Abstract: A turbomachine of an aircraft comprising an outer casing delimiting with an inner hub, a flow path of a gas stream in which is disposed a low-pressure turbine configured to rotationally drive a low-pressure shaft; said turbomachine comprising, in the direction of flow of the gas stream, a first propeller; and a second propeller downstream of the first propeller, the first propeller being rotationally driven by said low-pressure shaft and the second propeller being rotationally driven by an electric motor, the second propeller being further disposed at a distance between 1.5 and 4 cord lengths from the first propeller defined between the respective axes of shimming of each of the first and second propellers.

Abstract: A housing for an aircraft turbine engine includes an annular body extending around an axis A. The body includes a securing flange extending radially outwards at each of the axial ends thereof, and an abradable layer arranged inside the body. A method for manufacturing the housing includes a step b) of bonding the layer to the first surface, during which the housing is heated and compressed by a system present at least partially inside the housing. The method further includes, prior to step b), a step a) of mounting at least two tools between the flanges, each of the tools being located in a circumferential area outside the body so as to exert bearing forces in opposite axial directions on the flanges.

Abstract: An aircraft turbomachine having a centrifugal compressor, an annular combustion chamber, an annular casing extending around the chamber and delimiting an annular space (E) in which the chamber is situated, and a heat exchanger. The heat exchanger can include a first circuit supplied with exhaust gas from the turbomachine, and a second circuit connected by first and second volutes respectively to an outlet of the compressor and to the annular space. The first and second volutes can be positioned at an axial distance from one another, and the second volute can be connected to the annular space by a straightener which is situated at least in part outside the casing and which is integrated into an annular connecting pipe which connects the second volute to this casing.

Abstract: The present invention relates to a console for a seat unit comprising: an upper wall forming a table surface on which a passenger can place objects, a cavity, an articulated arm fixed to an internal face of the cavity, said articulated arm comprising a holding device for a portable electronic device provided with a screen, said articulated arm being movable between a stored position, in which the articulated arm is located inside the cavity, and a extended position in which the articulated arm is at least in part outside the cavity.

Abstract: A turbomachine having a contrarotating turbine for an aircraft, the turbomachine including a contrarotating turbine of which a first rotor is configured to rotate in a first direction of rotation and is connected to a first turbine shaft, and a second rotor configured to rotate in an opposite direction of rotation and connected to a second turbine shaft. The first rotor includes turbine wheels interleaved between turbine wheels of the second rotor, the turbomachine further including a planetary-type epicyclic mechanical reduction gear which has a sun gear driven in rotation by the second shaft, a ring gear driven in rotation by the first shaft, and a planet carrier attached to a first stator casing of the turbomachine located upstream of the contrarotating turbine with respect to a direction of flow of gas in the turbomachine. The turbomachine includes bearings for guiding the first and second shafts.

Abstract: The invention relates to an aircraft power architecture comprising a power transmission gearbox (12), located in a first compartment (30), a gas turbine (14), located in a second compartment (32) comprising a gas generator (18) and a free turbine (22) connected to the power transmission gearbox (12) by a power shaft (26) of the gas turbine (14), and an accessory gearbox (16), the gas turbine (14) being set into the main transmission gearbox (12), characterized in that the gas turbine (14) comprises a first electric machine (38), and in that the accessory gearbox (16) is placed in the first compartment (30) and comprises a second electric machine (42) configured to supply energy to the accessory equipment and to receive electrical energy transmitted via the first electric machine (38).

Abstract: A gas turbomachine combustion chamber includes bridges extending side by side to connect in one piece a radially inner wall and a radially outer wall towards a free end of the radially inner wall. The bridges, inner wall, and outer wall have an additive layer structure.

Abstract: An assembly including a nacelle element, such as a cover, a device for locking the nacelle element, a key for locking/unlocking the locking device and an indicator for indicating the position of the locking device. This assembly includes a control device which is configured to be actuated by the key when the key is withdrawn from the locking device so as to make the indicator move from a first state to a second state, the state of the indicator being representative of the locked or unlocked position of the locking device.

Abstract: Annular combustion chamber for an aircraft turbomachine, said chamber having two coaxial annular walls, an inner annular wall and an outer annular wall, respectively, which are connected upstream by an annular bottom wall of the chamber, wherein an injection device passes through an axis and comprises an air injection system and a frustoconical bowl which is flared downstream and has air passage openings, the chamber further having an annular deflector placed downstream of the bottom wall substantially parallel to the latter; and wherein the air injection system, the bottom wall, the deflector and the bowl are integrally formed.

Abstract: A system for disengaging a drive shaft from a bearing support. An assembly formed by the drive shaft and the bearing support is placed above a cylinder, which pushes on the drive shaft by passing by the central hole of the bearing support. The bearing support is driven upwards, but is blocked by a first and a second force transfer device, which allows the drive shaft to be disengaged from the bearing support.

Abstract: A turbine vane of a turbine engine such as a turboprop or a turbojet, the vane comprising a root supporting a blade, the vane including at least one air-circulation duct for cooling the air during operation, the duct including, at the vane root, an inlet portion for collecting the cooling air, the inlet portion extending away from a lower surface of the vane root opposite the blade. At least one inlet portion is provided with a helical element for rotating the cooling air in order to improve the cooling efficiency thereof.

Abstract: A method for refitting blade shrouds of a rotor wheel in an aircraft turbomachine is described. The rotor wheel has a disc bearing blades that each have an airfoil extending between a root and a shroud, the shroud of each blade having lateral edges including shapes complementary to the lateral edges of the shrouds of the adjacent blades. The lateral edges of the shrouds are interlocked in engagement with one another such that anti-wear coatings of these edges are in contact with one another in a desired interlocking engagement position, and at least one of the lateral edges of at least one of the shrouds being able to be disengaged from the lateral edge of an adjacent shroud in an undesired disengagement position. The method includes, when an undesired disengagement position is detected, a step of inserting a re-engagement device into the turbomachine.

Abstract: Method, navigation device and computer program product for assisting with the navigation of a vehicle equipped with a navigation device, comprising the following steps: acquiring (201) a priori values of kinematic variables of the navigation device, determining (202) respective current values of kinematic variables of the navigation device and a current uncertainty matrix representative of an uncertainty of the respective current values of the kinematic variables, based on respective previous values of the kinematic variables, determining (203) a correction, updating (204) the respective current values of the kinematic variables based on the correction and on the current uncertainty matrix.

Abstract: A method for manufacturing a turbomachine component made of composite material includes the step of creating a shaped fibrous preform by three-dimensional weaving of filaments to create a fibrous texture which is wound in several superposed layers onto a mandrel of a profile corresponding to that of the component that is to be manufactured, so as to obtain the fibrous preform of a shape corresponding to that of the component that is to be manufactured, the filaments being divided into warp filaments and weft filaments, the warp filaments being interwoven with the weft filaments in a three-dimensional weave. The method further includes the step of creating a stiffening zone in the fibrous preform, the stiffening zone being formed of at least one layer of warp filaments not interwoven with weft filaments that is interposed between the interwoven warp filaments and weft filaments.

Abstract: Method, navigation device and computer program product for assisting with the navigation of a vehicle equipped with a navigation device, comprising the following steps: acquiring a priori values of kinematic variables of the navigation device, determining (202) respective current values of the kinematic variables of the navigation device and a current uncertainty matrix representative of an uncertainty of the respective current values of the kinematic variables, based on respective previous values of the kinematic variables, a previous uncertainty matrix representative of an uncertainty of the respective previous values of the kinematic variables and a model of Earth's gravity experienced by the navigation device, the modeled gravity increasing with an altitude of the navigation device.

Abstract: Embodiments of the invention relate to a landing gear wherein it includes at least one static sealing point, the static sealing point comprising a gasket made of elastically deformable material held between a portion of a first part of the landing gear and a portion of a second part of the landing gear, the first and second parts of the landing gear comprising a steel substrate coated with a zinc and nickel alloy coating, including in mass percentage of the alloy between 12% and 18% and nickel, at most 0.5% of elements other than nickel and zinc, the balance being zinc, and the roughness Ra of the surface of the portions of the first and second parts of the landing gear being less than or equal to 1.6 ?m.

Abstract: A fan module for an aircraft turbine engine includes a fan having variable pitch blades and an oil transfer device configured to provide a transfer of oil between a stator and a rotor. The oil transfer device includes a stator ring having internal oil ducts, a shaft inserted into the ring and having internal oil ducts, and an annular support of the ring configured to deform elastically to allow movements of the ring in the radial direction. A plain bearing is located between the ring and the shaft, and roller bearings are mounted between the ring and the shaft on either side of the plain bearing.

Abstract: A method for making a wax model for the manufacture of a part, the part including at least one cavity, the method including: arranging at least two core elements made of a ceramic material into a wax mold, the core elements having at least partially a shape complementary to the cavity of the part to be manufactured, and injecting wax into the wax mold, around the core elements, so as to form the wax model, the two core elements being assembled prior to the injection of wax by at least one staple, the staple including a central portion from which two branches extend, each branch being fitted into an orifice formed in one of the core elements.

Abstract: Method for monitoring the braking of the wheels of an aircraft in which the braking of the wheels of the aircraft is controlled by a wheel braking controller actuating the wheel brakes of the aircraft based on both a deceleration regulation request and a thrust reverser deployment request.

Abstract: A splitter between a primary flow and a secondary flow of a dual flow turbomachine includes a single-piece structure including an outer annular wall, an inner annular wall, a radial annular wall and an inner annular baffle, defining a first cavity between the outer annular wall and the inner annular baffle, and a second cavity between the inner annular wall, the radial annular wall and the inner annular baffle.