ПАРОВАЯ ТУРБИНА

Изобретение относится к паровой турбине с валом, направленным вдоль оси турбины. Вдоль вала турбины предусмотрено множество ступеней турбины, охватывающих соответственно структуру направляющих лопаток и аксиально расположенную после нее систему рабочих лопаток. Достижимая в ступени турбины средняя степень реакции (r) составляет 5 - 70%. Степень реакции по меньшей мере двух ступеней турбины имеет различное значение. Такое выполнение паровой турбины позволит осуществить ее с хорошим коэффициентом полезного действия. 11 з.п.ф-лы, 3 ил.

ОХЛАЖДАЮЩИЙ ВЕНТИЛЯТОР С ЛОПАСТЬЮ ПОВЫШЕННОЙ ПРОЧНОСТИ

В изобретении раскрыто устройство циркулирования воздуха, содержащее вентилятор большого диаметра, имеющий усиленные лопасти вентилятора с сужающимся профилем аэродинамической поверхности. Лопасти вентилятора могут быть образованы посредством способа экструзионного формования, при этом каждая лопасть вентилятора может содержать усиленную заднюю кромку с выступающим вниз отбортованным краем. Преимущества выполнения каждой лопасти вентилятора с усиленной задней кромкой согласно настоящему изобретению приводят к улучшенным конструктивным характеристикам, так что задняя кромка каждой лопасти вентилятора может быть изготовлена более длинной с тем, чтобы улучшить аэродинамические характеристики. Кроме того, усиленная задняя кромка противодействует отрыву, который может произойти при экструзии, за счет добавления зоны изгиба с отбортованным краем к, по меньшей мере, одному дальнему концу задней кромки. Например, изогнутая часть задней кромки имеет тенденцию с большей легкостью обеспечить жесткое ...

ЛОПАТКА ДЛЯ ТУРБОМАШИНЫ И СПОСОБ ИЗГОТОВЛЕНИЯ УКАЗАННОЙ ЛОПАТКИ

Лопатка турбомашины из композитного материала с матрицей и армированием, содержащим сетку, включает переднюю и заднюю кромки, а также поверхности корытца и спинки, проходящие от передней до задней кромки. Сетка образует трехмерную конструкцию, которая проходит по большей части толщины лопатки между поверхностью корытца и поверхностью спинки и/или по большей части длины лопатки между передней кромкой и задней кромкой. Трехмерная конструкция армирования содержит стержни, соединенные между собой с помощью их соединительных узлов, заделанные в матрицу и имеющие шероховатую поверхность с выступающими зернами. Количество стержней, соединенных с соединительными узлами, является переменным. При изготовлении композитной лопатки получают армирование, содержащее сетку, вставляют армирование в форму для литья под давлением и осуществляют литье под давлением смолы в форму для литья под давлением для заполнения сетки. Затем полимеризуют смолу в форме для литья и извлекают композитную лопатку из формы ...

ЛОПАТКА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ С ПРАВИЛОМ МАКСИМАЛЬНОЙ ТОЛЩИНЫ С БОЛЬШИМ ЗАПАСОМ ПРОЧНОСТИ ПРИ ФЛАТТЕРЕ

Изобретение относится к лопатке ротора газотурбинного двигателя, в которой согласно изобретению соотношение между максимальной толщиной и хордой на 30% высоты лопатки составляет от 20% до 42% соотношения между максимальной толщиной и хордой в корневой части лопатки, соотношение между максимальной толщиной и хордой на 70% высоты лопатки составляет от 10% до 30% соотношения между максимальной толщиной и хордой в корневой части лопатки, соотношение между максимальной толщиной и хордой на 90% высоты лопатки составляет от 10% до 30% соотношения между максимальной толщиной и хордой в корневой части лопатки, соотношение между максимальной толщиной и хордой в концевой части лопатки составляет от 3% до 21% соотношения между максимальной толщиной и хордой в корневой части лопатки. Изобретение направлено на увеличение прочности лопатки. 3 н. и 8 з.п. ф-лы, 6 ил.

СОСТАВНАЯ ЛОПАТКА ОСЕВОЙ ТУРБОМАШИНЫ

Составная лопатка осевой турбомашины состоит из металлического хвостовика, радиального металлического стержня, прочно скрепленного с хвостовиком, бандажной полки, закрепленной на периферийном конце радиального стержня, дефлектора и профильной пустотелой оболочки, свободно установленных снаружи радиального стержня и упирающихся в бандажную полку при радиальном их перемещении вдоль радиального стержня. Профильная оболочка выполнена по длине составной, состоящей, по меньшей мере, из двух не соединенных между собой частей, свободно установленных на радиальном стержне снаружи дефлектора между хвостовиком и бандажной полкой, в которую упирается верхняя часть профильной оболочки. Все части профильной оболочки изготовлены из жаропрочных жаростойких керамических или композиционных материалов. Изобретение позволяет повысить надежность лопатки. 1 з.п. ф-лы, 1 ил.

УЗЕЛ АВИАЦИОННОГО ДВИГАТЕЛЯ И АВИАЦИОННЫЙ ДВИГАТЕЛЬ

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

ПРОФИЛИРОВАННАЯ КОНСТРУКЦИЯ ДЛЯ ЛЕТАТЕЛЬНОГО АППАРАТА ИЛИ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Изобретение относится к профилированной конструкции, удлиненной в направлении, в котором конструкция имеет длину, обдуваемую воздушным потоком, и поперечно к которому конструкция содержит переднюю кромку (164) и/или заднюю кромку, по меньшей мере одна из которых выполнена профилированной и имеет в указанном направлении удлинения зубчатости (28а), образованные следующими друг за другом зубцами (30) и впадинами (32). Вдоль профилированной(ых) передней кромки и/или задней кромки следующие друг за другом зубцы (30) и впадины (32) выполнены только на части указанной длины, обдуваемой воздушным потоком, на которой амплитуда и/или промежуток между зубцами меняются монотонно, за исключением нескольких зубцов, ближайших к каждому концу указанной части, при этом остальная часть (280) указанной длины является гладкой. Изобретение обеспечивает снижение шума, уменьшение аэродинамических потерь, снижение механических напряжений. 3 н. и 11 з.п. ф-лы, 25 ил.

КОРПУС КОМПРЕССОРА (ВАРИАНТЫ) И ЛОПАТКА РАБОЧЕГО КОЛЕСА КОМПРЕССОРА

Изобретение относится к газотурбинным двигателям, а более конкретно к их компрессорам. Корпус компрессора включает в себя выпуклую в осевом направлении внутреннюю поверхность, расположенную вокруг ряда лопаток рабочего колеса с радиальным зазором между ней и лопатками. Кромки при вершине лопаток дополняют собой контур корпуса, обеспечивая снижение потерь на кромках при вершине лопаток и уменьшая блокировку потока. 3 н. и 20 з.п. ф-лы, 6 ил.

ИЗДЕЛИЕ С МЕТАЛЛИЧЕСКИМ ОСНОВНЫМ ТЕЛОМ И СПОСОБ ЕГО ИЗГОТОВЛЕНИЯ

Изобретение относится к изделию с металлическим основным телом, выполненному в виде компоненты газовой турбины, в частности в виде лопатки. Изделие содержит металлическое основное тело по меньшей мере с одним проложенным внутри продольным каналом и множеством ответвляющихся от него поперечных каналов. На основном теле лежит снаружи покровный слой. Он служит в качестве защитного или адгезионного слоя. Обогащенный слой покрывает соответственно стенки продольного канала и соответственно поперечных каналов и части покровного слоя. Снаружи еще может быть предусмотрен керамический теплоизоляционный слой. Изобретение обеспечивает снижение расходов на изготовление, а также получение покрытия на внутренней поверхности поперечных каналов без неконтролируемого сужения поперечного сечения. 3 с. и 16 з.п. ф-лы, 4 ил.

ЛОПАТКА И ДВУГРАННЫЙ УГОЛ ЛОПАТКИ

Изобретение относится к лопатке компрессора, определенной в каждой из точек ее поверхности углом стреловидности и углом ν, содержащей: хвостовик; конец лопатки, причем расстояние между хвостовиком и концом, измеренное вдоль оси, называемой радиальной осью, перпендикулярной оси вращения компрессора, называется радиальной высотой (h); зону между хвостовиком и наконечником, первый участок которой имеет строго положительный угол ν передней кромки, и второй участок имеет строго отрицательный угол ν передней кромки; зону максимального угла ν, находящуюся вдоль указанной радиальной оси между r=0,25h и r=0,7h. Достигается оптимизация аэродинамических характеристик каждого лопаточного венца, определенная, в частности, запасом по помпажу и выходом, а также обеспечение или поддержание во время такой оптимизации механической прочности каждой лопатки. 2 н. и 6 з.п. ф-лы, 11 ил.

СПОСОБ ПРОФИЛИРОВАНИЯ ЗАМЕНЯЮЩЕЙ ЛОПАТКИ В КАЧЕСТВЕ ЗАМЕНЯЮЩЕЙ ЧАСТИ ДЛЯ СТАРОЙ ЛОПАТКИ ДЛЯ ТУРБОМАШИНЫ С ОСЕВЫМ НАПРАВЛЕНИЕМ ПОТОКА

Изобретение относится к способу изготовления заменяющей лопатки для турбомашины. Согласно указанному способу определяют геометрические характеристики контура ступицы и корпуса снабженного старой лопаткой проточного канала, а также осевое положение центра тяжести пера старой лопатки, которая с одной стороны зажата в ступице или в корпусе. Рассчитывают геометрические характеристики пера заменяющей лопатки, таким образом, что перо заменяющей лопатки на своей передней кромке наклоняется в направлении зажатия выше по потоку и обладает прямой стреловидностью. Задают близкую к зажатию область пера заменяющей лопатки, составляющую от 5% до 15% высоты заменяющей лопатки. Смещают участок пера заменяющей лопатки, расположенный за пределами указанной области, выше по потоку, пока осевое положение центра тяжести пера заменяющей лопатки не совпадет с осевым положением центра тяжести пера старой лопатки. В области от расположенной со стороны зажатия линии контура пера заменяющей лопатки до смещенного ...

СТАТОР КОМПРЕССОРА И КОМПРЕССОР ТУРБОМАШИНЫ

Статор предназначен для компрессора турбомашины с набором поворотных лопаток, каждая из которых установлена с возможностью поворота вокруг оси поворота в стенке, образующей поверхность вращения вокруг центральной продольной оси вращения турбомашины. Указанная ось поворота расположена с наклоном под определенным углом (α) к плоскости, перпендикулярной указанной центральной продольной оси. Согласно изобретению указанная стенка механически обработана локально напротив крайней части каждой лопатки для формирования сферического участка или кармана, центр которого расположен на оси поворота. Радиус участка определен таким образом, чтобы обеспечить во всем диапазоне функционирования лопатки минимальный заданный зазор между указанной крайней частью каждой лопатки и находящимся напротив нее сферическим участком, механически обработанным с учетом ограничения глубины механической обработки этого сферического участка. Предложен также компрессор турбомашины, содержащий вышеуказанный статор. Такое выполнение ...

ЛОПАТКА ТУРБИННОГО ДВИГАТЕЛЯ С ВСТРОЕННЫМ ХВОСТОВИКОМ, ИЗГОТОВЛЕННАЯ ИЗ КОМПОЗИЦИОННОГО МАТЕРИАЛА

При изготовлении лопатки турбинного двигателя из композиционного материала изготавливают волокнистую заготовку в виде единого цельного элемента посредством многослойного тканья. Волокнистая заготовка в продольном направлении, соответствующем продольному направлению изготавливаемой лопатки, содержит: первый комплект слоев нитей, для формирования пера, и второй комплект слоев нитей, связанных между собой для формирования хвостовика лопатки. Нити первого комплекта слоев нитей не связаны с нитями второго комплекта слоев нитей, при этом первый комплект слоев нитей пересекается нитями второго комплекта слоев нитей во второй части заготовки. Волокнистой заготовке придают форму для получения цельной волокнистой преформы, имеющей первую часть, образующую преформу пера, и вторую часть, образующую преформу хвостовика лопатки. Затем преформу уплотняют матрицей для получения лопатки из композиционного материала, содержащего волокнистый армирующий каркас, образованный преформой и уплотненный матрицей ...

СПОСОБ ОПТИМИЗАЦИИ ПРОФИЛЯ ЛОПАТКИ ИЗ КОМПОЗИЦИОННОГО МАТЕРИАЛА ДЛЯ ПОДВИЖНОГО КОЛЕСА ТУРБОМАШИНЫ И ЛОПАТКА, ИМЕЮЩАЯ КОМПЕНСИРУЕМЫЙ ВЫСТУП

При изготовлении лопатки из композиционного материала для подвижного колеса турбомашины сначала оптимизируют профиль лопатки посредством смещения центров тяжести различных сечений лопатки в тангенциальном и продольном направлениях. При этом компенсируют момент аэродинамической силы, действующий на нижнее сечение секций пера лопатки и нижнее сечение стойки лопатки, путем вычисления значения смещений для применения к центрам тяжести секций пера лопатки и стойки лопатки в тангенциальном и продольном направлениях для заданного режима вращения диска колеса турбомашины. После чего изготавливают лопатку с указанным оптимизированным профилем. Другое изобретение группы относится к лопатке из композиционного материала для подвижного колеса турбомашины, имеющей оптимизированный профиль и изготовленной указанным выше способом. При этом лопатка содержит перо, образованное множеством сечений лопатки, расположенных друг над другом в направлении, радиальном относительно колеса, замок, предназначенный для ...

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.

ЛОПАТКА ГАЗОТУРБИННОЙ УСТАНОВКИ

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

ЗАЩИТНЫЙ ЭЛЕМЕНТ ПЕРЕДНЕЙ КРОМКИ

Изобретение относится к области лопаток лопаточных машин и, в частности, к защитному элементу (32) передней кромки для лопатки (16) лопаточной машины, который проходит по высоте от нижнего конца (40) до верхнего конца (41), имеет наружную сторону (50), охватывающую переднюю кромку (18), и внутреннюю сторону (51), выполненную с возможностью крепления на корпусе (30) лопатки, и содержит пластинку (34) корытца, пластинку (36) спинки и центральный участок (38), соединяющий пластинку (34) корытца с пластинкой (36) спинки. Этот центральный участок (38) имеет толщину (е1), бóльшую, чем пластинка (34) корытца и пластинка (36) спинки, между наружной стороной (50) и внутренней стороной (51). Толщина (е1) центрального участка (38) является увеличивающейся со стабильным или возрастающим градиентом на первом сегменте (S1), начиная от нижнего конца защитного элемента (32) передней кромки, но имеет убывающий градиент за пределами указанного первого сегмента (S1). Обеспечивается защита лопатки при меньших ...

СПОСОБ ИЗГОТОВЛЕНИЯ МОНОБЛОЧНОГО ЛОПАТОЧНОГО КОЛЕСА РОТОРА И КОЛЕСО

Изобретение относится к области машиностроения, обработке резанием, конструкциям лопаточных колес. Способ заключается в том, что после черновой обработки заготовки колеса для формирования заготовок лопаток их обрабатывают фрезерованием посредством инструмента, совершающего последовательные проходы. Для повышения производительности обработки лопатки обрабатывают путем тангенциального фрезерования, проходы инструмента расположены, по существу, радиально по отношению к колесу. При этом после каждого прохода инструмент поворачивают на угол строки в тангенциальной плоскости по отношению к колесу. Моноблочное лопаточное колесо содержит лопатки со сторонами, образованными строками. Для достижения того же технического результата упомянутые строки расположены, по существу, продольно радиально. 2 н. и 8 з.п. ф-лы, 8 ил.

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Изобретение относится к металлическому изделию подверженному растрескиванию во время работы, например лопаточному элементу газотурбинного двигателя, и способу его изготовления. Изделие выполнено из металлического материала и содержит тело и концевую часть, выполненную заодно с телом. Концевая часть содержит полосу из металлического материала, по существу по всему поперечному сечению концевой части выполненную заодно с телом и проходящую внутрь него. Полоса находится под более высоким сжимающим напряжением, чем тело. Способ изготовления изделия заключается в том, что выбирают глубину полосы в концевой части и затем осуществляют деформацию валками на концевой части. Концевую часть подвергают деформации валками до тех пор, пока не будет достигнуто сжимающее напряжение по всему поперечному сечению концевой части. Изобретение позволяет создать в концевой части изделия полосу с остаточным сжимающим напряжением, препятствующую возникновению трещин. 2 н. и 14 з.п. ф-лы, 1 ил.

УЗЕЛ ПЕРА ЛОПАТКИ И ПОЛКИ ДЛЯ СВЕРХЗВУКОВОГО ПОТОКА

Узел пера лопатки и полки включает перо и полку, на поверхности которой установлено перо, причем поверхность полки имеет углубление между передней кромкой и задней кромкой пера лопатки. Наиболее глубокий сегмент углубления расположен в половине, выше по потоку, пера лопатки. Скелетная кривая является кривой, представляющей собой вариации скелетного угла пера лопатки в секущей плоскости, параллельной поверхности полки, в зависимости от положения вдоль оси колеса. Линеаризованная скелетная кривая является линеаризованным представлением скелетного угла в зависимости от положения вдоль оси колеса и представляет собой прямую линию, соединяющую точки, характеризующие скелетный угол при 10 и при 90% осевого размера пера лопатки от передней кромки, в непосредственной близости от полки скелетная кривая имеет приподнятый участок, лежащий над линеаризованной скелетной кривой. Плоскость, в которой смещение между скелетной кривой и линеаризованной скелетной кривой является максимальным, расположена ...

РЕЦИРКУЛЯРНОЕ УСТРОЙСТВО ДЛЯ ТУРБОКОМПРЕССОРОВ

В изобретении описано рециркуляционное устройство для турбокомпрессоров, имеющее кольцевую камеру, которая расположена в зоне свободных концов лопаток лопаточного венца и радиально граничит с основным проточным каналом, и множество расположенных в этой кольцевой камере и распределенных по ее окружности направляющих элементов, при этом в кольцевой камере в ее передней и/или задней частях предусмотрена возможность прохождения потока в ее окружном направлении, а направляющие элементы соединены, по меньшей мере, с одной стенкой кольцевой камеры, а в остальном консольно или свободно выступают в нее. Обращенные к основному проточному каналу вершины направляющих элементов лежат на и/или вблизи ограничивающей его контур линии и в осевом направлении перекрываются со свободными концами лопаток или в осевом направлении граничат с зоной, в которой расположены свободные концы лопаток. Это позволяет повысить порог помпажа и расширить область устойчивой работы турбокомпрессора без заметного снижения его ...

Рабочее колесо осевого компрессора газотурбинного двигателя

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ЛОПАТКА РОТОРА ВЕНТИЛЯТОРА РЕАКТИВНОГО ДВИГАТЕЛЯ ЛЕТАТЕЛЬНОГО АППАРАТА

Лопатка ротора вентилятора содержит корпус лопатки и оболочку, выполненную более жесткой, чем корпус лопатки. Корпус лопатки изготовлен из композитного материала и включает корыто, спинку, участок передней кромки, участок задней кромки, конец законцовки в продольном направлении и конец основания в продольном направлении, удерживаемый диском вентилятора. Оболочка проходит в продольном направлении корпуса лопатки, покрывает участок передней кромки и закреплена на нем. Оболочка включает в себя концевой участок законцовки, который расположен ближе к концевой стороне законцовки лопатки ротора вентилятора в продольном направлении, чем конец законцовки корпуса лопатки для покрытия конца законцовки корпуса лопатки. Концевой участок законцовки проходит от участка передней кромки к участку задней кромки корпуса лопатки, и заканчивается расположенным ниже по потоку концевым участком концевого участка законцовки. Расположенный ниже по потоку концевой участок концевого участка законцовки оболочки расположен ...

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ

Газотурбинный двигатель содержит корпус с установленным в подшипниках качения ротором, с размещенным на нем осевым компрессором, газовой турбиной, с их направляющими аппаратами, полость сгорания. Профиль направляющих лопаток осевого компрессора и газовой турбины выполнен в сечении, параллельном оси ротора, в форме треугольника, имеющий отношение длины от задней кромки до гребня треугольника к длине основания, равное 0,6÷0,7. Изобретение приводит к повышению КПД двигателя. 6 ил.

СПОСОБ ИЗГОТОВЛЕНИЯ ПОЛОЙ ЛОПАТКИ ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Способ может быть использован при изготовлении полой лопатки для газотурбинного двигателя, содержащий ножку и перо. Изготавливают две наружные детали путем ковки первичного элемента, образующего, по меньшей мере, перьевую часть наружной детали и, по меньшей мере, одного вторичного элемента, предназначенного для образования, по меньшей мере частично, нижней части наружной детали, относящейся к ножке. Соединяют каждый вторичный элемент с первичным элементом для получения наружной детали. Две наружные детали соединяют между собой диффузионной сваркой. Способ позволяет предварительно получить отдельные элементы наружных деталей, составляющих лопатку, каждый из которых имеет относительно равномерную толщину, с последующим их соединением, что обеспечивает снижение расходов по изготовлению полых лопаток за счет экономии металла и отказа от механической обработки. 6 з.п. ф-лы, 9 ил.

ТУРБИННАЯ ЛОПАТКА И ТУРБИННОЕ КОЛЕСО, СОДЕРЖАЩЕЕ ТУРБИННЫЕ ЛОПАТКИ

Турбинное колесо содержит турбинные лопатки (20 ), содержащие профильную часть (102), имеющую аэродинамическую форму. Данная аэродинамическая форма имеет номинальный профиль, соответствующий данным в приведенных в описании таблицах 1-11, в которых расстояния X, Y, Z и R выражены в дюймах, и в таблицах 1'-11', в которых расстояния X, Y, Z и R выражены в сантиметрах. Значения координат X и Y плавно соединены дугой радиуса R с образованием сечения профиля профильной части на каждом расстоянии Z. Сечения профиля на расстояниях Z плавно соединены друг с другом с образованием завершенной аэродинамической формы. Предложенный профиль профильной части лопатки обеспечивает в результате улучшение эффективности и несущей способности аэродинамической части и лопаток в целом. 2 н. и 8 з.п. ф-лы, 3 ил., 22 табл.

КОЛЬЦЕВОЙ ДИФФУЗОР ГАЗОВОЙ ТУРБИНЫ

Газовая турбина содержит диффузор выхлопа, расположенный по направлению потока ниже последней ступени турбины и включающий секцию прохождения струи и стойку. Секция прохождения струи содержит части первой и второй стенок, а стойка имеет переднюю кромку, проходящую между частью первой стенки и частью второй стенки. Передняя кромка стойки имеет первую и вторую части, причем вторая часть передней кромки расположена между первой частью передней кромки и частью второй стенки. Передняя кромка стойки также имеет третью прямолинейную часть, расположенную между первой и второй частями передней кромки. Первая часть передней кромки проходит на 20-40% расстояния между первой передней крайней точкой, в которой передняя кромка встречается с частью первой стенки, и второй передней крайней точкой, в которой передняя кромка встречается с частью второй стенки. Первая часть передней кромки наклонена к выпуску секции относительно направления нормали, перпендикулярного части первой стенки в первой передней ...

ЛОПАТКА КОМПРЕССОРА, ИМЕЮЩАЯ АЭРОДИНАМИЧЕСКУЮ ЧАСТЬ ЗАДАННОГО ПРОФИЛЯ, ЛОПАТКА КОМПРЕССОРА, ИМЕЮЩАЯ АЭРОДИНАМИЧЕСКУЮ ЧАСТЬ СО СТОРОНОЙ ПОНИЖЕННОГО ДАВЛЕНИЯ ЗАДАННОГО ПРОФИЛЯ, И КОМПРЕССОР

Лопатка компрессора имеет аэродинамическую часть заданного профиля по существу в соответствии со значениями X, Y и Z декартовой системы координат, представленными в масштабируемой таблице, выбранной из группы таблиц, состоящей из Таблиц 1-11, в которой значения X, Y и Z декартовой системы координат являются безразмерными значениями, приведенными с возможностью преобразования в размерные расстояния путем умножения значений X, Y и Z декартовой системы координат на некоторое число, при этом 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 полного профиля аэродинамической части. Достигается повышение общего коэффициента полезного действия, снижение вибраций и улучшение нагрузочной способности аэродинамической части. 3 н. и 17 з.п. ф-лы, 11 табл., 3 ил.

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РАБОЧАЯ ЛОПАТКА ПАРОВОЙ ТУРБИНЫ ДЛЯ СЕКЦИИ НИЗКОГО ДАВЛЕНИЯ ПАРОВОЙ ТУРБИНЫ

Рабочая лопатка (20) паровой турбины для секции низкого давления паровой турбины (10). Рабочая лопатка (20) паровой турбины содержит участок (42) аэродинамической поверхности. Секция (44) хвостовика прикреплена к одному концу участка (42) аэродинамической поверхности. Часть (40) в виде ласточкиного хвоста выступает от секции (44) хвостовика, а часть (40) в виде ласточкиного хвоста содержит скошенную часть (40) в виде ласточкиного хвоста с осевой заводкой, имеющую угол скоса, составляющий 19°. Секция (46) венца прикреплена к участку (42) аэродинамической поверхности на конце, противоположном секции (44) хвостовика. Бандажная полка (48) выполнена за одно целое в виде части секции (46) венца. Полка (50) прикреплена к промежуточной секции участка (42) аэродинамической поверхности между его концами. Рабочая лопатка (20) имеет площадь выходного кольцевого сечения, составляющую около 4,43 мили больше. Участок (42) аэродинамической поверхности имеет длину, составляющую около 68,1 см или больше.

ЛОПАТКА РАБОЧЕГО КОЛЕСА РОТОРА КОМПРЕССОРА НИЗКОГО ДАВЛЕНИЯ ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ

... 1. Лопатка снабжённого пазами диска рабочего колеса ротора компрессора низкого давления (КНД) турбореактивного двигателя (ТРД), включающего проточную часть, ограниченную по периферийному контуру корпусом двигателя, характеризующаяся тем, что содержит перо с радиальной осью и хвостовик с продольной осью и предназначена для установки в любой из пазов диска рабочего колеса второй ступени, для чего хвостовик лопатки имеет продольную ось, соосную или параллельную геометрической оси паза и образующую с осью ротора в проекции на условную осевую плоскость, нормальную к радиальной оси пера лопатки, угол установки хвостовика лопатки α, определённый в диапазоне α=(21÷27)°; a перо лопатки выполнено с переменной относительно оси ротора осевой закруткой, нарастающей с радиальным удалением от оси вала ротора с градиентом закрутки пера G, определённым в проекции на упомянутую условную осевую плоскость, определённым в диапазонегде α- проекция угла закрутки хорды корневого сечения пера лопатки относительно ...

ЛОПАТКА РАБОЧЕГО КОЛЕСА РОТОРА КОМПРЕССОРА НИЗКОГО ДАВЛЕНИЯ ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ

... 1. Лопатка снабжённого пазами диска рабочего колеса ротора компрессора низкого давления (КНД) турбореактивного двигателя (ТРД), включающего проточную часть, ограниченную по периферийному контуру корпусом двигателя, характеризующаяся тем, что содержит перо с радиальной осью и хвостовик с продольной осью и предназначена для установки в любой из пазов диска рабочего колеса третьей ступени, для чего хвостовик лопатки имеет продольную ось, соосную или параллельную геометрической оси паза и образующую с осью ротора в проекции на условную осевую плоскость, нормальную к радиальной оси пера лопатки, угол αустановки хвостовика лопатки, определённый в диапазоне α=(20,4÷29,8)°; a перо лопатки выполнено с переменной относительно оси ротора осевой закруткой, нарастающей с радиальным удалением от оси вала ротора с градиентом закрутки пера G, определённым в проекции на упомянутую условную осевую плоскость в диапазонегде α- проекция угла закрутки хорды корневого сечения пера лопатки относительно оси ротора ...

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Изобретение относится к области турбомашиностроения, а именно к конструкции лопатки турбомашины, в частности осевого компрессора газотурбинного двигателя. Лопатка турбомашины выполнена в виде пера с прикрепленными к нему входной и выходной кромками, выполненными из материала с пористой структурой. Соединение осуществляется при помощи пайки, сварки или лазерной наплавки. Толщина входной кромки в месте ее соединения с пером составляет 0,1-0,65, а выходной кромки - 0,1-0,35 от максимальной толщины пера. Входная или/и выходная кромки в продольном направлении могут быть выполнены из отдельных сегментов. Достигается увеличение надежности турбомашины и уменьшение её массы за счет локализации трещины в пористой структуре кромки и толщины места соединения кромки с пером, выбранной из условия сочетания максимальной прочности пера лопатки, места соединения пера и кромки, а также минимальной массы кромки. 2 з.п. ф-лы, 2 ил.

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... 1. Лопатка соплового аппарата с выпуклой и вогнутой сторонами профиля, отличающаяся тем, что профиль представляет собой сектор кругового кольца с несимметричными торцами, один из которых представляет собой прямой тангенциальный срез, образованный пересечением кольца с хордой, причем общая длина хорды меньше внутреннего диаметра кольца, а другой торец выполнен в форме ломаной линии, образованной пересечением кольца, как минимум, с двумя хордами разной длины, при этом большая из хорд имеет длину меньше внутреннего диаметра кольца.2. Лопатка по п.1, отличающаяся тем, что упомянутый торец с ломаной линией выполнен с, по меньшей мере, одной ступенькой.3. Лопатка по п.1, отличающаяся тем, что упомянутый торец с ломаной линией выполнен с, по меньшей мере, одной ступенчатой выемкой.

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

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

ЛОПАТКА РАБОЧЕГО КОЛЕСА КОМПРЕССОРА С ПЕРЕМЕННЫМ ЭЛЛИПТИЧЕСКИМ СОЕДИНЕНИЕМ

Рабочее колесо центробежного компрессора турбомашины имеет по меньшей мере одну лопатку (24), присоединенную к ступице (26) рабочего колеса посредством галтели (27). Лопатка продолжается вдоль хорды, образованной между передней кромкой (28) и задней кромкой лопатки. Галтель (27) имеет форму, которая изменяется непрерывно вдоль хорды. Галтель (27) представляет собой участок (Р) эллипса (Е1). Галтель (27) подвергается изменению посредством непрерывного изменения длин осей эллипса. Достигается оптимальное распределение механической прочности лопатки вдоль хорды. 3 н. и 3 з.п. ф-лы, 5 ил.

АНТИЭРОЗИЙНАЯ РАБОЧАЯ ЛОПАТКА ДЛЯ ПОСЛЕДНИХ СТУПЕНЕЙ ПАРОВЫХ КОНДЕНСАЦИОННЫХ ТУРБИН

Изобретение относится к области энергетического машиностроения и призвано защищать рабочие лопатки последних ступеней конденсационных паровых турбин от влажной паровой эрозии входных кромок указанных рабочих лопаток. Предложена рабочая лопатка для последних ступеней конденсационных паровых турбин, ее выпуклая поверхность в верхней части на длине, подверженной эрозийному износу, выполнена с продольно-ориентированным оребрением по направлению движения пара с высотой ребер h≥5 мм при шаге t≤2 мм и с толщиной ребер δ≤0,5 мм. Технический результат - снижение энергии силового взаимодействия капель влаги с обтекаемыми поверхностями, защита лопатки от эрозийного износа, повышение вибрационной надежности лопатки. 2 ил.

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Изобретение относится к машиностроению и может быть использовано при восстановлении геометрии профиля пера крупногабаритных лопаток вентилятора после летной эксплуатации при обнаружении загибов входной и выходной кромок лопаток. Лопатку устанавливают в базирующий разъемный стержень по оси z лопатки, содержащий контактные поверхности в виде профиля хвостовика, спинки и корыта лопатки, смазывают поверхности входной и выходной кромок профиля пера лопатки, нагревают одновременно базирующий стержень с лопаткой, нагрев лопатки осуществляют до температуры меньше температуры термообработки, причем лопатки и стержень выполнены из одинакового материала. Затем укладывают базирующий стержень с лопаткой в штамп и осуществляют правку лопатки в штампе в изотермических условиях с последующей выдержкой под давлением в течение времени релаксации напряжений в лопатке, после лопатку охлаждают совместно с базирующим стержнем. Выполняют контроль геометрии профиля пера лопатки после правки с последующим осуществлением ...

РАБОЧАЯ ЛОПАТКА ТУРБИНЫ

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Лопатка четвертой ступени ротора компрессора низкого давления турбореактивного двигателя, содержащего рабочее колесо с диском, наделенным пазами, и лопаточным венцом, имеющим решетку профилей пера с фронтальной линией. Лопатка содержит хвостовик и перо с выпукло-вогнутым профилем. Перо лопатки выполнено с углом γ установки профиля, определенным как угол между соединяющей входную и выходную кромки профиля хордой и фронтальной линией решетки лопаточного венца, имеющий в проекции на условную плоскость, перпендикулярную к оси пера, в корневом сечении профиля значение γ=(49,7÷57,7)°. Лопатка выполнена с переменным по высоте пера углом γ установки профиля пера относительно фронтальной линии решетки профилей лопаточного венца, убывающим с радиальным удалением от оси ротора с градиентом G=(152,3÷218,9) [град/м]. Перо лопатки выполнено с входной и выходной кромками, расходящимися к периферийному торцу с градиентом увеличения хорды G=(4,3÷6,2)·10[м/м]. Лопатка выполнена с отношением высоты h входной ...

ЛОПАТОЧНАЯ РЕШЕТКА ТУРБОМАШИНЫ

Изобретение относится к области турбостроения. Лопаточная решетка турбомашины содержит межпрофильные каналы и образующие их лопатки. Профильные сечения межпрофильных каналов в меридиональных сечениях имеют изогнутую среднюю линию, направленную своей передней точкой навстречу потоку рабочей среды и плавно от вершины изогнутости нисходящую к хорде, соединяющей оба конца средней линии, при монотонном уменьшении угла склонения между хордой и касательными к средней линии до нулевого значения при движении от передней точки к вершине изогнутости. Средняя линия профильных сечений расположена по одну сторону относительно хорды и сопряжена с последней в задней точке профиля под острым углом склонения, величина которого на участке от задней точки до вершины изогнутости сначала монотонно возрастает, а затем плавно убывает. Профиль лопатки выполнен симметричным относительно средней линии. Поток рабочей среды в средней области выходного сечения межпрофильного канала выровнен по скорости, так что оказываются ...

Способ нанесения покрытия на лопатки газотурбинного двигателя

Изобретение относится к области нанесения покрытий в вакууме электронно-лучевым способом, конкретно к контролю качества и скорости нанесения покрытий на изделия со сложным профилем, а именно на лопатки газотурбинного двигателя (ГТД). Способ включает обеспечение заданного температурного поля на поверхности лопаток ГТД посредством программатора, для настройки программы которого в вакуумную камеру напыления помещают контрольную кассету с лопатками ГТД, у замка и пера каждой из которых расположены контрольные термопары, и проводят нагрев упомянутых лопаток ГТД в вакуумной камере путем сканирования электронного луча электронно-лучевой пушки по поверхности лопаток ГТД. После подачи в вакуумную камеру напыления материала для нанесения покрытия осуществляют настройку программы программатора. Сканирование, проводимое при нагреве упомянутых лопаток, размещенных в контрольной кассете, осуществляют с изменением режима сканирования электронного луча по поверхности лопаток ГТД до достижения температур ...

УСТРОЙСТВО ДЛЯ СОЕДИНЕНИЯ ТРАЕКТОРНОЙ ПОТОКОВ АКСИАЛЬНО СОЕДИНЕННЫХ ТУРБИН

... 1. Устройство для соединения траекторий потоков аксиально соединенных турбин друг с другом, содержащее первую и вторую турбины (10, 12), аксиально соединенные друг с другом вдоль траектории потока, при этом поток текучей среды, проходящий вдоль первого участка (127) траектории потока вдоль первой турбины, выходит из первой турбины и поступает на второй участок (129) траектории потока вдоль второй турбины, причем турбины имеют соответствующие роторы (134, 136) и соединение (138) между первым и вторым роторами, предназначенное для соединения турбин друг с другом, внутренний кожух (152), проходящий между последней ступенью первой турбины и первой ступенью второй турбины и проходящий вокруг соединения между роторами, и расположенный над соединением между роторами, чтобы обеспечить изоляцию соединения между роторами от траектории потока и обеспечить по существу плавный переход потока текучей среды от первого участка траектории потока в первой турбине ко второму участку траектории потока во второй ...

СПОСОБ ИЗГОТОВЛЕНИЯ ПОЛОЙ ЛОПАТКИ ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

... 1. Способ изготовления полой лопатки (1) для газотурбинного двигателя, содержащей ножку (2) и перо (4), при этом способ содержит этап выполнения заготовки (14) лопатки, содержащей перьевую часть (18) и нижнюю часть (16), при этом этап выполнения заготовки (14) осуществляют таким образом, чтобы она содержала узел (28), по меньшей мере, из двух накладываемых друг на друга деталей (30, 32), соединяемых между собой при помощи диффузионной сварки, отличающийся тем, что этап выполнения указанной заготовки (14) лопатки содержит следующие операции: изготовление указанного узла (28), по меньшей мере, из двух накладываемых друг на друга деталей (30, 32), соединяемых при помощи диффузионной сварки, таким образом, чтобы этот узел образовал только указанную перьевую часть (18) заготовки (14); изготовление дополнительного элемента (34), предназначенного для образования полностью только указанной нижней части (16) заготовки (14); соединение дополнительного элемента (34) с узлом (28) для получения указанной ...

ЛОПАТКА РОТОРА ВЕНТИЛЯТОРА РЕАКТИВНОГО ДВИГАТЕЛЯ ЛЕТАТЕЛЬНОГО АППАРАТА

... 1. Лопатка ротора вентилятора для реактивного двигателя летательного аппарата, выполненная с возможностью установки в установочную канавку, образованную в диске вентилятора, и для отбора внешнего воздуха в канал, образованный кожухом двигателя, посредством вращения, как единое целое, с диском вентилятора, при этом лопатка ротора вентилятора содержит:корпус лопатки, изготовленный из композитного материала из смолы и волокон, и включающий в себя поверхность давления, обращенную к одной стороне в направлении вращения, и поверхность всасывания, обращенную к другой стороне в направлении вращения;установочный участок, предусмотренный на одной концевой стороне корпуса лопатки в ее продольном направлении и установленный в установочную канавку на диске вентилятора; иоболочку, выполненную более жесткой, чем корпус лопатки, и установленную на корпусе лопатки в состоянии, в котором оболочка покрывает участок передней кромки корпуса лопатки и в непосредственной близости к нему, причем участок передней ...

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, ...

ЛОПАТКА ВЕНТИЛЯТОРА ДЛЯ АВИАЦИОННОГО ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ С ИЗОГНУТЫМ ПРОФИЛЕМ В СЕЧЕНИЯХ НОЖКИ

... 1. Лопатка (4) вентилятора для авиационного турбореактивного двигателя, содержащая перо (6), аксиально проходящее между передней кромкой (18) и задней кромкой (20), и содержащая множество сечений пера (S), уложенных радиально между сечением ножки (S) и сечением головки (S), отличающаяся тем, что все сечения пера, заключенные между сечением ножки (S) и сечением пера (S), расположенным на радиальной высоте, соответствующей 30% общей радиальной высоты пера, имеют кривую средней линии, имеющей точку изгиба.2. Лопатка по п. 1, в которой точки изгиба кривых средней линии сечений пера, заключенных между сечением ножки и сечением пера, расположенным на радиальной высоте, соответствующей 30% общей радиальной высоты пера, расположены между 25% и 75% длины хорды лопатки, измеряемой от передней кромки в сторону задней кромки.3. Лопатка по п. 2, в которой точки изгиба кривых средней линии сечений пера, заключенных между сечением ножки и сечением пера, расположенным на радиальной высоте, соответствующей ...

СПОСОБ ИЗГОТОВЛЕНИЯ НАПРАВЛЯЮЩЕЙ ЛОПАТКИ, А ТАКЖЕ НАПРАВЛЯЮЩАЯ ЛОПАТКА

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

Способы изготовления лопаток турбомашин путем электроэрозионной обработки, лопатки и турбомашины

... 1. Лопатка для турбомашины, содержащая аэродинамическую часть, которая проходит в продольном направлении на определенную длину и имеет первый конец и второй конец, причем указанная аэродинамическая часть ограничена по бокам аэродинамической поверхностью и имеет внутреннюю полость, проходящую полностью вдоль всей указанной длины и ограниченную по бокам криволинейной поверхностью.2. Лопатка по п. 1, которая представляет собой статорную лопатку паровой турбины и содержит корневую часть, покрывающую часть и аэродинамическую часть, причем аэродинамическая часть проходит в продольном направлении на определенную длину и имеет первый конец и второй конец, при этом первый конец расположен смежно с корневой частью, а второй конец расположен смежно с покрывающей частью, причем аэродинамическая часть ограничена по бокам аэродинамической поверхностью и имеет внутреннюю полость, проходящую полностью вдоль всей указанной длины и ограниченную по бокам криволинейной поверхностью.3. Лопатка по п. 1, в которой ...

УСТРОЙСТВО, ПРЕПЯТСТВУЮЩЕЕ СРЫВУ ПОТОКА С КОНЦОВ ЛОПАТОК ТУРБОКОМПРЕССОРОВ

... 1. Компрессор, содержащий корпус с внутренней, в целом цилиндрической полостью, образующей основной канал для прохода газа, ротор по меньшей мере с одним рядом лопаток, по меньшей мере один ряд лопаток статора и выполненное в корпусе антисрывное устройство, включающее выполненный в корпусе рециркуляционный канал, в который во время работы компрессора с концов лопаток ротора попадает газ с низким количеством движения, который проходит через рециркуляционный канал и возвращается в основной канал для прохода газа в точке, расположенной на некотором расстоянии от точки, в которой газ попадает в рециркуляционный канал, который выполнен в виде кольцевого кармана с открытой внутренней стороной и множества расположенных в кольцевом кармане криволинейных направляющих лопаток, которые выступают радиально внутрь корпуса в направлении свободных концов, расположенных в открытом отверстии кармана или рядом с ним, и образуют в кольцевом кармане расположенный за лопатками кольцевой входной канал и/или ...

ПРОФИЛЬНАЯ ЛОПАТКА ТУРБОМАШИНЫ И ТУРБИНА С ТАКИМИ ЛОПАТКАМИ

... 1. Вращающаяся лопатка турбореактивного двигателя, подверженная воздействию продольного газового потока и содержащая множество сечений, расположенных вдоль линии (15) центров тяжести сечений лопатки, между основанием (16) и внешней кромкой (18) лопатки, которая ограничена в продольном направлении передней кромкой (20) и задней кромкой (22) и содержит в направлении вдоль радиальной оси (Z-Z) турбореактивного двигателя нижнюю часть (24), среднюю часть (26) и верхнюю часть (28), причем нижняя часть расположена в радиальном направлении между основанием (16) лопатки и нижней границей (30) средней части, а верхняя часть расположена в радиальном направлении между верхней границей (32) средней части и внешней кромкой (18) лопатки, отличающаяся тем, что линия (33) передней кромки в нижней части (24) имеет наклон в продольном направлении, линия передней кромки в средней части (26) имеет наклон назад в продольном направлении, линия передней кромки в верхней части (28) имеет наклон назад в продольном ...

ЛОПАТКА ВЕНТИЛЯТОРА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Изобретение относится к области машиностроения, а именно к лопаткам вентилятора с демпфером для гашения вибраций, в том числе и к длинным пустотелым широкохордным лопаткам вентиляторов. Лопатка вентилятора газотурбинного двигателя содержит основание, металлическую оболочку и несущие силовые элементы, установленные в полости внутри металлической оболочки. Она содержит концевую заглушку, с которой жестко соединены несущие силовые элементы, между основанием и концевой заглушкой установлены промежуточные перегородки, через отверстия в которых проходят несущие силовые элементы. Демпфирующий материал в виде сот, заполненных пустотелыми алюмосиликатными микросферами, размещен в полости между промежуточными перегородками и концевой заглушкой и промежуточной перегородкой. В отверстиях промежуточных перегородок могут быть установлены демпферы. Несущие силовые элементы могут быть смонтированы с предварительным натягом. Несущие силовые элементы могут быть выполнены прямоугольного поперечного сечения ...

ШИРОКОХОРДНАЯ ЛОПАТКА ВЕНТИЛЯТОРА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Широкохордная лопатка вентилятора газотурбинного двигателя состоит из металлической оболочки и несущих силовых элементов, установленных в полости внутри металлической оболочки и демпфирующего материала. Она содержит концевую заглушку, с которой жестко соединены несущие силовые элементы, которые выполнены из сплава титана с алюминием, полученного методом 3D-печати. При этом процентный состав титана от центра к периферии уменьшается от 100% до 0%, процентный состав алюминия возрастает от 0% до 100%а демпфирующий материал размещен в полости между несущими силовыми элементами. Несущие силовые элементы могут быть смонтированы с предварительным натягом. Несущие силовые элементы могут быть выполнены прямоугольного поперечного сечения. Несущие силовые элементы могут быть выполнены круглого поперечного сечения. Несущие силовые элементы могут быть выполнены четырехугольного поперечного сечения, при этом две стенки повторяют внутренний профиль участков оболочки, контактирующих с ними. Полости между ...

ЛОПАТКА ДЛЯ ПРОТОЧНОЙ МАШИНЫ

... 1. Лопатка для использования в проточной машине, подверженная при работе машины воздействию высокоскоростного газового потока (11), имеющая переднюю кромку (1), обращенную в сторону набегающего газового потока (11), и заднюю кромку (2), причем передняя кромка (1) имеет вогнутый участок (7), обеспечивающий в процессе работы образование зоны торможения набегающего газового потока (11) на некотором расстоянии перед внешней поверхностью (12) лопатки, в которой сформирован вогнутый участок (7), по меньшей мере существенно защищая эту поверхность (12) от набегающего газового потока, отличающаяся тем, что она имеет первую и вторую боковые стенки (3, 4), соединенные с образованием стыка (8), размещенного внутри вогнутого участка (7). 2. Лопатка по п.1, отличающаяся тем, что вогнутый участок (7) проходит от корневой части (5) лопатки к ее верхней части (6). 3. Лопатка по п.1 или 2, отличающаяся тем, что вогнутый участок (7) расположен в основном по центру передней кромки по отношению к первой и ...

КОНСТРУКЦИЯ ЛОПАТКИ СТАТОРА И ТУРБОВЕНТИЛЯТОРНЫЙ РЕАКТИВНЫЙ ДВИГАТЕЛЬ С ИСПОЛЬЗОВАНИЕМ ТАКОЙ ЛОПАТКИ

СПОСОБ ВОССТАНОВЛЕНИЯ ФОРМЫ ПОДВИЖНОЙ ЛОПАТКИ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ, ЛОПАТКА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ И ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ, СОДЕРЖАЩИЙ ТАКУЮ ЛОПАТКУ

... 1. Способ восстановления формы изношенной поверхности участка (14) поверхности подвижной лопатки (10) газотурбинного двигателя, изготовленной из титана или из титанового сплава, отличающийся тем, что наносят на упомянутый участок поверхности никель, кобальт или комбинацию двух этих металлов при помощи электролиза. ! 2. Способ по п.1, в соответствии с которым нанесение металла осуществляется при помощи электролиза с использованием тампона. ! 3. Способ по п.1, в соответствии с которым нанесение металла осуществляется при помощи электролиза в ванне. ! 4. Способ по п.1, в соответствии с которым нанесение металла осуществляется при помощи электролиза в микрованне. ! 5. Способ по п.1, в соответствии с которым нанесение металла осуществляется по меньшей мере в два этапа, сочетающих электролиз с использованием тампона и электролиз в ванне или в микрованне. ! 6. Способ по п.5, примененный к лопатке (10) вентилятора. ! 7. Способ по п.6, в соответствии с которым электролитический раствор содержит ...

ЛОПАТКА (ВАРИАНТЫ) И РОТОР ГАЗОТУРБИННОГО ДВИГАТЕЛЯ, И ТРУБОВЕНТИЛЯТОРНЫЙ ДВИГАТЕЛЬ

... 1. Лопасть для турбомашины, имеющей осевую линию, содержащая корневую часть, расположенную под углом к указанной осевой линии, и перо, отходящее от указанной корневой части, причем указанная корневая часть непосредственно примыкает к указанному перу. 2. Лопасть по п.1, отличающаяся тем, что представляет собой лопасть вентилятора турбовентиляторного двигателя. 3. Лопасть по п.1 или 2, отличающаяся тем, что указанная корневая часть имеет плавный профиль. 4. Лопасть по любому из пп.1-3, отличающаяся тем, что указанная корневая часть имеет участок увеличенного размера. 5. Лопасть по п.4, отличающаяся тем, что указанному участку увеличенного размера придан елочный профиль или профиль "ласточкина хвоста". 6. Лопасть по любому из предыдущих пунктов, отличающаяся тем, что дополнительно содержит средство для предотвращения смещения назад лопасти, установленной на диск. 7. Лопасть для турбомашины, имеющей продольную ось, содержащая: корневую часть, ориентированную по направлению указанной оси, и ...

СПОСОБ ОБРАЗОВАНИЯ ДЛИННОЙ ЛОПАТКИ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ, ИМЕЮЩЕЙ ГЛАВНУЮ СТЕНКУ С ТОНКОЙ ЧАСТЬЮ ОКОЛО ВЕРШИНЫ

Лопатка турбомашины, соответствующая турбомашина и способ изготовления турбинной лопатки

ТУРБИНА

Турбина, включающая неподвижный направляющий сопловой аппарат и, по меньшей мере, одно установленное на валу рабочее колесо с выполненными по его периметру профильными пазами, в которых установлены и зафиксированы своими соответственно профилированными основаниями рабочие лопатки с дугообразным поперечным сечением их рабочей части и содержащие нижнюю и верхнюю поперечные пластины, ограничивающие проточную часть турбины в радиальном направлении, отличающаяся тем, что ширина рабочей части лопатки по огибу в ее нижней половине составляет 0,6 и более длины ее рабочей части, для чего профилированные пазы на колесе турбины выполнены дугообразными с переменной площадью их поперечного сечения, при этом выступы и впадины боковых дугообразных стенок пазов образованы дугами разного радиуса, лежащими в параллельных плоскостях, а центры этих дуг находятся на двух параллельных прямых, расположенных в пределах рабочего колеса, ближе к задней торцевой плоскости колеса и тангенциально к оси вала турбины ...

РОТОРНАЯ ЛОПАТКА (ВАРИАНТЫ) И ТУРБОУСТАНОВКА

... 1. Роторная лопатка, имеющая номинальный профиль поверхности, по существу, соответствующий декартовым координатам X, Y и Z, приведенным в Таблице 1, причем координаты Х и Y обозначают выраженные в миллиметрах расстояния, которые при соединении плавными непрерывными дугами образуют сечения профиля аэродинамической части лопатки на каждом расстоянии Z, выраженном в миллиметрах, при этом указанные сечения профиля на расстояниях Z плавно соединены друг с другом с образованием полной формы аэродинамической части лопатки.2. Роторная лопатка по п.1, в которой указанный номинальный профиль поверхности, по существу соответствующий декартовым координатам X, Y и Z, приведенным в таблице 1, охватывает точки в пределах +/-1 мм в направлении по нормали к любой поверхности полной формы аэродинамической части лопатки.3. Роторная лопатка по п.1, в которой ее максимальная толщина (T), начиная от приблизительно 2,21% и до приблизительно 60% высоты роторной лопатки, описывается выражением T=-0,8646·h+1,1087 ...

ПРОФИЛИРОВАННАЯ КОНСТРУКЦИЯ ДЛЯ ЛЕТАТЕЛЬНОГО АППАРАТА ИЛИ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

ЛОПАСТЬ ОСЕВОГО ВЕНТИЛЯТОРА С ВОЛНООБРАЗНЫМ ПРОФИЛЕМ И ЗУБЦАМИ НА ЗАДНЕЙ КРОМКЕ

ТУРБИННАЯ ЛОПАТКА С ОБЕСПЫЛИВАЮЩИМ ОТВЕРСТИЕМ В ОСНОВАНИИ ЛОПАСТИ

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

СПОСОБ ИЗГОТОВЛЕНИЯ МОНОБЛОЧНОГО ЛОПАТОЧНОГО КОЛЕСА РОТОРА И КОЛЕСО

... 1. Способ изготовления моноблочного лопаточного колеса (8) ротора, отличающийся тем, что после грубой обработки заготовки колеса для формирования заготовок лопаток (6) обрабатывают заготовки лопаток путем тангенциального фрезерования посредством инструмента (17), осуществляющего последовательные проходы (22), по существу радиальные по отношению к колесу, при этом после каждого прохода инструмент поворачивают на угол фасок в тангенциальной плоскости по отношению к колесу. 2. Способ изготовления лопаточного колеса по п.1, отличающийся тем, что используют инструмент, содержащий коническую центральную часть (18), закругленную концевую часть (19) и закругленную соединительную часть (20) для соединения с осью вращения (21), сужающуюся к оси вращения (21), при этом центральная часть сужается в направлении к концевой части и к соединительной части, не образуя при этом стыков. 3. Способ изготовления лопаточного колеса по любому из п.1 или 2, отличающийся тем, что проходы осуществляют от обода (7 ...

ЛОПАТКА СТАТОРА ТУРБОМАШИНЫ

Лопатка статора турбомашины, содержащая перо с волнообразным выступом на корыте, о т л и ч а ю щ а я с я тем, что, с целью снижения потерь, волнообразный выступ расположен в прикорневой части пера, а его вершина расположена от входной кромки на расстоянии, составляющем 0,15-0,3 хорды пера.

ЛОПАТКА ОСЕВОЙ ТУРБОМАШИНЫ

Лопатка осевой турбомашины, содержащая перо, выполненное в виде тонкостенных обечаек переменной толщины, соединенных между собой у входной и выходной кромок, и размещенного между обечайками заполнителя, отличающаяся тем, что, с целью повышения надежности, поверхности обечаек в местах соединений очерчены в поперечном сечении дугами окружностей, центры радиусов которых у входной и выходной кромок расположены соответственно со стороны спинки и корыта пера, причем со стороны заполнителя обечайки соединены между собой ступенчато.

СПОСОБ ИЗГОТОВЛЕНИЯ ЛОПАТКИ КОМПРЕССОРА ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ И ЛОПАТКА

... 1. Способ изготовления полой моноблочной лопатки компрессора для газотурбинного двигателя, содержащей ножку (2), продолженную в радиальном направлении пером (4), отличающийся тем, что содержит следующие этапы: изготовление ковкой моноблочной лопатки (1) компрессора, содержащей множество вставок (26); и удаление вставок (26) для образования множества полостей (16, 116, 216) в указанной лопатке (1). 2. Способ по п.1, отличающийся тем, что этап выполнения ковкой лопатки (1), осуществляют таким образом, чтобы каждая из полостей (16, 116, 216), полученных в результате этапа удаления вставок (26), была расположена практически в радиальном направлении вдоль кривой центральной линии (18, 118, 218). 3. Способ по одному из п.1 или 2, отличающийся тем, что этап выполнения ковкой лопатки (1) осуществляют на основе исходного элемента (24), содержащего множество вставок (26). 4. Способ по п.1, отличающийся тем, что этап выполнения ковкой лопатки (1) осуществляют путем выполнения следующих операций: предварительная ...

ЛОПАТКА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ, ВЫПОЛНЕННАЯ ПО ПРАВИЛУ ПРОГИБА ПРОФИЛЯ ПЕРА, С БОЛЬШИМ ЗАПАСОМ ПО ФЛАТТЕРУ

ЛОПАТКА ДВОЙНОЙ КРИВИЗНЫ ДЛЯ НАПРАВЛЯЮЩЕГО АППАРАТА ТУРБОМАШИНЫ

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

Рабочая лопатка осевой турбомашины

Влагоулавливающее устройство

Лопатка рабочего колеса турбомашины

Профиль лопатки турбомашин

Перо лопатки турбомашины

Ступень турбомашины

Использование: втурбомашинах, работающих на низкопотенциальном паре, для установок с открытым контуром, служащих для опреснения воды и вырабатывания электроэнергии за счет излучаемой энергии солнца. Сущность изобретения: ступень снабжена магнитными опорами 9, 10, 11 и 12, выполненными из роторных и статорных частей 13 и 14, которые расположены с зазором между собой и изготовлены из постоянных магнитов, обращенных к зазору одноименными полюсами. Статорные части 14 закреплены в корпусе 1, а роторные части 13 размещены на бандажах 7 и 8 и промежуточных полочках 6 рабочих лопаток 5. Взаимодействие постоянных магнитов стабилизирует траектории вращения рабочих лопаток 5 ступени. 2 ил.

Лопатка осевой турбомашины

Изобретение относится к турбостроению. Цель - снижение концевых потерь. Лопатка осевой турбомашины имеет перо 1 с наклонными втулочным 2 и периферийным 3 сечениями в сторону корыта. Угол наклона втулочного и периферийного сечений равен 7 - 10°, а перо в средней части выполнено радиальным. Эффект наклона пера приводит к тому, что аэродинамическая смесь действует в плоскостях, расположенных под углом к поверхности тока. 1 ил.

Рабочее колесо осевой турбомашины

Паровая турбина

Двухпоточная паровая турбина

Лопатка рабочего колеса турбомашины

SCHAUFFELZUSAMMENSETZUNG FÜR EINE GASTURBINE MIT INTEGRIERTER KÜHLDÜSE

ZENTRIFUGAL-ODER HALBAXIALTURBOMASCHINEN

Minimale Leckströmung zwischen Schaufelspitze und gegenüberliegender Gehäusewand

Turbomaschine

KONSTRUKTIONSEINHEIT FUER STROEMUNGSTECHNISCHE EINRICHTUNGEN ODER MASCHINEN

BAUTEIL FÜR EINE STRÖMUNGSMASCHINE

Bauteil (20) für einen Gaskanal (5) einer Strömungsmaschine (1), wobei das Bauteil (20) eine gaskanalzugewandte Oberfläche (20.1) zum Entlangströmen eines in dem Gaskanal (5) geführten Gases aufweist, wobei die gaskanalzugewandte Oberfläche (20.1) zumindest bereichsweise mit einer Mehrzahl Strukturen (25), nämlich Erhebungen (45) und/oder Vertiefungen (35), geformt ist, und wobei eine jeweilige Struktur (25) in einer Ebene (30) der gaskanalzugewandten Oberfläche (20.1) genommen eine mittlere Weite D hat und senkrecht zu der Ebene (30) eine Erstreckung T hat, wobei die senkrechte Erstreckung T höchstens das 0,1-Fache der mittleren Weite D ausmacht, T ≤ 0,1·D.

Verfahren zur Verwindung einer Hohlschaufel fuer axial durchstroemte Turbomaschinen und Hohlschaufel zur Ausfuehrung des Verfahrens

Gepfeilte Turbomaschinenschaufel

VARIABLER TURBINENGEOMETRIE-TURBOLADER MIT INTERNEM BYPASS FÜR DEN ABGASSTROM

Einrichtung zum Herstellen von Turbinen- oder Verdichterschaufeln

Stroemungsmaschine mit ueberwiegend tangentialem Stroemungsverlauf im UEbergangsbereich des Laufrads an dessen Aussendurchmesser und in dem anschliessenden Bereich der zugehoerigen Leitelemente

Radialschaufelrad und Verfahren zum Herstellen eines Radialschaufelrads für ein Laufzeug eines Abgasturboladers

Radialschaufelrad, insbesondere Turbinenrad, für ein Laufzeug eines Abgasturboladers einer Brennkraftmaschine, mit einem einen Grundbereich (12) und einen Nabenbereich (14) aufweisenden Nabenkörper (10), an welchem eine Mehrzahl von Schaufeln (16) angeordnet sind, dadurch gekennzeichnet, dass• ein Verhältnis Veiner axialen Länge lwenigstens einer Schaufel (16) und eines zweifachen Radius Dder wenigstens einen Schaufel (16) zumindest 0,30;• ein Verhältnis Vder axialen Länge lder wenigstens einen Schaufel (16) und einer radialen Länge leiner Schaufelaustrittskante (20) der wenigstens einen Schaufel zumindest 1,40; und• ein Verhältnis Veines Radius Ddes Nabenbereichs (14) und eines Radius Deines radial äußeren Eckpunkts (C) der Schaufelaustrittskante (20) der wenigstens einen Schaufel (16) zumindest 0,30 beträgt.

TURBINENSCHAUFEL

SCHAUFEL FUER STROEMUNGSMASCHINEN

Strömungskörper

Device for generating energy, transporting gases, fluids, light fractions whilst meeting small space or low drive energy requirement has rotor profile formed on imaginary rotation ellipsoid

The device has a rotor (1) formed from a profile track on an imaginary rotation ellipsoid, whereby the position of the axis can be selected according to the application, i.e. with the short axis vertical or horizontal. The axis ratio is optional. The machine can be produced with two rotors in sequence, adjacent to each other or one above the other.

Components containing ceramic-based materials and coatings therefor

Hybrid components containing a ceramic material, in which detailed features of the components are formed of materials other than ceramic materials, yet result in a robust mechanical attachment of the ceramic and non-ceramic portions of the components. The components includes a first subcomponent formed of a ceramic-based material and a second subcomponent formed of a metallic material. The first subcomponent has a nub and the second subcomponent is separately formed and attached to the first subcomponent by casting the metallic material around the nub of the first subcomponent. The second subcomponent is attached to the first subcomponent by a compression fit and encapsulation of the second subcomponent on the nub of the first subcomponent. The nub has a compliant coating system that provides thermal expansion compliance between the metallic material of the second subcomponent and the ceramic-based material of the first subcomponent.

Axial flow turbine

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

Airfoil shape for a compressor blade

An article of manufacture having a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in TABLE 1. X and Y are distances in inches which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z in inches. The profile sections at the Z distances are joined smoothly with one another to form a complete airfoil shape.

Airfoil core shape for a turbomachine component

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

Turbine Blade Platform Undercut

A system and method of extending the useable life of a gas turbine blade is disclosed in which the gas turbine blade includes an undercut configuration designed to relieve mechanical and thermal stress imparted into the pedestal region of the airfoil trailing edge. The embodiments of the present invention include turbine blade configurations having different trailing edge undercut configurations as well as additional cooling supplied to the internal passages of the trailing edge region of the turbine blade.

Blade comprising pre-wired sections

The invention relates to a method for wiring suctions ((i), (i+1)) of a blade ( 1 ) for a blade row of a turbomachine. According to said method, a first central component (formula (II)) of a second section ((i+1)) is selected according to at least one central component ((formula (IV)), x CG(i) , r CG(i) ) of a first section ((i)), especially according to a formula (I) where: (i) is a variable of the first section; (i+1) is a variable of the second section; (formula (III)) is a central component of a section in the peripheral direction, preferably in the angular or radian measure; x CG is a central component of a section in the axial direction; r CG is a central component of a section in the radial direction; β is a graduated angle of a section; and (formula (V)) is a peripheral incline, preferably in the angular or radian measure. Θ CG  ( i + 1 ) = Θ CG  ( i ) + Arctan  [ 2 · ( x CG  ( i + 1 ) - x CG  ( i ) ) ( r CG  ( i + 1 ) + r CG  ( i ) ) · tan  ( β ( i + 1 ) + β ( i ) 2 ) ] ( I )  Θ CG  ( i + 1 ) ( II )  Θ CG ( III )  Θ CG  ( i ) ( IV )  Θ lean ( V )

Flow directing member for gas turbine engine

In a gas turbine engine, a flow directing member includes a platform supported on a rotor, a radially facing endwall, at least one axial surface extending radially inwardly from a junction with the endwall, an airfoil extending radially outwardly from the endwall, and a fluid flow directing feature. The fluid flow directing feature includes a groove extending axially into the axial surface and has radially inner and outer groove ends. The outer groove end defines an axially extending notch in the junction between the axial surface and the endwall and forms an opening in the endwall for directing a cooling fluid to the endwall. The groove further includes a first groove wall extending from the inner groove end to the outer groove end, and a second groove wall opposed from the first groove wall and extending from the inner groove end to the outer groove end.

Turbomachine component having an airfoil core shape

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

Aerofoil assembly

This invention relates to a stator aerofoil for a gas turbine engine, comprising: a body portion; at least one panel which forms at least part of one of either the pressure or section surface of the aerofoil; at least one internal chamber which is bounded by the body portion and panel; and, at least one elastomeric component within the at least one internal chamber, wherein the elastomeric component at least partially defines a cell within the internal chamber, wherein the cell is in fluid communication with the exterior of the aerofoil.

AIRFOIL STRUCTURE

An airfoil structure including a plurality of molded monolithic members that extend in a side-by-side manner along the spanwise direction of the airfoil structure. Each molded monolithic member is realized from continuous material that is molded to form a corrugated portion with top and bottom flange portions extending from the corrugated portion. As a result of the molding operation, the corrugated portion is integrally formed with and structurally joined to the top and bottom flange portions. The top flange portions of the side-by-side arrangement of molded monolithic members define a portion of the top surface of the airfoil structure. The bottom flange portions of the side-by-side arrangement of molded monolithic members define a portion of the bottom surface of the airfoil section. The corrugated portions of the side-by-side arrangement of monolithic members define internal support structures (preferably closed-cell cores) extending along the spanwise direction of the airfoil structure. The side-by-side arrangement of monolithic members structurally integrates portions of the top skin and bottom skin to support structures that extend between such top and bottom surface portions. Such structural integration provides enhanced structural resistance to torsional loading of the airfoil section due to the fact that such loads are resisted by the continuous material that makes up the both top and bottom surfaces as well as the support structures extending therebetween. 1. An apparatus comprising:an airfoil structure including a root, a tip, and a span extending along a spanwise direction between said root and said tip, said span having a top surface opposite a bottom surface, wherein at least part of said span is defined by a plurality of molded monolithic members that extend in a side-by-side manner along the spanwise direction, each molded monolithic member including a corrugated portion with a top flange portion and a bottom flange portion extending from the ...

IMPELLER AND METHOD FOR PRODUCING SAME

An impeller integrally includes a substrate and multiple impeller blades provided upright along a curved shape and circumferentially spaced apart on one face of the substrate. The impeller blades expand in diameter and each include an inner circumferential-side blade portion on the inner circumferential side of the substrate and extending in the rotary shaft core direction and have a tip side that is bent, and an outer circumferential-side blade portion on the outer circumferential side of the substrate and extending in the rotary shaft core direction. In the direction along the rotary shaft core, the length of the inner circumferential-side blade portion from the substrate side to the tip side is longer than the length of the outer circumferential-side blade portion from the substrate side to the tip side, and the blade thickness of the inner circumferential-side blade portion gradually decreases from the substrate side toward the tip side. 1. An impeller integrally comprising:a substrate; and a plurality of impeller blades that are provided upright along a curved shape and spaced apart in a circumferential direction on one face of the substrate,wherein the impeller blades extend in the circumferential direction while expanding in diameter from a rotary shaft core side,the impeller blades are each configured by an inner circumferential-side blade portion arranged so as to be located on an inner circumferential side of the substrate and oriented so as to extend in a direction along a rotary shaft core and have a tip side that is bent, and an outer circumferential-side blade portion arranged so as to be located on an outer circumferential side of the substrate and oriented so as to extend in the direction along the rotary shaft core,with respect to the direction along the rotary shaft core, the length of the inner circumferential-side blade portion from the substrate side to the tip side is longer than the length of the outer circumferential-side blade portion from ...

Process of welding a turbine blade, a process of welding a non-uniform article, and a welded turbine blade

A process of welding an article and a welded turbine blade are disclosed. The process includes fusion welding over a primary symmetry line determined from a center of gravity on a first side of the article or blade and fusion welding over the primary symmetry line determined from the center of gravity on a second side of the article or blade. The fusion treating includes multiple fusion treatments.

TURBINE BLADES AND METHODS OF FORMING MODIFIED TURBINE BLADES AND TURBINE ROTORS

Turbine blades and methods of forming modified turbine blades and turbine rotors for use in an engine are provided. In an embodiment, by way of example only, a turbine blade includes a platform and an airfoil. The platform includes a surface configured to define a portion of a flowpath, and the surface includes an initial contour configured to plastically deform into an intended final contour after an initial exposure of the blade to an operation of the engine. The airfoil extends from the platform. 1. A turbine blade for use in an engine , the turbine blade comprising:a platform including a surface configured to define a portion of a flowpath, the surface including an initial contour configured to plastically deform into an intended final contour after an initial exposure of the blade to an operation of the engine; andan airfoil extending from the platform.2. The turbine blade of claim 1 , wherein the operation of the engine includes a predetermined maximum speed and a predetermined maximum temperature.3. The turbine blade of claim 1 , wherein:the airfoil has an airfoil shape defined by a convex suction side wall, a concave pressure side wall, a leading edge, a trailing edge, a root and a tip, the convex suction side wall, the concave pressure side wall, and the tip each including an interior surface that defines an interior with the root, the trailing edge including a plurality of slots formed thereon.4. The turbine blade of claim 1 , wherein the initial contour is a depression.5. The turbine blade of claim 4 , wherein the depression is located adjacent to the concave pressure side wall.6. The turbine blade of claim 4 , wherein the depression is located adjacent to the convex suction side wall. The present application is a divisional application of U.S. Pat. No. 8,297,935 issued Oct. 30, 2012 (application Ser. No. 12/273,108, filed Nov. 18, 2008), the contents of which are herein incorporated by reference in their entirety.This invention was made with Government ...

Fan blade having internal rib break-edge

A fan blade has a main body extending between a leading edge and a trailing edge. Channels are formed into the main body from at least one open side. A plurality of ribs extend across the main body intermediate the channels. The fan blade has a dovetail, and an airfoil extends radially outwardly from the dovetail. The ribs having a thickness defined as measured from said leading edge toward said trailing edge. The ribs have break-edges at ends of the thickness that extend away from an outer face of the rib.

TURBOMACHINE BLADE OR VANE HAVING COMPLEMENTARY ASYMMETRICAL GEOMETRY

A turbomachine blade is made of composite material. The blade includes a first portion constituting at least an airfoil exhibiting two faces each connecting a leading edge to a trailing edge, the first portion forming a single part with at least one second portion present only on one of the faces of the airfoil. The second portion constitutes a portion of at least one of the following elements: a flowpath delimiting outer portion of an inner platform, an inner portion of an inner platform, a flowpath delimiting inner portion of an outer platform, and an outer portion of an outer platform. Portions of fiber reinforcements corresponding to the first and the second portions of the blade are at least partially mutually imbricated. Yarns of the first portion of the fiber reinforcement penetrate into the second portion of the fiber reinforcement. 1. A method for fabricating a turbomachine blade or vane made of composite material including a fiber reinforcement densified by a matrix , the method including:fabrication by three-dimensional weaving of a one-piece fiber blank,shaping of the fiber blank to obtain a one-piece fiber preform having a first portion constituting at least an airfoil preform, the airfoil exhibiting two faces, and at least one second portion present only on one of the faces of the airfoil preform, said second portion constituting a preform of a portion of at least one of the following elements: a preform of a flowpath delimiting outer portion of an inner platform, a preform of an inner portion of an inner platform, a preform of a flow path delimiting inner portion of an outer platform, and a preform of an outer portion of an outer platform,densification of the preform by a matrix to obtain a blade made composite material having a fiber reinforcement consisting of the preform and densified by the matrix, and forming a single part with at least a portion of one of the following elements: a flowpath delimiting outer portion of an inner platform, an inner ...

Erosion Resistant Helicopter Blade

A rotor blade including a main portion having a main portion leading edge shield; a main portion impact-resistant layer formed on a nose of the leading edge shield; a main portion erosion-resistant layer formed on a top and bottom of the main portion, rearward of the main portion impact-resistant layer; a main portion foil formed on the top and the bottom of the main portion rearward of the main portion erosion-resistant layer; a tip cap having a tip cap leading edge shield; a tip cap impact-resistant layer formed on a surface of the tip cap leading edge shield; and a tip cap foil formed on the top and the bottom of the tip cap rearward of the tip cap erosion-resistant layer. 1. A rotor blade comprising:a main portion having a main portion leading edge shield;a main portion impact-resistant layer formed on a nose of the leading edge shield;a main portion erosion-resistant layer formed on a top and bottom of the main portion, rearward of the main portion impact-resistant layer;a main portion foil formed on the top and the bottom of the main portion rearward of the main portion erosion-resistant layer;a tip cap having a tip cap leading edge shield;a tip cap impact-resistant layer formed on a surface of the tip cap leading edge shield; anda tip cap foil formed on the top and the bottom of the tip cap rearward of the tip cap erosion-resistant layer.2. The rotor blade of wherein:the main portion impact-resistant layer is a tungsten carbide layer.3. The rotor blade of wherein:the main portion erosion-resistant layer is a niobium layer.4. The rotor blade of wherein:the main portion erosion-resistant layer overlaps the main portion impact-resistant layer.5. The rotor blade of wherein:the main portion erosion-resistant layer underlaps the main portion impact-resistant layer.6. The rotor blade of wherein:the main portion foil is a steel foil.7. The rotor blade of wherein:the tip cap impact-resistant layer is a tungsten carbide layer.8. The rotor blade of wherein:the tip cap ...

Turbine of a turbomachine

A turbine of a turbomachine is provided and includes opposing endwalls defining a pathway for a fluid flow and a plurality of interleaved blade stages and nozzle stages arranged axially along the pathway. The plurality of the blade stages includes a last blade stage at a downstream end of the pathway and a next-to-last blade stage upstream from the last blade stage. The plurality of the nozzle stages includes a last nozzle stage between the last blade stage and the next-to-last blade stage and a next-to-last nozzle stage upstream from the next-to-last blade stage. At least one of the next-to-last blade stage and the next-to-last nozzle stage includes aerodynamic elements configured to interact with the fluid flow and to define a throat distribution producing a tip strong pressure profile in the fluid flow.

TURBINE BUCKET FOR USE IN GAS TURBINE ENGINES AND METHODS FOR FABRICATING THE SAME

A turbine bucket for use with a turbine engine. The turbine bucket includes an airfoil that extends between a root end and a tip end. The airfoil includes an outer wall that defines a cavity that extends from the root end to the tip end. The outer wall includes a first ceramic matrix composite (CMC) substrate that extends a first distance from the root end to the tip end. An inner wall is positioned within the cavity. The inner wall includes a second CMC substrate that extends a second distance from the root end towards the tip end that is different than the first distance. 1. A turbine bucket for use with a turbine engine , said turbine bucket comprising: an outer wall defining a cavity extending from said root end to said tip end, said outer wall comprising a first ceramic matrix composite (CMC) substrate extending a first distance from said root end to said tip end; and', 'an inner wall positioned within said cavity, said inner wall comprising a second CMC substrate extending a second distance from said root end towards said tip end that is different than said first distance., 'an airfoil extending between a root end and a tip end, said airfoil comprising2. A turbine bucket in accordance with claim 1 , wherein said first distance is greater than said second distance such that a tip cavity is defined at tip end to facilitate reducing a flow of air across said tip end.3. A turbine bucket in accordance with claim 2 , wherein said inner wall comprises an outer surface positioned at the second distance claim 2 , said outer wall comprises an inner surface claim 2 , said tip cavity defined between said inner surface and said outer surface.4. A turbine bucket in accordance with claim 3 , wherein said outer wall includes a pressure side and an opposite suction side claim 3 , said pressure side and said suction side extending between a leading edge and a trailing edge claim 3 , said outer surface of said inner wall extends continuously between said pressure side and said ...

Fluid Movement System and Method for Determining Impeller Blade Angles for Use Therewith

A fluid movement system that includes an impeller having a blade with a leading edge blade tip angle determined as a function of an increase in mass flow rate due to reinjection of flow from a flow stability device located proximate to the leading edge tip of the blade. In an exemplary method, the leading edge blade tip angle can be determined based on selecting a blade incidence level based on a mass flow gain versus flow coefficient curve. Blade leading edge tip angles determined in accordance with a method of the present invention are typically greater than blade leading edge tip angles determined using traditional methods. The greater blade leading edge tip angles can lead to more robust blades designs.

CROSS FLOW FAN AND AIR CONDITIONER

A cross flow fan and an air conditioner including the cross flow fan, where the cross flow fan includes a fixing member, and a plurality of blades fixed to an upper surface of the fixing member and spaced apart from each other in a circumferential direction. one or more of the blades include a protrusion protruding in a downward direction from a surface of the blade. 1. A cross flow fan , comprising:a fixing member; anda plurality of blades fixed to an upper surface of the fixing member and spaced apart from each other in a circumferential direction,wherein one or more of the blades have a protrusion at one end portion thereof, the protrusion protruding in a downward direction from a surface of the blade.2. The cross flow fan according to claim 1 , wherein the protrusion extends in a length direction of the one or more blades.3. The cross flow fan according to claim 1 , wherein the protrusion protrudes from an outer edge of the blade.4. The cross flow fan according to claim 1 , wherein an outer surface of the protrusion facing outwardly of the cross flow fan is smoothly connected to an outer edge surface of the blade.5. The cross flow fan according to claim 4 , wherein the outer surface of the protrusion is connected to the outer edge surface of the blade in the form of a curved surface.6. The cross flow fan according to claim 1 , wherein an inner surface of the protrusion facing inwardly of the cross flow fan is in a form of a flat surface inclined with respect to a lower surface of the blade at a predetermined angle.7. The cross flow fan according to claim 6 , wherein an angle defined by the inner surface of the protrusion and the lower surface of the cross flow fan is an acute angle.8. The cross flow fan according to claim 1 , wherein the protrusion is in the form of a curved shape protruding in the downward direction from a surface of the blade.9. The cross flow fan according to claim 1 , wherein the blade has a thickness that decreases in a direction from an ...

AEROFOILS

An aerofoil having a leading edge point within a leading edge region and a pressure surface with a profile wherein within the leading edge region the pressure surface profile has a local minimum. The local minimum reduces the loss which may be caused by high negative incidence on to the blade. 1. An aerofoil having a leading edge point within a leading edge region and a pressure surface with a profile wherein within the leading edge region the pressure surface profile has a local minimum.2. An aerofoil according to claim 1 , wherein the leading region extends along a fraction of the pressure surface length from the leading edge point also has a local maximum located further along the pressure surface length than the local minimum.3. An aerofoil according to claim 1 , wherein the leading edge region extends along a fraction of the pressure surface length from the leading edge point claim 1 , the fraction is less than 0.05 of the pressure surface length S.4. An aerofoil according to claim 3 , wherein the fraction is less than 0.02 of the pressure surface length S.5. An aerofoil according to claim 1 , wherein the local minimum is located at a pressure surface fraction of 0.01 of the pressure surface length from the leading edge point.6. An aerofoil according to claim 1 , wherein the peak displacement δp of the local minimum is between 10 and 40% of r claim 1 , where ris the radius of a circular leading edge.7. An aerofoil according to claim 1 , wherein the leading edge region extends along a fraction of the pressure surface length from the leading edge point claim 1 , the fraction is less than 0.05 of the pressure surface length Sand wherein the leading region also has a local maximum located further along the pressure surface length than the local minimum.8. An aerofoil according to claim 7 , wherein the peak displacement δp of the local minimum is between 10 and 40% of r claim 7 , where ris the radius of a circular leading edge.9. An aerofoil according to further ...

Turbine bucket airfoil profile

A turbine bucket is provided including a bucket airfoil having an airfoil shape, the bucket airfoil having a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table 1 wherein the Cartesian coordinate values of X, Y and Z are non-dimensional values from 0% to 100% convertible to dimensional distances in inches by multiplying the Cartesian coordinate values of X, Y and Z by a height of the airfoil in inches, and wherein X and Y are distances in inches which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z, the airfoil profile sections at Z distances being joined smoothly with one another to form a complete airfoil shape.

PROPELLER BLADE HAVING COMPLIANT ADHESIVE AT SPAR INTERFACE

A propeller blade includes a foam core, an adhesive layer formed on the core and a structural layer that covers at least a portion of the adhesive layer and that surrounds at least a portion of the foam core. 1. A propeller blade comprising:a foam core;an adhesive layer formed on the core; anda structural layer that covers at least a portion of the adhesive layer and that surrounds at least a portion of the foam core.2. The propeller blade of claim 1 , further comprising:a layer of fiberglass at least partially disposed between the foam core and the adhesive layer.3. The propeller blade of claim 1 , further comprising:a root disposed within the structural layer at an end of the propeller blade.4. The propeller blade of claim 4 , wherein the root surrounds a portion of the foam core.5. The propeller blade of claim 1 , wherein the foam core is formed of polyurethane foam claim 1 , polyisocyanurate foam claim 1 , or polymethacrylimide foam6. The propeller blade of claim 1 , wherein the adhesive layer is formed of an elastomeric material.7. The propeller blade of claim 1 , wherein the structural layer is formed of a resin-impregnated fiber material.8. The propeller blade of claim 1 , wherein the structural layer is formed of a resin-impregnated braided carbon fiber sheet.9. A method of forming a propeller blade comprising:forming a foam core;coating at least a portion of the form core with an adhesive layer; anddisposing a structural layer over at least a portion of the foam core such that it covers at least a portion of the adhesive layer.10. The method of claim 9 , wherein forming a foam core includes injecting a foam into a form having a fiberglass layer disposed therein.11. The method of claim 9 , further comprising:coupling a root to the foam core;wherein disposing includes disposing the structural layer over at least a portion of the root.12. The method of claim 9 , wherein the foam core is formed of polyurethane foam polyisocyanurate foam and polymethacrylimide ...

PROPELLER BLADE WITH LIGHTWEIGHT INSERT

A propeller blade includes a foam core and a structural layer that surrounds at least a portion of the foam core and includes a face side and a camber side is disclosed. The propeller blade also includes an insert disposed in the foam core in operable contact with the face side and the camber side of the structural layer. 1. A propeller blade comprising:a foam core;a structural layer that surrounds at least a portion of the foam core and includes a face side and a camber side; andan insert disposed in the foam core in operable contact with the face side and the camber side of the structural layer.2. The propeller blade of claim 1 , further comprising:a layer of fiberglass or carbon at least partially disposed between the foam core and the structural layer.3. The propeller blade of claim 1 , wherein the insert includes a face component and a camber component connected by a connecting member.4. The propeller blade of claim 3 , wherein the insert further includes another connecting member that connects the face component and the camber component and is separate from the connecting member.5. The propeller blade of claim 3 , wherein the connecting member is located at ends of one or both of the face component and the camber component.6. The propeller blade of claim 3 , wherein the connecting member is located at a location that is not at an end of one or both of the face component and the camber component.7. The propeller blade of claim 3 , wherein the connecting member is substantially perpendicular to the one or both of the face component and the camber component.8. The propeller blade of claim 1 , further comprising:a root disposed within the structural layer at an end of the propeller blade.9. The propeller blade of claim 8 , wherein a void exists between an end of the root and an end of the foam core.10. The propeller blade of claim 9 , wherein the end of the foam core is closer than an end of the insert to the end of the root.11. The propeller blade of claim 1 , ...

BLADING

A blading for a turbomachine, particularly for a gas turbine, wherein thickened areas and depressions formed and disposed on a lateral wall having a plurality of blades such that at least one depression is disposed on a blade pressure side and at least one thickened area is disposed on a blade suction side for each blade of the plurality of blades. 1. A blading for a turbomachine , in particular a gas turbine , in which thickened areas and depressions are formed on a lateral wall having a plurality of blades , comprising at least one depression disposed on a blade pressure side and at least one thickened area disposed on a blade suction side for each blade of the plurality of blades.2. The blading according to claim 1 , wherein the lateral wall is not contoured axially in front of and/or behind the thickened areas and/or depressions.3. The blading according claim 1 , wherein the depression and/or thickened area extends in the axial direction over at least 10% claim 1 , preferably at least 30% of the axial grid length.4. The blading according to wherein the thickened area disposed on the blade suction side and/or the depression disposed on the blade pressure side is disposed in the half of the blade grid facing the inlet side.5. The blading according to further comprising a second thickened area extending from the blade suction side to the blade pressure side in the half adjacent to the outlet side.6. The blading according to claim 5 , wherein the second thickened area is peripherally symmetrical or peripherally asymmetrical.7. The blading according wherein blade has a rounded root surface in the region of the transition from blade to lateral wall.8. The blading according wherein blade has a rounded root surface in the region of the transition from blade to lateral wall.9. The blading according to wherein the blade has a bent claim 1 , particularly orthogonal transition in the region of the transition from blade to lateral wall.10. The blading according to wherein ...

Water Impeller

The invention is a magnetically driven pump with a floating impeller and driven magnet, and the invention includes an impeller surface having geometric figures acting as the pumping bodies. 1. A rotatable pump impeller for use in a centrifugal pump comprising:an impeller surface;an axial center on said impeller surface, said impeller adapted for coupling to a rotatable driving means for rotation about said axial center;at least three geometric figures on said impeller surface, each geometric figure comprising a perimeter, each said perimeter defining an area of said impeller surface interior to said perimeter, said perimeter being raised above said impeller surface and said area interior to said perimeter, each geometric figure offset from said axial center and from every other geometric figure.2. The impeller of wherein each said geometric figure perimeter is substantially closed.3. The pump impeller of wherein each perimeter comprises a proximal portion closest to said axial center claim 2 , a distal portion furthest from said axial center claim 2 , a trailing portion between said proximal and distal portions clockwise from said proximal portion claim 2 , and a leading portion of said raised perimeter between said proximal and distal portions clockwise from said distal portion; wherein said leading portion has a first curvature and said trailing portion has a second curvature claim 2 , and said first and second curvatures are opposed.4. An impeller according to wherein said geometric figure perimeter have a substantially circular configuration.5. An impeller according to wherein said geometric figure perimeter have a substantially teardrop configuration.6. An impeller according to wherein said geometric figure perimeter have a substantially oval configuration.7. An impeller according to wherein each said raised perimeter monotonically decreases from said proximal to the distal portion.8. The impeller of wherein each of said geometric figures are substantially ...

TURBINE IMPELLER COMPRISING BLADE WITH SQUEALER TIP

A turbine impeller including a rotor; a blade extending from the rotor from a first end of the blade; and a squealer tip provided at a second end opposite to the first end of the blade, wherein at least one perforated portion penetrates through the squealer tip. 1. A turbine impeller comprising:a rotor;a blade extending from the rotor from a first end of the blade; anda squealer tip provided at a second end of the blade opposite to the first end,wherein at least one perforated portion penetrates through the squealer tip.2. The turbine impeller of claim 1 , wherein a first perforated portion of the at least one perforated portion penetrates through a portion of a pressure surface of the blade closer to a leading edge of the blade than to a trailing edge of the blade.3. The turbine impeller of claim 2 , wherein a fluid flows from the outside of the blade into the squealer tip via the first perforated portion of the pressure surface of the blade claim 2 , anda cross-sectional area of the first perforated portion of the pressure surface of the blade decreases in a direction from the outside of the blade to an interior of the squealer tip.4. The turbine impeller of claim 2 , wherein a second perforated portion of the at least one perforated portion penetrates through a portion of an absorbing surface of the blade closer to the trailing edge of the blade than to the leading edge of the blade.5. The turbine impeller of claim 4 , wherein a fluid flows from the outside of the blade into the squealer tip via the first perforated portion of the pressure surface of the blade and flows from the interior of the squealer tip to outside of the blade via the second perforated portion of the absorbing surface of the blade claim 4 , anda cross-sectional area of the first perforated portion of the pressure surface of the blade decreases in a direction from the outside of the blade to an interior of the squealer tip.6. The turbine impeller of claim 5 , whereina cross-sectional area of ...

ROTOR BLADE

A blade () having a fitting () for fastening to a hub and an aerodynamic element () having a main spar () surrounding a bushing () for transmitting centrifugal forces and connected to said fitting (). The blade includes a transmission member () for transmitting torsional forces and co-operating with a torsion box. The fitting () includes an annular recess () defining a first recess () and a second recess () that are separated by a top separator () and a bottom separator () locally obstructing the annular recess (), the blade () including a first ring () and a second ring () reversibly arranged respectively in the first recess () and in the second recess (), each ring () being fastened to the top separator () and to the bottom separator (), a ring () including a fastener pin () passing through said bushing () and said fitting () in co-operation with the other ring (). 1. A blade having a fitting for fastening to a hub and an aerodynamic element having a main spar extending span-wise from an attachment zone towards a distal zone , said attachment zone surrounding a bushing for transmitting centrifugal forces and connected to said fitting , said blade including a transmission member for transmitting torsional forces and provided with a first half-shell and a second half-shell together clamping on said attachment zone , each half-shell having a first portion co-operating with said bushing and a second portion co-operating with a torsion box surrounding said distal zone , said first portion being inserted in said fitting , wherein said fitting includes an annular recess defining a first recess and a second recess that are separated by a top separator and a bottom separator locally obstructing the annular recess , and the blade includes a first ring and a second ring reversibly arranged respectively in the first recess and in the second recess , each ring being fastened to the top separator and to the bottom separator , a ring including a fastener pin passing through said ...

EXHAUST-GAS TURBOCHARGER COMPONENT

An exhaust-gas turbocharger component () having at least one flow-conducting component surface () and having a discontinuity structure () which is formed on the component surface (). The discontinuity structure () has a multiplicity of punctiform depressions () which are arranged separately from one another on at least one part of the component surface (). 110. An exhaust-gas turbocharger component () with{'b': '11', 'at least one flow-conducting component surface () and'}{'b': 12', '11, 'a discontinuity structure () which is formed on the component surface (), wherein'}{'b': 12', '13', '11, 'the discontinuity structure () has a multiplicity of punctiform depressions () which are arranged separately from one another on at least one part of the component surface ().'}21311. The exhaust-gas turbocharger component as claimed in claim 1 , wherein the depressions () are provided in different numbers in a manner adapted to the respective component surface ().31311. The exhaust-gas turbocharger component as claimed in claim 1 , wherein the depressions () are provided in different arrangements in a manner adapted to the respective component surface ().413. The exhaust-gas turbocharger component as claimed in claim 1 , wherein the depressions () have different shapes.513. The exhaust-gas turbocharger component as claimed in claim 4 , wherein the depressions () are round claim 4 , elliptical or polygonal.61311. The exhaust-gas turbocharger component as claimed in claim 1 , wherein the depressions () have different depths in a manner adapted to the respective component surface ().713101. The exhaust-gas turbocharger component as claimed in claim 1 , wherein the depressions () are arranged on all the flow-conducting components () of an exhaust-gas turbocharger ().8. The exhaust-gas turbocharger component as claimed in claim 1 , wherein the depressions are manufactured by casting.912. The exhaust-gas turbocharger component as claimed in claim 1 , wherein the discontinuity ...

TURBOFAN FLOW PATH TRENCHES

An integrally bladed disk includes a rotor disk and circumferentially spaced first and second blades. The rotor disk has a rim the periphery of which forms a flow surface. The first and second blades extend integrally outward from the rim. The rim defines a trench in the flow surface between the first and second blades aft of a leading edge of the rim. The trench extends axially forward and rearward of a leading edge of the first blade. 1. An integrally bladed disk , comprising:a rotor disk having a rim the periphery of which forms a flow surface; andcircumferentially spaced first and second blades extending integrally outward from the rim, wherein the rim defines a trench in the flow surface between the first and second blades aft of a leading edge of the rim, wherein the trench extends axially forward and rearward of a leading edge of the first blade.2. The integrally bladed disk of claim 1 , wherein the trench extends axially to about a trailing edge of the first and second blades.3. The integrally bladed disk of claim 1 , wherein the trench has a depth of between about 0.005 inch (0.127 mm) and 0.060 inch (1.52 mm).4. The integrally bladed disk of claim 1 , wherein a base of each of the first and second blades has a fillet.5. The integrally bladed disk of claim 4 , wherein a deepest point of the trench circumferentially aligns with a location of highest stress concentration at a leading edge of the fillet.6. The integrally bladed disk of claim 4 , wherein the trench extends into at least one fillet.7. The integrally bladed disk of claim 6 , wherein the rim defines two or more axially spaced trenches claim 6 , each of the two or more trenches extends into the at least one fillet at different axially spaced locations.8. The integrally bladed disk of claim 1 , wherein the blades comprise two of a row of blades and a plurality of axially spaced trenches are disposed between each blade in the row.9. A gas turbine engine having a stress relief feature claim 1 , the ...

Propeller blade with reinforcing spars and boxes, and propeller comprising at least one such blade

The invention relates to a propeller blade which comprises a hollow casing () forming an extrados () and an intrados () which extend from a blade shank to a free end in the direction of the span, and a framework which is arranged in the hollow casing () and comprises a box spar (), having a plurality of soles () in surface contact with the hollow casing () so as to provide structural support for the hollow casing (), and at least two cavities () which are spaced apart in the direction of the chord (), wherein the propeller blade further comprises at least two reinforcing spars () which extend between the framework and the hollow casing. 1. A propeller blade comprising:{'b': 1', '5', '4', '2', '3', '25', '12', '20, 'a hollow casing () which has an outer surface forming an extrados () and an intrados (), each of which extends from a leading edge () to a trailing edge () in the direction of the chord () of the blade, and from a blade shank () to a free end () of the blade in the direction of the span of the blade,'}{'b': 1', '8, 'claim-text': [{'b': 9', '9', '10', '10', '1', '1, 'i': a', 'b', 'a', 'b, 'a plurality of soles (, , , ) in surface contact with the hollow casing () so as to provide structural support for the hollow casing (),'}, {'b': 15', '16', '25', '6', '7', '8', '1', '6', '7', '12', '21', '22, 'at least two cavities (, ) spaced apart from each other in the direction of the chord (), wherein the propeller blade further comprises at least two reinforcing spars (, ) which extend between said box spar () and the hollow casing (), and each reinforcing spar (, ) extends into the blade shank (), where it forms a root loop () adapted to receive a fixing pin ().'}], 'a framework which is arranged in the hollow casing () and comprises at least one spar, called the box spar (), having267. The blade as claimed in claim 1 , wherein the reinforcing spars ( claim 1 , ) are solid.3. The blade as claimed in claim 1 , wherein claim 1 , of the box spar and the reinforcing ...

LAMINATED ABUTMENT, A ROTOR PROVIDED WITH SUCH AN ABUTMENT, AND AN AIRCRAFT

A laminated abutment () comprising an inner strength member (), an outer strength member (), and an elastomeric member () comprising a stack of main layers () that are flexible and of secondary layers () that are rigid. The stack comprises at least one first main layer () and at least one second main layer (), each first main layer () presenting, in a lead-lag plane (PTRA), a first thickness (e) that varies between a minimum first thickness (e) on the pitch-variation axis (AXPAS) and a maximum first thickness (e) on the flapping axis (AXBAT), and each second main layer () presenting, in a lead-lag plane (PTRA), a second thickness (e) that varies between a minimum second thickness (e) on the flapping axis (AXBAT) and a maximum second thickness (e) on the pitch-variation axis (AXPAS). 111122212. A laminated abutment comprising an inner strength member and an outer strength member , said abutment further comprising an elastomeric member comprising a stack of main layers and of secondary layers , each main layer comprising a single flexible material and each secondary layer comprising a single rigid material , a strength member being able to perform rotary movements (ROTBAT , ROTTRA , ROTPAS) relative to the other strength member about a flapping axis (AXBAT) , about a lead-lag axis (AXTRA) , and about a pitch-variation axis (AXPAS) , said flapping axis (AXBAT) and said pitch-variation axis (AXPAS) having a geometrical center (CTR) in common , said pitch-variation axis (AXPAS) and said flapping axis (AXBAT) co-operating to define a lead-lag plane (PTRA) , said pitch-variation axis (AXPAS) and said lead-lag axis (AXTRA) co-operating to define a flapping plane (PBAT) , and said lead-lag axis (AXTRA) and said flapping axis (AXBAT) co-operating to define a plane (PPAS) of pitch-angle variation , each secondary layer being bonded to each adjacent main layer; wherein said stack comprises at least one first main layer and at least one second main layer , each first main layer ...

ROTOR BLADE AND AN AIRCRAFT

A rotor blade () provided with an outer covering () and with at least one load take-up spar () associated with at least one pin-receiving bushing (). Each spar () is a compartmentalized spar comprising centrifugal force take-up means () incorporated within a twisting stress take-up casing (), said centrifugal force take-up means () comprising at least two boxes (), each having said pin-receiving bushing () of the spar () passing therethrough, each box () including a closed retention belt () extending in the spanwise direction of the blade, the retention belt () of one box () surrounding the retention belt () of another box (), each retention belt () being provided with unidirectional fibers () wound around said pin-receiving bushing (), said casing () being provided with inclined fibers () presenting an angle relative to said unidirectional fibers (). 1. A rotor blade provided with an outer covering and with at least one load take-up spar associated with at least one pin-receiving bushing , the spar extending spanwise from an attachment zone towards a distal zone , said attachment zone of a spar having said at least one pin-receiving bushing of the spar passing therethrough , wherein each spar is a compartmentalized spar comprising centrifugal force take-up means incorporated within a twisting stress take-up casing , said centrifugal force take-up means comprising at least two boxes , each box having said pin-receiving bushing of the spar passing therethrough , each box forming a compartment of a spar , each box including a closed retention belt extending in the spanwise direction of the blade , the retention belt of one box of a spar surrounding the retention belt of another box of the spar , each retention belt being provided with unidirectional fibers wound around said pin-receiving bushing , said casing being provided with inclined fibers presenting an angle relative to said unidirectional fibers.2. A blade according to claim 1 , wherein said blade extends ...

SCREW TURBINE AND METHOD OF POWER GENERATION

A screw turbine comprising a helical turbine blade mounted for axial rotation, a mount associated with the helical turbine blade and mounting the helical turbine blade for axial rotation, and a generator associated with the helical turbine blade which converts energy imparted to the helical turbine blade to electricity, wherein the diameter of the helical turbine blade is less than the lead of the helical turbine blade and wherein said screw turbine is adapted to permit lateral exchange of fluid in use. 1. A screw turbine comprising:a helical turbine blade mounted for axial rotation;a mount associated with the helical turbine blade and mounting the helical turbine blade for axial rotation; anda generator associated with the helical turbine blade which converts energy imparted to the helical turbine blade to electricity,wherein the diameter of the helical turbine blade is less than the lead of the helical turbine blade and wherein said screw turbine is adapted to permit lateral exchange of fluid in use.21. A screw turbine according to claim , , wherein said helical turbine blade of said screw turbine is unsheathed in use to permit lateral exchange of fluid.3. A screw turbine according to claim 1 , wherein a ratio of diameter and lead of said helical turbine blade is 1:8.4. A screw turbine according to claim 1 , wherein said helical turbine blade has a lead angle of from about 50-75° claim 1 , for example of about 60-75°.5. A screw turbine according to claim 1 , wherein the helical turbine blade is an axleless helix.6. A screw turbine according to claim 1 , wherein said at least one mount includes a coupling provided with bearings which facilitates rotation of the helical turbine blade about its longitudinal axis.7. A screw turbine according to claim 1 , wherein a drive shaft associated with the blade is engaged with gearing that engages the generator to produce electricity claim 1 , the gearing preferably operating at low revolutions (revs) and high torque.8. A ...

RADIALLY COMPRESSIBLE AND EXPANDABLE ROTOR FOR A PUMP HAVING AN IMPELLER BLADE

The invention relates to a radially compressible and expandable rotor for a pump having at least one impeller blade, wherein the impeller blade has an impeller blade body whose material is elastically deformable as well as at least one stiffening strut which is at least partially embedded in the material of the impeller blade body. The struts are designed suitably in size, shape and arrangement and are integrated in suitable hollow spaces of the impeller blade body for stabilizing the impeller blade. Elements with tensile strength can additionally be provided. 115-. (canceled)16. A radially compressible and expandable rotor for a pump comprising at least one impeller blade , wherein the impeller blade has an impeller blade body whose material is elastically deformable as well as at least one stiffening strut which is at least partially embedded in the material of the impeller blade body , and wherein the impeller blade thickness between the pressure side and the suction side amounts to at least 80% more than the thickness of the at least one strut in the same direction.17. The rotor of claim 16 , wherein the impeller blade thickness between the pressure side and the suction side amounts to 100% of the thickness of the at least one strut in the same direction.18. The rotor of claim 17 , wherein the impeller blade thickness between the pressure side and the suction side is more than the largest thickness of a strut of said at least one strut.19. The rotor in accordance with claim 16 , wherein the thickness of the at least one impeller blade is constant in the region in which the at least one strut is positioned claim 16 , including the region disposed between the at least one strut in the direction transversely to the longitudinal direction of the at least one strut claim 16 , when viewed substantially perpendicular to the longitudinal direction of the at least one strut.20. The rotor in accordance with claim 16 , wherein the rotor adopts a first state without any ...

WIND GENERATOR HUB ASSEMBLY WITH HYBRID SAIL BLADES

A high solidity hub assembly with lightweight, high efficiency, hybrid airfoil/sail blades. The hub includes at three to eight, blade mounting surfaces each with a radially extending blade. The blades are identical and wedge shaped and evenly distributed radially around the hub. Gaps are created between each blade. Each blade has an airfoil profile with a curved outer skin layer with a large curved leading edge and a flat trailing edge, similar to an airfoil. The outer skin layer is secured along its leading edge and removeably attached along its trailing edge to a lightweight internal frame. Each blade includes a transversely aligned end cap with optional louvers. The blades are oriented so that their outer skin layers face downwind and rotate in the direction of their leading edges. When wind flows against the inside surface of the skin layers, the blades act as a sail to produce thrust. 1. A wind generator hub assembly , comprising:a. a hub assembly with at least three spatially aligned blade mounting surfaces; and,b. a lightweight hybrid airfoil/sail blade attached to each said mounting surface, said blade including an internal frame with a downward facing outer skin layer, said internal frame and said outer skin layer being configured so that said blade has an airfoil profile with curved leading edge and a substantially straight trailing edge and acts as a sail to capture wind energy, said outer skin layer is secured along its leading edge and removeably attached along its trailing edge so that said leading edge of said outer skin layer remains attached to said internal frame and said trailing edges selectively detaches from said internal frame when exposed to excessively high winds, said blade includes a rigid, transversely aligned end.2. The wind generator hub assembly claim 1 , as recited in claim 1 , further including said end cap include at least one louver that when open allows wind to travel through said end cap.3. The wind generator hub assembly claim 2 ...

SAFETY PROPELLER

A propeller () has a hub () with blades (). A safety member () is provided along at least a portion of the leading edge () of each blade (); and the blades () may incorporate anti-cavitation slots (). 1. A safety propeller , preferably for watercraft , of the type having a hub and a plurality of blades , where each blade has a leading edge extending from a proximal end of the blade adjacent the hub to a distal end spaced from the hub; anda respective safety member is provided over at least 50% of the leading edge of each blade, each safety member being of greater thickness and/or height than the leading edge, and at least a portion of the safety member extends from a thrust or driving face of the blade in the forward direction of rotation of the propeller.2. A propeller as claimed in claim 1 , wherein:for a propeller which rotates in a clockwise direction (when viewed from a trailing end of the hub), the safety member on each thrust or driving face of the blade appears to be directed towards the trailing end of the propeller.3. A propeller as claimed in claim 1 , wherein:the safety member extends over more than 50% of the length of the leading edge of the blade and is of a greater height so that the safety member will strike any obstruction before the balance of the leading edge not provided with the safety member.4. A propeller as claimed in claim 3 , wherein:the safety member is of a constant height over the leading edge, or of relatively increasing height towards the distal end of the blade.5. A propeller as claimed in claim 3 , wherein:the safety member located with a central axis substantially parallel to, or aligned with, the leading edge, or with the central axis increasingly leading the leading edge of the blade in the direction from the proximal end to the distal end of the blade.6. A propeller as claimed in claim 5 , wherein:the safety member is smoothly contoured into the adjacent portion of the blade.7. A propeller as claimed in claim 1 , wherein:the ...

Blade body and rotary machine

The blade body of the present invention is provided with a main body which has a dorsal face and a ventral face and also provided with a trailing edge portion which connects the dorsal face to the ventral face with a continuous curved face. The curved face of the trailing edge portion is gradually decreased in curvature radius from one of the dorsal face and the ventral face toward the rear end portion which is positioned most downstream in a direction at which a fluid flows, decreased to the greatest extent in curvature radius at the rear end portion, thereafter, gradually increased in curvature radius from the rear end portion toward the other of the dorsal face and the ventral face and arrives at the other of the dorsal face and the ventral face.

COMPOUND AIRFOIL

An airfoil is provided that has an arrangement that improves the lift of an airfoil and that include surface features that change the performance of the airfoil. Protrusions are provided on the top surface of the airfoil such that channels are formed between adjacent protrusions that affect the flow of air there through. In an additional respect, indentations can be provided on the bottom surface of the airfoil that affect the flow of air there through. 1. An airfoil for a vehicle , comprising:a top surface extending between a leading edge and a trailing edge on a top side of the airfoil and defining a chord length therebetween;a bottom surface extending between the leading edge and the trailing edge on a bottom side of the airfoil;a plurality of protrusions on the top surface of the air foil, wherein two adjacent protrusions define a channel therebetween that extends in the direction of the chord length;wherein each channel has a leading portion, a middle portion, and a trailing portion, the channel being sized and shaped such that the leading portion and the trailing portion are wider than the middle portion.2. An airfoil of claim 1 , wherein each channel extends along a majority of the top surface in the direction of the chord length.3. An airfoil of claim 1 , wherein the length of the leading portion of each channel is shorter than the middle portion in the direction of the chord length.4. An airfoil of claim 3 , wherein the length of the trailing portion of each channel is longer than the middle portion in the direction of the chord length.5. An airfoil of claim 1 , wherein each channel extends along less than half of the top surface in the direction of the chord length.6. An airfoil of claim 1 , wherein the walls of adjacent protrusions converge along a curved trajectory to define the leading portion of each channel.7. An airfoil of claim 1 , wherein the walls of adjacent protrusions diverge along a generally linear trajectory to define the trailing portion of ...

MOLTEN METAL IMPELLER

According to another embodiment, a molten metal impeller comprised of a generally cylindrical body including a plurality of passages extending from a top surface to a side wall and a cap member secured to a top surface of the body. The cap member is shaped cooperatively to overlay the cylindrical body. The cap member includes a first side seated on the top surface and a second opposed side. The top surface includes one of a notch or a protrusion and the first side of the cap member includes one of a notch or a protrusion oriented to mate with the notch or protrusion of the body. 1. A molten metal impeller comprised of a generally cylindrical body including a plurality of passages extending from a top surface to a side wall and a cap member secured to a top surface of the body , the cap member shaped cooperatively to overlay the body , the cap member including a first side seated on the top surface and a second opposed side , the top surface including one of a recess or protrusion and the first side of the cap member including one of a recess or protrusion oriented to cooperatively mate with the recess or protrusion of the body.2. The impeller of wherein the cap member first side includes a plurality of recesses or protrusions oriented to mate with a plurality of recesses or protrusions in the top surface of the body.3. The impeller of wherein the body is comprised of graphite.4. The impeller of wherein the cap member is comprised of ceramic.5. The impeller of wherein the body comprises a cylindrical base including a centrally disposed hub claim 1 , and a plurality of vanes disposed on the base and extending from the hub.6. The impeller of wherein the cap member includes a ring forming a central passage shaped to cooperatively overlay the hub and a plurality of fingers extending radially from the ring and shaped to cooperatively overlay the vanes of the body.7. The impeller of wherein the mated recesses and protrusions extend radially from an interior of the impeller ...

NICKEL BASE SUPERALLOY COMPOSITIONS BEING SUBSTANTIALLY FREE OF RHENIUM AND SUPERALLOY ARTICLES

A nickel base superalloy composition substantially free of rhenium includes, in percentages by weight: about 5-8 Cr; about 7-8 Co; about 1.3-2.2 Mo; about 4.75-6.75 W; about 6.0-7.0 Ta; if present, up to about 0.5 Ti; about 6.0-6.4 Al; about 0.15-0.6 Hf; if present, from about 0.03-0.06 C; if present, up to about 0.004 B; if present, one or more rare earths selected from Y, La, and Ce up to about 0.03 total, the balance including nickel and incidental impurities. The superalloy composition is able to provide sustained-peak low cycle fatigue and/or oxidation resistance properties comparable to second generation superalloy compositions including at least about 3 wt % rhenium. Superalloy articles incorporating the compositions include nozzles, shrouds, and splash plates for gas turbine engines. 1. A nickel base superalloy composition consisting of , in percentages by weight:about 5-6 Cr;about 7-8 Co;about 1.5 Mo;about 6 W;about 6.0-7.0 Ta;if present, up to about 0.5 Ti;about 6.0-6.4 Al;about 0.15-0.6 Hf;if present, from about 0.03-0.06 C;if present, up to about 0.004 B;if present, one or more rare earths selected from Y, La, and Ce up to about 0.03 total;wherein the superalloy composition is substantially free of Re and exhibits a sustained-peak low cycle fatigue resistance exceeding 10,500 cycles to failure at 2000° F./18 ksi APS;the balance of the superalloy composition being nickel and incidental impurities.2. The nickel base superalloy composition according to being characterized by a P-value of less than 3360 claim 1 , wherein the P-value is defined as: P=−200 Cr+80 Mo−20 Mo−250 Ti−50 (Ti×Ta)+15 Cb+200 W−14 W+30 Ta−1.5 Ta+2.5 Co+1200 Al −100 Al+100 Re+1000 Hf−2000 Hf+700 Hf−2000 V−500 C−15000 B−500 Zr.3. The nickel base superalloy composition according to being characterized by a P-value of less than 3050.4. The nickel base superalloy composition according to wherein the superalloy composition is in the form of a single crystal article.5. The nickel base ...

BLADE

A fan blade for a gas turbine engine has an aerofoil part and a root part. The root part includes a root former; the root former includes a zone of weakness, which reduces the ability of the root part to withstand an impact force. Thus, in an impact situation in which the fan blade has separated from the fan rotor and the fan blade has itself separated into fragments, the root part will fracture or buckle more easily than would be the case with conventional arrangements. This will lower the impact force of the root part upon the fan casing, thus permitting the fan casing to be designed to withstand lower impact forces. The fan casing can therefore be made lighter, and cheaper, than in conventional arrangements. 1. A blade for a turbine engine , comprising:an aerofoil part;a root part; anda root former, including a zone of weakness, said root former included in said root part.2. The blade according to claim 1 , in which the root part has a length claim 1 , and the zone of weakness is arranged to reduce the ability of the root part to withstand a force applied transversely across the length of the root part.3. The blade according to claim 1 , in which the root part is curved in the form of an arc claim 1 , and the zone of weakness reduces the ability of the root part to withstand a force applied transversely and radially across the arc.4. The blade according to claim 1 , in which the zone of weakness extends transversely across the root former.5. The blade according to claim 1 , in which the zone of weakness is in the form of a line of weakness.6. The blade according to claim 1 , in which the zone of weakness is in the form of a line of separation.7. The blade according to claim 1 , in which the root former includes a plurality of root former parts claim 1 , the zone of weakness is located between adjacent root former parts claim 1 , the zone of weakness includes a space defined between the adjacent root former parts claim 1 , and the space is in the form of a hole.8. ...

THREE-DIMENSIONALLY WOVEN COMPOSITE BLADE WTIH SPANWISE WEFT YARNS

A composite blade has a root and a tip in a spanwise direction and a leading edge and a trailing edge in a chordwise direction. The composite blade includes a three-dimensional woven preform having weft yarns and warp yarns. The weft yarns extend in the spanwise direction of the composite blade. The warp yarns interweave the weft yarns and extend in the chordwise direction of the blade. 1. A composite blade having a root and a tip in a spanwise direction and a leading edge and a trailing edge in a chordwise direction , the composite blade comprising: weft yarns extending in the spanwise direction of the composite blade; and', 'warp yarns interweaving the weft yarns and extending in the chordwise direction of the composite blade., 'a three-dimensionally woven preform comprising2. The composite blade of claim 1 , wherein the three-dimensionally woven preform further comprises stuffer yarns extending in the chordwise direction of the blade.3. The composite blade of claim 1 , wherein the weft yarns are generally straight and the warp yarns have a crimp angle.4. The composite blade of claim 1 , wherein the warp yarns interweave the weft yarns with a layer-to-layer angle interlock weave pattern.5. The composite blade of claim 1 , wherein the warp yarns interweave the weft yarns with a through-thickness angle interlock weave pattern.6. The composite blade of claim 1 , wherein the warp yarns interweave the weft yarns in the root with a first weave pattern and the warp yarns interweave the weft yarns in the tip with a second weave pattern.7. The composite blade of claim 1 , wherein the warp yarns vary in yarn size as measured by filament count in the spanwise direction to change a thickness of the three-dimensional woven airfoil.8. The composite blade of claim 1 , and further comprising:a first plurality of plies positioned between a pressure side of the composite blade and the three-dimensionally woven preform; anda second plurality of plies positioned between a suction ...

Transonic Blade

The present invention provides a transonic blade that concurrently achieves reduction in shock loss at a design point and improvement in stall margin in blades operating in a flow field of transonic speed or higher in an axial-flow rotating machine. A cross-sectional surface at each of spanwise positions of the blade is shifted parallel to a stagger line connecting a leading edge with a trailing edge of the blade. A stacking line is shifted toward an upstream side of working fluid. The stacking line connects together respective gravity center positions of blade cross-sectional surfaces at spanwise positions in a range from a hub cross-sectional surface joined to a rotating shaft or an outer circumferential side casing of a rotating machine to a tip cross-sectional surface lying at a position most remote from the hub cross-sectional surface in a spanwise direction. 1. A transonic blade used in a flow field through which an overall or partial flow passes at a transonic speed or higher ,wherein the transonic blade is formed such that:a cross-sectional surface at each of spanwise positions of the blade is shifted parallel to a line connecting a leading edge with a trailing edge of the blade;a stacking line is shifted toward an upstream side of working fluid, the stacking line connecting together gravity center positions of blade cross-sectional surfaces at respective spanwise positions in a range from a hub cross-sectional surface joined to a rotating shaft or an outer circumferential side casing of a rotating machine to a tip cross-sectional surface lying at a position most remote from the hub cross-sectional surface in a spanwise direction; anda maximum thickness position of the cross-sectional surface of the blade at each of the spanwise positions are shifted toward a trailing edge side of the blade in a range from the hub cross-sectional surface to the tip cross-sectional surface.2. A transonic blade used in a flow field through which an overall or partial flow ...

ALUMINIDE OR CHROMIDE COATINGS OF CAVITIES

Disclosed is a process for producing an alloyed, in particular multiple-alloyed aluminide or chromide layer on a component by alitizing or chromizing. First a green compact layer () consisting of a binder () and metal particles () is deposited on the component () to be coated and then alitizing or chromizing is carried out, binder and metal particles being deposited on the component separately from one another, first the binder and then the metal particles. A turbine component produced by this process is also disclosed. 1. A process for producing at least one alloyed aluminide and/or chromide layer on a component by alitizing and/or chromizing , wherein the process comprises depositing a green compact layer of a binder and metal particles on the component to be coated and thereafter carrying out alitizing and/or chromizing , the binder and the metal particles being deposited on the component separately from one another by first applying the binder to form a binder layer , and then depositing the metal particles on or in the binder layer.2. The process of claim 1 , wherein more than one alloyed aluminide and/or chromide layer is produced on the component.3. The process of claim 1 , wherein the binder is applied to the component by dip coating and then the metal particles are applied by dusting or embedding the component in a powder bed of the metal particles to be deposited.4. The process of claim 1 , wherein applying binder and depositing metal particles is repeated two or more times.5. The process of claim 1 , wherein the green compact layer is subjected to a heat treatment after the last metal particles have been applied.6. The process of claim 4 , wherein the green compact layer is subjected to a heat treatment after each time binder has been applied and metal particles have been deposited thereafter.7. The process of claim 4 , wherein the green compact layer is subjected to a heat treatment after applying binder and depositing metal particles has bee repeated ...

VERTICAL AXIS WIND TURBINE AIRFOIL

A vertical axis wind turbine airfoil is described. The wind turbine airfoil can include a leading edge, a trailing edge, an upper curved surface, a lower curved surface, and a centerline running between the upper surface and the lower surface and from the leading edge to the trailing edge. The airfoil can be configured so that the distance between the centerline and the upper surface is the same as the distance between the centerline and the lower surface at all points along the length of the airfoil. A plurality of such airfoils can be included in a vertical axis wind turbine. These airfoils can be vertically disposed and can rotate about a vertical axis. 1. (canceled)2. (canceled)3. (canceled)4. (canceled)5. (canceled)6. (canceled)7. (canceled)8. (canceled)9. (canceled)10. (canceled)11. (canceled)12. (canceled)13. (canceled)14. (canceled)15. (canceled)16. (canceled)17. (canceled)18. An aerodynamic regulator system for a wind turbine comprising:at least one aerodynamic regulator configured to rotate about a first axis and comprising a first moveable airfoil and a second moveable airfoil, wherein the first moveable airfoil and the second moveable airfoil are configured to move in different directions; andwherein movement of said first moveable airfoil and said second moveable airfoil slows the rotational speed of said wind turbine.19. The aerodynamic regulator system of claim 18 , wherein said first moveable airfoil and said second moveable airfoil are rotatable about a second axis wherein said second axis extends outwards from said first axis.20. The aerodynamic regulator system of claim 19 , wherein said aerodynamic regulator is configured to rotate said first moveable airfoil and said second moveable airfoil about said second axis in different directions.21. The aerodynamic regulator system of claim 20 , wherein rotation of said first and second moveable airfoils about said second axis changes the angle of attack of said first and second moveable airfoils and ...

MOVABLE BLADE FOR A TURBOMACHINE

A movable blade for a turbomachine, the blade including at its distal end a top portion including opposite first and second side edges and upstream and downstream sealing wipers. Each of the first and second side edges has a substantially Z-shaped profile between the upstream and downstream wipers. On a side of the first side edge, the platform includes a first outwardly projecting rim not connected to either of the upstream and downstream wipers. The platform includes a second outwardly projecting rim, the second rim including a downstream portion connected to the downstream wiper and extending along the second edge, and an upstream portion extending the downstream portion as far as the upstream wiper, and set back relative to the second side edge. 14-. (canceled)5. A movable blade for a turbomachine , the blade comprising at its distal end a top portion comprising:a platform defining an outside surface of a passage for gas passing through the turbomachine, and opposite first and second side edges; andupstream and downstream sealing wipers extending radially outwards from the platform;wherein each of the first and second side edges includes a substantially Z-shaped profile between the upstream and downstream wipers, which profile includes a first segment closer to the upstream wiper, an intermediate second segment, and a third segment closer to the downstream wiper, the first and third segments being substantially parallel to each other, and the second segment extending obliquely between the first and third segments;wherein the top portion includes, on a side of the first side edge, a first outwardly projecting rim that extends along the second segment of the first side edge, the first rim not being connected to either of the upstream and downstream wipers; andthe top portion includes, on a side of the second side edge, a second outwardly projecting rim, the second rim having a downstream portion connected to the downstream wiper and extending along the second and ...

MOVING BLADE AND TURBOMACHINE

A moving blade for a turbomachine, in particular an aircraft engine, is disclosed, having an inner shroud which has a front elongation for forming an axial overlap with an upstream guide blade, and on which at least one flow guide element for deflecting a leakage flow of a cooling air flow in the peripheral direction is situated. The at least one flow guide element is guided beyond a leading edge of the elongation. A turbomachine having a plurality of these types of moving blades is also disclosed. 1. A moving blade for a turbomachine comprising:an inner shroud having a front elongation for forming an axial overlap with an upstream guide blade;at least one flow guide element for deflecting a leakage flow of a cooling air flow being situated in a peripheral direction, the at least one flow guide element extending in an axial direction beyond a leading edge of the elongation.2. The moving blade as recited in wherein the flow guide element terminates in flush alignment with a hot gas side and with a cooling air side of the elongation.3. The moving blade as recited in wherein the flow guide element protrudes beyond a hot gas side and a cooling air side of the elongation in the radial direction.4. The moving blade as recited in wherein the flow guide element has a hot gas side section on a hot gas side of the elongation.5. The moving blade as recited in wherein the hot gas side is located over an entirety of an axial extension of the elongation.6. The moving blade as recited in wherein the hot gas side section has a constant height.7. The moving blade as recited in wherein the flow guide element has a cooling air side section on a cooling air side of the elongation.8. The moving blade as recited in wherein the cooling gas side is located over an entirety of an axial extension of the elongation.9. The moving blade as recited in wherein the cooling gas side section has a constant height.10. The moving blade as recited in wherein the flow guide element has a hot gas side ...

BLADE HAVING AN INTEGRATED COMPOSITE SPAR

A blade including a structure with aerodynamic profile including two opposite skins obtained by three-dimensional weaving of a fibrous reinforcement densified by a matrix, and a longeron including a fibrous reinforcement obtained by three-dimensional weaving and densified by a matrix, the longeron including a first part extending outside the structure with aerodynamic profile and designed to be connected to a drive hub in rotation of the blade and a second part arranged inside the structure with aerodynamic profile between the two skins. The second part of the longeron has a thickness substantially similar to that of the skins of the structure with aerodynamic profile. In addition, the fibrous reinforcement of the second part of the longeron has the same weaving armour as that of the reinforcement of the skins of the structure with aerodynamic profile. 1. A blade comprising:a structure with aerodynamic profile comprising two opposite skins obtained by three-dimensional weaving of a fibrous reinforcement densified by a matrix,a longeron comprising a fibrous reinforcement obtained by three-dimensional weaving and densified by a matrix, said longeron comprising a first part extending outside said structure with aerodynamic profile and designed to be connected to a drive hub in rotation of the blade and a second part arranged inside the structure with aerodynamic profile between the two skins,the second part of said longeron comprising one or two portions each having a thickness substantially similar to that of the skins of the structure with aerodynamic profile, andthe fibrous reinforcement of the second part of the longeron having the same weaving armour as that of the reinforcement of the skins of the structure with aerodynamic profile.2. The blade according to claim 1 , wherein the first part of the longeron extends underneath the second part of the longeron in increasing thickness.3. The blade according to claim 1 , wherein the variation in thickness between the ...

WIND TURBINE COMPONENT COMPRISING RADAR-ABSORBING MATERIAL

A method of making a wind turbine component of composite construction with enhanced radar absorbing properties is described. The method comprises making the component and then modifying the component by applying circuit analogue elements to a surface of the component. 1. A method of making a wind turbine component of composite construction with enhanced radar absorbing properties , wherein the method comprises making the component and then modifying the component by applying circuit analogue elements to a surface of the component.2. The method of claim 1 , wherein the step of making the component comprises joining two elements together to form at least part of the component and the step of modifying the component includes applying the circuit analogue elements to a region of the surface that bridges an interface between the two elements.3. The method of claim 1 , wherein the step of making the component comprises arranging in a mould one or more fibrous layers within a matrix material.4. The method of claim 4 , wherein the step of making the component comprises curing the matrix material.5. The method of claim 4 , wherein the step of making the component comprises providing a gel coat on a surface of the mould such that the resulting moulded component has a gel-coated outer surface claim 4 , and the step of modifying the component comprises applying the circuit analogue elements over the gel coated outer surface of the component.6. The method of claim 1 , wherein the step of making the component comprises arranging a ground plane in spaced apart relation with the surface of the component to which the circuit analogue elements are to be applied.7. The method of claim 1 , wherein the method comprises applying the circuit analogue elements as a repeating pattern on the surface of the component.8. The method of claim 1 , wherein the method comprises printing the circuit analogue elements on the surface of the component.9. The method of claim 1 , wherein the method ...

ELECTRICAL WIRING SYSTEM FOR A ROTOR HUB

The electrical harness system is configured for routing a harness between a rotor yoke and a rotor blade. A recess in an inboard cap member is configured to house a connector and an associated harness. As the recess extends along a radial path in the chordwise direction of the rotor blade, the harness is configured to lie within the recess. Operationally induced centrifugal forces promote positioning the slack of harness within the recess, while the slack in the harness remains available for relative dynamic movements between the rotor blade and the rotor yoke. Such a routing of the harness reduces aerodynamic drag and minimizes damage that could otherwise occur to the harness. 1. A rotor blade , comprising:an airfoil shaped body;a root end portion that is inboard from the airfoil shaped body;a cap member at the root end portion, the cap member having a recessed portion; anda first connector member coupled to a wall within the recessed portion.2. The rotor blade according to claim 1 , wherein the first connector member is located toward a leading edge portion of the recessed portion.3. The rotor blade according to claim 1 , wherein the wall has a surface that is approximately normal an adjacent inner surface of the recessed portion.4. The rotor blade according to claim 1 , wherein the first connector member has a threaded portion configured for receiving a second connector member.5. The rotor blade according to claim 1 , wherein the first connector member is an electrical power connector.6. The rotor blade according to claim 1 , wherein the first connector member is an electrical signal connector.7. The rotor blade according to claim 1 , wherein the first connector member is a hydraulic connector.8. The rotor blade according to claim 1 , wherein the recessed portion is configured to house a harness coupled to a second connector member claim 1 , the second connector member configured for mating to the first connector member.9. The rotor blade according to claim 1 , ...

BLADE OF A TURBOMACHINE, HAVING PASSIVE BOUNDARY LAYER CONTROL

A blade of a turbomachine is disclosed. A contour variation is provided on the suction side of the blade, where the contour variation has a negative step as viewed in a direction of flow. The step has a stepped surface extending perpendicularly to a contour of the suction side and the contour variation has a tangential surface which leads upstream tangentially on the contour of the suction side starting from a step edge. 17.-. (canceled)8. A blade of a turbomachine , comprising:a leading edge;a trailing edge;a suction side and a pressure side extending between the leading edge and the trailing edge; anda contour variation on the suction side, wherein the contour variation has a negative step as viewed in a direction of flow, wherein the step has a stepped surface extending perpendicularly to a contour of the suction side, and wherein the contour variation has a tangential surface which leads upstream tangentially on the contour of the suction side starting from a step edge.9. The blade according to claim 8 , wherein the contour variation is located at a distance from the leading edge which corresponds to approximately 50% to 85% of an axial chord length of the blade.10. The blade according to claim 8 , wherein the contour variation is located at a distance from the leading edge which corresponds to approximately 65% of an axial chord length of the blade.11. The blade according to claim 8 , wherein the step has a step height in a range of 0.1 mm to 0.4 mm.12. The blade according to claim 8 , wherein the step has a step height of 0.17 mm.13. The blade according to claim 8 , wherein the step has a step height of 0.15 mm.14. The blade according to claim 8 , wherein the contour variation is obtained by applying material to claim 8 , or removing material from claim 8 , the blade.15. The blade according to claim 8 , wherein the blade is a turbine blade of an aircraft engine. This application claims the priority of International Application No. PCT/DE2011/001066, filed May ...

PROPELLER BLADE WITH MODIFIED SPAR LAYUP

According to one embodiment, a propeller blade includes a foam core and a structural layer formed of multiple layers that surrounds at least a portion of the foam core. The structural layer includes a mid-thickness location defined between the foam core and an outer edge of the structural layer and the multiple layers include at least one unidirectional layer and at least one biased layer disposed asymmetrically about the mid-thickness location. 1. A propeller blade comprising:a foam core; anda structural layer formed of multiple layers that surrounds at least a portion of the foam core, the structural layer including a mid-thickness location defined between the foam core and an outer edge of the structural layer;wherein the multiple layers include at least one unidirectional layer and at least one biased layer disposed asymmetrically about the mid-thickness location.2. The propeller blade of claim 1 , wherein the layers include structural fibers disposed therein.3. The propeller blade of claim 2 , wherein the structural fibers in the unidirectional layer extend in a span wise direction along the propeller blade.4. The propeller blade of claim 2 , further comprising:additional fibers not contained in either the at least one unidirectional layer or the at least one biased layers that extend in a chord wise direction along the propeller blade.5. The propeller blade of claim 1 , further comprising:a layer of fiberglass at least partially disposed between the foam core and the structural layer.6. The propeller blade of claim 1 , further comprising:a root disposed within the structural layer at an end of the propeller blade.7. The propeller blade of claim 1 , wherein the foam core is formed of polyurethane foam claim 1 , polyisocyanurate foam claim 1 , or polymethacrylimide foam8. A method of forming a propeller blade comprising:forming a foam core; anddisposing a plurality of braided layers over at least a portion of the foam core to form a structural layer having mid- ...

COMPOSITE TURBINE BLADE AND METHOD OF MANUFACTURE

A composite turbine blade and a method of manufacture thereof is disclosed. The composite turbine blade comprises a turbine blade portion comprising a first material and a first tip plate comprising a second material. The turbine blade portion has an exterior wall and an interior wall surrounding a hollow interior cavity, and a top surface extending from the exterior wall to the interior wall bounding an orifice that is fluidly connected to the hollow interior cavity. The first tip plate may be attached to the turbine blade along the top surface and extending from proximate the exterior wall of the turbine blade across the orifice to cover the orifice. 1. A composite turbine blade comprising:a turbine blade portion comprising a first material, the turbine blade portion comprising an airfoil portion having a tip and a root and a platform portion attached to the airfoil portion at the root, the platform portion comprising an orifice passing therethrough; andan insert attached within the orifice of the platform portion, the insert comprising a second material.2. The composite turbine blade of claim 1 , wherein the second material comprises a different material than the first material.3. The composite turbine blade of claim 1 , wherein the second material comprises a material having a greater resistance to low cycle fatigue than the first material.4. The composite turbine blade of claim 1 , wherein the insert comprises a flange portion claim 1 , the flange portion mating with the platform portion of the composite turbine blade.5. The composite turbine blade of claim 1 , wherein the insert comprises a first tapered sidewall claim 1 , the first tapered sidewall of the insert mating with a second tapered sidewall of the orifice of the platform portion.6. The composite turbine blade of claim 2 , wherein the second material has a different grain orientation than the first material.7. The composite turbine blade of claim 2 , wherein the second material comprises a material ...

METHOD OF MAKING A METAL REINFORCING PIECE

A metal reinforcing piece for mounting on a leading edge or trailing edge of a composite blade for a turbine engine is made by shaping two metal sheets, positioning them on either side of a core, assembling the two sheets together around the core under a vacuum, shaping them on the core by hot isostatic compression, and cutting them to separate the reinforcing piece and release the core. A predetermined roughness is given to at least a portion of the surface of the core and is transferred to a corresponding portion of an inside surface of the reinforcing piece by the hot isostatic compression. 110-. (canceled)11. A method of making a metal reinforcing piece for mounting on a leading edge or trailing edge of a composite material blade of a turbine engine , the method comprising:shaping two metal sheets to approximate a final shape of the reinforcing piece that is to be made;positioning the two sheets on either side of a core reproducing inside shapes of suction and pressure sides of the reinforcing piece;assembling the two metal sheets together around the core in a sealed manner under a vacuum;shaping the sheets onto the core by hot isostatic compression;cutting the sheets to separate the reinforcing piece and release the core; andgiving non-zero predetermined roughness to at least a portion of the surface of the core and transferring the roughness to a corresponding portion of an inside surface of the reinforcing piece by hot isostatic compression.12. A method according to claim 11 , wherein at least a portion of the surface of the core has roughness Ra greater than 3 μm claim 11 , or in a range of 3.2 μm to 6.4 μm.13. A method according to claim 11 , wherein an outside surface of the core includes zones of different roughnesses.14. A method according to claim 11 , wherein the roughness is on an outside surface of the core and is obtained by milling claim 11 , by mechanical shot peening claim 11 , by sand blasting claim 11 , by laser machining claim 11 , by electro- ...

WIND TURBINE

A vertical axis wind turbine having multiple blades in which the root of each blade is oriented at an angle to a radial direction from a drive shaft and at least a portion each blade is swept back. The blades preferably have a symmetric partial aerofoil shape of generally V-shaped cross section of generally constant cross section. 1. A vertical axis wind turbine including:a. a shaft having a generally vertical axis of rotation; andb. a rotor comprising a plurality of blades extending from the shaft, each blade having a top portion with a leading edge and trailing edge and a bottom portion with a leading edge and trailing edge wherein the top portion and bottom portion are connected along at least a portion of their respective leading edges to form a leading edge of the blade and the respective trailing edges of the top portion and bottom portion diverge, and wherein the root of each blade is oriented at an angle to a radial direction from the shaft and at least a portion each blade is swept back.2. A vertical axis turbine as claimed in in which the ratio of the area of the upwind blades exposed to incident wind times the effective lever arm length is less than the area of the downwind blades exposed to the incident wind times the effective lever arm length.3. A vertical axis turbine as claimed in wherein the ratio is less than 0.9.4. A vertical axis turbine as claimed in wherein the ratio is less than 0.8.5. A vertical axis turbine as claimed in wherein the ratio is about 0.55.6. A vertical axis turbine as claimed in wherein the blades have a convex partial aerofoil shape.7. A vertical axis turbine as claimed in wherein the centre of camber is located towards the front of the aerofoil.8. A vertical axis turbine as claimed in wherein the blades have a generally V-shaped cross section.9. A vertical axis turbine as claimed in wherein the blade portions arc of substantially equal cross sectional length.10. A vertical axis turbine as claimed in wherein the blade portions ...

SPACE FRAME STRUCTURE

A structure comprising a space frame including a plurality of strut sections which cross at one or more locations within the space frame to define a node at each respective crossing location, wherein the strut sections are laminates comprising a plurality of ply layers, and wherein the ply layers of the crossing strut sections are interleaved at the node(s). Also a method of forming the structure. Also An aerofoil structure, comprising a space frame defining a cage adapted to receive upper and lower covers so as to form upper and lower aerodynamic surfaces of the aerofoil, and at least one spar extending generally spanwise across the structure, wherein the spar includes a spar web and upper and lower spar caps, and wherein the spar web is disposed on one of an interior or exterior side of the cage, and the spar caps are attached to the other of the interior or exterior side of the cage. 1. A structure comprising a space frame including a plurality of strut sections which cross at one or more locations within the space frame to define a node at each respective crossing location , wherein the strut sections are laminates comprising a plurality of ply layers , and wherein the ply layers of the crossing strut sections are interleaved at the node(s).2. A structure according to claim 1 , wherein at least one of the strut sections has a generally non-planar profile.3. A structure according to claim 2 , wherein the profile of the strut section flattens as it approaches a node.4. A structure according to claim 1 , wherein at least one of the node(s) is substantially planar.5. A structure according to claim 1 , wherein at least one of the strut sections has a termination and a generally non-planar profile which flattens as it approaches the termination.6. A structure according to claim 1 , wherein the space frame includes a composite material.7. A structure according to claim 1 , wherein the space frame has at least two integrally formed sides.8. A structure according to ...

ARCHIMDEAN SCREW APPARATUS

An Archimedean screw apparatus for use either in power generation or for the pumping or conveying of fluid material includes a screw body formed by a shaft and at least one helical flight that is located in a close-fitting channel structure. The screw body is rotatably mounted such that the flight is in frictional contact with in inner surface of the channel. The weight of the screw body is borne at least partially by the channel structure and is dissipated fully or partially along the length of channel structure. Hence, preferably the Screw both is either provided with only a single bearing located at one end of the shaft, the other end being left floating within the channel structure, or the screw body is not provided with any bearings and the shaft is connected directly to a drive mechanism or to part of a power generating apparatus. 1. An Archimedean screw apparatus comprising a screw body formed b a shaft and at least one helical flight located in a close-fitting channel structure , the screw body being rotatably mounted such that the flight is in frictional contact with an inner surface of the channel structure.2. An apparatus as claimed in claim 1 , wherein the screw body comprises a plurality of up to seven flights.3. An apparatus as claimed in claim 1 , wherein the channel structure comprises a trough claim 1 , an open channel claim 1 , a closed channel or a tube.4. An apparatus as claimed in claim 1 , wherein the screw body is provided with a single bearing located at one end of the shaft.5. An apparatus as claimed in claim 1 , wherein an other end of the screw body floats within the channel structure.6. An apparatus as claimed in claim 1 , wherein the screw body is not provided with any bearings.7. An apparatus as claimed in claim 1 , wherein the shaft is connected directly to a drive mechanism or to part of a power generating apparatus.8. An apparatus as claimed in claim 1 , comprising means retaining the screw body in position within the channel ...

Airfoil mateface sealing

An airfoil includes an airfoil working portion and an adjoining endwall. The endwall has a leading edge, a trailing edge, a first mateface, and a second mateface that collectively define a perimeter of the endwall, with the first and second matefaces arranged opposite one another. A forward zone is defined by the endwall that extends to the leading edge. The first and second matefaces are each oriented at an angle α II relative to radial in the forward zone. An aft zone is defined by the endwall that extends to the trailing edge. The first and second matefaces are each oriented at an angle α II relative to radial in the aft zone. The angles α II and α II are not equal. A middle zone is defined in between the forward and aft zones, and the first and second matefaces transition between the angles α II and α I in the middle zone.

COMPONENTS WITH COOLING CHANNELS AND METHODS OF MANUFACTURE

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

Airfoil cover system

An example airfoil for a gas turbine engine includes a body having a first surface extending from a first edge to a second edge and a cavity disposed in the body. A first cover is at least partially disposed within the cavity. The first cover includes a first portion cooperates with a corresponding second portion. A second cover covers the first cover and forms at least a portion of the first surface with the body. The first cover is disposed between the body and the second cover. The first cover and the second cover have a different coefficient of thermal expansion than the body.

PROPELLER BLADE WITH SPAR RIB

A propeller blade includes a foam core having a groove formed therein, a fibrous material filling at least a portion of the groove and a structural layer that surrounds the fibrous material and at least a portion of the foam core. 1. A propeller blade comprising:a foam core having a groove formed therein;a fibrous material filling at least a portion of the groove; anda structural layer that surrounds the fibrous material and at least a portion of the foam core.2. The propeller blade of claim 1 , further comprising:a layer of fiberglass or carbon at least partially disposed between the foam core and the fibrous material.3. The propeller blade of claim 1 , wherein the foam core includes a leading edge and a trailing edge and the fibrous material extends from the leading end to the trailing edge.4. The propeller blade of claim 1 , wherein the groove has a depth and the fibrous material fills the depth of the groove.5. The propeller blade of claim 1 , further comprising:a root disposed within the structural layer at an end of the propeller blade.6. The propeller blade of claim 1 , wherein the foam core is formed of one or more of: polyurethane (PU) claim 1 , polyisocyanurate claim 1 , and polymethacrylimide (PMI).7. The propeller blade of claim 1 , wherein the structural layer is formed of a resin-impregnated fiber material.8. The propeller blade of claim 7 , wherein the fibrous material is formed of the same material as the structural layer.9. The propeller blade of claim 1 , wherein the fibrous material is formed of carbon fibers.10. The propeller blade of claim 1 , wherein both the fibrous material and the structural layer have a resin disposed therein.11. A method of forming a propeller blade comprising:forming a foam core, the form core including a groove formed therein;disposing a fibrous material in the groove; andforming a structural layer that surrounds fibrous material and a portion of the foam core.1216. The method of claim claim 1 , further comprising: ...

METHOD FOR PRODUCING A PROTECTIVE LAYER FOR A ROTOR BLADE

A method for producing a protective layer for a rotor blade of a turbomachine is proposed. The rotor blade has a wear cap wall that in relation to the axis of the turbomachine protrudes radially towards the exterior of the turbomachine beyond a body of the rotor blade. The method enables the entire rotor blade to have a comparatively long service life and at the same time allows the wear cap wall to be thermally insulated in an especially simple manner. For the purpose, a profile producing a positive fit effective in the radial direction is inserted in a radially located region of the wear cap wall, a slip containing metal powder and/or a ceramic slip is applied to the region and the slip in sintered. 110.-. (canceled)11. A method for producing a protective layer for a rotor blade of a turbomachine having a wear tip wall projecting toward an outside of the turbomachine beyond a body of the rotor blade in a radial direction with respect to an axis of the turbomachine , comprising:introducing a profile producing a radial form fit into a radial region of the wear tip wall; andapplying a metal powder-containing slurry and/or a ceramic slurry to the region,wherein the slurry is sintered.12. The method as claimed in claim 11 , wherein the profile is cast and/or introduced by electric discharge machining.13. The method as claimed in claim 11 , wherein the slurry is dried before sintering claim 11 , and wherein shape of the dried slurry is machined.14. The method as claimed in claim 11 , wherein the slurry is water based and/or an organic binder is added to the slurry.15. The method as claimed in claim 11 , wherein a metal powder-containing slurry and a ceramic slurry are applied in sequence.16. The method as claimed in claim 11 , wherein the slurry comprises ceramic fibers or gauzes.17. The method as claimed in claim 11 , wherein the wear tip wall is arranged around a circumference of the rotor blade and forms a recess arranged radially toward the outside of the ...

High Volume Low Speed Fan

An HVLS fan system uses STOL technology for airfoils and angle of attack thus optimizing air movement efficiency and reducing drag. The HVLS fan system includes wingtip fence end caps to the airfoils for improving efficiency by reducing drag. The HVLS fan system also includes an interconnection of the airfoils to a securing plate thus providing a failsafe and reduced potential for damage or injury resulting from failure of the connection between the airfoil array and a drive unit such as an electric motor and associated gearing. 1. A high volume low speed fan system , comprising:a plurality of airfoils each having a generally elliptical top surface and a generally elliptical bottom surface wherein each airfoil is mounted to a flange on a central hub wherein the flange is angled between 7 and 10 degree from horizontal;a motor attached to a frame member wherein the motor is coupled to a drive mechanism for rotating the airfoils wherein the drive mechanism is rotatably coupled to the central hub; anda plurality of retaining members coupled to the central hub each passing through an opening of a safety frame coupled to the frame member wherein the retaining members pass through the opening of the safety frame without contacting the safety frame or the frame member wherein operational separation of the drive mechanism from the central hub does not result in separation of the central hub from the safety frame.2. The high volume low speed fan system of claim 1 , wherein the flange is angled between 8 and 9 degrees from horizontal.3. The high volume low speed fan system of claim 1 , wherein the flange is angled at 8 degrees from horizontal.4. The high volume low speed fan system of claim 1 , wherein the drive mechanism comprises:a drive shaft rotatably coupled to the motor; anda bushing coupled to the drive shaft wherein the bushing is housed in a cylindrical member and the cylindrical member is coupled to the central hub.5. The high volume low speed fan system of claim 4 , ...

PROCESS FOR PRODUCING A RUN-IN COATING

Disclosed is a process for producing a run-in coating () on a component of a turbomachine, in particular of a gas turbine. The run-in coating is applied and produced on the component of the turbomachine by a kinetic cold gas compacting process (K3). The invention also encompasses a run-in coating for a static or rotating component of a turbomachine and a static or rotating component of a turbomachine, in particular of a gas turbine, having at least one run-in coating. 1. A process for producing a run-in coating on a static or rotating component of a turbomachine , wherein the run-in coating is applied and produced on the component of the turbomachine by a kinetic cold gas compacting process (K3).2. The process of claim 1 , wherein the turbomachine is a gas turbine.3. The process of claim 1 , wherein prior to applying the run-in coating at least one bonding layer and/or at least one thermal barrier layer or titanium fire protection layer is applied to a surface of the static or rotating component to be coated with the run-in coating to form a composite run-in coating.4. The process of claim 3 , wherein the at least one bonding layer and/or at least one thermal barrier layer or titanium fire protection layer is produced by a thermal spraying process.5. The process of claim 4 , wherein the at least one bonding layer and/or at least one thermal barrier layer or titanium fire protection layer is produced by a kinetic cold gas compacting process (K3).6. The process of claim 1 , wherein after application of the run-in coating to the static or rotating component claim 1 , at least one of a structure and a contour is formed at least one of on and in the run-in coating.7. The process of claim 6 , wherein the at least one of a structure and a contour is formed by an electrochemical material removal process.8. The process of claim 7 , wherein the electrochemical material removal process comprises at least one of electrochemical machining (ECM) and precise/pulsed electrochemical ...

COMPOSITE AIRFOIL WITH INTEGRAL PLATFORM

A method of forming an airfoil with an integrated platform includes: a) providing an airfoil core; b) wrapping a first overwrap ply around the airfoil core; c) darting a first end of the first overwrap ply to allow the overwrap ply to extend perpendicular to the airfoil core to form a first platform; d) filling the darted parts filler plies; e) wrapping a second overwrap ply around the first overwrap ply; f) darting a first end of the second overwrap ply to allow the second overwrap ply to extend adjacent to the first overwrap ply to form the first platform; g) filling the darted parts of the second overwrap ply with one or more filler plies; and h) placing a cap ply in the shape of the platform adjacent to at least one of the first and second overwrap plies. An airfoil with an integrated platform is also disclosed. 1. A method of forming an airfoil with an integrated platform , the method comprising:a) providing an airfoil core;b) wrapping a first overwrap ply around the airfoil core;c) darting a first end of the first overwrap ply to allow the overwrap ply to extend at an angle to the airfoil core to form a first platform;d) filling the darted parts of the first overwrap ply with one or more filler plies;e) wrapping a second overwrap ply around the first overwrap ply;f) darting a first end of the second overwrap ply to allow the second overwrap ply to extend adjacent to the first overwrap ply to form the first platform;g) filling the darted parts of the second overwrap ply with one or more filler plies; andh) placing a cap ply in the shape of the platform adjacent to at least one of the first and second overwrap plies.2. The method of claim 1 , and further comprising:placing a cap ply adjacent to the first overwrap ply to form the surface of the platform flow path.3. The method of claim 1 , wherein steps (b)-(h) are repeated at least one time.4. The method of claim 1 , wherein step (c) further comprises:darting a second end of the first overwrap ply to allow it to ...

CORRUGATED PANEL FOR WIND POWER GENERATOR BLADE

A corrugated panel for a wind power generator blade is provided. The corrugated panel for a wind power generator blade includes: a plurality of wrinkles that are coupled to one surface or the other surface of the wind power generator blade having an airfoil transverse section and that are formed in a length direction of the blade, wherein a gap and a height at a predetermined position of the plurality of wrinkles each have a value of a predetermined ratio to a chord length of the airfoil transverse section of the blade at the predetermined position, when the plurality of wrinkles are coupled to the blade. 1. A corrugated panel for a wind power generator blade , the corrugated panel comprising:a plurality of wrinkles that are coupled to one surface or the other surface of the wind power generator blade having an airfoil transverse section and that are formed in a length direction of the blade,wherein a gap and a height at a predetermined position of the plurality of wrinkles each have a value of a predetermined ratio to a chord length of the airfoil transverse section of the blade at the predetermined position, when the plurality of wrinkles are coupled to the blade.2. The corrugated panel of claim 1 , wherein the corrugated panel covers an entire width of one surface of the blade from one end portion to the other end portion of the blade and is coupled to the entire width of one surface of the blade.3. The corrugated panel of claim 1 , wherein the corrugated panel is made of a material comprising one of a titanium alloy claim 1 , an aluminum alloy claim 1 , and a synthetic resin.4. A corrugated panel for a wind power generator blade claim 1 , the corrugated panel comprising:a panel body having one surface that is coupled to one surface or the other surface of the wind power generator blade having an airfoil transverse section; anda wrinkle portion that has a plurality of wrinkles in a length direction of the blade at the other surface of the panel body,wherein a gap ...

DECOUPLED COMPRESSOR BLADE OF A GAS TURBINE

The present invention relates to a compressor blade of a gas turbine having an airfoil made of a fiber-reinforced plastic, which is fastened by means of a blade root to a disk, as well as a metallic leading-edge element, which is arranged on the leading-edge side of the airfoil and partially encompasses the latter, with the leading-edge element itself being fastened to the disk. 1. Compressor blade of a gas turbine having an airfoil made of a fiber-reinforced plastic , which is fastened by means of a blade root to a disk , as well as a metallic leading-edge element , which is arranged on the leading-edge side of the airfoil and partially encompasses the latter , with the leading-edge element itself being fastened to the disk.2. Compressor blade in accordance with claim 1 , characterized in that the leading-edge element contacts the pressure-side and the suction-side surface of the airfoil with a pretension.3. Compressor blade in accordance with claim 2 , characterized in that the leading-edge element contacts the airfoil in the blade axis claim 2 , enabling a relative movement.4. Compressor blade in accordance with claim 1 , characterized in that the leading-edge element includes a first partial element and a second partial element which are connected to one another.5. Compressor blade in accordance with claim 4 , characterized in that the partial elements are connected to one another by means of a laser welding method or by means of positive connecting elements.6. Compressor blade in accordance with claim 1 , characterized in that between the leading-edge element and the airfoil at least one friction element is arranged.7. Compressor blade in accordance with claim 1 , characterized in that the leading-edge element is positively connected to the airfoil.8. Compressor blade in accordance with claim 1 , characterized in that at least one dampening element is arranged in the inflow-side area between the airfoil and the leading-edge element.9. Compressor blade in ...

BLADE CASCADE WITH SIDE WALL CONTOURS AND CONTINUOUS-FLOW MACHINE

A blade cascade for a continuous-flow machine having a non-axisymmetrical side wall contour, whereby the side wall contour has at least one pressure-side elevation and one suction-side depression, whose highest section and lowest section are located over an area 30% to 60% of an extension of the blades in the axial direction, and the axial positions of the outermost sections differ from each other at the maximum by 10% in the axial direction. 1. A blade cascade for a continuous-flow machine , comprising:at least one blade channel delimited in a circumferential direction by a pressure-side wall of a first blade and by an opposite suction-side wall of an adjacent second blade and delimited in a radial direction by two opposite delimiting walls,in the area of the pressure-side wall, at least one of the delimiting walls being provided with at least one pressure-side elevation and, in the area of the suction-side wall, the at least one delimiting wall having at least one suction-side depression, a highest section of the at least one elevation and a lowest section of the at least one depression being located over an area of 30% to 60% of an extension of the blades in an axial direction, and the axial positions of the highest and lowest sections differing from each other at the maximum by 10% in the axial direction.2. The blade cascade as recited in wherein claim 1 , as seen in the circumferential direction claim 1 , the elevation and the depression are at a distance from each other claim 1 , separated by a non-contoured delimiting wall section.3. The blade cascade as recited in wherein the elevation and the depression are arranged downstream from the leading edges of the blades and upstream from the trailing edges of the blades.4. The blade cascade as recited in wherein the highest section of the elevation is downstream from the lowest section of the depression.5. The blade cascade as recited in wherein claim 1 , in the area of the suction-side wall claim 1 , a second ...

BLADE FOR A ROTARY MACHINE

A blade for use in rotary machines such as a wind turbine blade has a plurality of dimples distributed over an area of a low pressure or suction surface of the blade which preferably extends from immediately adjacent to a leading edge of the blade at least partially towards a rear or trailing edge of the blade. 1. A blade having a plurality of dimples distributed over an area of at least one surface of the blade which extends from at or adjacent a leading edge of the blade at least partially towards a rear edge of the blade.2. A blade according to claim 1 , comprising an aerofoil blade.3. A blade as claimed in claim 1 , comprising a low pressure or suction surface and an opposed high pressure or pressure surface claim 1 , and the dimples are distributed over at least an area of the low pressure surface.4. A blade as claimed in claim 2 , in which the aerofoil is asymmetric.5. A blade as claimed in claim 1 , in which the leading edge is rounded and the trailing edge is sharp.6. A blade as claimed in claim 1 , wherein the area of dimples extends along at least a major part of the length of the blade.7. A blade as claimed in claim 1 , wherein the dimples are generally teardrop-shaped with a wider end facing towards the leading edge of the blade.8. A blade as claimed in claim 1 , wherein the dimples vary in distribution density.9. A blade as claimed in claim 1 , wherein the dimples vary in depth.10. A blade as claimed in claim 1 , wherein the dimples vary in cross section with depth.11. A blade as claimed in claim 1 , wherein the plurality of dimples comprises dimples of different size.12. A blade as claimed in claim 1 , wherein the trailing edge of the blade has a localised rearward or downstream extension.13. A blade as claimed in claim 12 , wherein the localised rearward extension is at a free end of the blade.14. A blade as claimed in claim 1 , wherein the blade is a rotary machine blade.15. A blade as claimed in claim 1 , wherein the blade is a wind turbine blade.16 ...

BLADE FOR A CONTINUOUS-FLOW MACHINE AND A CONTINUOUS-FLOW MACHINE

A blade for a continuous-flow machine is disclosed, especially an aircraft engine, whereby, starting from the middle section, the cross section of the blade tip is reduced with respect to the middle section, at least over a front partial section in the direction of the leading edge and over at least a rear section in the direction of the trailing edge, and a continuous-flow machine having at least one row of blades including such blades is also disclosed. 1. A runner blade for a continuous-flow machine , the runner blade comprising:a blade with a leading edge and a trailing edge opposite from it, both the leading edge and the trailing edge extending in a main direction of the runner blade and being at a distance from each other in a lengthwise direction of the runner blade, the blade having a pressure-side wall extending between the leading edge and the trailing edge as well as a suction-side wall opposite from the pressure-side wall, the blade having a tip delimiting the pressure-side wall and the suction-side wall in the main direction, and, starting from a middle section, a cross section of the blade tip is gradually reduced with respect to the middle section, in each case at least over a front partial section in the direction of the leading edge and at least over a rear partial section in the direction of the trailing edge.2. The runner blade as recited in wherein the cross section of the middle section is reduced with respect to at least one of the pressure-side wall and the suction-side wall.3. The runner blade as recited in wherein the cross sections of the middle section and of the front and rear partial sections are reduced on both sides.4. The runner blade as recited in wherein the middle section and the front and rear partial sections form a wing-like profile.5. The runner blade as recited in wherein the blade tip has a front end section and a rear end section transitioning into the leading edge and into the trailing edge respectively claim 1 , each of ...

HYBRID STRUCTURE AIRFOIL

A hybrid airfoil for a gas turbine engine is provided that includes a body and a panel. The body has a first side and a second side orientated opposite the first side. The first and second sides extend between a tip, a base, a leading edge and a trailing edge. The body includes a plurality of cavities disposed in the first side of the body, which cavities extend inwardly toward the second side. The cavities collectively form an opening. At least one rib is disposed between the cavities. A shelf is disposed around the opening. The panel is attached to the shelf first mounting surface and to the rib, and is sized to enclose the opening. The panel is a load bearing structure operable to transfer loads to the body and receive loads from the body. 1. An airfoil for a turbine engine , the airfoil comprising:an airfoil body extending between a first side and a second side and between a leading edge and a trailing edge, the body including a shelf that extends around an opening in the first side; anda panel attached to the body, the panel enclosing the opening and engaged with the shelf.2. The airfoil of claim 1 , wherein the shelf comprises an annular shelf surface that extends around the opening and engages the panel.3. The airfoil of claim 2 , wherein the shelf surface is substantially parallel with an outer surface of the body at the first side.4. The airfoil of claim 1 , further comprising a base adapted to attach the body to a rotor hub of the turbine engine claim 1 , wherein the body further extends from the base to a tip.5. The airfoil of claim 4 , wherein the shelf extends proximate the leading edge claim 4 , the trailing edge claim 4 , the base and the tip.6. The airfoil of claim 1 , wherein the panel comprises a load bearing structure adapted to transfer loads to the body and receive loads from the body.7. The airfoil of claim 1 , wherein the body includes a rib disposed within the opening and engaged with the panel.8. The airfoil of claim 7 , wherein the panel is ...

Heating Fan

A heating fan has a fastening means for fastening of said fan to a first surface. A stem connects the fastening means to a hub. The hub has a motor unit. The motor unit rotates a plurality of wings extending radially from said hub. The wings have upper and lower surfaces. Each wing has a leading edge and a trailing edge, defining upper and lower surfaces therebetween. At least one wing has electrical heating means integrated or connected to the upper surface. At least more than half the length of the leading and/or trailing edge is bent downwards relative to the surface of the wing. 1. Heating fan , where said fan comprises fastening means for fastening of said fan to a first surface , a stem connecting the fastening means to a hub , where said hub comprises a motor unit , where said motor unit rotates a plurality of wings extending radially from said hub , such that the wings have an upper and a lower surface , wherein each wing has a leading and trailing edge , defining upper and lower surfaces there between where at least one wing has electrical heating means integrated or connected to said upper surface , where at least more than half the length of the leading and/or trailing edge is bent relative to the surface of the wing.2. Heating fan according to wherein the wing profile in a cross section perpendicular to the radial extent of the wing relative to the hub claim 1 , comprises three distinct substantially linear sections claim 1 , a first linear section extending between 5 to 25 mm from the leading edge and a second linear section extending between 5 to 25 mm from the trailing edge claim 1 , and a third linear section arranged between the first and second sections claim 1 , where said third linear section has an extent between 35 and 110 mm claim 1 , where the first and second linear sections are angled 5° to 60° relative to the third section.3. Heating fan according to wherein the heating means is provided by metal leads having an Ohm-resistance of ...

Impeller for Centrifugal Pumps

An impeller of a centrifugal pump is provided. The impeller includes at least two blades () for conveying media containing solids. An impeller blade leading angle (β) is smaller than 0°. The blade angle (β) increases in a first section () until a value of 0° is reached. In a second section (), another increase occurs until a maximum value is reached. In a third section (), the blade angle (β) decreases again. 14. An impeller for centrifugal pumps having at least two blades () for conveying solids-containing media ,characterized in that{'sub': '1', 'b': 9', '10', '11, 'that the blade entry angle (β) is smaller than 0°, the blade angle β increasing in a first section () until it reaches a value of 0°, then increasing in a second section () up to a maximum value and decreasing in a third section ().'}2. The impeller as claimed in claim 1 , characterized in that the blade entry angle (β) is smaller than −10°.3910. The impeller as claimed in claim 1 , characterized in that the blade angle (β) increases with the same gradient in the first section () and second section ().4910. The impeller as claimed in claim 1 , characterized in that the blade angle (β) increases with a gradient of more than 0.35 in the first section () and/or second section ().511. The impeller as claimed in claim 1 , characterized in that claim 1 , from a reversal point claim 1 , the blade angle (β) decreases in a third section () to the blade exit angle (β).612. The impeller as claimed in claim 5 , characterized in that the blade angle (β) remains constant in a fourth section ().7. The impeller as claimed in claim 1 , characterized in that the impeller is configured as a radial wheel.8. The impeller as claimed in claim 1 , characterized in that the ratio of blade exit radius (R) to the blade entry radius (R) is smaller than 1.5.9612. The impeller as claimed in claim 1 , characterized in that the curvature radius of the blade entry edges () is equal to or smaller than the value of the blade thickness ...

Ni Base Alloy and Gas Turbine Blade and Gas Turbine Utilizing the Same

An Ni base alloy uses GTD-111 as a base to improve high-temperature strength while maintaining the weldability and corrosion resistance and a gas turbine blade utilizes the Ni base alloy. The Ni base alloy contains Al of 2.5 to 3.5%, Co of 1.5 to 5.5%, Cr of 11.8 to 13.8%, Mo of 0.4 to 1.4%, Ta of 3.0 to 5.0%, Ti of 5.1 to 6.1%, W of 3.3 to 4.3%, B of 0.01 to 0.02%, C of 0.08 to 0.12% in mass % and remainder containing Ni and inevitable impurities and does not substantially contain Nb. 1. An Ni base alloy containing Al of 2.5 to 3.5% , Co of 1.5 to 5.5% , Cr of 11.8 to 13.8% , Mo of 0.4 to 1.4% , Ta of 3.0 to 5.0% , Ti of 5.1 to 6.1% , W of 3.3 to 4.3% , B of 0.01 to 0.02% , C of 0.08 to 0.12% in mass % and remainder containing Ni and inevitable impurities.2. An Ni base alloy according to claim 1 , wherein the inevitable impurities contain Nb of 0 to 0.2% or less claim 1 , Hf of 0 to 2.0% or less claim 1 , Re of 0 to 0.5% or less claim 1 , Zr of 0 to 0.05% or less claim 1 , O of 0 to 0.005% or less claim 1 , N of 0 to 0.005% or less claim 1 , Si of 0 to 0.01% or less claim 1 , Mn of 0 to 0.02% or less claim 1 , P of 0 to 0.01% or less and S of 0 to 0.01% or less in mass %.3. An Ni base alloy according to claim 1 , wherein Nb is not substantially contained.4. An Ni base alloy according to claim 1 , wherein the Ni base alloy precipitates γ′-phase having NiAl which is intermetallic compound as representation in γ-phase which is matrix phase.5. An Ni base alloy according to claim 1 , wherein the Ni base alloy of precipitation strengthening type contains Al of 2.7 to 3.3% claim 1 , Co of 3.0 to 4.0% claim 1 , Cr of 12.3 to 13.3% claim 1 , Mo of 0.6 to 1.2% claim 1 , Ta of 3.5 to 4.5% claim 1 , Ti of 5.3 to 5.9% claim 1 , W of 3.6 to 4.0% claim 1 , B of 0.012 to 0.018% and C of 0.09 to 0.11% in mass %.6. A casting product utilizing the Ni base alloy according to .7. A gas turbine blade utilizing the casting product according to .8. A gas turbine utilizing the gas turbine ...

DEVICE FOR PRODUCING, REPAIRING AND/OR REPLACING A COMPONENT BY MEANS OF A POWDER THAT CAN BE SOLIDIFIED BY ENERGY RADIATION, METHOD AND COMPONENT PRODUCED ACCORDING TO SAID METHOD

The invention relates to a device for producing, repairing and/or replacing a component, particularly an aircraft component, by means of a powder that can be solidified by energy radiation of an energy radiation source, characterized in that the device comprises an application unit that is designed such that the powder can be applied onto an uneven surface by means of the application unit. 113.-. (canceled)14. A device for producing , repairing and/or replacing a component by means of a powder that can be solidified by energy radiation from an energy radiation source , wherein the device comprises an application unit which is configured so that it can apply powder to be solidified onto a non-planar surface.15. The device of claim 14 , wherein the application unit is configured so that it can be moved on a path along a contour or a contour profile of the non-planar surface.16. The device of claim 14 , wherein the application unit is present as a cylindrical blade or a cylindrical roll.17. The device of claim 14 , wherein the application unit is present as a scraper with a straight or planar shape.18. The device of claim 14 , wherein a shape of the application unit is adapted to a contour or a contour profile of the non-planar surface to apply a powder layer onto the non-planar surface.19. The device of claim 14 , wherein the application unit is configured so as to be replaceable.20. The device of claim 14 , wherein the application unit is connected to a drive unit for moving the application unit.21. The device of claim 20 , wherein the application unit is movable by the drive unit in one claim 20 , two or all three directions in space.22. The device of claim 20 , wherein the drive unit is connected to a control unit claim 20 , the control unit controlling the drive unit and the application unit connected thereto as a function of a predetermined path profile.23. The device of claim 22 , wherein the control unit comprises an NC controller.24. The device of claim 14 , ...

Multi-Part Modular Airfoil Section and Method of Attachment Between Parts

A fan system includes a motor, a rotatable hub, and a plurality of fan blades. Each of the fan blades includes a substantially rigid spine member, a resilient leading edge member, and a resilient trailing edge member. The leading edge member and trailing edge members are removably coupled with the spine member, such that different leading edge members and different trailing edge members may be chosen to customize the leading and trailing edges of the fan blades. Each fan blade may have more than one type of leading edge member or more than one type of trailing edge member. The leading edge member and trailing edge member may each be coupled with the spine member by urging the leading edge member and trailing edge member in a direction that is substantially perpendicular to the longitudinal axis defined by the spine member. 1. A fan blade , comprising: (i) a first end, wherein the first end is configured to be coupled with a fan hub,', '(ii) a second end, wherein the first fan blade portion has a length extending between the first end of the first fan blade portion and the second end of the first fan blade portion,', '(iii) a trailing edge, and', '(iv) a leading edge engagement portion; and, '(a) a first fan blade portion, wherein the first fan blade portion comprises'} (i) a first end,', '(ii) a second end, wherein the second fan blade portion has a length extending between the first end of the second fan blade portion and the second end of the second fan blade portion,', '(iii) a leading edge, and', '(iv) an engagement portion engaged with the leading edge engagement portion of the first fan blade portion;, '(b) a second fan blade portion secured to the first fan blade portion, wherein the second fan blade portion compriseswherein the first fan blade portion and the second fan blade portion together define an airfoil shape at one or more cross sections of the fan blade, wherein the trailing edge of the first fan blade portion defines a trailing edge of the airfoil ...

Composite Turbomachine Blade with In-Built Root

A turbine engine blade of composite material comprising fiber reinforcement obtained by multilayer weaving of yarns and densified by a matrix. The blade has a first portion constituting an airfoil forming a single piece with at least one second portion constituting a blade root. The fiber reinforcement portions corresponding to the first and second portions of the blade are mutually interleaved, at least in part, with yarns of the second fiber reinforcement portion penetrating into the first fiber reinforcement portion. 115-. (canceled)16. A method of fabricating a turbine engine blade out of composite material comprising fiber reinforcement densified by a matrix , the method comprising:making a fiber blank by multilayer weaving a single piece comprising, in the longitudinal direction of said fiber blank corresponding to the longitudinal direction of the blade to be fabricated: a first set of a plurality of yarn layers that are interlinked to form a first portion of the blank corresponding to an airfoil preform; and a second set of one or more yarn layers that are interlinked at least locally to form at least a second portion of the blank corresponding to a blade root preform; the yarns of the first set of yarn layers not being interlinked with the yarns of the second set of yarn layers, the first set of yarn layers having the yarns of the second set of yarn layers crossing therethrough in the second portion of the blank;shaping the fiber blank to obtain a one-piece fiber preform having said first portion forming an airfoil preform and at least said second portion forming a blade root preform; anddensifying the preform with a matrix in order to obtain a composite material blade having fiber reinforcement constituted by the preform and densified by the matrix, forming a single piece incorporating a blade root.17. The method according to claim 16 , wherein the fiber blank is woven with a second continuous set of yarn layers claim 16 , and the shaping of the fiber ...

IMPELLER FOR A CENTRIFUGAL PUMP

The present invention relates to a centrifugal pump, the impeller of which comprises a shroud (34) with at least one solid and rigid working vane (36), and at least one solid and rigid rear vane (38), the at least one working vane (36) having a leading edge region (46), a trailing edge region (48), a central region (C), a side edge, a pressure face (42) and a suction face (44), the at least one solid and rigid rear vane (38) having a trailing edge region, a side edge, a pressure face and a suction face. The trailing edge region (48) of the at least one working vane (36) is rounded by means of a rounding to have a thickness greater than that in the central region (C). 13638384850444648385038. An impeller for a centrifugal pump , the impeller comprising a hub () with at least one solid and rigid working vane () , the at least one solid and rigid working vane () having a leading edge region () , a trailing edge region () , a central region (C) , a thickness at the central region (C) , a side edge , a pressure face () , and a suction face () , the leading edge region () of the at least one solid and rigid working vane () being provided with a rounding or thickened part having a thickness greater than that in the central region (C) , wherein the trailing edge region () of the at least one solid and rigid working vane () is rounded by means of a rounding to have a thickness greater than that in the central region (C).2504438. The impeller as recited in claim 1 , wherein the rounding at the trailing edge region () is arranged on the pressure face () of the working vane ().3. The impeller as recited in claim 1 , wherein the rounding is mostly circular of its cross section.43850. The impeller as recited in claim 1 , wherein the thickness of the working vane () at its trailing edge region () is of the order of 1 claim 1 ,1*the thickness of the working vane at its central region C.5. The impeller as recited in claim 3 , wherein the rounding has a diameter of at least 1 claim 3 ...

Method of Manufacturing Complex Shaped Component

A method of forming a complex shaped part includes the steps of forming a polymer core by an additive manufacturing process. A metal is plated about surfaces of the polymer core, and the polymer core is removed, leaving hollows within a plate core. Metal powder is deposited within the hollows. An integral blade rotor is also disclosed. 1. A method of forming a complex shaped part including the steps of:(a) forming a polymer core by an additive manufacturing process;(b) plating a metal about surfaces of said polymer core;(c) removing said polymer core leaving hollows within a plating core; and(d) depositing metal powder within said hollows.2. The method as set forth in claim 1 , wherein a consolidation step occurs after the depositing of the metal powder into the hollows.3. The method as set forth in claim 2 , wherein the consolidation process is a hot isostatic pressurization process.4. The method as set forth in claim 1 , wherein said plating metal is a nickel based material.5. The method as set forth in claim 4 , wherein said metal powder is also a nickel based material.6. The method as set forth in claim 1 , wherein said complex shaped component is an integrally bladed rotor claim 1 , and said integrally bladed rotor having a hub and radially outwardly extending airfoils with said hollows being formed in both said hub and said airfoils.7. The method as set forth in claim 1 , wherein said plating occurs utilizing electroplating.8. The method as set forth in claim 1 , wherein said polymer core is removed in a furnace.9. The method of claim 8 , wherein said polymer core is melted claim 8 , disintegrated or evaporated in said furnace.10. The method as set forth in claim 1 , wherein said additive manufacturing process includes one of selective lithography analysis claim 1 , selective laser sintering claim 1 , fusion deposition of material or laminated object manufacturing.11. The method of wherein a computer model of the complex shaped component is utilized to control ...

TIP SHROUDED HIGH ASPECT RATIO COMPRESSOR STAGE

A gas turbine engine compressor stage includes a rotor. Compressor blades are supported by the rotor. The blades include an inner flow path surface each supporting an airfoil that has a chord that extends radially along a span to a tip. A shroud is supported at the tip and provides an outer flow path surface. The shroud provides a noncontiguous ring about the compressor stage. 1. A gas turbine engine compressor stage comprising:a rotor;compressor blades supported by the rotor, the blades include an inner flow path surface each supporting an airfoil that has a chord extending radially along a span to a tip, a shroud supported at the tip and providing an outer flow path surface, wherein the shroud provides a noncontiguous ring about the compressor stage.2. The compressor stage according to claim 1 , wherein the airfoils have an aspect ratio corresponding to the airfoil span to the airfoil chord claim 1 , the shroud extends the full chord claim 1 , wherein the aspect ratio is in a range of 1 to 5.3. The compressor stage according to claim 2 , wherein the aspect ratio is in a range of 1.4 to 3.0.4. The compressor stage according to claim 3 , wherein the aspect ratio is in a range of 1.6 to 2.8.5. The compressor stage according to claim 1 , wherein the shroud includes a sealing structure on a side of the shroud opposite of the outer flow path surface.6. The compressor stage according to claim 5 , wherein the sealing structure has labyrinth seals.7. The compressor stage according to claim 1 , wherein the blades each include a root received in a slot in the rotor.8. The compressor stage according to claim 1 , wherein shroud includes ring segments circumferentially spaced apart from one another and separated by gaps claim 1 , wherein multiple blades share a common ring segment.9. The compressor stage according to claim 1 , wherein blades are integrated with the rotor.10. A gas turbine engine comprising:engine static structure;first and second turbine sections rotatable ...

RADIAL TURBINE

A radial turbine comprising: a housing (), a rotor () mounted within the housing () on a shaft (); an inlet channel () for supplying a defined working medium generally tangentially in respect to the circumference of the rotor (); an outlet channel () located near the centre of the rotor (); the rotor () comprising working cavities () of a generally spiral shape starting generally tangentially to the circumference of the rotor (), for conducting the defined working medium from the inlet channel () to the outlet channel (). The working cavities () of the rotor are of generally rectangular cross-section, which has a width (W), oriented generally parallel to the axis of rotation of the rotor (), longer than the height (H) of said rectangular cross-section, oriented generally parallel to the radius of the rotor (), wherein the height (H) of the cavity cross-section is not greater than six times the thickness of a boundary layer of the defined working medium developed on the internal surface of the wider wall of the cavity () under normal working conditions of the turbine. 1. A radial turbine comprising:a housing;a rotor mounted within the housing on a shaft;an inlet channel for supplying a defined working medium generally tangentially in respect to a circumference of the rotor;an outlet channel located near a centre of the rotor;the rotor comprising working cavities having a spiral shape starting tangentially to a circumference of the rotor, for conducting the defined working medium from the inlet channel to the outlet channel;wherein the working cavities of the rotor have a rectangular cross-section, which has a width oriented parallel to an axis of rotation of the rotor, longer than a height of said rectangular cross-section, oriented generally parallel to a radius of the rotor, wherein the height of the cavity cross-section is not greater than six times a thickness of a boundary layer of the defined working medium developed on an internal surface of a wider wall of ...

Airfoil with Thickened Root and Fan and Engine Incorporating Same

In accordance with one aspect of the disclosure, an airfoil is disclosed. The airfoil may include a platform and a blade extending from the platform. The blade may have a root proximate the platform and a tip radially outward from the platform. The root may have a greater thickness than a cross-section at about a quarter-span of the blade or greater. 1. An airfoil , comprising:a platform; anda blade extending from the platform, where the blade has a root proximate the platform and a tip radially outward from the platform, the root having a greater thickness than a cross-section at about a quarter-span of the blade or greater.2. The airfoil of claim 1 , wherein the root of the blade has a thickness about twenty percent greater than a cross-section at about a quarter-span of the blade or greater.3. The airfoil of claim 1 , wherein the root of the blade includes about twenty-five percent of a radial height of the blade.4. The airfoil of claim 1 , further including a transition zone between the tip and the root of the blade claim 1 , the transition zone being aerodynamically smooth.5. The airfoil of claim 1 , further including a fillet joining the blade with the platform of the airfoil.6. The airfoil of claim 5 , wherein the fillet has a width that varies along an axial length of the blade.7. The airfoil of claim 6 , wherein the blade includes a leading edge claim 6 , a central portion claim 6 , and a trailing edge; the leading edge interacting with incoming airflow before other surfaces of the blade claim 6 , the trailing edge interacting with outgoing airflow claim 6 , and the central portion extending between the leading and trailing edges; and the leading and trailing edges of the blade have a steeper fillet than that of the central edge of the blade.8. A fan of a gas turbine engine claim 6 , comprising:a hub; anda plurality of airfoils radially extending from the hub, each airfoil having a platform and a blade radially extending from the platform, the blade having ...

Geared Turbofan Engine Having a Reduced Number of Fan Blades and Improved Acoustics

A rotor blade comprises an airfoil extending radially from a root section to a tip section and axially from a leading edge to a trailing edge, the leading and trailing edges defining a curvature therebetween. The curvature determines a relative exit angle at a relative span height between the root section and the tip section, based on an incident flow velocity at the leading edge of the airfoil and a rotational velocity at the relative span height. In operation of the rotor blade, the relative exit angle determines a substantially flat exit pressure ratio profile for relative span heights from 75% to 95%, wherein the exit pressure ratio profile is constant within a tolerance of 10% of a maximum value of the exit pressure ratio profile. 1. A rotor blade comprising:a leading edge, a trailing edge, a root section and a tip section; andan airfoil extending radially from the root section to the tip section and axially from the leading edge to the trailing edge, the leading and trailing edges defining a curvature therebetween;wherein, in operation of the rotor blade, the curvature determines a relative exit angle at a span height between the root section and the tip section, based on an incident flow velocity at the leading edge of the airfoil and a rotational velocity at the relative span height; andwherein, in operation of the rotor blade, the relative exit angle determines an exit pressure ratio profile that is substantially constant for relative span heights from 75% to 95%, within a tolerance of 10% of a maximum value of the exit pressure ratio profile.2. The rotor blade of claim 1 , wherein the exit pressure ratio profile is non-decreasing for relative span heights from 50% to 95%.3. The rotor blade of claim 1 , wherein the tolerance is 2% of the maximum value of the exit pressure ratio profile.4. The rotor blade of claim 1 , wherein the exit pressure ratio profile has an absolute value of at least 1.3 for each of the relative span heights from 75% to 95%.5. The ...

HYBRID FAN BLADES FOR JET ENGINES

A fan blade for a jet engine, such as a turbo fan jet engine is disclosed. The fan blade includes a root for coupling the fan blade to a rotating hub. The root is connected to a blade portion or an airfoil. The blade includes a leading edge and a trailing edge. The leading edge includes at least one of a protrusion or a recess along the leading edge to enhance the structural characteristics of the blade portion. To ensure that the recess or recesses and/or protrusion or protrusions do not adversely affect the aerodynamic qualities of the fan blade, the leading edge of the blade portion is covered by a protective sheath. The sheath may be fabricated from a titanium alloy and is typically adhered to the blade portion using an adhesive, such as an epoxy. 1. A fan blade for a jet engine , the fan blade comprising:a root connected to a blade portion, the blade portion including a leading edge and a trailing edge,the leading edge including at least one protrusion along the leading edge, the leading edge and the at least one protrusion being covered by a sheath.2. The fan blade of wherein at least the blade portion is fabricated from an aluminum-lithium alloy.3. The fan blade of wherein at least the blade portion is fabricated from an aluminum alloy.4. The fan blade of wherein the sheath is fabricated from a titanium alloy.5. The fan blade of wherein the leading edge further includes at least one recess claim 1 , the at least one recess being covered by the sheath.6. A fan blade for a jet engine claim 1 , the fan blade comprising:a root connected to a blade portion, the blade portion including a leading edge and a trailing edge,the leading edge including at least one recess along the leading edge, the leading edge and the at least one recess being covered by a sheath.7. The fan blade of wherein at least the blade portion is fabricated from an aluminum-lithium alloy.8. The fan blade of wherein at least the blade portion is fabricated from an aluminum alloy.9. The fan blade ...

Fan Blades with Protective Sheaths and Galvanic Shields

Light weight fan blades for turbofan jet engines are disclosed. The fan blades may be fabricated from an aluminum alloy. To enhance the hardness of the leading edge of the fan blade, a titanium sheath may be attached to the leading edges of the fan blades. To prevent galvanic coupling between the titanium and the aluminum, a polymeric liner may be disposed between the protective titanium sheath and the aluminum fan blade. The liner may be fabricated from a polymer material, such as a polyimide or another high performance polymer.

Gas turbine engine composite vane assembly and method for making same

A gas turbine engine composite vane assembly and method for making same are disclosed. The method includes providing at least two gas turbine engine airfoil composite preform components. The airfoil composite preform components are interlocked with a first locking component so that mating faces of the airfoil composite preform components face each other. A filler material is inserted between the mating surfaces of the airfoil composite preform components.

AIRCRAFT ENGINE COMPONENT WITH LOCALLY TAILORED MATERIALS

A method for making a component according to one embodiment of this disclosure includes modeling a response of the component to operating conditions. The model is then mapped to materials that would influence the response of the component to the operating conditions. A component is then fabricated using the materials such that the component includes a graded composition. 1. A method for making a component , comprising:modeling a response of a component to operating conditions;mapping the model to materials that would influence the response of the component to the operating conditions; andfabricating a component using the mapped materials, the component provided with a graded composition.2. The method as recited in claim 1 , wherein the model provides a response of the component at each of a plurality of regions of the component.3. The method as recited in claim 2 , wherein each of the plurality of regions of the component is mapped to a material that would influence the response of the component to the operating conditions.4. The method as recited in claim 3 , wherein the component is fabricated such that each of the plurality of regions of the component is fabricated using a respective one of the mapped materials.5. The method as recited in claim 1 , wherein the component is an airfoil of a gas turbine engine.6. The method as recited in claim 1 , wherein the component is fabricated using an additive manufacturing process.7. The method as recited in claim 1 , wherein the component is provided with a graded composition such that the component changes in composition in at least one direction.8. The method as recited in claim 1 , wherein the model is the result of a finite element analysis performed by a computing device.9. The method as recited in claim 1 , wherein the model indicates a thermal response of each of a plurality of regions of the component relative to the operating conditions.10. The method as recited in claim 9 , wherein each of the plurality of regions ...

METHOD OF FORMING HYBRID METAL CERAMIC COMPONENTS

A monolithic composite turbine component includes at least one first region of a first material and one second region of a second material formed by solid freeform fabrication (SFF). The first material may be a metal and the second material may be a ceramic or a ceramic matrix composite. Transition regions between the metal region and ceramic region are functionally graded regions to minimize internal stress during temperature fluctuations. 1. A monolithic composite turbine component comprising:at least one first region of a first material;at least one second region of a second material; andat least one functionally graded transition region of the first material and the second material between a first region and a second region of the component2. The component of wherein the first material is a metal.3. The component of wherein the second material is a ceramic or a ceramic matrix composite.4. The component of wherein the component is formed by a solid freeform (SFF) additive manufacturing process.5. The component of wherein the solid freeform (SFF) additive manufacturing process is selective laser melting (SLM).6. The component of wherein the component is a nozzle or a vane7. The component of wherein the component is a rotating airfoil claim 1 , bucket claim 1 , or blade.8. The component of wherein the component is a combustor panel claim 1 , combustor heat shield claim 1 , or combustor fuel nozzle.9. The component of wherein the metal is selected from the group consisting of nickel base claim 2 , iron base claim 2 , cobalt base superalloy claim 2 , titanium claim 2 , and titanium alloy.10. The component of wherein the ceramic is selected from the group consisting of silicon carbide claim 3 , silicon nitride claim 3 , silicon oxynitride claim 3 , and aluminum oxide.11. The component of wherein the ceramic matrix composite is selected from the group consisting of SiC/SiC claim 3 , C/SiC claim 3 , and SiC/Si.12. A method of forming a monolithic composite turbine ...

Gas Turbine Blade Array with Reduced Acoustic Output

In accordance with one aspect of the disclosure, a rotor for a gas turbine engine is disclosed. The rotor may include a rotor disk and a plurality of blade extending radially outward from the rotor disk. At least one of the blades may have a physical nonuniformity. The blades may be distributed about the rotor disk based on any physical nonuniformities of the blades to generate at least one decay-resistant harmonic. 1. A rotor for a gas turbine engine , comprising:a rotor disk; anda plurality of blades extending radially outward from the rotor disk, at least one blade having a physical nonuniformity, the blades being distributed about the rotor disk based on any physical nonuniformities of the blades to generate at least one decay-resistant harmonic.2. The rotor of claim 1 , wherein all the blades cooperate to generate a single dominant decay-resistant harmonic.3. The rotor of claim 1 , wherein each blade has a pitch angle claim 1 , a leading edge position claim 1 , and an angular offset relative a circumferentially adjacent blade claim 1 , the physical nonuniformity of any blade being at least one of the pitch angle claim 1 , the leading edge position claim 1 , and the angular offset.4. The rotor of claim 1 , wherein the blades are distributed to form a plurality of decay-resistant harmonics.5. The rotor of claim 1 , wherein an acoustic liner is positioned proximate the rotor within the gas turbine engine.6. The rotor of claim 5 , wherein the acoustic liner is configured to attenuate an amplitude of a dominant decay-resistant harmonic of the rotor.7. The rotor of claim 5 , wherein the acoustic liner is configured to attenuate an amplitude of a plurality of decay-resistant harmonics.8. The rotor of claim 1 , wherein the rotor is a fan of a gas turbine engine.9. A method of reducing combination tone noise in a rotor of a gas turbine engine claim 1 , comprising:distributing a plurality of blades about a rotor disk of the rotor such that the blades generate a decay- ...

COMPOSITE GAS TURBINE ENGINE COMPONENT WITH LATTICE

A method of assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, positioning an array of airfoils about an outer periphery of a hub, the hub being rotatable about an engine longitudinal axis of a gas turbine engine, the hub including a plurality of flanges, and each one of the airfoils including an airfoil section extending from a root section. An array of platforms are positioned about the outer periphery of the hub, each of the platforms including one or more slots defined by a plurality of platform flanges, and including the one or more slots receiving a respective one of the flanges of the hub. A plurality of retention pins are moved relative to the array of airfoils such each one of the retention pins extends through the flanges of the hub and through the plurality of platform flanges of a respective one of the platforms to mechanically attach a respective one of the platforms and the root section of a respective one of the airfoils to the hub. At least one of the airfoils and the platforms includes a plurality of composite layers that define an internal cavity, and includes a lattice structure in the internal cavity that extends between the plurality of composite layers. 1. A method of assembly for a gas turbine engine comprising:positioning an array of airfoils about an outer periphery of a hub, wherein the hub is rotatable about an engine longitudinal axis of a gas turbine engine, the hub includes a plurality of flanges, and each one of the airfoils includes an airfoil section extending from a root section;positioning an array of platforms about the outer periphery of the hub, wherein each of the platforms includes one or more slots defined by a plurality of platform flanges, and including the one or more slots receiving a respective one of the flanges of the hub;moving a plurality of retention pins relative to the array of airfoils such each one of the retention pins extends through the flanges of ...

LEADING EDGE SHIELD

The invention relates to the field of turbomachine blades, and more specifically to a leading-edge shield () for a turbomachine blade (), said leading-edge shield () including a pressure-side wing () and a suction-side wing (). Each pressure-side wing () and suction-side wing () extends heightwise from a bottom edge () to a top edge () and lengthwise from the leading edge () to a rear edge (), and they are connected together via the leading edge (), and wherein the rear edge () of the suction-side wing () presents at least one shaped segment (S, S) that is shaped in such a manner that the length of the suction-side wing () within this segment is shorter than the length (l, l, l, l, l, l) of the suction-side wing () at each end of this segment (S, S). 1. A leading-edge shield for a turbomachine blade , said leading-edge shield comprising a pressure-side wing and a suction-side wing , each extending heightwise from a bottom edge to a top edge and lengthwise from the leading edge to a respective rear edge , and connected together via the leading edge , and wherein the rear edge of the suction-side wing presents at least one shaped segment that is shaped in such a manner that the length of the suction-side wing within this segment is shorter than the length of the suction-side wing at each end of this segment.232. The leading-edge shield () according to claim 1 , wherein the at least one shaped segment of the rear edge of the suction-side wing is closer to the top edge than to the bottom edge of the suction-side wing.3. The leading-edge shield according to claim 1 , wherein the length of the suction-side wing within the at least one shaped segment is at least 10% shorter than the length of the suction-side wing at each end of the at least one shaped segment.4. The leading-edge shield according to claim 3 , wherein the length of the suction-side wing within the at least one shaped segment is at least 15% shorter than the length of the suction-side wing at each end of the ...

TURBOMACHINE BLADE HAVING A MAXIMUM THICKNESS LAW WITH HIGH FLUTTER MARGIN

The invention relates to a turbomachine rotor blade which is characterized in that:—the ratio between the maximum thickness and the chord at 30% of the height of the blade is between 20% and 42% of the ratio between the maximum thickness and the chord at the blade root,—the ratio between the maximum thickness and the chord at 70% of the height of the blade is between 10% and 30% of the ratio between the maximum thickness and the chord at the blade root,—the ratio between the maximum thickness and the chord at 90% of the height of the blade is between 10% and 30% of the ratio between the maximum thickness and the chord at the blade root,—the ratio between the maximum thickness and the chord at the blade head is between 3% and 21% of the ratio between the maximum thickness and the chord at the blade root. 1. A turbomachine rotor blade comprisinga plurality of sections stacked along an axis Z between a blade root and a blade tip,defining between them the height of the blade, a leading edge,', 'a trailing edge,', 'a pressure side, and', 'a suction side,, 'each section including'} a height between 0% corresponding to the blade root and 100% corresponding to the blade tip,', 'a chord defined by a length of a portion of a line connecting the leading edge and the trailing edge,', 'a maximum thickness defined by a maximum distance between the suction side and the pressure side, and', 'a ratio between the maximum thickness and the chord,, 'each section presenting'}whereina ratio of a section at 30% of the height of the blade is comprised between 20% and 42% of a ratio of a section at the blade root,a ratio of a section at 70% of the height of the blade is comprised between 10% and 30% of the ratio of the section at the blade root,a ratio of a section at 90% of the height of the blade is comprised between 10% and 30% of the ratio of the section at the blade root,a ratio of a section at the blade tip is comprised between 3% and 21% of the ratio of the section at the blade root. ...

INDUCER

The present invention relates to an inducer geometry which can optimize the behavior stability of cavitation in an inducer having a plurality of blades of the same geometry. In the inducer having a plurality of blades of the same geometry, a blade loading at a tip side in a front half of a blade is larger than that in a rear half of the blade; and when a blade angle from a circumferential direction of the inducer is expressed by β(degree) and a meridional distance is expressed by m (mm), an increase rate dβ/dm of the blade angle at the tip side is not less than 0.2 from a blade leading edge to a non-dimensional meridional location of 0.15, and the increase rate dβb/dm of the blade angle at a mid-span is not less than 0.25 from the blade leading edge to the non-dimensional meridional location of 0.15. 1. An inducer having a plurality of blades of the same geometry , characterized in that:a blade loading at a tip side in a front half of a blade is larger than that in a rear half of the blade; and{'sub': b', 'b, 'when a blade angle from a circumferential direction of the inducer is expressed by βb (degree) and a meridional distance is expressed by m (mm), an increase rate dβ/dm of the blade angle at the tip side is not less than 0.2 from a blade leading edge to a non-dimensional meridional location of 0.15, and the increase rate dβ/dm of the blade angle at a mid-span is not less than 0.25 from the blade leading edge to the non-dimensional meridional location of 0.15.'}2. The inducer according to claim 1 , characterized in that:{'sub': b', 'b, 'the increase rate dβ/dm of the blade angle at the tip side is in the range of 0.2 to 2.0 from the blade leading edge to the non-dimensional meridional location of 0.15, and the increase rate dβ/dm of the blade angle at the mid-span is in the range of 0.25 to 2.0 from the blade leading edge to the non-dimensional meridional location of 0.15.'}3. The inducer according to claim 1 , characterized in that:a blade geometry at the tip ...

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

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

GAS TURBINE ENGINE BLADE WITH INCREASED WALL THICKNESS ZONE IN THE TRAILING EDGE-HUB REGION

Airfoil outer wall thickness of a gas turbine engine blade is increased in the zone that is proximate the trailing edge and blade hub by forty to sixty percent (40-60%) greater than comparable greatest wall thickness anywhere else along the trailing edge from outboard that zone all the way to the blade tip. The increased thickness zone includes a transition zone that bridges the respective airfoil outer wall thicknesses proximate the hub and tip of the blade. Some embodiments also incorporate pedestals with compound curve fillets in the increased wall thickness zone. The increased thickness zone reduces blade cracking propensity and enhances service life. 1. A turbine engine blade , comprising:a hub, including a blade platform; an outer wall delimiting a pressure side, a suction side, a leading edge, and a trailing edge on an exterior surface thereof, an airfoil interior on an interior surface thereof, and an outer wall thickness between the respective interior and exterior surfaces;', 'a proximal end of the outer wall coupled to the blade platform from the leading edge to the trailing edge;', 'a distal end of the outer wall defining a blade tip; and', 'airfoil span defined between the proximal and distal ends thereof;, 'an elongated airfoil portion, havinga hub fillet circumscribing and joined to the airfoil outer wall exterior surface at the proximal end thereof, and joined to the blade platform; andan increased airfoil outer wall thickness zone, having a zone proximal end adjoining the blade platform and a zone distal end, wherein the zone outer wall thickness, excluding adjoining hub fillet thickness, is approximately forty to sixty percent (40-60%) greater along the trailing edge for approximately eight to ten percent (8-10%) of total airfoil span most proximate the hub than comparable greatest outer wall thickness anywhere else along the trailing edge from outboard the zone distal end all the way to the blade tip, with outer wall thickness at the zone distal ...

COMPOSITE AIRFOIL METAL LEADING EDGE ASSEMBLY

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

Frangible Sheath for a Fan Blade of a Gas Turbine Engine

In accordance with one aspect of the disclosure, a rotor for a gas turbine engine is disclosed. The rotor may include a rotor disk and a blade extending radially outward from the rotor disk. The blade may have a leading edge and a sheath may extend along the leading edge of the blade. The sheath may be formed from a brittle material. 1. A rotor for a gas turbine engine , comprising:a rotor disk; anda blade extending radially outward from the rotor disk, the blade having a leading edge; anda sheath extending along the leading edge of the blade, the sheath being formed from a brittle material.2. The rotor of claim 1 , wherein an adhesive joins the sheath to the leading edge of the blade.3. The rotor of claim 1 , wherein the brittle material is an aluminum alloy.4. The rotor of claim 1 , wherein the brittle material is a fiber-reinforced composite.5. The rotor of claim 1 , wherein the rotor is a fan of the gas turbine engine.6. The rotor of claim 1 , where the blade is made of aluminum.7. A gas turbine engine claim 1 , comprising:a fan section having a disk and a plurality of blades extending radially outward from the disk, each blade having a leading edge and a sheath extending along the leading edge of the blade, the sheath being formed from a brittle material;a compressor section downstream of the fan section;a combustor downstream of the compressor section; anda turbine section downstream of the combustor.8. The gas turbine engine of claim 7 , wherein an adhesive joins the sheath to the leading edge of the blade.9. The gas turbine engine of claim 7 , wherein the brittle material is an aluminum alloy.10. The gas turbine engine of claim 7 , wherein the brittle material is a fiber-reinforced composite.11. The gas turbine engine of claim 6 , wherein the blade is made of aluminum.12. A method of protecting a rotor of a gas turbine engine claim 6 , comprising:providing a rotor having a rotor disk and a plurality of blades extending radially outward from the rotor disk, ...

TURBINE ENGINE AND METHOD OF ASSEMBLING

A turbine engine that includes a stationary assembly, and a rotor assembly configured to rotate relative to the stationary assembly. The rotor assembly includes a plurality of unitary turbine and fan blades. Each unitary turbine and fan blade includes a single turbine airfoil, a single fan airfoil positioned radially outward from the single turbine airfoil, and a midspan shroud segment defined between the single turbine airfoil and the single fan airfoil. 1. A turbine engine comprising:a stationary assembly; and 'a plurality of unitary turbine and fan blades, wherein each unitary turbine and fan blade comprises a single turbine airfoil, a single fan airfoil positioned radially outward from said single turbine airfoil, and a midspan shroud segment extending between said single turbine airfoil and said single fan airfoil.', 'a rotor assembly configured to rotate relative to said stationary assembly, said rotor assembly comprising2. The turbine engine in accordance with claim 1 , wherein said single turbine airfoil is fabricated from a first material and said single fan airfoil is fabricated from a second material different from the first material.3. The turbine engine in accordance with claim 2 , wherein said midspan shroud segment is fabricated from a combination of the first material and the second material.4. The turbine engine in accordance with claim 1 , wherein said each unitary turbine and fan blade comprises at least one reinforcing support member extending through said single turbine airfoil claim 1 , across said midspan shroud segment claim 1 , and through said single fan airfoil.5. The turbine engine in accordance with claim 4 , wherein said at least one reinforcing support member extends longitudinally along thickest portions of said single turbine airfoil and said single fan airfoil.6. The turbine engine in accordance with claim 4 , wherein said at least one reinforcing support member comprises a first reinforcing support member and a second reinforcing ...

Airfoil for a rotary machine including a propellor assembly

In some embodiments, an airfoil comprises a proximal end; a distal end opposite said proximal end; a distal portion extending adjacent said distal end; an edge extending between said proximal end and said distal end; and a surface extending between said proximal end and said distal end, said edge and said surface defining a sweep and a cahedral through said distal portion, wherein the distal portion extends over an acoustically active portion of the airfoil.