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

FIELD: engines and pumps. SUBSTANCE: straightening unit includes two coaxial shells between which the blades passing mainly in radial direction are arranged. Diffuser includes two coaxial walls representing rotation bodies and connected to each other by means of radial partitions. One of the shells of straightening unit is made in the form of single part with one wall of diffuser, which represents rotation body. The other shell of straightening unit is connected to and fixed on the other diffuser wall representing rotation body. Blades of strengthening unit are rigidly connected through one end to one shell of straightening unit and are located at some distance with a small gap from the other shell on the other end. EFFECT: simpler manufacturing technology of the straightening unit. 15 cl, 8 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) 2 435 104 (13) C2 (51) МПК F23R 3/04 (2006.01) F01D 9/02 (2006.01) F04D 29/54 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ, ПАТЕНТАМ И ТОВАРНЫМ ЗНАКАМ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (21)(22) Заявка: 2007119785/06, 28.05.2007 (24) Дата начала отсчета срока действия патента: 28.05.2007 (73) Патентообладатель(и): СНЕКМА (FR) R U Приоритет(ы): (30) Конвенционный приоритет: 29.05.2006 FR 0604745 (72) Автор(ы): ДАГЕНЕ Люк Анри Клод (FR) (43) Дата публикации заявки: 10.12.2008 Бюл. № 34 2 4 3 5 1 0 4 (45) Опубликовано: 27.11.2011 Бюл. № 33 2 4 3 5 1 0 4 R U Адрес для переписки: 129090, Москва, ул. Б.Спасская, 25, стр.3, ООО "Юридическая фирма Городисский и Партнеры", пат.пов. А.В.Мицу, рег.№ 364 C 2 C 2 (56) Список документов, цитированных в отчете о поиске: SU 1032866 A1, 09.02.1995. RU 26840 U1, 20.12.2002. RU 44167 U1, 27.02.2005. EP 0942150 A1, 15.09.1999. US 5211003 A, 18.05.1993. US 5077967 A, 07.01.1992. (54) НАПРАВЛЯЮЩЕЕ УСТРОЙСТВО ДЛЯ ПОТОКА ВОЗДУХА НА ВХОДЕ В КАМЕРУ СГОРАНИЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ (57) Реферат: Направляющее устройство для потока воздуха на входе в камеру сгорания газотурбинного двигателя содержит ...

СПОСОБ ИЗГОТОВЛЕНИЯ ДИАФРАГМЫ СТАТОРА В ПАРОВОЙ ТУРБИНЕ

Способ изготовления диафрагмы статора включает этап подготовки двух колец, подвергаемых механической обработке отдельно для получения радиальных выемок. Радиальные выемки имеют профиль лопаток статорной ступени и пригодны для последующего помещения в них лопаток. Затем накладывают клейкую тканевую ленту на наружный и внутренний диаметры колец и впрыскивают в выемки соответствующее количество пастообразного паяльного материала. При этом контролируют давление впрыскивания так, чтобы пастообразный паяльный материал не поднимался свыше предела, обозначенного краями радиальных выемок, обращенных к каналу для потока пара. Кольца с помещенными в них лопатками подвергают операции пайки твердым припоем в печи в условиях вакуума. Изобретение позволяет уменьшить время изготовления и сборки диафрагмы статора. 5 з.п. ф-лы, 3 ил.

СПОСОБ ИЗГОТОВЛЕНИЯ СИСТЕМЫ, СОДЕРЖАЩЕЙ МНОЖЕСТВО ЛОПАТОК, УСТАНОВЛЕННЫХ В ПЛАТФОРМЕ

Изобретение относится к области металлургии, а именно, к изготовлению сектора газотурбинного двигателя. Способ изготовления сектора колеса газотурбинного двигателя (11), содержащего лопатки (9), установленные в полках (7, 8) лопаток включает изготовление лопаток (9) отдельно от полок (7, 8) лопаток; приготовление смеси металлического порошка с термопластическим связующим материалом; впрыскивание смеси в литейную форму для получения заготовок полок (7, 8) лопаток; удаление связующего материала из заготовок полок (7, 8) лопаток; соединение лопаток (9) с заготовками полок (7, 8) лопаток путем установки лопаток (9) между внутренней (8) и внешней (7) полками лопаток. Концы лопаток (9) устанавливают в ложементы, выполненные в полках (7, 8) лопаток; спекание с получением сектора колеса газотурбинного двигателя (11) соединенного сектора (11). Обеспечивается качественное соединение лопаток с полками лопаток колеса газотурбинного двигателя. 3 ил.

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

FIELD: engines and pumps. SUBSTANCE: aircraft turbine stator module ring has multiple through bores to accommodate stator vanes. Every said bore defines mid line extending between first edge for vane trailing edge location and second edge for vane leading edge location. Stator vane locating bore is aligned with mechanical load release through cutout made at said ring opposite and spaced from said first edge of the bore in direction of mid line. Other invention of the set relates to stator part including above described ring and multiple vanes, to aircraft turbine module with said stator part and to turbine engine with said module. EFFECT: lower probability of cracking at stator ring in the area of blade trailing edge. 13 cl, 7 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 527 809 C2 (51) МПК F04D 29/54 (2006.01) F01D 9/04 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ (21)(22) Заявка: ИЗОБРЕТЕНИЯ К ПАТЕНТУ 2011149631/06, 04.05.2010 (24) Дата начала отсчета срока действия патента: 04.05.2010 (72) Автор(ы): БЕРТОЛИ Винченцо (FR) (73) Патентообладатель(и): СНЕКМА (FR) Приоритет(ы): (30) Конвенционный приоритет: (43) Дата публикации заявки: 20.06.2013 Бюл. № 17 R U 07.05.2009 FR 0953055 (45) Опубликовано: 10.09.2014 Бюл. № 25 1724443 A1, 22.11.2006. GB 1100384 A, 24.01.1968. US 7329087 B2, 12.02.2008. US 2143466 A, 10.01.1939. SU 352030 A1, 21.09.1972 (85) Дата начала рассмотрения заявки PCT на национальной фазе: 07.12.2011 2 5 2 7 8 0 9 (56) Список документов, цитированных в отчете о поиске: WO 2008/111957 A1, 18.09.2008. EP 2 5 2 7 8 0 9 R U EP 2010/055998 (04.05.2010) C 2 C 2 (86) Заявка PCT: (87) Публикация заявки PCT: WO 2010/128025 (11.11.2010) Адрес для переписки: 129090, Москва, ул. Б. Спасская, 25, строение 3, ООО "Юридическая фирма Городисский и Партнеры" (54) ТУРБИННЫЙ ДВИГАТЕЛЬ ЛЕТАТЕЛЬНОГО АППАРАТА, ЕГО МОДУЛЬ, ЧАСТЬ СТАТОРА ДЛЯ ТАКОГО МОДУЛЯ, А ТАКЖЕ КОЛЬЦО ДЛЯ ТАКОГО СТАТОРА (57) Реферат: Кольцо статора модуля турбинного ...

ЛОПАТКИ СТАТОРА ОСЕВОГО ТУРБОКОМПРЕССОРА И СПОСОБ ПРОИЗВОДСТВА

... 1. Статор (24) осевой турбомашины, содержащий:- наружный кожух (28) с расположенными в ряд по окружности отверстиями и по меньшей мере одной внутренней кольцевой канавкой, предназначенной для фиксирования кольцевого слоя истираемого материала (36),- ряд лопаток (26) статора с полками (38), расположенными в отверстиях и закрепленными посредством одного или нескольких наплавленных валиков между полками (38) и отверстиями,отличающийся тем, что ряд отверстий и внутренняя канавка частично перекрываются, так что часть (40, 42) одного или нескольких наплавленных валиков расположена в осевом направлении во внутренней канавке.2. Статор (24) по п.1, отличающийся тем, что полки (38) лопаток (26) формируют выступ, образующий внутреннюю канавку.3. Статор (24) по одному из пп.1 и 2, отличающийся тем, что полки (38) проходят на участке дна внутренней канавки.4. Статор (24) по п.1, отличающийся тем, что отверстия проходят через кожух (28) и внутреннюю канавку.5. Статор (24) по п.1, отличающийся тем, что ...

УЗЕЛ ДИФФУЗОР-ВЫПРЯМИТЕЛЬ ДЛЯ ТУРБОМАШИНЫ

... 1. Узел диффузор-выпрямитель, предназначенный для подачи воздуха в кольцевую камеру сгорания в турбомашине, содержащий радиальный диффузор, содержащий два кольцевых фланца, соединенных лопатками со спиральным наклоном, и кольцевой выпрямитель, установленный на выходе диффузора и содержащий две цилиндрических обечайки, соединенные радиальными лопатками, отличающийся тем, что кольцевые фланцы диффузора и цилиндрические обечайки выпрямителя выполнены из листов, в которых лопатки закреплены пайкой. 2. Узел по п.1 отличающийся тем, что листы диффузора и выпрямителя содержат отверстия, в которые концы лопаток вставляются и припаиваются. 3. Узел по п.2, отличающийся тем, что в отверстиях фланцев и обечаек концы лопаток припаяны по всей длине их периметра. 4. Узел по п.2 отличающийся тем, что отверстия листов образованы посредством лазерной резки. 5. Узел по п.1 отличающийся тем, что радиально внутренние концы кольцевых фланцев диффузора соединены с кольцевыми стенками, которые снабжены крепежными ...

УЗЕЛ СОПЛА ДЛЯ ТУРБИНЫ

... 1. Узел сопла для турбины, содержащий лопатку (43) сопла, имеющую внутреннюю боковую стенку (44) и внешнюю боковую стенку (46) и частично образующую путь потока после ее установки в турбине; внешнее кольцо (20); разделитель потока (11), имеющий горизонтальное удлинение (21); границу перехода между внешним кольцом (20) и внешней боковой стенкой (46), имеющую (i) радиальный фиксатор (48), (76), (78) и/или (ii) охватываемую/охватывающую границу (54), (82), (84) перехода и/или (iii) охватывающую выемку (106) с фланцем в виде радиально выступающих охватываемых ступенек (108), как на передней, так и на задней кромке внешней боковой стенки (46); и границу перехода между горизонтальным удлинением (21) и внутренней боковой стенкой (44), имеющую (i) радиальный фиксатор (48), (76), (78) и/или (ii) охватываемую/охватывающую границу (54), (82), (84) перехода и/или (iii) охватывающую выемку (106) с фланцами в виде выступающих радиально охватываемых ступенек (108), как на передней, так и на задней кромке ...

РЕМОНТНЫЙ ЭЛЕМЕНТ ДЛЯ ЛОПАТОЧНОГО УЗЛА ГАЗОВОЙ ТУРБИНЫ И СПОСОБ РЕМОНТА ПОВРЕЖДЕННОЙ ЛОПАТКИ ЛОПАТОЧНОГО УЗЛА ГАЗОВОЙ ТУРБИНЫ

СОПЛО ТУРБИНЫ И СПОСОБ ЕГО ИЗГОТОВЛЕНИЯ

СПОСОБ СОЕДИНЕНИЯ ДВУХ ЛОПАТОК НАПРАВЛЯЮЩЕГО АППАРАТА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Steam turbine interstage steam flow ring production method - commences with U-section ring which has vanes pressed into EDM produced flange holes and web bored out to create flow passages

The finished ring consists of an inner ring (2), outer ring (3) and, spaced at intervals between them, aerofoil section vanes (4) which direct steam onto the turbine blades at the desired angle. To produce the ring a ring (5) having a U-section lying on its side with the opening (7) in the steam flow path is first cut out of plate. Profiled holes for securing the vanes (4) are produced in the flanges (52,53) by EDM using a wire electrode machine with automatic cycle control. The vanes may be a press fit in the holes or secured by adhesive. ADVANTAGE - Rings of higher strength and quality can be produced more simply in shorter throughput time and at lower production cost.

Vorrichtung zum Abstuetzen von Schaufeln in einer Turbomaschine

VERFAHREN ZUM VERSCHWEISSEN UNGLEICHER METALLE

Verfahren zum Reparieren bzw. Fertigen eines Bauteils

Die Erfindung betrifft ein Verfahren zum Reparieren eines Bauteils, insbesondere eines statorseitigen Bauteils einer Gasturbine wie eines Gehäuses oder eines Leitschaufelkranzes, wobei aus dem Bauteil ein beschädigter Abschnitt herausgetrennt wird, und wobei ein den beschädigten sowie herausgetrennten Abschnitt ersetzender Neuabschnitt mit dem Bauteil durch Schweißen fest verbunden wird. Erfindungsgemäß wird der beschädigte Abschnitt derart aus dem zu reparierenden Bauteil herausgetrennt, das die Länge einer Trennnaht und damit einer späteren Schweißnaht minimiert wird, wobei abhängig von der Materialdickenverteilung entlang der Trennnaht zur Bereitstellung einer möglichst gleichmäßigen Materialdicke entlang der späteren Schweißnaht vom Bauteil Material abgetragen wird und wobei nach dem Verbinden des Bauteils mit dem Neuabschnitt durch Laserpulverauftragschweißen zumindest das abgetragene Material erneuert wird.

Gehäusevorrichtung eines Strahltriebwerks

Es wird eine Gehäusevorrichtung (15) eines Strahltriebwerks mit einer radial inneren Gehäuseeinrichtung (21) und einer radial äußeren Gehäuseeinrichtung (23) vorgeschlagen. Die innere Gehäuseeinrichtung (21) ist mit einem radial inneren ringförmigen Gehäuseteil (27) und einem radial äußeren ringförmigen Gehäuseteil (29) ausgeführt. Die innere Gehäuseeinrichtung (21) weist mehrere sich im Wesentlichen in radialer Richtung (R) zwischen dem inneren Gehäuseteil (27) und dem äußeren Gehäuseteil (29) erstreckende Schaufeln (31) auf. Die radial äußere Gehäuseeinrichtung (23) ist mit einem radial inneren ringförmigen Gehäuseteil (35) und einem radial äußeren ringförmigen Gehäuseteil (37) ausgeführt. Die äußere Gehäuseeinrichtung (23) weist mehrere sich im Wesentlichen in radialer Richtung (R) zwischen dem inneren Gehäuseteil (35) und dem äußeren Gehäuseteil (37) erstreckende Schaufeln (39, 41) auf. Die innere Gehäuseeinrichtung (21) ist als ein erstes einstückiges Gussbauteil und die äußere Gehäuseeinrichtung ...

Improvements relating to stator blade assemblies for compressors or turbines

... 769,148. Turbine and compressor blade assemblies. GENERAL MOTORS CORPORATION. May 9, 1955, No. 13329/55. Class 110 (3). A compressor or turbine stator blade assem bly comprises inner and outer arcuate members of strip metal secured together, blades being inserted radially through slots in one of the members, each blade having a root in the form of a flat platform which is sandwiched between the two arcuate members to secure the blade against radial movement, and a shoulder which by interlocking engagement with a slot in one of the members, secures the blade against rotation. In a first embodiment the inner arcuate member 3 is of channel section as shown in Fig. 5, and has surfaces 8 and upstanding flanges 9, and is formed with apertures 10 of polygonal shape. A blade 12 having a platform 15 and a shouldered portion 17 is passed through each aperture 10, the shouldered portion 17 engaging the aperture 10 and being of corresponding shape prevents rotational movement of the blade. The outer ...

STEMMED BLADE FOR A FLOW-STRAIGHTENING STAGE OF A TURBOSHAFT ENGINE AND METHOD OF FIXING SAID BLADE

Mounting a turbine nozzle on a combustion chamber having CMC walls in a gas turbine

The gas turbine comprises an annular combustion chamber (10) having inner and outer walls (12, 13) made of ceramic matrix composite material, and a high pressure turbine nozzle (20) secured to a downstream end of the combustion chamber and comprising a plurality of stationary airfoils (21) extending between the inner and outer walls (22, 23) of an annular flow path (24) through the nozzle for the gas stream coming from the combustion chamber. The turbine nozzle (20) is made of ceramic matrix composite material and it is connected to the downstream end of the combustion chamber (10) by brazing.

Vane assembly for use in a fluid flow machine

... 1,101, 529. Compressor blade fixing. ROLLS-ROYCE Ltd. May 18, 1966, No. 22147/66. Heading F1T. Individual stator guide blades 31 in the intermediate pressure compressor of a gas turbine by-pass engine are mounted with an interference fit between an annular wall member 23 and an annular forward extension 30 from the inner wall 14 of the by-pass passage (18), Fig. 1 (not shown). Each blade 31 has platforms 32, 33 mounted in annular recesses 34, 35 in the walls 23, 30 respectively, the interference fit being provided by heating wall 30 at assembly so that shrinkage occurs on cooling. The blades may in addition be bonded to walls 23, 30 by epoxy resin or bolted to at least one of the walls. Wall 30 is secured to wall 14 by bolts 37. Annular wall 24, bolted to wall 23, carries a rotor bearing (21).

Improvements in and relating to elastic fluid turbines

... 143,376. British Thomson-Houston Co., (General Electric Co.). May 1, 1919. Casings.-A diaphragm for an elastic-fluid turbine comprises a web portion 10 and a ring portion 11 formed of a number of thin metal plates or laminµ spaced apart by plates 12 to form the fluid passages 13. The laminµ of both web and ring portions have spaced slots 16 stamped therein and are assembled with the slots in staggered relation to receive the straight portions of the plates 12, as shown in Fig. 7. Further laminµ are then placed in position over the curved portions of the plates 12 and the whole welded or soldered together to form the complete diaphragm.

Improvements relating to the attachment of stator vanes in an axial-flow compressor

... 841,971. Compressor blades. GENERAL MOTORS CORPORATION. Aug. 13, 1958 [Aug. 13, 1957], No. 26083/58. Classes 110 (1) and 110 (3). The reduced end portions 9 of axial flow compressor stator vanes 3 are brazed or welded into slots cut in the sheet metal channel section 13 and strip 11 welded to the channel. The channel 13 is axially located by the beading 17 on the sides of metal channel 15 welded to casing 1. The channel 13 is located radially by the strips 21 which are secured to the inner casing 19 and engage the edges 18, 22 at ground surfaces. In further embodiments, not shown, the beading 17 is replaced by spacer blocks secured to the inner surface of the channel section 15, which may be replaced by a pair of angles or the channel and inner casing may be formed in a single piece.

NOZZLE GUIDE VANE STRUCTURE FOR A GAS TURBINE ENGINE

Improvements in or relating to a method of securing turbine or compressor blades to supporting parts

... 833,638. Fixing turbine blades. ROLLSROYCE Ltd. Dec. 20, 1956 [Jan. 6, 1956], No. 603/56. Addition to 702,390. Class 110 (3). [Also in Group XXII] In a method of producing in the manufacture of turbo machines and the like a simple pressure butt-resistance welded joint between one part grooved to provide spaced lands and another part, as claimed in the parent Specification, the lands 12 are formed on a turbine or compressor rotor or stator 10, and the said other part is a turbine or compressor blade 11 of aerofoil section having greater thickness adjacent the mid-chord position than at the leading and trailing edges and the butting end of the blade is chamfered adjacent its leading and trailing edges 11a, 11b so that initially only the mid-section 11c of the butting end contacts the rotor, but during welding the whole section penetrates the lands. The leading edge chamfer 11a is at a greater angle but of less chordwise extent than the trailing edge chamfer. The grooving may be continuous ...

Improvements in and relating to turbine nozzle assemblies

... 865,198. Turbine nozzle assemblies. PARSONS & MARINE ENGINEERING TURBINE RESEARCH & DEVELOPMENT ASSOCIATION. Sept. 25, 1959 [Nov. 7, 1958], No. 35797/58. Class 110 (3). [Also in Group XXII] A turbine nozzle assembly, Fig. 4, is made by piercing indexing holes in a straight flat strip of metal, Fig. 2 (not shown), rolling the strip into a circular arc, Fig. 3 (not shown), forming slots 6 in the strip 4 to accommodate nozzle vanes 5, forming thin cuts 7 between the slots and adjacent edge of the strip, locating the ends of the nozzle vanes 5 in the slots, and pressing the said edge to close the slots on to the vanes. The indexing holes are drilled by the apparatus described in Specification 863,842 and the slots 6 are punched out using the holes to index the punch. Lower and upper strips 3, 4, Fig. 4, form shrouds which together with their vanes 5 are located between rings 1, 2, Fig. 1, and the vanes are welded to these rings. Finally the distorted edges of the strips are machined smooth.

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

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

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

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

STEAM TURBINE DIAPHRAGM

WELDED AIRFOIL BLADE STRUCTURE

DOUBLE ROW COMPRESSOR STATORS

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

TURBINE NOZZLE WITH IMPINGEMENT BAFFLE

A turbine nozzle apparatus includes: a vane (16) extending between inner and outer bands (12, 14), the interior of the vane (16) being open and communicating with an aperture (30) in the outer band (14), wherein the vane (16) and the bands (12, 14) are part of a monolithic whole of low-ductility material; a metallic baffle (42) inside the vane (16), the baffle (42) having upper (44) and lower ends (46) and a peripheral wall (48) including a plurality of impingement holes (56) defining an interior space, closed off by an end wall (50) at the lower end (46); and a metallic retainer (58) having a body (60) with a shape generally matching the shape of the aperture (30), the body (60) bearing against the upper end (44) of the impingement baffle (42) and being connected to the outer band (14) by a plurality of mechanical fasteners (86, 92).

Leitvorrichtung für Dampf- oder Gasturbinen.

Leitrad für Dampf- und Gasturbinen.

Verfahren zur Herstellung einer Leitvorrichtung für Dampf- oder Gasturbinen.

Leitvorrichtung für Dampf- und Gasturbinen.

Verfahren zur Herstellung von Leitapparaten.

Turbinenleitrad.

Leitvorrichtung für Kreiselmaschinen.

Beschaufelter Konstruktionsteil für Turbomaschinen.

Leitschaufel-Trägerkonstruktion und Verfahren zu ihrer Herstellung

Verfahren zur Herstellung von künstlichen Gebilden, wie Fäden, Garnen, Fasern und dergleichen, aus Viskose

Beschaufelter Statorteil einer Axial-Turbomaschine

Beschaufelter Zwischenboden für Turbinen und Verfahren zu dessen Herstellung

Leitschaufeldeckel für Dampf- oder Gasturbinen

Procédé pour souder au moins deux pièces en un alliage à base de nickel

Aubage et procédé pour sa fabrication

Geschweisster Leitschaufelträger für axial beaufschlagte Dampf- oder Gasturbinen

Axialturboverdichter

Verfahren zum Herstellen eines Turbinenleitapparates

Verfahren zur Herstellung eines Düsenkastens für Turbinen

Leitschaufelkranz für Turbomaschinen

PROCEDURE FOR MANUFACTURING A TURBINE IDLER.

Method for joining of blades of turbine with cover band element, involves providing multiple blades arranged on blade carrier, which have rivet shanks at blade tips

The method involves providing multiple blades (11) arranged on a blade carrier, which have rivet shanks (13) at the blade tips. The rivet shanks extend in radial direction. A cover band element (14) is provided which has through holes adapted to the rivet shanks of the blades. The cover band element is placed on the blade tips of the multiple blades. The rivet shanks are hot press formed with a tool (16). The required heating of the rivet shanks is generated by friction between the rivet shanks and the tool for the hot press forming of the rivet shanks.

EXHAUST HOUSING FOR A GAS TURBINE AS WELL AS PROCEDURE FOR MANUFACTURING SUCH A EXHAUST HOUSING.

Ein Abgasgehäuse für eine Gasturbine umfasst eine Tragstruktur, welche in konzentrischer Anordnung einen Aussenring und einen Innenring umfasst, die durch eine Mehrzahl von radial angeordneten Streben miteinander verbunden sind, sowie einen Strömungsliner (19), der innerhalb der Tragstruktur angebracht ist und in konzentrischer Anordnung einen Aussenliner (20) und einen Innenliner (22) umfasst, die durch eine Mehrzahl von radial angeordneten Linerrippen (21) miteinander verbunden sind, wobei die Linerrippen (21) jeweils die Streben der Tragstruktur umschliessen und der Aussenliner (20) und der Innenliner (22) innerhalb der Tragstruktur einen Ringkanal für die heissen Abgase der Gasturbine bilden, und wobei die Linerrippen (21) jeweils unter Ausbildung von axial im Aussenliner (20) und im Innenliner (22) bis zum Linerrand verlaufenden Schweissnähten (35, 36) aus einem in Strömungsrichtung vorderen Teil und einem Hinterteil (30) zusammengeschweisst sind. Eine höhere Betriebszuverlässigkeit ...

Inspection tool for use with the production of a nozzle segment.

Es wird ein Inspektionswerkzeug zur Verwendung bei der Herstellung eines Düsensegmentes, das in einer Turbinendüsengruppe verwendet wird, bereitgestellt. Das Inspektionswerkzeug umfasst eine Ausrichtungsplatte (202) und eine Positionierungsplatte (204), die mit der Ausrichtungsplatte verbunden ist, wobei die Ausrichtungsplatte mindestens eine Öffnung (208, 209) aufweist, die sich durch dieselbe an einer vorgegebenen Stelle erstreckt, die einer gewünschten Ausrichtung einer Übergangsstücknut entspricht, welche auf dem Düsensegment, das gerade geprüft wird, festgelegt ist.

APPLIANCE FOR FIXATION OF SECTORS OF GUIDING APPARATUS HOLDING BLADES ALONG THE ARC OF CIRCLE

Method and tooling for assembling guide vane stage

The invention relates to a method and tooling for assembling a guide vane stage (1). The guide vane stage (1) comprises an inner shroud (6) and an outer shroud which are coaxial and connected by radial blades (8). The method includes a step of keeping plates (19) pressed against the external surface of the inner shroud (6) so that the plates (19) in a fluidtight manner and at least partially cover the gaps (15) formed between openings (10) in the inner shroud (6) and the blades (8), and a step of applying a resin sealant (11) to the internal surface (12) of the inner shroud (6) so that the sealant (11) fills the gaps (15) and so that the radially internal ends of the blades (8) are embedded in the sealant (11).

STATOR?BLADE ASSEMBLY FOR TURBOMACHINES

Ring of guidance of gases for gas turbines in particular for engines of plane with exhaust nozzle

Manufactoring process applicable in particular to the assembly of wings of stator for compressors and turbines with axial flow

Similar filaments, wire, fibres, etc and products deriving from viscose and their manufacture

METHOD FOR PRODUCING METAL ALLOY PARTS OF COMPLEX SHAPE

La présente invention concerne un procédé d'obtention d'une aube de turbomachine comprenant une âme, une tête et un pied, le procédé comprenant : - une étape de réalisation d'une ébauche à partir d'au moins deux pièces (50, 51), au moins l'une d'elles étant une pièce massive, lesdites au moins deux pièces étant assemblées par une technique de raccordement sans fusion et, - une étape d'usinage de cette ébauche pour aboutir à une aube avec un profil défini.

METHOD AND DEVICE OPTIMIZED FOR CRIMPING A WORKPIECE SUPPORT AN ABRADABLE HAS A RADIALLY INNER WALL OF A SECTOR OF A TURBOMACHINERY BLADING

METHOD OF MAKING A TURBINE VANE SEGMENT OR COMPRESSOR STATOR IN COMPOSITE MATERIAL FOR TURBOMACHINE AND TURBINE OR COMPRESSOR INCORPORATING A DISPENSER OR A RECTIFIER SHAPE OF THESE SECTORS

METHOD OF ASSEMBLING TWO BLADES OF A TURBOMACHINE NOZZLE

FRICTION WELDING RECTIFIERS AND HOLLOW ARMS ON THE INTERMEDIATE CASING OF A TURBOMACHINE

Improvements with the compressors with axial flow

RING FOR STATOR OF TURBOMOTOR Of AIRCRAFT HAS SLITS OF MECHANICAL UNLOADING Of PADDLES.

Compressor for gas turbine engine

METHOD FOR REPAIRING OR MANUFACTURING A COMPONENT

The invention relates to a method for repairing a component (10), particularly a stator-side component of a gas turbine such a housing or a vane ring during which a damaged section is removed from the component, and a new section that replaces the damaged and thus removed section is joined to the component in a fixed manner by welding. According to the invention, the damaged section is removed from the component to be repaired in a manner that minimizes the length of a separating seam (14) and thus a subsequent weld seam (14). Material is removed from the component according to the distribution of material thickness along the separating seam (14) in order to provide a material thickness that is as uniform as possible along the subsequent weld seam. At least the removed material is replaced after the component has been joined to the new section by laser powder deposition welding.

Method For Producing A Compressor Stator Of An Axial Turbomachine

The invention relates to a method of producing a low-pressure compressor stator for an axial turbine engine. The stator comprises an external shroud with stubs and an annular row of stator blades extending radially towards the inside from the stubs. The method comprises the following stages: supply or production of a starting bar; bending of the bar so that it makes a circle, in order to form an unwrought external shroud; turning to form an axial annular wall delimited by annular fixing flanges; orbital friction-welding of a row of blades onto the stubs of the external shroud. The stubs are realized during a milling stage of the bar or of the external shroud, the milling being carried out before or after the bending stage. The shroud and the blades can be produced in titanium or in a thermoplastic polymer.

LOW HOOP STRESS TURBINE FRAME SUPPORT

A gas turbine frame has inner and outer annular bands, respectively, joined together by generally radially extending struts therebetween. A radially outer conical support arm extends radially outwardly from the outer band and a radially inner conical support arm extends radially inwardly from the inner band. Circumferentially spaced apart inner and outer openings are disposed in the inner and outer conical support arms, respectively. Each of the struts has at least one radially extending hollow passage which extends through the inner and outer bands. The frame is a single piece integral casting. The inner and outer conical support arms have an equal number of the inner and outer circumferentially spaced apart openings. The inner circumferentially spaced apart openings are equi-angularly spaced apart and the outer circumferentially spaced apart openings are equi-angularly spaced apart. Each pair of the inner and outer circumferentially spaced apart openings are linearly aligned with the ...

Welding method of turbine diaphragm

A welding method of a turbine diaphragm for joining a spacer having a plurality of nozzle blades to outer and inner wheels. The spacer and the outer or inner wheel are welded from both sides. At first or inlet side of the diaphragm where a motive fluid flows into the diaphragm, the spacer and the outer and inner wheel are welded so that weld depth from the first side will be at most 0.5 L, wherein L is a distance between the first side and the top of the nozzle blade. At the second or outlet side of the diaphragm, from which the motive fluid exits, the spacer and the outer and inner wheels are weld so that a welded depth from the second side will be at least 0.3 C, wherein C is height of the nozzle blades.

Method of soldering a compressor nozzle ring of a gas turbine

Titanium-based metal parts ( 1, 3 ) are soldered by using as a filler metal ( 7 ) an aluminium alloy containing magnesium and virtually no silicon. Application to the bonding of blades ( 3 ) to the inner shroud ( 1 ) of an aeronautical gas turbine engine compressor nozzle ring.

Exhaust Housing for a gas turbine and method for producing the same

СПОСОБ ИЗГОТОВЛЕНИЯ ДИАФРАГМЫ ПАРОВОЙ ТУРБИНЫ

Изобретение относится к области машиностроения и может быть использовано при изготовлении диафрагмы (1) внутреннего корпуса модуля низкого или среднего давления паровой турбины. Упомянутая диафрагма содержит внутренний обод (2) и наружный обод (3), а также лопатки (4), привариваемые своими концами к упомянутым ободам (2, 3). При этом перед сваркой осуществляют операцию автоматической механической обработки концов (6, 7) каждой лопатки (4) для выполнения периферийной выемки (12) вокруг каждой из упомянутых лопаток (6, 7). Размеры указанной выемки задают в зависимости от уровня рабочей нагрузки на диафрагму (1) в турбине и в зависимости от условий процесса сварки концов лопаток с ободами, причем соотношение ширины Р выемки к ее глубине Н устанавливают между 0,5 и 2. Использование изобретения позволяет повысить качество и надежность изготовления лопаток и диафрагмы. 2 н. и 3 з.п. ф-лы, 4 ил. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 580 254 C2 (51) МПК F01D 9/04 (2006.01) B23P 15/02 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ (21)(22) Заявка: ИЗОБРЕТЕНИЯ К ПАТЕНТУ 2013145507/02, 09.03.2012 (24) Дата начала отсчета срока действия патента: 09.03.2012 (72) Автор(ы): БЮГЕН, Арно (FR), ТРАВЕРС, Доминик (FR) (73) Патентообладатель(и): АЛЬСТОМ ТЕКНОЛОДЖИ ЛТД (CH) Приоритет(ы): (30) Конвенционный приоритет: (43) Дата публикации заявки: 20.04.2015 Бюл. № 11 R U 11.03.2011 FR 1152006 (45) Опубликовано: 10.04.2016 Бюл. № 10 (85) Дата начала рассмотрения заявки PCT на национальной фазе: 11.10.2013 (86) Заявка PCT: EP 2012/054172 (09.03.2012) (87) Публикация заявки PCT: 2 5 8 0 2 5 4 (56) Список документов, цитированных в отчете о поиске: (см. прод.) 2 5 8 0 2 5 4 R U Адрес для переписки: 129090, Москва, ул. Б. Спасская, 25, строение 3, ООО "Юридическая фирма Городисский и Партнеры" (54) СПОСОБ ИЗГОТОВЛЕНИЯ ДИАФРАГМЫ ПАРОВОЙ ТУРБИНЫ (57) Реферат: Изобретение относится к области выемки (12) вокруг каждой из упомянутых машиностроения и может быть ...

СТАТОР ОСЕВОЙ ТУРБОМАШИНЫ, СПОСОБ ЕГО ПРОИЗВОДСТВА И ТУРБОМАШИНА, СОДЕРЖАЩАЯ УКАЗАННЫЙ СТАТОР

FIELD: machine building. SUBSTANCE: stator of axial turbine machine contains an external casing and row of stator blades with flanges. The external casing has holes arranged by row along the circle, and internal ring groove to lock the rings layer of the abraded material. Flanges of the stator blades are located in holes of the external casing, and are secured by means of the deposited bead between the flanges and hole planes. Holes row and internal groove partially intercross such, that part of each deposited bead is located in the axial direction in the internal groove. Another invention of the group relates to the turbine machine containing the compressor and turbine, at that at least one of the stators of the compressor and/or turbine is made as above described. During production of the above described stator initially the external casing and blades are made, then the blade flanges are installed and welded in the external casing holes. Then layers of the abraded material are located in the internal groove. EFFECT: group of inventions reduces size of the turbine machine, simplifies its design and manufacturing. 21 cl, 6 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (51) МПК F01D 9/02 (13) 2 568 353 C2 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ (21)(22) Заявка: ИЗОБРЕТЕНИЯ К ПАТЕНТУ 2013151836/06, 21.11.2013 (24) Дата начала отсчета срока действия патента: 21.11.2013 (72) Автор(ы): Хавье ШАРЛЬ (BE) (73) Патентообладатель(и): ТЕКСПЕЙС АЕРО С.А. (BE) Приоритет(ы): (30) Конвенционный приоритет: (43) Дата публикации заявки: 27.05.2015 Бюл. № 15 R U 21.11.2012 EP 12193700.7 (45) Опубликовано: 20.11.2015 Бюл. № 32 C 2 2 5 6 8 3 5 3 (54) СТАТОР ОСЕВОЙ ТУРБОМАШИНЫ, СПОСОБ ЕГО ПРОИЗВОДСТВА И ТУРБОМАШИНА, СОДЕРЖАЩАЯ УКАЗАННЫЙ СТАТОР (57) Реферат: Статор осевой турбомашины содержит изобретение группы относится к турбомашине, наружный кожух и ряд лопаток статора с содержащей компрессор и турбину, причем, по полками. Наружный кожух имеет расположенные ...

РЕМОНТНЫЙ ЭЛЕМЕНТ ДЛЯ ЛОПАТОЧНОГО УЗЛА ГАЗОВОЙ ТУРБИНЫ И СПОСОБ РЕМОНТА ПОВРЕЖДЕННОЙ ЛОПАТКИ ЛОПАТОЧНОГО УЗЛА ГАЗОВОЙ ТУРБИНЫ

FIELD: turbines or turbomachines.SUBSTANCE: repair element (200) comprises an inner platform portion (201) configured to replace a portion of the inner platform of the damaged blade, an outer platform portion (202), made with possibility of replacement of damaged platform outer platform, and aerodynamic part (204), connecting sections (201, 202) of internal platform and external platform and made with possibility of replacement of front edge part or rear edge part of damaged blade. Inner platform (201) section is made so that it extends to edge of first ring or blade unit ring of damaged blade ring sector, and external platform section (202) is made so that it passes to the edge of the second blade or sector of the blade unit ring of the damaged blade. Inner platform section (201), outer platform section (202) and aerodynamic part (204) are made with provision of repair element (200) insertion into blade unit (100) by purely translational movement, performed along the direction of introduction and having a component parallel to axes (153) of first and second rings or sectors of rings (150, 160).EFFECT: disclosed is a repair element for a gas turbine blade assembly and a method of repairing a damaged blade of a gas turbine blade assembly.13 cl, 13 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 727 543 C2 (51) МПК F01D 5/00 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F01D 5/00 (2020.02) (21)(22) Заявка: 2018131456, 15.03.2017 (24) Дата начала отсчета срока действия патента: Дата регистрации: 22.07.2020 16.03.2016 IT 102016000027545 (43) Дата публикации заявки: 16.04.2020 Бюл. № 11 (45) Опубликовано: 22.07.2020 Бюл. № 21 (85) Дата начала рассмотрения заявки PCT на национальной фазе: 16.10.2018 (56) Список документов, цитированных в отчете о поиске: RU 2173389 C2, 10.09.2001. US 4805282 A, 21.02.1989. RU 2008438 C1, 28.02.1994. RU 2186260 C1, 27.07.2002. 2 7 2 7 5 4 3 (73) Патентообладатель(и): НУОВО ПИНЬОНЕ ...

СПОСОБ ПАЙКИ ЛОПАТОК СПРЯМЛЯЮЩЕГО АППАРАТА КОМПРЕССОРА ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ

Изобретение может быть использовано при соединении пайкой металлических деталей из сплава на основе титана. Пайку осуществляют при давлении газа менее 1·10-2 Па. В качестве металлического припоя используют алюминиевый сплав, содержащий магний, а количество кремния в нем ограничено 0,3 мас.%. Металлические детали представляют собой внутренний бандаж и множество лопаток спрямляющего аппарата компрессора турбореактивного двигателя. Лопатки распределены в окружном направлении и проходят в радиальном направлении от внутреннего бандажа к наружному бандажу, и каждая из них проходит сквозь соответствующее отверстие, выполненное во внутреннем бандаже. Изобретение позволяет повысить механическую и химическую стойкость соединения. 2 н.п. и 13 з.п. ф-лы, 3 ил.

СПОСОБ ИЗГОТОВЛЕНИЯ СПРЯМЛЯЮЩЕГО АППАРАТА И ПОЛУЧЕННЫЙ ТАКИМ СПОСОБОМ СПРЯМЛЯЮЩИЙ АППАРАТ

FIELD: process engineering. SUBSTANCE: in production of turbo machine straightener with rim provided with several stator blades, first plies of reinforcing part are wound. Mandrel that doubles as the die and has extending parts while said plies of reinforcing parts have long cutouts located opposite said extending parts. Prefabricated plate is placed on every extending part to wind the last reinforcing part plies thereon for form the mid workpiece. Resin is injected into closed die with said mid workpiece to polymerise the latter impregnated with resin. Then, polymerised workpiece is withdrawn from said die to weld the blade root base or flange to every plate. Another invention of the set relates to turbo machine straightener fabricated as described above. EFFECT: higher strength. 15 cl, 5 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 564 740 C2 (51) МПК F01D 9/04 (2006.01) F01D 5/28 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ (21)(22) Заявка: ИЗОБРЕТЕНИЯ К ПАТЕНТУ 2011108051/06, 03.03.2011 (24) Дата начала отсчета срока действия патента: 03.03.2011 (72) Автор(ы): ДЮШЕН Жорж (BE) (73) Патентообладатель(и): ТЕКСПЕЙС АЭРО С.А. (BE) Приоритет(ы): (30) Конвенционный приоритет: (43) Дата публикации заявки: 10.09.2012 Бюл. № 25 R U 02.04.2010 EP 10159064.4 (45) Опубликовано: 10.10.2015 Бюл. № 28 2 5 6 4 7 4 0 R U (54) СПОСОБ ИЗГОТОВЛЕНИЯ СПРЯМЛЯЮЩЕГО АППАРАТА И ПОЛУЧЕННЫЙ ТАКИМ СПОСОБОМ СПРЯМЛЯЮЩИЙ АППАРАТ (57) Реферат: При изготовлении композитного содержащую предварительную заготовку, спрямляющего аппарата турбомашины, впрыскивают смолу и полимеризуют имеющего обод, снабженный рядом статорных пропитанную смолой предварительную лопаток, наматывают на оправку первые слои заготовку. После чего извлекают из формы армирующей детали. Оправка служит формой и полимеризованную предварительную заготовку имеет выступающие части, а указанные первые и с помощью сварки закрепляют на каждой из слои армирующей детали имеют удлиненные пластинок основание ...

ТУРБОМАШИНА, СОПЛОВОЙ АППАРАТ КОТОРОЙ УСТАНОВЛЕН НА КАМЕРЕ СГОРАНИЯ СО СТЕНКАМИ ИЗ КОМПОЗИТНОГО МАТЕРИАЛА

FIELD: engines and pumps. ^ SUBSTANCE: turbo machine includes annular combustion chamber with internal and external walls made from composite material with ceramic matrix, and nozzle assembly of high pressure turbine, which is rigidly attached to rear edge of combustion chamber. Nozzle assembly of high pressure turbine includes wings of guide blades, which are located between internal wall and external wall of annular channel of gas flows coming from combustion chamber through nozzle assembly. Rear edges of internal and external walls of combustion chamber are continued to rear edge of nozzle assembly and form internal wall and external wall of annular channel of gas flow though nozzle assembly. Wings of guide blades are made from composite material with ceramic matrix and attached by applying the method of brazing to walls of annular combustion chamber. ^ EFFECT: invention allows simplifying turbo machine design, as well as decreasing its weight. ^ 22 cl, 10 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) 2 392 447 (13) C2 (51) МПК F01D 9/02 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ, ПАТЕНТАМ И ТОВАРНЫМ ЗНАКАМ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (21), (22) Заявка: 2005117831/06, 09.06.2005 (24) Дата начала отсчета срока действия патента: 09.06.2005 (73) Патентообладатель(и): СНЕКМА (FR) (43) Дата публикации заявки: 20.12.2006 2 3 9 2 4 4 7 (45) Опубликовано: 20.06.2010 Бюл. № 17 2 3 9 2 4 4 7 R U Адрес для переписки: 191186, Санкт-Петербург, а/я 230, "АРСПАТЕНТ", пат.пов. В.М.Рыбакову, рег. № 90 (54) ТУРБОМАШИНА, СОПЛОВОЙ АППАРАТ КОТОРОЙ УСТАНОВЛЕН НА КАМЕРЕ СГОРАНИЯ СО СТЕНКАМИ ИЗ КОМПОЗИТНОГО МАТЕРИАЛА сопловой аппарат. Задние края внутренней и внешней стенок камеры сгорания продолжены до заднего края соплового аппарата и образуют внутреннюю стенку и внешнюю стенку кольцевого канала течения газов через сопловой аппарат. Перья направляющих лопаток выполнены из композитного материала с керамической матрицей и прикреплены методом пайки к стенкам кольцевой ...

СПОСОБ И ИНСТРУМЕНТ ДЛЯ СБОРКИ СТУПЕНИ ВЫПРЯМЛЕНИЯ

... 1. Способ сборки ступени выпрямления (1), состоящей из соосных внутренней обечайки (6) и наружной обечайки (7), соединенных по существу радиальными лопатками (8), отличающийся тем, что он включает в себя следующие этапы, на которых:(а) вводят радиально наружные концы лопаток (8) в отверстия наружной обечайки (7) и во введении с зазором (15) радиально внутренних концов упомянутых лопаток (8) в отверстия (10) внутренней обечайки (6),(b) в закреплении наружных концов лопаток (8) на наружной обечайке (7), например, посредством сварки,(c) поддерживают пластины (19) с упором на наружную поверхность внутренней обечайки (6), так чтобы пластины (19) покрывали герметично и, по меньшей мере, частично зазоры (15), образованные между отверстиями (10) внутренней обечайки (6) и лопатками (8), при этом пластины (19) устанавливают по окружности между лопатками (8),(d) наносят заливочную смолу (11) на внутреннюю поверхность (12) внутренней обечайки (6), так чтобы смола (11) заполнила зазоры (15), а радиально ...

СПОСОБ ИЗГОТОВЛЕНИЯ КОМПОНЕНТА СТАТОРА ИЛИ РОТОРА

... 1. Способ изготовления компонента (1, 8, 11, 14, 16) статора или ротора, который имеет по меньшей мере одну направляющую поток газа и/или передающую усилия перегородку (3, 30), соединенную по меньшей мере с одним кольцевым элементом (2, 9, 12, 13, 17, 18), отличающийся тем, что край перегородки (3, 30) прочно приваривают лазерной сваркой к кольцевому элементу с противоположной в радиальном направлении стороны таким образом, что соединенные между собой участки перегородки и кольцевого элемента образуют соединение (4) Т-образной формы. 2. Способ по п.1, отличающийся тем, что кольцевой элемент соединяют с несколькими перегородками (3, 30), расположенными на расстоянии друг от друга в окружном направлении. 3. Способ по п.1 или 2, отличающийся тем, что кольцевые элементы (2) соединяют между собой в окружном направлении в одно общее кольцо. 4. Способ по п.1, отличающийся тем, что кольцевые элементы (2, 10, 12, 18) образуют внутреннее кольцо, при этом перегородки (3, 30) соединяют с кольцевым ...

ДИАФРАГМЕННОЕ КОЛЬЦО ДЛЯ СТАТОРНЫХ ЛОПАТОК, ТУРБОМАШИНА И СПОСОБ

1. Диафрагменное кольцо для статорных лопаток, содержащеестаторные лопатки, каждая из которых имеет корневую часть и радиальное гнездо под штифт,первое кольцо, имеющее пазы, каждый из которых соединен с корневой частью одной из указанных статорных лопаток,второе кольцо, имеющее радиальные сквозные проходы, каждый из которых совмещен с радиальным гнездом одной из статорных лопаток, иштифты, каждый из которых проходит из одного радиального сквозного прохода в радиальное гнездо одной из статорных лопаток.2. Диафрагменное кольцо по п. 1, в котором размеры поперечных сечений радиальных сквозных проходов обеспечивают возможность сквозного прохождения через них указанных штифтов и сопряжения с указанными штифтами, когда штифты проходят в радиальные гнезда.3. Диафрагменное кольцо по п. 1, в котором статорные лопатки не проходят во второе кольцо.4. Диафрагменное кольцо по п. 1, в котором наружная оконечная поверхность статорных лопаток сопряжена с внутренней поверхностью второго кольца.5. Диафрагменное кольцо по п. 1, в котором по меньшей мере один штифт присоединен путем пайки к одному из указанных проходов второго кольца.6. Диафрагменное кольцо по п. 1, в котором по меньшей мере один штифт присоединен путем пайки к радиальному гнезду одной статорной лопатки.7. Диафрагменное кольцо по п. 1, в котором корневая часть по меньшей мере одной статорной лопатки присоединена путем пайки к одному из указанных пазов.8. Диафрагменное кольцо по любому из пп. 5-7, в котором первое кольцо и второе кольцо разрезаны в радиальном направлении между указанными паяными соединениями с образованием тем самым сегментированного диафрагменного кольца для статорных лопаток.9. Диафр РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (51) МПК F01D 9/04 (13) 2014 139 125 A (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ЗАЯВКА НА ИЗОБРЕТЕНИЕ (21)(22) Заявка: 2014139125, 02.04.2013 (71) Заявитель(и): НУОВО ПИНЬОНЕ СРЛ (IT) Приоритет(ы): (30) Конвенционный приоритет: 06.04.2012 IT CO2012A000014 (85) ...

HERSTELLUNGSVERFAHREN FUER EIN TURBINENLEITRAD

Schaufelaufbau,z.B. fuer den Verdichter eines Gasturbinenstrahltriebwerks

Leitschaufelsegment und dieses enthaltende Axialströmungsfluidmaschine

Ein Leitschaufelsegment (11), das einen Teil eines Leitschaufelrings bildet und in dem eine Mehrzahl von Leitschaufeln (20) in einer Umfangsrichtung (Dc) verbunden sind, mit:einem äußeren Verbindungselement (50), das sich in der Umfangsrichtung (Dc) erstreckt und die Mehrzahl von Leitschaufeln (20) verbindet,Positionierungsvorrichtungen (61,62) zur Positionierung von Endleitschaufeln (20a,20b), die sich an beiden Enden in der Umfangsrichtung (Dc) der Mehrzahl von Leitschaufeln (20) befinden, in Bezug auf das äußere Verbindungselement (50), undRadialbeschränkungsabschnitten (38,39), die das äußere Verbindungselement (50) so begrenzen, dass es relativ unbeweglich zu einer in einer Radialrichtung (Dr) äußeren Seite des Leitschaufelrings in Bezug auf mindestens eine Leitschaufel (20) der Mehrzahl von Leitschaufeln (20) ist, wobeijede der Mehrzahl von Leitschaufeln (20) einen Leitschaufelhauptkörper (21), der sich in der Radialrichtung (Dr) erstreckt, und einen äußeren Verstärkungsrand (32), ...

Leitschaufelsegment und dieses enthaltende Axialströmungsfluidmaschine

Verfahren zum Befestigen eines Kranzes von Leitschaufeln im Gehaeuse von Axial-Stroemungsmaschinen

DÜSENSEGMENT FÜR EINE DAMPFTURBINE

Stator blade assemblies for axial flow compressors or turbines

... 1,100,384. Making turbines and compressors; welding by fusion. BRISTOL SIDDELEY ENGINES Ltd. 22 April, 1966 [10 March, 1965], No. 10274/65. Headings B3A and B3R. [Also in Division F1] The titanium stator assembly of an axial flow turbine or compressor is constructed by forming blade mounting slots in an inner wall 1 and then welding outer walls 2 and 3 to wall 1 at 5, 6, 7 and 8. The root portions of the blades 9 are inserted through the slots in wall 1 and caused to form slots in walls 2 and 3 by stabbing. The blades are then secured in position by circumferentially extending lines of electron beam welding 10. Each blade root may be sharpened to an arrow head to assist the stabbing, in which case the sharp portion is ground away prior to welding.

TURBINE ENGINE COMPONENTS

Method of fixing flow-straightening blades in a turboshaft engine

Turbine blade and diaphragm

In one aspect, an axial flow turbine blade comprises inner 22, and outer 23 platforms and a corner fillet 24 where the blade aerofoil 21 meets each platform. Each platform comprises a trailing edge 22A, 23A side edges 22B, 22C, 23B, 23C, and a curved leading edge 22D, 23D, that follows the edge of the corner fillet in the region of the leading edge 26 of the aerofoil. In another aspect, a diaphragm construction comprises an annulus of turbine blades having inner and outer platforms accommodated by apertures in inner and outer spacer rings. The platforms may be welded to the spacer rings by a method involving welding cover plates to the leading and trailing edges of the platforms, machining weld preparation slots, performing a welding process to fill the slots and secure the platforms to the spacer rings, and machining the cover plates to remove them from the diaphragm.

Welded stator vane assemblies for turbomachines

A method of manufacturing a welded stator vane assembly for a turbomachine. The method comprises the steps of providing radially spaced inner and outer slotted members 38, 40 and a plurality of vanes 14 each of which has end fittings 64, 66 at each end. The end fittings 64, 66 are slotted into the slots 48, 52 of the inner and outer members 38, 40 and the end fittings are welded to the members 38, 40. In effect, the slotted members 38, 40 define a plurality of "spacers" 70 between the slots 48, 52 interconnected by the metal of members 38, 40 that define the base of the slots 48, 52. Subsequently, this metal is machined away to expose the welds and leave two concentric rings interconnected by the vanes. The rings consist of alternately end fittings and "spacers". ...

Improvements in or relating to bladed or vaned structures for guiding the flow of a fluid

... 779,056. Turbine guide vane assembly. ROLLS-ROYCE, Ltd. Dec. 2, 1954 [Feb. 12, 1954], No. 4272/54. Class 110 (3). In a nozzle guide vane assembly such as a guide vane assembly of a gas turbine, the vanes are secured at each end in the adjacent parts of supporting structure and each vane has the securing connection at one end at least effected through flexible strip adapted to accommodate relative thermal expansion by bending. The guide vanes 12 which are disposed between rotor blading stages 10, 11 are mounted at their radially inner ends in apertures in rigid annular shrouds 13 and are welded thereto. The outer ends of the vanes pass through apertures in individual flexible strips 17 and are welded thereto. The strips extend beyond the leading and trailing edges of the vanes and are inclined in a direction parallel to the blade chords. The ends of the strips 17 are welded at their ends to annular members 18, 19, the member 18 having a radial flange by which it is bolted to the turbine ...

A turbine nozzle-ring assembly with guide blades and method for assembling the same

... 929,960. Making turbines; brazing. WERKSPOOR N.V. Aug. 26, 1960 [May 9, 1960], No. 29477/60. Classes 83 (2) and 83 (4). [Also in Group XXVI] A turbine nozzle ring assembly with guide blades has blades 3 having integral supports 4 engaging in recesses 5 in the supporting structure 1, 2 so that the supports 4 leave spaces 8, which are filled with brazing material. The brazing material also fills the clearance spaces between the supports 4 and the nozzle rings 1, 2 and between the blades 3 and nozzle rings. After inserting the blade 4 and the brazing material, the recesses 5 are closed by plates 6, 7 and the complete assembly placed in an oven for brazing. After brazing, heat treatment may be carried out in the same oven. The brazing material may be a silver-copper-nickel alloy or a nickel-chromium alloy.

PROCEDURE FOR REPAIRING A CONSTRUCTION UNIT

Сегментированный сопловой аппарат малорасходной паровой турбины на органическом рабочем теле

Полезная модель относится к области теплоэнергетики, а именно к конструкции малорасходных паровых турбин, и может быть использована в паротурбинных установках на органическом рабочем теле, как элемент конструкции статора турбины. Предлагаемый сегментированный сопловой аппарат содержит простые сегменты, размещенные по окружности встык и свариваемые с внешними и внутренними кольцами, крепящие сегментированный сопловой аппарат к статору турбины. Каждый сегмент имеет с двух сторон прилегающие поверхности со спрофилированными выборками, при сложении сегментов получается сегментированный диск с соплами, направленные под углом α1≤ 14° к плоскости вращения колеса турбины. Количество сегментов определяется количеством сопел nсегм= nсопел, рассчитанных при проектировании турбины. Технической проблемой, на решение которой направлена заявляемая полезная модель, является создание простой и надежной конструкции сегментированного соплового аппарата для малорасходной паровой турбины на органическом рабочем теле, не требующей специальной оснастки, высококвалифицированного персонала и высокой точности изготовления деталей. Технический результат достигается тем, что предлагаемая конструкция позволяет удовлетворить требования к малорасходным сопловым аппаратам, снизить трудоемкость при изготовлении и сборке соплового аппарата за счет доступности и простоты поверхностей, а также исключить осевой прогиб соплового аппарата за счет применения равнопрочных сварных швов с обеих сторон конструкции.В предлагаемой конструкции нет исходного профиля сопловой лопатки, так как отсутствует входная кромка, а сложение, как минимум, двух сегментов образует рабочий канал, который в совокупности с зазором между сегментированным сопловым аппаратом и рабочими лопатками образует рабочий канал, в котором происходит течение рабочего тела. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 205 426 U1 (51) МПК F01D 9/04 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ПОЛЕЗНОЙ МОДЕЛИ К ПАТЕНТУ ( ...

Method of sealing cooling holes

A method of sealing a gap between an aerofoil component ( 30 ) and a further component ( 32 ). The method comprises placing the aerofoil component ( 30 ) in close proximity with the further component ( 32 ) to define a gap ( 38 ) therebetween, applying a thermoplastic material ( 20 ) to the gap ( 38 ) in a molten phase and cooling the thermoplastic material ( 20 ) to set the thermoplastic material ( 20 ).

Turbine engine stator and method of assembly of the same

A turbine engine stator stage includes a plurality of vanes with each of the plurality of vanes having a camber angle. The plurality of vanes is arranged in a plurality of groups with each group including a pre-determined sequence of vanes. The ordering of vanes within each group is determined by the camber of the individual vanes. This results in an arrangement of vanes within the stator stage which can modify the flow characteristics of the air entering the stator stage to reduce the circumferential pressure variation in the flow region immediately downstream of the stator stage.

Bicast turbine engine components

A turbine blade assembly includes a turbine blade having a pressure sidewall and an opposed suction sidewall and a first snubber assembly associated with one of the pressure sidewall and the suction sidewall. The first snubber assembly includes a first base portion extending outwardly from the one of the pressure sidewall and the suction sidewall, and a first snubber portion. The first base portion is integrally cast with the turbine blade and includes first connection structure. The first snubber portion is bicast onto the first base portion and includes second connection structure that interacts with the first connection structure to substantially prevent separational movement between the first base portion and the first snubber portion.

GAS TURBINE ENGINE STATOR VANE ASSEMBLY WITH SPLIT SHROUD

A method of assembling gas turbine engine front architecture includes positioning a first shroud and a first shroud portion radially relative to one another. Multiple vanes are arranged circumferentially between the first shroud and the first shroud portion. A second shroud portion is secured to the first shroud portion about the vanes. The first and second shroud portions provide a second shroud. The vanes are mechanically isolated from the first and second shrouds. 1. A method of assembling gas turbine engine front architecture comprising the steps of:positioning a first shroud and a first shroud portion radially relative to one another;arranging multiple vanes circumferentially between the first shroud and the first shroud portion;securing a second shroud portion to the first shroud portion about the vanes, the first and second shroud portions providing a second shroud; andmechanically isolating the vanes from the first and second shrouds.2. The method according to claim 1 , wherein the first and second shrouds respectively correspond to inner and outer shrouds.3. The method according to claim 2 , wherein the arranging step includes inserting the vanes into first and second slots respectively provided in the outer and inner shrouds claim 2 , and comprising the step of applying a liquid sealant around a perimeter of the vanes and at least one of the shrouds claim 2 , and bonding and supporting the ends of vanes relative to one of the shrouds with the liquid sealant.4. The method according to claim 3 , wherein each blade includes outer and inner perimeters respectively received in the first and second slots claim 3 , and the arranging step includes providing gaps between the outer and the inner perimeters and the outer and inner shrouds at their respective first and second slots claim 3 , wherein the applying step includes laying the liquid sealant about at least one of the inner and outer perimeters within their respective gaps.5. The method according to claim 4 , ...

Gas turbine engine composite vane assembly and method for making same

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

COMBUSTOR INLET MIXING SYSTEM WITH SWIRLER VANES HAVING SLOTS

A combustor inlet mixing system () formed from a plurality of circumferentially spaced swirler vanes () extending radially outward from a nozzle hub. Each of the swirler vanes () may have a length () that extends downstream along at least a portion of the combustor inlet mixing system (), and may further have a thickness () that extends along a circumference of the nozzle hub. At least one of the swirler vanes () may further have at least one slot () cut entirely through the thickness () of a portion of the swirler vane (). The slot () may separate the swirler vane () from the nozzle hub along a portion of the length () of the swirler vane (). 117-. (canceled)18. A turbine engine , comprising:at least one combustor positioned upstream from a rotor assembly, wherein the rotor assembly includes at least one row of turbine blades extending radially outward from a rotor;a compressor positioned upstream from the at least one combustor;at least one compressor exhaust plenum extending between the compressor and the at least one combustor; andat least one combustor inlet mixing system formed from a plurality of circumferentially spaced swirler vanes extending radially outward from a nozzle hub, each of the plurality of swirler vanes having a length that extends downstream along at least a portion of the at least one combustor inlet mixing system and further having a thickness that extends along a circumference of the nozzle hub, wherein at least one swirler vane of the plurality of swirler vanes further has at least one slot cut entirely through the thickness of a portion of the at least one swirler vane, the at least one slot separating the at least one swirler vane from the nozzle hub along a portion of the length of the at least one swirler vane.19. The turbine engine of claim 18 , wherein the at least one slot is configured to add a layer of at least partially non-swirling air around the nozzle hub.20. The turbine engine of claim 18 , wherein the at least one slot is ...

Turbine blade, turbine, and method for producing turbine blade

A turbine blade disposed along a radial direction of a turbine includes: an airfoil portion positioned in a fluid flow passage of the turbine; and a shroud portion positioned on an inner side or an outer side of the airfoil portion in the radial direction, and having an opening with which an end portion of the airfoil portion is to be engaged. A clearance is formed between a wall surface forming the opening of the shroud portion and an outer peripheral surface of the end portion of the airfoil portion. The wall surface of the shroud portion and the outer peripheral surface of the airfoil portion are joined to each other. At least one of the shroud portion or the airfoil portion has a cooling hole formed thereon, the cooling hole having an opening into the clearance and being configured to supply the clearance with a cooling fluid.

POTTED STATOR VANE WITH METAL FILLET

A vane for use with a stator assembly is disclosed herein. The vane having: an upper mounting portion; a lower mounting portion; an airfoil body extending between the upper mounting portion and the lower mounting portion; a first integrally formed metal fillet located between the upper mounting portion and the airfoil body, the first integrally formed metal fillet defining a first mounting surface located between the upper mounting portion and first integrally formed metal fillet; and a second integrally formed metal fillet located between the lower mounting portion and the airfoil body, the second integrally formed metal fillet defining a second mounting surface located between the lower mounting portion and second integrally formed metal fillet. 1. A vane for use with a stator assembly , the vane comprising:an upper mounting portion;a lower mounting portion;an airfoil body extending between the upper mounting portion and the lower mounting portion;a first integrally formed metal fillet located between the upper mounting portion and the airfoil body, the first integrally formed metal fillet defining a first mounting surface located between the upper mounting portion and first integrally formed metal fillet; anda second integrally formed metal fillet located between the lower mounting portion and the airfoil body, the second integrally formed metal fillet defining a second mounting surface located between the lower mounting portion and second integrally formed metal fillet.2. The vane of claim 1 , wherein the first integrally formed metal fillet extends completely around a periphery of the airfoil and the second integrally formed metal fillet extends completely around a periphery of the airfoil.3. The vane of claim 2 , wherein the first mounting surface extends laterally away from the upper mounting portion and the second mounting surface extends laterally away from the lower mounting surface.4. The vane of claim 3 , wherein the first mounting surface further ...

VANE ASSEMBLY

A compressor for a gas turbine engine includes a plurality of rotating wheel assemblies, a plurality of static vane assemblies, and a case extending around the rotating wheel assemblies and static vane assemblies. The static vane assemblies include an inner band, an outer band, and a plurality of vanes extending between the inner and outer bands. 1. A vane ring segment for use in a gas turbine engine , the vane ring segment comprisingan inner band that extends around a portion of a central axis, the inner band including a radially-inner surface facing toward the central axis, a radially-outer surface facing away from the central axis, and a plurality of inner-band vane apertures extending through the radially-inner and radially-outer surfaces of the inner band,an outer band that extends around a portion of a central axis and that is radially spaced apart from the inner band, the outer band including a radially-inner surface facing toward the central axis, a radially-outer surface facing away from the central axis, and a plurality of outer-band vane apertures extending through the radially-inner and radially-outer surfaces of the outer band, anda plurality of vanes coupled to the inner and outer bands, wherein each vane extends radially outward through one of the plurality of outer-band vane apertures beyond the radially-outer surface of the outer band and each vane is bonded to the outer band by a first layer of braze.2. The vane ring segment of claim 1 , wherein each vane includes a body portion that extends from the radially-outer surface of the inner band to the radially-inner surface of the outer band and an outer attachment portion that extends radially outward from the body portion through one of the plurality of outer-band vane apertures beyond the radially-outer surface of the outer band and the first layer of braze is located between the outer attachment portion and the outer band.3. The vane ring segment of claim 2 , wherein a radial cross-section of the ...

Methods of manufacturing a tandem guide vane segment

Methods for manufacturing a tandem guide vane segment that includes an outer platform, a front guide vane, and a rear guide vane, wherein the front guide vane and the rear guide vane are arranged in a firmly fixated manner with respect to one another. One method includes manufacturing an integral front segment section that includes the front guide vane and a front section of the outer platform, manufacturing an integral rear segment section that includes the rear guide vane and a rear section of the outer platform, and connecting the two segment sections to each other.

MID-TURBINE FRAME AND GAS TURBINE ENGINE INCLUDING SAME

A mid-turbine frame for a gas turbine engine ducts gases between a high pressure turbine and a low pressure turbine. The mid-turbine frame may include an outer flowpath ring, an inner flowpath ring, and a plurality of vanes extending therebetween. The outer flowpath ring comprises a unitary structure, while the inner flowpath ring and the plurality of vanes comprises a plurality of segments. 1. A mid-turbine frame for a gas turbine engine comprising:a duct that extends between a high pressure turbine and a low pressure turbine, the duct comprising an outer flowpath ring and an inner flowpath ring;wherein the inner flowpath ring is situated radially inward of the outer flowpath ring;wherein the outer flowpath ring comprises a unitary structure; andwherein the inner flowpath ring comprises a plurality of segments that together form the inner flowpath ring.2. The mid-turbine frame of claim 1 , wherein each segment includes a first tenon that defines a first axial terminus of each segment and a second tenon that defines a second axial terminus of each segment.3. The mid-turbine frame of claim 1 , wherein the inner flowpath ring comprises a plurality of inner arcing surfaces claim 1 , each of the inner arcing surfaces carrying at least one vane that extends from radially outward from the inner arcing surface toward the outer flowpath ring.4. The mid-turbine frame of claim 1 , wherein each of the segments are formed as a unitary structure.5. The mid-turbine frame of claim 3 , wherein each vane includes a channel.6. The mid-turbine frame of claim 2 , wherein the first tenon of a first segment is joined to the second tenon of a second segment by a seal that clamps the first tenon and the second tenon together.7. The mid-turbine frame of claim 6 , wherein the seal comprises:a male seal structure having a body and a protruding member extending away from the body; anda receiving member having an aperture configured to receive the protruding member.8. The mid-turbine frame of ...

METHODS FOR MANUFACTURING A TURBINE NOZZLE WITH SINGLE CRYSTAL ALLOY NOZZLE SEGMENTS

Methods for manufacturing a turbine nozzle are provided. A plurality of nozzle segments is formed. Each nozzle segment comprises an endwall ring portion with at least one vane. The plurality of nozzle segments are connected to an annular endwall forming a segmented annular endwall concentric to the annular endwall with the at least one vane of each nozzle segment extending between the segmented annular endwall and the annular endwall. 1. A method for manufacturing a turbine nozzle comprising:forming a plurality of nozzle segments, each nozzle segment comprising an endwall ring portion with at least one vane, the at least one vane having a free end portion;connecting the free end portion of the at least one vane of each nozzle segment of the plurality of nozzle segments to an annular endwall, with the endwall ring portion of each nozzle segment of the plurality of nozzle segments forming a segmented annular endwall concentric to the annular endwall with the at least one vane of each nozzle segment extending between the segmented annular endwall and the annular endwall; andforming the annular endwall prior to the connecting by separately casting the annular endwall as one-piece,wherein the step of connecting the plurality of nozzle segments comprises brazing the free end portion of the at least one vane of each nozzle segment to the annular endwall.2. The method of claim 1 , wherein the step of forming a plurality of nozzle segments comprises forming the plurality of nozzle segments with a single crystal material.3. The method of claim 1 , wherein the step of forming a plurality of nozzle segments comprises forming by casting.4. The method of claim 1 , wherein the step of forming a plurality of nozzle segments comprises forming a plurality of singlet nozzle segments claim 1 , doublet nozzle segments claim 1 , triplet nozzle segments claim 1 , quadruplet nozzle segments claim 1 , or combinations thereof.5. The method of claim 1 , further comprising the step of ...

Gas turbine engine ceramic component assembly attachment

A gas turbine engine component assembly includes first and second portions, wherein at least one of the first and second portions is a ceramic material. The first portion includes an aperture having a first angled surface. The second portion is disposed within the aperture and includes a second angled surface adjacent to the first angled surface. The first and second angled surfaces lock the first and second portions to one another under a pulling load. A bonding material operatively secures the first and second angled surfaces to one another.

GAS TURBINE ENGINE COMPOSITE VANE ASSEMBLY AND METHOD FOR MAKING THE SAME

A gas turbine engine composite vane assembly and method for making same are disclosed. The method includes providing at least two gas turbine engine airfoil composite preform components. The airfoil composite preform components are interlocked with a first locking component so that mating faces of the airfoil composite preform components face each other. A filler material is inserted between the mating surfaces of the airfoil composite preform components. 120.-. (canceled)21. A method for forming a gas turbine engine component comprising:providing at least two gas turbine engine composite preform components,interlocking the airfoil composite preform components with a first locking component so that mating faces of the composite preform components face each other, andinserting a filler material between the mating faces of the composite preform components.22. The method of claim 21 , wherein one or more of the composite preform components are provided in a partially rigidized state.23. The method of claim 21 , wherein the composite preform components comprise woven fiber.24. The method of claim 21 , wherein the composite preform components comprise a first component including at least one connecting tab and an endwall including at least one endwall connecting tab claim 21 , and the interlocking comprises inserting the first locking component into a through hole in the at least one connecting tab and a through hole in the at least one endwall connecting tab.25. The method of claim 21 , wherein the composite preform components comprise a first component including at least one connecting tab providing the mating face of the first component claim 21 , and an endwall including at least one endwall connecting tab providing the mating face of the endwall claim 21 , and the inserting comprises inserting the filler material between the mating face of the at least one airfoil connecting tab and the mating face of the at least one endwall connecting tab.26. The method of claim ...

TURBINE ENGINE VANE ARRANGEMENT HAVING A PLURALITY OF INTERCONNECTED VANE ARRANGEMENT SEGMENTS

A turbine engine vane arrangement and a method for manufacturing a turbine engine vane arrangement are provided. The vane arrangement includes a plurality of vane arrangement segments arranged circumferentially around an axial centerline. Each of the vane arrangement segments includes an airfoil that extends radially between a first platform segment and a second platform segment. The first platform segment extends circumferentially between a first mate face and a second mate face. The first mate face of a first of the vane arrangement segments is bonded to the second mate face of a second of the vane arrangement segments. 1. A turbine engine vane arrangement , comprising:a plurality of vane arrangement segments arranged circumferentially around an axial centerline, the vane arrangement segments consisting of a plurality of base segments and a keystone segment;the base segments and the keystone segment each comprising an airfoil extending radially between a first platform segment and a second platform segment;the first platform segment extending circumferentially between a first mate face and a second mate face;the first mate face of a first of the base segments bonded to the second mate face of the keystone segment; andthe first platform segment comprising a width that extends circumferentially between the first mate face and the second mate face, the width of each of the base segments having a first value, and the width of the keystone segment having a second value that is different than the first value.2. The vane arrangement of claim 1 , wherein the first mate face of the first of the base segments is brazed to the second mate face of the keystone segment.3. The vane arrangement of claim 1 , wherein the first platform segment is configured as a radial inner platform segment claim 1 , and the second platform segment is configured as a radial outer platform segment.4. The vane arrangement of claim 1 , wherein the first platform segment is configured as a radial ...

Methods and Assemblies for Attaching Airfoils Within a Flow Path

Flow path assemblies for gas turbine engines are provided. For example, a flow path assembly comprises an inner wall; a unitary outer wall; and a plurality of nozzle airfoils having an inner end radially opposite an outer end. The unitary outer wall defines a plurality of outer pockets each configured for receipt of the outer end of one of the nozzle airfoils, and the inner wall includes defines a plurality of inner pockets each configured for receipt of the inner end of one of the plurality of nozzle airfoils. A portion of each inner pocket is defined by a forward inner wall segment and an aft inner wall segment. In another embodiment, a flow path assembly comprises an inner wall defining a plurality of bayonet slots that each receive a bayonet included with each of a plurality of nozzle airfoils that are integral with a unitary outer wall. 119.-. (canceled)20. A flow path assembly for a gas turbine engine , the flow path assembly comprising:an inner wall;an outer wall; anda plurality of nozzle airfoils, each nozzle airfoil having an inner end radially opposite an outer end,wherein the inner wall and the outer wall define a combustor of the combustion section,wherein the outer wall defines a plurality of outer pockets, each outer pocket configured for receipt of the outer end of one of the plurality of nozzle airfoils, andwherein the inner wall includes a forward segment and an aft segment, the inner wall defining a plurality of inner pockets such that a portion of each inner pocket is defined by the forward segment and a remaining portion of each inner pocket is defined by the aft segment, each inner pocket configured for receipt of the inner end of one of the plurality of nozzle airfoils such that a nozzle airfoil extends from each inner pocket to a respective outer pocket.21. The flow path assembly of claim 20 , wherein the plurality of outer pockets are defined along an area of an inner surface of the outer wall claim 20 , and wherein the outer wall is built up ...

METHOD OF ASSEMBLY OF BI-CAST TURBINE VANE

One aspect of the present disclosure includes a turbine vane assembly comprising a vane made from ceramic matrix composite material having an outer wall extending between a leading edge and a trailing edge and between a first end and an opposing second end; an endwall made at least partially from a ceramic matrix composite material configured to engage the first end of the vane; and a retaining region including corresponding bi-cast grooves formed adjacent the first end of the vane and a receiving aperture formed in the endwall; wherein a bond is formed in the retaining region to join the vane and endwall together. 120.-. (canceled)21. A method of assembling a gas turbine engine vane including an airfoil having an outer surface extending between a leading edge and a trailing edge and between a first end and a second end; a through slot extending between the first and second ends of the airfoil; and a spar slidingly engaged within the slot of the airfoil , the spar including a pair of extensions with at least one bi-cast groove formed on opposing ends thereof , wherein the extensions of the spar are configured to engage with corresponding apertures formed in a pair of opposing endwalls , the method comprising:inserting at least one of the extensions into one of the apertures of the end walls,injecting a pre-cursor into the at least one bi-cast groove,heating the pre-cursor to form a fixed connection.22. The method of claim 21 , wherein the pre-cursor is initially a powder or a liquid.23. The method of claim 21 , wherein a cross sectional shape of the spar substantially conforms with the cross sectional shape of the through slot and at least one of the pair of extensions protrudes from the through slot.24. The method of claim 21 , wherein heating the pre-cursor including forming a fixed connection between the spar relative to at least one of the endwalls while the airfoil remains slidingly free within the through slot.25. The method of claim 21 , wherein injecting the ...

METHOD AND APPARATUS FOR REPAIR OF A DIAPHRAGM OF A ROTARY MACHINE

A method of repairing a diaphragm of a rotary machine includes removing an initial steam path from the diaphragm. The initial steam path includes a plurality of initial partitions. Each initial partition is associated with an original trailing edge profile and an original axial length. The method also includes coupling a replacement steam path to the diaphragm. The replacement steam path includes a plurality of replacement partitions. Each replacement partition has a replacement axial length that is greater than the original axial length. 1. A method of repairing a diaphragm of a rotary machine , said method comprising:removing an initial steam path from the diaphragm, the initial steam path including a plurality of initial partitions, each initial partition associated with an original trailing edge profile and an original axial length; andcoupling a replacement steam path to the diaphragm, the replacement steam path including a plurality of replacement partitions, each replacement partition having a replacement axial length that is greater than the original axial length.2. The method of claim 1 , wherein said coupling the replacement steam path to the diaphragm comprises:coupling a replacement inner band of the replacement steam path to a radially inner web of the diaphragm; andcoupling a replacement outer band of the replacement steam path to a radially outer ring of the diaphragm.3. The method of claim 1 , wherein said coupling the replacement steam path to the diaphragm further comprises coupling the replacement steam path including each replacement partition having a preselected trailing edge profile that is substantially the original trailing edge profile.4. The method of claim 1 , wherein said coupling the replacement steam path to the diaphragm further comprises coupling the replacement steam path including each replacement partition having a preselected trailing edge profile that is other than substantially the original trailing edge profile.5. The method ...

DIFFUSER HAVING PLATFORM VANES

According to some aspects of the present disclosure, a diffuser for a centrifugal compressor is provided. The diffuser may comprise an outerband casing and an innerband casing. The outerband casing may comprise an annular flowpath boundary member that has a flowpath boundary surface. The flowpath boundary member may define a plurality of vane-receiving pockets spaced about a circumference of the member. The innerband casing may comprise an annular flowpath boundary member that has a flowpath boundary surface. The flowpath boundary member may comprise a plurality of vanes spaced about a circumference of the member. Each of said plurality of vanes may comprise a vane body that extends from the flowpath boundary surface, a platform head that has a lateral dimension normal to the length of the vane body greater than the lateral dimension of the vane body, and a fillet between the platform head and the vane body. The innerband casing may be positioned so that the platform head of each of the plurality of vanes is received in a respective vane-receiving pocket defined by the flowpath boundary member of the outerband casing. When received, the fillet of each of the plurality of vanes may be adjacent the flowpath boundary surface of the flowpath boundary member of said outerband casing. The flowpath boundary surfaces of each of said casings and said vanes define a fluid flowpath in said diffuser. 1. A diffuser for a centrifugal compressor comprising:an outerband casing comprising an annular flowpath boundary member having a flowpath boundary surface, said flowpath boundary member defining a plurality of vane-receiving pockets spaced about a circumference of the member; andan innerband casing comprising an annular flowpath boundary member having a flowpath boundary surface, said flowpath boundary member comprising a plurality of vanes spaced about a circumference of the member, each of said plurality of vanes comprising a vane body extending from the flowpath boundary ...

FLOW PATH ASSEMBLIES FOR GAS TURBINE ENGINES AND ASSEMBLY METHODS THEREFORE

Flow path assemblies and methods for forming such flow path assemblies for gas turbine engines are provided. For example, a method for assembling an airfoil with a boundary structure to form a flow path assembly is provided. The method includes machining an opening into the boundary structure. The opening is sized to receive an airfoil or other component. The method also includes machining a cutout into the boundary structure proximate the opening. A locking feature is inserted into the cutout. When the airfoil is inserted into the opening, the locking feature interlocks the airfoil with the boundary structure. To seal the airfoil with the boundary structure, the airfoil is pressed against or into the boundary structure. When the airfoil is pressed, the locking feature is compressed such that a seal is formed between the airfoil and the boundary structure to seal the flow path assembly. 1. A method for assembling an airfoil with a boundary structure , at least one of the airfoil and the boundary structure being formed from a composite material , the method comprising:inserting the airfoil defining a cutout into an opening defined by the boundary structure, the boundary structure defining a cutout, wherein when the airfoil is inserted into the opening, a locking ring is received within the cutout defined by the boundary structure and the cutout defined by the airfoil; andpressing the airfoil against the boundary structure such that the locking ring forms a seal between the airfoil and the boundary structure.2. The method of claim 1 , wherein the airfoil is formed of the composite material claim 1 , and wherein when the airfoil is inserted into the opening and pressed against the boundary structure claim 1 , the airfoil is in the green state.3. The method of claim 1 , wherein prior to inserting the airfoil into the opening claim 1 , the method further comprises:machining the opening into the boundary structure;machining the cutout into the airfoil; andmachining the ...

METHODS OF MANUFACTURING A TANDEM GUIDE VANE SEGMENT

Methods for manufacturing a tandem guide vane segment that includes an outer platform, a front guide vane, and a rear guide vane, wherein the front guide vane and the rear guide vane are arranged in a firmly fixated manner with respect to one another. One method includes manufacturing an integral front segment section that includes the front guide vane and a front section of the outer platform, manufacturing an integral rear segment section that includes the rear guide vane and a rear section of the outer platform, and connecting the two segment sections to each other. 1. A method for manufacturing a tandem guide vane segment that comprises an outer platform , a front guide vane , and a rear guide vane , wherein the front guide vane and the rear guide vane are arranged in a firmly fixated manner with respect to one another and form a gap in between them , wherein the method comprises:manufacturing an integral intermediate structure that has material areas that comprise the final outer platform, the final front guide vane, and the final rear guide vane, wherein, in the intermediate structure, the gap between the final front guide vane and the final rear guide vane is filled with material of the intermediate structure,further processing of the intermediate structure into the final tandem guide vane segment, wherein the further processing comprises manufacturing a gap in the intermediate structure for separating the front and the rear guide vane.2. The method according to claim 1 , wherein the manufacturing of the integral intermediate structure is performed by means of forging or casting.3. The method according to claim 1 , wherein the forming of the gap is performed by means of a milling or a cutting method.4. The method according to claim 1 , wherein the tandem guide vane segment to be manufactured additionally comprises an inner platform claim 1 , wherein the integral intermediate structure also comprises material areas that comprise the final inner platform.523-. ...

Integral half vane, ringcase, and id shroud

A vane stage includes a ringcase extending circumferentially about a center axis of the vane stage. The ringcase extends completely about the center axis to form a first ring. An inner shroud extends circumferentially about the center axis of the vane stage. The inner shroud extends completely about the center axis to form a second ring positioned radially within the ringcase relative the center axis. A plurality of stationary half vanes extend radially between the ringcase and the inner shroud, and are circumferentially spaced about the center axis. The plurality of stationary half vanes are integral with the ringcase and the inner shroud.

Turbine vane assembly with reinforced end wall joints

The present disclosure is related to turbine vane assemblies comprising ceramic matrix composite materials. The turbine vane assemblies further including reinforcements that strengthen joints in the turbine vane assemblies.

Jet pump system and method with improved efficency

The present disclosure is of a jet pump system, and reverse power generation system and other desirable applications consisting of an impeller with inlet vortex vanes and outlet vortex vanes. The inlet vortex vane induces rotational movement on mass entering the impeller inlet. The outlet vortex vanes remove swirl from mass exiting the impeller outlet. Embodiments include a jet pump system involving a pulley and belt which can allow for obstruction free movement of mass. In another embodiment the impeller is connected via an electromagnetic connection. In another embodiment the impeller acts as a rim-driven generator of electrical power. In another embodiment the drive pulley is a centrifugal clutch or uses a chain sprocket or tandem jet system in series.

STATOR ASSEMBLY FOR A GAS TURBINE ENGINE

A stator assembly comprises an annular platform () defining a boundary of an annulus and a plurality of guide vanes () for arranging in a circumferential array on the annular platform (). The platform () includes a circumferential array of slots each slot having, at a first end, walls converging () from an annulus facing side towards an opposite side of the platform and then extending parallel () towards a second end. One or more cooling holes () is arranged in a wall () of the slot. Each vane () includes a root portion () which converges in a first region () distal to the root end and then extends parallel towards the root end in a second region (). The root portion () is configured to engage in a slot of the platform () with the first region () abutting a convergent wall portion () of the slot and the second region () spaced from a parallel wall () of the slot. At least one bore () is provided in a parallel wall () of each slot and at least one bore () is also provided in the second region of the root portion (). The bores () in the two components are positioned to align when a vane () is engaged in a slot. The aligning bores () together are configured to receive a pin () for mechanically securing the vane in the slot. A braze joint () may be added between the root portion () and parallel slot walls 1. A stator assembly comprising at least one annular platform defining a boundary of an annulus and a plurality of guide vanes for arranging in a circumferential array on the annular platform;the platform including a circumferential array of slots each slot at a first end having walls converging from an annulus facing side to an opposite side of the platform and then extending parallel towards a second end;one or more cooling holes arranged in a wall of the slot;each vane including a root portion which converges in a first region distal to the root end and then extends parallel towards the root end in a second region, the root portion configured to engage in a slot of ...

WELDING METHOD AND SYSTEM

A welding method including the steps of: providing, in a first position, a weld target assembly having a welding region and a weld tab located adjacent the welding region, the weld tab changeable from a first physical configuration to a second physical configuration when subject to a first type of weld operation; performing a first weld process on the welding region and weld tab; moving the assembly from the first to a second position; providing safeguard apparatus for cooperating with the weld tab to: prevent the assembly from moving to the second position when the first weld procedure did not include the first type of weld operation; and permit the assembly be moved to the second position when the first weld procedure included the first type of weld operation. The assembly includes a sub-assembly having a foot member and blade portion welded together, thereby ultimately forming an output guide vane. 1. A welding method including the steps of:providing, in a first position, a weld target assembly having a welding region and a weld tab located adjacent the welding region, the weld tab being changeable from a first physical configuration to a second physical configuration when subject to a first type of weld operation;performing a first weld process on the welding region and on the weld tab;moving, after performing the first weld process, the assembly from the first position to a second position; (i) to prevent the assembly from being moved to the second position when the first weld procedure did not include the first type of weld operation; and', '(ii) to permit the assembly to be moved to the second position when the first weld procedure did include the first type of weld operation., 'providing a safeguard apparatus for cooperating with the weld tab2. A welding method according to further including the step of performing a second weld process claim 1 , including a second type of weld operation claim 1 , on another welding region of the weld target assembly when the ...

ADDITIVELY MANUFACTURED CONNECTION FOR A TURBINE NOZZLE

Turbine nozzles are provided for gas turbine engines. The turbine nozzle includes an arcuate inner band; an arcuate outer band; and a nozzle vane disposed between the arcuate inner band and the arcuate outer band. The radially inner end of the nozzle vane is attached to the arcuate inner band through an interlocking transition zone comprising a plurality of projections alternately extending from the radially inner end of the nozzle vane and the arcuate inner band, respectively, to undetachably couple the nozzle vane and the arcuate inner band. Optionally, the radially outer end of each nozzle vane is also attached to the arcuate outer band through an interlocking transition zone. 1. A turbine nozzle for a gas turbine engine , comprising:an arcuate inner band;an arcuate outer band; anda nozzle vane disposed between the arcuate inner band and the arcuate outer band, wherein a radially inner end of the nozzle vane is attached to the arcuate inner band through an interlocking transition zone comprising a plurality of projections alternately extending from the radially inner end of the nozzle vane and the arcuate inner band, respectively, to undetachably couple the nozzle vane and the arcuate inner band.2. The turbine nozzle of claim 1 , wherein the interlocking transition zone allows for variations in the relative motion between the radially inner end of the nozzle vane and the arcuate inner band.3. The turbine nozzle of claim 1 , wherein the interlocking transition zone defines a fluid channel claim 1 , the fluid channel being configured to provide fluid communication between the radially inner end of the nozzle vane and the arcuate inner band.4. The turbine nozzle of claim 1 , wherein the nozzle vane is constructed from a first metal alloy and the arcuate inner band is constructed from a second metal alloy that is different than the first metal alloy.5. The turbine nozzle of claim 1 , wherein a platform segment is embedded within an radially outer surface of the ...

Gas turbine engine ceramic component assembly attachment

A gas turbine engine component assembly includes first and second portions, wherein at least one of the first and second portions is a ceramic material. The first portion includes an aperture having a first angled surface. The second portion is disposed within the aperture and includes a second angled surface adjacent to the first angled surface. The first and second angled surfaces lock the first and second portions to one another under a pulling load. A bonding material operatively secures the first and second angled surfaces to one another.

METHOD FOR ASSEMBLING TURBOMACHINE PARTS AND ASSEMBLY USED DURING SUCH A METHOD

A method for assembly of a first turbomachine part with at least one second turbomachine part, including an injection of a vulcanisable elastomer, preferably a silicone that can be vulcanised at ambient temperature called an RTV silicone, in an injection zone at the junction between the first and the second parts; local heating of the injection zone so as to vulcanise the vulcanisable elastomer. 110-. (canceled)11. A method for assembly of a first turbomachine part with at least one second turbomachine part , comprising:injecting a vulcanisable elastomer in an injection zone at the junction between the first and the second parts, local heating of the injection zone so as to vulcanise the vulcanisable elastomer.12. The assembly method according to claim 11 , in which the first and the second parts are two annular parts configured to form a turbomachine guide vane after assembly.13. The method according to claim 11 , in which the local heating step is implemented using an assembly comprising a support for the first and the second parts and a local heating system comprising a heating zone and that is associated with the support such that the heating zone is facing the injection zone while the heating step is being performed.14. The method according to claim 13 , in which the assembly also acts as an assembly support used during the elastomer injecting.15. The method according to claim 11 , in which the local heating step is implemented with a heating element previously placed on the two parts of the turbomachine close to the injection zone.16. The method according to claim 11 , in which the local heating is implemented using a remote heating system.17. The method according to claim 16 , in which the remote heating system is selected among the group consisting of a hot air blowing system at the injection zone claim 16 , laser radiation and microwave radiation.18. The method according to claim 11 , in which the vulcanisable elastomer is a silicone that can be vulcanised ...

Method for manufacturing nozzle diaphragm and nozzle diaphragm

There is provided a method for manufacturing a nozzle diaphragm which can be assembled with high accuracy. The method for manufacturing a nozzle diaphragm includes a preparation step S 1 of preparing an inner ring, a plurality of nozzles, an outer shroud ring, and an outer ring, a ring installation step S 2 of installing the outer shroud ring, a nozzle installation step S 3 of installing the nozzles while an outer peripheral end portion of a nozzle main body is inserted into a through-hole formed in the outer shroud ring, an inner ring installation step S 4 of installing the inner ring, an outer ring installation step S 5 of installing the outer ring outside the outer shroud ring, and a welding step S 6 of welding the outer ring and the outer shroud ring to each other, and welding the inner ring and the inner shroud to each other in a circumferential direction.

COMPRESSOR AIRFOIL

Aspects of the present invention relate to systems and methods of a vane design utilizing welding techniques. The present invention concerns a method for preventing cracking within a vane assembly utilizing full penetration welding. Additional embodiments concern a vane design that, when assembled with another vane, comprises an axial slot that prevents cracking within a vane assembly. 1. A vane assembly for use in a welded vane assembly comprising:an inner shroud having a first inner sidewall, an opposing second inner sidewall, an inner radial thickness, an inner length, and an inner gas path surface, the first inner sidewall having a first inner groove and a second inner sidewall having a second inner groove;an airfoil extending radially outward from the inner gas path surface; andan outer shroud coupled to the airfoil and having a first outer sidewall, an opposing second outer sidewall, an outer radial thickness, an outer length, and has an outer gas path surface, the first outer sidewall having a first outer groove and a second outer sidewall having a second outer groove;wherein the first and second inner grooves and first and second outer grooves are positioned a distance along the inner and outer radial thicknesses, respectively.2. The vane of claim 1 , wherein the first and second inner grooves extend the inner length of the inner shroud.3. The vane of claim 1 , wherein the first and second outer grooves extend the outer length of the outer shroud.4. The vane of claim 1 , wherein the inner shroud claim 1 , airfoil and outer shroud are integrally fabricated together.5. The vane of claim 1 , wherein the outer gas path surface tapers at an angle.6. The vane of claim 1 , wherein the first and second inner grooves and first and second outer grooves are generally C-shaped.7. The vane of claim 1 , wherein the first and second inner grooves are positioned proximate a midpoint of the inner radial thickness.8. The vane of claim 1 , wherein the first and second outer ...

ARRANGEMENT FOR SECURING TURBINE BLADES

An arrangement for securing turbine blades in a groove of an inner housing of a steam turbine is provided. A caulking piece which is made of bi-metal is used between the base of the turbine guide blade and the groove. The component is designed as a casing, in particular as an inner casing of a steam turbine. 1. An arrangement for securing turbine blades , wherein the turbine blades have a turbine blade rootwhich is fitted into a slot in a component, wherein a caulking piece is arranged between the turbine blade root and the component,wherein the turbine blade is designed as a turbine guide vane,whereinthe caulking piece is a bimetal.2. The arrangement as claimed in claim 1 , wherein the component is an inner casing of a steam turbine. This application claims priority to PCT Application No. PCT/EP2014/068354, having a filing date of Aug. 29, 2014, based off of EP Application No. 13185739.3 having a filing date of Sep. 24, 2013, the entire contents of which are hereby incorporated by reference.The following relates to an arrangement for securing turbine blades, wherein the turbine blades have a turbine blade root which is fitted into a slot in a component, wherein a caulking piece is arranged between the turbine blade root and the component.Steam turbines, as an embodiment of a turbomachine, comprise a relatively large number of turbine blades which are divided into rotor blades and guide vanes. Rotor blades are arranged on the rotor and guide vanes are arranged in the casing. The guide vanes have a blade root which is installed in a corresponding slot in the casing. Installation is generally carried out by hand, which means that a technician inserts the individual steam turbine guide vanes into the corresponding slots by hand with a suitable force. In addition, a caulking piece is caulked between the blade root and the slot in order to exert an immobilizing force. Some guide vanes have shrouds which are twisted with respect to the root.Documents DE 32 36 021 A, EP 2 ...

Static blade for a turbine diaphragm and associated turbine diaphragm

Static blade for an axial flow turbine comprising an aerofoil portion having a leading edge, a trailing edge, a pressure side and a suction side and radially inner and outer reinforcement portions integral with said aerofoil portion. Each reinforcement portion closely follows the shape of the section of the aerofoil portion.

Stator Blades of an Axial Turbocompressor and Manufacturing Process

The present application relates to a stator of an axial turbomachine compressor having an external shell with a row of openings and an internal annular groove designed to hold an annular layer of abradable material forming a strip. The stator further includes a row of stator blades adjacent to the row of openings. The stator blades include platforms with surfaces matching the openings, the said platforms being fitted in the openings so as to mask them. Weld beads fix the platforms at their junctions with the openings. On one side of the row of stator blades a portion of the weld bead is situated axially in the internal annular groove. The present application also relates to methods of manufacturing the stator.

Geared gas turbine engine

A geared gas turbine engine comprises at least one compressor, at least one turbine, a fan, a gearbox and a support structure. At least one turbine is arranged to drive the fan via the gearbox. The fan is rotatably mounted in the support structure and the support structure supports the gearbox. The support structure includes a stator vane arrangement comprising a plurality of circumferentially spaced stator vanes extending radially between inner and outer annular walls. The inner annular wall has two axially spaced radially inwardly extending annular flanges and a plurality of circumferentially spaced buttresses extending axially between the annular flanges. The outer annular wall has two axially spaced radially outwardly extending annular flanges and a plurality of circumferentially buttresses extending axially between the annular flanges. The flanges and buttresses increase the strength of the support structure to carry the torque loads of the gearbox.

Method of Fabricating a Turbine or Compressor Guide Vane Sector Made of Composite Material for a Turbine Engine, and a Turbine or a Compressor Incorporating Such Guide Vane Sectors

Single-airfoil vanes each having an inner platform, an outer platform, and an airfoil are obtained by three-dimensionally weaving a fiber blank in a single piece, by shaping the fiber blank to obtain a single-piece fiber preform, and by densifying the preform with a matrix to obtain a vane of composite material forming a single piece with inner and outer platforms incorporated therein. A plurality of vanes is assembled together at an intermediate stage of densification to form a multi-airfoil composite material guide vane sector for a turbine nozzle or a compressor diffuser and the assembled-together vanes are bonded together. 118-. (canceled)19. A method of fabricating a turbine nozzle or a compressor diffuser for a turbine engine , the method comprising: forming a fiber blank by three-dimensional weaving, the blank being in the form of a strip and comprising a first segment with second and third segments extending the first segment at respective first and second longitudinal ends thereof, each of the second and third segments being split into two portions on either side of a zone of non-interlinking extending within the thickness and across the entire width of the strip;', 'forming a fiber preform for the vane to be made by laterally deploying the two portions of the second segment and the two portions of the third segment and shaping said portions so as to obtain inner and outer platform preforms, and by shaping the first segment so as to obtain an airfoil preform; and', 'at least partially densifying the fiber preform with a matrix in order to obtain and at least partially densified vane with inner an outer platforms incorporated therein; and, 'a) making a plurality of single-airfoil vane units, each vane having inner and outer platforms and an airfoil extending between the platforms and connected thereto, and being made by the bonding being performed by a process comprising at least one step selected from: a step of bonding by brazing and a step of bonding by ...

VANE ASSEMBLIES FOR GAS TURBINE ENGINES

A vane ring for a gas turbine engine includes an outer end wall and a plurality of spars coupled to the outer end wall. The vane ring further includes an inner end wall positioned radially inward of the outer end wall and coupled to the spars. The outer and inner end walls cooperate to form a flowpath. 1. A vane ring for use in a gas turbine engine , the vane ring comprisinga plurality of metal spars, each spar including a web section having an airfoil shape, a first end connector coupled to a radially outer portion of the web section, and a second end connector coupled to a radially inner portion of the web section,a plurality of inner end wall segments positioned to function as a continuous hoop and engage the metal spars, each inner end wall segment including a first flow surface positioned to guide expanding hot gases along a flow path through the gas turbine engine and at least one locator hole sized to receive the second end connector of the metal spars, anda unitary outer end wall forming a one-piece continuous hoop, the outer end wall comprising ceramic-matrix materials and including a second flow surface positioned to cooperate with the first flow surface of the inner end wall segments to form the flow path and a plurality of locator holes sized to receive the first end wall connectors to locate the metal spars circumferentially along the outer end wall.2. The vane ring of claim 1 , further comprising a plurality of outer end caps coupled to the first end connectors and positioned to engage a radially outer surface of the outer end wall.3. The vane ring of claim 2 , wherein the outer end caps are coupled to the first end connectors by one of welding claim 2 , brazing claim 2 , a bi-cast joint claim 2 , or a fastener.4. The vane ring of claim 2 , further comprising a plurality of inner end caps coupled to the second end connectors and positioned to engage a radially inner surface of the inner end wall segments claim 2 , the outer end caps and inner end caps ...

NOZZLE ASSEMBLY AND METHOD FOR FORMING NOZZLE ASSEMBLY

A nozzle assembly is disclosed, including a CMC nozzle shell, a nozzle spar, and an endwall. The CMC nozzle shell includes a CMC composition and an interior cavity. The nozzle spar is partially disposed within the interior cavity and includes a metallic composition, a cross-sectional conformation, a plurality of spacers protruding from the cross-sectional conformation, the plurality of spacers contacting the CMC nozzle shell, and a spar cap. The endwall includes at least one surface in lateral contact with the spar cap and maintains a lateral orientation of the CMC nozzle shell and the nozzle spar relative to the endwall. The lateral orientation maintains a predetermined throat area of the nozzle assembly. A method for forming the nozzle assembly includes inserting the nozzle spar into the interior cavity, rotating the CMC nozzle shell and the nozzle spar laterally relative to the endwall, and maintaining the lateral orientation. 1. A nozzle assembly , comprising: a CMC composition; and', 'an interior cavity having interior dimensions;, 'a ceramic matrix composite (CMC) nozzle shell, the CMC nozzle shell including a metallic composition;', 'a cross-sectional conformation including cross-sectional dimensions less than the interior dimensions;', 'a plurality of spacers protruding from the cross-sectional conformation, the plurality of spacers contacting the CMC nozzle shell; and', 'a spar cap; and, 'a nozzle spar partially disposed within the interior cavity, includingan endwall including at least one surface in lateral contact with the spar cap, the endwall maintaining a lateral orientation of the CMC nozzle shell and the nozzle spar relative to the endwall, the lateral orientation maintaining a predetermined throat area of the nozzle assembly.2. The nozzle assembly of claim 1 , wherein the endwall includes a first stanchion and a second stanchion extending from the endwall claim 1 , the at least one surface in lateral contact with the spar cap including a first ...

REPAIR MEMBER FOR A VANE ASSEMBLY OF A GAS TURBINE AND METHOD FOR REPAIRING A DAMAGED VANE OF A VANE ASSEMBLY OF A GAS TURBINE

The repair member includes an inner platform, an outer platform portion, and an airfoil portion connecting the inner and outer platform portions; the inner platform portion is configured so to reach an edge of a first ring or sector of ring of a vane assembly of the damaged vane, and the outer platform portion is configured so to reach an edge of a second ring or sector of ring of the a vane assembly of the damaged vane; the inner platform portion, the outer platform portion, and the airfoil portion are configured so to allow insertion of the repair member into the vane assembly by a pure translation movement being along an insertion direction and having a movement component parallel to an axes of the first and second rings or sectors of ring. 1. A repair member for a vane assembly of a gas turbine , wherein the vane assembly comprises an inner platform in the form of a first ring or sector of ring , an outer platform in the form of a second ring or sector of ring , and a plurality of vanes each consisting of a leading edge portion , a trailing edge portion and a body portion , the repair member being arranged for repairing a damaged vane of the plurality of vanes and comprising:an inner platform portion designed to replace part of the inner platform of the damaged vane,an outer platform portion designed to replace part of the outer platform of the damaged vane, andan airfoil portion connecting the inner and outer platform portions, designed to replace a leading edge portion or a trailing edge portion of the damaged vane;wherein the inner platform portion is configured so to reach an edge of the first ring or sector of ring,wherein the outer platform portion is configured so to reach an edge of the second ring or sector of ring, andwherein the inner platform portion the outer platform portion, and the airfoil portion are configured so to allow insertion of the repair member into the vane assembly by a pure translation movement being along an insertion direction and ...

TURBINE VANE ASSEMBLY WITH REINFORCED END WALL JOINTS

The present disclosure is related to turbine vane assemblies comprising ceramic matrix composite materials. The turbine vane assemblies further including reinforcements that strengthen joints in the turbine vane assemblies. 1. A turbine vane adapted for use in an aerospace gas turbine engine , the turbine vane comprisingan airfoil comprising ceramic matrix composite materials having ceramic-containing fibers infiltrated with ceramic matrix, the airfoil shaped to redirect hot gasses moving along a primary gas path within the aerospace gas turbine engine,an end wall comprising ceramic matrix composite materials having ceramic-containing fibers co-infiltrated with ceramic matrix along with the ceramic-containing fibers of the airfoil, the end wall including a panel that extends circumferentially from the airfoil about a central axis to define a boundary of the primary gas path and a rim that extends radially from the panel outside the primary gas path along an outer surface of the airfoil, andreinforcements configured to interconnect the airfoil and the end wall and strengthen a joint therebetween, the reinforcements arranged to extend between the outer surface of the airfoil and the rim of the end wall.2. The turbine vane of claim 1 , wherein the reinforcements are provided by tufted fibers pushed from one of the airfoil and the rim of the end wall into the other of the airfoil and the rim of the end wall.3. The turbine vane of claim 1 , wherein the reinforcements are provided by rods that extend through an outer surface of the airfoil facing away from a central cavity of the airfoil and an inner surface of the rim that faces the airfoil.4. The turbine vane of claim 3 , wherein the rods comprise ceramic-containing material.5. The turbine vane of claim 3 , wherein the rods are co-infiltrated with ceramic matrix along with the airfoil and the end wall.6. The turbine vane of claim 3 , wherein the airfoil is shaped to include a leading edge claim 3 , a trailing edge claim ...

TURBINE COMPONENTS AND METHODS OF MANUFACTURING

At least one turbine component for a gas turbine includes a base component formed by casting and an article. The base component includes a platform. The article on the upper surface of the platform is formed by additive manufacturing. The article has a proximal face sized and shaped to cover at least a portion of the upper surface of the platform of the turbine component and a contoured distal face opposite the proximal face. The contoured distal face has a contour surface serving as at least a portion of a hot gas path surface of the turbine component. The contour surface is arranged and disposed to provide a controlled flow pattern of a working fluid across the contour surface based on a mounting location of the turbine component in a turbine. Methods of manufacturing articles and turbine components are also disclosed. 1. A method of manufacturing comprising:additive manufacturing an article having a proximal face sized and shaped to cover at least a portion of an upper surface of a platform of a base component and a contoured distal face opposite the proximal face, the base component and the article together forming a turbine component, the contoured distal face of the article having a contour surface serving as at least a portion of a hot gas path surface of the turbine component, the contour surface being arranged and disposed to provide a controlled flow pattern of a working fluid across the contour surface based on a mounting location of the turbine component in a turbine.2. The method of further comprising brazing the article to the platform of the base component.3. The method of claim 1 , wherein the base component is formed by casting claim 1 , and the additive manufacturing of the article occurs directly on the platform of the base component.4. The method of further comprising casting the base component.5. The method of claim 1 , wherein the article includes at least a portion of at least one wall cooling feature.6. The method of claim 5 , wherein the ...

Methods and Features for Positioning a Flow Path Assembly within a Gas Turbine Engine

Flow path assemblies having features for positioning the assemblies within a gas turbine engine are provided. For example, a flow path assembly comprises an inner wall and a unitary outer wall that includes an integral combustion portion and turbine portion, the combustor portion extending through a combustion section of the gas turbine engine and the turbine portion extending through at least a first turbine stage of a turbine section of the gas turbine engine. The flow path assembly further comprises at least two positioning members for radially centering the flow path assembly within the gas turbine engine. The positioning members extend to the flow path assembly from one or more structures external to the flow path assembly, constrain the flow path assembly tangentially, and allow radial and axial movement of the flow path assembly. Other embodiments for positioning flow path assemblies also are provided. 1. A flow path assembly for a gas turbine engine , the flow path assembly comprising:an inner wall;a unitary outer wall including a combustor portion extending through a combustion section of the gas turbine engine and a turbine portion extending through at least a first turbine stage of a turbine section of the gas turbine engine, the combustor portion and the turbine portion being integrally formed as a single unitary structure; andat least two positioning members for radially centering the flow path assembly within the gas turbine engine, the positioning members extending to the flow path assembly from one or more structures external to the flow path assembly,wherein the positioning members constrain the flow path assembly tangentially, andwherein the positioning members allow radial and axial movement of the flow path assembly.2. The flow path assembly of claim 1 , further comprising a plurality of slots defined at an inner forward end of the flow path assembly claim 1 , wherein a plurality of positioning members extends axially into the slots claim 1 , ...

Methods and Assemblies for Attaching Airfoils within a Flow Path

Flow path assemblies for gas turbine engines are provided. For example, a flow path assembly comprises an inner wall; a unitary outer wall; and a plurality of nozzle airfoils having an inner end radially opposite an outer end. The unitary outer wall defines a plurality of outer pockets each configured for receipt of the outer end of one of the nozzle airfoils, and the inner wall includes defines a plurality of inner pockets each configured for receipt of the inner end of one of the plurality of nozzle airfoils. A portion of each inner pocket is defined by a forward inner wall segment and an aft inner wall segment. In another embodiment, a flow path assembly comprises an inner wall defining a plurality of bayonet slots that each receive a bayonet included with each of a plurality of nozzle airfoils that are integral with a unitary outer wall. 1. A flow path assembly for a gas turbine engine , the flow path assembly comprising:an inner wall;a unitary outer wall including a combustor portion extending through a combustion section of the gas turbine engine and a turbine portion extending through at least a first turbine stage of a turbine section of the gas turbine engine, the combustor portion and the turbine portion being integrally formed as a single unitary structure; anda plurality of nozzle airfoils, each nozzle airfoil having an inner end radially opposite an outer end,wherein the inner wall and the unitary outer wall define a combustor of the combustion section,wherein the unitary outer wall defines a plurality of outer pockets, each outer pocket configured for receipt of the outer end of one of the plurality of nozzle airfoils, andwherein the inner wall includes a forward segment and an aft segment, the inner wall defining a plurality of inner pockets such that a portion of each inner pocket is defined by the forward segment and a remaining portion of each inner pocket is defined by the aft segment, each inner pocket configured for receipt of the inner end of ...

Method of sealing cooling holes

A method of sealing a gap between an aerofoil component and a further component. The method comprises placing the aerofoil component in close proximity with the further component to define a gap therebetween, applying a thermoplastic material to the gap in a molten phase and cooling the thermoplastic material to set the thermoplastic material.

GROUP OF BLADE ROWS

A blade row group arrangeable in a main flow path of a fluid-flow machine and including N adjacent member blade rows firmly arranged relative to one another is provided. Here, a front member blade row with front blades as well as a rear member blade row with rear blades are provided in the meridional plane established by the axial direction and the radial direction, and the blade row group has two main flow path boundaries. It is provided that the profile of the blades of the member blade rows is firmly connected at at least one of the two main flow path boundaries to a blade root structure, where the blade root structure of the blades (i) of the front member blade row has at least one holding structure, and/or the blade root structure of the blades of the rear member blade row has at least one holding structure. 1. A blade row group arrangeable in a main flow path of a fluid-flow machine and including N adjacent member blade rows firmly arranged relative to one another in both the axial direction (x) and the circumferential direction , with the number N of the member blade rows being greater than/equal to 2 and (i) designating the running index with values between 1 and N , where a front member blade row with front blades (i) as well as a rear member blade row with rear blades (i+1) are provided in the meridional plane established by the axial direction (x) and the radial direction (r) , where the blade row group has two main flow path boundaries (HB) , whereinthe profile of the blades (i, i+1) of the member blade rows is firmly connected at at least one of the two main flow path boundaries (HB) to a blade root structure (F(i), F(i+1)), where the blade root structure (F(i)) of the blades (i) of the front member blade row has at least one holding structure (VF(i), HF(i)), and/or the blade root structure (F(i+1)) of the blades (i+1) of the rear member blade row has at least one holding structure (VF(i+1), HF(i+1)), and at least one of the holding structures (VF(i), ...

GROUP OF BLADE ROWS

A blade row group arrangeable in a main flow path of a fluid-flow machine and including N adjacent member blade rows firmly arranged relative to one another in both the axial direction and the circumferential direction is provided. Here, a front member blade row with front blades as well as a rear member blade row with rear blades are provided in the meridional plane established by the axial direction and the radial direction. The blade row group has two main flow path boundaries. It is provided that the blades of the member blade rows are fixed to the one main flow path boundary in the surrounding structure by means of a blade root structure, and that the blades of the member blade rows on the other main flow path boundary are each firmly connected to a base. 2. The blade row group in accordance with claim 1 , wherein the side of the base (F(i)) of the front member blade row facing downstream and the side of the base (F(i+1)) of the rear member blade row facing upstream adjoin each other along an unstepped contact surface (KF).3. The blade row group in accordance with claim 1 , wherein the side of the base (F(i)) of the front member blade row facing downstream and the side of the base (F(i+1)) of the rear member blade row facing upstream adjoin each other along a contact surface (KF) provided with a step extending in the radial direction claim 1 , so that a mutual seating between the bases (F(i) claim 1 , F(i+1)) of the member blade rows is assured.4. The blade row group in accordance with claim 1 , wherein a modular design of the bases (F(i) claim 1 , F(i+1)) is provided claim 1 , with the support (IT) being connected to the bases (F(i) claim 1 , F(i+1)) of the front and rear member blade rows claim 1 , thus securing the cohesion of the two bases (F(i) claim 1 , F(i+1)) claim 1 , with the support (IT) having a holding structure anchored in each base of the member blade rows claim 1 , i.e. a front holding structure (VHF) anchored in the base of the front member ...

Methods and Features for Positioning a Flow Path Assembly Within a Gas Turbine Engine

Flow path assemblies having features for positioning the assemblies within a gas turbine engine are provided. For example, a flow path assembly comprises an inner wall and a unitary outer wall that includes an integral combustion portion and turbine portion, the combustor portion extending through a combustion section of the gas turbine engine and the turbine portion extending through at least a first turbine stage of a turbine section of the gas turbine engine. The flow path assembly further comprises at least two positioning members for radially centering the flow path assembly within the gas turbine engine. The positioning members extend to the flow path assembly from one or more structures external to the flow path assembly, constrain the flow path assembly tangentially, and allow radial and axial movement of the flow path assembly. Other embodiments for positioning flow path assemblies also are provided.

Method and apparatus for maching workpiece

An apparatus includes an electrode assembly comprising a carriage having a plurality of electrode holders, the electrode holders being respectively configured to detachably receive a plurality of electrodes, the electrodes include a plurality of first electrodes and a plurality of second electrodes. The first electrodes are configured for rough machining a workpiece by electric discharging or wire electric discharging to remove material from the workpiece, the second electrodes are configured for finish machining the rough machined workpiece by electric discharging to remove material from the rough machined workpiece.

Integral Bent Housing for an Axial Turbomachine Compressor

The present application relates to a method for producing a housing for an axial turbomachine compressor having a metal sheet with annular rows of stator blades which are welded to the metal sheet. The method includes the steps of: (a) providing or producing a planar metal sheet with a step of machining in order to form blade stumps; (b) welding stator blades by means of friction to one of the planar faces of the metal sheet, the blades being arranged so as to form parallel straight rows of blades; (c) bending the metal sheet about a bending axis perpendicular to each row of blades, so as to form a half-tube with annular half-rows of blades which are axially spaced apart, and producing annular grooves by means of rolling; (d) welding annular flanges and axial flanges; (e) application of annular layers of abradable material. 1. A method for producing a housing of an axial turbomachine , the housing having a metal sheet and at least one annular row of stator blades , the method comprising the steps of:(a) providing a planar metal sheet;(b) welding stator blades to one of the planar faces of the metal sheet, the blades being arranged so as to form at least one row of blades; and(c) bending the metal sheet about a bending axis perpendicular to the row of blades;wherein the bending involves wrapping the blades inside the metal sheet, so as to form a housing with at least one angular tube portion as a result of the metal sheet, and with at least one angular portion of an annular row of blades.2. The method according to claim 1 , wherein during step (b) welding the blades claim 1 , the blades are welded by means of friction claim 1 , in accordance with a movement in the plane of the planar metal sheet.3. The method according to claim 1 , wherein following step (c) bending the metal sheet claim 1 , the metal sheet forms a tube with at least one annular row of blades.4. The method according to claim 1 , wherein the metal sheet forms an angular tube portion claim 1 , such as ...

ORC Turbine and Generator, And Method Of Making A Turbine

A turbine and a turbine-generator device for use in electricity generation. The turbine has a universal design and so may be relatively easily modified for use in connection with generators having a rated power output in the range of 50 KW to 5 MW. Such modifications are achieved, in part, through use of a modular turbine cartridge built up of discrete rotor and stator plates sized for the desired application with turbine brush seals chosen to accommodate radial rotor movements from the supported generator. The cartridge may be installed and removed from the turbine relatively easily for maintenance or rebuilding. The rotor housing is designed to be relatively easily machined to dimensions that meet desired operating parameters.

A METHOD OF FABRICATING AN AIRFOIL PREFORM, AN AIRFOIL, AND A NOZZLE SECTOR, BY SELECTIVE MELTING ON A BED OF POWDER

A method of fabricating an airfoil preform for a turbine engine by selective melting on a bed of powder, the preform including an airfoil and a removable support secured to the airfoil, the airfoil being fabricated layer by layer from a first edge to a second edge of the airfoil, the method including fabricating the removable support and the airfoil, the removable support being for securing to a fabrication platform and to a portion of a face of the airfoil situated near the first edge and facing the fabrication platform. The face of the airfoil facing the fabrication platform includes a flat extending away from the face, the flat being present over a portion of the face that is situated outside the first edge, the support being secured to the flat or both to the flat and to the portion of the face that is situated outside the first edge. 1. A method of fabricating an airfoil preform for a turbine engine by selective melting on a bed of powder , the preform comprising an airfoil and at least one removable support secured to the airfoil , the airfoil being fabricated layer by layer from a first edge of the airfoil corresponding to a leading edge or to a trailing edge of an airfoil to a second edge of the airfoil corresponding to a trailing edge or a leading edge of the airfoil , the method comprising fabricating the removable support and the airfoil , said removable support being for securing firstly to a fabrication platform and secondly to a portion of a face forming a pressure side or a suction side of the airfoil situated in the vicinity of the first edge of the airfoil and facing said fabrication platform;wherein the face forming a pressure side or a suction side of the airfoil and facing the fabrication platform includes a flat extending away from said face, the flat being present over a portion of said face that is situated outside the first edge of the airfoil, the removable support being secured to the flat or both to the flat and to the portion of said face ...

METHOD OF MANUFACTURING A BLADED STATOR ELEMENT FOR A TURBOMACHINE AND TOOL FOR CARRYING IT OUT

Methods of manufacturing a bladed stator element for a turbomachine include mounting a tool on a circumferential zone of an annular shell prior to welding vanes in the circumferential zone, welding radially outer ends of the vanes to the annular shell, dismantling the tool after welding the vanes in the circumferential zone, and repeatedly mounting and dismounting the tool on different circumferential zones of the annular shell so as to fix the vanes all around the annular shell. 1. A method of manufacturing a bladed stator element for a turbomachine , the bladed stator element having an annular shape around an axis and comprising an annular shell extending around the axis and comprising an annular fixing flange extending radially outwardly at each of its axial ends , and an annular row of vanes extending around the axis inside the annular shell , a radially outer end of each vane being fixed to the annular shell , the method comprising:welding the radially outer ends of the vanes to the annular shell;i) prior to welding one or more of the vanes in a circumferential zone (Z) of the annular shell, mounting a tool on the circumferential zone of the annular shell, between the annular fixing flanges of the annular shell, the tool being clamped axially against the annular fixing flanges so as to exert tensile forces in opposite axial directions on the annular fixing flanges; and(ii) after welding the one or more of the vanes, dismantling the tool,wherein the steps i) and ii) are repeated on different circumferential zones (Zi) of the annular shell so as to fix the vanes all around the annular shell.2. The method according to claim 1 , wherein in each step i) a first plate of the tool is applied against a radial face of one of the annular fixing flanges claim 1 , a second plate of the tool is applied against a radial face facing another of the annular fixing flanges claim 1 , and lengths of a plurality of cylinders connecting the first and second plates are increased to ...

TURBINE COMPONENTS AND METHODS OF MANUFACTURING

At least one turbine component for a gas turbine includes a base component formed by casting and an article. The base component includes a platform. The article on the upper surface of the platform is formed by additive manufacturing. The article has a proximal face sized and shaped to cover at least a portion of the upper surface of the platform of the turbine component and a contoured distal face opposite the proximal face. The contoured distal face has a contour surface serving as at least a portion of a hot gas path surface of the turbine component. The contour surface is arranged and disposed to provide a controlled flow pattern of a working fluid across the contour surface based on a clock mounting location of the turbine component in a turbine. 1. A turbine component for a gas turbine , the turbine component comprising:a base component formed by casting, the base component comprising a platform; andan article on an upper surface of the platform, the article being formed by additive manufacturing and having a proximal face sized and shaped to cover at least a portion of the upper surface of the platform of the turbine component and a contoured distal face opposite the proximal face, the contoured distal face having a contour surface serving as at least a portion of a hot gas path surface of the turbine component, the contour surface being arranged and disposed to provide a controlled flow pattern of a working fluid across the contour surface, the contour surface being customized based on a clock mounting location of the turbine component in a turbine.2. The turbine component of claim 1 , wherein the turbine component is one of a plurality of turbine components for the gas turbine claim 1 , each of the base components being cast from a single casting and the contoured distal face of each article being customized based on the flow pattern of the working fluid at the clock mounting location of the turbine component in the gas turbine.3. The turbine component of ...

TURBINE VANE ASSEMBLY WITH CERAMIC MATRIX COMPOSITE AIRFOIL AND FRICTION FIT METALLIC ATTACHMENT FEATURES

A turbine vane for a gas turbine engine incorporating a ceramic matrix composite airfoil is disclosed in this paper. The turbine vane includes an attachment unit configured to mount the ceramic matrix composite airfoil to other metallic components of the turbine vane. 1. A turbine vane comprising:an end wall comprising metallic materials and adapted to bound a primary gas path,an airfoil comprising ceramic matrix composite materials and aerodynamically shaped to redirect gasses that move along the primary gas path, the airfoil formed to include an external surface exposed to the primary gas path and an internal surface, opposite the external surface, that is shielded from the primary gas path, andan attachment unit configured to couple the airfoil to the end wall, the attachment unit shaped to include an airfoil-receiving channel that receives the airfoil and such that the attachment unit engages the external and internal surfaces of the airfoil with a friction fit so that the attachment unit is coupled via friction with the airfoil.2. The turbine vane of claim 1 , wherein the attachment unit includes an outer collar engaged with the exterior surface of the airfoil and an inner collar engaged with an interior surface of the airfoil.3. The turbine vane of claim 2 , wherein the outer collar includes a band that engages a portion of the exterior surface of the airfoil and a lip that engages along a radial end of the airfoil to radially locate the outer collar relative to the airfoil.4. The turbine vane of claim 3 , wherein the band and the lip of the outer collar form surfaces defining the airfoil receiving channel claim 3 , wherein the outer collar is coupled to the inner collar by a bond formed between the lip of the outer collar and the inner collar claim 3 , and wherein the bond is one of a diffusion weld bond and a braze bond.5. The turbine vane of claim 3 , wherein the inner collar includes a band that extends along a portion of the interior surface of the ...

Cmc airfoil joint

Joining an airfoil with a platform by mechanical keying can provide advantages in applications of ceramic materials, such as ceramic matrix composites.

FLOW PATH ASSEMBLIES FOR GAS TURBINE ENGINES AND ASSEMBLY METHODS THEREFORE

Flow path assemblies and methods for forming such flow path assemblies for gas turbine engines are provided. For example, a flow path assembly for a gas turbine engine has a boundary structure, an airfoil, and a locking feature. The boundary structure and the airfoil are formed from a composite material. The boundary structure defines an opening and a cutout proximate the opening, and the airfoil is sized to fit within the opening of the boundary structure. The locking feature is received within the cutout defined by the boundary structure to interlock the airfoil with the boundary structure. 1. A flow path assembly for a gas turbine engine , the flow path assembly comprising:a boundary structure formed from a composite material and defining an opening, the boundary structure further defining a cutout proximate the opening;an airfoil formed from a composite material and sized to fit within the opening of the boundary structure; anda locking feature received within the cutout defined by the boundary structure to interlock the airfoil with the boundary structure, the locking feature being located between the boundary structure and the airfoil.2. The flow path assembly of claim 1 , wherein the locking feature is a locking ring claim 1 , and wherein the opening defined by the boundary structure extends between an outer end and an inner end claim 1 , and wherein the locking ring extends toward the outer end of the opening and projects into the opening so as to define a recess between the locking ring and the boundary structure claim 1 , and wherein the airfoil comprises a locking portion that fills the recess to interlock the airfoil with the boundary structure.3. The flow path assembly of claim 2 , wherein the recess and the locking portion are V-shaped.4. The flow path assembly of claim 1 , wherein the locking feature is integrally formed with the airfoil.5. The flow path assembly of claim 4 , wherein the opening defined by the boundary structure extends between an ...

Turbine nozzle assembly

FIELD: machine building. SUBSTANCE: turbine nozzle assembly includes blade with internal and external side walls, external ring and flow divider with horizontal elongation. Boundary of transition between external ring and external side wall has enveloped/enveloping transition boundary or radial locking mechanism. Boundary of transition between horizontal elongation and internal side wall has radial locking mechanism or enveloped/enveloping transition boundary. Radial locking mechanism includes an enveloped step projecting along the axis out of the side wall and provided with a flange on its radial side in the form of the second enveloped step projecting along the axis to side wall. Enveloped/enveloping transition boundary includes radial enveloping groove on side wall and mating radial enveloped step. Where the enveloped/enveloping transition boundary is located on rear edge of side wall, it is connected by means of a butt weld located throughout the length of enveloped/enveloping transition boundary. Axial length of enveloped/enveloping transition boundary is approximately less than 1/4 of axial length of alignment between side wall and external ring or horizontal elongation. EFFECT: inventions allow reducing the distortion of steam flow trajectory in turbine and simplifying its assembly. 32 cl, 7 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) 2 465 467 (13) C2 (51) МПК F01D 25/24 (2006.01) F01D 3/02 (2006.01) F01D 9/02 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (21)(22) Заявка: 2007133831/06, 10.09.2007 (24) Дата начала отсчета срока действия патента: 10.09.2007 (72) Автор(ы): БЕРДЖИК Стивен С. (US), КРОЛЛ Томас В. (US) (73) Патентообладатель(и): ДЖЕНЕРАЛ ЭЛЕКТРИК КОМПАНИ (US) R U Приоритет(ы): (30) Конвенционный приоритет: 11.09.2006 US 11/518,708 (43) Дата публикации заявки: 20.03.2009 Бюл. № 8 2 4 6 5 4 6 7 (45) Опубликовано: 27.10.2012 Бюл. № 30 Адрес для переписки: 129090, Москва, ул.Б.Спасская, 25, стр.3, ООО " ...

Stator of gas-turbine axial-flow compressor

FIELD: mechanical engineering; gas turbines. SUBSTANCE: invention relates to stator of gas-turbine axial-flow compressor containing rigid outer ring equipment 2, rings 4a, 4b, 4c axially adjoining each other and arranged inside relative to said equipment 2 and carrying rims of fixed guide vanes 5. Said rings are formed by ring sectors 7 secured on equipment 2 whose inner wall limits aerodynamic channel passing compressed gaseous medium in outer direction differing in that ring sectors 7 are essentially soldered sectors formed by filler of cellular structure 8 enclosed between inner sheet 10 limiting said aerodynamic channel and outer sheet 9 and that tie with equipment 2 is provided only by means of outer sheet 9. EFFECT: improved reliability in operation. 5 cl, 2 dwg ÐÎÑÑÈÉÑÊÀß ÔÅÄÅÐÀÖÈß (19) RU (51) ÌÏÊ 7 (11) (13) 2 247 872 C1 F 04 D 29/54 ÔÅÄÅÐÀËÜÍÀß ÑËÓÆÁÀ ÏÎ ÈÍÒÅËËÅÊÒÓÀËÜÍÎÉ ÑÎÁÑÒÂÅÍÍÎÑÒÈ, ÏÀÒÅÍÒÀÌ È ÒÎÂÀÐÍÛÌ ÇÍÀÊÀÌ (12) ÎÏÈÑÀÍÈÅ ÈÇÎÁÐÅÒÅÍÈß Ê ÏÀÒÅÍÒÓ (21), (22) Çà âêà: 2003124062/06, 03.01.2002 (72) Àâòîð(û): ÊÀÐÎÍ Ñòåôàí (FR), ÄÅÁÅÍÅÉ Ïüåð (FR), ÃÅÐÓ Ôèëèïï (FR) (24) Äàòà íà÷àëà äåéñòâè ïàòåíòà: 03.01.2002 (30) Ïðèîðèòåò: 04.01.2001 FR 01/00060 (85) Äàòà ïåðåâîäà çà âêè PCT íà íàöèîíàëüíóþ ôàçó: 04.08.2003 (86) Çà âêà PCT: FR 02/00007 (03.01.2002) 2 2 4 7 8 7 2 R U Àäðåñ äë ïåðåïèñêè: 129010, Ìîñêâà, óë. Á.Ñïàññêà , 25, ñòð.3, ÎÎÎ "Þðèäè÷åñêà ôèðìà Ãîðîäèññêèé è Ïàðòíåðû", ïàò.ïîâ. Ã.Á. Åãîðîâîé (54) ÑÒÀÒÎÐ ÎÑÅÂÎÃÎ ÊÎÌÏÐÅÑÑÎÐÀ ÃÀÇÎÂÎÉ ÒÓÐÁÈÍÛ (57) Ðåôåðàò: íàðóæíîì íàïðàâëåíèè àýðîäèíàìè÷åñêèé êàíàë Èçîáðåòåíèå êàñàåòñ ñòàòîðà îñåâîãî äâèæåíè ñæàòîé ãàçîîáðàçíîé òåêó÷åé ñðåäû, êîìïðåññîðà ãàçîâîé òóðáèíû, ñîäåðæàùåãî îòëè÷àþùåãîñ òåì, ÷òî óïîì íóòûå êîëüöåâûå æåñòêóþ âíåøíþþ êîëüöåâóþ àðìàòóðó 2, ñåêòîðà 7 ïðåäñòàâë þò ñîáîé ïà íûå ñåêòîðà, ïðèìûêàþùèå äðóã ê äðóãó â îñåâîì íàïðàâëåíèè îáðàçîâàííûå çàïîëíèòåëåì ñîòîâîé ñòðóêòóðû 8, êîëüöà 4à, 4b, 4ñ, ðàñïîëàãàþùèåñ èçíóòðè ïî çàêëþ÷åííûì ìåæäó âíóòðåííèì ëèñòîì 10, îòíîøåíèþ ê óïîì íóòîé àðìàòóðå 2 è íåñóùèå ...

Method for producing parts of metal alloys of complex shape

FIELD: engine building. SUBSTANCE: invention relates to the field of production of solid blades of gas turbine engines, having a middle part, an edge/tip and a shank. A method includes a stage of producing a workpiece consisting of at least two parts, at least one of which is a solid part obtained by cutting a bar made of a metal alloy or ceramics-based material, while the specified at least two parts are assembled by diffusion connection without melting, and a stage of mechanical processing the resulting workpiece for the production of a blade with a set profile. EFFECT: use of the invention allows one to minimize the amount of waste material and reduce the cost of blade production. 13 cl, 2 tbl, 26 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 765 296 C2 (51) МПК B23P 15/04 (2006.01) F01D 5/12 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК B23P 15/04 (2021.08); F01D 5/12 (2021.08) (21)(22) Заявка: 2019128112, 12.03.2018 (24) Дата начала отсчета срока действия патента: (73) Патентообладатель(и): МЕКАХРОМ ФРАНС (FR) Дата регистрации: 28.01.2022 13.03.2017 FR 1752018 (43) Дата публикации заявки: 14.04.2021 Бюл. № 11 (56) Список документов, цитированных в отчете о поиске: US 3524712 A, 18.08.1970. RU 2380209 C1, 27.01.2010. FR 2981590 A1, 26.04.2013. WO 2014057208 A2, 17.04.2014. EP 1905956 A2, 02.04.2008. US 2015016972 A1, 15.01.2015. FR 2997885 А1, 16.05.2014. (45) Опубликовано: 28.01.2022 Бюл. № 4 (85) Дата начала рассмотрения заявки PCT на национальной фазе: 14.10.2019 2 7 6 5 2 9 6 Приоритет(ы): (30) Конвенционный приоритет: R U 12.03.2018 (72) Автор(ы): ДЕ ПОННАТ, Арнод (FR), МАРТИН, Оливье (FR) FR 2018/000051 (12.03.2018) C 2 C 2 (86) Заявка PCT: (87) Публикация заявки PCT: R U 2 7 6 5 2 9 6 WO 2018/167384 (20.09.2018) Адрес для переписки: 117342, Москва, ул. Миклухо-Маклая, 65, корп. 4, кв. 34, пат. пов. Стояченко И.Л. (54) СПОСОБ ПОЛУЧЕНИЯ ЧАСТЕЙ ИЗ СПЛАВОВ МЕТАЛЛОВ СЛОЖНОЙ ФОРМЫ (57) Реферат: ...

METHOD FOR PRODUCING A SYSTEM CONTAINING A LOT OF BLADES INSTALLED IN A PLATFORM

1. Способ изготовления системы (11), содержащей множество лопаток (9), установленных в платформах (7, 8), в которомизготавливают лопатки (9) отдельно от платформ (7, 8), причем эти лопатки (9) после их изготовления являются окончательно отделанными;приготавливают смесь металлического порошка с термопластическим связующим материалом;впрыскивают упомянутую смесь в литейную форму для получения заготовки платформы (7, 8);выполняют операцию удаления связующего материала из заготовки платформы (7, 8) перед соединением окончательно отделанных лопаток (9) с упомянутой заготовкой (7, 8);вставляют один конец окончательно отделанных лопаток (9) в ложемент, выполненный в заготовке платформы (7, 8), для того, чтобы соединить систему (11);выполняют операцию спекания системы, собранной из заготовки платформы (7, 8) с окончательно отделанными лопатками (9), для соединения системы (11).2. Способ по п.1, в котором упомянутая система (11) содержит множество окончательно отделанных лопаток (9), концы которых устанавливаются между внутренней платформой (8) и наружной платформой (7), причем дополнительно связывают концы окончательно отделанных лопаток (9) с заготовками платформ (7, 8) при помощи спекания металлического порошка. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (51) МПК B22F 7/06 (11) (13) 2011 147 659 A (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ЗАЯВКА НА ИЗОБРЕТЕНИЕ (21)(22) Заявка: 2011147659/02, 16.04.2010 (71) Заявитель(и): СНЕКМА (FR) Приоритет(ы): (30) Конвенционный приоритет: 24.04.2009 FR 0952707 (85) Дата начала рассмотрения заявки PCT на национальной фазе: 24.11.2011 R U (43) Дата публикации заявки: 27.05.2013 Бюл. № 15 (72) Автор(ы): БЕНАР Жан-Поль (FR), МЕНЖЕЛИН Ванесса (FR), МОТТЕН Жан-Батист (FR) (86) Заявка PCT: (87) Публикация заявки PCT: WO 2010/121966 (28.10.2010) R U (54) СПОСОБ ИЗГОТОВЛЕНИЯ СИСТЕМЫ, СОДЕРЖАЩЕЙ МНОЖЕСТВО ЛОПАТОК, УСТАНОВЛЕННЫХ В ПЛАТФОРМЕ (57) Формула изобретения 1. Способ изготовления системы (11), содержащей множество лопаток (9 ...

Stator blade segment and axial flow fluid machine with same

본 발명에 따른 정익 세그먼트는, 둘레방향으로 연장되어 복수의 정익을 연결하는 외측 연결 부재와, 상기 외측 연결 부재에 대하여, 복수의 정익 중에서 상기 둘레방향의 양단부에 위치하는 끝단 정익을 위치결정하기 위한 위치결정구와, 복수의 상기 정익 중 적어도 하나의 정익에 대하여, 상기 외측 연결 부재를 상기 정익환의 직경방향 외측으로 상대 이동 불가능하게 구속하는 직경방향 구속부를 구비하며, 복수의 상기 정익은 정익 본체와 외측 슈라우드를 갖고, 상기 외측 슈라우드에는, 복수의 상기 정익이 상기 둘레방향으로 나열하여 있는 상태로 서로 연이어서 상기 외측 연결 부재가 들어가는 홈이 형성되고, 상기 외측 연결 부재 중에서, 2개의 상기 끝단 정익의 상기 홈에 들어가는 부분에는, 상기 위치결정구가 관통 삽입되는 구멍이 형성되고, 2개의 상기 끝단 정익의 상기 홈의 바닥에는, 상기 구멍에 관통 삽입된 상기 위치결정구가 삽입되는 구멍부가 형성되고, 상기 위치결정구 중 적어도 한쪽은 제 1 원주부와, 상기 제 1 원주부의 중심축선에 대하여 편심되어 있는 제 2 원주부를 갖는 편심 위치결정구이다. A stator segment according to the present invention includes: an outer connecting member extending in a circumferential direction to connect a plurality of stator blades; and a stator for positioning the stator located at both ends in the circumferential direction of the plurality of stator blades And a radial restraining portion for restraining the outer connecting member relative to the radially outer side of the outer connecting member relative to at least one stator of the plurality of stator rods, Wherein a groove is formed in the outer shroud in the outer shroud in such a manner that a plurality of the stator is arranged in the circumferential direction so as to enter the outer connecting member, A hole is formed in the portion into which the positioning hole is inserted, , A hole portion through which the positioning hole inserted in the hole is inserted is formed in the bottom of the groove of the two end stator stitches, at least one of the positioning holes has a first circumferential portion, And a second circumferential portion that is eccentric with respect to the center axis of the second shaft portion.

Method and system for interfacing a ceramic matrix composite component to a metallic component

本发明提供了用于燃气涡轮发动机的翼型组件和从陶瓷基复合材料翼型组件传递负载至金属导叶组件支承结构的方法。翼型组件包括相对于燃气涡轮发动机的轴向方向的前端和后端。翼型组件还包括径向外端构件，其包括具有非压缩负载承载特征结构的径向向外面向端表面，所述特征结构从向外面向端表面沿径向向外延伸并且与外端构件一体地形成，所述特征结构构造成与形成在第一翼型组件支承结构的径向内表面中的互补特征结构相匹配，所述特征结构有选择地定位成正交于施加到翼型组件中的力。翼型组件还包括径向内端构件，以及在二者之间延伸的中空翼型本体，该翼型本体构造成接纳能在第一端处联接至第一翼型组件支承结构的支柱。

GAS TURBINE AXIAL COMPRESSOR STATOR

ÐÎÑÑÈÉÑÊÀß ÔÅÄÅÐÀÖÈß (19) RU (51) ÌÏÊ 7 (11) 2003 124 062 (13) A F 04 D 29/54 ÔÅÄÅÐÀËÜÍÀß ÑËÓÆÁÀ ÏÎ ÈÍÒÅËËÅÊÒÓÀËÜÍÎÉ ÑÎÁÑÒÂÅÍÍÎÑÒÈ, ÏÀÒÅÍÒÀÌ È ÒÎÂÀÐÍÛÌ ÇÍÀÊÀÌ (12) ÇÀßÂÊÀ ÍÀ ÈÇÎÁÐÅÒÅÍÈÅ (21), (22) Çà âêà: 2003124062/06, 03.01.2002 (71) Çà âèòåëü(è): ÑÍÅÊÌÀ ÌÎÒÅÐ (FR) (30) Ïðèîðèòåò: 04.01.2001 FR 01/00060 (43) Äàòà ïóáëèêàöèè çà âêè: 27.02.2005 Áþë. ¹ 6 (74) Ïàòåíòíûé ïîâåðåííûé: Åãîðîâà Ãàëèíà Áîðèñîâíà (86) Çà âêà PCT: FR 02/00007 (03.01.2002) Àäðåñ äë ïåðåïèñêè: 129010, Ìîñêâà, óë. Á.Ñïàññêà , 25, ñòð.3, ÎÎÎ "Þðèäè÷åñêà ôèðìà Ãîðîäèññêèé è Ïàðòíåðû", ïàò.ïîâ. Ã.Á. Åãîðîâîé R U Ôîðìóëà èçîáðåòåíè 1. Ñòàòîð îñåâîãî êîìïðåññîðà ãàçîâîé òóðáèíû, ñîäåðæàùèé æåñòêóþ âíåøíþþ êîëüöåâóþ àðìàòóðó (2), ïðèìûêàþùèå äðóã ê äðóãó â îñåâîì íàïðàâëåíèè êîëüöà (4à, 4b, 4ñ), ðàñïîëàãàþùèåñ èçíóòðè ïî îòíîøåíèþ ê óïîì íóòîé àðìàòóðå (2) è íåñóùèå íà ñåáå âåíöû íåïîäâèæíûõ íàïðàâë þùèõ ëîïàòîê (5), ïðè÷åì ýòè êîëüöà îáðàçîâàíû êîëüöåâûìè ñåêòîðàìè (7), çàêðåïëåííûìè íà àðìàòóðå (2), âíóòðåíí ñòåíêà êîòîðûõ îãðàíè÷èâàåò â íàðóæíîì íàïðàâëåíèè àýðîäèíàìè÷åñêèé êàíàë äâèæåíè ñæàòîé ãàçîîáðàçíîé òåêó÷åé ñðåäû, îòëè÷àþùèéñ òåì, ÷òî êîëüöåâûå ñåêòîðà (7) ïðåäñòàâë þò ñîáîé ïà íûå ñåêòîðà, îáðàçîâàííûå çàïîëíèòåëåì ñîòîâîé ñòðóêòóðû (8), çàêëþ÷åííûì ìåæäó âíóòðåííèì ëèñòîì (10), îãðàíè÷èâàþùèì óïîì íóòûé àýðîäèíàìè÷åñêèé êàíàë, è íàðóæíûì ëèñòîì (9), ïðè ýòîì ñâ çü ñ óïîì íóòîé àðìàòóðîé (2) îáåñïå÷èâàåòñ òîëüêî ïðè ïîìîùè íàðóæíîãî ëèñòà (9). 2. Ñòàòîð êîìïðåññîðà ïî ï.1, îòëè÷àþùèéñ òåì, ÷òî íàðóæíûé ëèñò (9) çàêðåïëåí íà àðìàòóðå (2) ïðè ïîìîùè áîëòîâ (14). 3. Ñòàòîð êîìïðåññîðà ïî ï.2, îòëè÷àþùèéñ òåì, ÷òî êàæäûé íàðóæíûé ëèñò (9) çàêðåïëåí íà àðìàòóðå (2) íà ñâîåì çàäíåì ïî ïîòîêó êîíöå (12) è íà ñâîåì ïåðåäíåì ïî ïîòîêó êîíöå (11) ïðè ïîìîùè ìíîæåñòâà áîëòîâ (14). 4. Ñòàòîð êîìïðåññîðà ïî ï.3, îòëè÷àþùèéñ òåì, ÷òî íàðóæíûé ëèñò (9) îòäåëåí îò àðìàòóðû (2) íåêîòîðûì ïðîñòðàíñòâîì â ïðîìåæóòêå ìåæäó åãî ïåðåäíèì ïî ïîòîêó êîíöîì (11) è åãî çàäíèì ïî ïîòîêó êîíöîì (12). 5. Ñòàòîð ...

Parting plate for steam turbine and manufacture thereof

(57)【要約】本公報は電子出願前の出願データであるた め要約のデータは記録されません。

SHOVEL AND BANDAGE WITH NEST FOR AXIAL TURBO MACHINE COMPRESSOR

А 2016119056 ко РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) РЦ (11) = (51) МПК РОО 900 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ЗАЯВКА НА ИЗОБРЕТЕНИЕ (21)(22) Заявка: 2016119056, 17.05.2016 (71) Заявитель(и): САФРАН АЭРО БУСТЕРС СА (ВЕ) Приоритет(ы): (30) Конвенционный приоритет: (72) Автор(ы): 21.05.2015 ВЕ 2015/5316 Жан-Франсуа КОРТЕКИСС (ВЕ) (43) Дата публикации заявки: 22.11.2017 Бюл. № 33 Адрес для переписки: 123242, Москва, пл. Кудринская, д. 1, а/я 35, "Михайлюк, Сороколат и партнеры - патентные поверенные" (54) ЛОПАТКА И БАНДАЖ С ГНЕЗДОМ ДЛЯ КОМПРЕССОРА ОСЕВОЙ ТУРБОМАШИНЫ (57) Формула изобретения 1. Лопаточный узел (34) осевой турбомашины, в частности компрессора (5; 6) осевой турбомашины (2), причем узел (34) включает в себя: - стенку (28; 30), которая предназначена для ограничения кольцевого потока (18; 20) турбомашины в радиальном направлении и которая содержит скрепляющее гнездо (32), - лопатку (24; 26), закрепленную в скрепляющем гнезде (32) и проходящую радиально относительно стенки (28; 30), и - скрепляющий слой (50), расположенный на стыке между лопаткой (24; 26) и гнездом (32), отличающийся тем, что место стыковки содержит неровности (48), находящиеся в контакте со скрепляющим слоем (50) с тем, чтобы обеспечить фиксацию за счет прилегания материала для закрепления лопатки (24; 26) в стенке (28; 30). 2. Лопаточный узел (34) по п. 1, отличающийся тем, что неровности (48), находящиеся в контакте со скрепляющим слоем (50), образованы на лопатке (24; 26) и/или внутри гнезда (32), ипредпочтительно неровности (48) распределены в аксиальном направлении. 3. Лопаточный узел (34) по п. 1, отличающийся тем, что лопатка (24; 26) содержит поверхность (54) давления и поверхность (56) разрежения, причем неровности (48) распределены по поверхности (54) давления и/или по поверхности (56) разрежения внутри гнезда (32). 4. Лопаточный узел (34) по любому из пи. 1-3, отличающийся тем, что лопатка (24; 26) содержит криволинейную поверхность (54) и/или выпуклую ...

Blade ring segment, turbomachine and method for producing same

本发明涉及一种用于流体机械（17）、尤其涡轮机的叶片环扇段（2、2’），所述叶片环扇段具有第一叶片（1）和第二叶片（1’），其中所述第一叶片（1）与所述第二叶片（1’）形状锁合地连接。

Compressor assembly and airfoil assembly

本发明涉及压缩机系统和翼型件组件。一种翼型件组件(44)，其包括：至少一个翼型件(50)，其具有前缘(51)和后缘(52)；带(60，80)，其沿带(60，80)与至少一个翼型件(50)之间的界面(70)的一部分刚性地联接于至少一个翼型件(50)，以用于对至少一个翼型件(50)提供至少一部分支撑；凹口(72)，其在至少一个翼型件(50)的前缘(51)或后缘(52)处位于带(60，80)中，且限定在带(60，80)与前缘(51)或后缘(52)之间的应力消除间隙(74)；和封闭件(76)，其阻止穿过凹口(72)的空气流。

Turbine machine nozzle assembly

一种用于涡轮机的喷嘴组件，包括：喷嘴叶片(43)，具有内侧壁(44)和外侧壁(46)，并且在组装时部分限定进入涡轮机的流路；外环(20)；分流器(11)，具有水平延伸部(21)；外环(20)和外侧壁(46)之间的界面，具有如下的至少一种(i)径向互锁件(48、76、78)；(ii)凸形/凹形界面(54、82、84)；或者(iii)凹形凹口(106)，在外侧壁(46)的导前和拖尾边缘处通过径向伸出的凸形台阶侧面(108)邻接；以及水平延伸部(21)和内侧壁(44)之间的界面，具有如下的至少一种(i)径向互锁件(48、76、78)；或者(ii)凸形/凹形界面(54、82、84)；或者(iii)凹形凹口(106)，在内侧壁(44)的导前和拖尾边缘处通过径向伸出的凸形台阶侧面(108)邻接。

Method for assembling turbomachine parts and assembly used in such method

Изобретение относится к способу сборки первой детали турбомашины по меньшей мере с одной второй деталью турбомашины, включающему в себя следующие этапы, на которых: нагнетают вулканизируемый эластомер, предпочтительно кремнийорганическое соединение (145), которое можно вулканизировать при температуре окружающей среды, именуемое кремнийорганическим соединением, вулканизируемым при комнатной температуре (кремнийорганическим ВКТ-соединением), в зоне (140) нагнетания на стыке между первой и второй деталями; локально нагревают зону (140) нагнетания, чтобы вулканизировать вулканизируемый эластомер. Изобретение также относится к узлу (20), применяемому при таком способе сборки. Техническим результатом изобретения является уменьшение длительности охлаждения готового изделия. 2 н. и 11 з.п. ф-лы, 5 ил. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 668 666 C2 (51) МПК B29C 45/16 (2006.01) F01D 5/00 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК B29C 45/16 (2018.05); F01D 5/00 (2018.05) (21)(22) Заявка: 2016105128, 17.07.2014 (24) Дата начала отсчета срока действия патента: (73) Патентообладатель(и): СНЕКМА (FR) Дата регистрации: 02.10.2018 (56) Список документов, цитированных в отчете о поиске: FR 2958323 A1, 07.10.2011. US 2009044906 A1, 19.02.2009. EP 2586989 A1, 01.05.2013. RU 99367 U1, 20.11.2010. 18.07.2013 FR 13 57078 (43) Дата публикации заявки: 21.08.2017 Бюл. № 24 (45) Опубликовано: 02.10.2018 Бюл. № 28 (86) Заявка PCT: C 2 C 2 (85) Дата начала рассмотрения заявки PCT на национальной фазе: 18.02.2016 FR 2014/051832 (17.07.2014) (87) Публикация заявки PCT: WO 2015/007994 (22.01.2015) 2 6 6 8 6 6 6 2 6 6 8 6 6 6 Приоритет(ы): (30) Конвенционный приоритет: R U R U 17.07.2014 (72) Автор(ы): ВАТЕН Эрик Раймон Жан (FR) Адрес для переписки: 129090, Москва, ул. Б.Спасская, 25, строение 3, ООО "Юридическая фирма Городисский и Партнеры" (54) СПОСОБ СБОРКИ ДЕТАЛЕЙ ТУРБОМАШИНЫ И УЗЕЛ, ПРИМЕНЯЕМЫЙ ПРИ ТАКОМ СПОСОБЕ (57) ...

Turbine blade, turbine, and method of manufacturing turbine blade

Method for repairing a component

The invention relates to a method for repairing a component (10), particularly a stator-side component of a gas turbine such a housing or a vane ring during which a damaged section is removed from the component, and a new section that replaces the damaged and thus removed section is joined to the component in a fixed manner by welding. According to the invention, the damaged section is removed from the component to be repaired in a manner that minimizes the length of a separating seam (14) and thus a subsequent weld seam (14). Material is removed from the component according to the distribution of material thickness along the separating seam (14) in order to provide a material thickness that is as uniform as possible along the subsequent weld seam. At least the removed material is replaced after the component has been joined to the new section by laser powder deposition welding.

METHOD FOR PRODUCING A STATOR OR ROTOR COMPONENT

ÐÎÑÑÈÉÑÊÀß ÔÅÄÅÐÀÖÈß (19) RU (51) ÌÏÊ 7 (11) 2004 119 044 (13) A B 23 K 1/00, B 23 P 15/00 ÔÅÄÅÐÀËÜÍÀß ÑËÓÆÁÀ ÏÎ ÈÍÒÅËËÅÊÒÓÀËÜÍÎÉ ÑÎÁÑÒÂÅÍÍÎÑÒÈ, ÏÀÒÅÍÒÀÌ È ÒÎÂÀÐÍÛÌ ÇÍÀÊÀÌ (12) ÇÀßÂÊÀ ÍÀ ÈÇÎÁÐÅÒÅÍÈÅ (21), (22) Çà âêà: 2004119044/02, 08.11.2002 (71) Çà âèòåëü(è): ÂÎËÜÂÎ ÀÝÐÎ ÊÎÐÏÎÐÅÉØÍ (SE) (30) Ïðèîðèòåò: 22.11.2001 SE 0103892-6 (43) Äàòà ïóáëèêàöèè çà âêè: 10.06.2005 Áþë. ¹ 16 (74) Ïàòåíòíûé ïîâåðåííûé: Âåñåëèöêà Èðèíà Àëåêñàíäðîâíà (86) Çà âêà PCT: SE 02/02029 (08.11.2002) Àäðåñ äë ïåðåïèñêè: 101000, Ìîñêâà, Ì.Çëàòîóñòèíñêèé ïåð., ä.10, êâ.15, "ÅÂÐÎÌÀÐÊÏÀÒ", È.À.Âåñåëèöêîé R U Ôîðìóëà èçîáðåòåíè 1. Ñïîñîá èçãîòîâëåíè êîìïîíåíòà (10, 12) ñòàòîðà èëè ðîòîðà, ó êîòîðîãî ïî ìåíüøåé ìåðå îäíà ëîïàòêà (2) îáëîïà÷åííîãî äèñêà èëè êîëüöà (1, 14) ñî ìíîæåñòâîì îáðàçóþùèõ ðàäèàëüíûå âûñòóïû ëîïàòîê (2) ñîåäèíåíà ïî ìåíüøåé ìåðå ñ îäíèì êîëüöåâûì ýëåìåíòîì (3, 11, 13), çàêëþ÷àþùèéñ â òîì, ÷òî ïîäãîòàâëèâàþò ñîåäèíèòåëüíûé ìàòåðèàë, çàïîëí þùèé çàçîð ìåæäó ïî ìåíüøåé ìåðå îäíîé ëîïàòêîé (2) è êîëüöåâûì ýëåìåíòîì (3, 11, 13), ëîïàòêó è êîëüöåâîé ýëåìåíò ñîáèðàþò äðóã ñ äðóãîì òàêèì îáðàçîì, ÷òî ïðè íàãðåâå ëîïàòêè è êîëüöåâîãî ýëåìåíòà è ðàñïîëîæåííîãî ìåæäó íèìè ñîåäèíèòåëüíîãî ìàòåðèàëà ìåæäó íèìè îáðàçóåòñ ñòûêîâîå ñîåäèíåíèå, à çàòåì íàãðåâàþò äî òåìïåðàòóðû ïëàâëåíè ñîåäèíèòåëüíîãî ìàòåðèàëà, êîòîðûé ïîñëå çàòâåðäåâàíè ïðî÷íî ñîåäèí åò ìåæäó ñîáîé ëîïàòêó è êîëüöåâîé ýëåìåíò, îòëè÷àþùèéñ òåì, ÷òî êîëüöåâîé ýëåìåíò è îáëîïà÷åííûé äèñê èëè êîëüöî èìåþò âíóòðåííþþ â ðàäèàëüíîì íàïðàâëåíèè ïîâåðõíîñòü (4), îñü ïî ìåíüøåé ìåðå ÷àñòè êîòîðîé íàêëîíåíà ê öåíòðàëüíîé îñü, à ôîðìà íàðóæíîé â ðàäèàëüíîì íàïðàâëåíèè ïîâåðõíîñòè (5) êîëüöåâîãî ýëåìåíòà è îáëîïà÷åííîãî äèñêà èëè êîëüöà ïî ñóùåñòâó ñîîòâåòñòâóåò ôîðìå íàêëîííîãî ó÷àñòêà âíóòðåííåé â ðàäèàëüíîì íàïðàâëåíèè ïîâåðõíîñòè, è êîëüöåâîé ýëåìåíò è îáëîïà÷åííûé äèñê èëè êîëüöî ñîåäèí þò ìåæäó ñîáîé ïóòåì èõ îòíîñèòåëüíîãî ïåðåìåùåíè â îñåâîì íàïðàâëåíèè äî óïîðà äðóã â äðóãà èõ íàêëîííûõ ïîâåðõíîñòåé. 2. Ñïîñîá ïî ï. 1, ...

Stationary blading for steam turbine, multistage steam turbine and method for manufacturing blade unit

FIELD: machine engineering. SUBSTANCE: stationary blading designed for the last stage of the steam turbine and containing guiding blade units which delimit the annular chamber and each of which contains an elongated blade portion. The mentioned elongated blade portion additionally has a longitudinal channel and an inner portion soldered to the first longitudinal end of the mentioned blade portion and having a through hole forming part of the annular chamber and an inner channel extending from the through hole to the longitudinal channel. The outer part which interacts with the mentioned steam turbine is soldered to the second longitudinal end of the mentioned blade portion. The outer part comprises an outer channel open to the surface of the mentioned steam turbine and to the mentioned longitudinal channel. EFFECT: simplicity in maintenance or replacement, reliable operation in the presence of wet steam, the design is simple and easier to manufacture. 11 cl, 11 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 631 852 C2 (51) МПК F01D 25/32 (2006.01) F01D 9/04 (2006.01) B23P 15/04 (2006.01) F01D 5/28 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (21)(22) Заявка: 2012153181, 11.12.2012 (24) Дата начала отсчета срока действия патента: 11.12.2012 Дата регистрации: (72) Автор(ы): ГРИЛЛИ Марко (IT), ДЖУСТИ Энрико (IT), ИМПАРАТО Энцо (IT) Приоритет(ы): (30) Конвенционный приоритет: (56) Список документов, цитированных в отчете о поиске: EP 1744018 A1, 17.01.2007. RU 12.12.2011 IT CO2011A000060 (45) Опубликовано: 26.09.2017 Бюл. № 27 2410226 C2, 27.01.2011. US 2011200430 A1, 18.08.2011. US 3881842 A, 06.05.1975. US 3724967 A, 03.04.1973. SU 1386719 A1, 07.04.1988. 2 6 3 1 8 5 2 R U (54) Направляющий лопаточный венец для паровой турбины, многоступенчатая паровая турбина и способ изготовления лопаточного узла (57) Реферат: Направляющий лопаточный венец, отверстия к продольному каналу. Ко второму предназначенный для последней ...

METHOD FOR PRODUCING A GUIDE BLADE FOR A TURBINE OR COMPRESSOR LEAD, AND GUIDE BLADE PRODUCED BY THE METHOD

Rotor or stator component forming method

Изобретение относится к сварке, в частности к способу изготовления компонента статора или ротора, и может найти применение в машиностроении для изготовления газовых турбин или в самолетостроении при изготовлении реактивных двигателей. Компонент (1) статора или ротора имеет, по меньшей мере, одну направляющую поток газа и/или передающую усилия перегородку (3). Перегородку (3) прочно приваривают лазерной сваркой, по меньшей мере, к одному кольцевому элементу (2) с противоположной в радиальном направлении стороны таким образом, что соединенные между собой участки перегородки и кольцевого элемента образуют соединение (4) Т-образной формы. В результате получают надежное соединение простым и производительным способом. 10 з.п. ф-лы, 6 ил. ÐÎÑÑÈÉÑÊÀß ÔÅÄÅÐÀÖÈß RU (19) (11) 2 287 418 (13) C2 (51) ÌÏÊ B23P 15/02 B23K 26/20 F04D 29/30 (2006.01) (2006.01) (2006.01) ÔÅÄÅÐÀËÜÍÀß ÑËÓÆÁÀ ÏÎ ÈÍÒÅËËÅÊÒÓÀËÜÍÎÉ ÑÎÁÑÒÂÅÍÍÎÑÒÈ, ÏÀÒÅÍÒÀÌ È ÒÎÂÀÐÍÛÌ ÇÍÀÊÀÌ (12) ÎÏÈÑÀÍÈÅ ÈÇÎÁÐÅÒÅÍÈß Ê ÏÀÒÅÍÒÓ (21), (22) Çà âêà: 2004109593/02, 14.08.2002 (72) Àâòîð(û): ËÓÍÄÃÐÅÍ ßí (SE) (24) Äàòà íà÷àëà îòñ÷åòà ñðîêà äåéñòâè ïàòåíòà: 14.08.2002 (73) Ïàòåíòîîáëàäàòåëü(è): ÂÎËÜÂÎ ÀÝÐÎ ÊÎÐÏÎÐÅÉØÍ (SE) R U (30) Êîíâåíöèîííûé ïðèîðèòåò: 29.08.2001 SE 0102883.6 (43) Äàòà ïóáëèêàöèè çà âêè: 10.06.2005 (45) Îïóáëèêîâàíî: 20.11.2006 Áþë. ¹ 32 2 2 8 7 4 1 8 (56) Ñïèñîê äîêóìåíòîâ, öèòèðîâàííûõ â îò÷åòå î ïîèñêå: RU 2120567 C1, 20.10.1998. SU 1116223 A1, 30.09.1984. SU 941148 A1, 07.07.1982. SU 1731975 A1, 07.05.1992. SU 166991 A, 21.01.1965. (85) Äàòà ïåðåâîäà çà âêè PCT íà íàöèîíàëüíóþ ôàçó: 29.03.2004 2 2 8 7 4 1 8 R U (87) Ïóáëèêàöè PCT: WO 03/020469 (13.03.2003) C 2 C 2 (86) Çà âêà PCT: SE 02/01453 (14.08.2002) Àäðåñ äë ïåðåïèñêè: 101000, Ìîñêâà, Ì.Çëàòîóñòèíñêèé ïåð., 10, êâ.15, "ÅÂÐÎÌÀÐÊÏÀÒ", ïàò.ïîâ. È.À.Âåñåëèöêîé, ðåã. ¹ 11 (54) ÑÏÎÑÎÁ ÈÇÃÎÒÎÂËÅÍÈß ÊÎÌÏÎÍÅÍÒÀ ÑÒÀÒÎÐÀ ÈËÈ ÐÎÒÎÐÀ (57) Ðåôåðàò: Èçîáðåòåíèå îòíîñèòñ ê ñâàðêå, â ÷àñòíîñòè ê ñïîñîáó èçãîòîâëåíè êîìïîíåíòà ñòàòîðà èëè ðîòîðà, è ìîæåò íàéòè ïðèìåíåíèå ...

RING FOR STATOR OF TURBINE ENGINE OF AIRCRAFT WITH CUTTINGS OF REMOVAL OF MECHANICAL LOAD OF BLADES

1. Кольцо (16) для статора модуля турбинного двигателя летательного аппарата, снабженного множеством сквозных отверстий (22), каждое из которых предназначено для расположения в нем лопатки статора (4); причем каждое отверстие определяет профиль (32), проходящий между первым краем (25), предназначенным для расположения задней кромки (26) лопатки, и вторым краем (27), предназначенным для расположения передней кромки (28) лопатки, отличающееся тем, что, по меньшей мере, с одним из вышеупомянутых отверстий (22) соединена прорезь (36) снятия механической нагрузки, выполненная сквозной на кольце и расположенная против и на удалении от упомянутого первого края (25) отверстия в направлении упомянутого профиля (32).2. Кольцо по п.1, отличающееся тем, что упомянутая прорезь (36) проходит по кривой линии (38) к отверстию.3. Кольцо по п.1, отличающееся тем, что упомянутое отверстие (22) содержит с одной и другой стороны профиля (32) часть внутренней поверхности (23) и часть внешней поверхности (24), которые соединяются на уровне упомянутых первого и второго краев (25, 27) отверстия, а также тем, что упомянутая прорезь (36) также проходит против и на удалении от участков (23а, 24а) частей внутренней поверхности и внешней поверхности (23, 24), которые соединяются на уровне упомянутого первого края (25) отверстия (22).4. Кольцо по п. 1, отличающееся тем, что упомянутая прорезь (36) в целом имеет U-образную или V-образную форму, внутри которой расположен упомянутый первый край (25) отверстия.5. Кольцо по п. 1, отличающееся тем, что упомянутая прорезь (36) заполнена материалом-наполнителем (40).6. Кольцо по любому из предшествующих пунктов, отличающееся тем, что оно образует непрерывную, по сущ РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (51) МПК F04D 29/54 (13) 2011 149 631 A (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ЗАЯВКА НА ИЗОБРЕТЕНИЕ (21)(22) Заявка: 2011149631/02, 04.05.2010 (71) Заявитель(и): СНЕКМА (FR) Приоритет(ы): (30) Конвенционный приоритет: (72) Автор( ...

Techniques and assemblies for joining components

The disclosure describes example techniques and assemblies for joining a first component and a second component. The techniques may include positioning the first and second component adjacent to each other to define a joint region between adjacent portions of the first component and the second component, the joint region being coated with an adhesion resistant coating. The techniques may also include positioning a braze material in the joint region, heating the braze material to form an at least softened material, and cooling the at least softened material to form a mechanical interlock including the braze material in the joint region joining the first and second components. The braze material does not metallurgically bond to the joint surface.

Steam Turbine Diaphragm and Manufacturing Method Thereof

내용 없음.

Patent RU2016119056A3

7 ВУ’” 2016119056” АЗ Дата публикации: 30.08.2019 Форма № 18 ИЗ,ПМ-2011 Федеральная служба по интеллектуальной собственности Федеральное государственное бюджетное учреждение 5 «Федеральный институт промышленной собственности» (ФИПС) ОТЧЕТ О ПОИСКЕ 1. . ИДЕНТИФИКАЦИЯ ЗАЯВКИ Регистрационный номер Дата подачи 2016119056/06(029903) 17.05.2016 Приоритет установлен по дате: [ ] подачи заявки [ ] поступления дополнительных материалов от к ранее поданной заявке № [ ] приоритета по первоначальной заявке № из которой данная заявка выделена [ ] подачи первоначальной заявки № из которой данная заявка выделена [ ] подачи ранее поданной заявки № [Х] подачи первой(ых) заявки(ок) в государстве-участнике Парижской конвенции (31) Номер первой(ых) заявки(ок) (32) Дата подачи первой(ых) заявки(ок) (33) Код страны 1. 2015/5316 21.05.2015 ВЕ Название изобретения (полезной модели): [Х] - как заявлено; [ ] - уточненное (см. Примечания) ЛОПАТКА И БАНДАЖ С ГНЕЗДОМ ДЛЯ КОМПРЕССОРА ОСЕВОЙ ТУРБОМАШИНЫ Заявитель: САФРАН АЭРО БУСТЕРС СА, ВЕ 2. ЕДИНСТВО ИЗОБРЕТЕНИЯ [Х] соблюдено [ ] не соблюдено. Пояснения: см. Примечания 3. ФОРМУЛА ИЗОБРЕТЕНИЯ: [Х] приняты во внимание все пункты (см. п см. Примечания [ ] приняты во внимание следующие пункты: р [ ] принята во внимание измененная формула изобретения (см. Примечания) 4. КЛАССИФИКАЦИЯ ОБЪЕКТА ИЗОБРЕТЕНИЯ (ПОЛЕЗНОЙ МОДЕЛИ) (Указываются индексы МПК и индикатор текущей версии) ЕО 9/04 (2006.01) Е04 29/54 (2006.01) 5. ОБЛАСТЬ ПОИСКА 5.1 Проверенный минимум документации РСТ (указывается индексами МПК) [ОТ 9/00-9/04, 25/00, 25/24, НО4О 29/00, 29/40, 29/52-29/54 5.2 Другая проверенная документация в той мере, в какой она включена в поисковые подборки: 5.3 Электронные базы данных, использованные при поиске (название базы, и если, возможно, поисковые термины): Езрасепеф, ]-Р]а( Ра, РАТЕМТЗСОРЕ, Рабеагсв, ОЗРТО 6. ДОКУМЕНТЫ, ОТНОСЯЩИЕСЯ К ПРЕДМЕТУ ПОИСКА Кате- Наименование документа с указанием (где необходимо) частей, Относится к гория* относящихся к ...

METHOD FOR ASSEMBLING TURBO MACHINE PARTS AND ASSEMBLY APPLICABLE WITH SUCH METHOD

А 2016105128 ко РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) 11 3 << < < х 3 } ха = (13 (50) МПК В29С 45/16 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ЗАЯВКА НА ИЗОБРЕТЕНИЕ (21)(22) Заявка: 2016105128, 17.07.2014 (71) Заявитель(и): СНЕКМА (ЕК) Приоритет(ы): (30) Конвенционный приоритет: (72) Автор(ы): 18.07.2013 ЕК 13 57078 ВАТЕН Эрик Раймон Жан (ЕВ) (43) Дата публикации заявки: 21.08.2017 Бюл. № 24 (85) Дата начала рассмотрения заявки РСТ на национальной фазе: 18.02.2016 (86) Заявка РСТ: ЕК 2014/051832 (17.07.2014) (87) Публикация заявки РСТ: УГО 2015/007994 (22.01.2015) Адрес для переписки: 129090, Москва, ул. Б.Спасская, 25, строение 3, ООО "Юридическая фирма Городисский и Партнеры" (54) СПОСОБ СБОРКИ ДЕТАЛЕЙ ТУРБОМАШИНЫ И УЗЕЛ, ПРИМЕНЯЕМЫЙ ПРИ ТАКОМ СПОСОБЕ (57) Формула изобретения 1. Способ сборки первой детали турбомашины, по меньшей мере, с одной второй деталью турбомашины, включающий в себя следующие этапы, на которых: - нагнетают вулканизируемый эластомер в зоне (140) нагнетания на стыке между первой и второй деталями; - локально нагревают зону (140) нагнетания, чтобы вулканизировать вулканизируемый эластомер. 2. Способ сборки по п. 1, в котором первая деталь и вторая деталь являются двумя кольцевыми деталями, которые будут образовывать направляющий аппарат (100) турбомашины после сборки. 3. Способ сборки по п. 1, в котором на этапе локального нагрева используют узел (20), содержащий опору (22) для первой и второй деталей и систему локального нагрева, которая содержит зону (27) нагрева и связана с опорой (22) так, что зона (27) нагрева обращена к зоне (140) нагнетания, когда проводят этап нагрева. 4. Способ сборки по п. 3, в котором узел (20) также действует как сборочная опора во время этапа нагнетания эластомера. 5. Способ сборки по п. 1, в котором на этапе локального нагрева используют нагревательный элемент, предварительно размещаемый на обеих деталях турбомашины вблизи зоны (140) нагнетания. Стр.: 1 па 86190191 0Сс У А 2016105128 ко 6. Способ ...

Turbine nozzle assembly

Combustor inlet mixing system with slotted swirler vanes

One-piece centred housing of an axial turbomachine compressor and associated manufacturing process

Device for guiding a stream of air entering a combustion chamber of a turbomachine

一种用于引导气流进入涡轮机燃烧腔的设备，其包括在扩散器(16)之后的导流板(14)，一个导流板的护罩(40)与扩散器的转子中的一个壁(34)一体化被形成，另一个导流板护罩(38)被添加和连接至扩散器的转子中的另一个壁(32)，导流板叶片(42)通过一端被固定至导流板中的一个护罩(40)以及在它们的另一端以小间隙(46)与另一个护罩(38)分开。

Device for fastening sectors of guide-vane assembly, guide-vane assembly and turbomachine

Изобретение относится к приспособлениям для крепления направляющих аппаратов в турбомашинах. Сектора (25) направляющего аппарата, на которых закреплены лопатки, располагаются рядом с корпусом (31) посредством уплотняющих секторов (32), чередующихся с упомянутыми секторами в осевом направлении, и снабжены поверхностями (37, 39), оказывающими сопротивление силам, и посредством этих поверхностей силы, прикладываемые к секторам распределительного аппарата, передаются на корпус. Внутренняя поверхность корпуса (31) сделана более гладкой и не оснащена никакими крюками, вследствие чего этот корпус (31) менее сложен в изготовлении и испытывает меньшие механические напряжения при креплении секторов направляющего аппарата, так что эти сектора направляющего аппарата можно устанавливать посредством только осевого движения. Изобретение обеспечивает упрощение конструкции корпуса и удобство сборки направляющего аппарата. 3 н. и 5 з.п. ф-лы. 2 ил.

METHOD FOR PRODUCING PARTS FROM ALLOYS OF METALS OF COMPLEX SHAPE

РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2019 128 112 A (51) МПК B23P 15/04 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ЗАЯВКА НА ИЗОБРЕТЕНИЕ (21)(22) Заявка: 2019128112, 12.03.2018 (71) Заявитель(и): МЕКАХРОМ ФРАНС (FR) Приоритет(ы): (30) Конвенционный приоритет: 13.03.2017 FR 1752018 (85) Дата начала рассмотрения заявки PCT на национальной фазе: 14.10.2019 R U (43) Дата публикации заявки: 14.04.2021 Бюл. № 11 (72) Автор(ы): ДЕ ПОННАТ, Арнод (FR), МАРТИН, Оливье (FR) (86) Заявка PCT: (87) Публикация заявки PCT: WO 2018/167384 (20.09.2018) R U (54) СПОСОБ ПОЛУЧЕНИЯ ЧАСТЕЙ ИЗ СПЛАВОВ МЕТАЛЛОВ СЛОЖНОЙ ФОРМЫ (57) Формула изобретения 1. Способ получения сплошных лопаток газотурбинных двигателей, имеющих среднюю часть (1), кромку/законцовку (3) и хвостовик (2), где способ включает: стадию производства заготовки, по меньшей мере, из двух частей (42, 61, 71, 50, 51, 64, 65, 80, 81, 90, 91, 92, 93), где, по меньшей мере, одна из них является твердой частью, указанные, по меньшей мере, две части собирают способом диффузионного соединения без плавления, и стадию механической обработки этой заготовки для производства лопатки с заданным профилем. 2. Способ по п. 1, где, по меньшей мере, две из указанных частей являются твердыми. 3. Способ по п. 1 или 2, где, по меньшей мере, одна из указанных частей (103, 104) получена с использованием порошка и сформирована непосредственно на указанной, по меньшей мере, одной твердой части. 4. Способ по любому из пп. 1-3, где указанная, по меньшей мере, одна твердая часть, является блоком, изготовленным из металлического сплава. 5. Способ по п. 4, где средняя часть (1) указанной лопатки получена из блока (61), вырезанного по периферии цилиндрической болванки (6) с кольцевым сечением, чтобы воспользоваться преимуществом ее радиуса кривизны при формировании внешней поверхности указанной средней части. 6. Способ по п. 4, где средняя часть (1) указанной лопатки получена из блока (42), Стр.: 1 A 2 0 1 9 1 2 8 1 1 2 A Адрес ...

Attachment of blades in holder of composite material by fixation

FIELD: instrument making. SUBSTANCE: invention relates to mechanical component assembly (1) for aircraft containing part (3) containing connected end; recess intended for fitting part (3), wherein said cavity (2) has wall with composite material with organic matrix; fixing composite material (4) containing a thermoplastic or thermosetting material with content of filler from 0 to 70 wt% and forming a mechanical and/or physical-chemical bond between said part (3) and socket (2) with a wall made from composite material with an organic matrix. EFFECT: achieving reduced weight, reduced production costs, simple and less costly repair, production of finished parts without any reworking after moulding. 15 cl, 2 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 583 183 C2 (51) МПК F04D 29/54 (2006.01) F01D 5/30 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ (21)(22) Заявка: ИЗОБРЕТЕНИЯ К ПАТЕНТУ 2013115872/06, 07.09.2011 (24) Дата начала отсчета срока действия патента: 07.09.2011 Приоритет(ы): (30) Конвенционный приоритет: (72) Автор(ы): РЕНАР Филипп (FR), ГРЕЛЕН Эрве (FR), БЕРАР Саша (BE) 16.09.2010 EP 10009660.1 (43) Дата публикации заявки: 27.10.2014 Бюл. № 30 R U (73) Патентообладатель(и): ТЕКСПЕЙС АЭРО С.А. (BE) (45) Опубликовано: 10.05.2016 Бюл. № 13 (85) Дата начала рассмотрения заявки PCT на национальной фазе: 16.04.2013 2 5 8 3 1 8 3 (56) Список документов, цитированных в отчете о поиске: RU 2317448 C2, 20.02.2008. EP 1081335 A2, 07.03.2001. DE 102009010613 A1, 02.09.2010. EP 0433111 A1, 19.06.1991. SU 1479705 A1, 15.05.1989. (86) Заявка PCT: 2 5 8 3 1 8 3 R U C 2 C 2 EP 2011/065445 (07.09.2011) (87) Публикация заявки PCT: WO 2012/034906 (22.03.2012) Адрес для переписки: 191002, Санкт-Петербург, а/я 5, ООО "Ляпунов и партнеры" (54) КРЕПЛЕНИЕ ЛОПАТКИ В ДЕРЖАТЕЛЕ ИЗ КОМПОЗИТНОГО МАТЕРИАЛА ПУТЕМ ФИКСАЦИИ (57) Реферат: Изобретение относится к механическому процентов и образующий механическую и/или сборочному узлу (1) для авиации, содержащему: ...

Method and tool for assemblage of straightening stage

FIELD: machine engineering. SUBSTANCE: invention relates to the method and tool for assemblage the straightening stage (1) including a coaxial inner shell (6) and outer shell connected by radial blades (8). The method consists of the stage for supporting plates (19) with abutment against the outer surface of the inner shell (6), so that the plates (19) cover hermetically and at least partially the gaps (15) formed between the inner shell (6) holes (10) and the blades (8), and the stage for applying a casting resin (11) on the inner surface (12) of the inner shell (6), so that the gaps (15) are filled with resin and the radial inner ends of the blades (8) are embedded in the resin (11). EFFECT: invention makes it possible to eliminate a stage which consists in injecting the resin by means of a syringe into said gaps before applying the filling resin onto the inner surface of this shell, this stage for injecting the resin, which is accurate and difficult to perform, is replaced by a simpler and much less long stage for placing and supporting the plates against the outer surface of the inner shell. 10 cl, 8 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 633 312 C2 (51) МПК F04D 29/54 (2006.01) F01D 9/04 (2006.01) B23P 15/04 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (21)(22) Заявка: 2014145861, 12.04.2013 (24) Дата начала отсчета срока действия патента: 12.04.2013 Дата регистрации: (72) Автор(ы): ШАРДОННЕ Ромен Рене Марсель (FR), ПЕТИ Патрик Жильбер (FR), МАРТИНЕ Ален (FR) Приоритет(ы): (30) Конвенционный приоритет: (56) Список документов, цитированных в отчете о поиске: EP 1219785 A1, 03.07.2002. RU 16.04.2012 FR 1253493 (45) Опубликовано: 11.10.2017 Бюл. № 29 2435037 C2, 27.11.2011. US 2010254804 A1, 07.10.2010. RU 2287418 C2, 20.11.2006. SU 1216917 A1, 20.12.2005. RU 2083430 C1, 10.07.1997. RU 2388501 C2, 10.05.2010. (85) Дата начала рассмотрения заявки PCT на национальной фазе: 17.11.2014 (86) Заявка PCT: FR ...

Method and apparatus for enhancing compressor performance

The present invention provides a non-axisymmetric flow member disposed in a duct of a gas turbine engine. The flow member narrows the area aft of a vane to reduce the cross sectional area through which a wake from the vane traverses.

Vane ring and method for producing the same

Die Erfindung betrifft einen Leitschaufelkranz einer Turbomaschine, insbesondere einer Gasturbine, mit einem Leitschaufelträgerring und mehreren am Leitschaufelträgerring befestigten Leitschaufeln. Erfindungsgemäß sind der Leitschaufelträgerring und die Leitschaufeln aus unterschiedlichen Werkstoffen hergestellt, wobei der Werkstoff, aus dem die Leitschaufeln hergestellt sind, höherwertiger ist als der Werkstoff, aus dem der Leitschaufelträgerring hergestellt ist. The invention relates to a vane ring of a turbomachine, in particular a gas turbine, with a vane support ring and a plurality of guide vanes attached to the vane support ring. According to the invention, the vane support ring and the guide vanes are made of different materials, wherein the material from which the vanes are made, is superior to the material from which the vane support ring is made.

Guide vane grid is blade rim or rim segment forming part of gas turbine with several fixed blades with ends fixed into and soldered in inner, outer ring segments; guide vanes are manufactured by powder metallurgical injection molding

The guide vane grid is a blade rim or rim segment and forms part of a gas turbine with several fixed blades (10) with ends fixed into an soldered in inner and outer ring segments. The guide vanes are manufactured by powder metallurgical injection molding. The guide vanes are manufactured based on a nickel material or based on a nickel alloy material. Independent claims are also included for the following: (A) a gas turbine with an inventive device (B) and a method of manufacturing an inventive device.