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Небесная энциклопедия

Космические корабли и станции, автоматические КА и методы их проектирования, бортовые комплексы управления, системы и средства жизнеобеспечения, особенности технологии производства ракетно-космических систем

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Мониторинг СМИ

Мониторинг СМИ и социальных сетей. Сканирование интернета, новостных сайтов, специализированных контентных площадок на базе мессенджеров. Гибкие настройки фильтров и первоначальных источников.

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Поддерживает ввод нескольких поисковых фраз (по одной на строку). При поиске обеспечивает поддержку морфологии русского и английского языка
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Применить Всего найдено 14576. Отображено 200.
21-01-2010 дата публикации

Leitschaufel eines Gebläses

Номер: DE602008000360D1
Принадлежит: SNECMA

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31-05-2007 дата публикации

Vorrichtung zur automatischen Aufnahme / Rückgabe eines Werkzeugs von einem / in einen Werkzeugstapel

Номер: DE602004002835T2
Принадлежит: SNECMA

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21-02-1980 дата публикации

Номер: DE0002319788B2

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24-10-1968 дата публикации

Speichereinrichtung

Номер: DE0001280930B
Автор: SAUVAN JACQUES, MAS LES

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24-04-2008 дата публикации

Rotorscheibe einer Turbomaschine

Номер: DE602005005254D1
Принадлежит: SNECMA

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28-03-2018 дата публикации

Turbomachine Propeller Blade Setting Device

Номер: GB0002508957B
Принадлежит: SNECMA

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06-07-2016 дата публикации

An annular casing for a turbine engine compressor

Номер: GB0002497644B
Автор: AUDE ABADIE, Aude Abadie
Принадлежит: SNECMA

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08-11-2017 дата публикации

Composite Turbine Engine Blade with Structural Reinforcement

Номер: GB0002507146B
Принадлежит: SNECMA

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08-09-1982 дата публикации

TEMPERATURE MEASURING ASSEMBLY FOR THE NOZZLE GUIDE RING OF A GAS TURBINE

Номер: GB0002032536B
Автор:

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31-03-1982 дата публикации

GAS TURBINE COMBUSTION CHAMBERS

Номер: GB0002003554B
Автор:
Принадлежит: SNECMA

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29-09-1982 дата публикации

GAS TURBINE STATOR CASING

Номер: GB0002033021B
Автор:

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26-08-2014 дата публикации

APPARATUS TO INJECT A MIXTURE OF AIR AND FUEL, COMBUSTION CHAMBER AND TURBINE ENGINE EQUIPPED WITH SUCH AN APPARATUS

Номер: CA0002582629C
Принадлежит: SNECMA

L'invention concerne un dispositif d'injection d'un mélange d'air et de carburant dans une chambre de combustion de turbomachine pour lequel l'alimentation en air est améliorée. L'invention concerne plus particulièrement un nouveau type de traversée coulissante.

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13-08-2013 дата публикации

MANUFACTURING PROCESS FOR AN INTERTWINED LAYER CONSISTING OF CERAMIC WIRING WITH A METAL MATRIX, DEVICE FOR IMPLEMENTING THE PROCESS AND THE LAYER OBTAINED THROUGH THE PROCESS

Номер: CA0002548630C
Принадлежит: SNECMA

... ²L'invention porte sur un procédé de fabrication d'une nappe liée comportant ²une pluralité de fils enduits (8), qui comprennent une fibre céramique (14) ²enrobée ²d'une gaine métallique (15), caractérisé par le fait que l'on dispose les fils ²les uns à ²côté des autres dans un même plan, et on soude les fils entre eux par points ²par ²soudage laser (13).²L'invention porte également sur le dispositif de mise en uvre du procédé et ²sur la nappe obtenue.²Le procédé s'applique à la fabrication de pièces dans le domaine des ²turbomachines aéronautiques.²² ...

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24-12-2013 дата публикации

IMPROVED TURBINE ENGINE DISTRIBUTOR

Номер: CA0002569564C
Принадлежит: SNECMA

... ²² La présente invention concerne un distributeur de turbine de turbomachine, ²notamment un secteur de distributeur comportant une plate-forme intérieure et ²une plate-forme extérieure, au moins une aube fixée entre lesdites plates-²formes, ²au moins une desdites plates-formes comportant au moins un flasque, ayant une ²première extrémité fixée sur la plate-forme et une seconde extrémité libre, ²ledit ²flasque comportant au moins un évidement d'assouplissement libre non-²débouchant.² ...

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04-09-2012 дата публикации

PROCESS FOR MANUFACTURING A TURBINE STATOR CASING

Номер: CA0002509486C
Принадлежит: SNECMA

L invention a pour objet un procédé de fabrication de carter de stator de turbine comprenant les opérations consistant à ménager entre les parois des parties (O1, O2, O3, E1, E2) d'un moule, une cavité de forme correspondant à celle de l'enveloppe dudit carter ; solidariser des noyaux solubles (10), à au moins une desdites parties de moule (E1), ces noyaux étant maintenus à distance de la paroi de cette partie de moule et matérialisant des espaces libres que l'on souhaite ménager à l'intérieur de ladite enveloppe ; mettre en place, entre les noyaux (10), des inserts (20) solubles matérialisant des chemins de circulation entre lesdits espaces libres ; remplir ladite cavité avec une poudre d'un alliage métallique (24) ; fritter cette poudre (24) par compression isostatique à chaud ; éliminer les noyaux (10) et les inserts (20) par dissolution ; et extraire l'enveloppe ainsi moulée. Application à la fabrication d'un carter de stator de turbine de turboréacteur d'avion.

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17-03-2015 дата публикации

NICKEL-BASED ALLOY

Номер: CA0002583140C

... ²²²L'invention concerne des alliages, ou superalliages, à base de nickel ²(Ni) comprenant essentiellement les éléments suivants dans les teneurs ²indiquées en pourcentages en poids:²Cr : 11,5 à 13,5 %; Co : 11,5 à 16,0 %; Mo : 3,4 à 5,0 %; W : 3,0 à ²5,0 %; Al : 2,2 à 3,2 %; Ti : 3,5 à 5,0 %; Nb : 0,5 à 2,0 %; Hf : 0,25 à ²0,35 %; Zr : 0 à 0,07 %; C : 0,015 à 0,030 %; B : 0,01 à 0,02 %; et Ni : ²complément à 100 %.²Utilisation pour la réalisation de disques de turbine ou de compresseur ²de turbomachines, selon des procédés de métallurgie des poudres.² ...

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08-07-2014 дата публикации

TURBINE ENGINE INLET CONE DEICING SYSTEM FOR AIRCRAFT

Номер: CA0002581540C
Принадлежит: SNECMA

L'invention se rapporte à un système de dégivrage (2) d'un cône d'entrée (4) de turbomoteur pour aéronef, comprenant des moyens de diffusion d'air (18) destinés à équiper le cône d'entrée du turbomoteur afin de lui délivrer de l'air chaud. Selon l'invention, il comporte également un circuit (20) d'évacuation de l'air de pressurisation d'au moins une enceinte-palier du turbomoteur, ce circuit communiquant avec les moyens de diffusion d'air pour pouvoir alimenter ces derniers en air chaud.

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16-09-2014 дата публикации

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

Номер: CA0002589925C
Принадлежит: SNECMA

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.

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09-06-2015 дата публикации

TURBOPROP ENGINE EQUIPPED WITH AN ADJUSTABLE BLADE DIRECTION ASSEMBLY

Номер: CA0002610056C
Автор: GALLET FRANCOIS
Принадлежит: SNECMA

... ²Turbopropulseur comportant un carter tournant muni de pales à ²orientation réglable permettant de gérer la poussée. Chaque pale est ²couplée, pour le réglage de son orientation à un organe d'actionnement ²d'un vérin annulaire porté par le carter tournant.²²²² ...

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10-03-2015 дата публикации

METHOD FOR REDUCING SPEED IN THE CASE OF THE BREAKAGE OF A TURBINE SHAFT IN A GAS TURBINE ENGINE

Номер: CA0002607244C
Автор: MONS CLAUDE MARCEL
Принадлежит: SNECMA

... ²² La présente invention porte sur une méthode pour réduire la vitesse de ²rotation, dans un moteur à turbine à gaz, d'une turbine comprenant un ²rotor (6) entraînant un arbre (8) et mobile en rotation à l'intérieur d'un ²stator (7), en cas de rupture dudit arbre. Cette méthode est caractérisée ²par le fait qu'elle consiste à mesurer (11) la température en un point sur ²une surface du stator (10A) située en aval du rotor (6), transmettre le ²signal de mesure à un moyen de commande (100) du freinage du rotor, ²ledit moyen de commande étant agencé pour commander le freinage du ²rotor lorsque la température atteint un seuil.²L'invention porte également sur le dispositif.² ...

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10-01-2012 дата публикации

TURBINE ENGINE HIGH-PRESSURE TURBINE STATOR AND ASSEMBLY PROCESS

Номер: CA0002500493C
Принадлежит: SNECMA

Procédé d'assemblage d'éléments sectorisés (14, 24) d'un stator annulaire (10) d'une turbine haute-pression de turbomachine autour d'un axe longitudinal (X-X) de ladite turbine, procédé selon lequel, on définit un motif de répartition angulaire des éléments du stator pour un secteur angulaire prédéterminé (.PSI.), le motif de répartition étant défini de façon à éviter un alignement radial entre des zones inter-secteurs (14a, 24a) des éléments du stator (14, 24) définies entre deux secteurs adjacents d'un même élément du stator, et à répéter ledit motif de répartition sur toute la circonférence du stator.

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31-03-2015 дата публикации

FUEL INJECTOR DEVICE FOR A TURBINE ENGINE

Номер: CA0002619352C
Принадлежит: SNECMA

Dispositif d'injection de carburant dans une turbomachine telle qu'un turboréacteur d'avion, comprenant une pompe haute-pression (10) alimentant un moyen (16) de réglage de débit dont la sortie est reliée par un clapet (20) de pressurisation et de coupure à un conduit (22) d'alimentation des injecteurs de carburant (18), le clapet (20) étant relié à l'entrée et à la sortie de la pompe (10) pour définir deux seuils de pressurisation du carburant, l'un servant pour le démarrage et le redémarrage de la turbomachine et l'autre pour le fonctionnement de la turbomachine à partir du régime de ralenti et pour la commande d'équipements à géométrie variable (24).

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31-12-2013 дата публикации

METHOD FOR PRODUCING A RIM AT THE FREE EDGE OF A VANE, VANE OBTAINED BY THIS METHOD AND TURBINE ENGINE EQUIPPED WITH SAID VANE

Номер: CA0002567885C
Автор: MONS CLAUDE, VIGNEAU JOEL
Принадлежит: SNECMA

... ²²Selon ce procédé, on fournit une aube présentant à son extrémité libre au ²moins une ²base de rebord, et on construit la partie saillante du rebord par le dépôt ²successif de ²couches sur la base , par réalisation des étapes suivantes :²- on active une source laser reliée à une tête optique focalisée sur un ²point de la ²surface du sommet de la base et une source de poudre reliée à une buse de ²projection, ce par quoi on forme un bain de fusion localisé au niveau dudit ²point, ²dans lequel est injecté la poudre d'où il en résulte la formation d'une ²surépaisseur ²localisée ; et²- on règle la tête optique et la buse sur un autre point adjacent à ladite ²surépaisseur ²et on retourne à l'étape précédente jusqu'à la formation d'une couche sur ²sensiblement toute la base.² ...

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23-07-2013 дата публикации

STAGE OF STRAIGHTENING BLADES ACTIVATED BY AN ELECTRICALLY-DRIVEN ROTATING CROWN

Номер: CA0002530137C
Автор: BOURU MICHEL ANDRE
Принадлежит: SNECMA

Etage d'aubes de redresseur à calage variable, lesdites aubes étant déplacées par une couronne d'actionnement rotative à centrage automatique. Le carter (11) comporte un rail annulaire (24) coaxial fixe, faisant saillie de sa surface externe et on prévoit trois équipages mobiles (26) espacés circonférenctiellement et assujettis à se déplacer le long dudit rail, chaque équipage mobile étant couplé à la couronne par un agencement de guidage radial.

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02-07-2013 дата публикации

AIRFOIL AXIAL RETAINING RING LOCKING DEVICE, THE ASSOCIATED ROTOR DISK AND RETENTION RING, AND THE ROTOR AND AIRCRAFT ENGINE COMPRISING THEM

Номер: CA0002557765C
Принадлежит: SNECMA

... ²Le dispositif d'immobilisation en rotation d'un anneau ²de rétention (20) sur un disque de rotor (10) comporte:²un premier (162, 262) et un deuxième (164, 264) ²crochets d'immobilisation successifs du disque (10),²au moins un taquet (302, 304, 306) dudit anneau de ²rétention (20), disposé près d'une fente (24) de l'anneau (20).²La position dudit au moins un taquet (302, 304, 306) ²sur ledit anneau de rétention (20) est telle que, lorsque ledit ²anneau de rétention (20) est en place dans une gorge (22) située ²sur le disque de rotor (10), ledit au moins un taquet (302, 304, ²306) est en butée contre ledit premier crochet d'immobilisation ²(162, 262), et la fente (24) est recouverte par ledit deuxième ²crochet d'immobilisation (164, 264).² ...

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28-01-2014 дата публикации

DEVICE FOR FASTENING RING SECTORS AROUND A TURBINE WHEEL OF A TURBINE ENGINE

Номер: CA0002582401C
Принадлежит: SNECMA SERVICES, SNECMA

... ²Dispositif de fixation de secteurs d'anneau autour d'une roue de ²turbine dans une turbomachine, chacun des secteurs d'anneau comprenant ²à son extrémité amont un rebord circonférentiel pouvant être maintenu sur ²un rail annulaire de carter par un organe annulaire de verrouillage, et à son ²extrémité aval une pièce pouvant venir en appui axial contre un élément fixe ²de la turbine pour empêcher le désengagement du rebord amont du secteur ²d'anneau de l'organe de verrouillage en cas d'usure importante du rail.²²² ...

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24-11-2009 дата публикации

TURBINE BLADE HAVING A COOLING AIR DEFLECTOR

Номер: CA0002444862C
Принадлежит: SNECMA

L'invention concerne une aube (11) pour turbine, l'aube présentant un pied (13) permettant de la rapporter dans une alvéole (14) d'un disque (12) de la turbine, l'aube possédant un circuit interne de refroidissement par air comprenant des moyens d'entrée d'air (15) situés sur le pied de l'aube et en regard de l'alvéole, et des moyens de sortie d'air. Le pied (13) de l'aube est pourvu de moyens (20) d'homogénéisation de la pression et de la température de l'air de refroidissement pénétrant dans les moyens d'entrée d'air.

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28-01-2014 дата публикации

TURBINE BLADE WITH IMPROVED COOLING CHARACTERISTICS AND USEFUL LIFE

Номер: CA0002569563C
Принадлежит: SNECMA

... ²²²²La présente invention concerne le domaine des aubes de turbine de ²turbomachine, notamment une aube de turbine (1), comportant une paroi ²intrados (2), une paroi extrados (3), au moins une première cavité radiale ²de bord de fuite (4), au moins une seconde cavité radiale (5) en amont de ²la cavité de bord de fuite (4), une paroi interne (6) séparant les cavités ²radiales (4 et 5) et comprenant au moins un canal (7) reliant les cavités (4 ²et 5) entre elles, ledit canal (7) étant orienté selon un axe (71) coupant la ²surface interne (42) de la paroi intrados (2).² ...

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20-09-2011 дата публикации

PRODUCTION OF TURBINES FOR TURBINE ENGINES HAVING BLADES WITH DIFFERENT ADJUSTED RESONANCE FREQUENCIES AND PROCESS FOR ADJUSTING THE RESONANCE FREQUENCY OF A TURBINE BLADE

Номер: CA0002457256C
Принадлежит: SNECMA

Pour la ou chaque roue mobile de la turbine, on utilise des aubes (40) comportant chacune au moins deux pales (42, 44) réunies à une plate-forme externe (46), une plate-forme interne (48) et un pied (60) communs. On utilise des aubes ayant un pied creux (60) dans lequel est formé un évidemment (62), et on confère volontairement aux pieds d'aubes (60) appartenant à une même roue et/ou aux pieds d'aubes appartenant à deux roues différentes des configurations différentes au niveau des évidements de leurs pieds pour ajuster les fréquences de résonance propres de ces aubes à des valeurs sensiblement différentes et réaliser ainsi un désaccordage entre aubes d'une même roue et/ou de deux roues différentes. L'invention est applicable aux turbines aéronautiques et industrielles.

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28-10-2014 дата публикации

COMPOSITE TURBINE ENGINE BLADE WITH METAL REINFORCEMENT

Номер: CA0002603003C
Принадлежит: SNECMA

L'invention concerne une aube de turbomachine comportant une surface aérodynamique (12) réalisée en un matériau composite qui s'étend selon une première direction (14) entre un bord d'attaque (16) et un bord de fuite et selon une deuxième direction entre un pied et un sommet de l'aube. L'aube comporte un renfort métallique plein (32) qui est collé au bord d'attaque (16) de la surface aérodynamique de l'aube, qui s'étend selon la première direction (14) au-delà du bord d'attaque de la surface aérodynamique et selon la deuxième direction entre le pied et le sommet, et qui comporte au moins un évidement (34a) destiné à absorber une partie au moins de l'énergie résultant de l'impact d'un corps étranger sur le bord d'attaque de l'aube.

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12-10-2010 дата публикации

EMERGENCY DEVICE FOR RESTARTING A TURBOJET IN AUTOROTATION

Номер: CA0002389780C
Принадлежит: SNECMA

Dispositif de secours au rallumage d'un turboréacteur en autorotation comportant une soufflante (2) entraînée par une turbine basse pression par l'intermédiaire d'un premier arbre (3), un compresseur entraîné par une turbine haute pression par l'intermédiaire d'un second arbre (4) disposé coaxialement par rapport au premier arbre, un différentiel (5) reliant les premier et second arbres en compensant leurs différences de vitesse de rotation lors du fonctionnement normal du turboréacteur, et un système de freinage (6) relié au différentiel afin de pouvoir ralentir ou bloquer celui-ci lorsque le turboréacteur s'éteint, de sorte que le premier arbre (3) entraîne alors le second arbre (4) pour que ce dernier atteigne un régime favorisant le rallumage du turboréacteur.

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05-01-2012 дата публикации

TURBINE ENGINE WITH NONSTREAMLINED IMPELLERS

Номер: US20120000177A1
Автор:
Принадлежит: SNECMA

The present invention relates to a turbine engine comprising two respectively upstream and downstream external impellers () that are nonstreamlined, coaxial and contrarotating. The engine is noteworthy in that the downstream impeller () is retractable so as to reduce its diameter. The blades () of the downstream impeller are mounted so as to pivot about a pivot (), the axis of which forms a nonzero angle, notably perpendicular, with the axis () of rotation of the impeller, the blades in the retracted position being tilted about the pivot (). 1. A turbine engine comprising two respectively upstream and downstream external impellers that are nonstreamlined , coaxial and contrarotating , wherein the downstream impeller is retractable so as to reduce its diameter and wherein it comprises a mechanism for actuating the downstream impeller in the retracted position with a braking means for braking the rotation of the impeller and a means for actuating the blades into a retracted position.2. The turbine engine as claimed in the preceding claim wherein the blades of the downstream impeller are mounted so as to pivot about a pivot , the axis of which forms a nonzero angle , notably perpendicular , with the axis of rotation of the impeller , the blades in the retracted position being tilted about the pivot.3. The turbine engine as claimed in claim 1 , wherein the actuation mechanism comprises a means for compensating for the braking of the downstream impeller with an acceleration of the upstream impeller.4. The turbine engine as claimed in claim 1 , wherein the means for actuating the blades comprises springs acting in opposition to the centrifugal force.5. The turbine engine as claimed in claim 1 , wherein the actuation means comprises cylinders.6. The turbine engine as claimed in claim 1 , wherein the brake is arranged between the turbine rotor driving the downstream impeller and the shaft of the downstream impeller so as to reduce the rotation speed of the impeller relative ...

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05-01-2012 дата публикации

Compressor casing with optimized cavities

Номер: US20120003085A1
Принадлежит: SNECMA SAS

The invention relates to a compressor for a turbine engine including a casing ( 4 ), at least one compressor stage consisting of a stationary blade ( 2 ) impeller and a mobile blade ( 1 ) impeller positioned upstream from said stationary blade ( 2 ) impeller, and cavities ( 5 ) made in said casing opposite the through-path of the mobile blades ( 1 ), said cavities having a length L 2 measured axially and being shifted upstream relative to the blades ( 1 ) so as to generate an overlap with a length L 1 , characterised in that the lengths L 1 and L 2 are respectively between 35% and 50% and between 80% and 90% of the axial chord C ax measured at the outer end of the blades ( 1 ), and in that the cavities ( 5 ) do not in communication with one another.

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19-01-2012 дата публикации

TURBOJET ENGINE WITH AN ELEMENT OF THE NACELLE ATTACHED TO THE INTERMEDIATE CASING

Номер: US20120011826A1
Принадлежит: SNECMA

The present invention relates to a turbojet engine with a fan () and a nacelle element () forming the secondary duct downstream of the fan, the turbojet engine comprising an intermediate casing () with an outer shroud () to which said nacelle element () is fixed by means of a connecting member, wherein the connecting member comprises a cylindrical element () secured to the shroud () while at the same time being capable of a rotational movement with respect to the axis of the engine and connected to said nacelle element () by a bayonet-type fitting. 1. A turbojet engine with a fan and a nacelle element forming the secondary duct downstream of the fan , the turbojet engine comprising an intermediate casing with an outer shroud to which said nacelle element is fixed by means of a connecting member , wherein the connecting member comprises a cylindrical element secured to the shroud while at the same time being capable of a rotational movement with respect to the axis of the engine and connected to said nacelle element by a bayonet-type fitting.2. The turbojet engine as claimed in the preceding claim , the bayonet-type fitting comprising at least one transverse groove extending over an arc of a circle , formed in one out of the cylindrical element and the nacelle element and collaborating with a radial flange portion formed in the other out of the cylindrical element and the nacelle element.3. The turbojet engine as claimed in claim 1 , a bearing being inserted between the cylindrical element and the outer shroud of the intermediate casing.4. The turbojet engine as claimed in claim 1 , the nacelle element being capable of axial movement once it has been detached from the cylindrical element.5. The turbojet engine as claimed in the preceding claim claim 1 , the nacelle element comprising a first guideway element capable of collaborating with a second guideway element that complements the first and is secured to a pylon by means of which the turbojet engine is attached to ...

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19-01-2012 дата публикации

TURBINE VANE WITH DUSTING HOLE AT THE BASE OF THE BLADE

Номер: US20120014810A1
Принадлежит: SNECMA

A cooled turbine vane for a turbine engine, that includes a blade mounted on a platform carried by a base, the blade including one or more cavities formed therein for cooling air circulation, the cavity extending along the trailing edge and being supplied with cooling air by a supply duct connecting an air intake located in a lower portion of the base and the cavity of the trailing edge by defining a bend within the base. The duct includes, on an axis substantially radial relative to the air intake a bell-shaped niche located under the platform, the niche being open at a top thereof via a dusting hole extending through the platform and being defined at a foot of the base by walls extending substantially radially from the platform to close the platform laterally. 114-. (canceled)15. A cooled turbine vane for a turbomachine comprising:a blade mounted on a platform supported by a root, the blade including one or more cavities cut into it for circulation of cooling air, the cavity extending along a trailing edge being supplied with cooling air by a supply duct connecting an air inlet situated in a lower part of the root to the trailing edge cavity, making a bend within the root;wherein the duct comprises, along an axis that is substantially radial with respect to the air inlet, a niche situated under the platform and in a shape of a bell, the niche opening at its top via a dusting hole that passes through the platform and being delimited inside the root by walls extending substantially radially from the platform to close the platform laterally.16. The turbine vane as claimed in claim 15 , further comprising at least one ventilation perforation for ventilating the platform claim 15 , the perforation being supplied with cooling air from the supply duct claim 15 , wherein the dusting hole coincides with the ventilation perforation of the platform.17. The turbine vane as claimed in claim 15 , in which at least one of the walls of the niche extending radially from the ...

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26-01-2012 дата публикации

WORKSHOP FOR PREPARING AIRCRAFT ENGINES FOR SHIPPING

Номер: US20120017431A1
Автор:
Принадлежит: SNECMA

A workshop for preparing aeroengines for shipping, the workshop including a station for fitting test and measurement mechanism on an engine, a mechanism for transporting the engine to test premises and for returning the engine to the workshop, a station for removing the test and measurement mechanism, a station for endoscopic inspection, a finishing station, and a shipping station. The engines are transported from one station to another by spreaders fastened on the engines and attached to hoists that are movable in translation along an overhead structure arranged in the workshop, each station being fitted with computer terminals for displaying and tracking the tasks to be performed on the engine in the corresponding station. 115-. (canceled)16. A workshop for preparing aeroengines for shipping , the workshop comprising:a plurality of stations including: a fitting station for fitting measurement and test means on each engine; a removal station for removing the measurement and test means; an inspection station for inspecting the engines by endoscopy; a finishing station for fitting final pieces of equipment on each engine; and a shipping station for fitting each engine on a shipping structure and for placing protective coverings on the engine, the plurality of stations being arranged in a fixed manner in the workshop along a predetermined path for the engines;conveyor and handling means for conveying and handling the engines, which means are carried by an overhead structure extending over the plurality of stations and including placing-and-taking means for placing engines on and for taking engines off stationary supports installed in the plurality of stations and configured to give direct access to all portions of the engine, and travel means for moving the engines from one station to another; andcontrol means for controlling the conveyor means so that a departure of an engine from any of the plurality of stations is followed substantially without delay by arrival of ...

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26-01-2012 дата публикации

DOUBLE-BODY GAS TURBINE ENGINE PROVIDED WITH AN INTER-SHAFT BEARING

Номер: US20120017603A1
Принадлежит: SNECMA

A double-body gas turbine engine, including a low-pressure body LP and a high-pressure body HP, rotatably mounted about a single shaft in a stationary casing, the low-pressure body LP including a compressor and a turbine connected by a low-pressure shaft LP. The low-pressure shaft LP is supported by an upstream LP bearing, a first downstream LP bearing, and an additional downstream LP bearing by the stationary casing, the high-pressure body being supported by an upstream HP bearing and a downstream HP bearing which is an inter-shaft bearing including an inner track rigidly connected to the HP turbine rotor and an outer track rigidly connected to the LP shaft. 18-. (canceled)9. A twin-spool gas turbine engine comprising:a low-pressure LP spool and a high-pressure HP spool, which spools are mounted so that they can rotate about a same axis in a fixed casing,the low-pressure LP spool including a compressor and a turbine which are connected by a low-pressure LP shaft, the low-pressure LP shaft being supported by an upstream LP bearing, a first downstream LP bearing, and an additional downstream LP bearing by the fixed casing,the high-pressure HP spool being supported by an upstream HP bearing and a downstream HP bearing,wherein the downstream HP bearing is an inter-shaft bearing comprising an inner raceway secured to the HP turbine rotor and an outer raceway secured to the LP shaft.10. The engine as claimed in claim 9 , in which the additional downstream LP bearing is located upstream of the downstream LP bearing.11. The engine as claimed in claim 9 , in which the additional downstream LP bearing is of a diameter greater than that of the downstream LP bearing.12. The engine as claimed in claim 11 , in which the additional downstream LP bearing is of a diameter greater than that of the downstream HP bearing.13. The engine as claimed in claim 12 , in which the additional downstream LP bearing is positioned axially between the downstream HP bearing and the downstream LP ...

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26-01-2012 дата публикации

ROTOR BLADE OF A GAS TURBINE ENGINE MADE OF COMPOSITE MATERIAL COMPRISING A CONNECTING YOKE, METHOD FOR MANUFACTURING THE BLADE

Номер: US20120018079A1
Автор: Gallet Francois
Принадлежит: SNECMA

A rotor blade of a turbine engine comprising a body made of composite material consisting of a fibre-reinforced thermosetting resin and a connecting yoke designed to be attached to a fastener of a rotor disc of the said turbine engine, the said blade characterized in that the connecting yoke comprises a metal reinforcement and a casing made of composite material encasing the said reinforcement. 1) Method for manufacturing a rotor blade of a gas turbine engine comprising a body made of composite material and a connecting yoke designed to be attached to a fastener of a rotor disc of the said turbine engine , the said method comprising the following steps:a step of installing fibres around a metal tube and a step of injecting thermosetting resin in order to form a casing of composite material around the metal tube, the tube comprising at least two sections having a high density of fibres and at least one section having a low density of fibres arranged between the two sections having a high density of fibres; anda step of machining the section having a low density of fibres so as to form a connecting yoke for the blade comprising two sections having a high density of fibres, forming a reinforcement of the connecting yoke, encased in composite material.2) Method according to claim 1 , in which claim 1 , the body of the blade made of composite material consisting of a fibre-reinforced thermosetting resin claim 1 , the method comprises a step of placing the fibres of the blade body around the metal tube in order to form the casing made of composite material around the metal tube.3) Method according to claim 1 , comprising a step of weaving fibres around the metal tube claim 1 , the weave comprising warp threads placed along the cord of the said blade and weft threads placed along the height of the blade. The field of the invention is that of gas turbine engines for aircraft. The invention relates more particularly to a device for fastening fan blades made of composite ...

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26-01-2012 дата публикации

DIFFUSER/RECTIFIER ASSEMBLY FOR A TURBINE ENGINE

Номер: US20120018543A1
Принадлежит: SNECMA

A diffuser-nozzle assembly for mounting at an outlet from a compressor in a turbomachine, the assembly including a nozzle including two substantially cylindrical walls, a radially inner wall and a radially outer wall that are connected together by radial vanes. The walls of the nozzle are extended downstream beyond the radial vanes and the radial spacing therebetween varies downstream from the vanes to be at a minimum substantially in register with the vanes and at a maximum between the vanes. 111-. (canceled)12. A diffuser-nozzle assembly for mounting at an outlet from a compressor in a turbomachine , the assembly comprising:a nozzle including two substantially cylindrical walls, a radially inner wall and a radially outer wall that are connected together by radial vanes,wherein the walls of the nozzle are extended downstream beyond the radial vanes and radial spacing therebetween varies circumferentially downstream from the vanes to be at a minimum substantially in register with the vanes and at a maximum between the vanes.13. A diffuser-nozzle assembly according to claim 12 , wherein the walls of the nozzle diverge going downstream from the vanes.14. A diffuser-nozzle assembly according to claim 12 , wherein at least one of the walls of the nozzle is corrugated or crenellated downstream from the vanes.15. A diffuser-nozzle assembly according to claim 12 , wherein both walls of the nozzle are corrugated or crenellated downstream from the vanes.16. A diffuser-nozzle assembly according to claim 14 , wherein the corrugations of the downstream portion(s) of the wall(s) of the nozzle are symmetrical relative to radial planes extending the vanes.17. A diffuser-nozzle assembly according to claim 14 , wherein the corrugations of the downstream portion(s) of the wall(s) of the nozzle are asymmetrical relative to radial planes extending the vanes.18. A diffuser-nozzle assembly according to claim 17 , wherein the corrugations extend helically around the axis of the compressor ...

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26-01-2012 дата публикации

ASSEMBLY BETWEEN A COMPRESSOR SHAFT TRUNNION AND A BEVEL GEAR FOR DRIVING AN ACCESSORY GEARBOX OF A TURBOMACHINE

Номер: US20120020774A1
Принадлежит: SNECMA

The invention relates to an assembly between a compressor shaft trunnion and a bevel gear for driving an accessory gearbox of a turbomachine. The assembly comprises a bevel gear, a compressor shaft trunnion arranged coaxially inside the bevel gear, torque transmission means between the compressor shaft trunnion and the bevel gear, a lock nut for preventing the bevel gear from moving axially on the compressor shaft trunnion, the lock nut being suitable for being screwed onto the bevel gear and including a shoulder projecting radially outwards and suitable for coming into axial abutment firstly upstream against a stop nut screwed onto an upstream end of the compressor shaft trunnion, and secondly downstream against a bearing surface of the compressor shaft trunnion, and an anti-rotation pin for preventing the lock nut from rotating. 1. An assembly between a compressor shaft trunnion and a bevel gear for driving an accessory gearbox of a turbomachine , the assembly comprising:a bevel gear centered on a longitudinal axis of the turbomachine;a compressor shaft trunnion disposed coaxially inside the bevel gear;torque transmission means between the compressor shaft trunnion and the bevel gear;a lock nut for preventing the bevel gear from moving axially on the compressor shaft trunnion, the lock nut being suitable for being screwed into the bevel gear on its inside and including a shoulder that projects radially outwards and that is suitable for coming into axial abutment firstly upstream against a stop nut screwed onto an upstream end of the compressor shaft trunnion, and secondly downstream against a bearing surface of said compressor shaft trunnion; andan anti-rotation pin mounted inside the compressor shaft trunnion to prevent the lock nut from rotating.2. An assembly according to claim 1 , further comprising means for preventing the anti-rotation pin from moving axially on the compressor shaft trunnion.3. An assembly according to claim 2 , wherein the anti-rotation pin ...

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26-01-2012 дата публикации

RETAINING RING

Номер: US20120020796A1
Принадлежит: SNECMA

A retaining ring presenting, on its periphery, a plurality of radial bores, each of the bores being able to receive a blade root. The retaining ring also includes at least one first radial opening traversed by a compression tube and at least one second radial opening, the at least first radial opening and the at least second radial opening being separated from each other by at least one of the bores of the plurality of radial bores. The ventilation device finds a particularly interesting application in the field of turbine engines including a pusher open rotor. 1. A retaining ring comprising:a plurality of radial bores arranged on a periphery of the retaining ring, each of said bores being configured to receive a blade root;at least one first radial opening traversed by a compression tube and at least one second radial opening, said at least one first radial opening and said at least one second radial opening being separated from each other by at least one of the bores of said plurality of radial bore.2. The retaining ring according to claim 1 , wherein said first radial opening is positioned upstream from said retaining ring claim 1 , and said second radial opening is positioned downstream from said retaining ring.3. The retaining ring according to claim 1 , comprising a mask covering at least one blade root claim 1 , said mask presenting a cavity to allow a flow to circulate under said retaining ring.4. The retaining ring according to claim 3 , wherein said mask forms an internal ring with a U-shaped section claim 3 , the upper part of said U-shaped section being covered by said retaining ring.5. The retaining ring according to claim 3 , wherein each root is covered by a mask.6. The retaining ring according to claim 1 , wherein for each radial bore claim 1 , a first opening and a second opening are placed opposite each other claim 1 , said first opening and said second opening are separated from each other by said radial bore.7. The retaining ring according to ...

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26-01-2012 дата публикации

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

Номер: US20120020799A1
Автор:
Принадлежит: SNECMA

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

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26-01-2012 дата публикации

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

Номер: US20120020800A1
Принадлежит: SNECMA

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

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26-01-2012 дата публикации

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

Номер: US20120020806A1
Автор:
Принадлежит: SNECMA

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

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02-02-2012 дата публикации

CIRCUIT AND A METHOD FOR FEEDING OIL TO ROLLING BEARINGS IN A TURBOMACHINE

Номер: US20120024631A1
Принадлежит: SNECMA

The invention relates to a circuit and a method for feeding oil to rolling bearings in a turbomachine, the circuit comprising an oil supply, an oil duct connecting the supply to lubrication enclosures, each containing at least one rolling bearing, an oil feed pump located between the supply and the lubrication enclosures, an oil recovery pump connected downstream from each lubrication enclosure, and a pressurizing valve placed in the duct between the oil feed pump and the lubrication enclosures, said pressurizing valve being adapted to open when the pressure at its inlet exceeds a predetermined threshold pressure; a branch connection duct connecting the inlet of the pressurization valve to one of the lubrication enclosures in order to feed it with oil; and means for feeding the lubrication enclosures at a small oil flow rate so long as the pressure at the inlet of the pressurization valve has not reached the predetermined threshold. 1. A circuit for feeding oil to rolling bearings in a turbomachine , the circuit comprising:an oil supply;an oil duct connecting the supply to a plurality of lubrication enclosures connected in parallel and each containing at least one rolling bearing;an oil feed pump disposed in the duct between the supply and the lubrication enclosures in order to take oil from the supply and deliver it to the rolling bearings; andan oil recovery pump connected downstream from each lubrication enclosure in order to recover the oil that has been fed to the rolling bearings and return it to the supply;wherein the circuit further comprises:a pressurizing valve placed in the duct between the oil feed pump and the lubrication enclosures, said pressurizing valve being adapted to open when the pressure at its inlet exceeds a predetermined threshold pressure;a branch connection duct connecting the inlet of the pressurization valve to one of the lubrication enclosures in order to feed it with oil; andmeans for feeding the lubrication enclosures at a small oil ...

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02-02-2012 дата публикации

Turbomachine blade, a rotor, a low pressure turbine, and a turbomachine fitted with such a blade

Номер: US20120027605A1
Принадлежит: SNECMA Propulsion Solide SA, SNECMA SAS

The invention relates to a turbomachine blade made of composite material and presenting a root with a bulb-shaped end suitable for engaging in a slot of a rotor disk. In characteristic manner, the end of the root of the blade is provided, beside one of its front faces, with a projecting portion having two symmetrical fins about the axial midplane of the root, each fin having a bearing face suitable for limiting tilting of the blade relative to the rotor disk about the axial direction.

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02-02-2012 дата публикации

METHOD FOR FABRICATING A THERMAL BARRIER COVERING A SUPERALLOY METAL SUBSTRATE, AND A THERMOMECHANICAL PART RESULTING FROM THIS FABRICATION METHOD

Номер: US20120028056A1
Принадлежит: SNECMA

A fabrication method of fabricating a thermal barrier covering a superalloy metal substrate, the thermal barrier including at least an underlayer and a ceramic layer, the method including: smoothing a surface state of the underlayer by at least one physicochemical and/or mechanical process prior to depositing the ceramic layer such that a number of defects presenting a peak-to-peak difference lower than or equal to 2 μm is at most five over any distance of 50 μm, and then depositing the ceramic layer. The method can be applied to turbine blades. 112-. (canceled)13. A fabrication method of fabricating a thermal barrier covering a superalloy metal substrate , the thermal barrier including at least an underlayer and a ceramic layer , the method comprising:smoothing a surface state of the underlayer by at least one physicochemical and/or mechanical process prior to depositing the ceramic layer such that a number of defects presenting a peak-to-peak difference greater than or equal to 2 μm is at most five over any distance of 50 μm; andthen depositing the ceramic layer.14. A fabrication method according to claim 13 , wherein the physicochemical and/or mechanical process gives rise to a surface state of the underlayer such that a number of defects presenting an amplitude greater than 1 μm relative to the mean position of a top face of the underlayer is at most five over any distance of 50 μm.15. A fabrication method according to claim 13 , wherein the physicochemical and/or mechanical process gives rise to a surface state of the underlayer such that roughness Ra of the underlayer is in a range of 0.05 μM to 3 μm.16. A fabrication method according to claim 13 , wherein the physicochemical and/or mechanical process gives rise to a surface state of the underlayer such that roughness Ra of the underlayer is in a range of 0.05 μm to 1 μm.17. A fabrication method according to claim 13 , wherein the physicochemical and/or mechanical process gives rise to a surface state of the ...

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09-02-2012 дата публикации

Sealing head for an installation for fluid tests on an aircraft turbine engine part

Номер: US20120031177A1
Принадлежит: SNECMA SAS

A sealing head for an installation configured to perform fluid tests on an aircraft turbine engine part. The sealing head includes a sealing element crossed by a passage for flowing of a gas flow, the sealing element including a sealing surface configured to come into contact with a part to be tested, at an aperture of the part to be tested configured to be fed with the gas flow. The sealing head includes a mechanism centering the sealing element relatively to the aperture of the part, the centering mechanism being firmly attached to the sealing element and protruding towards the front of the surface having an outer portion surrounding the centering mechanism.

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09-02-2012 дата публикации

METHOD OF DETECTING AN ICING STATE OR A NEED FOR MAINTENANCE IN A TURBOMACHINE FUEL CIRCUIT

Номер: US20120032809A1
Принадлежит: SNECMA

Detecting a state of icing in a turbomachine fuel circuit includes reading information representative of a first temperature Tof the fuel downstream from the metering unit, and of comparing the first temperature Twith a first reference temperature T. The method also includes detecting clogging of the filter unit. In the event of the temperature reading Tbeing less than the first reference temperature Tand of clogging being detected, a signal indicative of an icing state in the fuel circuit is issued. Information representative of a second fuel temperature Tin the tank may be read and compared with a reference temperature Tin order to confirm or contradict the icing state. A signal indicating a need for maintenance of the fuel circuit is issued when the temperature readings Tand Tare not less than the reference temperatures Tand Tand clogging has been detected. 17-. (canceled)8. A method of detecting a state of icing in a turbomachine fuel circuit , said circuit comprising at least a tank , a filter unit for filtering the fuel , a high pressure pump connected to the tank via the filter unit and a fuel metering unit connected to the outlet from the high pressure pump in order to control the flow rate of fuel for injection into the combustion chamber , the method comprising:{'sub': 1', '1', '01, 'reading information representative of a first temperature Tof the fuel in the fuel circuit downstream from the metering unit, and of comparing the first temperature Twith a first reference temperature T;'}detecting clogging of the filter unit; and{'sub': 1', '01, 'in the event of the temperature reading Tbeing less than the first reference temperature Tand of clogging being detected, issuing a signal indicative of an icing state in the fuel circuit.'}9. The method according to claim 8 , further comprising:{'sub': 2', '2', '02, 'reading information representative of a second temperature Tof the fuel in the tank, and of comparing the second temperature Twith a second reference ...

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16-02-2012 дата публикации

ROTATING INLET COWL FOR A TURBINE ENGINE, COMPRISING AN ECCENTRIC FORWARD END

Номер: US20120036827A1
Автор:
Принадлежит: SNECMA

A rotating inlet cowl for a turbine engine, having a rotation axis and for which the forward end is arranged to be eccentric relative to this rotation axis. Furthermore, a forward cone of the cowl is truncated by a truncation surface defining the forward end of the inlet cowl. 16-. (canceled)7. A rotating inlet cowl for a turbine engine , the cowl having a rotation axis and comprising:a forward cone defining a forward end of the inlet cowl,wherein the forward end is configured to be eccentric relative to the rotation axis of the inlet cowl, andwherein the forward cone is truncated by a truncation surface defining the forward end of the inlet cowl.8. A cowl according to claim 7 , wherein the forward cone is oblique with an axis inclined from the axis of rotation of the inlet cowl.9. A cowl according to claim 7 , wherein the forward cone has an axis parallel to and coincident with the inlet cowl rotation axis.10. A cowl according to claim 7 , wherein the truncation surface is approximately plane claim 7 , inclined relative to a plane orthogonal to the axis of rotation of the inlet cowl.11. A cowl according to claim 7 , wherein the front cone is forward from a rear shroud.12. A turbine engine claim 7 , or an aircraft engine claim 7 , comprising a rotating inlet cowl according to . This invention relates to the field of turbine engines, and more particularly to aircraft turbine engines, preferably of the turbojet type. More specifically, the invention relates to the rotating inlet cowl fitted on these turbine engines.Such a rotating inlet cowl is usually composed of two parts fixed to each other, the forward part in the form of a cone and the aft part in the form of a shroud. The aft end of the aft shroud is flush with the fan blade platforms in a known manner, and is in aerodynamic continuity with them and in front of them.The front cone has a forward end shaped like a cone tip centred on the axis of rotation of the inlet cowl, also corresponding to the longitudinal ...

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16-02-2012 дата публикации

ELBOW BRACKET FOR AERONAUTICAL EQUIPMENT

Номер: US20120037764A1
Автор:
Принадлежит: SNECMA

An elbow bracket for cables or pipes, supported by a structure of an aeronautical device, in which a reinforcing rib of the elbowed main plate is not assembled with the elbowed main plate by welding but by tightening following a twisting of end-pieces traversing slots of the main plate and causing an outer face of the plate to be trapped against an upper face of oblong slits partially cutting into the end-pieces. The support is faster to construct, and has sufficient resistance. 13-. (canceled)4. An elbow bracket for aeronautical equipment , comprising:a main plate including a side to be attached to a supporting structure and a side for carrying equipment, together with a reinforcing rib attached to inner faces of the sides by edges,wherein the reinforcing rib is separate from the main plate and includes end-pieces traversing slots of the plate, wherein the end-pieces are partially cut into by oblique slits extending partially in the slots to bottom portions of the slits and partially outside the slots, above outer faces of the sides, to emerging portions of the slits, andwherein the end-pieces are twisted at end portions and protrude beyond the slots when the end-pieces engage on the outer faces of the sides.5. An elbow bracket for aeronautical equipment according to claim 4 , wherein the slots are aligned and oblong claim 4 , lengthened in a direction of alignment of the slots claim 4 , and wherein plays in the slots and around the end-pieces are larger in the alignment direction than in the perpendicular direction.6. An elbow bracket for aeronautical equipment according to claim 4 , wherein there are three slots and three end-pieces claim 4 , two of which are on one of the sides and one of which is on the other side.7. An elbow bracket for aeronautical equipment according to claim 5 , wherein there are three slots and three end-pieces claim 5 , two of which are on one of the sides and one of which is on the other side. The invention relates to an elbow bracket ...

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16-02-2012 дата публикации

INTERMEDIATE CASING OF AIRCRAFT TURBOMACHINE INCLUDING STRUCTURAL CONNECTING ARMS WHICH PERFORM SEPARATE MECHANICAL AND AERODYNAMIC FUNCTIONS

Номер: US20120039710A1
Автор:
Принадлежит: SNECMA

A structural connecting arm for an intermediate casing of an aircraft turbomachine with a ducted fan, wherein the arm is configured to connect a hub and an outer ferrule of the casing, and including an aerodynamic outer surface manufactured such that the arm also forms an outlet guide vane. Multiple metal ties extend in the longitudinal direction of the arm, together with a shell made from composite material surrounding the ties and forming the aerodynamic outer surface. 19-. (canceled)10. A structural connecting arm for an intermediate casing of an aircraft turbomachine with a ducted fan , wherein the arm is configured to connect a hub and an outer ferrule of the intermediate casing , and comprising:an aerodynamic outer surface manufactured such that the arm also forms an outlet guide vane;multiple metal ties extending in a longitudinal direction of the arm, together with a shell made from a composite material surrounding the ties and forming the aerodynamic outer surface; andwherein at least one part of an inner space demarcated by the shell and traversed by the ties is filled by a filling material forming a support of the shell.11. A structural arm according to claim 10 , wherein each tie is separated from the shell made from composite material by the filling material.12. A structural arm according to claim 11 , wherein each tie is sunk in the filling material along an entire length of the shell.13. A structural arm according to claim 10 , wherein each tie extends claim 10 , in the longitudinal direction of the arm claim 10 , beyond the shell claim 10 , on either side of the shell.14. A structural arm according to claim 10 , wherein the ties support at their radially outer ends means for attaching the arm on the outer ferrule of the intermediate casing claim 10 , and support at their radially inner ends means for attaching the arm on the hub of the intermediate casing.15. A structural arm according to claim 14 , wherein the means for attaching the arm on the ...

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16-02-2012 дата публикации

METHOD FOR MANUFACTURING AN ASSEMBLY INCLUDING A PLURALITY OF BLADES MOUNTED IN A PLATFORM

Номер: US20120039738A1
Принадлежит: SNECMA

A method of manufacturing an assembly, including a plurality of blades mounted in a platform preparing a mixture of metal powder and thermoplastic binder, includes manufacturing the blades separately from the platform, and finishing the blades after the manufacturing. The method also includes injecting the mixture into a mold to obtain a platform blank, removing the binder from the platform blank prior to assembling the finished blades with the blank, inserting one end of the finished blades into a housing formed in the platform blank in order to assemble the assembly, and sintering the assembly comprising the platform blank and the finished blades to unify the assembly. 12.-. (canceled)3. A method of manufacturing an assembly comprising a plurality of blades mounted in a platform , comprising:manufacturing the blades separately from the platform, and finishing the blades after the manufacturing;preparing a mixture of metal powder and thermoplastic binder;injecting the mixture into a mold to obtain a platform blank;removing the binder from the platform blank prior to assembling the finished blades with said blank;inserting one end of the finished blades into a housing formed in the platform blank in order to assemble the assembly; andsintering the assembly comprising the platform blank and the finished blades to unify the assembly.4. The method as claimed in claim 3 , in which the assembly comprises a plurality of finished blades the ends of which are mounted between an inner platform and an outer platform claim 3 , and the ends of the finished blades are unified with the platform blanks by sintering the metal powder. The invention relates to the manufacture of a turbo machine blisk (bladed disk) sector in particular a fixed blisk sector.Hereinafter, a fixed blisk sector is known as a stator guide vanes assembly when it forms part of the compressor of the turbo machine and is known as a nozzle guide vane assembly when it forms part of the turbine of the turbo ...

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23-02-2012 дата публикации

method and device for feeding a turbomachine combustion chamber with a regulated flow of fuel

Номер: US20120042657A1
Принадлежит: SNECMA SAS

High-pressure fuel is supplied at a controlled rate to a combustion chamber via a position-controlled valve and a variable-restriction stop-and-pressurizing cut-off valve. A value representative of the real mass flow rate of fuel as delivered is calculated by a calculation unit on the basis of information representative of the pressure difference between the inlet and the outlet of the cut-off valve and of the flow section through the cut-off valve, e.g. as represented by the position X of the slide of the cut-off valve. The position-controlled valve has a variable position that is controlled by the calculation unit as a function of the difference between the calculated value representative of the real mass flow rate and a value representative of a desired mass flow rate.

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23-02-2012 дата публикации

UNDUCTED FAN WITH VARIABLE-PITCH BLADES FOR A TURBINE ENGINE

Номер: US20120045334A1
Принадлежит: SNECMA

A non-streamlined propeller including blades with a variable setting for a turbine engine, the blades of the propeller being rotatably mounted about the axes thereof in radial recesses of a rotor element, and each blade being supported by a plate held in a recess by a sectored ring, and including an inner portion mounted, by interlocking, in a groove of the plate, the ring sectors being inserted into the recess from the inside and locked by a nut screwed onto the plate. 111-. (canceled)12. An unducted fan comprising:variable-pitch blades for a turbine engine, the blades of the fan being rotatably mounted about the axes thereof in radial recesses of an annular rotor element and each blade being supported by a plate with a cylindrical body that is inserted from an outside into a radial recess of the rotor element and that is held in the recess by an annular ring mounted from an inside in the recess and applied on an inner edge of the recess by a bearing,wherein the ring is sectorized and comprises a radially inner portion inserted via interlocking of the ring sectors into an annular groove of an outer surface of the body of the plate.13. A fan according to claim 12 , wherein the ring comprises two sectors each having an angular extent of approximately 180° claim 12 , or three sectors each having an angular extent of approximately 120° claim 12 , or four sectors each having an angular extent of approximately 90°.14. A fan according to claim 12 , wherein the radially inner portion of the ring is mounted in the annular groove of the body of the plate with an adjustment at a very low assembly tolerance.15. A fan according to claim 12 , further comprising means for locking the ring in the groove of the body of the plate fixed from the inside of the recess onto the body of the plate.16. A fan according to claim 15 , wherein the ring comprises a tapered outer surface that widens towards an exterior and whereon is applied a tapered inner surface substantially complementary of ...

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23-02-2012 дата публикации

FAN ROTOR FOR AN AIRPLANE TURBOJET

Номер: US20120045341A1
Принадлежит: SNECMA

A fan rotor for an airplane turbojet including locking blade roots in the respective grooves of an airplane turbojet fan rotor. Each blade root is housed in a groove that is closed by a main latch and by an additional latch that is distinct from the main latch and that is spaced apart therefrom by a predetermined distance. 17-. (canceled)8. A fan rotor comprising:fan blades attached to a periphery of a wheel, each blade including a blade root engaged in a groove in the wheel and retained therein by a latch engaged in notches formed in a vicinity of and on either side of an upstream end of the corresponding groove to oppose movement of the blade root in an axial direction, each latch, as a main latch, being associated with an additional latch that is distinct from the main latch,wherein the additional latch is spaced apart from the main latch at a predetermined distance, the additional latch being situated between the main latch and an upstream end of the blade root engaged in the groove.9. A fan rotor according to claim 8 , wherein the additional latch is a thin wall engaged in corresponding notches formed in either side of the groove.10. A fan rotor according to claim 9 , wherein the thin wall is made of deformable material.11. A fan rotor according to claim 8 , wherein the additional latch is essentially constituted by a metal plate.12. A fan rotor according to claim 8 , wherein the additional latch includes a tongue extending upstream substantially perpendicularly to its surface and welded to an edge of the main latch.13. A fan rotor according to claim 8 , wherein a fastener tab projects upstream from the main latch for mounting a wedge inserted between the blade root and the bottom of the groove. The invention relates to a fan rotor for an airplane turbojet, and it relates more particularly to locking blade roots in their respective grooves. The invention relates in particular to improving the system of individual latches that enable the blades to be held in ...

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23-02-2012 дата публикации

Monitoring particles in a lubrication system

Номер: US20120046896A1
Принадлежит: SNECMA SAS

A process for monitoring a machine including moving pieces, a lubrication system and an electromagnetic sensor fitted with a magnet and two electrodes is disclosed. The sensor is capable of collecting particles present in the lubrication system between the electrodes. The monitoring process includes a step for obtaining measurements of resistance between the electrodes of the sensor, taken during an operating period of the machine; a step for determining from said measurements a first autoregressive mathematical model characterizing the resistance; a step for comparison between the first model and a predetermined reference model; and a step for working out an opinion on maintenance of the machine as a function of the comparison result.

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01-03-2012 дата публикации

MIXING SCREW FOR A FUEL INJECTOR IN A COMBUSTION CHAMBER OF A GAS TURBINE, AND CORRESPONDING COMBUSTION DEVICE

Номер: US20120047899A1
Автор:
Принадлежит: SNECMA

In a fuel injection device at a base of a combustion chamber, a mixing air inlet screw is manufactured with its peripheral holes extending only over a sector of a circle directed towards the air's propagation cone, wherein the remainder of the periphery of the screw is closed. It is then possible to reduce load loss of the injection system while obtaining an improved quality of the mix supplied to the chamber. Such a device can be used in gas turbines fitted with centrifugal compressors and in which the air flow towards the combustion chamber must therefore be made convergent. 110-. (canceled)11. A mixing screw for a fuel injector in a combustion chamber of a gas turbine , comprising:a general shape of a hollow cylinder fitted with at least one angular network of feed holes traversing the cylinder as far as the hollow,wherein the network is irregularly distributed over a circumference of the cylinder, and extends only over a sector of a circle.12. A fuel mixing screw according to claim 11 , wherein the sector of a circle is less than a half-circle.13. A fuel mixing screw according to claim 11 , wherein the sector of a circle is less than a quarter-circle.14. A fuel mixing screw according to claim 11 , wherein the feed holes are staggered in the cylinder's axial direction.15. A fuel mixing screw according to claim 14 , wherein there are three feed holes.16. A fuel mixing screw according to claim 11 , wherein the feed holes have different inclinations in their angular directions.17. A fuel mixing screw according to claim 11 , wherein the feed holes are at a relief angle claim 11 , with increasing widths in their angular directions claim 11 , moving in the outer radial direction.18. A fuel mixing screw according to claim 11 , further comprising at least one other angular network of feed holes traversing the cylinder as far as the hollow claim 11 , wherein the holes of the other network are distributed over a circle claim 11 , or over only a sector of a circle.19. A ...

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01-03-2012 дата публикации

SINGLE-PIECE BRACKET FOR AERONAUTICAL EQUIPMENT

Номер: US20120049029A1
Автор:
Принадлежит: SNECMA

A bracket for auxiliary equipment of a device, such as cables or pipes, includes a single metal plate that has been cut and then folded, forming elbow connections between the console, a support for attaching equipment, a flange for fastening to the device, and two intermediate ribs. The closed, or nearly closed, outline of the bracket and the parallel or convergent ribs form a trapezium and provide satisfactory transmission and satisfactory distribution of the efforts: the bracket is robust although it has been manufactured from quite a thin plate. The bracket, for example, finds application to aeronautics, to support, for example, a supply of electricity and fluids to engines. 15-. (canceled)6. A bracket for aeronautical equipment , comprising:a flange for connection to a supporting structure;a console for attaching equipment; andan intermediate reinforcing part;wherein the flange, the reinforcing part, and the console are formed from a single, folded metal plate,wherein the reinforcing part includes two ribs, each extending between the flange and the console, and attached to it by elbow connections, most of which are in oblique directions that are not perpendicular to a principal plane of extension of the flange or to a principal plane of extension of the console,wherein the console and the flange have no connection other than the ribs,wherein the ribs are connected on either side of the console, andwherein the ribs are flat and principal planes of extension of the flange and of the console are mutually perpendicular.7. A bracket for aeronautical equipment according to claim 6 , wherein the ribs extend in convergent directions claim 6 , forming an angle other than a straight angle.8. A bracket for aeronautical equipment according to claim 6 , wherein the flange is divided into two portions that are mutually separated and attached respectively to the ribs.9. A bracket for aeronautical equipment according to claim 8 , wherein the portions extend in divergent ...

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01-03-2012 дата публикации

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

Номер: US20120051895A1
Автор:
Принадлежит: SNECMA

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

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01-03-2012 дата публикации

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

Номер: US20120051931A1
Автор:
Принадлежит: SNECMA

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

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01-03-2012 дата публикации

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

Номер: US20120051932A1
Принадлежит: SNECMA

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

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01-03-2012 дата публикации

SHELL FOR AIRCRAFT TURBO-ENGINE STATOR WITH MECHANICAL BLADE LOAD TRANSFER SLITS

Номер: US20120051938A1
Автор: Bertoli Vincenzo
Принадлежит: SNECMA

A shell of a stator for a module of an aircraft turbo-engine including a plurality of through openings each configured to house a stator blade, each opening forming a skeleton extending between a first end configured to house the trailing edge of the blade and a second end configured to house the leading edge of the blade. At least one of the openings is associated with a mechanical load transfer slit passing through the shell, and arranged facing and at a distance from the first end of the opening along the direction of the skeleton. 113-. (canceled)14. A shell for a stator of an aircraft turbo-engine module , comprising:a plurality of through openings each configured to house a stator blade, each opening forming a skeleton extending between a first end configured to house the trailing edge of the blade and a second end configured to house the leading edge of the blade,wherein at least one of the openings is associated with a mechanical load transfer slit passing through the shell, and arranged facing and at a distance from the first end of the opening along the direction of the skeleton.15. A shell according to claim 14 , wherein the slit extends along a curved line towards the opening.16. A shell according to claim 14 , wherein the opening presents an intrados part and an extrados part on each side of the skeleton claim 14 , that come together at the first and second ends of the opening claim 14 , and wherein the slit also extends facing and at a distance from portions of the intrados and extrados parts that come together at the first end of the opening.17. A shell according to claim 14 , wherein the slit has a generally U or V shape and the first end of the opening is located inside the hollow of the U or V.18. A shell according to claim 14 , wherein the slit is filled in by an infill material.19. A shell according to claim 14 , forming a continuous approximately annular structure.20. A shell according to claims 14 , forming an approximately annular structure ...

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08-03-2012 дата публикации

FAN BLADES WITH CYCLIC SETTING

Номер: US20120055137A1
Принадлежит: SNECMA

A fan portion of a dual flow turbojet engine including a plurality of fan blades, a disk supporting the blades and configured to be rotated relative to a stator portion of the fan, along a longitudinal axis of the fan, and a system for setting the angle of attack associated with each fan blade, the systems configured such that the angle of attack of each blade varies according to a same setting law according to the angular position of the blade relative to the stator portion, along the longitudinal axis, the same setting law being periodic with a period of P=360°/n, where n is an integer at least equal to 1. 113.-. (canceled)14. A dual flow turbojet engine fan portion comprising:a plurality of fan blades;a disk to support the blades and configured to be rotated relative to a fan stator portion, along a longitudinal fan axis;a system for setting angles of attack associated with each fan blade, the systems configured so that an angle of attack of each blade varies according to a same setting law as a function of the angular position of the blade relative to the stator portion, along the longitudinal fan axis, the same setting law being periodic with period P=360°/n, with n corresponding to a whole number greater than or equal to 1.15. The fan portion according to claim 14 , wherein the system for setting angles of attack is steered passively by rotating the disk supporting the blades relative to the fan stator portion claim 14 , along the longitudinal axis of the fan.16. The fan portion according to claim 15 , wherein the passively steered system for setting the angle of attack comprises a lug arranged eccentrically on a foot of a concerned blade claim 15 , a first toothed wheel centered on the longitudinal fan axis and fastened to the fan stator portion claim 15 , and a second toothed wheel rotated along the longitudinal fan axis by the support disk claim 15 , and mounted freely rotating on the support disk along a wheel axis of rotation separate from the ...

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08-03-2012 дата публикации

TURBOMACHINE COMBUSTION CHAMBER COMPRISING IMPROVED MEANS OF AIR SUPPLY

Номер: US20120055164A1
Принадлежит: SNECMA

Annular combustion chamber () to be fitted on a turbomachine and comprising a chamber end wall (), a plurality of air and fuel injection systems () circumferentially distributed around an axis () of the combustion chamber () and mounted on said chamber end wall (), and, an air manifold () associated with each injection system (), comprising at least one wall () mounted on the chamber end wall () and projecting in the upstream direction to form an obstacle to a circumferential airflow around the axis () of the combustion chamber (), and an air inlet opening () formed at the upstream end of the air manifold () and opening radially outwards from an axis () of said injection system. 111-. (canceled)12. An annular combustion chamber comprising a chamber end wall arranged at the upstream end of the combustion chamber , and a plurality of air and fuel injection systems circumferentially distributed around an axis of the combustion chamber and mounted on said chamber end wall , said combustion chamber also comprising an air manifold associated with each injection system , said air manifold comprising at least one wall mounted on the chamber end wall and projecting in the upstream direction to form an obstacle to a circumferential airflow around the axis of the combustion chamber , and an air inlet opening formed at the upstream end of said air manifold and opening radially outwards from an axis of said injection system.13. The annular combustion chamber according to claim 12 , wherein when said air inlet opening of each air manifold is seen in projection in a transverse plane perpendicular to a tangential plane passing through said centre line of the corresponding injection system claim 12 , the part of said opening located radially outwards from said tangential plane has a larger opening area than the part of said opening that is located radially inwards from said tangential plane.14. The annular combustion chamber according to claim 12 , wherein each air manifold ...

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15-03-2012 дата публикации

CONTACTLESS ELECTRICAL CONNECTOR FOR AN INDUCTION SENSOR, AND SENSOR INCLUDING SUCH A CONNECTOR

Номер: US20120062214A1
Автор: Bares Jean Paul Yvon
Принадлежит: SNECMA

An electrical connector between an induction sensor and a cable for transmitting a signal provided by the sensor, including a current transformer with a primary coil and a mechanism for electrically coupling to the sensor, and a secondary coil with a mechanism for electrically coupling to the cable, the primary and secondary coils being attached together by a removable attachment mechanism. 111-. (canceled)12. An electrical connector between an induction sensor and an electrical cable for transmitting a signal supplied by the sensor , comprising:a current transformer including a primary coil and means for electrical connection to the sensor and a secondary coil including means for connecting to the cable,the primary and secondary coils being attached to one another by a removable attachment means.13. The connector as claimed in claim 12 , wherein the primary coil is secured to a board for mounting on the sensor.14. The connector as claimed in claim 12 , wherein the primary coil and the secondary coil are placed concentrically relative to one another.15. The connector as claimed in claim 13 , wherein the primary coil and the secondary coil are placed concentrically relative to one another.16. The connector as claimed in claim 15 , wherein the secondary coil is resting against a collar of the board.17. The connector as claimed in claim 16 , wherein the board is secured to a stem forming a core claim 16 , the primary and secondary coils being held together by a washer that is bolted onto the stem of the core and forming the removable attachment means.18. The connector as claimed in claim 12 , wherein the primary coil and the secondary coil are coaxial and superposed.19. The connector as claimed in claim 13 , wherein the primary coil and the secondary coil are coaxial and superposed.20. The connector as claimed in claim 19 , wherein the board is secured to a core whereof an extension forms a centering spindle for the secondary coil.21. The connector as claimed in claim ...

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15-03-2012 дата публикации

BUSHING FOR A VARIABLE SET BLADE

Номер: US20120063905A1
Принадлежит: SNECMA

A bushing for a variable set blade includes a cylindrical part configured to receive the blade, the root of the blade being able to rotate in the cylindrical part, and a base including: a first side able to be put in contact with a first edge of a circumferential groove of the ring, and/or a second side able to be put in contact with a second edge of the circumferential groove of the ring. At least one of the first side or second side of the base includes at least one bevel through which the first side or the second side is able to be put in contact with one of the first edge or second edge of the circumferential groove so as to block the bushing in rotation. The invention finds a particularly interesting application in the field of aircraft. 1. A bushing for a variable set blade comprising:a cylindrical part configured to receive said blade, a root of said blade being configured to rotate in said cylindrical part, and a first side configured to be put in contact with a first edge of a circumferential groove of a ring, and/or', 'a second side configured to be put in contact with a second edge of said circumferential groove of said ring,, 'a base comprisingwherein at least one of said first side or second side of said base comprises at least one bevel through which said first side or said second side is able to be put in contact with one of said first edge or second edge of said circumferential groove so as to block said bushing in rotation.2. The bushing for a variable set blade according to claim 1 , wherein said first side comprises a first bevel and a second bevel and/or said second side comprises a first bevel and a second bevel.3. The bushing for a variable set blade according to claim 2 , whereineach of said first bevel and second bevel of said first side presents a length on the order of 8 mm; andeach of said first bevel and second bevel of said second side presents a length on the order of 8 mm.4. The bushing for a variable set blade according to claim 1 , ...

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15-03-2012 дата публикации

CIRCUMFERENTIAL BLOCKING DEVICE OF CLAMP VANES FOR TURBINE ENGINE, WITH IMPROVED RADIAL DEPLOYMENT

Номер: US20120063907A1
Автор: Cloarec Yvon
Принадлежит: SNECMA

The invention relates to a device () for blockage of vanes in a circumferential groove () of a turbine engine disc, the device being intended to deploy in a radial direction to be locked in the circumferential groove open radially. According to the invention, the blocking device comprises a first and a second piece () designed to deploy in the radial direction by relative displacement of these pieces in the circumferential direction (). 13010. A blocking device () for vanes in a circumferential groove () of a turbine engine disc , said device being intended to deploy in a radial direction to be locked in said circumferential groove open radially ,{'b': 32', '34', '23, 'characterised in that said blocking device comprises a first and second piece (, ) designed to deploy in the radial direction by relative displacement of said pieces in the circumferential direction ().'}232345252a. The device as claimed in claim 1 , characterised in that at least one of said first and second pieces ( claim 1 , ) is equipped with a gripping member ( claim 1 , ) for being taken along relative to the other piece claim 1 , according to the circumferential direction.3. The device as claimed in claim 2 , characterised in that the gripping member is a tab or a cable.4525249a. The device as claimed in or claim 2 , characterised in that the piece equipped with the gripping member ( claim 2 , ) has a stop () in the circumferential direction.5323448. The device as claimed in claim 1 , characterised in that at least one of said first and second pieces ( claim 1 , ) has a ramp () for deployment of the device in the radial direction claim 1 , during relative displacement of said pieces in the circumferential direction.668703234. The device as claimed in claim 1 , characterised in that it is equipped with locking means ( claim 1 , ) of the two pieces ( claim 1 , ) in a position deployed radially.760305252a. A system () comprising two pieces intended to enter respectively in the constitution of two ...

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15-03-2012 дата публикации

Low-pressure turbine

Номер: US20120063914A1
Принадлежит: SNECMA SAS

A low-pressure turbine for a turbomachine including turbine disks with festooned annular flanges for fastening to a festooned annular flange of a drive cone connecting the turbine disks to a turbine shaft, solid portions of the festooned annular flanges of the turbine disks and of the drive cone being connected to peripheries of the turbine disks and of the drive cone, respectively, via concave filets that are asymmetrical.

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15-03-2012 дата публикации

A METHOD AND A DEVICE FOR COATING CERAMIC MATERIAL FIBERS IN METAL BY A LIQUID TECHNIQUE

Номер: US20120064254A1
Автор:
Принадлежит: SNECMA

A method of coating ceramic material fibers in metal using a liquid technique and a device implementing the method. The method maintains a charge of molten metal in levitation in a substantially spherical shape inside a crucible and causes a tensioned ceramic material fiber to travel at a predetermined speed between a bottom pulley and a top pulley disposed on either side of the crucible such that a portion of fiber is immersed in the charge to be covered in a metal coating. During coating, the portion of fiber that is immersed in the charge is shifted as a function of the remaining volume of the charge such that the instantaneous height of fiber that is immersed in the charge remains substantially constant throughout the coating operation. 18-. (canceled)9. A method of coating ceramic material fibers in metal using a liquid technique , the method comprising:maintaining a charge of molten metal in levitation in a substantially spherical shape inside a crucible;causing a tensioned ceramic material fiber to travel at a predetermined speed between a bottom pulley and a top pulley disposed on either side of the crucible such that a portion of fiber is immersed in the charge to be covered in a metal coating; andcausing the portion of fiber that is immersed in the charge to shift during coating as a function of the remaining volume of the charge such that instantaneous height of fiber that is immersed in the charge remains substantially constant throughout the coating.10. A method according to claim 9 , wherein the shifting of the portion of fiber that is immersed in the charge takes place in a direction that is substantially perpendicular to the travel direction of the fiber.11. A method according to claim 10 , wherein the portion of fiber that is immersed in the charge is shifted by at least one wheel interposed between one of the pulleys and the crucible and capable of shifting perpendicularly to the travel direction of the fiber.12. A method according to claim 9 , ...

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22-03-2012 дата публикации

BELLOWS TYPE SEALING DEVICE FOR PARTITION PENETRATION BY A CONNECTING ROD OF A TURBOPROP FAN BLADE ORIENTATION CONTROL SYSTEM

Номер: US20120070289A1
Принадлежит: SNECMA

The invention relates to a sealing device for the partition penetration of a connecting rod of a turboprop fan blade orientation control system. The device includes a bellows () of trunconical shape designed to allow passage of the connecting rod () and having, at its wider end, means () of attachment to the partition () to be sealed and, at its narrower end, an O-ring through which the connecting rod is free to slide. 12424262828222236384042404648303012323250505454626458a,ba,ba,ba,ba,ba,ba,b. A system for controlling the orientation of the fan blades of a turboprop comprising at least one set () of fan blades () with adjustable orientation , said set being integral in rotation with a rotating ring () connected mechanically with a rotating housing () , each blade of the set being coupled , for control of its orientation , to a blade root support () pivotably mounted on the rotating ring by means of a bevel gearset () made up of a first bevel gear () integral with the blade root support and centered on an axis () radial to the rotating ring and a second bevel gear () integral with the rotating ring , centered on an axis () tangential to said rotating ring , and bearing a counterweight () eccentric with respect to its axis of rotation , the system also comprising a cylinder () centered on the axis of rotation () of the rotating ring , rotating integrally with the rotating housing , and the rod () whereof is connected to each counterweight through radial connecting rods () and bellcranks () , the system also comprising , for each radial connecting rod , a bellows of trunconical shape through which passes the radial connecting rod and having , at its wider end , means ( , ) of attachment to a partition () to be sealed and , at its narrower end , an O-ring through which the radial connecting rod is free to slide.26258. A system according to claim 1 , wherein the wider end of the bellows includes an attachment clip () attached to the partition () to be sealed.3. A system ...

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22-03-2012 дата публикации

DEVICE FOR CONTROLLING THE PITCH OF FAN BLADES OF A TURBOPROP

Номер: US20120070290A1
Автор:
Принадлежит: SNECMA

A device for controlling pitch of fan blades of a turboprop including at least one set of adjustable-pitch fan blades secured to rotate with a rotary ring mechanically connected to a turbine rotor, each blade of the set being coupled for pitch adjustment to a synchronization ring. The device includes a rolling bearing including an inner cage slidably mounted on a turbine casing and connected to the rod of an actuator, and an outer cage that is mechanically connected to the synchronization ring by a plurality of connection arms connected to the actuator rod and hinge-mounted on the synchronization ring such that actuating the actuator causes the synchronization ring to move in turning about the longitudinal axis. 110-. (canceled)11. A device for controlling pitch of fan blades of a turboprop including at least one set of adjustable-pitch fan blades , the set being secured to rotate with a rotary ring centered on a longitudinal axis and mechanically connected to a turbine rotor , each blade of the set being coupled for pitch adjustment to a synchronization ring centered on the longitudinal axis , the device comprising:a rolling bearing including an inner cage slidably mounted on a turbine casing and connected to the rod of an actuator centered on the longitudinal axis; andan outer cage that is mechanically connected to the synchronization ring by a plurality of connection arms connected to the actuator rod and hinge-mounted on the synchronization ring such that actuating the actuator causes the synchronization ring to move in turning about the longitudinal axis.12. A device according to claim 11 , wherein each connection arm comprises an axial link connected to the outer cage of the rolling bearing claim 11 , a radial link connected to the synchronization ring claim 11 , and at least one bellcrank connecting the axial link to the radial link such that actuating the actuator causes the radial link to move in a direction that is substantially radial.13. A device ...

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22-03-2012 дата публикации

PROPELLER AND SYSTEM OF COUNTER-ROTATING PROPELLERS COMPRISING IMPROVED MEANS FOR LIMITING PITCH, AND A TURBINE ENGINE COMPRISING THEM

Номер: US20120070291A1
Автор:
Принадлежит: SNECMA

A propeller for an aircraft turbine engine, comprising means for adjusting the pitch of the blades comprising a hydraulic actuator borne by the rotor and including a cavity and a piston displaceable in said cavity and dividing said cavity into two chambers, as well as means for limiting the stroke of the piston in order to prevent displacement of said blades in a predetermined direction beyond a predetermined limiting pitch angle, and comprising hydraulic connection means with which both chambers may be put into communication with each other as soon as the pitch angle of the blades becomes equal to said predetermined limiting pitch angle during the displacement of said blades in said predetermined direction, these hydraulic connection means being borne by said rotor of the propeller. 1. A propeller for an aircraft turbine engine , comprising a propeller rotor , blades borne by said rotor , and adjustment means for adjusting the pitch of said blades , said adjustment means comprising at least one double acting hydraulic actuator borne by a said rotor and including at least one cavity and a corresponding piston housed and displaceable in said cavity for causing modification of the pitch angle of said blades , said piston dividing said cavity into two chambers , said adjustment means further comprising limitation means for limiting the stroke of each piston of each aforementioned actuator for preventing displacement of said blades into a predetermined direction , corresponding to a displacement of the piston in a direction orientated from a second chamber towards a first chamber of the cavity housing said piston , beyond a predetermined limiting pitch angle ,wherein said limitation means comprise hydraulic connection means allowing both chambers of each cavity of each aforementioned actuator to be put into fluidic communication with each other as soon as the pitch angle of said blades becomes equal to said predetermined limiting pitch angle during a displacement of ...

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22-03-2012 дата публикации

MOVABLE ACTUATOR DEVICE FOR CONTROLLING THE PITCH OF FAN BLADES OF A TURBOPROP

Номер: US20120070292A1
Принадлежит: SNECMA

A device for controlling pitch of fan blades of a turboprop including at least one set of adjustable-pitch fan blades secured to rotate with a rotary ring centered on the longitudinal axis, and mechanically connected to a turbine rotor, each blade of the set being coupled for pitch adjustment to a synchronization ring centered on the longitudinal axis. An actuator is secured to rotate with the turbine rotor and is mechanically connected to the synchronization ring by a plurality of connection arms connected to the actuator rod and hinge-mounted on the synchronization ring such that actuating the actuator causes the synchronization ring to move in turning about the longitudinal axis. 110-. (canceled)11. A device for controlling pitch of fan blades of a turboprop including at least one set of adjustable-pitch fan blades , the set being secured to rotate with a rotary ring centered on a longitudinal axis and mechanically connected to a turbine rotor , each blade of the set being coupled for pitch adjustment to a synchronization ring centered on the longitudinal axis , the device comprising:an actuator centered on the longitudinal axis, secured to rotate with the turbine rotor and mechanically connected to the synchronization ring by a plurality of connection arms each comprising an axial link connected to the actuator rod, a radial link connected to the synchronization ring, and at least one bellcrank connecting the axial link to the radial link such that actuating the actuator causes the synchronization ring to move in turning about the longitudinal axis.12. A device according to claim 11 , wherein each connection arm further includes another bellcrank secured to the rotary ring and connected firstly to the radial link and secondly to a tangential link fastened to the synchronization ring such that movement of the radial link in a direction that is substantially radial causes the synchronization ring to move in turning about the longitudinal axis.13. A device ...

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22-03-2012 дата публикации

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

Номер: US20120070298A1
Принадлежит: SNECMA

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y,Z′ given in Table 1, in which the coordinate Z′ is the quotient D/H where D is the distance of the point under consideration from a first reference plane P situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine. 1. An aerodynamic profile for a turbine blade , the profile being , when cold and in a non-coated state , substantially identical to a nominal profile determined by the Cartesian coordinates X ,Y ,Z′ given in Table 1 , in which the coordinate Z′ is the quotient D/H , where D is the distance of the point under consideration from a first reference X ,Y plane situated at the base of the nominal profile , and H is the height of said profile measured from said first reference plane that is the intersection of the stacking axis of the set of blades and the axisymmetric surface of the hub , out to a second reference plane that is the intersection of said stacking axis with the axisymmetric surface of the casing , the measurements D and H being taken radially relative to the axis of the turbine , while the coordinate X is measured in the axial direction of the turbine.2. The aerodynamic profile as claimed in claim 1 , which is defined within an envelope of ±1 mm in a direction normal to the surface of the nominal profile.314. The aerodynamic profile as claimed in claim 1 , wherein the coordinates X claim 1 ,Y of said profile () lie within a range of ±5% relative to the coordinates X claim 1 ,Y of the nominal profile.4. The aerodynamic profile as claimed in claim 1 , wherein the blade is a blade of a rotary wheel forming a portion of a rotor of a turbine.5. The ...

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29-03-2012 дата публикации

Turbomachine having an annular combustion chamber

Номер: US20120073259A1
Принадлежит: SNECMA SAS

A turbomachine including an annular combustion chamber, the combustion chamber presenting a connection flange at its downstream end for connection to an outer casing. The connection flange bears axially against the outer casing and is blocked in an axial direction by the upstream end of an inner casing of a high pressure turbine.

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29-03-2012 дата публикации

GAS TURBINE COMBUSTION CHAMBER MADE OF CMC MATERIAL AND SUBDIVIDED INTO SECTORS

Номер: US20120073306A1
Принадлежит: SNECMA PROPULSION SOLIDE

An assembled annular combustion chamber comprises an annular inner wall and an annular outer wall made of ceramic matrix composite material together with a chamber end wall connected to the inner and outer walls and provided with orifices for receiving injectors. Elastically-deformable link parts connect the inner wall and the outer wall of the chamber to inner and outer casings that are made of metal. The assembly formed by the inner wall, the outer wall, and the combustion chamber end wall is subdivided circumferentially into adjacent chamber sectors, each sector being made as a single piece of ceramic composite material and comprising an inner wall sector, an outer wall sector, and a chamber end wall sector. The link parts connect the inner metal casing and the outer metal casing respectively to each inner wall sector of the combustion chamber and to each outer wall sector of the chamber. The chamber end wall sectors are in contact with a one-piece ring to which they are connected. 1. An annular combustion chamber assembly for a gas turbine , the assembly comprising: an inner metal casing; an outer metal casing; an annular combustion chamber mounted between the inner and outer casings and comprising an annular inner wall and an annular outer wall of ceramic material together with a chamber end wall connected to the inner and outer walls and provided with orifices for receiving injectors; and elastically-deformable link parts supporting the combustion chamber between the inner metal casing and the outer metal casing; the assembly formed by the inner wall , the outer wall , and the end wall of the combustion chamber being subdivided circumferentially into adjacent chamber sectors , each comprising an inner wall sector , an outer wall sector , and a chamber end sector interconnecting the outer and inner wall sectors;wherein each chamber sector is made as a single piece of ceramic composite material, wherein elastically-deformable link parts connect the inner metal ...

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29-03-2012 дата публикации

METHOD AND DEVICE FOR MACHINING THE LEADING EDGE OF A TURBINE ENGINE BLADE

Номер: US20120077417A1
Принадлежит: SNECMA

A method and a device for machining the leading edge of a turbine engine blade by a machining center for which parameters are set is disclosed. The method includes: acquiring a 3D profile of the leading edge of the blade; calculating at least one characteristic of the leading edge from the 3D profile; comparing the value of the calculated characteristic with a known theoretical value of the characteristic to obtain an elementary difference for the characteristic; calculating at least one undulation of the leading edge between at least two consecutive elementary sections from the 3D profile; optimizing the elementary differences obtained as a function of the undulation; setting the parameters of the machining center as a function of the optimized elementary differences for the elementary sections to define machining passes over the leading edge; and machining the leading edge of the blade with the machining center with parameters set. 1. Method for machining the leading edge of a turbine engine blade by means of a machining centre for which parameters can be set , the method comprising the following steps:acquiring a 3D profile of the leading edge and the trailing edge of the blade, at least two elementary sections being defined on the profile of the blade over its height;calculating at least one characteristic of the leading edge from the 3D profile for each of the elementary sections, the said characteristic being chosen from the chord width of the blade, the radius of the leading edge, the pressure-side slope and the suction-side slope of the leading edge;for each given elementary section, comparing the value of the calculated characteristic with a known theoretical value of the said characteristic for the said given elementary section so as to obtain an elementary difference for the said characteristic for the said given elementary section;calculating at least one undulation of the leading edge between at least two consecutive elementary sections from the 3D ...

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05-04-2012 дата публикации

DEVICE AND METHOD FOR DETECTING A FAULT IN A LOW-PRESSURE FUEL PUMP OF A TURBOJET AND TURBOJET INCLUDING ONE SUCH DEVICE

Номер: US20120079832A1
Автор:
Принадлежит: SNECMA

A device for detecting a fault in a low-pressure fuel pump of a turbojet. The pump is driven by an accessory gearbox including a gear for mechanically driving the accessories. The device measures the vibration frequencies of the accessory gearbox and detects, from among the frequencies, at least one vibration frequency of the low-pressure fuel pump. The device allows a fault in the low-pressure fuel pump to be detected as soon as it occurs. 19-. (canceled)10. A device for detecting a failure of a low-pressure fuel pump of a jet engine including at least one rotary shaft that can rotate at different speeds , the pump being driven by the rotary shaft via an accessory relay box including a gear system for mechanically driving the accessories , the device comprising:means for measuring speed of rotation of the rotary shaft of the jet engine;means for measuring vibration frequencies of the accessory relay box; andmeans for detecting, from the detected frequencies, at least one normal vibration frequency of the low-pressure fuel pump at the measured rotation speed of the rotary shaft.11. A device according to claim 10 , in which the jet engine is a dual-body jet engine comprising a low-pressure body and a high-pressure body claim 10 , and the rotary shaft is the shaft of the high-pressure body of the jet engine.12. A device according to one of claim 10 , in which the means for measuring vibration frequencies of the accessory relay box comprises an accelerometer delivering a signal representative of the vibrations of the accessory relay box.13. A device according to claim 10 , in which the accessory relay box and its accessories are arranged so that the vibration frequencies of the accessories are all different from the vibration frequency of the low-pressure fuel pump.14. A jet engine comprising:at least one rotary shaft that can rotate at different speeds;a fuel circuit with a low-pressure fuel pump and a high-pressure fuel pump driven by the rotary shaft via an ...

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12-04-2012 дата публикации

METHOD FOR REDUCING THE DIAMETER OF AN OPENING

Номер: US20120084958A1
Автор:
Принадлежит: SNECMA

A method for reducing the diameter of an opening, including peening a perimeter of the opening. A method for correcting the permeability of a part including a plurality of openings for allowing a gaseous fluid to pass therethrough. The method identifies at least one opening with a diameter which exceeds a predetermined upper limit and reduces the excessive diameter by peening a perimeter of the opening. 111-. (canceled)12. A method for reducing a diameter of an opening , comprising:peening a perimeter of the opening by a tool, a contact end of which is spherical or approximately spherical or frustoconical.13. The method as claimed in claim 12 , in which the peening is carried out with a tool centered over the opening.14. The method as claimed in claim 13 , in which a contact end of the tool comprises a ball.15. The method as claimed in claim 12 , in which the perimeter of the opening is metallic.16. The method as claimed in claim 12 , in which the perimeter of the opening is composed of a refractory alloy.17. The method as claimed in claim 12 , in which the diameter of the opening is between 0.5 and 3 mm.18. A method for correcting permeability of a component including a plurality of through-openings for gaseous fluid claim 12 , the method comprising:identifying at least one opening, the diameter of which exceeds a predetermined upper limit; and{'claim-ref': {'@idref': 'CLM-00012', 'claim 12'}, 'reducing the excessive diameter by a method as claimed in .'}19. The method as claimed in claim 18 , further comprising a prior operation of checking permeability of the component.20. The method as claimed in claim 18 , in which the openings are cooling openings.21. The method as claimed in claim 18 , in which the component is part of a gas turbine hot section.22. The method as claimed in claim 18 , in which the component is part of a combustion chamber. The present invention relates to a method for reducing the diameter of an opening.The drilling of openings, in particular ...

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12-04-2012 дата публикации

METHOD OF TESTING A SYSTEM FOR PROTECTING A TURBOMACHINE AGAINST OVERSPEED WHILE STARTING

Номер: US20120085101A1
Автор:
Принадлежит: SNECMA

In a testing overspeed protection system: a) on receiving an order to start a turbomachine, an electronic regulation system (ERS) of the turbomachine sends an order to a control circuit of a fuel cutoff member to close the fuel cutoff member or to keep it in the closed position; b) the closed state of the FCM is verified on the basis of information transmitted to the ERS and representative of the position of the FCM; c) if the result of the verification in b) is positive, the ERS sends an order to the FCM control circuit to authorize opening of the FCM and enable the starting procedure to continue; and d) if the result of the verification in b) is negative, the ERS issues fault information concerning the overspeed protection system. 1. A method of testing a system for protecting a turbomachine against overspeed during starting of the turbomachine , the protection system including a fuel cutoff member and a circuit for controlling the cutoff member that is connected to an electronic regulation system of the engine to cause the cutoff member to close to interrupt or reduce the supply of fuel to a combustion chamber of the turbomachine in response to overspeed being detected , the method being comprising the following test sequence:a) on receiving an order to start the turbomachine, the electronic regulation system sending an order to the control circuit of the cutoff member to close the cutoff member or to keep it in the closed position;b) the electronic regulation verifying the closure state of the cutoff member on the basis of information received representative of the open or closed position of the cutoff member;c) when the result of the verification in step b) is positive, the electronic regulation system sending an order to the control circuit of the cutoff member to authorize opening of the cutoff member and continue with the procedure for starting the turbomachine; andd) when the result of the verification in step b) is negative, the electronic regulation ...

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12-04-2012 дата публикации

DEVICE FOR CHECKING A TURBOMACHINE ENGINE

Номер: US20120085156A1
Принадлежит: SNECMA

A device for non-destructive in situ inspection of parts of a turbine engine, the device including a stick carrying at its distal end a pivoting finger carrying at one of its ends a blade for supporting an inspection probe, and at its opposite end a skid for bearing against and/or catching on an element of the engine, the skid being movable in a direction parallel to the longitudinal axis of the finger. 111-. (canceled)12. A device for non-destructive in situ inspection of parts of an engine , the device comprising:a longitudinal stick including an inspection probe mounted at a distal end thereof; anda longitudinal finger pivotally mounted to the distal end of the stick, the finger carrying at a first end support means for supporting the inspection probe, and at a second end catch means for catching on an element of the engine, the catch means being movable in a direction parallel to the finger.13. A device according to claim 12 , wherein the support means comprises a blade of elongate shape that is pivotally mounted via one of its ends to the first end of the finger to pivot between a folded position in which it extends substantially parallel to the finger and a deployed position in which it extends substantially perpendicularly to the finger.14. A device according to claim 13 , wherein the blade is elastically deformable in bending.15. A device according to claim 13 , wherein the inspection probe is fastened to the free end of the blade.16. A device according to claim 13 , further comprising resilient return means urging the blade towards its folded position or its deployed position.17. A device according to claim 13 , further comprising at least one control cable for controlling pivoting of the blade from its deployed position to its folded position and/or from its folded position to its deployed position claim 13 , the control cable(s) extending along the stick.18. A device according to claim 12 , wherein the finger is pivotally mounted in its middle portion to ...

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12-04-2012 дата публикации

SUSPENSION FOR AN ENGINE ON AN AIRCRAFT STRUT INCLUDING A SUSPENSION ARCH

Номер: US20120085859A1
Автор:
Принадлежит: SNECMA

A suspension for an engine on an aircraft strut, the suspension comprising a beam with a plate provided with fixing means for fixing to said strut and a suspension arch linked to the beam by at least one pivot link whose axis is intended to be parallel to the axis of the engine, the suspension arch having, at each of its ends, fixing means for fixing to a case of the engine. 1. Rear suspension for an engine on an aircraft strut , the suspension comprising:a beam with a plate provided with fixing means for fixing to said strut anda suspension arch, having an open angle of between 25 and 40°, linked to the beam by at least one pivot link whose axis is intended to be parallel to the axis of the engine, the suspension arch having, at each of its ends , fixing means for fixing to a case of the engine.2. Suspension according to claim 1 , in which the suspension arch is linked to the beam by a single pivot link.3. Suspension according to claim 1 , in which the beam is in the form of a trapezoid claim 1 , preferably an isosceles trapezoid whose height extends radially relative to the axis of the engine.4. Suspension according to claim 1 , in which the suspension arch has claim 1 , at a first end claim 1 , at least one lug intended to be fixed to the case of the engine to form a pivot link whose axis is parallel to the engine axis.5. Suspension according to claim 1 , in which the suspension arch has claim 1 , at a second end claim 1 , at least one link rod intended to be fixed to the case of the engine.6. Suspension according to claim 1 , in which the suspension arch has an open angle equal to 33°.7. Suspension according to claim 1 , in which the ends of the suspension arch are arranged to be fixed to structural radial arms of the engine.8. Aircraft including at least one strut and an engine linked to said strut by a rear suspension according to . The present invention relates to the field of jet engines and targets the suspension of the latter on the structure of the ...

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12-04-2012 дата публикации

DEVICE FOR ACOUSTIC TREATMENT OF THE NOISE EMITTED BY A TURBOJET

Номер: US20120085861A1
Автор:
Принадлежит: SNECMA

The invention provides a device for acoustical treatment of the noise emitted by a bypass turbojet comprising a primary cowl having in an outer surface an inner annular acoustic treatment panel and a secondary cowl including in an inner surface an outer annular acoustic treatment panel arranged facing the inner panel. The inner and outer panels include respective central panel portions facing each other and extending axially over a common predetermined length, the length of the central panel portions lying in the range one-fifth to four-fifths of the total length of the panels, and the ratio between the acoustic resistances of the central panel portions being not less than 2. 1. A device for acoustically treating the noise emitted by a bypass turbojet , the device comprising:a primary cowl for surrounding the central core of the turbojet and including in an outer surface an internal annular acoustic treatment panel; anda secondary cowl surrounding the primary cowl to co-operate therewith to define an annular channel for passing a cold flow from the turbojet, the secondary cowl including in an inner surface of an outer annular acoustic treatment panel placed facing the inner panel and extending axially over the same length as the inner panel;wherein the inner and outer panels include respective central panel portions facing each other and extending axially over a common predetermined length, the length of the central panel portions lying in the range one-fifth to four-fifths of the total length of the panels, and the ratio between the acoustic resistances of the central panel portions being not less than 2.2. A device according to claim 1 , wherein the ratio between the acoustic resistance of the outer central panel portion to the acoustic resistance of the inner central panel portion lies in the range 2 to 9.3. A device according to claim 1 , wherein the acoustic resistance of the inner central panel portion lies in the range 0.3 ρc to 0.6 ρc.4. A device according ...

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12-04-2012 дата публикации

BENCH AND A METHOD FOR MAGNETOSCOPICALLY TESTING A TURBINE ENGINE SHAFT

Номер: US20120086441A1
Автор:
Принадлежит: SNECMA

A bench for magnetoscopically testing a tubular part, the bench including a tool of elongate shape for inserting inside the part and carrying an endoscopic mechanism for ultraviolet illumination of the inside surface of the part and for observing any defects of the part, and an indexing mechanism co-operating by mutual engagement with external references of the tool that are regularly distributed over at least a fraction of its length to control accurately the advance and the position of the tool inside the part. 113-. (canceled)14. A bench for magnetoscopic testing of a tubular part , the bench comprising:means for supporting the part and for turning the part;a tool of elongate shape carrying endoscopic means for ultraviolet illumination of an inside surface of the part and for observing any defects in the part; andmeans for supporting and guiding the tool to move in translation so as to be inserted inside the part, whereinthe tool includes a plurality of external references that are regularly distributed over at least a fraction of its length and that define regular steps for advancing the tool in translation along the longitudinal axis of the part, the means for supporting and guiding the tool including indexing means that co-operate with the references of the tool by mutual engagement to control accurately an advance and a position of the tool in the part.15. A bench according to claim 14 , wherein the tool includes a plurality of annular references extending around the longitudinal axis of the tool.16. A bench according to claim 15 , wherein the annular references are formed by applying annular marks or by forming annular grooves to an outside surface of the tool.17. A bench according to claim 14 , wherein the tool is generally cylindrical in shape.18. A bench according to claim 14 , wherein the tool is tubular and the endoscopic means extends inside the tool.19. A bench according to claim 14 , wherein the endoscopic means comprises ultraviolet light guide ...

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19-04-2012 дата публикации

Combustion chamber comprising a condensation-proof barrier on a regenerative circuit

Номер: US20120090292A1
Автор: Daniel Cornu
Принадлежит: SNECMA SAS

The invention concerns a combustion chamber ( 10 ) comprising a neck ( 15 ) downstream of the injection ( 11 ) of gases, and downstream of this neck a divergent section ( 20 ) whereof the outer face of the wall ( 30 ), when in operation, is cooled by a cooling system using a cryogenic product and surrounding this outer face. This divergent section ( 20 ), on the inner face ( 32 ) of its wall ( 30 ), comprises a coating ( 40 ) acting as temperature compensator so that the temperature of the inner face ( 42 ) of the coating ( 40 ) is higher than the condensation temperature of the combustion gases on this inner face ( 42 ) under operating conditions, such that no condensation is formed on this inner face ( 42 ).

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19-04-2012 дата публикации

UNDUCTED PROPELLER WITH VARIABLE PITCH BLADES FOR A TURBOMACHINE

Номер: US20120093652A1
Автор:
Принадлежит: SNECMA

A turbomachine including at least one unducted propeller with variable pitch blades, the blades being carried by respective plates mounted to pivot in radial housings of a rotor element and connected to a control ring that is driven in rotation about the axis of the turbomachine together with the rotor element, and that is movable in translation along the axis to pivot the plates about their axes, the control ring being centered and guided in rotation about the axis on a mechanism that is stationary in rotation and movable in translation along the axis by an actuator carried by the stator of the turbomachine. 17-. (canceled)8. A turbomachine comprising:at least one unducted propeller of variable pitch blades, the blades being carried by substantially cylindrical plates mounted to pivot about their respective axes in radial housings of an annular rotor element and connected via their radially inner ends to a control ring that is driven in rotation about the axis of the turbomachine together with a rotor element, and that is movable in translation along the axis to cause the plates to pivot about their respective axes, the control ring being centered and guided in rotation about the axis of the turbomachine on means for centering and guiding the control ring that is stationary in rotation and movable in translation along the axis by an actuator carried by a stator of the turbomachine,wherein the means for centering and guiding the control ring comprises an annular rail of substantially U-shaped section that extends around the axis of the turbomachine and that comprises two side walls that define between them an annular groove that is outwardly open, and in which the control ring is engaged, and further comprising rolling means being mounted on either side of the control ring between the ring and the side walls of the rail.9. A turbomachine according to claim 8 , wherein the rail includes two annular parts of common axis that are fastened to each other claim 8 , a ...

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19-04-2012 дата публикации

PROCESS FOR DEPOSITING A COATING FOR PROTECTION AGAINST OXIDATION AND AGAINST HOT CORROSION ON A SUPERALLOY SUBSTRATE, AND COATING OBTAINED

Номер: US20120094148A1
Принадлежит: SNECMA

Process for depositing a coating for protection against oxidation and against hot corrosion on a superalloy substrate, and coating obtained 1. Process for depositing a coating for protection against oxidation and against hot corrosion on a metallic superalloy substrate (1) , characterized by the fact that it comprises the deposition of the following successive layers on the substrate ,a first layer of aluminium and of at least one element capable of being alloyed with sulphur,a second layer of a material that isolates said element capable of being alloyed with sulphur.2. Process according to claim 1 , comprising a step of depositing a platinum third layer.3. Process according to claim 2 , the deposition of platinum being followed by a diffusion heat treatment.4. Process according to claim 1 , the aluminium and the element capable of being alloyed with sulphur being deposited via a thermochemical route.5. Process according to claim 1 , the element capable of being alloyed with sulphur being chosen from a reactive element such as zirconium claim 1 , hafnium claim 1 , yttrium or silicon or a rare earth element such as cerium or lanthanum or gadolinium.6. Process according to claim 1 , said isolating material being aluminium.7. Process according to claim 6 , the aluminium being deposited by CVD and especially by APVS.8. Process according to claim 2 , comprising the formation of a layer containing platinum and at least 35 wt % of nickel.9. Process according to claim 8 , the formation of said layer being obtained by depositing a layer of platinum followed by a layer of nickel via a thermochemical route.10. Coating for protection against oxidation and against hot corrosion of a metallic superalloy substrate comprising an additional layer and a diffused layer with a βNiAl phase claim 8 , the element platinum and at least one additional element from the group formed of the reactive elements zirconium claim 8 , hafnium claim 8 , yttrium or silicon and of the rare earth ...

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26-04-2012 дата публикации

DEVICE FOR SPACING ELECTRICAL HARNESSES IN A TURBOMACHINE

Номер: US20120097443A1
Принадлежит: SNECMA

The invention relates to a spacer device for spacing electrical harnesses in a turbomachine such as an airplane turboprop or turbojet, the device comprising clips fastened on a support. A first clip is mounted on a harness via retention means for preventing the device from moving relative to the harness, and the device includes a strip mounted around the first clip in order to close it after a harness has been mounted in the clip, the strip being fastened to the support in non-releasable manner. 1. A spacer device for spacing electrical harnesses in a turbomachine , such as an airplane turboprop or turbojet , the device comprising clamping spring clips for clamping onto the harnesses , said clips being mounted spaced apart from one another on a common support , wherein a first clip is mounted on a harness via retention means for preventing the device from moving relative to the harness , and wherein the device includes a strip mounted around the first clip to close it once a harness has been mounted in the clip , the strip being fastened in non-releasable manner on said support.2. A device according to claim 1 , wherein the strip comprises a U-shaped middle portion surrounding the first clip and two tabs formed at the ends of the U-shape and fastened to the above-mentioned support by rivets or by spot-welds.3. A device according to claim 1 , wherein the retention means comprise a sheath of plastics material surrounding the harness and having the clip mounted thereon.4. A device according to claim 3 , wherein the sheath is made of heat-shrink material.5. A device according to claim 1 , wherein the support is a plane or bent strip of rectangular shape.6. A device according to claim 1 , wherein the first clip is fastened to a middle portion of the support.7. A device according to claim 1 , wherein the other clips are fastened to the support and/or to the strip for closing the first clip.8. A turbomachine such as an airplane turboprop or turbojet including a casing ...

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26-04-2012 дата публикации

DEVICE FOR THE NONDESTRUCTIVE TEST OF A PART

Номер: US20120098531A1
Принадлежит: SNECMA

A device for testing to detect defects at a surface or at shallow depth in a part, for example a blade root for an airplane engine fan. The device includes a probe including a sensor, the probe being hinge-mounted to the end of a handle, a guide presenting a reference surface, and a mechanism adjusting the position of the guide parallel to an axis of the handle. 19-. (canceled)10. A device for non-destructive testing of a part by moving a sensor over a portion to be scanned , the device comprising:a probe including the sensor, the probe being hinge-mounted at an end of a handle;a guide presenting a reference surface; andmeans for adjusting a position of the guide in a direction parallel to an axis of the handle.11. A device according to claim 10 , wherein the guide is generally in a form of a sleeve that is coaxial with the handle from which the probe emerges claim 10 , one end of the sleeve presenting an annular front face constituting the reference surface.12. A device according to claim 11 , wherein the probe is hinge-mounted to a support to be configured to pivot about an axis perpendicular to the axis of the handle claim 11 , the support being installed in the sleeve at one end of the handle.13. A device according to claim 11 , wherein the handle includes a threaded segment having mounted thereon a nut that is secured to the sleeve.14. A device according to claim 13 , wherein the threaded segment is tubular.15. A device according to claim 14 , wherein an inside wall of the tubular threaded segment is lined with a tube projecting into the sleeve and constituting a portion of the support.16. A device according to claim 15 , wherein the tube constitutes a duct for passing electric wires claim 15 , the electric wires being connected to the sensor of the probe.17. A device according to claim 10 , wherein the sensor is an eddy current sensor.18. A device according to claim 10 , wherein the guide and the probe are respectively shaped to come into contact with an inner ...

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26-04-2012 дата публикации

AEROENGINE FAN CASING MADE OF COMPOSITE MATERIAL, AND A METHOD OF FABRICATING IT

Номер: US20120099981A1
Принадлежит: SNECMA

An aeroengine fan casing is made of composite material comprising fiber reinforcement densified by a matrix. The casing () is secured to at least one ring () or ring sector that is situated on the inside face of the casing and that is mechanically connected to an equipment fastener part () that is situated on the outside face of the casing, in particular for fastening an accessory gearbox. 1. An aeroengine fan casing made of composite material having fiber reinforcement densified by a matrix , the casing being secured to at least one ring or ring sector that is situated on the inside face of the casing and that is mechanically connected through the casing to at least one equipment fastener part situated on the outside face of the casing.2. A casing according to claim 1 , wherein the ring or ring sector is made of metal.3. A casing according to claim 1 , wherein the ring or ring sector is made of composite material having fiber reinforcement densified by a matrix.4. A method of fabricating a turbine casing according to claim 1 , the method comprising:making a fiber preform for the casing; anddensifying the fiber preform while shaped by means of a mold, densification comprising impregnating the preform with a liquid composition that is a precursor of the matrix of the composite material, and obtaining the matrix by transforming the precursor;the ring or ring sector being incorporated in the casing during the operations of shaping and densifying the casing preform.5. A method according to claim 4 , wherein the ring or ring sector is prefabricated and is put into place in the mold so as to be adjacent to the casing preform prior to impregnation.6. A method according to claim 4 , wherein a fiber structure forming the fiber reinforcement of a ring or a ring sector made of composite material is put into place in the mold so as to be adjacent to the casing preform claim 4 , and said fiber structure and the casing preform are co-densified.7. A gas turbine aeroengine ...

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26-04-2012 дата публикации

STATIONARY ACTUATOR DEVICE FOR CONTROLLING THE PITCH OF FAN BLADES OF A TURBOPROP

Номер: US20120099987A1
Принадлежит: SNECMA

A device for controlling pitch of fan blades of a turboprop including at least one set of variable-pitch fan blades constrained to rotate with a rotary ring centered on a longitudinal axis and mechanically connected to a turbine rotor, each blade coupled for pitch adjustment to a synchronization ring. A turntable is mounted via a rotary connection to a rod of an actuator that is secured to a stationary structural element of the turboprop, the turntable being mechanically connected to the synchronization ring by a plurality of connection arms hinge-mounted on the turntable and connected to the synchronization ring such that a longitudinal movement of the turntable under drive from the actuator causes the synchronization ring to turn about the longitudinal axis. 111-. (canceled)12. A device for controlling pitch of fan blades of a turboprop including at least one set of variable-pitch fan blades , the set being constrained to rotate with a rotary ring centered on a longitudinal axis and mechanically connected to a rotor of the turbine , each blade of the set being coupled for pitch adjustment to a synchronization ring centered on the longitudinal axis , the device comprising:a turntable centered on the longitudinal axis and mounted via a rotary connection to a rod of an actuator secured to a stationary structural element of the turboprop,the turntable being mechanically connected to the synchronization ring by a plurality of connection arms hinge-mounted to the turntable and connected to the synchronization ring such that longitudinal movement of the turntable under drive from the actuator causes the synchronization ring to turn about the longitudinal axis.13. A device according to claim 12 , wherein each connection arm comprises an axial link having one end fastened to the turntable by a connection that pivots about a tangential axis claim 12 , and a radial link having one end fastened to a free end of the axial link by a connection that pivots about a tangential ...

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26-04-2012 дата публикации

Drive mechanism for a pair of contra-rotating propellers through an epicyclic gear train

Номер: US20120099988A1
Принадлежит: SNECMA SAS

A turbine driving a planet gear and an epicyclic gear train and including a planet pinion cage and a ring driving two propellers in rotation, is connected to the planet gear through a flexible sleeve surrounding a turbine support shaft rather than through a support shaft itself to achieve a flexible assembly with a limit stop position in contact with the shaft to limit parasite internal forces applied to the epicyclic gear train without tolerating a loose assembly or breakage of the sleeve due to a condition of the limit stop after a clearance has been eliminated.

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26-04-2012 дата публикации

PROPELLER HUB

Номер: US20120099989A1
Принадлежит: SNECMA

A hub of a propeller with variable-pitch blades for a turbine engine, for example for a propfan engine. The hub of the propeller includes a polygonal ring with substantially radial cylindrical recesses distributed about a central axis of the ring for receiving the blades, a turbine rotor element of the turbine engine, and a securing flange that is attached to the ring so as to connect the ring to the rotor element. The hub further includes a plurality of back-up hooks inserted with clearance in openings, the back-up hooks being connected either to the ring or to the rotor element, and the openings being connected to the other one of the two. 110-. (canceled)11. A propeller hub with variable-pitch blades for a turboengine , comprising:a polygonal ring including substantially radial cylindrical recesses distributed about a central axis of the ring for receiving the blades;a turbine rotor element of the turboengine;a securing flange attached to the ring so as to connect the ring to the rotor element; anda plurality of back-up hooks inserted with clearance in openings, the back-up hooks being connected to either the ring or the rotor element, and the openings being connected to the other one of the two.12. The hub as claimed in claim 11 , wherein the back-up hooks are held by radial plates attached to the rotor element.13. The hub as claimed in claim 12 , wherein each radial plate bears at least two pins forming back-up hooks and that have a substantially circular cross section.14. The hub as claimed in claim 12 , wherein the back-up hooks have a non-round cross section claim 12 , as in the openings.15. The hub as claimed in claim 12 , wherein the securing flange is attached to a side of the ring claim 12 , and the plates are placed on a side axially opposite the ring.16. The hub as claimed in claim 12 , further comprising means for attaching the plates in a direction parallel to the central axis of the ring.17. The hub as claimed in claim 16 , wherein the plates ...

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26-04-2012 дата публикации

METHOD AND A DEVICE FOR MONITORING A SERVO-CONTROL LOOP OF AN ACTUATOR SYSTEM FOR ACTUATING VARIABLE-GEOMETRY COMPONENTS OF A TURBOJET

Номер: US20120101706A1
Принадлежит: SNECMA

A method for monitoring a servo-control loop () of an actuator system () for actuating variable-geometry components of a turbojet, said method comprising: 1322252122. A monitoring method for monitoring a servo-control loop () of an actuator system () for actuating variable-geometry components of a turbojet , the actuator system () comprising at least a servovalve () , and first and second actuators ( , ) , said method comprising:{'b': 20', '1', '17', '4', '14', '2, 'an estimation step (E) of estimating a plurality of monitoring parameters (m-m) from operating data (d-d) of the servo-control loop ();'}{'b': '30', 'an evaluation step (E) of evaluating a plurality of indicators (i_EVS, i_EVA, i_EVB, i_CID, i_CINT, i_EPA, i_EPB, i_SOMA, i_SOMB, and i_EWRAP) from the monitoring parameters;'}{'b': '40', 'an evaluation step (E) for evaluating at least one signature matrix, each signature matrix being representative of the values of at least some of the indicators; and'}{'b': '50', 'a detection and location step (E) of detecting and locating a degradation affecting the servo-control loop as a function of said at least one signature matrix;'}{'b': '20', 'claim-text': [{'b': 1', '2', '21', '22, 'a category of parameters representative of positions (VSV, VSV, VSVsel) of the actuators (, );'}, {'b': '25', 'a category of parameters representative of coefficients of autoregressive models used for predicting actuator positions as a function of a control current (iCMD) of the servovalve (); and'}, {'b': 25', '3, 'a category of parameters representative of the control current (iCMD) of the servovalve () or an integral current of the servo-control loop ().'}], 'wherein, during the estimation step (E), an estimate is made of a plurality of parameters selected from at least one of the following categories24713. A monitoring method according to claim 1 , wherein at least one of said parameters (m-m) is estimated from an autoregressive model depending on at least one exogenous variable ( ...

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03-05-2012 дата публикации

DEOILER LAYOUT

Номер: US20120102900A1
Принадлежит: SNECMA

A deoiler including a hub mounted on the vent shaft and a cover mounted on the hub between two stop faces which fix the cover avoiding the need to weld or bolt the cover. 16-. (canceled)7. A deoiler comprising:a hub including a cylindrically shaped sleeve for mounting on a shaft and with a hub plate extending beyond the sleeve; anda cover including a cover plate and a cylindrically shaped spacer surrounding the sleeve,wherein the hub plate includes a rim in which the spacer is engaged at one end, and the cover plate is engaged around the sleeve and retained in an elongation direction of the shaft by a first stop face forming part of the sleeve.8. A deoiler according to claim 7 , wherein the first stop face includes a cleat on the sleeve.9. A deoiler according to claim 7 , wherein the cover plate includes a projecting rim claim 7 , opposite the spacer and extending the spacer.10. A deoiler layout comprising:a turbine shaft;a vent shaft surrounded by the turbine shaft;{'claim-ref': {'@idref': 'CLM-00007', 'claim 7'}, 'the deoiler according to ;'}wherein the sleeve of the hub is mounted on the vent shaft, the sleeve is retained in an elongation direction of the shafts by an element of the vent shaft and an element of the turbine shaft, and the cover plate is held in place in the shaft elongation direction between the first stop plate forming part of the sleeve and a second stop plate being secured to either the sleeve or the turbine shaft.11. A deoiler layout according to claim 10 , wherein the second stop face forms part of a ring engaged on the sleeve.12. A deoiler layout according to claim 10 , wherein the second stop face is one end of the turbine shaft. The subject of this invention is the layout of a deoiler located at the end of the turbine shaft of a turbomachine.This type of deoiler (for which a description of a model is given in document EP 0 780 546 A) is placed at the exit from an airflow maintained in the machine to transport oil that lubricates bearings ...

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03-05-2012 дата публикации

DEVICE AND METHOD FOR POSITIONING VARIABLE-GEOMETRY EQUIPMENT FOR A TURBOMACHINE, USING A RELATIVE-MEASUREMENT JACK

Номер: US20120107088A1
Принадлежит: SNECMA

A device for controlling positioning of variable-geometry equipment of a turbomachine, including a computer, an actuator of variable geometry driven by the computer, and a drive train, the actuator including moving parts including a sensor for measuring its elongation, the drive train being connected at one of its ends to a point of attachment of the moving parts and at another end to a point of attachment of the equipment, the point of attachment moving under action of the actuator along a travel limited by an end stop and the drive train being elastically deformable under the action of the actuator when the point of attachment is against the end stop. An elongation instruction supplied by the computer to the moving parts is defined as a difference with respect to the value of the elongation of the moving parts that corresponds to contact between the point of attachment and the end stop. 19-. (canceled)10. A device for controlling positioning of an item of variable-geometry equipment of a turbomachine , comprising:a computer;an actuator for actuating the variable geometry driven by the computer and a kinematic system, the actuator comprising a movable assembly including a sensor for measuring its elongation, the kinematic system being connected at one of its ends to a coupling point of the movable assembly and at another end to an attachment point of the item of equipment, the attachment point moving under action of the actuator along a stroke limited by an abutment and the kinematic system configured to be deformed elastically under the action of the actuator when the attachment point is on the abutment,wherein an elongation setpoint supplied by the computer to the movable assembly is defined as a difference relative to a value of the elongation of the movable assembly that corresponds to a contact of the attachment point with the abutment.11. The device as claimed in claim 10 , wherein the computer detects the value of elongation of the movable assembly that ...

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03-05-2012 дата публикации

THERMAL PROTECTION COATING FOR A TURBINE-ENGINE PART, AND A METHOD OF MAKING IT

Номер: US20120107110A1
Автор:
Принадлежит: SNECMA

A thermal protection coating, in particular for a turbine-engine part (), the coating being deposited by thermal spraying onto the surface of the part () and including at least 80% by volume of hollow ceramic microbeads distributed in a metal alloy based on nickel or cobalt, it being possible for the coating to be deposited on a bonding layer () of metal alloy and for it to be covered in a layer () providing protection against erosion or against friction wear, or in a reflective layer that reflects thermal radiation. 1. A thermal protection coating comprising (i) at least 80% by volume of hollow ceramic microbeads , based on the total volume of the coating , and (ii) a nickel or cobalt metal alloy , wherein the microbeads are distributed in the metal alloy.2. The coating of claim 1 , comprising at least 90% by volume of hollow ceramic microbeads claim 1 , based on the total volume of the coating.3. The coating of claim 1 , wherein the coating has a thickness of less than or equal to 5 mm.4. The coating of claim 1 , wherein the microbeads have a diameter of 30 μm to 250 μm.5. The coating of claim 1 , wherein the metal alloy comprises aluminum.6. The coating of claim 1 , wherein the metal alloy comprises chromium.7. The coating of claim 1 , wherein the metal alloy comprises yttrium.8. A method of producing the coating of claim 1 , the method comprising:thermally spraying with a plasma torch a mixture of hollow ceramic microbeads and a powder of a nickel or cobalt metal alloy onto a surface of a part, wherein the powder and the ceramic microbeads are injected into a plasma cone of the plasma torch simultaneously and laterally, the powder upstream from the microbeads.9. The method of claim 8 , wherein the thermally spraying further comprises forming an alloy bonding layer on the surface of the part claim 8 , wherein the layer has a thickness of 50 μm to 200 μm claim 8 , and the layer comprises the same nickel or cobalt metal alloy as the thermal protection coating.10. ...

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03-05-2012 дата публикации

REINFORCED FAN BLADE SHIM

Номер: US20120107125A1
Принадлежит: SNECMA

A shim configured to be inserted between a fan blade root of a turbojet and a bottom of a compartment in which this root is housed, the compartment being delimited by a fan disk. The shim includes a metal stiffener including at least one external element made of an elastomer material, and including a support surface of the external element. The support surface includes at least one corrugated zone. 19-. (canceled)10. A shim configured to be inserted between a root of fan blade of a turbojet and a bottom of a compartment in which the root is housed , the compartment being delimited by a fan disk , the shim comprising:a metal stiffener including at least one external element made of an elastomer material, the stiffener including a support surface of the external element made of an elastomer material,wherein the support surface includes at least one corrugated zone.11. A shim according to claim 10 , wherein support surface includes two corrugated zones oriented in opposite directions.12. A shim according to claim 10 , wherein the shim is in a form of a strip extending along a longitudinal direction claim 10 , and the corrugated zone comprises a plurality of waves succeeding each other along the longitudinal direction.13. A shim according to claim 11 , wherein the shim is in a form of a strip extending along a longitudinal direction claim 11 , and the corrugated zone comprises a plurality of waves succeeding each other along the longitudinal direction.14. A shim according to claim 10 , wherein the external element is insert molded onto the metal stiffener.15. A shim according to claim 10 , wherein the metal stiffener is made of titanium.16. A turbojet fan comprising:a plurality of fan blades and a disk defining a plurality of compartments around its periphery,{'claim-ref': {'@idref': 'CLM-00010', 'claim 10'}, 'the root of each fan blade being housed in one of the compartments and a shim according to being inserted between the bottom of the compartment and said root.'}17 ...

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03-05-2012 дата публикации

METHOD AND A DEVICE FOR MONITORING A REDUNDANT MEASUREMENT SYSTEM

Номер: US20120109486A1
Принадлежит: SNECMA

The invention relates to monitoring a redundant measurement system of an aeroengine by means of an electronic unit. The monitoring comprises a step of obtaining first measurements of a physical magnitude measured in said engine, and a step of obtaining second measurements of said physical magnitude. According to the invention, the method includes a step of calculating detection residues as a function of differences between the first and second measurements, a step of determining a mean-jump flag, a step of determining a variance-jump flag, a step of determining a change-of-slope flag, and a step of generating a diagnostic notice as a function of said mean-jump flag, of said variance-jump flag, and of said change-of-slope flag. 1. A method of monitoring a redundant measurement system for an aeroengine , the method being executed by an electronic unit of said engine , said monitoring method comprising:a step of obtaining first measurements of a physical magnitude measured in said engine; anda step of obtaining second measurements of said physical magnitude;wherein the method comprises:a step of calculating detection residues as a function of differences between the first measurements and the second measurements;a step of determining a mean-jump flag representing a difference between the mean of the distribution of detection residues and the mean of a reference distribution;a step of determining a variance-jump flag representing a difference between the variance of the distribution of the detection residues and the variance of the reference distribution;a step of determining a change-of-slope flag representing a difference between the slope of the distribution of the detection residues and the slope of the reference distribution; anda step of generating a diagnostic notice as a function of said mean-jump flag, of said variance-jump flag, and of said change-of-slope flag.2. A monitoring method according to claim 1 , comprising:a step of obtaining modeled values of said ...

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17-05-2012 дата публикации

TURBOMACHINE NOZZLE COWL HAVING PATTERNS WITH LATERAL FINS FOR REDUCING JET NOISE

Номер: US20120118398A1
Автор:
Принадлежит: SNECMA

An annular cowl for a turbomachine nozzle, the cowl including a plurality of patterns arranged to extend a trailing edge of the cowl and circumferentially spaced apart from one another. Each pattern has an outline of substantially polygonal shape with a base formed by a portion of the trailing edge of the cowl and at least one vertex that is spaced downstream from the base and that is connected thereto by lateral edges, and in each of its lateral edges, each pattern includes at least one fin, each fin being inclined radially relative to the pattern in a plane that is inclined at an angle lying in the range 0° to 45° relative to a radial direction. 19-. (canceled)10: An annular cowl for a turbomachine nozzle , the cowl comprising:a plurality of patterns arranged to extend a trailing edge of the cowl and circumferentially spaced apart from one another, each pattern having an outline of substantially triangular shape with a base formed by a portion of the trailing edge of the cowl and a vertex that is a point of the outline that is spaced furthest downstream away from the base and that is connected thereto by two lateral edges,wherein each pattern includes, in each of its side edges, at least one fin, each fin being inclined radially relative to the pattern in a plane that is inclined at an angle lying in a range 0° to 45° relative to a radial direction.11: A cowl according to claim 10 , wherein each fin includes an upstream end connected to the lateral edge of the pattern that is spaced apart from the base of the pattern by a distance corresponding to at least 15% of the distance between the base and the vertex of the pattern claim 10 , and a downstream end connected to the lateral edge of the pattern that is spaced apart from the vertex of the pattern by a distance corresponding to at least 15% of the distance between the base and the vertex of the pattern.12: A cowl according to claim 10 , wherein the end of each fin that is furthest from the corresponding side edge ...

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17-05-2012 дата публикации

BEAM FOR SUSPENDING A TURBOSHAFT ENGINE FROM AN AIRCRAFT STRUCTURE

Номер: US20120119056A1
Принадлежит: SNECMA

A beam for suspending a turboshaft engine from an aircraft structure, including a first attachment mechanism configured to be secured to the aircraft structure and at least one second attachment mechanism configured to be secured to the engine. The beam is at least partially made from a metal-matrix composite material including reinforcing fibres. In one embodiment, the beam takes a form of a circle arc. 117-. (canceled)18. A beam for suspending a turboshaft engine from a structure of an aircraft , extending overall in a direction , and comprising:at least one first attachment means to be fixed to the aircraft structure;at least one second attachment means to be fixed to the engine; andwherein the beam is made at least in part of metal matrix composite incorporating reinforcing fibers, parallel to the direction.19. The beam as claimed in claim 18 , in a shape of an arc of a circle claim 18 , comprising two second attachment means each arranged at an end of the beam claim 18 , the first attachment means being formed between the two second attachment means claim 18 , in the middle.20. The beam as claimed in claim 19 , forming an arc of a circle subtending between 40° and 180°.21. The beam as claimed in claim 19 , comprising reinforcing fibers extending between two attachment means claim 19 , from one of the second attachment means to the other of the second attachment means claim 19 , the second attachment means being formed of clevises.22. The suspension beam as claimed in claim 18 , in a form of a ring claim 18 , the ring configured to encircle the turboshaft engine.23. The beam as claimed in claim 22 , in which the reinforcing fibers of the ring are arranged in rings that are concentric with the ring.24. The beam as claimed in claim 22 , in which the second attachment means claim 22 , formed of clevises claim 22 , is distributed about the periphery of the ring.25. The beam as claimed in claim 18 , further comprising:a mounting plate configured to accept fixing ...

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17-05-2012 дата публикации

BLADE RETENTION DISK

Номер: US20120121428A1
Принадлежит: SNECMA

A blade retention disk includes, on its external surface, a peripheral annular groove configured to receive a plurality of blade roots of the hammer attachment type, the peripheral annular groove including a loading orifice intended for the introduction of blade roots inside the peripheral annular groove. The retention disk also includes: a crown covering the peripheral annular groove; a crown centering device to center the crown, the crown centering device and the crown forming an air supply enclosure from the bottom of the peripheral annular groove; at least one orifice opening in the air supply enclosure, the orifice being configured to introduce cooling air into the air supply enclosure. The blade retention disk finds a direct application in the field of low pressure turbines for aircraft turbine engines. 1. A blade retention disk comprising , on its external surface , a peripheral annular groove configured to receive a plurality of blade roots of the hammer attachment type , said peripheral annular groove comprising a loading orifice configured for the introduction of blade roots inside said peripheral annular groove , said retention disk comprising:a crown covering said peripheral annular groove;a crown centering device configured to center said crown, said crown centering device and said crown forming an air supply enclosure from the bottom of said peripheral annular groove;at least one orifice opening in said air supply enclosure, said orifice being configured to introduce cooling air into said air supply enclosure.2. The blade retention disk according to claim 1 , wherein the crown centering device is formed by an annular spoiler axially extending the retention disk at a level of the upstream periphery of said retention disk.3. The blade retention disk according to claim 1 , wherein the crown comprises a hook configured to cooperate with the crown centering device.4. The blade retention disk according to claim 1 , wherein the crown is formed by a metal band ...

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17-05-2012 дата публикации

METHOD AND SYSTEM FOR CONTROLLING AIRCRAFT ENGINE STARTER/GENERATOR

Номер: US20120122631A1
Автор:
Принадлежит: SNECMA

A system and a method for controlling an aircraft engine starter/generator. The system includes an AGB of fixed gear ratio for coupling mechanically to a turbine shaft of the engine to enable the engine to be started, a gearbox having multiple gear ratios mechanically coupled to a gearwheel of the AGB, a starter/generator mechanically coupled to a gearwheel of the gearbox, and a controller to cause the gear ratio of the gearbox to be changed as a function of the mode of operation of the starter/generator. 1. A system for controlling a starter/generator of an aircraft engine , the system comprising:an AGB of fixed gear ratio for coupling mechanically to a turbine shaft of the engine, enabling the engine to be started;a starter/generator mechanically coupled to the AGB via a gearbox having multiple gear ratios and interposed between the AGB and the starter/generator; andmeans for controlling a change in the gear ratio of the gearbox as a function of the mode of operation of the starter/generator.2. A system according to claim 1 , wherein the gear ratio of the gearbox is changed as soon as the starter/generator passes from driving to driven relative to the turbine shaft.3. A system according to claim 1 , wherein the gear ratio of the gearbox is changed as soon as an electrical command of the starter/generator passes from one mode of operation to the other.4. A system according to claim 1 , wherein the means for changing the gear ratio of the gearbox include an electronic unit that controls an actuator device for actuating the gearbox.5. A system according to claim 4 , wherein the electronic unit for controlling the engine is connected to a sensor for sensing the speed of rotation of the turbine shaft.6. A system according to claim 4 , wherein the gearbox has parallel gearwheels.7. A system according to claim 4 , wherein the gearbox actuator device is an electrical claim 4 , hydraulic claim 4 , or pneumatic actuator.8. A system according to claim 1 , wherein the gear ...

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24-05-2012 дата публикации

METHOD FOR MANUFACTURING A FORGED PART WITH ADAPTIVE POLISHING

Номер: US20120124834A1
Принадлежит: SNECMA

A method for manufacturing a part by forging, including producing a semifinished part by precision forging and polishing the part by an abrasive strip, compliant geometric characteristics of the part to be obtained being predetermined in a theoretical model. The method includes: measuring the geometrical characteristics of the semifinished part after the forging operations and comparing the characteristics with the theoretical model; determining noncompliant areas on the surface of the part; determining the amount of material to be removed from each noncompliant area to make the area compliant; and polishing the part using the abrasive strip, controlling the strip so as to remove the amount of material from each noncompliant area. The method can be used for example for polishing turbine engine fan blades. 19-. (canceled)10. A method of manufacturing a component by forging , comprising:producing a semifinished component by precision forging and polishing the component using an abrasive belt, compliant geometric characteristics of the component to be obtained being determined in a theoretical model;measuring geometric characteristics of the semifinished component after the forging operations and comparing against the theoretical model;determining, on a surface of the component, those zones which are non-compliant;determining an amount of material to be removed in each non-compliant zone to make it compliant; andpolishing the component using the abrasive belt, controlling the belt so as to remove the amount of material in each non-compliant zone.11. The method as claimed in claim 10 , in which a plurality of measurement points is defined at a surface of the component claim 10 , the geometric characteristics of the semifinished component are measured at least some of the measurement points claim 10 , and removal of material by the abrasive belt is controlled at the measurement points based on a discrepancy between the geometric characteristics of the semifinished ...

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24-05-2012 дата публикации

SEALING SYSTEM WITH ANNULAR BRUSH SEAL

Номер: US20120126484A1
Принадлежит: SNECMA

The invention concerns a system for providing a seal between two volumes (V, V) delimited by a part (A) mounted to rotate about an axis (A) and a stator (), said system (S) including an annular brush seal () that comprises: 21181210. The sealing system claimed in the preceding claim , wherein the adjustment means include inclination members () mounted on and firmly fastened to the brush body () and associated with respective pressure members () adapted to act on said corresponding inclination members to modify the inclination α of the fibers ().311127. The sealing system claimed in the preceding claim , wherein the inclination members () and the corresponding pressure members () are distributed around a circumference of the brush seal ().411127. The sealing system claimed in the preceding claim , wherein the distribution of the inclination members () and the corresponding pressure members () is regular around a circumference of the brush seal ().5. The sealing system claimed in the claim 2 , wherein{'b': '11', 'each inclination member takes the form of a projecting tongue (); and'}{'b': 12', '12', '11', '8, 'each pressure member takes the form of a screw () the free end (A) of which is adapted to come into contact with the corresponding inclination member () firmly fastened to the brush body () to modify the angle α.'}61261189. The sealing system claimed in the claim 2 , wherein the pressure members () are mounted on the stator () so as to be able to operate on the corresponding inclination members () of the brush body () claim 2 , the latter being housed in a complementary groove () provided in the wall of the stator.781112. The sealing system claimed in the claim 1 , wherein the brush body () of the brush seal is formed from a deformable material adapted to accept the forces imposed by the means ( claim 1 , ) for adjusting the angle α.8. An aircraft engine including at least one sealing system (S) as specified in the . The present invention concerns a system for ...

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24-05-2012 дата публикации

NOZZLE STAGE FOR A TURBINE ENGINE

Номер: US20120128476A1
Принадлежит: SNECMA

A nozzle stage for a turbine engine includes two substantially cylindrical rings, respectively an inner ring and an outer ring, with, extending between them: vanes carried by a first one of the rings; vanes carried by the second ring; and spacer-forming vanes interconnecting the two rings; the vanes of the first and second rings being arranged in alternation. 113-. (canceled)14. A nozzle stage for a turbine engine , the stage comprising:two substantially cylindrical rings, of respectively an inner ring and an outer ring, with substantially radial vanes extending between them;a first series of vanes carried by a first one of the rings;a second series of vanes carried by the second ring; andconnection means interconnecting the two rings, the vanes of the first ring alternating with the vanes of the second ring.15. A nozzle stage according to claim 14 , wherein the vanes of each series are substantially regularly distributed around the longitudinal axis of the stage.16. A nozzle stage according to claim 14 , wherein the ring connection means comprises spacer vanes having their ends fastened to the first ring and to the second ring claim 14 , respectively.17. A nozzle stage according to claim 14 , wherein each ring is sectorized and comprises at least two annular sectors mounted end to end.18. A nozzle stage according to claim 16 , wherein each ring is sectorized and comprises at least two annular sectors mounted end to end claim 16 , and further comprising an annular rail mounted on the inner ring and including an outer annular groove in which an inner annular rim of each sector of the inner ring and the radially inner end of each spacer vane are engaged by sliding in the circumferential direction.19. A nozzle stage according to claim 18 , further comprising undulating spring-forming blades interposed between a bottom of the annular groove in the rail and both the sectors of the inner ring and roots of the spacer vanes.20. A nozzle stage according to claim 14 , wherein ...

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24-05-2012 дата публикации

ANTI-WEAR DEVICE FOR THE BLADES OF A TURBINE DISTRIBUTOR IN AN AERONAUTICAL TURBINE ENGINE

Номер: US20120128481A1
Принадлежит: SNECMA

A blade sector of a turbine distributor to be carried by a turbine casing of an aeronautical turbine engine, including a front connection mechanism and a rear connection mechanism, wherein the front connection mechanism is adapted so as to bear against a holder carried by the turbine casing. The sector further includes an anti-wear device formed by a piece of a metal material surrounding the front end of the front connection mechanism and provided between the front connection mechanism and the holder to ensure a sliding contact between the two parts. The anti-wear device is axially maintained in position on the blade sector using an attachment mechanism engaging with the front connection mechanism. 114-. (canceled)15: A turbine distributor blade segment configured to be supported by a turbine casing of an aeronautical turbo-engine , comprising:a front connection means and a rear connection means, the front connection means configured to rest on a support which is supported by the turbine casing;an anti-wear device including a part made of metal material that envelops a front end of the front connection means, and is interposed between the front connection means and the support, to provide sliding contact between the front connection means and the support,wherein the anti-wear device is retained axially in position on the blade segment by a securing means that co-operates with the front connection means.16: The blade segment as claimed in claim 15 , wherein the front connection means is in a form of a tongue that extends axially claim 15 , and an upper part of the device does not extend axially along an entire length of the front connection means.17: The blade segment as claimed in claim 15 , wherein the anti-wear device is retained in place on a front connection means by at least one pin that is fitted into a bore provided in a thickness of the connection means.18: The blade segment as claimed in claim 15 , wherein the anti-wear device extends along an entire length ...

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24-05-2012 дата публикации

OUTER SHELL SECTOR FOR A BLADED RING FOR AN AIRCRAFT TURBOMACHINE STATOR, INCLUDING VIBRATION DAMPING SHIMS

Номер: US20120128482A1
Принадлежит: SNECMA

An assembly forming an outer shell sector, for a bladed ring sector configured to be used on a compressor or turbine stator in an aircraft turbomachine, including a plurality of elementary sectors and vibration damping shims each of them being inserted between two elementary sectors associated with it. A profile of each vibration damping shim is approximately the same as a profile of the elementary sectors. 17-. (canceled)8. A bladed ring sector configured to be installed on a compressor stator of an aircraft turbomachine , comprising:an assembly forming an outer shell sector;an inner shell sector;a plurality of blades at a tangential spacing from each other and inserted between the assembly forming the outer shell sector and the inner shell sector, the blades being fixed to each assembly forming the outer shell sector and the inner shell sector; andthe assembly forming an outer shell sector comprising firstly a plurality of elementary sectors at a spacing from each other along a tangential direction of the assembly, and secondly vibration damping shims each of them being inserted between two elementary sectors associated with it, placed directly consecutively along the tangential direction,wherein a profile of each vibration damping shim is approximately a same as a profile of the elementary sectors.9. A sector according to claim 8 , wherein the shim is forced in contact with two parallel plane friction surfaces facing each other along the tangential direction and provided respectively on the two elementary sectors associated with the shim claim 8 , and wherein the shim has two complementary plane friction surfaces claim 8 , parallel to each other and cooperating with the two corresponding friction surfaces of the elementary sectors.10. A sector according to claim 8 , wherein the shim includes hooks to hold it in place on the compressor or turbine stator.11. A sector according to claim 8 , wherein the elementary sectors are separated from each other by radial slits ...

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24-05-2012 дата публикации

PRESTRESSING DEVICE HAVING CIRCUMFERENTIAL ACTIVITY

Номер: US20120128490A1
Автор:
Принадлежит: SNECMA

A prestressing device extending about an axis between a first plane and a second plane, the first plane being located at a predetermined distance from the second plane, each plane being perpendicular to the axis. The device includes at least three deformation areas. The deformation areas extend in the first plane. The device further includes, between each deformation area, a planar contact surface extending in the second plane. 17-. (canceled)8. A prestressing device which extends about an axis between a first plane substantially perpendicular to the axis and a second plane substantially parallel to the first plane and axially offset with respect to the first plane , the device comprising:at least three deformation zones, wherein the deformation zones extend within the first plane; andbetween each deformation zone, a planar contact surface extending within the second plane.9. The device as claimed in claim 8 , further comprising inclined transition zones between each deformation zone and the contact surfaces adjacent to the deformation zone.10. The device as claimed in claim 9 , wherein each inclined transition zone comprises a first rounded end part connected to one of the deformation zones claim 9 , a second end part connected to one of the planar contact surfaces claim 9 , and an essentially planar central part between the first and second end parts.11. The device as claimed in claim 8 , wherein each deformation zone comprises a first recess extending radially inward and a second recess extending radially outward.12. The device as claimed in claim 8 , wherein claim 8 , for each of the contact surfaces claim 8 , two successive planar contact surfaces are separated from each other by a predetermined distance.13. The device as claimed in claim 8 , further comprising a number of planar contact surfaces tabs between ten and twenty claim 8 , or between thirteen and seventeen.14. A rotor claim 8 , or an impeller rotor claim 8 , with no fairing claim 8 , for a turbine ...

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24-05-2012 дата публикации

PRESTRESSING DEVICE HAVING RADIAL ACTIVITY

Номер: US20120128503A1
Автор:
Принадлежит: SNECMA

A prestressing device extending about an axis between a first plane that is substantially perpendicular to the axis and a second plane that is substantially parallel to the first plane and axially offset relative to the latter. The device includes a first annular body extending in the first plane, and a first series of at least three tabs, each of which is connected to the first body and extending in an essentially radial direction, each of the tabs of the first series including a planar surface extending in the second plane. This arrangement ensures surface contact between the parts of the elements to be subjected to prestressing. The device can be used for example for a rotor, and for example for a rotor of an unducted propeller for a turbine engine. 19-. (canceled)10. A prestressing device which extends about an axis between a first plane substantially perpendicular to the axis and a second plane substantially parallel to the first plane and axially offset with respect to the first plane , comprising:a first annular body which extends within the first plane with a first series of at least three tabs extending in an essentially radial direction, each of the tabs of the first series comprising a planar surface extending within the second plane; anda second annular body extending within the first plane with a second series of at least three tabs extending in an essentially radial direction, in the opposite direction from the first series, each of the tabs of the second series comprising a planar surface extending within the second plane.11. The device as claimed in claim 10 , wherein the first annular body is an inner annular body claim 10 , the tabs of the first series extend outward claim 10 , the second annular body is an outer annular body arranged so as to extend around the inner annular body claim 10 , and the tabs of the second series extend inward.12. The device as claimed in claim 10 , wherein the distance between two successive tabs of each of the series ...

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31-05-2012 дата публикации

ROCKET ENGINE WITH CRYOGENIC PROPELLANTS

Номер: US20120131903A1
Автор:
Принадлежит: SNECMA

A cryogenic-propellant rocket engine includes: at least a first tank for a first liquid propellant; a second tank for a second liquid propellant; a third tank for an inert fluid; an axisymmetrical nozzle including a combustion chamber, a device for injecting first and second liquid propellants into the combustion chamber, a nozzle throat, and a divergent section; and a heater device including at least one duct for conveying the inert fluid and arranged outside the nozzle in immediate proximity thereof, but without making contact therewith, to recover energy of thermal radiation emitted when the rocket engine is in operation and to heat the inert fluid. 19-. (canceled)10: A cryogenic-propellant rocket engine comprising:at least a first tank for a first liquid propellant;a second tank for a second liquid propellant;a third tank for an inert fluid;an axisymmetrical nozzle including a combustion chamber, a device for injecting first and second liquid propellants into the combustion chamber, a nozzle throat, and a divergent section;a heater device including at least one duct for conveying the inert fluid and arranged outside the nozzle in an immediate proximity of nozzle, but without making contact with the nozzle, to recover energy of thermal radiation emitted when the rocket engine is in operation and to heat the inert fluid,the heater device comprising a plate in a form of a frustoconical sector that extends around the divergent section of the nozzle over an angle a lying in a range 30° to 360°.11: A rocket engine according to claim 10 , wherein the heater device comprises a metal structure in which the inert fluid for heating flows and a fine layer that is strongly absorbent from a thermal radiation point of view deposited at least on walls that face the nozzle constituting a radiant power source.12: A rocket engine according to claim 10 , wherein the plate of the heater device presents a thickness lying in a range 5 mm to 15 mm.13: A rocket engine according to claim ...

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31-05-2012 дата публикации

Method for producing martensitic steel with mixed hardening

Номер: US20120132326A1
Принадлежит: Aubert and Duval SA, SNECMA SAS

A method of producing a martensitic steel including a content of other metals such that it can be hardened by intermetallic compound and carbide precipitation, with an Al content of between 0.4% and 3%. The heat shaping temperature of a last heat shaping pass of the steel is lower than the solubility temperature of aluminum nitrides in the steel, and a treatment temperature for each potential heat treatment after the last heat shaping pass is lower than the solid-state solubility temperature of the aluminum nitrides in the steel.

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31-05-2012 дата публикации

LOW PRESSURE TURBINE FOR AN AIRCRAFT TURBOMACHINE, COMPRISING A SEGMENTED NOZZLE WITH AN IMPROVED DESIGN

Номер: US20120134788A1
Принадлежит: SNECMA

The invention relates to a low pressure turbine () of a turbomachine comprising a nozzle () belonging to an upstream stage (B) and a nozzle () belonging to a downstream stage (C), the nozzle () forming segments () having an external structure () comprising: 112622016262206262667066a. Low pressure turbine () of a turbomachine comprising a stator provided with a case () , a nozzle () belonging to an upstream stage (B) and a nozzle ( , ′) belonging to a downstream stage (C) , each of the two stages housed in the case also comprising a bladed mobile wheel arranged downstream from the nozzle of its stage , the nozzle () of the upstream stage being segmented so as to form angular nozzle segments () each with an external structure () from which nozzle blades () project radially inwards , characterised in that each external structure () comprises:{'b': 76', '79, 'a first tab () bearing radially on a hook on the stator () fixed to the case,'}{'b': 85', '87', '76', '85', '91', '2, 'a second tab () bearing radially on a reaming () in the case, said first and second tabs (, ) of the external structures of the segments delimiting an annular upstream cooling air circulation cavity (), jointly with said case (),'}{'b': 84', '85', '86', '92', '162', '62', '20', '93', '84', '2', '94', '91, 'a contact surface () extending in the downstream direction from the second tab () and lined with an abradable element () on the inside in contact with the mobile wheel of the upstream stage, said contact surface comprising a downstream end () bearing axially on the nozzle (, ′) of said downstream stage (C) and bearing radially on a stop () of the case or said nozzle in the downstream stage, the contact surfaces () of the external structures of the segments cooperating with said case (), to delimit an annular downstream cooling air circulation cavity () communicating with said upstream cavity ().'}2622062162a. Turbine according to claim 1 , characterised in that the nozzle (′) in the downstream ...

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31-05-2012 дата публикации

HUB FOR A PROPELLER HAVING VARIABLE PITCH BLADES

Номер: US20120134822A1
Принадлежит: SNECMA

A hub for a propeller having variable pitch blades for a turbine engine, for example for a propfan turbine engine. The propeller hub includes: a polygonal ring having substantially radial cylindrical housings distributed about a central axis of the ring for receiving the blades; a rotor element for the turbine of the turbine engine; a supporting flange attached to the ring so as to connect the ring to the rotor element; and a plurality of backup retaining members linked to the rotor element, and each of which includes at least one bearing surface opposite an outer surface of the ring with a radial spacing. 18-. (canceled)9. A propeller hub for a turbomachine with variable pitch blades , comprising:a turbomachine turbine rotor element;a polygonal ring fixed to the rotor element and including substantially radial cylindrical housings distributed around a central axis of the ring to receive the blades; anda plurality of backup retaining members connected to the rotor element, passing radially through the ring through openings in the ring and each including at least one bearing surface opposite an exterior surface of the ring with a radial spacing.10. The hub as claimed in claim 9 , wherein each backup retaining member includes an enlarged head with a bearing surface on each side of the opening.11. The hub as claimed in claim 9 , wherein each backup retaining member includes a dovetail-shape base positively interengaged with a complementary groove on the rotor element.12. The hub as claimed in claim 11 , wherein the base also includes at least one orifice receiving a bolt for limiting movement of the base relative to the rotor element in the direction of the groove.13. The hub as claimed in claim 11 , further comprising an annular stop ring disposed between the base and the rotor element to limit movement of the base relative to the rotor member in the direction of the groove.14. A propeller including a hub according to and blades received in the cylindrical housings.15 ...

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07-06-2012 дата публикации

DEVICE FOR FASTENING AN ELEMENT OF ELONGATE SHAPE TO A TURBINE ENGINE CASING

Номер: US20120137494A1
Принадлежит: SNECMA

A device for fastening an element of elongate shape, such as an electrical harness or a duct, to a casing of a turbine engine. The device includes a tubular portion mounted around the element and secured to at least one tab for fastening to the casing, the tubular portion being made of a heat-shrink plastics material and configured to be heated to be shrunk onto the element. 111-. (canceled)12. A device for fastening an element of elongate shape to a structure of a machine , the device comprising:a clamping portion for mounting around the element and secured to at least one tab for fastening to a casing,wherein the clamping portion is tubular and made of a heat-shrink plastics material, the clamping portion initially having an inside diameter that is greater than the outside diameter of the element so as to allow it to be mounted on the element, and configured to be heated to be shrunk onto the element.13. A device according to claim 12 , wherein the fastener tab is made of heat-shrink plastics material and is made integrally with the tubular portion.14. A device according to claim 12 , wherein the fastener tab includes a metal insert having a portion thereof embedded in the heat-shrink material of the tubular portion.15. A device according to claim 14 , wherein a portion of the fastener tab that is embedded in the heat-shrink material includes means for anchoring it in the material.16. A device according to claim 12 , wherein the fastener tab includes at least one orifice for passing a screw claim 12 , the orifice being oblong or elongate in shape in a transverse direction so as to accommodate assembly offsets between the device and the casing in this direction.17. A device according to claim 12 , wherein the tubular portion is not split in the longitudinal direction.18. An element of elongate shape claim 12 , which carries a fastener device according to .19. A turbine engine claim 12 , comprising at least one fastener device according to .20. A method of fastening ...

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07-06-2012 дата публикации

Combustion chamber for a turbomachine including improved air inlets

Номер: US20120137697A1
Принадлежит: SNECMA SAS

A combustion chamber for a turbomachine, including two coaxial walls including air inlets, each of which is configured such that its orthogonal projection, in a plane passing through the axis of the injection system closest to the inlet and perpendicular to an axial plane passing through this axis and through the axis of the combustion chamber, has an upstream edge of convex shape when seen from downstream.

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07-06-2012 дата публикации

SEGMENTED TURBINE RING FOR A TURBOMACHINE, AND TURBOMACHINE FITTED WITH SUCH A RING

Номер: US20120141257A1
Автор:
Принадлежит: SNECMA

The segmented ring has the gaseous hot flow of the turbomachine passing through it and comprises: 2. The ring as claimed in claim 1 , in which said cooling chamber is defined by two mutually-facing recesses formed in the radial faces of two assembled segments.3. The ring as claimed in the preceding claim claim 1 , in which the mutually-facing recesses that form the cooling chamber each have a substantially rectangular hollow profile claim 1 , with a bottom parallel to said lateral internal face of the corresponding segment claim 1 , two sides parallel to the transverse faces of the segments and in which a slot accommodating the lower sealing strip terminates claim 1 , and a main face set back parallel to the radial face of each of them.4. The ring as claimed in claim 3 , in which the bottom of the recesses is close to the slot accommodating the lower sealing strip with the external fresh air orifices opening into said recesses.5. The ring as claimed in claim 1 , in which the cooling chamber formed by the mutually-facing radial faces of two adjacent segments lies approximately in the central part thereof claim 1 , between their upstream and downstream transverse faces.6. The ring as claimed in claim 1 , in which said lower sealing strip that assembles two consecutive segments comprises several parts claim 1 , a front part between the upstream transverse face and the chamber claim 1 , a central part that hugs the periphery of said chamber claim 1 , and a rear part between the chamber and the downstream transverse face.7. The ring as claimed in claim 6 , in which the central part of each lower sealing strip comprises a straight sub-part attached into the bottom of the mutually-facing recesses and two sub-parts at a right angle claim 6 , one of the branches of each right angle running along the corresponding side of the chamber and the other branch being superposed with the relevant end of the respective front or rear part of the sealing strip.8. A turbomachine ...

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07-06-2012 дата публикации

VIBRATION-DAMPING SHIM FOR FAN BLADE

Номер: US20120141296A1
Автор:
Принадлежит: SNECMA

A vibration-damping shim configured to be interposed between a platform of a fan blade and a fan disk, including a radially external surface fitted with plates in contact with the fan blade platform, and a radially internal surface formed by an upstream surface, configured to be facing the disk, and a downstream surface separated from the upstream surface by a break in alignment. The upstream surface includes a zone protruding radially towards the interior, initiated at some distance from its upstream end. 17-. (canceled)8. A vibration-damping shim configured to be interposed between a platform of a fan blade and a fan disk , comprising:a radially external surface fitted with at least one plate in contact with the fan blade platform; anda radially internal surface formed by an upstream surface, configured to be facing the disk, and a downstream surface separated from the upstream surface by a break in alignment, wherein the upstream surface is located radially towards the interior relative to the downstream surface,wherein the upstream surface includes a zone protruding radially towards the interior, initiated at a distance from its upstream end.9. A damping shim according to claim 8 , further comprising an upstream plate in contact with the fan blade platform claim 8 , and a downstream plate in contact with the fan blade platform claim 8 , positioned respectively upstream and downstream relative to the break in alignment.10. A damping shim according to claim 9 , wherein the protruding zone is located radially perpendicular to the upstream contact plate.11. A damping shim according to claim 8 , further comprising a downstream end surface including a radially higher portion fitted with an axial stop plate.12. A damping shim according to claim 8 , wherein the protruding zone extends axially over approximately 40% to 70% of the upstream surface of the radially internal surface.13. A damping shim according to claim 8 , wherein the break in alignment includes one or more ...

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14-06-2012 дата публикации

ROCKET ENGINE WITH CRYOGENIC PROPELLANTS

Номер: US20120144797A1
Автор:
Принадлежит: SNECMA

A cyrogenic-propellant rocket engine includes: at least a first tank for a first liquid propellant; a second tank for a second liquid propellant; a third tank for an inert fluid; an axisymmetrical nozzle including a combustion chamber, a device for injecting first and second liquid propellants into the combustion chamber, a nozzle throat, and a divergent section; and a heater device including at least one duct for conveying the inert fluid and arranged outside the nozzle in immediate proximity thereof, but without making contact therewith, to recover energy of thermal radiation emitted when the rocket engine is in operation and to heat the inert fluid. 112-. (canceled)13. A cryogenic-propellant rocket engine comprising:at least a first tank for a first liquid propellant;a second tank for a second liquid propellant;a third tank for an inert fluid;an axisymmetrical nozzle including a combustion chamber, a device for injecting first and second liquid propellants into the combustion chamber, a nozzle throat, and a divergent section;a heater device including at least one duct for conveying the inert fluid and arranged outside the nozzle in immediate proximity of the nozzle, but without making contact with the nozzle, to recover energy of thermal radiation emitted when the rocket engine is in operation and to heat the inert fluid; andwherein the heater device comprises a metal structure in which the inert fluid for heating flows and a fine layer that is strongly absorbent from a thermal radiation point of view deposited at least on walls that face the nozzle constituting a radiant power source.14. A rocket engine according to claim 13 , wherein the heater device comprises at least one torus surrounding the nozzle.15. A rocket engine according to claim 13 , wherein the heater device comprises at least one tube wound helically around the divergent section of the nozzle.16. A rocket engine according to claim 14 , wherein the heater device comprises two to four tubes ...

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14-06-2012 дата публикации

Pylon for fixing an aircraft engine having unducted pusher propellers

Номер: US20120145824A1
Принадлежит: SNECMA SAS

A pylon for fixing an aircraft engine having unducted pusher propellers, the pylon ensuring the fixation of a propulsive system on the boattail of the aircraft, the pylon having a trailing edge, with an upper face and a lower face, for an airflow encountered by the pylon, wherein at least one of the two faces of the upper face and the lower face of the trailing edge is inclinable, at least in part.

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14-06-2012 дата публикации

SYSTEM FOR DETECTING AN EPHEMERAL EVENT ON A VANE IMPELLER OF AN AIRCRAFT ENGINE

Номер: US20120148400A1
Принадлежит: SNECMA

The invention relates to a method and a system for monitoring an ephemeral event on a vane impeller () of an aircraft engine (), including: 135. A system for monitoring an ephemeral event on a vane impeller () of an aircraft engine () , characterised in that it includes:{'b': 7', '1', '31', '3, 'acquisition means () for acquiring a time signal (S) relating to movable vanes () of the vane impeller (),'}{'b': 11', '1', '0', '3, 'processing means () for detecting a disturbance in said time signal (S) with respect to a reference time signal (S) corresponding to a same flight phase, the disturbed time signal being indicative of an ephemeral event on said vane impeller ().'}2113. The monitoring system according to claim 1 , characterised in that the processing means () are configured to analyse said disturbed time signal with respect to an angular reference to identify the impacted vane(s) of said vane impeller ().311. The monitoring system according to claim 1 , characterised in that the processing means () are configured to discriminate the nature of the ephemeral event by analysing the damping of said disturbance.49. The monitoring system according to claim 1 , characterised in that it includes alerting means () for indicating the occurrence of the ephemeral event.57713. The monitoring system according to claim 1 , characterised in that the acquisition means () include at least one tip-timing sensor () provided flush with said vane impeller () for acquiring a time signal indicative of a current passage time between the vanes.671732117. The monitoring system according to claim 5 , characterised in that the acquisition means () include a phonic wheel () integral with the vane impeller () and a top tour sensor () provided facing the phonic wheel () to provide an angular reference of the vanes of said vane impeller.7371. The monitoring system according to claim 5 , characterised in that the vane impeller () includes a vane having a singularity and in that the tip-timing ...

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21-06-2012 дата публикации

Turbojet including an automatically variable flow rate bleed circuit for cooling air

Номер: US20120151936A1
Принадлежит: SNECMA SAS

Bleeding cooling air to cool a subassembly, e.g. such as a turbine, with automatic adjustment of the air flow section as a function of the speed of the engine. According to the invention, a shutter element is fastened to co-operate with a bleed hole, with the material that constitutes either the shutter element or the wall in which the hole is formed being of a type in which it is possible to create eddy currents, and a magnet is mounted to move past said arrangement.

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21-06-2012 дата публикации

NONDESTRUCTIVE TEST OF A SEALING MEMBER

Номер: US20120153941A1
Автор:
Принадлежит: SNECMA

A device for inspecting an annular sealing wiper extending at the surface of a bladed-wheel drum of a rotor. The device includes a carriage including at least two guide wheels and carrying a probe situated in a location such that, when the carriage is in position, the probe is positioned facing an edge of the wiper for inspection and at a determined distance therefrom. 19-. (canceled)10. A device for inspecting an annular sealing wiper forming part of a sealing structure extending at a surface of a bladed-wheel drum , comprising:a carriage including at least two spaced-apart guide wheels configured to come into contact with at least a portion of the sealing structure that forms a circular rail; andwherein the carriage carries at least one probe situated at a location such that, when the wheels are in engagement with the circular rail, the probe is to be found positioned facing an edge of the wiper for inspection and at a determined distance therefrom.11. A device according to claim 10 , wherein the wheels are of grooved type claim 10 , and wherein an axis of the probe and bottoms of the grooves of the wheels are substantially coplanar claim 10 , the probe being situated between the two wheels.12. A device according to claim 10 , wherein the carriage includes lateral flanks extending on either side of an end of the probe that is configured to face the wiper for inspection.13. A device according to claim 11 , wherein the carriage includes at least one additional guide wheel situated in a plane parallel to the plane containing the other two guide wheels and configured to come into contact with the circular rail so as to stabilize orientation of the probe relative to the wiper for inspection.14. A device according to claim 12 , wherein the carriage includes at least one additional guide wheel situated in a plane parallel to the plane containing the other two guide wheels and configured to come into contact with the circular rail so as to stabilize orientation of the ...

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28-06-2012 дата публикации

COMPOSITE MATERIAL PART HAVING A CERAMIC MATRIX, AND METHOD FOR MANUFACTURING SAME

Номер: US20120164430A1
Принадлежит: SNECMA PROPULSION SOLIDE

In a composite material part having a ceramic matrix and including a fibrous reinforcement which is densified by a matrix consisting of a plurality of ceramic layers having a crack-diverting matrix interphase positioned between two adjacent ceramic matrix layers, the interphase includes a first phase made of a material capable of promoting the diversion of a crack reaching the interphase according to a first propagation mode in the transverse direction through one of the two ceramic matrix layers adjacent to the interphase, such that the propagation of the crack continues according to a second propagation mode along the interphase, and a second phase consisting of discrete contact pads that are distributed within the interphase and capable of promoting the diversion of the crack that propagates along the interphase according to the second propagation mode, such that the propagation of the crack is diverted and continues according to the first propagation mode through the other ceramic matrix layer that is adjacent to the interphase. 1. A composite material part having a ceramic matrix and comprising a fibrous reinforcement which is densified by a matrix consisting of a plurality of ceramic layers having a crack-diverting matrix interphase positioned between two adjacent ceramic matrix layers , wherein the interphase includes:a first phase made of a material capable of promoting the diversion of a crack reaching the interphase according to a first propagation mode in the transverse direction through one of the two ceramic matrix layers adjacent to the interphase, such that the propagation of the crack continues according to a second propagation mode along the interphase, anda second phase consisting of discrete contact pads that are distributed within the interphase and capable of promoting the diversion of the crack that propagates along the interphase according to the second propagation mode, such that the propagation of the crack is diverted and continues ...

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05-07-2012 дата публикации

ROCKET ENGINE WITH EXTENDABLE DIVERGENT

Номер: US20120167575A1
Принадлежит: SNECMA

A rocket engine with an extendible exit cone including an exhaust nozzle for gas from a combustion chamber, the nozzle presenting a longitudinal axis having a first portion defining a nozzle throat and a stationary first exit cone segment, at least one extendible second exit cone segment of section greater than the section of the stationary first exit cone segment, and an extension mechanism for extending the extendible second exit cone segment, the mechanism being located outside the first and second exit cone segments. A rigid thermal protection shield is interposed between the extension mechanism and the stationary first exit cone segment. The thermal protection shield presents a convex wall on its face facing towards the stationary first exit cone segment. 110-. (canceled)11. A rocket engine with an extendible exit cone , comprising:an exhaust nozzle for gas from a combustion chamber, the nozzle presenting a longitudinal axis and including a first portion defining a nozzle throat and a stationary first exit cone segment, at least one extendible second exit cone segment of section greater than the section of the stationary first exit cone segment, and an extension mechanism for extending the extendible second exit cone segment, the mechanism being located outside the first and second exit cone segments; anda rigid thermal protection shield interposed between the extension mechanism and the stationary first exit cone segment, andwherein the thermal protection shield presents a convex wall on its face facing towards the stationary first exit cone segment.12. A rocket engine according to claim 11 , wherein the thermal protection shield further comprises side fins on either side of the convex wall.13. A rocket engine according to claim 12 , further comprising a flexible thermal protection strip arranged between the rigid thermal protection shield and the extension mechanism.14. A rocket engine according to claim 13 , wherein the flexible thermal protection strip ...

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05-07-2012 дата публикации

CONTROLLING BLADE TIP CLEARANCES IN A TURBINE ENGINE

Номер: US20120167584A1
Автор: Philippot Vincent
Принадлежит: SNECMA

A turbine engine including a controller controlling clearance between tips of moving blades of a high-pressure turbine and an outer casing surrounding the blades, by cooling the outer casing by the impact of air taken from a high-pressure compressor stage of the engine, and by electric heating of top and bottom portions of the outer casing. 112-. (canceled)13. A turbine engine comprising:means for controlling clearance between tips of moving blades of a high-pressure turbine and an outer casing surrounding the blades, the control means comprising air-impact cooling means for cooling the outer casing by impact of air taken from a high-pressure compressor stage of the engine; andfirst electric heater means for heating a top portion of the outer casing and second electric heater means for heating a bottom portion of the outer casing, together with on/off control means for controlling the air-impact cooling means, and independent means for controlling the first and second electric heater means.14. A turbine engine according to claim 13 , wherein the air-impact cooling means comprises a ring carried by an outer casing and comprises projections that are axially spaced apart from one another with multiply-perforated strips installed therebetween for delivering the air taken from the high-pressure compressor.15. A turbine engine according to claim 13 , wherein the means for taking air from the high-pressure compressor comprises a valve for opening and closing delivery of air to the outer casing.16. A turbine engine according to claim 13 , wherein the cooling air is taken from a fourth stage of the high-pressure compressor at a rate that is about 0.7% of total air flow rate through the compressor.17. A turbine engine according to claim 13 , wherein the electric heater means comprises resistive circuits carried by the outer casing on the top and bottom portions thereof.18. A turbine engine according to claim 17 , wherein the resistive circuits are mounted in a vicinity of ...

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05-07-2012 дата публикации

CRYOGENIC TREATMENT OF MARTENSITIC STEEL WITH MIXED HARDENING

Номер: US20120168039A1
Автор: Ferrer Laurent
Принадлежит: SNECMA

The invention relates to a method for producing martensitic steel that comprises a content of other metals such that the steel can be hardened by an intermetallic compound and carbide precipitation, with an Al content of between 0.4% and 3%, comprising the following steps: 2. The method of claim 1 , wherein said the steel consists of:0.18 to 0.3% of C,5 to 7% of Co,2 to 5% of Cr,1 to 2% of Al,1 to 4% of Mo+W/2,traces to 0.3% of V,traces to 0.1% of Nb,traces to 50 ppm of B,10.5 to 15% of Ni with Ni≧7+3.5 Al,traces to 0.4% of Si,traces to 0.4% of Mn,traces to 500 ppm of Ca,traces to 500 ppm of at least one rare earth metal,traces to 500 ppm of Ti,traces to 50 ppm of 0 if developed from molten metal or to 200 ppm of O if developed through powder metallurgy,traces to 100 ppm of N,traces to 50 ppm of S,traces to 1% of Cu,traces to 200 ppm of P, anda remainder of Fe.3. The method of claim 2 ,wherein a content of C is from 0.200% to 0.250%,a content of Ni is from 12.00% to 14.00%,a content of Co is from 5.00% to 7.00%,a content of Cr is from 2.5% to 4.00%,a content of Al is from 1.30 to 1.70%,a content of Mo is from 1.00% to 2.00%.4. The method of claim 1 , wherein the time tis at least 1 hour.5. The method of claim 1 , wherein cooling the steel to approximately ambient temperature comprises cooling by quenching in a medium with a drasticity of at least a drasticity of air.6. The method of claim 1 , wherein cooling the steel in the cryogenic medium starts less than 70 hours after a surface temperature of the steel reaches 80° C.7. A piece made from a steel obtained by a method comprising the method of claim 1 , wherein a residual austenite level in the steel is less than 3%.8. A turbomachine transmission shaft made from a steel obtained by a process comprising the method of claim 1 , wherein a residual austenite level in said the steel is less than 3%.9. A steel obtained by a process comprising the method of claim 1 , wherein an average hardness of the steel is 575 Hv with ...

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12-07-2012 дата публикации

TOOLING AND A METHOD FOR HOT FORGING PIECES OF SHEET METAL

Номер: US20120174384A1
Принадлежит: SNECMA

The invention provides tooling and a method for hot forging pieces of sheet metal that are to form metal reinforcement mounted on the leading or trailing edge of a turbine engine blade, by means of bottom and top matrices each presenting a twisted elongate surface for use in shaping an initially plane piece of sheet metal, the shaping surface of the bottom matrix presenting a high portion, a low portion, and two end zones. The bottom matrix includes studs for positioning and guiding the piece, the studs being situated at the periphery of the corresponding shaping surface. 1. Tooling for hot forging pieces of sheet metal that are to form metal reinforcement mounted on the leading or trailing edge of a turbine engine blade , the tooling comprising bottom and top matrices each presenting a twisted elongate surface for use in shaping an initially plane piece of sheet metal , the shaping surface of the bottom matrix presenting a high portion , a low portion , and two end zones , wherein the bottom matrix includes at least three studs for positioning and guiding the piece , the studs being situated at the periphery of the corresponding shaping surface , the first and second studs being positioned respectively at each of the end zones of the shaping surface , a third stud being situated level with the low portion of the shaping surface.2. Tooling according to claim 1 , wherein three of the studs are all situated on the same side of a middle longitudinal axis of the shaping surface of the bottom matrix.3. Tooling according to claim 1 , wherein the height of the third stud is not less than the difference in height between the high portion and the low portion of the shaping surface.4. Tooling according to claim 1 , wherein at least one stud extends perpendicularly to the periphery of the shaping surface and presents a section that tapers towards the inside of the shaping surface claim 1 , so as to form a point or line bearing zone for the edge of the piece.5. Tooling ...

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12-07-2012 дата публикации

METHOD FOR MANUFACTURING A ONE-PIECE ANNULAR METAL PART HAVING A REINFORCING INSERT OF COMPOSITE MATERIAL

Номер: US20120175047A1
Автор:
Принадлежит: SNECMA

A method for manufacturing a one-piece annular metal part having a reinforcing insert of composite material. 2. The method as claimed in claim 1 , in which the metal wire and the composite fiber are cold wound at ambient temperature claim 1 , and in which the metal wire used is obtained by wire drawing and is made from the same metal as that of the coated fiber.3. The method as claimed in claim 1 , in which the metal wire and the composite fiber are wound substantially perpendicular to the axis of the rotating cylindrical mandrel.4. The method as claimed in claim 1 , in which the layers of composite fiber are arranged on a partial median area of the extent of the metal wire wound around the cylindrical mandrel claim 1 , and radially close thereto.5. The method as claimed in claim 1 , according to which the at least one layer of metal wires are held together by bonding means claim 1 , so as to obtain a blank (E) that can be handled until its introduction into said tool.6. The method as claimed in claim 1 , in which two transverse flanges are arranged on said cylindrical mandrel claim 1 , spaced in parallel to one another and between which the windings forming the layers of joined turns of metal wire are mounted.7. The method as claimed in claim 6 , in which recesses are made in the flanges claim 6 , for accommodating windings of metal wire claim 6 , and corresponding to changes in cross section of the annular part to be obtained.8. The method as claimed in claim 6 , in which pairs of annular flanges are provided claim 6 , being superimposed on the previously mounted flanges claim 6 , at each change in cross section of the part to be obtained in accordance with its production by the metal wire.9. The method as claimed in claim 6 , in which the cylindrical mandrel claim 6 , the annular flanges and the blank (E) having a metal wire and composite fiber claim 6 , are incorporated in the treatment tool. The present invention relates to a method for manufacturing a one- ...

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12-07-2012 дата публикации

TURBINE ENGINE WITH CONTRA-ROTATING NON-DUCTED PROPELLERS

Номер: US20120177493A1
Принадлежит: SNECMA

A turbine engine having two contrarotating unducted propellers, an upstream propeller () and a downstream propeller (), with a power turbine mounted axially between the two propellers (), the turbine comprising an outer rotor () constrained to rotate with the upstream propeller () and an inner rotor () driving rotation of an inlet shaft () of a step-down gearbox (), the gearbox having an outlet shaft () driving rotation of the rotor () of the upstream propeller () and an outlet shaft () driving rotation of the downstream propeller (). 112145052121450121456125462526456126614. A turbine engine having two contrarotating unducted propellers , an upstream propeller () and a downstream propeller () , the engine comprising a power turbine () connected via a step-down gearbox () to the rotors of the propellers ( , ) , the engine being characterized in that the power turbine () is mounted axially between the two propellers ( , ) and comprises an outer rotor () constrained to rotate with the upstream propeller () , and an inner rotor () driving rotation of an inlet shaft () of the gearbox () , the gearbox having an outlet shaft () driving rotation of the rotor () of the upstream propeller () and an outlet shaft () driving rotation of the downstream propeller ().25212. A turbine engine according to claim 1 , characterized in that the gearbox () is mounted axially upstream from the upstream propeller ().35060545856. A turbine engine according to or claim 1 , characterized in that the power turbine () comprises three stages in series claim 1 , each formed by an annular row of blades () carried by the inner rotor () and an annular row of blades () carried by the outer rotor ().4352687072748284767862. A turbine engine according to any one of to claims 1 , characterized in that the gearbox () comprises two epicyclic gear sets ( claims 1 , ) in which the sunwheels ( claims 1 , ) are constrained to rotate together and in which the planet carriers ( claims 1 , ) drive the upstream and ...

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