Abstract: A bypass turbojet engine including a low pressure shaft supported by at least two low pressure bearings, a high pressure shaft supported by at least two high pressure bearings, a fan shaft supported by at least two fan bearings, a reduction system coupling the low pressure shaft, with the fan shaft, enclosures housing the low pressure bearings, the high pressure bearings, the fan bearings and the reduction system, and a lubrication assembly including a closed oil circuit configured to supply oil to the enclosures in order to cool the bearings and the reduction system and at most five recovery pumps configured to recover oil from the enclosures.

Abstract: A fan of a turbine engine includes a disc provided with blades at the periphery thereof, the blades being mounted so as to pivot on the disc about a pivot axis, and a mechanism for changing the pitch of the blades. The mechanism is configured to adjust an angular position of each blade around the pivot axis. The angular position is in an angular setting range no greater than 20°.

Abstract: A multiflow turbojet engine generally includes an upstream fan driven by a gas generator having first and second coaxial compressors, an intake case forming a mounting for the rotors of the upstream fan and the first compressor, an inter-compressor case downstream from the intake case and forming a mounting for the rotors of the second compressor, and attachment means for thrust take-up control rods arranged in the inter-compressor case. The turbojet engine also includes a structural force shroud connecting the intake case to the inter-compressor case, of the and a floating first compressor case.

Abstract: Embodiments of the disclosure relate to a device for limiting overspeeding of a turbine shaft of a turbomachine, comprising structure for destroying the rotor blades of at least one turbine stage. The structure includes a casing and at least one projectile configured for projecting into the path of the rotor blades of the stage in order for the stage in question to be destroyed.

Abstract: A twin-spool turbojet engine includes: a low-pressure spool having a low-pressure compressor, a low-pressure turbine, and a low-pressure shaft connecting the low-pressure turbine to the low-pressure compressor; an exhaust casing through which exhaust gases flow downstream of the low-pressure turbine; and a low-pressure shaft thrust bearing having an inner ring and an outer ring, one of the rings being fastened on the low-pressure shaft and the other of the rings being fastened on the exhaust casing.

Abstract: A fan of a turbine engine includes a disc provided with blades at the periphery thereof, the blades being mounted so as to pivot on the disc about a pivot axis, and a mechanism for changing the pitch of the blades. The mechanism is configured to adjust an angular position of each blade around the pivot axis. The angular position is in an angular setting range no greater than 20°.

Abstract: Embodiments of the disclosure relate to a device for limiting overspeeding of a turbine shaft of a turbomachine, comprising structure for destroying the rotor blades of at least one turbine stage. The structure includes a casing and at least one projectile configured for projecting into the path of the rotor blades of the stage in order for the stage in question to be destroyed.

Abstract: A fan of a turbine engine, the fan including a fan disk provided with blades on its periphery, the blades being pivotally mounted on the disk around a pivoting axis and having a center of gravity, and a pitch change mechanism for the blades, each blade being configured so that its center of gravity is positioned on or at a small distance from a fictitious plane which passes through the pivoting axis of the blade and which is perpendicular to the axis of revolution of the fan when the blade is in a minimum drag position.

Abstract: The invention relates to a system for rotating a fan (2) of a turbojet engine (1) about a first rotation axis ?, including a reduction gear (3) made up of a planetary gearset placed at the centre of the fan (2), which includes: a central sun gear (31); an outer planet gear (33) attached to the fan (2); at least one satellite gear (32) arranged between the central sun gear (31) and the outer planet gear (33) in order to transmit a rotary movement between the central sun gear (31) and the outer planet gear (33); characterised in that the outer planet gear (33) includes an inner portion (33a) which meshes with the satellite gear (32), as well as an outer portion (33b) to which blades (21) of the fan (2) are directly attached.

Abstract: A bypass turbojet engine including a low pressure shaft supported by at least two low pressure bearings, a high pressure shaft supported by at least two high pressure bearings, a fan shaft supported by at least two fan bearings, a reduction system coupling the low pressure shaft, with the fan shaft, enclosures housing the low pressure bearings, the high pressure bearings, the fan bearings and the reduction system, and a lubrication assembly including a closed oil circuit configured to supply oil to the enclosures in order to cool the bearings and the reduction system and at most five recovery pumps configured to recover oil from the enclosures.

Abstract: A gas turbine engine, comprising a fan driven by a fan shaft and retractable axial retainer of the shaft mounted on a stator element and movable between a first operational position and a second non-operational position. The retainer may comprise a fluid supply configured for supplying at least one cavity with fluid. In some examples, the retainer is mounted movable or slidingly so as to generate movement of the retainer between the operation and non-operational positions.

Abstract: A turbofan including a low-pressure shaft; a high-pressure shaft; a fan shaft; a reducing mechanism that couples the low-pressure shaft to the fan shaft; and a maximum of six chambers that accommodate the bearings of the low-pressure shaft, the bearings of the high-pressure shaft and the bearings of the fan as well as the reducing mechanism.

Abstract: A multiflow turbojet engine generally includes an upstream fan driven by a gas generator having first and second coaxial compressors, an intake case forming a mounting for the rotors of the upstream fan and the first compressor, an inter-compressor case downstream from the intake case and forming a mounting for the rotors of the second compressor, and attachment means for thrust take-up control rods arranged in the inter-compressor case. The turbojet engine also includes a structural force shroud connecting the intake case to the inter-compressor case, of the and a floating first compressor case.

Abstract: A multi-flow turbojet engine generally includes an upstream ducted fan driven by a gas generator, the gas generator having a first and a second compressor, which are coaxial, an inlet case forming a support for the rotors of the upstream ducted fan and of the first compressor, an inter-compressor case downstream of the inlet casing and forming a support for the rotors of the second compressor, attachment means for thrust take-up rods, which attachment means are arranged on the inter-compressor case. The turbojet engine includes at least one axial stress transmission rod for connecting the inlet case to the inter-compressor case.

Abstract: A turbomachine module includes a rotor supporting variable pitch blades, and an inner annular shaft, a blade pitch control device, a device for feathering at least one blade comprising at least one gas generator placed between a body and a movable member. The control device and the feathering device are situated radially between the shaft and the blades, the movable member is coupled with the at least one blade and capable, under the action of the gas generated and/or released by the generator, of being moved into a reference position wherein it entails a feathered position to the at least one blade.

Abstract: According to the present invention there is provided a photovoltaic element, comprising a photovoltaic laminate; a structural substrate; wherein the photovoltaic element further comprises a plurality of hairs which are arranged to improve adhesion between said photovoltaic laminate and said structural substrate. There is also provided an intermediate product which is used in the manufacture of such a photovoltaic element and a method of manufacturing such a photovoltaic element.

Abstract: Aircraft turbine engine (10), comprising at least one first compressor, an annular combustion chamber (70) and at least one first turbine (46), which define a first flow duct (22) for a primary flow, characterised in that it comprises, between said combustion chamber (70) and said first turbine (46), a device (55, 55?) for discharging at least part of said primary flow.

Abstract: A fan of a turbine engine, the fan including a fan disk provided with blades on its periphery, the blades being pivotally mounted on the disk around a pivoting axis and having a center of gravity, and a pitch change mechanism for the blades, each blade being configured so that its center of gravity is positioned on or at a small distance from a fictitious plane which passes through the pivoting axis of the blade and which is perpendicular to the axis of revolution of the fan when the blade is in a minimum drag position.

Abstract: According to the present invention there is provided a photovoltaic element, comprising a photovoltaic laminate; a structural substrate; wherein the photovoltaic element further comprises a plurality of hairs which are arranged to improve adhesion between said photovoltaic laminate and said structural substrate. There is also provided an intermediate product which is used in the manufacture of such a photovoltaic element and a method of manufacturing such a photovoltaic element.