Abstract: An assistance method for assisting a pilot of a single-engined rotary-wing aircraft during a flight phase in autorotation, the aircraft including a hybrid power plant provided with a main engine, with an electric machine, with a main gearbox, and with an electrical energy storage device. The aircraft also includes a main rotor driven by the hybrid power plant. In the method, during a flight, operation of the main engine is monitored in order to detect a failure, in particular by means of a drop in power on the main rotor, then, when a failure of the main engine is detected, the electric machine is controlled to deliver auxiliary power We to the main rotor, making it possible to assist a pilot of the aircraft in performing the flight phase in autorotation following the failure.

Abstract: An assistance method for assisting a pilot of a single-engined rotary-wing aircraft during a flight phase in autorotation, the aircraft including a hybrid power plant provided with a main engine, with an electric machine, with a main gearbox, and with an electrical energy storage device. The aircraft also includes a main rotor driven by the hybrid power plant. In the method, during a flight, operation of the main engine is monitored in order to detect a failure, in particular by means of a drop in power on the main rotor, then, when a failure of the main engine is detected, the electric machine is controlled to deliver auxiliary power We to the main rotor, making it possible to assist a pilot of the aircraft in performing the flight phase in autorotation following the failure.

Abstract: A damper device (10) for a rotor (2) having a rotary ring (11) and a plurality of levers (12), each lever (12) extending from a first end (12?) connected to the ring (11) by a first mechanical connection (20) to a second end (12?) suitable for being hinged to a blade (4). The damper device (10) includes non-rotary damper means (30) secured to non-rotary fastener means (40) and to a non-rotary member (51) of an interface (50), the interface (50) being interposed between the non-rotary damper means (30) and the ring (11) to convert a rotary movement of a blade (4) transmitted to a lever (12) into a movement in translation acting on the damper means (30).

Abstract: A damper device (10) for a rotor (2) having a rotary ring (11) and a plurality of levers (12), each lever (12) extending from a first end (12?) connected to said ring (11) by a first mechanical connection (20) to a second end (12?) suitable for being hinged to a blade (4). The damper device (10) includes non-rotary damper means (30) secured to non-rotary fastener means (40) and to a non-rotary member (51) of an interface (50), said interface (50) being interposed between said non-rotary damper means (30) and said ring (11) to convert a rotary movement of a blade (4) transmitted to a lever (12) into a movement in translation acting on said damper means (30).

Abstract: A coaxial rotary coupling (11) includes a hollow inner tube (4) and a hollow outer tube (7), with inner and outer splines (12 and 13) in mutual engagement. The coupling (11) further includes: approach elements for longitudinally approaching the inner and outer tubes (4, 7) together with elements presenting a wedge arrangement for causing the inner and outer tubes (4, 7) to turn (23) relative to each other in opposite directions so as to limit tangential assembly slack; a soleplate (16) of the driven tube (7); and a flange (17); the driving tubes (4) having complementary splines (12, 13, 18) together with dogs including force-transformation ramps (31) that are distributed peripherally within the coupling (11).

Abstract: A coaxial rotary coupling (11) includes a hollow inner tube (4) and a hollow outer tube (7), with inner and outer splines (12 and 13) in mutual engagement. The coupling (11) further includes: approach elements for longitudinally approaching the inner and outer tubes (4, 7) together with elements presenting a wedge arrangement for causing the inner and outer tubes (4, 7) to turn (23) relative to each other in opposite directions so as to limit tangential assembly slack; a soleplate (16) of the driven tube (7); and a flange (17); the driving tubes (4) having complementary splines (12, 13, 18) together with dogs including force-transformation ramps (31) that are distributed peripherally within the coupling (11).

Abstract: The present invention relates to a method of controlling a drag damper for a helicopter blade under all operating configurations of the helicopter, the method consisting in: a) measuring the elongation of the damper; b) amplifying the measured signal and filtering out noise; c) isolating the natural response of the blade corresponding to the drag movement of the blade from the forced response imposed by a rotor driving the blades; and d) analyzing in real time the natural response and adjusting the damping as a function of the analysis.

Abstract: The present invention relates to a method of controlling a drag damper for a helicopter blade under all operating configurations of the helicopter, the method consisting in: a) measuring the elongation of the damper; b) amplifying the measured signal and filtering out noise; c) isolating the natural response of the blade corresponding to the drag movement of the blade from the forced response imposed by a rotor driving the blades; and d) analyzing in real time the natural response and adjusting the damping as a function of the analysis.

Abstract: A pitch control device for a blade of a tilting rotor of a convertible aircraft. The pitch control device is external to a rotor hub which is driven in rotation by a mast and connected to the latter by a mechanism for tilting the hub as a whole, each blade is connected to the hub by a coupling for retaining and hingeing the blade in pitch about its pitch change axis. The device includes a pitch change lever, integral in rotation with its blade root about its axis, a pitch control rod, a bellcrank pivoting about a pin on a drive plate integral in rotation with the mast, and to which the pitch control rod is hinged, the bellcrank being also hinged to a link rod connected to the pitch change lever.

Abstract: A constant velocity drive mechanism of a rotary-wing aircraft rotor, comprises a torque splitting differential mechanism, comprising a driving disc integral in rotation with a rotor mast and connected to each of two driven discs either side of the driving disc, by a least one connecting pin hinged to each of the discs by one of three ball joint connections. Each of the driven discs is connected to a hub of the rotor by at least one of two driving devices, each of which is also hinged to the hub, so as to drive the hub in rotation about a geometrical axis which can be inclined in all directions about the axis of rotation of the mast.

Abstract: A tilting rotor of a convertible aircraft comprises, between the mast and the hub, a constant velocity drive mechanism and pivoting arrangement comprising two gimbals, each of which is driven by the mast and pivots about one respectively of two diametral axes by two diametrically opposite bearings. Each gimbal is also hinged to the hub by ball joint connections diametrically opposite and centred respectively in a radial plane passing through the axis of the mast and through the diametral axis about which the other gimbal pivots. The two resulting trains transmitting torque between the mast and the hub have substantially the same torsional rigidity, and one at least of the components of each torque transmission train exhibits a flexibility in deformation about the axis of rotation of the hub.

Abstract: A drag damper for use on a rotary wing aircraft rotor comprises a body defining two variable volume chambers linked by a piston. The chambers are filled with fluid and are connected by a restriction port (35) and by a channel (32) of larger cross-section than the restriction port. A secondary piston slidable and pressure-tight piston is fitted in the channel and loaded by an elastic bias (34). The equivalent mass of this secondary piston (33) and of the fluid which it displaces, and the stiffness of the elastic bias (34) are such that the secondary piston is resonant in the channel (32) at the rotation frequency of the rotor, to filter the dynamic component at this frequency of stresses applied to the damper (20).

Abstract: A constant velocity drive mechanism of a rotary-wing aircraft rotor, comprises a torque splitting differential mechanism, comprising a driving disc integral in rotation with a rotor mast and connected to each of two driven discs either side of the driving disc, by a least one connecting pin hinged to each of the discs by one of three ball joint connections. Each of the driven discs is connected to a hub of the rotor by at least one of two driving devices, each of which is also hinged to the hub, so as to drive the hub in rotation about a geometrical axis which can be inclined in all directions about the axis of rotation of the mast.

Abstract: A pitch control device is provided for a blade of a tilting rotor of a convertible aircraft. The pitch control device is external to a rotor hub which is driven in rotation by a mast and connected to the latter by means for tilting the hub as a whole, and each blade of which is connected to the hub by a coupling for retaining and hingeing the blade in pitch about its pitch change axis.

Abstract: A tilting rotor of a convertible aircraft comprises, between the mast and the hub, a constant velocity drive mechanism and pivoting arrangement comprising two gimbals, each of which is driven by the mast and pivots about one respectively of two diametral axes by two diametrically opposite bearings. Each gimbal is also hinged to the hub by ball joint connections diametrically opposite and centred respectively in a radial plane passing through the axis of the mast and through the diametral axis about which the other gimbal pivots. The two resulting trains transmitting torque between the mast and the hub have substantially the same torsional rigidity, and one at least of the components of each torque transmission train exhibits a flexibility in deformation about the axis of rotation of the hub.

Abstract: A drag damper for use on a rotary wing aircraft rotor comprises a body defining two variable volume chambers linked by a piston. The chambers are filled with fluid and are connected by a restriction port (35) and by a channel (32) of larger cross-section than the restriction port. A secondary piston slidable and pressure-tight piston is fitted in the channel and loaded by an elastic bias (34). The equivalent mass of this secondary piston (33) and of the fluid which it displaces, and the stiffness of the elastic bias (34) are such that the secondary piston is resonant in the channel (32) at the rotation frequency of the rotor, to filter the dynamic component at this frequency of stresses applied to the damper (20).

Abstract: The rolling bearing 1 of this bearing assembly is of the relubricatable type and is mounted tightly on the shaft 4 by its inner race 2 in such a way as to be prevented from rotating and translating on the shaft, the latter for this purpose comprising a thrust shoulder 3. The outer race 6 of the rolling bearing is housed in a ball swivel with a spherical outer peripheral surface consisting of two half ball swivels 7 back to back and this ball swivel is itself mounted in the spherical housing 8 of a support consisting of two flanges 9, 10 arranged facing one another and slightly separated from one another in the axial direction prior to clamping. One 9 of these flanges may be fixed by lugs 11 to the structure of the craft and the other flange 10 may be fixed to the previous one using screws 12.

Abstract: The rolling bearing 1 of this bearing assembly is of the relubricatable type and is mounted tightly on the shaft 4 by its inner race 2 in such a way as to be prevented from rotating and translating on the shaft, the latter for this purpose comprising a thrust shoulder 3. The outer race 6 of the rolling bearing is housed in a ball swivel with a spherical outer peripheral surface consisting of two half ball swivels 7 back to back and this ball swivel is itself mounted in the spherical housing 8 of a support consisting of two flanges 9, 10 arranged facing one another and slightly separated from one another in the axial direction prior to clamping. One 9 of these flanges may be fixed by lugs 11 to the structure of the craft and the other flange 10 may be fixed to the previous one using screws 12.

Abstract: The anti-vibration suspension device comprises at least three oblique bars pivotably mounted each on the gearbox and on the structure, by means of a rigid lever supporting a vibrating weight and in turn pivotably mounted on this structure, and each lever is also linked to the structure by at least one torsion spring biased about a torsion axis substantially perpendicular to the radial plane defined by the axes of the rotor and the corresponding oblique bar.

Abstract: The present invention relates to a method and to a device for determining the state of a vibrating structure of a rotary wing aircraft, while the latter is operating. The device includes sensors arranged on the vibrating structure and capable of measuring the values of vibrational parameters, and a calculation unit receiving the values measured by the sensors and calculating, from these and from estimated values of the vibrational parameters, a correlation coefficient, comparing the latter with a predefined range of values and deducing from this comparison the state of the vibrating structure.