Abstract: A sub-assembly for a low-pressure compressor of an aircraft turbine engine includes a straightener provided with cantilevered vanes and a rotor hub having a cavity covered by an inner shroud opposite the vanes. Orifices are made in the inner shroud to allow an air flow to circulate in the downstream to upstream direction of the low-pressure compressor.

Abstract: An adapter is configured to be disposed on the inlet of an oil tank of an aircraft engine. The adapter includes a plug between the adapter and the exterior, and a flapper between the adapter and the oil tank. An adapter cavity located between them is set to negative pressure during flights. When the aircraft is back on the ground, this negative pressure has to be counteracted to open the plug, which can make the opening operation difficult. The adapter includes a lever that counteracts the negative pressure in the cavity of the adapter to allow easy opening of the plug with limited force.

Abstract: A system for cooling oil in an aircraft turbine engine includes an intermediate support casing configured to be located between a low-pressure compressor and a high-pressure compressor of the aircraft turbine engine. The system further includes a heat exchanger for cooling the oil by heat exchange with air. The heat exchanger is at least partially integrated into the intermediate support casing.

Abstract: A turbomachine of the type with an unducted fan, comprising: a separation nozzle splitting an air flow into a primary flow and a secondary flow; a compressor compressing the primary flow; and an air/oil heat exchanger; wherein the exchanger is positioned in a channel through which a tertiary flow passes, the tertiary flow being drawn from the secondary flow upstream of the exchanger and meeting at least one annular row of rotor blades of the compressor downstream of the exchanger.

Abstract: A group of stator vanes for a row of stator vanes of a stage of a turbomachine compressor, the vanes of the group of stator vanes being close together and affording easy assembly. The group of stator vanes include a platform of a first vane having an aperture for accommodating the head of a second vane, the outer surface of the platform of the first vane being in contact with the inner surface of the platform of the second vane.

Abstract: A method and system control the thrust of an aircraft turbomachine having a high bypass ratio by direct action on a variable-pitch system. The variable-pitch system varies the pitch of the vanes of a stator of a low-pressure compressor for the open-loop control of the thrust of the turbomachine. The method also provides closed-loop control of the pitch of the blades of a propeller based on a rotational speed of the propeller.

Abstract: A system for cooling oil in an aircraft turbine engine includes an intermediate support casing configured to be located between a low-pressure compressor and a high-pressure compressor of the aircraft turbine engine. The system further includes a heat exchanger for cooling the oil by heat exchange with air, wherein the heat exchanger is at least partially integrated into the intermediate support casing.

Abstract: Module (60) for attracting and detecting ferromagnetic debris in an oil flow from a turbomachine, the module (60) comprising: a permanent magnet (62); a bar (64), the bottom (66) of which extends radially and is wound around a coil (70). The coil (70) is able to detect the magnetic field generated by the magnet (62) and in particular its variations when a ferromagnetic particle comes into the vicinity of the magnet (62).

Abstract: A sub-assembly for a low-pressure compressor of an aircraft turbine engine includes a straightener provided with cantilevered vanes and a rotor hub having a cavity covered by an inner shroud opposite the vanes. Orifices are made in the inner shroud to allow an air flow to circulate in the downstream to upstream direction of the low-pressure compressor.

Abstract: A group of stator vanes for a row of stator vanes of a stage of a turbomachine compressor, the vanes of the group of stator vanes being close together and affording easy assembly. The group of stator vanes include a platform of a first vane having an aperture for accommodating the head of a second vane, the outer surface of the platform of the first vane being in contact with the inner surface of the platform of the second vane.

Abstract: A system for disengaging a drive shaft from a bearing support. An assembly formed by the drive shaft and the bearing support is placed above a cylinder, which pushes on the drive shaft by passing by the central hole of the bearing support. The bearing support is driven upwards, but is blocked by a first and a second force transfer device, which allows the drive shaft to be disengaged from the bearing support.

Abstract: A method for manufacturing a turbine engine vane a root connected to a blade extending in a longitudinal direction includes the steps of providing a root; and providing mold with a first cavity and a second cavity that together define a recess in which the vane is formed. The recess includes a first space in which the blade is formed and a second space in which the root is formed. The method further includes the steps of providing aluminum strips; positioning a fibrous reinforcement; arranging the vane root in the second space; and injecting a foam comprising aluminum or injecting an aluminum alloy into the first space of the recess of the mold such that the foam impregnates the fibrous reinforcement.

Abstract: A sub-assembly for a low-pressure compressor of an aircraft turbine engine includes a straightener provided with cantilevered vanes and a rotor hub having a cavity covered by an inner shroud opposite the vanes. Orifices are made in the inner shroud to allow an air flow to circulate in the downstream to upstream direction of the low-pressure compressor.

Abstract: A stator vane assembly for a compressor of an aircraft turbine engine includes an inner shroud, an outer shroud and stator vanes. The stator vanes are attached only to the inner shroud and are in non-immobilizing mechanical contact with the outer shroud.

Abstract: A compressor casing for a turbine engine, the casing comprising: a housing comprising a radially inner e surface; an upstream ferrule; and a downstream ferrule; each of the ferrules having a respective inner surface capable of defining an air stream, and each of the ferrules having an outer surface facing the inner surface of the housing, the ferrules being attached to the cantilevered housing and being axially arranged so as to be separated by an axial clearance. Additionally, a turbine engine comprising such a casing and a stage of variable-setting vanes.

Abstract: The invention relates to a sandblasting mask (50) as well as its use for the sandblasting of blade tips in a method comprising a step of positioning a plurality of sandblasting masks (50), each blade passing through an opening (58) of a mask and each blade tip extending projecting from the respective mask, the masks being assembled two by two preferably without play; and a step of sandblasting the protruding blade tips of the masks. The masks (50) are preferably produced by plastic additive manufacturing and coated with a polyurea-based coating.

Abstract: A turbomachine compressor module, comprising an annular array of structural arms defining inter-arm spaces between two circumferentially adjacent arms; and an annular array of stator vanes with variable orientation about a respective axis (A) and disposed at least partially within the inter-arm spaces. The arms are provided with flaps pivotable about a respective axis (B). Also, a turbomachine provided with such a module and a row of rotor blades directly downstream of the module.

Abstract: A sub-assembly for a low-pressure compressor of an aircraft turbine engine includes a straightener provided with cantilevered vanes and a rotor hub having a cavity covered by an inner shroud opposite the vanes. Orifices are made in the inner shroud to allow an air flow to circulate in the downstream to upstream direction of the low-pressure compressor.

Abstract: A system for disengaging a drive shaft from a bearing support. An assembly formed by the drive shaft and the bearing support is placed above a cylinder, which pushes on the drive shaft by passing by the central hole of the bearing support. The bearing support is driven upwards, but is blocked by a first and a second force transfer device, which allows the drive shaft to be disengaged from the bearing support.

Abstract: A tooling holds a blade during friction welding to a rotor element. The tooling includes a fixed jaw having a central part; two arms separated by a distance d for receiving at least one portion of the blade; and a fixed jaw body with bearing surfaces and a bore for receiving at least one portion of the blade. The bearing surfaces are configured to come into contact with the blade. A movable jaw includes a movable bearing surface that comes into contact with the blade, and pressure means that moves the movable jaw towards the fixed jaw body to press the blade.