Abstract: A fan module for an aircraft turbomachine includes a fan with a disc carrying fan blades and a rotor of an electric machine. The rotor has an annular shape and is mounted coaxially downstream of the fan. The module further includes an annular support for the rotor with a downstream end fixed to the rotor and with an upstream end fixed to the fan disc. The support is fixed to the disc by shear bolts configured to break when a torque transmitted by the disc to the support exceeds a predetermined threshold.

Abstract: A turbomachine includes a single ducted fan including a first shaft rotated by a second shaft via a speed reduction gearset, the second shaft being rotated by a third shaft of a turbine, the first shaft being guided in rotation with respect to a fixed structure via a first bearing and a second bearing placed upstream of the speed reduction gearset. The second shaft is guided in rotation with respect to the first shaft via a rolling bearing placed upstream of the speed reduction gearset, the rolling bearing comprising an outer ring housed in the first shaft, an inner ring connected to the second shaft and rolling elements arranged between the inner and outer rings.

Abstract: A pump that can be a gear pump that drives a rotating body of an aircraft turbine engine such as a fan shaft, is set in rotation always in the same direction, by virtue of a transmission including an oblique double gear formed by a toothed wheel on the same side as the rotating body and meshing with two pinions connected to a shaft driving the pump through free-wheels with identical driving directions, whereas the pinions have opposite directions of rotation.

Abstract: A drive shaft of an aircraft turbine engine includes a first portion and a second portion, and connecting means connecting the first and second portions and being configured to transmit a torque from the second portion to the first portion. The connecting means has at least one bellows with a first section having a diameter greater than the diameters of the first and second portions and second sections flanking the first section. The first section includes at least one fusible section with at least one through-hole and being configured to break when the value of a torque applied to the first portion exceeds a predetermined threshold value.

Abstract: A shaft assembly (22, 32) for an aircraft turbine engine (1), comprising a first outer shaft (32) and a second inner shaft (22), the first outer shaft being intended to be engaged axially on the second shaft and comprising inner longitudinal splines (34) for coupling with outer longitudinal splines (24) of the second shaft, characterised in that, when the first and second shafts are in the coupling position, the inner and outer splines are engaged in one another and cooperate with each other in an axial coupling area (Z), the inner and outer teeth being situated outside this coupling area.

Abstract: A speed reducer, particularly of an aircraft, includes a central sun gear mounted to rotate with a drive shaft about a rotation axis X, an annular ring gear coaxial with the X axis, and a plurality of planet gears arranged about the X axis and mounted to be movable on a planet carrier. The planet gears engages the sun gear and the annular ring gear, which extends around the planet gears. The speed reducer further includes a lubricant recovery device having a pipe for recovering lubricant discharged from the reducer by centrifuging. The recovery device has scoops mounted to pivot about a pivot axis P on a movable member of the speed reducer that rotates about the X axis between a deployed state and a rest state.

Abstract: The invention relates to a planetary gear reducer (21) for an aircraft turbine engine (1), the reducer (21) comprising a sun gear (30) and a ring gear (33) which are centred on a longitudinal axis X and a planet carrier (32) which carries at least one planet gear (31) rotatably mounted about a planet axis A parallel to the longitudinal axis X, the sun gear (30) being rotatable about the longitudinal axis, the planet gear (31) meshing both with the sun gear (30) and the ring gear (33), the planet carrier (32) being movable about the longitudinal axis and the ring gear (33) being rotationally fixed. According to the invention, the reducer (21) comprises an electric machine (50) integrated therewith and which comprises a rotor (51) mounted on the planet carrier (32) in such a way as to be rotated about the longitudinal axis X and a stator (52) mounted on the ring gear (33).

Abstract: An enclosure for a turbomachine includes a turbomachine drive shaft rotating about a longitudinal axis (X) by means of two roller bearings, an upstream bearing and a downstream bearing, each having an inner ring carried by the drive shaft. The two bearings share a single integral outer ring that has an upstream end and a downstream end connected to one another by a section of studs. The single outer ring is carried by an upstream base plate and a downstream base plate of a bearing support configured to be attached to a stationary structure of the turbomachine.

Abstract: A fan module for an aircraft turbomachine includes a fan with a disc carrying fan blades and a rotor of an electric machine. The rotor has an annular shape and is mounted coaxially downstream of the fan. The module further includes an annular support for the rotor with a downstream end fixed to the rotor and with an upstream end fixed to the fan disc. The support is fixed to the disc by shear bolts configured to break when a torque transmitted by the disc to the support exceeds a predetermined threshold.

Abstract: An aircraft turbomachine with a reduction gear has a first shaft and a second shaft having one same axis of rotation, the second shaft being rotationally driven by the first shaft via the reduction gear, the first shaft having elastically deformable means having bellows section(s) and being connected to the reduction gear by a connecting system likewise having elastically deformable means involving a hairpin or bellows section(s).

Abstract: A pump that can be a gear pump that drives a rotating body of an aircraft turbine engine such as a fan shaft, is set in rotation always in the same direction, by virtue of a transmission including an oblique double gear formed by a toothed wheel on the same side as the rotating body and meshing with two pinions connected to a shaft driving the pump through free-wheels with identical driving directions, whereas the pinions have opposite directions of rotation.

Abstract: A speed reducer, particularly of an aircraft, includes a central sun gear mounted to rotate with a drive shaft about a rotation axis X, an annular ring gear coaxial with the X axis, and a plurality of planet gears arranged about the X axis and mounted to be movable on a planet carrier. The planet gears engages the sun gear and the annular ring gear, which extends around the planet gears. The speed reducer further includes a lubricant recovery device having a pipe for recovering lubricant discharged from the reducer by centrifuging. The recovery device has scoops mounted to pivot about a pivot axis P on a movable member of the speed reducer that rotates about the X axis between a deployed state and a rest state.

Abstract: A turbine engine including a fan shaft driven by a turbine shaft with a rotational speed reduction device. A decoupling element is interposed between the reduction device and the fan shaft which operates to decouple the reduction device and the fan shaft. The reduction device is coupled to the fan shaft by a coupling having trapezoidal teeth of the curvic type which are used during the said decoupling operation. The decoupling element is configured to decouple the reduction device and the fan shaft in response to a predetermined resistive torque, referred to as the decoupling torque, which acts between the fan shaft and the reduction device.

Abstract: Dynamic sealing device for an aircraft turbomachine, comprising two rotors (3, 8) which are configured to rotate about the one same axis (A) and different speeds, the said rotors extending at least in part around one another and dynamic sealing means (14) being mounted between the said rotors, characterized in that: —the said sealing means comprise annular rings (38) extending about the said axis, these rings being mounted with the ability to slide on a first (3) of the said rotors, in a radial direction with respect to the said axis, and being able to bear radially via their external periphery against a second (8) of the said rotors which at least partially surrounds the said first rotor, and in that it further comprises: —oil discharge means comprising an annular cavity (44) formed in the said second rotor, the said annular cavity opening radially towards the inside in the vicinity of the said sealing means and being in fluidic communication with oil passage ports (45) formed in the bottom of the cavity and

Abstract: A drive shaft of an aircraft turbine engine includes a first portion and a second portion, and connecting means connecting the first and second portions and being configured to transmit a torque from the second portion to the first portion. The connecting means has at least one bellows with a first section having a diameter greater than the diameters of the first and second portions and second sections flanking the first section. The first section includes at least one fusible section with at least one through-hole and being configured to break when the value of a torque applied to the first portion exceeds a predetermined threshold value.

Abstract: A turbomachine having a fan shaft supported by a first bearing positioned downstream of the fan, the first bearing including an outer ring attached to an annular support secured to the stator. The turbomachine includes at least one gas feed duct leading into an enclosure positioned against the disc, the gas feed duct being adapted to be fed at a second end with pressurized gas taken from an airstream of a high-pressure compressor of the turbomachine, so that the gas applies an axial force towards upstream on the disc during some operating phases of the turbomachine. The turbomachine further includes a ferrule imperviously sealing the enclosure.

Abstract: A turbomachine includes a single ducted fan including a first shaft rotated by a second shaft via a speed reduction gearset, the second shaft being rotated by a third shaft of a turbine, the first shaft being guided in rotation with respect to a fixed structure via a first bearing and a second bearing placed upstream of the speed reduction gearset. The second shaft is guided in rotation with respect to the first shaft via a rolling bearing placed upstream of the speed reduction gearset, the rolling bearing comprising an outer ring housed in the first shaft, an inner ring connected to the second shaft and rolling elements arranged between the inner and outer rings.

Abstract: An aircraft turbomachine has a longitudinal axis A and includes a reduction gear, at least one turbine shaft, an upstream end of which is connected to an input shaft of a reduction gear, a fan, one shaft of which is connected to an output shaft of the reduction gear, and dynamic sealing means between said reduction gear input shaft and the fan shaft. The sealing means includes an annular cowling, the axial cross-section of which is substantially U-shaped An outer annular leg of the annular cowling is fastened to the fan shaft, and an inner annular leg supports at least one annular sealing joint that cooperates with the reduction gear input shaft. The cowling is elastically deformable in the radial direction in relation to the axis A.

Abstract: A turbine engine includes a fan. The fan includes a first shaft driven in rotation via a reduction gear, the first shaft being guided in rotation with respect to a support of a fixed structure via an upstream bearing and a downstream bearing, the downstream bearing guiding in rotation the first shaft with respect to a downstream branch of the support, the first shaft and the support together defining an enclosure. The reduction gear includes a ring gear fixed to the structure via a shell. The turbine engine also includes a lubricator composed of at least one pipe for conveying a liquid lubricant, wherein the pipe is located in the enclosure, the pipe passing through an orifice made in the shell and an opening made in the downstream branch.

Abstract: A multi-flow turbojet engine generally includes an upstream ducted fan driven by a gas generator, the gas generator having a first and a second compressor, which are coaxial, an inlet case forming a support for the rotors of the upstream ducted fan and of the first compressor, an inter-compressor case downstream of the inlet casing and forming a support for the rotors of the second compressor, attachment means for thrust take-up rods, which attachment means are arranged on the inter-compressor case. The turbojet engine includes at least one axial stress transmission rod for connecting the inlet case to the inter-compressor case.