Abstract: A transmission assembly for an aircraft, and a helicopter including such an assembly for controlling independently the speed of rotation of an engine and of a torque receiver. According to the invention, the assembly comprises a first inlet shaft (10a) configured to receive torque from a first engine (10), an outlet shaft (60a) configured to transmit torque to a torque receiver (60), a first transmission member (20) having at least two degrees of freedom and comprising first, second, and third movable portions, a first reversible electrical regulator machine (30), and a first reversible electrical balancing machine (40), wherein the inlet shaft (10a) is coupled to the first movable portion, the outlet shaft (60a) is coupled to the second movable portion, the first electrical regulator machine (30) is coupled to the third movable portion, and the first electrical balancing machine (40) is coupled in series with the inlet shaft or the outlet shaft (60a).

Abstract: A torque meter includes a casing and a part is configured to move in the casing in translation in a longitudinal direction under the effect of an axial thrust representative of the torque to be measured. The torque meter further includes a bearing surface connected to the casing via at least one of the ends thereof. The bearing surface extends at least partially in a plane that is substantially perpendicular to the longitudinal direction. One longitudinal end of the movable part is configured to be brought into contact with the bearing surface so that a longitudinal displacement of the movable part leads to a deformation of the bearing surface. The torque meter also has a device configured to measure the deformation of the bearing surface.

Abstract: A turboshaft engine for a helicopter, comprising a case inside which a gas generator and a turbine are accommodated, the turbine being mounted on a power shaft that extends along a longitudinal direction. The turboshaft engine further comprises means for removably mounting the power shaft into a reduction gearbox inside which at least one gear of a first reduction stage is accommodated. The means for removably mounting the power shaft include a pinion having a central bore, the shape of which is adapted to that of the power shaft in such a way that the pinion can slide over the power shaft; furthermore, the contour of the pinion is adapted to the shape of the gear of the first stage in such a way that the pinion can form a leading input pinion of the gearbox in said gear once the power shaft has been mounted in the reduction gearbox.

Abstract: An emergency start system for a turbine engine of an aircraft including at least one solid propellant gas generator, an electrically controlled ignition device, a computer connected to the ignition device, and at least one starter motor comprising a turbine for driving a shaft which is for coupling to a shaft of the turbine engine The outlet of the gases from the generator are connected to the inlet of the turbine of the starter motor.

Abstract: An annular combustion chamber (10) for a turbomachine (100), the combustion chamber presenting an axial direction (X), a radial direction, and an azimuth direction, and comprising a first annular wall (12) and a second annular wall (14), each wall delimiting at least a portion of the volume of the annular combustion chamber (10), the first and second walls (12, 14) presenting complementary fitting elements (12d, 14d), the first wall (12) presenting at least one first through hole (12f), while the second wall (14) presents at least one second through hole (14f), the combustion chamber (10) also having at least one pin (18) engaged in a pair of holes comprising a first hole (12f) and a second hole (14f), said pin (18) locking the fitting of the first and second walls (12, 14).

Abstract: The invention relates to a device for controlling an auxiliary engine (8) comprising a gas generator and a free turbine suitable for being able to be connected mechanically to the rotor (12) of a helicopter in order to supply it with thrust power, characterised in that said control device comprises a proportional-integral controller (30) having a proportional gain (Kp) and an integral gain (Ki), which are dependent on the rotation speed of said gas generator, said controller (30) being configured to receive an error signal representing a speed error of said free turbine, and to generate a signal (Sc) for correcting the drive speed of said gas generator obtained by adding a signal proportional to said error signal in accordance with said proportional gain (Kp), and an integrated signal (Si) resulting from the addition of a signal proportional to said error signal in accordance with said integral gain (Ki) and a memory signal (Sm), supplied by a feedback loop (31) of said integrated signal (Si), said memory sig

Abstract: A combustion chamber of a turbine engine, including at least one wall defining the combustion chamber and including an aperture for enabling a through-part to pass therethrough. The through-part includes, in the portion thereof located within the combustion chamber, at least one opening able to create an air film to cool the area downstream of the through-part.

Abstract: A constant-volume combustion system is for a turbine engine. The system includes a plurality of combustion chambers regularly distributed around a longitudinal axis; a toroidal manifold including a radially oriented outlet for supplying compressed air, from a compressor, to each combustion chamber; a toroidal exhaust pipe including a radially oriented inlet for collecting the combustion gases from the combustion chambers, the combustion chambers being radially positioned between the outlet of the manifold and the inlet of the exhaust pipe; and a timing device for each chamber for drawing in compressed air from the outlet of the manifold and ejecting combustion gas towards the exhaust pipe.

Abstract: The invention relates to a turbine wheel (34) comprising a plurality of vanes connected to an annular platform (24) carrying annular lips (36, 38). According to the invention, one of the upstream lip (38) and the downstream lip (36) is of a first type or of a second type, with the first type corresponding to one lip (36) having the upstream face (36a) which is concave curved and the downstream face (36b) which is convex curved and the second type corresponding to a lip (38) having the upstream (38a) and downstream (38b) faces which are substantially flat and mutually parallel.

Abstract: The invention relates to a wheel of a turbine engine comprising a plurality of radially extending vanes, one radially internal or external end of which is connected to an annular platform carrying annular lips (42, 44) extending from said platform in a direction opposite the vane in order to sealingly cooperate with a radially facing abradable material (28). According to the invention, the wheel comprises at least one lip (42, 44) having a concave curved upstream face and a convex curved downstream face or having a convex curved upstream face and a concave curved downstream face.

Abstract: An air supply plenum for an engine, the plenum being disposed upstream of an air intake of the engine, the air intake being provided in a casing of the engine. The air supply plenum includes a first lateral wall and a second lateral wall which together form a conduit in which an air flow flows as the engine functions. Each lateral wall includes a step which forms a transverse recess relative to the direction of the air flow in the air supply plenum, such that an aerodynamic separation occurs in the step when the engine is operating.

Abstract: The rapid assistance device applies to a free turbine engine of an aircraft having at least a first free turbine engine provided with a gas generator and associated with an electrical machine capable of operating both as a starter and as a generator, the first engine being capable of being put into a standby mode or into an unwanted shut-down mode, the electrical machine being powered from on on-board electrical energy power supply network. The device further includes at least one electrical energy storage member adapted to be electrically connected to the electrical machine associated with the first engine in order to provide a burst of assistance to the gas generator of that engine. The electrical energy storage member constitutes a non-rechargeable “primary” energy storage member that can be used once only.

Abstract: A propulsion unit, notably for an aircraft, including a turbomachine; a propeller that can be selectively coupled to the turbomachine; a rotary electrical machine able to drive at least the turbomachine, and selective coupling of the rotary electric machine to the propeller and/or the turbomachine that is able to couple only the propeller to the rotary electrical machine during a defined stage of operation of the propulsion unit and that is able to couple or not couple the rotary electric machine to a gas generator and/or a free turbine of the turbomachine.

Abstract: The invention relates to an architecture of a propulsion system of a multi-engine helicopter comprising turboshaft engines connected to a power transmission gearbox, characterised in that it comprises: at least one hybrid turboshaft engine (20) capable of operating in at least one standby mode during a stable cruise flight of the helicopter; at least two systems (30; 40) for controlling each hybrid turboshaft engine (20), each system (30; 40) comprising an electric machine (31; 41) connected to the hybrid turboshaft engine (20) and suitable for rotating the gas generator thereof, and at least one source (33; 43) of electrical power for said electric machine (31; 41), each reactivation system (30; 40) being configured such that it can drive said turboshaft engine (20) in at least one operating mode among a plurality of predetermined modes.

Abstract: A method of notifying an authorization to shut down completely an aircraft gas turbine engine, the method being applied after detecting that the engine has passed to an idling speed, and including a) an evaluation step of using a value of a first operating parameter of the engine to evaluate a value for a second parameter T45MG wherein a thermal behavior of a part of the engine that might be subjected to coking; b) a comparison step (E30) of comparing the value of the second parameter T45MG with a predefined threshold value T45thresh corresponding to a value of the second parameter that does not lead to coking of the part; and c) a notification step of notifying authorization to shut down completely the engine if the value of the second parameter T45MG is lower than the value of the predefined threshold T45thresh, else reiterating steps a) to c).

Abstract: A system for igniting a combustion chamber of a turboshaft engine, comprising: a plurality of start-up injectors which are suitable for injecting fuel into said chamber during a combustion-initiating phase; a circuit for supplying fuel to said start-up injectors, comprising a first sub-circuit, referred to as the primary start-up circuit, designed to supply fuel to some of said plurality of start-up injectors; a second sub-circuit, referred to as the secondary start-up circuit, designed to supply fuel to the other start-up injectors of said plurality.

Abstract: The present invention concerns a method for switching, by a local processing unit (1,2) of a flight control system of an aircraft, configured to control at least one local actuator, connected to at least one local sensor and connected via at least one link (3,4) to an opposite processing unit (2,1) configured to control at least one opposite actuator and be connected to at least one opposite sensor, said local processing unit (1,2) being further configured to be connected to backup communication means (13,14) enabling data exchanges between the local processing unit (1,2) and the opposite processing unit (2,1) in the case of failures of the links connecting same (3,4), said backup communication means comprising an array of sensors or actuators (13) and/or a secure onboard network for the avionics (14), comprising steps of: •—sending, to the opposite processing unit (2,1), acquisition data relative to the at least one local sensor and actuator data relative to the at least one local actuator, •—receiving, from

Abstract: A hydraulic torquemeter comprising a rod which is translatable in a longitudinal direction under the effect of axial thrust which is representative of the torque to be measured, and a hydraulic circuit in which a fluid flows. It also comprises a manometric capsule comprising a pressure chamber which is supplied with fluid by said hydraulic circuit, a sealed flexible membrane which is in contact with the fluid in the pressure chamber, extends in a plane which is substantially perpendicular to said rod, and is rigidly connected to one end of the rod in such a way that a longitudinal movement of the rod deforms the membrane, a drain port for the fluid in said pressure chamber, and a valve designed to adjust the pressure of the fluid in said pressure chamber according to the deformation of said membrane.

Abstract: The hybrid propulsion system for a multi-engine aircraft includes a plurality of free-turbine engines, each having a gas generator, among which at least a first engine, or hybrid engine, is suitable for operating in at least one standby mode during stabilized flight of the aircraft, while other engines of the plurality of engines operate alone during such stabilized flight. The hybrid engine is associated with first and second identical electric powertrains, each including a respective electrical machine capable of operating as a starter and as a generator, itself connected to a respective electronic power module, itself selectively connected to a specific electrical power supply network, such as an onboard network, and to a respective at least one electrical energy storage member. Each of the electric powertrains is adapted to deliver maximum power not less than half the total power needed for rapid reactivation of the hybrid engine.

Abstract: The invention relates to a removable reactivation pack for a turboshaft engine of a helicopter, comprising a gas generator equipped with a drive shaft, said turboshaft engine (6) being capable of operating in at least one standby mode during a stable flight of the helicopter, said removable pack comprising: a removable gearbox comprising a gearbox output shaft; controlled means for rotating said gearbox output shaft, referred to as reactivation means of said turboshaft engine; mechanical means for reversibly coupling said gearbox output shaft to said drive shaft of said gas generator.