The present invention is concerned with a heat exchanging apparatus for a gas turbine engine. The heat exchanging apparatus comprises a first end, a second end, and a plurality of radially offset portions extending between the first and second ends, each portion comprising at least two radially offset rows of fluid passageways configured to transport fluid between the first end and the second end. Each portion is fixed at the first end and is axially moveable at the second end with respect to adjacent portions.
F28F 9/02 - Boîtes de distributionPlaques d'extrémité
F28D 7/16 - Appareils échangeurs de chaleur comportant des ensembles de canalisations tubulaires fixes pour les deux sources de potentiel calorifique, ces sources étant en contact chacune avec un côté de la paroi d'une canalisation les canalisations étant espacées parallèlement
F28F 9/013 - Supports auxiliaires pour les éléments pour les tubes ou les assemblages de tubes
A rear engine mounting structure for connecting an engine system to an aircraft structure, the rear engine mounting structure comprising: an elongate body having a first end arranged in use for connection to an aft side of a turbine exhaust case, the elongate body comprising one or more mounting points for connection to the aircraft structure, wherein the one or more mounting points are spaced from the first end in an axial direction.
B64D 27/40 - Aménagements pour le montage de groupes moteurs sur aéronefs
B64D 33/04 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des sorties d'échappement ou des tuyères
F02C 7/20 - Montage ou bâti de l'ensemble fonctionnelDisposition permettant la dilatation calorifique ou le déplacement
F02K 1/04 - Montage d'un cône d'échappement dans la tubulure de jet
3.
TURBINE ENGINE CASING FABRICATION AND HEAT TREATMENT
The invention concerns a method of heat treatment for a component comprising a first portion formed of a first nickel-based alloy. The method comprises heating the component to a temperature from 920°C to 1120°C for less than about 0.75 hours.
C22F 1/10 - Modification de la structure physique des métaux ou alliages non ferreux par traitement thermique ou par travail à chaud ou à froid du nickel ou du cobalt ou de leurs alliages
B22F 5/00 - Fabrication de pièces ou d'objets à partir de poudres métalliques caractérisée par la forme particulière du produit à réaliser
B22F 10/64 - Traitement de pièces ou d'articles après leur formation par des moyens thermiques
B23K 15/00 - Soudage ou découpage par faisceau d'électrons
F01D 25/24 - Carcasses d'enveloppeÉléments de la carcasse, p. ex. diaphragmes, fixations
A casing structure for a gas turbine engine comprising a plurality of struts defining a plurality of annular circumferentially extending airflow passages and where the strut side surfaces comprise recesses.
A tool damper for a machine tool, such as a milling machine, is arranged around the tool stem of the machine or the tool holder of the machine and arranged to damper vibrations of the tool.
The invention concerns a compressor case for a gas turbine engine, the case comprising a generally cylindrical body formed of a plurality of adjacent cylindrical sub-sections being formed of Ti-6242 and Ti-64.
The invention concerns a method of turbine blade repair in which a damaged region (Di) is machined away (Rz) and repaired with a directed energy deposition process before being machined to a desired and predetermined aerodynamic profile.
B23P 6/00 - Remise en état ou réparation des objets
F01D 5/00 - AubesOrganes de support des aubesDispositifs de chauffage, de protection contre l'échauffement, de refroidissement, ou dispositifs contre les vibrations, portés par les aubes ou les organes de support
B22F 10/25 - Dépôt direct de particules métalliques, p. ex. dépôt direct de métal [DMD] ou mise en forme par laser [LENS]
A gas turbine engine has a core flow path and an outer flow path, the outer flow path positioned at a greater radial displacement from a rotational axis of the engine than the core flow path. The gas turbine engine also has a cavity to provide fluid communication between a bleed passage and a duct, the bleed passage for communicating air from the core flow path to the cavity and the duct for communicating air from the cavity to the outer flow path. The cavity has a downstream wall. The duct has a circumferentially extending opening that provides a passage through the downstream wall adjacent to a radially outer wall of the cavity.
F02C 3/13 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur ayant des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre des étages de différents rotors
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
F02K 3/075 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c.-à-d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p. ex. du type à double flux commande du rapport des débits des différents flux
The invention concerns an aircraft fuel heating arrangement comprising an exhaust cone for a gas turbine engine, the exhaust cone having an outer body defining an internal cone cavity, the internal cone cavity comprising one or more conduits passing therethrough for communicating fuel through the cavity, and wherein the outer body of the exhaust cone comprises at least one inlet and at least one outlet to allow exhaust gas to pass into through and out of the exhaust cone.
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
F02C 7/10 - Chauffage de l'air d'alimentation avant la combustion, p. ex. par les gaz d'échappement au moyen d'échangeurs de récupération de chaleur
F02K 1/04 - Montage d'un cône d'échappement dans la tubulure de jet
F02C 6/18 - Utilisation de la chaleur perdue dans les ensembles fonctionnels de turbines à gaz à l'extérieur des ensembles eux-mêmes, p. ex. ensembles fonctionnels de chauffage à turbine à gaz
F02C 3/22 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail utilisant un combustible, un oxydant ou un fluide de dilution particulier pour produire les produits de combustion le combustible ou l'oxydant étant gazeux aux température et pression normales
An outlet guide vane (OGV) structure for a gas turbine engine can comprise a plurality of radially extending guide vanes having inlets and outlets to allow cooling of a medium within the guide vanes.
A gas turbine engine comprises at least one radially extending bleed passage optionally in fluid communication with at least one generally circumferentially extending plenum. The passage has an upstream inlet in fluid communication with a bleed passage and an outlet for releasing air from the plenum. The upstream leading edge of the inlet or the downstream trailing edge of the inlet has a non-uniform profile.
F02C 6/08 - Ensembles fonctionnels de turbines à gaz délivrant un fluide de travail chauffé ou pressurisé à d'autres appareils, p. ex. sans sortie de puissance mécanique délivrant des gaz comprimés le gaz étant prélevés sur le compresseur de la turbine à gaz
09 - Appareils et instruments scientifiques et électriques
Produits et services
Laser welding machines; laser welding devices in the nature of laser welding machines; laser welding devices in the nature of
component parts for laser welding machines, namely, laser optics, vision systems, seam tracking systems, fixtures, robot dress
package, optical sensors, focus positioning system, control system for regulating process parameters, station human machine
interface, media panel for regulating, distributing, and measuring flows, temperature, and pressures, laser safety enclosures Laser equipment for non-medical purposes
09 - Appareils et instruments scientifiques et électriques
Produits et services
Laser welding machines; laser welding devices in the nature of laser welding machines; laser welding devices in the nature of
component parts for laser welding machines, namely, laser optics, vision systems, seam tracking systems, fixtures, robot dress
package, optical sensors, focus positioning system, control system for regulating process parameters, station human machine
interface, media panel for regulating, distributing, and measuring flows, temperature, and pressures, laser safety enclosures Laser equipment for non-medical purposes
The disclosure concerns a compressor blade for gas turbine engine. Specifically the blades of the compressor are modified according to predetermined requirements for both aerodynamic stability and fuel economy in multiple planes.
The invention concerns a tool damper for a machine tool, such as a milling machine, wherein the damper is arranged around the tool stem of the machine or the tool holder of the machine and arranged in use to damper vibrations of the tool.
A front centre body (FCB) structure for a geared turbofan engine comprises a plurality of vanes extending across the inlet duct to a low pressure compressor and integrates a heat exchanging arrangement to control the temperature of the gearbox of the turbofan engine.
A method of bonding a composite vane to at least one support. The method includes positioning an end of a vane within a support, causing an adhesive to flow between the vane and the support from an inlet to an outlet, and reversing the flow of adhesive from the outlet towards the inlet.
A milling machine includes a spindle which is arranged to receive a tool holder and in use to cause rotation of the tool holder within the spindle. A portion of the spindle housing surrounding the tool holder is provided with at least one pair of opposing damping units.
B23Q 1/70 - Éléments fixes ou mobiles supportant des broches de travail pour fixer des outils ou des pièces
B23Q 11/00 - Accessoires montés sur les machines-outils pour maintenir les outils ou les organes de la machine dans de bonnes conditions de travail ou pour refroidir les pièces travailléesDispositifs de sécurité spécialement combinés aux machines-outils, disposés dans ces machines ou spécialement conçus pour être utilisés en relation avec ces machines
F16F 7/108 - Amortisseurs de vibrationsAmortisseurs de chocs utilisant un effet d'inertie l'élément d'inertie étant monté de manière élastique sur des ressorts en matière plastique
F16F 15/08 - Suppression des vibrations dans les systèmes non rotatifs, p. ex. dans des systèmes alternatifsSuppression des vibrations dans les systèmes rotatifs par l'utilisation d'organes ne se déplaçant pas avec le système rotatif utilisant des moyens élastiques avec ressorts en caoutchouc
An airflow arrangement for a gas turbine engine in which air can be selectively directed to and from a bypass channel of the engine and into a duct communicating air into a compressor of a gas turbine engine.
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
F02C 6/08 - Ensembles fonctionnels de turbines à gaz délivrant un fluide de travail chauffé ou pressurisé à d'autres appareils, p. ex. sans sortie de puissance mécanique délivrant des gaz comprimés le gaz étant prélevés sur le compresseur de la turbine à gaz
F02K 3/075 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c.-à-d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p. ex. du type à double flux commande du rapport des débits des différents flux
F04D 29/54 - Moyens de guidage du fluide, p. ex. diffuseurs
An apparatus and corresponding method for machining at least one spline in an aeronautical component. The apparatus comprises a holder defining an axial direction about which the holder is moveable, and a cutting insert attachable to the holder. The apparatus further comprises a coupling for attaching the cutting insert to the holder about a second direction that is perpendicular to the axial direction of the holder. The holder may comprise a first alignment feature for orientating the cutting insert at a single orientation with respect to the holder. The cutting insert may comprise at least one cutting tooth, wherein each cutting tooth comprises a cutting edge. The cutting insert may comprise a first datum surface which is spaced from each cutting edge by a respective predetermined spacing.
B23B 27/10 - Outils de coupe avec une disposition particulière pour le refroidissement
B23B 27/16 - Outils de coupe sur lesquels les taillants ou éléments tranchants sont en matériaux particulier à éléments tranchants interchangeables, p. ex. pouvant être fixés par des brides
B23C 5/22 - Dispositifs pour la fixation des taillants ou des dents
B23D 3/02 - Machines à raboter ou à mortaiser taillant par déplacement relatif de l'outil et de la pièce à usiner selon une direction verticale ou oblique pour tailler des rainures
B23B 29/04 - Porte-outils pour un seul outil de coupe
B23C 5/28 - Caractéristiques se rapportant à la lubrification ou au refroidissement
An aircraft engine circulation system comprises a conduit arranged in use to communicate a lubricant to and from one or more bearings of an engine. The conduits define a space which comprises a material that is selected to change phase at a predetermined temperature.
A gas turbine engine comprising at least one radially extending bleed passage optionally in fluid communication with at least one generally circumferentially extending plenum. The passage has an upstream inlet in fluid communication with a bleed passage and an outlet for releasing air from the plenum. The upstream leading edge of the inlet or the downstream trailing edge of the inlet has a non-uniform profile.
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
The invention concerns a compressor blade for gas turbine engine. Specifically the blades of the compressor are modified according to predetermined requirements for both aerodynamic stability and fuel economy in multiple planes.
An outlet guide vane (OGV) structure for a gas turbine engine comprises a plurality of radially extending guide vanes (30), at least one guide vane comprises at least one inlet (35) allowing air to pass into at least one internal cavity (37) within the guide vane and at least one outlet (36) allowing air to exit the at least one internal cavity, wherein the inlet is arranged upstream of an outlet in a direction, in use, of airflow over the guide vane.
A turbine rear structure for a gas turbine engine includes a central hub and a circumferential outer ring coaxial with the central hub. The turbine rear structure further includes a plurality of guide vanes extending radially between the central hub and the circumferential outer ring, and an intermediate guide vane located in a space defined between adjacent guide vanes. The intermediate guide vane is located closer to one of the guide vanes than the other guide vane.
F01D 17/14 - Organes de commande terminaux disposés sur des parties du stator faisant varier l'aire effective de la section transversale des injecteurs ou tuyères de guidage
F01D 9/04 - InjecteursLogement des injecteursAubes de statorTuyères de guidage formant une couronne ou un secteur
The present invention relates to a front centre body (FCB) structure for a geared turbofan engine. The FCB comprises a plurality of vanes extending across the inlet duct to the low pressure compressor and integrates a heat exchanging arrangement to control the temperature of the gearbox of the turbofan engine.
A method of bonding a composite vane (1) to at least one support (2). The method comprises the steps of positioning an end of a vane (1) within a support (2), causing an adhesive to flow between the vane and the support from an inlet (3) to an outlet (4) and reversing the flow of adhesive from the outlet (4) towards the inlet (3).
A multi-axis milling machine comprising a spindle (3), the spindle arranged to receive a tool holder (3) and in use to cause rotation of the tool holder within the spindle, wherein a portion of the spindle housing (11) surrounding the tool holder (3) is provided with one or more damping units (12A, 12B).
B23Q 11/00 - Accessoires montés sur les machines-outils pour maintenir les outils ou les organes de la machine dans de bonnes conditions de travail ou pour refroidir les pièces travailléesDispositifs de sécurité spécialement combinés aux machines-outils, disposés dans ces machines ou spécialement conçus pour être utilisés en relation avec ces machines
B23Q 1/70 - Éléments fixes ou mobiles supportant des broches de travail pour fixer des outils ou des pièces
31.
AN APPARATUS, INSERT AND METHOD FOR MACHINING AN AERONAUTICAL COMPONENT
An apparatus and corresponding method for machining at least one spline 205 in an aeronautical component 200. The apparatus comprises a holder 15 defining an axial direction about which the holder 15 is moveable, and a cutting insert 20 attachable to the holder 15. The apparatus further comprises a coupling for attaching the cutting insert 20 to the holder 15 about a second direction that is perpendicular to the axial direction of the holder. The cutting insert 20 comprises at least one cutting tooth 55, wherein each cutting tooth 55 comprises a cutting edge 60. The cutting insert 20 comprises a first datum surface 64 which is spaced from each cutting edge 60 by a respective predetermined spacing SI.
B23D 3/02 - Machines à raboter ou à mortaiser taillant par déplacement relatif de l'outil et de la pièce à usiner selon une direction verticale ou oblique pour tailler des rainures
B23F 5/12 - Fabrication de dents d'engrenage droites, impliquant le déplacement d'un outil par rapport à la pièce à usiner avec un mouvement de roulage ou d'enveloppement par rapport aux dents à réaliser par rabotage ou mortaisage
B23Q 17/22 - Agencements sur les machines-outils pour indiquer ou mesurer pour indiquer ou mesurer la position réelle ou désirée de l'outil ou de la pièce
B23Q 17/09 - Agencements sur les machines-outils pour indiquer ou mesurer pour indiquer ou mesurer la pression de coupe ou l'état de l'outil de coupe, p. ex. aptitude à la coupe, charge sur l'outil
The invention concerns an airflow arrangement for a gas turbine engine in which air can be selectively directed to and from a bypass channel of the engine and into a duct communicating air into a compressor of a gas turbine engine.
F02C 6/08 - Ensembles fonctionnels de turbines à gaz délivrant un fluide de travail chauffé ou pressurisé à d'autres appareils, p. ex. sans sortie de puissance mécanique délivrant des gaz comprimés le gaz étant prélevés sur le compresseur de la turbine à gaz
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
F02K 3/075 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c.-à-d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p. ex. du type à double flux commande du rapport des débits des différents flux
F02C 3/13 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur ayant des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre des étages de différents rotors
A turbine exhaust casing (TEC) cooling arrangement for a gas turbine engine includes cooling the struts of a TEC using compressed air communicated from one of the engine's compressors.
The invention concerns a turbine exhaust casing (TEC) for a gas turbine engine in which portions of the inner surface of the casing against which exhaust gas flows are provided with recesses extending into the surfaces. The recesses are positioned proximate to the leading edges of struts which extend between an outer shroud and inner hub of the casing.
An aircraft engine circulation system comprising a multi-layered pipe or tube (11) arranged in use to communicate a lubricant to and from one or more bearings of an engine comprising a central pipe (12) for communicating the oil and a peripheral pipe (13) surrounding the pipe. The multi-layered pipe or tube defines a space (14) which comprises a material (15) that is selected to change phase at a predetermined temperature.
A gas turbine engine comprising at least one radially extending bleed passage in fluid communication with at least one generally circumferentially extending plenum. A plenum has an upstream end in fluid communication with a bleed passage and an outlet for releasing air from the plenum. A plenum further comprises a downstream surface defining a downstream closed end of the plenum and the downstream surface of one or more plenum is/are provided with an outwardly extending projection extending into the plenum.
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
A cutting tool comprises at least one cutting edge and at least one relief surface adjacent thereto. The relief surface is arranged to provide clearance behind the cutting edge between the cutting tool and a workpiece and comprises one or more grooves arranged to communicate, in use, a cutting media across the relief surface.
Heat treatment of an Alloy 282 which has been subjected to an initial solution annealing followed by cooling can be heat treated by heating the Alloy 282 to a temperature between 954° C. and 1010° C. until the gamma prime (γ′) phase is sufficiently dissolved, and cooling the Alloy 282 to a temperature a sufficiently low temperature, and at a sufficiently high cooling rate, to suppress gamma prime precipitation. A component such as a turbine exhaust case and a gas turbine engine made of said alloy can be heat treated in the above manner.
C22F 1/10 - Modification de la structure physique des métaux ou alliages non ferreux par traitement thermique ou par travail à chaud ou à froid du nickel ou du cobalt ou de leurs alliages
The invention concerns a transition duct for a multi-stage compressor of a gas turbine engine. Regions of the inner surface of the duct are provided with a predetermined and dissimilar surface roughness to optimise gas flow efficiency within the duct.
A wire dispenser for a laser metal wire deposition machine comprises a longitudinal duct for guiding a wire from a proximal end to a distal end of the duct. A nozzle unit is connected to the distal end of the duct and has a through bore for receiving the wire from the distal end of the duct and for discharging the wire adjacent to a laser metal wire deposition site. The nozzle unit includes a cooling circuit for a cooling liquid.
B23K 26/14 - Travail par rayon laser, p. ex. soudage, découpage ou perçage en utilisant un écoulement de fluide, p. ex. un jet de gaz, associé au faisceau laserBuses à cet effet
B23K 26/34 - Soudage au laser pour des finalités autres que l’assemblage
41.
TURBINE REAR STRUCTURES, CORRESPONDING GAS TURBINE ENGINE, AIRCRAFT AND METHOD OF MANUFACTURING
A turbine rear structure for a gas turbine engine which comprises primary and secondary exhaust gas guide vanes and wherein the second set of vanes are offset so as to be closer to one of the primary guide vanes than another.
The invention concerns a turbine exhaust casing (TEC) for a gas turbine engine in which portions of the inner surface of the casing against which exhaust gas flows are provided with recesses extending into the surfaces. The recesses are positioned proximate to the leading edges of struts which extend between an outer shroud and inner hub of the casing.
F01D 25/16 - Aménagement des paliersSupport ou montage des paliers dans les stators
F02C 3/04 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur
The invention concerns a turbine exhaust casing (TEC) cooling arrangement for a gas turbine engine. The arrangement involves cooling the struts of a TEC using compressed air communicated from one of the engine's compressors.
The present invention is concerned with a cutting tool comprising at least one cutting edge and at least one relief surface (10) adjacent thereto. The relief surface (10) is arranged to provide clearance behind the cutting edge between the cutting tool and a workpiece (20) and comprises one or more grooves (17) arranged to communicate, in use, a cutting media across the relief surface (10).
The present disclosure relates to a method for heat treatment of an Alloy 282 which has been subjected to an initial solution annealing followed by cooling, the method comprising the steps of heating the Alloy 282 to a temperature between 954°C and 1010°C until the gamma prime (γ') phase is sufficiently disclosed, and cooling the Alloy 282 to a temperature a temperature sufficiently low, and at a cooling rate sufficiently high, to suppress gamma prime precipitation. Furthermore said alloy being heat treated according to the method and a component, a turbine exhaust case and a gas turbine engine made of said alloy are disclosed.
C21D 9/50 - Traitement thermique, p. ex. recuit, durcissement, trempe ou revenu, adapté à des objets particuliersFours à cet effet pour joints de soudure
C22C 19/05 - Alliages à base de nickel ou de cobalt, seuls ou ensemble à base de nickel avec du chrome
C22F 1/10 - Modification de la structure physique des métaux ou alliages non ferreux par traitement thermique ou par travail à chaud ou à froid du nickel ou du cobalt ou de leurs alliages
46.
WIRE DISPENSER, FOR A LASER METAL WIRE DEPOSITION MACHINE, WITH A COOLING CIRCUIT; CORRESPONDING LASER METAL WIRE DEPOSITION MACHINE; METHOD OF PERFORMING LASER METAL WIRE DEPOSITION ON A WORKPIECE WITH SUCH WIRE DISPENSER
A wire dispenser (9) for a laser metal wire deposition machine comprises a longitudinal duct (8) for guiding a wire (4A, 4B) from a proximal end (81) to a distal end (82) of the duct (8). A nozzle unit (6) is connected to the distal end (82) of the duct (8) and has a through bore (51) for receiving the wire (4B) from the distal end (82) of the duct (8) and for discharging the wire (4B) adjacent to a laser metal wire deposition site (16). The nozzle unit (6) includes a cooling circuit (771) for a cooling liquid.
B23K 26/08 - Dispositifs comportant un mouvement relatif entre le faisceau laser et la pièce
B23K 26/14 - Travail par rayon laser, p. ex. soudage, découpage ou perçage en utilisant un écoulement de fluide, p. ex. un jet de gaz, associé au faisceau laserBuses à cet effet
B23K 26/34 - Soudage au laser pour des finalités autres que l’assemblage
The invention concerns a transition duct for a multi-stage compressor of a gas turbine engine. Regions of the inner surface of the duct are provided with a predetermined and dissimilar surface roughness to optimise gas flow efficiency within the duct.
F04D 29/54 - Moyens de guidage du fluide, p. ex. diffuseurs
F04D 29/66 - Lutte contre la cavitation, les tourbillons, le bruit, les vibrations ou phénomènes analoguesÉquilibrage
F04D 29/68 - Lutte contre la cavitation, les tourbillons, le bruit, les vibrations ou phénomènes analoguesÉquilibrage en agissant sur les couches limites
A method for joining two structural elements by welding, in particular by butt welding comprises forming a weld line joining the two structural elements; and adding material across the weld line, thereby forming one or more crack stoppers for limiting crack propagation along the weld line. The one or more crack stoppers each have a limited extension along the weld line as seen in relation to a length of the weld line. A structural system comprising two structural elements joined by the method is disclosed. The method may be applied, e.g., to components of aircraft engines.
B23K 31/02 - Procédés relevant de la présente sous-classe, spécialement adaptés à des objets ou des buts particuliers, mais non couverts par un seul des groupes principaux relatifs au brasage ou au soudage
F01D 25/28 - Dispositions pour le support ou le montage, p. ex. pour les carters de turbines
B23K 31/12 - Procédés relevant de la présente sous-classe, spécialement adaptés à des objets ou des buts particuliers, mais non couverts par un seul des groupes principaux relatifs à la recherche des propriétés, p. ex. de soudabilité, des matériaux
A method for joining two structural elements (26, 28) by welding, in particular by butt welding, is disclosed. The method comprises the following steps: - forming a weld line (30) joining the two structural elements; - adding material (32) across the weld line (30), thereby forming one or more crack stoppers (34) for limiting crack propagation along the weld line. The one or more crack stoppers (34) each having a limited extension along the weld line (30) as seen in relation to a length of the weld line. A structural system comprising two structural elements joined by the method is disclosed. Especially, the method may be applied to components of aircraft engines.
B23K 31/02 - Procédés relevant de la présente sous-classe, spécialement adaptés à des objets ou des buts particuliers, mais non couverts par un seul des groupes principaux relatifs au brasage ou au soudage
B23K 31/12 - Procédés relevant de la présente sous-classe, spécialement adaptés à des objets ou des buts particuliers, mais non couverts par un seul des groupes principaux relatifs à la recherche des propriétés, p. ex. de soudabilité, des matériaux
B23K 9/04 - Soudage pour d'autres buts que l'assemblage de pièces, p. ex. soudage de rechargement
B23P 6/04 - Réparation de pièces ou de produits métalliques brisés ou fissurés, p. ex. de pièces de fonderie
B23P 6/00 - Remise en état ou réparation des objets
The invention concerns a gas turbine engine component (37) comprising an outer ring structure (10), an inner ring structure (20) and a plurality of circumferentially spaced radial elements (15) for transferring loads between the inner ring structure (20) and the outer ring structure (10), wherein the outer ring structure (10) comprises a circular ring member (11) for withstanding an internal pressure during operation of the gas turbine engine. The invention is characterized in that the outer ring structure (10) comprises a first set of circumferentially distributed and substantially straight reinforcement ribs (16) for achieving radial stiffness, wherein the reinforcement ribs (16) form sides in a polygonal shape and wherein each of said reinforcement ribs (16) extends between two adjacent radial elements (15). The invention also concerns a gas turbine engine (1) comprising a component (37) of the above type.
A reliability range is determined for a parameter of a component in a machine subjected to life reducing loads during operation, comprising the steps of: acquiring, for each of a plurality of load sessions, at least one parameter value for said component; generating a distribution pattern containing said parameter values for the plurality of load sessions; assigning a reliability range for the distribution pattern, wherein parameter values outside said reliability range are considered as being unrealistic; analyzing the parameter values outside the reliability range to determine which of said parameter values outside the reliability range are confirmed to be unrealistic; and adjusting the reliability range if a ratio between the confirmed unrealistic parameter values and the considered unrealistic parameter values is outside a further range being predetermined for the ratio.
A set of load data for a selected point in time and resulting from the machine operation is received. The load data is provided from a first database comprising predefined machine conditions associated to different sets of load data for the machine. One of the predefined machine conditions that is most representative of the received set of load data is selected.
G01M 1/38 - Machines ou dispositifs combinés pour déterminer et corriger à la fois le balourd
G05B 13/02 - Systèmes de commande adaptatifs, c.-à-d. systèmes se réglant eux-mêmes automatiquement pour obtenir un rendement optimal suivant un critère prédéterminé électriques
G05B 15/02 - Systèmes commandés par un calculateur électriques
G06F 17/30 - Recherche documentaire; Structures de bases de données à cet effet
A first set of data is received relating to the operation of a first mechanical part. A plurality of steady state conditions for the first set of operational data is determined. A load history is determined for the first mechanical part based on the plurality of determined steady states and the first set of operational data. One of a plurality of predefined life consumption calculation models is selected based on a type of the first mechanical part and a position of a critical area at the first mechanical part. A level of life consumption for the critical area of the first mechanical part is determined based on the selected life consumption calculation model and the determined load history.
G07C 3/00 - Enregistrement ou indication de l'état ou du fonctionnement de machines ou d'autres appareils à l'exclusion des véhicules
G07C 5/08 - Enregistrement ou indication de données de marche autres que le temps de circulation, de fonctionnement, d'arrêt ou d'attente, avec ou sans enregistrement des temps de circulation, de fonctionnement, d'arrêt ou d'attente
B64D 45/00 - Indicateurs ou dispositifs de protection d'aéronefs, non prévus ailleurs
A life consumption of a component in a machine may be predicted. Load data may be received from a load session of the machine. A plurality of parameter sets may be accessed, each associated with a critical point of the component, which point is considered to have critical life consumption. For each critical point, life consumption may be calculated using a life consumption calculation model receiving the load data and the parameter sets as input. By selecting a plurality of critical points on the component, a more complete view is presented of how the different parts of the component are affected by the load session.
G07C 5/08 - Enregistrement ou indication de données de marche autres que le temps de circulation, de fonctionnement, d'arrêt ou d'attente, avec ou sans enregistrement des temps de circulation, de fonctionnement, d'arrêt ou d'attente
B64F 5/60 - Test ou inspection des composants ou des systèmes d'aéronefs
F01K 23/00 - Ensembles fonctionnels caractérisés par plus d'une machine motrice fournissant de l'énergie à l'extérieur de l'ensemble, ces machines motrices étant entraînées par des fluides différents
A supporting structure for a gas turbine engine includes an inner ring, an outer ring, and a plurality of circumferentially spaced, load carrying radial elements connecting the inner and outer rings. The radial elements have an airfoil shape with a leading edge directed towards the inlet side of the supporting structure, a trailing edge directed towards the outlet side of the supporting structure, and two opposite sides connecting the leading edge and the trailing edge. At least one of the radial elements includes a gas passage arrangement configured to lead a separate bleeding gas flow from the supporting structure. The gas passage arrangement includes a radially extending gas channel arranged inside the radial element and at least one opening in communication with the gas channel. The at least one opening is arranged at one of the two opposite sides of the radial element.
F02C 7/20 - Montage ou bâti de l'ensemble fonctionnelDisposition permettant la dilatation calorifique ou le déplacement
F01D 9/06 - Conduits d'admission du fluide à l'injecteur ou à l'organe analogue
F01D 25/16 - Aménagement des paliersSupport ou montage des paliers dans les stators
F02C 6/08 - Ensembles fonctionnels de turbines à gaz délivrant un fluide de travail chauffé ou pressurisé à d'autres appareils, p. ex. sans sortie de puissance mécanique délivrant des gaz comprimés le gaz étant prélevés sur le compresseur de la turbine à gaz
F04D 29/54 - Moyens de guidage du fluide, p. ex. diffuseurs
A supporting structure for a gas turbine engine comprises an inner ring, an outer ring, and a plurality of circumferentially spaced, load carrying radial elements connecting the inner and outer rings, said radial elements being configured to transfer loads between the inner ring and the outer ring, wherein a gas channel for a primary axial gas flow is defined between the inner and outer rings, wherein the supporting structure has an inlet side for primary gas flow entrance and an outlet side for primary gas outflow, wherein the radial elements have an airfoil shape with a leading edge directed towards the inlet side, a trailing edge directed towards the outlet side, and two opposite sides connecting the leading edge and the trailing edge, and wherein at least a first of said radial elements is connected to an adjacent part of the supporting structure via a weld joint that extends across the leading edge and circumferentially at least partly around the first radial element.
F01D 9/04 - InjecteursLogement des injecteursAubes de statorTuyères de guidage formant une couronne ou un secteur
F01D 25/24 - Carcasses d'enveloppeÉléments de la carcasse, p. ex. diaphragmes, fixations
F01D 25/28 - Dispositions pour le support ou le montage, p. ex. pour les carters de turbines
B23K 31/02 - Procédés relevant de la présente sous-classe, spécialement adaptés à des objets ou des buts particuliers, mais non couverts par un seul des groupes principaux relatifs au brasage ou au soudage
b) and that is intended to form part of one of the ring structures. The invention also concerns a gas turbine engine (1) comprising a component (37) manufactured according to the above method.
In a method for training a person while operating a vehicle, the vehicle has a control system for receiving vehicle operating commands from the person for controlling the vehicle. A calculation unit is provided for simulating a state of the vehicle and/or the environment to which the vehicle is subjected, the simulated state being a possible real state of the vehicle and/or the environment which is different from the actual state of the vehicle and/or the environment. The vehicle operating commands and the calculation unit are used for calculating vehicle command signals. The vehicle command signals are used for controlling the vehicle so as to cause the vehicle to respond to the vehicle operating commands in a way that corresponds to the state simulated by the calculation unit instead of the actual state of the vehicle and/or the environment.
G09B 9/02 - Simulateurs pour l'enseignement ou l'entraînement pour l'enseignement de la conduite des véhicules ou autres moyens de transport
G09B 9/44 - Simulateurs pour l'enseignement ou l'entraînement pour l'enseignement de la conduite des véhicules ou autres moyens de transport pour l'enseignement du pilotage des aéronefs, p. ex. bancs d'entraînement au pilotage sans visibilité assurant la simulation dans un aéronef réel qui vole à travers l'atmosphère sans limitation de sa trajectoire
G09B 19/16 - Conduite des véhicules ou autres moyens de transport
G06G 7/48 - Calculateurs analogiques pour des procédés, des systèmes ou des dispositifs spécifiques, p. ex. simulateurs
G09B 9/00 - Simulateurs pour l'enseignement ou l'entraînement
The present invention relates to a support structure (16) for a gas turbine engine (1). The support structure (16) has an axial extension in an axial direction (A) and a circumferential extension in a circumferential direction (C). Moreover, the support structure (16) comprises a plurality of tubular members (18, 20) of a first material type arranged in sequence in the circumferential direction (C). Each tubular member (18, 20) at least partially delimits a flow guiding passage extending at least partially in the axial direction (A). The support structure (16) comprises a leading portion (22) and a trailing portion (24) in the axial direction (A). Furthermore, the support structure (16) comprises a leading edge member (26) of a second material type, the leading edge member (26) being located at the leading portion (22). At least two of the tubular members (18, 20) are fixedly attached to a leading edge member (26). According to the present invention, the first material type is different from the second material type.
B23K 31/02 - Procédés relevant de la présente sous-classe, spécialement adaptés à des objets ou des buts particuliers, mais non couverts par un seul des groupes principaux relatifs au brasage ou au soudage
F01D 25/28 - Dispositions pour le support ou le montage, p. ex. pour les carters de turbines
F01D 9/04 - InjecteursLogement des injecteursAubes de statorTuyères de guidage formant une couronne ou un secteur
B23K 101/00 - Objets fabriqués par brasage, soudage ou découpage
a) axially facing in a direction towards the second position (31), wherein the load carrying members (32) are arranged to form a load carrying connection between the annular load transfer structure (23) and said elements (22) via the inner ring (20). The invention also concerns a gas turbine engine (1) comprising a component (27) of the above type.
A composite guide vane for a gas turbine structure is adapted to extend in a guide vane direction from a first housing towards a second housing of the gas turbine structure. The guide vane includes a guide vane length in the guide vane direction and the guide vane includes a first attachment portion with at least one first opening for attachment of the guide vane to the first housing. The first opening extends in a first opening direction which forms an angle with the guide vane direction. The guide vane includes a cover portion including a first material and a core portion which is at least partially enclosed by the cover portion.
A gas turbine structure includes a first housing and a second housing, one of the first and second housings being located around the other of the first and second housings such that a core flow passage is obtained between the first and second housings. The gas turbine structure further includes an elongate structural member extending in a structural member direction from the first housing to the second housing and the gas turbine structure further includes a fairing circumferentially enclosing at least a portion of the structural member.
A gas turbine structure includes a guide vane. The gas turbine structure (further includes a first housing and a second housing and the guide vane extends from the first housing to the second housing. The guide vane includes a leading edge and a trailing edge and the guide vane extends from the leading edge to the trailing edge along a mean camber line. The guide vane includes a first attachment structure, a first fastening arrangement and a second fastening arrangement. Moreover, the first attachment structure includes a stiffening member extending over at least a portion of a circumference of the first housing. Furthermore, the stiffening member includes a stiffening member center point as measured along the mean camber line.
The present invention relates to a method for determining a reliability range for a parameter of a component in a machine subjected to life reducing loads during operation, comprising the steps of: acquiring, for each of a plurality of load sessions, at least one parameter value for said component; generating a distribution pattern containing said parameter values for the plurality of load sessions; assigning a reliability range for the distribution pattern, wherein parameter values outside said reliability range are considered as being unrealistic; analyzing the parameter values outside the reliability range to determine which of said parameter values outside the reliability range are confirmed to be unrealistic; and adjusting the reliability range if a ratio between the confirmed unrealistic parameter values and the considered unrealistic parameter values is outside a further range being predetermined for the ratio.
The invention concerns a supporting structure (27) for a gas turbine engine (1), said structure (27) comprising an inner ring (20), an outer ring (21), and a plurality of circumferentially spaced, load carrying radial elements (22) connecting the inner and outer rings (20, 21), said radial elements (22) being configured to transfer loads between the inner ring (20) and the outer ring (21), wherein a gas channel for a primary axial gas flow is defined between the inner and outer rings (20, 21), wherein the supporting structure (27) has an inlet side for primary gas flow entrance and an outlet side for primary gas outflow, wherein the radial elements (22) have an airfoil shape with a leading edge (221) directed towards the inlet side, a trailing edge (222) directed towards the outlet side, and two opposite sides (223, 224) connecting the leading edge (221) and the trailing edge (222), and wherein at least a first of said radial elements (22a, 22b) is connected to an adjacent part (30) of the supporting structure (27) via a weld joint (50) that extends across the leading edge (221) and circumferentially at least partly around the first radial element (22a, 22b). The invention is characterized in that the first radial element (22a, 22b) is provided with a region of reduced stiffness (10) in the vicinity of the leading edge (221). The invention also concerns a gas turbine engine (1) comprising a supporting structure (27) of the above type. The invention further comprises a method of manufacturing a supporting structure (27) of the above type.
B23P 15/04 - Fabrication d'objets déterminés par des opérations non couvertes par une seule autre sous-classe ou un groupe de la présente sous-classe d'aubes de turbine ou d'organes équivalents, en plusieurs pièces
A method and system for predicting a life consumption of a component in a machine. The method comprises receiving load data from a load session of said machine, accessing a plurality of parameter sets, each associated with a critical point of said component, which point is considered to have critical life consumption, and for each critical point, calculating life consumption using a life consumption calculation model receiving said load data and said parameter sets as input. By selecting a plurality of critical points on the component, a more complete view is presented of how the different parts of the component are affected by the load session.
There is provided a method for determining a machine condition indicative of life consumption of a machine component subjected to loads during machine operation, comprising the steps of: receiving a set of load data, for a selected point in time, resulting from the machine operation; from a first database comprising predefined machine conditions associated to different sets of load data for the machine, selecting one of the predefined machine conditions, which is most representative of the received set of load data. There is also provided a computer program product comprising code for performing the method and a system configured to perform the method.
The present invention relates to a method for generating a simplified calculation model for use in predicting life consumption of a component subjected to loads during operation, comprising the steps of: receiving a first set of load input data resulting from a first set of load sessions during operation; calculating at least one of stresses, strains and temperatures for a critical area of said component by means of a numerical calculation model; predicting life consumption of said component based on said at least one of the numerically calculated stresses, strains and temperatures; and generating said simplified calculation model defining a relationship between load input data and predicted life consumption by means of: assigning a plurality of linear difference equations for said simplified calculation model; and calculating parameters of said plurality of linear difference equations based on said relationship between the first set of load input data and said numerically calculated predicted life consumption. The present invention also relates to a corresponding systems and computer program products thereof. Furthermore, the invention also relates to a method for predicting life consumption of a component and a corresponding system thereof.
The present invention relates to a method for determining a level of life consumption of a critical area of a first mechanical part, comprising the steps of receiving a first set of data relating to the operation of the first mechanical part, determining a plurality of steady state conditions for the first set of operational data, determining a load history for the first mechanical part based on the plurality of determined steady states and the first set of operational data, selecting one of a plurality of predefined life consumption calculation models based on the type of the first mechanical part and the position of the critical area at the first mechanical part, and determining the level of life consumption for the critical area of the first mechanical part based on the selected life consumption calculation model and the determined load history. The present invention also relates to a corresponding system and computer program product.
A turbine exhaust case has an outer housing to be secured within a gas turbine engine and a central hub. Struts extend between the outer housing and the central hub. The struts are formed at least in part of a first material. The central hub is formed at least in part of a second material.
F02C 7/00 - Caractéristiques, parties constitutives, détails ou accessoires non couverts dans, ou d'un intérêt plus général que, les groupes Entrées d'air pour ensembles fonctionnels de propulsion par réaction
F02C 7/32 - Aménagement, montage ou entraînement des auxiliaires
F02K 3/00 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant
F02K 99/00 - Matière non prévue dans les autres groupes de la présente sous-classe
72.
METHOD FOR APPLYING A TITANIUM ALLOY ON A SUBSTRATE
Method for applying a titanium alloy on a substrate (18) which method comprises the step of melting or depositing said titanium alloy on said substrate (18) and solidifying said deposited or molten titanium alloy. The method comprises the step of adding 0.01-0.4 weight % Boron to said titanium alloy before or during said step of melting, welding or depositing said titanium alloy on said substrate (18).
C23C 24/10 - Revêtement à partir de poudres inorganiques en utilisant la chaleur ou une pression et la chaleur avec formation d'une phase liquide intermédiaire dans la couche
The invention concerns a supporting structure (37, 37') for a gas turbine engine (1, 100), said structure (37, 37') comprising an inner ring (10), an outer ring (11), and a plurality of circumferentially spaced, load carrying radial elements (12) connecting the inner and outer rings (10, 11), said radial elements (12) being configured to transfer loads between the inner ring (10) and the outer ring (11), wherein a gas channel for a primary axial gas flow is defined between the inner and outer rings (10, 11), wherein the supporting structure (37, 37') has an inlet side for primary gas flow entrance and an outlet side for primary gas outflow, wherein the radial elements (12) have an airfoil shape with a leading edge (121) directed towards the inlet side, a trailing edge (122) directed towards the outlet side, and two opposite sides (123, 124) connecting the leading edge (121) and the trailing edge (122), wherein the locus of points midway between the two opposite sides (123, 124) forms a mean camber line of each radial element (12), wherein at least one of the radial elements (12) comprises a gas passage arrangement configured to lead a separate bleeding gas flow from the supporting structure (37, 37'). The invention is characterized in that the gas passage arrangement comprises a radially extending gas channel (32) arranged inside the radial element (12) and at least one opening (34) in communication with said gas channel (32), wherein the at least one opening (34) is arranged at one of the sides (123, 124) of the radial element (12). The invention also concerns a gas turbine engine (1, 100) comprising a supporting structure (37, 37') of the above type.
F01D 9/04 - InjecteursLogement des injecteursAubes de statorTuyères de guidage formant une couronne ou un secteur
F02C 6/04 - Ensembles fonctionnels de turbines à gaz délivrant un fluide de travail chauffé ou pressurisé à d'autres appareils, p. ex. sans sortie de puissance mécanique
The invention concerns a gas turbine engine component (27) comprising an outer ring (21), an inner ring (20), a plurality of circumferentially spaced elements (22) extending between the inner ring (20) and the outer ring (21), wherein an annular load transfer structure (23) extends circumferentially along an inner side of the inner ring (20) and also inwards in a radial direction of the component (27). The first portion (23a), at least along a part of the circumference, is inclined in the radial direction in relation to a second portion (23b), wherein the two inclined portions (23a, 23b) are connected in a connection zone (26) between the inclined portions (23a, 23b), and wherein a radial and/or axial position of the connection zone (26) varies along the circumference such that the radial/axial position of the connection zone in a location that circumferentially corresponds to a first element (22a), i.e. in a first circumferential location (A), is radially and/or axially different from the radial/axial position of the connection zone (26) in-between the first element (22a) and an adjacent second element (22b), i.e. in a second circumferential location (B).
The present invention relates to a containment assembly (16) adapted to circumscribe a casing (15) for a gas turbine engine (1). The casing (15) has an axial extension in an axial direction (A) and a circumferential extension in a circumferential direction (C). Moreover, the casing (15) comprises an inner casing surface (15') and an outer casing surface (15"). The inner casing surface (15') is adapted to form a part of a duct (6, 7) of the gas turbine engine (1). The containment assembly (16) comprises an enclosing assembly (20) adapted to at least partially enclose the outer casing surface (15") when mounted. The containment assembly (16) further comprises a resilient means (26) adapted to interact with the enclosing assembly (16). According to the present invention, the enclosing assembly (20) comprises at least two enclosing segments (22, 24) that are arranged at least partially in sequence in the circumferential direction (C) of the casing (15) when the containment assembly (16) circumscribes the casing (15).
F01D 21/04 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs sensibles à une position incorrecte du rotor par rapport au stator, p. ex. indiquant cette position
F01D 25/24 - Carcasses d'enveloppeÉléments de la carcasse, p. ex. diaphragmes, fixations
The invention concerns a gas turbine engine component (37) comprising an outer ring structure (10), an inner ring structure (20) and a plurality of circumferentially spaced radial elements (15) for transferring loads between the inner ring structure (20) and the outer ring structure (10), wherein the outer ring structure (10) comprises a circular ring member (11) for withstanding an internal pressure during operation of the gas turbine engine. The invention is characterized in that the outer ring structure (10) comprises a first set of circumferentially distributed and substantially straight reinforcement ribs (16) for achieving radial stiffness, wherein the reinforcement ribs (16) form sides in a polygonal shape and wherein each of said reinforcement ribs (16) extends between two adjacent radial elements (15). The invention also concerns a gas turbine engine (1) comprising a component (37) of the above type.
The invention concerns a method for manufacturing of a gas turbine engine component (37) comprising an outer ring structure (42, 47), an inner ring structure (41), and a plurality of circumferentially spaced elements (46, 46a, 46b) extending between the inner ring structure (41) and the outer ring structure (42), wherein a primary gas channel for axial gas flow is defined between the elements (46, 46a, 46b), and wherein the component (37) has an inlet side for gas entrance and an outlet side for gas outflow. The invention is characterized in that the method comprises the step of machining a one-piece metal blank as to form a one-piece part (47) comprising: a portion (46b) of each of said elements (46), wherein said portion (46b) relates to a portion of an extension length of the elements (46) between said ring structures (41, 42); and a ring-shaped member (42) that connects said element portions (46b) and that is intended to form part of one of the ring structures. The invention also concerns a gas turbine engine (1) comprising a component (37) manufactured according to the above method.
The present invention relates to a support structure (16) for a gas turbine engine (1). The support structure (16) has an axial extension in an axial direction (A) and a circumferential extension in a circumferential direction (C). Moreover, the support structure (16) comprises a plurality of tubular members (18, 20) of a first material type arranged in sequence in the circumferential direction (C). Each tubular member (18, 20) at least partially delimits a flow guiding passage extending at least partially in the axial direction (A). The support structure (16) comprises a leading portion (22) and a trailing portion (24) in the axial direction (A). Furthermore, the support structure (16) comprises a leading edge member (26) of a second material type, the leading edge member (26) being located at the leading portion (22). At least two of the tubular members (18, 20) are fixedly attached to a leading edge member (26). According to the present invention, the first material type is different from the second material type.
B23P 15/04 - Fabrication d'objets déterminés par des opérations non couvertes par une seule autre sous-classe ou un groupe de la présente sous-classe d'aubes de turbine ou d'organes équivalents, en plusieurs pièces
F01D 9/04 - InjecteursLogement des injecteursAubes de statorTuyères de guidage formant une couronne ou un secteur
The invention concerns a gas turbine engine component (27) comprising an outer ring (21), an inner ring (20), a plurality of circumferentially spaced elements (22) extending between the inner ring (20) and the outer ring (21), wherein a primary gas channel for axial gas flow is defined between the elements (22), wherein the component (27) has an inlet side for gas entrance and an outlet side for gas outflow, and an annular load transfer structure (23) positioned internally of the inner ring (20) for transferring loads between said elements (22) and a bearing structure (24) for a turbine shaft (11) positioned centrally in the component (27), wherein the annular load transfer structure (23) extends circumferentially along at least a part of an inner side of the inner ring (20) and inwards in a radial direction of the component (27), wherein the annular load transfer structure (23) has a first portion (23a) and a second portion (23b), and wherein the first portion (23a) is located closer to the inner ring (20) than the second portion (23b). The invention is characterized in that the first portion (23a) is radially inclined between a first position (30) in the vicinity of the inner ring (20) and an axially displaced second position (31) and wherein the second portion (23b) extends from the second position (31) and is inclined in relation to the first portion (23a), and wherein the annular load transfer structure (23) is provided with a plurality of circumferentially spaced load carrying members (32) arranged at a side of the first portion (23a) axially facing in a direction towards the second position (31), wherein the load carrying members (32) are arranged to form a load carrying connection between the annular load transfer structure (23) and said elements (22) via the inner ring (20). The invention also concerns a gas turbine engine (1) comprising a component (27) of the above type.
The present invention relates to a gas turbine structure (28) comprising a first housing (30) and a second housing (32), one of the first and second housings (30, 32) being located around the other of the first and second housings (30, 32) such that a core flow passage (34) is obtained between the first and second housings (30, 32). The gas turbine structure (28) further comprises an elongate structural member (36) extending in a structural member direction (DSM) from the first housing (30) to the second housing (32) and the gas turbine structure (28) further comprises a fairing (38) circumferentially enclosing at least a portion of the structural member (36).
The present invention relates to a composite guide vane (22) for a gas turbine structure (15). The guide vane (22) is adapted to extend in a guide vane direction (DGv) from a first housing (28) towards a second housing (20) of the gas turbine structure (15). Moreover, the guide vane comprises a guide vane length (LGv) in the guide vane direction (DGv) and the guide vane (22) comprises a first attachment portion (30) with at least one first opening (32) for attachment of the guide vane (22) to the first housing (28). The first opening (32) extends in a first opening direction which forms an angle with the guide vane direction (DGv). The guide vane (22) comprises a cover portion (34) comprising a first material and a core portion (36) which is at least partially enclosed by the cover portion (34).
The present invention relates to a gas turbine structure (15) comprising a guide vane (22). The gas turbine structure (15) further comprises a first housing (28) and a second housing (20) and the guide vane (22) extends from the first housing (28) to the second housing (20). The guide vane (22) comprises a leading edge (38) and a trailing edge (40) and the guide vane (22) extends from the leading edge (38) to the trailing edge (40) along a mean camber line (42). The guide vane (22) comprises a first attachment structure (31 ), a first fastening means (32) and a second fastening means (33). Moreover, the first attachment structure (31 ) comprises a stiffening member (50) extending over at least a portion of a circumference of the first housing (28). Furthermore, the stiffening member (50) comprises a stiffening member centre point (70) as measured along the mean camber line (42).
A gas turbine engine component includes a ring element and a plurality of circumferentially spaced load carrying vanes extending in a radial direction of the ring element. At least one of the vanes is made of a composite material. The at least one composite vane is fastened to the ring element. The component further includes first and second wall portions that are fixed in relation to the ring element and arranged on opposite sides of the composite vane. At least a first hole extends through the first wall portion, the composite vane and the second wall portion. The first hole is adapted to receive a first insertion member. The component is arranged such as to, with regard to the first hole, provide a tight fit for the first insertion member in the first wall portion as well as in the composite vane and to provide a loose fit with a radial play in the second wall portion.
A supporting structure for a gas turbine engine includes at least one annular member, and a plurality of circumferentially spaced elements that extend in a radial direction of the annular member and that are connected to the annular member. At least one of the circumferentially spaced elements has an airfoil shape in cross section. The annular member includes at least two substantially flat panel segments that are arranged side by side in a circumferential direction of the annular member. The panel segments are connected to each other in a connection region that extends along facing end parts of the two panel segments. The airfoil shaped element is connected to the annular member at the connection region. A mean camber line of the airfoil shaped element at least along a portion of its length is inclined in relation to an axial direction of the annular member, and the connection region at least along a portion of its length is inclined in relation to the axial direction of the annular member. The inclined portion of the connection region is directed substantially in parallel to the inclined portion of the mean camber line of the airfoil shaped element. A gas turbine engine including a supporting structure of the above type is also provided.
A gas turbine engine component providing a gas flow passage is provided, which gas flow passage surrounds an inner passage surface, an outer passage surface forming an outer delimitation of said gas flow passage. In an aspect, the trace of the inner passage surface in a cross-section perpendicular to a central axis of the component presents at least one substantially straight portion. In another aspect, where a plurality of circumferentially spaced blades extends between the inner and outer passage surfaces, two portions of the trace, in a cross-section perpendicular to a central axis of the component, of the outer passage surface between two adjacent blades are substantially straight and oriented in an angle in relation to each other so as to form a concavity in the gas flow passage.
F01D 25/16 - Aménagement des paliersSupport ou montage des paliers dans les stators
F01D 25/24 - Carcasses d'enveloppeÉléments de la carcasse, p. ex. diaphragmes, fixations
F02D 9/02 - Commande des moteurs par étranglement des conduits d'amenée de l'air ou du mélange air-combustible ou par étranglement des conduits d'échappement par étranglement des conduits d'amenée
F02M 35/10 - Tubulures d'admission de l'airSystèmes d'introduction
In a blade intended to be exposed to a gas flow at high speed during operation of a flow machine comprising the blade, the blade includes a front end designed to face towards the incoming gas flow and a rear end. The front end is provided with a concave area that is such that during operation a stagnation point for the incoming gas flow arises at a distance in front of an outer blade surface that defines the concave area and such that the outer blade surface is thereby at least partially protected from the incoming gas flow.
The invention relates to a device (20) for moving at least one moveable element (19) in a gas turbine engine (1) between a first and a second position. The device includes a linkage (21) that connects a pivotable annular member (18) with the moveable element (19) in such a way that the movement of the moveable element (19) between the two positions is accomplished when the annular member (18) is pivoted. The linkage (21) has a link member (22) connected to the moveable element (19) via a first articulation joint (23). The device also has a support member (24) that supports the link member (22) at a distance from the first articulation joint.
