Abstract: A set of rotor disks for a turbine engine, or an airplane turboprop or turbojet, the set including disks rigidly connected together by bolts and shaped at their outer peripheries with slots for mounting blade roots. Each disk includes a rim extending radially and not having any axial flanges, two consecutive disks being connected together with help of an axially-extending ferrule that is fastened to the rims of the disks and that is fitted with a holder plate for holding the blade roots of the downstream disk.

Abstract: A method of fabricating a turbine engine drum including at least two rotor disks, the method including: positioning the two disks on a common axis spaced apart from each other and in a predetermined angular position relative to each other; placing a wall forming a body of revolution between the two disks coaxially about their common axis; driving the wall in rotation about its axis; and moving the disks axially towards the wall so as to put the disks into contact with ends of the wall and thus welding the disks to the wall by inertial friction welding.

Abstract: A blade retention disk includes, on its external surface, a peripheral annular groove configured to receive a plurality of blade roots of the hammer attachment type, the peripheral annular groove including a loading orifice intended for the introduction of blade roots inside the peripheral annular groove. The retention disk also includes: a crown covering the peripheral annular groove; a crown centering device to center the crown, the crown centering device and the crown forming an air supply enclosure from the bottom of the peripheral annular groove; at least one orifice opening in the air supply enclosure, the orifice being configured to introduce cooling air into the air supply enclosure. The blade retention disk finds a direct application in the field of low pressure turbines for aircraft turbine engines.

Abstract: A fan for an aircraft turbine engine, including an intake cone including a balancing system including a plurality of mass accommodation holes and at least one balancing mass mounted in one of the holes. The balancing system further includes an attachment mechanism exerting pressure that flattens the balancing mass against the bottom of its mass accommodation hole, the accommodation hole being made blind in the intake cone, so as to open out inside the intake cone.

Abstract: A one-piece rotor spool for a gas turbine engine extending along the axis of the engine is disclosed. The spool includes an upstream portion and a downstream portion that are radially offset relative to one another. The radially outer downstream portion includes a plurality of housings for rotor blades formed in the outer surface of the radially outer downstream portion. The radially inner upstream portion includes an attachment mechanism designed to interact with a rotor ring in order to prevent a tangential movement of the rotor ring relative to the radially inner upstream portion.

Abstract: A deoiler including a hub mounted on the vent shaft and a cover mounted on the hub between two stop faces which fix the cover avoiding the need to weld or bolt the cover.

Abstract: A set of rotor disks for a turbine engine, or an airplane turboprop or turbojet, the set including disks rigidly connected together by bolts and shaped at their outer peripheries with slots for mounting blade roots. Each disk includes a rim extending radially and not having any axial flanges, two consecutive disks being connected together with help of an axially-extending ferrule that is fastened to the rims of the disks and that is fitted with a holder plate for holding the blade roots of the downstream disk.

Abstract: A method of fabricating a turbine engine drum including at least two rotor disks, the method including: positioning the two disks on a common axis spaced apart from each other and in a predetermined angular position relative to each other; placing a wall forming a body of revolution between the two disks coaxially about their common axis; driving the wall in rotation about its axis; and moving the disks axially towards the wall so as to put the disks into contact with ends of the wall and thus welding the disks to the wall by inertial friction welding.

Abstract: A machining method for machining peripheral slots in a disk of a turbine of a turbine engine, such as an airplane turboprop or turbojet, the method including: machining the slots by reaming using a substantially cylindrical reamer of length greater than an axial thickness of the disk, and extending substantially parallel to the axis of the disk, outlines of the slots being machined by moving the reamer perpendicularly to its axis along the profile of each slot.

Abstract: A bearing support journal for an internal shaft of a gas turbine engine including a device for retaining a sealing sleeve surrounding the shaft is disclosed. The retaining device includes a full radial annular flange formed upstream on an axial cylindrical part of the journal, the full flange being intended to butt axially against a radial annular flange of the sealing sleeve; and a tangential immobilizing device for immobilizing the sealing sleeve with respect to the journal, the tangential immobilizing device being created downstream on the axial cylindrical part of the journal.

Abstract: A blade retention disk includes, on its external surface, a peripheral annular groove configured to receive a plurality of blade roots of the hammer attachment type, the peripheral annular groove including a loading orifice intended for the introduction of blade roots inside the peripheral annular groove. The retention disk also includes: a crown covering the peripheral annular groove; a crown centering device to center the crown, the crown centering device and the crown forming an air supply enclosure from the bottom of the peripheral annular groove; at least one orifice opening in the air supply enclosure, the orifice being configured to introduce cooling air into the air supply enclosure. The blade retention disk finds a direct application in the field of low pressure turbines for aircraft turbine engines.

Abstract: A deoiler including a hub mounted on the vent shaft and a cover mounted on the hub between two stop faces which fix the cover avoiding the need to weld or bolt the cover.

Abstract: A one-piece rotor spool (20) for a gas turbine engine extending along the axis of the engine, the spool comprising an upstream portion (20A) and a downstream portion (20B) that are radially offset relative to one another, the radially outer portion (20B) comprising a plurality of housings (24) for rotor blades formed in the outer surface of said radially outer portion (20B), the spool (20) being characterized in that the radially inner portion (20A) comprises attachment means (23) designed to interact with a rotor ring (30) in order to prevent a tangential movement of the rotor ring (30) relative to the radially inner portion (20A).

Abstract: A fan for an aircraft turbine engine, including an intake cone including a balancing system including a plurality of mass accommodation holes and at least one balancing mass mounted in one of the holes. The balancing system further includes an attachment mechanism exerting pressure that flattens the balancing mass against the bottom of its mass accommodation hole, the accommodation hole being made blind in the intake cone, so as to open out inside the intake cone.

Abstract: A bearing support journal (100) for an internal shaft of a gas turbine engine comprising a means of retaining a sealing sleeve surrounding said shaft, the retaining means comprising: a full radial annular flange (132) formed upstream on an axial cylindrical part (130) of the journal (100), said full flange (132) being intended to butt axially against a radial annular flange of the sealing sleeve, and tangential means (135) of immobilizing the sealing sleeve with respect to the journal (100), these being created downstream on the axial cylindrical part (130) of the journal (100). The invention also relates to assembly of a journal (100) and of a sealing sleeve.