УСТРОЙСТВО КОНРОЛЯ ВИБРАЦИИ ЛОПАТОК ТУРБОМАШИН

Устройство контроля вибрации лопаток турбомашин, содержащее периферийный, оборотный, корневой датчики и электронный блок обработки сигналов датчиков, отличающееся тем, что в качестве периферийного, оборотного и корневого датчиков устройство снабжено индуктивными фазогенераторными датчиками контроля приближения токопроводящего объекта.

АМОРТИЗАТОР ДЛЯ ЛОПАТОК ГАЗОТУРБИННОГО ДВИГАТЕЛЯ, РОТОР ГАЗОТУРБИННОГО ДВИГАТЕЛЯ (ВАРИАНТЫ), КОМПРЕССОР ГАЗОТУРБИННОГО ДВИГАТЕЛЯ (ВАРИАНТЫ) И ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ (ВАРИАНТЫ)

Амортизатор для лопаток ротора компрессора газотурбинного двигателя. Конструкция амортизатора приспособлена для размещения между нижней гранью платформ двух смежных лопаток газотурбинного двигателя и ободом диска ротора, на котором установлены лопатки. Амортизатор содержит инерционный груз (11), основание (13), конструкция которого обеспечивает возможность его опоры на указанный обод, и пружину (12). Пружина соединяет инерционный груз с основанием. По меньшей мере инерционный груз (11) выполнен из композитного материала. Амортизатор содержит, но меньшей мере, на одном свободном торце основания или инерционного груза отрезок тонкой пластинки (14, 15), образующий упор или фиксирующий крючок, который оказывается прижатым, либо (14) к радиальному ребру жесткости лопатки, либо (15) к заднему краю обода диска. Обеспечивается возможность устанавливать амортизатор в труднодоступные места и производить его техническое обслуживание по месту установки при одновременном ограничении трения контактирующих ...

Демпфер для роторов турбомашин

Способ динамического уравновешивания на рабочей скорости роторов турбомашин

Паровой щит ступени турбины

Способ подбора лопаток турбин и компрессоров и установка для его осуществления

Устройство для вибрационного испытания лопатки турбомашины

Устройство для наладки приборов бесконтактного измерения колебаний лопаток турбомашины

DICHTUNGSTEIL, INSBESONDERE DICHTUNGSRING, FUER EIN GASTURBINENTRIEBWERK

A vibration damping system

A vibration damping system 8 wherein the system 8 comprises a magnetism generating medium 12 and a magnetism energy dissipating medium 16 whereby, in use, vibration of the magnetism generating medium 12 generates a magnetic field, the magnetism generating medium 12 and the magnetism energy dissipating medium 16 being so disposed with respect to each other that the magnetic field is then dissipated by the magnetism energy dissipating medium 16 thereby damping the vibrations of the magnetism generating medium 12.

A turbomachine blade

A turbomachine blade comprises an aerofoil portion 64 extending between a tip 66 and a root 62, and having a pressure surface (78, fig 6a) and a suction surface 80. At least one passageway (76, fig 7) extends through the aerofoil portion from an inlet 72 in the pressure surface to an outlet 74 in the suction surface. The outlet is located in a section of the aerofoil portion extending between the tip and 50% of the length of the aerofoil portion from the tip. The outlet may be between the tip and 25% of the length of the aerofoil portion from the tip. The inlet and outlet may be located at different points along the chord length of the aerofoil portion. The passageway may extend in a direction from a leading edge 68 to a trailing edge 70. A plurality of passageways may extend through the aerofoil portion, and may have a common inlet and a plurality of outlets. The blade may be a fan blade. A gas turbine engine has a rotor comprising a plurality of the blades.

APPARATUS AND METHOD FOR OPERATING A SUBSEA COMPRESSION SYSTEM IN A WELL STREAM

Method for operating a subsea compression system and a subsea compression system in a well stream are disclosed. The subsea compression system comprising a compressor (1) and an electromotor driven pump (2), wherein the compressor is operable for compression of gas and the pump is operable for pumping liquid from the compression system. Compressed gas is recycled from the compressor downstream side to the upstream side of the compressor via a turbo-expander unit (10) which is drivingly connected to a generator (17). The generator is configured to supply charging current to a battery (18) in response to recycling of compressed gas via the turbo-expander unit, wherein the pump is operable on electrical power that is supplied from the battery. Fig. 3 ...

COOLED ROTOR BLADE WITH VIBRATION DAMPING DEVICE

A rotor blade for a rotor assembly is provided that includes a root, an airfoil, and a damper. The airfoil has a length, a base, a tip, a first side wall, a second side wall and at least one cavity. The length extends the base and the tip. The at least one cavity is disposed between the side walls and the channel is defined by a first wall portion and a second wall portion. The damper, which is selectively received within the channel includes a first bearing surface, a second bearing surface, a forward surface and an aft surface, all of which extend lengthwise. At least one of the surfaces is shaped to form a lengthwise extending passage within the channel. The passage has a flow direction oriented along the length of the at least one surface to permit cooling air travel along the at least one surface in a lengthwise direction. According to one aspect of the present invention, the damper has an arcuate lengthwise extending centerline.

Schaufelanordnung für die Leiträder und Laufräder an Turbinen.

Methods and systems for monitoring the performance of rotor blades.

Die Erfindung betrifft ein System und ein Verfahren zur Überwachung der Funktionstüchtigkeit eines Rotors. Das System enthält ein Verarbeitungssubsystem, das eine Messmatrix auf der Basis mehrerer erster Resonanzfrequenz-Delta-Ankunftszeitvektoren, die einer Laufschaufel und einer ersten Sensorvorrichtung entsprechen, und mehrerer zweiter Resonanzfrequenz-Delta-Ankunftszeitvektoren generiert (709), der Laufschaufel und einer zweiten Sensorvorrichtung entsprechen, eine Resonanzmatrix auf der Basis der Messmatrix derart generiert (708), dass Einträge in der Resonanzmatrix im Wesentlichen linear unkorreliert und linear unabhängig sind, und ein Resonanzsignal unter Verwendung eines ersten Teilsatzes der Einträge der Resonanzmatrix generiert (712), wobei das Resonanzsignal im Wesentlichen gemeinsame Beobachtungen und Komponenten der mehreren ersten Resonanzfrequenz-Delta-Ankunftszeitvektoren und der mehreren zweiten Resonanzfrequenz-Delta-Ankunftszeitvektoren aufweist.

Methods and systems for monitoring the performance of rotor blades.

Die Erfindung betrifft ein System und ein Verfahren zur Überwachung der Funktionstüchtigkeit eines Rotors. Das System enthält ein Verarbeitungssubsystem, das eine Messmatrix auf der Basis mehrerer erster Resonanzfrequenz-Delta-Ankunftszeitvektoren, die einer Laufschaufel und einer ersten Sensorvorrichtung entsprechen, und mehrerer zweiter Resonanzfrequenz-Delta-Ankunftszeitvektoren generiert (709), der Laufschaufel und einer zweiten Sensorvorrichtung entsprechen, eine Resonanzmatrix auf der Basis der Messmatrix derart generiert (708), dass Einträge in der Resonanzmatrix im Wesentlichen linear unkorreliert und linear unabhängig sind, und ein Resonanzsignal unter Verwendung eines ersten Teilsatzes der Einträge der Resonanzmatrix generiert (712), wobei das Resonanzsignal im Wesentlichen gemeinsame Beobachtungen und Komponenten der mehreren ersten Resonanzfrequenz-Delta-Ankunftszeitvektoren und der mehreren zweiten Resonanzfrequenz-Delta-Ankunftszeitvektoren aufweist.

Methods and systems for monitoring the performance of rotor blades.

Die Erfindung betrifft ein System und ein Verfahren zur Überwachung der Funktionstüchtigkeit eines Rotors. Das System enthält ein Verarbeitungssubsystem, das mehrere Frequenzspitzenwerte, die zwei oder mehreren jeweiligen Signalfenstern entsprechen, durch iteratives Verschieben der zwei oder mehreren Signalfenster entlang von Delta-Ankunftszeitsignalen (206), die einer Laufschaufel in dem Rotor entsprechen, generiert, einen oder mehrere Resonanzfrequenz-Rotordrehzahlbereiche der Laufschaufel durch Identifizierung von Rotordrehzahlen, die einem Teilsatz der mehreren Frequenzspitzenwerte entsprechen, bestimmt (214) und die Laufschaufel überwacht, um eines oder mehrere Defekte in der Laufschaufel während der Resonanzfrequenz-Rotordrehzahlbereiche festzustellen.

Methods and systems for monitoring the performance of rotor blades.

Die Erfindung betrifft ein System und ein Verfahren zur Überwachung der Funktionstüchtigkeit eines Rotors. Das System enthält ein Verarbeitungssubsystem, das mehrere Frequenzspitzenwerte, die zwei oder mehreren jeweiligen Signalfenstern entsprechen, durch iteratives Verschieben der zwei oder mehreren Signalfenster entlang von Delta-Ankunftszeitsignalen (206), die einer Laufschaufel in dem Rotor entsprechen, generiert, einen oder mehrere Resonanzfrequenz-Rotordrehzahlbereiche der Laufschaufel durch Identifizierung von Rotordrehzahlen, die einem Teilsatz der mehreren Frequenzspitzenwerte entsprechen, bestimmt (214) und die Laufschaufel überwacht, um eines oder mehrere Defekte in der Laufschaufel während der Resonanzfrequenz-Rotordrehzahlbereiche festzustellen.

Fluid-flow machine for a supercharger with oscillation reduction.

Vorgeschlagen wird eine Strömungsmaschine für einen Abgasturbolader, mit einem Gehäuse (1), welches einen Strömungskanal für ein Fluid definiert, das die Strömungsmaschine durchströmt; einer in dem Gehäuse drehbar aufgenommenen Laufradnabe (2); mehreren auf der Nabe angeordneten Laufschaufeln (3) zum Umlenken einer Fluidströmung; und einer Leiteinrichtung mit Leitschaufeln (4) zum Umlenken einer Fluidströmung; und einem Störungsgenerator zur Induzierung einer Störung in der Fluidströmung, die eine durch die Leitschaufeln (4) induzierte Schwingung der Laufschaufeln (3) durch Überlagerung reduziert.

Fluid-flow machine for a supercharger with oscillation reduction.

Vorgeschlagen wird eine Strömungsmaschine für einen Abgasturbolader, mit einem Gehäuse (1), welches einen Strömungskanal für ein Fluid definiert, das die Strömungsmaschine durchströmt; einer in dem Gehäuse drehbar aufgenommenen Laufradnabe (2); mehreren auf der Nabe angeordneten Laufschaufeln (3) zum Umlenken einer Fluidströmung; und einer Leiteinrichtung mit Leitschaufeln (4) zum Umlenken einer Fluidströmung; und einem Störungsgenerator zur Induzierung einer Störung in der Fluidströmung, die eine durch die Leitschaufeln (4) induzierte Schwingung der Laufschaufeln (3) durch Überlagerung reduziert.

SHOCK ABSORBER Of PADDLES AND SEAL FOR TURBINES

SHOCK ABSORBER FOR PADDLES OF TURBOSHAFT ENGINES

La présente invention porte sur un amortisseur pour aubes de turbomachine, agencé pour être logé entre la face inférieure des plateformes de deux aubes adjacentes de turbomachine et la jante du disque de rotor sur lequel les aubes sont montées. Cet amortisseur est caractérisé par le fait qu'il comprend une masselotte (11), une semelle (13) et un ressort (12), le ressort reliant la masselotte à la semelle.

METHOD FOR INTRODUCING A DELIBERATE DETUNING AUBAGEE WHEEL IN A TURBOMACHINE

La présente invention se rapporte à un procédé (100) pour introduire un désaccordage volontaire dans une roue aubagée d'une turbomachine (10), comprenant les étapes suivantes : a) sélectionner un mode propre de vibration de la roue aubagée (23) à k diamètres nodaux, k étant un nombre entier naturel différent de zéro et, lorsque le nombre d'aubes N de la roue aubagée est un nombre pair, différent de N/2 ; b) déterminer le déplacement des aubes sur toute la circonférence de la roue aubagée pour chacune des deux ondes stationnaires de déformation de même fréquence qui combinées génèrent la déformée modale tournante de la roue aubagée au mode propre de vibration sélectionné ; c) à partir du déplacement aube par aube ainsi déterminé pour chacune des deux ondes stationnaires de déformation, déterminer les aubes pour lesquelles un ventre de vibration d'une première desdites ondes stationnaires de déformation correspond à un nœud de vibration de la deuxième onde stationnaire de déformation ; d) ...

DAMPING DEVICE

Turbine vane having insert supports

Damper pin having restoring force effect induced by centrifugal force

VIBRATION DAMPER FOR A TURBOMACHINE ROTOR VANE

A turbomachine rotor, comprising a disk carrying vanes, each vane comprising a blade linked by a platform to a root, recesses being defined between the platforms of the vanes and the disk, and vibration dampers being mounted in at least some of said recesses, each vibration damper comprising a first structural portion (102) configured to be in contact with a platform of which the vibrations are to be dampened, and a second mass portion (104) configured to carry out a function of damping these vibrations, characterised in that the second mass portion is in the form of a powder and the first structural portion is in the form of a box containing said powder.

LATERALLY REINFORCED VARIABLE PITCH ROTOR

A propulsion unit having reduced noise employs a rotor having a plurality of variable pitch blades. A tensioning joint mechanism is carried within each blade and a tension element is engaged by the tension joint mechanism. The tension element is configured to maintain tension between the blades in the rotor and the tensioning joint mechanism adapted to allow variation of pitch of the blades without altering tension in the tension element. 1. A propulsion unit having reduced noise comprising:a rotor having a plurality of variable pitch blades;a tensioning joint mechanism carried within each blade;a tension element engaged by the tension joint mechanism in each blade, said tension element configured to maintain tension between the blades in the rotor and said tensioning joint mechanism adapted to allow variation of pitch of the blades without altering tension in the tension element.2. The propulsion unit as defined in wherein the tension element has an airfoil shape.3. The propulsion unit as defined in wherein the tension element has a plurality of tension segments and an airfoil shaped fairing covering the tension segments.4. The propulsion unit as defined in wherein each blade has a cavity formed therein and the tension joint mechanism is installed in the cavity.5. The propulsion unit as defined in wherein the cavity has an elliptical shape.6. The propulsion unit as defined in wherein the tensioning joint mechanism comprises:a support bearing mounted to allow rotation of the blade relative to an axis for pitch change;a connector extending from the bearing and fixed relative to pitch change of the blade, said tension element attached to said connector.7. The propulsion unit as defined in wherein the tension element incorporates a tension segment and the tension segment is engaged by the connector with a clamping bushing.8. The propulsion unit as defined in wherein the support bearing is a ball bearing.9. A reduced noise aerodynamic rotor comprising:a plurality of ...

Angular sector of a stator for a turbine engine compressor, a turbine engine stator, and a turbine engine including such a sector

A stator angular sector for a turbine engine compressor including: an outer shroud and an inner shroud, and at least one vane extending radially between the shrouds. The outer shroud includes first and second mounting mechanisms for mounting the stator angular sector on a casing of the engine, which mechanisms are oriented parallel to the axis in opposite directions and connected together by an intermediate portion. The outer shroud includes at least one axial end portion extending from the intermediate portion with which at least one damper-forming insert is configured to come into contact, such that beyond a given value for amplitude of vibration of the end portion, the damper insert and the end portion are configured to move relative to each other to vary total moving mass moving with the end portion, thereby modifying vibratory behavior of the end portion.

Turbine airfoil incorporating modal frequency response tuning

A turbine airfoil includes an airfoil body and a generally hollow flow displacement element positioned in an interior portion of the airfoil body and extending along a span-wise extent thereof. The flow displacement element defines an inactive cavity therewithin. The flow displacement element is spaced from a pressure side wall and a suction side wall of the airfoil body to respectively define a first near-wall cooling flow channel and a second near-wall cooling flow channel. The flow displacement element includes an outer surface facing the near-wall cooling flow channels and an inner surface facing the inactive cavity. The inner surface facing the inactive cavity includes features configured to influence a mass and/or stiffness of the turbine airfoil, to thereby produce a predetermined modal frequency response of the turbine airfoil.

Control of low volumetric flow instabilities in steam turbines

Configuration (10) of the last stage of a steam turbine where rotor blades (2) rotate encircled by a vane carrier (1), such that a plurality of passages (20) are located in the vane carrier (1), such that a fluid is blown through these passages (20) forming a flow that impinges onto the rotor blades (2), the number of passages (20), the location of the passages (20) in the vane carrier (1) and the velocity of the flow impinging onto the rotor blades (2), being calculated in such a way that rotating flow instabilities in the rotor blades (2) when the steam turbine operates at low volumetric flow conditions are avoided.

Damping ring

Offenbart ist ein Ringelement 1 für eine Turbomaschine, insbesondere für eine Fluggasturbine, mit einem Ringelementgrundkörper 14, der zwei benachbart angeordnete Ringenden 16, 18 aufweist, wobei die Ringenden 16, 18 formschlüssig bezüglich einer Axialebene 26 miteinander verbunden sind. Weiterhin offenbart ist eine Turbomaschine mit zumindest einem derartigen Ringelement 1.

FIBRE-REINFORCED COMPOSITE PART WITH INCREASED VIBRATION RESISTANCE

VIBRATION DAMPER FOR STATOR BLADE ROW OF AXIAL-FLOW TURBO-MACHINE

УЗЕЛ ВЕНТИЛЯТОРНОЙ ЛОПАТКИ С АМОРТИЗАТОРОМ, АМОРТИЗАТОР ВЕНТИЛЯТОРНОЙ ЛОПАТКИ И СПОСОБ КАЛИБРОВКИ АМОРТИЗАТОРА

Узел вентиляторной лопатки газотурбинного двигателя с вентилятором и амортизатором вентиляторной лопатки и способ калибровки амортизатора. Вентиляторная лопатка содержит основание и платформу. Амортизатор вентиляторной лопатки выполнен с возможностью крепления в ложементе, образованном в нижней поверхности платформы. Ложемент содержит входной фланец, перпендикулярный основанию лопатки. Амортизатор содержит переднюю кромку, первая часть которой параллельна входному фланцу ложемента, а вторая часть наклонена относительно упомянутого входного фланца. Амортизатор содержит верхнюю поверхность, включающую верхнюю правую наклонную поверхность и верхнюю левую наклонную поверхность. Способ калибровки амортизатора узла вентиляторной лопатки газотурбинного двигателя с вентилятором и амортизатором вентиляторной лопатки с использованием резервного массового объема, включает следующие операции: определяют относительную массу Mref амортизатора; измеряют эффективную массу Meff амортизатора; отрезают резервный ...

УСИЛЕННАЯ ПРОКЛАДКА ЛОПАТКИ ВЕНТИЛЯТОРА

Прокладка для вставления между хвостом лопатки вентилятора турбореактивного двигателя и нижней частью отсека, в котором размещен этот хвост. Отсек ограничен диском вентилятора. Прокладка имеет металлический элемент жесткости, оснащенный, по меньшей мере, одним наружным элементом, выполненным из эластомерного материала, и содержащий несущую поверхность (134) этого наружного элемента. Несущая поверхность (134) содержит, по меньшей мере, одну волнистую зону (136). Достигается надежное удержание и демпфирование лопатки за счет улучшенного сцепления между элементом жесткости и наружным элементом. 3 н. и 6 з.п. ф-лы, 5 ил.

РАБОЧЕЕ КОЛЕСО КОМПРЕССОРА ТУРБОМАШИНЫ

Рабочее колесо компрессора турбомашины содержит диск с лопатками, расположенными друг за другом по его окружности, установленными с возможностью непосредственного взаимодействия между полками смежных лопаток. По меньшей мере, один демпфирующий элемент установлен под полками смежных лопаток с возможностью контакта с ними. Хвостовики лопаток зафиксированы в кольцевом пазе, выполненном в диске. Демпфирующий элемент выполнен в виде полого цилиндра, ограниченного сверху плоской пластиной с расположенными на ней двумя опорными выступами, а снизу кольцевой обечайкой с фаской. Элемент установлен с возможностью контакта с полками каждой из смежных лопаток через опорные выступы и с возможностью контакта поверхностей кольцевой обечайки с боковой поверхностью и/или днищем кольцевого паза. По бокам плоской пластины элемента выполнены срезы, параллельные, и/или срезы, перпендикулярные опорным выступам для фиксирования демпфирующего элемента в кольцевом пазе и между хвостовиками лопаток. Позволяет за ...

РОТОР ОСЕВОЙ МНОГОСТУПЕНЧАТОЙ ТУРБОМАШИНЫ

Ротор осевой многоступенчатой турбомашины, где каждая ступень содержит рабочее колесо, включающее диск с лопаточным венцом, отличающийся тем, что для любой пары рабочих колес ротора выполняется условие - значения частот колебаний лопаток совместно с диском по первой изгибной и первой крутильной формам колебаний для каждой пары рабочих колес во всем эксплуатационном диапазоне частоты вращения ротора, сравниваемые в едином масштабе, различны для каждой из диаметральных форм колебаний.

УСТРОЙСТВО КОНТРОЛЯ АМПЛИТУДЫ КОЛЕБАНИЙ БАНДАЖИРОВАННЫХ РАБОЧИХ ЛОПАТОК ТУРБОМАШИН ДИСКРЕТНО-ФАЗОВЫМ МЕТОДОМ

Устройство контроля амплитуды колебаний бандажированных рабочих лопаток турбомашин дискретно-фазовым методом, содержащее периферийный, корневой, оборотный датчики, связанные с регистрационным и анализирующим блоками, и возбудители периферийного датчика, размещенные на бандажных полках, отличающееся тем, что каждый возбудитель выполнен в виде линейного выступа или паза на внешней поверхности бандажной полки, а периферийный датчик выполнен в виде индуктивного фазогенераторного датчика.

КОНДЕНСАТОР ПАРОТУРБИННОЙ УСТАНОВКИ

... 1. Конденсатор паротурбинной установки, установленный с возможностью температурных вертикальных перемещений, например, на пружинных опорах, содержащий корпус конденсатора, по крайней мере, один соединительный патрубок, сообщающий корпус конденсатора с выхлопной частью турбины, отличающийся тем, что соединительный патрубок снабжен линзовыми компенсаторами, разнесенными по высоте патрубка, стенки патрубка над верхним компенсатором и под нижним компенсатором соединены изнутри с возможностью относительного перемещения стенок патрубка вдоль оси турбины жесткими в вертикальном направлении стяжками в виде шарнирно закрепленных на этих стенках жестких стержней или жестко скрепленных с этими стенками гибких пластин. 2. Конденсатор паротурбинной установки по п. 1, отличающийся тем, что соединительный патрубок снабжен, по крайней мере, одной распорной решеткой стержней, расположенной между стенками соединительного патрубка, размещенными между линзовыми компенсаторами. 3. Конденсатор паротурбинной ...

УСТРОЙСТВО ДЛЯ ЗАМЕРА АМПЛИТУД КОЛЕБАНИЙ РАБОЧИХ ЛОПАТОК ТУРБОМАШИНЫ ДИСКРЕТНО-ФАЗОВЫМ МЕТОДОМ

Устройство для замера амплитуд колебаний лопаток турбомашины дискретно-фазовым методом. Устройство содержит регистрационную и анализирующую аппаратуру МИК, периферийный, корневой датчики и датчик оборотной частоты. В этом устройстве сердечник периферийного датчика выполнен из ферромагнитного материала, поперечное сечение сердечника имеет форму вытянутого прямоугольника с отношением сторон 10 ≥ в/а ≥ 3, где а и в - соответственно ширина и длина поперечного сечения сердечника. Изобретение позволяет производить замеры колебаний рабочих лопаток, закрытых бондажными полками. При этом для замера тангенциальной и аксиальной составляющей амплитуды колебаний угол 30-60o между минимальной осью инерции сердечника периферийного датчика и осью турбины выбирается в диапазоне 30-60o а для замера только тангенциальной составляющей амплитуды угол 30-60o установки сердечника выбирается равным 0.

Электроннолучевое устройство для измерения параметров вибраций лопаток рабочего колеса турбомашины

Verfahren zum Auftragen eines Dämpfermaterials mit eingeschlossener Schicht

Gebläse mit aktiver Schwingungsdämpfung einer Rotorstufe

Gasturbinenmaschine

Eine Gasturbinenmaschine umfasst mindestens einen Schaufelcluster (10), in dem mehrere Schaufelblätter (12) an einem gemeinsamen Innenring (14) und/oder Außenring (16) angeordnet sind. Der Schaufelcluster (10) weist wenigstens einen Hohlraum (18) auf, wobei im Inneren des Hohlraums (18) mehrere Dämpfungskörper (20, 22, 24) vorgesehen sind, die sich unabhängig voneinander relativ zu den Wänden des Hohlraums (18) und relativ zueinander bewegen können.

Schwingungsdämpfer für rotierende Teile

Die vorgeschlagene Erfindung bezieht sich auf ein rotierendes Bauteil (1) für eine rotierende Maschine, insbesondere für eine Turbinenmaschine, das mit einem Schwingungsdämpfer (25) für Schwingungen des rotierenden Bauteils während des Betriebs der Maschine versehen ist, wobei der Schwingungsdämpfer (25) einen geschlossenen Hohlraum (2) im Körper des rotierenden Bauteils aufweist, wobei in dem geschlossenen Hohlraum (2) ein beweglicher fester Dämpfungskörper (3) sowie ein Dämpfungsmedium (4) untergebracht sind, wobei das Dämpfungsmedium (4) auf einem Metall basiert, das bei der Betriebstemperatur, bei der eine Dämpfung erwünscht ist, in flüssigem Zustand ist. Sie bezieht sich außerdem auf Turbomaschinen, die mit solchen rotierenden Bauteilen ausgestattet sind, sowie auf Verwendungen von Flüssigmetall für solche Schwingungsdämpfer.

Aktive Schwingungsdämpfung für eine Rotorstufe einer Turbomaschine

Turbomachinery blade

A turbomachinery blade 40 comprising a blade 44 and a disk, where the blade comprises pressure and suction sides 46, 48 and a superelastic alloy damping member 50 located between them and in contact with both. Preferably the pressure and suction sides are unconstrained at least at a first end such that at least the first end of the pressure and suction sides are able to move relative to one another. The superelastic alloy may be a titanium alloy and may be a shape memory alloy. The superelastic titanium alloy may comprise any of titanium niobium zirconium tin alloy and titanium niobium tantalum zirconium alloy. Preferably the first and second layers are spaced by a third layer in an axial direction and preferably the damping member is bonded to one or both the pressure and suctions sides.

A vibration damping system and a method of damping vibrations

METHOD OF MEASURING OR MONITORING THE STATE OF OSCILLATION OF THE MOVING BLADES OF AN AXIAL FLOW TURBOMACHINE WITHOUT PHYSICAL CONTACT AND APPARATUS FOR PERFORMING THE METHOD

Structural damper

Cooled rotor blade with vibration damping device

TURBINE BLADE NESTED SEAL DAMPER ASSEMBLY

A turbine blade damper seal assembly includes a seal and a damper that both abut a radially outermost non-gas path surface. The seal is fabricated from a plastically deformable material and nests within a recess of the damper. The damper is fabricated from a rigid material that absorbs vibrational energy generated during operation. The recess within the damper provides for both the damper and the seal to be positioned at the radially outermost non-gas path surface.

VIBRATION CONTROL DEVICE FOR AN AXIAL RETAINING RING FOR THE FAN BLADES OF A TURBOMACHINE

Dispositif d'amortissement des vibrations d'un anneau de rétention axiale des aubes de soufflante (16) d'une turbomachine, ces aubes étant destinées à être montées par leur pied (14) sur un disque rotatif (10) comportant une bride annulaire (18) s'étendant axialement et munie d'une pluralité de créneaux radiaux (20) destinés à venir en contact avec une pluralité de créneaux radiaux complémentaires (32) d'un anneau de rétention (30) destiné à être monté autour de la bride du disque, le dispositif comportant un élément de butée (38) en matériau élastomère destiné à venir se loger axialement entre deux créneaux (20) adjacents de la bride et deux créneaux complémentaires (32) adjacents de l'anneau de rétention et radialement entre la bride (18) du disque rotatif et l'anneau de rétention (30), cet élément de butée présentant des surfaces de contact (40, 44) destinées à venir en contact avec les créneaux adjacents, l'anneau de rétention et la bride du disque rotatif. Figure 1 ...

ROTOR ASSEMBLY

A rotor assembly is disclosed herein, The rotor assembly includes a hub disposed for rotation about an axis of rotation. The rotor assembly also includes at least one blade fixed with the hub for concurrent rotation with the hub. The at least one blade extends along a height axis radially outward from the hub with a root portion proximate to the hub at a first end of the height axis and an airfoil portion extending radially inward from a second end of the height axis opposite the first end. The rotor assembly also includes a sheath extending at least partially around the root portion. The sheath is pivotally engaged with the at least one blade to pivot about the height axis and allowed to weathervane to reduce noise.

AEROFOIL ARRAY FOR A GAS TURBINE WITH ANTI FLUTTERING MEANS

An aerofoil array for a gas turbine system has an inner annular platform (3a) and an outer annular platform (4a), which extend about a longitudinal axis (la) and radially delimit an annular channel (5) for a gas flow; the annular channel houses a plurality of aerofoils, arranged at a substantially constant angular pitch and comprising respective central portions (7a, 7b) and respective ends (8a, 8b) connected to the platforms (3a, 4a); the aerofoils are formed by two series of aerofoils (5a, 5b) having a different geometrical feature in order to intentionally vary the eigenfreguencies and arranged about the longitudinal axis (la) with a sequence that is regularly repeated all along the annular channel (5); even though the external geometry of the aerofoils (5a, 5b) is varied, the cross-sections (9a, 9b) remain unchanged in the central portions (7a, 7b), at any given radius with respect to the longitudinal axis (la).

A METHOD FOR INTRODUCING A VOLUNTARY DESACCORDAGE ON A WHEEL AUBAGEE OF TURBOSHAFT ENGINE COILS AUBAGEE PRESENTING A VOLUNTARY DESACCORDAGE

Amortisseur d'aubes et joint d'étanchéité pour turbines.

L'invention concerne un amortisseur d'aubes et un joint d'étanchéité pour turbines. Un organe d'étanchéité 48 est disposé entre le carter 22 d'un moteur à turbine et plusieurs aubes fixes 24, 26. Cet organe d'étanchéité 48 comporte une section centrale flexible 52 qui est comprimée dans la position d'installation. L'invention est utilisée pour améliorer l'étanchéité à l'air extérieur et amortir en même temps les vibrations des aubes fixes dans une turbine.

ENHANCED BUCKET VIBRATION SYSTEM

GAS TURBINE AIRFOIL CLOCKING

The first stage of vanes (16) and second stage of vanes (24) each contain the same number of vanes. The second stage of vanes are located such that the wake flow (38) from the first stage of vanes falls on or near the leading edge, after passing through the stage of rotating blades.

GAS TURBINE MACHINE COMPRISING A DAMPED BLADE CLUSTER

A gas turbine machine comprises at least one blade cluster (10), in which several blades (12) are arranged on a common inner ring (14) and/or outer ring (16). The blade cluster (10) has at least one hollow space (18), inside which several damping elements (20, 22, 24) are provided that can be independently moved relative to the walls of the hollow space (18) and relative to one another.

Damper with varying thickness for a blade

A blade for a gas turbine engine. The blade having: a root; a platform located between the root and the blade, wherein the platform defines a cavity; a damper seal received in the cavity, the damper seal having a main body portion that extends along a major axis of the damper seal between a first end portion and an opposing second end portion of the damper seal, the first end portion and the second end portion each extend towards the root when the damper seal is located in the cavity and wherein the damper seal has a variable thickness along at least a portion of a minor axis of the damper seal that extends between opposite peripheral edges of the main body portion.

Rotor assembly with wear member

A rotor assembly having a wear member secured to an inner surface of the outer wall of the flow path. The wear member is made of material abradable by the blades. The wear member is located upstream of the blades. A downstream end of the wear member is abradably shaped by the blades upon rotation. An inner surface of the wear member is directed radially inwardly along a direction of flow in the flow path for deflecting a boundary layer of the flow into the annular blade path. A gas turbine engine and a method reducing tip vortices in a rotor assembly are also discussed.

Structures and methods for forcing coupling of flow fields of adjacent bladed elements of turbomachines, and turbomachines incorporating the same

Turbomachines having close-coupling flow guides (CCFGs) that are designed and configured to closely-couple flow fields of adjacent bladed elements. In some embodiments, the CCFGs may be located in regions extending between the adjacent bladed elements, described herein as coupling avoidance zones, where conventional turbomachine design would suggest no structure should be added. In yet other embodiments, CCFGs are located upstream and/or downstream of rows of blades coupled to the bladed elements, including overlapping one of more of the rows of blades, to improve flow coupling and machine performance. Methods of designing turbomachines to incorporate CCFGs are also provided.

Blade arrangement

A blade arrangement 31 includes an array of radially extending blades 20 , which may for example comprise a fan of a gas turbine engine for an aircraft. The blades 20 are mounted for rotation about a central axis X-X. The blade arrangement 31 further includes a damping arrangement 32 comprising means 34 for inducing an axi-symmetric magnetic field whose axis of symmetry coincides with the central axis X-X of rotation of the blades 20 . The damping arrangement 32 is configured such that when the magnetic field is induced, any movement of the blade 20 other than pure rotation about the central axis results in the magnetic field causing a force to be exerted on the blade 20 , the force resisting such movement. The damping arrangement may be provided with means for inducing the magnetic field only when there is an increased likelihood of vibration of the blades, for example when a foreign body has entered the air intake of the engine.

Rotary machine

There is provided a rotary machine including a rotary shaft configured to rotate around an axis; rotor blades; a casing surrounding the rotor blades radially outside the rotor blades, and in which a recessed portion accommodates tips of the rotor blades; a sealing portion extending from one of a bottom portion of the recessed portion and the tip of the rotor blade, and having a clearance with the other; and a variable breaker installed in the casing and is capable of being displaced between a protrusion position where the variable breaker protrudes into the recessed portion and an accommodation position where the variable breaker is accommodated in the casing.

GROOVED SEAL ARRANGEMENT FOR TURBINE ENGINE

Stator assembly for a gas turbine engine

MODULAR NOZZLE RING FOR A TURBINE STAGE OF A CONTINUOUS FLOW MACHINE

СПОСОБ СНИЖЕНИЯ ВИБРАЦИИ В РАБОЧИХ ЛОПАТКАХ ТУРБОМАШИНЫ

Изобретение предназначено для использования в турбомашиностроении. Способ снижения вибрации в рабочих лопатках турбомашины заключается в том, что проводят тензометрирование лопаток отдельного рабочего колеса на работающей турбомашине, по его результатам определяют наиболее опасную резонансную частоту колебаний лопаток рабочего колеса. Берут два комплекта лопаток, предназначенных для сборки рабочего колеса, устанавливают поочередно каждую лопатку на вибростенд, определяют для каждой значение выявленной при тензометрировании опасной собственной частоты колебаний. Определяют для двух комплектов среднее значение опасной собственной частоты, строят график среднеквадратичного отклонения значений собственных частот колебаний каждой лопатки. По графику определяют значение средней опасной собственной частоты и разделяют все лопатки на четыре равные группы, две из которых располагаются ближе к значению средней опасной собственной частоты, а оставшиеся две дальше. Собирают два новых комплекта лопаток ...

УСТРОЙСТВО ДЕМПФИРОВАНИЯ КОЛЕБАНИЙ РАБОЧИХ КОЛЕС БЛИСКОВОГО ТИПА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Изобретение относится к демпферам для гашения вибраций рабочих колес газотурбинных двигателей, а именно к устройствам демпфирования колебаний рабочих колес. Устройство демпфирования колебаний рабочих колес блискового типа газотурбинного двигателя представляет собой упругое кольцо, установленное с натягом на внутренней поверхности обода блиска. Устройство демпфирования колебаний установлено на цилиндрической или конической поверхности обода блиска и снабжено замковым соединением и упором. Между замковым соединением и упором имеется зазор, при этом замковое соединение выполнено с возможностью расширения в радиальном направлении под действием центробежных сил до уменьшения величины зазора до нулевого значения. Изобретение позволяет уменьшить неравномерность контактного давления и повысить эффективность демпфирования колебаний за счет исключения возможности «закусывания» демпфера при высоких частотах вращения ротора. 2 з.п. ф-лы, 3 ил.

Способ уравновешивания роторов

Устройство для автоматической балансировки

Устройство для определения статических моментов и групповой замены лопаток турбомашин

Schwingungsdämpfungsvorrichtung sowie Verfahren zur Schwingungsdämpfung zur aktiven Dämpfung von Schwingungen eines Bauteils

Die Erfindung bezieht sich auf eine Schwingungsdämpfungsvorrichtung und ein Verfahren zur aktiven Dämpfung von Schwingungen eines Bauteils (1) mit zumindest einem auf das Bauteil (1) aufgebrachten Piezo-Element, dadurch gekennzeichnet, dass die Piezo-Elemente (2a; 2b) mit einer elektrischen Schaltung (3) verbunden sind, welche zur Aufnahme der bei einer Schwingung des Bauteils (1) durch die Verformung eines Sensorelementes induzierten Messsignals, zur Verarbeitung des Messsignals und zur Rückleitung eines phasenversetzten Stroms an die Piezo-Elemente (2a; 2b) ausgebildet sind.

ROTATIONSMASCHINE

Eine Rotationsmaschine umfasst eine Rotationswelle, die so ausgestaltet ist, dass sie sich um eine Achse dreht, und eine Schaufelreihe mit einer Vielzahl von Schaufeln in Intervallen in der Umfangsrichtung der Achse. Jede der Schaufeln enthält ein Faserlaminat (9), das durch Laminieren einer Vielzahl von Faserlagen (11) erhalten wird, und ein Harz, das zur Bildung einer äußeren Form der Schaufel durch Imprägnieren des Faserlaminats (9) verwendet wird. Mindestens zwei der Schaufeln in der Schaufelreihe haben Faserlaminate (9) mit verschiedenen Strukturen.

CHANNEL WALL FOR A BLOWER OF A GAS TURBINE ENGINE

Cooled rotor blade with vibration damping device

TURBOFAN NACELLE ASSEMBLY WITH FLOW DISRUPTOR

... ²A turbofan engine is disclosed which includes a nacelle assembly, having an ²interior ²surface for directing airflow, and a flow disruptor positioned on the interior ²surface ²upstream of the fan, the flow disruptor extending towards the axis a height ²greater than ²the anticipated boundary layer height of the airflow. A turbofan engine which ²includes ²an array of circumferentially disposed flow disruptors extending from a fan ²case inner ²surface is also disclosed. A method of mitigating fan flutter in a gas turbine ²engine by ²generating a circumferential asymmetrically in the airflow, upstream of the ²fan, is also ²described.² ...

BLADE DAMPER AND METHOD FOR MAKING SAME

A turbine engine (10) blade damper (50) for damping vibrations of blades (12) in a turbine engine (10). The blade damper (50) includes an elongate body (52) extending between a forward end (54) and a rearward end (56) opposite the forward end (54). The body (52) is sized and shaped for receipt within a gap (46) formed between adjacent platforms (22) of the blades (12) so the body (52) frictionally engages the adjacent platforms (22) to dampen vibrations of the blades (12) and to prevent air from passing through the corresponding gap (46) during engine (10) operation. The damper (50) also includes a retainer (70) mounted on at least one of the forward and rearward ends (56) of the body (52). The retainer (70) is sized and shaped for receipt within a recess (72) formed in at least one of the adjacent blades (12) to hold the body (52) between the blades (12). The retainer (70) is sized and shaped to prevent air from passing between the blades (12) and the retainer (70) during engine (10) operation ...

ROTOR ASSEMBLY AND ROTATING MACHINE

A rotor assembly provided with: a rotor disc; a plurality of rotor blades secured to the rotor disc and extending radially outward in a radial direction of the rotor disc; and at least one rolling element configured to be able to roll on a curved surface facing inward in the radial direction of the rotor disc.

BLADE DAMPER AND METHOD FOR MAKING SAME

A turbine engine (10) blade damper (50) for damping vibrations of blades (12) in a turbine engine (10). The blade damper (50) includes an elongate body (52) extending between a forward end (54) and a rearward end (56) opposite the forward end (54). The body (52) is sized and shaped for receipt within a gap (46) formed between adjacent platforms (22) of the blades (12) so the body (52) frictionally engages the adjacent platforms (22) to dampen vibrations of the blades (12) and to prevent air from passing through the corresponding gap (46) during engine (10) operation. The damper (50) also includes a retainer (70) mounted on at least one of the forward and rearward ends (56) of the body (52). The retainer (70) is sized and shaped for receipt within a recess (72) formed in at least one of the adjacent blades (12) to hold the body (52) between the blades (12). The retainer (70) is sized and shaped to prevent air from passing between the blades (12) and the retainer (70) during engine (10) operation ...

FAN BLADE ANTI-FRETTING INSERT

DAMPER DISPOSITION MOUNTED BETWEEN ROTOR VANES

Damper disposition (8) mounted between two neighbouring vanes (5) and which includes free inners (22) in housings (21) delimited by a radially external wall (16) slanted outwardly in the direction of the respective vane (5). The centrifugal forces produced by rotation of the disk of the rotor (4) move the inners (22) outside the housings (21) and forcibly apply them to the vanes (5) so as to dampen the vibrations, especially when considerable friction is produced against the outer wall (16). Application for turbo-engines and especially for the large vanes of input blowers.

COIL OF ROTOR OF TURBOSHAFT ENGINE

DAMPING DEVICE

PROCESS AND DEVICE OF MEASUREMENT AND CONTROL WITHOUT CONTACT OF the VIBRATORY STATE OF the MOBILE WINGS Of an AXIAL TURBOSHAFT ENGINE

BLADE OR VANE FOR A TURBOMACHINE AND AXIAL TURBOMACHINE

The present invention relates to a blade ( 100 ) or vane for a turbomachine, having at least one impulse element housing ( 1.1, 1.2 ) with a first impact cavity ( 10 ), in which an impulse element ( 11 ) is arranged with play of movement, wherein the impulse element housing has at least one second impact cavity ( 20 ), which is in alignment with the first impact cavity in a first matrix direction (A) and in which an impulse element ( 21 ) is arranged with play of movement, and has at least one third impact cavity ( 30 ), which is in alignment with the first impact cavity in a second matrix direction (B) crosswise to the first matrix direction and in which an impulse element ( 31 ) is arranged with play of movement.

Turbine blades having damper pin slot features

A turbine blade includes an airfoil that extends radially between a root end and a tip end, a platform coupled to the root end, and a shank that extends radially inwardly from the platform. The shank includes a cover plate. The cover plate includes an outer surface, an opposite inner surface, and a contoured face that at least partially defines a damper pin slot. The contoured face extends from the outer surface to a first blend edge. The cover plate also includes a blended surface that extends from the first blend edge to a second blend edge. The second blend edge intersects with the inner surface.

Turbomachine flow path having circumferentially varying outer periphery

A turbomachine includes an annular flow path section between a plurality of radially extending stator vanes and a plurality of radially extending rotor blades. At least a first portion of the flow path section has a circumferentially varying outer periphery which includes a recess circumferentially alternating with axisymmetric surfaces.

NANOCELLULAR FOAM DAMPER

A machine includes a section that defines a target vibrational mode to dampen and a nanocellular foam damper that includes interconnected ligaments in a cellular structure. The interconnected ligaments have an average ligament size defined with respect to a vibrational loss modulus of the nanocellular foam damper and the target vibrational mode. Also disclosed is a method of damping vibration.

Damped airfoil for a gas turbine engine

An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil section extending in a spanwise direction, extending between a leading edge and a trailing edge in a chordwise direction, and extending in a thickness direction between a pressure side and a suction side. The airfoil section has a main body and a first skin. The main body includes a plurality of ribs defining a plurality of internal channels. The first skin is attached to the main body to enclose the plurality of internal channels such that the main body and the first skin cooperate to define the pressure and suction sides. A damper has at least one layer of damping material sandwiched between the first skin and the plurality of ribs. A method of forming a gas turbine engine component is also disclosed.

WIND TURBINE BLADE COMPRISING ONE OR MORE OSCILLATION DAMPERS

The invention relates to a wind turbine blade comprising one or more oscillation dampers for damping oscillations or vibrations of the wind turbine blade. The first damper parts being rigidly connected to the blade or being a part of the blade. The dampers further comprise second damper parts, wherein the first damper part surfaces and the second damper part surfaces are arranged to move relatively to each other during the oscillations. Even further the dampers comprise a load transferring coupling, coupling the first damper part surfaces and the second damper part surfaces, so that the relative movement results in a oscillation-damping dissipation of kinetic energy. The invention further relates to a wind turbine, an oscillation damper, a method for damping oscillations of a wind turbine blade and use hereof.

DAMPER STACKS FOR TURBOMACHINE ROTOR BLADES

Damper stacks, rotor blades, and turbomachines are provided. A rotor blade includes a main body including a shank and an airfoil extending radially outwardly from the shank. The rotor blade further includes a platform surrounding the main body, the platform comprising a slash face. The rotor blade further includes a damper stack disposed at the slash face and extending generally along an axial direction. The damper stack includes a plurality of damper pins, each of the plurality of damper pins being in contact with a neighboring damper pin. 1. A rotor blade for a turbomachine , the rotor blade comprising:a main body comprising a shank and an airfoil extending radially outwardly from the shank;a platform surrounding the main body, the platform comprising a slash face; anda damper stack disposed at the slash face and extending generally along an axial direction, the damper stack comprising a plurality of damper pins, each of the plurality of damper pins in contact with a neighboring damper pin.2. The rotor blade of claim 1 , wherein the plurality of damper pins comprises a first damper pin and a second damper pin claim 1 , wherein each of the plurality of damper pins extends between a first end and a second end claim 1 , and wherein the first end of the first damper pin contacts the second end of the second damper pin.3. The rotor blade of claim 2 , wherein the first end of the first damper pin has an outward spherical shape claim 2 , and the second end of the second damper pin has an inward spherical shape.4. The rotor blade of claim 2 , wherein a length of each of the plurality of damper pins is defined between the first end and the second end of the damper pin claim 2 , and wherein the length of the first damper pin is different from the length of the second damper pin.5. The rotor blade of claim 1 , wherein the damper stack further comprises a wire extending through each of the plurality of damper pins.6. The rotor blade of claim 1 , wherein a groove is defined in ...

Assembly for turbomachine

A turbomachine assembly includes a casing, first and second rotors, and a damper. The first rotor includes a disk and blades and is movable in rotation relative to the casing. The second rotor is movable relative to the casing around a longitudinal axis. The damper damps a movement of the first rotor relative to the second rotor. The damper includes first to third parts. The first part bears on the first rotor in a first area extending over a first angular sector around the longitudinal axis and applies a first centrifugal force on the first rotor. The second part bears on the first rotor in a second area that is smaller than the first angular sector and extends over a second angular sector around the longitudinal axis. The third part bears on the second rotor and applies a second centrifugal force on the second rotor.

Annular stator blade platform for low pressure turbine in a gas turbine engine

A damper pin for a bucket damper slot in a turbine includes slot insertion ends (12) shaped to fit into the bucket damper slot, and at least a first scallop section (14) formed or machined between the slot insertion ends and shaped to receive a bucket shank pocket radial contour at bucket Hi-C. A second scallop section (16) may also be formed or machined diametrically opposed and anti-symmetrical to the first scallop section between the slot insertion ends.

Rotary machine having non-uniform blade and vane spacing

A system, including a rotary machine including: a stator, a rotor configured to rotate relative to the stator, wherein the rotor comprises a plurality of blades having a non-uniform spacing about a circumference of the rotor.

Ring element and turbomachine having such a ring element

A ring element for a turbomachine, in particular for an aircraft gas turbine, is disclosed. The ring element has a ring element main body that has two adjacently arranged ring ends, the ring ends being connected to one another in a form-locking manner with respect to an axial plane. Also disclosed is a turbomachine having at least one such ring element.

Vibration damping blade for fluid

The vibration damping blade for fluid of the present invention has an integrally formed wedge damper, in which a thickness h(x) at a distance x from an imaginary line outside of an outer edge is h(x)=εx n (where ε is a positive constant, and n is a real number of 1 or more). As a result, it is possible to offer a vibration damping blade for fluid which can be easily manufactured, and which obtains damping effects across a wide range of frequency regions without disturbing the flow of fluid.

Alternate shroud width to provide mistuning on compressor stator clusters

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

BLADE

A blade having a root portion and an aerofoil portion, wherein the aerofoil portion has a tip remote from the root portion, and a leading edge and a trailing edge, and wherein the tip of the aerofoil portion has a set-back portion extending from the leading edge or the trailing edge of the aerofoil portion part way towards the respective other edge and set back from the remainder of the tip of the aerofoil portion towards the root portion. 1. A blade comprising a root portion and an aerofoil portion , wherein the aerofoil portion has a tip remote from the root portion , and a leading and a trailing edge , and wherein the tip of the aerofoil portion has a set-back portion extending from the leading edge or the trailing edge of the aerofoil portion part way towards the respective other edge and set back from the remainder of the tip of the aerofoil portion towards the root portion.2. A blade as claimed in claim 1 , wherein the set-back portion in the tip is serrated.3. A blade as claimed in claim 2 , wherein the serrated set-back portion has shaped serration slots which extend not aligned with the circumferential direction of motion of the tip when the blade is rotating in use in a fan.4. A blade as claimed in claim 3 , wherein the serration slots are approximately perpendicular to the surface of the tip which is flow-washed when the blade is rotating in use in a fan.5. A blade as claimed in claim 3 , wherein the serration slots are 2 mm deep.6. A fan having a plurality of blades as claimed in and a fan casing around the tips of the blades claim 1 , wherein as a result of the set-back portions in the tips the tip clearance area between tips and fan casing is changed by at least 1% of fan area as compared with a case in which the set-back portions in the tips were omitted.7. (canceled)8. A blade as claimed in claim 4 , wherein the serration slots are 2 mm deep.9. A fan having a plurality of blades as claimed in and a fan casing around the tips of the blades claim 2 , ...

ROTOR BEHAVIOUR DETERMINATION

A method of resolving vibration behaviour of a rotor assembly having a plurality of rotor blades. The method including generating a computational model of the rotor by discretization of the rotor geometry and assigning parameter values representative of a plurality of physical characteristics of the rotor. An artificial parameter feature is applied to each rotor blade such that the parameter values for the artificial parameter feature substantially depart from the physical characteristics of the rotor. A vibration response is calculated for the rotor or blade thereof for the applied artificial parameter features and compared to a predetermined vibration response so as to determine a value of the artificial parameter feature which results in a calculated vibration response that substantially matches the predetermined vibration response. 1. A method of resolving vibration behaviour of a rotor assembly having a plurality of components , the method comprising:obtaining a measured vibration response for the rotor;generating a computational model of the rotor by discretization of the rotor geometry and assigning parameter values representative of a plurality of physical characteristics of the rotor;applying a localised artificial parameter feature to each component of the rotor such that the parameter value for said artificial parameter for each component departs fromthe physical characteristics of said rotor;calculating a vibration response for the rotor or component thereof for said applied artificial parameter features;comparing the calculated vibration response against the measured vibration response so as to determine a value of the artificial parameter feature which results in a calculated vibration response that substantially matches the measured vibration response; and,modifying one or more components of the rotor in dependence on the determined artificial parameter values.2. A method according to claim 1 , wherein the artificial parameter feature comprises an ...

BLADE ARRANGEMENT AND ASSOCIATED GAS TURBINE

A blade arrangement with a rotor and a plurality of blades which are distributed in a ring along the circumference of the rotor is provided. Two immediately adjacent blades of the ring form a blade pair, between the blades of which a damping element is arranged, and wherein the respective damping element comes into contact with the two blades of the blade pair assigned to them during a rotation of the rotor about a rotor axis as a result of a centrifugal force which acts in the radial direction. In order to bring about frequency detuning of the oscillation properties of blades, as a result of which machining of the turbine blade becomes unnecessary, it is proposed that the blade ring has at least two blade pairs with different damping elements. 14-. (canceled)5. A blade arrangement , comprising:a rotor; anda plurality of blades which are distributed in a ring along the circumference of the rotor and comprise respectively in succession a blade root, a platform and a blade airfoil, wherein two immediately adjacent blades of the ring form a blade pair and are assigned at least one damping element,wherein each respective damping element comes into contact with the platforms of the two blades of the blade pair assigned to it during a rotation of the rotor about a rotor axis as a result of a centrifugal force acting in the radial direction,wherein, for adjusting the natural frequencies of the blades, the blade ring includes at least two blade pairs with different damping elements and each blade of the ring is assigned to two blade pairs and wherein two or more groups of blade pairs are provided, within each group the damping elements are in each case identical and the damping elements differ from group to group,wherein a majority of the blade pairs or each blade pair of the first group has an adjacent blade pair of the first group and an adjacent blade pair of the second group, orwherein a first group, a second group and a third group of blade pairs are provided wherein a ...

Unknown

A method for vibration damping of at least one blade of a turbomachine, wherein initially at least one damping element is arranged on the blade such that it can move in the axial direction, employs a damping element having a larger permeability constant than the blade (μ>μ), and then a magnetic field acting in the radial direction is generated at least temporarily during rotation of a rotor hub of the turbomachine in order to adjust the mass of the damping element in real time. A damping device includes, for example, a ferromagnetic damping element as well as a magnetic field source, and a turbomachine. 11. A method for vibration damping of at least one blade () of a turbomachine comprising the steps:{'b': 18', '1', '1, 'sub': 'rD', 'arranging at least one damping element () at the blade () such that it can move in the axial direction, said damping element having a larger permeability constant (μ) than the blade (); and'}{'b': '26', 'generating a magnetic field () acting in the radial direction at least temporarily during rotation of a rotor hub of a turbomachine.'}226. The method according to claim 1 , wherein the magnetic field () is switched on and off during the rotation.3. The method according to claim 1 , wherein a magnetic field strength is varied as a function of the speed of rotation of the rotor hub.42618. The method according to claim 1 , wherein the magnetic field () acts radially inward on the damping element ().52618. The method according to claim 1 , wherein the magnetic field () acts radially outward on the damping element ().62618. The method according to claim 1 , wherein the magnetic field () rotates together with the damping element ().726. The method according to claim 1 , wherein the magnetic field () is fixed in position.8181. The method according to claim 1 , wherein the damping element () is arranged close to the blade mount when the blade () is braced in a blade mount subject to tolerances.9211812018. A damping device () for vibration ...

Turbine assembly

According to one aspect of the invention, a turbine assembly includes an airfoil extending from a blade and a dovetail located on a lower portion of the blade, wherein the dovetail has a dovetail contact surface. The turbine assembly also includes a member with a slot configured to couple to the airfoil via the dovetail, the slot having a slot contact surface to contact the dovetail contact surface, wherein the dovetail contact surface is reduced by a relief to alter a fundamental frequency of an assembly of the blade and member.

MAGNETICALLY-COUPLED DAMPER FOR TURBOMACHINERY

A system, method, and apparatus for damping vibration in a rotor supported by primary bearings are provided. The system includes a magnetic coupling configured to magnetically engage a rotor supported by one or more primary bearings, and a piston coupled to the magnetic coupling. The system also includes a damper engaging the piston and configured to damp the rotor, wherein the damper is substantially non-load bearing. 1. A damper system for a rotor , comprising:a magnetic coupling configured to magnetically engage a rotor supported by one or more primary bearings;a piston coupled to the magnetic coupling; anda damper engaging the piston and configured to damp the rotor, wherein the damper is substantially non-load bearing.2. The damper system of claim 1 , wherein the damper includes an eddy current damper.3. The damper system of claim 2 , wherein the eddy current damper includes a housing claim 2 , and one of the housing and the piston includes a magnet and the other includes a conductive material claim 2 , such that movement of the piston in the housing induces eddy currents in the conductive material to resist relative movement between the piston and the housing.4. The damper system of claim 1 , wherein the damper includes a dashpot damper.5. The damper system of claim 1 , wherein the damper includes a sealed housing disposed around the rotor claim 1 , wherein the piston and at least a part of the magnetic coupling are disposed in the sealed housing.6. The damper system of claim 5 , wherein the piston defines orifices extending radially therethrough so as to communicate a first space at least partially defined between a radial-outside of the piston and the sealed housing with a second space at least partially defined between a radial inside of the piston and the sealed housing claim 5 , the sealed housing being substantially filled with a damping fluid claim 5 , such that radial movement of the piston forces the damping fluid through the orifices.7. The damper ...

Blade Wedge Attachment

A rotor includes a disk that has slots circumferentially arranged around its periphery. Blades include respective roots that are mounted in respective ones of the slots. The roots are smaller than the slots such that there are circumferential gaps between the roots and circumferential sides of the slots. Wedges are respectively located within the circumferential gaps. The wedges are free floating with regard to the blades and the disk.

Airfoil with break-way, free-floating damper member

An airfoil includes an airfoil body that has a leading edge and a trailing edge and a first sidewall and a second sidewall that is spaced apart from the first sidewall. The first sidewall and the second sidewall join the leading edge and the trailing edge and at least partially define a cavity in the airfoil body. A damper member is enclosed in the cavity and is free-floating within the cavity.

METHOD FOR DAMPING A GAS-TURBINE BLADE, AND VIBRATION DAMPER FOR IMPLEMENTING SAME

A turbine wheel with optimal-mass dampers to dampen a predetermined resonance in context of vibration of a turbine, particularly a low-speed turbine, while assisting in flexibility of adapting to bearing surfaces of recesses of the dampers, by separating mass and flexibility functions by a flexible portion for clamping against the platform, and a mass portion for controlling frictional forces. A damper includes a plate and a counterweight. The plate is stamped from a metal sheet that is substantially thinner than that of the counterweight. The plate includes a wall configured to flexibly contact a platform of the blade of the wheel, while at least partially surrounding a surface of the counterweight. The damper can be used in particular for a wheel of a turbine of a turbine engine, of a fan, or of a BP compressor having mounted blades. 17-. (canceled)8. A damping method for blades mounted on low speed wheel disks of a gas turbine , the turbine including housings recesses under a platform of a blade , configured to receive vibration dampers , the method comprising:carrying out in an independent way a flexible portion clamped against the platform and a mass portion for concentrating efforts so as to direct friction forces against the platform via the clamping action, coupling both parts together in a reversible way, the coupling of both parts being made by surrounding at least partially the mass portion through at least one clamping area of the flexible portion against the platform, the flexible portion being sufficiently flexible to be configured to a required contact level; andinserting the dampers in two parts within the housing recesses being dedicated.9. A vibration damper for implementing the blade damping method according to claim 8 , comprising:a plate and at least one counterweight,the plate being stamped from a metal sheet that is substantially thinner than the counterweight one, andthe plate including a wall configured to flexibly contact at least one ...

Damping means for damping a blade movement of a turbomachine

A damper () for damping a blade movement of a turbomachine (), and to a method for producing the damper (). The damper () has at least one side surface (′) which can be brought into frictional contact with a friction surface of the turbomachine () in order to damp a blade movement. The side surfaces (′) are asymmetrically convex in shape. 111-. (canceled)12. A damper for damping a blade movement of a turbomachine , the damper comprising:at least one side surface intended to damp the blade movement by frictional contact with a friction surface of the turbomachine, the side surface being asymmetrically convex in shape.13. The damper as recited in wherein the side surface has at least two zones of different radii of curvature.14. The damper as recited in wherein a zone of the side surface radially farther away from a rotor axis of the turbomachine has a smaller radius of curvature than a zone of the side surface radially closer to the rotor axis.15. The damper as recited in wherein the damper has a triangular or polygonal shape in a cross section normal to a rotor axis of the turbomachine.16. The damper as recited in further comprising an anti-rotation device.17. The damper as recited in further comprising a fastener for limiting movement of the damper.18. The damper as recited in wherein the fastener limits movement in a direction of a rotor axis of the turbomachine.19. A turbomachine comprising:a rotor;at least one blade; and{'claim-ref': {'@idref': 'CLM-00012', 'claim 12'}, 'the damper as recited in .'}20. The turbomachine as recited in wherein the blade is a rotor blade coupled to the rotor.21. The turbomachine as recited in wherein the blade has an airfoil and a shroud segment at the end of the airfoil distal from the rotor claim 19 , the shroud segment having a pocket at least partially defining a cavity claim 19 , the damper being disposed in the cavity.22. The turbomachine as recited in wherein the cavity is a closed cavity.23. The turbomachine as recited in ...

METHOD FOR THE GENERATIVE PRODUCTION OF A COMPONENT WITH AN INTEGRATED DAMPING ELEMENT FOR A TURBOMACHINE, AND A COMPONENT PRODUCED IN A GENERATIVE MANNER WITH AN INTEGRATED DAMPING ELEMENT FOR A TURBOMACHINE

The invention provides a method for the generative manufacture of a component () with integrated damping for a turbomachine, in particular a gas turbine, having the following method steps: building up the component () in a generative manner, and introducing a damping material () into the component () during the method step of the generative building up of the component (). The invention further provides a component () with integrated damping for a turbomachine, in particular a gas turbine, wherein the component () is built up in a generative manner, and the component () has a damping material () which is introduced into the component () during the generative building up of the component (). The invention further provides a turbomachine, in particular a gas turbine, comprising such a component (). 111.-. (canceled)12. A method for the generative manufacture of a component with integrated damping for a turbomachine , wherein the method comprises building up the component in a generative manner , and introducing a damping material into the component during the generative buildup of the component.13. The method of claim 12 , wherein an unsolidified base material of the component is introduced as damping material.14. The method of claim 12 , wherein the component is built up at least in portions thereof with a cavity.15. The method of claim 14 , wherein the damping material is introduced into the cavity.16. The method of claim 14 , wherein during the generative buildup of the component an integrated supporting and/or cooling structure is built up in the cavity.17. The method of claim 15 , wherein during the generative buildup of the component an integrated supporting and/or cooling structure is built up in the cavity.18. The method of claim 17 , wherein the supporting structure is built up at least in portions thereof with a cavity into which the damping material is introduced.19. The method of claim 14 , wherein during the generative buildup of the component a cooling ...

Vibration damper

Vibration damping is important with regard to such components as hollow turbine blades in gas turbine engines. Traditionally damping has occurred through damping elements secured at the root or tip of such blades. Such damping is not optimised and results in potential problems with wear in operational life. By providing a tube of deformable material which can be located within a hollow cavity it is possible to provide an element which through friction engagement can absorb vibration energy and therefore damp such vibration. The tube incorporates a number of cuts and/or grooves in an appropriate pattern in order to define a deformation profile once the tube is expanded in location. The tube is secured in position internally upon an expandable element which is typically an inflatable device. Once in position the tube is retained in its expanded deformable profile and the engagement between the tube and the hollow cavity wall surface results in energy absorption through vibration episodes. It is also possible to provide a tube formed from a shape memory alloy which will expand of its own right in location to engage the hollow cavity wall surfaces for energy absorption during vibration episodes.

System and Method for Vibration Isolation

In accordance with one embodiment of the present disclosure, a system includes a first housing, a second housing, a seal, and a spring system. The first housing includes a first volume of fluid. The first housing is capable of connecting to a first element and to a second element, and is also capable of reducing an amount of movement transferred from the first element to the second element. The second housing is connected to the first housing. The second housing includes a second volume of fluid and a volume of gas. The first volume of fluid is in fluid communication with the second volume of fluid. The seal is capable of separating the second volume of fluid from the volume of gas. The spring system is capable of applying pressure to the first volume of fluid and the second volume of fluid. 1. An aircraft , comprising:a rotor comprising a plurality of aircraft blades operable to revolve around an axis; a first portion operable to couple to a fuselage of the aircraft;', 'a second portion operable to couple to the fuselage of the aircraft;', 'a moveable portion coupled to the first portion and the second portion, the moveable portion operable to couple to the rotor; and', 'a first volume of fluid, wherein the first housing is operable to reduce an amount of movement transferred from the rotor to the fuselage of the aircraft by transferring a portion of the first volume of fluid from the second portion of the first housing to the first portion of the first housing through the moveable portion;, 'a first housing comprisinga second housing coupled to the first housing, the second housing comprising a second volume of fluid and a volume of gas, the first volume of fluid being in fluid communication with the second volume of fluid;a rubber rolling seal positioned within the second housing, the rubber rolling seal operable to separate the second volume of fluid from the volume of gas; anda mechanical spring positioned within the second housing, the mechanical spring ...

Filled static structure for axial-flow machine

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

Flat Bottom Damper Pin For Turbine Blades

A damper system for the buckets of a gas turbine engine. The damper system may include damper pins having a generally rounded top portion and a generally flat bottom portion along substantially the entire length thereof. The generally flat bottom portion may allow addition or removal of material to or from the pin in order to achieve an optimal dynamic weight ratio. 1. A damper pin for a bucket damper slot in at least one of two adjoining buckets installed in a rotor in a turbine , the damper pin comprising:a. a generally rounded top portion, the rounded top portion configured to contact said adjoining buckets prior to full-speed rotation of said rotor; andb. a generally flat bottom portion.2. The damper pin of claim 1 , wherein the damper pin is substantially symmetrical in cross section.3. The damper pin of claim 1 , wherein the damper pin includes bossed ends.4. The damper pin of claim 1 , wherein the generally flat bottom portion extends across substantially the entire length of the damper pin.5. The damper pin of claim 1 , wherein the generally rounded top portion comprises a fully round portion.6. The damper pin of claim 1 , wherein the damper pin includes a Murphy proofing tab on the bottom thereof7. The damper pin of claim 1 , wherein said generally rounded top portion and said generally flat bottom portion are joined at a pair of rounded or beveled edges.8. A turbine engine comprising:a. a rotor having a generally circular periphery;b. a plurality of buckets mounted about the periphery of the rotor, each of said buckets including an airfoil extending outward from a platform; andc. a damper for damping vibrations of said plurality of buckets, the damper including an elongate body, the elongate body being sized and shaped for receipt within a bucket damper slot formed in undercuts of said plurality of buckets so that the elongate body frictionally engages said undercuts to dampen vibrations of said buckets, said damper being generally symmetrical in cross ...

TURBINE

A turbine includes a shaft configured to rotate about a rotor axis; a pair of rotating blade rows, the pair of rotating blade rows including a pair of disks that extend radially outward from the shaft and are disposed at an interval in a direction of the rotor axis, each one of the pair of rotating blade rows including a plurality of rotating blades arranged in a circumferential direction on an outer peripheral end of the disk; and a pair of stator vane rows disposed in a one-to-one manner on a first side of the pair of rotating blade rows in the direction of the rotor axis, each one of the pair of stator vane rows including a plurality of stator vanes arranged in the circumferential direction, wherein a number of the rotating blades on each one of the pair of rotating blade rows is the same, and a number of the stator vanes on each one of the pair of stator vane rows is the same. 1. A turbine , comprising:a shaft configured to rotate about a rotor axis;a pair of rotating blade rows, the pair of rotating blade rows including a pair of disks that extend radially outward from the shaft and are disposed at an interval in a direction of the rotor axis, each one of the pair of rotating blade rows including a plurality of rotating blades arranged in a circumferential direction on an outer peripheral end of the disk; anda pair of stator vane rows disposed in a one-to-one manner on a first side of the pair of rotating blade rows in the direction of the rotor axis, each one of the pair of stator vane rows including a plurality of stator vanes arranged in the circumferential direction, whereina number of the rotating blades on each one of the pair of rotating blade rows is the same, and a number of the stator vanes on each one of the pair of stator vane rows is the same.2. The turbine according to claim 1 , whereinthe number of the stator vanes ranges from 30% to 70% of the number of the rotating blades.3. The turbine according to claim 1 , further comprising:an attachment ...

Cavity Sealing

A method of sealing one or more openings provided in a wall of an aerofoil for a gas turbine engine, the aerofoil comprising at least one cavity which is at least partly filled with a vibration damping material, the method comprising steps to provide a metallic material onto the wall of the aerofoil in order to cover the opening and bond the metallic material to the wall of the aerofoil to seal the opening. 1. A method of sealing one or more openings provided in a wall of an aerofoil for a gas turbine engine , the aerofoil comprising at least one cavity which is at least partly filled with a vibration damping material , the method comprising steps to:a) provide a metallic material onto the wall of the aerofoil in order to cover the opening; and,b) bond the metallic material to the wall of the aerofoil to seal the opening.2. A method as claimed in , wherein step a) of comprises steps to overlap the metallic material and the wall such that a weld interface between the metallic material and the wall is generally co-planar with the wall of the aerofoil.3. A method as claimed in claim 1 , wherein the metallic material comprises a metal or alloy comprising either or both of substantially identical chemical composition or mechanical properties to the wall of the aerofoil.4. A method as claimed in claim 1 , wherein the method comprises steps to form a planar section on the wall of the aerofoil in an area adjacent to or surrounding the opening.5. A method as claimed in claim 4 , wherein the steps to form a planar section on the wall of the aerofoil comprises further steps to remove at least a portion of the wall from the aerofoil.6. A method as claimed in claim 1 , wherein the metallic material comprises two or more layers.7. A method as claimed in claim 1 , wherein step b) of comprises steps to heat and plasticise at least a portion of the aerofoil wall and one or more layers of the metallic material.8. A method as claimed in claim 1 , wherein the method comprises steps to ...

Bearing device for load reduction

A bearing assembly for a gas turbine engine comprises a bearing; a bearing bracket, which holds the bearing and is secured by a predetermined breaking device on a connecting element, which can be connected or is connected to a support structure of the gas turbine engine; and a clutch for transmitting a torque from a first clutch element connected in a fixed manner to the rotor of the bearing to a second clutch element supported on the bearing bracket, wherein the clutch elements are spaced apart when the predetermined breaking device is intact and can be brought into contact with one another by destruction of the predetermined breaking device. A gas turbine engine and a method are furthermore provided.

COMPRESSOR BLADE HAVING ORGANIC VIBRATION STIFFENER

A compressor blade of a gas turbine includes a root member; an airfoil that is disposed on the root member and includes a first interior wall and a second interior wall forming a hollow space defined between the first and second interior walls; and an organic vibration stiffener (OVS) formed on at least one of the first interior wall and the second interior wall. The OVS is formed by 3D printing performed with respect to a surface of the at least one of the first interior wall and the second interior wall and includes an uneven surface formed on at least part of the at least one of the first interior wall and the second interior wall. The OVS may include a protruded or recessed portion protruding from or recessed into at least part of the at least one of the first interior wall and the second interior wall. 1. A compressor blade of a gas turbine , comprising:a root member;an airfoil that is disposed on the root member and includes a first interior wall and a second interior wall forming a hollow space defined between the first and second interior walls; andan organic vibration stiffener (OVS) formed on at least one of the first interior wall and the second interior wall.2. The compressor blade according to claim 1 , wherein the OVS is formed by 3D printing performed with respect to a surface of the at least one of the first interior wall and the second interior wall.3. The compressor blade according to claim 1 , wherein the OVS includes an uneven surface formed on at least part of the at least one of the first interior wall and the second interior wall.4. The compressor blade according to claim 1 , wherein the OVS includes a protruded portion protruding from at least part of the at least one of the first interior wall and the second interior wall.5. The compressor blade according to claim 1 , wherein the OVS includes a recessed portion recessed into at least part of the at least one of the first interior wall and the second interior wall.6. The compressor blade ...

METHOD FOR ALTERING THE LAW OF TWIST OF THE AERODYNAMIC SURFACE OF A GAS TURBINE ENGINE FAN BLADE

A method of altering the twisting relationship for the aerodynamic surface of a fan blade of a gas turbine engine, wherein the following steps are performed: establishing, for a portion of the aerodynamic surface of the fan blade, an alteration relationship defined by variation of a pitch angle of the blade as a function of radial height along the blade, the alteration relationship including alterations that are each defined by a height along with the radial height of the fan blade and by an amplitude; and applying the alteration relationship as established in this way to an initial twisting relationship of the fan blade so as to obtain an altered twisting relationship for the fan blade, the initial twisting relationship being defined by a polynomial for the radial height of the fan blade as a function of its pitch angle. 1. A method of altering the twisting relationship for the aerodynamic surface of a fan blade of a gas turbine engine , the method comprising:establishing, for a portion of the aerodynamic surface of the fan blade, an alteration relationship defined by variation of a pitch angle of the blade as a function of radial height along the blade, said alteration relationship comprising alterations that are each defined by a height along with the radial height of the fan blade and by an amplitude; andapplying the alteration relationship as established in this way to an initial twisting relationship of the blade so as to obtain an altered twisting relationship for the fan blade, said initial twisting relationship being defined by a polynomial for the radial height of the fan blade as a function of its pitch angle.2. The method according to claim 1 , wherein the alteration relationship is defined in such a manner as to be zero and to have a derivative of zero at least one end of the portion of the aerodynamic surface of the fan blade.3. The method according to claim 1 , wherein the alteration relationship is also defined in such a manner that the amplitude of ...

MECHANICAL COMPONENT FOR THERMAL TURBO MACHINERY

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

METHOD FOR DETUNING A ROTOR-BLADE CASCADE

A method for detuning a rotor-blade cascade of a turbomachine having a plurality of rotor blades includes: a) establishing at least one target natural frequency for at least one vibration mode; b) setting up a value table having discrete mass values and radial centre-of-gravity positions, and determining respective natural frequency; c) measuring the mass and radial centre-of-gravity position of one of the rotor blades; d) determining an actual natural frequency by interpolating the measured mass and radial centre-of-gravity position in the value table; e) if actual natural frequency is outside a tolerance around target natural frequency, selecting a value pair that at least approximates target natural frequency, and removing material from the rotor blade in such a way that mass and radial centre-of-gravity position correspond to the value pair; f) repeating steps c) to e) until actual natural frequency is within the tolerance around target natural frequency. 1. A method for detuning a rotor-blade cascade , comprising a multiplicity of rotor blades , of a turbomachine , the method comprising:{'sub': 'F,S', 'a) establishing for each of the rotor blades of the rotor-blade cascade at least one setpoint natural frequency νwhich the rotor blade has for at least one predetermined oscillation mode during normal operation of the turbomachine under the effect of centrifugal force, such that the oscillation load of the rotor-blade cascade under the centrifugal force lies below a tolerance limit;'}{'sub': F', 'S', 'S', 'F', 'S, 'b) compiling a value table ν(m, r) with selected value pairs of discrete mass values m and radial center-of-mass positions r, which result from variations of the nominal geometry of the rotor blade, and determining the respective natural frequency νof the predetermined oscillation mode under the centrifugal force for each selected value pair m and r;'}{'sub': I', 'S,I, 'c) measuring the mass mand the radial center-of-mass position rof one of the rotor ...

TURBOMACHINE BLADE

A turbomachine blade having a main body () that includes a plane, first attachment surface () having a first rim contour (), a cover () that has a plane, second attachment surface () having a second rim contour () that is welded to the first attachment surface, and a tuning body configuration having at least one tuning body (-) for contacting an inner wall () of the cover by impact therewith, a gap (s) being formed between the first and the second rim contour. 1. A turbomachine blade comprising:a main body including a plane, first attachment surface having a first rim contour; anda cover including a plane, second attachment surface having a second rim contour and welded to the first attachment surface; anda tuning body configuration having at least one tuning body for contacting an inner wall of the cover by impact therewith, a gap being formed between the first and second rim contour.2. The turbomachine blade as recited in wherein the cover is disposed on the main body in a way that allows the cover to extend freely therefrom.3. The turbomachine blade as recited in wherein at least one tuning body of the tuning body configuration is accommodated in a cavity completely formed in the cover.4. The turbomachine blade as recited in wherein the cover is a multipart cover.5. The turbomachine blade as recited in wherein at least one tuning body of the tuning body configuration is accommodated in a cavity formed at least in sections thereof in the main body.6. The turbomachine blade as recited in further comprising at least two cavities having a different shape or different volume; at least one tuning body of the tuning body configuration being accommodated in at least one of these cavities.7. The turbomachine blade as recited in wherein at least one tuning body of the tuning body configuration is accommodated in a cavity; a circumferentially extending groove being formed in the first or second attachment surface between the cavity and the second rim contour.8. The ...

Low energy wake stage

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

MISTUNING OF TURBINE BLADES WITH ONE OR MORE INTERNAL CAVITIES

A bladed rotor system includes first and second sets of blades with respective airfoils each having at least one internal cavity. The airfoils of both the first and second sets of blades have identical outer shapes defined by an outer surface of an outer wall of the respective airfoils. The airfoils of the first set of blades are distinguished from the airfoils of the second set of blades by a geometry and/or position of at the least one internal cavity, which is unique to blades of a given set. The natural frequency of a blade of the first set differs from the natural frequency of a blade of the second set by a predetermined amount. The blades of the first set and the second set are alternately arranged in a periodic fashion in said circumferential row, to provide a frequency mistuning to stabilize flutter of the blades. 1. A bladed rotor system for a turbomachine , comprising:a circumferential row of blades mounted on a rotor disc, each blade comprising an airfoil having an outer wall delimiting an airfoil interior, the airfoil interior comprising one or more internal cavities, the airfoils of both the first and second sets of blades have identical outer shapes defined by an outer surface of the outer wall of the respective airfoils, and', 'the airfoils of the first set of blades are distinguished from the airfoils of the second set of blades by a geometry and/or position of at least one internal cavity, which is unique to blades of a given set,, 'the row of blades comprising a first set of blades and a second set of blades, whereinwhereby the natural frequency of a blade of the first set differs from the natural frequency of a blade of the second set by a predetermined amount, andwherein blades of the first set and the second set are alternately arranged in a periodic fashion in said circumferential row, to provide a frequency mistuning to stabilize flutter of the blades.2. The bladed rotor system according to claim 1 , wherein an outer wall thickness of the ...

DAMPING DEVICE

The invention relates to an assembly () for a turbomachine comprising: 1. A turbomachine comprising:a first rotor module comprising a first blade, the first blade having a first length;a second rotor module connected to the first rotor module and comprising a second blade, the second blade having a second length, the second length being smaller than the first length; anda damping device extending with at least one component along a turbomachine longitudinal axis, the damping device being annular while extending circumferentially around the turbomachine longitudinal axis, the damping device comprising a first radial external surface supported with friction against the first rotor module, as well as a second radial external surface supported with friction against the second rotor module, so as to couple the first rotor module with the second rotor module in order to damp vibrational movements of the first rotor module relative to the second rotor module during operation.2. The assembly of claim 1 , wherein the damping device is an annular tab claim 1 , a cross section of the damping device being shaped like a V claim 1 , a first external surface of a first branch of the damping device forming the first radial external surface claim 1 , a second external surface of a second branch of the damping device forming the second radial external surface.3. The assembly of claim 1 , wherein:the first rotor module comprises a disk centered on the turbomachine longitudinal axis;the first blade is mounted on an external periphery of the disk, the first blade thus extending from the external periphery of the disk, the first blade further comprising an airfoil, a platform, a support and a root, the root being embedded in a housing of the disk, the first radial external surface being supported with friction on a radially internal surface of the platform; andthe second rotor module comprises a ferrule, the ferrule comprising a circumferential extension extending toward the platform of ...

BLADED ROTOR SYSTEM AND CORRESPONDING METHOD OF SERVICING

A bladed rotor system for a turbomachine includes a circumferential row of blades mounted on a rotor disc, and includes a plurality of under-platform dampers. Each damper is located between adjacent blade platforms. The plurality of dampers includes a first set of dampers and a second set of dampers. The dampers of the first set are distinguished from the dampers of the second set by a cross-sectional material distribution in the damper that is unique to the respective set. Dampers of the first set and the second set are positioned alternately in a periodic fashion in a circumferential direction, to provide a frequency mistuning to stabilize flutter of the blades. 1. A bladed rotor system for a turbomachine , comprising: a platform;', 'a root extending radially inward from the platform for mounting the blade to the rotor disc; and', 'an airfoil extending span-wise radially outward from the platform;, 'a circumferential row of blades mounted on a rotor disc, each blade comprisingwherein platforms of adjacent blades align circumferentially to define an inner diameter boundary for a working fluid flow path; anda plurality of dampers, each damper being located between adjacent platforms;wherein the plurality of dampers comprise a first set of dampers and a second set of dampers, wherein the dampers of the first set are distinguished from the dampers of the second set by a cross-sectional material distribution in the damper that is unique to the respective set, andwherein dampers of the first set and the second set are positioned alternately in a periodic fashion in a circumferential direction, to provide a frequency mistuning to stabilize flutter of the blades.2. The bladed rotor system according to claim 1 ,wherein the dampers of the first set are solid,and the dampers of the second set are hollow, each defining an internal cavity therewithin.3. The bladed rotor system according to claim 2 , wherein the dampers of the first set and the dampers of the second set are ...

UNSHROUDED TURBOMACHINE IMPELLER WITH IMPROVED RIGIDITY

An unshrouded turbomachine impeller is disclosed. The impeller comprises a hub and a plurality of sequentially arranged blades. Each blade extends from a blade root at the hub to a blade tip and is comprised of a first blade edge and a second blade edge. A flow vane is formed between each pair of neighboring blades. A connection member extends across each flow vane between neighboring blades and rigidly or monolithically connects a first modal displacement region of a first one of the pair of neighboring blades to a second modal displacement region of a second one of the pair of neighboring blades. 1. An unshrouded turbomachine impeller comprising:a rotation axis;a hub;a plurality of sequentially arranged blades, each blade extending from a blade root at the hub to a blade tip and comprised of a first blade edge and a second blade edge, the first blade edge and the second blade edge extending from the hub to the blade tip; anda flow vane arranged between each pair of neighboring blades;wherein a connection member extending across each flow vane between pairs of neighboring blades connects a first modal displacement region at a certain frequency of a first one of said pair of neighboring blades to a second modal displacement region at said frequency of a second one of said pair of neighboring blades; andwherein each connection member has a first end rigidly or monolithically connected to a pressure side of a first one of said pair of neighboring blades and a second end rigidly or monolithically connected to a suction side of a second one of said pair of neighboring second blades, or vice versa.2. The turbomachine impeller of claim 1 , wherein the first blade edge is located at a first radial distance from the rotation axis and the second blade edge is located at a second radial distance from the rotation axis the first radial distance being smaller than the second radial distance.3. The turbomachine impeller of claim 1 , wherein the first modal displacement region is ...

COMPRESSOR ROTOR WITH COATED BLADES

A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades. 1. A compressor rotor for a gas turbine engine , the compressor rotor comprising blades extending a span length from a central hub , the blades including circumferentially alternating first and second blades having airfoils with corresponding geometric profiles , each of the airfoils including a leading edge , a trailing edge , a root , a tip and a mid-span region between the root and the tip along the span , the airfoil of the first blades having a coating on a first portion of the first blade adjacent the root with a root coating thickness , and the coating being provided on a second portion adjacent the tip of the first blade with a tip coating thickness , the root coating thickness being greater than the tip coating thickness , the coating defining a first coating structure providing the first blade with a first natural vibration frequency different from a second natural vibration frequency of the second blade.2. The compressor rotor as defined in claim 1 , wherein the coating is provided on the mid-span region of the airfoil of the first blades.3. The compressor rotor as defined in claim 2 , wherein the coating on the mid-span region has a mid-span coating thickness claim 2 , the root coating thickness being greater than the mid-span coating thickness.4. The compressor rotor as defined in claim 1 , wherein the airfoil of the second blade is free of coating.5. The compressor rotor as defined in claim 1 , wherein the airfoil of the second blade has a coating defining a second coating structure claim 1 , the coating being provided on ...

Fan Blade Damping Device

An airfoil for a gas turbine engine and method of manufacture of the airfoil are disclosed. The airfoil may comprise a first side extending axially from a leading edge to a trailing edge and extending radially from a base to a tip, a second side opposite to the first side, a pocket disposed in the first side, a filler disposed in the pocket, and a preloaded spring disposed within the filler. 1. An airfoil for a gas turbine engine , comprising:a first side extending axially from a leading edge to a trailing edge and extending radially from a base to a tip;a second side opposite to the first side;a pocket disposed in the first side;a filler disposed in the pocket; anda preloaded spring disposed within the filler.2. The airfoil of claim 1 , wherein the preloaded spring exerts a force within the pocket for resisting flutter of the airfoil and for damping vibratory response of the airfoil.3. The airfoil of claim 1 , wherein the filler surrounds and encapsulates the preloaded spring claim 1 , and is formed to fill the pocket.4. The airfoil of claim 3 , wherein the filler dampens a vibratory response of the airfoil claim 3 , prevents decompression of the preloaded spring claim 3 , imparts force exerted from the preloaded spring to the airfoil claim 3 , and prevents contact of the preloaded spring with the pocket.5. The airfoil of claim 1 , wherein the filler is bonded to the pocket of the first side claim 1 , and the second side is bonded to the first side and the filler.6. The airfoil of claim 1 , wherein the preloaded spring and the filler are configured to change an inherent frequency of the airfoil to a predetermined frequency outside a range of resonant frequencies of the airfoil.7. The airfoil of claim 6 , wherein at least one of a composition of the filler claim 6 , thickness of the filler claim 6 , durometer of the filler claim 6 , thickness of the preloaded spring claim 6 , preload of the preloaded spring claim 6 , stiffness of the preloaded spring claim 6 , and ...

COMPOSITE OUTLET GUIDE VANE WITH METAL FASTENER FOR A TURBOMACHINE

The invention describes an outlet guide vane () for a fan () of a turbomachine, comprising: 1. An outlet guide vane for a fan module of a turbomachine , the outlet guide vane extending along a radial axis , the radial axis extending along a direction perpendicular to a central axis of symmetry of the fan module and passing therethrough , the fan module comprising a hub including an upstream flange and a downstream flange , the outlet guide vane comprising:a blade made of composite material comprising a fiber reinforcement densified by a matrix and extending along the radial axis, said blade having a radial inner end and a radial outer end,a first metal fastener portion fastened to the radial inner end of the blade,at least a first platform extending transversely with respect to the blade in the vicinity of its radial inner end,wherein said first metal fastener portion is configured to be mounted on and fastened, on the one hand, to the upstream flange and, on the other hand, to the downstream flange of the hub, and in that the first platform is monolithic with the first metal fastener portion.2. The outlet guide vane according to claim 1 , further comprising a second platform extending transversely with respect to the blade in the vicinity of the radial outer end of the blade.3. The outlet guide vane according to claim 1 , further comprising a second fastener portion fastened to the radial outer end of the blade claim 1 , said second portion being configured to be mounted on and fastened to a casing claim 1 , said second portion being metallic or made of a composite material comprising a fiber reinforcement densified by a matrix.4. The outlet guide vane according to taken in combination claim 2 , wherein the second fastener portion is metallic and the second platform is metallic and monolithic with the second metal fastener portion.5. The outlet guide vane according to claim 3 , wherein the second fastener portion is metallic claim 3 , and further comprising a ...

REDUCTION OF EQUALLY SPACED TURBINE NOZZLE VANE EXCITATION

A reduction in excitation amplitudes affecting turbine blade durability in a turbine nozzle assembly having a plurality of vanes and turbine blades, includes: identifying a turbine blade design of the turbine nozzle assembly; performing a modal model analysis of at least one of the turbine blades in the turbine blade design; reducing aerodynamic impact by ensuring that each of the turbine blades is free of aero-excitation from an upstream flow at the vanes in an operating speed range; identifying blade natural frequencies with respect to the nozzle vanes; and modifying a trailing edge of at least one of the vanes to reduce the excitation amplitudes. 1. Method for reducing excitation amplitudes affecting turbine blade durability in a turbine nozzle assembly having a plurality of vanes and turbine blades , comprising the steps of:identifying a turbine blade design of said turbine nozzle assembly;performing a modal model analysis of at least one of said turbine blades in said turbine blade design using a computer;reducing aerodynamic impact by ensuring that each of said turbine blades is free of aero-excitation from an upstream flow at said vanes in an operating speed range;identifying blade natural frequencies with respect to said nozzle vanes using said computer; anddetermining at least one modification to a trailing edge of at least one of said vanes to reduce said excitation amplitudes.2. The method of claim 1 , wherein said determining comprises altering an angle at which a flow of gas enters said turbine blades and interrupts energy build up.3. The method of claim 1 , wherein said determining comprises performing a CFD analysis to determine a vane exit angle resulting in maximum pressure perturbance and minimizing P(ω)4. The method of claim 1 , further comprising guiding a modification of said vane exit angle in a direction of shifting blade pressure loading toward a leading edge of said at least one turbine blade away from a blade anti-node.5. The method of ...

MISTUNED CONCENTRIC AIRFOIL ASSEMBLY AND METHOD OF MISTUNING SAME

An airfoil assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, an annular shroud having a radially inner face and a radially outer face opposing the radially inner face, a radially inner array of airfoils extending from the radially inner face, and a radially outer array of airfoils extending from the radially outer face. The radially inner array of airfoils are configured to guide flow within a radially inner bypass flow passage, the radially inner bypass flow passage bypassing and being radially outward of a compressor section. At least one, but fewer than each, airfoil of the radially inner array of airfoils is circumferentially aligned with a corresponding airfoil in the radially outer array of airfoils, and the remaining airfoils in the radially inner array of airfoils are misaligned with the airfoils of the radially outer array of airfoils. A method of reducing a vibratory response of airfoils is also disclosed. 1. An airfoil assembly for a gas turbine engine , comprising:an annular shroud having a radially inner face and a radially outer face opposing the radially inner face;a radially inner array of airfoils extending from the radially inner face;a radially outer array of airfoils extending from the radially outer face;wherein the radially inner array of airfoils are configured to guide flow within a radially inner bypass flow passage, the radially inner bypass flow passage bypassing and being radially outward of a compressor section; andwherein at least one, but fewer than each, airfoil of the radially inner array of airfoils is circumferentially aligned with a corresponding airfoil in the radially outer array of airfoils, and the remaining airfoils in the radially inner array of airfoils are misaligned with the airfoils of the radially outer array of airfoils.2. The assembly of claim 1 , wherein one airfoil of the radially inner array of airfoils is circumferentially aligned with a corresponding ...

Rotor blade for a turbomachine and turbomachine

The invention relates to a blade ( 2 ) of a turbomachine, in particular a rotor blade of a gas turbine, which has a variable transition radius (Rv 1 ) in the vicinity of at least one platform overhang ( 16 ), and to a turbomachine having at least one such blade ( 2 ).

UNDERPLATFORM DAMPING MEMBERS AND METHODS FOR TURBOCHARGER ASSEMBLIES

Damping members for turbocharger assemblies, methods for providing turbocharger assemblies, and turbocharger assemblies are described herein. The damping members include bodies having shapes to fit between a recess extending into a rotor disk of a turbocharger and laterally protruding shoulders of platforms in neighboring blades of the turbocharger. The bodies dampen vibrations of the blades during rotation of the blades. The damping members may include a variety of shapes, such as a sheet, a wedge, a tapered pin, a cylindrical pin, a bent sheet, or another shape. 1. A damping member comprising:a body having a shape to fit between a recess extending into a rotor disk of a turbocharger and laterally protruding shoulders of platforms in neighboring blades of the turbocharger, the body configured to dampen vibrations of the blades during rotation of the blades and the rotor disk by engaging under surfaces of the shoulders of the platforms that oppose the rotor disk of at least one of the blades.2. The damping member of claim 1 , wherein the body has the shape to concurrently engage the under surfaces of the shoulders of the platforms of the blades when the platforms are misaligned with respect to the rotor disk.3. The damping member of claim 1 , wherein the shape of the body is sized to fit between the recess in the rotor disk and the laterally protruding shoulders of the platforms in the neighboring blades that do not include shanks vertically extending between dovetails of the blades and the platforms of the blades.4. The damping member of claim 1 , wherein the body has the shape of a sheet having opposite parallel first and second sides and opposite parallel first and second edges claim 1 , wherein the first and second edges intersect the first side at curved interfaces that are configured to concurrently engage the under surfaces of the shoulders of the platforms in the blades.5. The damping member of claim 1 , wherein the body has the shape of a wedge comprising a ...

ASSEMBLING METHOD OF A BUCKET AND A FIXTURE FOR A BUCKET FOR A TURBINE BLADE

Disclosed herein is a fixture for a bucket for a turbine blade. The fixture includes a rotor wheel that includes a plurality of dovetail grooves. A platform seat is formed between the dovetail grooves and provided with a first insertion groove in a circumferential direction. A bucket includes a platform that inserts into one of the dovetail grooves. A base platform disposed on an upper surface of the platform includes a second insertion groove at a position facing the first insertion groove. 1. A fixture for a bucket for a turbine blade , comprising: a plurality of dovetail grooves extending in axial direction of a rotor and disposed to be spaced apart from each other along a circumferential direction of the rotor, and', 'a plurality of platform seats disposed between the dovetail grooves, each including a first insertion groove extending in the circumferential direction;, 'a rotor wheel that includes'} a platform that inserts into one of the dovetail grooves, and', 'a base platform disposed on an upper surface of the platform, the base platform including a second insertion groove at a position that faces the first insertion groove;, 'a bucket including'}a first fixture including in an internal area defined by the first and second insertion grooves;a second fixture operable to be inserted into the internal area formed in the first and second insertion grooves of a final bucket to fix the final bucket among buckets mounted on the rotor wheel; andan auxiliary fixture operable to be inserted into an opening hole formed in the base platform of the final bucket and facing the second fixture.2. The fixture of claim 1 , wherein the first insertion groove is disposed at a leading position corresponding to a front end with respect to the axial direction in an upper surface of the platform seat.3. The fixture of claim 1 , wherein the first insertion groove is disposed at a position between a front end with respect to the axial direction in an upper surface of the platform ...

SPLIT RING SPRING DAMPERS FOR GAS TURBINE ROTOR ASSEMBLIES

A spring damper includes a split ring body. The split ring body defines a center and a circular gap separating opposed first and second end portions of the split ring body. The first and second end portions are connected by a split ring body segment that is evenly spaced from the center. At least one of the first and second end portions is unevenly spaced from the center in relation to the segment that is evenly spaced with respect to the center. 1. A spring damper , comprising:a split ring body defining a center and a circular gap separating opposed first and second end portions of the split ring body, wherein the first and second end portions are connected by a segment of the split ring body that is evenly spaced from the center, wherein at least one of the first and second end portions is unevenly spaced from the center in relation to the segment that is evenly spaced.2. A damper as recited in claim 1 , wherein an end of the first end portion is spaced radially outward from the center in relation to the evenly spaced segment.3. A damper as recited in claim 1 , wherein ends of the first and second end portions are spaced radially outward from the center in relation to the evenly spaced segment.4. A damper as recited in claim 1 , wherein the evenly spaced segment is offset from the center by a uniform radius.5. A damper as recited in claim 1 , wherein the split ring body has an elliptical shape.6. A damper as recited in claim 1 , wherein the spring damper has an unloaded configuration wherein ends of the first and second end portions extend radially outward in relation to the evenly spaced segment and are separated by an unloaded gap width.7. A damper as recited in claim 6 , wherein the spring damper has a statically-loaded configuration wherein ends of the first and second end portions are spaced inwards in relation to the evenly spaced segment.8. A damper as recited in claim 6 , wherein ends of the end portions are separated by a statically loaded gap with a ...

BLADED GAS TURBINE ROTOR

The present invention relates to a bladed rotor for a gas turbine, with a body and a plurality of rotating blades where contact regions of a cylindrical surface of the body and front faces of the rotating blades are joined together cohesively. At least one of these contact regions and/or at least one of these front faces has at least one blind hole, in which at least one impulse body is arranged with play of movement and which is sealed and rotating blades are joined with the cylindrical surface. Alternatively, the body and rotating blades are manufactured integrally with one another by machining, and at least one blind hole extends from an inner surface of the body facing away from the blade out to one of the rotating blades, at least one impulse body being disposed with play of movement in this hole, and this hole is sealed. 1. A bladed rotor for a gas turbine , comprising:{'b': 10', '20', '30', '12', '13', '11', '21', '31', '40', '50', '41, 'a basic body () and a plurality of rotating blades (, ), wherein contact regions (, ) of a cylindrical surface () of the basic body and front faces (, ) of the rotating blades are joined together cohesively, and at least one of these contact regions and/or at least one of these front faces has at least one blind hole (), in which at least one impulse body () is arranged with play of movement and which is sealed (), in addition to the joining of the rotating blades with the cylindrical surface.'}2. The bladed rotor according to claim 1 , wherein the contact regions and front faces are joined together by friction welding.3. The bladed rotor according to claim 1 , wherein the basic body and rotating blades are manufactured integrally with one another by machining.4. The bladed rotor according to claim 1 , wherein at least one of the blind holes is sealed cohesively by welding.541. The bladed rotor according to claim 1 , wherein at least one of the blind holes is sealed by a separate cover ().6. The bladed rotor according to ...

STATOR SHROUD WITH MECHANICAL RETENTION

A stator assembly for a gas turbine engine includes an arcuate shroud including a shroud pocket, the shroud pocket having a shroud slot extending therethrough. A stator vane is insertable into the shroud pocket and includes a vane slot extending therethrough. A strap extends through the shroud slot and the vane slot to retain the vane to the shroud. A gas turbine engine includes a combustor and a stator and case assembly in in fluid communication with the combustor. The stator and case assembly includes a case defining a working fluid flowpath for the gas turbine engine and a stator assembly secured at the case. 1. A stator assembly for a gas turbine engine , comprising:an arcuate shroud including a shroud pocket, the shroud pocket having a shroud slot extending therethrough;a stator vane insertable into the shroud pocket and including a vane slot extending therethrough; anda strap extending through the shroud slot and the vane slot to retain the vane to the shroud.2. The stator assembly of further comprising a volume of potting disposed at the shroud pocket to retain the stator vane thereat.3. The stator assembly of claim 2 , wherein the potting is a rubber material.4. The stator assembly of claim 2 , wherein the potting comprises a grommet disposed at the shroud pocket.5. The stator assembly of claim 1 , wherein the shroud pocket includes a pocket sidewall and a pocket base.6. The stator assembly of claim 5 , wherein the shroud slot extends through the pocket sidewall.7. The stator assembly of claim 1 , wherein the stator vane is inserted in two shroud pockets of two shrouds claim 1 , with a strap extending through a vane slot and a pocket slot at each shroud of the two shrouds.8. A stator and case assembly for a gas turbine engine comprising:a case defining a working fluid flowpath for the gas turbine engine; and an arcuate shroud including a shroud pocket, the shroud pocket having a shroud slot extending therethrough;', 'a stator vane insertable into the shroud ...

FLUID PULSE DEVICE AND METHOD OF EXCITING GAS TURBINE ENGINE TUROMACHINERY COMPONENTS

A fluid pulse device includes a piezo actuated valve, a piezo injector, a delivery system, and a control system. The piezo actuated valve has a valve body, a tube, and a piston assembly. The valve body defines a first opening, a second opening, and a valve cavity that is disposed between the first opening and the second opening. The tube extends from the valve body. The piston assembly includes a piston head and a stem that is connected to the piston head. The piezo injector has an injector tip that extends through the first opening. The delivery system is fluidly connected to the piezo injector. The control system is operatively connected to the piezo injector and is configured to actuate the piezo injector to provide a fluid pulse. 1. A piezo actuated valve provided with a fluid pulse device , comprising:a valve body extending along a first axis between a first end and a second end, the valve body defining a first opening disposed proximate the first end, a second opening disposed proximate the second end, and a valve cavity disposed between the first opening and the second opening;a tube extending from the valve body along a second axis disposed transverse to the first axis; a piston head slidably disposed within the valve cavity; and', 'a stem extending from the piston head and extending through the second opening and slidably received within the tube; and, 'a piston assembly disposed within the valve body comprisinga piezo injector having an injector tip extending through the first opening.2. The piezo actuated valve of claim 1 , wherein the tube defines a passageway extending between an inlet and an outlet.3. The piezo actuated valve of claim 2 , wherein the stem defines an aperture.4. The piezo actuated valve of claim 3 , wherein the piston assembly is movable between a first position in which the stem inhibits fluid flow through the passageway and a second position in which the aperture is at least partially disposed within the passageway to facilitate fluid ...

Blade platform with damper restraint

A gas turbine engine, having: a disk; a plurality of blades secured to the disk, each of the plurality of blades having a platform located between a root portion and an airfoil portion of the blade, wherein the platform of one of the plurality of blades is configured to define a cavity with a platform of an adjacent blade that is secured to the disk; a damper seal located in the cavity and positioned adjacent to a gap defined by edges of the platforms of the blades; and a damper restraint located on an interior surface of each platform, wherein the damper restraint extends into the cavity and is a raised feature configured to contact a peripheral edge portion of a damper seal when it is adjacent to the gap defined by the platforms of the blades.

Tuned airfoil assembly

An airfoil assembly may include a shroud and an airfoil. The shroud may include a first attachment arm, a second attachment arm, and a shroud rail extending from a first surface of the shroud. A first channel may be defined between the first attachment arm, the first surface, and the shroud rail and a second channel may be defined between the second attachment arm, the first surface, and the shroud rail. The airfoil may extend from a second surface of the shroud opposite the first surface. In various embodiments, a height of the shroud rail, as measured from the first surface of the shroud, is non-uniform.

CENTRIFUGAL FAN AND AIR-CONDITIONING APPARATUS

A centrifugal fan includes a main plate fixed to a rotating shaft of a drive unit, a shroud having an air suction port, and a plurality of blades disposed between the main plate and the shroud. A rib for reinforcement is formed in each of the blades. Deformation due to an external pressure can be suppressed, the main plate, the shroud, and the blades are firmly fixed to each other, and a decrease in reliability, generation of noise, and the like due to deformation are suppressed. 1. A centrifugal fan comprising:a main plate fixed to a rotating shaft of a drive unit;a shroud including an air suction port; anda plurality of blades disposed between the main plate and the shroud, wherein a rib is formed in each of the blades.2. The centrifugal fan of claim 1 , wherein the rib is formed so as to extend between a surface on which the shroud is in contact with each of the blades and a surface on which the main plate is in contact with each of the blades.3. The centrifugal fan of claim 1 , wherein the height of the rib is partially varied.4. The centrifugal fan of claim 1 , wherein the thickness of the rib is partially varied.5. The centrifugal fan of claim 1 , wherein each of the blades includes a main blade fixed to the main plate and the shroud claim 1 , and a blade cover that is assembled with the main blade so as to have an internal space between the blade cover and the main blade and thereby forms each of the blades.6. The centrifugal fan of claim 5 , wherein the main blade includes a protruding portion that is fitted in a recessed portion formed in the shroud and whose top surface serves as a contact surface to be in contact with the shroud claim 5 , and the rib is formed in the space so as to be continuous with a side wall surface of the protruding portion.7. The centrifugal fan of claim 5 , wherein the rib is formed in a shape which prevents the main blade and the blade cover from deforming toward the space.8. The centrifugal fan of claim 7 , wherein the rib is ...

Variable stator vane structure of axial compressor

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

Turbomachine blade

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

Tangential Blade Root Neck Conic

A root extending from a platform of an airfoil is disclosed. The root may include a first portion having a generally cylindrical shape, and a second portion extending from the first portion to the platform. The second portion may have a circumference larger than a circumference of the first portion. 1. A root extending from a platform of an airfoil , comprising:a first portion having a generally cylindrical shape; anda second portion extending from the first portion to the platform, the second portion having a circumference larger than a circumference of the first portion.2. The root of claim 1 , wherein the second portion increases in cross-sectional area from the first portion to the platform.3. The root of claim 1 , wherein the second portion has a generally conical shape.4. The root of claim 1 , wherein the second portion increases a frequency at which lower modes of the airfoil occur.5. The root of claim 1 , further comprising a third portion between the first portion and the second portion claim 1 , the third portion having a generally arcuate shape.6. The root of claim 1 , further comprising a fourth portion between the platform and the second portion claim 1 , the fourth portion having a generally arcuate shape.7. The root of claim 1 , wherein the root is shaped to fit within a groove of a disk claim 1 , and wherein the circumference of the first portion is determined by an upper wall of the groove.8. The root of claim 7 , further comprising an end portion extending from the first portion in a direction away from the platform claim 7 , the end portion having a bearing surface which is in contact with a bearing surface of the groove.9. The root of claim 1 , wherein the first and second portions are aligned with a central axis of the airfoil.10. A gas turbine engine claim 1 , comprising:a compressor section;a combustor section downstream of the compressor section; and a platform, and', a cylindrical neck, and', 'a segment gradually increasing in cross- ...

FLUTTER-RESISTANT BLADE

An aircraft engine having a fan, the fan having at least one flutter-resistant blade, the blade having a leading edge (LE), a trailing edge (TE), a midchord (MC), a minimum radial height r, a maximum radial height r, and a radial extent between rand r, wherein, at every point along the radial extent of the blade, the blade has a modeshape value Vfor a blade first vibratory mode defined as 2. The aircraft engine of claim 1 , wherein the first set of blades consists of a single blade.3. The aircraft engine of claim 1 , wherein the first set of blades consists of 50% or more of the plurality of blades.4. The aircraft engine of claim 1 , wherein the first set of blades consists of 75% or more of the plurality of blades.5. The aircraft engine of claim 1 , wherein the first set of blades consists of 90% or more of the plurality of blades.6. The aircraft engine of claim 1 , wherein the modeshape value Vof each blade of the first set of blades is from 0 to 1.0.7. The aircraft engine of claim 1 , wherein the modeshape value Vof each blade of the first set of blades is from 0 to 0.5.8. The aircraft engine of claim 1 , wherein the modeshape value Vof each blade of the first set of blades is from 0 to 0.2.9. The aircraft engine of claim 1 , wherein the modeshape value Vapplies to at least 85% of the radial extent of each blade in the first set of blades.10. The aircraft engine of claim 1 , wherein the modeshape value Vapplies to at least 90% of the radial extent of each blade in the first set of blades.11. The aircraft engine of claim 1 , wherein the modeshape value Vapplies to at least 95% of the radial extent of each blade in the first set of blades.12. The aircraft engine of claim 1 , wherein the modeshape value Vapplies to at least 99% of the radial extent of each blade in the first set of blades.13. An aircraft having at least one aircraft engine according to . This specification is based upon and claims the benefit of priority from UK Patent Application Number GB2014744.3 ...

System and Method for Airfoil Vibration Control

A system for airfoil vibration control is generally provided. The system includes an airfoil including a ferromagnetic material, and a static structure including an electromagnet adjacent to the ferromagnetic material of the airfoil. A method for controlling vibration at an airfoil of a turbo machine is further provided. The method includes placing a ferromagnetic material at the airfoil, placing an electromagnet at a static structure adjacent to the ferromagnetic material at the airfoil, and applying an electromagnetic force to the ferromagnetic material at the airfoil via the electromagnet at the static structure. 1. A system for airfoil vibration control , the system comprising:an airfoil comprising a ferromagnetic material; anda static structure comprising an electromagnet adjacent to the ferromagnetic material of the airfoil.2. The system of claim 1 , wherein the static structure comprises a stator airfoil directly adjacent to the airfoil.3. The system of claim 2 , wherein the electromagnet is disposed at a trailing edge tip of the static structure comprising the stator airfoil.4. The system of claim 1 , wherein the static structure comprises a casing surrounding the airfoil.5. The system of claim 4 , wherein the electromagnet is disposed at the static structure comprising the casing at least partially forward of a leading edge of the airfoil.6. The system of claim 1 , wherein the airfoil comprises a first material at least partially surrounding the ferromagnetic material.7. The system of claim 6 , wherein the ferromagnetic material comprises a plurality of particles within the first material of the airfoil.8. The system of claim 1 , wherein the airfoil comprises a first material comprising a composite material.9. The system of claim 1 , wherein the ferromagnetic material defines a plurality of structures defining one or more cross sectional areas at the airfoil.10. The system of claim 1 , wherein the ferromagnetic material is disposed between a tip and ...

INTERNALLY DAMPED AIRFOILED COMPONENT AND METHOD

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

VIBRATIONAL DAMPENING ELEMENTS

A vibrational dampening element is attached to a component and configured to adjust the amplitude of oscillations of the component. The vibrational dampening element includes a mass. The mass includes a main body and a member extending from the main body. A casing that encapsulates the mass. A fluidic chamber defined between the mass and the casing. A first fluidic portion is disposed between a first side of the mass and the casing. The first fluidic portion includes a first accumulator portion directly neighboring the member. A second fluidic portion is disposed between a second side of the mass and the casing. The second fluidic portion includes a second accumulator portion directly neighboring the member. The first accumulator portion is in fluid communication with the second accumulator portion. The vibrational dampening element further includes a primary passage that extends between the first fluidic portion and the second fluidic portion. 1. A vibrational dampening element attached to a turbine component and configured to adjust the amplitude of oscillations of the turbine component , the vibrational dampening element comprising:a mass having a main body and a member extending from the main body;a casing encapsulating the mass;a fluidic chamber defined between the mass and the casing and filled with a fluid;a first fluidic portion of the fluidic chamber disposed between a first side of the mass and the casing, wherein the first fluidic portion includes a first accumulator portion that extends along the member;a second fluidic portion disposed between a second side of the mass and the casing, wherein the second fluidic portion includes a second accumulator portion that extends along the member, and wherein the first accumulator portion is in fluid communication with the second accumulator portion; anda primary passage that extends between the first fluidic portion and the second fluidic portion.2. The vibrational dampening element as in claim 1 , wherein the ...

METHOD FOR PRODUCING A VIBRATION-DAMPING STRUCTURE COMBINATION FOR DAMPING VIBRATIONS OF MOVABLE MASSE

A method for producing a vibration-damping structure combination for damping vibrations for movable masses, having a first structure and a further structure, the further structure movable within a stop surface defined by a first structure surface of the first structure. The method includes a) providing the first structure, having the first structure surface and which defines a coating surface of a coating at least in some sections; b) coating the first structure surface of the first structure with the coating, the coating surface of the coating being applied such that a cavity is formed; c) filling the cavity with the filler; d) curing the filler until the further structure having a further structure surface is formed, which lies against the coating surface; and e) removing the coating, the further structure thus being movable relative to the first structure within the stop surface defined by the first structure surface. 1. A method for producing a vibration-damping structure combination for damping vibrations for movable masses , having a first structure and a further structure , the further structure being movable within an abutment surface defined by a first structure surface of the first structure , the method comprising:a) providing the first structure, which comprises the first structure surface and at least in sections determines a coating surface of a coating;b) coating the first structure surface of the first structure with the coating, the coating surface of the coating being applied so as to form at least one cavity;c) filling the cavity with a filler;d) curing the filler until the further structure having a further structure surface, which bears on the coating surface is formed; ande) removing the coating so that the further structure is movable relative to the first structure within the abutment surface defined by the first structure surface.2. The method as claimed in claim 1 ,wherein a lattice structure is at least partially provided as the first ...

MISTUNED ROTOR FOR GAS TURBINE ENGINE

A rotor for a gas turbine engine. The rotor includes blades circumferentially distributed around a hub. The blades have airfoils with a span defined between a root and tip, a chord defined between a leading edge and a trailing edge, and a thickness defined between a pressure side surface and suction side surface. The blades include first blades and second blades. The airfoil of the first blades has a first thickness distribution defining a first natural vibration frequency of the airfoils of the first blades. The airfoil of the second blades has a second thickness distribution defining a second natural vibration frequency different than the first natural vibration frequency. The first thickness distribution is different than the second thickness distribution along a radially-inner half of the span, and the first thickness distribution matches the second thickness distribution along a radially-outer half of the span. 1. A rotor for a gas turbine engine , the rotor comprising blades circumferentially distributed around a hub , the blades having airfoils with a span defined between a root and tip of the airfoils , the airfoils having a chord defined between a leading edge and a trailing edge of the airfoils , the airfoils having a thickness defined between a pressure side surface and suction side surface of the airfoils , the blades including first blades and second blades interleaved about the rotor , the airfoil of the first blades having a first thickness distribution along the span defining a first natural vibration frequency of the airfoils of the first blades , the airfoil of the second blades having a second thickness distribution along the span defining a second natural vibration frequency different than the first natural vibration frequency , the first thickness distribution being different than the second thickness distribution along a radially-inner half of the span , and the first thickness distribution matching the second thickness distribution along a ...

COMPRESSOR ROTOR WITH COATED BLADES

A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades. 1. A compressor rotor for a gas turbine engine , the rotor comprising blades circumferentially distributed around and extending a span length from a central hub , the blades including alternating first and second blades having airfoils with a leading edge , a trailing edge , a root , a tip and a mid-span region midway between the root and the tip along the span , the airfoils of the first and second blades having corresponding geometric profiles , the airfoil of the first blades having a coating defining a first coating structure , the coating being provided on at least a portion of the first blade adjacent the root and having a root coating thickness , the mid-span region of the first blade having a mid-span thickness , the coating being provided on a portion adjacent the tip of the first blade and having a tip coating thickness , the root coating thickness being greater than at least one of the tip coating thickness and a coating thickness of the airfoil of the first blade at the mid-span region , the first coating structure of the first blade selected to provide the first blade with a first natural vibration frequency different from a second natural vibration frequency of the second blade.2. The compressor rotor as defined in claim 1 , wherein the airfoil of the second blade is free of coating.3. The compressor rotor as defined in claim 1 , wherein the airfoil of the second blade has a coating defining a second coating structure claim 1 , the coating being provided on at least a portion adjacent the root of the second blade and having ...

FIRST-STAGE STATIONARY VANE OF GAS TURBINE AND GAS TURBINE

A first-stage stationary vane of a gas turbine includes: a vane portion including a pressure surface and a suction surface; a shroud wall portion which connects to an end portion of the vane portion and which forms a flow passage wall; a pressure-surface side fillet portion disposed on a corner portion formed by the pressure surface and a wall surface of the shroud wall portion; and a suction-surface side fillet portion disposed on a corner portion formed by the suction surface and the wall surface of the shroud wall portion. The pressure-surface side fillet portion and the suction-surface side fillet portion are separated at a leading-edge side of the vane portion so as not to connect to each other. 1. A first-stage stationary vane of a gas turbine , comprising:a vane portion including a pressure surface and a suction surface;a shroud wall portion which connects to an end portion of the vane portion and which forms a flow passage wall;a pressure-surface side fillet portion disposed on a corner portion formed by the pressure surface and a wall surface of the shroud wall portion; anda suction-surface side fillet portion disposed on a corner portion formed by the suction surface and the wall surface of the shroud wall portion,wherein the pressure-surface side fillet portion and the suction-surface side fillet portion are separated at a leading-edge side of the vane portion so as not to connect to each other.2. The first-stage stationary vane of a gas turbine according to claim 1 ,wherein an upstream-side end portion of the vane portion includes an upstream-side end surface which connects the pressure surface and the suction surface, andwherein the upstream-side end surface includes a flat surface which connects to the shroud wall portion.3. The first-stage stationary vane of a gas turbine according to claim 2 ,wherein an upstream-side end surface of the pressure-surface side fillet portion and an upstream-side end surface of the suction-surface side fillet portion are ...

Aircraft gas turbine engine with shock-absorbing element for fan blade loss

An aircraft gas-turbine engine having a fan mounted on a fan shaft and including several fan blades, where the fan shaft is mounted on an engine structure by means of an upstream fan bearing, when seen in the flow direction, and a downstream bearing, where at least one predetermined breaking point is provided in the area of the fan bearing, characterized in that at least one shock-absorbing element designed as metal fabric damper is arranged at a radial gap to have a parallel effect relative to the predetermined breaking point. 1. An aircraft gas-turbine engine having a fan mounted on a fan shaft and including several fan blades , where the fan shaft is mounted on an engine structure by means of an upstream fan bearing , when seen in the flow direction , and a downstream bearing , where at least one predetermined breaking point is provided in the area of the fan bearing , wherein at least one shock-absorbing element designed as metal fabric damper is arranged at a radial gap to have a parallel effect relative to the predetermined breaking point.2. The aircraft gas-turbine engine in accordance with claim 1 , wherein the predetermined breaking point and the shock-absorbing element are arranged on the radially outer side of the fan bearing facing the engine structure.3. The aircraft gas-turbine engine in accordance with claim 1 , wherein the shock-absorbing element is designed as a combination of two shock-absorbing elements with differing damping characteristics.4. The aircraft gas-turbine engine in accordance with claim 1 , wherein the shock-absorbing element is designed as a one-piece shock-absorbing element claim 1 , which has two areas of differing damping characteristics.5. The aircraft gas-turbine engine in accordance with claim 1 , wherein the shock-absorbing element is designed as a combination of a shock-absorbing element with a plastically deforming structural element.6. The aircraft gas-turbine engine in accordance with claim 1 , wherein the radial gap is ...

Turbine blade, turbine, and method of tuning natural frequency of turbine blade

A turbine blade includes: a platform; an airfoil portion extending from the platform in a blade height direction and having a pressure surface and a suction surface extending between a leading edge and a trailing edge; a blade root portion positioned opposite to the airfoil portion across the platform in the blade height direction and having a bearing surface; and a shank positioned between the platform and the blade root portion. The shank has a cross-section which is perpendicular to the blade height direction of the airfoil portion, and in which a line segment connecting a widthwise center position of a leading-edge-side end portion of the shank and a widthwise center position of a trailing-edge-side end portion of the shank is sloped to a center line between a pressure-surface-side contour of the blade root portion and a suction-surface-side contour of the blade root portion.

ROTATING MACHINE AND SEAL RING

A rotating machine includes a rotor which is configured to rotate about an axis, a stator which faces the rotor in a radial direction, and a plurality of seal fins which are provided so as to be protruded from one of the rotor and the stator toward the other of the rotor and the stator in the radial direction and are arranged with gaps therebetween in an axial direction, and one of the rotor and the stator includes an acoustic space which is formed as a hollow portion inside the one of the rotor and the stator, and a communication hole which allows communication between the acoustic space and a part between adjacent seal fins of the plurality of seal fins, the adjacent seal fins being adjacent to each other in the axial direction. 1. A rotating machine comprising:a rotor which is configured to rotate about an axis;a stator which faces the rotor in a radial direction; anda plurality of seal fins which are provided so as to be protruded from one of the rotor and the stator toward the other of the rotor and the stator in the radial direction and are arranged with gaps therebetween in an axial direction, an acoustic space which is formed as a hollow portion inside the one of the rotor and the stator; and', 'a communication hole which allows communication between the acoustic space and a part between adjacent seal fins of the plurality of seal fins, the adjacent seal fins being adjacent to each other in the axial direction., 'wherein one of the rotor and the stator includes2. The rotating machine according to claim 1 , wherein:the stator includes a seal ring main body which has a tubular shape centering on the axis, wherein the plurality of seal fins are provided on an inner circumferential surface of the seal ring main body; andthe acoustic space is formed inside the seal ring main body.3. The rotating machine according to claim 2 , wherein a plurality of the acoustic spaces are formed inside the seal ring main body and arranged with gaps therebetween in a ...

ROTATING BODY PROVIDED WITH BLADES

A rotating body includes a rotating body core, and a plurality of blades provided at an outer or inner circumference of the rotating body core at equal intervals in a circumferential direction, and connected circumferentially via an annular connection portion provided separately from the rotating body core. A resonance frequency under a two nodal diameter number mode of the rotating body is lower than or equal to a rotational secondary harmonic frequency with respect to a rated rotation speed. Where Nrepresents an order of a maximum mistuned component among order components of circumferential distribution of mass, rigidity or natural frequency of the blades, arrangement of the blades satisfies N≧5, and has order components each having ratio less than 1/2, in which the ratio is obtained by dividing the order component by the magnitude of the component of the order N. 1. A rotating body comprising:a rotating body core; anda plurality of blades provided at an outer circumference or an inner circumference of the rotating body core at equal intervals in a circumferential direction, the plurality of blades forming a grouped blade structure in which the blades are connected over the entire circumference via an annular connection portion provided separately from the rotating body core, whereina resonance frequency under a two nodal diameter number mode of the rotating body is lower than or equal to a rotational secondary harmonic frequency with respect to a rated rotation speed of the rotating body, and{'sub': d', 'd', 'd, 'where an order of a maximum mistuned component is defined as Namong order components of mass distribution, rigidity distribution, or natural frequency distribution of the plurality of blades in the circumferential direction, the plurality of blades are arranged so as to satisfy N≧5, and arranged so as to have order components each having a ratio less than 1/2, in which the ratio is obtained by dividing the order component by a magnitude of the component ...

METHOD FOR PROFILING A TURBINE ROTOR BLADE

A method for profiling a turbine rotor blade for an axial flow machine, having the following steps: providing a geometric model of a blade profile, having a camber line of a profile section of the turbine rotor blade; determining boundary conditions for a flow flowing around the turbine rotor blade; changing the camber line such that the flow which is adjusted by the boundary conditions produces the maximum of the difference of the isentropic mach number between the pressure side and the suction side of the turbine rotor blade in a blade section which extends from the blade trailing edge in the direction towards the blade leading edge and the length of which is 65% of the length S of the blade chord. 116.-. (canceled)17. A method for profiling a turbine rotor blade for an axial flow machine , comprising:providing a geometrical model of a blade profile, which has a mean camber line of a profile section of the turbine rotor blade;determining boundary conditions for a flow flowing around the turbine rotor blade;changing the mean camber line in such a way that the flow that is established by the boundary conditions produces the maximum of the difference of the isentropic Mach number between the pressure side and the suction side of the turbine rotor blade in a blade portion that extends from the blade trailing edge in the direction of the blade leading edge and the length of which is 65% of the length S of the blade chord,wherein the mean camber line is formed by a first fourth-degree polynomial, which describes the mean camber line from the blade leading edge to an extreme point, and a second fourth-degree polynomial, which describes the mean camber line from the extreme point to the blade trailing edge, andwherein the extreme point is the point of the mean camber line that is at the maximum distance from the blade chord.18. The method as claimed in claim 17 ,{'sub': S1', '1, 'wherein the first polynomial is formed by using a leading-edge mean camber-line angle, which ...

Damping device for turbine blade assembly and turbine blade assembly having the same

A damping device for a turbine blade assembly is provided. The damping device may include a damper slot formed by first and second slots being axially opposite to each other under respective platforms of adjacent first and second turbine blades of a plurality of turbine blades radially disposed along circumferential surfaces of turbine rotor disks, and a damper pin disposed in the damper slot, wherein the damper pin disposed in the damper slot partially protrudes out of an open inlet of the first slot such that a height of the damper pin protruding out of the open inlet of the first slot is smaller than a gap distance between the first and second turbine blades.

Method to Reduce Noise and Vibration in a Jointed Wind Turbine Blade, and Associated Wind Turbine Blade

A method to reduce noise and vibration between separate blade segments of a jointed wind turbine rotor blade includes determining an actual offset at a chord-wise joint line between the shell members of the first and second blade segments at a load condition on the jointed wind turbine rotor blade, wherein the offset is any one or combination of a flap-wise offset, a twist-wise offset, or a yawl-wise offset. The method defines a modified configuration of the joint structure at a no-load condition on the wind turbine rotor blade that compensates at least partially for the actual offset at the load condition, and the first and second blade segments are connected with the modified configuration of the joint structure. 1. A method to reduce noise and vibration generated by joint structure configuration between a first blade segment and a second blade segment of a jointed wind turbine rotor blade , the first and second blade segments each comprising a shell member , the method comprising:determining an actual offset at a chord-wise joint line between the shell members of the first and second blade segments at a load condition on the jointed wind turbine rotor blade, wherein the offset is any one or combination of a flap-wise offset generated by a flap-wise force, a twist-wise offset generated by a twist-wise force, or a yawl-wise offset generated by a yawl force acting on the first blade segment;defining a modified configuration of the joint structure at a no-load condition on the wind turbine rotor blade that compensates at least partially for the actual offset at the load condition;connecting the first and second blade segments with the modified configuration of the joint structure; andwherein at the load condition, the modified configuration of the joint structure at least partially reduces the actual offset between the shell members of the first and second blade segments.2. The method of claim 1 , wherein the joint structure comprises a beam structure extending span- ...

GAS TURBINE ENGINE AIRFOILS HAVING MULTIMODAL THICKNESS DISTRIBUTIONS

Gas turbine engine (GTE) airfoils, such as rotor and turbofan blades, having multimodal thickness distributions include an airfoil tip, and an airfoil root opposite the airfoil tip in a spanwise direction. The GTE airfoil has a first, second and third locally-thickened region, with the first locally-thickened region defined at the airfoil root. A maximum thickness of each chord between the airfoil root and the airfoil tip transitions toward the leading edge between the first locally-thickened region and the second locally-thickened region, and the third locally-thickened region extends in the spanwise direction. A chord line that extends through the third locally-thickened region contains a first local thickness maxima and a second local thickness maxima interspersed with at least two local thickness minima, and the first local thickness maxima is defined by the third locally-thickened region and is greater than the second local thickness maxima. 1. A gas turbine engine airfoil , comprising:an airfoil tip;an airfoil root opposite the airfoil tip in a spanwise direction, with a span 0% at the root and 100% at the tip;a leading edge;a trailing edge spaced from the leading edge in a chordwise direction; anda first locally-thickened region, a second locally-thickened region, and a third locally-thickened region, the first locally-thickened region defined at the airfoil root,wherein a maximum thickness of each chord between the airfoil root and the airfoil tip transitions toward the leading edge between the first locally-thickened region and the second locally-thickened region, the third locally-thickened region extends in the spanwise direction and is defined between 40% to 80% of the span, and a chord line that extends through the third locally-thickened region contains a first local thickness maxima and a second local thickness maxima interspersed with at least two local thickness minima, and the first local thickness maxima is defined by the third locally-thickened ...

BLADE OR VANE FOR A GAS TURBINE ENGINE

A blade or vane for a gas turbine engine comprising a pressure surface and a suction surface. The pressure surface or the suction surface comprises a roughness zone which is configured to provide greater flow resistance in a direction along the blade or vane than in a direction across the blade or vane. 1. A blade or vane for a gas turbine engine comprising a pressure surface and a suction surface , wherein at least one of the pressure surface and the suction surface comprises a roughness zone which is configured to provide greater flow resistance in a direction (y) along the blade or vane than in a direction (x) across the blade or vane.2. A blade or vane for a gas turbine engine as claimed in claim 1 , wherein the roughness zone is elongate and extends generally along the blade or vane.3. A blade or vane for a gas turbine engine as claimed in claim 1 , wherein the roughness zone comprises a plurality of elongate roughened areas which are spaced apart in a direction along the blade or vane and extend generally in a flow direction across the blade or vane.4. A blade or vane for a gas turbine engine as claimed in claim 3 , wherein the roughened areas are substantially continuous along their length.5. A blade or vane for a gas turbine engine as claimed in claim 3 , wherein a plurality of smooth areas are formed between the roughened areas claim 3 , wherein the smooth areas generally extend in the flow direction across the blade or vane claim 3 , wherein optionally the smooth areas are polished areas and the roughened areas are unpolished areas.6. A blade or vane for a gas turbine engine as claimed in claim 3 , wherein:the roughened areas are areas which have been machined or processed to provide increased roughness compared to a remainder of the suction or pressure surface; orthe roughened areas comprise a roughened coating or roughened element which is applied to the blade or vane.7. A blade or vane for a gas turbine engine as claimed in claim 1 , wherein the ...

Aerofoil array for a gas turbine with anti fluttering means

An aerofoil array for a gas turbine system has an inner annular platform and an outer annular platform, which extend about a longitudinal axis and radially delimit an annular channel for a gas flow; the annular channel houses a plurality of aerofoils, arranged at a substantially constant angular pitch and comprising respective central portions and respective ends connected to the platforms; the aerofoils are formed by two series of aerofoils having a different geometrical feature in order to intentionally vary the eigenfrequencies and arranged about the longitudinal axis with a sequence that is regularly repeated all along the annular channel; even though the external geometry of the aerofoils is varied, the cross-sections remain unchanged in the central portions, at any given radius with respect to the longitudinal axis.

ADDITIVELY MANUFACTURED ROTOR BLADES AND COMPONENTS

A rotor blade formed via additive manufacturing is provided. The rotor blade includes an airfoil and a coupled component. The airfoil includes a plurality of fused layers of a first material formed via additive manufacturing and defines a leading edge and a tip at a distal end. The coupled component includes a plurality of fused layers of a second material formed via additive manufacturing. An interlocking transition zone includes a plurality of projections alternately extending from the airfoil and the coupled component, respectively, to undetachably couple the airfoil and the coupled component. 1. A rotor blade , comprising:an airfoil comprising from a plurality of fused layers of a first material formed via additive manufacturing, wherein the airfoil defines a leading edge and a tip at a distal end;a coupled component comprising from a plurality of fused layers of a second material formed via additive manufacturing; andan interlocking transition zone comprising a plurality of projections alternately extending from the airfoil and the coupled component, respectively, to undetachably couple the airfoil and the coupled component.2. The rotor blade of claim 1 , wherein the coupled component is a protective rotor blade tip claim 1 , and wherein the interlocking transition zone is disposed at the tip of the airfoil.3. The rotor blade of claim 2 , wherein each of the plurality of projections allows relative motion between the airfoil and the protective rotor blade tip claim 2 , such that the protective rotor blade tip is retractable when the rotor blade contacts an object.4. The rotor blade of claim 1 , wherein the coupled component is a protective leading edge component claim 1 , and wherein the interlocking transition zone is disposed at the leading edge of the airfoil.5. The rotor blade of claim 4 , wherein at least a portion of adjacent projections between the airfoil and the protective leading edge component prevent relative motion between the airfoil and the ...

AIRFOIL FOR AXIAL FLOW MACHINE

An airfoil for an axial flow machine includes: an airfoil body extending in a radial direction; a platform (an end wall) provided at an end portion of the airfoil body in the radial direction, the end wall being formed into a plate shape as a wall of a channel in which the airfoil body is installed and which supports the airfoil body; and at least one convex portion formed so as to protrude from a back surface of the platform in a direction away from the airfoil body. The convex portion is formed integrally with a portion for generating a node of a primary vibration mode when an edge portion of the platform vibrates as a free end of the primary vibration mode. 1. An airfoil for an axial flow machine , comprising:an airfoil body extending in a radial direction;an end wall provided at an end portion of the airfoil body in the radial direction, the end wall being formed into a plate shape as a wall of a channel in which the airfoil body is installed and which supports the airfoil body; andat least one convex portion formed so as to protrude from a back surface of the end wall in a direction away from the airfoil body, whereinthe convex portion is formed integrally with a portion for generating a node of a primary vibration mode when an edge portion of the end wall vibrates as a free end of the primary vibration mode and raises a natural frequency of the primary vibration mode.2. The airfoil according to claim 1 , whereinthe convex portion is separated from the edge portion of the end wall.3. The airfoil according to claim 1 , whereinthe convex portion extends toward a portion corresponding to an antinode of the primary vibration mode at the edge portion of the end wall.4. The airfoil according to claim 2 , whereinthe convex portion extends toward a portion corresponding to an antinode of the primary vibration mode at the edge portion of the end wall.5. The airfoil according to claim 1 , whereinthe convex portion is provided individually for each of a plurality of ...

GAS TURBINE ENGINE BLADES WITH AIRFOIL PLUGS FOR SELECTED TUNING

A rotor for use in a gas turbine engine includes a wheel and a plurality of blades. The wheel is arrange about an axis of the gas turbine engine for rotation. The plurality of blades are arranged around the wheel and extend radially outward from the wheel to interact with gases flowing through the engine. 1. A rotor for use in a gas turbine engine , the rotor comprisinga wheel arranged around an axis,a first blade comprising a first material and extending radially outwardly away from the wheel relative to the axis to allow the first blade to interact with gases surrounding the rotor, the first blade having a first external surface, a second external surface opposite the first external surface to define a leading edge, trailing edge, pressure side, and suction side of the first blade, and a first hole that extends through the first external surface and into the first blade in a direction normal to the first external surface, anda first plug located in the first hole and having a first outer surface flush with the first external surface of the first blade such that the first outer surface is exposed to the gases surrounding the rotor, and the first plug comprises a second material that is different than the first material of the first blade.2. The rotor of claim 1 , wherein the first hole is a through hole that extends entirely through the first external surface and the second external surface and wherein the first plug includes a second outer surface that is flush with the second external surface and exposed to the gases surrounding the rotor.3. The rotor of claim 2 , wherein the first blade includes a base located adjacent the wheel and a blade tip spaced apart radially outward from the base and the first hole is located at the blade tip.4. The rotor of claim 3 , wherein the first hole is located at one of the leading edge and the trailing edge of the first blade.5. The rotor of claim 4 , wherein the first material comprises metallic materials and the second ...

GAS TURBINE ENGINE AIRFOIL

A component for a gas turbine engine includes a platform that has a radially inner side and a radially outer side. A root portion extends from the radially inner portion of the platform. An airfoil extends from the radially outer side of the platform. The airfoil includes a pressure side that extends between a leading edge and a trailing edge. A suction side extends between the leading edge and the trailing edge. A bowed tip portion extends perpendicular to a mid-camber line of the airfoil. 1. A component for a gas turbine engine comprising:a platform having a radially inner side and a radially outer side;a root portion extending from the radially inner portion of the platform; and a pressure side extending between a leading edge and a trailing edge;', 'a suction side extending between the leading edge and the trailing edge; and', 'a bowed tip portion extending perpendicular to a mid-camber line of the airfoil., 'an airfoil extending from the radially outer side of the platform, the airfoil including2. The component of claim 1 , wherein the bowed tip portion extends in a circumferential direction and an axial direction.3. The component of claim 2 , wherein the bowed tip portion beings at between 75% and 85% of a span of the airfoil.4. The component of claim 3 , wherein the bowed tip portion begins at 80% of the span of the airfoil.5. The component of claim 1 , wherein said bowed tip portion is bowed between 3 and 19.6 degrees perpendicular to the mid-camber line in a circumferential direction.6. The component of claim 5 , wherein the bowed surface is bowed between 3 and 19.6 degrees perpendicular to the mid-camber line in an axial direction.7. The component of claim 6 , wherein the bowed surface is bowed by the same degree in both the circumferential direction and the axial direction.8. The component of claim 6 , wherein the bowed tip portion follows a curvilinear profile beginning at 3 degrees and increasing to 19.6 degrees.9. The component of claim 6 , wherein a ...

ROTOR BLADE WITH REDUCED ACOUSTIC RESPONSE

A turbine engine rotor blade having a trailing edge (A) with a modified surface state () enabling the flow speed passing around the blade to be altered so as to modify the acoustic interaction against structural elements () that interact with the flow downstream from the rotor blade. 1. A turbine engine rotor having a plurality of rotor blades of wake that impacts against structural elements mounted downstream , each of said rotor blades having a determined zone between a trailing edge and at most 50% of the chord of the blade with a surface state that is modified , wherein said determined zones of two consecutive blades are of different lengths , thereby enabling the speed of the flow over said blades to be altered in such a manner as to modify the acoustic interaction against structural elements interacting with said flow downstream from said rotor blade.2. A rotor according to claim 1 , wherein said determined zone is present over the entire height of the blade.3. A rotor according to claim 1 , wherein said modified surface state is present on the suction side of the blade.4. A rotor according to claim 3 , wherein said modified surface state comprises applying a textured paint or adhesively bonding a material acting as a rough skin claim 3 , thereby slowing down the flow on approaching said trailing edge.5. A rotor according to claim 1 , wherein said modified surface state is present on the pressure side of the blade.6. A rotor according to claim 5 , wherein said modified surface state comprises applying a textured paint or adhesively bonding a material providing a surface that is more effective from an aerodynamic point of view claim 5 , so as to accelerate the flow on approaching said trailing edge.7. A rotor according to claim 6 , wherein said modified surface state presents a “shark's skin” type surface state. The present invention relates to the general field of rotor blades of a turbine engine, and more particularly to rotor blades having a wake that ...

Bladed rotor disk including anti-vibratory feature

A rotor disk includes a ring shaped rotor body defining a radially inward opening, rims protrude radially outward from the rotor body, and outwardly facing rotor blade retention slots are defined between circumferentially adjacent rims. Each slot is operable to receive and retain a corresponding rotor blade, and each rim of the rims includes an anti-vibratory feature. The anti-vibratory feature includes a structure defining an isogrid pattern intruding into a surface of the rim.

HOLLOW FAN BLADE CONSTRAINED LAYER DAMPER

A fan blade having a damping system is provided. The fan blade comprising: a fan blade body having one or more compartments within the fan blade body; a damping material located within at least one of the one or more compartments; one or more partial ribs originating at the fan blade body and extending into at least one of the one or more compartments, wherein each of the one or more partial ribs terminate at a distal end; and one or more damping plates, wherein each of the one or more damping plates are attached to a distal end of a partial rib, wherein the damping material is located between the damping plate and the fan blade body. 1. A fan blade having a damping system , the fan blade comprising:a fan blade body having one or more compartments within the fan blade body;a damping material located within at least one of the one or more compartments;one or more partial ribs originating at the fan blade body and extending into at least one of the one or more compartments, wherein each of the one or more partial ribs terminate at a distal end; andone or more damping plates, wherein each of the one or more damping plates are attached to a distal end of a partial rib, wherein the damping material is located between the damping plate and the fan blade body.2. The fan blade of claim 1 , wherein the one or more compartments includes a first compartment proximate a root end of the fan blade body and a second compartment proximate the tip end of the fan blade body.3. The fan blade of claim 2 , wherein the damping material and each of the one or more damping plates are located in the first compartment.4. The fan blade of claim 2 , wherein the first compartment includes a first sub-compartment located proximate a leading edge of the blade and a second sub-compartment located proximate the trailing edge of the blade.5. The fan blade of claim 4 , wherein the damping material and each of the one or more damping plates are located in the first sub-compartment.6. The fan blade of ...

PULSE BODY MODULE WITH BEARING PROTECTION

An impulse body module a turbine stage of a gas turbine includes a holder component of one-piece construction, with a base and surrounding side walls arranged at the base, wherein the side walls and the base define a holding chamber; an insert component, of one-piece construction, which is inserted into the holding chamber of the holder component, wherein the holder component and the insert component accommodated therein define a number of mutually separated cavities, and wherein an impulse body, in particular a sphere, is accommodated in each cavity; and a closure component of one-piece construction, which is joined to the holder component in a material-bonded manner where the holding chamber is closed and the insert component is surrounded by the holder component and the closure component. The holder component has at least one base projection in the region of its base, which extends away from the holding chamber. 1. An impulse body module for a turbomachine , in particular a turbine stage of a gas turbine , comprising:a holder component of one-piece construction, with a base and surrounding side walls arranged at the base, wherein the side walls and the base define a holding chamber,an insert component of one-piece construction, which is inserted into the holding chamber of the holder component, wherein the holder component and the insert component accommodated therein are constructed in such a way that they together define a plurality of mutually separated cavities and wherein an impulse body, in particular a sphere, is accommodated in each cavity, anda closure component of one-piece construction, which is joined to the holder component in a material-bonded manner in such a way that the holding chamber is closed and the insert component is surrounded by the holder component and the closure component,wherein the holder component has at least one base projection in the region of its base or of the closure component, which extends away from the holding chamber.2. ...

Turnbuckle dampening links

An elastomeric dampening link including two or more interconnected dampening bushings mounted around corresponding ones of two or more adjustable length turnbuckles linked to devices and elastomeric dampening link includes at least one bar connecting adjacent ones of the dampening bushings. Clamping bands of clamp may be clamped around each of the dampening bushings. Bar may be in tension between the dampening bushings. Each of the turnbuckles may include a rod disposed in a corresponding one of the dampening bushings and having distal hollow internally threaded first end, first eyelet mounted on the first externally threaded shank adjustably screwed into internally threaded first end. One of the bushings may be substantially solid, have a rectangular slot extending inwardly from an annular surface of bushing, and the rod may have six sided surface with two opposite sides slidingly engaging and fitting snugly in the slot. Turnbuckles and links may be used to actuate variable stator vanes.

Vibration-Damped Composite Airfoils and Manufacture Methods

A turbine engine component () comprises a fiber structure () forming at least a portion of an airfoil (). A matrix () embeds the fiber structure. A carbon nanotube filler () is in the matrix. 1100. A turbine engine component () comprising:{'b': 125', '126', '102, 'a fiber structure (, ) forming at least a portion of an airfoil ();'}{'b': '128', 'a matrix () embedding the fiber structure; and'}{'b': '130', 'a carbon nanotube filler () in the matrix.'}2. The component of wherein:{'b': '130', 'the carbon nanotube filler () in the matrix exists through a thickness of at least 3 plies of the fiber structure.'}3. The component of wherein:the fiber structure forms at least 30% by volume of a composite portion of the component.4. The component of wherein:the fiber structure forms 45-65% by volume of a composite portion of the component.5. The component of wherein:the airfoil is an airfoil of a turbine engine blade.6. The component of wherein:the airfoil is an airfoil of a turbofan engine fan blade.7. The component of wherein:the airfoil is an airfoil of a turbine engine vane.8. The component of wherein:the airfoil is an airfoil of a turbofan engine fan vane.9. The component of wherein:the fiber structure comprises at least 50% carbon fiber by weight.10. The component of wherein:the fiber structure comprises one or more woven members.11. The component of wherein:the matrix comprises a cured resin.12. The component of wherein:the carbon nanotube filler has a content of 0.05-0.49% in the matrix by weight.13. The component of wherein:the carbon nanotube filler has a characteristic diameter of 0.5 nanometer to 5 nanometers; andthe carbon nanotube filler has a characteristic length of 10 nanometers to 100 nanometers.14. The component of wherein:{'b': '130', 'the carbon nanotube filler () in the matrix is in a multi-ply thickness of the fiber structure, inter-ply and intra ply.'}15. The component of wherein:{'b': 130', '124', '123, 'the carbon nanotube filler () in the matrix is ...

Damping inlay for turbine blades

The invention concerns a turbine blade comprising a surface, a recess within the surface, and a damping inlay within the recess. The damping inlay comprises a chamber with a damping material, for example particles. The damping inlay should substantially maintain the aerodynamic profile of the blade to enable normal operation. A further embodiment of the invention describes the method of manufacture of a turbine blade with a damping inlay. The method comprises the steps of manufacturing a turbine blade having a surface and a recess in the surface, and providing one or more damping inlays within the recess such that the damping inlay substantially maintains the aerodynamic profile of the blade, the damping inlay comprising a chamber and a damping material disposed within the chamber.

TURBINE BLISK INCLUDING CERAMIC MATRIX COMPOSITE BLADES AND METHODS OF MANUFACTURE

In some embodiments, an apparatus includes a disk, a coupling member and a set of blades. The coupling member has a first surface and a second surface, and defines a set of openings between the first surface and the second surface. The first surface is configured to be coupled to the outer surface of the disk. A portion of each blade from the set of blades is disposed within an opening from the set of openings when the first surface of the coupling member is coupled to the outer surface of the disk such that the blade is coupled to the disk. 1. A turbine disk assembly adapted for use in a gas turbine engine , the assembly comprisinga disk comprising metallic materials and forming an outer surface, the outer surface being a continuous circumferential surface that extends around a central axis that defines an axial direction parallel thereto and a radial direction perpendicular thereto,a blade comprising ceramic matrix composite materials, the blade including a root arranged radially outward of the outer surface formed by the disk and an airfoil extending radially outward from the root, anda ring coupled to the disk and extending around the outer surface formed by the disk, the ring formed to include an opening that receives at least a portion of the root included in the blade to couple the blade to the ring while at least a portion of the airfoil included in the blade is arranged outside of the opening radially outward of the ring to interact with gasses radially outward of the ring.2. The assembly of claim 1 , wherein a diameter of the outer surface formed by the disk is greater than a radially-inner surface of the ring when the disk and the ring are at the same temperature so that the disk and the ring are interference fit with one another.3. The assembly of claim 2 , wherein the ring is bonded to the disk via a braze joint arranged between the radially-inner surface of the ring and the outer surface formed by the disk.4. The assembly of claim 3 , wherein the root ...

Gas turbine engine

A gas turbine engine includes in axial flow series a fan rotor, a series of outlet guide vanes for guiding flow from the fan rotor; and a bifurcation. The gas turbine engine further includes a substantially annular fluid passageway extending from fan to the bifurcation, the outlet guide vanes being positioned within the passageway. The passageway includes a profiled region positioned upstream of the bifurcation, the profiled region including a first circumferential portion positioned adjacent a second circumferential portion, the first circumferential portion having a first average radial thickness and a second circumferential portion having a second average radial thickness, the first average radial thickness being smaller than second average radial thickness. The profiled region of the passageway is configured to modify the flow through the passageway so as to improve uniformity of a static pressure field from immediately upstream of the bifurcation to just downstream of the fan rotor.

Combination for sealing a gap between turbomachine blades and for reducing vibrations of the turbomachine blades

A combination including a seal (-) for sealing a gap (s) between blade platforms () of two adjacent blades of a turbomachine and a reducer (-) for reducing vibrations of at least one of the blades, the seal including at least one rib () having a rib thickness and at least one wall () having a wall thickness that is smaller than the rib thickness and/or the reducer including a tuning-element guide housing () having at least one cavity () in which at least one tuning element () is disposed with play for impacting contact with the tuning-element guide housing. 110-. (canceled)11. A combination comprising:a seal for sealing at least one gap between blade platforms of two adjacent blades of a turbomachine and a reducer for reducing vibrations of at least one of the blades, the seal including at least one rib having a rib thickness and at least one wall having a wall thickness that is smaller than the rib thickness or the reducer including a tuning-element guide housing having at least one cavity in which at least one tuning element is disposed with play for impacting contact with the tuning-element guide housing.12. The combination as recited in wherein the seal and the reducer are interconnected.13. The combination as recited in wherein the seal and the reducer are interconnected by a web.14. The combination as recited in wherein the seal and the reducer and formed in one piece or integrally with each other.15. The combination as recited in wherein the seal and the reducer are unconnected.16. The combination as recited in wherein the seal or the reducer is at least partially produced using a generative manufacturing process or is at least partially made of a nickel alloy.17. The combination as recited in wherein the combination has the tuning-element guide and the tuning-element guide housing is at least partially produced using the generative manufacturing process or is at least partially made of the nickel alloy.18. The combination as recited in wherein the seal ...

AEROFOIL

An aerofoil having a leading edge and a trailing edge, the leading edge including a plurality of slits extending toward the trailing edge, such that the leading edge is defined by alternating peaks and troughs. Each peak extends in a generally spanwise direction and defines a peak width, each peak being separated from an adjacent peak in the spanwise direction by a trough. Each trough extends in the generally spanwise direction and is spaced in a chordwise direction from the peak, each trough defining a trough width. A ratio of the peak width to the trough width is between 4:1 and 10:1. 1. An aerofoil having a leading edge and a trailing edge , the leading edge comprising a plurality of slits extending toward the trailing edge , such that the leading edge is defined by alternating peaks and troughs;each peak and defining a peak width extending in a generally spanwise direction, each peak being separated from an adjacent peak in the spanwise direction by a trough;each trough being spaced in a chordwise direction from the peak, each trough defining a trough width extending in the spanwise direction; wherein a ratio of the peak width to the trough width is between 4:1 and 10:1.2. An aerofoil according to claim 1 , wherein each slit comprises a generally chordwise extending side surface provided at each end of each peak claim 1 , interconnecting each peak with an adjacent trough.3. An aerofoil according to claim 1 , wherein each peak defines a generally spanwise extending end surface.4. An aerofoil according to claim 2 , wherein each side surface extends orthogonally to each end surface.5. An aerofoil according to claim 1 , wherein each peak comprises a generally chevron shaped end surface.6. An aerofoil according to claim 1 , wherein each side surface comprises a convex curve extending in a generally chordwise direction.7. An aerofoil according to claim 1 , wherein opposing side surfaces of each slit are angled inwardly toward one another in a downstream direction.8. ...

MODIFIED J TYPE CANTILEVERED VANE AND GAS TURBINE HAVING THE SAME

A modified J-type cantilevered vane reduces rubbing of an airfoil and enhances the vibration stability of a vane hub and the fluidity of fluid. The cantilevered vane includes a root attachment inserted into a slot formed in an inner circumferential surface of a casing of a gas turbine, and an airfoil vertically extending from the root attachment to a predetermined height. The airfoil includes a linear part extending in a radial direction from the inner circumferential surface of the casing toward a rotating shaft of the gas turbine, and a curved part integrally formed with the linear part, the curved part including a bend beginning from one end of the linear part and curving in a circumferential direction of the rotating shaft. Overall, the linear part and the curved part are biased in the circumferential direction, since the airfoil has a J-shaped X-axis profile and a C-shaped Y-axis profile. 1. A cantilevered vane comprising:a root attachment configured to be inserted into a slot formed in an inner circumferential surface of a casing of a gas turbine; and a linear part extending in a radial direction from the inner circumferential surface of the casing toward a rotating shaft of the gas turbine, and', 'a curved part integrally formed with the linear part, the curved part including a bend beginning from one end of the linear part and curving in a circumferential direction of the rotating shaft,, 'an airfoil vertically extending from the root attachment to a predetermined height, the airfoil includingwherein the linear part and the curved part are biased in an axial direction of the rotating shaft.2. The cantilevered vane according to claim 1 , further comprising a rounded structure formed at a junction between the root attachment and the airfoil.3. The cantilevered vane according to claim 2 , wherein the rounded structure has a curvature radius corresponding to a length ranging from 10% to 35% of a width of the root attachment.4. The cantilevered vane according to ...

STRUCTURE FOR MITIGATING VIBRATORY MODES OF COUNTER-ROTATING ENGINE ROTORS

A gas turbine engine including a first rotor assembly is generally provided. The first rotor assembly includes outer drum and an outer drum airfoil. The outer drum airfoil is coupled to the outer drum and extended inward along a radial direction. A damper structure is coupled to one or more of the outer drum or the outer drum airfoil. 1. A gas turbine engine , comprising:a first rotor assembly comprising an outer drum and an outer drum airfoil, wherein the outer drum airfoil is coupled to the outer drum and extended inward along a radial direction, and further comprising a damper structure is coupled to one or more of the outer drum or the outer drum airfoil.2. The gas turbine engine of claim 1 , wherein the damper structure defines a substantially annular ring claim 1 , the annular ring defining springing properties generating an outward force along the radial direction.3. The gas turbine engine of claim 2 , wherein the damper structure defines an axial split through the annular ring such as to define a first end and a second end of the damper structure.4. The gas turbine engine of claim 2 , wherein the damper structure defines a first portion extended generally co-directional to a portion of the outer drum or the outer drum airfoil to which the damper structure is coupled.5. The gas turbine engine of claim 4 , wherein the damper structure defines one or more radii at the first portion extended at least partially inward along the radial direction claim 4 , wherein the one or more radii enable springing of the damper structure in response to an axial load onto the damper structure.6. The gas turbine engine of claim 1 , wherein the damper structure is disposed at an inner radius of the outer drum.7. The gas turbine engine of claim 6 , wherein the damper structure is disposed outward along the radial direction of the outer drum airfoil.8. The gas turbine engine of claim 7 , wherein the outer drum airfoil defines an arm extended at least partially along the radial ...

ROTOR ASSEMBLY WITH ACTIVE DAMPING FOR GAS TURBINE ENGINES

An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil section that extends from a root section. The airfoil section extends between a leading edge and a trailing edge in a chordwise direction and extends between a tip portion and the root section in a radial direction. The airfoil section defines a pressure side and a suction side separated in a thickness direction, and the airfoil section includes a metallic sheath that defines an internal cavity receiving a composite core. The root section defines at least one bore dimensioned to receive a retention pin. At least one damping element is received in the internal cavity and that selectively causes the airfoil section to stiffen. 1. An airfoil for a gas turbine engine comprising:an airfoil section extending from a root section;wherein the airfoil section extends between a leading edge and a trailing edge in a chordwise direction and extends between a tip portion and the root section in a radial direction, the airfoil section defines a pressure side and a suction side separated in a thickness direction, and the airfoil section includes a metallic sheath that defines an internal cavity receiving a composite core;wherein the root section defines at least one bore dimensioned to receive a retention pin; andat least one damping element received in the internal cavity and that selectively causes the airfoil section to stiffen.2. The airfoil as recited in claim 1 , wherein the at least one damping element comprises a piezeoelectic material.3. The airfoil as recited in claim 2 , wherein the at least one damping element abuts against the core.4. The airfoil as recited in claim 2 , further comprising at least one sensor received in the internal cavity claim 2 , wherein the at least one damping element is responsive to a control signal based on information from the at least one sensor.5. The airfoil as recited in claim 1 , wherein:the sheath includes a ...

MISTUNED AIRFOIL ASSEMBLIES

A mistuned airfoil assembly for a gas turbine engine is disclosed. The mistuned airfoil assembly may comprise a hub and airfoils extending radially from the hub. The airfoils may consist of first airfoils and at least one second airfoil. The first airfoils may be formed from a first titanium alloy and the at least one second airfoil may be formed from a second titanium alloy. The first titanium alloy and the second titanium alloy may have different natural frequencies. 1. A method of manufacturing a mistuned airfoil assembly for a gas turbine engine , comprising:providing a hub;assembling a plurality of first airfoils to the hub;assembling at least one second airfoil to the hub;wherein the plurality of first airfoils and the at least one second airfoil are formed from different alloys and have different natural frequencies, the at least one second airfoil configured with a natural frequency to prevent staged flutter vibration of the mistuned airfoil assembly; andwherein the plurality of first airfoils and the at least one second airfoil have different densities.2. The method of claim 1 , wherein the at least one second airfoil is a plurality of second airfoils claim 1 , and the plurality of first airfoils and the plurality of second airfoils are arranged in an alternating sequence about the hub.3. The method of claim 1 , wherein the at least one at least one second airfoil consists of a single second airfoil.4. The method of claim 1 , wherein first airfoils and the at least one second airfoil have identical geometries.5. The method of claim 1 , wherein the plurality of first airfoils are each formed from a first titanium alloy and the at least one second airfoil is formed from a second titanium alloy claim 1 , the first titanium alloy and the second titanium alloy having different natural frequencies.6. The method of claim 5 , wherein a difference between a natural frequency of the first titanium alloy and a natural frequency of the at least one second titanium ...

Alternating Vane Asymmetry

A vane assembly for a gas turbine engine may include a plurality of vanes being arranged in vane groupings symmetrically spaced circumferentially from each other. Each vane grouping may include at least a first and a second vane. The at least first and second vanes may be spaced from each other at a first pitch. Each vane grouping may be spaced from each other at a second pitch. The first pitch may be dissimilar from the second pitch. 1. A vane assembly for a gas turbine engine , the vane assembly comprising:a plurality of vanes being arranged in vane groupings symmetrically spaced circumferentially from each other, each vane grouping including at least a first and a second vane, the at least first and second vanes being spaced from each other at a first pitch, each vane grouping being spaced from each other at a second pitch, the first pitch being dissimilar from the second pitch.2. The vane assembly of claim 1 , wherein the first pitch is less than the second pitch.3. The vane assembly of claim 1 , wherein the at least first vane has an airfoil shape that is dissimilar to an airfoil shape of the at least second vane.4. The vane assembly of claim 1 , wherein the at least first vane is offset axially downstream from the at least second vane.5. The vane assembly of claim 4 , wherein each vane grouping further includes at least a third vane being spaced from the at least second vane at a third pitch claim 4 , the third pitch being dissimilar from both the first pitch and the second pitch.6. A gas turbine engine claim 4 , the engine comprising:a compressor;a combustor downstream of the compressor; anda turbine downstream of the combustor, one of the compressor and the turbine including a vane assembly, the vane assembly including a plurality of vanes being arranged in vane groupings symmetrically spaced circumferentially from each other, each vane grouping including at least a first and a second vane, the at least first and second vanes being spaced from each other at ...

TURBOMACHINE BLADE SYSTEM

A turbomachine blade system, in particular for a compressor or turbine stage of a gas turbine, which includes at least one blade, in particular a moving or guide blade, and at least one moving body for reducing the vibrations of this blade, at least one area of a guide for guiding the body and/or at least one area of a supporting structure for resiliently mounting the body and/or at least one area of the body being or becoming generatively manufactured together with at least one area of the blade, in particular of a vane and/or blade root and/or a shroud situated thereon. 1. A turbomachine blade system comprising:at least one blade; andat least one moving body for reducing the vibrations of the blade, at least one area of a guide for guiding the body or at least one area of a supporting structure for resiliently mounting the body or at least one area of the body being generatively manufactured together with at least one area of the blade.2. The turbomachine blade system as recited in wherein the body is an impulse body for reducing the vibrations of the blade through multiple impact contacts.3. The turbomachine blade system as recited in wherein the generatively manufactured area of the blade has a contact surface for contacting the body.4. The turbomachine blade system as recited in wherein the generatively manufactured area of the blade claim 1 , the generatively manufactured area of the guide claim 1 , the generatively manufactured area of the supporting structure or the generatively manufactured area of the body includes metal.5. The turbomachine blade system as recited in wherein the generatively manufactured area of the blade claim 1 , the generatively manufactured area of the guide claim 1 , the generatively manufactured area of the supporting structure or the generatively manufactured area of the body is manufactured from metal powder.6. The turbomachine blade system as recited in wherein the guide has an open or closed hollow space claim 1 , the body being ...

TURBINE BLADE RAIL DAMPER

A device for damping of vibratory energy in the blades of rotor assemblies during operation where the blades have a shroud attached thereto with at least one sealing rail extending radially outward from the shroud to an outer diameter surface. A damper element is attached to the turbine blade sealing rail extending radially inward from the rail outer diameter surface along rail sides to maintain the damper element out of the flow of gas and positioned at a radial location on the blade for damping. 1. A device for damping vibratory energy in a rotor assembly during operation , comprising:a first turbine blade having a shroud with a first sealing rail, the first sealing rail having a first slot at a radial face of the first sealing rail, wherein the first slot has radially inner and radially outer surfaces that are generally perpendicular to the radial face of the first sealing rail;a second turbine blade adjacent the first blade and having a shroud with a second sealing rail, the second sealing rail having a second slot at a radial face of the second sealing rail such that the first slot is adjacent and opposing the second slot, wherein the second slot has radially inner and radially outer surfaces that are generally perpendicular to the radial face of the second sealing rail, and wherein the radial face of the first sealing rail abuts the radial face of the second sealing rail; anda damper element positioned in and extending between the first and second slots.2. The device of claim 1 , wherein the damper element is made from metal or ceramic.3. The device of claim 1 , wherein the first and second slots at the radial faces of the first and second turbine blades are positioned between the shroud and an outer surface of the first and second sealing rails to keep the damper element out of the flow of gas.4. The device of claim 1 , wherein the damper element is generally “U” shaped claim 1 , and wherein the “U” has a flat center portion that engages an end face-of the ...

MISTUNED FAN

A compressor rotor for a gas turbine engine is described which includes sets of blades having different airfoil thickness distributions, each including a frequency modifier forming a thickness differential relative to a baseline blade thickness. The frequency modifiers provide different natural vibration frequencies for each of the blades, and facilitate modifying natural vibration frequency separation between adjacent blades of the compressor rotor. 1. A mistuned compressor rotor assembly for a gas turbine engine , the mistuned compressor rotor assembly comprising a hub to which a plurality of airfoil blades are mounted , the airfoil blades each having an airfoil selected from at least first and second airfoil types and arranged as generally alternating with one another around the circumference of the rotor , the first airfoils having an airfoil thickness less than an airfoil thickness of the second airfoils at a first selected span of the respective blades , and the second airfoils having an airfoil thickness less than an airfoil thickness of the first airfoil at a second selected span of the respective blades different from the first selected span.2. The mistuned compressor rotor assembly of claim 1 , wherein the first and second airfoils have substantially identical thickness distribution profiles but for in regions immediately adjacent the first and second selected spans.3. The mistuned compressor rotor assembly of claim 1 , wherein the first airfoil thickness at the first selected span at least partially provides the first airfoil blade with a lower natural vibration frequency than the second airfoil blade.4. The mistuned compressor rotor assembly of claim 3 , wherein the second airfoil thickness at the second selected span at least partially provides the second airfoil blade with a higher natural vibration frequency than the first airfoil blade.5. The mistuned compressor rotor assembly of claim 4 , wherein the second selected span corresponds in use to a span ...

VIBRATION DAMPER FOR A TURBOMACHINE ROTOR VANE

A turbomachine rotor has a disk carrying vanes, each vane having a blade linked by a platform to a root. For at least one vane, a recess is defined between the platform and the disk, and a vibration damper is mounted in the recess. The vibration damper includes a first structural portion configured to contact the platform of which the vibrations are to be dampened, and a second mass portion configured to dampen these vibrations. The second mass portion is a powder and the first structural portion is a box containing the powder. 1. A rotor for a turbomachine , comprising: a disk carrying vanes , each vane comprising a blade connected by a platform to a root , wherein for at least one of the vanes , a recess is defined between the platform and the disk , and a vibration damper is mounted in the recess , the vibration damper comprising a first structural portion configured to contact the platform , and a second mass portion configured to dampen vibrations , wherein the second mass portion is a powder and the first structural portion is a box containing the powder.2. The rotor according to claim 1 , wherein the box is closed in a sealed manner to prevent the powder from escaping from the box.3. The rotor according to claim 1 , wherein the powder occupies an entirety of an internal volume of the box.4. The rotor according to claim 1 , wherein the powder occupies less than an entirety of an internal volume of the box.5. The rotor according to claim 1 , wherein the box has a parallelpiped shape.6. The rotor according to claim 1 , wherein the box and the powder are made of a same metallic material.7. The rotor according to claim 1 , wherein the box is a product of melting a powder identical to that contained in the box.8. The rotor according to claim 1 , wherein the box comprises a plurality of fallen edges configured to ensure that the box remains in the recess when the turbomachine is stopped.9. A turbomachine comprising the rotor according to .10. The rotor according to ...

Systems and Methods to Control Combustion Dynamic Frequencies

Systems and methods for frequency separation in a gas turbine engine are provided herein. The systems and methods for frequency separation in a gas turbine engine may include determining a hot gas path natural frequency, determining a combustion dynamic frequency, and modifying a compressor discharge temperature to separate the combustion dynamic frequency from the hot gas path natural frequency.

Blade comprising a shank, provided with a depressed portion

The shank of a blade includes a first side and a second side provided with a depression. By concentrating this depression on a single side, a shank that is more rigid and resistant with an equal lightening is obtained, with the stress concentrations depending on the depth of the depression but being absent from first flat side. The vibrations and bending deformations are also reduced.

MISTUNED CONCENTRIC AIRFOIL ASSEMBLY AND METHOD OF MISTUNING SAME

An example method of reducing a vibratory response of airfoils that support a shroud includes circumferentially misaligning some of the airfoils in a radially inner array of airfoils with all airfoils of a radially outer array of airfoils, and circumferentially aligning at least one of the airfoils in the radially inner array of airfoils with all airfoils of the radially outer array of airfoils. 1. An airfoil assembly , comprising:an annular shroud having a radially inner face and a radially outer face opposing the radially inner face;a radially inner array of airfoils extending from the radially inner face; anda radially outer array of airfoils extending from the radially outer face, wherein at least one airfoil of the radially inner array is circumferentially aligned with a corresponding airfoil in the radially outer array, and at least one airfoil of the radially inner array is circumferentially misaligned with the airfoils of the radially outer array.2. The assembly of claim 1 , wherein one airfoil of the radially inner array is circumferentially aligned with a corresponding airfoil in the radially outer array at a twelve o'clock position claim 1 , and the remaining airfoils in the radially inner array are misaligned with the airfoils of the radially outer array.3. The assembly of claim 1 , wherein the radially inner array of airfoils are configured to guide flow within a radially inner bypass flow passage of a gas turbine engine claim 1 , and the radially outer array of airfoils are configured to guide flow within a radially outer bypass flow passage of the gas turbine engine.4. The assembly of claim 3 , wherein the radially inner bypass flow passage and the radially outer bypass flow passage are both radially outside a core flow passage of the gas turbine engine.5. The assembly of claim 1 , wherein each of the radially inner array of airfoils at or above a horizontal midline is circumferentially aligned with a corresponding airfoil in the radially outer when ...

SQUEEZE FILM DAMPER FOR A GAS TURBINE ENGINE

Squeeze film damping systems and methods therefore are provided that include features for optimizing the damping response to vibrational loads experienced by a rotary component of a gas turbine engine. In one exemplary aspect, a damping system actively controls a dynamic sleeve to adjust the damping response. In particular, the dynamic sleeve is disposed within a chamber defined by a damper housing. The damping system controls the damper gap by translating the dynamic sleeve. When the dynamic sleeve is translated, a variable damper gap is varied, allowing for fluid to squeeze into or out of the damper gap, thereby adjusting the damping response to the vibration of the rotary component. 120-. (canceled)21. A damping system for damping vibrations of a rotary component operatively coupled thereto , the damping system comprising:a damper housing having an inclined surface and defining a chamber;a dynamic sleeve movable within the chamber, the dynamic sleeve having an inclined surface complementary to the inclined surface of the damper housing, the inclined surface of the dynamic sleeve and the inclined surface of the damper housing defining a damper gap therebetween, andwherein when the sleeve is moved within the chamber, the damper gap is varied such that a damping response of the damping system is varied.22. The damping system of claim 21 , wherein the dynamic sleeve extends between a first end and a second end along an axial direction and has a first end portion at the first end claim 21 , and wherein the dynamic sleeve has a first wall extending between the first end and the second end claim 21 , the inclined surface of the dynamic sleeve forms a portion of the first wall claim 21 , and wherein the first wall along the first end portion is not inclined with respect to the axial direction.23. The damping system of claim 22 , further comprising:one or more seals operable to engage the first end portion of the dynamic sleeve throughout the movement of the dynamic ...

BALANCER

A rotating object balancer may include an insert configured for securing to an object and having a chamber for placement of balancing elements and a curable solution, a rotational system configured to rotate the object at high-speed, a monitoring system for monitoring vibrations of the object rotating at high-speed, and a curing tool configured for curing the curable solution when the object has reached a balanced state. A method of balancing may be performed by the balancer and a balanced object may be created by the balancer. 1. A rotating object balancer , comprising:an insert configured for securing to an object and having a chamber for placement of balancing elements and a curable solution;a rotational system configured to rotate the object at high-speed;a monitoring system for monitoring vibrations of the object rotating at high-speed; anda curing tool configured for curing the curable solution when the object has reached a balanced state.2. The balancer of claim 1 , wherein the curing tool comprises an ultraviolet light emitting device.3. The balancer of claim 2 , wherein the curing tool comprises a heating element.4. The balancer of claim 1 , wherein the object is a propeller.5. The balancer of claim 4 , wherein the insert is configured for engaging the hub of the propeller.6. The balancer of claim 1 , wherein the rotational system comprises a universal switch configured to isolate vibrations from a drive system so as to avoid affecting vibration of the object.7. The balancer of claim 6 , wherein the universal switch comprises a drive receiving portion claim 6 , a central portion pivotable relative to the drive receiving portion about a first axis claim 6 , and a top portion pivotable relative to the central portion about a second axis.8. The balancer of claim 7 , wherein the drive receiving portion defines a longitudinal axis and the first axis and second axis are substantially perpendicular to the longitudinal axis.9. The balancer of claim 8 , wherein the ...