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

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

СПОСОБ ИЗГОТОВЛЕНИЯ ИЗДЕЛИЯ ИЗ КОМПОЗИЦИОННОГО МАТЕРИАЛА

Изобретение относится к способу изготовления изделия из композиционного материала, содержащего полимерную матрицу, армированную волокнистой структурой. Способ включает в себя этапы, на которых волокнистую структуру укладывают на опору, образующую формовочную поверхность, закрывают ее контрформой и уплотняют структуру посредством сближения поверхности контрформы и поверхности опоры. Опора содержит цилиндрическую часть и стенку, ориентированную радиально относительно цилиндрической части. Контрформа содержит подвижные относительно друг друга две части, которые перемещают соответственно: одну в направлении оси цилиндрической части и другую в направлении радиальной стенки опоры. Способ применяют, в частности, для изготовления картера вентилятора турбореактивного двигателя. Изобретение обеспечивает повышение физико-механических свойств получаемых изделий. 9 з.п. ф-лы, 4 ил.

КОРПУС ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ И ТУРБОРЕАКТИВНЫЙ ДВИГАТЕЛЬ, СОДЕРЖАЩИЙ ТАКИЕ КОРПУСА

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

ОБЛОПАЧЕННЫЙ ЭЛЕМЕНТ ДЛЯ ТУРБОМАШИНЫ И ТУРБОМАШИНА

Облопаченный элемент турбомашины содержит набор лопаток с множеством лопаток, смещенных относительно друг друга в боковом направлении, и вихрегенераторы, расположенные выше по потоку от указанного набора лопаток в аксиальном направлении, перпендикулярном указанному боковому направлению. Выше по потоку от конца каждой лопатки расположена группа из множества вихрегенераторов, причем в каждой группе вихрегенероторы взаимно смещены и вбок, и аксиально. Каждая группа вихрегенераторов имеет по меньшей мере три вихрегенератора, причем вихрегенераторы каждой группы, по существу, параллельны. Другое изобретение группы относится к турбомашине, содержащей указанный выше облопаченный элемент. Группа изобретений позволяет снизить срыв потока на стороне всасывания лопатки и обеспечить при этом низкое аэродинамическое сопротивление. 2 н. и 8 з.п. ф-лы, 8 ил.

КОМПОЗИТНЫЙ КОРПУС ДЛЯ КОМПРЕССОРА ОСЕВОЙ ТУРБОМАШИНЫ, ПОЛУЧЕННЫЙ ДВУХКОМПОНЕНТНЫМ ЛИТЬЕВЫМ ФОРМОВАНИЕМ

Изобретение относится к сегментированному композитному корпусу компрессора осевой турбомашины. Каждый сегмент 18, 20 образуется из первого полимерного материала и содержит по меньшей мере одну рабочую поверхность 28, образованную из второго полимерного материала, подвергающегося двухкомпонентному литьевому формованию с первым полимерным материалом сегмента. Рабочая поверхность может представлять собой поверхность контакта с лопаткой. В этом случае профиль рабочей поверхности имеет выступ и изготавливается из эластомерного материала. Рабочая поверхность может также представлять собой внутреннюю поверхность 28, предназначенную для сцепления с истираемым материалом. В этом случае в качестве материала может использоваться силикон, чтобы обеспечить сцепление истираемого материала с силиконовым основанием. Рабочая поверхность может также представлять собой боковую поверхность в передней или задней части корпуса, при этом эта поверхность предназначена для контакта с соответствующей фиксированной ...

КОРПУС КОМПРЕССОРА С ПОЛОСТЯМИ С ОПТИМИЗИРОВАННОЙ РЕГУЛИРОВКОЙ

Компрессор для турбомашины, содержащий корпус, по меньшей мере одну ступень компрессора, образованную колесом с неподвижными лопатками и колесом с подвижными лопатками, полости, выполненные в толще корпуса и расположенные по окружности корпуса напротив подвижных лопаток. Полости выполнены удлиненной формы в основном направлении ориентации и закрыты в направлении выше по потоку и ниже по потоку соответственно расположенной выше по потоку стороной и расположенной ниже по потоку стороной, пересечения которых с корпусом образуют расположенную выше по потоку границу и расположенную ниже по потоку границу. Полости смещены относительно подвижных лопаток таким образом, чтобы выступать в направлении выше по потоку от колеса с подвижными лопатками, перекрывая их расположенный выше по потоку конец, расположенная ниже по потоку граница этих полостей ориентирована параллельно хорде в верхней части подвижной лопатки. Изобретение направлено на улучшение аэродинамических характеристик. 2 н. и 3 з.п. ф-лы ...

СПОСОБ ЦИРКУЛЯЦИИ ВОЗДУХА В КОМПРЕССОРЕ ТУРБИНЫ, КОНСТРУКЦИЯ КОМПРЕССОРА С ПРИМЕНЕНИЕМ ЭТОГО СПОСОБА, СТУПЕНЬ КОМПРЕССОРА И КОМПРЕССОР, СОДЕРЖАЩИЕ ТАКУЮ КОНСТРУКЦИЮ, И АВИАЦИОННЫЙ ДВИГАТЕЛЬ, ОБОРУДОВАННЫЙ ТАКИМ КОМПРЕССОРОМ

Группа изобретений относится к компрессорам турбин, в частности авиационных двигателей, и обеспечивает повышение степени сжатия ступени компрессора без ухудшения его производительности и предела помпажа, а также снижает зазор в газовоздушном тракте между лопатками колеса для устранения паразитных воздушных потоков внутри компрессора. Указанный технический результат достигается в ступени компрессора, содержащего несколько ступеней компрессора, содержащих подвижное и неподвижное колеса, наружный и внутренний картеры, формирующие газовоздушный тракт, включающий в себя «наружный тракт» и «внутренний тракт». Способ содержит операцию, в процессе которой воздух всасывают во внутреннем тракте и направляют его на неподвижную лопатку неподвижного колеса, и операцию, во время которой воздух отбирают в упомянутой неподвижной лопатке и направляют его наружу компрессора. 8 н. и 39 з.п. ф-лы, 9 ил.

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

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

ДАТЧИК ТУРБУЛЕНТНОСТИ КОМПРЕССОРА ТУРБОМАШИНЫ

Согласно настоящему изобретению предложена система измерения турбулентности потока (18) турбомашины, в частности компрессора турбомашины. Система (30) содержит: первый приемный элемент (47) с первым датчиком (52) давления и первым отверстием (48); второй приемный элемент (54) со вторым датчиком (58) давления и вторым отверстием (56), выполненным под наклоном относительно первого отверстия (48); и датчик (53) температуры. Система (30) предназначена для вычисления по меньшей мере двух компонентов направления скорости потока на основании данных от датчиков (52; 58) давления и датчика (53) температуры. Отверстия выполнены в хвостовике лопатки, на передней кромке на уровне внутренней оболочки. 3 н. и 16 з.п. ф-лы, 5 ил.

КОМПРЕССОРНОЕ УСТРОЙСТВО ГАЗОВОЙ ТУРБИНЫ И КОРПУСНОЙ ЭЛЕМЕНТ КОМПРЕССОРА

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

КОНСТРУКЦИЯ ЛОПАТКИ СТАТОРА И ТУРБОВЕНТИЛЯТОРНЫЙ РЕАКТИВНЫЙ ДВИГАТЕЛЬ С ИСПОЛЬЗОВАНИЕМ ТАКОЙ ЛОПАТКИ

НАПРАВЛЯЮЩАЯ ЛОПАТКАВЕНТИЛЯТОРА, ВЫПОЛНЕННАЯ ИЗ ТРЕХМЕРНОГО КОМПОЗИЦИОННОГО МАТЕРИАЛА

... 1. Способ изготовления направляющей лопатки, отличающийся тем, что он включает: ! а) изготовление волоконной преформы (1) посредством трехмерного переплетения одной детали, причем преформа содержит первую часть (10), продолжающуюся вдоль продольной оси (А) и образующую преформу для аэродинамического профиля лопатки, и расположенную на продольном конце (11) первой части (10) вторую часть (20), образующую преформу для платформы лопатки, причем вторая часть (20) выполнена в виде первого слоя (40) и второго слоя (50), обращенного к первому слою и отделенного от первого слоя (40) посредством разделения без разрезания при изготовлении преформы (1); ! b) сгибание первого и второго слоев (40, 50) таким образом, что каждый из них расположен в плоскости, перпендикулярной продольной оси, по существу, симметрично друг другу относительно первой части (10), и таким образом, что первый участок (R1) первого слоя (40) перекрывает второй участок (R2) второго слоя (50) перед передней кромкой (15) первой части ...

ЛОПАСТИ НАГНЕТАТЕЛЬНОГО ВЕНТИЛЯТОРА С ОСЕВЫМ ПОТОКОМ

УГЛОВОЙ СЕКТОР СТАТОРА КОМПРЕССОРА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ, СТАТОР ГАЗОТУРБИННОГО ДВИГАТЕЛЯ И ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ, СОДЕРЖАЩИЙ ТАКОЙ СЕКТОР

ДАТЧИК ТУРБУЛЕНТНОСТИ КОМПРЕССОРА ТУРБОМАШИНЫ

НАГНЕТАТЕЛЬНОЕ УСТРОЙСТВО И НАРУЖНЫЙ БЛОК КОНДИЦИОНЕРА ВОЗДУХА, СОДЕРЖАЩИЙ ЕГО

ВЕНТИЛЯТОР И ВХОДНАЯ НАПРАВЛЯЮЩАЯ РЕШЕТКА ДЛЯ ВЕНТИЛЯТОРА

ВЕНТИЛЯТОР

... 1. Вентилятор, в частности, для применения в технике кондиционирования и технике охлаждения, состоящий из корпуса, крыльчатки, которая установлена с возможностью вращения вокруг центральной оси, приводного двигателя и защитной решетки, отличающийся тем, что в целиком изготовленном из синтетического материала корпусе (1) интегрированы струйные элементы подающего сопла (1.1), имеющая форму цилиндра направляющая (1.2) потока, спрямляющие лопасти (1.3) и диффузор (1.5), причем спрямляющие лопасти (1.3) одновременно предназначены для присоединения крепления (1.4) двигателя к наружному контуру, и на корпусе (1) имеется устройство, на котором может быть установлена защитная решетка (3), в частности, в виде ячеистой решетки, причем радиально наружу лежащие концы (2.3) лопастей (2.1) крыльчатки (2) выполнены в виде специальных струйных элементов (винглетов), которые имеют особенно малый зазор с боковой стенкой (1, 2), причем приводной двигатель (5) представлен двигателем с внешним ротором, который ...

НАПРАВЛЯЮЩАЯ ЛОПАТКА, ИЗГОТОВЛЕННАЯ ИЗ КОМПОЗИЦИОННОГО МАТЕРИАЛА, СОДЕРЖАЩАЯ РАЗНЕСЕННЫЕ ДРУГ ОТ ДРУГА КРЕПЕЖНЫЕ ЛАПЫ, ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

ГОТОВОЕ ИЗДЕЛИЕ, ИМЕЮЩЕЕ АЭРОДИНАМИЧЕСКУЮ ЧАСТЬ СО СТОРОНОЙ ПОНИЖЕННОГО ДАВЛЕНИЯ ЗАДАННОГО ПРОФИЛЯ, И КОМПРЕССОР

... 1. Готовое изделие, имеющее аэродинамическую часть со стороной пониженного давления заданного профиля, по существу, в соответствии со значениями X, Y и Z декартовой системы координат стороны пониженного давления, приведенными в масштабируемой таблице, которая выбрана из группы таблиц, состоящей из Таблиц 1-11, и в которой значения X, Y и Z декартовой системы координат являются безразмерными значениями, преобразуемыми в размерные расстояния путем умножения значений X, Y и Z декартовой системы координат на некоторое число, причем координаты X и Y представляют собой координаты, которые, будучи соединенными непрерывными дугами, задают сечения профиля аэродинамической части на каждой высоте Z, при этом сечения профиля аэродинамической части на каждой высоте Z соединены друг с другом с формированием окончательной формы стороны пониженного давления аэродинамической части, причем значения X, Y и Z координат являются масштабируемыми в зависимости от указанного числа для получения по меньшей мере ...

ЛОПАТКА ОСЕВОЙ ТУРБОМАШИНЫ С ПЛАТФОРМАМИ, ИМЕЮЩИМИ УГЛОВОЙ ПРОФИЛЬ

... 1. Лопатка (28) статора осевой турбомашины (2), предназначенная для установки на втулку на кольцевом ряде аналогичных лопаток, при этом лопатка (28) содержит платформу (32, 132, 232, 332, 432, 532) со средствами (34, 134, 234, 334, 434, 534) крепления к втулке, позволяющую выполнять угловую регулировку лопатки, передний край (38, 138, 238, 338, 438, 538), задний край (38, 138, 238, 338, 438, 538) и два противоположных боковых края (36, 136, 236, 336, 436, 536);отличающаяся тем, чтоформа платформы содержит на каждом из своих боковых краев угловой профиль, выполненный с возможностью сопряжения со смежным краем платформы аналогичной смежной лопатки (28), для обеспечения углового позиционирования лопатки по меньшей мере в одном направлении вращения.2. Лопатка (28) по п. 1, отличающаяся тем, что угловой профиль содержит по меньшей мере одну контактную часть (48, 148, 248, 348, 448, 548), по существу, наклоненную к оси (14) турбомашины под углом, большим либо равным 20°, предпочтительно большим ...

НАПРАВЛЯЮЩИЙ АППАРАТ КОМПРЕССОРА ДЛЯ ТУРБОМАШИНЫ

... 1. Разделенный на сектора направляющий аппарат (3) компрессора для турбомашины, содержащий скрепленные сектора (5), образующие два концентрических кольца - внешнее (8) и внутреннее (10), - между которыми размещены лопатки (12) с их передними кромками (15) и задними кромками (16), находящимися близко, соответственно, с поперечными сторонами (17, 18) колец, внешнее кольцо (8) которого снаружи снабжено средством крепления (7) с внешним корпусом (6) приема упомянутых секторов,отличающийся тем, что упомянутое средство крепления (7) смещено в осевом направлении относительно задней поперечной стороны (18) внешнего кольца (8) и размещено для восприятия статических усилий между корпусом и направляющим аппаратом; причем упомянутое средство крепления (7) к внешнему корпусу содержит, относительно направления потока, проходящего через лопатки, передний (19) угловой периферический выступающий край зацепления, расположенный на уровне расположенной выше по потоку поперечной стороны (17) внешнего кольца ...

ОСЕВОЙ ВЕНТИЛЯТОР И СПОСОБ ИЗГОТОВЛЕНИЯ ОСЕВОГО ВЕНТИЛЯТОРА

... 1. Осевой вентилятор с мотором (1), на котором со стороны ротора закреплена крыльчатка (23, 24), от втулки которой (23) отходят лопасти вентилятора (24), имеющие передний и задний канты (26, 27), и с подвеской (2), посредством которой мотор (1) закреплен на корпусе (3), и которая имеет по меньшей мере один состоящий из плоскостного материала элемент подкоса (4, 5, 8, 43), который соединяет мотор (1) с корпусом (3) и в направлении потока воздуха располагается приблизительно в положении "на ребре", отличающийся тем, что элемент подкоса (4, 5, 8, 43) на части своей длины снабжен по меньшей мере одним вырезом (7), образованным путем высечки в плоскостном материале.2. Осевой вентилятор по п. 1, отличающийся тем, что элемент подкоса (4, 5, 8, 43) образован металлической листовой деталью.3. Осевой вентилятор по п. 1, отличающийся тем, что по меньшей мере от одного края выреза (7) отходит по меньшей мере одна опорная деталь (34, 35), которая предпочтительно выполнена в виде одной детали с элементом ...

НАПРАВЛЯЮЩИЙ АППАРАТ КОМПРЕССОРА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ С УПЛОТНЕНИЕМ

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

Schutzgehäuse für Ventilator

Cooling fan has air guide on air inlet side of frame with curved surface, whereby diameter of exterior of air guide is greater than frame's diameter, bearing with hub seat and several ribs and blade wheel

The fan has an annular frame (20) with an air inlet side and an air outlet side, whereby at least one object to be cooled can be arranged on the air outlet side, an air guide (21) formed on the air inlet side of the frame with a curved surface, whereby the diameter of the exterior of the air guide is greater than the frame's diameter, a bearing (22) with a hub seat (221 and several ribs (222) and a blade wheel (23).

Assembly for axial flow turbomachine in two-phase jet engine, has guide vanes and diffuser connected with each other through connecting portions, where guide vanes are provided with radially inner stator hub and radially outer housing ring

The assembly has guide vanes (5a, 5b) provided along a flow direction of a diffuser (6). The diffuser is provided with a compressor at the guide vanes. The guide vanes and the diffuser are produced separately. The guide vanes and the diffuser are connected with each other through connecting portions (81, 82). The guide vanes are provided with a radially inner stator hub (52) and a radially outer housing ring (51). The diffuser comprises a radially inner diffuser ring (62) and a radially outer diffuser ring (61).

KONSTRUKTIONSEINHEIT FUER STROEMUNGSTECHNISCHE EINRICHTUNGEN ODER MASCHINEN

Lüfter mit einem im Kunststoffgehäuse integrierten Ventilatormodul

Lüfter (100) zum Lüften und Trocknen von Gebäuden und Räumen, wobei der Lüfter (100) ein kunststoffaufweisendes oder aus Kunststoff bestehendes Lüftergehäuse (10) aufweist, wobei der Lüfter (100) ferner ein im Lüftergehäuse (10) angeordnetes Ventilatormodul (11) mit einem Ventilator (12) zur Luftförderung aufweist, dadurch gekennzeichnet, dass das Ventilatormodul (11) ein metallaufweisendes oder aus Metall bestehendes Ventilatorgehäuse (13) aufweist, welches den Ventilator (12) umfänglich umschließt.

Diaphragmaaufbau eines Verdichters.

Kompressor und Turbinenstrahltriebwerk

LEITGITTER FUER EINEN AXIALVENTILATOR

Gasturbine und Verfahren zum Ändern der aerodynamischen Gestalt einer Gasturbinenschaufel

Ein Verfahren zum Verändern der aerodynamischen Gestalt einer Schaufel (22, 24) einer Gasturbine sieht vor, diese mit einer Beschichtung (42) aus einer Formgedächtnis-Legierung zu versehen. Über Kanäle (32) wird Heißluft (36) zur beschichteten Schaufel (22, 24) transportiert, um eine Phasenänderung der Beschichtung (42) und eine aerodynamische Gestaltänderung der Schaufel (22, 24) herbeizuführen.

Turboverdichter und Verfahren zum Betrieb eines Turboverdichters

Verbesserter Verdichter in Axialbauart

Ein verbesserter Verdichter in Axialbauart für ein Gasturbinentriebwerk, wobei der Verdichter zumindest einen Rotor mit Laufschaufeln und ein über seine axiale Länge mehrteilig aufgebautes Verdichtergehäuse mit Leitschaufeln aufweist, ist dadurch gekennzeichnet, dass ein steuer- bzw. regelbares Gehäusekühlsystem zur thermischen Anpassung des Verdichtergehäuses an den Rotor vorgesehen ist.

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

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

Strömungsarbeitsmaschine mit Schaufelreihengruppe

... tsmaschine mit einem durch eine rotierende Welle (2) und ein Gehäuse (1) gebildeten Hauptströmungspfad, in dem zumindest eine Anordnung rotierender, einen Rotor bildender und dem Fluid Energie zuführender Schaufeln vorgesehen ist und in dem in Hauptströmungsrichtung benachbart zum Rotor eine Anordnuesehen ist, wobei die Rotorschaufeln an einer rotierenden Welle (2) befestigt sind und zumindest eine der Anordnungen des Rotors und des Stators durch eine Schaufelreihengruppe gebildet wird, wobei die zumindest eine Schaufelreihengruppe aus mehreren in Hauptströmungsrichtung benachbart angeordneten Rotorschaufelreihen oder Statorschaufelreihen selben Typs als Mitgliedsschaufelreihen besteht, wobei der Flächenquerschnittverlauf des Hauptströmungspfades in mindestens einer eine Rotoranordnung und eine Statoranordnung umfassenden Stufe zu einem überhöhten Rotor-Stator-Einschnürungsverhältnis QRS führt, welches folgender Beziehung genügt: [0,2 + (KT - 0,45)] < QRS wobei QRS nach folgender Formel ...

Blower e.g. diagonal blower, for cooling of electronic device, has conducting element with air passage opening, and check valve proided at exhaust side of air passage opening for partially closing air passage opening in closing position

The blower has a blower housing with an air inlet opening, an air outlet opening and a conducting element (30), which is provided in an area of the air inlet opening. The conducting element is designed in such a manner that the conducting element partially converts twist of the air, escaping from the blower during operation, into static pressure. The conducting element has an air passage opening, and a check valve (72) is provided at an exhaust side of the air passage opening for partially closing the air passage opening in a closing position.

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

A motor and a handheld product having a motor

Filled static structure for axial-flow machine

AXIAL GUIDE BAFFLE ASSEMBLY FOR COMPRESSORS

Vorticity control in a gas turbine engine aerofoil

Stator for turbomachines

A stator vane assembly for a turbomachine

Stator assembly including bleed ports for turbine engine compressor

Motor mounting assembly for an axial fan

Compressor shroud comprising a sealing element provided with a structure for driving and deflecting discharge air

The invention relates to a compressor shroud for an aircraft turbomachine, said shroud (20) being arranged between two rotating bladed wheels (16) and radially perpendicularly to a stator (12), and comprising a sealing device (30) comprising at least one sealing element (32), one of which is a downstream sealing element (32a) whereon a downstream-projecting structure (44) for driving and deviating air is provided, said structure being designed to axially straighten the discharge air coming from the sealing element (32a).

Vane support in a gas turbine engine

Vacuum pump

A turbo-molecular vacuum pump 10, comprising rotor 14 and stator 18 components in a housing 12, which has an inlet 24 and outlet 26 side. The rotor components are coupled to a drive shaft 20 to be driven about a longitudinal axis 22, and the drive shaft is coupled to the housing via bearing means 31, 31 to allow it relative rotary movement. Each of the stator components has a series of stator blades (32, fig 2 + 3) extending radially from an inner portion at the longitudinal axis to an outer portion (34), where each of the stator blades being angled with respect to a plane defined by the inner portion. A spacer 28 locates and couples the stator components relative to the housing, with the outer portion of at least one of the stator components comprising a resilient portion 34 that cooperates with the spacer. Preferably, the resilient portions are at the outer tip of the stator blade. Preferably, the pump includes a stator stack comprising a plurality of spacers interposed between adjacent ...

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

Variable pitch guide vane made of composite materials

A compressor variable stator vane assembly comprises a blade 50 and at least one pivot 20, 21. Each pivot comprises an internal pivot element 40a, 40b, 41a, 41b and a pivot cap (30, 31, figure 4). The blade and internal pivot element(s) are each made from a composite material, and at least one contact surface of each pivot cap is metallic. Each internal pivot element is assembled on a cleat 52, 53, the cleat being made in a single piece from the same material as the blade. The blade and internal pivot element(s) may be made from a long-fibre pre-impregnated 2D or woven 3D composite, and internal pivot element(s) being short fibres in an organic matrix respectively. The pivot cap may be fully metallic. The internal pivot element may be injected on, or glued onto, the cleat. A method of manufacturing the vane may include assembling each internal pivot element on a cleat, and then assembling a cap on each internal pivot element.

Casing section

Improvements relating to cased screw fans

... 631,231. Screw fans. AEREX, Ltd., REEVES, H. C., and HAIGH, F. S. Dec. 10, 1947, No. 32554. [Class 110(i)] The delivery guide vanes a have straight leading edges d which in a radially outward direction are inclined rearwardly relatively to the direction of rotation of the impeller. The vanes are of laminar section, rounded at the leading edge and tapered at the trailing edge. Alternatively, they may be of solid or hollow aerofoil shape. The vanes are attached at their outer and inner ends to the casing and fairing respectively, either by eyed lugs or by welding. The trailing edges of the impeller blades b and the leading edges d of the guide vanes are oppositely inclined, so that they pass each other progressively, thus reducing noise due to syren effect.

Improvements in or relating to axial-flow compressors

... 710,282. Axial-flow compressors. ROLLSROYCE, Ltd. Sept. 15, 1952 [Sept. 14, 1951], No. 21687/51. Addition to 621,544. Classes 110(1) and 110(3) The axial-flow compressor stator casing described in the parent Specification is formed in a number of axially-abutting sections, the plane of division between a pair of abutting sections 18, 19 being at right angles to the casing axis and each section having at its abutting end a reinforcing web 26 which in addition to having a working channel wall portion 11b and a surface 11a for locating stator blades 13, affords the abutment surface 24. the sections being fastened together by bolts or, as shown, studs 28 which extend through the webs. The sections are. in addition, split on a longitudinally extending plane. The blades 13 may be located radially by the engagement of ribs 16 with a web 12, as in the parent Specification. Alternatively, locating surfaces 11c may be provided.

Columnar air moving devices, systems and methods

Vacuum pump and vacuum pump set for evacuating a semiconductor processing chamber

A vacuum pump, a set of vacuum pumps and method for evacuating a semiconductor processing chamber is disclosed. The pump is configured for mounting to a semiconductor processing chamber to evacuate it to between 1 mbar and 5 X10-2 mbar. The pump comprises: a rotor having a plurality of angled blades 30 arranged along a helical path from an inlet 16 to an outlet 18, mounted within a stator having a plurality of perforated elements 50-53 intersecting the helical path. The perforations 38 allow molecules to travel along the helical path. The rotor is mounted on a magnetically levitated bearing. The perforated elements located towards an inlet of the pump comprise a transparency of more than 40% and the perforated elements located towards an outlet comprise a transparency of more than 30%. The set comprises the described pump, a roots blower and a primary vacuum pump. The method comprises attaching the described pump to a processing chamber by a conduit of less than 2m and remotely locating ...

Fans for ventilation

A fan for ventilation such as a building ventilation extractor comprising a housing 22, an impeller 44, a motor 48 arranged to drive the impeller to drive air from an inlet to an outlet 24, the impeller located centrally in a flow duct, and at least one inlet guide vane 70 for guiding flow approaching the impeller. Also disclosed is: a fan having the impeller located axially wholly within a tapered duct portion of the duct; A fan having at least one downstream guide vane being at least partly curved and arranged to reduce swirl in downstream flow; A fan wherein the motor is located in a duct and having a motor cover, the motor and cover secured to a support by a common fixing; A ventilation system arranged to push air out of a building through an exhaust duct; A data communication system having a link from a plurality of fan units to a controller to enable common fan speed boost; A system to control the fans based on weather data; A data communications system having a data store remote ...

DISC FOR A MOLECULAR VACUUM PUMP

... 1275386 Molecular vacuum pumps BALZERS PATENT UND BETEILIGUNGS A G 18 Jan 1971 [10 April 1970] 2434/71 Heading F1C A rotor disc or, as shown, a stator disc for a molecular vacuum pump, comprises a hollow region formed between parallel walls 3, 4, Figs. 1A and 1B, e.g. of sheet titanium, each provided with an annular array of radial slits 5 having inwardly bent lips 6, the slits being angularly staggered relative to each other so that there is no axial line of sight through the disc. The stator disc shown is secured to the pump housing by peripheral flange 1. The rotor disc has no peripheral flange, and the opposed lips are connected in pairs, e.g. by welding, Figs. 2A to 2D (not shown).

Turbomachine

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

Compressor shroud comprising a sealing element provided with a structure for entraining and diverting discharge air

Vorticity control in a gas turbine engine

Turbine engine

Air moving device with bypass intake

STATOR DISK FOR TURBO-MOLECULAR PUMP

PROCEDURE FOR THE PRODUCTION OF COMPOUND SHAPED PARTS WITH A PP MATRIX

COMPRESSOR DIAPHRAGM ASSEMBLY

GUIDANCE DEVICE FOR AN INLET AIR FLOW TO A COMBUSTION CHAMBER IN A TURBINE ENGINE

Dispositif de guidage d'un flux d'air à l'entrée d'une chambre de combustion dans une turbomachine, comprenant un redresseur (14) suivi d'un diffuseur (16), une des viroles (40) du redresseur étant formée d'une seule pièce avec une paroi de révolution (34) du diffuseur, l'autre des viroles (38) du redresseur étant rapportée et fixée sur l'autre paroi de révolution (32) du diffuseur, et les aubes (42) du redresseur étant solidaires par une extrémité d'une virole (40) du redresseur et écartées d'un jeu faible (46) de l'autre virole (38) à leur autre extrémité.

GUIDANCE DEVICE FOR AN INLET AIR FLOW TO A COMBUSTION CHAMBER IN A TURBINE ENGINE

Dispositif de guidage d'un flux d'air à l'entrée d'une chambre de combustion dans une turbomachine, comprenant un redresseur (14) suivi d'un diffuseur (16), une des viroles (40) du redresseur étant formée d'une seule pièce avec une paroi de révolution (34) du diffuseur, l'autre des viroles (38) du redresseur étant rapportée et fixée sur l'autre paroi de révolution (32) du diffuseur, et les aubes (42) du redresseur étant solidaires par une extrémité d'une virole (40) du redresseur et écartées d'un jeu faible (46) de l'autre virole (38) à leur autre extrémité.

COLUMNAR AIR MOVING DEVICES, SYSTEMS AND METHOD

A columnar air moving device can comprise separately formed modular stator vanes in a stator vane assembly. The stator vanes can be arranged in a radial pattern, and can direct air in an axial direction. The modular stator vanes, as well as other components of the stator vane assembly, can be replaced, adjusted, and/or removed from the columnar air moving device.

COLUMNAR AIR MOVING DEVICES, SYSTEMS AND METHOD

A columnar air moving device can comprise sepa-rately formed modular stator vanes in a stator vane assembly. The stator vanes can be arranged in a radial pattern, and can direct air in an axial direction. The modular stator vanes, as well as other com-ponents of the stator vane assembly, can be replaced, adjusted, and/ or removed from the columnar air moving device.

WET COMPRESSION APPARATUS AND METHOD

... ²²²This wet compression invention with a vaporizable fluid mist demonstrates ²major performance improvements over the relevant art in achieving a high ²degree of saturation, providing sensible cooling, strongly reducing the ²temperature increase due to compression work, reducing excess diluent air flow ²for downstream combustion, reducing compression noise, and increasing the ²achievable compressor pressure ratio . These improvements are obtained by one ²or more of: high mist or overspray from a) progressive axial injection of ²vaporizable fluid along the streamwise compression flow path, and b) ²transverse vaporizable fluid delivery from stators, rotors, perforated tubes, ²and/or duct walls, matching the gaseous fluid flow distribution across the ²compressor stream; c) reducing the compressor cross-sectional flow area of ²downstream compressor stages relative to up-stream stages, and d) increasing ²the rate of downstream vaporizable fluid injection relative to the rate of ²upstream injection ...

DEVICE FOR ATTACHING A STATIONARY BLADE IN A TURBINE ENGINE ANNULAR CRANKCASE, TURBOJET CONTAINING THE DEVICE AND PROCESS FOR ASSEMBLING THE BLADE

La présente invention porte sur un dispositif de fixation d'une aube fixe de turbomachine dans un carter annulaire (20) de la turbomachine, l'aube comprenant une tête formée d'une plateforme (14) avec deux bords (141, 142), l'un formant un bord amont (141) et l'autre formant un bord aval (142), une pale (11) et un pied (13), le carter annulaire (20) comprenant une gorge circonférentielle (23) ménagée dans sa paroi interne avec un flanc amont (21) et un flanc aval (22), l'aube étant retenue par engagement de la plateforme (14) dans la gorge (23), le bord amont contre le flanc amont (21) et le bord aval (142) contre le flanc aval (22) de la gorge (23),le flanc aval (22) de la gorge (23) étant en forme de rainure (22) d'ouverture axiale vers l'amont, le bord aval de la plateforme (23) étant engagé dans ledit flanc aval (22) et un moyen de fixation (40) étant rapporté pour maintenir la plateforme (14) dans la gorge à l'amont, caractérisé par le fait que l'aube comporte une portion de plateforme ...

FAN ASSEMBLY HAVING FLOW RECIRCULATION CIRCUIT WITH ROTATING AIRFOILS

There is disclosed a fan assembly including a fan rotor including a hub and fan blades. The fan blades have a leading edge and a trailing edge. A fan stator downstream of the fan rotor relative to a direction of an airflow through the fan assembly. The fan stator includes vanes extending between radially inner ends and radially outer ends. A flow recirculation circuit has an inlet downstream of the vanes of the fan stator and an outlet upstream of the vanes. A recirculation rotor has a plurality of airfoils circumferentially distributed around the axis and located in the flow recirculation circuit. The recirculation rotor is rotatable about the axis within the recirculation circuit. A method of operating the fan assembly is also disclosed.

DOUBLE ROW COMPRESSOR STATORS

A method of manufacturing a compressor stator having: a first stator blade with a first leading edge and a first trailing edge; a second stator blade disposed a circumferential distance from the first stator blade, the second stator blade having a second leading edge disposed an axial distance from the first leading edge and a second trailing edge disposed an axial distance from the first trailing edge; the method comprising: using additive manufacturing to deposit and fuse together progressive layers of metal material commencing at a substrate to form the first stator blade, the second stator blade, at least one intermediate support structure disposed between the first stator blade and the second stator blade, and at least one primary support structure disposed between the substrate and at least one of: the first stator blade; and the second stator blade; and removing the primary support structure and the intermediate support structure.

ANGULAR SECTOR OF THE DOWNSTREAM GUIDE VANES FOR A TURBINE ENGINE COMPRESSOR, TURBINE ENGINE DOWNSTREAM GUIDE VANES AND TURBINE ENGINE INCLUDING SUCH A SECTOR

La présente invention concerne un secteur angulaire de redresseur (18) pour compresseur de turbomachine comportant une virole externe (24), une virole interne (22), et au moins une pale (26) s'étendant radialement entre les dites viroles (22, 24), dans lequel la virole externe (24) comporte des premier et second moyens de montage (44, 46) sur un carter (20) de la turbomachine, orientés en direction axiale selon des sens opposés et reliés entre eux par une portion intermédiaire (41). La virole externe comporte en outre au moins une portion d'extrémité axiale (36) s'étendant depuis ladite portion intermédiaire, avec laquelle au moins un élément rapporté (601) formant amortisseur est apte à venir en contact, de sorte qu'au-delà d'une valeur donnée de l'amplitude de vibration de la portion d'extrémité (36), l'élément amortisseur (601) et la portion d'extrémité (36) sont aptes à se déplacer relativement l'un par rapport à l'autre de manière à faire varier la masse totale en mouvement qui est ...

FIXED GUIDE VANE ASSEMBLY SEPARATED INTO SECTORS FOR A TURBOMACHINE COMPRESSOR

The invention relates to a fixed guide vane assembly comprising a circular casing composed of at least two parts (11) and supporting sectors (5), each composed of an inner segment (7) and an outer segment (6) connected by vanes (8). The parts of the casing and the sectors of the guide vane assembly are held together by a system forming a slide and a slider. Anti-rotation means (30) are provided to prevent the sectors from rotating with respect to the casing.

FIBER REINFORCED AIRFOIL

An airfoil and a method of manufacturing an airfoil may be provided, where the airfoil comprises a core and a shell. The core comprises core ceramic fibers extending along a span of the airfoil. The shell surrounds the core and includes shell ceramic fibers. Substantially all of the core ceramic fibers are arranged in a radial direction. The airfoil may also be a ceramic matrix composite formed by infiltrating the core and the shell with a matrix material.

AFT ENGINE FOR AN AIRCRAFT

A propulsion system for an aircraft is provided having an aft engine configured to be mounted to the aircraft at an aft end of the aircraft. The aft engine includes a fan rotatable about a central axis of the aft engine having a plurality of fan blades attached to a fan shaft. The aft engine also includes a nacelle encircling the plurality of fan blades and a structural support system for mounting the aft engine to the aircraft. The structural support system extends from the fuselage of the aircraft, through the fan shaft, and to the nacelle when the aft engine is mounted to the aircraft. The aft engine may increase a net thrust of the aircraft when mounted to the aircraft.

PERFORATED DRUM OF A COMPRESSOR OF AN AXIAL TURBINE ENGINE

A rotor, in particular, a drum of a low-pressure compressor of a turbojet aero engine has an outer annular wall delimiting a primary annular flow of the turbine engine, sealing devices with two rubbing strips or annular ribs formed on the wall. The rubbing strips cooperate by abrasion with inner shrouds. In addition, the annular wall includes rows of intake orifices for leakages which are arranged between each pair of rubbing strips in order to aspirate the recirculation leakages there. A collector for leakages is formed inside the rotor by means of a composite web, then evacuates the parasitic flow downstream of the turbine engine via the central shaft.

PLATED POLYMER COMPRESSOR

Plated polymeric gas turbine engine parts and methods for fabricating lightweight plated polymeric gas turbine engine parts are disclosed. The parts include a polymeric substrate plated with one or more metal layers. The polymeric material of the polymeric substrate may be structurally reinforced with materials that may include carbon, metal, or glass. The polymeric substrate may also include a plurality of layers to form a composite layup structure.

CERAMIC-ENCAPSULATED THERMOPOLYMER PATTERN OR SUPPORT WITH METALLIC PLATING

A method for fabricating a ceramic component is disclosed. The method may comprise: 1) forming a polymer template having a shape that is an inverse of a shape of the ceramic component, 2) placing the polymer template in a mold; 3) injecting the polymer template with a ceramic slurry, 4) firing the ceramic slurry at a temperature to produce a green body, and 5) sintering the green body at an elevated temperature to provide the ceramic component.

STATOR VANE

Aube (1) statorique pour rotor d'une turbomachine et incluant au moins une portion (10) en matériau apte à la super-élasticité. La portion (10) en matériau apte à la super-élasticité est agencée pour entrer en résonance à un régime prédéterminé de la turbomachine notamment lors d'un régime de la turbomachine typique de la phase de croisière de l'aéronef.

COMPRESSOR AFT ROTOR RIM COOLING FOR HIGH OPR (T3) ENGINE

In one aspect, the present disclosure is directed to a cooling circuit for a gas turbine engine. The cooling circuit includes a rotor blade having a connection portion and a rotor disc having a first axial side and a second axial side. The rotor disc defines a connection slot and a cooling passage extending between the first axial side and the second axial side. The connection slot receives the connection portion to couple the rotor blade to the rotor disc. Cooling air flows through the cooling passage.

ANGULAR SECTOR OF THE DOWNSTREAM GUIDE VANES FOR A TURBINE ENGINE COMPRESSOR, TURBINE ENGINE DOWNSTREAM GUIDE VANES AND TURBINE ENGINE INCLUDING SUCH A SECTOR

La présente invention concerne un secteur angulaire de redresseur (18) pour compresseur de turbomachine comportant une virole externe (24), une virole interne (22), et au moins une pale (26) s'étendant radialement entre les dites viroles (22, 24), dans lequel la virole externe (24) comporte des premier et second moyens de montage (44, 46) sur un carter (20) de la turbomachine, orientés en direction axiale selon des sens opposés et reliés entre eux par une portion intermédiaire (41). La virole externe comporte en outre au moins une portion d'extrémité axiale (36) s'étendant depuis ladite portion intermédiaire, avec laquelle au moins un élément rapporté (601) formant amortisseur est apte à venir en contact, de sorte qu'au-delà d'une valeur donnée de l'amplitude de vibration de la portion d'extrémité (36), l'élément amortisseur (601) et la portion d'extrémité (36) sont aptes à se déplacer relativement l'un par rapport à l'autre de manière à faire varier la masse totale en mouvement qui est ...

COMPOSITE BLADE MADE BY ADDITIVE MANUFACTURING

The invention relates to a blade of low pressure rectifier axial turbomachine. The blade can also be a rotor blade and/or a turbine blade. The blade comprises a composite material with a matrix and a reinforcement that comprises a mesh forming a three dimensional structure with a plurality of rods that describe a three-dimensional mesh based on polyhedrons. The three-dimensional structure extends over the majority of the thickness of the blade between the pressure side surface and the suction side surface and/or the majority of the length of the blade between the leading edge and the trailing edge. The rods of the reinforcement are bonded to each other and are distributed throughout the volume between the pressure side surface and the suction side surface of the blade. The rods form a three-dimensional mesh occupying the entire blade. The invention also relates to an iterative method for manufacturing a blade by additional layer manufacturing.

COMPRESSOR ROTOR AND STATOR BLADE WHEEL ASSEMBLIES

Provided is a compressor for use in a gas turbine engine, capable of preventing the creation of rust on an inner surface of the compressor casing, without complicating the assembling process. The casing of the compressor accommodates rotor and stator blade wheels. The stator blade wheels are supported on the inner surface of the casing through outer flanges thereof. Seal rings are provided at inner surface portions of the casing opposing the radially outward ends of the rotor blade wheels. The inner surface of the casing is covered by the seal rings and the outer flanges of the stator blade wheels.

Axial-Kompressor-Stator.

Stator de turbo-machine à plusieurs étages et à écoulement axial

Leitrad für Achsialverdichter oder Achsialturbinen.

Schaufelträger in axialen Kreiselmaschinen.

Compresseur axial à aubes.

Vielstufiger Axialverdichter mit Drallschaufeln am Austritt

Disque pour pompe à vide turbomoléculaire

Palettatura di un compressore assiale rotativo

Turbine engine compressor stator

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

compressor nozzle stage for a turbine engine

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

Stationary vane unit of rotary machine, method of producing the same, and method of connecting the same

The stationary vane unit of a rotary machine includes: a first band member that extends in the circumferential direction and comes into contact with the outer shrouds of the plurality of stationary vane members from one side thereof in the main axial direction in which a central axis extends; a second band member that extends in the circumferential direction and comes into contact with the outer shrouds of the plurality of stationary vane members from the other side thereof in the main axial direction; and a fastening member that fastens the first band member and the second band member to each other so that the outer shrouds of the plurality of stationary vane members are connected to each other.

Recirculation fan and wind-guiding device thereof

A recirculation fan includes a casing, a covering member, a wind-guiding device, a passive impeller, and an active impeller. The covering member is coupled with the casing to define an accommodation space. The wind-guiding device is disposed on the covering member, and includes a wind-guiding cover and a magnetoresistive structure. The magnetoresistive structure is disposed on the covering member and the wind-guiding cover. The passive impeller is disposed within the accommodation space. The active impeller is disposed within the accommodation space and located beside the passive impeller for generating a wind to drive rotation of the passive impeller and the wind-guiding cover. In response to a magnetic torque resulted from a magnetic vortex of the magnetoresistive structure, a rotating speed of the wind-guiding cover is slowed down.

Blade cascade and turbomachine

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

Method of manufacturing an airfoil

Disclosed is a method of manufacturing an airfoil. The method includes establishing an Argon (Ar)-free environment, providing a bed within the Argon free environment, providing a set of data instructions for manufacturing the airfoil, and providing a powdered Nickel (Ni)-based alloy on the bed. In one example, the powdered Nickel (Ni)-based alloy consists essentially of about 4.8 wt. % Iron (Fe), about 21 wt. % Chromium (Cr), about 8.6 wt. % Molybdenum (Mo), about 0.07 wt. % Titanium (Ti), about 0.40% Aluminum (Al), about 5.01 wt. % Niobium (Nb), about 0.03 wt. % Carbon (C), about 0.14 wt. % Silicon (Si), and a balance Nickel (Ni). The method further includes fusing the powdered Nickel (Ni)-based alloy with an electron beam with reference to the data instructions to form the airfoil.

Method and tool for use with compressors

A tool for drilling, tapping, and back spot facing a hole in dynamoelectric machines includes a servo motor cutting device and a drill unit skid. The servo motor cutting device includes a cutting tool. The drill unit skid engages a hook fit slot in a case of the dynamoelectric machine and supports the servo motor cutting device. The drill unit skid includes hook fit slides configured to guide the drill unit skid along the hook fit slot and act as a stop in a radial direction. The tool includes a laser positioning device configured to detect the position of the cutting tool, and a micro switch configured to stop the servo motor cutting device when the micro switch activates. The servo motor cutting device rotates the cutting tool to create the hole, and a rotor of the dynamoelectric machine is left in place during drilling, tapping and back spot facing.

Turbomachine with variable-pitch vortex generator

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

Filled static structure for axial-flow machine

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

COOLING SYSTEM FOR STREAMLINED AIRFLOW

A cooling system includes a fan and a system component. The fan includes a plurality of fan blades and configured to rotate in a fan direction. The system component is located downstream of the fan, and includes a cutout for passing of airflow from the fan, and a bridge spanning the cutout. The bridge includes a center section and at least one arm section extending from the center section to an edge of the cutout along a curved path offset towards the fan direction. 1. A system for providing streamlined airflow , comprising:a fan comprising a plurality of fan blades and configured to rotate in a fan direction; and a center section; and', 'at least one arm section extending from the center section to an edge of the cutout along a curved path offset towards the fan direction;, 'a system component located downstream of the fan, comprising a cutout for passing of airflow from the fan, and a bridge spanning the cutout, wherein the bridge comprises2. The system of claim 1 , further comprising a spinner faring that extends from the center section of the system component.3. The system of claim 1 , further comprising a reverse blade faring extends from each of the at least one arm section.4. The system of claim 1 , wherein each of the plurality of fan blades comprises a leading fan edge facing towards the fan direction and a trailing fan edge facing against the fan direction claim 1 , wherein each of the at least one arm section comprises a leading arm edge facing the fan direction and a trailing arm edge facing against the fan direction.5. The system of claim 4 , wherein the leading arm edge follows a convex path from the center section to the edge of the cutout.6. The system of claim 4 , wherein the trailing arm edge follows a concave path from the center section to the edge of the cutout.7. The system of claim 4 , wherein the leading arm edge and trailing arm edge have an approximately equivalent shape.8. The system of claim 4 , wherein the leading arm edge of each of the ...

Systems and methods for inspecting blades or vanes in turbomachinery

A scanner for inspecting a component is disclosed. The scanner may include a main body having a clamp attachable to the component, a first arm moveably attached to the main body, a first probe attached to the first arm, a first axial actuator in mechanical communication with the first arm, a second arm moveably attached to the main body, a second probe attached to the second arm, a second axial actuator in mechanical communication with the second arm, and a radial actuator in mechanical communication with the first arm and the second arm.

MECHANICAL COMPONENT FOR THERMAL TURBO MACHINERY

A mechanical component for thermal turbo machinery, such as a steam or gas turbine, includes a base part and at least one additional device being mechanically coupled to the base part in order to influence the vibration characteristic of the base part during operation of the turbo machine. High-Cycle Fatigue at part-load can be reduced by enabling the mechanical coupling between the base part and the at least one additional device to change with the temperature of the at least one additional device. 1. Mechanical component for thermal turbo machinery , comprising a base part , and at least one additional device being mechanically coupled to said part in order to influence a vibration characteristic of said part during operation of the turbo machine , wherein a mechanical coupling between said part and said at least one additional device changes with a temperature of said at least one additional device.2. Component as claimed in claim 1 , wherein said at least one additional device is a device claim 1 , which changes with temperature its form and position relative to said base part in order to establish an additional mechanical contact between said part and said at least one additional device within a predetermined temperature range.3. Component as claimed in claim 2 , wherein said at least one additional device is a bi-metallic device.4. Component as claimed in claim 2 , wherein said at least one additional device is a shape-memory-alloy device.5. Component as claimed in claim 2 , wherein said additional mechanical contact is a stiffening contact claim 2 , which mechanically stiffens said part.6. Component as claimed in claim 2 , wherein said additional mechanical contact is a friction contact claim 2 , which dampens vibrations in said part.7. Component as claimed in claim 2 , wherein said at least one additional device has the form of a longitudinal beam or curved plate claim 2 , which is fixedly connected at both ends to said part claim 2 , such that it ...

TURBOMACHINE COMPONENT OR COLLECTION OF COMPONENTS AND ASSOCIATED TURBOMACHINE

The present invention relates to a turbomachine component () or collection of components comprising at least a first and a second blade (I, E) and a platform () from which the blades (I, E) extend, characterized in that the platform (), between the pressure face of the first blade (I) and the suction face of the second blade (E) has a non-axisymmetric surface (S) defining a plurality of fins () of substantially triangular section extending downstream of a leading edge (BA) of each of the blades (I, E), each fin () being associated with a leading position and a trailing position on the surface (S), between which positions the fin () extends, such that: the leading position is situated at between 5% and 35% length relative to a chord of the blade (I, E) extending from a leading edge (BA) to a trailing edge (BF) of the blade (I, E);—the further a fin () is from the suction face of the second blade (E), the further the leading position of said fin () is axially from the leading edge (BA) of the blades (I, E). 1. A part or set of parts of a turbomachine comprising:at least first and second blades, and a platform from which the blades extend, wherein the leading position is located at between 5% and 35% of a relative length of a chord of a blade extending from a leading edge to a trailing edge of the blade;', 'the more a fin is separated from the extrados of the second blade, the more the leading position of said fin is axially separated from the leading edge of the blades., 'the platform has, between an intrados of the first blade and an extrados of the second blade, a non-axisymmetric surface defining a plurality of fins with a substantially triangular section extending downstream of a leading edge of each of the blades, each fin being associated with a leading position and a trailing position on the surface between which the fin extends, such that2. The part or set of parts according to claim 1 , wherein each fin has a width comprised between 5% and 20% of a distance ...

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 ...

Attachment Faces for Clamped Turbine Stator of a Gas Turbine Engine

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

GAS TURBINE COMPRESSOR PASSIVE CLEARANCE CONTROL

A gas turbine engine is disclosed having a turbine, one or more hydrocarbon gas combustors, and a compressor. The compressor has a rotor assembly with one or more rotor blade rows extending radially outward from an inner wheel disk. The compressor also has a stator assembly with one or more stator vane rows extending radially inward from an inner casing and positioned between adjacent rotor blade rows. The inner casing extends circumferentially around the rotor assembly and is constructed from at least one low-alpha metal alloy. 1. A compressor for a gas turbine , comprising:a rotor assembly comprising one or more rotor blade rows comprising circumferentially spaced-apart rotor blades, each rotor blade extending radially outward from an inner wheel disk;a stator assembly comprising one or more stator vane rows comprising circumferentially spaced-apart stator vanes extending radially inward from an inner casing, each stator vane row positioned between adjacent rotor blade rows, the inner casing extending circumferentially around the rotor assembly thereby forming a plurality of inner flow paths defined by the rotor blades cooperating with the stator vanes, the rotor blades exhibiting a hot running rotor tip clearance and a cold build rotor tip clearance; andwherein said inner casing comprises at least one low-alpha metal alloy.2. The compressor according to wherein the at least one low-alpha metal alloy exhibits a coefficient of thermal expansion in the range of about 12 microns/meter/degrees Kelvin or less.3. The compressor according to wherein the inner casing comprises a low-alpha metal alloy having an alpha less than the alpha of the rotor blades.4. The compressor according to wherein the at least one low-alpha metal alloy is selected from the group consisting of aluminum claim 1 , iron claim 1 , nickel claim 1 , titanium claim 1 , cobalt claim 1 , niobium claim 1 , iron claim 1 , carbon claim 1 , chromium or mixtures thereof.5. The compressor according to ...

Stator vane assembly for a gas turbine engine

A gas turbine engine has a stator vane assembly. The stator vane assembly includes an inner diameter shroud, an outer diameter shroud located radially outward from the inner diameter shroud, a vane extending radially outward from the first inner diameter shroud to the outer diameter shroud. The wedge clip is positioned horizontally through the vane to prevent the vane from being dislodged from the stator vane assembly.

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.

PROTECTIVE EDGE FOR A BLADE AND METHOD OF MANUFACTURING SAID EDGE

A method of manufacturing a protective edge for a blade, wherein a protective edge () made of an anodizable metal is provided, and that protective edge () undergoes a micro-arc oxidation electrolytic treatment. Protective edge () manufactured using said method. 1. A method of manufacturing a protective edge intended to form a leading edge for a blade , wherein:a protective edge made of an anodizable metal is provided, the protective edge having an outer face and an internal face opposed to each other, the outer face being configured to define an aerodynamic surface of the blade, and the internal face being configured to be glued on a body of the blade to form the leading edge of the blade, and that said protective edge undergoes a micro-arc oxidation electrolytic treatment that is applied to both the inner face and the outer face of the protective edge.2. The method according to claim 1 , wherein the protective edge has an outer face defining in part the aerodynamic surface of the blade claim 1 , and wherein the outer face is polished after the micro-arc oxidation electrolytic treatment.3. The method according to claim 1 , wherein the protective edge is glued on a blade body.4. The method according to claim 1 , wherein the micro-arc oxidation electrolytic treatment comprises the following steps:the protective edge is immersed in an electrolytic bath, the protective edge forming a first electrode,a second electrode is immersed in the electrolytic bath, anda voltage is applied to the first and second electrodes.5. The method according to claim 4 , wherein when the voltage is applied claim 4 , a current is imposed whose intensity has pulses.6. The method according to claim 4 , wherein the second electrode is arranged in the electrolytic bath facing the inner face of the protective edge.7. The method according to claim 1 , wherein the protective edge is made of titanium or titanium alloy.8. A protective edge manufactured using the method according to .9. A blade ...

SPLITTER NOSE OF A LOW-PRESSURE COMPRESSOR OF AN AXIAL TURBOMACHINE WITH ANNULAR DEICING CONDUIT

The invention comprises a splitter nose of an axial turbomachine, in particular a compressor, the splitter nose comprising: an annular casing an annular conduit for de-icing a separation edge of the splitter nose; the conduit is connected to the casing only in a first zone in the region of a hot air inlet and in a second zone located in a position diametrically opposite the inlet, or forming relative to the axis of the turbomachine an angle α less than 30° with respect to the position so as to allow expansion deformations of the conduit. The invention also comprises a compressor and a turbomachine comprising such a splitter nose. 1. A splitter nose of an axial turbomachine , said splitter nose comprising:an annular casing that forms an annular cavity and a circular separation edge of an air flow of the turbomachine; andan annular conduit that is arranged in the annular cavity, the annular conduit being configured to de-ice the separation edge by circulation of hot air in the cavity, and the conduit comprises an air inlet that is structured and operable to be connected to a hot air supply pipe of the turbomachine, the air inlet forming a first zone; whereinthe conduit is connected to the casing only in a second zone, diametrically opposite the air inlet, and in the region of the air inlet, the second zone forming an angular portion of the annular conduit that is less than 60° so as to allow expansion deformations of the conduit.2. The splitter nose of claim 1 , wherein the second zone forms an angular portion of the annular conduit that is at most 30°.3. The splitter nose of claim 1 , wherein the annular casing comprises an internal surface that delimits the cavity claim 1 , that is free from fixation in contact with the annular conduit over at least 120°.4. The splitter nose of further comprising a flange joining the second zone of the annular conduit to the annular casing claim 1 , the flange extending radially and axially.5. The splitter nose of claim 1 , wherein ...

FAN ASSEMBLY

A fan assembly includes a first fan and a second fan spaced from the first fan. The first fan includes a first central portion and a plurality of first fan blades being supported by the first central portion. The second fan includes a second central portion and a plurality of second fan blades being supported by the second central portion. The fan assembly also includes a single drive unit configured to selectively operate both of the first and second fans. The single drive unit is spaced from at least one of the first and second fans to indirectly operate at least one of the fans. 1. A fan assembly comprising:a first fan including a first central portion disposed along a first axis and a plurality of first fan blades being supported by the first central portion and spaced radially away from the first axis;a second fan spaced from the first fan, with the second fan including a second central portion disposed along a second axis and a plurality of second fan blades being supported by the second central portion and spaced radially away from the second axis; anda single drive unit configured to selectively operate both of the first and second fans, with the single drive unit being spaced from at least one of the first and second fans to indirectly operate at least one of the fans.2. The assembly as set forth in wherein the single drive unit is coupled to one of the first and second central portions to directly operate the corresponding one of the fans claim 1 , and the single drive unit is spaced from the other one of the first and second fans to indirectly operate the other one of the fans.3. The assembly as set forth in wherein one of the first and second fans includes a plurality of magnets claim 2 , with the single drive unit interacting with the magnets in a spaced apart relationship when indirectly operating the respective one of the first and second fans which causes one of the plurality of first and second fan blades to rotate about the first and second axes ...

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

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

POWER EQUIPMENT COOLING SYSTEM

A method of cooling blower components may include rotating a fan assembly () responsive to operation of a motor (). The rotation of the fan assembly may create an underpressure region in an inlet portion () and an overpressure region at a main channel () of the fan assembly (). The method may further include drawing air into the inlet portion () through a control unit housing portion () of a housing () of the blower () to cool a control unit () of the blower (), drawing air into the inlet portion () through a battery compartment () to cool a battery () of the blower (), and pushing air from the main channel () through a motor housing () to cool the motor (). 1. A blower comprising:a housing including a handle portion;a motor provided in a motor housing;a fan assembly operably coupled to the motor to force air through a blower tube responsive to operation of the motor; andcontrol circuitry provided in a control unit housing portion of the housing to selectively apply power to the motor for operation of the motor,wherein the fan assembly draws air into an inlet portion to be expelled at an outlet of the blower tube responsive to operation of the motor, andwherein the control unit housing portion includes at least one inlet aperture in the housing to draw cooling air to cool the control unit prior to expelling the cooling air into the inlet portion via at least one outlet aperture.2. The blower of claim 1 , wherein the motor housing further comprises at least one motor housing inlet aperture to enable a portion of the air passing through the fan assembly to enter the motor housing to cool the motor.3. The blower of claim 2 , wherein the portion of the air that enters the motor housing moves in a direction parallel to a tube axis of the blower tube prior to exiting the motor housing into the blower tube.4. The blower of claim 2 , wherein the at least one motor housing inlet aperture is located between a fan blade and a stator blade of the fan assembly.5. The blower of ...

WORK PIECE HAVING SELF-SUPPORTING GUSSET AND METHOD RELATED THERETO

A structure includes a first body section that has a wall that spans in a vertical direction. The wall has a relatively thin thickness with respect to a length and a width of the wall. A second body section is arranged next to, but spaced apart from, the first body section. A gusset connects the first body section and the second body section. The gusset extends obliquely from the wall of the first body section with respect to the vertical direction such that the gusset is self-supporting. The first body section has a geometry that corresponds to an end-use component exclusive of the gusset. 1. A structure comprising:a first body section including a wall spanning in a vertical direction and having a relatively thin thickness with respect a length and a width of the wall;a second body section next to, but spaced apart from, the first body section; anda gusset connecting the first body section and the second body section, the gusset extending obliquely from the wall of the first body section with respect to the vertical direction such that the gusset is self-supporting, and the first body section having a geometry corresponding to an end-use component exclusive of the gusset.2. The structure as recited in claim 1 , wherein the first body section is an airfoil.3. The structure as recited in claim 1 , where the first body section and the second body section are airfoils.4. The structure as recited in claim 1 , wherein the gusset extends at an angle of 45°±5° with respect to the vertical direction.5. The structure as recited in claim 1 , wherein the gusset extends at an angle of no greater than 70° with respect to the vertical direction.6. The structure as recited in claim 1 , wherein the gusset includes a first section and a second section oriented perpendicularly to the first section.7. The structure as recited in claim 6 , wherein the first section is directly connected to the first body section and the second section is directly connected to the second body section.8. ...

Vane with Sealed Lattice in a Shroud of an Axial Turbomachine Compressor

The present application relates to a stator of an axial turbomachine compressor. The stator includes a circular wall, such as an internal shroud, with a guiding surface in order to guide the primary flow of the turbomachine. The stator further includes a circular row of stator vanes, each of them including an airfoil which extends radially in the primary flow of the turbomachine, and a securing portion. The securing portion of the vane includes a lattice which has rods and which is secured and/or sealed in the shroud in order to fix the vanes to the shroud via the lattices. The stator includes a joint of abradable material which is arranged inside the internal shroud, and in which the lattice is secured in order to ensure retention, a fixing between the vane and the internal shroud. The vane is produced by means of additive production.

NACELLE ASSEMBLY

The present disclosure relates generally to a fan nacelle assembly circumferentially surrounding a fan section, the fan nacelle assembly including an inner wall including an inner wall axial length, and an outer wall an outer wall axial length, wherein the outer wall axial length is greater than the inner wall axial length. 1. A gas turbine engine including a fan section , the fan section comprising:a plurality of fan blades;a plurality of guide vanes positioned aft of the plurality of fan blades;a fan containment case circumferentially surrounding the plurality of fan blades and the plurality of guide vanes, the fan casing comprising a first casing coupling member located on an outer surface forward of the plurality of fan blades, and a second casing coupling member located on the outer surface aft of the plurality of guide vanes; and an inner wall including an inner aft end, wherein the inner aft end is operably coupled to the first casing coupling member;', 'an outer wall including an outer wall aft end, wherein the outer wall aft end is positioned aft of the first casing coupling member., 'a fan nacelle assembly circumferentially surrounding the fan casing, the fan nacelle assembly comprising2. The gas turbine of engine of claim 1 , wherein the outer wall is monolithic.3. The gas turbine engine of claim 1 , further comprising an engine accessory affixed to the outer surface of the fan containment case claim 1 , wherein the engine accessory includes an accessory axial length.4. The gas turbine engine of claim 3 , wherein the outer wall further comprises an access opening disposed therein claim 3 , wherein the access opening is located adjacent to the engine accessory.5. The gas turbine engine of claim 3 , wherein the fan nacelle assembly further comprises a fan compartment cowling panel operably coupled to the outer wall aft end and the second casing coupling member claim 3 , wherein the fan cowl panel includes a panel axial length.6. The gas turbine engine of ...

Perforated Drum of a Compressor of an Axial Turbine Engine

A rotor, in particular a drum of a low-pressure compressor of a turbojet aero engine, is disclosed. The rotor includes an outer annular wall delimiting a primary annular flow of the turbine engine, sealing devices with two rubbing strips or annular ribs formed on the wall. The rubbing strips cooperate by abrasion with inner shrouds. In addition, the annular wall includes rows of intake orifices for leakages which are arranged between each pair of rubbing strips in order to aspirate the recirculation leakages there. A plenum for leakages is formed inside the rotor by means of a composite partitioning, then evacuates the parasitic flow downstream of the turbine engine via the central shaft. 1. A rotor of an axial turbine engine , the rotor comprising:a rotation axis;an outer annular wall around the rotation axis; anda sealing device formed on the wall, 'at least one intake orifice for leakages, the intake orifice being disposed within the sealing device in order to divert the leakages therefrom and to evacuate the leakages axially beyond the rotor.', 'wherein the annular wall comprises2. The rotor of claim 1 , wherein the wall surrounds an annular space which communicates with the at least one intake orifice.3. The rotor of claim 1 , further comprising:at least one annular row of rotor blades which is carried by the annular wall and arranged upstream of at least one orifice and of the sealing device.4. The rotor of claim 1 , wherein the wall comprises:several orifices forming at least one circular row or several circular rows.5. The rotor of claim 1 , further comprising:dynamic balancing elements to compensate for the presence of the intake orifice.6. The rotor of claim 1 , wherein the annular wall comprises:escape openings directed upstream, which communicate with escape piercings of a fan disc.7. The rotor of claim 1 , wherein the wall comprises:a revolution profile with a portion extending principally axially and a portion extending principally radially, at least ...

ASSEMBLY WITH MISTAKE PROOF BAYONETED LUG

An assembly for rotational equipment with an axial centerline. The assembly includes a first component and a second component. Each of the components extends circumferentially around and axially along the centerline. The first component includes a flange and a lug aperture extending axially through the flange. The second component includes a mount base and a bayoneted lug on the mount base. The mount base is configured to axially engage the flange where the bayoneted lug extends through the lug aperture. A fastener secures the components together. The fastener projects axially into a fastener aperture in the mount base for an axial length that is less than or equal to an axial length of the bayoneted lug. 1. An assembly for rotational equipment with an axial centerline , comprising:a plurality of components including a first component and a second component, each of the components extending circumferentially around and axially along the centerline;the first component including a flange and a lug aperture extending axially through the flange;the second component including a mount base and a bayoneted lug on the mount base, and the mount base configured to axially engage the flange where the bayoneted lug extends through the lug aperture; anda fastener securing the components together, the fastener projecting axially into a fastener aperture in the mount base for an axial length that is less than or equal to an axial length of the bayoneted lug.2. The assembly of claim 1 , wherein the axial length that the fastener projects into the fastener aperture is less than the axial length of the bayoneted lug.3. The assembly of claim 1 , whereinthe bayoneted lug comprises a lug base and a lug bayonet;the lug base projects axially out from the mount base;the lug bayonet projects laterally out from the lug base; andthe flange is configured to be received within a channel axially between the lug bayonet and the mount base.4. The assembly of claim 3 , wherein the lug bayonet ...

METHOD FOR IMPROVING TURBINE COMPRESSOR PERFORMANCE

A method and device for retrofitting a gas turbine engine for improved hot day performance are disclosed. The method can include removing a first selected stator bladerow from the plurality of compressor stages, the first selected stator bladerow having a first inlet swirl angle and including a first plurality of fixed stator vanes. Each stator vane of the first plurality of fixed stator vanes can have a first stator vane angle. The method can also include providing a first improved stator bladerow to replace the first selected stator bladerow. The first improved stator bladerow can have a second plurality of fixed stator vanes, each having a second stator vane angle smaller than the first stator vane angle. The method can also include replacing the first selected stator bladerow with the first improved stator bladerow to produce an increased pressure ratio and flow rate compared to the first selected stator bladerow. 1. A method for retrofitting a gas turbine engine for improved hot day performance , the gas turbine engine having a plurality of compressor stages , each compressor stage of the plurality of compressor stages having a rotor bladerow and a stator bladerow , the method comprising:removing a first selected stator bladerow from the plurality of compressor stages, the first selected stator bladerow having a first inlet swirl angle and including a first plurality of fixed stator vanes, each stator vane of the first plurality of fixed stator vanes having a first stator vane angle referenced to a central axis of the gas turbine engine;providing a first improved stator bladerow to replace the first selected stator bladerow, the first improved stator bladerow having a second plurality of fixed stator vanes, each stator vane of the second plurality of fixed stator vanes having a second stator vane angle smaller than the first stator vane angle;installing the first improved stator bladerow in place of the first selected stator bladerow, the first improved stator ...

SHIELD FOR A TURBINE ENGINE AIRFOIL

An airfoil for a turbine engine can include a body having an outer wall defining a pressure side and a suction side, and the body can extend axially between a truncated nose and a trailing edge. A shield can be positioned upstream of the nose to define a leading edge for the airfoil. 1. An airfoil for a turbine engine comprising:a body having an outer wall defining a pressure side and a suction side opposite the pressure side, the outer wall enclosing an interior and extending axially between a truncated nose and a trailing edge; anda shield positioned upstream of the nose to define a leading edge for the airfoil, with the shield spaced from the nose to define a gap between the body and the shield.2. The airfoil of wherein the shield further comprises an apex that defines the leading edge.3. The airfoil of wherein the shield further comprises an upstream surface and a downstream surface claim 1 , where the downstream surface confronts the nose.4. The airfoil of wherein the shield further comprises at least one internal cooling passage.5. The airfoil of wherein the shield further comprises at least one film hole extending from the at least one internal cooling passage to one of the upstream surface or downstream surface.6. The airfoil of wherein the shield further comprises at least one film hole extending from the downstream surface to the upstream surface.7. The airfoil of wherein the body further comprises at least one cooling hole fluidly coupling the interior of the body to the gap.8. The airfoil of wherein the body further comprises at least one cooling hole fluidly coupling the interior of the body to the gap.9. The airfoil of wherein the gap is continuous.10. The airfoil of wherein the gap comprises multiple discrete gaps.11. The airfoil of wherein a span-wise extent of the shield is greater than or equal to a span of the nose.12. The airfoil of wherein the shield has at least a portion being made of a material with a higher temperature capability than the ...

Geared gas turbine engine with reduced fan noise

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

Turbine airfoil with passive morphing structure

A turbine engine airfoil apparatus, including an airfoil defined by a plurality of airfoil sections arrayed along a stacking axis that extends between a root and a tip, wherein at least two of the airfoil sections spaced apart from each other have differing airfoil section thermal expansion properties.

Endwall Controuring

An airfoil array may include an endwall, and a plurality of airfoils radially projecting from the endwall. Each airfoil may have a first side and an opposite second side extending axially in chord between a leading edge and a trailing edge. The airfoils may be circumferentially spaced apart on the endwall thereby defining a plurality of flow passages between adjacent airfoils. The airfoil array may further include a convex profiled region extending from the endwall adjacent to the first side of at least one of said plurality of airfoils near the leading edge of the at least one of said plurality of airfoils, and a concave profiled region in the endwall and extending across said at least one of said plurality of flow passages.

COMPRESSOR STATOR

A stator that has vanes extending between a first end and a second end along a span and from a leading edge to a trailing edge along a chord length, the vanes having a first end portion extending from the first end to about 30% of the span to a first location, a chord ratio of the chord length at the first end to the chord length at the first location greater than or equal to 1.1, a throat ratio of a width of a throat between two adjacent vanes at the first location to a width of the throat at the first end is greater than or equal to 1.3, a sweep angle difference between a maximum sweep angle of the leading edge along the first end portion and a minimum sweep angle of the leading edge along the first end portion is at least 15 degrees. 1. A stator having a central axis , the stator comprising: vanes circumferentially distributed around the central axis , the vanes extending between a first end and a second end along a span and from a leading edge to a trailing edge along a chord length , the vanes having a first end portion extending from the first end to about 30% of the span to a first location , a chord ratio of the chord length at the first end to the chord length at the first location greater than or equal to 1.1 , a throat ratio of a width of a throat between two adjacent vanes at the first location to a width of the throat at the first end is greater than or equal to 1.3 , a sweep angle difference between a maximum sweep angle of the leading edge along the first end portion and a minimum sweep angle of the leading edge along the first end portion is at least 15 degrees.2. The stator of claim 1 , wherein the first end is a radially inner end of the vane.3. The stator of claim 1 , wherein the first end is a radially outer end of the vane.4. The stator of claim 1 , wherein each of the vanes has a second end portion extending from the second end along about 30% of the span to a second location claim 1 , the chord ratio of the chord length at the second end to ...

Axial Flow Compressor, Gas Turbine Including the Same, and Stator Blade of Axial Flow Compressor

An axial flow compressor includes multiple rotor blade rows configured to include multiple rotor blades and multiple stator blade rows configured to include multiple stator blades, the multiple rotor blades and the multiple stator blades being arranged in an annular channel through which a working fluid flows. A portion of at least one wall surface on an inner peripheral side and an outer peripheral side of the annular channel, the portion being at an arrangement portion where at least any one blade row of the rotor blade rows and the stator blade rows is located, has a protruding portion such that downstream side part of the portion is curved so as to further protrude to the annular channel than upstream side part of the portion. Each blade of the blade row located at the protruding portion of the wall surface is configured such that an increase rate in a wall surface direction of a blade outlet angle in a blade end portion on the side of the wall surface having the protruding portion is greater than an increase rate in the wall surface direction of a blade outlet angle in a blade height intermediate portion. 1. An axial flow compressor comprising:multiple rotor blade rows configured to include multiple rotor blades and multiple stator blade rows configured to include multiple stator blades, the multiple rotor blades and the multiple stator blades being arranged inside an annular channel through which a working fluid flows,wherein a portion of at least one wall surface on an inner peripheral side and an outer peripheral side of the annular channel, the portion being at an arrangement portion where at least any one blade row of the rotor blade rows and the stator blade rows is located, has a protruding portion such that downstream side part of the portion is curved so as to further protrude to the annular channel than upstream side part of the portion, andwherein each blade of the blade row located at the protruding portion of the wall surface is configured such ...

FAN ASSEMBLY

A fan assembly includes a body and a nozzle connected to the body. The body includes an impeller, a motor housing, and a motor located within the motor housing for driving the impeller. The impeller has a substantially conical hub connected to the motor, a plurality of curved vanes connected to the outer surface of the hub, and an annular vane connected to the base of the hub. The motor housing has an annular groove for receiving at least the tip of the annular vane, and an annular lip forming a drip edge extending about the motor housing, the annular lip defining an outer peripheral wall of the groove. 1. A fan assembly comprising a body and a nozzle connected to the body , the body comprising:an impeller;a motor housing; anda motor located within the motor housing for driving the impeller;wherein the impeller comprises a substantially conical hub connected to the motor, a plurality of curved vanes connected to the outer surface of the hub, and an annular vane connected to the base of the hub,and wherein the motor housing comprises an annular groove for receiving at least the tip of the annular vane, and an annular lip forming a drip edge extending about the motor housing, the annular lip defining an outer peripheral wall of the groove.2. The fan assembly of claim 1 , wherein the annular vane is substantially cylindrical in shape.3. The fan assembly of claim 1 , wherein the annular groove is formed in substantially frustoconical wall of the motor housing which is located adjacent to the hub of the impeller.4. The fan assembly of claim 1 , wherein the annular lip does not protrude downwardly from the motor housing beyond the hub of the impeller. This application claims the priority of United Kingdom Application No. 1413426.6, filed Jul. 29, 2014, the entire contents of which are incorporated herein by reference.The present invention relates to a fan assembly. In a preferred embodiment, the present invention provides a humidifying apparatus for generating a flow of ...

ARTICLE FOR HIGH TEMPERATURE SERVICE

An article includes a substrate that is substantially opaque to visible light and a coating disposed on the substrate. The coating includes a coating material having an inherent index of refraction, wherein the coating has an effective index of refraction that is less than the inherent index of refraction, and wherein the effective index of refraction is less than 1.8. 1. An article , comprising:a substrate that is substantially opaque to visible light; anda coating disposed on the substrate, wherein the coating comprises a coating material having an inherent index of refraction, wherein the coating has an effective index of refraction that is less than the inherent index of refraction, and wherein the effective index of refraction is less than 1.8, and wherein the coating comprises a plurality of columnar structures oriented such that a longitudinal axis of a columnar structure forms an angle with respect to a direction tangential to the substrate that is less than 90 degrees.2. The article of claim 1 , wherein the substrate comprises a metallic material claim 1 , a ceramic material claim 1 , or an intermetallic material.3. The article of claim 1 , wherein the substrate comprises a titanium alloy claim 1 , a superalloy claim 1 , or a ceramic-matrix composite.4. The article of claim 1 , wherein the article comprises a component for a turbine assembly.5. The article of claim 4 , wherein the component is a compressor blade or compressor vane.6. The article of claim 1 , wherein the coating comprises a plurality of columnar structures oriented such that a longitudinal axis of a columnar structure forms an angle with respect to a direction tangential to the substrate that is less than 90 degrees.7. The article of claim 1 , wherein the angle is less than 80 degrees.8. The article of claim 1 , wherein the angle is less than 60 degrees.9. The article of claim 1 , wherein the plurality of columnar structures has a nominal intercolumnar spacing of less than about 5 ...

ASSEMBLY AND METHOD FOR INFLUENCING FLOW THROUGH A FAN OF A GAS TURBINE ENGINE

Assemblies and methods for providing injection air to influence flow in a flow passage defined by a fan of a gas turbine engine are disclosed. In one embodiment, the method comprises: receiving air into an interior of a nose cone; increasing the pressure of the air in the interior of the nose cone and directing the pressurized air; and discharging the air into the flow passage defined by the fan. 1. A fan assembly for a turbofan engine , the fan assembly comprising:a fan comprising a plurality of circumferentially distributed fan blades extending from a hub, the fan blades and the hub defining a flow passage through which ambient air is propelled;a nose cone disposed upstream of the hub and having an interior in fluid communication with the ambient air; anda pump at least partially housed in the interior of the nose cone, the pump being configured to, using the ambient air in the nose cone, drive injection air into the flow passage defined by the fan blades and the hub to influence flow in the flow passage.2. The fan assembly as defined in claim 1 , wherein the hub defines one or more injection passages for directing injection air from the pump to one or more locations in the flow passage.3. The fan assembly as defined in claim 3 , wherein the one or more injection passages comprise one or more respective openings formed in an outer surface of the hub.4. The fan assembly as defined in claim 1 , wherein the pump comprises a plurality of rotor blades secured for common rotation with the fan.5. The fan assembly as defined in claim 4 , wherein the pump comprises a circular array of the rotor blades.6. The fan assembly as defined in claim 4 , wherein the plurality of rotor blades are secured to an inner surface of the outer wall of the nose cone.7. The fan assembly as defined in claim 5 , wherein the pump comprises a first stage including a first circular array of the rotor blades and a second stage including a second circular array of the rotor blades claim 5 , the ...

HEAT DISSIPATION MODULE

A heat dissipation module is provided, including a frame, a base, and a fan. The frame has a depressed portion at an inner edge thereof. The base is connected to the frame and has a bottom surface. The fan is movably disposed on the base and has a plurality of blades. The end of an upper edge of the blades has a first height with respect to the bottom surface. The depressed portion has a second height with respect to the bottom surface, wherein the first height is less than the second height. 1. A heat dissipation module , comprising:a frame, having an outward expansion portion at an inner edge of the frame and adjacent to an air inlet of the heat dissipation module, wherein the outward expansion portion forms a depressed structure and has a curved surface or a sloping surface at the inner edge;a base, connected to the frame and having a bottom surface; anda fan, movably disposed on the base and rotatable with respect to the frame, wherein the fan has a plurality of blades, the end of an upper edge of each of the blades has a first height with respect to the bottom surface, and the outward expansion portion has a second height with respect to the bottom surface, wherein the first height is less than the second height.2. The heat dissipation module as claimed in claim 1 , wherein the end of the upper edge of each of the blades forms a curved structure.3. The heat dissipation module as claimed in claim 1 , wherein the upper edges of the blades are parallel to the bottom surface.4. The heat dissipation module as claimed in claim 1 , wherein the frame has a plurality of stationary vanes disposed on an air outlet of the frame and connected to the base claim 1 , a first angle is formed between the upper edges of the stationary vanes and the bottom surface claim 1 , and a second angle is formed between the lower edges of the stationary vanes and the bottom surface claim 1 , wherein the first angle and the second angle are acute angles.5. The heat dissipation module as ...

VANE ACTUATION SYSTEM FOR A GAS TURBINE ENGINE

A compressor for use in a gas turbine engine is disclosed herein. The compressor includes a case, vanes, and a vane actuation system. The vanes are arranged circumferentially adjacent to one another about a central axis inside the case and each one of the vanes is configured for rotation about a vane axis generally perpendicular to the central axis. The vane actuation system is configured to cause rotation of at least one of the vanes about the vane axis. 1. A compressor comprisinga case extending circumferentially about a central axis,a plurality of vanes arranged circumferentially adjacent to one another about the central axis inside the case, each one of the vanes configured for rotation about a vane axis generally perpendicular to the central axis, anda vane actuation system including an actuator and an actuation ring coupled to one of the vanes and arranged radially inward of a portion of the case relative to the central axis, the actuator configured to drive movement of the actuation ring to cause rotation of the one of the vanes about the vane axis.2. The compressor of claim 1 , wherein the vane actuation system includes a first actuation arm arranged radially outward of the portion of the case relative to the central axis and coupled to the actuator and the actuation ring to cause movement of the first actuation arm driven by the actuator to drive movement of the actuation ring.3. The compressor of claim 2 , wherein the vane actuation system includes a second actuation arm arranged radially inward of the portion of the case relative to the central axis and coupled to the actuation ring and the one of the vanes to cause movement of the second actuation arm driven by the actuation ring to drive rotation of the one of the vanes about the vane axis.4. The compressor of claim 3 , wherein the actuation ring is formed to include a groove sized to receive a follower and the follower is coupled directly to the second actuation arm such that movement of the actuation ...

TURBOMACHINE FAN FRAME COMPRISING IMPROVED ATTACHMENT MEANS

A turbomachine fan frame including a central hub, a plurality of radial vanes installed on the hub and distributed around the hub, each vane including an inner radial end root at which the vane is fixed to the hub, a plurality of flow stream reconstitution platforms, each platform being installed between the roots of two adjacent vanes and radially at a distance from the hub, wherein each platform is fixed to the hub by at least one support element associated with the platform, the support element located between the associated platform and the hub and located between the roots of two adjacent vanes. 1. A turbomachine fan frame comprising:a central hub, and an outer wheel forming structural elements of the fan frame,a plurality of radial vanes installed on the hub and on the wheel that are distributed around the hub, each vane comprising a radial end at which the vane is fixed to one and/or the other of the structural elements,a plurality of flow stream reconstitution platforms, each platform being installed between the radial ends of two adjacent vanes and radially at a distance from, said structural element to which the platform is fixedat least one support element associated with each platform that is located between the associated platform and the structural element and that is arranged between said radial ends of two adjacent vanes,wherein the support element comprises two cleats for attachment of the support element onto the structural element that are located in a first plane parallel to the principal axis of the structural element and comprises a plate located in a second plane approximately parallel to and radially offset from said first plane on which the associated platform will be mounted.2. The turbomachine fan frame comprising:a central hub,a plurality of radial vanes installed on the hub and distributed around the hub, each vane comprising an inner radial end root at which the vane is fixed to the hub,a plurality of flow stream reconstitution ...

OIL COOLING SYSTEMS FOR A GAS TURBINE ENGINE

A heat exchanger assembly for a gas turbine engine that includes an outer engine case. The heat exchanger assembly includes at least one cooling channel, the at least one cooling channel is configured to receive a flow of fluid to be cooled. At least one first coolant flow duct that is configured to receive a flow of a first coolant, wherein the at least one cooling channel is disposed between a first inlet and a first outlet. The heat exchanger assembly further include at least one second coolant flow duct that is configured to receive a flow of a second coolant, wherein the at least one cooling channel is disposed between a second inlet and a second outlet. 1. A heat exchanger assembly for a gas turbine engine comprising an outer engine case , said heat exchanger assembly comprising:at least one cooling channel adjacent the outer engine case, said at least one cooling channel configured to receive a flow of fluid to be cooled;at least one first coolant flow duct configured to receive a flow of a first coolant from a first inlet to a first outlet, wherein said at least one cooling channel disposed between said first inlet and said first outlet; andat least one second coolant flow duct configured to receive a flow of a second coolant from a second inlet to a second outlet, wherein said at least one cooling channel disposed between said second inlet and said second outlet.2. The heat exchanger assembly in accordance with further comprising at least one ejector disposed downstream of said at least one cooling channel claim 1 , said at least one ejector configured to selectively receive a flow of high pressure fluid and draw the second coolant flow through said at least one second coolant flow duct.3. The heat exchanger assembly in accordance with further comprising a filter positioned between said at least one cooling channel and said at least one ejector claim 1 , said filter configured to remove particulates entrained within the second coolant flow.4. The heat ...

PUMP WITH AXIAL CONDUIT

In one aspect, a pump for moving air includes an inlet, an outlet, an outer housing adapted to couple to an inflatable device, and an inner housing located within the outer housing. An air conduit is defined between the inner housing and the outer housing. A motor is at least partly positioned within the inner housing, and a plurality of vanes are positioned within the air conduit. 1. An inflatable device comprising: a housing including an inlet fluidly coupled to ambient and an outlet fluidly coupled to an inflatable bladder, the housing defining an air conduit;', 'a valve assembly including a valve configured to fluidly couple the outlet of fluid controller to the inflatable device, the valve including a self-sealing diaphragm assembly configured to seal the outlet;', 'an electromechanical device configured to act on the self-sealing diaphragm assembly to open the valve;', 'a pump including a motor and an impeller located within the housing, the pump configured for moving air from the inlet through the air conduit to the outlet; and, 'a fluid controller includingwherein the inflatable device includes an inflatable bladder;wherein a majority of the fluid controller is positioned within a profile of the inflatable bladder in a mounted position and orientation, andwherein in the same mounted position and orientation of the fluid controller, the fluid controller is configured to electromechanically open the valve via the electromechanical device to permit air to exit the inflatable bladder through the fluid controller and to energize the pump to provide air to the inflatable bladder through the fluid controller.2. The inflatable device of claim 1 , further comprising at least one vane positioned within the air conduit claim 1 , wherein the at least one vane includes a sweep.3. The inflatable device of claim 2 , wherein the fluid controller includes an axis claim 2 , wherein the pump moves air through the air conduit parallel to the axis claim 2 , and wherein the at ...

AERODYNAMIC ARM FOR AN AIRCRAFT TURBINE ENGINE CASING

An aerodynamic arm for an aircraft turbine engine casing includes a tubular outer shell having a generally elongate shape extending substantially along an axis. The shell has axial ends configured to be connected to a turbine engine casing. An electrically conductive core extends inside the shell and has ends configured to electrically connect to the ends of the shell. An insulating material occupies a space between the core and the shell. 1. A casing aerodynamic arm for an aircraft turbine engine , comprising:an outer tubular shell having a generally elongated shape extending substantially along an axis (A-A), the shell comprising axial ends configured to connect to a casing of the turbine engine;an electrically conductive core extending inside the shell and having electrical connection ends at each of the ends of the shell; andan insulating material configured to occupy a space provided between the core and the shell, wherein for any cross-section of the aerodynamic arm in a plane perpendicular to the axis (A-A) of the shell, a maximum thickness (Ep) of the aerodynamic arm is between 2 mm and 10 mm, and a chord length (L) of the aerodynamic arm is between 30 mm and 150 mm.2. The aerodynamic arm according to claim 1 , wherein the insulating material has a minimum thickness (e) between 0.6 mm and 1.5 mm.3. The aerodynamic arm according to claim 1 , wherein a thickness of the core is between 1 mm and 5 mm.4. The aerodynamic arm according to claim 1 , wherein the ends of the core are configured to be connected to electrical conductors by one of mechanical connections and welds.5112a. The aerodynamic arm according to claim 1 , wherein each of the axial ends of the shell comprises a flange () configured to connect to the casing of the turbine engine.6. The aerodynamic arm according to claim 1 , wherein the insulating material is configured to withstand temperatures of up to 200° C. and is made from one of a liquid insulator and an organic insulating powder polymerised ...

STATOR BLADE UNIT FOR A TURBOMOLECULAR PUMP

The present invention provides a stator blade unit for a turbomolecular pump comprising an array of polymer stator blades, a turbomolecular pump including such a stator blade unit, and to a method of assembling a stator blade unit for a turbomolecular pump. 1. A stator blade unit for a turbomolecular pump comprising:an array of polymer and/or moulded stator blades comprising an inner rim and an outer rim adjoined to the array of stator blades, wherein the outer rim comprises an integrally formed spacer for preventing clashing between the stator blades and an adjacent rotor, and configured to engage with an axially adjoining stator array within a stator stack and wherein the stator blade unit comprises at least two sections, in the form of single unitary structures, arranged such that stator blades of one section are alternately arranged with stator blades of the at least one other section.2. The stator blade unit according to wherein the stator blades axially overlap.3. The stator blade unit according to wherein the inner rim is configured to couple with an inner rim of another stator blade unit.4. The stator blade unit according to wherein the stator blade unit is injection moulded or additive manufactured claim 1 , preferably an injection moulded polymer and/or metal stator blade unit.5. The stator blade unit according to wherein the stator blade unit comprises polymer stator blades each comprising one or more inner polymer layers and an outermost polymer layer claim 1 , wherein the outermost polymer layer comprises a polymer that is less hard than the polymer forming the one or more inner layers.6. The stator blade unit according to wherein the stator blade unit is substantially semi-annular.7. The stator blade unit according to wherein the stator blade unit is configured to mate with a second substantially semi-annular stator blade array to form an annular stator blade array.8. The stator blade unit according to wherein each stator blade comprises an inlet-side ...

FIRE-FIGHT VENTILATOR WITH OVALISED AIR JET

A fire fighting blower () comprises a propeller () coaxially mounted in a tubular casing () for generating an axial airflow. The blower comprises an airflow guiding device () in the casing () for obtaining a concentrated air jet having a substantially ovalized section, it comprises within the tubular casing () an assembly of first deflectors () for concentrating the axial airflow and generating a concentrated axial air jet and an assembly of second deflectors () for generating an air stream deflected from the concentrated axial air jet so that the concentrated air jet having an ovalized section is a combination of the concentrated axial air jet and the deflected air stream. 1. A fire fighting blower comprising a propeller coaxially mounted in a tubular casing for generating an axial airflow , and an airflow guiding device for obtaining a concentrated air jet having a substantially ovalized section , wherein the fire fighting blower comprises within said tubular casing an assembly of first deflectors for concentrating said axial airflow and generating a concentrated axial air jet and an assembly of second deflectors , including a first right deflector which directs a part of the airflow in the direction of rotation of said propeller and a second direction reversing deflector , angularly spaced from and upstream of the first right deflector in the direction of rotation of said propeller , which directs this part of the airflow in the opposite direction of rotation of said propeller for generating an air stream deflected from the concentrated axial air jet so that said concentrated air jet having an ovalized section is a combination of said concentrated axial air jet and said deflected air stream.2. The blower according to claim 1 , wherein said right deflector and said direction reversing deflector are angularly spaced relative to each other by an angle between 90° and 180°.3. The blower according to claim 2 , wherein said right deflector and said direction reversing ...

INLET ASSEMBLY FOR AN AIRCRAFT AFT FAN

The present disclosure is directed to a low-distortion inlet assembly for reducing airflow swirl distortion entering an aft fan mounted to a fuselage of an aircraft. Further, the inlet assembly includes a plurality of structural members, such as struts or strakes, mounted at predetermined locations around a circumference of the fan shaft of the fan at the inlet. The predetermined location(s) may be determined as a function of swirl distortion entering the inlet. As such, the structural member(s) are configured to reduce swirl distortion of the airflow entering the fan. In some embodiments, the inlet assembly may also include inlet guide vanes. In alternative embodiments, the inlet assembly may be absent of inlet guide vanes. 1. A boundary layer ingestion fan assembly for mounting to an aft end of a fuselage of an aircraft , the boundary layer ingestion fan assembly comprising:a fan rotatable about a central axis of the boundary layer ingestion fan, the fan comprising a plurality of fan blades rotatable about a fan shaft;a nacelle surrounding the plurality of fan blades of the fan, the nacelle defining an inlet with the fuselage of the aircraft, the inlet extending substantially around the fuselage of the aircraft when the boundary layer ingestion fan is mounted at the aft end of the aircraft; one or more inlet guide vanes configured within the inlet; and', 'one or more structural members mounted at predetermined radial locations around a circumference of the fan shaft of the fan at the inlet, the structural members configured to reduce a swirl distortion entering the inlet of the fan., 'a low-distortion inlet assembly configured with the inlet, the inlet assembly comprising2. The boundary layer ingestion fan assembly of claim 1 , wherein the one or more structural members comprise at least one of a strut or a strake.3. The boundary layer ingestion fan assembly of claim 1 , wherein one or more of the structural members are integrated with at least one of the nacelle ...

INLET ASSEMBLY FOR AN AIRCRAFT AFT FAN

The present disclosure is directed to a low-distortion inlet assembly for reducing airflow swirl distortion entering an aft fan mounted to a fuselage of an aircraft. Further, the inlet assembly includes a plurality of structural members mounted at one or more predetermined locations around a circumference of the fan shaft of the fan. More specifically, the predetermined location(s) has a swirl distortion exceeding a predetermined threshold. Further, the inlet assembly includes at least one airflow modifying element configured within the inlet so as to reduce swirl distortion entering the fan. 1. A boundary layer ingestion fan assembly for mounting to an aft end of a fuselage of an aircraft , the boundary layer ingestion fan assembly comprising:a fan rotatable about a central axis of the boundary layer ingestion fan, the fan comprising a plurality of fan blades rotatable about a fan shaft;a nacelle surrounding the plurality of fan blades of the fan, the nacelle defining an inlet with the fuselage of the aircraft, the inlet extending substantially around the fuselage of the aircraft when the boundary layer ingestion fan is mounted at the aft end of the aircraft; one or more structural members mounted at predetermined locations around a circumference of the fan shaft of the fan within the inlet, the predetermined locations comprising a swirl distortion exceeding a predetermined threshold; and', 'at least one airflow modifying element configured within the inlet so as to reduce swirl distortion entering the fan., 'a low-distortion inlet assembly mounted within the inlet, the inlet assembly comprising2. The boundary layer ingestion fan assembly of claim 1 , wherein the one or more structural members comprise at least one of an inlet guide vane or a strut.3. The boundary layer ingestion fan assembly of claim 2 , further comprising a plurality of inlet guide vanes placed in groups at the predetermined locations around the circumference of the fan shaft.4. The boundary ...

Mounting arrangement for aerofoil body

The present disclosure provides an aerofoil body comprising a root portion and a tip portion, each having a pressure surface and a suction surface. The pressure surface and suction surface of the root and/or tip portion each comprise a respective ridge portion. Each ridge portion has an inclined first face and an oppositely inclined second face. Each ridge portion may be, for example, a triangular or semi-circular prism.

METHODS AND SYSTEMS FOR CONTROLLING TURBOCOMPRESSORS

A method for regulating a turbocompressor to prevent surge, is described. The method comprises the following steps: providing at least one surge limit line of turbocompressor; continuously determining an actual value of a corrected speed of the turbocompressor; continuously determining at least a maximum admissible pressure ratio on the surge limit line, corresponding to the actual value of the corrected speed; continuously determining an actual pressure ratio; acting upon an antisurge arrangement, if the actual pressure ratio is equal to or higher than the maximum admissible pressure ratio. 1. A method for regulating a turbocompressor to prevent surge , comprising the following steps:providing at least one surge limit line and/or at least one choke line of said turbocompressor;determining continuously an operating point of the compressor measuring a processing gas temperature at a compressor inlet, the rotary speed of the compressor, a delivery pressure value and a suction pressure value;continuously determining an actual value of a corrected speed of the turbocompressor, the corrected speed being proportional to a ratio of the rotary speed to a square root of the processing gas temperature;continuously determining at least a maximum admissible pressure ratio on said surge limit line and/or at least a minimum admissible pressure ratio on said choke line, corresponding to the actual value of the corrected speed;continuously determining an actual pressure ratio, equal to a ratio between the delivery pressure value and the suction pressure value; andacting upon an antisurge arrangement to recirculate a fraction of a the compressed gas in the compressor through a suction line, if the actual pressure ratio is equal to or higher than the maximum admissible pressure ratio or equal to or lower than the minimum admissible pressure ratio.2. The method of claim 1 , further comprising the step of calculating a surge parameter defined as a ratio ofthe maximum admissible ...

Integrated Environmental Control System Manifold

A compressor intermediate case for a gas turbine engine includes a plurality of intermediate case struts joining the compressor intermediate case to an inner engine structure. Each strut of the plurality of intermediate case struts includes a leading edge. A turning scoop is disposed at the leading edge of each strut of the plurality of intermediate case struts. A plurality of diffusers extends radially outwardly from the compressor intermediate case so that each diffuser of the plurality of diffusers engages with a corresponding turning scoop. A substantially annular structural fire wall extends radially outwardly from the compressor intermediate case. An environmental control system manifold is disposed on the compressor intermediate case. The environmental control system manifold includes an exit port. 1. An intermediate case for a gas turbine engine compressor , the intermediate case comprising:a plurality of intermediate case struts joining the intermediate case to an inner engine structure, each strut of the plurality of intermediate case struts including a leading edge;a turning scoop being disposed at the leading edge of each strut of the plurality of intermediate case struts;a plurality of diffusers extending radially outwardly from the intermediate case, each diffuser of the plurality of diffusers being engaged with a corresponding turning scoop;a substantially annular structural fire wall extending radially outwardly from the intermediate case; andan environmental control system manifold being disposed on the intermediate case, the environmental control system manifold including an exit port.225. The intermediate case of claim 1 , further including a non-structural fairing extending radially outwardly from the intermediate case claim 1 , the non-structural fairing disposed upstream of the annular structural fire wall to define a . bleed duct therebetween.3252525. The intermediate case of claim 2 , further including a . stability bleed valve in operable ...

APPARATUS AND METHOD FOR WATER AND ICE FLOW MANAGEMENT IN A GAS TURBINE ENGINE

A compressor of a gas turbine engine includes at least one rotor, and at least one stator, wherein the at least one stator includes a water channel wall located on a surface of the stator, wherein the water channel wall is configured to direct water away from a center line of the turbine engine and is located an offset distance from an outside wall of the compressor, wherein the offset distance is zero percent to twenty percent of the span of the stator. 1. A stator of a compressor of a turbine engine , the stator having a water channel wall located on a surface of the stator , wherein the water channel wall is configured to direct water away from a center line of the turbine engine and is located an offset distance from an outside wall of the compressor , wherein the offset distance is zero percent to twenty percent of a span of the stator.2. The stator of claim 1 , wherein the water channel wall is integrally formed with the stator.3. The stator of claim 1 , wherein the stator is disposed upstream of a bleed port of the turbine engine.4. The stator of claim 1 , wherein the stator is disposed upstream of a combustor of the turbine engine.5. The stator of claim 4 , wherein the stator is an exit guide vane.6. The stator of claim 1 , wherein the water channel wall is disposed on a pressure side of the stator.7. The stator of claim 1 , wherein the water channel wall includes an outer edge angled towards the outside wall.8. The stator of claim 1 , wherein the water channel wall includes an inner edge angled towards the outside wall.9. A method to extract water within a compressor of a turbine engine claim 1 , the method comprising:introducing an airflow to the turbine engine;directing the airflow towards a surface of a stator of the turbine engine; andcapturing a water content of the airflow via a water channel wall located on a surface of the stator of the turbine engine, wherein the water channel wall is located an offset distance from an outside wall of the ...

SYSTEM FOR REMOVING HEAT FROM TURBOMACHINERY COMPONENTS

A turbomachinery component for a turbomachine includes feed manifold, a return manifold and a sidewall. The feed manifold is configured to receive a coolant stream therein and includes a plurality of feed plenums. The return manifold includes a plurality of return plenums. The sidewall defines a plurality of feed channels and a plurality of return channels therein. The sidewall further includes an inner surface and an outer surface opposite the inner surface. Each feed channel is in fluid communication with at least one of the feed plenums. Each return channel is in fluid communication with at least one of the return plenums. The sidewall further at least partially defines a plurality of microchannels. Each microchannel is in fluid communication with one of the feed channels and one of the return channels. 1. A turbomachinery component comprising:a feed manifold configured to receive a coolant stream therein, said feed manifold comprising a plurality of feed plenums;a return manifold comprising a plurality of return plenums; and an inner surface; and', 'an outer surface opposite said inner surface;, 'a sidewall defining a plurality of feed channels and a plurality of return channels therein, said sidewall comprisingwherein each feed channel of said plurality of feed channels is in fluid communication with at least one feed plenum of said plurality of feed plenums, wherein each return channel of said plurality of return channels is in fluid communication with at least one return plenum of said plurality of return plenums, and wherein said sidewall at least partially defines a first plurality of microchannels adjacent said outer surface, each microchannel of said first plurality of microchannels in fluid communication with a respective feed channel of said plurality of feed channels and a respective return channel of said plurality of return channels.2. The turbomachinery component in accordance with claim 1 , wherein said outer surface comprises at least one pre- ...

AIRFOIL SYSTEMS AND METHODS OF ASSEMBLY

An airfoil assembly includes an airfoil body extending from a root to a tip defining a longitudinal axis therebetween. The airfoil body includes a leading edge between the root and the tip. A sheath is direct deposited on the airfoil body. The sheath includes at least one metallic material layer conforming to a surface of the airfoil body. In accordance with another aspect, a method for assembling an airfoil assembly includes directly depositing a plurality of material layers on an airfoil body to form a sheath. In accordance with some embodiments, the method includes partially curing the airfoil body. 1. An airfoil assembly comprising:an airfoil body extending from a root to a tip defining a longitudinal axis therebetween, wherein the airfoil body includes a leading edge between the root and the tip; anda sheath direct deposited on the airfoil body, wherein the sheath includes at least one metallic material layer conforming to a surface of the airfoil body.2. An airfoil as recited in claim 1 , wherein the sheath is direct deposited on the leading edge of the airfoil body.3. An airfoil as recited in claim 1 , wherein the airfoil body includes a composite material.4. An airfoil as recited in claim 1 , wherein the sheath defines an internal pocket that includes a lattice structure.5. An airfoil as recited in claim 1 , wherein the sheath includes at least one of a composite or fiberglass structure bonded in between layers of the sheath.6. An airfoil as recited in claim 1 , wherein the sheath includes a plurality of layers.7. An airfoil as recited in claim 6 , wherein the layers are alternating material layers.8. An airfoil as recited in claim 6 , wherein an exterior layer includes a material of a higher erosion resistance than an interior layer.9. An airfoil as recited in claim 6 , wherein a first layer in direct contact with the airfoil body includes a material having a lower deposition temperature than layers exterior to the first layer.10. A method for assembling an ...

STRUT FOR AN AIRCRAFT ENGINE

A strut for a gas turbine engine includes a body defining a first side and an opposite second side. The first side is spaced from the second side along the circumferential direction of the gas turbine engine. Additionally, the body includes an inner section, a middle section, and an outer section. Each of the inner, middle, and outer sections are arranged in series order along a span of the strut and define a thickness between the first and second sides of the strut. The thickness of the middle section is greater than the thicknesses of the inner section and of the outer section to increase the strut's resistance to buckling in response to forces exerted thereon. 1. A strut for a gas turbine engine defining a circumferential direction , the strut defining a span and comprising:a body defining a first side and an opposite second side, the first side spaced from the second side along the circumferential direction, the body comprising an inner section, a middle section, and an outer section arranged in series order along the span of the strut, the inner section, middle section, and outer section each defining a thickness between the first and second sides, the thickness of the middle section being greater than the thicknesses of the inner section and the outer section.2. The strut of claim 1 , wherein the strut is formed of a composite material.3. The strut of claim 1 , wherein the thickness of middle section is at least about ten percent (10%) greater than the thicknesses of the inner section and of the outer section.4. The strut of claim 1 , wherein the thickness of middle section is at least about fifteen percent (15%) greater than the thicknesses of the inner section and of the middle section.5. The strut of claim 1 , wherein the inner section comprises an inner thirty percent of the span of the strut claim 1 , wherein the middle section comprises a middle forty percent of the span of the strut claim 1 , and wherein the outer section comprises the outer thirty ...

Fan Module including Coaxial Counter Rotating Fans

A fan module for a vehicle engine cooling system includes a pair of co-axial, counter-rotating, axial flow fans. Each fan is supported on and driven by a dedicated downstream motor, and each motor is supported by a dedicated shroud. The shroud that supports the first motor includes a barrel, a motor carrier that is surrounded by the barrel, and vanes that extend between the barrel and the motor carrier. For each vane, a line that extends between the vane nose and the vane tail is angled relative to the fan rotational axis, and the angle is selected to minimize air flow losses through the shroud.

BRIDGE BRACKET FOR VARIABLE-PITCH VANE SYSTEM

A variable-pitch vane system for a gas turbine engine includes a plurality of vanes and a synchronization ring assembly operably connected to the plurality of vanes. The synchronization ring assembly includes a first synchronization ring, a second synchronization ring and a bridge bracket connecting the first synchronization ring to the second synchronization ring. The bridge bracket includes a first face sheet, a second face sheet, a honeycomb core located between the first face sheet and the second face sheet, a first attachment feature located at a first end of the bridge bracket at which the first synchronization ring is secured, and a second attachment feature located at a second end of the bridge bracket at which the second synchronization ring is secured. 1. A bridge bracket for a variable-pitch vane system of a gas turbine engine , comprising:a first face sheet;a second face sheet;a honeycomb core disposed between the first face sheet and the second face sheet; andone or more attachment features disposed at opposing ends of the bridge bracket, the one or more attachment features configured for securing the bridge bracket to a synchronization ring of the gas turbine engine.2. The bridge bracket of claim 1 , wherein the one or more attachment features comprise one or more spools extending through the bridge bracket from first face sheet to the second face sheet.3. The bridge bracket of claim 2 , wherein the one or more spools are secured in the bridge bracket by one of orbital machining or electron beam welding.4. The bridge bracket of claim 1 , wherein one or more of the first face sheet claim 1 , the second face sheet and the honeycomb core are formed from a metallic material.5. The bridge bracket of claim 1 , further comprising a reinforcing edge portion secured to the bridge bracket.6. The bridge bracket of claim 5 , wherein the reinforcing edge portion extends over the one or more attachment features.7. The bridge bracket of claim 5 , wherein the ...

AIRFOIL SYSTEMS AND METHODS OF ASSEMBLY

An airfoil assembly includes an airfoil body extending from a root to a tip defining a longitudinal axis therebetween. The airfoil body includes a leading edge between the root and the tip. A sheath is direct deposited on the airfoil body. The sheath includes at least one metallic material layer conforming to a surface of the airfoil body. In accordance with another aspect, a method for assembling an airfoil assembly includes directly depositing a plurality of material layers on an airfoil body to form a sheath. In accordance with some embodiments, the method includes partially curing the airfoil body. 1. A method for assembling an airfoil assembly comprising:directly depositing at least one material layer on an airfoil body to form a sheath.2. A method as recited in claim 1 , further comprising partially curing the airfoil body.3. A method as recited in claim 1 , wherein the at least one material layer is one of a plurality of material layers claim 1 , the method further comprising ball milling at least one of the material layers prior to depositing an adjacent one of the material layers.4. A method as recited in claim 1 , wherein directly depositing the at least one material layer includes directly depositing at least one of material layers of alternating materials claim 1 , or groups of material layers of alternating materials.5. A method as recited in claim 1 , wherein the at least one material layer is one of a plurality of material layers claim 1 , the method further comprising bonding at least one of a composite or fiberglass structure between adjacent material layers of the sheath.6. A method as recited in claim 1 , wherein directly depositing the at least one material layer on the airfoil body includes depositing the material layer using a micro plasma spray process. This application is divisional application of U.S. patent application Ser. No. 15/235,291 filed Aug. 12, 2016, the contents of which are incorporated by reference herein in their entirety.The ...

VARIABLE VANE DEVICES CONTAINING ROTATIONALLY-DRIVEN TRANSLATING VANE STRUCTURES AND METHODS FOR THE PRODUCTION THEREOF

Variable vane devices containing rotationally-driven translating vane structures are provided, as are methods for fabricating variable vane devices. In one embodiment, the variable vane device includes a flow assembly having a centerline, an annular flow passage extending through the flow assembly, cam mechanisms, and rotationally-driven translating vane structures coupled to the flow assembly and rotatable relative thereto. The translating vane structures include vane bodies positioned within the annular flow passage and angularly spaced about the centerline. During operation of the variable vane device, the cam mechanisms adjust translational positions of the vane bodies within the annular flow passage in conjunction with rotation of the translating vane structures relative to the flow assembly. By virtue of the translational movement of the translating vane structures, a reduction in the clearances between the vane bodies and neighboring flow assembly surfaces can be realized to reduce end gap leakage and boost device performance. 1. A variable vane device , comprising:a flow assembly having a centerline and an annular endwall partially bounding the flow passage;an annular flow passage extending through the flow assembly;rotationally-driven translating vane structures coupled to the flow assembly and rotatable relative thereto, the rotationally-driven translating vane structures having an angular Range of Motion (ROM) and including vane bodies positioned within the annular flow passage and angularly spaced about the centerline, wherein edge portions of the vane bodies are separated from the annular endwall by radial clearances; andcam mechanisms coupled to the flow assembly and to the rotationally-driven translating vane structures, the cam mechanisms adjusting translational positions of the vane bodies within the annular flow passage as the rotationally-driven translating vane structures rotate relative to the flow assembly, and such that an average value of the ...

ENGINE COMPONENT WITH POROUS TRENCH

An apparatus and method for cooling an engine component, such as an airfoil, including a wall to separate a hot flow from a cooling fluid flow. The component can include at least one trench disposed in a hot surface. The trench can be fed with the cooling fluid flow to cool the engine component along the hot surface with the cooling fluid flow. 1. A component for a turbine engine , which generates a hot flow and provides a cooling fluid flow , the component comprising:a wall defining an interior and separating the hot flow from the cooling fluid flow and having a hot surface facing the hot flow and a cooling surface facing the cooling fluid flow;at least one trench disposed in the hot surface;at least one hole in the wall fluidly coupling the interior to the trench; anda porous material at least partially filling the trench.2. The component of wherein the at least one trench includes multiple trenches.3. The component of wherein the multiple trenches are separate from each other.4. The component of wherein the multiple trenches are parallel to each other.5. The component of wherein the multiple trenches are arranged in a row.6. The component of wherein the multiple trenches include multiple discrete trenches.7. The component of wherein the multiple discrete trenches are arranged in one of a row or parallel to one another.8. The component of wherein the trench includes an arcuate profile defining an apex and the at least one hole is disposed on the apex.9. The component of wherein the at least one hole includes multiple holes.10. The component of wherein the multiples holes are arranged in a row.11. The component of further comprising a solid element extending at least partially along the trench.12. The component of wherein the solid element is shaped to direct a flow of cooling fluid passing through the porous material.13. An airfoil for a turbine engine claim 11 , the airfoil comprising:a wall bounding an interior and defining a pressure side and a suction side ...

Airfoil for a turbine engine with porous rib

An apparatus and method for cooling an engine airfoil, including a wall bounding an interior extending axially between a leading edge and a trailing edge and radially between a root and a tip. A cooling circuit it located within the interior having full-length ribs and partial-length ribs to define the cooling circuit, with the partial length ribs defining a turn.

AIRFOIL FOR A TURBINE ENGINE

A method and apparatus for an airfoil in a gas turbine engine can include an outer surface bounding an interior and spanning from a root to a tip. At least one flow channel can be defined among one or more full-length and partial-length ribs to further define an air flow channel within the airfoil. The air flow channel can have at least one tip turn at the partial-length rib, having at least one hole, for example a film hole, in a portion of the tip. 1. An airfoil for a turbine engine comprising:a wall bounding an interior and defining a pressure sidewall and a suction sidewall, extending chord-wise from a leading edge to a trailing edge and span-wise from a root to a tip defining a tip surface;an air flow channel located within the interior and having a portion located adjacent the tip; andat least one hole extending through the tip and having an inlet fluidly coupled to the air flow channel, an outlet located in the tip surface, and a curvilinear passage fluidly coupling the inlet to the outlet, and the passage adjacent the outlet has a centerline approaching the tip surface at an angle less than 60 degrees.2. The airfoil of wherein the angle is greater than 10 degrees.3. The airfoil of wherein the portion of the air flow channel is a tip turn.4. The airfoil of wherein the outlet is downstream of the inlet.5. The airfoil of wherein the at least one hole comprises a plurality of holes.6. The airfoil of wherein the tip comprises a tip floor defining at least a portion of the tip surface.7. The airfoil of wherein the outlet is located in the tip floor.8. The airfoil of wherein the curvilinear passage extends through the tip floor.9. The airfoil of wherein the tip further comprises at least one tip rail extending above the tip floor and defining at least a portion of the tip surface.10. The airfoil of wherein the outlet is located in the at least one tip rail.11. The airfoil of wherein the passage extends through the tip rail.12. The airfoil of wherein the at least ...

GAS-TURBINE ENGINE COMPOSITE COMPONENTS WITH INTEGRAL 3-D WOVEN OFF-AXIS REINFORCEMENT

A gas turbine engine and a composite apparatus is disclosed. The gas turbine engine includes a composite apparatus body formed from a composite material, the composite material including a three dimensional preform, the three dimensional preform including a plurality of warp fibers disposed in a warp direction in a first plane, a plurality of fill fibers disposed in a fill direction, wherein the fill direction is perpendicular to the warp direction in the first plane, a plurality of z-yarn fibers disposed in a z-yarn direction, wherein the z-yarn direction intersects the warp direction through the first plane, and a plurality of bias fibers disposed in a bias direction, wherein the bias direction is not aligned with the warp direction and the fill direction. 1. A fan blade comprising:a fan blade body formed from a composite material, the composite material including a plurality of warp fibers disposed in a warp direction in a first plane;', 'a plurality of filling fibers disposed in a fill direction, wherein the fill direction', 'is perpendicular to the warp direction in the first plane;', 'a plurality of z-yarn fibers disposed in a z-yarn direction, wherein the z-yarn direction intersects the warp direction through the first plane; and', 'a plurality of bias fibers disposed in a bias direction, wherein the bias direction is not aligned with the warp direction and the fill direction., 'a three dimensional preform, the three dimensional preform including2. The fan blade of claim 1 , wherein the bias direction is disposed in the first plane.3. The fan blade of claim 1 , wherein the fan blade body includes a root portion and an airfoil portion.4. The fan blade of claim 3 , wherein the root portion is thicker than the airfoil portion.5. The fan blade of claim 1 , wherein the fan blade body is formed from a plurality of composite layers.6. A gas turbine engine comprising: a plurality of warp fibers disposed in a warp direction in a first plane;', 'a plurality of filling ...

ROTOR BLADE AND AXIAL FLOW ROTARY MACHINE

A rotor blade () provided in an axial flow compressor () including a rotating shaft, a casing (), a diffuser portion () provided on a downstream side of the casing to communicate with a flow path (C) of the casing and form an annular shape and configured to define a diffuser flow path (DC) in which a cross-sectional area of the flow path expands toward the downstream side, a plurality of stator vane rows (), and rotor blade rows () performing compression of a gas. A plurality of rotor blades are spaced apart from each other in a circumferential direction, and constitute a final rotor blade row (A) positioned on a most downstream side among the rotor blade rows, and include a blade portion () having a larger deflection angle on a hub side and a chip side than at a central portion in a blade height direction. 116-. (canceled)17. A rotor blade provided in an axial flow rotary machine , whereinthe axial flow rotary machine includes:a rotating shaft which extends in the direction of the axis and rotates about the axis;a casing which supports the rotating shaft to be rotatable from an outer circumferential side and defines a flow path of the fluid between the rotating shaft and the casing;a diffuser portion provided on a downstream side of the casing to communicate with the flow path and form an annular shape about the axis and configured to define a diffuser flow path in which a cross-sectional area of the flow path expands toward the downstream side;a plurality of stator vane rows provided in a direction of the axis and protruding from the casing toward a radial inner side of the axis; anda plurality of rotor blade rows provided adjacent to the stator vane rows in the direction of the axis and configured to perform compressing or pressure-feeding of the fluid,the plurality of rotor blades are provided to be spaced apart from each other in a circumferential direction of an axis and configured to constitute a final rotor blade row positioned on a most downstream side of a ...

Methods, apparatus, computer programs, and non-transitory computer readable storage mediums for repairing aerofoils of gas turbine engines

Methods, apparatus, computer programs, and non-transitory computer readable storage mediums for repairing an aerofoil of a gas turbine engine A method of repairing an aerofoil of a gas turbine engine, the method comprising: controlling machining of at least one of: a first aerofoil and a second aerofoil; the machining causing an increase in a throat area defined by the first aerofoil, the second aerofoil, a first platform and a second platform, the first aerofoil and the second aerofoil being arranged to couple to the first platform and to the second platform; and controlling provision of a coating to at least one of the first platform and the second platform subsequent to controlling machining to reduce the throat area.

HIGH EFFICIENCY SELF-CONTAINED MODULAR TURBINE ENGINE POWER GENERATOR

A high efficiency self-contained modular turbine engine unit for generating power includes a housing defining an air intake and an exhaust port. A turbine engine is positioned and operable within the housing. The turbine engine includes a drive shaft a compressor rotor assembly, a compressor vane assembly, a combustor and diffuser assembly, a turbine vane assembly and a turbine rotor assembly. The combustor and diffuser assembly is a one-piece unit defining a shroud extending forwardly therefrom and a plurality of combustion flow channels extending rearwardly and radially inwardly thereby forming a flowpath angle in the range from about 15° to about 35° with the drive shaft. An igniter is positioned in each flow channel to ignite a fuel/oxygen mixture introduced into the compressor rotor assembly. External components required for operation of turbine engine are mounted within the housing. 1. A high efficiency self-contained modular turbine engine unit for generating power , the modular turbine engine unit comprising:a housing having a housing frame, a top panel, a bottom panel, a first side panel, a second side panel, a third side panel and a fourth side panel, each of the panels being removeably secured to the housing frame;an air intake defined in the housing; andan exhaust port defined in the housing; a drive shaft defining a drive shaft centerline,', 'at least one compressor rotor assembly mounted on the drive shaft,', 'at least one compressor vane assembly mounted on the drive shaft proximate to and downstream from the at least one compressor rotor assembly,', 'a combustor and diffuser assembly mounted on the drive shaft proximate to and downstream from the at least one compressor vane assembly, the combustor and diffuser assembly comprising a one-piece unit defining a shroud extending forwardly therefrom to define a flowpath for compressed air exiting the at least one compressor vane assembly, and a plurality of combustion flow channels extending rearwardly ...

SEAL ASSEMBLY FOR ROTATIONAL EQUIPMENT

Assemblies are provided for rotational equipment. One of these assemblies includes a rotor disk structure, a stator structure and a seal assembly. The rotor disk structure includes a rotor disk and a seal land circumscribing the rotor disk. The stator structure circumscribes the seal land. The seal assembly is configured for sealing a gap between the stator structure and the seal land, where the seal assembly includes a non-contact seal. 1. An assembly for rotational equipment , the assembly comprising:a rotor disk structure including a rotor disk and a seal land circumscribing the rotor disk;a stator structure circumscribing the seal land; anda seal assembly configured for sealing a gap between the stator structure and the seal land, wherein the seal assembly includes a non-contact seal.2. The assembly of claim 1 , wherein the non-contact seal is a hydrostatic non-contact seal.3. The assembly of claim 1 , wherein the non-contact seal comprises:an annular base;a plurality of shoes arranged around and radially adjacent the seal land; anda plurality of spring elements, each of the spring elements radially between and connecting a respective one of the shoes to the base.4. The assembly of claim 1 , wherein the seal land is an outer hub of the rotor disk structure.5. The assembly of claim 1 , wherein the rotor disk includes an annular counterweight mass and an annular web extending radially inward to the counterweight mass.6. The assembly of claim 1 , wherein the seal land includes a cylindrical outer surface claim 1 , and the gap extends radially between the stator structure and the outer surface.7. The assembly of claim 6 , wherein the seal land comprises an axially extending annular flange which forms the outer surface.8. The assembly of claim 1 , further comprising:a bladed rotor assembly including a second rotor disk structure; anda linkage extending axially between and connected to the rotor disk structure and the second rotor disk structure.9. The assembly of ...

FAN FRAME

A fan frame includes a metal base and a frame body. The metal base has a center portion, a plurality of supporting portions with strip shapes and a plurality of wing portions. Each of the supporting portions has a head end, a middle part and a tail end. The head end connects to a peripheral part of the center portion. The wing portions extend outwardly at an angle from the middle parts of the supporting portions, respectively. The frame body has a frame wall. The wing portions are partially at least covered by the frame body. Another fan frame is also disclosed. 1. A fan frame , comprising:a metal base having a center portion, a plurality of supporting portions with strip shapes and a plurality of wing portions, wherein each of the supporting portions has a head end, a middle part and a tail end, the head end connects to a peripheral part of the center portion, the wing portions extend outwardly at an angle from the middle parts of the supporting portions, respectively; anda frame body having a frame wall, wherein the wing portions are at least partially covered by the frame body.2. The fan frame according to claim 1 , wherein the angle is an obtuse angle.3. The fan frame according to claim 1 , wherein frame body is made of plastic.4. The fan frame according to claim 1 , wherein the wing portions are flat structures.5. The fan frame according to claim 1 , wherein the frame wall is made of metal claim 1 , the supporting portions radially extend from the frame wall inwardly and connect to the center portion.6. The fan frame according to claim 1 , wherein the center portion claim 1 , the supporting portions and the wing portions are integrally formed as one piece.7. The fan frame according to claim 1 , further comprising:a peripheral connection portion, wherein the peripheral connection portion is an annular structure, the tail ends of the supporting portions connect to the peripheral connection portion.8. The fan frame according to claim 7 , wherein the center portion ...

Method of forming gas turbine engine components

A method of forming a gas turbine engine component according to an example of the present disclosure includes, among other things, attaching a cover skin to an airfoil body, the airfoil body and the cover skin cooperating to establish pressure and suction sides of an airfoil, positioning the airfoil between first and second dies of a deforming station, heating the airfoil body to a first predefined temperature threshold between the first and second dies, and moving the first die relative to the second die to hold the airfoil between the first and second dies subsequent to the heating step, and then deforming the airfoil between the first and second dies.

Hot corrosion-resistant coatings for gas turbine components

A gas turbine component for use in a gas turbine engine includes a substrate a ceramic-based thermal barrier coating (TBC), and a diffusion chromide bond coat between the base material and the TBC. A thermally grown oxide (TGO) layer can be formed on the bond coat prior to application of the TBC. The TBC and the TGO include a common metal oxide. The oxide can be sacrificially in use and soluble in a molten sulfate salt, make the coating system particularly suitable for use in a marine environment.

Fluid flow machine with high performance

A fluid flow machine, includes a main flow path formed by a hub and a housing, an arrangement of rotating blades in the main flow path to supply energy to the fluid, forming a rotor assembly group, an arrangement of resting blades arranged adjacent to the rotor in the main flow path, forming a stator assembly group, wherein respectively one rotor assembly group and one stator assembly group adjacent thereto form a stage of the fluid flow machine. In at least one stage, the averaged blade profile angles β RI , β RE of the rotor assembly group, the averaged blade profile angles β SI , β SE of the stator assembly group, as well as the course of the cross-sectional areas of the main flow path in a downstream direction are selected such that they fulfill a certain relationship to increase the realizable level of efficient work to be supplied.

FAN AND FRAME STRUCTURE THEREOF

A fan includes an impeller, a drive device, and a frame structure. The driving device drives the impeller to rotate, and the frame structure accommodates the driving device. The frame structure includes a frame body, a first frame, a second frame, a base and a plurality of supports. The frame body has first engaging portions and second engaging portions, which are disposed annularly at the upper end and the lower end of the frame body, respectively. The first frame is mounted on the frame body and disposed corresponding to the first engaging portions. The second frame is mounted on the frame body and disposed corresponding to the second engaging portions. The base is disposed within the frame body and located adjacent to the second frame. The supports are disposed around the base. The material of the frame body is different from that of the first and second frames. 1. A frame structure , comprising:a frame body having a plurality of first engaging portions and a plurality of second engaging portions, wherein the first engaging portions are disposed annularly at an upper end of the frame body, and the second engaging portions are disposed annularly at a lower end of the frame body;a first frame mounted on the upper end of the frame body and disposed corresponding to the first engaging portions, wherein a part of the first frame is embedded in the first engaging portions;a second frame mounted on the lower end of the frame body and disposed corresponding to the second engaging portions;a base disposed within the frame body and located adjacent to the second frame; anda plurality of supports disposed around a periphery of the base, wherein two ends of each of the supports are connected to the base and an inner wall of the frame body, respectively;wherein a material of the frame body is different from that of the first and second frames.2. The frame structure of claim 1 , wherein the frame body is made of metal or alloy claim 1 , and the first and second frames claim 1 ...

INNER SHROUD AND ORIENTABLE VANE OF AN AXIAL TURBOMACHINE COMPRESSOR

An assembly for the compressor stator of a turbomachine. The assembly comprises: a shroud, in various instances an inner shroud, that is axially divided into two parts; a pocket formed in the shroud; a bearing located in the pocket; and an orientable vane pivotably mounted in the bearing about a pivot axis. The shroud comprises an axial interface separating the parts that is axially offset from the pivot axis of the orientable vane. The invention also provides a process for assembling the assembly. 1. A stator assembly for an axial turbomachine , said stator assembly comprising:a shroud that is axially divided into two parts by an axial interface separating the two parts;a pocket formed in the shroud;a bearing located in the pocket; andan orientable vane pivotably mounted in the bearing about a pivot axis that is axially remote from the axial interface.2. The stator assembly according to claim 1 , wherein the bearing provides a seal between the orientable vane and the shroud claim 1 , the bearing wholly filling the pocket.3. The stator assembly according to claim 1 , wherein the separating interface axially delimits the bearing claim 1 , one of the two parts comprising a flat circular surface in contact with the bearing.4. The stator assembly according to claim 1 , wherein it comprises a one-piece outer shroud on which the orientable vane is mounted.5. The stator assembly according to claim 1 , wherein the bearing is longer axially than wide in circumference; and its width is greater than its radial thickness.6. The stator assembly according to claim 1 , wherein the pocket comprises a sealed base that is in contact with the bearing.7. The stator assembly according to claim 1 , wherein the bearing has two parallel lateral faces claim 1 , the parallel lateral faces extending over at least a half of the axial length of the bearing.8. The stator assembly according to claim 1 , wherein the pocket is wholly formed in one of the two parts.9. The stator assembly according ...

Columnar air moving devices, systems and methods

An air moving device includes a housing member, an impeller assembly, and a nozzle assembly. The nozzle assembly can include one or more angled vanes set an angle with respect to a central axis of the air moving device. The air moving device can output a column of moving air having an oblong and/or rectangular cross-section. A dispersion pattern of the column of moving air upon the floor of an enclosure in which the air moving device is installed can have an oblong and/or rectangular shape. The dimensions of the dispersion pattern may be varied by moving the air moving device toward or away from the floor, and/or by changing the angles of the stator vanes within the nozzle assembly.

FAN FRAME

A fan frame includes a frame body and a metal base. The frame body has a frame wall. The metal base includes a center portion and a plurality of static blades. The static blades radially extend from the center portion outwardly and connect to the frame wall. Another fan frame includes a base and a frame body. The frame body includes a metal frame wall and a plurality of static blades. The static blades radially extend from the metal frame wall inwardly and connect to the base. Another fan frame includes a frame body, a base and a plurality of static blades. The frame body has a frame wall. One end of each static blade connects to the base, and the other end of each static blade connects to the frame wall. At least one part of the static blades is made of metal. 1. A fan frame , comprising:a frame body having a frame wall; and a center portion, and', 'a plurality of static blades radially extending from the center portion outwardly and connecting to the frame wall., 'a metal base comprising2. The fan frame according to claim 1 , wherein each of the static blades comprises a supporting portion and a wing portion extending outwardly from the supporting portion.3. The fan frame according to claim 2 , wherein the frame body covers the supporting portions and at least one part of the wing portions.4. The fan frame according to claim 1 , further comprising:a peripheral connection portion, wherein the static blades connect to the frame wall through the peripheral connection portion.5. The fan frame according to claim 4 , wherein the center portion claim 4 , the peripheral connection portion and the static blades are integrally formed as one piece.6. The fan frame according to claim 4 , wherein a height of the frame wall is greater than a height of the peripheral connection portion.7. The fan frame according to claim 6 , wherein a height of the peripheral connection portion is greater than a thickness of the center portion.8. A fan frame claim 6 , comprising:a base; and at ...

Attachment Faces for Clamped Turbine Stator of a Gas Turbine Engine

An airfoil fairing shell for a gas turbine engine includes an airfoil section between an outer vane endwall and an inner vane endwall, at least one of the outer vane endwall and the inner vane endwall including a radial attachment face, a suction side tangential attachment face, a pressure side tangential attachment face, and an axial attachment face.

Gas Turbine Manufacturing Method

In a gas turbine manufacturing method for manufacturing a derivative gas turbine having a different cycle from a reference gas turbine including a reference compressor, a compressor of the derivative gas turbine is designed to add at least one additional stage further on an upstream side than a last stage of the reference compressor and on a downstream side of a bleed slit of a bleed chamber of the reference compressor, the compressor of the derivative gas turbine is manufactured on the basis of the design, and the derivative gas turbine is manufactured. Consequently, it is possible to manufacture a gas turbine that can secure a surge margin of a compressor with respect to fluctuation in the composition of fuel. 1. A gas turbine manufacturing method for manufacturing a derivative gas turbine having a different cycle from a reference gas turbine including a reference compressor , the gas turbine manufacturing method comprising:designing a compressor of the derivative gas turbine to add at least one additional stage further on an upstream side than a last stage of the reference compressor and on a downstream side of a bleed slit of a bleed chamber of the reference compressor; andmanufacturing the compressor of the derivative gas turbine on the basis of the design and manufacturing the derivative gas turbine.2. A gas turbine manufacturing method for manufacturing a derivative gas turbine having a different cycle from a reference gas turbine including a reference compressor , the gas turbine manufacturing method comprising:designing a compressor of the derivative gas turbine to add at least one additional stage on an upstream side of a last stage of the reference compressor and in a region where an inner diameter and an outer diameter of a compressor channel are fixed; andmanufacturing the compressor of the derivative gas turbine on the basis of the design and manufacturing the derivative gas turbine.3. A gas turbine manufacturing method for manufacturing a derivative ...

VENTILATION FAN HAVING AIR BEARING SYSTEM

A ventilation fan includes a shaft assembly, a rotor, a motor housing, and a bearing housing. The shaft assembly defines a first port and a second port. The rotor has a rotor first portion disposed about a first portion of the shaft assembly and a rotor second portion that extends from the rotor first portion. The rotor first portion defines a rotor port. The motor housing is disposed about a second portion of the shaft assembly. The motor housing defines a housing port. The bearing housing is operatively connected to the motor housing. The bearing housing is disposed about a third portion of the shaft assembly. 1. A ventilation fan , comprising:a shaft assembly extending along an axis and disposed within a housing assembly, the shaft assembly including a first shaft having a first shaft first portion, defining a first port, and a first shaft second portion, defining a second port, operatively connected to a second shaft having a second shaft first portion, defining a third port, and a second shaft second portion;a rotor having a rotor first portion disposed about the first shaft first portion and a rotor second portion extending from the rotor first portion, the rotor first portion defining a rotor port;a motor housing disposed about the first shaft second portion, the motor housing having a housing arm, a housing leg extending from the housing arm, a housing extension extending from the housing leg, and a vane platform extending from the housing extension;a first opening being defined between respective ends of the rotor second portion and the vane platform; anda bearing housing operatively connected to the motor housing and disposed about the second shaft second portion, the bearing housing having a bearing arm, a bearing leg extending from the bearing arm, and a bearing extension extending from the bearing leg.2. The ventilation fan of claim 1 , further comprising:a first journal air bearing that rotatably supports the shaft assembly, the first journal air ...

ANGULAR STATOR SECTOR FOR A TURBOMACHINE COMPRESSOR COMPRISING A BRUSH SEAL

The main object of the invention is an angular stator sector () for a turbomachine compressor, comprising an outer shroud and an inner shroud (S) arranged coaxially one inside the other, and at least one vane (P) extending radially between the outer shroud and the inner shroud (S) and connected to the latter by its radial ends, characterised in that the inner shroud (S) comprises at least one brush seal () forming an obstacle to the recirculation of the downstream gases upstream of the inner shroud (S). 2. Compressor according to claim 1 , wherein the inner shroud comprises a brush seal on its downstream axial end.3. Compressor according to claim 1 , wherein the inner shroud comprises a brush seal on its upstream axial end.4. Compressor according to claim 1 , wherein the inner shroud comprises a brush seal on its radial inner end.5. Compressor according to claim 1 , wherein the inner shroud of the angular stator sector comprises a brush seal fastened on the radial inner end of the inner shroud claim 1 , extending substantially in contact with the rotor shroud claim 1 , and wherein one or several lips are arranged on the radially outer portion of the rotor shroud by being axially separated from said brush seal.6. Compressor according to claim 1 , wherein the inner shroud of the angular stator sector comprises a first brush seal fastened on the downstream axial end of the inner shroud and/or a second brush seal fastened on the upstream axial end of the inner shroud claim 1 , the first brush seal and/or the second brush seal extending from the inner shroud respectively to the downstream rotor platform and/or the upstream rotor platform according to respectively a first angle and/or a second angle in relation to the axis of rotation of the turbomachine in such a way as to form a substantially continuous evolution of the inner wall of the aerodynamic stream at the passage between the inner shroud and the downstream rotor platform and/or at the passage between the inner ...

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 wall located in a gas turbine engine , the wall comprising:first and second surfaces; and [ a longitudinal first side; and', 'a longitudinal second side opposite the longitudinal first side, wherein the longitudinal first side comprises a first passage wall and a second passage wall, and wherein the second longitudinal side comprises a third passage wall and a fourth passage wall;, 'a metering section commencing at the inlet, the metering section comprising, 'a diffusing section in communication with the metering section and terminating at the outlet;', the first and second passage walls are straight and intersect to form a vertex; and', 'the third passage wall is curved, extending from the first passage wall; and', 'the fourth passage wall is curved, extending from the second passage wall to intersect the third passage wall., 'wherein, viewing the metering section from the diffusion section], 'a cooling passage extending between an inlet at the first surface and an outlet at the second surface, the cooling passage comprising2. The wall of claim 1 , wherein the longitudinal first side is an upstream side of the cooling passage.3. The wall of claim 1 , wherein the longitudinal first side is a downstream side of the cooling passage.4. The wall of claim 1 , wherein the first passage wall and the second passage wall intersect to form an angle at ...

AIRFOIL WITH SEAL BETWEEN ENDWALL AND AIRFOIL SECTION

An airfoil includes an endwall section and an airfoil section that defines, at least in part, an airfoil profile. The airfoil section includes an internal passage and a rib that sub-divides the internal passage. At least one of the rib or the endwall section includes a seal cavity. A seal is disposed in the seal cavity. 1. An airfoil comprising:an endwall section;an airfoil section defining, at least in part, an airfoil profile, the airfoil section including an internal passage and a rib sub-dividing the internal passage, at least one of the rib or the endwall section including a seal cavity; anda seal disposed in the seal cavity.2. The airfoil as recited in claim 1 , wherein the seal is rigidly attached with the other one of the rib or the endwall section.3. The airfoil as recited in claim 1 , wherein the seal cavity is in the rib claim 1 , and the seal cavity opens radially.4. The airfoil as recited in claim 3 , wherein the seal is rigidly attached with the endwall section.5. The airfoil as recited in claim 1 , wherein the rib is radially elongated.6. The airfoil as recited in claim 5 , wherein the seal cavity is in a radial face of the rib.7. The airfoil as recited in claim 6 , wherein the seal cavity is in an enlarged radial end of the rib.8. The airfoil as recited in claim 1 , wherein the airfoil section is formed of ceramic and the seal is formed of metal.9. A gas turbine engine comprising:a compressor section;a combustor in fluid communication with the compressor section; anda turbine section in fluid communication with the combustor, an airfoil section defining, at least in part, an airfoil profile, the airfoil section including an internal passage and a rib sub-dividing the internal passage, at least one of the rib or the endwall section including a seal cavity, and', 'a seal disposed in the seal cavity., 'at least one of the turbine section or the compressor section including an airfoil having an endwall section,'}10. The gas turbine engine as recited in ...

Variable stator vane structure of axial compressor

In an axial compressor including a row of rotor blades ( 70 ) provided on a rotational shaft ( 20 ) around a central axial line of the rotational shaft at a prescribed pitch, and a row of stator vanes ( 40 ) provided on a casing around the central axial line at a prescribed pitch so as to adjoin the row of rotor blades on an upstream or downstream side thereof, the rotor blades each extend along a radial line (R) emanating from the central axial line, and the stator vanes each extend along a slanted line (I) that is slanted with respect to a corresponding radial line in a circumferential direction.

Turbomachine blade

A blade for a turbomachine, in particular a compressor or turbine stage of a gas turbine, having at least one matrix having a first impact chamber ( 10 ) in which at least one impulse element ( 11 ) is disposed with play, at least one second impact chamber ( 20 ) whose volumetric centroid is offset from a volumetric centroid of the first impact chamber ( 10 ) along a first matrix axis (A) and in which at least one impulse element ( 21 ) is disposed with play, and at least one third impact chamber ( 30 ) whose volumetric centroid is offset from the volumetric centroid of the first impact chamber ( 10 ) along a second matrix axis (B) transversely to the first matrix axis (A) and in which at least one impulse element ( 31 ) is disposed with play, the first matrix axis (A) and an axis of rotation (R) of the turbomachine forming an angle of at least 60° and no more than 120° is provided.

Axial Turbomachine Blade with Platforms Having an Angular Profile

The present application relates to the stator blades of an axial turbomachine intended to be fitted on a ferrule in an annular row of identical blades. Each blade includes a platform with opposed lateral edges and a means of attachment to the ferrule. The contour of the platform has, at each of the lateral edges, an angular profile designed to mate with the contiguous edge of the platform of an adjacent blade. This feature enables the angular position or pitch of the blade to be set. The present application also relates to a stator comprising the blade row. Mechanical clearances J 1, J 2 and J 3 and the means of clamping cause an average angular orientation error. The blades are made with a compensatory angle to offset the angular orientation error.

Switched reluctance motor assembly

Disclosed herein is a switched reluctance motor assembly, including: a rotating shaft forming a rotating center of a motor; a rotor part rotatably coupled on the rotating shaft; a front part mounted over the rotor part to support a first bearing part of the rotating shaft; a diffuser part having a plurality of integrated guide vanes mounted at an outer side thereof while being coupled with the axial upper portion of the front part; and an impeller part coupled with the axial upper portion of the diffuser and coupled with the rotating shaft. According to the preferred embodiments of the present invention, it is possible to reduce the noise generated at the time of driving the motor by manufacturing the guide vanes mounted in the diffuser of the switched reluctance motor so as to be vertically integrated.

HEART ASSIST DEVICE WITH EXPANDABLE IMPELLER PUMP

An impeller includes a hub and a blade supported by the hub. The impeller has a stored configuration in which the blade is compressed so that its distal end moves towards the hub, and a deployed configuration in which the blade extends away from the hub. The impeller may be part of a pump for pumping fluids, such as blood, and may include a cannula having a proximal portion with a fixed diameter, and a distal portion with an expandable diameter. The impeller may reside in the expandable portion of the cannula. The cannula may have a compressed diameter which allows it to be inserted percutaneously into a patient. Once at a desired location, the expandable portion of the cannula may be expanded and the impeller expanded to the deployed configuration. A flexible drive shaft may extend through the cannula for rotationally driving the impeller within the patient. 1. (canceled)2. A catheter pump comprising: a non-expandable portion;', 'an expandable portion extending between a proximal end and a distal end;', 'a nose grommet;', 'a plurality of inlet struts extending between the nose grommet and the expandable portion distal end; and', 'a plurality of discharge struts extending between the non-expandable portion and the expandable portion proximal end; and, 'a cannula comprisingan impeller disposed in the cannula, the impeller sized and shaped to be inserted into a heart of a patient.3. The catheter pump of claim 2 , wherein the plurality of inlet struts are configured to prevent obstructions from entering the cannula.4. The catheter pump of claim 2 , wherein the plurality of discharge struts are configured to act as stationary stator blades and remove swirl velocity from a discharge flow of the impeller.5. The catheter pump of claim 2 , wherein the plurality of discharge struts are flat linear elements.6. The catheter pump of claim 2 , wherein the plurality of discharge struts have an airfoil cross-section.7. The catheter pump of claim 2 , wherein the expandable portion ...

GAS TURBINE ENGINE COMPONENT HAVING ENGINEERED VASCULAR STRUCTURE

A component according to an exemplary aspect of the present disclosure includes, among other things, a wall and a hollow vascular engineered lattice structure formed inside of the wall. The hollow vascular engineered lattice structure has an inlet hole and an outlet hole that communicate fluid into and out of the hollow vascular structure. The hollow vascular engineered lattice structure further has at least one resupply inlet hole between the inlet hole and the outlet hole to communicate additional fluid into the hollow vascular engineered lattice structure. 1. A component , comprising:a wall; anda hollow vascular engineered lattice structure formed inside of said wall, said hollow vascular engineered lattice structure having an inlet hole and an outlet hole that communicate fluid into and out of said hollow vascular structure, said hollow vascular engineered lattice structure further having at least one resupply inlet hole between said inlet hole and said outlet hole to communicate additional fluid into said hollow vascular engineered lattice structure.2. The component as recited in claim 1 , wherein said hollow vascular engineered lattice structure includes hollow passages that extend through one or more nodes and branches of said hollow vascular engineered lattice structure.3. The component as recited in claim 2 , wherein said at least one resupply inlet hole is configured to communicate fluid into one of said nodes.4. The component as recited in claim 2 , wherein said one or more nodes and branches are one of (1) uniformly distributed throughout said hollow vascular engineered lattice structure and (2) non-uniformly distributed throughout said hollow vascular engineered lattice structure.5. The component as recited in claim 2 , wherein said branches are one of (1) orthogonal to said nodes and (2) non-orthogonal to said nodes.6. The component as recited in claim 1 , wherein said at least one resupply inlet hole is inclined at a non-zero angle relative to a ...

ARRANGEMENT WITH A VACUUM PUMP AND METHOD OF COMPENSATING MAGNETIC FIELD PRODUCED BY MAGNETIC INTERFERENCE FIELD OF AT LEAST ONE VACUUM PUMP COMPONENT

An arrangement includes a vacuum pump having a rotor, and a drive unit for driving the rotor and having at least one magnetic interference field-generating component and at least one compensation coil for compensating the magnetic interference field generated by the at least one component. 1. An arrangement , comprising a vacuum pump including a rotor , and drive means for driving the rotor , at least one magnetic interference field-generating component; and at least one compensation coil for compensating the magnetic interference field generated by the at least one component of the vacuum pump.2. An arrangement according to claim 1 , comprising altogether three compensation coils for compensating the magnetic interference field generated by the at least one component of the vacuum pump.3. An arrangement according to claim 2 , wherein magnetic fields of the three compensation coils are oriented in three spatial directions.4. An arrangement according to claim 3 , wherein the magnetic fields of the three compensation coils are oriented at an angle of 90° to each other.5. An arrangement according to claim 1 , wherein the at least one compensation coil is formed of at least one winding of an electrical conductor.6. An arrangement according to claim 1 , further comprising at least one sensor for at least one sizing and measuring the magnetic interference field or phase reference of the at least one component.7. An arrangement according to claim 6 , further comprising a current source for the at least one compensation coil; and a compensation device for controlling current in the at least one compensation coil dependent on the at least one of sizing and measuring the magnetic interference field of the at least one component.8. An arrangement according to claim 7 , wherein the compensation device and the at least one compensation coil are formed as separately controlled components.9. An arrangement according to claim 1 , wherein the drive means comprises an electric motor ...

COMPRESSOR CASING COMPRISING CAVITIES WITH OPTIMISED SETTING

A compressor for a turbine engine, including: a casing, at least one compressor stage including an impeller having stationary blades and an impeller having moving blades positioned downstream of the stationary blade impeller, and cavities in a thickness of the casing that are disposed along a circumference of the casing opposite the moving blades. The cavities, which are elongate and extend along a main direction of orientation, are closed upstream and downstream by upstream and downstream faces respectively, and an upstream border and a downstream border are formed at the intersections between same and the casing. The cavities are offset in relation to the moving blades to overlap the moving blade impeller in the upstream portion, thereby covering the upstream end thereof. The downstream border of the cavities is oriented parallel to the chord at the head of the moving blade. 15-. (canceled)6. A compressor for a turbine engine comprising:a casing;at least one compressor stage including a fixed-vane wheel and a movable-blade wheel positioned downstream of the fixed-vane wheel; andcavities hollowed out, so as not to communicate with one another, in a thickness of the casing from its internal face and disposed parallel to one another on a circumference of the casing opposite a passage path of the movable blades,the cavities having an elongate shape in a principal orientation direction and being closed towards an upstream side and towards a downstream side by an upstream face and by a downstream face respectively, intersections of which with the casing forming an upstream boundary and a downstream boundary respectively,the cavities being offset with respect to the movable blades to project towards the upstream side of the movable-blade wheel while covering the upstream end thereof,wherein the downstream boundary of the cavities is oriented parallel to a chord at a head of the movable blade.7. A compressor according to claim 6 , wherein a direction of orientation of the ...

Full-span forward swept airfoils for gas turbine engines

Rotor of a gas turbine engines having a rotor hub and a plurality of blades extending from the rotor hub, wherein each blade has a full-span forward sweep along a leading edge of the blade that starts at an airfoil root of the blade at the hub and extends to a blade tip, wherein a sweep of a blade is a percentage of a root axial chord length of the respective blade.

FAN EXIT STATOR ASSEMBLY

A fan exit stator assembly of a gas turbine engine may include a radially inward shroud defining a first slot, a radially outward shroud, and a vane. The vane may include a base portion extending through the first slot and a tip portion coupled to the radially outward shroud. The vane may also include a retainer plate disposed radially inward of the radially inward shroud that projects in a circumferential direction from the base portion of the vane. A circumferential plate dimension of the retainer plate in the circumferential direction is greater than a first slot radius in the circumferential direction of the first slot, thus preventing radially outward movement of the vane relative to the radially inward shroud. 1. A fan exit stator assembly of a gas turbine engine , the fan exit stator assembly comprising:a radially inward shroud defining a first slot;a radially outward shroud; anda vane comprising a base portion extending through the first slot and a tip portion coupled to the radially outward shroud, wherein the vane comprises a retainer plate disposed radially inward of the radially inward shroud and projects in a circumferential direction from the base portion of the vane, wherein a circumferential plate dimension of the retainer plate in the circumferential direction is greater than a first slot radius in the circumferential direction of the first slot.2. The fan exit stator assembly of claim 1 , wherein the retainer plate is configured to engage a radially inward surface of the radially inward shroud adjacent the first slot to prevent radially outward movement of the vane relative to the radially inward shroud.3. The fan exit stator assembly of claim 2 , wherein an elastomeric material is disposed between the retainer plate and the radially inward surface of the radially inward shroud.4. The fan exit stator assembly of claim 1 , wherein the retainer plate projects circumferentially from opposing sides of the base portion of the vane claim 1 , wherein the ...

AXIAL TURBO MACHINE

An axial turbo machine includes: a plurality of blades that constitutes a moving blade row or a stationary blade row; an end wall to which the plurality of blades is fixed and which forms a channel of a fluid together with the blades; and at least one concave portion that is locally formed in a region located between the adjacent blades or on the upstream side of a front edge of the blade on a surface of the end wall and that controls a secondary flow of the fluid. 1. An axial turbo machine , comprising:a plurality of blades that constitutes a moving blade row or a stationary blade row;an end wall to which the plurality of blades is fixed and which forms a channel of a fluid together with the blades; andat least one concave portion that is locally formed in a region located between the adjacent blades or on an upstream side of a front edge of the blade on a surface of the end wall and that controls a flowing direction of a secondary flow of the fluid,wherein only the at least one concave portion is locally provided in the region on the surface of the end wall.2. The axial turbo machine according to claim 1 , whereina bending portion that continuously connects a surface of the blade with the surface of the end wall is provided at a corner that has been formed by fixing of the blade to the end wall, anda boundary between the region and another region includes an edge part on the end wall side at the bending portion.3. The axial turbo machine according to claim 1 , whereinthe blades form a throat of the channel, andthe concave portion is located on an upstream side of the throat.4. The axial turbo machine according to claim 2 , whereinthe blades form a throat of the channel, andthe concave portion is located on an upstream side of the throat.5. The axial turbo machine according to claim 1 , whereinthe concave portion has a shape of at least one of a circular shape, an elliptical shape, a fan shape and a rectangular shape.6. The axial turbo machine according to claim 2 ...

GAS TURBINE COMPRESSOR STAGE

The present invention relates to a compressor stage for a gas turbine, in particular, an aircraft engine, having a row of rotating blades () and a row of guide vanes (), which is adjacent downstream, wherein the choke point σ and the aspect ratio AR, which is defined by the quotient between average channel height (h) and average chord length (l), satisfy the condition 134. A compressor stage for a gas turbine aircraft engine , having a row of rotating blades () and a row of guide vanes () , which is adjacent downstream , wherein the choke point σ and the aspect ratio AR , which is defined by the quotient between average channel height (h) and average chord length l , satisfy the condition{'br': None, 'b': 1', '33, 'i': '·AR', 'sub': 'ax', 'σ>−.+5.16.'}2. The compressor stage according to claim 1 , wherein the aspect ratio ARis greater than 0.5 and/or less than 2.5.3. The compressor stage according to claim 1 , wherein the compressor stage is configured and arranged in a gas turbine having at least one compressor.4. The compressor stage according to claim 1 , wherein a total pressure ratio Π of at least one of the compressors amounts to at least 40.5. The compressor stage according to claim 1 , wherein the compressor stage is configured and arranged in an aircraft engine having a gas turbine.6. The compressor stage according to claim 1 , wherein a by-pass ratio BPR of the aircraft engine is at least 10.734. A method for configuring at least one compressor stage of at least one compressor of a gas turbine aircraft engine claim 1 , having a row of rotating blades () and a row of guide vanes () claim 1 , which is adjacent downstream claim 1 , comprising the step of:{'sub': ax', 'ax, 'claim-text': {'br': None, 'i': '·AR', 'sub': 'ax', 'σ>−1.33+5.16.'}, 'aerodynamically configuring the compressor stage so that the choke point σ and the aspect ratio AR, which is defined by the quotient between average channel height (h) and average chord length l, satisfy the condition'}8. ...

Compressor variable vane assembly

A variable vane assembly for a gas turbine engine compressor and method of manufacturing same is described. A plurality of projections on the inner and/or outer shroud protrude into the annular gas path, each projection being at least partially circumferentially disposed between two variable vanes and located adjacent the overhang portion thereof. The projections have an angled planar surface that is substantially parallel to a plane defined by a terminal edge of the overhang portion of the variable vanes when pivoted through a vane pivot arc.

FAN EXIT STATOR ASSEMBLY RETENTION SYSTEM

A retention system for a stator vane assembly including a stator vane having a radially inner end and a radially outer end. Also included is an outer diameter shroud coupled to the radially outer end of the stator vane. Further included is an inner diameter shroud coupled to the radially inner end of the stator vane. Yet further included is a flange of the outer diameter shroud coupled to a frame member with a mechanical fastener. Also included is an inner shroud flange extending radially inwardly and defining a radial recess, the radial recess allowing radial movement of the radially inner end of the stator vane. 1. A retention system for a stator vane assembly comprising:a stator vane having a radially inner end and a radially outer end;an outer diameter shroud coupled to the radially outer end of the stator vane;an inner diameter shroud coupled to the radially inner end of the stator vane;a flange of the outer diameter shroud coupled to a frame member with a mechanical fastener; andan inner shroud flange extending radially inwardly and defining a radial recess, the radial recess allowing radial movement of the radially inner end of the stator vane.2. The retention system of claim 1 , further comprising:a slot defined by the stator vane proximate the radially inner end of the stator vane; anda retainer bar insertable in the slot, the retainer bar located on a radially inner side of the inner diameter shroud to prevent withdrawal of the stator vane from the inner diameter shroud.3. The retention system of claim 2 , wherein the retainer bar has a primarily rectangular cross-section.4. The retention system of claim 1 , wherein the radially inner end of the stator vane is a base portion that includes a width that is greater than a width of the remainder of the stator vane.5. The retention system of claim 1 , further comprising:a slot defined by the base of the stator vane; anda retainer bar insertable in the slot, the retainer bar located on a radially inner side of ...

SEGMENTED STATOR VANE

A vane assembly may include a vane comprised of a composite material. An outer shroud segment may be disposed at radially outer end of the vane. An inner shroud may define an aperture. A radially inner end of the vane may be disposed within the aperture. 1. A vane assembly , comprising:a vane comprising a composite material;an outer shroud segment disposed at radially outer end of the vane; andan inner shroud defining an aperture, wherein a radially inner end of the vane is disposed within the aperture.2. The vane assembly of claim 1 , wherein the outer shroud segment comprises the same composite material as the vane.3. The vane assembly of claim 2 , wherein the outer shroud segment is integrally formed with the vane.4. The vane assembly of claim 1 , wherein the composite material includes at least one of carbon fiber claim 1 , glass fiber claim 1 , aramid fiber or para-aramid fiber impregnated with a resin.5. The vane assembly of claim 1 , wherein the vane is cantilever mounted to the inner shroud.6. The vane assembly of claim 5 , further comprising a flexible material disposed within the aperture and between the vane and the inner shroud.7. The vane assembly of claim 6 , further comprising an abradable material formed on an inner diameter surface of the inner shroud.8. The vane assembly of claim 1 , wherein the inner shroud comprises at least one of a composite material or a metal.9. A gas turbine engine claim 1 , comprising: a vane comprising a composite material,', 'an outer shroud segment disposed at radially outer end of the vane, and', 'an inner shroud defining an aperture, wherein a radially inner end of the vane is disposed within the aperture., 'an engine section comprising at least one of a compressor section or a fan section, the engine section comprising10. The gas turbine engine of claim 9 , wherein the outer shroud segment comprises the same composite material as the vane.11. The gas turbine engine of claim 10 , wherein the outer shroud segment is ...

GEARED TURBOFAN FRONT CENTER BODY THERMAL MANAGEMENT

An assembly for use in a gas turbine engine includes a center body support section including inner and outer annular walls, a plurality of struts, and a heat shield. The outer annular wall is disposed radially outward of the inner annular wall and a plurality of struts connect the inner and outer annular walls. The heat shield is disposed radially inward of and connected to the inner annular wall, such that a cavity is formed between the heat shield and the inner annular wall. The cavity is open to an air flow at a forward face of the center body support section. The inner annular wall and heat shield include first and second forward edges, respectively. The first and second forward edges are aligned axially. 1. An assembly for use with a gas turbine engine , the assembly comprising: an inner annular wall having a first forward edge;', 'an outer annular wall disposed radially outward of the inner annular wall;', 'a plurality of struts connecting the inner and outer annular walls; and', 'a heat shield having a second forward edge, the heat shield being disposed radially inward of and connected to the inner annular wall, such that a cavity is formed between the heat shield and the inner annular wall, wherein the cavity is open to an air flow at a forward face of the center body support section, and wherein the first forward edge of the inner annular wall and the second forward edge of the heat shield are aligned axially., 'a center body support section comprising2. The assembly of claim 1 , wherein the plurality of struts includes hollow struts open to the cavity formed between the inner annular wall and the heat shield.3. The assembly of claim 2 , wherein the hollow struts are open to an outer circumferential surface of the outer annular wall claim 2 , and wherein a passageway extends through each of the hollow struts from the cavity to the outer circumferential surface.4. The assembly of and further comprising:a fan section, in which the center body support section ...

ANTI-ROTATION STATOR VANE ASSEMBLY

A vane assembly may comprise an engine case and an anti-rotation lug coupled to the engine case. The anti-rotation lug may have a forward end and an aft end. A vane cluster may be supported within the engine case. The vane cluster may include an outer shroud with a first slot defined by a forward flange of the outer shroud and with a second slot defined by an aft flange of the outer shroud. The second slot may be configured to receive the aft end of the anti-rotation lug from first direction and wherein the aft flange may be configured to block receipt of the anti-rotation lug from a second direction, which is opposite the first direction. 1. A vane assembly , comprising:an engine case;an anti-rotation lug coupled to the engine case, the anti-rotation lug having a forward end and an aft end; anda vane cluster supported within the engine case, wherein the vane cluster includes an outer shroud with a first slot defined by a forward flange of the outer shroud and with a second slot defined by an aft flange of the outer shroud;wherein the second slot is configured to receive the aft end of the anti-rotation lug from first direction and wherein the aft flange is configured to block receipt of the anti-rotation lug from a second direction, which is opposite the first direction.2. The vane assembly of claim 1 , wherein a width of the second slot is less than a width of the anti-rotation lug.3. The vane assembly of claim 1 , wherein the second slot includes a tapered opening.4. The vane assembly of claim 3 , wherein the aft end of the anti-rotation lug includes a tapered geometry and is configured to fit into the tapered opening of the second slot.5. The vane assembly of claim 1 , wherein an aft edge of the aft flange forms an aft wall of the second slot.6. The vane assembly of claim 1 , wherein the vane cluster is positioned adjacent to a radially inner surface of the engine case.7. The vane assembly of claim 1 , wherein the first direction is directed from the forward ...

AIRFOIL SINGLETS

Composite airfoil singlet, includes airfoil extending from base to tip of airfoil, integrally formed with no more than one outer platform at tip and/or no more than one inner platform at base. Parallel composite plies or woven fibers extend through airfoil and through outer and/or inner platforms. Outer and/or inner curved sections extend between outer and/or inner platforms and airfoil respectively. Assembly includes circular row of the composite airfoil singlets depending radially inwardly from and mounted to an outer shroud or casing. Outer and/or inner fasteners may secure outer and inner platforms to outer shroud or casing and an inner shroud respectively and include shanks extending substantially perpendicularly from outer and inner fastening plates though platform holes in outer and inner platforms and through outer and inner holes in outer shroud or casing and inner shroud respectively. Nuts are screwed on threaded ends of shanks. 1. A composite airfoil singlet comprising:an airfoil extending from a base to a tip of the airfoil, andthe airfoil integrally formed with no more than one outer platform at the tip and/or no more than one inner platform at the base.2. The singlet as claimed in claim 1 , further comprising:the airfoil extending longitudinally or radially from the base to the tip,the outer platform extending transversely or circumferentially from the tip in a right hand or a clockwise direction or a left hand or a counter-clockwise direction, andthe inner platform extending transversely or circumferentially from the base in the right hand or the clockwise direction or the left hand or the counter-clockwise direction.3. The singlet as claimed in claim 2 , further comprising parallel composite plies or woven fibers extending through the airfoil and through the outer and/or the inner platforms.4. The singlet as claimed in claim 3 , further comprising outer and/or inner curved sections extending between the outer and/or inner platforms and the airfoil ...

INTERNALLY DAMPED AIRFOILED COMPONENT AND METHOD

An airfoiled component comprises: a root section, an airfoil section, a damper pocket enclosed within a portion of the airfoil section, and a damper. The airfoil section includes a suction sidewall and a pressure sidewall each extending chordwise between a leading edge and a trailing edge, and extending spanwise between the root section and an airfoil tip. The damper includes a fixed end unified with a damper mounting surface, and a free end extending into the damper pocket from the damper mounting surface. 1. An airfoiled component for a turbine engine , the component comprising:a root section;an airfoil section including a suction sidewall and a pressure sidewall each extending chordwise between a leading edge and a trailing edge, and extending spanwise between the root section and an airfoil tip;a damper pocket enclosed within a portion of the airfoil section, the damper pocket including a damper mounting surface; anda damper including a fixed end unified with the damper mounting surface, and a free end extending into the damper pocket from the damper mounting surface.2. The airfoiled component of claim 1 , wherein the fixed end of the damper is unified with the damper mounting surface using an additive manufacturing process.3. The airfoiled component of claim 1 , wherein the airfoil section comprises:a plurality of stacked component wall build layers forming the suction sidewall and the pressure sidewall.4. The airfoiled component of claim 3 , wherein the damper comprises:a plurality of stacked damper build layers forming the fixed end and the free end.5. The airfoiled component of claim 1 , wherein the airfoil section includes a first airfoil alloy composition claim 1 , and the damper includes a first damper alloy composition.6. The airfoiled component of claim 5 , wherein the first airfoil alloy composition is substantially different from the first damper alloy composition.7. The airfoiled component of claim 5 , wherein the damper also includes a second damper ...