ПРИВОД ВСПОМОГАТЕЛЬНОГО ОБОРУДОВАНИЯ

Изобретение относится к приводу вспомогательного оборудования, такого как топливный насос или смазочный насос в газотурбинном двигателе. Привод содержит электродвигатель и характеризуется тем, что дополнительно содержит воздушную турбину (22), связанную с электродвигателем. Эта воздушная турбина предназначена для питания воздушным потоком, отбираемым от компрессора газотурбинного двигателя, и участвует в приведении в действие указанного вспомогательного оборудования. Привод содержит также клапан регулирования расхода воздуха, отбираемого от компрессора, причем во время фазы запуска газотурбинного двигателя этот клапан находится в закрытом положении. Такое выполнение привода позволит повысить его надежность. 12 з.п. ф-лы, 10 ил.

ПАРОВАЯ ТУРБИНА

ПАРОВАЯ ТУРБИНА

СТУПЕНЬ ТЕПЛОТУРБИНЫ

Использование: в турбинах на тепловых электростанциях, автомобильном и водном транспорте, в авиации. Сущность: ступень турбины содержит установленные в корпусе входные сопла, ротор с рабочими лопатками, последние установлены на диске и имеют периферийный обод. Выхлопная полость выполнена кольцевой, а примыкающий к ней торец корпуса выполнен в виде диска с выходными окнами и перемычками между ними. Выходные окна сообщают межлопаточные каналы с выхлопной полостью, перемычки расположены против сопел. 7 ил.

ПАРОВАЯ ТОРОИДАЛЬНАЯ ТУРБИНА А.М.РЕПИНА

Использование: в турбостроении. Сущность изобретения: паровая паровозная турбина содержит корпус, ротор, диафрагмы, причем ротор в виде облопаченного диска с закрепленными на вершинах лопаток цельного кольца, несущего дополнительные ряды рабочих лопаток, причем в статоре направление потока пара организовано по винтовой тороидальной линии с подачей пара на рабочие лопатки многократно. 1 с.п. ф-лы, 4 з.п. ф-ды. 6 ил.

ПАРОВАЯ ТУРБИНА (ВАРИАНТЫ)

... 1. Паровая турбина, содержащая:первую ступень, содержащую диафрагму, имеющую внутреннюю перемычку, наружное кольцо и расположенные между ними неподвижные лопатки, при этом наружное кольцо имеет выступающий элемент, проходящий в осевом направлении вниз по потоку и расположенный над концевыми участками рабочих лопаток, образуя часть ступени турбины, причем рабочие лопатки имеют верхнюю по потоку сторону и нижнюю по потоку сторону, а пар протекает через ступень в первом направлении от верхней по потоку стороны к нижней по потоку стороне рабочих лопаток,по меньшей мере один паз, выполненный на поверхности выступающего элемента и предназначенный для отвода части пара в паровом тракте перед рабочими лопатками ступени турбины и направления указанной части пара в обход рабочих лопаток указанной первой ступени, иуплотнительное устройство, расположенное на выступающем элементе в первом местоположении и предназначенное для уплотнения концевых частей рабочих лопаток,причем указанный по меньшей мере ...

ПРИВОД ВСПОМОГАТЕЛЬНОГО ОБОРУДОВАНИЯ

... 1. Привод топливного насоса или смазочного насоса в газотурбинном двигателе (10), содержащий электродвигатель (21), снабженный статором (43) и ротором (42), отличающийся тем, что дополнительно содержит воздушную турбину (22), снабженную корпусом (35) и вращающимся блоком (30, 36, 50) и выполненную, для участия в приведении указанного насоса, с возможностью питания воздушным потоком, отбираемым от компрессора (12) газотурбинного двигателя (10). 2. Привод по п.1, отличающийся тем, что дополнительно содержит клапан (25) регулирования расхода воздуха, отбираемого от компрессора (12), причем этот клапан находится в закрытом положении во время фазы запуска газотурбинного двигателя (10) и в открытом положении после завершения запуска. 3. Привод по п.2, отличающийся тем, что выполнен с возможностью отбора воздуха от компрессора (12) с расходом, достаточным для обеспечения, в случае отсутствия электропитания или отказа указанного электродвигателя (21), функционирования насоса (13) с приводом только ...

Многоступенчатая осевая турбина

Парциальная турбина

Осевая турбина

Отсек ступеней низкого давления паровой турбины

AXIAL-FLOW TURBOMACHINE

Осевая турбомашина

TURBINE WHEEL HAVING BUCKETS OR BLADES MACHINED INTO THE OUTER CIRCUMFERENCE OF THE WHEEL

Impulse turbine design

An impulse turbine arrangement comprises annular offset vaneless ducts 12 and 13 which position the nozzle (guide vane) rows 8, 9 at a larger radius from the rotor axis than the rotor blades 5. The guide vanes are designed to operate in periodically reversing flows, and rotor blades may have an unconventionally high turning angle of 70 degrees to obtain peak efficiency. Performance may be further enhanced by the incorporation of a boundary layer blowing system into the guide vane design. The system may include blowing holes or slots as well as a compressor to raise the blowing pressure and/or mass flow rate.

Improvements relating to impulse turbines

... 1,149,616. Boundary layer control. ESCHER WYSS A.G. 20 April, 1966 [22 April, 1965], No. 17277/66. Heading F2R. [Also in Division F1] To improve the efficiency of a gas or steam turbine of impulse type having guide blades 11 directing working fluid on to rotor blades 21, the circumferentially extending surfaces SW 2i and SW 2a extending between the rotor blade roots and tips to define the flow passage through the rotor 20 are of radii with respect to the corresponding surfaces SW 1i and SW 1a of the guide blade ring 10 such as to extend into the boundary layer flow zones of the fluid delivered by the guide blades. The radial offset h, Fig. 3, between the adjacent rotor and guide ring surfaces is between 0À5 and 2 times where # 0 is the thickness of the boundary layer adjacent the guide ring surfaces, #S is the axial gap between the guide ring 10 and rotor 20 and 1 is the angle between the direction of fluid flow at the guide ring outlet and a plane normal to the rotational axis of the turbine ...

Mehrdruck-Dampfturbine zur Krafterzeugung

Mehrdruck-Dampfturbine zur Krafterzeugung, die für jeden der Dampfströme mindestens einen eigenen Düsenkasten (11', 11'') vorsieht, welche gemeinsam auf die Laufschaufelreihe (5) des Turbinenlaufrades (4) ausblasen und wirken; die Düsenkästen (11', 11'') sind, bezogen auf die Turbinenwelle (1), sektoral angeordnet, wobei die Düsenwinkel (18', 18'') der Düsenkästen derart angepasst sind, dass sich minimierte Verluste ergeben.

Mehrdruck-Dampfturbine zur Krafterzeugung

Mehrdruck-Dampfturbine zur Krafterzeugung, die für jeden derDampfströme mindestens einen eigene Düsenkasten (11', 11'')vorsieht, welche gemeinsam auf die Laufschaufelreihe (5) mitGleichdruck-Beschaufelung des einzigen Turbinenlaufrades (4)ausblasen und wirken; die Düsenkästen (11) sind, bezogen auf dieTurbinenwelle (1), sektoral und/oder konzentrisch angeordnet,wobei die Düsenwinkel des Düsenkastens (18) bzw. dieRadiusabstände (27) zur Turbinenwellenachse (17) derartangepasst sind, dass sich minimierte Verluste ergeben.

IMPULSE TURBINES

INLET CASING FOR STEAM TURBINE.

Improvements in turbines

This invention relates to a turbine for extracting energy from an oscillating working fluid. The turbine includes a housing defining a flow passage for the working fluid. An energy conversion unit is disposed in the housing. Flow control means is selectively movable to occlude a predetermined portion of the flow passage such that the working fluid is directed to act on a certain section of the energy conversion unit. Figure 6 21 24 1 025 28 NOZZLE SEGMENT 27 CHECK 26 OZZLE SEGMENT Figure 6 13NOZZLES(1800) SWING CHFCK Figure 7 ...

WIND TURBINE

SUPERSONIC NOZZLE UNIT

POWERED AUGMENTED FLUID TURBINES

A powered augmented fluid turbine for generating electricity from a fluid in motion comprising: a central annular ducted channel extending between an inlet distribution header and an outlet distribution header, the channel comprising a converging section configured to accelerate the fluid received at the inlet distribution header, a turbine assembly for generating electricity, and a diffuser section configured to decelerate the fluid before it exits at the outlet distribution header; a recycle line for transporting the exiting fluid to the inlet distribution header in a closed-loop configuration, the recycle line comprising a recycle line propulsor controllable by a recycle line controller and a recycle line heat exchanger; and a compressed fluid distribution line configured to pressurize the fluid in motion by transporting a compressed fluid from a compressed fluid source to the inlet and outlet distribution headers, the compressed fluid distribution line controllable by at least one pressure ...

ASSIST AND BACK-UP FOR ELECTRICAL AUXILIARY DRIVE

L'invention concerne un système d'entraînement d'un accessoire, tel qu'une pompe à carburant ou une pompe de lubrification, dans un turbomoteur , ledit système comportant un moteur électrique , caractérisé en ce qu'il comporte en outre une turbine à air associée audit moteur électrique . Ladite turbine à air est susceptible d'être alimentée par un débit d'air prélevé dans un compresseur dudit turbomoteur et participe à l'entraînement de l'accessoire. Il comporte en outre une vanne de réglage du débit d'air prélevé au compresseur qui est en position fermée lors de la phase de démarrage du turbomoteur.

GEARED AXIAL MULTISTAGE EXPANDER DEVICE, SYSTEM AND METHOD

Method, system and axial multistage expander including a casing and a plurality of stages. A stage includes a stator part connected to the casing and having plural statoric airfoils, and a rotor part configured to rotate relative to the stator part and having plural rotoric airfoils. The axial multistage expander also includes a support mechanism connected to the casing and configured to rotatably support the rotor part. Rotoric airfoils of at least one stage of the plurality of stages are configured to rotate with a speed different from rotoric airfoils of the other stages. The stator part, the rotor part and the support mechanism of the plurality of stages are provided inside the casing.

MIXER-EJECTOR TURBINE WITH ANNULAR AIRFOILS

Example embodiments are directed to fluid turbines that include a turbine shroud, a rotor and an ejector shroud. The turbine shroud includes an inlet, an outlet, a leading edge and a trialing edge. The leading edge of the turbine shroud can be round and the trialing edge of the turbine shroud can include linear faceted segments. The rotor can be disposed within the turbine shroud and can define a rotor plane. The turbine shroud can provide a first portion of a fluid stream to the rotor plane via the inlet of the turbine shroud. The ejector shroud can provide a second portion of the fluid stream to the outlet of the turbine shroud via an open area. An example method of operating a fluid turbine is also provided.

METHOD OF DRIVING A TURBINE IN ROTATION BY MEANS OF A JET DEVICE

... 2121029 9307361 PCTABS00021 In a method for using a turbine at variable speeds and power levels, a Venturi-effect injector (14) is placed upstream of a turbine (12) and converts a pressure command into a mixed fluid mass flow command. Sensors (19) and a regulation means (50) enable the speed of the turbine (12) to be adapted to a reference speed.

Umlaufender Teil mehrstufiger Dampf- oder Gasturbinen.

Verfahren zur Herstellung mehrstufiger, axialer Freistrahl-Dampf- oder Gasturbinen verschiedener Leistungen in Gruppen.

Verfahren zur Herstellung von für den Betrieb mit hochgespanntem Dampf ausbaufähigen, mehrstufigen Dampf- oder Gasturbinen mit Scheibenrädern.

Dampfturbinenanlage mit einem durch eine ursprünglich vorhandene Dampfturbinenanlage gebildeten Teil und wenigstens einer diesem Teil später vorgeschalteten Turbine für Hochdruckdampf.

Mehrstufige Hochdruck-Dampfturbine.

Verfahren zum Umändern von mehrstufigen Dampf- oder Gasturbinen für hohen und niederen Druck.

Dampfturbinenanlage mit einem durch eine ursprünglich vorhandene Dampfturbinenanlage gebildeten Teil und wenigstens einer diesem Teil später vorgeschalteten Turbine für Hochdruckdampf.

Dampfturbinenanlage mit einem durch eine ursprünglich vorhandene Dampfturbinenanlage gebildeten Teil und wenigstens einer diesem Teil später vorgeschalteten Turbine mit Hochdruckdampf.

Höchstdruckdampfturbine.

Mehrstufige Scheibenradturbine.

Druckstufenturbine mit mehr als einer Druckstufe im Hochdruckteil.

Mehrstufige Druckturbine für Dampf oder Gas.

Dampfturbine für sehr hohe Überhitzung.

Verfahren zur Herstellung von basisch substituierten Ketonen

Mehrstufige, axial beaufschlagte Gleichdruckturbine

Gleichdruck-Turbinenstufe

Verfahren zum Betrieb einer mit Festdruck arbeitenden Dampfturbinenanlage

Turbina a gas

Endstufe einer Gas- bzw. Dampfturbine

Dampfturbine für den Einsatz als Schnellstartturbine

TURBINE AXIALER BAUART.

WINDTURBINE.

TURBINE.

Turbine of a turbocharger and turbocharger axially.

Axialturbine (1) eines Turboladers zur Entspannung eines Mediums, mit einem Laufschaufeln (3) aufweisenden Turbinenrotor (2); mit einem radial aussen an die Laufschaufeln (3) angrenzenden Deckring (12); mit einem in Strömungsrichtung des zu entspannenden Mediums gesehen stromaufwärts der Laufschaufeln (3) positionierten, verstellbare Leitschaufeln (7) aufweisen Leitapparat (6), wobei die Leitschaufeln (7) in einem stromaufwärts des Deckrings (12) positionierten Leitring (8) gelagert sind; mit einem in Strömungsrichtung des zu entspannenden Mediums gesehen stromabwärts der Laufschaufeln (3) positionierten Diffusor (5); wobei ein Hohlraum (13), der sich radial aussen an den Deckring (12) und den Leitring (8) anschliesst, mit einem vom Diffusor (5) begrenzten Strömungskanal (16) derart gekoppelt ist, dass der Hohlraum (13) in den vom Diffusor (5) begrenzten Strömungskanal (16) entlüftbar ist.

Turbomachine shovel.

Die Schaufeln einer Turbomaschine umfassen Schaufelblätter (21, 31), welche derart gebogen sind, dass der Neigungswinkel (?), den die Auffädelungslinie des Schaufelblattes mit der Radialrichtung der Maschine aufweist und der in Drehrichtung (?) gemessen wird, über der Höhe des Strömungskanals (s) variiert und von der Nabe (2) zum Gehäuse (3) kleiner wird.

Steam turbine.

Steam turbine.

STEAM TURBINE.

Steam turbine with high pressure or gas

WIND TURBINES

IMPULSE TURBINE ROTOR DRUM AND IMPROVEMENT TO THESE TURBINES.

Rotary steam engine

COLD SUSPENSION OF THE FLOW CHANNEL OF A TURBOJET ENGINE BY MEANS OF CONNECTING RODS, AND CAPS ON THE EXHAUST CASE RADIAL

TURBINE STAGE

TURBINE HOUSING FOR DISTRICT HEATING

Internal combustion turbine that can operate at very high temperatures

STEAM TURBINE POWER GENERATING PLANT AND STEAM TURBINE

Blade for axial fluid machine having projecting portion at the tip and root of the blade

A blade for an axial fluid machine comprises an effective blade portion having a root portion and a tip portion and projecting blade portions. The projecting blade portions are defined by axis reference lines extending to an upstream side with respect to a fluid flow from at least one of the root portion and the tip portion of the effective blade portion and by axes obliquely extending from ends of the axis reference lines toward a leading edge of the effective blade portion. The projecting blade portions are formed continuously to and integrally with the leading edge of the effective blade portion so that the effective blade portion and the projecting blade portions have substantially the same maximum blade thickness.

INTERNAL COMBUSTION ENGINE

An internal combustion engine includes an exhaust gas passage, a water-cooled cylinder head and a turbocharger. The turbocharger includes: a compressor impeller; an axial flow turbine wheel coupled to the compressor impeller through a rotational shaft; a bearing that supports a portion of the rotational shaft located between the compressor impeller and the turbine wheel; and a housing that houses at least the compressor impeller and the bearing among the compressor impeller, the bearing and the turbine wheel. The turbine wheel is coupled to the rotational shaft such that the outlet of turbine blades of the turbine wheel is located on the side of the compressor impeller. The housing is fastened to the cylinder head, directly or with a first gasket interposed between the housing and the cylinder head, such that the turbine wheel is opposed to the cylinder head.

ROTATING VANE SEAL WITH COOLING AIR PASSAGES

An inner diameter vane seal for a gas turbine engine comprises an annular, ring-like body having inner and outer diameter rims, forward and aft faces and an air passage. The outer diameter rim extends circumferentially for engaging inner diameter ends of stator vanes. The inner diameter rim extends circumferentially and is spaced radially from the inner outer diameter rim. The forward and aft faces extend radially between the outer diameter rim and the inner diameter rim. The air passage extends from the forward face to the aft face between the inner and outer diameter rims.

Low-pressure steam turbine and method for operating thereof

The invention relates to a multi-stage low-pressure steam turbine (10) and a method for operating thereof. The steam turbine (10) comprises a last stage (18) in which the leading edge (20) of each vane (14) of the last stage (18) is skewed so as to form a W shaped K-distribution across the span (36) of the vanes (14). This shape enables efficient last stage (18) operation at low last stage exit velocities. The invention further includes a method for operating such a steam turbine (10) at a last stage exit velocity between 125 m/s and 150 m/s.

Gas turbine system and method of manufacturing

TURBINE TYPE OPERATION DEVICE

PURPOSE: To operate an operation body by means of the suction force of a vacuum generating device, by providing rotary vanes and an restricting body in a casing and attaching the operation body to a rotary shaft, in the device in the caption used for a household cleaner, etc. CONSTITUTION: A restricting body 4 for regulating rotary vanes 2 and an air flow direction to the rotary vanes 2 in a slant direction is provided in a cashing 1 connected to a vacuum generating device. The rotary shaft 2a of the rotary vanes 2 is fitted to an operating body 10 such as a cutter, etc. The operating body can be thereby operated by using the air flow generated by the suction force of the vacuum generating device. COPYRIGHT: (C)1989,JPO&Japio ...

РЕГЕНЕРАЦИЯ ЭНЕРГИИ

Изобретение относится к способу регулируемой регенерации энергии реакции окисления, при которой образуется газовый поток, каковую реакцию осуществляют в реакторе окисления непрерывного действия, в который подают газообразный окислитель. Способ включает: (a) нагревание газового потока до температуры по меньшей мере 800°C; (b) направление газового потока на ступень турбины внутреннего сгорания с открытым циклом, в которой имеется турбинное колесо, соединенное с компрессором, каковой компрессор сжимает газообразный окислитель, подаваемый в реактор; (c) регулирование давления на ступени турбины; (d) поддержание давления на ступени турбины в диапазоне больше минимальной величины, соответствующей энергетической потребности компрессора на сжатие газообразного окислителя, подаваемого в реактор окисления, и меньше максимальной величины, определяемой пределами газовой турбины по мощности или давлению, путем добавления газа в газовый поток; (e) обеспечение расширительного устройства или вспомогательного ...

ДВУХЪЯРУСНЫЙ ЦИЛИНДР НИЗКОГО ДАВЛЕНИЯ КОНДЕНСАЦИОННОЙ ПАРОВОЙ ТУРБИНЫ

Цилиндр низкого давления содержит одноярусные и двухъярусные ступени, состоящие из сопловых и рабочих лопаток, формирующих проточные части нижнего и верхнего ярусов цилиндра. Лопаточный аппарат верхнего яруса представляет собой проточную часть самостоятельной турбины, надстроенной над турбиной нижнего яруса и связанной с ней. В двухъярусных ступенях рабочие лопатки верхнего яруса кинематически связаны с рабочими лопатками нижнего яруса в области интегральных бандажей последних сплошными коническими кольцами, представляющими собой корневые узлы крепления рабочих лопаток верхнего яруса. Изобретение позволяет снизить потери, связанные с веерностью двухъярусных лопаток, и перетечки между ярусами, а также снизить трудоемкость изготовления двухъярусных ступеней. 1 з.п. ф-лы, 4 ил.

СОПЛОВОЙ АППАРАТ АКТИВНОЙ ТУРБИНЫ

Сопловой аппарат активной турбины содержит сопло, имеющее разгонный участок и выходной участок, в котором выходное сечение сопла на плоскости косого среза имеет средний радиус изгиба, равный среднему радиусу рабочей решетки колеса турбины. В сопле разгонный участок выполнен из осесимметричного сопла и изогнутого расширяющегося канала. Ширина канала постоянна и равна диаметру выходного сечения осесимметричного сопла. Высота канала выполнена плавно увеличивающейся от диаметра выходного сечения осесимметричного сопла до максимального значения на выходном участке соплового аппарата. Ось осесимметричного сопла выполнена по касательной к окружности среднего радиуса изгиба расширяющегося канала. Выходное сечение осесимметричного сопла не превышает высоты лопаток рабочей решетки турбины. Поверхность расширяющегося канала образована двумя цилиндрическими коаксиальными поверхностями со средним радиусом изгиба, равным среднему радиусу выходного сечения сопла на плоскости косого среза, и двумя сопряженными ...

УСТРОЙСТВО ДИСТАНЦИОННОГО КОНТРОЛЯ ПАРАМЕТРОВ ГАЗОРАСПРЕДЕЛИТЕЛЬНЫХ ПУНКТОВ

Устройство дистанционного контроля параметров газораспределительных пунктов, содержащее технологический блок с датчиками давления газа на входном трубопроводе, узел очистки газа с датчиком давления на его выходе, узел редуцирования с датчиком давления, счетчик расхода газа с корректором, датчик загазованности и датчик открытия дверей, отличающееся тем, что в устройство дополнительно введены блок сравнения, компаратор входного давления, компаратор узла очистки газа, компаратор узла редуцирования, задатчик контроля входного давления, задатчик контроля перепада давлений узла очистки газа, задатчик контроля узла редуцирования, регистр расхода газа, задатчик контроля расхода газа, блок контроля расхода газа, элемент ИЛИ, контроллер сотовой связи с аварийным и набором информационных входов и выходов, усилитель и клапан отключения подачи газа, причем выход датчика давления газа на входном трубопроводе соединен с первым входом блока сравнения, первым входом компаратора входного давления и информационным ...

Теплофикационная паротурбинная установка с охладителем основного конденсата на линии его рециркуляции

Область использования: теплоэнергетика. Теплофикационная паротурбинная установка ТПТУ содержит теплофикационную паровую турбину ТПТ 10, обогреваемую отборным паром указанной турбины тепловую сеть, оборудованную сетевым насосом СН 70; конденсатор К 20 выхлопного пара турбины, оборудованный установленным на линии 30 возврата основного конденсата в теплофикационный цикл конденсатным насосом КН 40; последовательно включенные по охлаждающей стороне поверхностный охладитель ПОЭ 90 для конденсации выхлопного пара парового эжектора ЭЖ 80, создающего вакуум в указанном конденсаторе, и поверхностный охладитель ПОУ 100 для конденсации пара из концевых уплотнений турбины; включаемую при высоких теплофикационных нагрузках для поддержания уровня воды в К 20 линию 300 рециркуляции отводимой к указанным охладителям части основного конденсата после его нагрева в них; установленный на отводной линии 200 от ПОУ 100 переключатель ПП 400 потока охлаждающей среды для ее подачи в зависимости от выбранного режима ...

АКТИВНАЯ ПАРОВАЯ ТУРБИНА

Использование: в теплоэнергетике, в частности в активных паровых турбинах. Сущность изобретения: активная паровая турбина содержит несколько ступеней, каждая из которых имеет диафрагму с решеткой неподвижных лопаток, снабженных бандажом с уплотнительным элементом, и размещенную за ней решетку подвижных лопаток, установленных на роторе. Ротор имеет барабан, а подвижные лопатки закреплены непосредственно на барабане. Каждая из них имеет пяту, при этом на торце пяты каждой лопатки со стороны диафрагмы и на ответном торце диафрагмы выполнены скосы, образующие между собой зазор для подачи утечек в основной поток пара в осевом направлении. Возможно, кроме того, выполнение в пяте каждой подвижной лопатки сквозного канала, параллельного оси ротора, или сообщенных каналов в диафрагме и неподвижных лопатках, а также наклонных каналов в бандаже неподвижных лопаток. 3 з. п. ф-лы, 9 ил.

ТЕПЛООБМЕННИК

МАЛОШУМНАЯ ТУРБИНА ДЛЯ РЕДУКТОРНОГО ТУРБОВЕНТИЛЯТОРНОГО ДВИГАТЕЛЯ

... 1. Газотурбинный двигатель, содержащий:вентилятор, компрессорную секцию, содержащую компрессор с частью низкого давления и частью высокого давления, секцию камеры сгорания и турбину с частью низкого давления, а также понижающий редуктор, обеспечивающий уменьшение скорости указанного вентилятора относительно входной скорости указанного вентилятора;указанная часть низкого давления указанной турбины имеет определенное число лопаток турбины в каждом из множества рядов указанной части турбины, при этом указанные лопатки турбины низкого давления предназначены для работы, по меньшей мере, некоторое время, с определенной угловой скоростью вращения, и при этом указанное количество лопаток и указанная угловая скорость вращения таковы, что следующая формула справедлива, по меньшей мере, для одного из рядов лопаток турбины низкого давления(число лопаток × скорость) / 60≥5500; при этомуказанная угловая скорость вращения представляет собой скорость при заходе на посадку, выраженную в оборотах в минуту ...

Проточная часть паровой турбины

Решетка турбомашины

Парциальная турбина

Strömungsmaschine mit Hochdruck- und Niederdruck-Schaufelbereich

System für den Antrieb einer Pumpe in einer Turbomaschine

Improvements in steam or gas turbines

... 212,508. Fairweather, M. G. C., (Erste Br³nner Maschinen-Fabriks-Ges.). Feb. 21, 1924. Axial-flow type.-Full-admission diaphragms in multi-stage disc-wheel high-pressure turbines are arranged in groups a, b, c of two or more, the guide passages f, f<1>, f<2> of each group being of the same cross-section, while the cross-section of the passages increases from group to group.

Improvements in or relating to Turbines.

... 9054. Shann, M. Churchill-. April 11. Axial-flow type; blades; plant.-The inlet nozzles L of a turbine A, which may be arranged on the shaft O of a gas engine N, the exhaust from which may provide the working- fluid, are pitched apart by an interval exceeding the sum of the width of the nozzle and the pitch of the blades C. After acting on the moving blades C, which are secured by rivets to rings A, B and increase in radial dimensions towards the outlet side, the working-fluid discharges against a ring of fixed reaction blades C<2>.

Improvements in and relating to Elastic Fluid Turbines.

... 6658. Warwick Machinery Co., [General Electric Co.]. March 16. Axial-flow type. - A turbine having the highpressure stage 1, the intermediate stage 2, and a part 3 of the lowpressure stage, of the reaction type is provided with final stages 5 of the impulse type. Specification 14,163/11 is referred to.

Improvements in or relating to stationary blading for turbines

... 636,860. Steaam turbine blades. WESTINGHOUSE ELECTRIC INTERNATIONAL CO. May 24, 1948, No. 13974. Convention date, June 11, 1947. [Class 110(iii)] A fabricated lowpressure stator blade ring is made in two or more segments. Each segment comprises inner and outer arcuate members 19, 18, providing a radially-divergent passage between them, and blades 15 welded to both members. Each blade is formed of sheet metal in two pieces welded together along the leading and trailing edges, or, in one piece welded at the trailing edge. The outer arcuate member may be made solid or fabricated from shaped wall parts 22...25 welded together.

Improvements in turbines

... 613,780. Turbines. TYE, W. D., MacPHAIL, D. C., and DAVIES, D. I. T. P. LLEWELYN-. July 2, 1946, No. 19803. [Class 110 (iii)] A turbine stator comprises an end inlet member 1, an end outlet member 8 and intervening nozzle members 21 and casing members 12. The members have spigot-and-socket engagements and are clamped together by bolts 10. The member 8 has a number of radial lugs 9 and an exhaust deflecting member 16 fits over the lugs and is secured in place by screws 18. Bearings 19, 26 for the rotor are' provided in the first nozzle member and the exhaust deflecting member. The bladed wheels 23 are splined upon the shaft 20. The final nozzle member has a full circle of nozzle passages: the others have nozzle passages over segments that increase from the inlet end. A conical guide member 27 covers the bearing 26.

PROCEDURE, IN ORDER TO PROPEL A TURBINE OF MEANS OF A STRAHLAPPARATES

TURBOMACHINE STAGE WITH DECREASED FLOW LOSS.

TURBOMACHINE, IN PARTICULAR STEAM TURBINE WITH HIGH STEAM INLET TEMPERATURE

Improved wind turbine

A multi-stage axial flow turbine adapted to operate at low steam temperatures

An axial flow turbine for generation of electrical power having multiple stages and configured for operation at low absolute pressure with the motive fluid being steam. The turbine is nominally ten stages with the first stage having a partial admission inlet, each subsequent stage increasing the amount of steam admission until complete admission is achieved towards the final stages. Each stage has blisks built as a single piece and the steam passages built into the periphery of the blisks. C~ p1 Figure 2 ...

AIR TURBINE

ASSIST AND BACK-UP FOR ELECTRICAL AUXILIARY DRIVE

L'invention concerne un système d'entraînement d'un accessoire, tel qu'une pompe à carburant (13) ou une pompe de lubrification, dans un turbomoteur (10), ledit système comportant un moteur électrique (21), caractérisé en ce qu'il comporte en outre une turbine à air (20) associée audit moteur électrique (21). Ladite turbine à air est susceptible d'être alimentée par un débit d'air prélevé dans un compresseur (12) dudit turbomoteur (10) et participe à l'entraînement de l'accessoire. Il comporte en outre une vanne de réglage (25) du débit d'air prélevé au compresseur qui est en position fermée lors de la phase de démarrage du turbomoteur (10).

TURBINE ENGINE GUIDE VANE AND ARRAYS THEREOF

An exit guide vane array for a turbine engine includes a set of guide vanes 28 having a solidity and defining fluid flow passages 74 with a chordwisely converging forward portion 80. The high solidity and convergent passage portion 80 resist fluid separation. The vanes may also cooperate with each other to restrict an observer's line of sight to planes upstream of the vane array.

Blade comprising pre-wired sections

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

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

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

Turbine of a turbomachine

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

Non axis-symmetric stator vane endwall contour

An airfoil comprises pressure and suction surfaces extending axially from a leading edge to a trailing edge and radially from a root section to a tip section, defining a mean span therebetween. An inner endwall defines an inner endwall contour extending axially and circumferentially from the root section, and an outer endwall defines an outer endwall contour extending axially and circumferentially from the tip section. The inner and outer endwall contours are defined by varying radial deviations from circumferentially uniform nominal inner and outer radii, where one of the radial deviations varies axially and circumferentially by at least three percent of a mean span of the airfoil.

AEROFOILS

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

Turbine bucket airfoil profile

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

Water Impeller

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

SCREW TURBINE AND METHOD OF POWER GENERATION

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

Turbine Wheel Arrangement For A Gas Turbine

The invention relates to a turbine wheel arrangement for a gas turbine comprising two successive turbine wheels () rotating in opposite directions, wherein the first turbine wheel () comprises flow channels () of a Laval cross-sectional shape, distributed over the circumference and having radially inner gas inlets () and radially further out gas outlets (), in each case with a substantially tangential flow direction component, wherein the gas outlets () act an the second axially or radially acting turbine wheel () in the direction of flow. This has the effect that the thermal energy and compressive energy of the gas at the nozzle inlet is largely converted into flow energy at the outlet of the turbine stage. The rotational speeds of the two rotors coupled to the turbine wheels can be set as desired, allowing the operational states of the two systems to be optimally set without adjusting systems. 1. A gas turbine comprising a first turbine wheel and a second turbine wheel rotating in the opposite direction which is impacted by the first turbine wheel , further comprising an inner shaft on the outside of which multiple first blade rows of a multi-stage axial compressor and the second turbine wheel are mounted , further comprising a hollow shaft on which the first turbine wheel and multiple second blade rows of the axial compressor are mounted , which are arranged alternatingly in the axial direction relative to the first blade rows.2. The gas turbine according to claim 1 , wherein the first turbine wheel comprises flow channels distributed over the circumference thereof claim 1 , which include gas outlets of a laval cross-sectional shape.3. The gas turbine according to claim 2 , wherein the gas outlets include a substantially tangential flow direction component.4. The gas turbine according to claim 3 , wherein the gas outlets of the first turbine wheel are oriented at an angle of 10° to 30° relative to the tangential direction.5. The gas turbine according to claim 1 , ...

TURBOMACHINE

A turbomachine including at least one blade-row group that is arranged in the main flow path and at least two rows of blades that are adjacent to each other in the main flow direction, each row having a plurality of blades, whereby the trailing edges of the blades of the upstream row of blades and the leading edges of the blades of the downstream row of blades are arranged at an axial edge distance that decreases from the center of the main flow path in the direction of at least one main flow limiter. 1. A turbomachine comprising:at least one blade-row group arranged in a main flow path and including an upstream row of blades and a downstream row of further blades adjacent to each other in a main flow direction, trailing edges of the blades of the upstream row and leading edges of the further blades of the downstream row being arranged at an axial edge distance decreases from a center of the main flow path in a direction of at least one main flow limiter.2. The turbomachine as recited in wherein the edge distance decreases steadily in the direction of a minimum at the main flow limiter.3. The turbomachine as recited in wherein the edge distance decreases to a minimum and then increases in the direction of the main flow limiter.4. The turbomachine as recited in wherein the edge distance at the main flow limiter is equal to or smaller than the edge distance in the center of the main flow path.5. The turbomachine as recited in wherein the edge distance decreases to a minimum and then remains constant up to the main flow limiter.6. The turbomachine as recited in wherein the trailing edges of the blades of the upstream row have a linear configuration and the leading edges of the blades of the downstream row have a curved configuration.7. The turbomachine as recited in wherein the trailing edges of the blades of the upstream row have a curved configuration and the leading edges of the further blades of the downstream row have a rectilinear configuration.8. The ...

TURBOMACHINE

A turbomachine including at least one blade-row group that is arranged in the main flow path and at least two rows of blades that are adjacent to each other in the main flow direction, each row having a plurality of blades (), whereby a narrow cross section and a degree of overlap between the blades of the upstream row of blades and the blades of the downstream row of blades vary starting at the center of the main flow path in the direction of at least one main flow limiter. 1. A turbomachine comprising:at least one blade-row group arranged in a main flow path including an upstream row of blades and a downstream row of further blades adjacent to each other in a main flow direction, a narrowest cross section and a degree of overlap between the blades of the upstream row and the further blades of the downstream row varying starting at a center of the main flow path in the direction of at least one main flow limiter.2. The turbomachine as recited in wherein the narrowest cross section increases steadily in the direction of the main flow limiter.3. The turbomachine as recited in wherein the narrowest cross section increases in the direction of the main flow limiter and then decreases.4. The turbomachine as recited in wherein the narrowest cross section increases in the direction of the main flow limiter and then remains constant.5. The turbomachine as recited in wherein the narrowest cross section decreases steadily in the direction of the main flow limiter.6. The turbomachine as recited in wherein the narrowest cross section decreases in the direction of the main flow limiter and then increases.7. The turbomachine as recited in wherein the narrowest cross section decreases in the direction of the main flow limiter and then remains constant.8. The turbomachine as recited in wherein the degree of overlap increases steadily in the direction of the main flow limiter.9. The turbomachine as recited in wherein the degree of overlap increases in the direction of the main flow ...

STEAM TURBINE

A steam turbine in an embodiment includes: an inner casing surrounding a turbine rotor; an outer casing composed of an upper half side outer casing and a lower half side outer casing and an annular diffuser through which steam passed through a turbine stage is discharged radially outward. A vertical distance H from a axis of a turbine rotor to an inner wall of the upper half side outer casing an outermost diameter D of final stage rotor blades a blade height B of the final stage rotor blade and a distance W between inner walls in vertical and horizontal directions to the axis of the turbine rotor 23, at a bottom portion of the lower half side outer casing forming a discharge port satisfy (H−D/2)/B≦1.7 and (W−D)/2B≧2. 1. A steam turbine having a path to guide steam having passed through a final turbine stage to a condenser provided below , the steam turbine , comprising:an inner casing surrounding a turbine rotor and composed of an upper half side inner casing and a lower half side inner casing;an outer casing surrounding the inner casing and composed of an upper half side outer casing and a lower half side outer casing; andan annular diffuser that is provided on a downstream side of the final turbine stage and through which the steam having passed through the final turbine stage is discharged radially outward, wherein [{'br': None, 'i': H−D/', 'B≦, '(2)/1.7 \u2003\u2003Expression (1)'}, {'br': None, 'i': W−D', 'B≧, '()/22 \u2003\u2003Expression (2)'}], 'a vertical distance H from a center axis of the turbine rotor to an inner wall of the upper half side outer casing, an outermost diameter D of final stage rotor blades forming the final turbine stage, a blade height B of the final stage rotor blade, and a distance W between inner walls in vertical and horizontal directions to the center axis of the turbine rotor, at a bottom portion of the lower half side outer casing forming a discharge port through which the steam is discharged to the condenser, satisfy the ...

Blade body and rotary machine

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

COMPOUND AIRFOIL

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

FLUID TURBINE WITH VORTEX GENERATORS

The present disclosure relates to fluid turbines having a turbine shroud assembly formed with mixing elements (e.g., both inwardly and outwardly curving elements) having airfoil cross sections. These airfoils form ringed airfoil shapes that provide a means of controlling the flow of fluid over the rotor assembly or over portions of the rotor assembly. The fluid dynamic performance of the ringed airfoils directly affects the performance of the turbine rotor assembly. The mass and surface area of the shrouds result in load forces on support structures. By delaying or eliminating the separation of the boundary layer over the ringed airfoils, boundary layer energizing members (e.g., vortex generators, flow control ports) on the ringed airfoils increase the power output of the fluid turbine system and allow for relatively shorter chord-length airfoil cross sections and therefore reduced mass and surface area of the shroud assemblies. 1. A shrouded fluid turbine system comprising:a rotor assembly;a turbine shroud assembly disposed about the rotor assembly, the turbine shroud having a low pressure side and a high pressure side, the low pressure side in fluid communication with the rotor assembly; andat least one boundary layer energizing member associated with the turbine shroud assembly, the at least one boundary layer energizing member configured and dimensioned to alter a fluid boundary layer over a surface of the turbine shroud assembly to alter the performance of the fluid turbine system.2. The system of claim 1 , wherein the at least one boundary layer energizing member is positioned proximal to a leading edge of the turbine shroud assembly.3. The system of further comprising a first plurality of boundary layer energizing members and a second plurality of boundary layer energizing members;wherein the first plurality of boundary layer energizing members are positioned proximal to a leading edge of the turbine shroud assembly; andwherein the second plurality of ...

MOVING BLADE AND TURBOMACHINE

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

WIND TURBINE

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

ARCHIMDEAN SCREW APPARATUS

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

Airfoil mateface sealing

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

High Volume Low Speed Fan

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

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

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

Impeller for Centrifugal Pumps

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

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

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

IMPELLER FOR A CENTRIFUGAL PUMP

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

Impeller and Electric-Motor Driven Water Pump Having the Same

An impeller is comprised of a hub configured to be rotated on a central axis, a shroud formed to be opposed to the hub in a direction of the central axis and having a central opening serving as a fluid inlet, and a plurality of circumferentially-equidistant spaced blades interleaved between the hub and the shroud. When a mating face of the shroud with each of the blades is divided into a radially inward region and a radially outward region, and a mating face of each of the blades with the shroud is divided into a radially inward region and a radially outward region, a given weld range is set only in the radially inward region of the mating face of the shroud with each of the blades and the radially inward region of the mating face of each of the blades with the shroud.

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

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

TURBOMACHINE BLADE HAVING A MAXIMUM THICKNESS LAW WITH HIGH FLUTTER MARGIN

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

OPTIMIZED AERODYNAMIC PROFILE FOR A TURBINE BLADE, IN PARTICULAR FOR A ROTARY WHEEL OF THE FOURTH STAGE OF A TURBINE

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P0 situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P1. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine. 1. An aerodynamic profile for a turbine blade , the profile being , when cold and in a non-coated state , substantially identical to a nominal profile determined by the Cartesian coordinates X ,Y ,Zadim given in Table 1 , in which the coordinate Zadim is the quotient D/H , where D is the distance of the point under consideration from a first reference X ,Y plane situated at the base of the nominal profile , and H is the height of said profile measured from said first reference plane that is the intersection of the stacking axis of the set of blades and the axisymmetric surface of the hub , out to a second reference plane that is the intersection of said stacking axis with the axisymmetric surface of the casing , the measurements D and H being taken radially relative to the axis of the turbine , while the coordinate X is measured in the axial direction of the turbine.2. The aerodynamic profile as claimed in claim 1 , wherein said profile is defined within an envelope of ±1 mm in a direction normal to the surface of the nominal profile.3. The aerodynamic profile as claimed in claim 1 , wherein the coordinates X claim 1 ,Y of said profile lie within a range of ±5% relative to the coordinates X claim 1 ,Y of the nominal profile.4. The aerodynamic profile as claimed in claim 1 , wherein the blade is a blade of a rotary wheel forming a portion of a rotor of ...

OPTIMIZED AERODYNAMIC PROFILE FOR A TURBINE VANE, IN PARTICULAR FOR A NOZZLE OF THE FIRST STAGE OF A TURBINE

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine. 1. An aerodynamic profile for a turbine vane , the profile being , when cold and in a non-coated state , substantially identical to a nominal profile determined by the Cartesian coordinates X ,Y ,Zadim given in Table 1 , in which the coordinate Zadim is the quotient D/H , where D is the distance of the point under consideration from a reference X ,Y plane situated at the base of the nominal profile , and H is the height of said profile measured from said reference plane out to the end of the vane , the measurements D and H being taken radially relative to the axis of the turbine , while the coordinate X is measured in the axial direction of the turbine.2. The aerodynamic profile as claimed in claim 1 , wherein said profile is defined within an envelope of ±1 mm in a direction normal to the surface of the nominal profile.3. The aerodynamic profile as claimed in claim 1 , wherein the coordinates X claim 1 ,Y of said profile lie within a range of ±5% relative to the coordinates X claim 1 ,Y of the nominal profile.4. The aerodynamic profile as claimed in claim 1 , wherein the vane is a nozzle vane forming a part of a stator of a turbine.5. The aerodynamic profile as claimed in claim 4 , wherein the vane is a nozzle vane of the first stage of the turbine.6. The aerodynamic profile as claimed in claim 4 , wherein the vane is a vane of the first stage nozzle ...

OPTIMIZED AERODYNAMIC PROFILE FOR A TURBINE VANE, IN PARTICULAR FOR A NOZZLE OF THE SECOND STAGE OF A TURBINE

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P0 situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P1. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine. 1. An aerodynamic profile for a turbine vane , the profile being , when cold and in a non-coated state , substantially identical to a nominal profile determined by the Cartesian coordinates X ,Y ,Zadim given in Table 1 , in which the coordinate Zadim is the quotient D/H , where D is the distance of the point under consideration from a reference X ,Y plane situated at the base of the nominal profile , and H is the height of said profile measured from said reference plane out to the end of the vane , the measurements D and H being taken radially relative to the axis of the turbine , while the coordinate X is measured in the axial direction of the turbine.2. The aerodynamic profile as claimed in claim 1 , wherein said profile is defined within an envelope of ±1 mm in a direction normal to the surface of the nominal profile.3. The aerodynamic profile as claimed in claim 1 , wherein the coordinates X claim 1 ,Y of said profile lie within a range of ±5% relative to the coordinates X claim 1 ,Y of the nominal profile.4. The aerodynamic profile as claimed in claim 1 , wherein the vane is a nozzle vane forming a part of a stator of a turbine.5. The aerodynamic profile as claimed in claim 4 , wherein the vane is a nozzle vane of the second stage of the turbine.6. The aerodynamic profile as claimed in claim 4 , wherein the vane is a vane of the second stage ...

OPTIMIZED AERODYNAMIC PROFILE FOR A TURBINE BLADE, IN PARTICULAR FOR A ROTARY WHEEL OF THE SEVENTH STAGE OF A TURBINE

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P0 situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P1. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine. 1. An aerodynamic profile for a turbine blade , the profile being , when cold and in a non-coated state , substantially identical to a nominal profile determined by the Cartesian coordinates X ,Y ,Zadim given in Table 1 , in which the coordinate Zadim is the quotient D/H , where D is the distance of the point under consideration from a first reference X ,Y plane situated at the base of the nominal profile , and H is the height of said profile measured from said first reference plane that is the intersection of the stacking axis of the set of blades and the axisymmetric surface of the hub , out to a second reference plane that is the intersection of said stacking axis with the axisymmetric surface of the casing , the measurements D and H being taken radially relative to the axis of the turbine , while the coordinate X is measured in the axial direction of the turbine.2. The aerodynamic profile as claimed in claim 1 , wherein said profile is defined within an envelope of ±1 mm in a direction normal to the surface of the nominal profile.3. The aerodynamic profile as claimed in claim 1 , wherein the coordinates X claim 1 ,Y of said profile lie within a range of ±5% relative to the coordinates X claim 1 ,Y of the nominal profile.4. The aerodynamic profile as claimed in claim 1 , wherein the blade is a blade of a rotary wheel forming a portion of a rotor of ...

OPTIMIZED AERODYNAMIC PROFILE FOR A TURBINE BLADE, IN PARTICULAR FOR A ROTARY WHEEL OF THE SECOND STAGE OF A TURBINE

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine. 1. An aerodynamic profile for a turbine blade , the profile being , when cold and in a non-coated state , substantially identical to a nominal profile determined by the Cartesian coordinates X ,Y ,Zadim given in Table 1 , in which the coordinate Zadim is the quotient D/H , where D is the distance of the point under consideration from a first reference X ,Y plane situated at the base of the nominal profile , and H is the height of said profile measured from said first reference plane that is the intersection of the stacking axis of the set of blades and the axisymmetric surface of the hub , out to a second reference plane that is the intersection of said stacking axis with the axisymmetric surface of the casing , the measurements D and H being taken radially relative to the axis of the turbine , while the coordinate X is measured in the axial direction of the turbine.2. The aerodynamic profile as claimed in claim 1 , wherein said profile is defined within an envelope of ±1 mm in a direction normal to the surface of the nominal profile.3. The aerodynamic profile as claimed in claim 1 , wherein the coordinates X claim 1 ,Y of said profile lie within a range of ±5% relative to the coordinates X claim 1 ,Y of the nominal profile.4. The aerodynamic profile as claimed in claim 1 , wherein the blade is a blade of a rotary wheel forming a portion of a rotor of a ...

OPTIMIZED AERODYNAMIC PROFILE FOR A TURBINE BLADE, IN PARTICULAR FOR A ROTARY WHEEL OF THE FIFTH STAGE OF A TURBINE

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine. 1. An aerodynamic profile for a turbine blade , the profile being , when cold and in a non-coated state , substantially identical to a nominal profile determined by the Cartesian coordinates X ,Y ,Zadim given in Table 1 , in which the coordinate Zadim is the quotient D/H , where D is the distance of the point under consideration from a first reference X ,Y plane situated at the base of the nominal profile , and H is the height of said profile measured from said first reference plane that is the intersection of the stacking axis of the set of blades and the axisymmetric surface of the hub , out to a second reference plane that is the intersection of said stacking axis with the axisymmetric surface of the casing , the measurements D and H being taken radially relative to the axis of the turbine , while the coordinate X is measured in the axial direction of the turbine.2. The aerodynamic profile as claimed in claim 1 , wherein said profile is defined within an envelope of ±1 mm in a direction normal to the surface of the nominal profile.3. The aerodynamic profile as claimed in claim 1 , wherein the coordinates X claim 1 ,Y of said profile lie within a range of ±5% relative to the coordinates X claim 1 ,Y of the nominal profile.4. The aerodynamic profile as claimed in claim 1 , wherein the blade is a blade of a rotary wheel forming a portion of a rotor of a ...

OPTIMIZED AERODYNAMIC PROFILE FOR A TURBINE BLADE, IN PARTICULAR FOR A ROTARY WHEEL OF THE SIXTH STAGE OF A TURBINE

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine. 1. An aerodynamic profile for a turbine blade , the profile being , when cold and in a non-coated state , substantially identical to a nominal profile determined by the Cartesian coordinates X ,Y ,Zadim given in Table 1 , in which the coordinate Zadim is the quotient D/H , where D is the distance of the point under consideration from a first reference X ,Y plane situated at the base of the nominal profile , and H is the height of said profile measured from said first reference plane that is the intersection of the stacking axis of the set of blades and the axisymmetric surface of the hub , out to a second reference plane that is the intersection of said stacking axis with the axisymmetric surface of the casing , the measurements D and H being taken radially relative to the axis of the turbine , while the coordinate X is measured in the axial direction of the turbine.2. The aerodynamic profile as claimed in claim 1 , wherein said profile is defined within an envelope of ±1 mm in a direction normal to the surface of the nominal profile.3. The aerodynamic profile as claimed in claim 1 , wherein the coordinates X claim 1 ,Y of said profile lie within a range of ±5% relative to the coordinates X claim 1 ,Y of the nominal profile.4. The aerodynamic profile as claimed in claim 1 , wherein the blade is a blade of a rotary wheel forming a portion of a rotor of a ...

OPTIMIZED AERODYNAMIC PROFILE FOR A TURBINE BLADE, IN PARTICULAR FOR A ROTARY WHEEL OF THE FIRST STAGE OF A TURBINE

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine. 1. An aerodynamic profile for a turbine blade , the profile being , when cold and in a non-coated state , substantially identical to a nominal profile determined by the Cartesian coordinates X ,Y ,Zadim given in Table 1 , in which the coordinate Zadim is the quotient D/H , where D is the distance of the point under consideration from a first reference X ,Y plane situated at the base of the nominal profile , and H is the height of said profile measured from said first reference plane that is the intersection of the stacking axis of the set of blades and the axisymmetric surface of the hub , out to a second reference plane that is the intersection of said stacking axis with the axisymmetric surface of the casing , the measurements D and H being taken radially relative to the axis of the turbine , while the coordinate X is measured in the axial direction of the turbine.2. The aerodynamic profile as claimed in claim 1 , wherein said profile is defined within an envelope of ±1 mm in a direction normal to the surface of the nominal profile.3. The aerodynamic profile as claimed in claim 1 , wherein the coordinates X claim 1 ,Y of said profile lie within a range of ±5% relative to the coordinates X claim 1 ,Y of the nominal profile.4. The aerodynamic profile as claimed in claim 1 , wherein the blade is a blade of a rotary wheel forming a portion of a rotor of a ...

TURBINE STAGE PLATFORM WITH ENDWALL CONTOURING INCORPORATING WAVY MATE FACE

A turbine stage includes a first airfoil and a second airfoil extending respectively from a first platform and a second platform that form an endwall for a flow passage. The endwall has a nominal surface that is axisymmetric about an axis of the turbine stage. The endwall further includes at least one contoured region that is non-axisymmetric with respect to the axis. The at least one contoured region extends from the first platform to the second platform across a platform splitline. The global maximum variation in elevation ΔEW of the endwall is at least 3% of an axial chord length L of the airfoils on the endwall. The maximum variation in elevation ΔMF at the mate facea of the platforms lies in the range 15-60% ΔEW. 1. A turbine stage comprising:a first airfoil extending from a first platform and a second airfoil spaced circumferentially from the first airfoil and extending from a second platform,wherein a flow passage of a working medium is defined between the first and second airfoils, the first and second platforms defining an endwall for said flow passage,wherein the first and second platforms comprise respective mate faces that interface along a platform splitline,wherein the endwall has a nominal surface that is axisymmetric about an axis of the turbine stage, the endwall further comprising at least one contoured region that is non-axisymmetric with respect to said axis, the at least one contoured region extending from the first platform to the second platform across the platform splitline,wherein a global maximum variation in elevation ΔEW of the endwall is at least 3% of an axial chord length L of the airfoils on said endwall, andwherein a maximum variation in elevation ΔMF at any of said mate faces lies in the range 15-60% ΔEW.2. The turbine stage according to claim 1 ,wherein the at least one contoured region comprises a bulge, andwherein the global maximum variation in elevation ΔEW of the endwall is defined by a height h of a peak of the bulge in ...

CENTRIFUGAL PUMP IMPELLERS

A centrifugal pump impeller includes front and back shrouds and a plurality of pumping vanes therebetween, each pumping vane having a leading edge in the region of an impeller inlet and a trailing edge, the front shroud has an arcuate inner face in the region of the impeller inlet, the arcuate inner face having a radius of curvature (Rin the range from 0.05 to 0.16 of the outer diameter of the impeller (D) The back shroud includes an inner main face and a nose having a curved profile with a nose apex in the region of the central axis which extends towards the front shroud, there being a curved transition region between the inner main face and the nose. Fis the radius of curvature of the transition region and the ratio F/Dis from 0.32 to 0.65. Other ratios of various dimensions of the impeller are also described. 14-. (canceled)5. An impeller which includes a front shroud and a back shroud , the back shroud including a back face and an inner main face with an outer peripheral edge and a central axis , a plurality of pumping vanes projecting from the inner main face of the back shroud to the front shroud , the pumping vanes being disposed in spaced apart relation on the inner main face providing a discharge passageway between adjacent pumping vanes , each pumping vane including a leading edge portion in the region of the central axis and a trailing edge portion in the region of the peripheral edge , the back shroud further including a nose having a curved profile with a nose apex in the region of the central axis which extends towards the front shroud , there being a curved transition region between the inner main face and the nose , wherein Fis the radius of curvature of the transition region and Dis the diameter of the impeller , and the ratio F/Dbeing from 0.20 to 0.75 , wherein one or more of the passageways have associated therewith one or more discharge guide vanes the or each discharge guide vanes being located at a main face of at least one of the shrouds.6. ...

GAS TURBINE ENGINE AIRFOIL

An airfoil for a turbine engine includes an airfoil that has pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a leading edge dihedral and a span position. The leading edge dihedral is negative from the 0% span position to the 100% span position. A positive dihedral corresponds to suction side-leaning, and a negative dihedral corresponds to pressure side-leaning. The airfoil is a fan blade for a gas turbine engine. The airfoil has a relationship between a trailing edge dihedral and a span position. The trailing edge dihedral is positive from the 0% span position to the 100% span position. A positive dihedral corresponds to suction side-leaning and a negative dihedral corresponds to pressure side-leaning. 1. An airfoil for a turbine engine comprising:an airfoil having pressure and suction sides extending in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip, wherein the airfoil has a relationship between a leading edge dihedral and a span position, the leading edge dihedral negative from the 0% span position to the 100% span position, wherein a positive dihedral corresponds to suction side-leaning, and a negative dihedral corresponds to pressure side-leaning;wherein the airfoil is a fan blade for a gas turbine engine; andwherein the airfoil has a relationship between a trailing edge dihedral and a span position, the trailing edge dihedral positive from the 0% span position to the 100% span position, wherein a positive dihedral corresponds to suction side-leaning, and a negative dihedral corresponds to pressure side-leaning.2. The airfoil according to claim 1 , wherein the leading edge dihedral at the 0% span position is in the range of −3° to −12°.3. The airfoil according to claim 2 , wherein the leading edge dihedral at the 0% span position is about −4°.4. ...

INTEGRATED STRUT AND IGV CONFIGURATION

A strut and IGV configuration in a gas turbine engine positioned at an upstream of a rotor includes a plurality of radial struts, for example for bearing engine loads, and a plurality of inlet guide vanes positioned axially spaced apart from the struts. The number of inlet guide vanes is greater than the number of struts. The struts are circumferentially aligned with the inlet guide vanes. 1. A method of providing an aircraft gas turbine engine , the method comprising:a) providing a plurality of circumferentially-spaced struts radially extending across an inlet flow passage leading to an engine rotor;b) providing a plurality of variable inlet guide vanes between the struts and the rotor, the number of variable inlet guide vanes being greater than the number of struts;c) circumferentially positioning the variable inlet guide vanes to allow the struts to circumferentially align with a respective one of the variable inlet guide vanes; andd) adjusting a position of a rotation axis of the respective variable inlet guide vanes such that in use a flow direction of air passing around each strut forms a wake which is then substantially redirected by a variable inlet guide vane when the variable inlet guide vane is in a maximum setting angle.2. The method as defined in further comprising a step of determining a chord length of the respective variable inlet guide vanes in a range of 10% to 200% of an axial gap between the struts and the variable inlet guide vanes.3. The method as defined in further comprising a step of determining a chord length of the respective variable inlet guide vanes in a range of 30% to 100% of an axial gap between the struts and the variable inlet guide vanes.4. The method as defined in further comprising a step of determining a chord length of said one of the variable inlet guide vanes circumferentially aligned with the respective struts claim 1 , greater or smaller than a chord length of the remaining variable inlet guide vanes.5. The method as ...

Fan blade platform spacer mounting

In a featured embodiment, a fan rotor comprises a platform. Clevises extend radially inwardly of the platform. Each clevis has an aperture. A hub has hub lugs positioned intermediate spaced ends of the clevises, and apertures. A pin extends through the apertures in the hub and the clevises to connect the platform to the hub. The apertures in the clevises are formed to have an inner surface for supporting the pin. A method of forming a fan blade platform is also disclosed.

ROTOR BLADE WITH A CONIC SPLINE FILLET AT AN INTERSECTION BETWEEN A PLATFORM AND A NECK

A rotor blade for a turbine engine includes an airfoil that is connected to a base. The base includes a platform, a neck and a fillet. The fillet extends along at least a portion of an intersection between the platform and the neck. The fillet has a radius that substantially continuously changes as the fillet extends from the platform to the neck. 1. A rotor blade for a turbine engine , comprising:an airfoil connected to a base;the base including a platform, a neck and a fillet that extends along at least a portion of an intersection between the platform and the neck;wherein the fillet has a radius that substantially continuously changes as the fillet extends from the platform to the neck.2. The rotor blade of claim 1 , wherein the radius increases as the fillet extends from the platform to the neck.3. The rotor blade of claim 1 , wherein a cross-sectional geometry of the fillet changes as the fillet extends along the intersection.4. The rotor blade of claim 1 , wherein the fillet extends along substantially an entire length of the intersection.5. The rotor blade of claim 1 , whereinthe fillet comprises a first fillet, and the neck further includes a second fillet that extends along a portion of the intersection between the first fillet and an end of the intersection; andthe second fillet has a substantially constant radius.6. The rotor blade of claim 1 , wherein the fillet extends at least along a concave portion of the intersection.7. The rotor blade of claim 1 , wherein the fillet is located on a suction side of the base.8. The rotor blade of claim 1 , wherein the fillet is located within a pocket of the base.9. The rotor blade of claim 1 , wherein the platform includes an under platform surface with a substantially flat cross-sectional geometry that extends from an edge of the platform to the fillet.10. The rotor blade of claim 1 , wherein the airfoil and the platform are configured for a turbine section of the turbine engine.11. The rotor blade of claim 1 , ...

Rotating circular airfoil and propeller system

A rotating circular airfoil system () that creates lift when rotates about its axis () is disclosed. This circular airfoil is symmetrical about its axis and has a body-sectional-shape () similar to those of modern airplane wings. It is hollow in the middle to create a circular inner edge (), which is the leading edge of the airfoil while its outer edge () is the trailing edge. Angle of attack () is symmetrically provided. Pairs of fins () are attached or embodied onto the top and the bottom of the airfoil. These sectional-tee-shape fins run radiately from the inner edge of the circular airfoil. Spokes () from the center of the circular airfoil attached to its inner edge enable it to be rotated by a turn-shaft that aligns to the airfoil axis through center hole (). When this circular airfoil rotates about its axis in the air, the surrounding air is contact by the fins and forced to travel from the inner edge to the outer edge of the circular airfoil, as a result of the centrifugal action. While the air continuously travels from the leading edge () to the trailing edge () of the airfoil, it creates the well known lift force () of the airfoil principal. This circular airfoil system also operates in water or other fluids to create lift or propelling force. 1. A rotating circular airfoil and propeller system comprises of a main circular body with a circular hollow at its center to provide the uniform inner edges of the circular airfoil where its cross-sectional shape begins as the leading edge of the airfoil toward the outer edge which serves as trailing edge of the airfoil , pairs of tee-shape fins radiate from the inner edge to the outer edge at the top and the bottom of the circular airfoil , set of spokes from the center of the airfoil are attached to the said main circular body with the provision at the center for the attachment of a turn shaft to rotate the said circular airfoil and propeller system.2. A rotating circular airfoil and propeller system of claim 1 , ...

IMPELLER FOR AN EXHAUST GAS TURBOCHARGER

An impeller for an exhaust gas turbocharger may include a hub main body and blades arranged thereon. The hub main body may be configured as a polygon with a number of segments that may be tilted with respect to one another, the number of the segments corresponding to a number of the blades. Alternatively, the hub main body may have a main surface that faces the blades and undulates in a circumferential direction, a number of the undulations corresponding to the number of the blades. 1. An impeller for an exhaust gas turbocharger , comprising: the hub main body is configured as a polygon with a number of segments that are tilted with respect to one another, the number of the segments corresponding to a number of the blades; or', 'the hub main body has a main surface that faces the blades and undulates in a circumferential direction, a number of undulations corresponding to the number of the blades., 'a hub main body and blades arranged thereon, wherein one of2. An impeller according to claim 1 , wherein the hub main body is configured as a polygon claim 1 , each of the number of segments having a main surface of straight cross section radially on an outside.3. An impeller according to claim 1 , wherein the hub main body is configured as a polygon claim 1 , and a transition from a segment into an associated blade is rounded.4. An impeller according to claim 3 , wherein the rounded transition is formed by way of a material addition to the main surface of the respective segment.5. An impeller according to claim 1 , wherein the hub main body has a main surface claim 1 , and a transition from the main surface into an associated blade is arranged in a region of an undulation peak.6. An impeller according to claim 5 , wherein the transition is rounded.7. An impeller according to claim 5 , wherein the transition into the main surface is formed by way of a tangent applied to an undulation slope.8. An impeller according to claim 1 , wherein the hub main body has a back that ...

IMPELLER ASSEMBLY OF FLUID ROTARY MACHINE AND MANUFACTURING METHOD THEREOF

Provided is a method of manufacturing an impeller assembly, the method including providing an impeller including: a rotary shaft; a base portion radially extending outward from the rotary shaft; and a plurality of blades extending radially outward from the rotary shaft and disposed on the base portion, each of the plurality of blades provided apart from one another in a circumferential direction around the rotary shaft; providing a mold in an area between the plurality of blades; and forming a shroud covering upper portions of the plurality of blades and an upper portion of the mold, wherein the forming the shroud comprises applying a melted metal on the upper portions of the plurality of blades and the upper portion of the mold.

TURBINE WHEEL, RADIAL TURBINE, AND SUPERCHARGER

Suction surfaces of blades of this radial turbine each have: a leading edge side of blade tip including a leading edge and the boundary between the suction surface and the tip; and a trailing edge side of blade tip including a trailing edge and the boundary between the suction surface and the tip. The leading edge side of blade tip forms a concave curved surface which is recessed towards the side opposite to the rotation side in a radial view. The trailing edge side of blade tip forms a convex curved surface which protrudes towards the rotation side in a radial view. 1. A turbine wheel comprising:a disk which has a shape rotationally symmetrical about an axis and a diameter which gradually decreases from a front side which is one side in an axial direction in which the axis extends toward a rear side which is the other side;a plurality of blades which are fixed to an outer peripheral surface of the disk at intervals in a circumferential direction D with respect to the axis,wherein each of the blades includesa leading edge which extends in a direction including an axial component from a portion on the front side of the disk and faces a radially outer side with respect to the axis,a trailing edge which extends in a direction including a radial component with respect to the axis from a portion on the rear side of the disk and faces the rear side,a pressure surface and a suction surface which extend from the leading edge to the trailing edge and face sides opposite to each other,a tip which forms an edge on a side far from the outer peripheral surface,wherein the suction surface includes a leading edge side of blade tip including a boundary between the suction surface and the tip and the leading edge and a trailing edge side of blade tip including a boundary between the suction surface and the tip and the trailing edge,wherein the leading edge side of blade tip forms a concave curved surface which is recessed to an counterrotation side from the suction surface toward ...

PROPELLER FAN

A propeller fan includes a shaft disposed on a rotation axis and a blade disposed adjacent to an outer circumferential surface of the shaft. The blade has a leading edge and a trailing edge. At least one of the leading edge and the trailing edge has a notch. The notch includes a pair of side edge-parts forming an acute included angle and bottom edge-part located between the pair of side edge-parts. The bottom edge-part includes at least one protrusion having an obtuse included angle. 1. A propeller fan comprising:a shaft disposed on a rotation axis; anda blade disposed adjacent to an outer circumferential surface of the shaft, the blade having a leading edge and a trailing edge,wherein at least one of the leading edge and the trailing edge has a notch,wherein the notch includes a pair of side edge-parts forming an acute included angle and bottom edge-part located between the pair of side edge-parts,wherein the bottom edge-part includes at least one first protrusion having an obtuse included angle, andwherein the first protrusion is located in middle part of the notch in a radial direction of the blade.2. (canceled)3. The propeller fan of claim 1 , wherein the first protrusion includes one or more arcs.4. The propeller fan of claim 1 ,wherein the notch includes two recesses arranged on opposite sides of the first protrusion, andwherein at least one of the two recesses has an obtuse included angle.5. The propeller fan of claim 1 , wherein the at least one first protrusion of the bottom edge-part comprises a plurality of first protrusions.6. A propeller fan comprising:a shaft disposed on a rotation axis; anda blade disposed adjacent to an outer circumferential surface of the shaft, the blade having a leading edge and a trailing edge,wherein at least one of the leading edge and the trailing edge has a notch,wherein the notch includes a pair of side edge-parts forming an acute included angle and bottom edge-part located between the pair of side edge-parts, andwherein the ...

MORPHING TRAILING EDGE DEVICE FOR AN AIRFOIL

A morphing trailing edge device for an airfoil includes a skin member, an actuator, and a torsion element. The skin member is configured to extend on a surface of a trailing edge region of an airfoil and includes a load introduction point within a reinforced area. The skin member further includes a stiffening member arranged essentially perpendicular to the load introduction point, wherein the load introduction point, the reinforced area and the stiffening member are integrated into the skin member. The actuator is configured to drive the torsion element with an actuation load. The torsion element is configured to translate the actuation load to the load introduction point, so that the actuation load morphs the trailing edge region up- or downwardly relative to a horizontal plane. A morphing airfoil for an aircraft, an aircraft with a morphing airfoil and a method for manufacturing a morphing airfoil are also described. 1. A morphing trailing edge device for an airfoil , the device comprising:a skin member;an actuator; anda torsion element,wherein the skin member is configured to extend on a surface of a trailing edge region of an airfoil and comprises a load introduction point within a reinforced area,wherein the skin member further comprises a stiffening member arranged essentially perpendicular to the load introduction point,wherein the load introduction point, the reinforced area and the stiffening member are integrated into the skin member,wherein the actuator is configured to drive the torsion element with an actuation load, andwherein the torsion element is configured to translate the actuation load to the load introduction point, so that the actuation load morphs the trailing edge region up- or downwardly relative to a horizontal plane.2. The morphing trailing edge device according to claim 1 , wherein the trailing edge region is a flap region claim 1 , a spoiler region claim 1 , a vertical stabilizer region claim 1 , a horizontal stabilizer region claim 1 , ...

GAS TURBINE SEALING BAND ARRANGEMENT HAVING AN UNDERLAP SEAL

A sealing band arrangement for a gas turbine including first and second adjoining rotor disks each including a disk arm having a slot. The sealing band arrangement includes at least one first seal strip segment located within the slots, wherein the seal strip segment includes first and second ends. The sealing band arrangement also includes a tab section that extends from the first end in order to form an underlap seal with an adjacent second seal strip segment. The underlap seal enables a thickness of the first and second ends to be substantially equivalent to a thickness of the first seal strip segment in order to improve wear life of the seal strip segment. The sealing band arrangement further includes at least one wide portion formed in the first seal strip segment wherein the wide portion is wider than a remaining portion of the first seal strip segment. 1. A sealing band arrangement for a gas turbine , wherein the gas turbine includes first and second adjoining rotor disks each including a disk arm wherein the disk arms are separated by a disk arm gap , comprising:at least one first seal strip segment located within the disk arm gap, wherein the seal strip segment includes first and second ends; anda tab section extending from the first end for forming an underlap seal with an adjacent second seal strip segment wherein a thickness of the first and second ends is substantially equivalent to a thickness of the first seal strip segment.2. The sealing band arrangement according to claim 1 , wherein the first end is separated from the second seal strip segment by a seal strip gap and the tab section seals the seal strip gap.3. The sealing band arrangement according to claim 1 , wherein the tab section is unistructurally formed with the first end.4. The sealing band arrangement according to claim 1 , wherein the tab section is approximately 9.5 mm wide.5. The sealing band arrangement according to claim 1 , wherein the first and second ends and the first seal strip ...

IMPELLER FOR FUEL PUMP OF VEHICLE

Provided is an impeller for a fuel pump of a vehicle capable of decreasing a magnitude of high frequency fluid noise due to high speed rotation of the impeller by upper and lower blades of impeller blades positioned between upper and lower casings of the fuel pump and coupled to a shaft of a driving motor to deliver a fuel by rotational force so as to have asymmetrical angles based on the center of a thickness of an impeller body in sucking the fuel from a fuel tank and supplying the fuel to an engine of an internal combustion engine. 1. An impeller for a fuel pump of a vehicle comprising:an impeller body having a disk shape and having a shaft fixing hole at the center thereof so as to penetrate therethrough so that a shaft of a driving motor is inserted thereinto and coupled thereto; anda plurality of blades formed at predetermined intervals along an outer circumferential surface of the impeller body and formed in an outward direction of the circumferential surface,wherein each of the blades includes an upper blade formed at an upper side of the impeller body in a thickness direction and a lower blade formed at a lower side of the impeller body in the thickness direction, andan angle of the upper blade is larger than that of the lower blade.2. The impeller for a fuel pump of a vehicle of claim 1 , wherein the angle of the upper blade is larger than that of the lower blade by 3 to 5 degrees.3. The impeller for a fuel pump of a vehicle of claim 2 , wherein a sum of the angle of the upper blade and the angle of the lower blade is 90 to 100 degrees.4. The impeller for a fuel pump of a vehicle of claim 1 , wherein a height of the upper blade is the same as that of the lower blade.5. The impeller for a fuel pump of a vehicle of claim 1 , further comprising a side ring formed on outer circumferential surfaces of the plurality of blades so as to form blade chambers allowing discharge and introduction of a fuel to be made at upper and lower sides of the blade claim 1 , ...

UNI-DIRECTIONAL AXIAL TURBINE BLADE ASSEMBLY

A uni-directional axial turbine blade assembly comprises a central disc member coupled to an elongated shaft of a motor or generator, a rotational support member mounted on outer circumference of the central disc member, an outer ring mounted on outer circumference of the rotational support member and a plurality of blade members spaced equidistantly, extending radially outward from the outer ring. Each blade member includes a straight leading edge, a curved trailing edge, an exterior surface including a first region having a curved surface, a second region having a flat surface and a convex surface extending from the flat surface and terminating to a radially inward portion and an interior surface including a third region having a concave surface. The preferred embodiment renders the formation of a unique blade configuration that facilitates uni-directional rotation of the blade assembly irrespective of changes in flow direction of fluid. 1. An axial turbine blade assembly for propellers , the axial turbine blade assembly comprising: a central disc member coupled to an elongated shaft of a motor or generator;a rotational support member mounted on outer circumference of the central disc member; an outer ring mounted on outer circumference of the rotational support member; and a plurality of blade members extending radially outward from the outer ring, each of the plurality of blade members having an exterior surface, an interior surface, a leading edge and a trailing edge; the exterior surface includes a first region having a curved surface along the leading edge, a second region having a flat surface along the trailing edge, the flat surface being attached proximate the central disc member and a convex surface extending from the flat surface along the trailing edge and terminating to a radially inward portion; and the interior surface includes a third region having a concave surface proximate the trailing edge;whereby the curved surface along the leading edge, the ...

AIRFOIL WITH RIB HAVING CONNECTOR ARMS

An airfoil includes an airfoil wall that defines a leading end, a trailing end, and first and second sides that join the leading end and the trailing end. A rib connects the first and second sides of the airfoil wall. The rib defines a tube portion that circumscribes a rib passage, and first and second connector arms that solely join the tube portion to, respectively, the first and second sides of the airfoil wall. 1. An airfoil comprising:an airfoil wall defining a leading end, a trailing end, and first and second sides joining the leading end and the trailing end; and a tube portion circumscribing a rib passage, and', 'first and second connector arms solely joining the tube portion to, respectively, the first and second sides of the airfoil wall., 'a rib connecting the first and second sides of the airfoil wall, the rib defining'}2. The airfoil as recited in claim 1 , wherein the airfoil wall and the rib bound a cooling channel there between claim 1 , and the cooling channel is flow isolated from the rib passage.3. The airfoil as recited in claim 1 , wherein the airfoil wall and the rib bound a cooling channel there between claim 1 , and the rib includes at least one cooling aperture connecting the cooling channel and the rib passage.4. The airfoil as recited in claim 1 , wherein the tube portion includes forward and aft walls and first and second side walls joining the forward and aft walls claim 1 , and the first connector arm projects from the first side wall and the second connector arm projects from the second side wall.5. The airfoil as recited in claim 4 , wherein the airfoil wall and the rib bound a cooling channel there between claim 4 , and at least one of the first and second side walls includes at least one cooling aperture connecting the rib passage and the cooling channel6. The airfoil as recited in claim 4 , wherein the airfoil wall and the rib bound a cooling channel there between claim 4 , at least one of the first and second side walls includes ...

ROTARY AIRFOIL

A rotary airfoil in a gas turbine engine is provided. The airfoil includes opposed pressure and suction sides joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. An axial component of a center of gravity of a cross-section taken chordally toward the tip of the airfoil being upstream relative to an axial component of a center of gravity of a cross-section taken chordally toward the root of the airfoil. A method for forming such blade is also presented. 1. A rotary airfoil in a gas turbine engine , the airfoil comprising:opposed pressure and suction sides joined together at chordally opposite leading and trailing edges, the pressure and suction sides extending in a span direction from a root to a tip; andan axial component of a center of gravity of a cross-section taken chordally toward the tip of the airfoil being upstream relative to an axial component of a center of gravity of a cross-section taken chordally toward the root of the airfoil.2. The rotary airfoil of claim 1 , a thickness is defined across the airfoil between the pressure and suction sides claim 1 , the thickness having a chord-wise distribution and a maximum thickness claim 1 , a thick region of a given cross-section being defined by a thickness at at least 85% of the maximum thickness and at most between −15% and +15% from the maximum thickness claim 1 , a thick region of the cross-section toward the root being chordally longer than a thick region of the cross-section toward the tip.3. The rotary airfoil of claim 2 , wherein the thick region of the cross-section toward the root is comprised between about 30% and 60% of the chord taken form the leading edge.4. The rotary airfoil of claim 3 , wherein the thick region of the cross-section toward the tip is comprised between about 30% and 45% of the chord taken form the leading edge.5. The rotary airfoil of claim 1 , wherein a thickness is defined across the ...

Turbine

A turbine includes a turbine rotor, a stationary body that covers the turbine rotor, and a diffuser provided on an outlet side of the stationary body. Last-stage moving blades of the turbine rotor include blade sections and covers provided at distal ends of the blade sections. The diffuser is formed such that an outer circumferential surface of an inlet section is small in diameter with respect to an inner circumferential surface of an outlet section of the stationary body and a circumferential wall section of the inlet section at least partially overlaps the covers in a radial direction when viewed from the axial direction. An annular gap space between the stationary body and the covers faces a space on an outer side of an outer circumferential surface of the diffuser when viewed from the axial direction. 1. A turbine comprising:a turbine rotor formed by providing, in an axial direction, a plurality of stages of moving blade rows including pluralities of moving blades arranged in a circumferential direction;a stationary body that covers the turbine rotor; anda diffuser provided on an outlet side of the stationary body, whereinlast-stage moving blades of the turbine rotor include blade sections and covers provided at distal ends of the blade sections, the covers of the blade sections adjacent to one another being coupled to configure an annular shape,the diffuser is formed such that an outer circumferential surface of an inlet section is small in diameter with respect to an inner circumferential surface of an outlet section of the stationary body and a circumferential wall section of the inlet section at least partially overlaps the covers in a radial direction when viewed from the axial direction, andan annular gap space between the stationary body and the covers faces a space on an outer side of an outer circumferential surface of the diffuser when viewed from the axial direction.2. The turbine according to claim 1 , wherein the diffuser is formed such that the ...

STEAM TURBINE

A steam turbine according to an embodiment includes: rotor blade cascades each made up at a turbine rotor an inner casing where the turbine rotor is provided to penetrate; an outer casing surrounding the inner casing stationary blade cascades each made up at an inner side of the inner casing and an annular diffuser provided at a downstream side of a final turbine stage, formed by a steam guide and a bearing cone and discharging steam toward outside in a radial direction. An enlarged inclination angle θ1 of an inner surface of a diaphragm outer ring at the final turbine stage relative to a turbine rotor axial direction is an enlarged inclination angle θ2 of an inner surface at an inlet of the steam guide relative to the turbine rotor axial direction or more. 1. A steam turbine , comprising:a turbine rotor;rotor blade cascades each made up by implanting plural rotor blades in a circumferential direction to the turbine rotor;an inner casing where the turbine rotor including the rotor blade cascades is provided to penetrate;an outer casing surrounding the inner casing;stationary blade cascades each made up by attaching plural stationary blades in the circumferential direction between diaphragm outer rings and diaphragm inner rings provided at an inner side of the inner casing, and disposed alternately with the rotor blade cascades in a turbine rotor axial direction; andan annular diffuser provided at a downstream side of a final turbine stage from among turbine stages each made up by the stationary blade cascade and the rotor blade cascade at immediately downstream of the stationary blade cascade, formed by a steam guide and a bearing cone at an inner side of the steam guide, and discharging steam passing through the final turbine stage toward outside in a radial direction,wherein an enlarged inclination angle θ1 of an inner surface of the diaphragm outer ring where an outer periphery of the stationary blade at the final turbine stage is attached relative to the turbine ...

WHEEL OF A FLUID FLOW MACHINE

A blade wheel of a turbomachine, which blade wheel has a multiplicity of blades which are suitable and provided for extending radially in a flow path of the turbomachine, wherein the blades form a blade entry angle and a blade exit angle. Provision is made whereby the blade wheel forms N blocks of blades, where N≥2, wherein the blades of a block have in each case the same blade entry angle and the same blade exit angle, and the blades of at least two mutually adjacent blocks have a different blade entry angle and/or a different blade exit angle. According to a further aspect of the invention, partial gaps that the blades form in relation to an adjacent flow path boundary are varied in mutually adjacent blocks. 128.-. (canceled)30. The blade wheel according to claim 29 , wherein the blades of at least two mutually adjacent blocks have a different blade entry angle and a different blade exit angle by virtue of the fact that the blades of the blocks claim 29 , in the case of identical shaping of the blades claim 29 , form a different stagger angle α.31. The blade wheel according to claim 29 , wherein the blades of at least two mutually adjacent blocks have a different blade entry angle or a different blade exit angle by virtue of the fact that the blades of the blocks have a different shape.32. The blade wheel according to claim 29 , wherein at least two of the blocks have a different extent angle in a circumferential direction claim 29 , wherein the blocks with different extent angle have a different number of blades.33. The blade wheel according to claim 29 , wherein the blades of a block are opened in relation to a nominal blade setting and the blades of an adjacent block are closed in relation to the nominal blade setting.38. The blade wheel arrangement according to claim 37 , wherein the second blade wheel and the third blade wheel are formed as blade wheels claim 37 , wherein the two blade wheels form the same number of N blocks of blades claim 37 , where N≥2.39. ...

POWER TURBINE VANE AIRFOIL PROFILE

A power turbine includes a first stage vane having an airfoil with a cold un-coated nominal profile substantially in accordance with at least an intermediate portion of the Cartesian coordinate values of X, Y and Z set forth in Table 2. The X and Y values are distances, which when smoothly connected by an appropriate continuing curve, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form a complete airfoil shape. 1. A turbine vane of a gas turbine engine having a gaspath , the turbine vane comprising an airfoil having an intermediate portion contained within the gaspath and defined by a nominal un-coated profile substantially in accordance with Cartesian coordinate values of X , Y , and Z of Sections 2 to 8 set forth in Table 2 , wherein the point of origin of the orthogonally related axes X , Y and Z is located at an intersection of a centerline of the gas turbine engine and a stacking line of the turbine vane , the Z values are radial distances measured along the stacking line , the X and Y are coordinate values defining the profile at each distance Z , wherein X and Y values define a set of points for each Z value which when connected by smooth continuing arcs define an airfoil profile section , the profile sections at the Z distances being joined smoothly with one another to form an airfoil shape of the intermediate portion.2. The turbine vane as defined in claim 1 , wherein the turbine vane is a power turbine vane of the gas turbine engine.3. The turbine vane as defined in claim 2 , wherein the power turbine vane is a first stage power turbine vane of a multi-stage power turbine.4. The turbine vane as defined in claim 1 , wherein the turbine vane has a manufacturing tolerance of ±0.009 inches in a direction perpendicular to the airfoil.5. A turbine vane for a gas turbine engine having a gaspath claim 1 , the turbine vane having an intermediate airfoil portion contained within the ...

AIRFOIL SHAPE FOR A COMPRESSOR

An article of manufacture having a nominal airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in a scalable table identified as TABLE 1, wherein the Cartesian coordinate values of X, Y, and Z are non-dimensional values convertible to dimensional distances by multiplying the Cartesian coordinate values of X, Y, and Z by a number, and wherein X and Y are coordinates which, when connected by continuing arcs, define airfoil profile sections at each Z height, the airfoil profile sections at each Z height being joined with one another to form a complete airfoil shape. The resulting article may be used as a stator vane in a compressor. 1. An article of manufacture comprising a nominal airfoil profile substantially in accordance with Cartesian coordinate values of X , Y , and Z set forth in a scalable table identified as TABLE 1 , wherein the Cartesian coordinate values of X , Y , and Z are non-dimensional values convertible to dimensional distances by multiplying the Cartesian coordinate values of X , Y , and Z by a number; and wherein X and Y are coordinates which , when connected by continuing arcs , define airfoil profile sections at each Z height , the airfoil profile sections at each Z height being joined smoothly with one another to form a complete airfoil shape.2. The article of manufacture according to claim 1 , wherein the article of manufacture comprises an airfoil.3. The article of manufacture according to claim 1 , wherein the article of manufacture comprises a stator vane configured for use with a compressor.4. The article of manufacture according to claim 1 , wherein the airfoil shape lies in an envelope within at least one of:+/−5% of a chord length in a direction normal to an airfoil surface location; and+/−0.25 inches in a direction normal to an airfoil surface location.5. The article of manufacture according to claim 1 , wherein the number claim 1 , used to convert the non-dimensional values to ...

STEAM TURBINE

A steam turbine according to an embodiment of the present invention includes: a rotor configured to rotate about an axis; a casing which houses the rotor rotatable; and a first stage including a first-stage stationary vane fixed to an inner wall portion of the casing and a first-stage rotor blade fixed to the rotor at downstream of the first-stage stationary vane. The rotor includes a first cavity having a concave shape and being formed on a portion facing the first-stage stationary vane, the first cavity being in communication with an inner space defined between the inner wall portion and the rotor at upstream of the first-stage stationary vane. The first-stage stationary vane includes a first-stage through hole which is in communication with the first cavity and which is formed through the first-stage stationary vane in a radial direction. 1. A steam turbine , comprising:a rotor configured to rotate about an axis;a casing which houses the rotor rotatably; anda first stage including a first-stage stationary vane fixed to an inner wall portion of the casing and a first-stage rotor blade fixed to the rotor at downstream of the first-stage stationary vane,wherein the rotor includes a first cavity having a concave shape and being formed on a portion facing the first-stage stationary vane, the first cavity being in communication with an inner space defined between the inner wall portion and the rotor at upstream of the first-stage stationary vane,wherein the first-stage stationary vane includes a first-stage through hole which is in communication with the first cavity and which is formed through the first-stage stationary vane in a radial direction, andwherein the steam turbine is configured such that steam introduced from the first cavity via an inlet opening of the first-stage through hole flows through the first-stage through hole.2. The steam turbine according to claim 1 , wherein the first stage is positioned upstream of a final stage of the steam turbine.3. The ...

TURBOMACHINERY

A turbomachine () configured to compress supercritical carbon dioxide is shown. The turbomachine comprises, in fluid flow series, an inlet (), an inducerless radial impeller () having a plurality of blades, and a fully vaneless diffuser (). The inlet is radially flared to induce a radial component in flow prior to an entry to the impeller. 1. A turbomachine configured to compress supercritical carbon dioxide , the turbomachine comprising , in fluid flow series:an inlet;an inducerless radial impeller having a plurality of blades; anda fully vaneless diffuser;wherein the inlet is radially flared to induce a radial component in flow prior to an entry to the impeller.2. The turbomachine of claim 1 , in which a hub hade angle of the impeller at the entry thereto (γ) is from 50 to 70 degrees.3. The turbomachine of claim 2 , in which said hade angle (γ) is 60 degrees.4. The turbomachine of claim 1 , in which each of the plurality of blades is a backswept blade.5. The turbomachine of claim 4 , in which each of the plurality of blades have a blade exit angle (χ) of from −50 to −70 degrees.6. The turbomachine of claim 5 , in which each of the plurality of blades have a blade exit angle (χ) of −60 degrees.7. The turbomachine of claim 1 , in which the plurality of blades comprises:a set of main blades; anda set of splitter blades.8. The turbomachine of claim 1 , in which a meridional chord length of the splitter blades (c) is 70 percent of a meridional chord length of the main blades (c).9. The turbomachine of claim 7 , in which the impeller comprises one splitter blade for each main blade.10. The turbomachine of claim 1 , in which the radius of the inlet (r) is from 25 to 50 percent of the radius of the impeller (r).11. The turbomachine of claim 10 , in which the radius of the inlet (r) is from 30 to 50 percent of the radius of the impeller (r).12. The turbomachine of claim 1 , in which the diffuser has an annulus height ratio (b/b) of 1.13. The turbomachine of claim 1 , in ...

TURBOMACHINERY

A turbomachine () configured to compress supercritical carbon dioxide is shown. The turbomachine comprises, in fluid flow series, an inlet (), an inducerless radial impeller () having a plurality of backswept blades () each of which have a blade exit angle (χ) of from −50 to −70 degrees, and a fully vaneless diffuser (). 1. A turbomachine configured to compress supercritical carbon dioxide , the turbomachine comprising , in fluid flow series:an inlet;{'sub': '2', 'an inducerless radial impeller having a plurality of backswept blades each of which have a blade exit angle (χ) of from −50 to −70 degrees; and'}a fully vaneless diffuser.2. The turbomachine of claim 1 , in which the backswept blades have a blade exit angle (χ) of −60 degrees.3. The turbomachine of claim 1 , in which the inlet is radially flared to induce a radial component in flow prior to an entry to the impeller.4. The turbomachine of claim 3 , in which a hub hade angle of the impeller at the entry thereto (γ) is from 50 to 70 degrees.5. The turbomachine of claim 4 , in which said hade angle (γ) is 60 degrees.6. The turbomachine of claim 1 , in which each of the plurality of blades is a backswept blade.7. The turbomachine of claim 1 , in which the plurality of blades comprises:a set of main blades; anda set of splitter blades.8. The turbomachine of claim 7 , in which a meridional chord length of the splitter blades (c) is 70 percent of a meridional chord length of the main blades (c).9. The turbomachine of claim 7 , in which the impeller comprises one splitter blade for each main blade.10. The turbomachine of claim 1 , in which the radius of the inlet (r) is from 25 to 50 percent of the radius of the impeller (r).11. The turbomachine of claim 10 , in which the radius of the inlet (r) is from 30 to 50 percent of the radius of the impeller (r).12. The turbomachine of claim 1 , in which the diffuser has an annulus height ratio (b/b) of 1.13. The turbomachine of claim 1 , in which the radius of the diffuser ...

TURBINE BLADE ANGEL WING WITH PUMPING FEATURES

A gas turbine engine, including: a plurality of blades () assembled into an annular row of blades and partly defining a hot gas path () and a cooling fluid path (), wherein the cooling fluid path extends from a rotor cavity () to the hot gas path; an angel wing assembly () disposed on a side () of a base () of the row of blades; and pumping features () distributed about the angel wing assembly configured to impart, at a narrowest gap () of the cooling fluid path, motion to a flow of cooling fluid flowing there through. The plurality of pumping features, the angel wing assembly, and the base of the row of blades are effective to produce a helical motion to the flow of cooling fluid as it enters the hot gas path. 1. A gas turbine engine , comprising:a plurality of blades assembled into an annular row of blades about a gas turbine engine longitudinal axis and partly defining both a hot gas path and a cooling fluid path, wherein the cooling fluid path extends from a rotor cavity, past a side of a radially inward base of the row of blades where the side is upstream with respect to a flow of hot gases in the hot gas path, and leads to the hot gas path;an angel wing assembly disposed on the side of the base of the row of blades; anda plurality of pumping features distributed about the angel wing assembly configured to impart, at a narrowest gap of the cooling fluid path defined by the angel wing, motion to a flow of cooling fluid flowing there through,wherein the plurality of pumping features, the angel wing assembly, and the base of the row of blades are effective to produce a helical motion about the gas turbine engine longitudinal axis to the flow of cooling fluid as it enters the hot gas path.2. The gas turbine engine of claim 1 , wherein with respect to the gas turbine engine longitudinal axis the plurality of pumping features are integral to a portion of the angel wing assembly radially inward of and axially adjacent to an opposing surface claim 1 , wherein the ...

POWER TURBINE BLADE AIRFOIL PROFILE

A power turbine includes a first stage blade having an airfoil with a cold un-coated nominal profile substantially in accordance with at least an intermediate portion of the Cartesian coordinate values of X, Y and Z set forth in Table 2. The X and Y values are distances, which when smoothly connected by an appropriate continuing curve, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form a complete airfoil shape. 1. A turbine blade of a gas turbine engine having a gaspath , the turbine blade comprising an airfoil having an intermediate portion contained within the gaspath and defined by a nominal un-coated profile substantially in accordance with Cartesian coordinate values of X , Y , and Z of Sections 2 to 8 set forth in Table 2 , wherein the point of origin of the orthogonally related axes X , Y and Z is located at an intersection of a centerline of the gas turbine engine and a stacking line of the turbine blade , the Z values are radial distances measured along the stacking line , the X and Y are coordinate values defining the profile at each distance Z , wherein X and Y values define a set of points for each Z value which when connected by smooth continuing arcs define an airfoil profile section , the profile sections at the Z distances being joined smoothly with one another to form an airfoil shape of the intermediate portion.2. The turbine blade as defined in claim 1 , wherein the turbine blade is a power turbine blade of the gas turbine engine.3. The turbine blade as defined in claim 2 , wherein the power turbine blade is a first stage power turbine blade of a multi-stage power turbine.4. The turbine blade as defined in claim 1 , wherein the turbine blade has a manufacturing tolerance of ±0.015 inches in a direction perpendicular to the airfoil.5. A turbine blade for a gas turbine engine having a gaspath claim 1 , the turbine blade having an intermediate airfoil portion ...

FLUTTER-RESISTANT BLADE

An aircraft engine having a turbine, the turbine 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 of the blades 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 of the blades 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 of the blades 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 of the blades 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 ...

Stress relieving feature in gas turbine blade platform

A gas turbine engine blade ( 10 ) having: an airfoil ( 12 ) having a leading edge ( 14 ), a trailing edge ( 16 ), a pressure side ( 22 ), and a suction side ( 24 ); an inner platform ( 30 ) associated with a base ( 18 ) of the airfoil, wherein the inner platform includes a platform aft face ( 34 ) spanning between an platform pressure side face ( 36 ) and a platform suction side face ( 38 ); and a trailing edge undercut ( 70 ) across the aft face. The trailing edge undercut tapers to zero penetration before reaching the platform suction side face. In a side cross section, a leading portion ( 100 ) of the trailing edge undercut is defined by a leading radius ( 102 ), an upper portion ( 104 ) is defined by an upper radius ( 106 ), and a lower portion ( 108 ) is defined by a lower radius ( 110 ). For at least one location along the aft face, values for all three radii are distinct.

ATTACHMENT OF COMPOSITE ARTICLE

A composite article including composite component extending heightwise from a component base to a component tip and lengthwise between spaced apart component first and second edges. Component plies having widthwise spaced apart ply first and second sides and ply edges therebetween. Component mounted on a spar which includes a shank extending heightwise into the composite component, tab at upper end of shank and substantially or fully embedded in the composite component, and tab tip. Ply edges of at least a first portion of the plies directly or indirectly contacting or pressing against the tab. Ply edges of at least a second portion of the plies may directly or indirectly contact or press against the tab tip. Ply edges of first portion may press against one or more indented or recessed surfaces in the tab. The composite article may be a composite blade or vane including a composite airfoil. 1. A composite article comprising:a composite component extending heightwise from a component base to a component tip and lengthwise between spaced apart component first and second edges,the composite component including plies having widthwise spaced apart ply first and second sides and ply edges therebetween,the composite component mounted on a spar including a shank extending heightwise from below the component base up through the component base into the composite component,a tab at an upper end of shank and substantially or fully embedded in the composite component,the tab including heightwise spaced apart tab base and tab tip, andthe ply edges of at least a first portion of the plies directly or indirectly contacting or pressing against the tab.2. The composite article as claimed in further comprising hooking means for hooking some of the plies of the composite airfoil claim 1 ,the hooking means disposed along widthwise spaced apart tab first and second sides of the tab,the hooking means including hooks disposed along lengthwise spaced apart tab first and second edges, andthe ...

PROPELLER

A propeller having a central post to which one or more blades are connected. The blades are disposed and configured to pull air in from the propeller's sides toward the propeller's axis of rotation to create pressure in an area in the vicinity of the center of the propeller's rotating axis for generating thrust. 1. A propeller comprising:a central post coincident with a rotational axis;one or more blades each having a distal end and a proximate end;each of the one or more blades comprising a top section, a bottom section, and a side section, the side section disposed at or toward the distal end;the top section and bottom section of each of the one or more blades connected at the proximate end to and extending radially outward from the central post;a gap between the central post connections of the top section and bottom section of at least one of the one or more blades wherein air is compressed upon rotation of the propeller; andthe one or more blades disposed and configured to pull air inward from the blade's side sections toward the propeller's axis of rotation and from the propeller's front to its back.2. The propeller of wherein the blades are in loop form claim 1 , wherein each of the blades in loop form spins in the same plane of rotation.3. The propeller of wherein at least one of the blade sections exhibits a non-zero blade angle.4. The propeller of wherein the cross-section of at least one of the top section or bottom section of the one or more blades is shaped substantially like an airfoil.5. The propeller of comprising at least one pair of blades opposing one another about the rotational axis.6. The propeller of wherein the pair of opposing blades form a single claim 5 , contiguous loop.7. The propeller of wherein the pair of blades has a first blade intersecting the central post at a first blade intersection and a second opposing blade intersecting the central post at a second blade intersection; andwherein the angle of the first blade intersection with ...

COMPRESSOR BLADE FOR GAS TURBINE ENGINE

A compressor blade for a gas turbine engine is disclosed. The compressor blade may have a root configured to engage a hub of the gas turbine engine, and an airfoil radially extending a distance from the root to a tip. The airfoil may have a suction side, a pressure side, a leading edge connecting the suction and pressure sides, and a trailing edge connecting the suction and pressure sides opposite the leading edge. The distance that the airfoil extends from the root to the tip may be divided into a plurality of radially adjacent regions. At least one, but not all, of the plurality of radially adjacent regions away from the base and the tip may have a substantially constant thickness. 1. A compressor blade for a gas turbine engine , comprising:a root configured to engage a hub of the gas turbine engine; andan airfoil having a suction side, a pressure side, a leading edge connecting the suction and pressure sides, and a trailing edge connecting the suction and pressure sides opposite the leading edge, the airfoil radially extending a distance from the root to a tip, 'the distance that the airfoil extends from the root to the tip is divided into a plurality of radially adjacent regions; and', 'whereinat least one, but not all, of the plurality of radially adjacent regions away from the base and away from the tip have a substantially constant thickness.2. The compressor blade of claim 1 , wherein the substantially constant thickness of the at least one of the plurality of radially adjacent regions extends along a line substantially perpendicular to a chord of the airfoil at a general midpoint of the airfoil.3. The compressor blade of claim 2 , wherein at least one of the leading edge and the trailing edge of the airfoil also has a substantially constant thickness throughout the at least one of the plurality of radially adjacent regions.4. The compressor blade of claim 1 , wherein the at least one of the plurality of radially adjacent regions extends about 10-40% of a ...

STEAM TURBINE

A steam turbine in an embodiment includes a casing, a turbine rotor, a rotor blade cascade, a stationary blade cascade, and a spray unit, the rotor blade cascade and the stationary blade cascade being each arranged at a plurality of stages alternately in an axial direction of the turbine rotor. The turbine rotor is housed inside the casing. In the rotor blade cascade, a plurality of rotor blades are arranged in a circumferential direction of the turbine rotor. In the stationary blade cascade, a plurality of stationary blades are arranged in the circumferential direction of the turbine rotor between a diaphragm inner ring and a diaphragm outer ring. The spray unit sprays spray water to a space located upstream from a rotor blade cascade at a final stage among the rotor blade cascades at the plurality of stages inside the casing. 1. A steam turbine , comprising:a casing;a turbine rotor housed inside the casing;a rotor blade cascade having a plurality of rotor blades arranged in a circumferential direction of the turbine rotor; 'the rotor blade cascade and the stationary blade cascade being each arranged at a plurality of stages alternately in an axial direction of the turbine rotor; and', 'a stationary blade cascade having a plurality of stationary blades arranged in the circumferential direction of the turbine rotor between a diaphragm inner ring and a diaphragm outer ring,'}a spray unit for spraying spray water to a space located upstream from a rotor blade cascade at a final stage among the rotor blade cascades at the plurality of stages inside the casing.2. The steam turbine according to claim 1 ,wherein the spray unit sprays the spray water to a space demarcated by the diaphragm outer ring and the casing.3. The steam turbine according to claim 2 ,wherein the spray unit sprays the spray water to a space demarcated by the diaphragm outer ring provided in a stationary blade cascade at a final stage among the stationary blade cascades at the plurality of stages and ...

MULTI-FLUID CARGO PUMPS

A submerged electrical pump for liquefied hydrocarbon gasses that is adapted for use encompassing a range of different temperatures and viscosities is disclosed. Notable elements of some embodiments of multiple fluid pumps include bearings and bearing liners made from the same material, a motor larger enough to pump the most vicious and dense fluid, extra thick bearing liners, and a trial-and-error process for choosing other pump design specifications, such as impeller wear rings, bushings, and other critical radial clearances. 1. Turbomachinery for multiple liquefied hydrocarbon gasses , comprising:a bearing at an interface between a rotor and a stator;an inner bearing liner affixed to the rotor;an outer bearing liner affixed to the stator; andwherein the bearing, the inner bearing liner, and the outer bearing liner all consist of the same material; the bearing is disposed between the inner bearing liner and the outer bearing liner; and the bearing is lubricated with one of the multiple liquefied hydrocarbon gasses.2. The turbomachinery of claim 1 , wherein the inner bearing liner and outer bearing liner each comprise a thickness that is thick enough to prevent at least one of the inner bearing liner or the outer bearing liner from yielding under pressure applied to an outer liner surface at a coldest potential operating temperature.3. The turbomachinery of claim 1 , wherein at least one of a thickness of at least one of the inner bearing liner or the outer bearing liner claim 1 , or a size of the bearing is determined based on expansion or contraction of the same material under varying temperature conditions or varying pressure conditions associated with the multiple liquefied hydrocarbon gasses.4. The turbomachinery of claim 3 , wherein the varying temperature conditions are based on different boiling points of the multiple liquefied hydrocarbon gasses.5. The turbomachinery of claim 1 , wherein the same material of the bearing claim 1 , the inner bearing liner ...

COMPRESSOR AND GAS TURBINE INCLUDING THE SAME

A compressor includes a compressor screen mounted on the outer circumferential surface of a compressor disk or the inner wall of a compressor casing. The compressor screen is positioned between a row of compressor blades and a row of compressor vanes to stabilize airflow by regulating the amount of air flowing in the compressor. The compressor screen is formed of annular plate having openings through which the compressed air flows, arranged in a pattern to smoothly regulate the flow of air. The compressor screen may be variously configured, for example, to facilitate installation or to guide the flow of air using a tapered surface or a directionality of the openings or a flow guide. A gas turbine includes a combustor, a turbine, and a compressor employing the compressor screen. 1. A compressor comprising:a compressor casing in which a plurality of compressor disks are installed;a plurality of compressor blades fixed to an outer circumferential surface of each compressor disk so as to be arranged in rows facing an inner circumferential surface of the compressor casing;a plurality of compressor vanes fixed to the inner circumferential surface of the compressor casing and arranged in rows interlacing the compressor blades rows; anda compressor screen disposed between adjacently arranged rows of the compressor vanes and the compressor blades.2. The compressor according to claim 1 , wherein the compressor screen includes an annular plate through which a plurality of openings are formed according to a pattern.3. The compressor according to claim 2 , wherein the openings are formed obliquely to have directionality.4. The compressor according to claim 2 , wherein the openings increase in size toward a mounting surface for receiving the compressor screen.5. The compressor according to claim 2 , wherein the compressor screen comprises a flow guide formed on a downstream side of the compressor screen claim 2 , the flow guide extending from an edge of each opening to direct a ...

TURBINE NOZZLES WITH SLIP JOINTS AND METHODS FOR THE PRODUCTION THEREOF

Embodiments of a turbine nozzle are provided, as are embodiments of methods for the manufacture of turbine nozzles. In one embodiment, the turbine nozzle includes a support ring and a slip joint ring, which is substantially concentric with the support ring and radially spaced apart therefrom. The slip joint ring has a plurality of slots therein. A plurality of vanes is fixedly coupled to the support ring and extends radially therefrom into the plurality of slots. A plurality of radial slip joints is formed between the plurality of vanes and the plurality slots. Each slip joint extends around a different one of the plurality of vanes to permit relative radial movement between the plurality of vanes and the slip joint ring during operation of the turbine nozzle. 1. A turbine nozzle , comprising:a support ring;a slip joint ring substantially concentric with the support ring and radially spaced apart therefrom, the slip joint ring having a plurality of circumferentially-spaced slots therein;a plurality of vanes fixedly coupled to the support ring and extending radially therefrom into the plurality of circumferentially-spaced slots; anda plurality of radial slip joints formed between the plurality of vanes and the plurality of circumferentially-spaced slots, each slip joint extending around a different one of the plurality of vanes to permit relative radial movement between the plurality of vanes and the slip joint ring during operation of the turbine nozzle.2. The turbine nozzle of wherein the support ring circumscribes the slip joint ring.3. The turbine nozzle of wherein the support ring is formed from a plurality of arched ring segments.4. The turbine nozzle of wherein the plurality of arched ring segments are bonded along a plurality of bond lines each located between a different pair of the plurality of vanes claim 3 , as taken in a radial direction.5. The turbine nozzle of wherein each arched ring segments is integrally formed with at least one of the plurality of ...

BLADE OF FAN OR COMPRESSOR

To provide a blade of a fan or compressor that is reduced in loss by enlarging a laminar flow region over a blade surface. A blade according to the present disclosure is divided into a subsonic region where the relative Mach number of the inlet air flow during rated operation of a turbofan engine is lower than 0.8 and a transonic region where the relative Mach number is equal to or higher than 0.8. Provided that a parameter (δ) defined according to δ=(βin−β)/(βin−βex) is referred to as a blade surface angle change rate where β denotes an angle formed by a tangent to the blade surface and the axial direction of the turbofan engine, βin denotes the blade surface angle at the leading edge of the blade, and the βex denotes the blade surface angle at the trailing edge, in each of the subsonic region and the transonic region, the minimum value of the blade surface angle change rate on the pressure surface, an upper limit value of the blade surface angle change rate at a predetermined axial location along the chord on the pressure surface, and an upper limit value and a lower limit value of the blade surface angle change rate at a predetermined axial location along the chord on the suction surface are defined. 1wherein the blade is divided into a subsonic region and a transonic region in a height direction, a relative Mach number of an air flow flowing to the blade during rated operation of the turbofan engine being lower than 0.8 in the subsonic region and equal to or higher than 0.8 in the transonic region,a cross section of the blade at each location in the height direction is formed by a concave pressure surface and a convex suction surface each of which extends between a leading edge and a trailing edge of the blade, andin the cross section, {'br': None, 'δ=(βin−β)/(βin−βex)\u2003\u2003formula (1)'}, 'provided that an angle formed by a tangent at a point on the pressure surface or suction surface and an axial direction of the turbofan engine is referred to as a blade ...

GAS TURBINE ENGINE AIRFOIL MISTUNING

A circumferential airfoil array that provides a total airfoil count for the stage and that includes first and second sets of airfoils with different vibrational frequencies. First and second arcuate regions are arranged opposite one another and third and fourth arcuate regions are arranged opposite one another. The first set of airfoils has first primary, first secondary and first tertiary airfoil groups. The second set of airfoils has second primary, second secondary and second tertiary airfoil groups. The first primary airfoil group and second primary airfoil group are respectively arranged in the first and second arcuate regions. The first secondary airfoil group and second secondary airfoil group are arranged in the third arcuate region. The first tertiary airfoil group and the second tertiary airfoil group are arranged in the fourth arcuate region. The first and second arcuate regions provide greater than 50% of the total airfoil count. 1. A stage for a gas turbine engine comprising:a circumferential airfoil array providing a total airfoil count for the stage and that includes first and second sets of airfoils, the airfoil array has first, second, third and fourth arcuate regions, the first and second arcuate regions arranged opposite one another, and the third and fourth arcuate regions arranged opposite one another, the first set of airfoils includes a different vibrational frequency than the second set of airfoils, the first set of airfoils has first primary, first secondary and first tertiary airfoil groups, the second set of airfoils has second primary, second secondary and second tertiary airfoil groups, the first primary airfoil group and second primary airfoil group respectively arranged in the first and second arcuate regions, the first secondary airfoil group and second secondary airfoil group arranged in the third arcuate region, and the first tertiary airfoil group and the second tertiary airfoil group arranged in the fourth arcuate region, the ...

POWER TURBINE VANE AIRFOIL PROFILE

A power turbine includes a first stage vane having an airfoil with a cold nominal profile substantially in accordance with at least an intermediate portion of the Cartesian coordinate values of X, Y and Z set forth in Table 2. The X and Y values are distances, which when smoothly connected by an appropriate continuing curve, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form a complete airfoil shape. 1. A turbine vane of a gas turbine engine having a gaspath , the turbine vane comprising an airfoil having an intermediate portion contained within the gaspath and defined by a nominal un-coated profile substantially in accordance with Cartesian coordinate values of X , Y , and Z of Sections 2 to 9 set forth in Table 2 , wherein the point of origin of the orthogonally related axes X , Y and Z is located at an intersection of a centerline of the gas turbine engine and a stacking line of the turbine vane , the Z values are radial distances measured along the stacking line , the X and Y are coordinate values defining the profile at each distance Z , wherein X and Y values define a set of points for each Z value which when connected by smooth continuing arcs define an airfoil profile section , the profile sections at the Z distances being joined smoothly with one another to form an airfoil shape of the intermediate portion.2. The turbine vane as defined in claim 1 , wherein the turbine vane is a power turbine vane of the gas turbine engine.3. The turbine vane as defined in claim 2 , wherein the power turbine vane is a first stage power turbine vane of a multi-stage power turbine.4. The turbine vane as defined in claim 1 , wherein the turbine vane has a manufacturing tolerance of ±0.009 inches in a direction perpendicular to the airfoil.5. A turbine vane for a gas turbine engine having a gaspath claim 1 , the turbine vane having an intermediate airfoil portion contained within the gaspath and ...

TURBINE BUCKET WITH ENDWALL CONTOUR AND AIRFOIL PROFILE

Turbine frequency tuning, fluid dynamic efficiency, and performance can be improved using an airfoil profile and/or an endwall contour including at least one of a pressure side bump, a pressure side leading edge bump, or a suction side trough. In particular, by including two endwall bumps on the pressure side and a trough on the suction side combined with a particular airfoil profile, performance can be further improved. 1. A turbomachine bucket comprising:a base;an airfoil supported by the base, the airfoil including opposed first and second ends with the first end at the base, the airfoil further including opposed pressure and suction sidewalls extending in chord between opposed leading and trailing edges and extending in span between the first and second ends of the airfoil;an endwall on the base and connected to the first end of the airfoil, the endwall including opposed endwall leading and trailing edges extending substantially circumferentially between opposed pressure and suction splitlines, a distance between the pressure and suction splitlines being substantially equal to a pitch;opposed leading edge and trailing edge regions of the endwall each extending from the respective endwall leading and trailing edges to about half way therebetween;pressure side and suction side regions of the endwall extending from the respective pressure and suction sidewalls of the airfoil;at least one pressure side feature in the endwall in the pressure side region, each respective pressure side feature including at least one of a trough or a bump relative to a nominal surface of the endwall;at least one suction side feature in the endwall in the suction side region, each respective suction side feature including at least one of a trough or a bump relative to the nominal surface of the endwall; andat least one leading edge feature in the endwall including at least one of a trough or a bump relative to the nominal surface in the leading edge region.2. The turbomachine bucket of ...

Tunnel Power Turbine System To Generate Potential Energy From waste Kinetic Energy

A system for generating energy from waste includes a generator device and an ejector device integrated in a pipe line unit. The generator device includes nozzle venture inlets. The ejector device is coupled with the generator device and includes a slit venture outlet to restore any velocity pressure loss in the pipe line unit and eliminate any back pressure buildup in the generator device. 117-. (canceled)18. A system for transforming waste energy into potential energy comprising three devices , one of which is an internal device and two are external , a system's large body consists of; combining chamber , holds an internal tunnel power device , the chamber has two additional inlets that joins two external devices: shrouded funnel device and temperature dispersing device;wherein further, the system have a vital component ejector contraption with venturi within and passage that extends through ejector body;wherein further the said three devices individually and collectively, when joined, compound waste energy with air or wind or more waste energy or any combination thereof, increases composites volume and enhances velocity from system's body exit opening into a tunnel that rotates the impellor which drives generator/compressor thereof;furthermore, the said three devices are:a) a tunnel power device having an external inlet opening passage that extends through ejector contraption body to outlet opening in system's large body combining chamber, located between two external devices inlet openings and large body main exit outlet opening above, passage has fixed defuser plates;b) a wind powered device having a hemispherical cup device configured to drive a fan below in funnel inlet passage and an exit outlet opening at the end of tunnel that connects systems large body combining chamber located above the tunnel power device outlet opening and has a similar ejector contraption and fixed defuser plates; andc) a temperature control dispersing device having a similar ejector ...

TURBINE BUCKET HAVING NON-AXISYMMETRIC BASE CONTOUR

Various embodiments of the invention include turbine buckets and systems employing such buckets. Various particular embodiments include a turbine bucket having: an airfoil having: a suction side; a pressure side opposing the suction side; a leading edge spanning between the pressure side and the suction side; and a trailing edge opposing the leading edge and spanning between the pressure side and the suction side; and a base connected with a first end of the airfoil along the suction side, pressure side, trailing edge and the leading edge, the base including a non-axisymmetric contour proximate a junction between the base and the airfoil. 1. A turbine bucket comprising:an airfoil having: a suction side; a pressure side opposing the suction side; a leading edge spanning between the pressure side and the suction side; and a trailing edge opposing the leading edge and spanning between the pressure side and the suction side; anda base connected with a first end of the airfoil along the suction side, pressure side, trailing edge and the leading edge, the base including a non-axisymmetric contour proximate a junction between the base and the airfoil.2. The turbine bucket of claim 1 , further comprising a fillet connecting a surface of the base to a surface of the airfoil.3. The turbine bucket of claim 1 , wherein the turbine bucket includes at least one of a first stage bucket or a second stage bucket.4. The turbine bucket of claim 1 , wherein the base directs flow of a working fluid across a passage trough proximate the suction side of the airfoil.5. The turbine bucket of claim 1 , wherein the base is radially inboard of the airfoil.6. The turbine bucket of claim 1 , wherein the non-axisymmetric contour includes a thickened area.7. The turbine bucket of claim 6 , wherein the thickened area has an apex at approximately 0 percent to approximately 5 percent of an axial chord length of the endwall claim 6 , and wherein the thickened area has an approximately zero percent to ...

TURBINE ENGINE BLADE INCLUDING STRUCTURAL REINFORCEMENT ADHESIVELY BONDED USING AN ADHESIVE BOND OF INCREASED TOUGHNESS

A turbine machine blade has an aerodynamic surface that is made of organic matrix composite material reinforced by fibers and metal structural reinforcement that is adhesively bonded by an epoxy adhesive bond on the leading edge, which is of matching shape, and that presents along its entire height a section that is substantially V-shaped with a base extended by two lateral flanks of respective profiles that become thinner at free ends going towards the trailing edge. In order to increase the toughness of the epoxy adhesive bond in the event of the epoxy adhesive bond cracking, the epoxy adhesive bond includes a reinforcing sheet of elastomeric polymer enabling the reinforcing sheet to be torn into two portions, the elastomeric polymer having the following properties at 23° C.: Young's modulus E≈10 MPa; stress at rupture σ>10 MPa; strain at rupture ε>80%. 1. A turbine machine blade having an aerodynamic surface extending along a first direction between a leading edge and a trailing edge , and along a second direction that is substantially perpendicular to said first direction between a blade root and a blade tip , said aerodynamic surface being made of organic matrix composite material reinforced by fibers , the blade also including metal structural reinforcement that is adhesively bonded by an epoxy adhesive bond on said leading edge , which is of matching shape , and that presents along its entire height a section that is substantially V-shaped with a base extended by two lateral flanks of respective profiles that become thinner at free ends going towards said trailing edge , wherein , in order to increase the toughness of the epoxy adhesive bond in the event of the epoxy adhesive bond cracking , said epoxy adhesive bond includes a reinforcing sheet of elastomeric polymer enabling the reinforcing sheet to be torn into two portions , said elastomeric polymer having the following properties at 23° C.: Young's modulus E≈10 MPa; stress at rupture σ>10 MPa; strain at ...

IMPELLER AND AXIAL FLOW FAN

An impeller includes: a boss portion driven to rotate by a motor; and a plurality of rotating blades projecting radially from the boss portion in a direction in which a diameter increases from a rotational axis of the motor and generating airflow in an axial direction of the rotational axis. The rotating blades have an S-shaped radial cross section in which an inner peripheral side portion is protruded with respect to the airflow and an outer peripheral side portion is recessed with respect to the airflow, and a recess-shaped portion of the rotating blades has a distribution of a radius of curvature value such that the radius of curvature value gradually decreases toward a blade trailing edge portion from a blade leading edge portion and a rate of the gradual reduction becomes smaller toward the blade trailing edge portion. 1. An impeller comprising:a boss portion driven to rotate by a motor; anda plurality of rotating blades projecting radially from the boss portion in a direction in which a diameter increases from a rotational axis of the motor and generating airflow in an axial direction of the rotational axis, whereinthe rotating blades each have an S-shaped radial cross section in which an inner peripheral side portion is protruded with respect to the airflow and an outer peripheral side portion is recessed with respect to the airflow, anda recess-shaped portion of the rotating blades has a distribution of a radius of curvature value such that the radius of curvature value gradually decreases toward a blade trailing edge portion from a blade leading edge portion.2. The impeller according to claim 1 , whereinthe rotating blades are inclined toward an upstream side of the airflow in the blade leading edge portion with an angle of inclination becoming smaller toward the blade trailing edge portion and are inclined toward a downstream side of the airflow in the blade trailing edge portion.3. An axial flow fan comprising:{'claim-ref': {'@idref': 'CLM-00001', 'claim ...

TURBINE BLADE AIRFOIL AND TIP SHROUD

A turbine blade airfoil (R, M, T, ) comprising an outer surface shape defined by Cartesian coordinates of successive transverse profiles at radial increments as set forth in Tables 1a to 1k, wherein each table defines a transverse sectional profile characterized by a smooth curve connecting the coordinates, and the surface shape comprises a smooth surface connecting the sectional profiles. The blade may include a tip shroud with edge profiles defined by Cartesian coordinates set forth in Table 2a and 2b. A gusset/fillet may be provided between the blade airfoil and the tip shroud, with a planar diagonal surface over most of a diagonal bracing area of the gusset/fillet. 1. A turbine blade airfoil comprising an outer surface shape defined by Cartesian coordinate values of X and Y at successive radial increments as set forth in Tables 1a to 1k , wherein each said table defines a transverse sectional profile characterized by a smooth curve connecting the X and Y coordinates , and the surface shape comprises a smooth surface connecting the sectional profiles.2. The turbine blade airfoil of claim 1 , wherein the Cartesian coordinate values are absolute values in inches claim 1 , and a manufacturing tolerance of the smooth surface is +/−0.050 inches measured normal to said surface.3. The turbine blade airfoil of claim 1 , wherein the Cartesian coordinate values are relative values claim 1 , and a manufacturing tolerance of the smooth surface is a relative value of +/−0.050 inches measured normal to said surface.4. The turbine blade airfoil of claim 1 , further comprising a tip shroud comprising an axially forward edge profile defined by Cartesian coordinate values of X and Y set forth in Table 2a claim 1 , and an axially aft edge profile defined by Cartesian coordinate values of X and Y set forth in table 2b.5. The turbine blade airfoil of claim 4 , wherein the Cartesian coordinate values are absolute values in inches claim 4 , and a manufacturing tolerance of the axially ...

TURBINE ASSEMBLY

The invention relates to a turbine in which a bypass-passage extends through a base member of a stationary vane to join seal cavities of adjacent rotating blade rows so that seal flow passing between a casing and shrouds of the rotating blades at least partially bypasses the turbine main flow passage. 1. A turbine assembly comprising:a rotor with a rotational axis; a first rotating blade row having a plurality of circumferentially adjacent first rotating blades each with a first shroud, at a first rotating blade end that is distal from the rotor and blade airfoils extending into the flow passage;', 'a first sealing means located between the casing and the first shroud;', 'a stationary vane row axially adjacent and downstream of the first rotating blade row, having a plurality of circumferentially adjacent stationary vanes each with a base member and a vane airfoil extending into the flow passage;', 'a second rotating blade row in the flow passage axially adjacent and downstream of the stationary vane row having a plurality of circumferentially adjacent second rotating blades each with a second shroud, at a second rotating blade end that is distal from the rotor; and', 'a second sealing means located between the casing and the second shroud,, 'a casing enclosing the rotor to form a flow passage therebetween;'}wherein a first cavity is formed by the first shroud, the sealing means and the base member and a second cavity is formed by the second shroud, the base member and the second sealing means, anda bypass-passage extending from a first end at the first cavity to a second end at the second cavity so as to bypass the flow passage and the vane airfoil.2. The turbine assembly of claim 1 , wherein each of the vane airfoils has a leading edge wherein the first end of the bypass-passage is located at a point of the first cavity circumferentially between the leading edges of two of the circumferentially adjacent vanes airfoils.3. The turbine assembly of claim 1 , wherein ...

ROTOR OF ROTARY MACHINE AND ROTARY MACHINE

A rotor () of a rotary machine (T) according to the invention includes a plurality of rotor members (, and ) which are joined to each other in the axial direction in which the axis (P) extends, and among the plurality of rotor members ( and ), the first rotor member () in hydraulic fluid injection portions (and ) of a passageway () is formed of Ni-based alloy so that the inside thereof is hollow throughout the entire length in the axial direction. 113-. (canceled)14. A rotor of the steam turbine that is formed to have rotor members which are formed of different materials from each other and which are joined to each other in the axial direction , the rotor comprising:a first rotor member; anda second rotor member,wherein the first rotor member is formed of an Ni-based alloy,wherein the first rotor member includes a guide portion through which a working fluid is guided,wherein the first rotor member includes a rotor vane through which the working fluid passes,wherein the second rotor member is formed of a material which is more easily molded than the Ni-based alloy,wherein the second rotor member is joined to the first rotor member on the downstream side of flow of the working fluid,wherein the second rotor member includes a discharge portion through which the working fluid is discharged,wherein the second rotor member includes a rotor vane, through which the working fluid passes, on a joint side with the first rotor member and includes a sealing portion from the outside and a bearing on a side opposite to the joint side,wherein the first rotor member and a joint portion of the second rotor member with the first rotor member are disposed in the same pressure range,wherein the number of vane rows of the first rotor member is greater than the number of vane rows of the second rotor member, and{'sup': −6', '−6', '−6', '−6, 'wherein an average linear expansion coefficient of the first rotor member in a temperature range of room temperature to 700° C. which becomes an ...

HVATA-HYBRID VERTICAL AXIS TURBINE ASSEMBLY OPERABLE UNDER OMNI-DIRECTIONAL FLOW FOR POWER GENERATING SYSTEMS

A hybrid, vertical axis helical turbine assembly capable of providing unidirectional rotation under an omni-directional low speed obverse fluid flow (gas or liquid), respectively gas flow is disclosed. The assembly comprises a minimum (but not limited to) of three hybrid (airfoil enhanced helical vane profile) wings, which are substantially spaced from the vertical axis (Z) and circumferentially spaced from one another. Each hybrid wing is fixed to the center hub in a rigid position by two or more arms, which are symmetrically located from each other in conjunction with the hub's horizontal axis (X). The hybrid, vertical axis helical turbine assembly provides high torque at very low wind speed because of the absolute symmetric airfoil enhanced helical vane profile wing, which design maintain adaptive lift to drag ratio over the one rotational revolution time line in coincidence of the upwind direction. These characteristics make the hybrid, vertical axis helical turbine assembly suitable for urban off grid and grid tie applications in low wind speed areas and areas of reputable wind turbulence. 1. A Hybrid , Vertical Axis helical Turbine Assembly (HVATA) capable of providing unidirectional rotation under an omni-directional low speed obverse fluid flow (gas or liquid) , respectively gas flow is disclosed.2. A HVATA according to claim 1 , further comprising high torque at very low wind speed because of the absolute symmetric airfoil enhanced helical vane profile wing.3. A hybrid vertical axis wind turbine assembly according to claim 2 , wherein the turbine wings are substantially spaced from the vertical axis (Z) and circumferentially spaced from one another on 60 degree for the entire height of the HVATA.4. A wind turbine according to claim 3 , wherein the turbine wings are positioned in absolute symmetric airfoil enhanced helical vane profile claim 3 , which design maintain adaptive lift to drag ratio over the one rotational revolution time line in coincidence of ...

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 side 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 ratio of leading edge radius to maximum thickness is between 0.40 (r/T) and 0.45 (r/T) and a ratio of leading edge radius to axial chord length is between 0.17 (r/b) and 0.27 (r/b). 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 side 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', {'sub': max', 'max', 'x', 'x, 'a ratio of leading edge radius to maximum thickness of between 0.40 (r/T) and 0.45 (r/T) and a ratio of leading edge radius to axial chord length of between 0.17 (r/b) and 0.27 (r/b).'}], 'an airfoil extending from the radially outer side of the platform, the airfoil including2. The component of claim 1 , wherein the ratio of the leading edge radius to maximum chord thickness at 0% span is between approximately 0.43-0.47 (r/T).3. The component of claim 2 , wherein the ratio the leading edge radius to maximum chord thickness at 25% span is between approximately 0.41-0.45 (r/T).4. The component of claim 3 , wherein the ratio of the leading edge radius to maximum chord thickness at 50% span is between approximately 0.41-0.45 (r/T).5. The component of claim 4 , wherein the ratio of the leading edge radius to maximum chord thickness at 75% span is between approximately 0.42-0.47 (r/T).6. The component of claim 5 , wherein the ratio of the leading edge radius to maximum chord thickness at 100% span is between ...

TURBINE BLADE

A turbine blade can shorten the assembly time of a multistage turbine section of a gas turbine and reduce assembly errors such as misalignment between adjacent turbine blades, while minimizing the occurrence of a secondary vortex. The turbine blade includes a main end wall; a first turbine airfoil that is connected on a hub side to the main end wall and has a leading edge and a trailing edge; and a second turbine airfoil that is connected on a hub side to the main end wall and has a leading edge and a trailing edge, wherein the main wall is integrally formed with each of the first and second turbine airfoils. The second turbine airfoil is disposed on the main end wall so as to face the first turbine airfoil, and the main end wall is disposed between the first turbine airfoil and the second turbine airfoil. 1. A turbine blade comprising:a main end wall;a first turbine airfoil that is connected on a hub side to the main end wall and has a leading edge and a trailing edge; anda second turbine airfoil that is connected on a hub side to the main end wall and has a leading edge and a trailing edge,wherein the main wall is integrally formed with each of the first and second turbine airfoils.2. The turbine blade according to claim 1 , wherein the second turbine airfoil is disposed on the main end wall so as to face the first turbine airfoil.3. The turbine blade according to claim 1 , wherein the main end wall is disposed between the first turbine airfoil and the second turbine airfoil.4. The turbine blade according to claim 3 , wherein the main end wall disposed between the first turbine airfoil and the second turbine airfoil include an inclined portion that is inclined in a pitch-wise direction that is a direction toward the second turbine airfoil from the first turbine airfoil.5. The turbine blade according to claim 3 , wherein the main end wall disposed between the first turbine airfoil and the second turbine airfoil includes a curved portion having a streamlined profile ...

TURBINE WHEEL

A turbine wheel, in particular in a charging device for use in an internal combustion engine, is specified, wherein the turbine wheel () comprises a plurality of blades () on a hub () that forms a rear wall (), wherein adjacent blades () form an inlet surface () having two leading edges () and an outlet surface () having two trailing edges () and situated substantially axially inward, wherein a surface () of a blade () is configurable by way of an angle (T) and a length (Z) of a plurality of curvatures () situated next to one another between the leading edge () and the trailing edge (), wherein, for each of the curvatures (), the angle (T) of the leading edge () initially increases or remains constant and then decreases as the length (Z) increases so as to form a maximum () 110121614121820222426120303820243038200404040200. A turbine wheel , in particular in a charging device for use in an internal combustion engine , wherein the turbine wheel () comprises a plurality of blades () on a hub () that forms a rear wall () , wherein adjacent blades () form an inlet surface () having two leading edges () and an outlet surface () having two trailing edges () and situated substantially axially inward , wherein a surface () of a blade () is configurable by way of an angle (T) and a length (Z) of a plurality of curvatures ( , . . . , ) situated next to one another between the leading edge () and the trailing edge () , wherein , for each of the curvatures ( , . . . , ) , the angle (T) of the leading edge () initially increases or remains constant and then decreases as the length (Z) increases so as to form a maximum ( , ′ , ″) , the angle (T) being formed as a polar angle and the length being formed along an axis of rotation (Z) around a direction of rotation (R) , a zero point of the angle (T) being chosen to increase along the leading edge () towards the direction of rotation (R) , and the length (Z) being normalized along the axis of rotation (Z).2303812162812. The turbine ...

ROTOR BLADE OF A WIND TURBINE

A rotor blade of a wind turbine with a raked tip portion is provided. The tip portion is raked, i.e. curved, in a rotor blade plane comprising the tip base chord and a line which is parallel to the pitch axis of the rotor blade. Additionally, the orientation of the chords with reference to the chord at the tip base changes between the tip base and the tip. In other words, the trailing edge of the tip portion is twisted. Optionally, the tip portion is additionally swept out of the rotor blade plane, which is characterized by a cant angle. 1. A rotor blade of a wind turbine , comprisinga blade body extending between a root portion of the rotor blade and a tip portion of the rotor blade, a trailing edge and a leading edge,wherein the rotor blade is arranged and prepared for being mounted to a hub of the wind turbine and for being pitched about a pitch axis, andwherein a rotor blade plane is defined by the plane comprising the chord at the tip base and a line which is parallel to the pitch axis, wherein the tip base is the part of the rotor blade at which the tip portion joins the blade body,wherein the trailing edge of the tip portion has a curved shape as projected on the rotor blade plane such that a trailing edge sweep angle increases from the tip base to the tip of the rotor blade, andwherein orientation of the chords with reference to the chord at the tip base changes between the tip base and the tip such that a chord tilt angle changes between the tip base and the tip.2. The rotor blade according to claim 1 , wherein the trailing edge sweep angle increases by at least 2 degrees from the tip base to the tip.3. The rotor blade according to claim 1 , wherein the trailing edge sweep angle increases by at most 40 degrees from the tip base to the tip.4. The rotor blade according to claim 1 , wherein the chord tilt angle varies by at least 2 degrees between the tip base and the tip.5. The rotor blade according to claim 1 , wherein the chord tilt angle varies by at most ...

Sealing Device and Turbo Machine

The turbo machine includes a rotor with a rotating shaft , a stator enclosing the rotor and a sealing device installed in a clearance passage defined between the rotor and the stator , the sealing device controlling a leakage flow from the clearance passage B. The sealing device includes a plurality of sealing fins disposed on at least one of the rotor and the stator , and arranged in an axial direction of the rotor . The sealing device further includes at least one deceleration controlling member provided on the rotational side. The deceleration controlling member projects toward a chamber defined between the sealing fins and is configured to control a reduction in the velocity of the leakage flow B in the chamber in the rotational direction of the rotor 1. A turbo machine , comprising:a rotor with a rotating shaft;a stator enclosing the rotor; anda sealing device installed in a clearance passage defined between the rotor and the stator, the sealing device controlling a leakage flow from the clearance passage;wherein the sealing device includesa plurality of sealing fins disposed on at least one of the rotor and the stator, the plurality of sealing fins being arranged in an axial direction of the rotor, andat least one deceleration controlling member provided on a rotational side, the at least one deceleration controlling member projecting toward a chamber defined between the sealing fins, the at least one deceleration controlling member being configured to control a reduction in the velocity of the leakage flow in the chamber in a rotational direction of the rotor.2. The turbo machine according to claim 1 ,wherein each of the plurality of sealing fins is a fin that projects from the rotor, andthe at least one deceleration controlling member is a rib that projects from one of the plurality of sealing fins toward a downstream of a leakage flow in an axial direction of the rotor.3. The turbo machine according to claim 2 ,wherein, among the plurality of sealing fins, ...

VANE UNIT AND STEAM TURBINE

Provided are a vane unit and a steam turbine. A vane unit in which an outer ring and an inner ring are connected by a plurality of vanes arranged at predetermined intervals in a circumferential direction is provided with: a steam outer ring inlet portion provided in a first heating chamber of the outer ring; a steam outer ring outlet portion provided in the first heating chamber of the outer ring so as to be separated from the steam outer ring inlet portion in the circumferential direction; and a first steam passage that makes the steam outer ring inlet portion and the steam outer ring outlet portion communicate with each other in the first heating chamber of the outer ring. Thus, steam is efficiently used, and erosion due to wet steam is suppressed and a decrease in thermal efficiency is suppressed. 1. A vane unit in which an outer ring and an inner ring are connected by a plurality of vanes arranged at predetermined intervals in a circumferential direction , the vane unit comprising:a steam outer ring inlet portion provided in a cavity portion of the outer ring;a steam outer ring outlet portion provided in the cavity portion of the outer ring so as to be separated from the steam outer ring inlet portion in the circumferential direction; anda first steam passage making the steam outer ring inlet portion and the steam outer ring outlet portion communicate with each other in the cavity portion of the outer ring, whereinthe first steam passage is arranged along an inner circumferential side in the cavity portion of the outer ring.2. (canceled)3. The vane unit according to claim 1 , whereinthe first steam passage is formed by a tube.4. The vane unit according to claim 1 , whereinthe steam outer ring inlet portion includes an outer ring inlet header formed by partitioning a part of the cavity portion by a pair of inlet partition plates, and a steam supply port provided in the outer ring and communicating with the outer ring inlet header, andthe steam outer ring outlet ...

AIRFOIL WITH ENERGY ABSORBING EDGE GUARD

An edge guard apparatus for an airfoil includes: a body having a nose with spaced-apart first and second wings extending therefrom, the body defining a cavity between the first and second wings; and an energy absorbing structure disposed in the cavity. 1. An edge guard apparatus for an airfoil , comprising:a body having a nose with spaced-apart first and second wings extending therefrom, the body defining a cavity between the first and second wings; andan energy absorbing structure disposed in the cavity.2. The apparatus of wherein the energy absorbing structure comprises a cellular structure.3. The apparatus of wherein the energy absorbing structure comprises a plurality of tubes arranged in a side-by-side configuration.4. The apparatus of wherein the tubes extend in a spanwise direction of the body.5. The apparatus of wherein at least one of the tubes extends over only a portion of a span of the body.6. The apparatus of wherein a greater number of tubes are disposed near a first end of the body and a fewer number of tubes are disposed near a second end of the body.7. The apparatus of where at least one of the tubes is bonded to another one of the tubes.8. The apparatus of wherein the tubes are separated by voids claim 3 , the voids being filled with an adhesive.9. The apparatus of wherein the tubes comprise a metal alloy.10. The apparatus of wherein the body comprises a sheet stock material.11. An airfoil apparatus claim 1 , comprising:an airfoil having convex and concave sides extending between a leading edge and a trailing edge;a body having a nose with spaced-apart first and second wings extending therefrom, the body defining, in cooperation with the leading edge of the airfoil, a cavity; andan energy absorbing structure disposed in the cavity.12. The apparatus of wherein the energy absorbing structure comprises a cellular structure.13. The apparatus of wherein the energy absorbing structure comprises a plurality of tubes arranged in a side-by-side ...

GAS TURBINE ENGINE AIRFOIL

An airfoil for a turbine engine includes an airfoil that has pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a leading edge dihedral and a span position. The leading edge dihedral is negative from the 0% span position to the 100% span position. A positive dihedral corresponds to suction side-leaning, and a negative dihedral corresponds to pressure side-leaning. The airfoil has a relationship between a trailing edge dihedral and a span position. The trailing edge dihedral is positive from the 0% span position to the 100% span position. A positive dihedral corresponds to suction side-leaning and a negative dihedral corresponds to pressure side-leaning. The airfoil includes at least one of a least negative dihedral in a range of 5-15% span position, a least negative dihedral in a range of 0-10% span position and a least positive trailing edge dihedral in a 40%-55% span position. 1. An airfoil for a turbine engine comprising:an airfoil having pressure and suction sides extending in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip, wherein the airfoil has a relationship between a leading edge dihedral and a span position, the leading edge dihedral negative from the 0% span position to the 100% span position, wherein a positive dihedral corresponds to suction side-leaning, and a negative dihedral corresponds to pressure side-leaning;wherein the airfoil has a relationship between a trailing edge dihedral and a span position, the trailing edge dihedral positive from the 0% span position to the 100% span position, wherein a positive dihedral corresponds to suction side-leaning, and a negative dihedral corresponds to pressure side-leaning; andwherein the airfoil includes one of:a least negative dihedral is in a range of 5-15% span position;a least negative ...

NOISE REDUCING EXTENSION PLATE FOR ROTOR BLADE IN WIND TURBINE

Rotor blade assemblies and methods for constructing rotor blade assemblies are provided. A rotor blade assembly may include a rotor blade having exterior surfaces defining a pressure side, a suction side, a leading edge and a trailing edge each extending between a tip and a root, the rotor blade defining a span and a chord. The rotor blade assembly further includes an extension plate mounted to one of the pressure side or the suction side, the extension plate extending in the chord-wise direction between a first end and a second end, the second end extending beyond the trailing edge. The rotor blade assembly further includes a filler substrate provided on an inner surface of the extension plate and the trailing edge, the filler substrate tapering from the trailing edge towards the second end. 1. A rotor blade assembly , comprising:a rotor blade having exterior surfaces defining a pressure side, a suction side, a leading edge and a trailing edge each extending between a tip and a root, the rotor blade defining a span and a chord;an extension plate mounted to one of the pressure side or the suction side, the extension plate extending in the chord-wise direction between a first end and a second end, the second end extending beyond the trailing edge; and,a filler substrate provided on an inner surface of the extension plate and the trailing edge, the filler substrate tapering from the trailing edge towards the second end.2. The rotor blade assembly of claim 1 , wherein the extension plate has a thickness of less than 2 millimeters.3. The rotor blade assembly of claim 1 , wherein the extension plate has a thickness of approximately 1 millimeter.4. The rotor blade assembly of claim 1 , wherein an extension portion of the extension plate which extends beyond the trailing edge has a width in the chord-wise direction of between approximately 5 times and approximately 20 times a thickness of the trailing edge.5. The rotor blade assembly of claim 1 , wherein an extension ...

ENGINE AIRFOILS AND METHODS FOR REDUCING AIRFOIL FLUTTER

A method is provided for designing an airfoil. The method includes considering a baseline airfoil having a first camber distribution and a first aerodynamic efficiency; reducing the first camber distribution to result in a reduced camber airfoil with a second camber distribution and a second aerodynamic efficiency such that the second aerodynamic efficiency is approximately equal to the first aerodynamic efficiency; and producing the airfoil with the second camber distribution 1. A method for designing an airfoil , the method comprising the steps of:considering a baseline airfoil having a first camber distribution and a first aerodynamic efficiency;reducing the first camber distribution to result in a reduced camber airfoil with a second camber distribution and a second aerodynamic efficiency, wherein the second aerodynamic efficiency is approximately equal to the first aerodynamic efficiency; andproducing the airfoil with the second camber distribution.2. The method of claim 1 , wherein the baseline airfoil has a first chord length claim 1 , and wherein the producing step further includes producing the airfoil with the first chord length.3. The method of claim 2 , wherein the baseline airfoil has a first thickness distribution claim 2 , and wherein the producing step further includes producing the airfoil with the first thickness distribution.4. The method of claim 3 , wherein the baseline airfoil has a first leading edge angle and a first trailing edge angle claim 3 , and wherein the producing step further includes producing the airfoil with the first leading edge angle and the first trailing edge angle.5. The method of claim 1 , wherein the baseline airfoil has a leading edge and a trailing edge claim 1 , and wherein the reducing step includes reducing the camber distribution while maintaining the leading edge and the trailing edge.6. The method of claim 1 , wherein the baseline airfoil has a first twist to flex ratio claim 1 , and wherein the reducing step ...

Rotor With Blades Secured by Woven Fiber Attachment System

A propulsive engine rotor configured for operation within a gas turbine includes a plurality of spaced airfoil blades circumferentially affixed to an outer hoop-style rim. At least two of the airfoil blades have an integral root configured to extend either through or at least radially inwardly from the rim. A woven fiber system provides that the blade roots are interdigitally wrapped by a plurality of woven fibers before impregnation and/or encapsulation of the fibers and wrapped roots within a ceramic or other composite matrix material to form a composite ring. The composite ring defines the interior body of the rotor, includes a bore through which passes the rotational axis of the rotor, and has a lower mass than the rotor rim. The woven fiber interface between the roots increases the tensile load capacity of the airfoil blades relative to the composite ring, and increases the self-sustaining radius of the rotor. 1. A propulsive engine rotor comprising:a rim with an annular composite ring and a plurality of circumferentially offset airfoil blades extending outwardly therefrom and integrally formed blade roots extending radially inward therefrom, wherein:a plurality of fiber filaments are interdigitally woven about and secured to the roots, the fiber filaments being encapsulated within the composite ring for securing the blades thereto.2. The engine rotor of claim 1 , wherein the rotors claim 1 , roots claim 1 , and rim are formed of a heat durable metal alloy.3. The engine rotor of claim 1 , wherein the composite ring is a ceramic material.4. The engine rotor of claim 1 , wherein at least one root has a bulb-shaped body claim 1 , and has a neck portion of a smaller circumferential dimension than the root body.5. The engine rotor of claim 1 , wherein a portion of the woven fibers is interdigitally secured directly to the neck of the at least one root.6. The engine rotor of claim 5 , wherein a portion of the woven fibers forms complete circular rings of fiber ...

Lifting Foil

A lifting foil having a configuration with a leading course and trailing course which is rotated about an axis of rotation into a fluid. 1. A lifting foil configuration which is rotated about an axis of rotation into a fluid , comprising:a rotating shaft connected to a drive mechanism;a plurality of lifting foils equally spaced and connected to said rotating shaft, wherein said drive mechanism being disposed adjacent a first end of said rotating shaft relative to each said foil being disposed adjacent a second end of said rotating shaft, and with respect to said rotating shaft each said foil includes a first inwardly disposed configuration having a first inwardly disposed trailing course with a positive camber having a first inwardly disposed port end connected to said rotating shaft and a first inwardly disposed starboard margin, said first inwardly disposed trailing course extending sideward between said first inwardly disposed port end and said first inwardly disposed starboard margin and being responsive to fluid flow over said first inwardly disposed trailing course for generating a first fluid reaction force having a first lifting component, a first outwardly disposed configuration having a first outwardly disposed leading course with a positive camber having an outwardly disposed port end connected to said rotating shaft and an outwardly disposed starboard margin, said first outwardly disposed leading course being positioned outward of said first inwardly trailing course and extending sideward between said outwardly disposed port end and said outwardly disposed starboard margin and being responsive to fluid flow over said first outwardly disposed leading course for generating a second fluid reaction force having a second lifting component parallel and additive to said first fluid reaction force, a starboard flow guide extending vertically between said inwardly disposed starboard margin and said outwardly disposed starboard margin, having a progressively ...

TURBINE ENGINE FLOW PATH

A turbine engine casing flow-path segment that is locally diffusing, followed by a flow-path segment contracting in the vicinity of a fan blade. This contraction accelerates the fluid flow axially forward of the fan blade leading edge at the tip and converges with the linear flow-path aft of the fan blade leading edge but forward of the fan blade trailing edge. More diffused fluid flow results in increased flow capacity of the fan, and increased fan efficiency. 1. In a fluid propulsion system comprising:a shroud defining the outer boundary of a flow path, the flow path extending continuously downstream from an inlet of the shroud to an outlet;a plurality of blades oriented radially about an axis within the flow path, wherein the plurality of blades, the shroud and the flow path are coaxial with the axis,each of the blades having a leading edge, a trailing edge, and a blade tip, the blade tip extending from the leading edge downstream to the trailing edge; the constant portion having a constant radial displacement from the axis as the constant portion extends downstream;', 'the expanded portion having an continuously increasing radial displacement from the axis as the expanded portion extends downstream; and,', 'the contracted portion having a continuously decreasing radial displacement from the axis as the contracted portion extends downstream along the axis;, 'the shroud having a constant portion followed downstream and connected to an expanded portion followed downstream and connected to a contracted portion;'}the improvement wherein the expanded portion is proximate and upstream of each of the blades and the contracted portion terminates axially between the leading tip point and trailing tip point of the blade tip.2. The system of wherein the outer boundary of the flow path at the expanding portion is inclined θ with respect to the axis and the outer boundary of the flow path at the contracted portion is declined α with respect to the axis claim 1 , wherein α is ...

WIND TURBINE BLADE HAVING TWISTED SPAR WEB

A radially twisted wind turbine blade (), including a radially twisted spar web (), wherein a center () of a radial cross section of a base () of the blade rotates within a plane of rotation (), and wherein at a base (), a radial cross section of the spar web forms a first angle () with the plane of rotation, and at a tip () a radial cross section of the spar web forms a second angle () with the plane of rotation different than the first angle. 1. A radially twisted wind turbine blade , comprising a radially twisted spar web , wherein a center of a radial cross section of a base of the blade rotates within a plane of rotation , and wherein at a base , a radial cross section of the spar web forms a first angle with the plane of rotation , and at a tip a radial cross section of the spar web forms a second angle with the plane of rotation different than the first angle.2. The wind turbine blade of claim 1 , wherein the first angle is less than the second angle.3. The wind turbine blade of claim 2 , wherein the first angle is less than 45 degrees claim 2 , and the second angle is greater than 45 degrees.4. The wind turbine blade of claim 1 , wherein a twist of the spar web is curvilinear from a base of the spar web to a tip of the spar web.5. The wind turbine blade of claim 1 , wherein from a base of the spar web to a tip of the spar web the spar web is located between a pressure side spar cap and a suction side spar cap to maximize an average length of the spar web between the pressure side spar cap and the suction side spar cap.6. The wind turbine blade of claim 1 , wherein a tangent of a leading edge of the blade forms right angles with a tangent line of a pressure side of the blade and a tangent line of a suction side of the blade claim 1 , and the spar web spans between a tangent point of the pressure side tangent line and a tangent point of the suction side tangent line.7. The wind turbine blade of claim 1 , further comprising spar caps secured to ends of the spar ...

GUIDE VANE ASSEMBLY FOR A ROTARY MACHINE AND METHODS OF ASSEMBLING THE SAME

A blade includes an airfoil, a stationary portion coupled to a radially inner end of the airfoil, and a leakage flow guide vane assembly coupled to the stationary portion. The leakage flow guide vane assembly includes a plurality of passages defined therein. The passages are oriented to induce a swirl velocity to a working fluid flowing through the passages. 1. A blade comprising:an airfoil;a stationary portion coupled to a radially inner end of said airfoil; anda leakage flow guide vane assembly coupled to said stationary portion, said leakage flow guide vane assembly comprising a plurality of passages defined therein, said plurality of passages oriented to induce a swirl velocity to a working fluid flowing through said passages.2. A blade in accordance with claim 1 , wherein said leakage flow guide vane assembly comprises a plurality of guide vanes extending axially from a first end to an opposite claim 1 , free second end claim 1 , wherein said first end is coupled to a downstream end of said stationary portion.3. A blade in accordance with claim 2 , wherein each said guide vane comprises a first portion that extends radially outward from a bottom surface of said stationary portion claim 2 , and a second portion that extends circumferentially with respect to said first portion.4. A blade in accordance with claim 3 , wherein said second portion of at least one of said each guide vane extends circumferentially at a predetermined angle with respect to said first portion of at least one of said guide vanes and overlaps said first portion of an adjacent guide vane in a radial direction.5. A blade in accordance with claim 2 , wherein each said guide vane is coupled to said stationary portion by at least one of a welding process claim 2 , a brazing process claim 2 , and a bonding process.6. A blade in accordance with claim 2 , wherein each said guide vane is integrally formed with said fixed blade using at least one of an additive manufacturing process and a machining ...

TURBINE ENGINE ASSEMBLY AND METHOD OF MANUFACTURING

A turbine engine assembly is provided. The assembly includes a low-pressure turbine assembly including a first turbine section configured to rotate in a first rotational direction at a first rotational speed, and a second turbine section configured to rotate in a second rotational direction at a second rotational speed. The second rotational direction is opposite the first rotational direction and the second rotational speed is lower than the first rotational speed. The assembly also includes a first drive shaft coupled to the first turbine section, and a fan assembly including a first fan section coupled to the first drive shaft such that the first fan section rotates in the first rotational direction at the first rotational speed, and a second fan section coupled to the second turbine section such that the second fan section rotates in the second rotational direction at the second rotational speed. 1. A turbine engine assembly comprising: a first turbine section configured to rotate in a first rotational direction at a first rotational speed; and', 'a second turbine section configured to rotate in a second rotational direction at a second rotational speed, the second rotational direction opposite the first rotational direction and the second rotational speed lower than the first rotational speed;, 'a low-pressure turbine assembly comprisinga first drive shaft coupled to said first turbine section; and a first fan section coupled to said first drive shaft such that said first fan section is configured to rotate in the first rotational direction at the first rotational speed; and', 'a second fan section coupled to said second turbine section such that said second fan section is configured to rotate in the second rotational direction at the second rotational speed., 'a fan assembly comprising2. The turbine engine assembly in accordance with further comprising a second drive shaft coupled to said second turbine section claim 1 , said second fan section coupled to said ...

END SUPPORTED HELICAL TURBINE

The claimed invention relates to a turbine adapted to extract energy from the velocity of a streaming fluid such as wind, steam, tidal streams and water waves. The invented turbine is arranged with its axis of turbine rotation directed at substantially right angles to the current direction of the streaming fluid and comprising a kind of self-supported blade body which is rotationally symmetric and constructed by rotor blades integrated transversely and supported two by two, allowing the fluid to flow through the turbine with less turbulence compared to other types of turbines equipped with separate rotor blades. 1. A turbine adapted to production of useful energy from the motion of a streaming fluid and arranged at substantially right angles of an axis of turbine rotation to the current direction of the streaming fluid , comprisingat least one turbine roller bearing comprising a rotatable bearing housing and a non-rotatable bearing housing, and exhibiting a centre point and a centre line passing through said centre point; andat least one hub arranged in fixed attachment to the rotatable bearing housing and to a supporting structure arranged in fixed attachment to the non-rotatable bearing housing; anda blade body located in whole or in part in the fluid and arranged in attachment to the hub, wherein the motion of the streaming fluid makes the blade body feasible to rotate around the axis of turbine rotation which coincides with the centre line passing through a point identical to said centre point, and comprisinga plurality of rotor blades each of which extends continuously in the axial and radial direction of a helix exhibiting an axis coinciding with the axis of turbine rotation, and exhibits a handedness around the axis of turbine rotation and in a plane normal to the helix axis is provided with a cross section exhibiting a centre line normal to said cross section and a wing profile which is provided with two end sections, wherein a first end section is provided ...