Abstract: A device for centering and guiding a shaft of an aircraft turbine engines. The device includes an outer ring of a rolling bearing, the ring extending about a main axis and having orifices arranged around the axis; an annular bearing support extending about the main axis and at least partially about the ring, the support having orifices and openings arranged about the axis; and a series of studs for linking the ring to the support, the studs being distributed about the main axis and having elongation axes substantially parallel to the main axis, the body of each of the studs passing through one of the openings. Some of the studs, have bodies that pass with a first positive clearance and a second virtually zero clearance through the first openings, the first and second clearances being configured so that the device has different displacement amplitudes.

Abstract: A device for centering and guiding a shaft of an aircraft turbine engine, the device including an outer ring of a rolling bearing, the ring extending about an axis; an annular bearing support extending about the axis and at least partially about the ring; at least one series of connecting elements for linking the ring to the support, the connecting elements being interposed between an inner cylindrical rim of the ring and an outer cylindrical rim of the support that extends about the inner rim, and each having a first radially outer end for linking to the outer cylindrical rim and a second radially inner end for linking to the inner cylindrical rim, the ring, the support and the connecting elements being integrally formed, the connecting elements being enclosed in an annular housing that is radially delimited by the rims and laterally enclosed by an annular web.

Abstract: A device for centering and guiding a shaft of an aircraft turbine engine is provided. The device includes an outer ring of a rolling bearing extending about an axis and having orifices, an annular bearing support extending about the axis and at least partially about the ring, the support having orifices and a series of studs for connecting the ring to the support. The studs can be distributed about the axis and extend parallel to the axis. The ends of first studs can be engaged without clearance in the orifices, and ends of second studs can be engaged with clearances in the orifices of the ring and/or of the support, the clearances being configured so that the device has different stiffnesses in at least two directions perpendicular to the axis.

Abstract: An impeller (230) for a planet carrier of a planetary gear speed reducer of a turbomachine, is configured intended to be rotatably secured to the planet carrier and to be rotated about an axis A of the speed reducer. The impeller has an annular shape about the axis and includes lubrication means (43, 45, 238), in particular for lubricating bearings of planet gears of the speed reducer. The lubrication means include an annular cavity (238) situated at the inner periphery of the impeller. The impeller includes an inner peripheral wall (246) closing the cavity (238) in the radial direction, and the impeller includes an annular port (248) that extends around the axis and that opens in the axial direction into the cavity in order to supply it with lubricating oil.

Abstract: A scoop for an aircraft turbine engine, this scoop including a body which is movable between at least two positions, including a first position in which it is configured to capture part of a first air stream flowing in a first direction, wherein the body is movable into a second position in which it is configured to capture part of a second air stream flowing in a second direction which is opposite to the first direction, and wherein the body is mounted to move freely between the at least two positions so as to automatically adopt the first position when the first air stream flows, and to automatically adopt the second position when the second air stream flows.

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 device for centring and rotationally guiding a turbomachine shaft includes a rolling-element bearing outer ring; a bearing support having a radially inner annular surface surrounding the outer ring; a connection structure connecting the outer ring to the bearing support and including an elastically deformable structure; a first cavity formed between the radially inner annular surface and the outer ring to receive a first damping fluid film; at least one second cavity formed between the bearing support and the connection structure which is arranged radially outside the radially inner annular surface to receive a second respective damping fluid film. Such an arrangement allows for the coexistence of damping fluid films having very different damping properties.

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: The invention relates to a turbine engine module, comprising: - a fan shaft (6), - a power shaft (15) rotating the fan shaft via a speed reducer (21) arranged in a lubrication chamber (22), - a lubrication system (50) for the speed reducer (21) having a main circuit (51), which is closed, for supplying the lubrication chamber, and an auxiliary circuit (60), which is closed, for supplying the lubrication chamber when the main circuit is inactive, the auxiliary circuit comprising an auxiliary pump (61) driven by an electric motor (63), and - an electric machine (65) configured to power the electric motor (63), the electric machine having a rotor connected to the fan shaft in order to be rotated and a stator mounted on a stationary member at least partially supporting the electric machine.

Abstract: A reducing assembly includes a planet carrier that is flexible and fixed, borne by a casing which surrounds it by two distinct connections with this casing, with a device for applying a torsional prestress to the planet carrier between these two connections. Prestressing the flexible planet carrier makes it possible to limit its level of mechanical stress when it is in service, in order to simplify its design and dimensioning.

Abstract: A dual-flow turbojet engine having a central shaft surrounded by a high-pressure body which rotate about the same longitudinal axis while being independent in rotation, and including a fan driven by the central pressure shaft; a high-pressure compressor and a high-pressure turbine mounted on the high-pressure body; an inter-turbine casing; a low-pressure turbine mounted on a low-pressure rotor surrounding the central shaft; an exhaust casing on which an output cone is mounted; a reduction gear with which the low-pressure rotor drives the central pressure shaft; two bearings mounted on the exhaust casing and respectively receiving the central shaft and the low-pressure rotor; a bearing mounted on the inter-turbine casing and receiving the low-pressure rotor.

Abstract: The invention relates to a turbomachine for an aircraft, comprising: a first rotor comprising a first shaft, a second rotor comprising a second shaft, a mechanical reduction gearing having an epicyclic gear set comprising a sun gear connected to the second shaft, a ring gear connected to the first shaft, and planet gears located between the sun gear and the ring gear and borne by a planet carrier attached to a stator of the turbomachine, rolling element bearings for guiding said first shaft and second shaft in rotation, an annular gutter which extends around the ring gear of the reduction gearing and which is configured to recover oil for lubricating the reduction gearing that is sprayed by centrifugal action out from the ring gear during operation, and an annular bearing support which is attached, with the gutter, to a stator of the turbomachine and which supports at least one of said bearings, characterized in that it also comprises: at least one device for conveying oil recovered by said gutter, which d

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: Aircraft turbomachine comprising a casing, a fan, a compressor and a turbine and an epicyclic gear train comprising an input driven in rotation by the turbine, a first output stage configured to drive in rotation the compressor and a second output stage coupled to the first output stage and configured to drive in rotation the fan, the compressor being driven in rotation by the ring gear of the first output stage.

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: An aircraft turbine engine includes a turbine shaft having a first axis of rotation, a propulsion propeller having a second axis of rotation parallel to and spaced from the first axis, and a mechanical reduction gear coupled to the turbine shaft and rotating the propeller. The reduction gear has a sun gear connected to the turbine shaft, a ring gear, and at least two planet gears, each including a first external toothing that is meshed with an external toothing of the sun gear. A second external toothing is located within the ring gear and is meshed with an internal toothing of the ring gear.

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: An impeller (230) for a planet carrier of a planetary gear speed reducer of a turbomachine, is configured intended to be rotatably secured to the planet carrier and to be rotated about an axis A of the speed reducer. The impeller has an annular shape about the axis and includes lubrication means (43, 45, 238), in particular for lubricating bearings of planet gears of the speed reducer. The lubrication means include an annular cavity (238) situated at the inner periphery of the impeller. The impeller includes an inner peripheral wall (246) closing the cavity (238) in the radial direction, and the impeller includes an annular port (248) that extends around the axis and that opens in the axial direction into the cavity in order to supply it with lubricating oil.

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).