Abstract: A turbomachine assembly and, in particular, a low-pressure compressor of an aircraft turbojet engine includes an annular row of upstream vanes with trailing edges extending radially from an upstream support; an annular row of downstream vanes with leading edges axially facing the trailing edges and extending radially from a downstream support; an annular passageway delimited by the upstream support and the downstream support. The downstream support has a profile with: an upstream portion delimiting the annular passageway forming an annular slide, a downstream portion axially at the level of downstream vanes, and a connecting arc connecting the upstream portion to the downstream portion. The connecting arc is arranged downstream of the leading edges.

Abstract: A turbomachine assembly and, in particular, a low-pressure compressor of an aircraft turbojet engine includes an annular row of upstream vanes with trailing edges extending radially from an upstream support; an annular row of downstream vanes with leading edges axially facing the trailing edges and extending radially from a downstream support; an annular passageway delimited by the upstream support and the downstream support. The downstream support has a profile with: an upstream portion delimiting the annular passageway forming an annular slide, a downstream portion axially at the level of downstream vanes, and a connecting arc connecting the upstream portion to the downstream portion. The connecting arc is arranged downstream of the leading edges.

Abstract: The present application relates to a rotor drum of a low-pressure compressor of an axial turbomachine. The drum includes a wall generally symmetrical in revolution about an axis and following a generally curved profile. The wall is configured to support a plurality of blade rows. A first blade row is formed by an annular platform integrally formed with the wall at the peak of its profile in relation to the axis; and a second blade row directly downstream of the first; and a third blade row directly upstream of the first; are formed by one or more blade-retaining grooves formed on the wall. The first blade row and the drum form an integral assembly, eliminating certain vibrations. Anchoring the blades is hybrid or mixed. The rotor may be mounted in a housing with ring-shaped stators.

Abstract: A rectifier stage of an axial turbine engine compressor is placed directly downstream of the input blower of a turbo-reactor or turboshaft engine compressor. The blade roots of the rectifier stage are housed inside the inner shroud by insertion into respective slots. The blade roots include at least one opening protruding from the slot of the shroud, so as to receive a spiral ring having at least two loops so as to prevent the blade roots from escaping from the respective slots of the inner shroud, and so as to form a connection between the blade roots.

Abstract: The present invention relates to a turbine engine rectifier stage comprising a plurality of fixed vanes connecting an inner ferrule support 19 to an outer ferrule support 14, said inner ferrule support and said outer ferrule support defining an aerodynamic vein and comprising a set of contiguously assembled sectors 20,15 respectively, the plurality of fixed vanes comprising so-called structural vanes 5 surrounding so-called non-structural or aero vanes 6 so as to form boxes 7, said structural vanes 5 being rigidly attached to the ends of the sectors of the inner ferrule support 20 by means of attachment elements 17 in order to impart increased stiffness to the boxes 7.

Abstract: The present invention relates to a method for manufacturing a turbomachine rectifier comprising an outer ring (5) and a plurality of stator vanes (2), said vanes comprising a platform (9) and a blade (12), wherein said method comprises at least the following steps: a) first folds (7) are draped over a core comprising perforations and having the shape of the aerodynamic flow; b) optionally, the first folds (7) are cut along incisions (8) placed opposite the perforations of the core; c) the blades (12) of the vanes or, alternatively, the blades (12) of the vane preforms, are inserted through the incisions (8) of the first folds (7) and through the perforations of the core; d) last folds (10) are draped over the platforms (9) thus forming a rectifier preform; e) a resin is injected into a closed mold with the preform and the resin-impregnated preform is polymerized; f) a molded piece, essentially having the shape and dimensions of said rectifier, is retrieved from the mold.

Abstract: The present invention relates to a turbine engine rectifier stage comprising a plurality of fixed vanes connecting an inner ferrule support 19 to an outer ferrule support 14, said inner ferrule support and said outer ferrule support defining an aerodynamic vein and comprising a set of contiguously assembled sectors 20,15 respectively, the plurality of fixed vanes comprising so-called structural vanes 5 surrounding so-called non-structural or aero vanes 6 so as to form boxes 7, said structural vanes 5 being rigidly attached to the ends of the sectors of the inner ferrule support 20 by means of attachment elements 17 in order to impart increased stiffness to the boxes 7.

Abstract: A rectifier stage of an axial turbine engine compressor is placed directly downstream of the input blower of a turbo-reactor or turboshaft engine compressor. The blade roots of the rectifier stage are housed inside the inner shroud by insertion into respective slots. The blade roots include at least one opening protruding from the slot of the shroud, so as to receive a spiral ring having at least two loops so as to prevent the blade roots from escaping from the respective slots of the inner shroud, and so as to form a connection between the blade roots.

Abstract: The present invention relates to a method for manufacturing a turbomachine rectifier comprising an outer ring (5) and a plurality of stator vanes (2), said vanes comprising a platform (9) and a blade (12), wherein said method comprises at least the following steps: a) first folds (7) are draped over a core comprising perforations and having the shape of the aerodynamic flow; b) optionally, the first folds (7) are cut along incisions (8) placed opposite the perforations of the core; c) the blades (12) of the vanes or, alternatively, the blades (12) of the vane preforms, are inserted through the incisions (8) of the first folds (7) and through the perforations of the core; d) last folds (10) are draped over the platforms (9) thus forming a rectifier preform; e) a resin is injected into a closed mold with the preform and the resin-impregnated preform is polymerized; f) a molded piece, essentially having the shape and dimensions of said rectifier, is retrieved from the mold.