Real Time Touch



new TOP 200 Companies filing patents this week

new Companies with the Most Patent Filings (2010+)




Real Time Touch

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.

ARCHIVE: New 2018 2017 2016 2015 2014 2013 2012 2011 2010 2009 | Company Directory "S" | Safran Aircraft Engines-related inventors


 new patent  Safran Aircraft Engines

. . ... Safran Aircraft Engines

 new patent  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

 new patent  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

 new patent  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

Safran Aircraft Engines

. . ... Safran Aircraft Engines

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

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

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

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

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

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

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

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

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 / #the invention relates to a means (27) for controlling a system for changing the pitch of blades of a turbine engine propeller, the control means (27) comprising a fixed member (28) and a member (29) which is movable in translation along a longitudinal axis (x) relative to the fixed member (28), and an anti-rotation device (33) configured so as to prevent the rotation of the movable member (29) relative to the fixed member (28) about the axis (x).

Safran Aircraft Engines

. . ... 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 / #the present invention relates to a control system (10) of a turbopropeller (1) including:

Safran Aircraft Engines

. . ... 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 / #a turbine engine, comprising at least one drive shaft having an axis of rotation a and configured to drive a fan wheel by means of an epicyclic reduction gear train, said reduction gear having:

Safran Aircraft Engines

. . ... 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

01/18/18 / #20180016903

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

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

01/11/18 / #20180010799

Fuel injection system for aircraft turbomachine, comprising a variable section air through duct

An assembly includes an injection system and an injector for an aircraft turbomachine combustion chamber. The system includes an aerodynamic bowl including a first end widening toward the downstream end and centred on a central axis of the injection system, this also including a central body along which a film of fuel is intended to flow in the downstream direction. ... Safran Aircraft Engines

01/11/18 / #20180010613

Fan blade

A blade including at least one web and a vane having a leading edge and a trailing edge, wherein, for at least one aerofoil of the vane in the vicinity of the web, a maximum sweep angle associated with a position along a chord of the aerofoil extending from the leading edge to the trailing edge of the vane corresponding to a relative chord length of at least 50%.. . ... Safran Aircraft Engines

01/11/18 / #20180010478

System for controlling variable pitch blades for a turbine engine

The invention relates to a system for controlling variable pitch blades for a turbine engine, comprising an annular row of variable pitch blades extending about an axis (a) and each comprising a blade connected at the radially outer end thereof to a pivot (20) that defines a substantially radial axis of rotation of the blade and which is connected by a lever (34) to control means (40a, 40b) extending about said axis. The invention is characterized in that said control means include first links (40a) supported by said pivots and second links (40b) extending between said first links, said first and second links extending substantially along a same circumference of said axis and being connected to one another and to actuation means (56).. ... Safran Aircraft Engines

01/11/18 / #20180010462

Fitted platform for a turbine engine fan, and a method of fabricating it

The invention provides a fitted platform (1) for positioning between two adjacent blades of an aviation turbine engine fan, said platform comprising a flow passage wall (10) made of composite material having a central portion (16) and first and second margins (18) each extending in a longitudinal direction of said wall, each margin extending over a determined distance (d) from the central portion (16) in a transverse direction of said wall, said flow passage wall comprising fiber reinforcement densified by a matrix, the platform being characterized in that the fiber reinforcement present in the central portion (16) presents three-dimensional weaving, and in that the fiber reinforcement present in the first and second margins (18) presents two-dimensional weaving, at least in part. The invention also provides a fan module, a turbine engine, and a method of fabricating such a platform.. ... Safran Aircraft Engines

01/11/18 / #20180009522

Turboprop

A turboprop including a propeller including a blade extending in a direction, which also includes a root, a leading edge, a trailing edge, and a wing tip, and an inner air stream channel, wherein the inner air stream channel includes an inlet located at the root of the blade and an outlet leading to the trailing edge of the blade transversely directed in relation to the main elongation direction, such that an inner stream of air flowing in the inner air stream channel by entering via the inlet adjacent to the root of the blade is discharged via the outlet adjacent to the trailing edge of the blade by forming a stream of blown air that moves away from the trailing edge in a direction which is transverse to the main elongation direction and which has a component in the direction of a skeleton line of the blade at the trailing edge.. . ... Safran Aircraft Engines

01/04/18 / #the invention relates to a method for producing a part by means of a laser beam, with a nozzle (1) that sprays a metal powder towards a substrate (5). initially, the trajectory of the nozzle is defined in a pre-determined manner, and then, during the production of the part (7):

Safran Aircraft Engines

. . ... Safran Aircraft Engines

01/04/18 / #20180003616

Peeling test device

A device and a method for peeling tests, in order to test the peeling resistance of coupons each formed of a support and an adhesive. The device comprises: (i) a frame comprising rollers with parallel axes designed to maintain the coupon supported while guiding movement of the latter, (ii) a traction device comprising a vertical jack linked to an attachment element comprising a loop configured in order to cause detachment of the adhesive from the surface of the support, (iii) a device for measuring the force exerted by the jack in order to pull the loop during peeling, and (iv) a coupon, complex in shape, such as one derived from a reinforced vane. ... Safran Aircraft Engines

01/04/18 / #20180003385

Sealing device between an injection system and a fuel injection nozzle of an aircraft turbine engine

An arrangement for an aircraft turbine engine combustion chamber including an injection system and a fuel injector is provided. The injection system includes an injector nozzle guide, the inner surface of which delimits an opening for centering the nozzle, which includes an outer casing. ... Safran Aircraft Engines

01/04/18 / #20180003193

Variable pitch vane control ring bush retention foil and turbojet containing same

A flexible metal foil can be fixed on the inside part of a variable pitch vane control ring of a turbojet compressor, so as to close all through holes in which lever pins and their surrounding bushes are housed. As a result, the foil can retain the shank of the bush if the bush breaks. ... Safran Aircraft Engines

12/28/17 / #20170371327

State controller for a system driven by a command

A state monitor for monitoring the state of a system and including a calculator and a memory; the system being controlled by a command defining a plurality of modes of operation of the system, each mode of operation corresponding to applying a command of constant value; the memory containing a set of stored state matrices representing, for each mode of operation of the system, the value of the projection of its state in time; and the calculator being configured, during operation of the system, to determine estimated values for the state of the system at a given instant with the state functions and of its state at an earlier instant.. . ... Safran Aircraft Engines

12/28/17 / #20170370859

Phantom intended for use in quality control of tomographic images

A phantom for use in quality control of tomographic images, the phantom including a cylindrical plate made of a uniform material having a density d1, with two cylinders being inserted in the plate, the cylinders being made out of uniform materials having different densities d2, d3, the density of one of the cylinders being greater than the density d1 of the plate, and the density of the other cylinder being less than the density d1 of the plate, and including a first series of pairs of holes of different diameters drilled in the plate, the axes of the holes of the first series being oriented axially relative to an axis of revolution of the plate, and the holes in a given pair being spaced apart from each other by a distance equal to their diameter.. . ... Safran Aircraft Engines

12/28/17 / #20170370325

Deployable grille with fins for aircraft turbine engine thrust-reversal system

A deployable grille with fins for a thrust-reversal system for an aircraft turbine engine. The grille can adopt a rest position and a deployed, active position wherein the fins are axially spaced further apart than in the rest position. ... Safran Aircraft Engines

12/28/17 / #20170369153

Turbomachine with multi-diameter propeller

A turbomachine including at least two unducted propellers, one of which is an upstream propeller and one a downstream propeller, the upstream propeller including a plurality of blades, at least one first blade of which has a different length from that of a second blade.. . ... Safran Aircraft Engines

12/21/17 / #20170363290

Air intake ring for a turbomachine combustion chamber injection system and method of atomizing fuel in an injection system comprising said air intake ring

A system for improving fuel-air mixing inside an injection system of a turbomachine combustion chamber. An air intake ring has an annular deflection wall, or venturi, having an internal profile provided with a discontinuity inducing an increase in the radius (φ) of the internal profile downstream of the discontinuity. ... Safran Aircraft Engines

12/21/17 / #20170362950

Turbomachine blade fitted with an elastomer gasket

A turbomachine blade (10) comprising a body (11) and an elastomer gasket (12) fastened to said body (11).. . ... Safran Aircraft Engines

12/21/17 / #20170362946

Aerodynamic link in part of a turbine engine

The invention relates to part of a turbine engine comprising two arms passing through a stream of the turbine engine, wherein each arm comprises an outer surface and an aerodynamic linking device. The aerodynamic linking device comprises fairings extending between the two arms, compressible interface means interposed between the fairings and means for retaining the fairings in place by pressure in relation to the arms, which compress the interface means.. ... Safran Aircraft Engines

12/14/17 / #20170356306

Device for recovering lubrication oil ejected by centrifugal effect in a turbine engine

A device for recovering oil injected by centrifugal effect in a turbine engine, comprising a substantially circular ring around an axis, the ring comprising a first part forming a basin surrounding the axis and having an opening turned radially inward, so as to recover the oil injected radially across from the opening, and a second part forming a substantially toroidal chamber, radially open outward at a low point, so as to allow oil to escape, a passage being arranged between the basin and the chamber substantially over the entire circumference around the axis so as to cause the oil recovered by the basin to enter the chamber, wherein a rim is arranged between a radially inner bottom of the chamber and the passage, so as to contain the oil accumulating in this radially inner bottom of the chamber, and in some embodiments, at a high point.. . ... Safran Aircraft Engines

12/14/17 / #20170355117

Device and a method for fabricating a part by injection molding

An injection molding device for fabricating a part, the device including an injection mold, the mold being formed of a support and a countermold that are distinct and that define between them a mold cavity presenting a first shape; and at least one injection device for injecting a fluid material into the mold cavity; the device including deformation elements configured to modify the shape of the mold cavity into a second shape distinct from the first shape.. . ... Safran Aircraft Engines

12/07/17 / #20170352205

Validation tool for an aircraft engine monitoring system

A tool for validation of a system for monitoring at least one piece of equipment in an aircraft engine, also comprising a computer configured to: collect observation data related to the equipment, calculate a current value of at least one quality indicator on a current quantity of observation data, estimate the probability that the current value of the quality indicator reaches a predetermined reliability criterion, thus forming a probabilistic reliability law, and estimate a minimum quantity of observation data from the probabilistic reliability law, starting from which the value of the quality indicator reaches a predetermined reliability criterion with a probability exceeding a predetermined value.. . ... Safran Aircraft Engines

12/07/17 / #20170348808

Method for repairing a fan casing

A method for repairing a fan casing in which the inner surface is damaged, includes attaching a reinforcement element to the fan casing, the reinforcement element being attached to the outer surface of the fan casing opposite the damage.. . ... Safran Aircraft Engines

11/30/17 / #20170343033

System for repairing a fastener equipping a reactor wall

An attachment system intended to equip a wall (17), the system including a nut intended to receive a screw (31) of which the orientation is normal to the wall (17), the screw (31) passing through an element such as an outer panel (32) in order to attach the element to the wall (17). The attachment system comprises a socket (26) having a threaded cylindrical outer face (34) intended to be screwed into a hole (29) passing through the wall (17) and having dimensions greater than the dimensions of the fastener that the repair socket (26) replaces, the socket (26) carrying, in the central region of same, a nut (28) receiving the attachment screw (32).. ... Safran Aircraft Engines

11/30/17 / #20170342524

A titanium-based intermetallic alloy

A titanium-based intermetallic alloy includes, in atomic percent, 16% to 26% al, 18% to 28% nb, 0% to 3% of a metal m selected from mo, w, hf, and v, 0.1% to 2% of si, 0% to 2% of ta, 1% to 4% of zr, with the condition fe+ni≦400 ppm, the balance being ti, the alloy also presenting an al/nb ratio in atomic percent lying in the range 1.05 to 1.15.. . ... Safran Aircraft Engines

11/30/17 / #20170341732

An assembly of two parts, one of which is made of composite material, the parts being assembled together by a mechanical anchor element

An assembly of two parts, one of the parts being made of composite material with fiber reinforcement obtained from a fiber preform made by three-dimensional weaving and densified with a matrix, the assembly including a mechanical anchor element secured to one of the parts and inserted inside the other part.. . ... Safran Aircraft Engines

11/23/17 / #the invention relates to a turbine engine, comprising:

Safran Aircraft Engines

. . ... Safran Aircraft Engines

11/23/17 / #20170335766

Arrangements for drawing in air and trapping foreign bodies in an aircraft propulsion assembly

The invention relates to an arrangement, in a pod of an aircraft propulsion assembly, for drawing in air and trapping foreign bodies. Said arrangement includes a main air inlet duct (11) separating into, on one hand, a channel (13) for leading air to a compressor and, on the other hand, a bypass channel (12) capable of trapping foreign bodies (5) that enter said main duct (11). ... Safran Aircraft Engines

11/23/17 / #20170335709

Rotor vane with active clearance control, rotary assembly and operating method thereof

The invention relates to a motor vane for a turbine engine, comprising a body (170) locally defining a blade provided at the radially outer end with a root (33), characterised in that it also comprises at least one sealing element (39) extending beyond the radially outer end of the root and connected to an area of the root by means of a movable mechanical link (37).. . ... Safran Aircraft Engines

11/23/17 / #20170334152

Mold assembly for resin transfer molding

A mold assembly includes an injection mold configured to receive a woven preform. The preform has a principal direction with at least one edge extending substantially along the principal direction. ... Safran Aircraft Engines

11/16/17 / #20170329744

Method for producing mechanical devices comprising several assembled identical parts

The invention relates to a method for producing a plurality of mechanical devices, in which each mechanical device comprises a defined number n of identical parts to be assembled, the parts to be assembled having been produced according to a set of specifications including at least one compliance specification, the parts that meet the compliance specification being compliant parts and the parts that do not meet the compliance specification being non-compliant parts, characterized in that production is controlled in such a way that the number of mechanical devices containing a number of non-compliant parts strictly higher than a threshold value n1 are in a proportion less than or equal to a proportion p1, the proportion p1 being non-zero and strictly lower than 1. The invention also relates to a method for repairing a mechanical device that has been produced with such a production method.. ... Safran Aircraft Engines

11/16/17 / #20170329317

Method of manufacturing parts based on analysis of statistical indicators in a situation of diminished control

The invention pertains to a method of manufacturing a population of parts produced with a manufacturing device, based on the analysis of at least one statistical indicator representative of a characteristic dimension of the parts, according to which: a) a sample comprising a number n of parts is collected from among the parts produced with the manufacturing device; b) the characteristic dimension of each part of the sample is measured, and a measured value of the statistical indicator is calculated for the sample; c) a mathematical expectation of the proportion of parts which are noncompliant with respect to a specification regarding the characteristic dimension is calculated, said calculation being performed on the basis of the measured value of the statistical indicator for the sample collected and of the number n of parts of the sample; d) the mathematical expectation of the proportion of parts that are noncompliant calculated is compared with a threshold value of proportion of noncompliant parts; e) the manufacture of the parts is steered as a function of the results of the comparison of step d).. . ... Safran Aircraft Engines

11/16/17 / #20170328404

Plain self-centering bearing

The present disclosure relates to a mechanical assembly of two mechanical parts rotatable relative to each other and enabling a self-centering fluid bearing to be obtained; it comprises a first part provided with a cylindrical cavity, a second part (34) having at least one cylindrical portion engaged in the cylindrical cavity of the first part, a gap separating the cylindrical portion and the wall of the cylindrical cavity so as to allow relative movement in rotation between the first part and the second part (34), and a lubricant distribution network (37, 38) configured for feeding said gap with a fluid lubricant so as to form a fluid bearing; a first surface (34s) selected from the inside surface of the cylindrical cavity of the first part and the outside surface of the cylindrical portion of the second part is provided with at least two lubricant admission orifices (39a, 39b) that are spaced apart from each other by not less than 120° about the main axis (f) of the first surface (34s), and the first surface (34s) also presents at least one circumferential groove (40a) extending circumferentially from the vicinity of a first lubricant admission orifice (39a) over at least 100° and in the direction of rotation of the second of said surfaces relative to the first surface (34s).. . ... Safran Aircraft Engines

11/16/17 / #20170328393

A holder device for being present at the surface of a part made of composite material

A holder device for being present at the surface of a first part made of composite material in order to enable a second part to be held to the first part, the holder device including a body made of composite material including a fiber structure and a matrix present in the pores of the fiber structure; and a holder member including a base and holder tabs extending from either side of the base, the holder tabs being for defining a holding zone for holding the second part to the first part, the base of the holder device being present in the body, and the holder tabs projecting through a surface of the body.. . ... Safran Aircraft Engines

11/16/17 / #20170328379

Vane for turbomachinery, such as an aircraft turbojet or turbofan engine or an aircraft turboprop engine

A vane for turbomachinery, such as, for example, an aircraft turbojet or turbofan engine, or an aircraft turboprop engine. The vane includes: (i) a first deicing fluid flow circuit inside the vane; (ii) a second deicing fluid flow circuit inside the vane; and (iii) a selector for directing the majority of the fluid towards the first circuit when the turbomachinery is in a first operating state, and for directing the majority of the fluid towards the second circuit when the turbomachinery is in a second operating state.. ... Safran Aircraft Engines

11/16/17 / #20170328229

Turbine blade having an end cap

The invention relates to a turbine blade (1) of a turbine engine including an upper surface (11), a lower surface (12), a leading edge (13), a trailing edge (14), and a squealer tip (2) at the top thereof, wherein said squealer tip (2) is defined by a rim (2a) and comprises at least one inner rib (3) that is spaced apart from the rim (2a) defining the squealer tip (2). The invention is characterized in that said inner rib (3) is shaped to define, inside the squealer tip, a cavity (4) inside of which the passage of leak flows (5) is limited, wherein an upstream opening (131) is made in the rim (2a) at the leading edge (13), and a downstream opening (141) is made in the rim (2a) at the trailing edge (14).. ... Safran Aircraft Engines

11/16/17 / #20170328227

Turbine assembly of an aircraft turbine engine

The present invention relates to a turbine assembly (10) of a turbine engine (1), comprising at least: a first bladed rotor (12), a bladed stator (13) and a second bladed rotor (14) arranged in series, the rotors (12, 14) being mounted on a shaft (2); a sealing plate (20) extending between the stator (13) and the shaft (2) and separating a first recess (c1) arranged between the first rotor (12) and the stator (13), from a second recess (c2) arranged between the stator (13) and the second rotor (14); and pressure-reducing means (300, 31) positioned inside the first recess (c1), the assembly being characterised in that said pressure-reducing means (300, 31) comprise a plurality of substantially radial recompression fins (300) extending into the first recess (c1).. . ... Safran Aircraft Engines

11/16/17 / #20170328225

Method for friction-welding a blade to a turbomachine vane, including a surfacing process

According to the invention, a blade is friction-welded to a rotor disk of a turbomachine, the disk comprising a projecting block having an outer surface to which the blade is to be welded. To this end: a surfacing process is carried out on at least a part of the periphery of the block, in the region of said outer surface; the outer surface of the block and the surfacing are machined in order to level same; and friction-welding is then carried out between the surfaced outer surface of the block and the blade.. ... Safran Aircraft Engines

11/16/17 / #20170328222

Method for manufacturing a turbine engine blade including a tip provided with a complex well

A method for manufacturing a turbine engine blade (25) comprising a pressure side and a suction side separated from one another by an inner space for the circulation of cooling air, the blade (25) comprising a tip (s) with a closing wall (29) joining the pressure side and suction side walls in the region of this tip (s) in order to define a well shape, the closing wall including through-holes. The closing wall (29) obtained by moulding has a considerable nominal thickness with pits (36, 37) locally reducing this thickness at each through-hole in order to facilitate the removal by chemical etching of alumina rods defining the holes. ... Safran Aircraft Engines

11/16/17 / #20170326757

Composite blade comprising a platform equipped with a stiffener

A fiber preform for a turbine engine blade and also a single-piece blade suitable for being formed using such a preform, a rotor wheel, and a turbine engine including such a blade, the fiber preform being 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), a first transverse segment extending transversely from the junction between the first and second longitudinal segments and suitable for forming a first platform (23), and a first stiffener strip extending downwards from the distal edge of the first transverse portion and suitable for forming a first platform stiffener (25).. . ... Safran Aircraft Engines

11/09/17 / #20170321604

Device for de-icing a splitter nose of an aviation turbine engine

A device for de-icing a splitter nose of an aviation turbine engine, the device including a splitter nose having an outer annular wall defining the inside of the bypass stream flow channel and an inner annular wall defining an inlet of the core stream flow channel, and an inner shroud mounted at its upstream end on the inner annular wall of the splitter nose and designed to have inlet guide vanes fastened thereto, the splitter nose and the inner shroud defining an annular volume. The device includes an annular deflector positioned inside the annular volume so as to subdivide the annular volume into a first annular cavity and a second annular cavity, the second annular cavity being defined between the annular deflector and the outer annular wall of the splitter nose.. ... Safran Aircraft Engines

11/09/17 / #20170321303

A method of fabricating three-dimensional parts out of an alloy of aluminum and titanium

A method of fabricating a sintered three-dimensional part, the method including: preparing an injection composition including a binder and a powder of a titanium-based alloy including aluminum and/or chromium; injecting the composition into a cavity of a mold to obtain a blank; eliminating the binder present in the blank; a first step of sintering the powder, the powder subjected to a first pressure higher than or equal to 1 mbar to obtain a preform of the part; and a second sintering step during which a second pressure, which is lower than the first pressure, is imposed, the duration for which the second pressure is applied being selected so that the content by weight of aluminum and/or chromium in a layer having a thickness of 200 μm situated at the surface of the preform does not vary by more than 5% in relative value due to the second sintering step.. . ... Safran Aircraft Engines

11/09/17 / #20170320174

Method for producing a turbine engine part

The production method comprises the steps for producing a preform by selective melting, the preform comprising an assembly surface to be brazed to the part to be repaired and containing a brazing material, and then assembling the preform to the turbine engine part by diffusion brazing. The thermal amplitude of the main transformation peak (a1) of the brazing material used to make the preform must at least be twice that of each of the respective thermal amplitudes of the secondary transformation peaks (a2, a3) of this brazing material.. ... Safran Aircraft Engines

10/26/17 / #20170307511

Peeling test coupon

A coupon suitable for peeling tests, derived from a vane and comprising: (i) a portion of blade that comprises a frontside surface, a backside surface and a leading edge and/or trailing edge and (ii) a vane reinforcement that covers and is glued to at least a part of the frontside surface, a part of the backside surface and which extends beyond the leading and/or trailing edge. The reinforcement is split over the entire length of the leading edge and/or trailing edge such that the reinforcement is separated into two plates separate from one another and facing each other on either side of the slit beyond the leading and/or trailing edge. ... Safran Aircraft Engines

10/26/17 / #20170306773

Simplified pitch actuation system for a turbine engine propeller

Pitch actuation system for a turbine engine propeller, comprising an actuator, a movable part of which is designed to be connected to blades of the propeller so as to rotate said blades relative to blade pitch-setting axes, characterised in that the actuator is an electromechanical actuator and comprises first electrical means for controlling blade pitch, which means comprise at least two electric motors for driving a common rotor, and a transmission screw rotated by said common rotor, and in that the system further comprises a nut, through which said transmission screw passes and which is designed to cooperate with the blades so as to move them.. . ... Safran Aircraft Engines

10/12/17 / #20170292795

Panel for heat exchange and improved noise reduction for a turbomachine

A heat exchange and noise reduction panel the panel for an aircraft comprising: an external surface intended to be swept by an airflow and from which fins extend along a first and a second main predetermined direction; cavities forming helmholtz resonators, linked to the first ends of channels for the passage of air, the second ends of which communicate with said airflow, such that said channels form necks, referred to as helmholtz resonators, extending substantially along the first direction; and at least one oil flow chamber extending between said external surface and said at least one cavity, and intended to discharge the thermal energy carried by the oil, characterized in that wherein said channels are formed, at least in part, inside said fins.. . ... Safran Aircraft Engines

10/12/17 / #20170292526

Sealing system and turbopump comprising such a system

A sealing system for at least one floatingly mounted ring in a support for a rotary shaft rotatable about an axis directed along a longitudinal direction provides dynamic sealing between the shaft and the support. The system includes a coupling part interposed between the ring and the support, whereby the ring is secured to the support. ... Safran Aircraft Engines

10/12/17 / #20170291713

Turbine engline, such as for example an aircraft turbojet engine or a turboprop engine

The invention relates to a turbine engine provided with an element (3), comprising a wall (11) and at least one load-bearing member (17) extending substantially perpendicularly relative to the wall (11), with said member (17) being intended to be attached onto a mounting (18) used for the attachment thereof onto an aircraft structural part, characterized in that a thermal protection member (23) surrounds said member (17), with said thermal protection member (23) comprising a base flexibly supported on the wall (11) of the element (3), with said base matching the shape of said wall and at least one covering part which surrounds said load-bearing member.. . ... Safran Aircraft Engines

10/05/17 / #a control system (10) for controlling the pitch of a propeller (50), the system comprising a propeller shaft (60), a blade swivel device (20) having a rotary control element (22) suitable for placing the blades (52) in an angular position corresponding to a desired propeller pitch, and a transmission (12) presenting an outlet member coupled in rotation with the rotary control element (22) of the blade swivel device (20).

Safran Aircraft Engines

. . ... Safran Aircraft Engines

10/05/17 / #20170284889

Method for inspecting a connection seal between two parts

A method for checking a connection seal between two elements of a part, includes dipping the part to be checked in a penetrant having a compound suitable for reacting to light excitation; cutting the part at the connection to be checked; and checking for the presence of penetrant under a light capable of exciting the penetrant.. . ... Safran Aircraft Engines

10/05/17 / #20170284417

Output director vane for an aircraft turbine engine, with an improved lubricant cooling function using a heat conduction matrix housed in an inner duct of the vane

A guide vane for a dual flow aircraft turbine engine, the aerodynamic part of the vane including an inner duct for lubricant cooling extending in a main direction and being partly bounded by a pressure side wall and a suction side wall of the vane. A heat conduction matrix is lodged in the duct, and presents main heat transfer wings extending parallel to the direction, and laid out in staggered rows.. ... Safran Aircraft Engines

10/05/17 / #20170284408

Supply of air to an air-conditioning circuit of an aircraft cabin

Aircraft turboprop engine (110), comprising at least one low-pressure body (12) and one high-pressure body (14), the low-pressure body driving a thrust propeller by means of a gearbox (16), the turboprop engine further comprising means for supplying air to an air-conditioning circuit (36) of an aircraft cabin, said supply means comprising a load compressor (60), a rotor (61) of which is coupled to said low-pressure body by means of said gearbox, characterised in that said supply means further comprise means for controlling the rotation speed of the rotor (61) of said compressor (60), which means comprise an electric motor (70), and a mechanical differential (72) for coupling a first output shaft (74) of said electric motor to a second output shaft (76) of said gearbox and to said compressor rotor (61), such that the rotation speed (v3) of said rotor depends on the respective rotation speeds (v1, v2) of said first and second output shafts.. . ... Safran Aircraft Engines

10/05/17 / #20170284306

Aircraft turbine engine comprising a discharge device

Aircraft turbine engine (10), comprising at least one first compressor, an annular combustion chamber (70) and at least one first turbine (46), which define a first flow duct (22) for a primary flow, characterised in that it comprises, between said combustion chamber (70) and said first turbine (46), a device (55, 55′) for discharging at least part of said primary flow.. . ... Safran Aircraft Engines

10/05/17 / #20170284205

Variable pitch bladed disc

A variable-pitch bladed disc including a plurality of blades, each at a variable pitch in relation to a rotation axis of the blade and each having a root, the plurality of blades including at least one first blade and at least one second blade, a plurality of rotor connecting shafts, each having a root and a tip, each root being mounted at the tip of a corresponding rotor connecting shaft by way of a pivot so as to allow each blade to rotate about the blade rotation axis, the first blade having a first blade inclination, such that the first blade is inclined in a fixed manner with respect to the blade rotation axis of the first blade, and the second blade having a second blade inclination different from the first blade inclination.. . ... Safran Aircraft Engines








ARCHIVE: New 2018 2017 2016 2015 2014 2013 2012 2011 2010 2009



###

This listing is an abstract for educational and research purposes is only meant as a recent sample of applications filed, not a comprehensive history. Freshpatents.com is not affiliated or associated with Safran Aircraft Engines in any way and there may be associated servicemarks. This data is also published to the public by the USPTO and available for free on their website. Note that there may be alternative spellings for Safran Aircraft Engines with additional patents listed. Browse our Agent directory for other possible listings. Page by FreshPatents.com

###