine for a coolant flow, whereby a trailing edge ejection region is provided for cooling said trailing edge, said trailing edge ejection region comprising a trailing edge passage of said multi-pass serpentine running essentially parallel to said trailing edge and being connected over its entire length with a pressure side bleed. An optimized cooling is achieved by mainly determining the cooling flow from the trailing edge passage to the pressure side bleed by means of a staggered field of pins, which is provided between said pressure side bleed and said trailing edge passage, with the lateral dimension of said pins increasing in coolant flow direction.

Gas turbine engine airfoil internal cooling features

An airfoil for a gas turbine engine includes spaced apart pressure and suction walls joined at leading and trailing edges to provide an airfoil. Intermediate walls interconnect the pressure and suction walls to provide cooling passageways. The cooing passageways have interior pressure and suction surfaces that are respectively provided on the pressure and suction walls. Trip strips include a chevron-shaped trip strip that is provided on at least one of the interior pressure and suction surfaces.

Gas Turbine, Gas Turbine Blade, and Manufacturing Method of Gas Turbine Blade

A gas turbine blade includes a hollow-blade-form portion formed by a leading edge on an upstream side of an working fluid of a gas turbine in a flow direction, a trailing edge on a downstream side of the working fluid in the flow direction, and a suction surface and a pressure surface reaching the trailing edge from the leading edge, and a shank portion for supporting the blade form portion. The blade also includes a partition for connecting the suction surface and the pressure surface in a hollow region of the blade-form portion, coolant paths formed by the partition, the suction surface, and the pressure surface, an impingement cooling hole formed in the partition for dividing the first path that is a flow path nearest to the leading edge side among the coolant paths and the second path adjacent to the first path, and first and second converter portions.

TURBINE ROTOR BLADE PLATFORM COOLING

A cooling arrangement in a platform in a rotor blade or a sidewall in a stator blade in a turbine of a combustion turbine engine is described. The cooling arrangement may include: a cooling chamber configured to pass coolant from an inlet to an outlet; and a rib positioned within the cooling chamber. The rib may partially divide the cooling chamber so to form a switchback. The rib may be canted with respect to the cooling chamber such that the switchback has an ever narrowing channel. 1. A cooling arrangement in one of a sidewall of a stator blade and a platform in a rotor blade in a turbine of a combustion turbine engine , the cooling arrangement comprising:a cooling chamber configured to pass coolant from an inlet to an outlet; anda rib positioned within the cooling chamber, the rib partially dividing the cooling chamber so to form a switchback;wherein the rib is canted with respect to the cooling chamber such that the switchback comprises an ever narrowing channel.2. The cooling arrangement according to claim 1 , wherein ever narrowing channel comprises a channel that narrows at a constant rate as the channel extends from the inlet to the outlet of the cooling chamber.3. The cooling arrangement according to claim 1 , wherein the ever narrowing channel comprises a channel that narrows at a constant rate along both flanks of the rib.4. The cooling arrangement according to claim 3 , wherein the switchback comprises a pass positioned on each flank of the rib: an upstream pass disposed on a flank of the rib that coincides with the inlet; and a downstream pass disposed on a flank of the rib that coincides with the outlet; andwherein, between the upstream pass and the downstream pass, the switchback comprises a turn section that defines a turn of approximately 180°.5. The cooling arrangement according to claim 4 , wherein the cooling chamber comprises a first edge and a second edge claim 4 , wherein the second edge opposes the first edge across the cooling chamber; ...

Turbine blade of a gas turbine with swirl-generating element and method for its manufacture

The present invention relates to a turbine blade of a gas turbine with an airfoil arranged on a blade root and having at least one cooling air duct running in the longitudinal direction of the turbine blade, arranged inside the turbine blade and extending through the blade root, characterized in that at least one swirl-generating element is arranged in the transitional area between blade root and airfoil in the cooling air duct, with the swirl-generating element including an outer ring and several swirl-generating stator vanes arranged thereon, which are connected to a centric area, as well as to a method for its manufacture. 1. Turbine blade of a gas turbine with an airfoil arranged on a blade root and having at least one cooling air duct running in the longitudinal direction of the turbine blade , arranged inside the turbine blade and extending through the blade root , characterized in that at least one swirl-generating element is arranged in the transitional area between blade root and airfoil in the cooling air duct , with the swirl-generating element including an outer ring and several swirl-generating stator vanes arranged thereon , which are connected to a centric area.2. Turbine blade in accordance with claim 1 , characterized in that inside the cooling air duct at least one fixing element is provided that can be positively engaged with the swirl-generating element.3. Turbine blade in accordance with claim 1 , characterized in that the outer ring is dimensioned such that it can be inserted into the cooling air duct from the blade root.4. Turbine blade in accordance with claim 1 , characterized in that the swirl-generating element is connected to the turbine blade in one piece.5. Turbine blade in accordance with claim 1 , characterized in that the swirl-generating element is designed in the form of a guide vane arrangement imparting a swirl motion to the cooling airflow.6. Turbine blade in accordance with claim 1 , characterized in that the swirl-generating ...

GAS TURBINE ENGINE COMPONENT

An internally cooled gas turbine engine component has a line of cooling air discharge holes, an internal cooling channel, an internal feed cavity for feeding cooling air from the channel to the discharge holes, and flow disrupting pedestals arranged in rows. A method of configuring the component includes: 1. A method of configuring an internally cooled gas turbine engine component , the component having a line of cooling air discharge holes , an internal cooling channel forward of and extending substantially parallel to the line of discharge holes , and an internal feed cavity between the channel and the line of discharge holes for feeding cooling air from the channel to the discharge holes , the component further having a plurality of flow disrupting pedestals extending between opposing sides of the feed cavity , the pedestals being arranged in a number N of rows which extend substantially parallel to the line of discharge holes , the first row being at the entrance from the channel to the feed cavity , the Nrow being at the exit from the feed cavity to the discharge holes , the remaining rows being spaced therebetween , and the pedestals being spaced apart from each other within each row , the method including:determining an angle α of the direction of cooling air flow into the first row;{'sup': 'th', 'determining an angle β of the direction of cooling air flow from the Nrow;'}defining a change in angle φ of the direction of cooling air flow between rows as φ=(β−α)/N; and{'sup': th', 'th', 'th, 'positioning the pedestals such that a line extending forward from the centre of each pedestal in the irow at an angle {α+φ(i−1)} intersects the (i−1)row at a location which is midway between two neighbouring pedestals of the (i−1)row, i being an integer from 2 to N.'}2. A method according to claim 1 , wherein the rows are spaced substantially equal distances apart.3. A method according to claim 1 , wherein the method further includes:{'sup': th', 'th, 'determining the ...

METHOD AND COOLING SYSTEM FOR COOLING BLADES OF AT LEAST ONE BLADE ROW IN A ROTARY FLOW MACHINE

A method and a cooling system for cooling blades of at least one blade row in a rotary flow machine includes an axial flow channel which is radially limited on the inside by a rotor unit and at the outside by at least one stationary component, the blades are arranged at the rotary unit and provide a shrouded blade tip facing radially to said stationary component. Pressurized cooling air is fed through from radially outside towards the tip of each of said blades in the at least one blade row, and the pressurized cooling air enters the blades through at least one opening at the shrouded blades' tip. 1. A method for cooling blades of at least one blade row in a rotary flow machine , comprising an axial flow channel which is radially limited on the inside by a rotor unit and at the outside by at least one stationary component , said blades are arranged at the rotary unit and provide a shrouded blade tip facing radially to said stationary component , wherein the pressurized cooling air is fed through from radially outside towards the tip of each of said blades in the at least one blade row , and said pressurized cooling air enters the blades through at least one opening at the shrouded blades' tip.2. The method according to claim 1 , wherein the pressurized cooling air is fed through the stationary component surrounding said at least one blade row radially and entering a cavity enclosed by the stationary component and shrouded tips of the blades in the at least one blade row.3. The method according to claim 2 , wherein the pressurized cooling air is fed into the cavity through at least one claim 2 , stationary component sided opening such that a static pressure prevails within said cavity which is higher than a total relative pressure of a flow in the axial flow channel at a leading edge of the blades in the at least one blade row.4. A cooling system for cooling blades of at least one blade row in a rotary flow machine comprising an axial flow channel which is radially ...

GAS TURBINE ENGINE COMPONENTS WITH LATERAL AND FORWARD SWEEP FILM COOLING HOLES

An engine component includes a body having an internal surface and an external surface, the internal surface at least partially defining an internal cooling circuit. The engine component further includes a plurality of cooling holes formed in the body and extending between the internal cooling circuit and the external surface of the body. The plurality of cooling holes includes a first cooling hole with forward diffusion and lateral diffusion. 1. An engine component , comprising:a body having an internal surface and an external surface, the internal surface at least partially defining an internal cooling circuit; anda plurality of cooling holes formed in the body and extending between the internal cooling circuit and the external surface of the body, the plurality of cooling holes including a first cooling hole with forward diffusion and lateral diffusion.2. The engine component of claim 1 , wherein the first cooling hole includes an inlet at the internal cooling circuit claim 1 , a metering section extending from the inlet claim 1 , a first exit portion extending from the metering section claim 1 , a second exit portion extending from the first exit portion claim 1 , and an outlet defined on the external surface and fluidly coupled to the second exit portion.3. The engine component of claim 2 , wherein the first exit portion extends at a first angle relative to the metering section and the second exit portion extends at a second angle relative to the metering section claim 2 , the second angle being greater than the first angle.4. The engine component of claim 2 , wherein the first exit portion is curved with a first radius of curvature and the second exit portion is curved with a second radius of curvature claim 2 , the second radius of curvature being less than the first radius of curvature.5. The engine component of claim 2 , wherein the first exit portion is curved with a first radius of curvature and the second exit portion is curved with a second radius of ...

METHOD FOR CHECKING COOLING HOLES OF A GAS TURBINE BLADE

A method for checking cooling holes of a gas turbine blade, including in sequence: step 1: injecting a liquid plastic material into a cooling hole of a gas turbine blade; step 2: stopping the flow of the liquid plastic material, and curing the liquid plastic material to form a model; step 3: separating the model from the gas turbine blade; step 4: scanning the model. By scanning the model, the cooling holes of the gas turbine blade can be checked, so as to obtain more data about the position, dimensions and shape thereof, while improving the accuracy and efficiency of measurement. 1. A method for checking at least one of positions , angles , dimensions , and shapes of cooling holes which pass between an outer surface and an interior cavity of a gas turbine blade , comprising in sequence:step 1: injecting a liquid plastic material into at least one cooling hole of the gas turbine blade;step 2: stopping the flow of the liquid plastic material, and curing the liquid plastic material to form a model of the at least one cooling hole;step 3: separating the model from the gas turbine blade;step 4: scanning the model to determine at least one of the positions, angles, dimensions, and shapes of the model of the at least one cooling hole.2. The method as claimed in claim 1 , further comprising after step 4:step 5: comparing data about the model, obtained by scanning, with data about a standard model of the gas turbine blade.3. The method as claimed in claim 1 , further comprising performing steps 1 to 4 on multiple cooling holes at the same time.4. The method as claimed in claim 2 , further comprising obtaining the data about the standard model by performing steps 1 to 4 in sequence.5. The method as claimed in claim 1 , further comprising before step 1:step 11: covering an inner wall of the cooling hole with a layer of mold-release coating for improving the separating of the model.6. The method as claimed in claim 3 , wherein step 3 comprises claim 3 , in sequence:step 31: ...

FILM COOLING DIFFUSER HOLE

An airfoil for a gas turbine engine is disclosed. In various embodiments, the airfoil includes a cooling passage; an outer wall separating a core flow path from the cooling passage; a diffuser in fluid communication with the cooling passage and opening into the core flow path, the diffuser being characterized by a linear ridge on a downstream end of the diffuser; and a thermal barrier coating covering the outer wall and the linear ridge. 1. An airfoil for a gas turbine engine , comprising:a cooling passage;an outer wall separating a core flow path from the cooling passage;a diffuser in fluid communication with the cooling passage and opening into the core flow path, the diffuser being characterized by a linear ridge on a downstream end of the diffuser; and 'wherein the linear ridge includes an upstream facing side that is characterized by a height extending a first distance in a direction normal to the outer wall, the height having a value within a range of between five one-hundredths and seventy-five one-hundredths of a depth of the cooling passage, the depth extending between a first wall and a second wall that define the cooling passage and in the direction normal to the outer wall.', 'a thermal barrier coating covering the outer wall and the linear ridge,'}2. The airfoil of claim 1 , wherein the diffuser defines a rectangular shape in the direction normal to the outer wall.3. The airfoil of claim 2 , wherein the linear ridge extends perpendicular to the cooling passage along the downstream end of the diffuser.4. The airfoil of claim 3 , wherein the thermal barrier coating includes a first portion upstream of the linear ridge claim 3 , the first portion extending from the cooling passage and being characterized by a first radius of curvature.5. The airfoil of claim 4 , wherein the thermal barrier coating includes a second portion claim 4 , the second portion extending from the first portion and over the linear ridge and being characterized by a second radius of ...

AIRFOIL TIP POCKET WITH AUGMENTATION FEATURES

A component for a gas turbine engine includes, among other things, an airfoil that includes a pressure sidewall and a suction sidewall that meet together at both a leading edge and a trailing edge, the airfoil extending radially from a platform to a tip, a tip pocket formed in the tip and terminating prior to the trailing edge, and one or more heat transfer augmentation devices formed in the tip pocket. 1. A component for a gas turbine engine comprising:an airfoil extending in a chordwise direction between a leading edge and a trailing edge and in a thickness direction between a pressure sidewall and a suction sidewall that meet together at both the leading edge and the trailing edge, the airfoil extending in a radial direction from a platform to a tip;a tip pocket formed in the tip and terminating prior to the trailing edge;wherein the tip pocket includes a suction side lip, a pressure side lip, a leading edge lip and a trailing edge lip that each extend outwardly in the radial direction from a floor to the tip;one or more heat transfer augmentation devices formed in the tip pocket, the one or more heat transfer augmentation devices including at least one rib extending from one of the suction side lip and the pressure side lip such that a wall of the at least one rib is spaced apart from another one of the suction side lip and the pressure side lip;wherein the at least one rib extends outwardly in the radial direction from the floor towards the tip such that a length of the at least one rib is slanted at an acute angle in the chordwise direction toward either the leading edge lip or the trailing edge lip relative to the floor; anda plurality of cooling holes defined in the floor that fluidly connect the tip pocket to at least one internal cooling cavity formed inside the airfoil.2. The component as recited in claim 1 , wherein at least one of the plurality of cooling holes is angled relative to the floor.3. The component as recited in claim 1 , wherein the at least ...

GAS TURBINE ENGINE COMPONENT HAVING TRANSVERSELY ANGLED IMPINGEMENT RIBS

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a body portion that includes a first wall spaced apart from a second wall and disposed about a centerline axis. At least one rib extends between the first wall and the second wall. The at least one rib extends along a rib axis that is transversely angled relative to the centerline axis. At least one impingement hole extends through the at least one rib. 1. A component for a gas turbine engine , comprising:a body portion that includes a first wall spaced apart from a second wall and disposed about a centerline axis;at least one rib that extends between said first wall and said second wall, wherein said at least one rib extends along a rib axis that is transversely angled relative to said centerline axis; andat least one impingement hole that extends through said at least one rib.2. The component as recited in claim 1 , wherein said body portion is an airfoil of one of a blade and a vane.3. The component as recited in claim 1 , wherein said body portion is part of one of a blade outer air seal (BOAS) and a combustor liner.4. The component as recited in claim 1 , comprising a cooling circuit disposed within said body portion and including at least a first cavity and a second cavity in fluid communication with said first cavity.5. The component as recited in claim 1 , wherein said first wall is a suction side wall and said second wall is a pressure side wall.6. The component as recited in claim 1 , wherein said at least one impingement hole is oriented toward said first wall.7. The component as recited in claim 1 , wherein said at least one impingement hole is oriented toward said second wall.8. The component as recited in claim 1 , wherein said at least one rib includes a first impingement hole that is oriented toward said first wall and a second impingement hole that is oriented toward said second wall.9. The component as recited in claim 1 , ...

GAS TURBINE ENGINE COMPONENT COOLING WITH INTERLEAVED FACING TRIP STRIPS

A gas turbine engine component includes first and second walls spaced apart from one another to provide a cooling passage. First and second trip strips are respectively provided on the first and second walls and arranged to face one another. The first and second trip strips are arranged in an interleaved fashion with respect to one another in a direction. 1. A gas turbine engine component comprising:first and second walls spaced apart from one another to provide a cooling passage, first and second trip strips respectively provided on the first and second walls and arranged to face one another, the first and second trip strips arranged in an interleaved fashion with respect to one another in a direction.2. The gas turbine engine component according to claim 1 , wherein the gas turbine engine component is an airfoil claim 1 , and the direction is a radial direction of the airfoil.3. The gas turbine engine component according to claim 2 , wherein the cooling passage is provided near a leading edge of the airfoil.4. The gas turbine engine component according to claim 3 , wherein trip strips are chevron trip strips arranged asymmetrically claim 3 , an apex of the chevron trip strips shifted within the cooling passage toward a leading edge.5. The gas turbine engine component according to claim 2 , wherein the cooling passage is provided near a trailing edge of the airfoil.6. The gas turbine engine component according to claim 1 , wherein the trip strips are chevron trip strips.7. The gas turbine engine component according to claim 1 , wherein the trip strips extend from an inner surface a distance e claim 1 , and the first and second walls respectively include first and second inner surfaces that are spaced a distance H from one another claim 1 , an e/H ratio is provided in the range of 0.05-0.40.8. The gas turbine engine component according to claim 7 , wherein the trip strips are spaced from an opposing surface a distance in the range 0.035-0.045 inch (0.89-1.14 mm) ...

GAS TURBINE ENGINE SHAPED FILM COOLING HOLE

A component for a gas turbine engine includes a wall that adjoins an interior cooling passage and provides an exterior surface. A film cooling hole fluidly connects the interior cooling passage and the exterior surface. The film cooling passage includes inlet and outlet passages that fluidly interconnect and adjoin one another in a misaligned non-line of sight relationship. 1. A component for a gas turbine engine comprising:a wall adjoining an interior cooling passage and providing an exterior surface, and a film cooling hole fluidly connecting the interior cooling passage and the exterior surface, the film cooling passage including inlet and outlet passages fluidly interconnecting and adjoining one another in a misaligned non-line of sight relationship.2. The component according to claim 1 , wherein the inlet and outlet passages are generally linear.3. The component according to claim 2 , wherein are arranged at an angle relative to one another.4. The component according to claim 3 , wherein the angle is acute.5. The component according to claim 1 , wherein the outlet passage provides a diffuser shape.6. The component according to claim 1 , wherein the inlet passage provides a metering section having a cross-sectional area that is less than a cross-sectional area of the outlet portion.7. The component according to claim 6 , wherein the inlet passage includes first and second metering portions claim 6 , the second metering portion adjoining the outlet passage and including a length L and a diameter D having an L/D ratio of greater than 1.8. The component according to claim 8 , wherein the L/D ratio is greater than 3.9. The component according to claim 1 , wherein the film cooling hole is additively manufactured.10. The component according to claim 1 , wherein the component is one of an airfoil combustor BOAS and platform.11. A method of manufacturing airfoil component for a gas turbine engine claim 1 , comprising:depositing multiple layers of a powdered metal onto ...

METHOD FOR DETECTING A COMPROMISED COMPONENT

A method for providing visually detectable changes to a surface that has been subjected to a temperature in excess of a predetermined temperature. A coating is applied to the surface, wherein the coating will melt when the predetermined temperature has been reached. Centrifugal forces acting on the melted coating will cause it to be displaced such that the disturbed surface is visibly detectable upon inspection after solidifying. 1. A method for determining if a component having a coating thereon has been compromised , the method comprising the steps of:a) visually inspecting the coating; andb) determining that the component has been compromised if the coating is displaced.2. The method of claim 1 , wherein the coating comprises a metallic coating.3. The method of claim 2 , wherein the metallic coating comprises NiCoCrAlY.4. The method of claim 2 , wherein the metallic coating comprises more than one layer.5. The method of claim 1 , wherein the component comprises an turbine blade in a turbine engine.6. The method of claim 5 , wherein the component comprises a high pressure turbine blade in a turbine engine.7. The method of claim 1 , further comprising the step of:c) determining that the component should be serviced if it is determined at step (b) that the component has been compromised.8. The method of claim 1 , wherein the displaced coating is rumpled.9. The method of claim 1 , wherein the displaced coating exhibits evidence of having flowed.10. The method of claim 2 , wherein the metallic coating is at least partially covered by a thermal barrier coating.11. The method of claim 1 , wherein the component includes at least one cooling hole formed therein and the displaced coating has flowed into at least one of the at least one cooling holes.12. A method for determining if a component has been operated above a predetermined temperature claim 1 , the method comprising the steps of:a) applying a coating to the component;b) visually inspecting the coating; andc) ...

Turbine airfoil with ambient cooling system

A turbine airfoil usable in a turbine engine and having at least one ambient air cooling system is disclosed. At least a portion of the cooling system may include one or more cooling channels configured to receive ambient air at about atmospheric pressure. The ambient air cooling system may have a tip static pressure to ambient pressure ratio of at least 0.5, and in at least one embodiment, may include a tip static pressure to ambient pressure ratio of between about 0.5 and about 3.0. The cooling system may also be configured such that an under root slot chamber in the root is large to minimize supply air velocity. One or more cooling channels of the ambient air cooling system may terminate at an outlet at the tip such that the outlet is aligned with inner surfaces forming the at least one cooling channel in the airfoil to facilitate high mass flow.

METHOD FOR MAKING GAS TURBINE ENGINE CERAMIC MATRIX COMPOSITE AIRFOIL

A method for making a gas turbine engine ceramic matrix composite airfoil is disclosed. The method includes fabricating an airfoil preform that has a slotted forward end and a continuous trailing end. The slotted forward end of the airfoil preform is coupled to an airfoil core insert. A ceramic matrix composite covering is applied to cover the slots of the airfoil perform. The continuous trailing end of the airfoil preform is removed to expose the slots. A gas turbine engine airfoil is also disclosed. 1. A method comprising:fabricating an airfoil preform having a slotted forward end and a continuous trailing end;coupling the slotted forward end of the airfoil preform to an airfoil core insert;applying a ceramic matrix composite covering to cover the slots of the airfoil preform; andremoving the continuous trailing end of the airfoil preform to expose the slots.2. The method of in which the airfoil preform comprises a monolithic ceramic.3. The method of in which the slots are substantially perpendicular to the continuous trailing end of the airfoil preform.4. The method of in which the slots extend through the thickness of the airfoil preform.5. The method of in which the slots are formed by machining material from the airfoil preform.6. The method of in which the coupling comprises capturing the slotted forward end of the airfoil preform in a spanwise groove in the trailing end of the airfoil core insert.7. A method comprising:forming a ceramic material airfoil core having a spanwise extending delivery member at a leading end, a spanwise extending sacrificial member at a trailing end, and a plurality of channel-defining members extending chordwise between the spanwise extending delivery member and the spanwise extending sacrificial member, wherein adjacent ones of the plurality of channel-defining members define therebetween cooling channels in the ceramic material airfoil core;applying a ceramic fiber cover to the ceramic material airfoil core to encapsulate the ...

TURBINE BLADE

A turbine vane for a rotary turbomachine is described having a turbine blade which is delimited by a concave pressure-side wall and a convex suction-side wall which are connected in the region of a vane front edge which can be assigned to the turbine blade and enclose a cavity which extends in the longitudinal extent of the vane front edge and is delimited on the inner wall by the pressure-side wall and the suction-side wall in the region of the vane front edge and by an intermediate wall which extends in the longitudinal direction to the vane front edge and connects the suction-side wall and the pressure-side wall on the inner wall. The disclosed vane is distinguished by the fact that the intermediate wall has a perforation at least in sections in the connecting region to the suction-side wall and/or pressure-side wall, in order to increase the elasticity of the intermediate wall. 1. A turbine blade for a rotating turbomachine; the turbine blade comprising:having a blade airfoil which is bounded by a concave pressure side wall and a convex suction side wall which are connected in the region of a blade leading edge which can be assigned to the blade airfoil, and which enclose a cavity which extends in the longitudinal extent of the blade leading edge and is delimited internally by the pressure side wall and suction side wall in the region of the blade leading edge and by an intermediate wall which extends in the longitudinal direction to the blade leading edge and connects the suction side wall and pressure side wall internally, whereinthe intermediate wall has, at least in sections, a perforation in a connection region to the suction side wall and/or pressure side wall in order to increase the elasticity of the intermediate wall in the connection region.2. The turbine blade as claimed in claim 1 , wherein the perforation comprises a row of cylindrical holes.3. The turbine blade as claimed in claim 1 , wherein the perforation comprises a row of longitudinal holes or ...

Turbine rotor blade and gas turbine

A second outer surface (49ab) of a top plate (49) is recessed from a first outer surface (49aa) of the top plate (49) in the direction away from the inner peripheral surface (34a) of a turbine casing (34) so that a step (50) is formed between the second outer surface (49ab) and the first outer surface (49aa). At least part of the discharge opening (53B) of a cooling hole (53) is disposed in the second outer surface (49ab). The cooling hole (53) extends so as to be tilted relative to the second outer surface (49ab) so that the cooling hole (53) discharges a cooling medium to the upstream side of a combustion gas flowing between the second outer surface (49ab) and the inner peripheral surface (34a) of the turbine casing (34).

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

A method of manufacturing a cooled gas turbine component includes forming a core with an outer surface. The outer surface includes a core feature. The method also includes casting an outer wall of an airfoil about the core. The outer wall has an exterior surface and an interior surface. The interior surface includes a shaped inlet portion that corresponds to the core feature. Moreover, the method includes forming an outlet portion through the outer wall to fluidly connect the outlet portion to the shaped inlet portion. The shaped inlet portion and the outlet portion cooperatively define a cooling aperture through the outer wall. 1. A method of manufacturing a cooled gas component for a turbomachine , the method comprising:forming a core with an outer surface, the outer surface including a core feature that tapers to reduce in width as the core feature projects from the core;casting an outer wall about the core including casting a portion of the outer wall that covers the core feature, the portion including an exterior surface and an interior surface, the interior surface defining a shaped inlet portion that is cast to inversely correspond to the core feature; andafter casting the outer wall, forming an outlet portion through the outer wall by progressively removing material through the portion of the outer wall in a direction from the exterior surface toward the interior surface and toward the shaped inlet portion to fluidly connect the outlet portion to the shaped inlet portion, the shaped inlet portion and the outlet portion cooperatively defining a cooling aperture through the outer wall.2. The method of claim 1 , further comprising adjusting the outlet portion by changing a dimension of the outlet portion.3. The method of claim 1 , further comprising:re-casting the outer wall; andforming an adjusted outlet portion through the re-cast outer wall.4. The method of claim 3 , wherein forming the outlet portion through the outer wall includes forming the outlet ...

ENGINE COMPONENT WITH SET OF COOLING HOLES

An apparatus and method an engine component for a turbine engine comprising an outer wall bounding an interior and defining a pressure side and an opposing suction side, with both sides extending between a leading edge and a trailing edge to define a chord-wise direction, and extending between a root and a tip to define a span-wise direction, at least one cooling passage located within the interior, a set of cooling holes having an inlet fluidly coupled to the cooling passage, an outlet located on one of the pressure side or suction side, with a connecting passage fluidly coupling the inlet to the outlet. 130-. (canceled)31. An airfoil for a turbine engine comprising:an outer wall bounding an interior and defining a pressure side and an opposing suction side, with both sides extending between a leading edge and a trailing edge to define a chord-wise direction, and extending between a root and a tip to define a span-wise direction;at least one cooling passage located within the interior;a first set of cooling holes having first inlet fluidly coupled to the cooling passage, a first outlet located on the pressure side, with a first connecting passage having a curvilinear centerline fluidly coupling the first inlet to the first outlet, and the first connecting passage having a portion extending along the suction side;a second set of cooling holes having second inlet fluidly coupled to the cooling passage, a second outlet located on the suction side, with a second connecting passage having a curvilinear centerline fluidly coupling the second inlet to the second outlet, and the second connecting passage having a portion extending along the pressure side; anda third set of cooling holes having a third outlet proximate the leading edge of the outer wall.321. The airfoil of claim wherein the at least one cooling passage is multiple cooling passages separated by an interior wall and the inlet is located along the interior wall.331. The airfoil of claim wherein at least one of ...

COMPONENT FOR A GAS TURBINE ENGINE WITH A FILM HOLE

A component is provided and comprises at least one wall comprising a first and a second surface. At least one film cooling hole extends through the wall between the first and second surfaces and has an outlet region at the second surface. The film cooling hole includes a first expansion section being a side diffusion portion and a second expansion section being a layback diffusion portion, wherein the side diffusion portion is upstream and spaced from the layback diffusion portion. 1. A component for a gas turbine engine comprising:a hot side exposed to a hot air flow;a cool side exposed to a cooling air flow;a film hole passage extending between the cool side and the hot side with an inlet on the cool side and an outlet on the hot side, the film hole passage defining a diameter, the film hole passage further defining a side diffusion portion defining a side diffusion length between a start of the side diffusion portion and the outlet, and a layback diffusion portion defining a layback length between a start of the layback diffusion portion and the outlet, wherein the side diffusion length is greater than the layback diffusion length.2. The component of wherein the side diffusion portion is upstream and spaced from the layback diffusion portion.3. The component of wherein the side diffusion portion defines a side diffusion angle claim 1 , α claim 1 , relative to a centerline for the film hole passage claim 1 , and the side diffusion angle is less than 12.5 degrees.4. The component of wherein the layback diffusion portion defines a layback diffusion angle claim 3 , ß claim 3 , relative to a centerline for the film hole passage claim 3 , and the layback diffusion angle is less than 12 degrees.5. The component of wherein the layback diffusion length is less than 4 times the diameter.6. The component of wherein the layback diffusion length is equal to or less than zero.7. The component of wherein the layback diffusion length is less than four times the diameter and the ...

ADDITIVE MANUFACTURING METHOD FOR THE ADDITION OF FEATURES WITHIN COOLING HOLES

A method for forming a diffusion cooling hole in a substrate includes removing material from the substrate to form a metering section having an inlet on a first side of the substrate and removing material from the substrate to form a diffusing section that extends between the metering section and an outlet located on a second side of the substrate generally opposite the first side. The method also includes forming a feature on a substrate surface within one of the metering section and the diffusing section. Forming the feature includes depositing a material on the substrate surface and selectively heating the material to join the material with the substrate surface and form the feature. 1. A method for forming a diffusion cooling hole in a substrate , the method comprising:removing material from the substrate to form a metering section having an inlet on a first side of the substrate;removing material from the substrate to form a diffusing section that extends between the metering section and an outlet located on a second side of the substrate generally opposite the first side; depositing a material on the substrate surface;', 'selectively heating the material to join the material with the substrate surface and form the feature., 'forming a feature on a substrate surface within one of the metering section and the diffusing section comprising2. The method of claim 1 , wherein the steps of removing material from the substrate to form a metering section and removing material from the substrate to form a diffusing section are performed by a technique selected from the group consisting of casting claim 1 , drilling claim 1 , laser drilling claim 1 , machining claim 1 , electrical discharge machining and combinations thereof.3. The method of claim 1 , wherein the substrate surface on which the feature is formed is located within the metering section.4. The method of claim 3 , wherein the feature obscures a line of sight between the inlet and the outlet.5. The method of ...

COOLING HOLE INSPECTION SYSTEM

An inspection system includes a thermographic sensor configured to capture thermographic data of a component having holes as a fluid is pulsed toward the holes, and one or more processors configured to temporally process the thermographic data to calculate temporal scores and spatial scores for the corresponding holes. The scores can be used to obtain a reference dataset and a test dataset. A performance score can be assigned to the component based on the difference between the datasets. 1. A method of performance scoring airflow through a turbine engine component having a first surface and a second surface spaced from the first surface , and a plurality of film holes with inlets formed in the second surface fluidly coupled to outlets formed in the first surface , the method comprising:generating a reference dataset for at least a subset of the plurality of film holes, the reference dataset comprising a reference airflow value indicative of the airflow through the film holes for each of the film holes in the subset of the plurality of film holes, where the reference airflow value is a normalized value;determining a test dataset for a turbine engine component comprising test airflow values for at least some of the film holes in the subset of the plurality of film holes;calculating a difference between the test dataset and the reference dataset; andcalculating a performance score for the turbine engine component based on the difference.2. The method of wherein the normalized value is determined from an aggregate of measured airflow values for multiple turbine engine components.3. The method of wherein the subset of the plurality of film holes comprises all of the plurality of film holes.4. The method of wherein the first surface is an exterior bounding an interior defined by the second surface claim 1 , wherein the interior comprises multiple cooling air circuits and the plurality of film holes comprises different groups of film holes claim 1 , with each grouping ...

GAS TURBINE ENGINE HIGH LIFT AIRFOIL COOLING IN STAGNATION ZONE

An airfoil for a gas turbine engine includes pressure and suction side walls joined to one another at leading and trailing edges. A stagnation line is located on the pressure side wall aft of the leading edge. A cooling passage is provided between the pressure and suction side walls. Forward-facing cooling holes are provided adjacent to the stagnation line on the pressure side wall and oriented toward the leading edge. 1. An airfoil for a gas turbine engine , comprising:pressure and suction side walls joined to one another at leading and trailing edges, a stagnation line located on the pressure side wall aft of the leading edge, a cooling passage provided between the pressure and suction side walls, and forward-facing cooling holes provided adjacent to the stagnation line on the pressure side wall and oriented toward the leading edge.2. The airfoil according to claim 1 , comprising shower head cooling holes clustered about the leading edge claim 1 , the forward-facing cooling holes spaced aft of the shower head cooling holes.3. The airfoil according to claim 2 , wherein the shower head cooling holes include a cluster of three rows of holes extending in a radial direction claim 2 , the three rows including a first row extending along the leading edge and second and third rows respectively arranged adjacent to and on opposing sides of the first row.4. The airfoil according to claim 1 , wherein the forward-facing cooling holes extend from a midspan of the airfoil to a tip.5. The airfoil according to claim 1 , comprising aft-facing cooling holes provided adjacent to the stagnation line on the suction side wall and oriented toward the trailing edge.6. The airfoil according to claim 5 , wherein the aft-facing cooling holes are aft of the stagnation line and oriented toward a tip of the airfoil.7. The airfoil according to claim 1 , wherein the stagnation line overlaps the leading edge.8. The airfoil according to claim 1 , wherein the forward-facing cooling holes are ...

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.

COOLING HOLE FOR A GAS TURBINE ENGINE COMPONENT

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a wall having an internal surface and an outer skin, a cooling hole having an inlet extending from the internal surface and merging into a metering section, and a diffusion section downstream of the metering section that extends to an outlet located at the outer skin. The diffusion section of the cooling hole includes a first side diffusion angle, a second side diffusion angle and a downstream diffusion angle at a downstream surface of the diffusion section, the downstream diffusion angle being less than the first side diffusion angle and the second side diffusion angle. 1. A component for a gas turbine engine , comprising:a wall having an internal surface and an outer skin;a cooling hole having an inlet extending from said internal surface and merging into a metering section, and a diffusion section downstream of said metering section that extends to an outlet located at said outer skin; andwherein said diffusion section of said cooling hole includes a first side diffusion angle, a second side diffusion angle and a downstream diffusion angle at a downstream surface of said diffusion section, said downstream diffusion angle being less than said first side diffusion angle and said second side diffusion angle.2. The component as recited in claim 1 , wherein said wall is part of a vane.3. The component as recited in claim 1 , wherein said wall is part of a blade.4. The component as recited in claim 1 , wherein said wall is part of a blade outer air seal (BOAS).5. The component as recited in claim 1 , wherein said diffusion section includes a first side surface that diverges in a first axial direction from an axis of said metering section and a second side surface that diverges in a second axial direction from said axis.6. The component as recited in claim 5 , wherein said first side surface and said second side surface diverge at said first and ...

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 ...

MANUFACTURING METHODS FOR MULTI-LOBED COOLING HOLES

A method for producing a diffusion cooling hole extending between a wall having a first wall surface and a second wall surface includes forming a cooling hole inlet at the first wall surface, forming a cooling hole outlet at the second wall surface, forming a metering section downstream from the inlet and forming a multi-lobed diffusing section between the metering section and the outlet. The inlet, outlet, metering section and multi-lobed diffusing section are formed by laser drilling, particle beam machining, fluid jet guided laser machining, mechanical machining, masking and combinations thereof. 1. A method for producing a diffusion cooling hole extending between a wall having a first wall surface and a second wall surface , the method comprising:forming a cooling hole inlet at the first wall surface;forming a cooling hole outlet at the second wall surface;forming a metering section downstream from the inlet; andforming a multi-lobed diffusing section between the metering section and the outlet,wherein the inlet, outlet, metering section and multi-lobed diffusing section are formed by a technique selected from the group consisting of laser drilling, particle beam machining, fluid jet guided laser machining, mechanical machining, masking and combinations thereof.2. The method of claim 1 , wherein the wall comprises a metal or superalloy substrate.3. The method of claim 1 , wherein the second wall surface comprises a coating claim 1 , and wherein at least a portion of the cooling hole extends through the coating.4. The method of claim 3 , wherein the coating comprises:a bond coating; anda thermal barrier coating.5. The method of claim 4 , wherein a portion of the diffusing section is located within the coating.6. The method of claim 5 , wherein the entire diffusing section is located within the coating.7. The method of claim 6 , wherein a portion of the metering section is located within the coating.8. The method of claim 1 , wherein the inlet and metering section ...

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

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

POROUS SPACE FILLERS FOR CERAMIC MATRIX COMPOSITES

A space filler for forming a fibrous preform may comprise an additively manufactured ceramic material. The additively manufactured ceramic material may define a plurality of pores. A shape of the additively manufactured ceramic material may complement a shape of a void formed by fibrous regions of the fibrous preform. 1. A space filler for forming a fibrous preform , the space filler , comprising:a ceramic material defining a plurality of pores, wherein the ceramic material is formed by additive manufacturing,wherein a shape of the ceramic material complements a shape of a void formed in a fibrous region of the fibrous preform.2. The space filler of claim 1 , wherein a first volume of the ceramic material at an end of the space filler is less than a second volume of the ceramic material at an interior section of the space filler.3. The space filler of claim 2 , wherein the interior section of the space filler corresponds to an interior portion of the fibrous preform.4. The space filler of claim 1 , wherein the plurality of pores form a plurality of interconnected channels through the ceramic material.5. The space filler of claim 1 , wherein the ceramic material forms a generally triangular prism.6. The space filler of claim 1 , further comprising a first end having a first height claim 1 , and a second end having a second height less than the first height.7. The space filler of claim 1 , further comprising an interior section having a first height that is less than a second height of an end of the space filler.8. The space filler of claim 1 , further comprising a concave surface.9. A ceramic matrix composite component for a gas turbine engine claim 1 , comprising:a first fibrous region;a second fibrous region adjacent to the first fibrous region; anda space filler disposed in a void defined, at least partially, by the first fibrous region and the second fibrous region, wherein the space filler comprises a ceramic material formed by additive manufacturing, the ...

COOLED GAS TURBINE ENGINE COMPONENT

A gas turbine component having a cooling passage is disclosed. In one form, the passage is oriented as a turned passage capable of reversing direction of flow, such as a turned cooling hole. The gas turbine engine component can include a layered structure having cooling flow throughout a region of the component. The cooling hole can be in communication with a space in the layered structure. The gas turbine engine component can be a cast article where a mold can be constructed to produce the cooling hole having a turn. 1. An apparatus comprising:a cooled gas turbine engine component having a wall forming a boundary of an internal passage used for conveyance of a cooling fluid; anda cooling hole extending between a hot-side and a cold-side of the cooled gas turbine engine component having a first end oriented to receive cooling fluid from the internal passage and a second end having an outlet capable of discharging the cooling fluid from the gas turbine engine component, the cooling hole having opposing sides routed along a curvilinear path.2. The apparatus of claim 1 , wherein the cooled gas turbine engine component is a multi-walled cooled component claim 1 , and wherein the internal passage is situated between a hot-side wall and a cold-side wall of an inter-wall passage.3. The apparatus of claim 2 , wherein the curvilinear path of the cooling hole is near a leading edge of the multi-wall cooled component.4. The apparatus of claim 3 , wherein the inter-wall passage includes a plurality of pedestals claim 3 , and wherein the cooling hole is substantially free of pedestals.5. The apparatus of claim 1 , wherein the cooled gas turbine engine component includes a construction to permit transpiration cooling claim 1 , and wherein the cooling hole includes a plurality of cooling holes in flow communication with a transpiration cooling passage.6. The apparatus of claim 5 , wherein the plurality of cooling holes include outlets in a leading edge region of the cooled gas ...

COOLED TURBINE GUIDE VANE OR BLADE FOR A TURBOMACHINE

A turbine airfoil for a turbomachine is provided. The airfoil includes a suction side wall and a pressure side wall bordering an airfoil cavity, which receives a cooling fluid for cooling the airfoil. The suction side wall includes one or more protrusions extending inside the cavity. The number of protrusions on the suction side may be higher than the number of protrusions on the pressure side. The density of protrusions on the suction side may be higher than the density of protrusions on the pressure side and/or the surface of protrusions on the suction side may be larger than the surface of protrusions on the pressure side, so that heat transfer from the suction side to the cooling fluid is higher compared to heat transfer from the pressure side to the cooling fluid during operation of the turbomachine. 1. A turbine airfoil comprising a blade or a vane for a turbomachine , the airfoil comprisinga suction side wall and a pressure side wall bordering an airfoil cavity, which is adapted to be flowed through by a cooling fluid for cooling of the side walls and therefore of the airfoil,wherein the suction side wall comprises at least one protrusion extending therefrom inside the airfoil cavity, wherein the number of the at least one protrusion on the suction side wall is higher than the number of protrusions on the pressure side wall, the density of the at least one protrusion on the suction side wall is higher than the density of protrusions on the pressure side wall and/or the surface of the at least one protrusion on the suction side wall is larger than the surface of protrusions on the pressure side wall, so that the heat transfer from the suction side wall to the cooling fluid is higher compared to the heat transfer from the pressure side wall to the cooling fluid during the operation of the turbomachine such that an excess of the heat transfer from the suction side wall is generated.2. The turbine airfoil according to claim 1 , wherein at least one of the ...

TURBINE COMPONENT AND METHODS OF ASSEMBLING THE SAME

A turbine component is provided. The turbine component includes an airfoil having a first surface and a second surface. A thermal barrier coating is coupled to the second surface, wherein the thermal barrier coating includes a first portion, a second portion and a trench defined between the first and second portions. A channel is coupled in flow communication to the first surface and the trench, wherein the channel includes a first sidewall and a second sidewall opposite of the first sidewall. The first and second sidewalls extend from the first surface and toward the trench at an angle. The turbine component includes a cover coupled to the second surface, wherein the cover includes a first end coupled to the first portion and a second end extending into the trench and spaced from the second portion. 1. A turbine component comprising:an airfoil comprising a first surface and a second surface;a thermal barrier coating coupled to said second surface and comprising a first portion, a second portion and a trench defined between said first and second portions;a channel coupled in flow communication to said first surface and said trench, said film channel comprising a first sidewall and a second sidewall opposite of said first sidewall, said first and second sidewalls extending from said first surface and toward said trench at an angle; anda cover coupled to said second surface and comprising a first end coupled to said first portion and a second end extending into said trench and spaced from said second portion.2. The turbine component of claim 1 , wherein said first portion comprises a straight end and said second portion comprises an angled end.3. The turbine component of claim 1 , wherein said first portion comprises a straight end and said second portion comprises an angled end that is substantially aligned with said second sidewall.490. The turbine component of claim 3 , wherein said angle of said angled end and second sidewall is less than about degrees.5. The ...

GAS TURBINE ENGINE AIRFOIL COOLING CIRCUIT

An airfoil for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil body and a cooling circuit disposed inside the airfoil body and including a leading edge cavity with a first portion extending from a radially inner wall to a radially outer wall of the airfoil body and a second portion that extends from a leading edge inner wall to a trailing edge inner wall of the airfoil body. The cooling circuit is configured to communicate cooling airflow through the first portion and the second portion prior to exiting the leading edge cavity into a second cavity of the cooling circuit.

Airfoil platform impingement cooling holes

An airfoil structure for a gas turbine engine includes an airfoil which includes a leading edge and a trailing edge. A platform is located adjacent a first end of the airfoil and includes a core passage that extends through the platform, a mate-face for engaging an adjacent airfoil structure and a set of impingement cooling holes in communication with the core passage that extend through the mate-face adjacent the trialing edge of the airfoil.

GAS TURBINE ENGINE COMPONENT HAVING SHAPED PEDESTALS

A component according to an exemplary aspect of the present disclosure includes, among other things, a first wall, a second wall and at least one row of shaped pedestals extending between the first wall and the second wall. The at least one row of shaped pedestals includes a first set of C-shaped pedestals and a second set of C-shaped pedestals adjacent to the first set of C-shaped pedestals.

HEAT TRANSFER DEVICE AND RELATED TURBINE AIRFOIL

Various embodiments include a heat transfer device, while other embodiments include a turbine component. In some cases, the device can include: a body portion having an inner surface and an outer surface, the inner surface defining an inner region; at least one aperture in the body portion, the at least one aperture positioned to direct fluid from the inner region through the body portion; and at least one fluid receiving feature formed in the outer surface of the body portion, the at least one fluid receiving feature positioned to receive post-impingement fluid from the at least one aperture, wherein the at least one aperture does not define any portion of the at least one fluid receiving feature, and the at least one fluid receiving feature segregates relatively higher velocity post-impingement fluid from relatively lower velocity fluid within an impingement cross-flow region. 1. A device , comprising:a body portion having an inner surface and an outer surface, the inner surface defining an inner region;at least one aperture in the body portion, the at least one aperture positioned to direct fluid from the inner region through the body portion; andat least one fluid receiving feature formed in the outer surface of the body portion, the at least one fluid receiving feature positioned to receive post-impingement fluid from the at least one aperture;wherein the at least one aperture does not define any portion of the at least one fluid receiving feature, and the at least one fluid receiving feature segregates relatively higher velocity post-impingement fluid from relatively lower velocity fluid within an impingement cross-flow region.2. The device of claim 1 , wherein the at least one fluid receiving feature further comprises a fluid directing feature.3. The device of claim 2 , wherein the fluid directing feature is formed within the fluid receiving feature.4. The device of claim 2 , wherein the fluid directing feature comprises a turning vane positioned within an ...

AIRFOIL WITH TIP RAIL COOLING

An apparatus and method for cooling an airfoil tip for a turbine engine can include a blade, such as a cooled turbine blade, having a tip rail extending beyond a tip wall () enclosing an interior for the airfoil at the tip. A plurality of film-holes can be provided in the tip rail. A flow of cooling fluid can be provided through the film-holes from the interior of the airfoil to cool the tip of the airfoil. 1. An airfoil for a turbine engine , the airfoil comprising:a body defining an interior, and extending axially between a leading edge and a trailing edge to define a chord-wise direction and radially between a root and a tip to define a span-wise direction, which terminates in a tip wall and a tip rail extending from the tip wall;at least one cooling passage formed in the interior;at least one cooling cavity provided within the tip rail and comprising at least one cooling conduit defining a flow path having a centerline intersecting with a first surface of the cooling cavity and fluidly coupled to the cooling passage; andat least one film-hole non-aligned in the chord-wise direction with the at least one cooling conduit having an inlet fluidly coupled to the at least one cooling cavity at a second surface opposite the first surface and an outlet provided on an exterior surface of the tip rail.2. The airfoil of wherein the at least one cooling cavity comprises multiple cooling cavities.3. The airfoil of wherein the at least one cooling conduit comprises multiple cooling conduits.4. The airfoil of wherein the at least one cooling conduit comprises a curved cooling conduit.5. The airfoil of wherein the at least one cooling conduit comprises multiple cooling conduits.6. The airfoil of wherein the at least one cooling conduit comprises multiple cooling conduits.7. The airfoil of further comprising a plurality of film-holes provided along a distal end of the tip rail.8. The airfoil of wherein the exterior surface comprises an outer wall and the outlet is fluidly ...

Gas turbine engine component cooling circuit

A method of manufacturing a component for a gas turbine engine according to an exemplary aspect of the present disclosure includes forming the component with a first manufacturing technique to include a first cavity and a second cavity, and a microcircuit in fluid communication with the first cavity, the component including an outer wall and a ribbed portion or a bulged portion extending from the outer wall. The exemplary method includes forming a plunged hole with a second manufacturing technique different from the first manufacturing technique to intersect at least a portion of the microcircuit and extend into the ribbed portion or the bulged portion.

SYSTEM AND METHOD OF FABRICATING AND REPAIRING A GAS TURBINE COMPONENT

A method of fabricating and repairing a gas turbine component having a plurality of cooling holes defined therein is provided. The method includes determining a parameter of a first cooling hole defined in the gas turbine component, and generating a tool path for forming a protective cap around the first cooling hole. The tool path is based at least partially on the parameter of the first cooling hole. The method also includes directing a robotic device to follow the tool path, and discharging successive layers of ceramic slurry towards the gas turbine component as the tool path is followed such that the protective cap is formed around the first cooling hole. 1. A method of fabricating and repairing a gas turbine component having a plurality of cooling holes defined therein , said method comprising:determining a parameter of a first cooling hole defined in the gas turbine component;generating a tool path for forming a protective cap around the first cooling hole, the tool path based at least partially on the parameter of the first cooling hole;directing a robotic device to follow the tool path; anddischarging successive layers of ceramic slurry towards the gas turbine component as the tool path is followed such that the protective cap is formed around the first cooling hole.2. The method in accordance with claim 1 , wherein determining a parameter comprises determining at least one of a size of the first cooling hole claim 1 , an edge profile of the first cooling hole claim 1 , or a location of the first cooling hole on the gas turbine component.3. The method in accordance with claim 1 , wherein determining a parameter comprises conducting a non-destructive inspection of the gas turbine component.4. The method in accordance with claim 1 , wherein generating a tool path comprises determining an arrangement of a plurality of individual layers for forming a three-dimensional representation of the protective cap claim 1 , wherein the plurality of individual layers ...

COOLING PASSAGES FOR GAS TURBINE ENGINE COMPONENT

A gas turbine engine component includes a wall that has an inner surface and an outer surface. An inlet is defined by the inner surface. At least one non-rectangular slot is defined by the outer surface and includes at least one protrusion extending into the slot. A slot passage fluidly connects the inlet to the at least one non-rectangular slot. The slot passage comprises an inlet portion that extends through the wall from the inlet to an intermediate portion. An outlet portion extends through the wall from the intermediate portion to the at least one non-rectangular slot. 1. A gas turbine engine component comprising:a wall having an inner surface and an outer surface;at inlet defined by the inner surface;at least one non-rectangular slot defined by the outer surface and including at least one protrusion extending into the slot;a slot passage fluidly connecting the inlet to the at least one non-rectangular slot, the slot passage comprising:an inlet portion extending through the wall from the inlet to an intermediate portion; andan outlet portion extending through the wall from the intermediate portion to the at least one non-rectangular slot.2. The gas turbine engine component of claim 1 , wherein that at least one protrusion extends from a downstream side of the at least one non-rectangular slot.3. The gas turbine engine component of claim 1 , wherein that at least one protrusion extends from an upstream side of the at least one non-rectangular slot.4. The gas turbine engine component of claim 1 , wherein the at least one protrusion includes a plurality of protrusions extending into the at least one non-rectangular slot from at least one of a downstream side and an upstream side.5. The gas turbine engine component of claim 1 , wherein the at least one non-rectangular slot includes a plurality of non-rectangular slots spaced radially from each other and each of the plurality of non-rectangular slots includes a corresponding slot passage.6. The gas turbine engine ...

COOLING STRUCTURE FOR TURBINE AIRFOIL

A structure for cooling a turbine airfoil includes: a cooling passage formed between a first airfoil wall curved as to be concave relative to a high-temperature gas passage and a second airfoil wall curved so as to be convex relative to the high-temperature gas passage; lattice structure bodies each formed by stacking a plurality of ribs in a lattice pattern; a partition body provided between the adjacent lattice structure bodies; a cooling medium discharge port for discharging a cooling medium within the cooling passage to the outside; and an exposed wall portion formed as a portion of the second airfoil wall extending beyond the cooling medium discharge port to the outside. At outlet portions of the lattice structure bodies adjacent to each other across the partition body, adjacent first rib sets and second rib sets are inclined in opposite directions relative to the partition body, respectively. 1. A cooling structure for cooling a turbine airfoil of a turbine driven by high-temperature gas , the cooling structure comprising:a cooling passage formed between a first airfoil wall of the turbine airfoil that is curved so as to be concave relative to a passage for the high-temperature gas and a second airfoil wall of the turbine airfoil that is curved so as to be convex relative to the passage for the high-temperature gas;a plurality of lattice structure bodies each including a first rib set composed of a plurality of first ribs extending linearly and provided on a wall surface of the first airfoil wall that faces the cooling passage, and a second rib set composed of a plurality of second ribs extending linearly and provided on a wall surface of the second wall that faces the cooling passage, the second rib set being stacked on the first rib set so as to form a lattice pattern;a partition body provided between the adjacent two lattice structure bodies and configured to close a passage formed in each rib set;a cooling medium discharge port provided at a downstream end ...

Turbine vane with dust tolerant cooling system

A turbine vane includes an airfoil that extends from an inner diameter to an outer diameter, and from a leading edge to a trailing edge. The turbine vane includes an inner platform coupled to the airfoil at the inner diameter. The turbine vane includes a cooling system defined in the airfoil including a first conduit in proximity to the leading edge to cool the leading edge and a second conduit to cool the trailing edge. The first conduit has an inlet at the outer diameter to receive a cooling fluid and an outlet portion that is defined at least partially through the inner platform. The first conduit includes a plurality of cooling features that extend between a first surface and a second surface of the first conduit, and the first surface of the first conduit is opposite the leading edge.

MODULATED TURBINE COMPONENT COOLING

Features and methods for modulating a flow of cooling fluid to gas turbine engine components are provided. In one embodiment, an airfoil is provided having a flow modulation insert for modulating a flow of cooling fluid received in a cavity of a body of the airfoil. In another embodiment, a shroud is provided comprising a cooling channel for a flow of cooling fluid and an insert that varies in position to modulate the flow of cooling fluid through the cooling channel. In yet another embodiment, a method for operating a gas turbine engine having a cooling circuit for cooling one or more components of the gas turbine engine comprises increasing power provided to the engine and decreasing power provided to the engine to modulate a position of a flow modulation insert located in the cooling circuit and thereby modulate the flow of cooling fluid through the cooling circuit. 119.-. (canceled)20. A method for operating a gas turbine engine , the gas turbine engine including a cooling circuit for providing a flow of cooling fluid to one or more components of the gas turbine engine , the method comprising:increasing power provided to the gas turbine engine; anddecreasing power provided to the gas turbine engine,wherein increasing and decreasing the power provided to the gas turbine engine modulates a position of a flow modulation insert located in the cooling circuit to modulate the flow of cooling fluid through the cooling circuit.21. The method of claim 20 , wherein the flow modulation insert deflects toward an open position when increasing power provided to the gas turbine engine claim 20 , and wherein the flow modulation insert deflects toward a closed position when decreasing power provided to the gas turbine engine.22. The method of claim 20 , wherein the cooling circuit comprises a shroud positioned radially adjacent a plurality of turbine rotor blades claim 20 , and wherein the shroud comprises:a cooling channel for a flow of cooling fluid, the cooling channel having ...

Ceramic and refractory metal core assembly

A core assembly for forming a cast component includes a refractory metal core and a ceramic core element. The refractory metal core includes first and second ends and sides extending from the first end to the second end. The ceramic core element includes a slot positioned between first and second lands, each land having an inner surface facing the slot and an adjacent outer surface. The first end of the refractory metal core is secured within the slot with an adhesive, and the refractory metal core extends from the ceramic core element in both a longitudinal and a transverse direction. The slot, lands, and refractory metal core form a core assembly providing access paths to the sides of the refractory metal core. Surplus adhesive is removed from the refractory metal core via the access paths. Investment casting provides the component with an internal passage and an internal cooling circuit.

LOW PRESSURE LOSS COOLED BLADE

A rotor blade comprises a root section, an airfoil section, a leading edge cooling cavity, an intermediate cooling cavity, and a trailing edge cooling cavity. The leading edge, intermediate, and trailing edge cooling cavities each extend spanwise through the airfoil section from a coolant inlet passage in the root section, and each terminate proximate the airfoil tip. 1. A rotor blade comprising:a root section including a coolant inlet passage;an airfoil section including a suction sidewall and a pressure sidewall each extending chordwise between a leading edge and a trailing edge, and extending spanwise between the root section and an airfoil tip;a leading edge cooling cavity extending spanwise through the airfoil section from the coolant inlet passage and terminating proximate the airfoil tip;an intermediate cooling cavity disposed aft of the leading edge cavity and extending spanwise through the airfoil section from the coolant inlet passage and terminating proximate the airfoil tip; anda trailing edge cooling cavity disposed aft of the intermediate cavity and extending spanwise through the airfoil section from the coolant inlet passage, and terminating proximate the airfoil tip.2. The rotor blade of claim 1 , further comprising:a first rib disposed between the leading edge cooling cavity and the intermediate cooling cavity, the first rib extending spanwise through substantially all of the airfoil section between the root section and the airfoil tip.3. The rotor blade of claim 2 , wherein the first rib includes at least one crossover hole disposed proximate the airfoil tip.4. The rotor blade of claim 2 , wherein a portion of the leading edge cooling cavity in the airfoil section is bounded by the airfoil leading edge claim 2 , the first rib claim 2 , and at least one of the suction sidewall and the pressure sidewall.5. The rotor blade of claim 2 , further comprising:a second rib disposed between the intermediate cooling cavity and the trailing edge cooling cavity ...

MULTI-LOBED COOLING HOLE

A gas turbine engine component subjected to a flow of high temperature gas includes a wall having first and second surfaces and a cooling hole extending through the wall. The cooling hole includes an inlet located at the first surface, an outlet located at the second surface, a metering section extending downstream from the inlet, and a diffusing section extending from the metering section to the outlet. The diffusing section includes a first lobe diverging longitudinally and laterally from the metering section and having a first downstream end adjacent the outlet and spaced from the inlet by a first distance, a second lobe diverging longitudinally from the metering section and having a second downstream end adjacent the outlet and spaced from the inlet by a second distance different from the first, and a transition region positioned between the lobes, the transition region having a third downstream end adjacent the outlet. 1. A gas turbine engine component subjected to a flow of high temperature gas , the component comprising:a wall having first and second wall surfaces; and an inlet located at the first wall surface;', 'an outlet located at the second wall surface;', 'a metering section extending downstream from the inlet; and', a first lobe diverging longitudinally and laterally from the metering section and having a first downstream end adjacent the outlet and spaced from the inlet by a first distance;', 'a second lobe diverging longitudinally from the metering section and having a second downstream end adjacent the outlet and spaced from the inlet by a second distance different from the first distance; and', 'a first transition region positioned between the first and second lobes, the first transition region comprising a third downstream end adjacent the outlet., 'a diffusing section extending from the metering section to the outlet and comprising], 'a cooling hole extending through the wall and comprising2. The component of claim 1 , wherein the metering ...

TURBINE ROTOR BLADE WITH COUPON HAVING CORRUGATED SURFACE(S)

A turbine rotor blade includes an airfoil body having a leading edge, a trailing edge and a smooth outer surface. A cutout is included within at least one of the leading edge and the trailing edge, the cutout removing a predetermined area of the airfoil body. A coupon is coupled in the cutout to replace the predetermined area of the airfoil body. The coupon includes a first corrugated surface on at least a portion of an outer surface thereof. The coupon allows for the addition of advantageous wake mixing and cooling efficiencies to preexisting blades. 1. A turbine rotor blade comprising:an airfoil body having a leading edge, a trailing edge and a smooth outer surface;a cutout within at least one of the leading edge and the trailing edge, the cutout removing a predetermined area of the airfoil body; anda coupon coupled in the cutout to replace the predetermined area of the airfoil body, the coupon including a first corrugated surface on at least a portion of an outer surface thereof.2. The turbine rotor blade of claim 1 , wherein the airfoil body includes an airfoil coolant flow passage therein claim 1 , the cutout exposing the airfoil coolant flow passage claim 1 , andwherein the coupon includes a first coupon coolant flow passage configured to fluidly mate with the airfoil coolant flow passage.3. The turbine rotor blade of claim 2 , wherein the coupon includes a plurality of flow passages extending from an interior surface thereof to the outer surface thereof.4. The turbine rotor blade of claim 2 , wherein the cutout is positioned in the trailing edge claim 2 , and the coupon includes a pin bank within the first coupon coolant flow passage.5. The turbine rotor blade of claim 2 , wherein the airfoil body includes an airfoil radially extending chamber therein and an airfoil rib partitioning the airfoil radially extending chamber to form the airfoil coolant flow passage therein claim 2 , the cutout exposing the airfoil rib and the airfoil coolant flow passage claim 2 ...

BLADE WITH PARALLEL CORRUGATED SURFACES ON INNER AND OUTER SURFACES

A blade includes an airfoil body defined by a concave pressure side outer wall and a convex suction side outer wall that connect along leading and trailing edges and, therebetween, form a radially extending chamber for receiving the flow of a coolant. The airfoil body has an inner surface facing the radially extending chamber and an outer surface, a first corrugated surface on a portion of the outer surface, and a second corrugated surface on the inner surface paralleling the first corrugated surface. The corrugated surface on the outer surface of the airfoil provides wake mixing. The blade may also include an integrally formed impingement cooling structure having a third corrugated surface parallel to the second corrugated surface, which is made possible through additive manufacturing. The impingement cooling structure so formed provides improved cooling of the blade. 1. A blade comprising:an airfoil body defined by a concave pressure side outer wall and a convex suction side outer wall that connect along leading and trailing edges and, therebetween, form a radially extending chamber for receiving the flow of a coolant, the airfoil body having an outer surface and an inner surface facing the radially extending chamber;a first corrugated surface on a portion of the outer surface;a second corrugated surface on the inner surface paralleling the first corrugated surface; andan impingement cooling structure positioned within the radially extending chamber, the impingement cooling structure including a portion of an outer surface thereof having a third corrugated surface paralleling the second corrugated surface on the inner surface of the airfoil body.2. The blade of claim 1 , further comprising a plurality of internal supports positioning the impingement cooling structure relative to the radially extending chamber claim 1 , and wherein the airfoil body and the impingement sleeve include a plurality of integral material layers that also include the plurality of internal ...

FORMING COOLING PASSAGES IN THERMAL BARRIER COATED, COMBUSTION TURBINE SUPERALLOY COMPONENTS

Delamination of thermal barrier coatings (“TBC's”) () from superalloy substrates () of components () for turbine engines (), such as engine blades (), vanes (), or castings in transitions (), is inhibited during subsequent cooling passage () formation. Partially completed cooling passages (), which have skewed passage paths that end at a terminus (), which is laterally offset from the passage entrance (), are formed in the superalloy component () prior to application of the TBC layer(s) (). The skewed, laterally offset path of each partially completed cooling passage () establishes an overhanging shield layer () of superalloy material that protects the TBC layer () during completion of the cooling passage (). 1. A method for forming a cooling passage in a thermal barrier coated , superalloy component for a combustion turbine engine , the cooling passage having a passage path , including an inlet and an outlet , comprising:forming a partially completed cooling passage in a surface of a superalloy component for a combustion turbine engine, the partially completed cooling passage having an entrance formed in the surface, corresponding to a cooling passage inlet or outlet, and a skewed passage path within the superalloy component having a terminus distal and laterally offset from the entrance in the surface, and an overhanging shield layer of superalloy material interposed between the terminus and the surface proximate the entrance;applying a thermal barrier coating over the surface and the entrance;providing an ablation device for ablating thermal barrier coating and superalloy material;aligning the ablation device proximate the entrance;ablating thermal barrier coating material from the partially completed cooling passage with the ablation device, reaching the terminus; andcompleting the cooling passage by ablating superalloy material out of the skewed passage path along a the partially completed cooling passage from the terminus to the other of the cooling passage ...

GAS TURBINE

A gas turbine denitrifies combustion gas by using a denitrification catalyst and ammonia as a reducing agent, the gas turbine includes: a turbine provided with turbine blades, the turbine blades being exposed to the combustion gas reaching a temperature higher than an average value in a temperature distribution of the combustion gas, and a compressor configured to supply the turbine blades with a cooling air and the ammonia, wherein the gas turbine is configured to lower the temperature of the turbine blades by supplying the turbine blades with the ammonia and the cooling air. 1. A gas turbine denitrifying combustion gas by using a denitrification catalyst and ammonia as a reducing agent , the gas turbine comprising:a turbine provided with turbine blades, the turbine blades being exposed to combustion gas having a temperature higher than an average value in a temperature distribution of the combustion gas, anda compressor configured to supply the turbine blades with a cooling air and the ammonia, whereinthe gas turbine is configured to lower the temperature of the turbine blades by supplying the turbine blades with the ammonia and the cooling air supplied from the compressor.2. The gas turbine according to claim 1 , wherein the gas turbine is configured such that the ammonia supplied to the turbine blades is supplied to a combustor as a fuel after cooling the turbine blades.3. The gas turbine according to claim 1 , wherein the gas turbine is configured such that the ammonia supplied to the turbine blades is injected into a combustion gas flow path after cooling the turbine blades and is discharged from the turbine after mixing with the combustion gas.4. The gas turbine according to claim 1 , wherein the gas turbine is configured such that the ammonia is supplied to the turbine blades in a liquid state.5. The gas turbine according to claim 1 , wherein the gas turbine is configured such that the ammonia is mixed in advance with the cooling air and is supplied to the ...

Gas turbine engines and methods associated therewith

A method of forming a gas turbine engine component, the method including forming a plurality of cooling apertures in a preform structure of the component, the plurality of cooling apertures of the preform structure comprising a first cooling aperture and a second cooling aperture, wherein cross-sectional shapes of the first and second cooling apertures of the preform structure are different from one another, as measured in a same relative plane; and applying a coating to at least a portion of the preform structure to form the component, wherein a cross-sectional shape of the first and second cooling apertures of the component are approximately the same as one another, as measured in the same relative plane.

COMPONENT CORE WITH SHAPED EDGES

A gas turbine engine according to an example of the present disclosure includes, among other things, a plurality of blade outer air seals and a plurality of airfoils, at least one of seals and airfoils including at least one cooling passage. The cooling passage includes a first wall and an opposed second wall bounding the cooling passage, a surface contour of the first wall having a plurality of first surface features and a surface contour of the second wall having a plurality of second surface features. The first surface features and the second surface features are arranged such that a width of the cooling passage varies along a length of the cooling passage defined by the first surface features and the second surface features. The first surface features have a first profile, and the second surface features have a second, different profile. A casting core for forming cooling passages in an aircraft component is also disclosed. 1. A gas turbine engine , comprising:a plurality of blade outer air seals each arranged circumferentially about an axis to define a flow path;a plurality of airfoils spaced from said plurality of blade outer air seals, each of said plurality of airfoils including an airfoil section extending from a platform; a first wall and an opposed second wall bounding said at least one cooling passage, a surface contour of said first wall having a plurality of first surface features and a surface contour of said second wall having a plurality of second surface features;', 'wherein said plurality of first surface features and said plurality of second surface features are arranged such that a width of said cooling passage varies along a length of said cooling passage defined by said plurality of first surface features and said plurality of second surface features, said plurality of first surface features having a first profile, and said plurality of second surface features having a second, different profile; and', 'wherein said plurality of first surface ...

COOLING HOLE FOR A GAS TURBINE ENGINE COMPONENT

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a wall having an internal surface, an outer skin and a cooling hole having an inlet extending from the internal surface and merging into a metering section, and a diffusion section downstream of the metering section that extends to an outlet located at the outer skin. 1. A component for a gas turbine engine , comprising:a wall having an internal surface and an outer skin; anda cooling hole having an inlet extending from said internal surface and merging into a metering section, and a diffusion section downstream of said metering section that extends to an outlet located at said outer skin;wherein the outlet includes a leading edge and a trailing edge, and the leading edge is downstream of a downstream end of the metering section.2. The component as recited in claim 1 , wherein the trailing edge is a straight edge.3. The component as recited in claim 1 , wherein the metering section has a substantially constant flow area from the inlet to the diffusion section.4. The component as recited in claim 1 , wherein the diffusion section includes a first lateral edge including a first curvature.5. The component as recited in claim 4 , wherein the diffusion section includes a second lateral edge laterally opposite the first lateral edge and including a second curvature.6. The component as recited in claim 5 , wherein the trailing edge is a straight edge.7. The component as recited in claim 6 , wherein the metering section extends along a central axis claim 6 , and the trailing edge defines a right angle relative to the central axis.8. The component as recited in claim 7 , wherein the diffusion section includes a raised transition portion disposed between the first lateral edge and the second lateral edge.9. The component as recited in claim 7 , wherein the metering section has a substantially constant flow area from the inlet to the diffusion section. ...

AIRFOIL COOLING CIRCUITS

An airfoil includes leading and trailing edges; first and second sides extending from the leading edge to the trailing edge, each side having an exterior surface; a core passage located between the first and second sides and the leading and trailing edges; and a wall structure located between the core passage and the exterior surface of the first side. The wall structure includes a plurality of cooling fluid inlets communicating with the core passage for receiving cooling fluid from the core passage, a plurality of cooling fluid outlets on the exterior surface of the first side for expelling cooling fluid and forming a cooling film along the exterior surface of the first side, and a plurality of cooling passages communicating with the plurality of cooling fluid inlets and the plurality of cooling fluid outlets. At least a portion of one cooling passage extends between adjacent cooling fluid outlets. 1. A refractory metal core comprising: a downstream end wall;', 'an upstream end wall opposite the downstream end wall;', 'a first sidewall connecting the downstream end wall to the upstream end wall;', 'a second sidewall connecting the downstream end wall to the upstream end wall, the second sidewall opposite the first sidewall;', 'a plurality of primary curved tabs located between the downstream end wall and the upstream end wall and aligned between the first sidewall and the second sidewall, the plurality of primary curved tabs extending outwardly from the refractory metal sheet in a first direction;', 'a plurality of primary openings in the refractory metal sheet, wherein each one of the plurality of primary openings is proximate to one of the plurality of primary curved tabs;', 'a plurality of secondary curved tabs proximate the downstream endwall, the plurality of secondary curved tabs extending outwardly from the refractory metal sheet; and', 'a plurality of secondary openings in the refractory metal sheet located between the plurality of primary curved tabs and ...

Method of Manufacturing Conductive Film Holes

A method for applying a coating to a substrate having a plurality of holes. The method comprises: applying a braze material to a substrate having a plurality of holes; heating the substrate to melt the braze material to form a melt; cooling the substrate to solidify the melt to form plugs in the respective holes; applying a coating to the substrate; and further heating the substrate to melt the plugs. 1. A method for applying a coating to a substrate having a plurality of holes , the method comprising:applying a braze material to a substrate having a plurality of holes;heating the substrate to melt the braze material to form a melt;cooling the substrate to solidify the melt to form plugs in the respective holes; 'during the applying of the coating, the coating preferentially deposits away from the plugs so that the deposition leaves a gap or thinning in the coating over the exposed plug surface; and', 'applying a coating to the substrate, whereinfurther heating the substrate to melt the plugs.2. The method of the applying the coating comprises:applying a metallic bondcoat; andapplying a ceramic coating atop the bondcoat.3. The method of wherein the applying the braze paste comprises:applying the braze material to portions of a surface of the substrate aside the hole but not plugging the holes.4. The method of wherein:the braze material is an Au—Cu braze material.5. (canceled)6. The method of wherein:the braze material is applied via a braze paste.7. The method of wherein:the braze paste comprises a water-based gel binder.8. The method of wherein:the braze paste comprises particles of an alloy having a by weight composition of Au, Cu, and no more than 5% all other elements total, if any.9. The method of wherein:a ratio of Au to Cu in said alloy is between 67/33 and 56/44.10. The method of wherein:said particles of said alloy form at least 95% of alloy in the braze paste.11. The method of wherein:said water-based gel binder forms 8% to 25% by weight of the braze paste ...

TRAILING EDGE COOLING ARRANGEMENT FOR AN AIRFOIL OF A GAS TURBINE ENGINE

An airfoil () including an internal cooling circuit () to direct cooling fluid () through an interior of the airfoil. The airfoil also includes a trailing edge (′) defining plugged holes () and cooling holes () along a radial direction (). A selective group () of the cooling and plugged holes are based on a reduced cooling fluid requirement at the trailing edge (′) resulting from an improved thermal barrier coating (). A process () is also provided including removing () an airfoil from service in a gas turbine engine and adding or improving () a thermal barrier coating on the airfoil. The process also includes selectively () plugging holes in the airfoil in response to a reduced cooling fluid required through the airfoil as a result of the added or improved thermal barrier coating. 1. An airfoil for a gas turbine engine , the airfoil comprising:an internal cooling circuit to direct cooling fluid through an interior of the airfoil; anda trailing edge comprising a plurality of cooling holes along a radial direction, said cooling holes coupled to the internal cooling circuit; andplugs blocking a selective group of the plurality of holes.2. The airfoil of claim 1 , wherein locations of the plugs are responsive to a cooling requirement at the trailing edge.3. The airfoil of claim 2 , further comprising:a thermal barrier coating on an outer surface of the airfoil;wherein locations of the plugs are responsive to a reduced cooling requirement at the trailing edge resulting from the thermal barrier coating.4. The airfoil of claim 3 , wherein the selective group has a non-uniform distribution in the radial direction along the trailing edge.5. The airfoil of claim 3 , wherein the selective group has a uniform distribution in the radial direction along the trailing edge.6. The airfoil of claim 2 , wherein a mainstream fluid incident on the airfoil has an incident temperature profile in the radial direction claim 2 , and wherein said cooling requirement is responsive to the ...

METHOD FOR COOLING A TURBO-ENGINE COMPONENT AND TURBO-ENGINE COMPONENT

Disclosed is a turbo-engine component and a method for cooling a turbo-engine component. The method includes guiding a working fluid flow along a hot gas side surface of a wall of the component and in a main working fluid flow direction, discharging a coolant discharge flow at the hot gas side surface from a coolant discharge duct provided in the wall, and supplying a coolant supply flow to the coolant discharge duct and through a coolant supply path. The method also includes discharging the coolant supply flow into the coolant discharge duct as a free jet oriented across a cross section of the coolant discharge duct, and directing the free jet onto an inner surface section of the coolant discharge duct, thus effecting impingement cooling of the inner surface section. 1. A method for cooling a turbo-engine component , the method comprising:guiding a working fluid flow along a hot gas side surface of a wall of the component and in a main working fluid flow direction,discharging a coolant discharge flow at the hot gas side surface from a coolant discharge duct provided in the wall,supplying a coolant supply flow to the coolant discharge duct and through a coolant supply path,discharging the coolant supply flow into the coolant discharge duct as a free jet oriented across a cross section of the coolant discharge duct, anddirecting the free jet onto an inner surface section of the coolant discharge duct, thus effecting impingement cooling of the inner surface section.2. The method according to claim 1 , comprising:guiding the coolant supply flow through a means for generating a free jet and discharging the free jet from said means for generating.3. The method according to claim 1 , comprising:discharging the coolant discharge flow in a direction inclined with respect to a normal of the hot gas side surface at: discharge location, whereby the coolant discharge duct is inclined with respect to said normal thus having a first inner surface section disposed towards the hot ...

Article, airfoil component and method for forming article

An article is disclosed including a manifold, an article wall, a post-impingement cavity and a plurality of post-impingement partitions. The manifold includes an impingement wall defining a plenum and a plurality of impingement apertures. The article wall includes a plurality of external apertures. The post-impingement cavity is disposed between the manifold and the article wall, and is arranged to receive a fluid from the plenum through the plurality of impingement apertures and exhaust the fluid through the plurality of external apertures. The plurality of post-impingement partitions divide the post-impingement cavity into a plurality of sub-cavities, and hermetically separate the plurality of sub-cavities from one another. The impingement wall, article wall and plurality of post-impingement partitions are integrally formed as a single, continuous article. The article may be an airfoil component. A method for forming the article includes forming a single, continuous object by an additive manufacturing technique.

OBTUSE ANGLE CHEVRON TRIP STRIP

An airfoil includes a cooling air passage for receiving a cooling air flow. A chevron including a first rib and a second rib extends from a common tip is disposed within the cooling passage for generating a turbulent flow to improve heat transfer. The chevron includes an angle between the first rib and the second rib that is greater than 90 degrees. 1. An airfoil comprising:a cooling air passage receiving a cooling air flow; anda chevron including a first rib and a second rib extending from a common tip, wherein an angle between the first rib and the second rib is greater than ninety (90) degrees.2. The airfoil as recited in claim 1 , wherein the angle between the first rib and the second rib is less than one-hundred-eighty (180) degrees.3. The airfoil as recited in claim 1 , wherein the angle between the first rib and the second rib is between about ninety-five (95) degrees and one-hundred-seventy-five (175) degrees.4. The airfoil as recited in claim 1 , wherein the first rib and second rib include a uniform height above a surface of the cooling channel.5. The airfoil as recited in claim 1 , wherein the first rib and the second rib include an increasing height from the tip towards a first end of the first rib and a second end of the second rib.6. The airfoil as recited in claim 1 , wherein the first rib includes a first end and the second rib includes a second end and a height of the chevron increases from the first end toward the second end.7. The airfoil as recited in claim 1 , wherein the first rib includes a first end and the second rib includes a second end and a height of the chevron decrease from the first end toward the second end.8. The airfoil as recited in claim 1 , wherein the first rib and the second rib are of a common length.9. The airfoil as recited in claim 1 , wherein the first rib and the second rib are of unequal lengths.10. The airfoil as recited in claim 1 , wherein the tip is pointed into cooling air flow.11. The airfoil as recited in claim 1 ...

Gas turbine engine end-wall component

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

COOLING ARRANGEMENT INCLUDING OVERLAPPING DIFFUSERS

A gas turbine engine component according to an example of the present disclosure includes a wall extending in a thickness direction between first and second wall surfaces. The first wall surface bounds an internal cavity. The wall includes a plurality of cooling passages. Each of the cooling passages extend in a first direction between an inlet port and an outlet port coupled to a respective diffuser, and the inlet port coupled to the internal cavity along the first wall surface. Sidewalls of adjacent diffusers are conjoined to establish a common diffuser region interconnecting the diffusers and a common outlet along the second wall surface. A method of cooling a gas turbine engine component is also disclosed. 1. A gas turbine engine component comprising:a wall extending in a thickness direction between first and second wall surfaces, the first wall surface bounding an internal cavity, the wall including a plurality of cooling passages, each of the cooling passages extending in a first direction between an inlet port and an outlet port coupled to a respective diffuser, and the inlet port coupled to the internal cavity along the first wall surface;wherein sidewalls of adjacent diffusers are conjoined to establish a common diffuser region interconnecting the diffusers and a common outlet along the second wall surface; andwherein each of the cooling passages has a minimum cross-sectional area, a total of the minimum cross-sectional area of all of the cooling passages establishes a combined cross-sectional area, the common outlet establishes an outlet cross-sectional area, and an area ratio of the combined cross-sectional area to the outlet cross-sectional area is equal to or greater than 0.15, and the area ratio is less than or equal to 0.40.2. The gas turbine engine component as recited in claim 1 , wherein the plurality of cooling passages includes at least three cooling passages distributed in a second direction perpendicular to the first direction and the thickness ...

AIRFOIL LEADING EDGE COOLING CHANNELS

An airfoil cooling system may be provided. An airfoil may have a pressure side and a suction side that are separated by a leading edge. The leading edge may pass through a stagnation point of the airfoil. A spar may have an outer surface comprising standoffs. The standoffs may define a cooling channel that extends across the leading edge on the outer surface of the spar, from the pressure side to the suction side. The cooling channel may have a first portion and a second portion defined by the standoffs. The first portion may be located closer to a base or a tip of the airfoil than the second portion. The spar may further comprise an inlet on the pressure side or the suction side. The inlet may be configured to convey a cooling fluid from a passageway located inside of the spar to the cooling channel. 1. A blade or a stator for use in a gas turbine engine , the blade or the stator comprising:an airfoil having a first side and a second side that are separated by a leading edge, the leading edge passing through a stagnation point of the airfoil, the airfoil comprising:a cover sheet; anda spar having an outer surface, the spar comprising a plurality of standoffs configured to receive the cover sheet, wherein two of the standoffs define a cooling channel that extends across the leading edge on the outer surface of the spar, from the first side to the second side, the cooling channel having a first portion defined by the two of the standoffs and a second portion defined by the two of the standoffs, the first portion offset from the second portion in a spanwise direction, the spar further comprising an inlet on the first side, the inlet configured to convey a cooling fluid from a passageway located inside of the spar to the cooling channel, wherein the cooling channel is configured to convey the cooling fluid from the inlet, across the leading edge, and toward an outlet on the second side, wherein the outlet is at least partially defined by the cover sheet.2. The airfoil ...

METHOD FOR CREATING A FILM COOLED ARTICLE FOR A GAS TURBINE ENGINE

A method for finishing a film cooled article includes providing a film cooled article including at least one inner cooling plenum and at least one opening connecting the inner cooling plenum to an exterior surface of the film cooled article, positioning a machining element in contact with the exterior surface of the film cooled article, automatically moving the machining element along the exterior surface while maintaining contact between the machining tool and the surface, identifying an actual position of at least one film opening based on sensory feedback from the machining element using a controller, removing material from the exterior surface at the at least one film opening using the machining element, thereby creating a depression at the at least one film opening. 1. A finishing apparatus for a film cooled article comprising:a central control machine including a computerized controller;at least one articulating device controlled by said central control machine;a machining tool mounted to said articulating device, such that said articulating device is operable to move said machining tool;at least one of a touch sensor apparatus and a visual sensor apparatus mounted to said machining tool and communicatively coupled to the computerized controller; andwherein said computerized controller stores instructions operable to cause said finishing apparatus to perform the steps of:positioning a machining element in contact with an exterior surface of a film cooled article;automatically moving said machining element along said exterior surface while maintaining contact between the machining element and the surface;identifying an actual position of at least one film opening based on sensory feedback from the machining element using a controller while maintaining contact between the machining element and the surface;removing material from said exterior surface at said film opening using said machining element, thereby creating a depression at said at least one film opening ...

DIRECTIONAL COOLING ARRANGEMENT FOR AIRFOILS

An airfoil according to an example of the present disclosure includes, among other things, an internal wall and an external wall. The external wall defines pressure and suction sides between a leading edge and a trailing edge, and the airfoil section defines a mean camber line that extends between the leading and trailing edges to bisect a thickness of the airfoil section. A first cavity and a second cavity are separated by the internal wall. The second cavity is bounded by the external wall at the leading edge. At least one crossover passage within the internal wall connects the first cavity to the second cavity. The crossover passage defines a passage axis. The passage axis defines a passage angle with respect to the mean camber line such that the passage axis extends transversely from the mean camber line to intersect a surface of the second cavity. 1. An airfoil for a gas turbine engine , comprising:an airfoil section extending in a spanwise direction from a platform, the airfoil section having an internal wall and an external wall, the external wall defining pressure and suction sides between a leading edge and a trailing edge, and the airfoil section defining a mean camber line extending between the leading and trailing edges to bisect a thickness of the airfoil section;a first cavity and a second cavity separated by the internal wall, the second cavity bounded by the external wall at the leading edge;at least one crossover passage within the internal wall that connects the first cavity to the second cavity; andwherein the at least one crossover passage defines a passage axis, the passage axis defines a passage angle with respect to the mean camber line such that the passage axis extends transversely from the mean camber line to intersect a surface of the second cavity, and the passage angle is between 15 degrees and 90 degrees.2. The airfoil as recited in claim 1 , wherein the airfoil is a turbine airfoil claim 1 , and the passage axis intersects the suction ...

Airfoil with trailing edge rounding

An airfoil for a gas turbine engine includes a substrate portion extending from an airfoil leading edge to an airfoil trailing edge portion. The airfoil trailing edge portion includes a flared portion wherein a substrate portion thickness increases along a camber line of the airfoil, and a trailing edge defined as a full constant radius extending from a pressure side of the airfoil to a suction side of the airfoil. A coating portion includes a coating applied over at least a portion of the substrate portion.

Combustion device and gas turbine engine system

The combustion device includes: a compressor that compresses combustion air; a combustor that combusts the compressed combustion air and fuel ammonia; and an ammonia injector that injects the fuel ammonia into the combustion air during or before compression of the combustion air by the compressor and cools the combustion air.

CERAMIC MATRIX COMPOSITE COMPONENTS WITH HEAT TRANSFER AUGMENTATION FEATURES

An airfoil assembly for use in a turbine of a gas turbine engine includes an airfoil that extends radially relative to an axis. The airfoil includes an inner surface that defines a cooling cavity that extends radially into the airfoil and an outer surface that defines a leading edge, a trailing edge, a pressure side, and a section side of the airfoil. The airfoil assembly further includes features for increasing the heat transfer coefficient of the airfoil. 1. An airfoil assembly for use in a gas turbine engine , the airfoil assembly comprisinga ceramic matrix composite airfoil that extends axially relative to an axis, the ceramic matrix composite airfoil including an inner surface that defines an airfoil shaped cooling cavity that extends radially into the ceramic matrix composite airfoil and an outer surface that defines a leading edge, a trailing edge, a pressure side, and a section side of the ceramic matrix composite airfoil,a guide structure that extends axially relative to the airfoil, the cooling cavity defined between the inner surface of the airfoil and an outer surface of the guide structure, anda plurality of ceramic matrix composite pins embedded in the ceramic matrix composite airfoil, the plurality of ceramic matrix composite pins protrude away from the inner surface partway into the cooling cavity to increase a heat transfer coefficient of the ceramic matrix composite airfoil.2. The airfoil assembly of claim 1 , wherein the plurality of ceramic matrix composite pins include through-thickness reinforcement pins.3. The airfoil assembly of claim 1 , wherein the ceramic matrix composite airfoil includes a body and a layer of environmental barrier coating coupled with the body claim 1 , the body defines the inner surface and the environmental barrier coating defines the outer surface of the ceramic matrix composite airfoil claim 1 , each of the plurality of ceramic matrix composite pins extends into the body toward the environmental barrier coating ...

Crossover cooled airfoil trailing edge

A cooling circuit for a turbine bucket having an airfoil portion includes a trailing edge cooling circuit portion provided with a first radially outwardly directed inlet passage intermediate leading and trailing edges of the airfoil portion of the bucket, extending from a platform portion of the bucket to a location adjacent a radially outer tip of the bucket, and connecting to a second radially inwardly directed passage extending from a location adjacent the radially outer tip to a location adjacent the platform portion. The second radially inwardly directed passage connects to a third trailing edge region passage, and a plurality of crossover passages connect a radially outer half of the second radially inwardly directed passage to a radially outer half of the third trailing edge region passage.

COOLING STRUCTURE FOR HOT-GAS PATH COMPONENTS WITH METHODS OF FABRICATION

Embodiments of the present disclosure provide components for hot gas path (HGP) components and methods of forming the same. A structure according to the present disclosure can include: an HGP component extending radially from a rotor axis of a turbomachine, the HGP component including a tapered edge; a plurality of first passages in fluid communication with a preliminary cooling zone of the HGP component, and extending through a sidewall positioned between the preliminary cooling zone and the tapered edge; and a plurality of second passages extending through at least the tapered edge, wherein each of the plurality of second passages is in fluid communication with the flow path for the operative fluid and at least one passage of the plurality of first passages, and wherein at least one of the plurality of second passages is radially displaced from each passage of the plurality of first passages. 1. A cooling structure comprising:a hot gas path (HGP) component configured to be positioned within a flow path of an operative fluid and extending radially from a rotor axis of a turbomachine, the HGP component including a tapered edge;a plurality of first passages in fluid communication with a preliminary cooling zone of the HGP component, and extending through a positioned between the preliminary cooling zone and the tapered edge; anda plurality of second passages extending through at least the tapered edge, wherein each of the plurality of second passages is in fluid communication with the flow path for the operative fluid and at least one passage of the plurality of first passages, and wherein at least one of the plurality of second passages is radially displaced from each passage of the plurality of first passages.2. The cooling structure of claim 1 , wherein each second passage is non-coaxial with each first passage.3. The cooling structure of claim 1 , wherein at least one of the plurality of first passages and at least one of the plurality of second passages is in ...

COOLING HOLE WITH ENHANCED FLOW ATTACHMENT

A gas turbine engine component includes a wall having first and second wall surfaces, a cooling hole extending through the wall and a convexity. The cooling hole includes an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. The diffusing section includes a first lobe diverging longitudinally and laterally from the metering section and a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section. The convexity is located near the outlet. 1. A gas turbine engine component comprising:a wall having first and second wall surfaces; an inlet located at the first wall surface;', 'an outlet located at the second wall surface;', 'a metering section extending downstream from the inlet; and', a first lobe diverging longitudinally and laterally from the metering section; and', 'a second lobe adjacent the first lobe and diverging longitudinally and laterally from the metering section; and, 'a diffusing section extending from the metering section to the outlet, the diffusing section comprising], 'a cooling hole extending through the wall and comprisinga convexity located near the outlet.2. The component of claim 1 , further comprising:an indentation located downstream from the convexity on the second wall surface and downstream from the outlet.3. The component of claim 2 , wherein the indentation has a lateral width that is greater than or equal to a lateral width of the outlet.4. The component of claim 2 , wherein the indentation has a lateral width that is between about 100% and about 150% of the lateral width of the convexity.5. The component of claim 2 , wherein the indentation has a longitudinal length that is substantially equal to a longitudinal length of the convexity.6. The component of claim 1 , wherein the diffusing section further comprises:an interlobe region ...

AEROFOIL

An aerofoil component of a gas turbine engine has an aerofoil portion which spans, in use, a working gas annulus of the engine. The aerofoil portion has a pressure side outer wall and a suction side outer wall, each extending from the leading edge to the trailing edge of the aerofoil portion. The aerofoil portion further has one or more main passages which extend in the annulus-spanning direction of the aerofoil portion and which receive, in use, a flow of coolant. The aerofoil portion further has one or more suction wall passages which extend in the annulus-spanning direction of the aerofoil portion and which receive, in use, a flow of coolant, each suction wall passage being bounded on opposing first sides by the suction side outer wall and an inner wall of the aerofoil portion, the inner wall separating the suction wall passages from the main passages. 1. An aerofoil component of a gas turbine engine , the component having an aerofoil portion which spans , in use , a working gas annulus of the engine , the aerofoil portion having:a pressure side outer wall and a suction side outer wall which respectively define the external pressure side and suction side aerofoil surfaces of the aerofoil portion, each outer wall extending from the leading edge to the trailing edge of the aerofoil portion;one or more main passages which extend in the annulus-spanning direction of the aerofoil portion and which receive, in use, a flow of coolant therethrough;one or more suction wall passages which extend in the annulus-spanning direction of the aerofoil portion and which receive, in use, a flow of coolant therethrough, each suction wall passage being bounded on opposing first sides by the suction side outer wall and an inner wall of the aerofoil portion, the inner wall separating the suction wall passages from the main passages; anda plurality of dividing walls which extend between the suction side outer wall and the inner wall, each suction wall passage being bounded on opposing ...

ENGINE COMPONENT WITH HOLLOW TURBULATORS

An apparatus and method for cooling an engine component, such as an airfoil, for a turbine engine including an outer wall separating a cooling fluid flow from a hot fluid flow. A cooling circuit including a cooling passage having opposing sidewalls can be provided in the engine component. At least one turbulator can be provided between the opposing sidewalls, and can include a conduit having an inlet and an outlet passing through the at least one turbulator. 1. An airfoil for a turbine engine , the airfoil comprising:an outer wall defining a pressure side and a suction side extending axially between a leading edge and a trailing edge defining a chord-wise direction and extending radially between a root and a tip defining a span-wise direction;a cooling circuit located within the airfoil and comprising at least one cooling passage including a first sidewall and a second sidewall; andat least one turbulator provided in the at least one cooling passage extending between the first sidewall and the second sidewall; andat least one conduit having an inlet and an outlet, fluidly coupled to the cooling circuit, and at least partially passing through the at least one turbulator.2. The airfoil of wherein the at least one turbulator comprises a plurality of turbulators.3. The airfoil of wherein the at least one conduit includes a plurality of conduits passing through each of the plurality of turbulators.4. The airfoil of wherein the inlet is provided on one of an upstream surface or a downstream surface of the at least one turbulator claim 1 , or on one of the first or second sidewalls.5. The airfoil of wherein the outlet is provided on one of an upstream surface or a downstream surface of the at least one turbulator claim 1 , or on one of the first or second sidewalls.6. The airfoil of further comprising a film hole provided in the outer wall and fluidly coupled to the at least one cooling passage wherein the film hole forms the outlet.7. The airfoil of wherein the at least ...