An air intake system for an aircraft, which is switchable between a performance mode and a filtered mode, includes a duct forming filtered air inlet slits. The air intake system also includes interconnected gills adjacent to the filtered air inlet slits. The gills are movable between various gill positions including a closed position substantially covering the filtered air inlet slits and an open position substantially exposing the filtered air inlet slits. The air intake system also includes an actuator configured to move the gills into the closed position in the performance mode and the open position in the filtered mode.

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.

A gas turbine engine, has: a shaft rotatable about a central axis and engaged at an end thereof to a rotatable load via a spline; a reference tube circumferentially extending around the shaft, the reference tube secured at a first end to the shaft for rotation therewith and a second end free relative to the shaft for measuring a deformation of the shaft, the reference tube defining at least one tube aperture therethrough; an oil nozzle fluidly connected to a source of lubricant, the oil nozzle defining an exit flow axis intersecting the at least one tube aperture for injecting oil through the reference tube, the shaft defining at least one shaft aperture extending therethrough, the oil nozzle aligned with the spline via the at least one tube aperture and the at least one shaft aperture.

The invention is directed to an aerial vehicle with a hybrid drive unit (10) and with a rotor unit (1, 1′) wherein the hybrid drive unit (10) comprises at least a combustion engine (11), a generator (12) and a first electric motor (7) and the rotor unit (1, 1′) comprises a first rotor (1), wherein the combustion engine (11) is configured to drive the generator (12) to produce electricity, the generator (12) is coupled to the first electric motor (7) in such a way that the first electric motor (7) is feedable with electricity from the generator (12). The rotor unit (1, 1′) comprises a second rotor (1) and the hybrid drive unit (10) comprises a second electric motor (7′), wherein the generator (12) is coupled to the second electric motor (7′) in such a way that the second electric motor (7′) is feedable with electricity from the generator (12), and wherein the first rotor (1) is driven by the first electric motor (7) and the second rotor (1′) is driven by the second electric motor (7′).

A multi-engine system for an aircraft, has: a first engine having a first output shaft, a first core shaft, and a first electric machine drivingly engaged by the first output shaft or the first core shaft; a second engine having a second output shaft, a second core shaft, and a second electric machine drivingly engaged to the second core shaft; a reduction gearbox drivingly engaged by the first output shaft and by the second output shaft for driving a common load; and a transmission path between the first engine and the second engine, the transmission path being independent from the reduction gearbox and being one or more of: a torque-transfer connection between the second core shaft and the first core shaft or the first output shaft via a coupling gearbox, and an electrical connection between the generator and the electric motor to transmit electrical power to the electric motor.

Methods and systems of operating a gas turbine engine in a low-power condition are provided. In one embodiment, the method includes supplying fuel to a combustor by supplying fuel to a first fuel manifolds and a second fuel manifold of the gas turbine engine. The method also includes, while supplying fuel to the combustor by supplying fuel to the first fuel manifold: stopping supplying fuel to the second fuel manifold; and using a pump to drive gas into the second fuel manifold to flush fuel in the second fuel manifold into the combustor and hinder coking in the second fuel manifold and associated fuel nozzles.

A method for stopping an engine of a rotorcraft in overspeed, the rotorcraft comprising at least one engine, the engine comprising a gas generator and a power assembly, the power assembly comprising at least one power turbine rotated by gases originating from the gas generator, the power assembly comprising at least one power shaft rotationally secured to the power turbine, the power assembly rotating about a longitudinal axis at a speed referred to as the “speed of rotation”. The method comprises steps consisting in measuring a current value of the speed of rotation, determining a time derivative of the current value of the speed of rotation, referred to as the “current derivative $\left(\frac{dN2i}{dt}\right)”,$
and automatically stopping the engine when the current derivative $\left(\frac{dN2i}{dt}\right)$
changes sign.

An inertial particle separator (IPS) for a gas turbine engine, has: inner and outer walls extending about a central axis, an inlet defined between the inner and outer walls and oriented axially; swirling vanes extending at least radially between the inner and outer walls and circumferentially distributed around the central axis, the swirling vanes configured for inducing a circumferential component in an airflow flowing between the swirling vanes; a plenum between the inner and outer walls downstream of the swirling vanes, the plenum circumferentially extending about the central axis, the outer wall converging toward the central axis in a direction of the airflow; and a splitter radially between the inner and outer walls downstream of the plenum and circumferentially extending around the central axis, a particle outlet radially between the splitter and the outer wall, an air outlet radially between the inner wall and the splitter.

One embodiment includes a multistage infrared suppression exhaust system for an aircraft, including: a stage one including a first exhaust conduit to receive a first exhaust air flow at a first temperature-pressure product T.sub.1P.sub.1, a second exhaust conduit to receive a second exhaust air flow at a second temperature-pressure product T.sub.2P.sub.2, and a flow integrator mechanically configured to mix the first exhaust air flow with the second exhaust air flow in an integration chamber while preventing back flow into the second exhaust conduit; and a stage two including a stage two cooling airflow to cool the mixed first and second exhaust air flows.

Methods and systems of operating a gas turbine engine in a low-power condition are provided. In one embodiment, the method includes supplying fuel to a combustor by supplying fuel to a first fuel manifold and a second fuel manifold of the gas turbine engine. The method also includes, while supplying fuel to the combustor by supplying fuel to the first fuel manifold: stopping supplying fuel to the second fuel manifold; and supplying pressurized air to the second fuel manifold to flush fuel in the second fuel manifold into the combustor and hinder coking in the second fuel manifold and associated fuel nozzles.