Abstract: Stator structure (22) extending around an axis of the turbomachine and comprising: —a support (50) having an inner surface centred on the axis and—a flange (60) defining an air chamber (A2) and having an outer surface centred on the axis, the support (50) extending around the flange (60) such that the inner and outer surfaces are opposite to each other, the structure (22) defining an oil circuit and an air circuit which are formed by upstream channels (64, 65) and downstream channels (54, 55), —each upstream channel (64, 65) defining an outer opening in the outer surface, —each downstream channel (54, 55) defining an inner opening in the inner surface, each circuit being oriented between the outer and inner openings in a direction comprising a component radial to the axis.

Abstract: An auxiliary tank for an aircraft turbine engine is provided, and in conjunction with a pump and associated auxiliary lubrication circuit, supplies oil to a reducer when, for example, a phase of free rotation of the fan is detected. Thus, the reducer is always lubricated, even during the phases of free rotation of the fan, thus ensuring a longer life of the reducer gears. The auxiliary tank is arranged with respect to the reducer so that oil is recovered from the reducer at least in part by the auxiliary tank.

Abstract: The present invention relates to a shaft for a propulsion system configured to rotate a reducing mechanism about a rotational axis, the shaft comprising: —a first end configured to engage with an input gear of the reducing mechanism, —a first bellows and a second bellows, the first bellows and the second bellows being rotationally symmetrical about the rotational axis, the first bellows extending between the first end and the second bellows, and —a frustoconical body mechanically connecting the first bellows and the second bellows.

Abstract: Stator structure (22) extending around an axis of the turbomachine and comprising: —a support (50) having an inner surface centred on the axis and —a flange (60) defining an air chamber (A2) and having an outer surface centred on the axis, the support (50) extending around the flange (60) such that the inner and outer surfaces are opposite to each other, the structure (22) defining an oil circuit and an air circuit which are formed by upstream channels (64, 65) and downstream channels (54, 55), —each upstream channel (64, 65) defining an outer opening in the outer surface, —each downstream channel (54, 55) defining an inner opening in the inner surface, each circuit being oriented between the outer and inner openings in a direction comprising a component radial to the axis.

Abstract: An auxiliary tank for an aircraft turbine engine has a curved general shape, and includes a cylindrical or frustoconical radially outer wall and three cylindrical or frustoconical radially inner walls arranged facing the radially outer wall, the three radially inner walls having a middle wall and two side walls arranged to either side of the middle wall. The inner volume of the tank forms a U-shape, the branches of which are formed by lateral volume portions between the side walls and the radially outer wall, and the base of which is formed by a middle volume portion between the middle wall and the radially outer wall.

Abstract: An assembly for an aircraft turbine engine, includes a fan, a speed reduction gear and an intermediate casing including an internal annular structure that houses an epicyclic gear train equipped with a planet carrier. The assembly includes a secondary retention device for the ring gear of the planet carrier, the device including: first projections that project radially inwards from the internal annular structure, each first projection having a recess that is open radially inwards, axially downwards and circumferentially in a first direction; and second projections that project radially outwards from the ring gear of the planet carrier and cooperate with the first projections in such a way that each second projection is partially accommodated in the recess of the first projection with which it is associated and downstream therefrom.

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: Device for distributing oil from a rolling bearing (8) for an aircraft turbine engine, comprising: - a rolling bearing (8), an oil distributor ring (5) and - an annular track (26) of a dynamic seal (22), characterised in that it further comprises a nut (16) screwed on to a thread (5d) of the distributor ring and bearing axially against an axial end of the inner ring so as to clamp it axially, and in that the annular track is configured to bear axially against the distributor ring and comprises rotating locking elements (27) engaging with the additional elements (28)of the nut.

Abstract: The invention relates to an assembly for holding an epicyclic gear train in a turbomachine comprising an annular housing in which are engaged a first annular part and a second annular part which are locked against rotation in the housing by first annular interlocking means and second annular interlocking means, respectively, wherein the first annular interlocking means are configured to allow a first circumferential clearance between the first annular part and the housing, the second annular interlocking means are configured to allow a second circumferential clearance between the second annular part and the housing, and the first circumferential clearance is strictly less than the second circumferential clearance.

Abstract: The present invention relates to a shaft for a propulsion system configured to rotate a reducing mechanism about a rotational axis, the shaft comprising: —a first end configured to engage with an input gear of the reducing mechanism, —a first bellows and a second bellows, the first bellows and the second bellows being rotationally symmetrical about the rotational axis, the first bellows extending between the first end and the second bellows, and —a frustoconical body mechanically connecting the first bellows and the second bellows.

Abstract: The invention concerns an aircraft turbine engine (1), comprising a gas generator comprising at least one annular gas flow duct (2), the duct (2) being defined by two annular housings, respectively an external housing (3b) and an internal housing (3a), extending one around the other and connected together by at least one tubular arm (5) for the passage of a lubricating oil line (7). According to the invention, the line (7) comprises a first fixed section (14) secured to the external housing (3b), a second fixed section (16) secured to a piece of equipment (4) of the turbomachine capable of moving or vibrating during operation relative to the housings (3a, 3b). and an intermediate section (18) connecting the first and second sections (14, 16), this intermediate section (18) having a generally elongate shape and comprising longitudinal ends engaged and capable of swiveling and/or sliding in ends of the first and second sections (14, 16).

Abstract: A planetary reduction gear (10) for an aircraft turbine engine, said reduction gear comprising a rotatable sun gear, a rotatable ring gear (9) and satellite gears meshed with the sun gear and the ring gear and carried by a planet carrier intended to be fixed to a stator, the reduction gear further comprising an annular trough for recovering and channelling lubrication oil that is mounted around the ring gear, characterised in that the reduction gear further comprises an annular deflector (22) fixed to the ring gear and configured to route the oil exiting radially towards the outside of the ring gear up to the trough by virtue of centrifugal forces, the trough comprising an annular chamber (18) which is axially remote relative to a median plane (P) substantially passing through the centre of the ring gear, and an annular duct (20) located on a side of said chamber and emerging in said chamber.

Abstract: The invention concerns a device for centring and guiding in rotation a rotating part rotating about a turbomachine axis, comprising a rolling bearing comprising an outer ring, a bearing support surrounding the outer ring, an annular attachment flange mounted on the bearing support and a flexible portion connecting the outer ring to the attachment flange and comprising first arms extending from the attachment flange in a first direction with a first radial component towards the inside, second arms extending from the outer ring of the rolling bearing in a second direction with a second radial component opposite the first radial component, towards the outside, and an annular base that the first arms and the second arms join; the first arms alternating circumferentially with the second arms.

Abstract: An aircraft propulsion assembly including a fan shaft which is connected to a turbine shaft by means of a reduction gear. The guiding of the fan shaft is carried out by a first bearing including cylindrical rollers and a second beating including two rows of balls with oblique contact.

Abstract: The invention concerns a device for centring and guiding in rotation a rotating part rotating about a turbomachine axis, comprising a rolling bearing comprising an outer ring, a bearing support surrounding the outer ring, an annular attachment flange mounted on the bearing support and a flexible portion connecting the outer ring to the attachment flange and comprising first arms extending from the attachment flange in a first direction with a first radial component towards the inside, second arms extending from the outer ring of the rolling bearing in a second direction with a second radial component opposite the first radial component, towards the outside, and an annular base that the first arms and the second arms join; the first arms alternating circumferentially with the second arms.

Abstract: An auxiliary tank for an aircraft turbine engine, has a curved general shape, and includes a cylindrical or frustoconical radially outer wall and three cylindrical or frustoconical radially inner walls arranged facing the radially outer wall, the three radially inner walls having a middle wall and two side walls arranged to either side of the middle wall. The inner volume of the tank forms a U-shape, the branches of which are formed by lateral volume portions between the side walls and the radially outer wall, and the base of which is formed by a middle volume portion between the middle wall and the radially outer wall.

Abstract: The present disclosure describes a lubrication oil piping gutter of an aircraft turbomachine. The lubrication oil piping gutter can be used for a sun gear train reducer of the turbomachine. The gutter may include two lateral annular walls connected at their outer periphery by an annular bottom wall. In some embodiments, the annular bottom wall may include at least one oil evacuation opening in which is located a deflector configured to facilitate the flow of oil radially from the inside to the outside of the gutter.

Abstract: The invention relates to an assembly for holding (42) an epicyclic gear train in a turbomachine comprising an annular housing (50) in which are engaged a first annular part (44) and a second annular part (46) which are locked against rotation in the housing (50) by first annular interlocking means and second annular interlocking means, respectively, wherein: the first annular interlocking means are configured to allow a first circumferential clearance between the first annular part (44) and the housing (50); the second annular interlocking means are configured to allow a second circumferential clearance between the second annular part (48) and the housing (50); the first circumferential clearance is strictly less than the second circumferential clearance.

Abstract: The present disclosure describes a lubrication oil piping gutter of an aircraft turbomachine. The lubrication oil piping gutter can be used for a sun gear train reducer of the turbomachine. The gutter may include two lateral annular walls connected at their outer periphery by an annular bottom wall. In some embodiments, the annular bottom wall may include at least one oil evacuation opening in which is located a deflector configured to facilitate the flow of oil radially from the inside to the outside of the gutter.