Eurocopter

Airbus Helicopters

Bernard Certain

Louis Delorme

Francois-Xavier Filias

Jeremy Jouve

David Ayache

François-Xavier Filias

Nicolas Lamarche

Jérémy Jouve

A rotorcraft tail rotor comprising at least two blade elements, each blade element being suitable for pivoting about a collective pitch variation axis Z in order to vary the collective pitch of each blade element of the tail rotor, each blade element including at least one compensation weight comprising a projection emerging substantially perpendicularly to a main inertia axis of the blade element, the main inertia axis being parallel to a longitudinal direction of the blade element. In the invention, the tail rotor is wherein the compensation weight includes a deformable portion that is movable relative to the projection in a plane P that is parallel to the main inertia axis of the blade element.

Rotorcraft tail rotor, a rotorcraft fitted with such a tail rotor, and a method of statically and/or dynamically balancing a rotorcraft tail rotor

A method of stopping an engine of a rotorcraft in overspeed, the engine comprising a gas generator and a power assembly. When the engine is in operation, a relationship is established giving a limit derivative that varies as a function of the speed of rotation of the power assembly. The speed of rotation, referred to as the 'current speed', reached by the power assembly is measured and the time derivative of the speed of rotation is determined and referred to as the 'current derivative'. The engine is stopped automatically when the limit derivative corresponding to the current speed as determined by the relationship is less than or equal to the current derivative.

Method of stopping a rotorcraft engine in overspeed, and a system and a rotorcraft associated therewith

A non-ducted tail rotor comprising a hub and at least five blades, each of the blades extending to a free end that, in rotation, describes a circle presenting a given 'maximum' radius. The hub is rigidly secured to a drive shaft, the hub comprising a hollow circular body centered on an axis of rotation, the body presenting an 'internal' radius between the axis of rotation and an outer periphery of the body, the internal radius lying in the range 0.2 times to 0.4 times the maximum radius, each blade presenting a root extended by an airfoil element projecting from the body, extending from the root to the free end, the root being swivel-hinged to the outer periphery about a pitch axis.

Rotary wing aircraft having a non-ducted tail rotor with at least five blades

A method of stopping an engine of a rotorcraft in overspeed, the engine comprising a gas generator and a power assembly. When the engine is in operation, the engine is automatically stopped when the following three conditions are satisfied simultaneously: a torque (Tq) measured on the power assembly is below a predetermined torque threshold (Tq); and a speed of rotation referred to as a 'first speed of rotation (N)' of the gas generator is above a threshold referred to as a 'first speed threshold (S)'; and a speed of rotation referred to as a &#x201c;second speed of rotation (N)&#x201d; of the power assembly is above a threshold referred to as a &#x201c;second speed threshold (S)&#x201d;.

A buoyancy system () for an aircraft (), the buoyancy system () being provided with at least one inflatable float (). The buoyancy system () has at least one inflator () and at least one actuator () interposed between the inflator () and a float (), the actuator () having a cylinder () and a rod () partially received in the cylinder (). The rod () is secured to a piston () defining a first chamber () within the cylinder () and in fluid flow communication with the inflator (), and a second chamber () within the rod () and in fluid flow communication with the float (), and the piston () has a channel () to put the first chamber () into fluid flow communication with the second chamber (), the deployment device () having a shutter () for shutting the channel ().

Buoyancy system for an aircraft

A method of enabling an autopilot () to cause a rotorcraft () to follow a path. At least one guide mode (G) relative to at least one progression axis (P, R, V, Y) of the rotorcraft () is selected by the rotorcraft pilot. Said selection causes the selected guide mode (G) to be inhibited () and causes a path setpoint (C) to be acquired () from the pilot of the rotorcraft () operating a manual control member () for controlling the progression of the rotorcraft (). The path setpoints (C) relating to other guide modes (G) of the rotorcraft () that continue to be engaged are conserved in their initial states and the autopilot () adapts the commands relating to the progression axes (P, R, V, Y) relating to these other guide modes (G).

Automatic flight control method for a rotorcraft enabling the rotorcraft to maintain a path by tracking manual flight controls

An aircraft () having a rotary wing () and turboshaft engines () for driving said rotary wing (). The aircraft then includes two main engines () that are identical, each capable of operating at at least one specific rating associated with a main power (maxTOP, OEIcont), and a secondary engine () capable of operating at at least one specific rating by delivering secondary power (maxTOP', OEIcont&#x2032;) proportional to the corresponding main power (maxTOP, OEIcont) in application of a coefficient of proportionality (k) less than or equal to 0.5, said aircraft having a control system () for driving the rotary wing by causing each main engine () to operate continuously throughout a flight, and by using the secondary engine () as a supplement during at least one predetermined specific stage of flight.

Rotary wing aircraft having two main engines together with a less powerful secondary engine, and a corresponding method

A rotorcraft ducted rotor comprising a rotary assembly arranged in a passage to rotate about an axis of symmetry (AX). The rotary assembly () has a plurality of blades (), each fastened to a hub (), each blade () complies with a twisting relationship defining an angle of twist lying in the range zero degrees inclusive todegrees inclusive. Spanwise, each blade () comprises a first zone () followed by a second zone (), which second zone presents rearward sweep, being provided with a second trailing edge () situated downstream from a first trailing edge () of the first zone (). Each first zone () includes a root () connected to the hub () by a fastener device () having a rolling bearing () without clearance and a conical laminated abutment ().

Ducted rotor for an aircraft and a rotorcraft

A buoyancy system () includes at least one float (), a deployment assembly () having at least one deployment member () for deploying the float (), an engagement controller () which activates the deployment assembly (), and at least two immersion sensors () for issuing an order for automatic deployment of the float () to the deployment assembly (). The deployment assembly () is provided with a memory () containing a pre-established list of events. The deployment assembly () deploys each float () when a predetermined event occurs.

Method of controlling a buoyancy system for an aircraft, a buoyancy system implementing said method, and an aircraft

The present invention relates to a power plant () having a single engine () together with both a main gearbox (MGB) suitable for driving the rotary wing () of a helicopter () and a tail gearbox (TGB) suitable for driving an anti-torque rotor () of a helicopter (). The power plant () also includes a first electric motor () mechanically connected to said main gearbox (MGB) in order to be capable of driving said main gearbox (MGB), and a second electric motor () mechanically connected to said tail gearbox (TGB) in order to be capable of driving said tail gearbox (TGB).

Power plant, a helicopter including such a power plant, and a method implemented by said power plant

A hybrid engine installation () includes drive element () suitable for driving a mechanical element (BTP, BTA) in rotation. In addition, the hybrid engine installation is remarkable in that it includes at least one gas turbine () and at least one electric motor () mechanically linked to the drive element () to drive it in rotation.

Hybrid engine installation and a method of controlling such an engine installation

The present invention relates to a method and an associated device for automatically jettisoning an external load () that is reversibly suspended from a release-hook () of an aircraft fuselage () by a sling (), which method comprises:

Method of jettisoning an external load carried by an aircraft, and an associated device

Apparatus for regulating the flow rate of fuel to a turboshaft engine in acceleration or in deceleration, the engine having a free turbine and a core engine, the apparatus comprising sensors transmitting information to regulator means, said information relating: to a first speed of rotation NTL of said free turbine; to a second speed of rotation Ng of said engine's gas generator; to an internal temperature Tof the gas at the inlet to the free turbine; to the external pressure; and to the external temperature. The apparatus further comprises control means activated by said regulator means to actuate a fuel metering system of the engine. In addition, the regulator means evaluates the flow rate of fuel to be supplied to the engine in acceleration or in deceleration on the basis of at least one optimum regulation relationship, said regulation relationship determining a main modulated flow rate as a function of a modulated speed of rotation.

Method for regulating the flow rate of fuel to a turboshaft engine in acceleration or in deceleration

A method and a device for obtaining a predictive vertical speed of a rotorcraft, the device constituting a predictive vertical speed indicator () that includes at least:

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Aix en Provence, France

Bernard Certain