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Safran Aircraft Engines patents


Recent patent applications related to Safran Aircraft Engines. Safran Aircraft Engines is listed as an Agent/Assignee. Note: Safran Aircraft Engines may have other listings under different names/spellings. We're not affiliated with Safran Aircraft Engines, we're just tracking patents.

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System for controlling variable-setting blades for a turbine engine

The invention relates to a system for controlling variable-setting blades (14) for a turbine engine, which includes at least one control ring (36) mounted rotatably mobile about an annular casing (16) with axis of revolution a, at least one annular row of variable-setting blades (14) extending substantially radially relative to said axis a and connected to said at least one control ring so that a rotation of the ring about the casing rotates the blades about substantially radial axes b, and means (40) for actuating said at least one ring in order to rotate same about the casing, characterised in that said actuation means are connected to said at least one ring by linking means comprising a shaft (52) extending along an axis c which is substantially radial relative to said axis a and mounted rotatably mobile about said axis c on the housing.. . ... Safran Aircraft Engines

Preform take-up in a jacquard loom

Jacquard loom (100) for producing a woven preform (102) from a plurality of warp yarns and a plurality of weft yarns, said loom comprising a device (106) for taking up the preform when it is being produced, in order to move it along an axis (x) as it is being formed, which axis is substantially parallel to a production direction for the preform, characterised in that said loom also comprises means (105) for rotating the preform, substantially about said axis.. . ... Safran Aircraft Engines

Device and a method for repairing a hole in a part

A repair device for repairing a hole in a part includes a resin tank and an injector endpiece connected to the tank. The endpiece is designed to be inserted in the hole. ... Safran Aircraft Engines

Method for producing a part consisting of a composite material

A method includes fabricating a part out of composite material including fiber reinforcement densified by a metal matrix.. . ... Safran Aircraft Engines

Method for the non-destructive testing of a casing

. . A method for the non-destructive testing of the heating of a part made from polymer material, the method comprising the following steps: a) carrying out a measurement by infrared spectroscopy on a part to be tested and extracting therefrom at least one of absorbance values and transmittance values according to a spatial frequency; and b) from the measurement of at least one of absorbance and transmittance, determining the period of time during which said region of the part to be tested has been subjected to a given heating temperature and determining said heating temperature, using a reference database comprising at least one of absorbance measurements and transmittance measurements, the measurements established over a plurality of reference samples made from polymer material that have been subjected to a given temperature during a given period of time.. . ... Safran Aircraft Engines

An assembly by mechanical connection including at least one part made of composite material

An assembly includes a first part made of composite material and a second part, which parts are held one against the other by at least one fastener system having a fastener element with a head from which there extends a shank. The fastener system further includes a bushing, the bushing including a collar presenting an inside face and an outside face of conical shape. ... Safran Aircraft Engines

Anti-icing system for a turbine engine vane

An anti-icing system for a turbine engine vane extending between an outer casing and an inner casing defining a passage, the system includes an injector device for injecting a jet of air into the passage upstream from the vane, the injector device including a first injection orifice for injecting the jet of air, the first orifice passing through one member selected from the outer casing and the inner casing, and being located in such a manner that, at a first speed of the engine wherein a stream of air flows generally in the passage along a first direction, the first orifice and a leading edge of the vane are substantially in alignment along the first direction, the first orifice is spaced apart from the junction between the leading edge of the vane and the member through which the first orifice passes by a distance of at least 10 mm, and preferably of about 15 mm.. . ... Safran Aircraft Engines

Transmission assembly comprising a transmission member and an oil distribution system

A transmission assembly including a transmission member and an oil distribution system. The transmission member includes a rotary pivot, and a pivot portion for pivoting about the rotary pivot; the oil distribution system is to receive oil feed and to transfer it to an oil reception chamber of the rotary pivot; the rotary pivot includes injection orifices putting the oil reception chamber into fluid flow communication with the gap between the rotary pivot and the pivot portion to form a fluid bearing; and the transmission assembly is for injecting oil into the gap at a first pressure into an outer portion of the gap, and at a second pressure into an inner portion of the gap, the second injection pressure being different from the first injection pressure.. ... Safran Aircraft Engines

Device for attaching manifolds for cooling the casing of a turbine-engine turbine by air jets

The invention relates to a device (3) for attaching manifolds (1) for cooling the casing (2) of a preferably low-pressure turbine of a turbine engine by air jets, comprising a mounting (4) of said manifolds, shaped such as to keep said manifolds spaced apart from one another and a plurality of elements (5) for supporting said manifold mounting (4), each supporting element being attached to the casing, and connected to said manifold mounting (4) by connection means (6, 6′). Said device is characterised in that it comprises n cooling manifolds (1) and n−1 supporting elements (5), each supporting element (5) being arranged between two adjacent cooling manifolds.. ... Safran Aircraft Engines

Method for manufacturing parts made by powder metallurgy comprising the application of a coating

Method for manufacturing a turbine engine part, said method comprising a step (101) of producing said part by powder metallurgy using a material forming the substrate of said part, then a finishing operation comprising at least one first step (103), in which a determined material is deposited onto at least one surface (s1) of the substrate of said part after the powder metallurgy production step (101), and a second step (104) corresponding to a heat treatment operation, so as to form a smooth coating for said surface (s1), characterised in that said determined material is a metal material, so as to form a metal coating.. . ... Safran Aircraft Engines

Bent combustion chamber from a turbine engine

The invention relates to a turbine engine combustion chamber including: an outer annular housing; a flame tube (20) connected to the outer housing, said flame tube (20) including an inner annular wall (20b) and an outer annular wall (20a) that define a first radial inlet portion of the flame tube and a second axial outlet portion of the flame tube, the flame tube also including a chamber base (30) located at the inlet of the flame tube (20); and a fuel injection system (40′) configured to inject fuel into the flame tube via the inlet of the flame tube. The injection system includes an injector axis (aa′), parallel to the first portion, and an air manifold (40′d) configured to move air towards twists in the injection system (40′). ... Safran Aircraft Engines

Aircraft including a streamlined rear thruster with an input stator having movable flaps

An aircraft including a fuselage and a propulsion assembly. The propulsion assembly includes at least one fan rotor placed behind the fuselage as an extension thereof along a longitudinal axis, and a nacelle which forms a fairing of the at least one fan rotor through which at least one air flow passes. ... Safran Aircraft Engines

Stage of variable-pitch blades for a turbine engine, turbine engine and associated installation method

A stage of variable-pitch vanes for a turbine engine includes a plurality of vanes. Each vane has a blade with a first radially internal frusto-conical surface. ... Safran Aircraft Engines

Aircraft comprising a turbine engine incorporated into the rear fuselage comprising a system for blocking the fans

An aircraft comprising a fuselage and propelled by a turbine engine having two coaxial and contrarotating fans, the turbine engine comprising a power turbine having two contrarotating rotors, one of which drives a fan upstream from the turbine, the other a fan downstream from the turbine, each fan comprising a ring of blades, and the assembly of the fans and the power turbine being incorporated at the rear of the fuselage, in the extension of same. The aircraft comprises, for at least one of the fans, a device for blocking the rotation of the fan and a device configured to modify the pitch of the blades of the fan in such a way as to make it operate as a flow straightener with respect to the other fan.. ... Safran Aircraft Engines

07/26/18 / #20180209380

Aircraft comprising two contra-rotating fans to the rear of the fuselage, with spacing of the blades of the downstream fan

The invention relates to an aircraft comprising a fuselage (1), which is propelled by a turbine engine with two coaxial fans, namely an upstream fan (7) and a downstream fan (8), driven by two contra-rotating rotors (5, 6) of a power turbine (3). The two fans (7, 8) and the turbine (3) are integrated into a nacelle (14) which projects downstream from the fuselage (1) and through which air flows. ... Safran Aircraft Engines

07/26/18 / #20180209378

Aircraft propulsion assembly comprising a thrust reverser

An aircraft propulsion assembly, including a turbine engine comprising at least one gas generator configured to generate a main flow, which is supplied by a central jet to at least one power turbine, the central jet being surrounded by an outer fairing, and the power turbine driving, on the periphery thereof, at least one fan rotor. The aircraft propulsion assembly comprises first movable means which are arranged so as to divert at least some of the main flow from the central jet to the outside of the outer fairing and preferably upstream of the turbine engine so as to generate thrust reversal. ... Safran Aircraft Engines

07/26/18 / #20180209346

Isolation of an aircraft turbo engine tank in case of a fire by closing a valve sensitive to the release of an extinguishing agent

A method for isolating a tank of combustible fluid from a downstream portion of a fluid supply system for a turbine engine is provided. The supply system includes the tank and a cutoff valve located between the tank and the downstream portion. ... Safran Aircraft Engines

07/26/18 / #20180209344

Fuel metering device protected against icing

A fuel metering unit comprising a movable element including at least one fuel flow section opening upstream towards a fuel supply duct and opening downstream towards a utilization duct through a metering slot with a flared profile having a narrow flow section and a wide flow section, the movable element able to be moved with respect to a stationary element between a low flow rate position and a high flow rate position, the metering slot made in the stationary element or in the movable element and its obstruction obtained by covering the slot with a wall of the movable element or of the stationary element.. . ... Safran Aircraft Engines

07/26/18 / #20180209302

Assembly comprising a locked securing stud

An assembly including a securing stud held by a locking key that rotates with the stud and is provided with bearing surfaces for sitting on an edge of one of the parts to be assembled. Sufficient clamping torque can be introduced without risk of damage and at the same time avoiding excessive screwing of the stud.. ... Safran Aircraft Engines

07/26/18 / #20180209294

Aircraft comprising a turbine engine incorporated into the rear fuselage with variable supply

The invention concerns an aircraft propelled by a turbine engine having contrarotating fans (7, 8), the turbine engine being incorporated at the rear of a fuselage (1) of the aircraft, in the extension of same and comprising at least two gas generators (2a, 2b) that supply, via a shared central stream (4), a power turbine (3), the turbine (3) comprising two contrarotating rotors (5, 6) for driving two fans (7,8) disposed downstream from the gas generators (2a, 2b), said aircraft comprising means (15) arranged for separating the gas flow in the power turbine (3) into at least two concentric streams (16, 17) and a device comprising first means for distributing the gas flow (21-24) between said streams (16, 17) from the central stream (4), the first distribution means being configured to be able to open or close the supply of at least one so-called sealable stream (16) of the streams (16, 17) of the power turbine (3).. . ... Safran Aircraft Engines

07/26/18 / #20180209291

Aircraft turbine-engine module casing, comprising a heat pipe associated with a sealing ring surrounding a movable impeller of the module

An aircraft turbine-engine module casing including an external module casing and at least one sealing ring intended to surround a movable impeller of the module and arranged radially towards the inside with respect to the external casing. The casing includes at least one capillary heat pipe, a first end which is fixed to the sealing ring, and a second end which, opposite to the first, is fixed to a casing element arranged radially towards the outside with respect to the ring.. ... Safran Aircraft Engines

07/26/18 / #20180208322

Aircraft comprising a propulsion assembly including a fan on the rear of the fuselage

The present invention relates to an aircraft comprising a fuselage and a thruster downstream of the fuselage. The thruster includes a power turbine, located inside a main flow jet, and at least one fan, located inside a secondary flow jet and mechanically driven by the power turbine. ... Safran Aircraft Engines

07/19/18 / #20180202458

Rotary assembly of an aeronautical turbomachine comprising an added-on fan blade platform

. . A rotary assembly of an aviation turbine engine, includes a fan disk having at least one tooth and at least one platform mounted on the tooth of the fan disk. The tooth of the fan disk includes a tab extending the tooth axially upstream, and the platform includes a locking ring at its upstream end for receiving the tab of the tooth of the fan disk. ... Safran Aircraft Engines

07/12/18 / #20180195465

Bleed flow duct for a turbomachine comprising a passively actuated variable section vbv grating

. . The invention relates to a hub (2) of an intermediate casing (1) for a bypass turbomachine comprising: —a bleed stream duct (18), —a bleed valve, comprising a mobile door at the inlet orifice to the bleed stream duct (18), —a set of bleed vanes (22) which are mounted with the ability to rotate about a pivot (26) in the bleed stream duct (18) between an open configuration in which a flow of air coming from the inlet orifice (4) passes between the bleed vanes (22) and a closed configuration, the pivot (26) for each bleed vane (22) being closer to its leading edge (ba) than to its trailing edge (bf).. . ... Safran Aircraft Engines

07/12/18 / #20180195416

Turbine engine exhaust casing with improved lifetime

The invention relates to an exhaust casing for a turbine engine of an aircraft, including: a collar (4), a hub (5), hollow arms (63) connecting said collar (4) to said hub (5), and an opening piece (3) which is located on the collar (4) and is suitable for being connected to an outlet tube (30) of a transient-operation valve of the turbine engine. In said exhaust casing, said opening piece (3) is attached onto the collar (4) and forms an extension of a hollow arm (63) such that an air flow (7) leaving the outlet tube (30) of the transient-operation valve penetrates the hollow arm (63) and flows into the hub (5).. ... Safran Aircraft Engines

07/12/18 / #20180195411

Assembly for turbine

An assembly for a turbine of a turbine engine, including a casing and an annular duct surrounding the casing, which can be connected to a device for supplying cooling air, and having a radially inner annular wall provided with openings arranged opposite the casing in order to cool same by the impact of cooling-air jets. The casing has a plurality of axial grooves including first grooves and second grooves arranged in alternation, and the openings are distributed in a plurality of annular rows in which any pair of consecutive annular rows is such that the openings of one of the annular rows of the pair are centered relative to the first grooves while the openings of the other annular row of the pair are centered relative to the second grooves.. ... Safran Aircraft Engines

07/12/18 / #20180193920

Method for manufacturing a blade comprising a bathtub tip integrating a small wall

The invention relates to a method for manufacturing a turbine engine blade (16) including an active-surface wall (17) and a passive-surface wall (18) separated from one another, this blade (16) including a tip that has a closing wall grouping together the active-surface (17) and passive-surface (18) walls in the region of this tip to define the bottom (23) of a bathtub tip shape located at the tip of the blade, the method comprising a moulding step implementing a core defining the bathtub tip shape. According to the invention, there is a step of adding metal to the bottom (23) of the bathtub tip by means of a direct laser additive manufacturing (clad) method, to deposit material onto the bottom of the bathtub tip to form therein an inner partition (28) supported by the bottom thereof (23).. ... Safran Aircraft Engines

07/12/18 / #20180193906

Installation for manufacturing a part by implementing a bridgman method

An installation for manufacturing a part by implementation of a bridgman method includes in particular a mold intended to receive a melted material and a thermal screen movable with respect to the mold intended to be positioned in front of the solidification front during the directional solidification.. . ... Safran Aircraft Engines

07/05/18 / #20180187553

Turbine blade comprising an improved trailing-edge

A turbomachine blade including a hollow body defining a cavity, and including a downstream trailing edge. The blade further includes at least one hole which communicates with the cavity and which opens downstream onto the trailing edge. ... Safran Aircraft Engines

07/05/18 / #20180185905

Casting tree and method of assembly

A field of lost pattern casting, and more particularly to a casting tree for lost pattern casting, and also to its method of assembly is provided. The casting tree includes at least one part support, at least one pattern, and at least one first male-female connection connecting the pattern to the part support. ... Safran Aircraft Engines

06/28/18 / #20180179898

Nozzle sector for a turbine engine with differentially cooled blades

A nozzle sector for a turbine engine. The nozzle sector includes a radially outer platform, a radially inner platform, a first end blade, a second end blade and at least one first central blade between the end blades in a circumferential direction of the platforms. ... Safran Aircraft Engines

06/28/18 / #20180178280

Knockout method and machine for a cluster of lost-pattern castings

A knockout method for knocking out a cluster (30) of lost-pattern metal castings (32), the cluster of castings being formed in a shell (1), wherein at least one knife is moved by means of a machine without making contact with the cluster in such a manner that the knife engages the shell, breaks it into a plurality of fragments, and detaches at least a portion of the shell from the cluster; a machine for performing the method.. . ... Safran Aircraft Engines

06/28/18 / #20180178274

Shell mold for a sector of a 360º-set of guide vanes

A shell mold (100) for a guide vane sector comprising a first platform and a second platform that are mutually coaxial and between which there extends at least one airfoil, the shell mold having a first platform molding portion (112) suitable for forming the first platform, a second platform molding portion (114) suitable for forming the second platform, an airfoil molding portion (116) suitable for forming the airfoil, a fastener molding portion (118) suitable for forming a fastener projecting from the first platform towards a first edge of the first platform, a main casting duct (120), a distribution duct (122) connecting the main casting duct (120) to the fastener molding portion (118), and an auxiliary duct (124) connecting the main casting duct (120) to the first platform molding portion (112) beside said edge.. . ... Safran Aircraft Engines

06/28/18 / #20180178243

Device for coating a turbomachine annular casing

A device applies a coating to a surface of an turbomachine annular casing, wherein the casing has an abradable layer obtained by polymerising a resin. The device includes first support means and, optionally second automated means movable relative to the first support means. ... Safran Aircraft Engines

06/21/18 / #20180172111

Tool for balancing a turbine engine module

. . The invention relates to tooling for balancing a turbine engine module (10) in a balancing machine, said module having at least one stator housing (14) and a rotor (16) having a shaft (18) with a longitudinal axis a and at least one blade stage (20) surrounded by said stator housing (14), said tooling having at least a balancing frame (14), having rotor (16) guide bearings, first and second means (30, 32) designed to be attached to said stator housing (14), third and fourth means (34, 36) provided on said frame (24), to attach said first and second means (30, 32) to said frame, fifth means for transporting the frame (24), and sixth means (84, 94) for supporting the frame, provided on said frame (24) and cooperating equally well with the balancing machine and with the fifth means for transporting the frame (24).. . ... Safran Aircraft Engines

06/21/18 / #20180171932

Assembly comprising an exhaust case and a downstream rotationally symmetrical part

The invention relates to an assembly comprising: a turbomachine exhaust case (110) that includes an external sleeve and an internal sleeve inside the former, both sleeves extending concentrically about a turbomachine axis, and also includes a plurality of arms extending radially between the sleeves; and an annular part (130) that is centered about the axis, is mounted on one sleeve of the exhaust case, and is located downstream of the exhaust case in the direction in which the air flows inside the turbo-machine; the assembly is characterized in that the annular part and the sleeve of the exhaust case on which the annular part is mounted each have a circumferential thread (131, 115), said threads cooperating with each other in order to allow the annular part to be screwed onto the sleeve of the exhaust case.. . ... Safran Aircraft Engines

06/21/18 / #20180171803

Method for manufacturing a turbomachine fan having a reduced noise level at multiple rotational frequencies of said turbomachine

The invention relates to a method (100) for manufacturing a turbomachine fan comprising a plurality of blades mounted on a disc extending along a longitudinal axis, said method comprising the following steps of: —measuring (102) at least one structural parameter of each of the blades in the cold state, and, for each of the blades, estimating (103) at least one structural parameter in operation relating to a blade from the structural parameter(s) measured on said blade in the cold state, —determining (104) an optimal sequence of the blades around the disc from the structural parameters estimated in operation for each of the blades, and —mounting (105) the blades on the disc in the optimal sequence thus determined.. . ... Safran Aircraft Engines

06/21/18 / #20180170580

Method for monitoring an aircraft engine operating in a given environment

The present invention relates to method for monitoring an engine (1) of an aircraft (2) operating in a given environment. The invention is characterized in that it comprises the implementation, via means for data processing (31), of the steps of: (a) receiving a sequence of n-tuples (x1-exec . ... Safran Aircraft Engines

06/21/18 / #20180170522

System for electromechanical pitch actuation for a turbine engine propeller

A pitch actuation system for a turbine engine propeller includes an actuator with movable part configured to rotate the blades of the propeller relative to the blade pitch axes. The actuator includes a transmission screw that is rotatable and movable in translation along a longitudinal axis, and a nut that engages the screw to move in translation along the longitudinal axis to adjust the pitch of the propeller blades. ... Safran Aircraft Engines

06/21/18 / #20180169972

Method and tooling for shaping a fan casing

A method of shaping the profile of a fan casing having an inside surface, the method including placing the casing around a surface of revolution of a drum of shaping tooling; interposing at least one bladder that is inflatable under the action of a fluid under pressure between a portion of the inside surface of the casing and the drum, the bladder extending over all or part of the surface of revolution of the drum; stoving the assembly including the casing, the tooling, and the at least one bladder at a predetermined temperature; and during the stoving, applying isostatic pressure via the at least one bladder so as to impart a cylindrical profile to the portion of the inside surface of the casing facing the at least one bladder.. . ... Safran Aircraft Engines

06/14/18 / #20180164150

Method and device for determining the vibration of rotor blades

. . A method for determining the vibration of turbomachine rotor blades, including steps of measuring, via one or more sensors, the variation in the minimum distance between each sensor and the top of each blade along a radial axis of the rotor, between successive rotations of each blade in front of each sensor, a minimum distance value being obtained on each passage of each blade in front of each sensor, in order to deduce therefrom a variation in the lengths of the blades along the radial axis; and, using directly, as such, the variation in the length of the blades along the radial axis in a model of the deformation of the blades, in order to deduce therefrom characteristics of one or more vibrational modes of the rotating blades. A turbomachine can be equipped with a device implementing this method.. ... Safran Aircraft Engines

06/14/18 / #20180163850

Reduction gear having an epicyclic gear train for a turbine engine

A reduction gear having an epicyclic gear train for a turbine engine, in particular of an aircraft, comprising: a planetary shaft having an axis of rotation a; a ring gear having an axis a extending around said planetary shaft; planet gears distributed around said axis a, which mesh with said ring gear and the planetary shaft; and a planet carrier including members for supporting bearings of the planet gears, having axes of rotation b, which are evenly distributed around the axis a, as well as a part holding each supporting member substantially by the middle thereof along the axis b thereof, wherein said supporting members are made as a single part having means for supplying lubricating oil to said bearings. A method for assembling said reduction gear is also provided.. ... Safran Aircraft Engines

06/14/18 / #20180163556

Assembly of turbine engine parts comprising a fan blade having an integrated platform, and corresponding turbine engine

The invention relates to an assembly of turbine engine (50) parts having a longitudinal axis (x), the assembly comprising at least two adjacent fan blades made of a 3d woven composite material, each blade (1) comprising a root (3), a vane (2) extending from the root (3), and an integrated platform (4) inserted between the root (3) and the vane (2), the platform (4) extending on either side of the vane (2) and forming a pressure face platform portion (4i) and a suction face platform portion (4e), the pressure face and suction face platform portions of adjacent blades being connected by a fin (14) extending at least radially at the outer side of the pressure face and suction face platform portions (4i, 4e), relative to the longitudinal axis (x), having a substantially triangular transverse cross section and extending axially and so as to upper overlap the longitudinal edges (11a, 11b) of the pressure face and suction face platform portions (4i, 4e) of the adjacent blades.. . ... Safran Aircraft Engines

06/07/18 / #20180156654

Oil level sensor

The oil level sensor is associated with a tank containing oil and comprises a float movable along guide means and which can float on the oil of the tank, so as to move with the oil level, a permanent magnet movable with the float, and an electronic card provided with magnetic switches sensitive to said magnet. The guide means of the spherical float are positioned about the float and the contacts thereof with said float are limited to three substantially linear zones.. ... Safran Aircraft Engines

06/07/18 / #20180156239

Turbine engine part with non-axisymmetric surface

A turbine engine part or set of parts including at least first and second obstacles each having a leading edge and a trailing edge, and a platform from which the obstacles extend; wherein the platform has, between the pressure side of the first obstacle and the suction side of the second obstacle a non-axisymmetric surface defining at least one fin with a substantially triangular cross-section, each fin being associated with a leading position and a trailing position on the surface between which the fin extends, such that the leading position is upstream of each of the leading edges; the trailing position is downstream of each of the leading edges.. . ... Safran Aircraft Engines

06/07/18 / #20180156237

Turbine engine flow guide vane with removable attachment

The invention relates to a guide vane intended to be mounted in a turbine engine between an inner shroud (17) and an outer shroud (16), comprising a longitudinal straightening body (41) for an air flow extending between a first end (42) intended to be positioned at the inner shroud (17) and a second end (44) intended to be positioned at the outer shroud (16), the longitudinal straightening body having an aerodynamic profile defined by a leading edge (41a) and a trailing edge (41b) in the flow direction of the air flow, and by a camber line (41c) extending from the leading edge (41a) to the trailing edge (41b). It further comprises a first attachment heel (43) and a second attachment heel (45) positioned in the continuation of the longitudinal body (41) at the first end (42) and the second end (44) respectively, the first (43) and second (45) attachment heels being planar elements arranged parallel with respect to one another, each attachment heel (43; 45) being arranged at a distance from the leading edge (41a) and from the trailing edge (41b).. ... Safran Aircraft Engines

06/07/18 / #20180156232

Composite blade, comprising a leading-edge reinforcement made of another material

The invention relates to a blade made of composite material, provided with a reinforcement (17) made of a stronger material with regard to a leading edge (5) likely to be struck by solid objects, which comprises, under the aerodynamic portion of the blade, extensions (18, 19) attached to the blade root. According to the invention, the extensions (18, 19) are dissymmetric, the extension (18) located on the suction side being generally further back from the leading edge (5) than the extension (19) on the pressure side.. ... Safran Aircraft Engines

06/07/18 / #20180156139

Hydromechanical cutoff device with hysteresis for a turbomachine lubrification system

The invention relates to a lubrication system for turbomachine. The hydromechanical cutoff device is configured to close when a rotation speed of a turbomachine shaft reduces and becomes lower than a first threshold, and wherein the hydromechanical cutoff device is configured to open when the rotation speed of the shaft increases and becomes higher than a second threshold higher than the first threshold.. ... Safran Aircraft Engines

06/07/18 / #20180156070

Turbine for turbine engine

A turbine for a turbine engine is disclosed, comprising a casing and a rotor comprising blades, the radially external periphery of which comprises at least one first wiper extending radially outwards, sealing means extending radially around the blades and comprising a ring made from abradable material, the radially external ends of the first wipers being engaged in a groove in said ring made from abradable material so as to form a labyrinth-type seal, wherein said ring is formed by a plurality of contiguous annular sectors, each sector comprising at least one fixing member cooperating with at least one complementary attachment flange of the casing so as to provide mounting of each sector on the casing by axial movement of said sector with respect to the casing, the casing being configured so as to allow the mounting of the sectors in the casing solely by axial movement of said sectors.. . ... Safran Aircraft Engines

06/07/18 / #20180156068

A turbine ring assembly

A turbine ring assembly includes both a plurality of ring sectors made of ceramic matrix composite material forming a turbine ring, and also a ring support structure. Each ring sector includes a portion forming an annular base with an inner face defining the inside space of the turbine ring and an outer face from which an attachment portion of the ring sector extends for attaching it to the ring support structure. ... Safran Aircraft Engines

06/07/18 / #20180156066

Oil tank comprising an oil level control device

An oil tank for a turbine engine comprising a closed enclosure having the shape of an arc of circle adapted to receive oil, with the enclosure having a lower portion and an upper portion at a distance from each other, and an oil level control device in the enclosure, wherein the oil level control device comprises a first sensor positioned in the lower portion of the enclosure so as to control the oil level, and a second sensor separate from the first sensor and positioned in the upper portion of the enclosure so as to control the oil level.. . ... Safran Aircraft Engines

06/07/18 / #20180155041

Turbine engine suspension device

A suspension device for suspending, for example, a turbine engine to a pylon, the device comprising a first unit interposed between a first lug and a second lug of a second unit, the first unit having a bore for passing an axle running through first and second bushings mounted respectively in said first and second lugs; and a clamping unit interacting with the axle. The device further comprising a third bushing mounted in the second lug and having radial centering means which interact with the complementary centering means of a head section of the axle; and a fourth axially slidably mounted bushing, biased axially by the clamping unit, and comprising radial centering means which interact with complementary centering means of the first bushing.. ... Safran Aircraft Engines

06/07/18 / #20180154427

Method for producing a pattern for lost pattern casting

A fabrication method for fabricating a pattern for lost pattern casting, comprising at least one insert (10), providing at least two pattern portions (12, 14), said at least two portions being made of a material that can be eliminated, and assembling said at least two pattern portions together around at least a portion of said at least one insert in sealed manner so as to make said pattern (16).. . ... Safran Aircraft Engines

05/31/18 / #20180149621

Method of non-destructive inspection of a weld bead

. . A non-destructive method for inspecting a weld bead (18) connecting together two parts (14, 16), a longitudinal direction (x) of the weld bead (18) extending along the interface between the two parts (14, 16), the method comprising providing an emitter (10) and a receiver (12) and taking at least one measurement of a signal emitted by the emitter (10) and received by the receiver (12) after passing through the weld bead (18), wherein the emitter (10) and the receiver (12) are positioned relative to the weld bead (18) in such a manner that the plane containing the axis (a1) of the emitter and the axis (a2) of the receiver is substantially parallel to the longitudinal direction (x).. . ... Safran Aircraft Engines

05/31/18 / #20180149544

Modular calibration rotor for a horizontal balancer

A calibration rotor for a horizontal balancer, configured to be driven by a driver. The calibration rotor includes a main barrel that has a longitudinal axis and the periphery of which includes points for attaching balance weights that are evenly distributed in an axial and angular manner about the axis. ... Safran Aircraft Engines

05/31/18 / #20180149034

Turbine ring assembly supported by flanges

A turbine ring assembly includes both a plurality of ring sectors made of ceramic matrix composite material forming a turbine ring and also a ring support structure having first and second annular flanges, each ring sector having first and second tabs, the tabs of each ring sector being held between the two annular flanges of the ring support structure. Each of the first and second tabs of the ring sectors has an annular groove. ... Safran Aircraft Engines

05/31/18 / #20180149033

Turbomachine case comprising an acoustic structure and an abradable element

A case of a turbomachine including an inner wall of revolution coaxial with a principal longitudinal axis of the turbomachine, an acoustic insulation structure installed inside the inner wall of revolution that has an annular principal shape centred on the longitudinal axis, and an abradable annular element, wherein the abradable element is fixed on an inner annular face of the acoustic insulation structure.. . ... Safran Aircraft Engines

05/24/18 / #20180142626

Butterfly valve for bleeding a compressor for an aircraft turbine engine

. . The invention relates to a butterfly valve (24) for bleeding a compressor for an aircraft turbine engine, the valve including a valve body (32), a butterfly (36), and a device (42) for controlling the angular position of the butterfly, the device (42) including a mobile actuation member (64) connected to the butterfly by a link (70), the member (64) being subjected: to a first adjustable pressure force (f1) applied by air from the compressor, the first force (f1) returning the butterfly (36) to a closed position; and to a second mechanical force (f2) returning the butterfly (36) to an open position, and coming from an aerodynamic torque (c) applied by the air to the butterfly (36), of which the axis of rotation (38) is off-centre relative to the butterfly.. . ... Safran Aircraft Engines

05/24/18 / #20180142563

Annular wall of a combustion chamber with optimised cooling

An annular turbine engine combustion chamber wall including air admission orifices to create zones of steep temperature gradient, and cooling orifices to enable the air flowing on the cold side to penetrate to the hot side in order to form a film of cooling air along the annular wall, the annular wall being further includes, in the zones of steep temperature gradient, multi-perforation holes having respective bends of an angle α greater than 90°, the angle α being measured between an inlet axis ae and an outlet axis as of the multi-perforation hole, the outlet axis of the multi-perforation hole being inclined at an angle θ3 relative to the normal n to the annular wall through which the multi-perforation holes with bends are formed, in a “gyration” direction that is at most perpendicular to the axial flow direction d of the combustion gas.. . ... Safran Aircraft Engines

05/24/18 / #20180142562

Turbine engine comprising a lobed mixer having scoops

The invention relates to a lobed mixer (8) for arranging on the downstream end of a hood (4) separating the two co-axial streams, respectively inner and outer, the mixer (8) being shaped so as to have a least one peripheral succession of lobes (20, 21) which are generally radially oriented in relation to a longitudinal axis (ll′) of the mixer, each lobe (20, 21) forming a channel extending mainly along the longitudinal axis and comprising at least one peripheral succession of baffles (26) from the first stream to the second stream and/or from the second stream to the first stream, arranged on said lobes (20, 21).. . ... Safran Aircraft Engines

05/17/18 / #20180135561

Testing method

. . The invention relates to the field of technical testing, and more particularly to a method of testing a machine, the method comprising: at least one step (s101) of determining a plurality of operating points for said machine, each operating point being defined by a duration and a specific value of at least one operating parameter of the machine; a step of calculating a set of distances between pairs of operating points; a step (s106) of selecting an optimum sequence of operating points by applying an algorithm for solving the traveling salesman problem to said set of distances between pairs of operating points; and a step (s107) of controlling the operation of said machine according to said optimum sequence of operating points.. . ... Safran Aircraft Engines

05/17/18 / #20180135462

Device for uncoupling first and second parts of a turbine engine

Device for uncoupling first and second parts of a turbine engine, the parts extending around an axis a, wherein the first part comprises an annular row of axial through-openings that extend around the axis a, and in that the second part comprises an annular row of axial fusible lugs that pass axially through the openings and include threaded portions onto which a nut having the axis a and intended for axially clamping the parts is screwed.. . ... Safran Aircraft Engines

05/17/18 / #20180135436

Blade having platforms including inserts

A fiber preform for a turbine engine blade and a single-piece blade suitable for being made by means of such a preform, a bladed wheel, and a turbine engine including such a blade; the fiber preform obtained by three-dimensional weaving comprises a first longitudinal segment (41) suitable for forming a blade root, a second longitudinal segment (42) extending the first longitudinal segment (41) upwards and suitable for forming an airfoil portion, a first transverse segment (51) extending transversely from the junction between the first and second longitudinal segments (41, 42), and suitable for forming a first platform; the first transverse segment (51) includes at least one non-interlinked portion comprising a top strip (51b) and a bottom strip (51a), and at least one insert (61) is arranged between the top and bottom strips (51b, 51a) of the non-interlinked portion of the first transverse segment (51).. . ... Safran Aircraft Engines

05/17/18 / #20180135433

Turbine for a turbine engine

A turbine for a turbine engine, the turbine comprising a rotor including blades the radially external periphery of which includes at least one first wiper radially extending outwards, sealing means radially extending about the blades and including a ring made of abradable material, with the radially external ends of the wipers cooperating with said ring made of abradable material so as to form a labyrinth-type seal, wherein said ring includes at least one first portion axially extending upstream of the first wiper and a second portion, different from the first portion, axially extending downstream of the first wiper, with the first portion and/or the second portion including a groove, wherein the first wiper has been inserted, with said groove being defined by the first portion and by the second portion.. . ... Safran Aircraft Engines

05/17/18 / #20180134402

Bipartite cradle with slide for turbomachine

A cradle for supporting an aircraft turbine engine, said cradle comprising an attachment interface for attaching a gas generator of the turbine engine. The cradle is produced in at least two portions comprising: an upper half-cradle, which is designed to be attached to a wing of the aircraft, a lower half-cradle, which is movable between a position in which it is connected to the upper half-cradle and a position in which it is disconnected from the upper half-cradle, and which comprises at least a portion of the attachment interface for attaching the gas generator, a guide configured for slidably guiding the lower half-cradle between the connected and disconnected positions thereof, and at least one lock for locking the lower half-cradle in the connected position thereof.. ... Safran Aircraft Engines

05/17/18 / #20180133791

Method for knocking out a foundry core, and method for manufacturing by moulding including such a method

A method for knocking out a foundry core confined in an internal cavity in a part at the end of a casting operation, in particular a lost-wax casting operation, includes at least a primary chemical knocking-out step. During the primary chemical knowing-out step, the part is subjected to a chemical solution to dissolve the core, in a sealed enclosure. ... Safran Aircraft Engines

05/10/18 / #20180128289

Hydraulic circuit with controlled recirculation circuit

The invention relates to a hydraulic circuit (10) for an aircraft turboprop comprising a hydraulic fluid tank (16), a pump (14), a component (12) that is supplied with fluid pressurised by the pump (14) and that is selectively put into operation, and a fluid recirculation circuit (20) between the pump discharge (14) and the tank (16) characterised in that it comprises a valve (22) located in the recirculation circuit (20), that is capable of closing the recirculation circuit (20) when the component (12) is not in operation and is capable of opening the recirculation circuit (20) when the component is in operation.. . ... Safran Aircraft Engines

05/10/18 / #20180128183

Turbine engine with a pair of contrarotating propellers placed upstream of the gas generator

Engine comprising a propeller unit with a pair of contrarotating propellers (31, 32), a gas generator (5) supplying a power turbine (53), the pair of propellers being rotationally driven by the shaft (53a) of the power turbine via a speed reduction gearbox, the axis of rotation (xx) of the pair of propellers not being coaxial with that (yy) of the power turbine, the speed reduction gearbox comprising a differential gearset (7) and a first stage (6) comprising a simple gearset connecting the turbine shaft (53a) and the differential gearset (7), the engine air intake comprising an air intake duct (11), the air intake duct (11) being in the shape of a lobe adjacent to the assembly formed by the simple gearset and the differential gearset (7).. . ... Safran Aircraft Engines

05/10/18 / #20180128180

Support providing a complete connection between a turbine shaft and a degassing pipe of a turbojet

The invention relates to a support providing a complete connection between a degassing pipe (3) and a turbine shaft (2), said support including a plurality of outer contact spans (41a) intended to bear on the inner walls of the turbine shaft to secure the degassing pipe with respect thereto, characterized in that the different spans are each bordered by at least one elastomer insert (41g) which contributes to the protection of the turbine shaft during insertion of the support therein.. . ... Safran Aircraft Engines

05/10/18 / #20180128173

Turbine engine fan module including a turbine engine inlet cone de-icing system, and a de-icing method

The invention relates to an aviation turbine engine fan module including a de-icing system (10) for de-icing an inlet cone (1) and comprising a sheath (30) placed inside an inside space defined upstream by the inlet cone, said sheath comprising a first duct (38) having at least one hot air admission orifice (42), said first duct being configured to convey hot air from a bearing enclosure (22) of the engine towards a wall of the inlet cone in order to heat it from the inside, the sheath further comprising a second duct (40) having at least one outlet situated downstream from the admission orifice of the first duct, said second duct being configured to discharge air from the first duct towards the downstream end of the engine. The invention also provides a method of de-icing a turbine engine inlet cone.. ... Safran Aircraft Engines

05/10/18 / #20180128120

Connecting assembly for cooling the turbine of a turbine engine

A connecting assembly comprising an air distribution housing, between an air inlet passage and at least one duct connected with the housing by at least one bush mounted on an orifice of a wall of the housing. The wall of the housing and an inner wall of the bush are connected by a fillet having a radius which is maximum over an angular sector.. ... Safran Aircraft Engines

05/10/18 / #20180127084

Unducted-fan aircraft engine including a propeller comprising vanes with roots outside the nacelle and covered by detachable covers

The invention relates to an unducted-fan aircraft engine comprising a generally cylindrical nacelle through which a primary jet flows, the nacelle bearing a fan rotor comprising variable-pitch vanes (33) located radially outside the nacelle in order to be traversed by a secondary jet (31) flowing longitudinally around the nacelle. The rotor comprises a hub bearing variable-pitch vane supports each carrying one vane (33), each vane (33) comprising a blade extending from a root that is used to removably attach same to a base of the associated support (34). ... Safran Aircraft Engines

05/10/18 / #20180126468

Method for machining an attachment flange of an aircraft turbomachine case

A method of machining an attachment flange of an aircraft turbomachine case, the method also including a system to machine the two opposite surfaces of the flange, and including a shape follower machining module, this module being designed to follow the shape of the flange and including a first structure equipped with a first machining tool, and a second structure equipped with a second machining tool, the flange fitting between the structures such that its two opposite surfaces are machined by tools, the module also including shape follower elements carried by the structures; and a device to drive movement of the machining module along a circumferential direction of the flange.. . ... Safran Aircraft Engines

05/03/18 / #20180119682

Device and method for regulating flow rate

A flow rate regulator device is provided, including an upstream chamber, a downstream chamber, a plurality of electrically conductive capillary ducts providing parallel fluid flow connections between the upstream chamber and the downstream chamber, first and second electrical terminals configured to be connected to an electric current source, and at least one electric switch configured to connect one or more of the capillary ducts selectively between the electrical terminals. A system for feeding propellant gas to a space electric thruster is also provided, including at least one such flow rate regulator device to regulate a propellant gas flow rate. ... Safran Aircraft Engines

05/03/18 / #20180119575

Turbine engine unit for lubricating a bearing holder

Provided, including an inter-turbine housing that includes a hub including a bearing bolder, a ferrule extending around and at a distance from the hub, at least one arm extending radially between the hub and the ferrule, and at least one lead-through for lubricating the bearing holder. The lead-through includes a first pipe having an end portion that can be screwed onto the hub so as to place the first pipe in fluid communication with the bearing holder, an intermediate portion secured to the end portion placed inside the arm when the end portion is screwed onto the hub, and a clamping portion secured to the end portion and rotatable by a clamping tool. ... Safran Aircraft Engines

05/03/18 / #20180119551

Reinforcement for the leading edge of a turbine engine blade

A turbine engine blade comprising an aerodynamic surface that extends in a first direction between a leading edge and a trailing edge, and in a second direction substantially perpendicular to the first direction between a root and a tip of the blade, and a leading-edge reinforcement comprising a fin partly covering the aerodynamic surface of the blade, characterised in that the fin has a radially outer edge arranged in the vicinity of the tip of the blade and extending between the leading edge and the trailing edge, this radially outer edge comprising an upstream point fitting flush with the tip of the blade at the leading edge and a so-called downstream point separated from the tip of the blade.. . ... Safran Aircraft Engines

05/03/18 / #20180119550

Blade comprising lands with a stiffener

A preform for a turbine engine blade, the preform comprising a main fiber preform obtained by three-dimensional weaving and comprising a first longitudinal segment suitable for forming a blade root (21), a second longitudinal segment extending the first longitudinal segment upwards, and suitable for forming an airfoil portion (22), and a first transverse segment extending transversely from the junction between the first and second longitudinal segments, and suitable for forming a first platform (23), wherein the preform also includes at least one stiffener (40) fitted on the main fiber preform along at least a portion of the distal edge of the first transverse segment.. . ... Safran Aircraft Engines

05/03/18 / #20180118358

Assembly between an aircraft pylon and a turbine engine

An assembly between an aircraft structural pylon and an aircraft turbine engine is disclosed, with the assembly comprising a beam intended to be attached to the turbine engine and wherein a knuckle intended for the installation of a pad integral with the pylon is mounted, with the beam comprising suspension lugs each including a bore for the passage of a shaft intended to further go through a bore formed in the pylon to connect the beam with the pylon.. . ... Safran Aircraft Engines

05/03/18 / #20180117807

Method of fabricating a blade platform out of composite material with integrated gaskets for a turbine engine fan

There is provided a method of fabricating a blade platform out of composite material with integrated gaskets for a turbine engine fan, the platform including a base and a stiffener, the method including using three-dimensional weaving to make a single-piece fiber blank with a plurality of longitudinal yarn layers extending in a direction corresponding to a longitudinal direction of the base of the platform; shaping the fiber blank to form a fiber preform having a first portion forming a base preform and a second portion forming a stiffener preform; positioning platform gaskets at side margins of the first portion of the fiber preform forming a base preform; placing the fiber preform with the gaskets in an injection mold; injecting resin into the injection mold; compacting the assembly; heating the injection mold to solidify the resin; and unmolding the resulting platform.. . ... Safran Aircraft Engines

04/19/18 / #20180105278

Turbine engine having horizontally offset axes

The invention relates to an aircraft propulsion assembly comprising a cradle receiving a turbine engine comprising at least one propeller having a longitudinal axis of rotation, a gas turbine engine having a longitudinal axis of rotation offset from the axis, and a reduction gear by means of which said propeller receives drive power from said engine, wherein the propeller and the gas turbine engine are designed such that axes and are offset from one another within said cradle at least by a given value in a transverse direction, the axis of the gas turbine engine being transversely closer to a proximal lateral side of the cradle than to an opposite distal lateral side of the cradle in order to create a lateral space between said engine and said distal lateral side of the cradle, thereby forming at least one region for installing equipments, components or accessories of said turbine engine.. . ... Safran Aircraft Engines

04/12/18 / #20180101825

System for pooling data relating to aircraft engines

A system for pooling observation data relating to aircraft engines includes a receiver adapted for recovering the observation data from distinct entities, a processor adapted for describing the observation data in a metric space by transforming them into measurable observation states, and a database adapted for storing therein the observation states.. . ... Safran Aircraft Engines

04/12/18 / #20180101635

System and method for product data management and 3d model visualization of electrical wiring design and specifications

A method for generating a three-dimensional (3d) computer model of an assembly that includes wiring routing, which includes creating a part data structure that defines a part in a virtual product management system. The part data structure includes a plurality of nodes that define at least 3d part design data, 3d wiring routing design data and wiring routing annotation data of the part. ... Safran Aircraft Engines

04/12/18 / #20180100516

Vane comprising an assembled platform and blade

A blade portion of a vane is assembled to a platform portion by insertion into a cavity in the latter, in a lateral direction from an opening in the cavity, where the cavity possesses a back wall over at least a portion of its surface area. The bonding surface area between the portions of the vane is increased. ... Safran Aircraft Engines

04/12/18 / #20180100400

Blade equipped with platforms comprising a retaining leg

A preform for a turbine engine blade, the preform being obtained by three-dimensional weaving and comprising a first longitudinal segment (31) suitable for forming at least a portion of a blade root, a second longitudinal segment (32) extending the first longitudinal segment (31) upwards, and suitable for forming at least a portion of a stilt portion, a third longitudinal segment (33) extending the second longitudinal segment (32) upwards, and suitable for forming an airfoil portion, a first transverse segment (34) extending transversely from the junction between the second and third longitudinal segments (32, 33), and suitable for forming a first platform, and a first oblique segment (36) extending from the junction between the first and second longitudinal segments (31, 32) to the first transverse segment (34), and suitable for forming a retaining leg (26) for the first platform.. . ... Safran Aircraft Engines

04/05/18 / #20180094584

Cooling of the oil circuit of a turbine engine

The invention relates to a turbine engine, such as a turbojet engine or a turboprop engine of an aeroplane, including at least one oil circuit (8) and cooling means (16) for cooling the oil of said circuit (8), the cooling means (16) including a refrigerant circuit (17) provided with a first heat exchanger (18) capable of exchanging heat between the refrigerant and the air and forming a condenser, a second heat exchanger (19) capable of exchanging heat between the refrigerant and the oil of the oil circuit and forming an evaporator, a pressure reducer (20), a compressor (21) and first regulator means (31) capable of regulating the pressure of the refrigerant entering the first exchanger (18).. . ... Safran Aircraft Engines

04/05/18 / #20180094535

Gyratory-effect flow deflector of a discharge valve system, discharge valve system and turbine engine comprising such a discharge valve system

A discharge valve system of a bypass turbine engine compressor includes a flow deflector. The flow deflector has a wall provided with a plurality of ejection channels configured to discharge a discharge airflow from the compressor in a duct of the turbine engine in which an airflow circulates. ... Safran Aircraft Engines

04/05/18 / #20180094526

Rotor disk comprising a variable thickness web

A disk of a rotor including an annular radial web, a radially central hub located at the inner radial end of the web and a rim located at the outer radial end of the web, the web including an upstream face and a downstream face, and a plurality of orifices through which bolts pass for the attachment of at least one annular flange forming part of another adjacent rotor disk on either the upstream face or the downstream face of the web, or on both faces. The upstream face and/or the downstream face of the web includes a globally annular shaped indentation, with a bottom set back along the axial direction inwards into the web, and that extends radially outwards from the hub of the disk towards the rim, and that surrounds a radially inner part of each of the orifices of the web, at a distance.. ... Safran Aircraft Engines

03/29/18 / #20180087405

Turbine ring assembly that can be set while cold

A turbine ring assembly comprises a plurality of ring sectors made of ceramic matrix composite material forming a turbine ring and a ring support structure including first and second annular flanges. In section, each ring sector presents a k-shape having an annular base-forming portion with an inside face defining the inside face of the turbine ring and an outside face with first and second s-shaped tabs projecting therefrom. ... Safran Aircraft Engines

03/29/18 / #20180087401

Turbine ring assembly comprising a cooling air distribution element

A turbine ring assembly includes a plurality of ring segments and a ring support structure, the ring assembly further including, for each ring segment, a cooling distribution element fixed to the ring support structure and positioned in a first cavity delimited between the turbine ring and the ring support structure.. . ... Safran Aircraft Engines

03/29/18 / #20180087400

Turbine ring assembly comprising a cooling air distribution element

A turbine ring assembly includes a plurality of ring segments and a ring support structure, the ring assembly further including, for each ring segment, a cooling distribution element fixed to the ring support structure and positioned in a first cavity delimited between the turbine ring and the ring support structure.. . ... Safran Aircraft Engines

03/29/18 / #20180087399

Turbine ring assembly comprising a cooling air distribution element

A turbine ring assembly includes a plurality of ring segments and a ring support structure, the ring assembly further including, for each ring segment, a cooling distribution element fixed to the ring support structure and positioned in a first cavity delimited between the turbine ring and the ring support structure.. . ... Safran Aircraft Engines

03/29/18 / #20180087392

Turbomachine provided with a vane sector and a cooling circuit

A turbomachine including at least one stator vane sector (10) and a fluid distribution circuit (22), the stator vane sector comprising at least one vane (12), a fluid inlet (25a), a fluid outlet (25b), and a channel (24a) providing fluid flow connection between the fluid inlet and the fluid outlet while extending at least in part in the vane (12), the vane and the channel being adapted, to enable heat to be exchanged between a hot fluid passing through the channel and a stream of cold air passing through the vane sector, the fluid distribution circuit (22) presenting a feed pipe (22a) and a recovery pipe (22b), the fluid inlet (25a) being in fluid flow connection with a branch tapping (23a) of the feed pipe (22a) while the fluid outlet (25b) is in fluid flow connection with a branch tapping (23b) of the recovery pipe (22b).. . ... Safran Aircraft Engines

03/29/18 / #20180087389

Blisk comprising a hub having a recessed face on which a filling member is mounted

A fan blisk for a turbomachine, the blisk comprising a hub delimited by an upstream face and a downstream face in addition to an outer peripheral face and a revolving inner peripheral face delimiting an inner opening, the hub carrying blades each having a leading edge and a trailing edge, the hub and the blades forming a one-piece assembly. The upstream face and/or the downstream face is offset, being located along the axis of rotation (ax) between the leading edges and the trailing edges of the blades, the blisk (bladed disk) comprising a filling member mounted on the offset face, the filling member comprising an inner centring ferrule engaging in the inner face of the hub, a radial portion resting against the offset face and an outer ferrule extending the outer peripheral face of the hub.. ... Safran Aircraft Engines

03/29/18 / #20180087386

Fan blisk for aircraft turbomachine

The invention relates to a fan blink (28) for an aircraft turbomachine, comprising a hub, an annular platform (42) and fan blades (44) arranged projecting from the annual platform. It also comprises a mechanical discharge slit (52) from a trailing edge (49) of the fan blade, associated with at least one of the fan blades (44), for the case of ingestion of a bird, the slit being made on the annular platform (42) going around the trailing edge (49).. ... Safran Aircraft Engines

03/29/18 / #20180085856

Device for fabricating annular pieces by selectively melting powder, the device including a powder wiper

A device (10) for fabricating annular pieces by selectively melting powder, the device comprising an inner annular wall (12) and an outer annular wall (14) that are concentric and that define an annular powder deposition zone (a), and a powder dispenser (16) movable in rotation about the axis (x) of the inner and outer annular walls (12, 14), the powder dispenser (16) including a wiper (18) extending between the inner annular wall (12) and the outer annular wall (14) and forming an angle (α) with the radial direction (r) of the inner and outer annular walls (12, 14).. . ... Safran Aircraft Engines

03/22/18 / #20180080408

Device with gratings for ejecting microjets in order to reduce the jet noise of a turbine engine

A device for reducing the jet noise of a turbine engine includes an outer cover having an inside wall defining the outside of an annular passage for passing a bypass stream from the engine, the wall of the outer cover including a plurality of microjet circuits, each including intakes for taking a gas stream from the bypass stream flow passage and leading to a single feed duct, which in turn opens out into the trailing edge of the outer cover via at least one ejection grating suitable for splitting the intake gas stream into a plurality of gas streams of right sections of dimensions less than a right section of the feed duct.. . ... Safran Aircraft Engines

03/22/18 / #20180080385

Assembly for passing an electrical harness into a turbine engine

The invention relates to an assembly for passing an electrical harness through a wall (28), comprising a tubular metal end piece (34) passing right through the wall (28) and housing the electrical harness, and a sleeve (38, 64, 66) made from heat-shrinkable material extending around an end part (40, 50) of the tubular end piece (34) and of the electrical harness. The assembly comprises means (42, 44, 46, 76) for extracting heat from the tubular end piece that are arranged on the side of the end piece (34) surrounded by the heat-shrinkable sleeve (38, 64, 66).. ... Safran Aircraft Engines

03/22/18 / #20180080345

Reinforced exhaust casing and manufacturing method

The invention relates to an exhaust casing (10) of a turbine engine for an aircraft which extends along an axis and which comprises a central hub (20), an annular outer shroud (30) and arms (40) which connect the central hub (20) to the outer shroud (30), at least one yoke (50) for attaching the exhaust casing (10) to the turbine engine being located on the outer shroud (30) and forming at least one lug (51) extending in a plane perpendicular to the axis and protruding toward the exterior of said outer shroud (30), characterized in that the outer shroud (30) comprises ribs (60) which form a constant excess of said outer shroud (30), which are located on either side of said at least one lug (51) of said at least one yoke (50), and which are aligned with said at least one lug (51).. . ... Safran Aircraft Engines

03/22/18 / #20180080344

A turbine ring assembly comprising a plurality of ring sectors made of ceramic matrix composite material

A turbine ring assembly includes a ring support structure and a plurality of ring sectors made of ceramic matrix composite material, each ring sector having a portion forming an annular base with an inside face defining the inside face of the turbine ring, and an outside face from which there extends a wall defining an internal housing in which a holder member made of metal material is present, the holder member being connected to the ring support structure and including a body from which elastically deformable holder elements extend inside the internal housing on either side of the body, the holder elements bearing against the wall.. . ... Safran Aircraft Engines

03/22/18 / #20180080343

A turbine ring assembly comprising a plurality of ring sectors made of ceramic matrix composite material

A turbine ring assembly includes a plurality of ring sectors made of ceramic matrix composite material, together with a ring support structure, each ring sector having a portion forming an annular base with an inner face defining the inner face of the turbine ring and an outer face from which there project at least two tab-forming portions, the ring support structure having at least two attachment tabs extending radially, the tabs of each ring sector gripping the attachment tabs of the ring support structure at least at the radially-inner ends of the attachment tabs.. . ... Safran Aircraft Engines

03/22/18 / #20180080337

Discharge flow duct of a turbine engine comprising a vbv grating with variable setting

A hub of an intermediate casing for a dual-flow turbine engine includes a discharge flow duct extending between an inner shroud and an outer shroud of the hub, the discharge flow duct leading into the secondary flow space through an outlet opening formed in the outer shroud, the outlet opening included in a discharge plane substantially tangential to the outer shroud; and discharge fins including an upstream fin and a downstream fin. An upstream acute angle between the discharge plane and the skeleton line of the upstream fin is smaller than a downstream acute angle between the discharge plane and the skeleton line of the downstream fin.. ... Safran Aircraft Engines

03/22/18 / #20180080332

Intermediate casing guide vane wheel

An ogv wheel comprising guide vanes made of polymer matrix composite material reinforced by fibers, each having a vane root and a vane tip, the vane roots being fastened on a hub of the wheel by first connection means and the vane tip being fastened on an outer shroud of the wheel by second connection means, the first connection means including a bearing plane secured to the hub and a first backing plate for securing to the hub, the vane roots being sandwiched between the bearing plane and the first backing plate, and the second connection means including a second backing plate for securing to the shroud, the vane tip being sandwiched between the shroud and the second backing plate.. . ... Safran Aircraft Engines

03/15/18 / #20180073384

Aircraft turbine engine with planetary or epicyclic gear train

Aircraft turbine engine comprising a low-pressure spool that comprises a low-pressure shaft (24), means (44) for taking off power from said low-pressure shaft, and a fan (28) that is driven by said low-pressure shaft by means of a reduction gear (32), said reduction gear comprising at least one first element (50) that is connected to said low-pressure shaft for conjoint rotation, at least one second element (56) that is connected to said fan for conjoint rotation, and at least one third element (52) that is connected to a stator casing of the turbine engine, characterised in that said at least one third element is connected to said stator casing by disengageable connection means (60), and comprising at least one member that can move from a first position in which said at least one third element is fixedly connected to said stator casing into a second position in which said at least one third element is separated from said stator casing and is free to rotate about said longitudinal axis.. . ... Safran Aircraft Engines

03/15/18 / #20180073373

Ceramic core for a multl-cavity turbine blade

A ceramic core used for fabricating a hollow turbine blade for a turbine engine by using the lost-wax casting technique and shaped to constitute the cavities of the blade as a single element, includes, in order to feed the insides of these cavities jointly with cooling air, core portions that are to form first and second lateral cavities and that are connected to a core portion that is to form at least one central cavity, firstly in the core root via at least two ceramic junctions, and secondly at various heights up the core via a plurality of other ceramic junctions of positioning that defines the thickness of the internal partitions of the blade, while also ensuring additional cooling air for predetermined critical zones of the first and second lateral cavities.. . ... Safran Aircraft Engines

03/08/18 / #20180066581

Assembly for an aircraft turbine engine comprising a fan casing equipped with an acoustic liner incorporating a fan casing stiffener

The present invention relates to an assembly (20) for an aircraft turbine engine comprising a fan casing (14) having an inner surface (14b), at least one acoustic panel (26) fastened using fastening elements (48, 54) to the inner surface of the fan casing, and at least one circumferential stiffener (40) of the fan casing (14). According to the invention, the fastening elements (48, 54) connect the fan casing (14) to the stiffener (40) incorporated with the acoustic panel (26).. ... Safran Aircraft Engines

03/08/18 / #20180066540

Intermediate casing for a turbomachine turbine

A turbine comprising an intermediate casing axially inserted between an upstream high pressure turbine casing and a downstream low pressure turbine casing and comprising an outer annular shroud from which an annular flange radially extends, characterized in that the downstream end of the high pressure turbine casing and the upstream end of the low pressure turbine casing are attached on the radial annular flange of the intermediate casing.. . ... Safran Aircraft Engines

03/08/18 / #20180066536

Compressor stage

The invention relates to the field of compressors, and specifically a compressor stage (100) comprising at least a casing (101) delimiting an air passage (2), a stator (102) comprising a plurality of guide vanes (103) arranged radially around a central axis (x) in the air passage (2), and a rotor (104) suitable for rotating about the central axis (x) relative to the stator (102) and comprising a plurality of blades (105) arranged radially around the central axis (x) in the air passage (2) downstream from the guide vanes (103). Each blade (105) of the rotor (104) extends from a blade root (105a) to a blade tip (105b) further away from the central axis (x) than the blade root (105a) and presents radial clearance (j) between the blade tip (105b) and the casing (101). ... Safran Aircraft Engines

03/08/18 / #20180066528

Turbine rotor with air separation ferrules for cooling of blade and disk coupling portions, for a turbomachine

A turbine rotor is fitted to a turbomachine and includes disks each containing a coupling portion with recesses each holding a root of a blade, and coupled to one another by a first annular ferrule attached to one of them close to the recesses, and second rotationally coupled ferrules, which are respectively coupled radially to the disks, which each consists of at least two semi-annular sectors, and which form with the associated disk a space which communicates with the recesses of the coupling portion of this latter disk. In addition, each first ferrule includes through-holes enabling air to enter this space, and then the recesses, intended to cool the coupling portion which couples its disk and the blade roots.. ... Safran Aircraft Engines

03/08/18 / #20180065730

A radial shaft device for controlling the pitch of fan blades of a turbine engine having an un-ducted fan

A device for controlling pitch of fan blades of a turbine engine including an un-ducted fan, the device including: at least one set of fan blades of adjustable pitch, the set being constrained to rotate with a rotary ring centered on a longitudinal axis and mechanically connected to a turbine rotor, each blade of the set being mounted on a blade root support that is pivotally mounted on the rotary ring; and at least one radial control shaft adjusting pitch of at least two adjacent blades of the set, the control shaft being constrained to rotate with the rotary ring and being configured to pivot about an axis of the shaft, being coupled to the blade root supports of the at least two blades of the set to adjust their pitch via a transmission system including eccentrics connected together by at least one connecting rod.. . ... Safran Aircraft Engines

03/08/18 / #20180065727

Aircraft propulsion unit comprising an unducted-fan turbine engine and an attachment pylon

A propulsion assembly for aircraft, the assembly including a turbojet having at least one unducted propulsion propeller; and an attachment pylon for attaching the turbojet to a structural element of the aircraft, the pylon being positioned on the turbojet upstream from the propeller and having an airfoil extending transversely between a leading edge and a trailing edge, the trailing edge of the airfoil of the pylon includes a cutout extending longitudinally over a fraction of the trailing edge facing at least a portion of the propeller, the cutout being configured to increase locally the distance between the trailing edge and the propeller, the cutout presenting an outline having a curved shape presenting at least two points of inflection.. . ... Safran Aircraft Engines

03/01/18 / #20180059037

Method of fabricating a reference blade for calibrating tomographic inspection, and a resulting reference blade

. . A method of fabricating a reference blade for calibrating non-destructive inspection by tomography of real blades of similar shapes and dimensions, including making a three-dimensional blank out of resin, creating housings in the thickness of the blank at predetermined locations, and introducing in each of the housings a cylinder including an artificial defect or a real defect in order to obtain the reference blade.. . ... Safran Aircraft Engines

03/01/18 / #20180058973

Method for detecting a fluid leak in a turbomachine and fluid distribution system

A method for detecting a high temperature fluid leak in a turbomachine. The turbomachine includes a source of high temperature pressurized fluid, at least one fluid distribution line suitable for distributing said high temperature fluid, and a turbomachine compartment wherein the distribution line is at least partially housed. ... Safran Aircraft Engines

03/01/18 / #20180058260

Turbine engine with an oil guiding device and method for disassembling the turbine engine

A turbine engine is provided with a longitudinal rotation axis having at least one shaft with a radial axis, in particular a pitch change system for the blades of a propeller, said shaft traversing a radial passage of a substantially cylindrical case around the longitudinal axis. The turbine engine includes an annular oil guiding device around the radial shaft. ... Safran Aircraft Engines

03/01/18 / #20180058258

Turbomachine vane provided with a structure reducing the risk of cracks

A vane (54) for a turbomachine, comprises an aerodynamic profile (61) formed from a body (70) made from a first material, and an add-on part (74) fixed to the body (70) by brazing and made from a second material. The add-on part (74) is housed in a recess (72) formed in a median part of an end part (73) of the body (70), between extreme parts (73a, 73b) of the end part, such that the add-on part (74) forms a median portion of the leading edge (62) of the aerodynamic profile (61). ... Safran Aircraft Engines

03/01/18 / #20180056406

Tooling for machining a groove of a turbine engine casing

The invention relates to tooling (24) for machining an annular groove of a turbine engine casing, wherein said tooling (24) comprises a machining tool (25), a baseplate (33), first means of positioning (28) the machining tool (25) in relation to the baseplate (33) along a first axis (y) forming a radial axis, second means of positioning (30) the machining tool (25) in relation to the baseplate (33) along a second axis (x) perpendicular to the first axis (y), wherein said second axis (x) extends along the axis of the groove and of the annular casing and third means of positioning capable of positioning the baseplate (33) axially and radially in relation to the groove of the casing.. . ... Safran Aircraft Engines

02/22/18 / #20180051881

Turbomachine combustion chamber comprising an airflow guide device of specific shape

A combustion chamber for a turbomachine that includes a chamber end wall and a plurality of air and fuel injection systems distributed circumferentially about an axis of the combustion chamber. The combustion chamber includes, associated with each injection system, a guide device for guiding an airflow including at least one wall mounted on the injection system and projecting in the upstream direction, one wall acting as an obstacle to a circumferential flow of air around the axis. ... Safran Aircraft Engines

02/22/18 / #20180051640

Method of detecting a malfunction of a valve in a turboshaft engine

A method of monitoring a valve in a turboshaft engine, said valve switching, by closing and/or opening, in response to a control instruction sent at a determined instant, said method comprising calculating a first form of a time signal from the change in a status variable of said turboshaft engine reacting to a switching of said valve, applying a signature test of the switching of the valve to a form of said signal, wherein the method further comprises defining a time interval after sending said control instruction to perform said signature test; acquiring one or more parameters other than the switching of the valve; modelling a signal of said time signal in response to a change in said other parameter(s) to calculate its change; and calculating said second form of the signal is calculated from the first form of the signal by subtracting therefrom the change in the signal calculated from the change in said other parameter(s), over said time interval following a control instruction.. . ... Safran Aircraft Engines

02/22/18 / #20180051591

Turbine ring assembly

A turbine ring assembly includes ring sectors made of ceramic matrix composite forming a ring and a ring support structure. Each sector includes an annular base with, in a radial direction, an inside face and an outside face from which extend two attachment tabs held between two radial tabs of the structure. ... Safran Aircraft Engines

02/22/18 / #20180051590

Turbine ring assembly

A turbine ring assembly includes ring sectors forming a turbine ring and a ring support structure, each ring sector having, in a section plane defined by an axial direction and a radial direction of the turbine ring, a portion forming an annular base with, in the radial direction, an inside face defining the inside face of the turbine ring and an outside face from which a first and a second attachment tab protrude, the ring support structure having a central annulus from which a first and a second radial tab protrude, between which the first and second attachment tabs of each ring sector are held. The first radial tab comprises a one-piece annular flange that is fastened in a removable manner to the central annulus of the ring support structure.. ... Safran Aircraft Engines

02/22/18 / #20180051581

Turbine ring assembly

A turbine ring assembly includes both a plurality of cmc ring sectors forming a turbine ring and a ring support structure, each ring sector having a portion forming an annular base that presents an outside face in the radial direction of the turbine ring, with first and second attachment tabs projecting therefrom in the radial direction, each attachment tab presenting an end that is free, each ring sector having third and fourth attachment tabs, each extending in the axial direction of the turbine ring between the free end of the first attachment tab and the free end of the second attachment tab. Each ring sector is fastened to the ring support structure by a bolt having a bolt head bearing against the ring support structure and a thread co-operating with tapping formed in a plate, the plate co-operating with the third and fourth attachment tabs.. ... Safran Aircraft Engines

02/15/18 / #20180043991

Pitch change system equipped with means for supplying fluid to a control means and corresponding turbine engine

. . A system is configured to change the pitch of blades of at least one turbine engine propeller provided with multiple blades. The system includes a control means acting on a connecting mechanism connected to the blades of the propeller. ... Safran Aircraft Engines

02/15/18 / #20180043990

Pitch change module for turbine engine and corresponding turbine engine

The invention relates to turbine engine module (1) including a case (9) rotating around a longitudinal axis (x) and carrying a propeller having a plurality of blades, a stationary case (15) comprising a cylindrical wall (16) extending between an inner wall (17) and an outer wall (18) of the rotating case (9), and a system (26) for changing the pitch of the blades (14) of the propeller. The wall (16) is connected downstream to a first substantially frustoconical wall (42) and upstream to a second substantially frustoconical wall (41), a first rolling bearing (19) being inserted respectively downstream directly between a radially outer face (21) of the inner wall (17) and a radially inner face (23) of the first frustoconical wall, and a second rolling bearing (19′) inserted downstream directly between the radially outer face (21) of the inner wall (17) and an inner face (43) of the second frustoconical wall (41).. ... Safran Aircraft Engines

02/15/18 / #20180043989

Pitch-change system equipped with means for adjusting blade pitch and corresponding turbine engine

A system changes the pitch of blades of at least one turbine engine propeller provided with a plurality of blades. The system includes a link mechanism connected to the propeller blades at a first interface, and a control means acting on the link mechanism and having a body movable in translation along a longitudinal axis. ... Safran Aircraft Engines

02/15/18 / #20180043423

Method for high temperature forging of a preformed metal part, and shaping equipment suitable for forging

A forging method serving to use shaping tooling suitable for high temperature forging of a preformed metal part having angular twist undercuts (49) in its final shape, the method comprising placing the preformed metal part on a movable central insert (44) of the tooling and blocking it in the tooling (40), and forming side fins of the preformed metal part (30) in their final shape by moving a movable top first die and the movable central insert in a common direction towards a stationary bottom die, the movable central insert including at least two cutaway zones (20, 52) for eliminating the angular twist undercuts and thus enabling the preformed metal part in its final shape to be dislodged in a single extraction direction.. . ... Safran Aircraft Engines

02/08/18 / #20180039957

System for assisting with the production or repair of a turbomachine

The invention relates to a method for assisting with the production or repair of a turbomachine. The method comprises a step (91) of displaying first production or repair information about the turbomachine on a portable display device. ... Safran Aircraft Engines

02/08/18 / #20180038235

Turbine engine air guide assembly with improved aerodynamic performance

A turbine engine assembly including an air flow guide assembly, including at least one guide vane and at least one structural arm, the vane and arm extending radially about an axis. The arm includes an upstream end portion having a guide vane profile and including a leading edge aligned with that of the vane; a downstream portion; and an intermediate portion including an upper surface extending between an upstream end point and a downstream end point. ... Safran Aircraft Engines

02/08/18 / #20180036914

Method for manufacturing a turbomachine blade made of composite material

A method of fabricating a turbine engine blade out of composite material including fiber reinforcement densified by a matrix, the method including using multilayer weaving to make a first fiber that has a first portion forming a blade root preform and extended by a second portion, the second portion forming a tenon preform; using multilayer weaving to make a second fiber preform, the second preform including a first portion made up of two skins defining between them an internal housing, the first portion forming an airfoil preform, and a second portion extending from an outside surface of the skins, the second portion forming a platform preform; assembling the first preform with the second preform in the non-consolidated state by engaging the second portion of the first preform in the internal housing; and co-densifying the first and second preforms as assembled together in this way to obtain a turbine engine blade.. . ... Safran Aircraft Engines

02/01/18 / #20180031365

Method of inspecting the thickness of a part of hollow shape

A thickness inspection method inspects the thickness of a part having a hollow shape by using tooling enabling a counter-shape to be molded that matches said hollow shape. The method includes putting the part into place on a support secured to the tooling, locking the part in place, and filling the hollow shape with a molding material in order to form the counter-shape. ... Safran Aircraft Engines

02/01/18 / #20180031240

Flame-holder device

A flame-holder device for a reheat channel of a turbojet, the device including an arm in the form of a trough defining a cavity and a heat shield fastened in the cavity of the arm. The flame-holder device further includes a fastener plate including a first leg integrally formed with the fastener plate and a second leg removably mounted on the plate, the arm being fastened to the first and second legs via fastener members.. ... Safran Aircraft Engines

02/01/18 / #20180030852

Aircraft comprising a turbojet engine integrated into the rear fuselage comprising a fairing allowing the ejection of blades

The invention relates to an aircraft comprising a fuselage, flight control surfaces and a turbojet engine (20) integrated into the rear of said fuselage in the extension thereof, the turbojet engine (12) comprising two gas generators (22) that supply, via a common central duct (30), a power turbine (32) comprising two counter-rotating rotors (34, 36) respectively driving two upstream (38) and downstream (40) coaxial and counter-rotating fans each comprising a ring of vanes (42, 44), the set of fans (38, 40) being integrated into a fairing (46) of the turbojet engine (20) formed at the rear of the fuselage (12), characterised in that at least said fairing (46) is axially arranged behind the flight control surfaces and comprises an upstream section (50), surrounding the upstream fan (38), configured to be radially traversed by at least one fragment (43) of a vane (42) of the upstream fan (38) in the event of the breakage of a vane (42) of said upstream fan (38) and the ejection of said at least one fragment (43).. . ... Safran Aircraft Engines

02/01/18 / #20180030849

Device for the individual adjustment of a plurality of variable-pitch radial stator vanes in a turbomachine

A device for adjusting the pitch of at least one annular row of stator vanes for a turbine engine module. The device includes a first control ring mounted to rotate freely about an axis of the turbine engine. ... Safran Aircraft Engines

02/01/18 / #20180030843

Guide assembly with optimised aerodynamic performance

The invention relates to a turbine engine air flow guide assembly including: a structural arm (30); and a guide vane (21) on the lower surface of the structural arm, comprising a leading edge (22), a trailing edge (23), and a camber line (24), said vane and arm extending radially about an axis (x-x) of the turbine engine and defining therebetween an air flow channel. The structural arm (30) comprises: an upstream end (31) having a guide vane profile (21) and comprising a leading edge (32) aligned with that of the vane; and a shoulder (35) located on the lower surface of the arm, defining a neck in the channel. ... Safran Aircraft Engines

01/25/18 / #20180023406

Intermediate case for an aircraft turbomachine made from a single casting with a lubricant duct

. . The invention relates to an intermediate case (25) for a twin spool turbomachine for an aircraft, comprising a hub (26), an outer shell (23) and outlet guide vanes (24) installed at their ends on the hub and on the outer shell, and each of at least some of the outlet guide vanes (24) performing a heat exchanger function and comprising a lubricant passage (50a, 50b) designed to be cooled by the fan flow (58) following an outer surface of the outlet guide vane. According to the invention, the case also comprises at least one lubricant duct (55) passing along a circumferential direction of the hub (26) and at least part of which is made from a single casting with the hub, the lubricant duct (55) having at least one lateral opening communicating with the lubricant passage (50a, 50b) of at least one of the vanes (24).. ... Safran Aircraft Engines

01/25/18 / #20180023405

Corrosion protection plug for filling an attachment opening, and system including said plug

A plug for preventing corrosion of an attachment opening, includes a bottom surface including an opening for receiving a head of a tightening device; a top surface; and a substantially cylindrical side wall extending between the bottom surface and the top surface. The side wall includes a first side area, wherein the plug is formed of a elastically deformable material and has a diameter at rest and resiliency enabling the plug to block, and be held in, a top portion of the attachment opening via resilient change in shape; and the top surface includes at least one first blind hole, the shape of which is such that, when the plug is subjected to an elastic deformation that causes a ridge to appear between the top surface and the first side area of the plug, the ridge is removed by elastic deformation of the blind hole.. ... Safran Aircraft Engines

01/25/18 / #20180022475

Hall effect thruster and a space vehicle including such a thruster

A hall effect thruster arranged inside a wall and including a magnetic circuit and an electric circuit including an anode, a first cathode, and a voltage source. The magnetic circuit and the electric circuit are arranged in such a manner as to generate magnetic and electric fields around the wall. ... Safran Aircraft Engines

01/18/18 / #20180017961

Method, system and computer program for learning phase of an acoustic or vibratory analysis of a machine

A method of analysis of the state of operation of a machine including a learning step supplementing a reference database with one or more thresholds for one or more indicators calculated on the basis of signals delivered by a sensor associated with the machine, the learning step including the following operations implemented by a computer processing unit; an acquisition of signals characteristic of normal operation and of abnormal operation of the machine: of each of the signals characteristic of normal operation, formation of at least one so-called deviation signal by implementing a mathematical operation having as attributes the signal characteristic of normal operation and one of the signals characteristic of normal or abnormal operation other than the signal characteristic of the normal operation; for each of the deviation signals, calculation of an indicator; determination of an indicator threshold representative of a limit between normal operation and abnormal operation of the machine.. . ... Safran Aircraft Engines

01/18/18 / #20180016940

Optimized aerodynamic profile for an arm of a structural casing of a turbine, and structural casing having such an arm

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the cartesian coordinates x,y, zadim given in table 1, in which the coordinate zadim is the quotient d/h where d is the distance of the point under consideration from a first reference plane p0 situated at the base of the nominal profile, and h is the height of said profile measured from the first reference plane to a second reference plane p1. The measurements d and h are taken radially relative to the axis of the turbine, while the x coordinate is measured in the axial direction of the turbine.. ... Safran Aircraft Engines

01/18/18 / #20180016931

Assembly for controlling variable pitch vanes in a turbine engine

An assembly, in particular for controlling variable pitch vanes in a turbine engine, comprising an actuating ring surrounding a casing of the turbine engine and connected by rods to variable pitch vanes, in addition to a driving means for rotating the actuating ring around the casing. The assembly includes a slidingly connected passive element, one end of which is connected by a sliding pivoting link on the actuating ring and a second end is connected by a ball-joint link to the casing.. ... Safran Aircraft Engines

01/18/18 / #20180016929

Nut for axially locking a bearing ring in a turbomachine

A nut for a turbine engine, in particular for axially locking a bearing race. The nut comprises a thread for screwing onto a part of the turbine engine, and a lock for ensuring the nut cannot rotate relative to the engine part. ... Safran Aircraft Engines

01/18/18 / #20180016926

Optimized aerodynamic profile for an arm of a structural casing of a turbine, and structural casing having such an arm

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the cartesian coordinates x,y, zadim given in table 1, in which the coordinate zadim is the quotient d/h where d is the distance of the point under consideration from a first reference plane p0 situated at the base of the nominal profile, and h is the height of said profile measured from the first reference plane to a second reference plane p1. The measurements d and h are taken radially relative to the axis of the turbine, while the x coordinate is measured in the axial direction of the turbine.. ... Safran Aircraft Engines

01/18/18 / #20180016925

Optimized aerodynamic profile for a turbine vane, in particular for a nozzle of the fourth stage of a turbine

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the cartesian coordinates x,y, zadim given in table 1, in which the coordinate zadim is the quotient d/h where d is the distance of the point under consideration from a first reference plane p0 situated at the base of the nominal profile, and h is the height of said profile measured from the first reference plane to a second reference plane p1. The measurements d and h are taken radially relative to the axis of the turbine, while the x coordinate is measured in the axial direction of the turbine.. ... Safran Aircraft Engines

01/18/18 / #20180016913

Optimized aerodynamic profile for a turbine blade, in particular for a rotary wheel of the third stage of a turbine

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the cartesian coordinates x,y, zadim given in table 1, in which the coordinate zadim is the quotient d/h where d is the distance of the point under consideration from a first reference plane p0 situated at the base of the nominal profile, and h is the height of said profile measured from the first reference plane to a second reference plane p1. The measurements d and h are taken radially relative to the axis of the turbine, while the x coordinate is measured in the axial direction of the turbine.. ... Safran Aircraft Engines

01/18/18 / #20180016911

Optimized aerodynamic profile for a turbine blade, in particular for a rotary wheel of the first stage of a turbine

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the cartesian coordinates x,y, zadim given in table 1, in which the coordinate zadim is the quotient d/h where d is the distance of the point under consideration from a first reference plane 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.. ... Safran Aircraft Engines

01/18/18 / #20180016910

Optimized aerodynamic profile for a turbine blade, in particular for a rotary wheel of the sixth stage of a turbine

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the cartesian coordinates x,y, zadim given in table 1, in which the coordinate zadim is the quotient d/h where d is the distance of the point under consideration from a first reference plane 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.. ... Safran Aircraft Engines

01/18/18 / #20180016909

Optimized aerodynamic profile for a turbine blade, in particular for a rotary wheel of the fifth stage of a turbine

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the cartesian coordinates x,y, zadim given in table 1, in which the coordinate zadim is the quotient d/h where d is the distance of the point under consideration from a first reference plane 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.. ... Safran Aircraft Engines

01/18/18 / #20180016908

Optimized aerodynamic profile for a turbine blade, in particular for a rotary wheel of the second stage of a turbine

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the cartesian coordinates x,y, zadim given in table 1, in which the coordinate zadim is the quotient d/h where d is the distance of the point under consideration from a first reference plane 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.. ... Safran Aircraft Engines

01/18/18 / #20180016907

Optimized aerodynamic profile for a turbine blade, in particular for a rotary wheel of the seventh stage of a turbine

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the cartesian coordinates x,y, zadim given in table 1, in which the coordinate zadim is the quotient d/h where d is the distance of the point under consideration from a first reference plane p0 situated at the base of the nominal profile, and h is the height of said profile measured from the first reference plane to a second reference plane p1. The measurements d and h are taken radially relative to the axis of the turbine, while the x coordinate is measured in the axial direction of the turbine.. ... Safran Aircraft Engines

01/18/18 / #20180016906

Optimized aerodynamic profile for a turbine vane, in particular for a nozzle of the second stage of a turbine

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the cartesian coordinates x,y, zadim given in table 1, in which the coordinate zadim is the quotient d/h where d is the distance of the point under consideration from a first reference plane p0 situated at the base of the nominal profile, and h is the height of said profile measured from the first reference plane to a second reference plane p1. The measurements d and h are taken radially relative to the axis of the turbine, while the x coordinate is measured in the axial direction of the turbine.. ... Safran Aircraft Engines

01/18/18 / #20180016905

Optimized aerodynamic profile for a turbine vane, in particular for a nozzle of the sixth stage of a turbine

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the cartesian coordinates x,y, zadim given in table 1, in which the coordinate zadim is the quotient d/h where d is the distance of the point under consideration from a first reference plane 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.. ... Safran Aircraft Engines

01/18/18 / #20180016904

Optimized aerodynamic profile for a turbine vane, in particular for a nozzle of the first stage of a turbine

When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the cartesian coordinates x,y, zadim given in table 1, in which the coordinate zadim is the quotient d/h where d is the distance of the point under consideration from a first reference plane 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.. ... Safran Aircraft Engines








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