Abstract: A tensile testing machine comprising a test specimen whose elongation is to be measured along a tensile axis, slide plates, an intermediate plate, and first and second parallel guide rods, which freely guide the slide plates axially past them.

Abstract: A pre-vaporisation tube for a turbine engine combustion chamber includes a main body defining a first inner duct configured to have an injector mounted therein. The tube includes a first end attached to a wall of the chamber, and at least two end pieces are arranged at a second end of the body and define second inner ducts. The end pieces include first portions and second portions, respectively. The second portions each include two coaxial cylindrical walls which are inner and outer coaxial cylindrical walls, respectively, and which define an annular cavity therebetween. The inner wall defines an inner passage and has first openings for fluid communication between the passage and the annular cavity.

Abstract: A modular tray, for the additive manufacturing of a part with an axis of revolution on a powder bed, includes: a shaft-mounted circular module including a shaft provided with a circular tray at one of the ends thereof, the shaft and the circular tray being concentric; and a main support module including, in one face, a cavity configured for receiving the shaft-mounted circular module, the shaft being completely inserted in the cavity. The assembly of the shaft-mounted circular module and of the main support module define a planar top surface that is at least partly formed by the circular tray of the shaft-mounted circular module.

Abstract: A combustion assembly for a gas turbine includes a flame tube and a fuel supply including a flow rate limiter supplying an injector. The flow rate limiter and the fuel injector are formed in one piece.

Abstract: A propulsion system for an aircraft includes a rotor and a nacelle fairing that extends around the rotor in relation to an axis. The nacelle fairing includes an upstream portion forming an inlet section of the nacelle fairing as well as a downstream portion, a downstream end of which forms an outlet section of the nacelle fairing. The downstream portion includes radially inner and outer walls, both of which are made of a deformable shape memory material. The wall has independently actuatable piston actuator mechanisms, each actuator mechanism being actuatable independently of the others and being designed to cooperate with means built into an inner surface of the wall to deform the wall in a radial direction in relation to the axis under the effect of a predetermined displacement command.

Abstract: In order to reduce the weight of a hybrid propulsion system for a multi-rotor rotary-wing aircraft, the system comprises at least one inverter configured to supply power in parallel to multiple electric motors intended to drive the corresponding propellers of the system.

Abstract: The present disclosure relates to an architecture of a propulsion system of a multi-engine helicopter comprising turboshaft engines connected to a power transmission gearbox, characterized in that it comprises: at least one hybrid turboshaft engine capable of operating in at least one standby mode during a stable cruise flight of the helicopter; at least two systems for controlling each hybrid turboshaft engine, each system comprising an electric machine connected to the hybrid turboshaft engine and suitable for rotating the gas generator thereof, and at least one source of electrical power for the electric machine, each reactivation system being configured such that it can drive the turboshaft engine in at least one operating mode among a plurality of predetermined modes.

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 method for manufacturing a blade with a first portion and a second portion, the method includes forming the first portion that includes forming a model of the first portion from removable material, forming a first shell mould from the model of the first portion, and forming the single-crystal or columnar first portion m a first metal alloy in the first shell mould from a single-crystal seed, and forming the second portion in which the second portion is formed on the first portion, and in which the first portion and the second portion are made from different materials, the second portion being polycrystalline and formed from a second metal alloy. The blade includes a single-crystal or columnar first portion made from a first metal alloy and a polycrystalline second portion made from the second metal alloy different from the first metal alloy.

Abstract: Rolling bearing with three contact points, wherein the inner raceway, the third contact point being located in a sector of the outer raceway delimited by a proximal end of the outer raceway on the one hand, and by the rolling plane on the other hand, characterized in that the outer ring comprises a drain provided in the outer raceway, and opening from an outer surface of the outer ring, said drain opening eccentrically on the outer raceway with respect to the rolling plane, in a sector of the outer raceway delimited by the rolling plane on the one hand, and a distal end of the outer raceway on the other hand, the drain and the third contact point being disjoint.

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).

Abstract: A hybrid propulsion chain for an aircraft, the hybrid propulsion chain comprising a plurality of propulsion rotors connected to an electrical distribution module by a plurality of electrical connections, the electrical distribution module being connected, on the one hand, to a non-propulsion turbine engine via an electrical generation system and, on the other hand, to an electric battery, each propulsion rotor comprising a stator member and at least one rotor shaft which is configured to be rotated with respect to the stator member when the stator member is electrically powered, the hybrid propulsion chain comprising an auxiliary mechanical drive system mechanically connected to the non-propulsion turbine engine, the auxiliary mechanical drive system comprising a plurality of mechanical connections for mechanically rotating at least one rotor shaft of each propulsion rotor.

Abstract: A blade of a high-pressure turbine of a turboshaft engine, the blade including an airfoil extending in a spanwise direction, terminating in an apex and having a suction wall and a pressure wall joined by a leading edge and joined by a trailing edge. The blade further includes an internal cooling circuit having only an upstream duct and a central chamber for cooling the blade by circulating air. The upstream duct and the central chamber are separately supplied with air. The upstream duct being dedicated to the cooling of the leading edge and the suction wall, and the central chamber being dedicated to the cooling of the pressure wall and the trailing edge and being provided with bridge elements each connecting the pressure wall and the suction wall.

Abstract: A turbomachine comprises a rotating spool comprising a drive shaft delivering mechanical power. The turbomachine comprises an electromagnetic pump mechanically decoupled from the drive shaft. The electromagnetic pump comprises at least one stator delimiting an annular internal volume in which is present a rotor able to drive a fluid, a plurality of magnets distributed annularly on the rotor and at least one plurality of coils distributed annularly inside the rotor. The coils of the plurality of coils face magnets along an axial direction.

Abstract: A stator vane of a turbine of a gas turbine engine, including an outer platform and an inner platform between which there extends an outer wall forming an outer skin, wherein it includes an inner wall, forming an inner skin, facing the outer wall so as to define an inter-skin cavity between the outer wall and the inner wall, the inner wall including a plurality of cooling orifices for impingement cooling of the outer wall, the outer wall and inner wall being produced by additive manufacturing.

Abstract: The invention relates to a device for the rapid reactivation of a helicopter turbine engine (6), characterised in that it comprises a pneumatic turbine (7) mechanically connected to said turbine engine (6) so as to be able to rotate it and ensure reactivation thereof; a pneumatic storage (9) connected to said pneumatic turbine (7) by means of a pneumatic circuit (10) for supplying pressurised gas to said pneumatic turbine (7); a controlled fast-opening pneumatic valve (11) arranged on the pneumatic circuit (10) between said storage (9) and said pneumatic turbine (7) and suitable for being on demand placed at least in an open position in which the gas can supply said pneumatic turbine (7), or in a closed position in which said pneumatic turbine (7) is no longer supplied with pressurised gas.

Abstract: A blade for a turbine engine includes an aerodynamic airfoil extending radially outwards, a blade tip, a blade root, and a platform connected to the root by an upper end. A maximum radial distance between the blade tip and the platform defines a maximum extent of the bearing surface of the airfoil. The blade is partially covered by a protective strip to combat oxidation and corrosion. The strip extends between a lower limit situated on at least one portion of the platform and an upper limit situated on the blade root. The blade tip is not covered by the protective strip.

Abstract: A process for manufacturing a turbomachine part coated with a thermal barrier, includes manufacturing the part by additive manufacture; electrophoretic depositing the part of a layer including particles of a ceramic material; consolidating the layer by heat treatment to obtain a ceramic coating.

Abstract: A constant volume combustion system for a turbomachine comprises a plurality of combustion chambers distributed annularly about an axis, each combustion chamber comprising an inlet orifice and an outlet orifice, a selective shut-off element movable in rotation relative to the combustion chambers, the selective shut-off element comprising a shroud facing the inlet and outlet orifices of the combustion chambers. The shroud including on a first annular portion at least one intake aperture intended to cooperate with the inlet orifice of each combustion chamber during the rotation of the selective shut-off element and on a second annular portion at least one exhaust aperture intended to cooperate with the outlet orifice of each combustion chamber during the rotation of the selective shut-off element. Each combustion chamber comprises a fuel injection device whose opening and closing are synchronized by the shut-off element.

Abstract: A method for checking the maximum available power of a turbine engine of an aircraft equipped with two turbine engines configured to operate in parallel and together to supply a necessary power to the aircraft during a flight phase includes: placing one of the turbine engines in a maximum take-off power regime, and adjusting a power supplied by the other turbine engine, such that the turbine engines continue to supply the necessary power to the aircraft during the flight phase; determining a power supplied by the turbine engine placed in the maximum take-off power regime, and processing the supplied power determined in this way, in order to deduce a piece of information relating to the maximum available power.