Abstract: A process for using an air input of a turboreactor nacelle of an aircraft, comprising an air input lip which comprises at least one fixed portion and at least one portion which can be moved between a first position, in which the air input lip has an aerodynamic profile so as to guide the internal air flow over the internal wall in order to promote a thrust phase, and a second position, in which the portion is displaced in relation to the fixed portion so that the air input lip has a second radial thickness in the second position which is less than the first radial thickness in the first position so as to promote a reverse thrust phase.

Abstract: An aeronautical thruster having a longitudinal axis and having a hub, an upstream annular row and a downstream annular row which are spaced apart from each another along the longitudinal axis, the upstream annular row being rotatable around the longitudinal axis, and the downstream annular row having a first blade and a second blade each extending in a radial direction to define a radial dimension between the hub and a radially outer end of the corresponding blade. When observed angularly around the longitudinal axis, the second blade can be positioned closer to an angular position at 3 o'clock or an angular position at 9 o'clock than the first blade, and the radial dimension of the first blade can be greater than the radial dimension of the second blade.

Abstract: An aeronautical thruster having a longitudinal axis and having a hub and at least two annular rows of unducted blades including an upstream annular row and a downstream annular row which are spaced apart from one another along the longitudinal axis, the upstream annular row being rotatable around the longitudinal axis, and the downstream annular row comprising a series of blades including a first blade and a second blade each extending in a radial direction from the hub to define a radial dimension between the hub and a radially outer end of the corresponding blade, wherein the radial dimension of the first blade is greater than the radial dimension of the second blade.

Abstract: The invention relates to a composite blade (5) for a turbine engine rotor, for example, an unducted propeller, comprising a skin (6) made of woven fibres forming the outer profile of the blade and an attachment (11) with cylindrical geometry in the direction of the span emerging from the blade root (9), which attachment is intended to retain the blade (5) on a hub of the rotor, characterised in that it further comprises a spar (16) having a hollow tube structure made of braided carbon fibres, fixed to the attachment (11) and extending inside the skin (6) over at least part of the span of the blade (5). The invention also relates to a propeller comprising said blade and to a method for manufacturing said blade.

Abstract: The present invention relates to a blade comprising: —a structure made of composite material; —a blade root attachment part comprising a wall delimiting a cavity, a first opening formed in the wall and a second opening located under the blade root portion, the structure made of composite material extending through the first opening; —two locking parts configured to axially abut against a shoulder of the attachment part and to bear against the blade root portion; and —a cover for compressing the blade root portion against the locking parts.

Abstract: A turbomachine of an aircraft comprising an outer casing delimiting with an inner hub, a flow path of a gas stream in which is disposed a low-pressure turbine configured to rotationally drive a low-pressure shaft; said turbomachine comprising, in the direction of flow of the gas stream, a first propeller; and a second propeller downstream of the first propeller, the first propeller being rotationally driven by said low-pressure shaft and the second propeller being rotationally driven by an electric motor, the second propeller being further disposed at a distance between 1.5 and 4 cord lengths from the first propeller defined between the respective axes of shimming of each of the first and second propellers.

Abstract: A propulsion assembly for an aircraft comprising a dual-flow turbine engine equipped with a fan, an aerodynamic outer shroud acting as a nacelle as well as a mounting pylon, the propulsion assembly having a secondary flow path defined by an outer radial defining surface formed by the shroud, the turbine engine including stator vanes, and the mounting pylon comprising a part housed in the secondary flow path, referred to as upstream part. According to the invention, the upstream part of the pylon extends radially from the inner radial defining surface, along a radial pylon height strictly less than a total radial height of the secondary flow path, and the upstream part of the pylon extends in the downstream direction from a root part of one of the stator vanes.

Abstract: An aircraft turbojet engine extending along an X axis and comprising a blower configured to provide a reverse thrust and a nacelle comprising an air intake which comprises at least one deflection member movably mounted between a deployed position in which the deflection member projects from the inner wall or from the lip of the air intake in a radially inward direction of deployment facing the X axis or in a longitudinal direction of deployment with respect to the X axis, in order to allow a release of the reverse air flow from the inner wall to support the reverse thrust phase, and a retracted position in which the air intake has an aerodynamic profile so as to guide the internal air flow along the inner wall in order to support the thrust phase.

Abstract: The present invention concerns a turbomachine hub, intended to be mounted so as to be able to rotate about a longitudinal axis (X) of the turbomachine, the hub comprising a main body (1) arranged around the longitudinal axis (X), an outer surface of which has a plurality of recesses; a plurality of platforms (3) comprising an outer surface delimited by a circular outer edge of radius r, each platform (3) being arranged in a corresponding recess of the plurality of recesses of the main body (1), and intended to receive a variable pitch blade (2), the pitch of which is variable according to a pitch change axis (Z), the platform (3) being centred and able to rotate about the pitch change axis (Z); characterised in that, for each platform, at least one part of the outer surface of the platform (3) and the main body (1) of the hub comprises a curvature defined by a same sphere portion of radius R and centre C, the at least one part being situated at the outer edge of the platform (3), in a junction zone between th

Abstract: Disclosed is a method for using an aircraft turbojet engine comprising an air inlet comprising a plurality of rectifier vanes, each rectifier vane being mounted such that it can move between a retracted position to assist the thrust phase and a deployed position in which the rectifier vane protrudes from the inner wall in a radially inward direction in order to rectify the reverse air flow of the inner wall to assist a thrust-reverse phase, in which method at least one rectifier vane is in the retracted position during a turbojet engine thrust phase, the method comprising, during a thrust-reverse phase of the turbojet engine, a step of moving the rectifier vane to the deployed position.

Abstract: The invention relates to a composite blade (5) for a turbine engine rotor, for example, an unducted propeller, comprising a skin (6) made of woven fibres forming the outer profile of the blade and an attachment (11) with cylindrical geometry in the direction of the span emerging from the blade root (9), which attachment is intended to retain the blade (5) on a hub of the rotor, characterised in that it further comprises a spar (16) having a hollow tube structure made of braided carbon fibres, fixed to the attachment (11) and extending inside the skin (6) over at least part of the span of the blade (5). The invention also relates to a propeller comprising said blade and to a method for manufacturing said blade.

Abstract: The present invention relates to a blade comprising: - a structure made of composite material; - a blade root attachment part comprising a wall delimiting a cavity, a first opening formed in the wall and a second opening located under the blade root portion, the structure made of composite material extending through the first opening; - two locking parts configured to axially abut against a shoulder of the attachment part and to bear against the blade root portion; and - a cover for compressing the blade root portion against the locking parts.

Abstract: The invention relates to a nubomachine with a longitudinal axis, comprising two, respectively upstream (122) and downstream, coaxial outer propellers (122), characterised in that at least some of the blades (148) of the upstream propeller (122) comprise at least one internal air circulation chimney (150) that communicates with air-bleeding openings (152) in tire boundary layers of the blades (148), and communicates with air outflow openings (158) on the radially outer end thereof, the air-bleeding openings (152) leading to opening inlets (152a) on tire passive surfaces (156) of the blades (148), the inlets (152a) of the air-bleeding openings being radially arranged in an area (H1) contained between 10% and 45% of the radial dimension (H2) of the blades (148), measured above turd from the radial height of the blades for which the tangent of the leading edge (138) of the blades is orthogonal to the longitudinal axis, and the inlets (152a) of the air bleeding openings being arranged in an area contained between

Abstract: The invention relates to a bypass turbomachine (2) for an aircraft, comprising a gas generator (5) and a ducted fan (4) comprising variable pitch blades (18) configured to take a reverse thrust position driving a reverse flow (24) of air within a secondary duct (16), the gas generator (5) being connected to a fan casing (3) by a stator blade assembly (40) that passes through the secondary duct, first openings (28) for letting in air from the reverse flow being located on an outer casing (17) at least partially internally delimiting the outer duct, and second openings (29) for letting said air out being located on an inner casing (14) at least partially externally delimiting an inner duct (12). The first openings are located within a plane that is perpendicular to a longitudinal axis (C) of the turbomachine and passes substantially through the middle of the blade assembly.

Abstract: The present invention concerns a turbomachine hub, intended to be mounted so as to be able to rotate about a longitudinal axis (X) of the turbomachine, the hub comprising a main body (1) arranged around the longitudinal axis (X), an outer surface of which has a plurality of recesses; a plurality of platforms (3) comprising an outer surface delimited by a circular outer edge of radius r, each platform (3) being arranged in a corresponding recess of the plurality of recesses of the main body (1), and intended to receive a variable pitch blade (2), the pitch of which is variable according to a pitch change axis (Z), the platform (3) being centred and able to rotate about the pitch change axis (Z); characterised in that, for each platform, at least one part of the outer surface of the platform (3) and the main body (1) of the hub comprises a curvature defined by a same sphere portion of radius R and centre C, the at least one part being situated at the outer edge of the platform (3), in a junction zone between th

Abstract: An aircraft turbojet engine extending along an X axis and comprising a blower configured to provide a reverse thrust and a nacelle comprising an air intake which comprises at least one deflection member movably mounted between a deployed position in which the deflection member projects from the inner wall or from the lip of the air intake in a radially inward direction of deployment facing the X axis or in a longitudinal direction of deployment with respect to the X axis, in order to allow a release of the reverse air flow from the inner wall to support the reverse thrust phase, and a retracted position in which the air intake has an aerodynamic profile so as to guide the internal air flow along the inner wall in order to support the thrust phase.

Abstract: A process for using an air input of a turboreactor nacelle of an aircraft, comprising an air input lip which comprises at least one fixed portion and at least one portion which can be moved between a first position, in which the air input lip has an aerodynamic profile so as to guide the internal air flow over the internal wall in order to promote a thrust phase, and a second position, in which the portion is displaced in relation to the fixed portion so that the air input lip has a second radial thickness in the second position which is less than the first radial thickness in the first position so as to promote a reverse thrust phase.

Abstract: Disclosed is a method for using an aircraft turbojet engine comprising an air inlet comprising a plurality of rectifier vanes, each rectifier vane being mounted such that it can move between a retracted position to assist the thrust phase and a deployed position in which the rectifier vane protrudes from the inner wall in a radially inward direction in order to rectify the reverse air flow of the inner wall to assist a thrust-reverse phase, in which method at least one rectifier vane is in the retracted position during a turbojet engine thrust phase, the method comprising, during a thrust-reverse phase of the turbojet engine, a step of moving the rectifier vane to the deployed position.

Abstract: The invention relates to a turbomachine of an aircraft comprising an outer casing (2) delimiting with an inner hub (3), a flow path (1) of a gas stream in which is disposed a low-pressure turbine configured to rotationally drive a low-pressure shaft; said turbomachine comprising, in the direction of flow of the gas stream, a first propeller (31); and a second propeller (32) downstream of the first propeller, the first propeller (31) being rotationally driven by said low-pressure shaft and the second propeller being rotationally driven by an electric motor (70), the second propeller (32) being further disposed at a distance between 1.5 and 4 cord lengths (LC1) from the first propeller (31) defined between the respective axes of shimming (A31, A32) of each of the first and second propellers.

Abstract: The invention relates to a vane (7) comprising: a structure made of composite material, a vane (7) root (16) attachment (9) comprising a base having a radial outer face (20) and in which a slot (24) configured to receive the vane (7) root (16) is formed and two platforms (26), extending on either side of the blade (18) opposite the radial outer face (20), and two stiffening parts (30), added and fixed on the radial outer face (20) on either side of the stilt (17) so as to tightly abut against the stilt (17).