Abstract: Devices for distributing oil to a rolling bearing for an aircraft turbine engine include a rolling bearing including two rings, respectively an inner ring and an outer ring, an oil distribution ring configured to be mounted on a turbine engine shaft, said distribution ring including a first outer cylindrical surface for mounting the inner ring of the bearing, an oil recovery scoop supplying a lubricating circuit of the bearing, and an annular track of a dynamic seal. The distribution ring and the track are formed by a single-piece body, and the lubricating circuit is formed in said body and extends into the distribution ring and the track.

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: A fan module for an aircraft turbine engine, the module including a fan having variable-pitch blades and an oil transfer device for transferring oil between a stator and a rotor, the device including a stator ring having a cylindrical inner surface and internal oil ducts and a shaft engaged in the ring and having a cylindrical outer surface and internal oil ducts; the device includes a plain bearing located between said cylindrical inner surface and the cylindrical outer surface and defined by a single band; and rolling bearings mounted between the ring and the shaft, on either side of the plain bearing, each of these rolling bearings having one of the rings thereof integrated in one of the elements selected from the ring and the shaft.

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: 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: A bypass turbomachine is disclosed. The turbomachine includes a longitudinal axis having an upstream blower impeller and a downstream air-flow straightening assembly of a secondary annular passage delimited radially on the inside by a radially inner shroud and radially on the outside by a radially outer shroud, blades extending between the radially inner and outer shrouds and being attached at a first end portion to the radially inner shroud and at a second end portion to the radially outer shroud, the blades including a useful portion extending between the first and second end portions and defining a lower face and an upper face. For each blade, in a plane perpendicular to the longitudinal axis, each aerodynamically useful portion is bent into a C-shape in the circumferential direction.

Abstract: An aircraft turbine engine having a primary air flow path with low-pressure and high-pressure compressors, a secondary air flow path which is located around the primary path and runs coaxially thereto, the turbine engine including vanes distributed about a main axis of the turbine engine. A pressurized air circuit draws air between the low-pressure compressor and the high-pressure compressor or in the high-pressure compressor and supplies at least one component located close to a main axis of the turbine engine. The pressurized air circuit includes a heat exchanger between the stream of pressurized air and the stream of air flowing in the secondary path, the heat exchanger being arranged in at least one of the straightening vanes, where a heat exchanger pipe is arranged, the pipe having a pressurized-air inlet and a pressurized-air outlet that are located at the same radial end of the vane.

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: A turbomachine blade including a body that extends mainly in a plane defined by a main axis and a longitudinal direction, which is defined by a lower surface wall, an upper surface wall, a leading edge located at a first longitudinal end of the body and a trailing edge located at a second longitudinal end of the body, wherein the body of the blade includes a plurality of first pipes that extend mainly along the direction of the main axis, for circulation of a gas flow, and a plurality of second pipes that extend mainly along the longitudinal direction, for circulation of a second gas flow.

Abstract: Devices for distributing oil from a rolling bearing for an aircraft turbine engine include a rolling bearing including two rings, respectively an inner ring and an outer ring, an oil distribution ring configured to be mounted on a turbine engine shaft, said distribution ring including a first outer cylindrical surface for mounting the inner ring of the bearing, an oil recovery scoop supplying a lubricating circuit of the bearing, and an annular track of a dynamic seal. The distribution ring and the track are formed by a single-piece body, and the lubricating circuit is formed in said body and extends into the distribution ring and the track.

Abstract: A turbine engine of an aircraft includes: a primary air flow duct; a secondary air flow duct which is located around the primary duct, the secondary duct including a stator including a plurality of blades distributed around a main axis of the turbine engine and inter-blade platforms located between radially internal ends or between radially external ends of two adjacent blades, each platform including a wall partially delimiting the secondary duct; and a fluid circuit which includes a heat exchanger formed by at least one of the platforms. The platform includes a line that has an inlet port of the fluid and a fluid outlet port. The fluid circuit includes a distributor associated with each port of the at least one platform with the rest of the fluid circuit and of which each distributor is axially offset with respect to the platform along the main axis.

Abstract: A turbomachine blade including a body that extends mainly in a plane defined by a main axis B and a longitudinal direction, which is defined by a lower surface wall, an upper surface wall, a leading edge located at a first longitudinal end of the body and a trailing edge located at a second longitudinal end of the body, wherein the body of the blade includes a plurality of first pipes that extend mainly along the direction of the main axis B, for circulation of a gas flow, and a plurality of second pipes that extend mainly along the longitudinal direction, for circulation of a second gas flow.

Abstract: The invention relates to an intermediate case (25) for a twin spool turbomachine for an aircraft, comprising a hub (26), an outer shell (23) and outlet guide vanes (24) installed at their ends on the hub and on the outer shell, and each of at least some of the outlet guide vanes (24) performing a heat exchanger function and comprising a lubricant passage (50a, 50b) designed to be cooled by the fan flow (58) following an outer surface of the outlet guide vane. According to the invention, the case also comprises at least one lubricant duct (55) passing along a circumferential direction of the hub (26) and at least part of which is made from a single casting with the hub, the lubricant duct (55) having at least one lateral opening communicating with the lubricant passage (50a, 50b) of at least one of the vanes (24).

Abstract: The invention relates to a guide vane for a bypass aircraft turbomachine, its aerodynamic part comprising a first lubricant cooling interior passage in which heat transfer structures are arranged and a second lubricant cooling interior passage in which heat transfer structures are arranged, the aerodynamic part comprising a bent area connecting a lubricant output end of the first interior passage to a lubricant input end of the second passage, the bent area extending along a curved generatrix and being partly delimited by the intrados wall and the extrados wall of the vane. According to the invention, the bent area comprises one or more lubricant guide(s) arranged between the intrados and extrados walls of the vane, and each extending substantially parallel to the curved generatrix of the bent area.

Abstract: A guide vane arranged in an air flow from a fan of an twin-spool aircraft engine, the aerodynamic part of the vane including an internal lubricant cooling passage partly delimited by an extrados wall and an extrados wall of the vane. The passage is equipped with a heat conduction matrix compressed between the walls and separating a first lubricant circulation space from a second lubricant circulation space. Furthermore, the matrix defines firstly first contact elements of the intrados wall formed in the first space and between which the lubricant from the first space will circulate, and secondly second contact elements of the extrados wall formed in the second space and between which the lubricant from the first space will circulate.

Abstract: A guide vane located in a fan air flow in an aircraft twin-spool turbomachine, the vane being made with an extrados body and an intrados body between which there is a thermal conduction matrix. Furthermore, the attachment devices between the two spools are arranged outside the aerodynamic part of the vane.

Abstract: The invention proposes a turbine engine (10) of an aircraft including a primary air flow duct (16) and a secondary air flow duct (20) which is located around the primary duct (16), the secondary duct (20) including a stator (52) including a plurality of blades distributed around a main axis (A) of the turbine engine (10) and which includes inter-blade platforms (80) each one of which is located between the radially internal ends (56) or between the radially external ends (58) of two adjacent blades (54), each platform (80) including a wall (82) partially delimiting the secondary duct (20), and including a fluid circuit (40), which includes a heat exchanger (44) formed by at least one of the platforms (80), characterised in that the platform (80) includes a line (84) that has an inlet port (90) of the fluid and a fluid outlet port (92), and in that the fluid circuit (40) includes a distributor (104) associated with each port (90, 92) of said at least one platform (80) with the rest of the fluid circuit (40)

Abstract: A ring of vanes for an aircraft turbine engine, the ring presenting an axial direction and a radial direction and including a first annular sheath presenting an inside surface and a second annular sheath presenting an inside surface facing the first inside surface of the first annular sheath, the first and second sheaths being coaxial and defining between them a flow passage for a gas stream. The ring further includes both partially annular acoustic treatment panels including resonant cavities and also vanes extending in the radial direction between the first and second sheaths. Each vane includes a radial airfoil with at least two attachment tabs at each radial end of the airfoil fastened to the first or second sheath, and the inside surface of at least one of the first and second annular sheaths and the corresponding attachment tabs are covered by acoustic treatment panels arranged between the airfoils.

Abstract: A turbine engine is provided with a longitudinal rotation axis having at least one shaft with a radial axis, in particular a pitch change system for the blades of a propeller, said shaft traversing a radial passage of a substantially cylindrical case around the longitudinal axis. The turbine engine includes an annular oil guiding device around the radial shaft. The device has first and second annular parts nested in one another and secured to one another by hooping. The first part is secured by hooping to the radial shaft, and the second part is configured so as to be separated from the first part under the action of a force oriented along the radial axis and exerted on the second part by a member of the turbine engine forming a stop during a radial movement of the radial shaft.

Abstract: A guide vane for a twin-spool aircraft turbomachine has an aerodynamic part that includes an internal lubricant cooling passage extending along a principal lubricant flow direction. The aerodynamic part is made in a single piece and also includes heat transfer fins arranged in the passage connecting the intrados and extrados walls and extending approximately parallel to the direction, these fins being distributed in successive rows along the principal direction and made such that for two rows of staggered directly consecutive fins, a first row includes fins forming a positive acute angle A1 with a dummy reference plane, while a second row includes fins forming a negative acute angle A2 with this plane.