ТЕПЛООБМЕННИК ДЛЯ КОНТУРА ВОЗДУШНОГО ОХЛАЖДЕНИЯ ТУРБИНЫ

Способ подачи охлаждающего воздуха в горячие зоны турбореактивного двигателя, содержащего последовательно от входа к выходу компрессор, диффузор, камеру сгорания, направляющий сопловой аппарат и турбину, приводящую в действие указанный компрессор, заключается в том, что отбирают расход воздуха в воздушном потоке, подаваемом компрессором, охлаждаемом в теплообменнике, расположенном радиально снаружи камеры сгорания. Далее этот воздух направляют радиально внутрь через неподвижные лопатки направляющего соплового аппарата и обдувают рабочее колесо турбины. Расход охлаждающего воздуха отбирают в зоне нижней части камеры, окружающей диффузор. Неподвижные лопатки направляющего соплового аппарата обдувают вторым расходом воздуха, отбираемого в нижней части камеры. Изобретение позволяет уменьшить загрязнение теплообменника и повысить эффективность охлаждения направляющего соплового аппарата. 2 н. и 2 з.п. ф-лы, 2 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.

СПОСОБ УПРАВЛЕНИЯ ПРОЦЕССОМ ОХЛАЖДЕНИЯ КОМПОНЕНТОВ ТУРБИНЫ

Изобретение относится к энергетике. Способ управления процессом охлаждения компонентов турбины, при котором во время фазы туманного охлаждения для охлаждения компонентов турбины используется разбавленный водяным туманом воздушный поток. В частности, фазе туманного охлаждения предшествует фаза воздушного охлаждения, во время которой для охлаждения компонентов турбины используется воздушный поток. При этом для процесса охлаждения задается один неизменный временной градиент температуры, причем плотность воздушного потока устанавливается посредством положения управляемого регулировочного клапана, и осуществляется переход из фазы воздушного охлаждения в фазу туманного охлаждения, когда достигнута максимальная плотность воздушного потока и, в частности, когда регулировочный клапан полностью открыт. Изобретение позволяет улучшить процесс принудительного охлаждения компонентов турбины. 4 з.п. ф-лы, 2 ил.

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ (ВАРИАНТЫ) И СПОСОБ ОХЛАЖДЕНИЯ РАЗМЕЩЕННЫХ ВНУТРИ НЕГО ДЕТАЛЕЙ

Газотурбинный двигатель снабжен компрессорной секцией сжатия входящего воздуха в воздух высокого давления и воздух промежуточного давления, секцией камеры сгорания топлива со сжатым воздухом, имеющей сообщение с указанной компрессорной секцией, турбинной секцией, связанной с указанной секцией камеры сгорания с возможностью сообщения с газообразными продуктами сгорания из нее. Турбинная секция содержит турбинную лопатку, имеющую вершину, расположенную в области более низкого давления, чем указанное промежуточное давление. Кольцевая полость расположена выше компрессорной секции по направлению потока и содержит отвод, сообщающийся с воздухом промежуточного давления. В кольцевой полости установлена с возможностью преобразования динамического напора указанного воздуха промежуточного давления в повышенное статическое давление этого воздуха перегородка. Газотурбинный двигатель снабжен каналом, имеющим входное отверстие сообщения с полостью и выходное отверстие, имеющее сообщение с турбинной лопаткой ...

ТУРБОАГРЕГАТ

Изобретение относится к энергетическому машиностроению. Турбоагрегат содержит корпус с установленным внутри него на подшипниках валом. На валу закреплено, по крайней мере, одно расширительное рабочее колесо. Подшипники выполнены несмазываемыми из полимерных композиционных материалов. В расширительном рабочем колесе и в валу выполнены каналы. Выполненные каналы сообщают проточную часть расширительного рабочего колеса с зазорами, образованными валом и подшипниками. Изобретение направлено на упрощение конструкции и снижение массогабаритных характеристик турбоагрегата. 1 ил.

ГИПЕРЗВУКОВОЙ САМОЛЕТ С КОМБИНИРОВАННОЙ СИЛОВОЙ УСТАНОВКОЙ И СИСТЕМА ОХЛАЖДЕНИЯ ТУРБИНЫ ВЫСОКОГО ДАВЛЕНИЯ ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ ТАКОГО САМОЛЕТА

Группа изобретений относится к гиперзвуковым самолетам. Гиперзвуковой самолет с комбинированной силовой установкой содержит фюзеляж, складываемые консоли крыла, два маршевых комбинированных двигателя, два маршевых ракетных двигателя, складывающиеся консоли переднего горизонтального оперения и кабину пилотов. Каждый маршевый двигатель имеет две ступени - турбовентиляторный двигатель и турбореактивный двигатель. В передней части фюзеляжа располагается обтекатель, внутри которого находятся двигатели бокового и вертикального разворота. На обтекателе расположены передние интерцепторы. В хвостовой части фюзеляжа располагается центральный газовод с кольцевым основанием, на котором установлен промежуточный газовод, который снабжен направляющими лопатками. На центральном газоводе установлен корпус привода промежуточного газовода. Турбореактивный двигатель имеет компрессор, турбину высокого давления и турбину низкого давления, которые расположены по внешней окружности корпуса турбореактивного двигателя ...

ОПОРА ТУРБИНЫ НИЗКОГО ДАВЛЕНИЯ

Изобретение относится к опорам турбин газотурбинных двигателей авиационного и наземного применения. Опора турбины низкого давления выполнена с радиальными силовыми стойками, размещенными в обтекателях, установленных в газовом тракте турбины. Обтекатели силовых стоек опоры выполнены с передней, средней и задней полостями и с симметричным наружным профилем, с углом наклона оси симметрии наружного профиля обтекателя относительно радиальной плоскости, равным 10…50 градусов. Силовые стойки установлены в средней полости обтекателя, причем средняя полость каждого обтекателя выполнена с возможностью прохода охлаждающего воздуха для охлаждения силовых стоек, установленных в указанной полости. Силовые стойки выполнены с цилиндрической внутренней поверхностью, образующей внутреннюю полость, и с цилиндрическими и плоскими поверхностями наружной поверхности, с образованием щелевых воздушных полостей между внешними плоскими боковыми поверхностями силовой стойки и внутренними плоскими боковыми поверхностями ...

СПОСОБ И УСТРОЙСТВО ДЛЯ ОХЛАЖДЕНИЯ УПЛОТНЕНИЯ ДЛЯМАШИННОГО ОБОРУДОВАНИЯ

Способ для охлаждения уплотнения турбинного вала, расположенного в стенке камеры. Уплотнение нагревается горячим сжатым паром, который просачивается через лабиринт в камеру, и внутренним трением. Способ содержит этапы: а) обеспечения камеры, в которой расположено уплотнение и в которую просачивается горячий сжатый пар, b) впрыска холодной жидкости в камеру, и с) охлаждения и конденсации горячего сжатого пара в камере. Способ применяется в силовых установках, которые включают в себя испаритель, конденсатор и циркуляционный насос. Такие способ и устройство позволят снизить износ на поверхностях уплотнения. 2 н. и 8 з.п. ф-лы, 6 ил.

СИСТЕМА ПОДАЧИ ОХЛАЖДАЮЩЕГО ВОЗДУХА В ГАЗОВУЮ ТУРБИНУ

Система подачи охлаждающего воздуха в газовую турбину, в которой охлаждающий воздух от источника высокого давления поступает внутрь газовой турбины и перемещается в радиальные ускорители. Радиальные ускорители вызывают тангенциальное ускорение воздуха в направлении окружного движения поверхности ротора. После ускорения охлаждающего воздуха до окружной скорости ротора он входит в радиальные отверстия. Затем охлаждающий воздух выпускается в полый ротор при соответственно уменьшенном радиусе выхода. Система содержит лабиринтное уплотнение в сочетании с щеточным уплотнением для отделения камеры для подачи воздуха в радиальные отверстия от камеры, которая сообщается с передним пространством ротора в 1-й ступени турбины, посредством образования промежуточной камеры, которая предотвращает смешивание потока вследствие утечки от осевого компрессора с потоком охлаждающего воздуха от ускорителей. Лабиринтное уплотнение в сочетании с щеточным уплотнением приспособлено для отделения нагнетательной стороны ...

УСТРОЙСТВО ПОДВОДА ОХЛАЖДАЮЩЕГО ВОЗДУХА К НАДРОТОРНЫМ ВСТАВКАМ

Полезная модель относится к области турбостроения, в частности к устройствам подвода охлаждающего воздуха к надроторным вставкам в газотурбинном двигателе, и может быть использована в статорах высокотемпературных турбин газотурбинных двигателей авиационного и наземного применения. Техническим результатом заявляемого устройства является повышение эффективного использования охлаждающего воздуха для охлаждения надроторной вставки, что позволяет снизить его потребное количество, при этом достигается минимизация неплотностей за счет герметизирующего элемента, установлено в соосных выборках промежуточного элемента и надроторной вставки, через которые охлаждающий воздух перетекает в газовоздушный тракт, не совершив необходимой работы по охлаждению уплотнения. Технический результат достигается тем, что в устройстве подвода охлаждающего воздуха к надроторным вставкам, содержащее промежуточный элемент, вмонтированный в силовой корпус, в промежуточный элемент вмонтирована надроторная вставка, имеющая ...

СОПЛОВОЙ АППАРАТ ТУРБИНЫ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Сопловой аппарат турбины высокого давления, содержащий закрепленные в наружном корпусе 1 пустотелые лопатки 2 с внутренней разделительной перегородкой 3 и двумя радиальными ребрами 4, 7 на нижней полке, направленными к оси турбины, причем переднее ребро 4, являющееся продолжением внутренней перегородки 3 и закрепленное на верхнем опорном фланце 5 корпуса внутреннего 6, ограничивает совместно с другим радиальным ребром 7, расположенным на внутренней поверхности нижней полки в районе выходной кромки 8 лопатки, серединную часть нижней полки 9 с выполненными в ней отверстиями 10, выходящими в проточную часть и сообщенными на входе с полостью 11, сформированной внутренней поверхностью серединной части нижней полки лопатки и отстоящей от нее с необходимым зазором обечайкой 12, герметично скрепленной по всему периметру с ребрами 4, 7 и торцевыми стыками лопатки 14, отличающийся тем, что внутренняя поверхность серединной части нижней полки 9 выполнена гладкой, а в обечайке 12 выполнены отверстия ...

Двухпоточный цилиндр среднего давления паровой турбины

Изобретение относится к области теплоэнергетического машиностроения и может быть использовано при модернизации действующего оборудования и создании новых турбин. Предложен двухпоточный цилиндр среднего давления паровой турбины, включающий наружный и внутренний корпусы, ротор с дисками и рабочими лопатками проточной части прямого и обратного потоков, направляющие лопатки первых ступеней прямого и обратного потоков, диафрагмы вторых ступеней прямого и обратного потоков, кольцевое экранирующее тело, установленное в центральной части внутреннего корпуса, и обойму, расположенную осесимметрично внутри экранирующего тела и снабженную кольцевыми камерами, соединенными между собой и имеющими отверстия на внутренней и торцевых стенках обоймы, трубопровод подачи охлаждающего пара от внешнего источника в обойму, при этом в диафрагмах вторых ступеней прямого и обратного потоков выполнены кольцевые камеры и установлены форсунки, в направляющих лопатках диафрагм вторых ступеней обоих потоков выполнены ...

ТУРБИНА ПРЕИМУЩЕСТВЕННО ДЛЯ ЖИДКОСТНОГО РАКЕТНОГО ДВИГАТЕЛЯ

Турбина преимущественно для жидкостного ракетного двигателя содержит корпус, в котором выполнены рабочая камера, канал подвода рабочего тела, сообщенный с камерой, полость охлаждения, сообщенный с ней канал подвода охлаждающей среды. В торцевой части рабочей камеры со стороны канала подвода рабочего тела расположено средство для подачи охлаждающей среды на рабочее колесо, выполненное в виде кольцевой камеры, сообщенной с кольцевой щелью, расположенной в корпусе, и с каналом подвода охлаждающей среды. Турбина содержит также втулку, выполненную из материала с теплопроводностью, большей, чем теплопроводность материала корпуса, и расположенную коаксиально рабочему колесу. Втулка закреплена в корпусе с образованием охлаждающей полости, расположенной на ее наружной поверхности, выполнена длиной, большей, чем длина рабочего колеса, и установлена с выступом за его габариты в сторону канала подвода рабочего тела. Наружные поверхности кольцевой камеры и кольцевой щели средства для подачи охлаждающей ...

Паротурбинная установка с охлаждением элементов проточной части цилиндра низкого давления на малорасходных режимах

Изобретение относится к области энергомашиностроения, в частности паротурбостроения, и может быть использовано при проектировании и модернизации проточных частей цилиндров низкого давления паровых турбин тепловых и атомных станций. Паротурбинная установка включает паровую турбину, содержащую цилиндр низкого давления с паровпускным патрубком и проточной частью, в которой размещен ротор с рабочими лопатками, диафрагмы с направляющими лопатками и выхлопной патрубок. По меньшей мере одна диафрагма снабжена полыми направляющими лопатками и ободом с кольцевой камерой, причем каждая полая направляющая лопатка имеет по меньшей мере одну внутреннюю полость с каналами для выпуска охлаждающего влажного пара, соединяющими внутреннюю полость с проточной частью. Внутренняя полость выполнена на высоте 0,15-0,25 высоты каждой полой направляющей лопатки в ее периферийной зоне со стороны выходной кромки, а каналы для выпуска охлаждающего влажного пара выполнены в виде сквозных щелей. Кольцевая камера сообщена ...

СПОСОБ КОНТРОЛЯ ОХЛАЖДАЕМОЙ СТЕНКИ С ОТВЕРСТИЯМИ

Изобретение может быть использовано в энергомашиностроении и авиадвигателестроении. Способ контроля охлаждаемой стенки с отверстиями заключается в обдуве газом стенки, контактирующей с расплавом кристаллизующегося металла, с температурой, меньшей температуры кристаллизации расплава. Одновременно вдувают воздух в поток газа и измеряют местные толщины корки образовавшегося на ее поверхности металла. Контакт с расплавом кристаллизующегося металла осуществляют с наружной стороны стенки, температуру вдуваемого воздуха поддерживают выше температуры основного потока газа, но ниже или равной температуре кристаллизации расплава и об эффективности охлаждения судят по коэффициенту теплоотдачи, определяемому по формуле, приведенной в тексте описания. Изобретение позволяет повысить точность и снизить трудоемкость контроля. 1 ил.

СПОСОБ ИЗГОТОВЛЕНИЯ ЛОПАТКИ РОТОРА ТУРБИНЫ

Настоящее изобретение относится к способу изготовления лопатки ротора турбины. Способ изготовления лопатки ротора турбины с использованием ковочного сплава на основе Ni содержит этап размягчения, включающий этап горячей ковки и этап охлаждения, заключающийся в обеспечении повышения содержания γ'-фазы, не когерентной с γ-фазой, которая представляет собой матричную фазу в ковочном сплаве на основе Ni; первый этап обработки, заключающийся в формировании по меньшей мере двух элементов конструкции, составляющих лопатку ротора, с использованием ковочного сплава на основе Ni, осуществляемый после этапа размягчения; второй этап обработки, заключающийся в формировании элементов охлаждающей структуры в каждом из элементов конструкции в виде канала прохождения охлаждающего потока; и третий этап обработки, заключающийся во взаимном соединении элементов конструкции при помощи сварки трением с перемешиванием; причем содержание γ'-фазы в ковочном сплаве на основе Ni составляет при температуре не ниже ...

СИСТЕМА ОХЛАЖДЕНИЯ НЕПОДВИЖНО УСТАНОВЛЕННОГО СТЯЖНОГО КОЛЬЦА ГАЗОВОЙ ТУРБИНЫ

... 1. Фиксированное стяжное кольцо (4), расположенное вокруг газового окна (14) прохождения горячих газов газовой турбины (1) и окруженное зафиксированной кольцевой полостью (2), образующей кольцевую камеру охлаждения (26), в которую выходит, по меньшей мере, одно впускное отверстие подачи охлаждающего воздуха (28); при этом вышеуказанное стяжное кольцо (4) состоит из множества кольцевых сегментов (16), отличающееся тем, что каждый кольцевой сегмент (16) включает в себя внутреннюю систему верхнего охлаждения (А) и внутреннюю систему нижнего охлаждения (В, В'), независимую от внутренней системы верхнего охлаждения (А), радиально смещенную относительно нее и включает в себя, по меньшей мере, одно впускное отверстие подачи охлаждающего воздуха (4) из камеры охлаждения (26). 2. Стяжное кольцо (4) по п.1, отличающееся тем, что система верхнего охлаждения (А) каждого кольцевого сегмента (16) включает в себя, по меньшей мере, одну внутреннюю первую полость (32), которая располагается под углом между ...

СПОСОБ РАСХОЛАЖИВАНИЯ ПАРОВОЙ ТУРБИНЫ

Способ относится к теплоэнергетике, в частности к методам принудительного воздушного или комбинированного расхолаживания паровых турбин, преимущественно с двухкорпусным вариантом выполнения цилиндров высокого и среднего давления при их останове и выводе в ремонт. Изобретение обеспечивает интенсификацию охлаждения проточной части цилиндров высокого и среднего давления (ЦВД и ЦСД), повышение надежности путем уменьшения вероятности возникновения тепловых ударов и равномерность остывания всех участков цилиндров турбины. Для этого к потокам воздуха, пропущенного через межкорпусные пространства (МКЦ), добавляют дополнительный воздух, подаваемый в ЦВД и ЦСД отдельными потоками через камеры отборов, причем поддержание на заданном уровне температурного состояния металла фланцевых соединений, наружных корпусов и относительного укорочения ротора осуществляют перераспределением расходов воздуха, подаваемых через камеры отборов, с одновременным изменением расходов воздуха на фланцевые соединения и в ...

КОЛЬЦЕВОЙ КОРПУС СТАТОРА ГАЗОВОЙ ТУРБИНЫ И УСТРОЙСТВО ОХЛАЖДЕНИЯ КОЛЬЦЕВОГО КОРПУСА

... 1. Устройство охлаждения кольцевого корпуса статора, окружающего канал (6) прохождения горячих газов в газовой турбине, причем указанный кольцевой корпус содержит кольцевые сегменты (10), прикрепленные при помощи передних и задних систем крепления к сегментам (12) перемычки, образующим неподвижную поддерживающую перемычку, окружающую указанный кольцевой корпус и ограничивающую, по меньшей мере, одну расположенную между ними кольцевую напорную полость (36), в которую выходит, по меньшей мере, одно отверстие (38) подачи охлаждающего воздуха, причем стенки каждого из кольцевых сегментов (10) содержат отверстия (40) выпуска воздуха, соединяющие напорную полость (36) с каналом (6) прохождения горячих газов, отличающееся тем, что дополнительно содержит средства (48, 50, 54, 56) направления воздуха, поступающего из течей в системах крепления между кольцевыми сегментами (10) и сегментами (12) перемычки, к, по меньшей мере, одному из аксиальных краев (22, 32) кольцевых сегментов (10), находящемуся ...

СПОСОБ УПРАВЛЕНИЯ ПРОЦЕССОМ ОХЛАЖДЕНИЯ КОМПОНЕНТОВ ТУРБИНЫ

... 1. Способ управления процессом охлаждения компонентов (8, 10, 12) турбины, в частности вала паровой турбины,при котором во время фазы (P4) туманного охлаждения для охлаждения компонентов (8, 10, 12) турбины используется разбавленный водяным туманом воздушный поток,при котором фазе (P4) туманного охлаждения предшествует фаза (P3) воздушного охлаждения, во время которой для охлаждения компонентов (8, 10, 12) турбины используется воздушный поток,при котором во время фазы (P3) воздушного охлаждения и во время фазы (P4) туманного охлаждения для процесса охлаждения задается один неизменный временной градиент температуры,при котором задается временной градиент температуры, равный примерно 10 К/ч,при котором для задания градиента температуры во время фазы (P3) воздушного охлаждения регулируется плотность воздушного потока, а во время фазы (P4) туманного охлаждения количество добавленного в воздушный поток водяного тумана,при котором плотность воздушного потока устанавливается посредством положения ...

ГАЗОВАЯ СИЛОВАЯ ТУРБИНА

Газовая силовая турбина с магнитной опорой, имеющая вал, электромагнитные подшипники и страховочные подшипники качения, отличающаяся тем, что вал силовой турбины выполнен двойным, состоящим из внешнего и внутреннего валов с воздушной межвальной полостью между ними, причем на переднем хвостовике внутреннего вала установлены диски турбины, а на заднем хвостовике - упругая муфта передачи полезной мощности, при этом на наружной поверхности внешнего вала установлены роторные элементы электромагнитных подшипников, а страховочные подшипники качения установлены на переднем и заднем хвостовиках внешнего вала, причем внешний вал зафиксирован относительно внутреннего вала в окружном направлении шлицами на заднем хвостовике, в радиальном направлении - радиальными ребрами, расположенными в межвальной воздушной полости, а в осевом направлении - опорными буртами, расположенными в межвальной воздушной полости, а также фланцем упругой муфты, при этом межвальная воздушная полость на входе через отверстия ...

Воздушная система газотурбинного двигателя

Изобретение относится к воздушным системам газотурбинных двигателей. Воздушная система газотурбинного двигателя содержит полости наддува опор и предмасляные полости компрессора низкого давления и компрессора высокого давления, полость наддува опор и предмасляные полости турбины, питающий воздуховод. Полости наддува опор компрессора низкого давления и компрессора высокого давления и полость наддува опор турбины воздуховодами сообщены и друг с другом, и с внутривальной зоной, образованной валом турбины низкого давления, и с межвальной зоной, образованной компрессором высокого давления и валом турбины низкого давления, а через уплотнения с газовоздушным трактом двигателя и с одноименными предмасляными полостями и полостями маслосистемы. Воздушная система снабжена, по меньшей мере, одним дополнительным воздуховодом, вход которого сообщен с внутривальной зоной, а выход непосредственно с предмасляными полостями турбины. При этом площадь проходного сечения дополнительного воздуховода превышает ...

СПОСОБ ИЗГОТОВЛЕНИЯ ЛОПАТКИ РОТОРА ТУРБИНЫ

Настоящее изобретение относится к способу изготовления лопатки ротора турбины. Способ изготовления лопатки ротора турбины с использованием ковочного сплава на основе Ni содержит этап размягчения, включающий этап горячей ковки и этап охлаждения, заключающийся в обеспечении повышения содержания γ'-фазы, не когерентной с γ-фазой, которая представляет собой матричную фазу в ковочном сплаве на основе Ni; первый этап обработки, заключающийся в формировании по меньшей мере двух элементов конструкции, составляющих лопатку ротора, с использованием ковочного сплава на основе Ni, осуществляемый после этапа размягчения; второй этап обработки, заключающийся в формировании элементов охлаждающей структуры в каждом из элементов конструкции в виде канала прохождения охлаждающего потока; и третий этап обработки, заключающийся во взаимном соединении элементов конструкции при помощи сварки трением с перемешиванием; причем содержание γ'-фазы в ковочном сплаве на основе Ni составляет при температуре не ниже ...

УСТРОЙСТВО ДЛЯ ОХЛАЖДЕНИЯ ПАЗОВ ДИСКА РОТОРА В ТУРБОМАШИНЕ, ИМЕЮЩЕЕ ДВА ПОТОКА ПОДАВАЕМОГО ВОЗДУХА

... 1. Устройство для охлаждения пазов в диске ротора турбомашины, содержащее: ! диск (4) ротора, содержащий: ! на его периферии множество, по существу, осевых пазов (10), которые распределены с равными интервалами вокруг оси (X-X) вращения диска; и ! кольцевой фланец (14), проходящий перед расположенной выше по потоку радиальной поверхностью (16) диска; ! множество лопаток (2), каждая из которых имеет хвостовик (12), установленный в соответствующем пазу в диске ротора; ! удерживающее кольцо (22), имеющее конец (23), который установлен у расположенной выше по потоку радиальной поверхности (16) диска, и кольцевой фланец (24), который проходит перед указанной расположенной выше по потоку радиальной поверхностью диска и размещен вокруг фланца (14) диска, при этом он взаимодействует с ним так, что между ними остается кольцевое пространство (28), образующее полость для диффузии охлаждающего воздуха, причем диффузионная полость открывается на ее расположенном ниже по потоку конце в нижнюю часть каждого ...

Выхлопной патрубок паровой турбины

ВЫХЛЬПНОЙ ПАТРУБОК ПАРОВОЙ ТУРБИНЫ, по авт. св. № 877090, отличающийся тем, что, с целью повышения надежности путем уменьшения эрозии лопаток , к внутренней поверхности стенки прикреплен дополнительный коллектор с автономным подводом пара, сообщенный с проточной частью кольцевой щелью, на выходе которой установлен генератор акустических колебаний в виде струны или пластины . (Л 4 СО ОО ю ...

Устройство для охлаждения ротора паровой турбины

Изобретение относится к турбостроению, может быть использовано при проектировании и модернизации паровых турбин и позволяет повысить эффективность охлаждения ротора и его эксплуатационную надежность. Устройство содержит основной и дополнительный распределительные коллекторы 3 и 18 с патрубками подвода охлаждающего пара, к которым прикреплены отводящие патрубки 21 и 22, пропущенные соответственно через отверстия 7 в гребне 6 обоймы 5 концевого уплотнения и отверстия 23, 24, 25 и 26 в обойме 17, ободе 27, лопатках 28 и теле 29 диафрагмы 15 второй ступени 16. Выходы патрубков 21 и 22 расположены в камерах 8 и 14, куда поступает охлаждающий пар из коллекторов 3 и 18. В камере 8 охлаждающий пар смешивается с паром из первой ступени 12 и охлаждает ротор 10 в зоне думмиса 9 и диска 11 первой ступени 12. Пар из камеры 14 разделяется на два потока, один из которых через разгрузочные отверстия 13 поступает в камеру 8, а другой - в диафрагменные уплотнения 31 диафрагмы 15, охлаждая ротор 10 в зоне ...

Охлаждаемая диафрагма

Паросиловая установка

Изобретение относится к теплоэнергетике и позволяет повысить надежность и экономичность процесса расхолаживания. Всасывающая камера . (К) 9 эжектора 7 расхолаживания трубопроводом 14 сообщена с атмосферой, а выхлопная К 8 подключена к коллектору 13 отсоса неконденсирующихся газов. Такое выполнение позволяет интенсифицировать охлаждение фланцев и шпилек за счет использования энергии перепада давлений между К с одновременным охлаждением внутренних полостей цилиндра. 1 ил. 28 34 333 8 17 (Л 00 OCI 00 sj со 8 17 J3 12 ...

Способ охлаждения статора газовой турбины обоймовой конструкции

Статор газовой турбины

Паротурбинная установка

Устройство для расхолаживания паровой турбины

SCHAUFFELZUSAMMENSETZUNG FÜR EINE GASTURBINE MIT INTEGRIERTER KÜHLDÜSE

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.

DICHTUNGSTEIL, INSBESONDERE DICHTUNGSRING, FUER EIN GASTURBINENTRIEBWERK

Gasturbinenanlage, insbesondere Kleingasturbinenanlage

REGELUNG DES SCHAUFELSPRITZENSPIELS FÜR EINE GASTURBINE.

Kuehleinrichtung fuer Verbrennungsturbinen

Turbinenseitiges Lagerträger-und Kühlsystem

Ein Turbolader (300) weist eine drehbare Welle (312) auf, die durch ein Lagergehäuse (303) verläuft, mit einem turbinenseitigen Lagerträger (580) mit einem Tragriegel (600). Ein turbinenseitiger Lagerträger (580) kann den Tragriegel (600) mit radialen Trägern (610) und (620) umfassen, die mit jeder gegenüberliegenden Seite (492) und (494) des Hohlraums (496) des Lagergehäuses (303) verbunden sind, wobei Öl in Ausnehmungen (640) und (650) über den radialen Trägern (610) und (620) strömen kann. Jeder radiale Träger (610) und (620) des Tragriegels (600) ist vorzugsweise in die entsprechende Seite (492) und (494) integriert. Ein turbinenseitiger Lagerträger (580) kann auch den Tragriegel (600) mit einem Stangenabschnitt (622) umfassen, der einen unteren Verbindungstragriegel (612) bildet, der die Böden beider Lagerträger (490) und (580) unterstützt, die von oben durch das Lagergehäuse (303) gehalten werden. Jeder Tragriegel (600) ist vorzugsweise einteilig mit dem turbinenseitigen Lagerträger ...

GASTURBINE UND GASTURBINENBETRIEBSVERFAHREN

Eine Gasturbine wird mit einem Abgasdiffusor 5, in dem ein Abgasströmungsweg Pe zum Zirkulieren von Abgas von einer Turbine gebildet ist, und einer Kühlvorrichtung 6 zum Kühlen einer Struktur bereitgestellt, die zu dem Abgasströmungsweg Pe im Abgasdiffusor 5 gerichtet ist. Die Kühlvorrichtung 6 weist ein Führungsteil 7, in dem ein Führungsströmungsweg Pg zum Zirkulieren eines Kühlmediums gebildet ist, und das das Kühlmedium zu der Struktur führt, und ein Umschaltteil 8 auf, das in der Lage ist, zwischen einem ersten Zustand, bei dem eine Fließgeschwindigkeit des Kühlmediums, das durch den Führungsströmungsweg Pg strömt, eine erste Fließgeschwindigkeit ist, die einer Fließgeschwindigkeit während eines Nennbetriebs entspricht, und einem zweiten Zustand, bei dem die Fließgeschiwindigkeit des Kühlmediums eine zweite Fließgeschwindigkeit ist, die höher als die erste Fließgeschwindigkeit ist, umzuschalten.

Kuehleinrichtung fuer Gasturbinen, insbesondere fuer mit Schweroel betriebene Gasturbinen

Verfahren und Vorrichtung für die Kühlung von Gasturbinenleitschaufeln

Turbinengehäuse mit einem Hilfsflansch

Ein Turbinengehäuse (300) kann eine Außenfläche (345) mit einem Hilfsflansch (340) und eine Innenfläche (375) mit einer Wärmeableiteinrichtung (370) enthalten, die benachbart zu dem Hilfsflansch (340) positioniert ist.

Gekühlte Schaufel für eine Gasturbine

Eine gekühlte Schaufel (10) für eine Gasturbine weist ein von einem Schaufelfuß (12) und einem Schaufelschaft (25) ausgehendes Schaufelblatt (11) mit einer Vorderkante (19) und einer Hinterkante (20) sowie innerhalb des Schaufelblatts (11) eine Mehrzahl von sich in radialer Richtung erstreckenden, strömungsmäßig hintereinander geschalteten Kühlkanälen (13, 14, 15) auf, von denen ein erster Kühlkanal (13) entlang der Vorderkante (19) und ein zweiter Kühlkanal (15) entlang der Hinterkante (20) vom Schaufelfuß (12) zur Spitze des Schaufelblatts (11) von einem Hauptstrom eines Kühlmediums durchströmt werden und der Ausgang des ersten Kühlkanals (13) über einen ersten Umlenkbereich (17), einen zwischen dem ersten und zweiten Kühlkanal (13, 15) angeordneten dritten Kühlkanal (14) und einen zweiten Umlenkbereich (18) mit dem Eingang des zweiten Kühlkanals (15) in Verbindung steht. Es sind erste Mittel (22, 23, 24) vorgesehen, durch welche dem vom dritten Kühlkanal (14) in den zweiten Kühlkanal ...

Controlled thermal and acoustic insulation structure for machine housing e.g. power station turbine housing, has isolation chambers that are filled or emptied with insulation materials or hot and cold mediums

The insulation structure has isolation chambers (6-9) which are enclosed with machine housing (1). The chambers are filled or emptied with insulation materials or hot and cold mediums by pumping and suction processes.

Vorrichtung zur Kühlung von Flanschen bei Dampfturbinen

ROTIERENDE MASCHINE

AUSGLEICHSVENTIL MIT VARIABLEM QUERSCHNITT

STRÖMUNGSMASCHINE MIT EINER KÜHLBAREN ANORDNUNG VON WANDELEMENTEN UND VERFAHREN ZUR KÜHLUNG EINER ANORDNUNG VON WANDELEMENTEN

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 innerer Kühlung

Einrichtung zur Mischung verschiedener Gasstroeme in einer Gasturbine

Wellenanordnung für eine dampfgekühlte Gasturbine

Mischvorrichtung für geschmolzene Substanzen mit hohen Temperaturen

Plattformkühlanordnung für den Leitschaufelkranz einer Gasturbine

Bei einer Plattformkühlanordnung für den der Brennkammer nachgeschalteten Leitschaufelkranz einer Gasturbine sind am Umfang eine oder mehrere parallele Reihe(n) von kontinuierlich oder in Gruppen angeordneten Kühlluftausblaskanälen (10) angeordnet. Die Kühlluftausblaskanäle sind, bezogen auf die Umfangsrichtung (15), in einem Winkel (alpha) winklig angeordnet, um an der Oberfläche der Plattform (2) eine Wirbelstruktur zu erzeugen, die zum einen die Vermischung der Kühlluftstrahlen (11) mit dem Heißgasstrom (8) vermindert und zum anderen auch die vollständige Kühlung des stromabwärts einer Grenzschichttrennlinie (13) liegenden Bereichs der Grenzschichtablösung (12) bis hin zur Saugseite (14) der benachbarten Leitschaufel (1) gewährleistet.

Gasturbine

Gasturbinenstatorschaufelsegment mit einem zweigeteilten Hohlraum

Gasturbine

Zentrierung eines Bauteils im Innern einer Welle

Improvements in gas turbines

... 679,916. Gas turbines. FEILDEN, G. B. R., and RUSTON and HORNSBY, Ltd. April 27, 1950 [April 29, 1949], No. 11454/49. Class 110(iii) A casing construction for an axial flow gas turbine comprises inner and outer cylindrical casings 11, 10 connected together at the ends remote from the turbine rotor, the free end of the outer casing having an extension 19 with inwardly projecting webs 12, 22 on which the segments of the shroud ring 23 are supported. The inlet volue 27 is embedded in lagging 12 between the inner and outer casings 11, 10 and the inner wall of the volute is supported by means of a ring member 26 which comprises an annular plate 26d attached to the end of the inner casing 11, an intermediate thin-walled cylindrical portion 26c which is relatively long comparea with its thickness to permit expansion, an annular web 26b and a further cylindrical portion 26a which forms the inner wall of the nozzle annulus. The free end of the inner casing 11 has an inward flange 11a to which is ...

Improvements in or relating to internal combustion turbine motors

... 693,682. Gas-turbine plant. SHARMA, D. N. Oct. 25, 1949, No. 27349/49. Class 110(iii) A gas turbine plant comprising a multi-stage compressor 7, 8. driven by a turbine 2 through a shaft 16 has a hollow cylindrical stationary shaft 19 surrounding the shaft 16. Compressor blades 3 mounted on the shaft 16 are adapted to supply cooling air to the bearings and turbine blades. The compressor 7, 8 may be of the axial flow or centrifugal type. The cooling air is sucked into the shaft 19 through apertures 20 and is forced by the compressor blades 3 in the directions indicated by the arrows z so as to cool the bearings 12, 14 and 15 and the turbine blades by means of the passages 6. Power may be taken from the shaft through reduction gearing 11.

Turbocharger

A turbocharger comprises a turbine rotor 20 and a compressor rotor 24 mounted on a shaft 30 and a housing 12, 14 defining a volute about each impeller. The housing includes a wall 25 between the impellers which comprises a ceramic portion 26 adjacent the turbine and a metal portion 28 adjacent the compressor and contiguous with the ceramic portion. A ball bearing 32 is mounted in the metal portion of the wall and supports the shaft, and a ball bearing is mounted in the housing 14 and supports the compressor. The bearings are lubricated by oil from sump 40 via delivery line 42 and 44. ...

Boundary layer control using electroformed microporous material

Boundary layer control of a structural element in a fluid stream is achieved by the following operations:providing in such structural element at least one region equipped with micro porous structure by an electroforming technique;having a fluid stream flow through the external surface of the said at least one region, inwards or outwards with respect to the environment in which that element is placed.

Gas turbine cooling air flow control using shaped memory metal valve

A cooling air flow control device for a gas turbine engine comprises a cooling passage 76 defined within a component 50 and a shaped memory metal valve 114 disposed in the passage 76 to regulate, in use, the flow rate of a cooling air flow supplied through the passage 76 wherein the valve 114 operates by changing shape to control the cooling air flow rate in response to the temperature of the component 50. The valve 114 may be a disc which dilates (fig 4A) or distorts (fig 4B), or a plate (122, fig 5A,5B) or series of plates (122, fig 6A,6B) which distort. A plurality of valves (114, fig 8) which dilate or distort at different temperatures may be arranged in series in the passage 76. The component 50 may be a blade, vane disc, casing or bearing chamber or heat exchanger. The valve 114 may switch shape at a given temperature, or change shape gradually over a temperature range.

Vorticity control in a gas turbine engine aerofoil

Gas turbine engine cooling apparatus

Gas turbine engine temperature modulated cooling flow

Dual generator gas turbine genset; Rotor cooling

Gas turbomachinery electricity generation apparatus includes a gas turbomachinery arrangement having an associated rotary drive take-off, an electricity generating arrangement which includes a first generator stage including a first generator rotor and generator stator arrangement; a second generator stage including a second generator rotor and generator stator arrangement, at least one of the first and second generator stage rotors is driven by the rotary drive take-off.

Gas turbine secondary air to seal shaft bearing lubricant

A gas turbine engine comprises an inner shaft 8 connecting a low pressure compressor and a low pressure turbine, an outer shaft 7, coaxial with the inner shaft 8, connecting a high pressure compressor HC and a high pressure turbine HT, and a pair of axially spaced bearing boxes 21, (25, fig 1), each accommodating a bearing 5f, 6f (5r. 6r, fig 3), for supporting an end of one of the shafts 7, 8. An air passage 41, 45, (48, fig 3) transfers air, drawn form the high pressure compressor HC, to a seal section provided at each of the bearing boxes 21 (25, fig 1), and a high pressure introduction turbine 62 is provided in the air passage 41 and attached to the outer shaft 7 for rotation therewith. The high pressure compressor HC may be of the centrifugal type.

Cooling system for turbine exhaust assembly

A cooling system for a turbine exhaust assembly 10 comprises an annular case 20, 21, a flow-path ring 12 and a splash plate or baffle 32. The flow-path ring 12 is coaxially disposed within the annular case 20, 21, and the splash plate or baffle 32 extends axially between the annular case 20, 21 and the flow-path ring 12. A plurality of cooling fluid apertures 31 are formed in the annular case 20, 21 in order to provide cooling fluid flow onto the splash plate or baffle 32. A plurality of impingement holes 33 are formed in the splash plate or baffle 32 in order to provide impingement cooling flow onto the flow-path ring 12. The flow-path ring may be provided with film cooling apertures 41 along the aft portion and/or the full axial length thereof. A blocker door 28 may be provided outside of the annular case 20, 21, downstream of the apertures 31, to increase a relative pressure of the cooling fluid flow. The annular case 20, 21 may comprise an inner turbine exhaust case or an outer turbine ...

A cooling arrangement

AN AXIALLY FLEXIBLE RADIALLY STIFF RETAINING RING FOR SEALING IN A GAS TURBINE ENGINE

Shroud hanger with diffused cooling passage

Two stage turbocharger with cooling arrangement

The high pressure compressor wheel 21 of a two stage turbocharger assembly is cooled by charge air bled from the charge air flowpath 2 downstream of the aftercooler 5. A wastegate 40 may be arranged across the high pressure turbine 22 and operated by an actuator 41 which is operable by the pressure of the charge air in the cooling flowpath 30. The cooling airflow may be blocked to open the wastegate 40 and opened to close the wastegate 40 so that cooling air is supplied only while the high pressure compressor wheel 21 is under load.

COOLING AIR PRESSURE CONTROL IN A GAS TURBINE ENGINE

A bearing arrangement for the main shaft of a steam or gas turbine

... 957,616. Cooling bearings. LICENTIA PATENT - VERWALTUNGS - G.m.b.H. Feb. 8, 1961 [Feb. 8, 1960], No. 4682/61. Heading F2A. A bearing arrangement for the main shaft 12 of a turbine comprises a bearing bracket 11 which is integral with the turbine casing 10, wherein a rotating fan-type part 21 attached to the main shaft end 12 produces a circulation 23 of coolant such as air axially through aperture 20 around the bearing housing 15 to outlets 24.

Improvements in internal combustion turbine plant

... 595,348. Cooling bearings. BAUMANN, K., and METROPOLITAN-VICKERS ELECTRICAL CO., Ltd. Sept. 22, 1941, No. 12274. [Class 12 (i)] [Also in Group XXVI] In a gas turbine plant, for the propulsion of aircraft, air taken from an intermediate stage of the compressor is used for cooling the shaft bearing near the turbine disc and the chamber enclosing the bearing is itself enclosed in a chamber to which is admitted air taken from the compressor delivery. This latter air passes to a space between the turbine rotor and a cover disc and thence to the turbine blade stream. The turbine disc stub shaft 3 and the compressor rotor 21a are coupled by a tubular shaft 21 which includes a flanged part 15 bolted to the shaft 21 and held in place by a locking bolt 18 as described in Specification 582,123. [Group XXIII]. The part 15 and the stub shaft 3 have a number of interengaging splines 16. The inner ball-race is held between bushes 48, 50, surrounding a bush 44 and is clamped by a flanged end nut 52. The ...

External cooling arrangement for the casings of hot steam and gas turbines

... 740,024. Turbine casings. AKT.-GES. BROWN, BOVERIE & CIE. Feb. 4, 1954 [Feb. 5, 1953], No. 3297/54. Class 110(3) A cooling arrange- ,ment for the casing 3a, 3b of a steam or gas turbine comprises ducts 7 formed by corrugated half-tubes which are welded externally to the casing and along which a cooling medium is passed. Said ducts are connected at their ends to larger, corrugated ducts 8 also welded to the casing, the ducts 8 being provided respectively with a steam or air inlet pipe 9 and an outlet pipe 10.

COOLING SYSTEM FOR TURBOCHARGER

Improvements in or relating to gas turbines

... 1,126,469. Turbines. ENGLISH ELECTRIC CO. Ltd. Sept. 24, 1965 [Sept. 24, 1964], No.21631/67. Divided out of 1,126,467. Headings F1G and F1T. In a gas turbine the stator blades 11, 13 are supported concentrically within outer casing 17 by frusto-conical members 20, 40 having a relatively thin control portion 21, 41 and thick end rings 22, 23 and 42, 43. The thin portions 21, 41 enable the blades 11, 13 and rings 23, 43 to expand radially on heating, with the blades maintained concentric with the outer casing. Cooling air is delivered through conduit 59, chamber 60 and passages 63, 72 to circumferential spaces 65, 73, whereafter it flows into the main stream to provide a cooling film over the platforms 29, 49 and shroud rings 27, 47. Additionally, air emerges into the main stream through a cooling space enclosed by a sheet metal cover 71.

An exhaust gas driven turbo-blower

... 513,849. Centrifugal pumps. MASCHINENFABRIK AUGSBURG-NURNBERG AKT.-GES. March 22, 1938, No. 8770. Convention date. March 23, 1937. [Class 110 (i)] In an exhaust gas driven turbo-blower. spaced bearings 4, 5 are used to separate the revolving hollow shaft 6 carrying the turbine wheel 7 and the stationary shaft 3 secured in the blower housing 1. The ball races of the bearing 5 are rigidly secured to the shafts 3 and 6 so as to locate the rotor whilst the outer race of the bearing 4 permits linear movement of the shaft 6 due to heat expansion. The bearing 5 is so arranged that no bending stresses are applied to the shaft 6 by the blower wheel 20. The shaft 3 is kept cool by means of passages 8, 9, 10, 11 through which cooling fluid passes and by means of plates 12, 13 which prevent radiant heat from the wheel 7 reaching it. If desired, the outer race of the bearing 4 may also be mounted in a bush of non-conducting material. The shaft 6 is prevented from overheating by water injected on to ...

Air cooled bearing

A component for a gas turbine engine

A component 100 for a gas turbine engine comprises first 22 and second 24 opposing aerodynamic surfaces with a trailing edge surface 36 therebetween. The trailing edge surface forms a tapered trailing edge section 35 of the component and the trailing edge surface is provided with one or more cooling openings 38 for ejecting a cooling fluid flow. The cooling openings may be defined by a bore with an axis oblique to the trailing edge surface, and the cooling holes may be provided adjacent a trailing edge of the component. An internal angle may be formed between a pressure surface and the trailing edge surface, and the angel may be between 160 and 175 degrees. The trailing edge surface may be straight, concave or S-Shaped in cross section in a plane perpendicular to the longitudinal axis of the component. Also claimed is a gas turbine engine comprising the component as described. The object of the invention is to prevent blockage of cooling holes and limit disruption of aerodynamic flow over ...

GAS TURBINE ENGINE CLEARANCE CONTROL

Improvements in and relating to gas turbines

... 883,646. Gas-turbine plant. B.M.W. TRIEBWERKBAU G.m.b.H. March 23, 1960 [March 28, 1959], No. 10290/60. Class 110 (3). [Also in Group XXXIV] In a gas-turbine plant comprising reduction gearing between the turbine shaft and the load, the fuel, before entering the combustion chamber, is used to cool and lubricate the turbine and gearing. A compressor 1 draws in air at a and discharges it through an annular combustion chamber 2, with a central atomizing disc 17, to a turbine 3 driving the compressor. Reduction gearing 4 is provided to connect the turbine to the load 6. Fuel from a tank 7 is pumped by the lubricating pump 8 into a header pipe 12 and sprayed by nozzles 13 on to the members of the gear that need cooling. The suction head generated by the compressor 1 draws some of the fuel through the bearings 14 and discharges it to the combustion chamber 2. The remainder of the fuel collects in the sump 11 and is then fed through the regulating means 9 and shaft 16 to the atomizing disc 17.

Turbine units

... 822,173. Air turbine units. UNITED AIRCRAFT CORPORATION. Dec. 20, 1955 [Dec. 29, 1954], No. 36474/55. Class 110 (3). [Also in Group XXXIV] In an air turbine unit, a turbine rotor is mounted on a shaft and the exhaust from the turbine passes through an exhaust pipe, the whole bearing structure in which the shaft is mounted being located within the exhaust pipe in heat exchange relation with the turbine exhaust whereby the bearing structure is cooled. The turbine unit shown comprises a turbine rotor 26 and a fan 18 mounted on a shaft 20 and disposed within a housing 12, 14, 16. Air from the compressor of a jet propulsion engine is supplied to the turbine inlet 30 and in expanding through the turbine the air becomes cooled, the cooled air passing through outlet 37 to compartments of an aircraft. The turbine shaft 20 is mounted in bearings 52, 54 which are disposed in a bearing sleeve 46 which in turn is disposed within a cylindrical portion 48. The portion 48 is supported by three radial webs ...

GAS TURBINE ENGINE

... 1519590 Gas turbine engines; cooling ROLLS-ROYCE Ltd 1 Dec 1975 [11 Nov 1974] 48662/74 Heading F1G The invention relates to a method of cooling a wall of a gas turbine engine, the wall defining a duct 8 through which combustion gases flow. A chamber 12 is defined outwardly of the wall 11 and a partition 20 is disposed within the chamber to define an inlet region 12A, to which cooling air is supplied through inlet 16, and an outlet region 12B from which the cooling air finally discharges through outlet 17. The partition is in the form of a corrugated member which defines projections 20A which project towards the outer surface 11B of the wall 11 and discharge orifices 15 are formed in the projections through which cooling air discharges on to the surface 11B, the air then passing to the spaces 21 between the projections, finally discharging through outlets 17 which open into the inlet guide vanes 8 from which the air discharges through outlets along the trailing edge thereof. In Figs. 2 and ...

MEANS FOR COOLING A SURFACE BY THE IMPINGEMENT OF A COOLING FLUID

Turbine engine including a heat exchanger formed in a platform

A turbine engine 10 of an aircraft including a primary air flow duct 16 and a secondary air flow duct 20. The secondary duct including a stator 52 including a plurality of blades 54 distributed around a main axis (A) of the turbine engine, and which includes inter-blade platforms (80, fig 2), each one of which is located between the radially internal ends (56, fig 2) or between the radially external ends (58, fig 2) of two adjacent blades. Each platform including a wall (82, fig 2) partially delimiting the secondary duct, and including a fluid circuit 40, which includes a heat exchanger 44 formed by at least one of the platforms, characterised in that the platform includes a tube (84, fig 3) that has a fluid inlet port (90, fig 3) and a fluid outlet port (92, fig 3). The fluid circuit includes a distributor (104, fig 8) associated with each port of at least one platform with the rest of the fluid circuit and of which each distributor is axially offset with respect to said platform along ...

SAFETY DIVICE FUER A PLANT TO THE SUPPLY FROM BLAST FURNACE GASES TO A TURBINE

Internal insulation for high-temperature steam turbines

ABGASTURBOLADER

TIFF 00000006.TIF 297 212 ...

IMPELLER OF A TURBINE WITH SHOVEL AND AT LEAST ONE COOLING DUCT

GAP COOLING BETWEEN COMBUSTION CHAMBER WALL AND TURBINE WALL OF A GAS TURBINE SYSTEM

INNERE ISOLATION FUER HOCHTEMPERATUR-DAMPFTURBINEN

SECURITY CONCEPT FOR A STEAM TURBINE

INTERNAL ISOLATION FOR HIGH-TEMPERATURE STEAM TURBINES

GAS TURBINE JET ENGINE WITH A HIGH PRESSURE COMPRESSOR.

COOLING OF THE SIDE PANELS OF TUBINENLEITAPPARATSEGMENTEN

Apparatus and method for cooling a combustor cap

A combustor includes an end cap having a perforated downstream plate and a combustion chamber downstream of the downstream plate. A plenum is in fluid communication with the downstream plate and supplies a cooling medium to the combustion chamber through the perforations in the downstream plate. A method for cooling a combustor includes flowing a cooling medium into a combustor end cap and impinging the cooling medium on a downstream plate in the combustor end cap. The method further includes flowing the cooling medium into a combustion chamber through perforations in the downstream plate.

Gas Turbine Nozzle Arrangement and Gas Turbine

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

Cooled gas turbine engine member

A gas turbine engine is disclosed having a vane disposed in a flow path of a gas turbine engine component, for example a gas turbine engine compressor. The vane is in thermal contact with a heat tube that extends through a wall of the engine component and into a space in which a thermal fluid passes. The thermal fluid can be at a different temperature than the vane such that heat is transferred between the two. In one embodiment the vane forms part of an intercooler for a compressor of the gas turbine engine. The vane can have a fin disposed at the end of the heat tube to facilitate a heat transfer.

Gas turbine engine

A gas turbine engine including a segment of an annular guide vane assembly is provided. When the engine is used, the segment directs hot combustion gases onto rotor blades of the engine. The segment includes a platform disposed at a side of the segment radially inward/outward with respect to the axis of rotation of the engine. The platform has a trailing edge portion downstream with respect to the flow of hot combustion gases through the segment, the trailing edge portion includes a rail that extends radially inwardly/outwardly from the trailing edge portion. The engine also includes a support and cooling arrangement for supporting the segment and directing a cooling fluid to cool the segment. The arrangement is located radially inward/outward of the platform, and includes a flange part that extends radially outwardly/inwardly from the arrangement. The arrangement further includes a leaf seal and a retaining pin.

Passive cooling system for a turbomachine

A turbomachine includes a housing having an outer surface and an inner surface that defines an interior portion. The housing includes a fluid plenum. A rotating member is arranged within the housing. The rotating member includes at least one bucket having a base portion and a tip portion. A stationary member is mounted to the inner surface of the housing adjacent the tip portion of the at least one bucket. At least one fluid passage passes through at least a portion of the stationary member. The at least one fluid passage includes a fluid inlet fluidly coupled to the fluid plenum and a fluid outlet exposed to the interior portion. The fluid outlet being configured and disposed to direct a flow of fluid toward the tip portion of the at least one bucket.

Turbine combustion system cooling scoop

A scoop ( 54 ) over a coolant inlet hole ( 48 ) in an outer wall ( 40 B) of a double-walled tubular structure ( 40 A, 40 B) of a gas turbine engine component ( 26, 28 ). The scoop redirects a coolant flow ( 37 ) into the hole. The leading edge ( 56, 58 ) of the scoop has a central projection ( 56 ) or tongue that overhangs the coolant inlet hole, and a curved undercut ( 58 ) on each side of the tongue between the tongue and a generally C-shaped or generally U-shaped attachment base ( 53 ) of the scoop. A partial scoop ( 62 ) may be cooperatively positioned with the scoop ( 54 ).

Systems and methods for cooling high pressure and intermediate pressure sections of a steam turbine

The present application provides a section cooling system for a steam turbine to limit a leakage flow therethrough. The section cooling system may include a first pressure flow extraction from a first section to a shaft packing location between the first section and a second section and a rotor aperture extending towards the first section. The first pressure flow extraction diverts the leakage flow from the first section into the rotor aperture so as to limit the leakage flow to the second section.

Seal plate with cooling passage

A seal assembly for a gas turbine engine includes a carbon seal and a seal plate. The seal plate has a contact face configured to slidably engage the carbon seal. The seal plate has a surface disposed on an opposing side of the seal plate from the contact face. The surface forms a portion of a single passage that extends an entire length of the seal plate. The continuous cooling provided by single passage along the seal plate allows for a more uniform temperature profile along the contact face of seal plate.

TURBINE BLADE TIP CLEARANCE CONTROL

An aircraft engine for use in a low-bypass turbofan application has a high pressure turbine having a blade and an engine casing disposed about the blade. A shield is disposed around the casing adjacent to the blade to create an area between the shield and the casing. A gate selectively controls entry of cooling air into the area and may be closed if the engine is maneuvering and may be open if cruising. 1. An aircraft engine for use in a fighter jet , said aircraft engine comprising;a high pressure turbine having a blade,an engine casing disposed about said blade,a shield disposed around said casing adjacent to said blade and creating an area between said shield and said casing,a gate for selectively controlling entry of cooling air into said area wherein said gate may be closed if said engine is maneuvering and wherein said gate may be open if cruising.2. The aircraft engine of wherein gate may be partially open if said engine is being operated in a steady state.3. The aircraft engine of wherein said gate is built into a front of said shield.4. The aircraft engine of wherein said gate comprises an opening and a strap having a slot claim 1 , said strap being movable relative to said opening such that said slot and said opening may be in register with each other.5. The aircraft engine of wherein said opening is disposed in front of said shield.6. The aircraft engine of wherein said opening has a race therein for holding said strap.7. The aircraft engine of wherein said strap moves within said race for moving said slots of said strap into and out of register with said openings.8. The aircraft engine of wherein said shield and said strap form an annulus.9. The aircraft engine of wherein said cooling air is fan air.10. A cooling system for an aircraft engine for use in a fighter jet claim 1 , the aircraft engine having a high pressure turbine having a blade and an engine casing disposed about said blade claim 1 , said cooling system comprising;a shield for disposal ...

AEROFOIL COOLING ARRANGEMENT

An aerofoil typically for a blade or vane for a gas turbine engine comprises a pressure wall and a suction wall, at least one of the pressure and suction walls comprise corrugations and a coolant hole on an inner surface, the corrugations define a downstream surface and the coolant hole having an inlet defined in the downstream surface. 1. An aerofoil for a gas turbine engine ,the aerofoil comprises a pressure wall and a suction wall,at least one of the pressure and suction walls comprise corrugations and a coolant hole on an inner surface,the corrugations define a downstream surface with respect to a coolant flow and the coolant hole having an inlet defined in the downstream surface.2. The aerofoil of wherein the corrugations define an upstream surface.3. The aerofoil of wherein the wall defines a passage therein and the coolant hole connects between the inlet and the passage.4. The aerofoil of wherein the coolant flows through the aerofoil and the downstream surface is angled between 20° and 90° relative to the direction of coolant flow.5. The aerofoil of wherein the coolant flows through the aerofoil and the downstream surface is angled α is approximately 45° relative to the direction of coolant flow.6. The aerofoil of wherein the coolant flows through the aerofoil and the upstream surface is angled θ between 20° and 70° relative to the direction of coolant flow.7. The aerofoil of wherein the coolant flows through the aerofoil and the upstream surface is angled θ is approximately 45° relative to the direction of coolant flow.8. The aerofoil of wherein the corrugations define any one or more the sectional profiles generally sinusoidal claim 1 , rounded triangular claim 1 , trapezoidal or saw-toothed.9. The aerofoil of wherein the corrugations define a peak and a cross-section of the passage has an apex which is adjacent the peak.10. The aerofoil of wherein the corrugations define an amplitude or height of the corrugations approximately 0.7 mm.11. The aerofoil of ...

STEAM TURBINE ARRANGEMENT OF A THREE CASING SUPERCRITICAL STEAM TURBINE

A steam turbine arrangement is provided and includes a first steam turbine with a first steam outlet, wherein the first steam turbine exhibiting a first sealing leakage, a second steam turbine exhibiting a second sealing leakage, and a third steam turbine with a cooling steam inlet, a functional device and a cooling arrangement. The cooling arrangement is coupled to the cooling steam inlet. The cooling arrangement is adapted for guiding a cooling steam to the functional device for cooling purposes. A steam pipe is coupled to the first steam turbine and to the second steam turbine such that a first steam being provided by the first sealing leakage and a second steam being provided by the second sealing leakage is gathered to a cooling steam in the steam pipe. The steam pipe is coupled to the cooling steam inlet such that the cooling steam is injectable to the cooling arrangement. 1. A steam turbine arrangement , comprisinga first steam turbine with a first steam outlet, the first steam turbine exhibiting a first sealing leakage;a second steam turbine exhibiting a second sealing leakage;a third steam turbine with a cooling steam inlet, a functional device and a cooling arrangement;a steam pipe, coupled to the first steam turbine and to the second steam turbine such that a first steam provided by the first sealing leakage and a second steam provided by the second sealing leakage is gathered to a cooling steam in the steam pipe; andan extraction control valve,wherein the cooling arrangement is coupled to the cooling steam inlet,wherein the cooling arrangement is adapted for guiding a cooling steam to the functional device for cooling purposes,wherein the steam pipe is coupled to the cooling steam inlet such that the cooling steam is injectable to the cooling arrangement, andwherein the extraction control valve which is coupled between the first steam outlet of the first steam turbine and the steam pipe such that an extraction of a first control steam from the first ...

TURBINE BLADE AND GAS TURBINE HAVING THE SAME

A turbine blade of the present invention includes a blade body mounted on a rotor body so as to extend outward from the rotor body in a radial direction of the rotor body and a chip shroud fixed to the outside of the blade body in the radial direction. A cooling passage which extends in the radial direction of the rotor body and in which a cooling medium circulates is formed in the blade body. The chip shroud includes a shroud body where a recess opened to the outside in the radial direction and communicating with the cooling passage is formed on an outer peripheral end face, and a plug that includes a plurality of plug pieces closing an opening of the recess in cooperation with each other by being inserted into mounting grooves formed on side surfaces of the recess. 1. A turbine blade comprising:a blade body that is mounted on a rotor body so as to extend outward from the rotor body in a radial direction of the rotor body; anda chip shroud that is fixed to the outside of the blade body in the radial direction,wherein a cooling passage which extends in the radial direction of the rotor body and in which a cooling medium circulates is formed in the blade body, andthe chip shroud includes a shroud body where a recess opened to the outside in the radial direction and communicating with the cooling passage is formed on an outer peripheral end face, and a plug that includes a plurality of plug pieces closing an opening of the recess in cooperation with each other by being inserted into mounting grooves formed on side surfaces of the recess.2. The turbine blade according to claim 1 ,wherein the recess extends in a direction along the outer peripheral end face as a longitudinal direction thereof,the mounting grooves are formed on the pair of side surfaces along the longitudinal direction, andthe plurality of plug pieces close the opening of the recess by lining up in the longitudinal direction so as to come into contact with each other.3. The turbine blade according to ...

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.

GAS TURBINE OUTER CASE ACTIVE AMBIENT COOLING INCLUDING AIR EXHAUST INTO A SUB-AMBIENT REGION OF EXHAUST FLOW

A gas turbine engine including an outer case extending circumferentially around the central longitudinal axis. A cooling channel is associated with the outer surface of the outer case, the cooling channel having a channel inlet and a channel outlet. An air duct is provided including an inlet end in fluid communication with the channel outlet and an outlet end in fluid communication with an exhaust gas flow from a turbine section of the gas turbine engine. An exit structure is located at the air duct outlet end, and the exit structure provides a sub-ambient pressure at the air duct outlet end to induce a flow from the air duct inlet end to the air duct outlet end. 1. A gas turbine engine comprising:an outer case defining a central longitudinal axis, and an outer surface of said outer case extending circumferentially around the central longitudinal axis;a cooling channel associated with said outer surface of said outer case, said cooling channel having a channel inlet and a channel outlet;an air duct including an inlet end in fluid communication with the channel outlet and an outlet end in fluid communication with an exhaust gas flow from a turbine section of said gas turbine engine; andan exit structure at said air duct outlet end, said exit structure providing a sub-ambient pressure at said air duct outlet end to induce a flow from said air duct inlet end to said air duct outlet end.2. The gas turbine engine of claim 1 , wherein said exit structure interacts with a portion of said exhaust gas flow passing over said exit structure for effecting a reduced pressure at said outlet end to draw air from said cooling channel into said air duct.3. The gas turbine engine of claim 2 , wherein said exit structure partially covers said air duct outlet end claim 2 , shielding an upstream side of said outlet end and defining a downstream facing opening adjacent a downstream side of said outlet end.4. The gas turbine engine of claim 1 , wherein said exit structure produces a jet ...

COMPONENT HAVING COOLING CHANNEL WITH HOURGLASS CROSS SECTION

A cooling channel (B, -) cools inner surfaces () of exterior walls () of a component (). Interior side surfaces () of the channel converge to a waist (W), forming an hourglass shaped transverse profile (). The inner surfaces () may have fins () aligned with the coolant flow (). The fins may have a transverse profile (A, B) highest at mid-width of the inner surfaces (). Turbulators () may be provided on the side surfaces () of the channel, and may urge the coolant flow toward the inner surfaces (). Each turbulator () may have a peak () that defines the waist of the cooling channel. Each turbulator may have a convex upstream side (). These elements increase coolant flow in the corners (C) of the channel to more uniformly and efficiently cool the exterior walls (). 1. A component comprising an interior cooling channel , the cooling channel further comprising:first and second inner surfaces of respective first and second exterior walls of the component; andfirst and second side surfaces spanning between the inner surfaces,wherein a transverse section of the channel has an hourglass-shaped profile in which the side surfaces taper toward each other to a waist that is narrower than a width of each of the first and second inner surfaces; andwherein an overall direction of a coolant flow in the channel is normal to the hourglass-shaped profile.2. The component of claim 1 , wherein the first and second inner surfaces are parallel to respective first and second portions of exterior surfaces of the respective exterior walls.3. The component of claim 1 , wherein the first and second exterior walls are respectively pressure and suction sides of a turbine airfoil.4. The component of claim 1 , wherein the waist comprises a width of 80% or less than the width of at least one of the inner surfaces.5. The component of claim 1 , wherein the each of the side surfaces has a taper angle in the profile of at least −1 degrees toward the waist relative to a straight line between ...

AMBIENT AIR COOLING ARRANGEMENT HAVING A PRE-SWIRLER FOR GAS TURBINE ENGINE BLADE COOLING

A gas turbine engine including: an ambient-air cooling circuit () having a cooling channel () disposed in a turbine blade () and in fluid communication with a source () of ambient air: and an pre-swirler (), the pre-swirler having: an inner shroud (); an outer shroud (); and a plurality of guide vanes (), each spanning from the inner shroud to the outer shroud. Circumferentially adjacent guide vanes () define respective nozzles () there between. Forces created by a rotation of the turbine blade motivate ambient air through the cooling circuit. The pre-swirler is configured to impart swirl to ambient air drawn through the nozzles and to direct the swirled ambient air toward a base of the turbine blade. The end walls () of the pre-swirler may be contoured. 1. A gas turbine engine , comprising:an ambient-air cooling circuit comprising a cooling channel disposed in a turbine blade and in fluid communication with a source of ambient air that provides cooling fluid: and an inner shroud;', 'an outer shroud; and', 'a plurality of guide vanes, each spanning from the inner shroud to the outer shroud,', 'wherein circumferentially adjacent guide vanes define respective nozzles there between, the nozzles defining a portion of the cooling circuit, each nozzle defined by a pressure side of a first guide vane, a suction side of the adjacent guide vane, an outer end wall defined by the outer shroud, and an inner end wall defined by the inner shroud;, 'a pre-swirler, comprisingwherein forces created by a rotation of the turbine blade motivate the cooling fluid through the cooling circuit; andwherein the pre-swirler is configured to impart swirl to the cooling fluid drawn through the nozzles and to direct the swirled cooling fluid toward a base of the turbine blade.2. The gas turbine engine of claim 1 , wherein the inner shroud is formed as a monolithic body.3. The gas turbine engine of claim 1 , wherein the outer shroud is formed as a monolithic body.4. The gas turbine engine of ...

BLADE FOR A TURBO MACHINE

A blade for a turbomachine, for example a gas turbine, is provided. The blade is arranged on a turbine rotor of the gas turbine. The blade includes a root portion having two narrow sides and two broad sides, a cooling air supply passage in the root portion, and a cooling air bleed which is arranged in the root portion and is in fluid connection with the cooling air supply passage. The cooling air bleed includes a nozzle on one of the narrow sides of the root portion, wherein the nozzle is formed by a hole and wherein an axial direction of the hole is inclined upward between 92° and 135° with respect to a longitudinal direction of the blade. 113-. (canceled)14. Blade for a turbomachine , comprising:a root portion with two narrow sides and two broad sides;a cooling air supply passage in the root portion; anda cooling air bleed arranged in the root portion and in fluid connection with the cooling air supply passage;wherein the cooling air bleed comprises a nozzle on one of the narrow sides of the root portion, and wherein the nozzle is formed by a hole,wherein the blade further comprises an upper blade platform and a lower blade platform,wherein the upper blade platform and the lower blade platform are embodied as parts of a labyrinth-sealing when assembled in the turbomachine, andwherein the nozzle is arranged between the upper blade platform and the lower blade platform, andwherein an axial direction of the hole is inclined upward between 92° and 135° with respect to a longitudinal direction of the blade.15. The blade according to claim 14 , wherein the hole of the nozzle is machined into the root portion.16. The blade according to claim 14 , wherein the nozzle is arranged on a front surface of the blade.17. The blade according to claim 14 , wherein the nozzle is arranged for generating an air flow which is directed towards a platform region of an adjacent nozzle guide vane when assembled in the turbomachine.18. The blade according to claim 17 , wherein the air flow ...

TURBINE BLADE FOR A GAS TURBINE

A turbine blade for a gas turbine is provided. The quantity of coolant flowing off the rear edge thereof is set relatively simply and exactly directly upon casting the turbine blade, without reworking the cast turbine blade with respect to the setting of coolant consumption being necessary. Raised areas are situated on the inner surfaces of the intake side wall or pressure side wall, between which a throttle element is present, by means of which the quantity of coolant flowing out is set. This arrangement allows a core tool to be produced simply, by means of which the casting cores required for casting the turbine blade are produced having the desired precision in great quantities. 16-. (canceled)7. A turbine blade for a gas turbine , comprising:a main blade part, around which a hot gas flows and which comprises a suction-side wall and a pressure-side wall which extend in the direction of flow of the hot gas from a common leading edge to a trailing edge,wherein an opening for blowing out a coolant which cools the main blade part beforehand is arranged on the trailing edge, which opening is fluidically connected to a cavity arranged in the main blade part by means of a channel,wherein the channel is also delimited by an inwardly facing face of the suction-side wall and by an inwardly facing face of the pressure-side wall and a throttling element is provided for setting the quantity of coolant emerging from the opening,wherein the throttling element comprises two elevations, upstream in relation to the throughflow direction of the channel, of the opening, which are arranged offset in relation to one another, as seen in the throughflow direction of the cooling channel, and between which there is arranged the minimum throughflow cross section of the channel, andwherein the two throttling elements are each arranged each on one of the two inwardly facing faces.8. The turbine blade as claimed in claim 7 , wherein a first elevation which is arranged on the inwardly facing ...

AEROFOIL BLADE OR VANE

An aerofoil blade or vane for the turbine of a gas turbine engine is provided. The blade or vane includes an aerofoil portion which, in use, extends radially across a working gas annulus of the engine. A coolant inlet is formed at an end of the aerofoil portion for the entry of a cooling air flow into the portion. A corresponding coolant exhaust is formed at the trailing edge of the aerofoil portion for spent cooling air to flow from the portion. A passage within the aerofoil portion connects the inlet to the exhaust. A fence within the aerofoil portion extends radially and forwardly from a start position at the end of the aerofoil portion adjacent the trailing edge to an end position. The passage forms a loop which extends along one side of the fence, wraps around the end position, and extends along the other side of the fence. 1. An aerofoil blade or vane for the turbine of a gas turbine engine , the blade or vane including:an aerofoil portion which, in use, extends radially across a working gas annulus of the engine, a coolant inlet being formed at an end of the aerofoil portion for entry of a flow of cooling air into the aerofoil portion, a corresponding coolant exhaust being formed at the trailing edge of the aerofoil portion for the flow of spent cooling air from the aerofoil portion, and a passage within the aerofoil portion connecting the inlet to the exhaust, anda fence within the aerofoil portion, the fence extending radially and forwardly from a start position at said end of the aerofoil portion adjacent the trailing edge to an end position;wherein the passage forms a loop which extends along one side of the fence, wraps around the end position, and extends along the other side of the fence to connect the inlet to the exhaust.2. An aerofoil blade or vane according to which has a plurality of coolant inlets formed at the end of the aerofoil portion for entry of respective flows of cooling air into the aerofoil portion claim 1 , a plurality of corresponding ...

Methods and systems for cooling a transition nozzle

A transition nozzle for use with a turbine assembly is provided. The transition nozzle includes a liner defining a combustion chamber therein, a wrapper circumscribing the liner such that a cooling duct is defined between the wrapper and the liner, a cooling fluid inlet configured to supply a cooling fluid to the cooling duct, and a plurality of ribs coupled between the liner and the wrapper such that a plurality of cooling channels are defined in the cooling duct.

Turbine assembly and method for reducing fluid flow between turbine components

According to one aspect of the invention, a turbine assembly includes a stator and a rotor adjacent to the stator. The turbine assembly also includes a passage formed in a projection from the rotor to form a fluid curtain between the rotor and stator, wherein the fluid curtain reduces a flow of fluid into a hot gas path.

WAVE-DRIVEN BLOWER AND ELECTRIC MOTOR/GENERATOR

The present invention provides a wave-driven blower and a wave-driven generator. Embodiments can accelerate fluid flows with waves requiring with no or minimal moving parts. The waves driving the flow can be surface-thermal waves on the walls of the device. The velocity of the surface-thermal wave entrains the fluid near the surface and imparts a velocity to the fluid. Other types of waves can generate fluid flow. These other waves can be produced by variations in chemical composition, ionic concentration, chemical potential, total pressure, partial pressure and surface texture. Operating as a generator, the device extracts energy from a flowing fluid to amplify wave motions. The wave motions in turn generate electrical power or some other form of useable power 1. An apparatus for affecting fluid flow , comprising: (a) a first surface comprising a plurality of regions , each region characterized by a controllable property; (b) a control system configured to control the regions of the first surface such that the controllable properties change in a manner such that the regions in the first surface exert a force along the first surface on a fluid proximal the first surface.2. An apparatus as in claim 1 , wherein the control system is configured to change the controllable properties such that the controllable properties establish a wave moving along the first surface.3. An apparatus as in claim 1 , wherein the controllable property comprises the temperature of the region.4. An apparatus as in claim 1 , wherein the controllable property comprises the surface texture of the region.5. An apparatus as in claim 1 , wherein the controllable property comprises permeability of the region to fluid pressure.6. An apparatus as in claim 1 , wherein the controllable property comprises heat transfer between the region and fluid proximal the region.7. An apparatus as in claim 6 , wherein the controllable property comprises one or more of: ohmic heating of the region claim 6 , ...

TURBINE VANE HOLLOW INNER RAIL

A guide vane device for a turbine has an inner platform with a through hole forming a fluid channel for a cooling fluid, wherein the inner platform extends in a circumferential direction around a shaft of the turbine. The guide vane device further includes a hollow aerofoil with a cooling opening for exchanging the cooling fluid passing the through hole into or from the hollow aerofoil, wherein the hollow aerofoil is fixed to a first surface of the inner platform, and a rail with a recess with a cooling fluid passage forming a passage for the cooling fluid to the through hole, wherein the rail is fixed to a second surface of the inner platform and the rail extends along the second surface in the circumferential direction around the shaft. The cooling fluid passage has in the circumferential direction at least the dimension of the through hole. 18.-. (canceled)9. Guide vane device for a turbine , the guide vane device comprising:an inner platform with a through hole forming a fluid channel for a cooling fluid, wherein the inner platform extends in a circumferential direction around a shaft of the turbine,a hollow aerofoil with a cooling opening for exchanging the cooling fluid passing the through hole into or from the hollow aerofoil, wherein the hollow aerofoil is fixed to a first surface of the inner platform, anda rail comprising a recess with a cooling fluid passage forming a passage for the cooling fluid to the through hole, wherein the rail is fixed to a second surface of the inner platform and the rail extends along the second surface in the circumferential direction around the shaft,wherein the cooling fluid passage comprises in the circumferential direction at least the dimension of the through hole.10. The guide vane device of claim 9 , wherein the recess is larger than the through hole.11. The guide vane device of claim 9 , wherein the rail is integrally formed with the inner platform.12. The guide vane device of claim 9 , wherein the hollow aerofoil is ...

Nozzle guide vane with cooled platform for a gas turbine

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

Set of rotor disks for a turbine engine

A set of rotor disks for a turbine engine, or an airplane turboprop or turbojet, the set including disks rigidly connected together by bolts and shaped at their outer peripheries with slots for mounting blade roots. Each disk includes a rim extending radially and not having any axial flanges, two consecutive disks being connected together with help of an axially-extending ferrule that is fastened to the rims of the disks and that is fitted with a holder plate for holding the blade roots of the downstream disk.

GAS TURBINE BLADE, MANUFACTURING METHOD THEREFOR, AND GAS TURBINE USING TURBINE BLADE

Gas turbine blades which simplify the formation of cooling channels provided inside the turbine blades while simultaneously avoiding loss of turbine blade strength and rigidity due to forming of the cooling channels. Cooling channels provided in the interior of a gas turbine blade include a plurality of straight channel-like base-side elongated holes extending in a longitudinal direction at a base side of the turbine blade, a plurality of straight channel-like tip-side elongated holes extending in a longitudinal direction at a tip side of the turbine blade, and a plurality of communicating hollow portions interposed at connection portions between the two types of elongated holes to allow the two types of elongated holes to communicate with each other, and have larger cross-sectional areas than the channel cross-sectional areas of both elongated holes. The communicating hollow portions are formed to match the position of a platform portion of the turbine blade. 1. A gas turbine blade in which cooling channels are formed inside the turbine blade , and the turbine blade is cooled by causing cooling air to circulate through the cooling channels , whereinthe cooling channels comprise:a plurality of straight channel-like base-side elongated holes that extend in a longitudinal direction at a base side of the turbine blade,a plurality of straight channel-like tip-side elongated holes that extend in a longitudinal direction at a tip side of the turbine blade, anda plurality of communicating hollow portions each having a spherical or spheroidal shape and interposed at respective connection portions that connect the base-side elongated holes and respective corresponding tip-side elongated holes to individually allow the two types of elongated holes to communicate with each other, and the respective communicating hollow portions having larger cross-sectional areas than the channel cross-sectional areas of the two types of elongated holes.2. A gas turbine blade according to ...

GAS TURBINE BLADE

A gas turbine blade including a root, an airfoil with a leading edge and a trailing edge, a radial outer tip, and a pressure side and a suction side between the leading edge and the trailing edge, and a cooling air channel system extending from an air inlet opening in the root throughout the airfoil to a plurality of air outlets at the pressure side and the leading edge of the top of the tip of the airfoil. 113-. (canceled)14. A gas turbine blade , comprising:a root;an airfoil with a leading edge and a trailing edge, a radial outer tip, and a pressure side and a suction side between the leading edge and the trailing edge; anda cooling air channel system extending from an air inlet opening in the root throughout the airfoil to a plurality of air outlets at the pressure side and the leading edge of the top of the tip of the airfoil,wherein a number of air outlets per area near the leading edge of the tip is higher than the average number of air outlets per area in the top of the tip,wherein a concentration of air outlets at the top of the tip of the airfoil is higher on the pressure side than on the suction side, andwherein the plurality of air outlets closest to the trailing edge are larger in air cross section than the plurality of air outlets in the middle between the leading edge and the trailing edge.15. The gas turbine blade according to claim 14 , wherein the plurality of air outlets at the leading edge form a group of air outlets arranged at the leading edge of the tip.16. The gas turbine blade according to claim 15 , wherein the shortest distance between the group and an air outlet on the pressure side closest to the group is larger than the diameter of the group.17. The gas turbine blade according to claim 14 , wherein the plurality of air outlets at the pressure side of the top of the tip are arranged in a row completely inside a rib at the pressure side of the tip leaving the thickness of the rib untouched.18. The gas turbine blade according to claim 14 , ...

System and Method for Cooling Gas Turbine Components

The present application provides a cooling system for a gas turbine. The cooling system may include a source of CO, a stator blade cooling system positioned about a casing of the gas turbine and in communication with the source of COand a number of stator blades, and a rotor blade cooling system positioned about a rotor shaft of the gas turbine and in communication with the source of COand a number of rotor blades. A first portion of a flow of COmay flow through the stator blade cooling system and returns to the source of COin a first closed loop and a second portion of the flow of COmay flow through the rotor blade cooling system and returns to the source of COin a second closed loop. 1. A cooling system for a gas turbine , comprising{'sub': '2', 'a source of CO;'}{'sub': '2', 'a stator blade cooling system positioned about a casing of the gas turbine and in communication with the source of COand a plurality of stator blades; and'}{'sub': '2', 'a rotor blade cooling system positioned about a rotor shaft of the gas turbine and in communication with the source of COand a plurality of rotor blades;'}{'sub': 2', '2', '2', '2, 'wherein a first portion of a flow of COflows through the stator blade cooling system and returns to the source of COin a first closed loop and wherein a second portion of the flow of COflows through the rotor blade cooling system and returns to the source of COin a second closed loop.'}2. The cooling system of claim 1 , wherein the stator blade cooling system comprises a plurality of stator blades with an internal stator plenum therein.3. The cooling system of claim 2 , wherein the stator blade cooling system comprises a stator coolant inlet plenum in communication with the source of COand the plurality of stator blades.4. The cooling system of claim 2 , wherein the internal stator plenum comprises a plurality of inlet flow channels leading to a cap plenum and a plurality of outlet flow channels.5. The cooling system of claim 2 , wherein the ...

GAS TURBINE BLADE AND GAS TURBINE HAVING THE SAME

Provided is a gas turbine blade capable of improving the heat-conducting capacity of a serpentine channel. In a gas turbine blade () including a serpentine channel in which a plurality of cooling channels (), extending from the base end side to the distal end side of the blade, are provided from the leading edge to the trailing edge of the blade, at least two of these cooling channels () being connected in a folded manner at the base end or distal end, the serpentine channel is formed such that the channel cross-sectional area becomes sequentially smaller from the cooling channel () provided at the extreme upstream side of the serpentine channel to the cooling channel () provided at the extreme downstream side. 1. A gas turbine blade comprising a serpentine channel in which a plurality of cooling channels , extending from the base end to the distal end of the blade , are provided from the leading edge to the trailing edge of the blade , at least two of these cooling channels being connected in a folded manner at the base end or the distal end ,a first wall portion that partitions a first cooling channel located at the leading edge side and a second cooling channel located adjacent to the trailing edge side of the first cooling channel;a second wall portion that partitions the second cooling channel and a third cooling channel located adjacent to the trailing edge side of the second cooling channel; anda third wall portion that partitions the third cooling channel and a fourth cooling channel located adjacent to the trailing edge side of the third cooling channel;wherein the serpentine channel is formed by the second to fourth cooling channels such that the second cooling channel is provided at the extreme downstream side;the first wall portion and the third wall portion are arranged such that the distance therebetween becomes greater from the pressure side towards the suction side of the blade;the second wall portion extends substantially parallel to the third wall ...

Gas turbomachine including a counter-flow cooling system and method

A gas turbomachine includes a casing assembly surrounding a portion of the gas turbomachine and a counter-flow cooling system arranged within the casing. The counter-flow cooling system is configured and disposed to guide cooling fluid through the casing assembly in a first axial direction and return cooling fluid through the casing assembly in a second axial direction that is opposite the first axial direction.

COOLING DEVICE FOR A JET ENGINE

The present invention relates to a cooling device for a jet engine having an axial compressor with several compressor stages including a rotor with rotor blades, a stator with stator vanes and an annulus. In order to reduce the temperature of the components at the outlet of the high-pressure compressor by simple measures and hence to increase the efficiency of a jet engine, a slot-like branch opening surrounding the rotor for a cooling airflow diverted from the main airflow into a first cavity upstream of the rotor is provided upstream of the last compressor stage of the axial compressor, with passage openings being arranged in the rotor for passing on the diverted cooling airflow from the first cavity into a second cavity downstream of the rotor. 1. Cooling device for a jet engine having an axial compressor with several compressor stages including a rotor with rotor blades , a stator with stator vanes and an annulus , characterized in that upstream of the last compressor stage of the axial compressor , a slot-like branch opening surrounding the rotor is provided for a cooling airflow diverted from the main airflow into a first cavity upstream of the rotor and that passage openings are arranged in the rotor for passing on the diverted cooling airflow from the first cavity into a second cavity downstream of the rotor.2. Cooling device in accordance with claim 1 , characterized in that the passage openings pass circumferentially/axially through the rotor and are arranged concentrically to the rotor axis.3. Cooling device in accordance with claim 1 , characterized in that a heat shield with sealing lips and heat shield openings is provided in the first cavity between the rotor and the stator and that the cooling airflow is passed through the heat shield openings and through passage openings on the axial blade mountings of the rotor blades into the second cavity.4. Cooling device in accordance with claim 1 , characterized in that the slot-like branch opening is provided ...

Near-Wall Serpentine Cooled Turbine Airfoil

Certain exemplary embodiments can provide a serpentine coolant flow path formed by inner walls in a cavity between pressure and suction side walls of a turbine airfoil and/or can be adapted to provide cooling matched to the heating topography of the airfoil, minimize differential thermal expansion, revive the coolant, and/or minimize the flow volume needed. 1. A turbine airfoil comprising:a pressure side wall and a suction side wall connected to each other along leading and trailing edges;a cavity disposed between the pressure and suction side walls;a continuous serpentine cooling flow path formed by first and second inner walls in the cavity; [ a) a cooling inlet channel that extends span-wise along at least a portion of the pressure side wall;', 'b) a forward pressure side near-wall channel along a forward portion of the pressure side wall;', 'c) a leading edge near-wall channel;', 'd) a forward suction side near-wall channel along a forward portion of the suction side wall; and', 'e) a loop channel routed between the first and second inner walls;, 'the continuous serpentine cooling flow path routes a coolant flow in the following sequence as seen in a transverse section of the airfoil, 'the cooling inlet channel, as seen in the transverse section, is adjacent to the pressure side wall at position between 30% and 70% of a chord length from the leading edge of the airfoil;', 'the first inner wall comprises a first end joined to an inner surface of the pressure side wall at a position between 50% and 75% of a chord length from the leading edge; and', 'the first inner wall extends span-wise along at least a portion of the airfoil., 'wherein2. The turbine airfoil of claim 1 , wherein the continuous serpentine cooling flow path routes the coolant flow through an intermediate suction side near-wall channel along an intermediate portion of the suction side wall.3. The turbine airfoil of claim 1 , wherein the continuous serpentine cooling flow path routes the coolant flow ...

WIND TURBINE HAVING HELIPLATFORM ARRANGEMENT AND METHOD OF USING SAME

A wind turbine for generating electrical energy may include a tower, a nacelle at the top of the tower, and a rotor coupled to a generator within the nacelle. The wind turbine further includes a cooler including a spoiler and at least one cooler panel projecting above a roof of the nacelle. A heliplatform includes a support structure extending from the nacelle and at least partially integrated with the cooler. The wind turbine may also include a crane coupled to the nacelle and configured to move between a first stowed position underneath the nacelle roof and a second operational position. In the operational position, the crane is selectively positionable over the heliplatform. A method of using the wind turbine and crane is also disclosed. 1. A wind turbine , comprising:a tower;a nacelle disposed adjacent a top of the tower and including a nacelle roof; a rotor including a hub and at least one wind turbine blade, the rotor operatively coupled to a generator housed within the nacelle;a cooler including a spoiler and at least one cooler panel projecting above the nacelle roof, the cooler being operable to remove heat from an interior of the nacelle; anda heliplatform including a support structure extending from the nacelle and at least partially integrated with the cooler.2. The wind turbine according to claim 1 , wherein the platform includes a front portion extending forward of the at least one cooler panel claim 1 , and a rear portion disposed rearward of the at least one cooler panel.3. The wind turbine according to claim 2 , wherein the at least one cooler panel includes a plurality of cooler panels claim 2 , and the front portion of the platform extends through a gap formed between two of the cooler panels.4. The wind turbine according to claim 2 , wherein the at least one cooler panel includes a plurality of cooler panels claim 2 , and one of the cooler panels located adjacent to the front portion of the platform is pivotally coupled to one or more of the ...

Wind Turbine Generator System

According to the present invention, a wind turbine generator system is provided which can not only remove the influence of salt damage in case the system is established off-shore, but even if the facility becomes larger, which can also cool equipment and the generator provided in the tower and can reduce the possibility of decreasing power generation efficiency. The wind turbine generator system of the present invention comprising a rotor having a hub and blades; a generator connected with the rotor by way of a main shaft connected with the hub; a nacelle which contains at least the generator and supports the rotor pivotally by way of the main shaft; a tower on a top of which the nacelle is supported, and opposite to the top the tower is fixed to a base, wherein a heat exchanger is provided at the tower close to the base and cooling medium passes through the heat exchanger by way of a pipe arrangement, and thereby the heat of the cooling medium and the heat of air inside the tower are exchanged and the air inside the tower is cooled. 1. A wind turbine generator system comprising:a rotor having a hub and blades;a generator connected with the rotor by way of a main shaft connected with the hub;a nacelle which contains at least the generator and supports the rotor pivotally by way of the main shaft; anda tower on a top of which the nacelle is supported, and opposite to the top the tower is fixed to a base;wherein a heat exchanger is provided at the tower close to the base andwherein cooling medium passes through the heat exchanger by way of a pipe arrangement, and thereby the heat of the cooling medium and the heat of air inside the tower are exchanged and the air inside the tower is cooled.2. A wind turbine generator system according to claim 1 , wherein the wind turbine generator system is provided off-shore claim 1 , and the cooling medium is sea water.3. A wind turbine generator system according to claim 2 , wherein a pump to pump up the sea water from the sea is ...

GAS TURBINE ENGINE TURBINE VANE AIRFOIL PROFILE

A turbine vane for a gas turbine engine includes inner and outer platforms joined by a radially extending airfoil. The airfoil includes leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface. The inner and outer platforms respectively include inner and outer sets of film cooling holes. One of the inner and outer sets of film cooling holes are formed in substantial conformance with platform cooling hole locations described by one of the sets of Cartesian coordinates set forth in Tables 1 and 2. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate. The cooling holes with Cartesian coordinates in Tables 1 and 2 have a diametrical surface tolerance relative to the specified coordinates of 0.200 inches (5.08 mm). 1. A turbine vane for a gas turbine engine comprising:inner and outer platforms joined by a radially extending airfoil, the airfoil including leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface; andwherein the inner and outer platforms respectively include inner and outer sets of film cooling holes, wherein one of the inner and outer sets of film cooling holes are formed in substantial conformance with platform cooling hole locations described by one of the sets of Cartesian coordinates set forth in Tables 1 and 2, the Cartesian coordinates provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate, and the cooling holes with Cartesian coordinates in Tables 1 and 2 have a diametrical surface tolerance relative to the specified coordinates of 0.200 inches (5.08 mm).2. The turbine vane according to claim 1 , wherein the turbine vane is a first stage turbine vane.3. The turbine vane according to claim 1 , wherein the zero-coordinate corresponds to a TOBI pin hole.4. The turbine vane ...

Gas turbine engine turbine vane airfoil profile

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

GAS TURBINE ENGINE COMPONENT HAVING PLATFORM COOLING CHANNEL

A component for a gas turbine engine according to an exemplary embodiment of the present disclosure can include a platform having an outer surface and an inner surface. A cover plate can be positioned adjacent to the outer surface of the platform. The outer surface of the platform can include a pocket and the cover plate is positioned relative to the pocket to establish a platform cooling channel therebetween. 1. A component for a gas turbine engine , comprising:a platform having an outer surface and an inner surface; anda cover plate positioned adjacent to said outer surface of said platform, wherein said outer surface of said platform includes a pocket and said cover plate is positioned relative to said pocket to establish a platform cooling channel therebetween.2. The component as recited in claim 1 , wherein said platform is an inner diameter platform.3. The component as recited in claim 1 , wherein the component is a turbine vane.4. The component as recited in claim 1 , wherein at least a portion of said pocket is exposed to establish said platform cooling channel.5. The component as recited in claim 4 , wherein said portion of said pocket is a side opening of said pocket that faces a mate face of said platform.6. The component as recited in claim 1 , wherein said pocket is a cast feature of said platform.7. The component as recited in claim 1 , wherein said platform cooling channel is bound by said cover plate and said pocket on all but a single side.8. The component as recited in claim 1 , wherein said platform cooling channel extends adjacent to a pressure side of an airfoil that extends from the platform.9. The component as recited in claim 1 , wherein a pocket wall extends between said pocket and a slot of a mate face of said platform.10. The component as recited in claim 1 , wherein said pocket is enclosed by said cover plate to establish said platform cooling channel.11. The component as recited in claim 1 , wherein said platform cooling channel is a ...

SEAL APPARATUS OF TURBINE AND THERMAL POWER SYSTEM

A sealing device for a turbine has a sealing member provided in a gap between a rotor and a stator arranged to surround the rotor, and a fluid path provided within the stator, to introduce, into the stator, a cooling medium used to cool stationary blades extending radially inward from the stator, and to flow the cooling medium at least to an upstream side of the sealing member. 1. A sealing device for a turbine comprising:a sealing member provided in a gap between a rotor and a stator arranged to surround the rotor; anda fluid path provided within the stator, to introduce, into the stator, a cooling medium used to cool stationary blades extending radially inward from the stator, and to flow the cooling medium at least to an upstream side of the sealing member.2. The sealing device of claim 1 ,wherein the fluid path comprises:a first hole configured to take in the cooling medium used to cool the corresponding stationary blade; anda second hole configured to flow the cooling medium at least into the upstream side of the sealing member.3. The sealing device of claim 2 ,wherein the sealing member has a plurality of sealing fins arranged in an axial direction, andthe second hole is provided between the sealing fins in first and second stages on the upstream side of the sealing fins.4. The sealing device of claim 3 ,wherein a plurality of second holes are provided, and a part of the holes is provided between the sealing fins in stages following the second stage on the upstream side of the sealing fins.5. The sealing device of claim 3 ,wherein the second hole is provided between two sealing fins adjacent to each other, andan interval between these two sealing fins is narrowed around the second hole.6. The sealing device of claim 3 ,wherein the sealing fins are provided on an outer circumferential surface of the stationary blades, and on a surface of the stator facing rotor blades extending radially outward from an outer circumferential surface of the rotor.7. The sealing ...

AIRCRAFT TAIL REGION WITH A COOLING SYSTEM INSTALLED IN AIRCRAFT TAIL REGION

An aircraft tail region including a cooling system installed in the aircraft tail region. The cooling system comprises a cooler, which forms a section of an outer skin of the aircraft tail region, and includes coolant channels allowing a flow of ambient air therethrough, and extending from a first surface of the cooler to a second surface of the cooler. The cooling system also includes a fan system, which is adapted to convey ambient air through the coolant channels of the cooler at least in specified operating phases of the cooling system, and a first opening, which is formed in the outer skin of the aircraft tail region, and which allows, in conveying operation of the fan system, ambient air which is supplied through the coolant channels of the cooler into an interior of the aircraft tail region to be discharged back into the aircraft environment. 1. An aircraft tail region with a cooling system installed in the aircraft tail region , the cooling system comprising:a cooler forming a section of an outer skin of the aircraft tail region, and including coolant channels allowing a flow ambient air therethrough and extending from a first surface of the cooler to a second surface of the cooler,a fan system adapted to convey ambient air through the coolant channels of the cooler at least in specified operating phases of the cooling system,a first opening formed in the outer skin of the aircraft tail region, and which allows, in conveying operation of the fan system, ambient air which is supplied through the coolant channels of the cooler into an interior of the aircraft tail region to be discharged back into the aircraft environment.2. The aircraft tail region according to claim 1 , wherein the cooler forms a section of an outer skin of the aircraft tail region claim 1 , said section being arranged at a first distance from a transom of the aircraft tail region claim 1 , and the fan system being arranged at a second distance from the transom of the aircraft tail region ...

GAS TURBINE APPARATUS

Provided is a gas turbine apparatus including: a turbine unit comprising an output shaft; a cooling gas generation unit, comprising a rotation shaft, which receives power from the output shaft through the rotation shaft and generates a compressed cooling gas; a first duct unit which transfers the generated compressed cooling gas to the turbine unit; a clutch unit which controls a power transfer connection between the output shaft and the rotation shaft; and a control unit which controls the transferring of the generated compressed cooling gas. 1. A gas turbine apparatus comprising:a turbine unit comprising an output shaft;a cooling gas generation unit, comprising a rotation shaft, which receives power from the output shaft through the rotation shaft and generates a compressed cooling gasa first duct unit which transfers the generated compressed cooling gas to the turbine unit;a clutch unit which controls a power transfer connection between the output shaft and the rotation shaft; anda control unit which controls the transferring of the generated compressed cooling gas.2. The gas turbine apparatus of claim 1 , wherein the turbine unit comprises at least one expanding stage portion.3. The gas turbine apparatus of claim 1 , wherein at least one temperature sensor is disposed at the turbine unit.4. The gas turbine apparatus of claim 3 , wherein the at least one temperature sensor measures a temperature of a shaft of a rotor claim 3 , a temperature of a blade claim 3 , and a temperature of a vane claim 3 , and transmits the measured temperature to the control unit.5. The gas turbine apparatus of claim 1 , wherein the cooling gas generation unit comprises at least one compressor that receives power from the rotation shaft claim 1 , and generates the compressed cooling gas.6. The gas turbine apparatus of claim 5 , wherein the at least one compressor comprises a plurality of compressors claim 5 ,wherein the plurality of compressors generate compressed cooling gases having ...

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 HEAT EXCHANGERS AND METHODS OF ASSEMBLING THE SAME

A heat exchanger assembly for use in a gas turbine engine includes a bypass valve and at least one body portion. The body portion includes at least one de-congealing inlet channel in flow communication with the bypass valve, a plurality of cooling channels in flow communication with the bypass valve and the at least one de-congealing inlet channel, and at least one de-congealing outlet channel in flow communication with the bypass valve and the at least one de-congealing inlet channel. The bypass valve is configured to deliver a fluid between the at least one de-congealing inlet channel and the plurality of cooling channels during a first mode of operation to facilitate reducing a temperature of the fluid. The bypass valve is further configured to deliver the fluid between the at least one de-congealing inlet channel and the at least one de-congealing outlet channel during a second mode of operation. 1. A heat exchanger assembly for use in a gas turbine engine including a core gas turbine engine having an axis of rotation , a splitter circumscribing the core gas turbine engine , a fan assembly positioned upstream of the core gas turbine engine , a fan casing substantially circumscribing the fan assembly , and a bypass duct that is defined between the fan casing and the splitter , said heat exchanger assembly comprising:a bypass valve; and at least one de-congealing inlet channel in flow communication with said bypass valve;', 'a plurality of cooling channels in flow communication with said bypass valve and said at least one de-congealing inlet channel, wherein said bypass valve is configured to deliver a fluid between said at least one de-congealing inlet channel and said plurality of cooling channels during a first mode of operation to facilitate reducing a temperature of the fluid; and', 'at least one de-congealing outlet channel in flow communication with said bypass valve and said at least one de-congealing inlet channel, wherein said bypass valve is configured ...

BUCKET ASSEMBLY FOR TURBOMACHINE

Bucket assemblies are provided. The bucket assembly includes a shank, and an airfoil positioned radially outward of the shank. The bucket assembly further includes a main cooling circuit defined in the airfoil and the shank, the main cooling circuit comprising seven passages, each of the seven passages fluidly connected with an adjacent one of the seven passages. A maximum rotation number in each of the seven passages is less than or equal to approximately 0.4. 1. A bucket assembly , comprising:a shank;an airfoil positioned radially outward of the shank;a main cooling circuit defined in the airfoil and the shank, the main cooling circuit comprising seven passages, each of the seven passages fluidly connected with an adjacent one of the seven passages,wherein a maximum rotation number in each of the seven passages is less than or equal to approximately 0.4.2. The bucket assembly of claim 1 , wherein the rotation number in each of the seven passages is in a range between approximately 0.01 and approximately 0.35.3. The bucket assembly of claim 1 , wherein a first passage of the seven passages has a maximum cross-sectional area in the airfoil that is in a range between approximately 10% and approximately 20% less than a maximum cross-sectional area in the airfoil of another passage of the seven passages.4. The bucket assembly of claim 1 , further comprising an exhaust passage defined in the airfoil and in fluid communication with a seventh passage of the seven passages.5. The bucket assembly of claim 1 , further comprising a forward cooling circuit and an aft cooling circuit are each defined in the airfoil and the shank.6. The bucket assembly of claim 1 , wherein the main cooling circuit consists of seven passages.7. The bucket assembly of claim 1 , further comprising a platform positioned radially between the shank and the airfoil.8. The bucket assembly of claim 1 , wherein the airfoil has exterior surfaces defining a pressure side and a suction side extending between ...

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. 1. A method for casting a component , the method comprising:forming a refractory metal core having first and second ends and first and second sides, each side extending from the first end to the second end; an upstream end;', 'a downstream end;', 'a first side extending from the upstream end to the downstream end;', 'a second side extending from the upstream end to the downstream end and generally opposite the first; and', 'a slot positioned between a first land and a second land for receiving the first end of the refractory metal core, each land having an inner surface facing the slot and an outer surface adjacent the inner surface;, 'forming a ceramic core element, the ceramic core element comprisingsecuring the first end of the refractory metal core into the slot of the ceramic core with an adhesive, wherein the slot and the first and second lands of the ceramic core element and the first end and first and second sides of the refractory metal core form a core assembly that provides access paths to the first and second sides of the refractory metal core near the first end, and ...

AIRFOIL WITH VARIABLE TRIP STRIP HEIGHT

An airfoil component for a gas turbine engine includes an airfoil extending from a platform. At least one of the airfoil and the platform includes a cooling passage defined by a surface. A chevron-shaped trip strip extends from the surface into the cooling passage at a trip strip height along a length. The trip strip height varies along the length. A turbine vane for a gas turbine engine includes inner and outer platforms. A cooling passage is provided in the inner platform. The cooling passage is provided by first and second radially extending legs spaced circumferentially apart from one another and joined to one another by a circumferential passage. A pair of airfoils extend radially from the same inner platform. A trip strip extends from the surface into the circumferential passage at a trip strip height along a length. The trip strip height varying along the length. 1. An airfoil component for a gas turbine engine comprising:an airfoil extending from a platform, at least one of the airfoil and the platform including a cooling passage defined by a surface; anda chevron-shaped trip strip extending from the surface into the cooling passage at a trip strip height along a length, the trip strip height varying along the length.2. The airfoil component according to claim 1 , wherein the length is provided by multiple zones claim 1 , the height varying between the zones.3. The airfoil component according to claim 2 , wherein the multiple zones include first claim 2 , second and third cooling passage heights claim 2 , and the trip strip includes first claim 2 , second and third trip strip heights respectively within the first claim 2 , second and third zones.4. The airfoil component according to claim 1 , wherein the chevrons are provided by first and second legs joined to one another at an apex to provide the chevron-shape.5. The airfoil component according to claim 3 , wherein a trip strip portion within each of the multiple zone includes a p/e ratio claim 3 , wherein ...

COMPONENT AND METHODS OF FABRICATING AND COATING A COMPONENT

A component is disclosed. The component comprises a substrate comprising an outer surface and an inner surface, where the inner surface defines at least one hollow, interior space, where the outer surface defines one or more grooves, and where each of the one or more grooves extends at least partially along the surface of the substrate and has a base. One or more access holes extend through the base of a respective groove to place the groove in fluid communication with respective ones of the at least one hollow interior space. The component further comprises a coating disposed over at least a portion of the substrate surface, where the coating comprises one or more layers. At least one of the layers defines one or more permeable slots, such that the respective layer does not completely bridge each of the one or more grooves. The grooves and the coating together define one or more channels for cooling the component. 1. A component comprisinga substrate comprising an outer surface and an inner surface, wherein the inner surface defines at least one hollow, interior space, wherein the outer surface defines one or more grooves, wherein each of the one or more grooves extends at least partially along the surface of the substrate and has an opening defined at an outer surface of the substrate and a base, 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; anda coating disposed over at least a portion of the surface of the substrate, wherein the coating comprises one or more layers, wherein at least one of the layers comprises a structural coating having a non-homogenous microstructure and defines one or more permeable slots, such that the coating does not completely bridge each of the one or more grooves, and wherein the grooves and the coating together define one or more channels within the substrate for cooling ...

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 ...

GAS TURBINE ENGINE AIRFOIL SQUEALER POCKET COOLING HOLE CONFIGURATION

A gas turbine engine airfoil includes a body that provides an exterior airfoil surface that extends in a radial direction to a tip. The exterior surface has a leading edge in a forward direction and a trailing edge in an aft direction. The tip includes a squealer pocket that has a recess surface. A cooling passage is arranged in the body. Each of the cooling holes extends from an inlet at the cooling passage to an outlet at the recessed surface. The inlet and outlet are arranged at an angle in an angular direction relative to the recessed surface. The angular direction is toward at least one of the forward and aft directions. 1. A gas turbine engine airfoil comprising:a body that provides an exterior airfoil surface that extends in a radial direction to a tip, the exterior surface has a leading edge in a forward direction and a trailing edge in an aft direction, the tip includes a squealer pocket that has a recess surface, a cooling passage is arranged in the body, and each of the cooling holes extend from an inlet at the cooling passage to an outlet at the recessed surface, the inlet and outlet are arranged at an angle in an angular direction relative to the recessed surface, the angular direction is toward at least one of the forward and aft directions.2. The airfoil according to claim 1 , wherein the angular direction of at least one of the cooling holes is toward the forward direction.3. The airfoil according to claim 1 , wherein the angular direction of at least one of the cooling holes is toward the aft direction.4. The airfoil according to claim 1 , wherein the angular direction of at least one of the cooling holes is toward the forward direction and the angular direction of at least another one of the holes is toward the aft direction.5. The airfoil according to claim 4 , wherein the exterior airfoil surface includes pressure and suction side joined at the leading and trailing edges claim 4 , wherein the angular directions of one set of cooling holes nearest ...

TURBINE BLADE

A turbine blade includes an airfoil having an internal cooling circuit. The cooling circuit includes a body cooling passage with at least one turn, and a tip cooling channel that forms a cooling barrier to thermally isolate the turn from at least a portion of the exterior surface of the airfoil. 1. A turbine blade comprising:an airfoil extending between a root and a tip, and having a pressure sidewall and a suction sidewall joined together to define a leading edge and a trailing edge;a body cooling passage located within the airfoil and having at least one tip turn located proximate the tip; anda tip cooling channel extending along the tip and enveloping the at least one tip turn to form a cooling barrier between the at least one tip turn and an exterior surface of the airfoil on all sides of the at least one tip turn.2. The turbine blade of claim 1 , wherein the body cooling passage further comprises a serpentine cooling passage.3. The turbine blade of claim 1 , wherein the tip cooling channel is in fluid communication with the body cooling passage downstream of the at least one tip turn.4. The turbine blade of and further comprising a plurality of tip turns located proximate the tip claim 1 , wherein the tip cooling channel envelops the tip turns to form a cooling barrier between the tip turns and an exterior surface of the airfoil on all sides of the tip turns.5. The turbine blade of and further comprising a plurality of body cooling passages located within the airfoil claim 1 , wherein each of the plurality of body cooling passages has at least one tip turn located proximate the tip.6. The turbine blade of and further comprising a trailing edge cooling channel having at least one film hole located along the trailing edge and fluidly coupled to the tip cooling channel.7. A turbine blade comprising:an airfoil extending between a root and a tip, and having a pressure sidewall and a suction sidewall joined together to define a leading edge and a trailing edge;a body ...

GAS TURBINE BLADE

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

AXIAL TRANSFER TUBE

What is described is a transfer tube for use with an airfoil of a gas turbine engine coupled to a platform. The transfer tube includes a main body having a first end defining an inlet configured to receive a flow of fluid and a second end defining an outlet for the flow of fluid, the main body having a curved section. The transfer tube also includes a first mating face coupled to the first end of the main body. The transfer tube also includes a second mating face coupled to the second end of the main body. At least one of the first mating face or the second mating face is configured to be coupled to a platform body of the outer diameter platform. 1. A transfer tube for use with an airfoil of a gas turbine engine coupled to a platform , comprising:a main body having a first end defining an inlet configured to receive a flow of fluid and a second end defining an outlet for the flow of fluid, the main body having a curved section;a first mating face coupled to the first end of the main body; anda second mating face coupled to the second end of the main body,wherein at least one of the first mating face or the second mating face is configured to be coupled to a platform body of the platform.2. The transfer tube of claim 1 , wherein the curved section has an angle of at least 20 degrees.3. The transfer tube of claim 2 , wherein the angle is between 80 degrees and 100 degrees.4. The transfer tube of claim 1 , further comprising a first flange coupled to the first end and a second flange coupled to the second end claim 1 , the first flange defining the first mating face and the second flange defining the second mating face.5. The transfer tube of claim 1 , wherein the transfer tube is configured to allow fluid to flow between a hook of the platform and the platform body of the platform.6. The transfer tube of claim 5 , wherein the platform is an outer diameter platform claim 5 , the first mating face is configured to be coupled to the hook such that the inlet receives the ...

GAS TURBINE COOL-DOWN PHASE OPERATION METHODS

The application describes a method of operating a gas turbine during a cool-down phase. The gas turbine provides a compressor, a combustor downstream of the compressor, and a turbine downstream of the combustor, with the turbine providing a turbine vane carrier. The method includes feeding a flow of cooling air from the compressor to the turbine vane carrier, measuring a temperature of the flow of cooling air and measuring a temperature of the turbine vane carrier. In the method, the flow of cooling air is fed at a first flow rate when the temperature of the turbine vane carrier is lower than the temperature of the cooling air, and the flow of cooling air is fed at a second flow rate when the temperature of the turbine vane carrier is higher than the temperature of the cooling air, wherein the first flow rate is higher than the second flow rate. 1. A method of operating a gas turbine during a cool-down phase , the gas turbine providing a compressor , a combustor downstream of the compressor , and a turbine downstream of the combustor , the turbine providing a turbine vane carrier , the method comprising:feeding a flow of cooling air from the compressor to the turbine vane carrier;measuring a temperature of the flow of cooling air; andmeasuring a temperature of the turbine vane carrier;wherein the flow of cooling air is fed at a first flow rate when the temperature of the turbine vane carrier is lower than the temperature of the cooling air; andwherein the flow of cooling air is fed at a second flow rate when the temperature of the turbine vane carrier is higher than the temperature of the cooling air,wherein the first flow rate is higher than the second flow rate.2. The method of claim 1 , wherein the flow of cooling air is fed from the compressor to the turbine vane carrier through a cooling unit.3. The method of claim 2 , wherein the flow of cooling air is fed from the compressor to the turbine vane carrier through a once-through cooler.4. The method of claim 2 , ...

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 , ...

CONTOURED BLADE OUTER AIR SEAL FOR A GAS TURBINE ENGINE

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

ARRANGEMENT FOR A TURBOMACHINE

An arrangement for a turbomachine is provided. The arrangement includes a vane for directing a hot gas during the operation of the turbomachine, a stator ring for securing the vane, a heat shield for protecting the stator ring from the hot gas flow wherein the heat shield is arranged in downstream direction of the hot gas flow in front of the stator ring characterized in that the heat shield comprises a plurality of channels formed therein for directing a cooling air. 19-. (canceled)10. An arrangement for a turbomachine comprising ,a vane for directing a hot gas during the operation of the turbomachine,a stator ring for securing the vane,a heat shield for protecting the stator ring from the hot gas flow wherein the heat shield is arranged in downstream direction of the hot gas flow in front of the stator ring,wherein the heat shield comprises a plurality of channels formed therein for directing a cooling air, the channels being arranged such that the cooling air is released into a hot gas flow path of the hot gas flow.11. The arrangement for the turbomachine according to claim 10 ,wherein the cooling air from the plurality of channels in the heat shield is directed into the hot gas flow path.12. The arrangement for the turbomachine according to claim 10 ,wherein the vane comprises an airfoil portion extending in a direction radial to an axis of rotation of a rotor of the turbomachine, and a root portion mounted on the stator ring.13. The arrangement for the turbomachine according to claim 12 , further comprisinga second stator ring for fixing a head portion of the vane, wherein the head portion of the vane is at an opposing end of the root portion.14. The arrangement for the turbomachine according to claim 13 ,wherein the second stator ring is located in a radially outward direction from the first stator ring with respect to the axis of rotation of the rotor of the turbomachine.15. The arrangement for the turbomachine according to claim 13 ,wherein the second stator ...

GAS TURBINE ENGINE MULTI-VANED STATOR COOLING CONFIGURATION

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

INTERIOR COOLING CONFIGURATIONS FOR TURBINE ROTOR BLADES

A turbine rotor blade that includes an interior cooling configuration having a section configuration that includes: a main channel divided into three non-overlapping segments in which an upstream segment connects to a downstream segment via a transition segment positioned therebetween; and one or more branching channels extending from the main channel via connections each makes to the transition segment. The transition segment includes a variable cross-sectional flow area that accommodates a main channel flow area reduction occurring between the upstream segment and the downstream segment. The one or more branching channels having a total branching channel flow area. The section configuration is configured according to a section channel ratio that is defined as the main channel flow area reduction divided by the total branching channel flow area, with the value of the section channel ratio being configured according a desired coolant flow characteristic. 1. A rotor blade for use in a turbine of a gas turbine that includes:an airfoil defined between a concave pressure face and a laterally opposed convex suction face, wherein the pressure face and the suction face extend axially between opposite leading and trailing edges and radially between an outboard tip and an inboard end that attaches to a root configured to couple the rotor blade to a rotor disc; a main channel divided into three non-overlapping segments in which an upstream segment connects to a downstream segment via a transition segment positioned therebetween, wherein the transition segment comprising a variable cross-sectional flow area that accommodates a main channel flow area reduction occurring between the upstream segment and the downstream segment; and', 'one or more branching channels extending from the main channel via connections each of the one or more branching channels makes to the transition segment, wherein the one or more branching channels comprise a total branching channel flow area;, 'an ...

Gas turbine blade

A gas turbine blade includes a plurality of guide ribs spaced apart from each other in a movement direction of a cooling fluid in order to guide movement of the cooling fluid flowing along a cooling passage formed in the turbine blade, and an opening formed in each of the guide ribs in order to guide a portion of the cooling fluid to a bottom of the cooling passage between the guide ribs or to side walls adjacent to the bottom.

MODULAR ANNULAR HEAT EXCHANGER

An annular duct including a modular annular heat exchanger for a gas turbine engine is provided, where the modular annular heat exchanger includes a plurality of radial modules in circumferentially adjacent arrangement. Each radial module includes a cooled fluid inlet plenum segment, a plurality of blades, and a cooled fluid outlet plenum segment. The plurality of blades is configured in circumferentially adjacent arrangement and defines an angular space that is conformal between each circumferentially adjacent blade. The cooled fluid inlet plenum segment, the plurality of blades, and the cooled fluid outlet plenum segment are in serial axial flow arrangement and define an internal cooled fluid flowpath and an external cooling fluid flowpath parallel to the internal cooled fluid flowpath. Each radial module further includes an inner annular ring segment and an outer annular ring segment. The inner annular ring segment and the outer annular ring segment define a plurality of blade retainers. The blade retainers define an axial, radial, and circumferential position of the blades, the cooled fluid inlet plenum segment, and the cooled fluid outlet plenum segment. 1. An annular duct including a modular annular heat exchanger for a gas turbine engine , comprising: a cooled fluid inlet plenum segment;', 'a plurality of blades, configured in circumferentially adjacent arrangement and defining an angular space that is conformal between each circumferentially adjacent blade;', 'a cooled fluid outlet plenum segment, wherein the cooled fluid inlet plenum segment, the plurality of blades, and the cooled fluid outlet plenum segment are in serial axial flow arrangement and define an internal cooled fluid flowpath and an external cooling fluid flowpath parallel to the internal cooled fluid flowpath;', 'an inner annular ring segment; and', 'an outer annular ring segment, wherein the inner annular ring segment and the outer annular ring segment define a plurality of blade retainers, ...

TURBINE COMPONENT WITH SHAPED COOLING PINS

A cooling circuit to receive a cooling fluid includes at least one shaped cooling pin disposed in the cooling circuit. The at least one shaped cooling pin has a first end and a second end extending along an axis. The first end has a first curved surface defined by a minor diameter and the second end has a second curved surface defined by a major diameter. The first curved surface is to be upstream in the cooling fluid and the minor diameter is less than the major diameter. 1. A cooling circuit adapted to receive a cooling fluid , comprising:at least one shaped cooling pin disposed in the cooling circuit, the at least one shaped cooling pin having a first end and a second end extending along an axis, the first end having a first curved surface defined by a minor diameter and the second end having a second curved surface defined by a major diameter, the first curved surface is configured to be upstream in the cooling fluid and the minor diameter is less than the major diameter.2. The cooling circuit of claim 1 , wherein the first curved surface is spaced apart from the second curved surface by a length.3. The cooling circuit of claim 1 , wherein the first curved surface and the second curved surface are interconnected by a pair of surfaces defined by a pair of planes substantially tangent to a respective one of the first curved surface and the second curved surface.4. The cooling circuit of claim 2 , wherein claim 2 , in cross-section claim 2 , the second curved surface is defined by a first circle having a first center point and the first curved surface is defined by a second circle having a second center point claim 2 , and the length is defined between the first center point and the second center point.5. The cooling circuit of claim 4 , wherein the at least one shaped cooling pin comprises a plurality of shaped cooling pins that are arranged in a pattern that includes at least one row of a first sub-plurality of the plurality of shaped cooling pins and at least ...

ENGINE COMPONENT

A combustion engine component is disclosed. The combustion engine component comprises a body that includes a first surface in operative thermal communication with a hot combustion gas, and a second surface in operative fluid communication with a cooling fluid. Also, as disclosed in greater detail below, the second surface includes a first surface contour feature configured to increase a contact angle of a liquid on the second surface. 1. A combustion engine component , comprising:a body comprising a first surface in operative thermal communication with a hot combustion gas, and a second surface in operative fluid communication with a cooling fluid,wherein the second surface includes a first surface contour feature configured to increase a contact angle of a liquid on the second surface.2. The combustion engine component of claim 1 , wherein the first surface contour feature includes a plurality of first surface projections disposed on the second surface and individually configured as conical claim 1 , or as truncated spheres or truncated spheroids3. The combustion engine component of claim 1 , wherein the first surface contour feature includes a plurality of first surface projections disposed on the second surface including a first characteristic dimension of less than 0.02 inches (0.51 mm).4. The combustion engine component of claim 3 , wherein the first characteristic dimension is at least 0.00001 inches (0.25 μm).5. The combustion engine component of claim 3 , wherein the first characteristic dimension is selected from a height claim 3 , a width claim 3 , or a spacing between adjacent surface projections.6. The combustion engine component of claim 1 , wherein the first surface contour feature of the second surface includes a surface roughness Ra of 0.00005-0.01 inches (1.3 μm-0.25 mm) claim 1 , and RΔa of 0.00001-0.005 inches (0.25 μm-0.13 mm).7. The combustion engine component of claim 1 , wherein the second surface further includes a second surface contour ...

HOUSING STRUCTURE FOR A TURBOMACHINE, TURBOMACHINE AND METHOD FOR COOLING A HOUSING PORTION OF A HOUSING STRUCTURE OF A TURBOMACHINE

The present invention relates to a housing structure for a turbomachine, a turbomachine, and a method for cooling a housing portion of a housing structure of a turbomachine, wherein the housing structure has an outer housing wall, which is formed by at least one housing part, and an inner wall, wherein the inner wall is arranged in the radial direction inside of the housing wall and spaced apart from the housing wall and is designed to bound the main flow channel, at least partially, wherein the housing structure comprises a cooling air channel for cooling a housing portion to be cooled, and has an upstream end and a downstream end, wherein the cooling air channel extends, at least partially, between the housing wall and the inner wall. 1. A housing structure for a turbomachine , wherein the housing structure surrounds annularly , at least partially , a main flow channel of the turbomachine , in which rotating blades and guide vanes are arranged , and to delimit the turbomachine with respect to the surroundings , wherein the housing structure comprises an outer housing wall , which is formed by at least one housing part , and an inner wall , and at least one cooling air channel , which extends , at least partially , in a volume between the housing wall and the inner wall , for cooling a housing portion to be cooled of the housing wall ,wherein the housing wall delimits the turbomachine with respect to the surroundings,wherein the inner wall is arranged in the radial direction inside of the housing wall and spaced apart from the housing wall in the radial direction, and bounds, at least partially, the main flow channel together with the rotating blades and the guide vanes arranged therein,wherein, in relation to a main flow direction, with which a main flow passes through the main flow channel, the housing portion to be cooled has an upstream end and a downstream end,wherein the cooling air channel has a cooling air channel inlet for feeding cooling air into the ...

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 ...

TURBINE TIP SHROUD ASSEMBLY WITH PLURAL SHROUD SEGMENTS HAVING INTER-SEGMENT SEAL ARRANGEMENT

A shroud assembly for a gas turbine engine includes a plurality of shroud segments that are attached to a shroud support with an inter-segment joint defined between shroud segments. The shroud assembly also includes a cooling flow path cooperatively defined by the shroud support and the first shroud segment. The cooling flow path includes an internal cooling passage within the shroud segments. The cooling flow path includes an outlet chamber configured to receive flow from the internal cooling passage. The shroud assembly additionally includes a seal arrangement that extends across the inter-segment joint. The seal arrangement, the first shroud segment, and the second shroud segment cooperatively define a seal chamber that is enclosed. 1. A shroud assembly for a gas turbine engine comprising:a shroud support that extends arcuately about an axis;a plurality of shroud segments that are attached to the shroud support and that are arranged annularly about the axis at different circumferential positions with respect to the axis, the plurality of shroud segments including a first shroud segment and a second shroud segment, an inter-segment joint defined circumferentially between the first and second shroud segments;a seal arrangement that extends circumferentially across the inter-segment joint; andthe seal arrangement, the first shroud segment, and the second shroud segment cooperatively defining a seal chamber that is enclosed.2. The shroud assembly of claim 1 , wherein the intersegment joint includes a leading edge and a trailing edge that are separated apart at a distance along the axis; andwherein the seal chamber is disposed proximate the trailing edge and is spaced apart at a distance from the leading edge.3. The shroud assembly of claim 1 , wherein the seal arrangement includes a first sealing member and a second sealing member that are arranged in-series with the seal chamber separating the first sealing member and the sealing member apart at a distance.4. The ...

COOLING CONFIGURATION FOR A GAS TURBINE ENGINE AIRFOIL

A gas turbine engine airfoil includes an outer wall including a suction side, a pressure side, a leading edge, and a trailing edge, the outer wall defining an interior chamber of the airfoil. The airfoil further includes cooling structure provided in the interior chamber. The cooling structure defines an interior cooling cavity and includes a plurality of cooling fluid outlet holes, at least one of which is in communication with a pressure side cooling circuit and at least one of which is in communication with a suction side cooling circuit. At least one of the pressure and suction side cooling circuits includes: a plurality of rows of airfoils, wherein radially adjacent airfoils within a row define segments of cooling channels. Outlets of the segments in one row align aerodynamically with inlets of segments in an adjacent downstream row such that the cooling channels have a serpentine shape. 1. A gas turbine engine airfoil comprising:an outer wall including a radially inner end, a radially outer end, a suction side, a pressure side, a leading edge, and a trailing edge, the outer wall defining an interior chamber of the airfoil;cooling structure provided in the interior chamber, the cooling structure:located closer to the leading edge of the outer wall than to the trailing edge of the outer wall;defining an interior cooling cavity; andincluding a plurality of cooling fluid outlet holes, at least one of the outlet holes in communication with a pressure side cooling circuit and discharging cooling fluid from the interior cooling cavity of the cooling structure at least partially in a direction toward the leading edge of the outer wall, and at least one of the outlet holes in communication with a suction side cooling circuit and discharging cooling fluid from the interior cooling cavity of the cooling structure at least partially in a direction toward the leading edge of the outer wall;each of the pressure side cooling circuit and the suction side cooling circuit ...

Turbine engine component with an insert

A component for a turbine engine comprises a wall with a surface along which a hot airflow passes, a second surface along which a cooling airflow passes, and an insert mounted to the wall wherein the material used for the insert can have a higher temperature capability than that of the wall.

COMPLIANT ROTATABLE INTER-STAGE TURBINE SEAL

Compliant bellow seal may be axially disposed between first and second cooling plates bounding first and second cooling passages between cooling plates and first and second stage disks in turbine. Bellow seal includes two or more convolutions with oppositely facing forward and aft sealing surfaces, which may be flat, on forward and aft annular sealing walls and cylindrical annular outer and inner contact and sealing surfaces on and facing radially outwardly or inwardly from one of the convolutions. A snake bellow seal embodiment may have at least two of the convolutions being full convolutions of unequal width and a forwardmost partial convolution including the sealing wall. 1. A compliant bellow seal comprising:two or more convolutions circumscribed about an axis of rotation,oppositely facing forward and aft sealing surfaces on axially spaced apart forward and aft annular legs or sealing walls, andcylindrical annular outer and inner contact and sealing surfaces on and facing radially outwardly or inwardly with respect to the axis of rotation from one of the convolutions.2. The bellow seal as claimed in claim 1 , further comprising the outer contact and sealing surface being located on a radially outwardly extending cylindrical extension on one of the convolutions.3. The bellow seal as claimed in claim 1 , further comprising the forward and aft sealing surfaces being flat.4. The bellow seal as claimed in claim 3 , further comprising the outer contact and sealing surface being located on a radially outwardly extending cylindrical extension on one of the convolutions.5. The bellow seal as claimed in claim 1 , further comprising:the bellow seal being a snake bellow seal,at least two of the convolutions being full convolutions of unequal width, anda forwardmost partial convolution including the forward annular leg or sealing wall.6. The bellow seal as claimed in claim 5 , further comprising the outer contact and sealing surface being located on a radially inwardly ...

STEAM TURBINE COOLING UNIT

A steam turbine cooling unit for a steam turbine includes a coolant steam path provided to penetrate a casing (an outer casing and an inner casing) along a superheated steam supply tube to reach a gap; and a coolant steam supplying unit configured to supply coolant steam flowing through the coolant steam path along the superheated steam supply tube to reach the gap, and having a pressure higher than and a temperature lower than those of superheated steam to be supplied by the superheated steam supply tube. This configuration provides improved cooling efficiency. 1. A steam turbine cooling unit for a steam turbine that includes a rotor which is a rotating body extending along an axial center of rotations of the rotor , a casing configured to house the rotor , a steam path provided between the rotor and the casing in an extending direction of the rotor , a steam nozzle unit attached to the casing with a gap formed between an outer surface of the steam nozzle unit and an outer circumferential surface of the rotor , the gap having an annular shape surrounding the outer circumference of the rotor and communicating with the steam path , the steam nozzle unit including a steam nozzle chamber having an annular shape formed along internal of the steam nozzle unit and an opening facing the extending direction of the rotor from the steam nozzle chamber to communicate with the steam path , and a superheated steam supply tube to which superheated steam is supplied , the superheated steam supply tube being provided to penetrate the casing from external of the casing to communicate with the steam nozzle chamber in the steam nozzle unit , a coolant steam path provided to penetrate the casing along the superheated steam supply tube to reach the gap; and', 'a coolant steam supplying unit configured to supply coolant steam flowing through the coolant steam path along the superheated steam supply tube to reach the gap, the coolant steam having a pressure higher than and a temperature ...

GAS TURBINE ENGINE WITH THERMOELECTRIC INTERCOOLER

A gas turbine engine includes a compressor, a cooling source, and a thermoelectric intercooler adapted for selective operation in response to operational states of the gas turbine engine. 1. A gas turbine engine for generating drive from combustion of fuel , comprisinga compressor including a plurality of rotating stages each adapted to compress air,a cooling source adapted to provide coolant to the compressor of the gas turbine engine, anda thermoelectric intercooler located axially between rotating stages of the compressor along a central engine axis, the thermoelectric intercooler including a compressed air passageway fluidly coupled to the compressor to pass compressed air of the compressor therethrough, a coolant passageway fluidly coupled to the cooling source to pass coolant of the cooling source therethrough, and a thermoelectric section configured in thermal communication with each of the compressed air passageway and the coolant passageway,wherein the thermoelectric section is disposed between the compressed air passageway and the coolant passageway.2. The gas turbine engine of claim 1 , further comprising a controller configured to determine an operational state of the gas turbine engine and to selectively apply voltage across the thermoelectric section based on the operational state of the gas turbine engine.3. The gas turbine engine of claim 2 , wherein the cooling source is one of a fuel system of the gas turbine engine and a cooling air stream.4. The gas turbine engine of claim 2 , wherein the gas turbine engine is configured to provide propulsion for an aircraft and the operational state of the gas turbine engine includes one of ground idle claim 2 , takeoff claim 2 , climb claim 2 , cruise claim 2 , and flight idle.5. The gas turbine engine of claim 4 , wherein the controller is configured to apply voltage across the thermoelectric section to direct current through the thermoelectric section in a first direction in response to determination that the ...

DEVICE AND METHOD FOR MONITORING THE LIFETIME OF A HYDRAULIC APPARATUS OF AN AIRCRAFT

The invention relates to a device for monitoring the lifetime of at least one hydraulic apparatus of an aircraft that is subject to ventilations in hydraulic pressure during flight, comprising an interface for receiving measurement data which are representative of hydraulic pressure (P). The invention is characterised in that the device comprises a processing device, comprising a means for detecting a pressure (P) load (SOLL) of a damaging nature, which load is defined by the fact that the pressure (P) comprises a pressure increase (ΔP) that is greater than a predetermined damage threshold (S), followed by a pressure decrease (ΔP) that is greater than the threshold (S), a means for calculating a pressure variation magnitude that is equal to the maximum increase (ΔP) and the maximum decrease (ΔP), a means for projecting the magnitude onto a decreasing curve or straight line of a damage model in order to determine the permissible number of loads corresponding to the magnitude, a means for calculating a potential damage ratio that is equal to a number of reference loads divided by the permissible number, a means for increasing a count of accumulated ratios by said ratio. 1. A device for monitoring the lifetime of at least one hydraulic apparatus of an aircraft subjected to variations of hydraulic pressure in flight , the device comprising an interface for receiving measurement data representative of the hydraulic pressure of the at least one hydraulic apparatus as a function of flight time ,the at least one hydraulic apparatus comprising:a processing device comprising a detector for detecting, based on the measurement data, a pressure load of a damaging nature, defined by the fact that the pressure comprises a pressure increase, greater than a predetermined damage threshold greater than zero, followed by a pressure reduction greater than the predetermined damage threshold,a calculator for calculating a pressure variation amplitude, equal to the maximum of the absolute ...

SHROUD BLOCK SEGMENT FOR A GAS TURBINE

A shroud block segment for a gas turbine includes a main body having a leading portion, a trailing portion, a first side portion and an opposing second side portion that extend axially between the leading portion and the trailing portion. The main body further includes an arcuate combustion gas side, an opposing back side and a cooling chamber defined in the back side. A cooling plenum and an exhaust passage are defined within the main body where the exhaust passage provides for fluid communication out of the cooling plenum. An insert opening extends within the main body through the back side towards the cooling plenum. A cooling flow insert is disposed within the insert opening. The cooling flow insert comprises a plurality of cooling flow passages that provide for fluid communication between the cooling chamber and the cooling plenum. 1. A shroud block segment , comprising:a. a main body having a leading portion, a trailing portion, a first side portion and an opposing second side portion that extend axially between the leading portion and the trailing portion, an arcuate combustion gas side, an opposing back side and a cooling chamber defined in the back side;b. a cooling plenum defined within the main body;c. an exhaust passage defined within the main body, wherein the exhaust passage provides for fluid communication out of the cooling plenum;d. an insert opening that extends within the main body through the back side towards the cooling plenum; ande. a cooling flow insert disposed within the insert opening, wherein the cooling flow insert comprises a plurality of cooling flow passages that provide for fluid communication between the cooling chamber and the cooling plenum.2. The shroud block segment as in claim 1 , wherein the cooling passages are offset with respect to the exhaust passages.3. The shroud block assembly as in claim 1 , wherein the cooling passages are arranged in one of a triangular or circular pattern within the cooling flow insert.4. The shroud ...

GAS TURBINE WITH HIGH-PRESSURE TURBINE COOLING SYSTEM

The present invention relates to a gas turbine with a turbine stator wheel, which is fitted with stator vanes and includes a ring segment-shaped vane root, where the stator vanes are designed hollow and have a vane interior which can be supplied with cooling air, where a ring-shaped sealing element of an inter-stage seal is arranged radially on the inside, relative to an engine axis, on the vane root, where in the vane root at least one outflow duct is provided, characterized in that between the vane root and the sealing element an annular space extending substantially in the axial direction is formed, into which the outflow duct issues and which discharges into the area of the inter-stage seal. 1. Gas turbine with a turbine stator wheel , which is fitted with stator vanes and includes a ring segment-shaped vane root , where the stator vanes are designed hollow and have a vane interior which can be supplied with cooling air , where a ring-shaped sealing element of an inter-stage seal is arranged radially on the inside , relative to an engine axis , on the vane root , where in the vane root at least one outflow duct is provided , wherein between the vane root and the sealing element an annular space extending substantially in the axial direction is formed , into which the outflow duct issues and which discharges into the area of the inter-stage seal.2. Gas turbine in accordance with claim 1 , wherein the annular space has a substantially constant width.3. Gas turbine in accordance with claim 1 , wherein the annular space ribs extending in the flow direction are arranged.4. Gas turbine in accordance with claim 3 , wherein the ribs are arranged on the sealing element.5. Gas turbine in accordance with claim 1 , wherein the discharge area of the annular space is designed for axial outflow of the cooling air.6. Gas turbine in accordance with claim 1 , wherein the discharge area of the annular space is designed such that the outflow of the cooling air takes place at an ...

Pressure casing of a turbomachine

The invention relates to a pressure casing, which includes a plurality of casing shells which are connected in a pressure-tight manner in a parting plane by means of a flange. The casing shells are pressed together with sealing effect in the parting plane in the region of the flange by means of at least one threaded bolt which extends in a through hole through the flange perpendicularly to the parting plane. Reduced temperature differences between the flange and the connecting bolts of the flanged joint are achieved by the at least one threaded bolt being charged with a heat transfer medium over a part of its length. The heat transfer medium is supplied via holes extending through the flange.

COOLING SYSTEM OF A WIND TURBINE

A cooling system of a wind turbine is provided which includes a generator (), a housing portion (), and a hub () which is rotatable with respect to the housing portion () and drivingly connected to the generator (). The hub () supports at least one blade (). The housing portion () and the hub () enclose a cavity for accommodating at least said generator. Heat exchange means () are provided on the outer surface of said hub () for transferring at least a part of heat produced by said generator () in operation to the outside. 1. A cooling system of a wind turbine comprising:{'b': '10', 'a generator ();'}{'b': '1', 'a housing portion (); and'}{'b': 2', '1', '10', '2', '4, 'a hub () rotatable with respect to the housing portion () and drivingly connected to the generator (), said hub () supporting at least one blade ();'}{'b': 1', '2', '10, 'wherein said housing portion () and said hub () enclose a cavity for accommodating at least said generator(); and'}{'b': 2', '10, 'wherein heat exchange means are provided on an outer surface of said hub () for transferring at least a part of heat produced by said generator () in operation to outside.'}2. The cooling system as claimed in claim 1 ,{'b': 1', '2, 'wherein said cavity defined by said housing portion () and said rotatable hub () is substantially sealed or sealable to the outside.'}3. The cooling system as claimed in any of and claim 1 ,{'b': 6', '2', '2', '6, 'wherein said heat exchange means comprise at least one fin () mounted at one end thereof on the outer surface of said hub () with the other end thereof extending radially from said surface of said hub () and forming a free end of said at least one fin ().'}4. The cooling system as claimed in claim 3 ,{'b': 6', '2', '2, 'wherein said at least one fin () is arranged on the outer surface of said hub () on a line or curve which is at least partially inclined or skewed with respect to a rotational axis of said hub ().'}562262. The cooling system as claimed in any of and ...

Orifice element for turbine stator and/or rotor vanes

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

TURBINE ENGINE BLADE WITH IMPROVED COOLING

A turbine blade including a root carrying an impeller terminated by a tip in the form of a squealer tip. This impeller also includes a serpentine median circuit, including a first radial pipe collecting air at the root and that is connected by a first bend to a second radial pipe that is connected by a second bend to a third radial pipe, a cavity under the squealer tip running along the pressure side wall, extending from a central region of the tip to the trailing edge, and a radial central pipe collecting air at the root extending between at least two of the three pipes of the median circuit and directly supplying the cavity under the squealer tip. 114-. (canceled)15. A turbine vane of a turbomachine , for being mounted about an axis of rotation on a rotor disc rotating about an axis of rotation , comprising a root for mounting thereof in a cell of the disc , and a hollow blade extending from the root in a radial spanwise direction and terminating in a top forming a bathtub , the blade comprising a lower surface wall and an upper surface wall , as well as a leading edge , a trailing edge and a top wall delimiting a bottom of the bathtub , and with which the lower surface wall is connected to the upper surface wall , said blade also comprising:a paper clip-type median circuit, including a first radial duct collecting air at the root and which is connected through a first bend to a second radial duct which is connected through a second bend to a third radial duct;an under-bathtub cavity located on the side of the lower surface wall and the top wall and which extends from a central region of the top to the trailing edge;a central radial duct located on the side of the lower surface wall and which collects air at the root and extends between at least two of the three ducts of the median circuit and directly feeds the under-bathtub cavity.16. The vane of claim 15 , wherein one end of the third duct and at least part of the first bend are located between the under- ...

SEAL RUNNER FLOW DAMPER

A gas turbine engine component includes a first member, a second member rotatable relative to the first member about an axis, and a seal assembly that includes a seal supported by the first member and a seal runner that rotates with the second member relative to the seal. The seal runner includes at least one internal passage to direct cooling fluid flow through the seal runner. A restriction is associated with the at least one internal passage to restrict flow through the at least one internal passage in response to an engine condition. 1. A gas turbine engine component comprising:a first member;a second member rotatable relative to the first member about an axis;a seal assembly that includes a seal supported by the first member and a seal runner that rotates with the second member relative to the seal, and wherein the seal runner includes at least one internal passage to direct cooling fluid flow through the seal runner; anda restriction associated with the at least one internal passage to restrict flow through the at least one internal passage in response to an engine condition.2. The gas turbine engine component according to claim 1 , wherein the restriction comprises a resilient member.3. The gas turbine engine component according to claim 2 , including a damper associated with the resilient member claim 2 , wherein the resilient member moves the damper from an initial position to a restricted position to reduce a flow rate of fluid flowing through the at least one internal passage in response to the engine condition.4. The gas turbine engine component according to claim 3 , wherein the second member rotates a first speed under a low-speed engine operating condition and rotates at a second speed higher than the first speed under a high-speed engine operating condition claim 3 , and wherein the engine condition corresponds to the high-speed engine operating condition such that fluid exits an outlet of the at least one internal passage at a first flow rate when ...

GAS TURBINE ENGINE COMPONENT COOLING CHANNELS

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

Intercooled Compressor for a Gas Turbine Engine

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

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

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

METHOD AND APPARATUS FOR HANDLING PRE-DIFFUSER AIRFLOW FOR USE IN ADJUSTING A TEMPERATURE PROFILE

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

METHOD AND APPARATUS FOR COLLECTING PRE-DIFFUSER AIRFLOW AND ROUTING IT TO COMBUSTOR PRE-SWIRLERS

A gas turbine engine is provided that includes a compressor section, a combustor section with a multiple of pre-swirlers, a diffuser case module, and a manifold. The diffuser case module includes a multiple of struts within an annular flow path from said compressor section to said combustor section, wherein at least one of said multiple of struts defines a mid-span pre-diffuser inlet in communication with said annular flow path. The manifold is in communication with said mid-span pre-diffuser inlet and said multiple of pre-swirlers. 1. A gas turbine engine comprising:a compressor section;a combustor section with a multiple of pre-swirlers;a diffuser case module with a multiple of struts within an annular flow path from said compressor section to said combustor section, at least one of said multiple of struts defines a mid-span pre-diffuser inlet in communication with said annular flow path; anda manifold in communication with said mid-span pre-diffuser inlet and said multiple of pre-swirlers.2. The gas turbine engine as recited in claim 1 , wherein said manifold communicates a temperature tailored airflow.3. The gas turbine engine as recited in claim 1 , wherein said manifold communicates a temperature tailored airflow thru a heat exchanger.4. The gas turbine engine as recited in claim 3 , wherein said manifold communicates said temperature tailored airflow from said heat exchanger as buffer air.5. The gas turbine engine as recited in claim 4 , wherein said buffer air is communicated thru a buffer passage to said bearing compartment.6. The gas turbine engine as recited in claim 1 , wherein said mid-span pre-diffuser inlet supplies a temperature tailored airflow into said manifold.7. The gas turbine engine as recited in claim 1 , wherein said mid-span pre-diffuser inlet is located radially outboard of a central span section of said multiple of struts.8. The gas turbine engine as recited in claim 1 , wherein said mid-span pre-diffuser inlet is located radially inboard ...

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 ...

COOLING SYSTEM FOR A GASPATH COMPONENT OF A GAS POWERED TURBINE

A gaspath component includes a flowpath body. A cooling plenum is disposed within the flowpath body. The cooling plenum includes a first region configured to receive a cooling flow and a second region configured to expel the cooling flow from the flowpath body. A metering obstruction is positioned between the first region and the second region and is configured to meter a flow of coolant through the cooling plenum. 1. A gaspath component comprising:a flowpath body;a cooling plenum disposed within said flowpath body, the cooling plenum including a first region configured to receive a cooling flow and a second region configured to expel the cooling flow from the flowpath body; anda metering obstruction positioned between the first region and the second region and configured to meter a flow of coolant through the cooling plenum.2. The gaspath component of claim 1 , wherein the gaspath component is a stator vane claim 1 , and wherein the flowpath body extends between a first platform and a second platform.3. The gaspath component of claim 1 , wherein the metering obstruction is a wall segment including at least one restricted opening.4. The gaspath component of claim 3 , wherein the at least one restricted opening includes a plurality of metering holes.5. The gaspath component of claim 4 , wherein the plurality of metering holes are evenly distributed across the metering obstruction.6. The gaspath component of claim 4 , wherein the plurality of metering holes are unevenly distributed across the metering obstruction.7. The gaspath component of claim 6 , wherein a density of the metering holes at an end of the metering obstruction farthest from an opening for receiving the cooling flow is greater than a density of the metering holes at an end of the metering obstruction closest to the opening for receiving the cooling flow.8. The gaspath component of claim 1 , wherein the metering obstruction is integral to the flowpath body.9. The gaspath component of claim 1 , wherein ...

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

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

COMPRESSOR BLEED COOLING SYSTEM FOR MID-FRAME TORQUE DISCS DOWNSTREAM FROM A COMPRESSOR ASSEMBLY IN A GAS TURBINE ENGINE

A cooling system configured to cool aspects of the turbine engine between a compressor and a turbine assembly is disclosed. In at least one embodiment, the cooling system may include one or more mid-frame cooling channels extending from an inlet through one or more mid-frame torque discs positioned downstream of the compressor and upstream of the turbine assembly. The inlet may be positioned to receive compressor bleed air. The mid-frame cooling channel may be positioned in a radially outer portion of the mid-frame torque disc to provide cooling to outer aspects of the mid-frame torque disc such that conventional, low cost materials may be used to form the mid-frame torque disc rather than high cost materials with capacity to withstand higher temperatures. The cooling fluid routed through the mid-frame cooling channel in the mid-frame torque disc may be exhausted into a cooling system () for the downstream turbine assembly. 113.-. (canceled)14. A cooling system for a turbine engine comprising:a compressor formed from a plurality of stages positioned within a compressor chamber, each of the plurality of stages includes a set of radially extending compressor blades; andat least one mid-frame cooling channel extending from an inlet through at least one mid-frame torque disc positioned downstream of the compressor and upstream of a turbine assembly,wherein the at least one mid-frame torque disc is formed from a torque disc rim positioned radially outward of a torque disc hub,wherein the torque disc rim and the torque disc hub are separated by a torque disc web having an axially extending width that is less than both the torque disc rim and the torque disc hub, andwherein the at least one mid-frame cooling channel is positioned in the torque disc rim.15. The cooling system of claim 14 , further comprising at least one cooling fluid supply bleed circuit with a bleed inlet placing the at least one bleed circuit in fluid communication with the compressor chamber for ...

GAS TURBINE ENGINE COMPONENT COOLING CAVITY WITH VORTEX PROMOTING FEATURES

A component according to an exemplary aspect of the present disclosure includes, among other things, a body, a wall extending inside of the body and a plurality of vortex promoting features arranged in a helical pattern along the wall. 1. A method , comprising:additively manufacturing a casting article layer-by-layer using an additive manufacturing process; andcasting a component to include a plurality of vortex promoting features arranged in a helical pattern along a wall of the component; using the casting article.2. The method as recited in claim 1 , wherein the component is made of a ceramic material.3. The method as recited in claim 1 , wherein the component is made of a refractory metal.4. The method as recited in claim 1 , wherein the casting article includes a body having a plurality of indents extending into the body claim 1 , and the indents are arranged in a helical pattern to form the plurality of vortex promoting features.5. The method as recited in claim 1 , wherein the step of manufacturing includes building the component to include the plurality of vortex promoting features layer-by-layer using an additive manufacturing process.6. The method as recited in claim 5 , wherein the casting article is a casting core.7. The method as recited in claim 1 , wherein the casting article is a casting core.8. The method as recited in claim 1 , wherein the additive manufacturing step includes building the casting article layer by layer by delivering a powdered metal to a build platform claim 1 , and melting a layer of powdered material.9. The method as recited in claim 8 , comprising:moving the build platform;adding a second layer of powdered material; andmelting the second layer.10. The method as recited in claim 8 , wherein the melting is performed by a laser.11. The method as recited in claim 8 , wherein the melting is performed by an electron beam melting device.12. The method as recited in claim 8 , wherein the powdered material is a ceramic material.13. The ...

HOT SECTION DUAL WALL COMPONENT ANTI-BLOCKAGE SYSTEM

A system for a hot section dual wall component in a gas turbine engine is used to avoid blockage by minimizing particulate deposits. The system includes impingement apertures formed in first wall of a cooling passageway of the dual wall component, and posts included on a second wall of the cooling passageway. The impingement apertures and the posts are respectively aligned opposite each other in the cooling passageway in operative cooperation so that working fluid exhausting into the cooling passageway from the impingement apertures in a first direction is directed to flow in a second direction along the cooling passageway to provide a laminar flow of the working fluid through the cooling passageway in order to minimize deposition of particles. 1. A system comprising:a hot section component of a gas turbine;a dual wall included in the hot section component, the dual wall including a first wall and a second wall, the first wall and the second wall disposed adjacently to define a cooling passageway therebetween;the first wall formed to include a series of impingement apertures providing fluid communication between the cooling passageway and a source of cooling fluid external to the cooling passageway; anda series of posts, each of the respective posts extending from the second wall toward a respective one of the series of impingement apertures, the posts sized and positioned to receive and direct a flow of fluid into the cooling passageway.2. The system of claim 1 , wherein a cross-sectional area of a base of each of the posts is greater than a cross sectional area of a tip of each of the posts claim 1 , the base coupled with second wall claim 1 , and one of the tip and a sidewall of the post aligned with the respective one of the series of impingement apertures.3. The system of claim 2 , wherein the sidewall is tapered between the tip and the base.4. The system of claim 3 , wherein the sidewall is a planar surface between the tip and the base.5. The system of claim 3 ...

GUIDE VANE FOR A TURBOMACHINE

A guide vane for a turbomachine, having a blade airfoil and at least one platform, to which the blade airfoil is connected. A cooling channel system is provided for cooling the platform and the blade airfoil. The platform has, on the side thereof facing the airfoil, at least one sealing lip for sealing to a rotating system of the turbomachine. At least one cooling channel extends through the sealing lip, which cooling channel forms part of the cooling channel system. 1. A guide vane for a turbomachine , comprising:a vane blade and at least one platform to which the vane blade is connected,a cooling duct system for cooling the platform and the vane blade,wherein the platform, for sealing in relation to a rotating system of the turbomachine, on the side thereof that faces away from the blade comprises at least one seal lip that projects from the platform,wherein at least one cooling duct which forms part of the cooling duct system extends through the seal lip,a plurality of cooling ducts, the exit openings of said cooling ducts being disposed in the free end of the seal lip, andwherein the cooling ducts, in each case proceeding from a free end of the seal lip, extend through the seal lip in at least a substantially radial direction.2. The guide vane as claimed in claim 1 ,wherein the at least one seal lip is disposed in a peripheral region of the platform that is forward or rearward in the axial flow direction, and from the side of the platform that faces away from the blade projects in at least a substantially radial manner such that said seal lip defines a terminating face of the platform that extends so as to be largely parallel to the radial direction.3. The guide vane as claimed in claim 1 ,wherein at least one connecting duct which connects the plurality of cooling ducts to one another and forms part of the cooling duct system is provided so as to be spaced apart from the free end of the seal lip.4. The guide vane as claimed in claim 1 ,wherein the platform is ...

GAS TURBINE SHROUD COOLING

A shroud segment for a casing of gas turbine includes a body configured for attachment to the casing proximate a localized critical process location within the casing. The body has a leading edge, a trailing edge, and two side edges. The critical process location is located between the leading edge and the trailing edge when the body is attached to the casing. A cooling passage is defined in the body along one of the side edges with one of an inlet or an outlet proximate the critical process location. The cooling passage is configured large enough to cool the one side edge adjacent the cooling passage to a desired level during operation of the gas turbine. The critical process locations may be related to temperatures, pressures or other measurable features of the gas turbine environment when in use. 1. A shroud segment for a casing of gas turbine comprising:a body configured for attachment to the casing proximate a localized critical process location within the casing, the body having a leading edge, a trailing edge, and two side edges, the body having a first surface for facing the casing and a second surface opposite the first surface for facing a hot gas path, the critical process location being located between the leading edge and the trailing edge when the body is attached to the casing; andat least two cooling passages defined in the body related to one of the side edges, a first of the cooling passages having an inlet and extending to an outlet, one of the inlet or the outlet being adjacent the critical process location, a second of the cooling passages having an inlet and extending to an outlet, one of the inlet or the outlet being adjacent the critical process location, the first and second cooling passages being configured large enough to cool the one side edge to a desired level during operation of the gas turbine.2. The shroud segment of claim 1 , wherein the first cooling passage has a first portion along the leading edge and a second portion along the ...

Gas turbine engine with vane having a cooling inlet

An apparatus and method of cooling a hot portion of a gas turbine engine, such as a multi-stage compressor of a gas turbine engine, by reducing an operating air temperature in a space between a seal and a blade post of adjacent stages by routing cooling air through an inlet in the vane, passing the routed cooling air through the vane, and emitting the routed cooling air into the space.

OIL SLINGER WITH CONVECTIVE COOLING OF RADIAL SURFACE

Oil slinger systems include a seal runner comprising an annular radial member having a radius (R) and an outer axially extending member having an axial length (L) such that a ratio (L/R) is between 0.8 and 1.4, the annular radial member disposed at a first angle with respect to the outer axially extending member, a heat shield in mechanical communication with the seal runner, and a volume bounded by an outer face of the annular radial member and an inner face of the heat shield. Methods of radial convective cooling include pumping a cooling liquid through the oil slinger system and convectively cooling the oil slinger. 1. An oil slinger system comprising:a seal runner comprising an annular radial member having a radius (R), an inner axially extending member, and an outer axially extending member having an axial length (L) such that a ratio (L/R) of the axial length to the radius is between 0.8 and 1.4, the annular radial member disposed at a first angle with respect to the outer axially extending member and the annular radial member disposed at a second angle with respect to the inner axially extending member;a heat shield in mechanical communication with the seal runner; anda volume bounded by an outer face of the annular radial member and an inner face of the heat shield;wherein the inner axially extending member comprises a radial oil passage configured to permit a lubricating oil to pass through the inner axially extending member to an inner face of the annular radial member such that the lubricating oil is driven by centrifugal force in the radially outward direction along the inner face of the annular radial member from the inner axially extending member to the outer axially extending member.2. The oil slinger system according to claim 1 , wherein the inner axially extending member comprises a lower oil passage configured to receive the lubricating oil claim 1 , wherein the lubricating oil passes from the lower oil passage and through the inner axially ...

ELECTRIC MOTOR HAVING AN INTEGRATED COOLING SYSTEM AND METHODS OF COOLING AN ELECTRIC MOTOR

The present disclosure pertains to electric machines such as electric propulsion systems for aircraft that integrated cooling systems, and methods of cooling such an electric machine. Exemplary electric machines include an electric motor that has a stator, a rotor, and a drive shaft operably coupled to the rotor. Exemplary electric machines further include a motor cooling conduit that defines a pathway for conveying a cooling fluid through or around at least a portion of the electric motor, a casing assembly that circumferentially surrounds at least a portion of the electric motor, a casing assembly conduit integrally formed within at least a portion of the casing assembly which defines a pathway for conveying the cooling fluid through the at least a portion of the casing assembly, and a pump or compressor operably coupled to the drive shaft and configured to circulate the cooling fluid through the motor cooling conduit and the casing assembly conduit. 1. An electric machine having an integrated cooling system , the electric machine comprising:an electric motor comprising a stator, a rotor, and a drive shaft operably coupled to the rotor;a motor cooling conduit that defines a pathway for conveying a cooling fluid through or around at least a portion of the electric motor, the motor cooling conduit having a thermally conductive relationship with the at least a portion of the electric motor; an annular casing;', 'a plurality of support members extending radially from the annular casing and circumferentially spaced relative to the annular casing; and', 'a casing assembly conduit integrally formed within at least a portion of the casing assembly, the casing assembly conduit defining a pathway for conveying the cooling fluid through the at least a portion of the casing assembly, the casing assembly circumferentially surrounding at least a portion of the electric motor, and the casing assembly conduit having a thermally conductive relationship with an external surface of ...

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 ...

AUGMENTED COOLING SYSTEM

An apparatus and method for cooling a dual, walled component is disclosed herein. An augmented cooling system according to the present disclosure includes transporting a cooling fluid through one wall of a cooling pathway formed between two opposing spaced apart walls of the dual walled component. The cooling fluid can be deflected away from one wall of the cooling pathway with a first trip strip as the cooling fluid traverses along the cooling pathway. The cooling fluid can be deflected away from the opposing wall of the cooling pathway with a second trip strip as the cooling fluid continues traversing along the cooling pathway. The cooling fluid can then be discharged from the cooling pathway through the opposing wall of the dual walled component. 1. A cooling system comprising:a component having an inner wall and an outer wall spaced apart from one another;a plurality of pedestals extending between the inner and outer walls;a plurality of inner trip strips projecting from the inner wall towards the outer wall at a predetermined height;a plurality of outer trip strips projecting from the outer wall towards the inner wall at a predetermined height,wherein one of either an inner trip strip or an outer trip strip extends between adjacent pedestals;at least one inlet through aperture formed in the inner wall of the component operable for transporting a cooling fluid into a space between the inner and outer walls of the component; anda plurality of outlet through apertures formed in the outer wall of the component operable for transporting the cooling fluid out of the space between the inner and the outer walls of the component;wherein at least one of the inlet through apertures and outlet through apertures is located in one of an inner well and an outer well, respectively, wherein the inner well is bounded on all sides by a plurality of inner trip strips, and wherein the outer well is bounded on all sides by a plurality of outer trip strips.2. The cooling system of ...

IMPINGEMENT COOLING OF TURBINE BLADES OR VANES

A turbine assembly having a hollow aerofoil and impingement device, the aerofoil having a first side wall from leading to trailing edge and a cavity arranged a distance to an inner surface of the cavity for impingement cooling and a flow channel for cooling medium from the leading to trailing edge, the impingement device has first and second pieces arranged side by side, the second piece downstream of the first forming a first flow passage providing passage from one side of the aerofoil towards an opposite side. A blocking element is arranged in the flow channel between the second piece and first side wall at a suction side for blocking flow of cooling medium from leading to trailing edge denying access to a section of the flow channel downstream of the blocking element while directing cooling medium in the first flow passage away from the suction side towards pressure side. 1. A turbine assembly comprising: wherein the hollow aerofoil has at least a first side wall extending from a leading edge towards a trailing edge of the hollow aerofoil and at least a cavity in which in an assembled state of the at least one impingement device in the hollow aerofoil the at least one impingement device is arranged with a predetermined distance in respect to an inner surface of the cavity for impingement cooling of the at least one inner surface and to form a flow channel for a cooling medium extending from the leading edge towards the trailing edge and', 'wherein the at least one impingement device comprises a first piece and a second piece being arranged side by side in an axial direction with the second piece being located viewed in the axial direction downstream of the first piece and with an axial distance in respect to each other forming a first flow passage providing a passage from one side of the aerofoil towards an opposite side of the aerofoil, and, 'a basically hollow aerofoil and at least an impingement device,'}at least a first blocking element, which is arranged in ...

COMPRESSOR TURBINE WHEEL

A turbine wheel for a gas turbine engine including a compressor impeller and a radial inflow turbine integral to or attached to the compressor impeller is provided. A compressor turbine wheel including features to increase surface area of a surface of the compressor impeller and/or the radial inflow turbine and/or a passage to flow air between the compressor impeller and the radial inflow turbine is further provided. Methods for cooling radial inflow turbines integral to compressor impellers are further provided. 1. A compressor turbine wheel for a gas turbine engine , the compressor turbine wheel rotatable about a forward-aft axis passing through a forward-aft axial wheel bore at a center of the compressor turbine wheel , the compressor turbine wheel comprising:a compressor impeller comprising a forward impeller vane inlet, a radial impeller vane outlet configured to flow compressed air radially outward to a diffuser, an aft inner surface extending from the forward impeller vane inlet to the radial impeller vane outlet, and inner vane side walls projecting forward from the aft inner surface; anda radial inflow turbine aft of the compressor impeller and comprising a radial turbine vane inlet configured to receive combustion gases from a combustor, an aft turbine vane outlet, a forward inner turbine surface extending from the radial turbine vane inlet to the aft turbine vane outlet, and inner turbine vane side walls projecting aft from the forward inner turbine surface; andwherein the radial inflow turbine is integral to the compressor impeller.2. The compressor turbine wheel of claim 1 , wherein the aft inner surface and/or the inner vane side walls comprise a rib configured to increase surface area of the aft inner surface and/or the inner vane side walls.3. The compressor turbine wheel of claim 1 , wherein the forward inner turbine surface and/or the inner turbine vane side walls comprise a pin claim 1 , a rib claim 1 , a vane claim 1 , and/or a turbulator ...

FORMING COOLING PASSAGES IN COMBUSTION TURBINE SUPERALLOY CASTINGS

Cooling passages () are formed in components for combustion turbine engines, such as blades (), vanes (), ring segments () or castings in transitions (), during investment casting, through use of ceramic shell inserts () within the casting mold (). Ceramic posts () formed in the ceramic shell insert () have profiles conforming to corresponding profiles of partially completed cooling passages (). Posts () are removed after superalloy component casting, forming the partially completed cooling passages, which are subsequently completed by removing remaining superalloy material along the cooling passage path. 1. A method for forming a cooling passage in an investment cast , superalloy component for a combustion turbine engine , comprising:providing a wax injection mold defining a mold cavity, whose mold cavity surface conforms to a corresponding surface profile of a component for a combustion turbine engine;providing a ceramic shell insert, having an insert surface profile and at least one ceramic post projecting from the insert surface, which in combination conform to a corresponding surface profile of a partially completed cooling passage in the engine component;inserting the ceramic shell insert into the wax injection mold, the shell insert surface and each ceramic post forming part of the mold cavity surface, each post projecting into the mold cavity;filling the mold cavity with wax, enveloping each post therein;hardening the wax, creating a wax pattern that embeds the ceramic shell insert and each post therein;removing the wax injection mold, the hardened wax pattern, along with the insert surface and each ceramic post conforming to the component surface profile;enveloping the hardened wax pattern and embedded ceramic shell insert in ceramic slurry;firing the ceramic slurry, thereby hardening the slurry into an outer ceramic shell, which is joined to the ceramic shell insert, and eliminating the wax, the joined outer ceramic shell and ceramic shell insert defining ...

COOLING STRUCTURE FOR STATIONARY BLADE

Embodiments of the present disclosure provide a cooling structure for a stationary blade. The cooling structure can include: an airfoil having a cooling circuit therein; an endwall coupled to a radial end of the airfoil; a chamber positioned within the endwall for receiving a cooling fluid from the cooling circuit, wherein the cooling fluid absorbs heat from the endwall, and a temperature of the cooling fluid in an upstream region is lower than a temperature of the cooling fluid in a downstream region; a first passage within the endwall fluidly connecting the upstream region of the chamber to a wheel space positioned between the endwall and the turbine wheel; and a second passage within the endwall fluidly connecting the downstream region of the chamber to the wheel space. 1. A cooling structure for a stationary blade , the cooling structure comprising:an airfoil having a cooling circuit therein;an endwall coupled to a radial end of the airfoil, relative to a rotor axis of a turbomachine;a chamber positioned within the endwall for receiving a cooling fluid from the cooling circuit and including an upstream region and a downstream region therein, wherein the cooling fluid absorbs heat from the endwall, and a temperature of the cooling fluid in the upstream region is lower than a temperature of the cooling fluid in the downstream region;a first passage within the endwall fluidly connecting the upstream region of the chamber to a wheel space positioned between the endwall and the turbine wheel, wherein a first portion of the cooling fluid in the upstream region passes through the first passage; anda second passage within the endwall fluidly connecting the downstream region of the chamber to the wheel space, wherein a second portion of the cooling fluid in the downstream region passes through the second passage, and a remainder portion of the cooling fluid bypasses the first passage and the second passage without entering the wheel space.2. The cooling structure of ...

COOLING STRUCTURE FOR STATIONARY BLADE

Embodiments of the present disclosure provide a cooling structure for a stationary blade, including: an endwall coupled to a radial end of an airfoil; a chamber positioned within the endwall and radially displaced from a radially outer end of the trailing edge of the airfoil, wherein the chamber includes a pair of opposing chamber walls, one of the pair of opposing chamber walls being positioned proximal to the pressure side surface of the airfoil and the other of the pair of opposing chamber walls being positioned proximal to the suction side surface and the trailing edge of the airfoil, and wherein the cooling fluid in the chamber is in thermal communication with least a portion of the endwall positioned proximal to the pressure side surface and the trailing edge of the airfoil; and a plurality of thermally conductive fixtures positioned within the chamber. 1. A cooling structure for a stationary blade , comprising:an endwall coupled to a radial end of an airfoil relative to a rotor axis of a turbomachine, the airfoil including a pressure side surface, a suction side surface, a leading edge, and a trailing edge;a chamber positioned within the endwall and radially displaced from the radial end of the trailing edge of the airfoil, the chamber receiving a cooling fluid from a cooling fluid source, wherein the chamber includes a pair of opposing chamber walls, one of the pair of opposing chamber walls being positioned proximal to the pressure side surface of the airfoil and the other of the pair of opposing chamber walls being positioned proximal to the suction side surface and the trailing edge of the airfoil, and wherein the cooling fluid in the chamber is in thermal communication with least a portion of the endwall positioned proximal to the pressure side surface and the trailing edge of the airfoil; anda plurality of thermally conductive fixtures positioned within the chamber and distributed substantially uniformly throughout the chamber.2. The cooling structure ...