Abstract: The present invention relates to a device (1) for controlling an electric aircraft-propelling assembly, said propelling assembly comprising a propeller (3) and at least one electric motor (4) that is powered by an electric supply voltage and that delivers a torque and a rotation speed to drive the propeller (3). The control device (1) comprises at least a unit (11) for measuring an electric supply voltage, and a control unit (12) suitable for making a signal delivered to the electric motor vary as a function of said electric supply voltage, with a view to making the rotation speed of the propeller vary.

Abstract: A fuel supply circuit of an aircraft engine includes a centrifugal pump mechanically coupled with an engine shaft delivering mechanical power. The circuit further includes at least one electromagnetic pump including at least one stator delimiting an annular internal volume in which is present a rotor able to drive a fluid, a plurality of magnets annularly distributed on the rotor and at least a plurality of coils annularly distributed inside the stator face-to-face with the magnets. The rotor is connected to the engine shaft by a one-way clutching element.

Abstract: A turbine for a turbomachine of longitudinal axis including an alternating arrangement of annular rows of movable blades and of fixed blades and a radially inner annular cavity formed radially inside the movable and fixed blades, and a supply circuit for supplying cooling air to the inner annular cavity, the downstream end of the supply circuit comprising an inner annular row of orifices and an outer annular row of orifices opening into the radially inner annular cavity. The turbine may also include means for controlling the flow rate of supply air to the orifices of the inner and outer annular rows of orifices.

Abstract: A propulsion assembly for a hybrid aircraft including a first and a second engine each having a gas generator and a free turbine, a main rotor coupled to the free turbine of the first and second engines, the first engine including a first electric machine and a second electric machine of lower power than the first electric machine, one of the first or of the second electric machine being able to be coupled to the gas generator and to set the gas generator in rotation during a start phase of the engine, and being further able to be coupled to the free turbine in order to generate electrical energy after the start phase, the other of the first or of the second electric machine being coupled to the gas generator only.

Abstract: A method for manufacturing a flame tube for a turbomachine, the flame tube extending about an axis and comprising an annular radially internal wall and an annular radially external wall, connected to each other by an end wall or head wall, the internal wall, the external wall and the end wall defining an internal volume, at least one part of the said end wall forming a double wall comprising a first part and a second part connected to each other and spaced apart from each other so as to delimit a flow channel for a flow of cooling air opening into the said internal volume, the said flow channel comprising at least one air-inlet opening, the first and second parts of the double wall being connected by connecting zones or bridges extending into the flow channel for the cooling-air flow, the flame tube being manufactured by additive manufacturing.

Abstract: A hydraulic torque measurement device for an aircraft engine unit includes a hermetically sealed enclosure having a rigid wall and a deformable wall that define an internal volume. The device further includes at least one pressure measuring element capable of measuring the pressure prevailing in the internal volume. The hydraulic torque measuring device allows a service state in which the internal volume is filled exclusively with an incompressible liquid, and in which the deformable wall is configured to be subjected to a pressure force. In some embodiments, the hydraulic torque measurement device is suitable for use with a gearbox.

Abstract: A system for pumping and metering a fluid for a turbine engine, which system includes at least one pump for the fluid and an electronic computer configured to determine the flow rate of the fluid to be delivered to the turbine engine, the pumping and metering system being wherein it further includes a first electric motor and a second electric motor, which are each configured to drive the at least one pump, and in that the electronic computer includes a first control loop for controlling at least the first electric motor and a second control loop for controlling at least the second electric motor.

Abstract: A rotor wheel for an aircraft turbine engine has a disc with a main axis and cells at its outer periphery. The cells extend along the axis, and each has a bottom and two side flanks. Vanes are mounted in the cells of the disc, each vane including a blade connected by a platform to a root mounted in one of the cells. Each root includes, at its radially inner end, a lobe with a first axial end having a circumferential notch and a second axial end having a radially inward facing stop configured to axially bear on a first face of the disc. An annular ring engages the notches of the vanes and is axially clamped against a second face of the disc. Each lobe has a radially inward facing projecting bulb configured to radially bear on the surface of the bottom of the corresponding cell.

Abstract: Aircraft turbomachine including a centrifugal compressor, a combustion chamber, the combustion chamber being supplied by the compressor via a diffuser and via a straightener, and a heat exchanger, the exchanger including a first circuit, supplied with exhaust gas from the turbomachine, and a second circuit, which are connected by volutes on the one hand to an outlet of the diffuser and on the other hand to an inlet of the straightener, the volutes having reversed winding directions such that their connection ports to the exchanger are independent of one another and are substantially diametrically opposed, and such that the minimum cross section of each duct is situated at a larger cross section of the other duct.

Abstract: The invention relates to a propulsion system (1, 1) for an aircraft, comprising a rotor (2) and a nacelle failing (3) that extends around said rotor in relation to an axis (X) and includes an upstream portion (10) forming an inlet section (BA) of the nacelle fairing (3) as well as a downstream portion (20), a downstream end (21) of which forms an outlet section (BF) of the nacelle fairing (3); and characterized in that the downstream portion (20) has a radially inner wall (20a) and a radially outer wall (20b), both of which are made of a deformable shape memory material, and in that the downstream end (21) includes pneumatic or hydraulic actuators (23, 23?) extending in different consecutive angular sectors about said axis (X), each actuator being independently actuatable and being configured to deform, in a direction that extends radially in relation to said axis (X) and is centered angularly in relation to its angular sector, under the effect of a predetermined control pressure.

Abstract: A liquid fuel supply system for an aircraft engine includes a fuel tank, a suction duct connected to the fuel tank and located higher than the fuel tank, an electric pump, a supply pump that is mechanically driven by an accessory gear box and an outlet of the supply pump being connected to a fuel supply circuit of the engine, and an air expulsion drain. The electric pump is in communication with the suction duct independently of the supply pump, the electric pump is in communication with the air expulsion drain, and the supply pump is in communication with the suction duct independently of the electric pump.

Abstract: A hybrid propulsion unit for an aircraft with multi-rotor rotary wings includes an electrical generator driven by an internal combustion engine, a rectifier configured to convert an AC current sent by the electrical generator into DC current, a DC-AC converter, an electrical network connecting the rectifier to the converter and including a high-voltage DC current bus, electric motors powered by propeller converters coupled to the electric motors, electrical energy storage connected to the electrical network, the electrical storage including at least one primary storage element and at least one secondary storage element.

Abstract: An aircraft turbomachine has a gas generator that includes, along a longitudinal axis (X), at least one compressor, a combustion chamber, and at least one turbine. The turbomachine further includes at least one duct for supplying liquid to at least one member chosen from an oil jet and a fuel injector. The duct (20) has rectilinear portions and bent portions and includes at least one region in which the liquid is liable to coke. The turbomachine also includes at least one turbulence element in the at least one region in the duct.

Abstract: The present invention relates to a metal powder for a powder bed additive manufacturing process, the metal powder comprising a nickel-based alloy comprising at least 0.05% carbon, at least 14.25% cobalt, at least 14% chromium, at least 4% aluminium, at least 3.9% molybdenum, at least 3% titanium, at most 0.5% iron, at least 0.012% boron, at most 0.060% zirconium, at most 0.150% manganese, at most 0.2% silicon, at most 0.1% copper, at most 0.5 ppm bismuth, at most 5 ppm silver, at most 5 ppm lead, at most 25 ppm sulphur, at most 200 ppm oxygen, and at most 60 ppm nitrogen.

Abstract: Propulsion system for a helicopter comprising a main engine, a main rotor, a main gearbox including an output mechanically connected to the main rotor, a reduction gearbox mechanically coupled between the main engine and a first input of the main gearbox, and an assistance device. The assistance device comprises a first electric machine mechanically coupled to the reduction gearbox and configured to operate as an electric generator to take off energy produced by the main engine, and a second electric machine mechanically coupled to a second input of the main gearbox, the second electric machine being supplied with electrical power by the first electric machine and configured to operate as an electric motor to deliver additional mechanical power to the main gearbox.

Abstract: A propulsion unit having a propeller for an aircraft including a nacelle; a propeller mounted in the nacelle so as to be capable of rotating about a longitudinal axis of rotation, the propeller having blades mounted by a root so as to be capable of pivoting between a deployed position, in which they extend radially relative to the axis of rotation, and a folded position, in which they are longitudinally received against the nacelle; drive means that rotate the propeller; indexing means for stopping the propeller in at least one indexed angular position (? i) relative to the nacelle; the propulsion unit wherein the indexing means consist of a stepping electric motor including a rotor that is coupled to the propeller.

Abstract: A process for three-dimensional printing of a workpiece (10), comprising a succession of steps for producing a layer (18) of the workpiece (10) by means of a nozzle (28) which has an output cross-section along a path (32) of the nozzle (28), wherein, during at least one step for producing a layer (18), use is made of at least one nozzle (28), an output portion (42) of which has an output cross-section which is variable between a first maximum cross-section and a second minimum cross-section, and wherein the cross-section of at least one part of an output portion (42) of the nozzle is varied along at least one portion (32a, 32b) of the path (32).

Abstract: A device for controlling an airflow guiding system comprising: at least one vane movable in rotation about an axis of rotation between a first angle and a second angle, an actuator comprising a body inside which a piston is mounted in translation secured to an actuation rod, and a control rod comprising a downstream end connected to the axis of the vane, the actuator being configured to drive the control rod in movement between a first end position and a second end position of a nominal operating range and, in the event of breakdown of said device, to perform an over-stroke of the actuation rod into a safety position. The device comprises a drive mechanism linking an upstream end of the actuation rod to an upstream end of the control rod, opposite to the downstream end.

Abstract: Disclosed is a turbomachine, comprising a gas generator (13) equipped with a first shaft (18), at least one reversible electrical machine (11), a free turbine (12) provided with a second shaft (17) and caused to rotate by a gas flow generated by the gas generator (13), an accessory gear box (14) and at least one accessory (15, 16). Said at least one electrical machine (11) is mechanically coupled to said second mechanical shaft (17) via the accessory gear box (14) during all phases of operation of the turbomachine (10), the accessory gear box (14) is coupled to the at least one accessory (15, 16) and the turbomachine (10) further comprises a single mechanical coupling means (20) for mechanically coupling said first mechanical shaft (18) to the accessory gear box (14) in a first configuration and mechanically uncoupling said first mechanical shaft (18) from the accessory gear box (14) in a second configuration.

Abstract: A system (11) for balancing at least one parameter to be balanced of an electric motor of a propulsion system (1), in particular of an aircraft, includes at least two electric motors (3, 4) and a propulsion member (2) driven in rotation by said electric motors. The balancing system is configured to calculate a correction of the speed setpoint (Corr_Cons_VI, Corr_Cons_V2) as a function of a correction factor (F1, F2) of the speed setpoint depending on a parameter (P1, P2) of the associated electric motor that is intended to be balanced and on a speed setpoint (Cons_VH) of the propulsion member (2).