Abstract: A turbojet engine, having a primary duct and a secondary duct surrounding the primary duct, a row of guide vanes extending in the secondary duct downstream of a fan, at least one conduit extending in the secondary duct downstream of the guide vanes and housed in a profiled casing, and at least one air sampling scoop located in the region of the casing and equipped with a control valve. Where at least one guide vane located in the vicinity of the casing has a variable geometry zone, the geometry of which can be adjusted during operation.

Abstract: An intermediate casing includes an inner wall having an outer surface extending along a longitudinal axis, the outer surface having a first profile in a first plane, an outer wall having an inner surface a second profile in the first plane, and first through fourth arms each extending radially. The outer and inner surfaces and the first and second arms delimit a first cavity, the first cavity having a first area and the outer and inner surfaces being separated by a first distance. The outer and inner surfaces and the third and fourth arms delimit a second cavity, the second cavity having a second area and the outer and inner surfaces being separated by a second distance. The first and second profiles are such that the first area is substantially identical to the second area and the first distance is different from the second distance.

Abstract: A method including at least two turbomachines disposed on either side of a main axis of the aircraft, the turbomachines each including a low-pressure shaft and a set of at least three sensors for controlling the operation of the turbomachine. The method including the following steps: for each turbomachine, acquiring the data of the set of at least three sensors, for each turbomachine, normalizing the first pressure with respect to the speed of the low-pressure shaft and with respect to the second pressure, comparing the normalized first pressure of a turbomachine with the normalized first pressure of the at least one other turbomachine of the aircraft, detecting the presence of a crosswind based on the result of the comparison.

Abstract: A module including a device for changing the pitch of the vanes about the pitch axes thereof, the device including a hydraulic actuator capable of rotating about the longitudinal axis and configured to rotate the vanes about their pitch axes, a pump for supplying fluid to the hydraulic actuator, including a transmission shaft and a placed around the transmission shaft and rotating as one with the drive shaft, and an electric machine including a rotor rotating as one with the transmission shaft and a stator rotationally fixed relative to the longitudinal axis. The stator can be connected to a power supply cable passing through a passage provided in the third element of the speed reducing gear.

Abstract: Turbine shroud assembly comprising sections (10) made from CMC and forming a shroud (1) and a support structure (3), each section having a base (12) with a radially internal face (12a) and a radially external face (12b), from which there extend in a projecting manner an upstream attachment lug (14) and a downstream attachment lug (16), the support structure comprising a collar (31), from which there radially extend in a projecting manner towards the shroud an upstream radial flange (32) and a downstream radial flange (36), by which the lugs of each section of the shroud are retained, the shroud (1) being retained by axial pins (119, 120) which cooperate, on the one hand, with the upstream radial flange, via first and second annular end plates (33, 34), and directly with the downstream radial flange and, on the other hand, with the upstream and downstream attachment lugs, respectively.

Abstract: A distributor for a turbine of a turbomachine. The distributor includes a radially inner platform, a blade and a root which is coated with a track made of an abradable material. According to the invention, the distributor includes a leakage gas reintroduction duct. The duct includes an inlet through the track made of an abradable material, an outlet opening through a downstream surface of the root, and a duct intermediate portion which extends from the inlet up to the outlet. The duct intermediate portion is orientated with a tangential component.

Abstract: Movable ring assembly for a turbomachine rotor, the movable ring including an outer peripheral toothing, each tooth of the outer peripheral toothing having an inner face intended to be positioned facing a blade root of the rotor, and an outer face opposite to the inner face and intended to be positioned facing an axial retaining hook arranged on the blade root. Furthermore, the movable ring assembly includes at least one protective element arranged on the outer face of at least one tooth of the outer peripheral toothing.

Abstract: A method for producing casing for an aircraft turbine engine includes a first step of manufacturing an annular shell having an axis A made of a composite material in a mold that includes an annular cavity that receives the preform. A resin is injected into the cavity and subsequently polymerized. During a second step, at least one axial abutment means is arranged inside the shell, and during a third step, the annular cartridge made of an abradable material is arranged inside the annular shell in contact with the axial abutment means. A fourth step includes bonding the annular cartridge to the first inner annular surface. The first and second steps are simultaneous, and the at least one axial abutment means is molded with the preform in the mold.

Abstract: A double-flow turbomachine has an exhaust casing through which a hot primary flow passes and surrounded by a cold secondary flow, and a cone carried by the exhaust casing. The exhaust casing has a hollow radial arm passing through the primary flow to convey air into the cone in order to form a cooling flow for components located in the cone. The cone has an outer wall and an inner wall which runs alongside the outer wall and is radially spaced therefrom to delimit a thermal insulation space through which the cooling flow passes from upstream to downstream before it is discharged through an outlet opening at the end of the conc. It has an inner cover surrounding a component to channel the cooling flow so that it runs along the component to cool it before passing through the thermal insulation space to be discharged there.

Abstract: A turbomachine part of an annular assembly includes a structural body and a connecting edge integral with the structural body. Each connecting edge includes at least one groove intended to receive a sealing tab. Each connecting edge is made of composite material including a fibrous reinforcement consisting of randomly-oriented short fibres, the reinforcement being densified by a ceramic matrix.

Abstract: The invention proposes an aeronautical turbine engine assembly comprising an upstream casing (55) to which guide blading (48a) is fastened, and a downstream casing (58) to which a sealing element (62) provided with an abradable material for rotor blading is fastened. This assembly further comprises a shroud ring (66) placed between the upstream casing and the downstream casing and fastening means (68) for detachably fastening the shroud ring. In order to be fastened to the upstream casing, the guide blading (48a) of the turbine engine is mounted on a downstream hook (480b) of the upstream casing, without being hooked onto the shroud ring (66), and the downstream casing (58) has an upstream hook with which the sealing element (62) is engaged in order to be fastened to the downstream casing, or the shroud ring has an upstream hook on which the sealing element (62) is mounted so as to be fastened to the downstream casing.

Abstract: A method for producing an aerodynamic element for an aircraft, including a wall which is covered, at least partially, by a printed film including a plurality of ribs and/or grooves, the method including: providing a raw film made of a deformable material which is devoid of grooves and ribs; providing a pressure plate which includes a face provided with ribs and/or grooves complementary to the ribs and/or grooves of the printed film; placing said raw film on the wall of the element; positioning the pressure plate on the raw film, the printed side of the pressure plate facing the raw film; and a forming step during which the raw film is bonded with the wall of the component and during which the raw film is shaped by cooperation with the pressure plate to obtain the printed film including the ribs and/or the grooves.

Abstract: A method for monitoring a status of a reduction gear of a gas turbine includes the steps of: obtaining measurements of parameters accomplished during an operating phase of the gas turbine, these parameters including temperatures of a lubricating oil of the reduction gear at the inlet and at the outlet of the reduction gear, a parameter representing a speed of the gas turbine, as well as at least one context parameter; selecting measurements; normalizing temperatures of the lubricating oil using the measurements of the context parameter; evaluating a thermal efficiency of the reduction gear by using a physical model defining the thermal efficiency based on a difference between the temperature of the lubricating oil; and determining a status of the reduction gear depending on a step of comparing the evaluated thermal efficiency with respect to a reference signature.

Abstract: A module including a device for changing the pitch of the vanes about the pitch axes thereof, the device including a hydraulic actuator capable of rotating about the longitudinal axis and configured to rotate the vanes about their pitch axes, a pump for supplying fluid to the hydraulic actuator, including a transmission shaft and a placed around the transmission shaft and rotating as one with the drive shaft, and an electric machine including a rotor rotating as one with the transmission shaft and a stator rotationally fixed relative to the longitudinal axis. The stator can be connected to a power supply cable passing through a passage provided in the third element of the speed reducing gear.

Abstract: The present invention relates to an assembly (12) for a turbomachine, comprising: —an outer casing (15) made of half-shells; —an inner casing (14); —a plurality of guide vanes (13) mounted between the outer casing (15) and the inner casing (14), —a structural casing comprising a flange (23) which is radially inside the inner casing (14); and—a system (24) for blocking the inner casing (14) on the structural casing, comprising a generally annular tab (25) fixed to one of the inner casing (14) and the structural casing and a complementary chute (26) fixed to the other of the inner casing (14) and the structural casing, the chute (26) being configured to receive the tab (25) and to block it along the axis (X) with respect to the structural casing.

Abstract: System for controlling a device provided with at least one equipment item, the control system comprising at least one sensor capable of measuring operating quantities of the device, and two specific channels for measuring a redundant sensor for each equipment item, at least one embedded control means configured to determine a command intended for at least one equipment item as a function of the at least one measurement of the operating quantity of the device and a selected measurement of the controlled equipment item, the control system further comprising a calculating means configured to determine the selected measurement of the at least one controlled equipment item, by means of a Kalman filter observer, as a function of the at least one measurement of the operating quantity of the device and the measurements of the two specific channels for measuring a redundant sensor for the controlled equipment item.

Abstract: A method for manufacturing an aeronautical part utilizes injection molding. After injecting the prepared mixtures to obtain two green blanks, an assembly area of at least one of these two blanks is heated. The blanks are assembled and then debinding is carried out. A sintering treatment is then carried out.

Abstract: A ventilation ring for a bearing support member of an aircraft turbine engine includes two or more tubular walls and one or more spacer walls. The two or more tubular walls extend opposite each other, and one or more of the tubular walls includes a main portion and a base, the base having a thickness greater than a thickness of the main portion. The one or more spacer walls connect the two or more tubular walls. The one or more spacer walls have one or more apertures extending into the base and a rim around the aperture. An entirety of the rim has a thickness greater than a thickness of the spacer wall at a distance from the rim.

Abstract: A rotor of a turbine extending around a longitudinal axis includes a rotor disc bearing at its periphery one or more rotor blades, each blade including an airfoil provided at its radially outer end with a platform equipped with an upstream lip and a downstream lip, the platform having a front lateral edge and a back lateral edge with respect to the normal direction of rotation of the rotor about the longitudinal axis, and a suction face portion which extends from the front lateral edge to the suction face of the airfoil. The upstream lip and the downstream lip each includes sawtooth-shaped cut-outs such as to form a front tooth and a back tooth which are contiguous, in that the back tooth has a cutting edge and in that the front tooth of each lip is disposed on the suction face portion of the platform.

Abstract: A method for controlling a turbojet engine thrust reverser during an aborted aircraft takeoff, the thrust reverser including doors movable between a stowed position, an overstowed position and a deployed position; door actuators to move the doors between the stowed, overstowed and deployed positions; a device for locking the doors in the stowed position, moveable between a locking position and an unlocking position; and a lock actuator to move the locking device between the locking and unlocking positions. The method includes decreasing the engine speed of the turbojet engine by following a setpoint value below a first engine speed threshold value at which the aerodynamic forces being exerted on the doors are equal to the forces developed by the door actuators; controlling the door actuators to bring the doors into the overstowed position; controlling the lock actuator to bring the locking device into the unlocking position.