Abstract: To increase the inertia of a sealing lip of a blade for an aircraft turbine engine, and thus improve the service life of such a sealing lip, the sealing lip is conformed so as to have a trough in the outer surface thereof and a corresponding boss in the inner surface thereof, the trough and the boss being defined based on a connection cross section of the sealing lip to a blade body, and being formed at a distance from a free axial end of the sealing lip.

Abstract: To increase the inertia of a sealing lip of a blade for an aircraft turbine engine, and thus improve the service life of such a sealing lip, the sealing lip is conformed so as to have a trough in the outer surface thereof and a corresponding boss in the inner surface thereof, the trough and the boss being defined based on a connection cross section of the sealing lip to a blade body, and being formed at a distance from a free axial end of the sealing lip.

Abstract: A turbine engine has a blade with lower and upper surfaces. The blade has a root at a radially external end and includes a transverse head section in a plane perpendicular to a radial direction of the blade, taken at the radially external end with a first center of gravity. The root has a second center of gravity defined in a plane parallel to the transverse head section and transversely offset from the first center of gravity. The second center of gravity is defined in a predetermined zone at least partly demarcated by a V that is open towards the lower surface and includes a peak, the orthogonal projection of which on the transverse head section is located on the first center of gravity.

Abstract: A turbomachine rotor blade, said blade having, at the distal end thereof, a heel comprising: a platform (2) having a first edge (201) on the lower side and a second edge (202) on the upper side, at least one sealing member having a first end portion (301) on the lower side and a second end portion (302) on the upper side, said sealing member having a sealing top that extends radially outwards from said platform (2) between said first and second end portions (301, 302), characterized in that, for at least one sealing member, the heel (105) comprises, on at least at one of the edges (201, 202), a portion forming a bowl (5) extending along the end portion (301, 302) of the sealing member which corresponds to the edge (201, 202), the portion forming the bowl (5) being suitable for receiving a deposit of anti-wear material.

Abstract: A blank casting for producing a turbine engine rotor blade, with a vane connected by a platform to a root, is provided. The root has an axial machining allowance intended to be removed by machining from each of its axial extremities. At least one of the axial extremities of the root includes at least one transverse channel that extends over the entire transverse width of the root and whose depth is greater than the machining allowance. A process for obtaining the blade by machining this blank casting is also provided.

Abstract: A turbine engine blade includes an airfoil connected by a platform to a middle radial wall extending axially and prolonged radially inwards by a blade root that is mountable in a slot of a disk, and the platform and the middle radial wall defining two lateral cavities situated on either side of the middle radial wall and opening out circumferentially to receive scaling members. The cavities are open downstream with two fins projecting on either side of the middle radial wall at downstream end of the middle radial wall, the fins coming to bear against two adjacent teeth of the disk between which the slot is formed. The downstream end of the middle radial wall includes a setback extending between at least one rib of the platform and the fins.

Abstract: A turbomachine vane for cooperating with a vane retention disk including a plurality of peripheral teeth, the vane including a shank including a neck and a bulb for cooperating with a socket of the disk so as to radially hold the shank in the socket, the socket being a space bounded by two successive teeth of the disk, a pole on top of the neck, including two side walls facing the active surface side of the vane and two side walls facing the passive surface side of the vane, each wall having a rib at a lower end of the pole, the rib extending in the direction of the side wall facing the wall and providing a function of holding the vane when the vane is retained in the disk.

Abstract: A turbomachine rotor blade, said blade having, at the distal end thereof, a heel comprising: a platform (2) having a first edge (201) on the lower side and a second edge (202) on the upper side, at least one sealing member having a first end portion (301) on the lower side and a second end portion (302) on the upper side, said sealing member having a sealing top that extends radially outwards from said platform (2) between said first and second end portions (301, 302), characterised in that, for at least one sealing member, the heel (105) comprises, on at least at one of the edges (201, 202), a portion forming a bowl (5) extending along the end portion (301, 302) of the sealing member which corresponds to the edge (201, 202), the portion forming the bowl (5) being suitable for receiving a deposit of anti-wear material.

Abstract: A turbomachine vane for cooperating with a vane retention disk including a plurality of peripheral teeth, the vane including a shank including a neck and a bulb for cooperating with a socket of the disk so as to radially hold the shank in the socket, the socket being a space bounded by two successive teeth of the disk, a pole on top of the neck, including two side walls facing the active surface side of the vane and two side walls facing the passive surface side of the vane, each wall having a rib at a lower end of the pole, the rib extending in the direction of the side wall facing the wall and providing a function of holding the vane when the vane is retained in the disk.

Abstract: The invention relates to a turbine engine blade (1) comprising an airfoil (2) connected by a platform (3) to a middle radial wall (4) extending axially and prolonged radially inwards by a blade root (5) for mounting in a slot (6) of a disk (7), the platform (3) and said middle wall (4) defining two lateral cavities (16) situated on either side of said middle wall (4) and opening out circumferentially for the purpose of receiving sealing members. Said cavities (16) are open downstream with two fins (21) projecting on either side of said middle wall (4) at its downstream end, said fins (21) coming to bear against two adjacent teeth (19) of the disk (7) between which the slot (6) is formed. The downstream end (22) of the middle wall (4) including a setback (24) extending between said rib (20) and the fins (21).

Abstract: The invention relates to a blank casting for producing a turbine engine rotor blade, with a vane connected by a platform (4?) to a root (2?), said root having an axial machining allowance intended to be removed by machining from each of its axial extremities, characterised in that at least one of the axial extremities of the root (2?) comprises at least one transverse channel (14?) that extends over the entire transverse width of the root and whose depth is greater than said machining allowance. The invention also relates to the process for obtaining the blade by machining this blank casting.

Abstract: A device for injecting a mixture of air and fuel into a combustion chamber of a turbomachine is disclosed. The injection device has symmetry of revolution about an axis and includes radial swirl inducers of outside diameter D2, a bowl spaced axially from the radial swirl inducers, and a support ring for supporting the injection device, which is arranged around the bowl. The bowl has a cylindrical wall extended by a flared wall, the flared wall being provided with air intake holes. The support ring includes a first cylindrical portion of outside diameter D3, connected to a second cylindrical portion of outside diameter D4 spaced axially from the first cylindrical portion, the first cylindrical portion being connected to the cylindrical wall of the bowl by a plurality of support tabs.

Abstract: The present invention is concerned with an annular combustion chamber of a gas turbine engine with an external annular wall and an internal annular wall, including an upstream fairing for separating the gas stream at the inlet of the chamber into a combustion stream and a bypass stream which bypasses the inlet of the chamber, the fairing including an annular wall forming a cap, which includes a downstream portion for fastening to a wall of the chamber and an upstream portion forming an edge of the flow cross section for the combustion stream, wherein the upstream portion is continued into at least one additional downstream portion for fastening to the wall of the chamber.