Abstract: A turbine arrangement comprising a journal having a flange carrying a disc that is attached to the flange with bolts, wherein it includes a stationary circular trough surrounding the flange for collecting oil capable of passing radially between the disc and the flange carrying this disc.

Abstract: A double flow turbojet including: a low pressure compressor; a series of casings downstream of this low pressure compressor to delimit a primary flow path for circulating a primary stream, and including an upstream edge delimiting an inlet opening; a high pressure compressor in the primary flow path; a shroud surrounding the series of casings to delimit a flow path for circulating an intermediate stream, and having an upstream edge delimiting a circular inlet opening situated upstream of the high pressure compressor; a secondary flow path casing surrounding the shroud to delimit a secondary flow path for circulating a secondary stream; an exhaust casing including radial arms collecting the air coming from the intermediate flow path.

Abstract: A method for optimizing the energy management and reducing the greenhouse gas emissions of a complex aeronautical assembly comprising at least one aircraft and an auxiliary power unit (APU). The method analyzing, in a centralized manner outside the aeronautical assembly, data from the aeronautical assembly to compare at least one state of a parameter of the assembly with a predetermined optimal state of the parameter. The data measured by sensors of the aeronautical assembly are collected. The collected data is transmitted to a digital processing and analysis platform. The data is processed by the platform implementing machine learning algorithms. Information relating to the processed data is displayed on a dashboard accessible via different terminals. A real-time alert is generated in the case of detection of an anomaly in the aeronautical assembly.

Abstract: A ventilated chamber (19), subject to an irruption of air at a pressure higher than that in a surrounding chamber (20) through labyrinth seals separating them and evacuation of excess air through an evacuation duct (30), comprises a motor-driven compressor (32) in the evacuation duct to facilitate suction and evacuation, even when the engine speed does not allow a large overpressure in the chamber (20).

Abstract: A mechanical decoupler (15) at the inlet to a turbomachine is positioned on the outside of an intake casing, where radiating arms (27) meet an external casing (13) so as to partially unload a low-pressure shaft when a significant out-of-balance appears. Because it is positioned a long way from the bearing, the decoupler (15) can be designed with a greater degree of freedom at a location where there is more space available and where layout constraints are less of an issue. More specifically, it is housed in a cavity (30) of the external casing (13) which opens onto the flow path (5).

Abstract: The invention relates to a turbojet bearing housing including a fixed envelope traversed by a rotor, this envelope including a cylindrical end surrounding a seal ensuring the leak tightness of the cylindrical end of the envelope with the rotor, this envelope being equipped with a shell screwed onto the cylindrical end thereof, this shell including radial channels emerging opposite the seal while being arranged so that an air stream traversing the seal towards the housing mainly comes from these radial channels.

Abstract: A turbojet includes an inlet casing, a low pressure compressor, a high pressure spool, a low pressure turbine, and an exhaust casing. The inlet casing and the low pressure compressor define an air inlet channel that divides downstream into a primary flow channel and a secondary flow channel. A low pressure shaft is supported by a first roller bearing, borne by the inlet casing, a second bearing borne by the intermediate casing, a third bearing borne by the exhaust casing, and an additional ball bearing borne by the intermediate casing and arranged between the first bearing and the second bearing. Such a configuration of bearings makes it possible to withstand an increased level of load and to be compatible with an increase in the rate of dilution.

Abstract: An aircraft propulsion assembly, including a turbine engine comprising at least one gas generator configured to generate a main flow, which is supplied by a central jet to at least one power turbine, the central jet being surrounded by an outer fairing, and the power turbine driving, on the periphery thereof, at least one fan rotor. The aircraft propulsion assembly comprises first movable means which are arranged so as to divert at least some of the main flow from the central jet to the outside of the outer fairing and preferably upstream of the turbine engine so as to generate thrust reversal. An aircraft which uses the propulsion assembly, particularly on the rear tip of the fuselage of the aircraft.

Abstract: The present invention relates to an aircraft comprising a fuselage (1) and a propulsion unit at the rear of the fuselage, the propulsion unit comprising at least one fan rotor (7, 8), a nacelle (14) fairing the fan and at least one connection means (15) connecting the nacelle to the fuselage, the fan being rotated by the energy supplied by at least one gas-turbine gas generator (2a, 2b) housed in the fuselage, said gas generator comprising auxiliary equipment cooled by a cooling circuit. The aircraft is characterised in that said cooling circuit comprises at least one heat exchanger exchanging heat with the ambient air housed in one of said connection means (15) and/or in said nacelle (14). The cooling circuit optionally comprises also a heat exchanger exchanging heat with the ambient air, housed in the tail unit.

Abstract: A ventilated chamber (19), subject to an irruption of air at a pressure higher than that in a surrounding chamber (20) through labyrinth seals separating them and evacuation of excess air through an evacuation duct (30), comprises a motor-driven compressor (32) in the evacuation duct to facilitate suction and evacuation, even when the engine speed does not allow a large overpressure in the chamber (20).

Abstract: The invention relates to a turbomachine (1), such as an aircraft turbojet or turboprop, extending along an axis (X), comprising a primary flow jet (2) comprising at least one compressor (4, 5), at least one combustion chamber (6) and at least one turbine (7, 8), and a secondary flow jet (3) located radially outside the primary flow jet (2), at least one duct (18) for the flow of cooling air, said duct (18) comprising an air inlet opening into the primary flow jet (2), at the compressor (4, 5), and an air outlet opening into the primary flow jet (2), at the turbine (7, 8).

Abstract: A turbojet includes an inlet casing, a low pressure compressor, a high pressure spool, a low pressure turbine, and an exhaust casing. The inlet casing and the low pressure compressor define an air inlet channel that divides downstream into a primary flow channel and a secondary flow channel. A low pressure shaft is supported by a first roller bearing, borne by the inlet casing, a second bearing borne by the intermediate casing, a third bearing borne by the exhaust casing, and an additional ball bearing borne by the intermediate casing and arranged between the first bearing and the second bearing. Such a configuration of bearings makes it possible to withstand an increased level of load and to be compatible with an increase in the rate of dilution.

Abstract: The invention relates to a turbojet bearing housing including a fixed envelope traversed by a rotor, this envelope including a cylindrical end surrounding a seal ensuring the leak tightness of the cylindrical end of the envelope with the rotor, this envelope being equipped with a shell screwed onto the cylindrical end thereof, this shell including radial channels emerging opposite the seal while being arranged so that an air stream traversing the seal towards the housing mainly comes from these radial channels.

Abstract: The invention relates to a double flow turbojet including: a low pressure compressor; a series of casings downstream of this low pressure compressor to delimit a primary flow path for circulating a primary stream, and including an upstream edge delimiting an inlet opening; a high pressure compressor in the primary flow path; a shroud surrounding the series of casings to delimit a flow path for circulating an intermediate stream, and having an upstream edge delimiting a circular inlet opening situated upstream of the high pressure compressor; a secondary flow path casing surrounding the shroud to delimit a secondary flow path for circulating a secondary stream; an exhaust casing including radial arms collecting the air coming from the intermediate flow path.

Abstract: A mechanical decoupler (15) at the inlet to a turbomachine is positioned on the outside of an intake casing, where radiating arms (27) meet an external casing (13) so as to partially unload a low-pressure shaft when a significant out-of-balance appears. Because it is positioned a long way from the bearing, the decoupler (15) can be designed with a greater degree of freedom at a location where there is more space available and where layout constraints are less of an issue. More specifically, it is housed in a cavity (30) of the external casing (13) which opens onto the flow path (5).

Abstract: A turbine engine impeller comprises a disc bearing vanes and an annular plate. The roots of the vanes are fitted in grooves and the root of each vane is connected to a blade by an integrated platform. The platforms of said vanes cooperate to form a radially inner wall of a duct for a flow passing through the blades. The annular plate comprises a transverse annular portion forming an axial stop of the roots of the vanes. The integrated platform of each vane has no spoiler and stops axially to the right of transverse faces of the root. The transverse annular portion of said annular plate bears, between the integrated platforms and the disc, an annular spoiler configured to extend in an axial space between the impeller and an adjacent module of the turbine engine so that said annular plate bears annular knife-edge seals of a labyrinth sealing joint.

Abstract: The present invention relates to an aircraft comprising a fuselage (1), a propulsion unit (7) at the rear tip of the fuselage and at least two engines (3, 5) each driving an engine shaft (32, 52), the propulsion unit including at least one propeller (71, 73) driven mechanically by the engine shafts (32, 52), characterized in that the propulsion unit (7) includes two propellers (71, 73), each of the propellers being driven by a propeller shaft (72, 74), the two propeller shafts being driven by the engine shafts (32, 52) via a first differential epicyclic gear train (40). Specifically, the input of the first differential gear train (40) is connected to the output of a second differential epicyclic gear train (60), the second differential gear train (60) including two inputs (61c2, 62c2), each of said two inputs being connected mechanically to one of said engine shafts (32, 52).

Abstract: A rotor for a turbine engine includes a disk having cavities called primary cavities at its periphery; a plurality of blades each having a root of which the lower part is composed of a bulb locked axially in the primary cavities; a plurality of added platforms, each being arranged between two consecutive blades, wherein the platforms have: a substantially straight plate and a bulb extending radially under the plate, the bulb being locked axially in the secondary cavities arranged at the periphery of the disk, the secondary cavities being positioned between two consecutive primary cavities; a spoiler extending in the axial direction, the spoiler forming an annular sector facing at least two consecutive blades.

Abstract: The present invention relates to an aircraft comprising a fuselage (1), a propulsion unit (7) at the rear tip of the fuselage and at least two engines (3, 5) each driving an engine shaft (32, 52), the propulsion unit including at least one propeller (71, 73) driven mechanically by the engine shafts (32, 52), characterised in that the propulsion unit (7) includes two propellers (71, 73), each of the propellers being driven by a propeller shaft (72, 74), the two propeller shafts being driven by the engine shafts (32, 52) via a first differential epicyclic gear train (40). Specifically, the input of the first differential gear train (40) is connected to the output of a second differential epicyclic gear train (60), the second differential gear train (60) including two inputs (61c2, 62c2), each of said two inputs being connected mechanically to one of said engine shafts (32, 52).

Abstract: The invention relates to a turbomachine rotor including a disk and at least one blade wherein the root comprises a bulb accommodated into a groove associated with the disk, wherein the bulb includes a slot that opens up radially inwards at its lower face, and in that the rotor includes a retaining part extending partly into the groove projecting radially outwards from the bottom face of the groove, of which a radially outer end of the retaining part fits into the slot of the bulb and bears radially inwards on the inner wall of the slot of the bulb.