Abstract: A turbojet includes in succession, from upstream to downstream in the flow direction of a primary air stream: a low-pressure casing, an intermediate casing, and a high-pressure casing that are longitudinally aligned and which jointly delimit an internal annular passage for the circulation of the air stream. The intermediate casing delimits a portion of the annular passage called a gooseneck and includes: an air discharge system able to extract air from the air stream circulating in a first zone of the gooseneck and to discharge it outside the passage, and an air extraction system able to extract air circulating in a second zone of the gooseneck for an air conditioning system, the second air extraction zone being located downstream of the first air extraction zone.

Abstract: A stator assembly including plural stator vanes distributed around an axis of revolution of the stator assembly, a chord of the stator vane, taken at a root of the stator vane, not overlapping, in the direction of the axis of revolution, a chord of an adjacent stator vane, taken at a root of the adjacent stator vane, and a chord of the stator vane, taken at a tip of the stator vane, overlapping, in the direction of the axis of revolution, a chord of the adjacent stator vane, taken at a tip of the adjacent stator vane.

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: The invention relates to a straightening assembly (28) comprising two radially inner and outer coaxial shells (34) between which extend vanes (36) made of composite material, fixed to a first end portion on the radially inner shell and to a second end portion on the radially outer shell, characterized in that: for each vane (36), in a plane perpendicular to the axis of the radially inner and outer shells (34), a straight line passing through a junction between said first and second end portions and the useful part forms an angle ? with a radius of the radially inner shell, passing through the junction between said first end portion and the useful part of said vane (36), such that 0°<??30°; the radially inner shell has a diameter ranging from 1,000 mm to 1,600 mm; and in that the radially outer shell has a diameter ranging from 2,000 to 2,800 mm; the number of vanes (36) ranges from twenty-five to forty-five.

Abstract: A method determines the angular setting (?_2) of an annular row of stator vanes arranged downstream of a pusher propeller of a propulsion system with a longitudinal axis. The annular row of stator vanes receives an air flow having a velocity (V2) including a longitudinal component (Viz) and a tangential component (V_i?) associated with the velocity of gyration generated by the pusher propeller. The method includes the steps of establishing a theoretical model of the pusher propeller using a power (P1) and a mechanical speed (N1) associated with said pusher propeller, and flight conditions that include a velocity of the airflow incident on the pusher propeller, the altitude of said propulsion system and ambient temperature. The method further includes the step of determining an angular setting (?_1) of the pusher propeller from the theoretical model.

Abstract: A turbomachine part or assembly of parts includes a first blade and a second blade, and platform from which the first blade and the second blade extend, The platform has, between a pressure side of the first blade and a suction side of the second blade a non-axisymmetric surface defining at least one fin, and a transverse section of the fin is asymmetric. The transverse section includes a first and a second oblique faces joining at a dorsal edge, the first oblique face is oriented toward the pressure side of the first blade, the second oblique face is oriented toward the suction side of the second blade, and the first oblique face has a steeper slope than the second oblique face.

Abstract: The invention relates to an aircraft turbine engine (101) which comprises, downstream of the low pressure compressor of its low pressure body (104) and upstream of its combustion chamber (102), an electric motor (206) which is configured to rotate a rotor blade ring so as to generate a flow of air. Moreover, the rotational speed of the electric machine rotor is independent of the rotational speed of the compressor rotor.

Abstract: The invention relates to a turbomachine assembly (1) comprising a fan (2) and a booster drum-type part (3), the fan (2) comprising:—blades (20) comprising an airfoil (23) and an extension (30) mounted on and attached to the trailing edge (25) of the airfoil (23),—a fan (2) disc (10) and—a series of inter-blade platforms (16), the extension (30) of each blade (20) extending beyond the downstream face (14) of the fan (2) disc (10) in the direction of the upstream edge (4) of the part (3) and covering, at least partially, the cavity (6) between the fan (2) and the part (3).

Abstract: The invention relates to a straightening assembly for an air flow of a turbomachine, comprising two respectively coaxial radially inner and outer shrouds (30), between which extend the vanes (32) made of composite material attached, at a first end portion (42) thereof, to the radially inner shroud and at a second end portion (44) thereof, to the radially outer shroud (30), with the vanes (32) comprising a useful portion extending between the two end portions (42, 44). According to the invention, for each vane (32) in a plane perpendicular to the axis of the radially inner and outer shrouds (30), a straight line passing through a junction between said first and second end portions (42, 44) and the useful portion forms, with a radius of the radially inner shroud passing through the junction between said first end portion and the useful portion of said vane (32), an angle ? such that 0°<??90°.

Abstract: The invention relates to a turbomachine assembly (1) comprising a fan (2) and a booster drum-type part (3), the fan (2) comprising:—blades (20) comprising an airfoil (23) and an extension (30) mounted on and attached to the trailing edge (25) of the airfoil (23),—a fan (2) disc (10) and—a series of inter-blade platforms (16), the extension (30) of each blade (20) extending beyond the downstream face (14) of the fan (2) disc (10) in the direction of the upstream edge (4) of the part (3) and covering, at least partially, the cavity (6) between the fan (2) and the part (3).

Abstract: Part comprising a wall which comprises a first zone (541), a first zone (541) and the second zone (542), a network of riblets being formed on the first zone (541), the second zone (542) and also on the transition zone (54t) so as to reduce the drag of the part when a flow of air flows along said wall; the height, the width and the spacing of the riblets formed on the transition zone (54t) changing along said transition zone (54t) so as to pass from the height, width and spacing of the riblets formed on the first zone at a first end of the transition zone to the height, width and spacing of the riblets formed on the second zone (542) at a second end of the transition zone (54t), the transition zone (54t) comprising a central portion on which the riblets comprise on one hand the height and the width that are respectively equal to the height and width of the riblets on the first zone (541), and on the other hand a spacing equal to the spacing of the riblets of the second zone (542).

Abstract: A turbine engine part or set of parts including at least first and second obstacles each having a leading edge and a trailing edge, and a platform from which the obstacles extend; wherein the platform has, between the pressure side of the first obstacle and the suction side of the second obstacle a non-axisymmetric surface defining at least one fin with a substantially triangular cross-section, each fin being associated with a leading position and a trailing position on the surface between which the fin extends, such that the leading position is upstream of each of the leading edges; the trailing position is downstream of each of the leading edges.

Abstract: Blade made of composite material, with an integrated platform, for an aircraft turbine engine, including a root connected to an aerofoil and having a connection platform between the root and the aerofoil, the root having a generally elongate shape along an axis (A) of elongation, the aerofoil having a curved cross-sectional shape and including a pressure face and a suction face, the platform including a first portion extending on the pressure-face side of the aerofoil and including a first longitudinal free edge and a second opposite longitudinal edge for connecting to the rest of the blade, the platform including a second portion extending on the suction-face side of the aerofoil and comprising a third longitudinal free edge and an opposite fourth longitudinal edge for connecting to the rest of the blade, wherein the first longitudinal free edge is concave and the third longitudinal free edge is convex, and, for any plane (P, P?) perpendicular to the axis of elongation and intersecting the platform, the dist

Abstract: Part comprising a wall which comprises a first zone (541), a first zone (541) and the second zone (542), a network of riblets being formed on the first zone (541), the second zone (542) and also on the transition zone (54t) so as to reduce the drag of the part when a flow of air flows along said wall; the height, the width and the spacing of the riblets formed on the transition zone (54t) changing along said transition zone (54t) so as to pass from the height, width and spacing of the riblets formed on the first zone at a first end of the transition zone to the height, width and spacing of the riblets formed on the second zone (542) at a second end of the transition zone (54t), the transition zone (54t) comprising a central portion on which the riblets comprise on one hand the height and the width that are respectively equal to the height and width of the riblets on the first zone (541), and on the other hand a spacing equal to the spacing of the riblets of the second zone (542).

Abstract: The invention relates to a straightening assembly (28) comprising two radially inner and outer coaxial shells (34) between which extend vanes (36) made of composite material, fixed to a first end portion on the radially inner shell and to a second end portion on the radially outer shell, characterized in that: for each vane (36), in a plane perpendicular to the axis of the radially inner and outer shells (34), a straight line passing through a junction between said first and second end portions and the useful part forms an angle ? with a radius of the radially inner shell, passing through the junction between said first end portion and the useful part of said vane (36), such that 0°<??30°; the radially inner shell has a diameter ranging from 1,000 mm to 1,600 mm; and in that the radially outer shell has a diameter ranging from 2,000 to 2,800 mm; the number of vanes (36) ranges from twenty-five to forty-five.

Abstract: The invention relates to a straightening assembly for an air flow of a turbomachine, comprising two respectively coaxial radially inner and outer shrouds (30), between which extend the vanes (32) made of composite material attached, at a first end portion (42) thereof, to the radially inner shroud and at a second end portion (44) thereof, to the radially outer shroud (30), with the vanes (32) comprising a useful portion extending between the two end portions (42, 44). According to the invention, for each vane (32) in a plane perpendicular to the axis of the radially inner and outer shrouds (30), a straight line passing through a junction between said first and second end portions (42, 44) and the useful portion forms, with a radius of the radially inner shroud passing through the junction between said first end portion and the useful portion of said vane (32), an angle ? such that 0°<??90°.

Abstract: Blade made of composite material, with an integrated platform, for an aircraft turbine engine, including a root connected to an aerofoil and having a connection platform between the root and the aerofoil, the root having a generally elongate shape along an axis (A) of elongation, the aerofoil having a curved cross-sectional shape and including a pressure face and a suction face, the platform including a first portion extending on the pressure-face side of the aerofoil and including a first longitudinal free edge and a second opposite longitudinal edge for connecting to the rest of the blade, the platform including a second portion extending on the suction-face side of the aerofoil and comprising a third longitudinal free edge and an opposite fourth longitudinal edge for connecting to the rest of the blade, wherein the first longitudinal free edge is concave and the third longitudinal free edge is convex, and, for any plane (P, P?) perpendicular to the axis of elongation and intersecting the platform, the dist