Abstract: A turbomachine includes a single ducted fan including a first shaft rotated by a second shaft via a speed reduction gearset, the second shaft being rotated by a third shaft of a turbine, the first shaft being guided in rotation with respect to a fixed structure via a first bearing and a second bearing placed upstream of the speed reduction gearset. The second shaft is guided in rotation with respect to the first shaft via a rolling bearing placed upstream of the speed reduction gearset, the rolling bearing comprising an outer ring housed in the first shaft, an inner ring connected to the second shaft and rolling elements arranged between the inner and outer rings.

Abstract: A method for regulating the temperature of the exhaust gases of a turbomachine, the method including regulation of the injection of fuel into a combustion chamber of the turbomachine so that the turbomachine generates a target thrust; regulation of the injection of mechanical power by an electric motor onto a shaft driven in rotation by a turbine, the electric motor being activated when a clearance between a casing and the blades of the turbine exceeds a threshold value.

Abstract: A drive shaft of an aircraft turbine engine includes a first portion and a second portion, and connecting means connecting the first and second portions and being configured to transmit a torque from the second portion to the first portion. The connecting means has at least one bellows with a first section having a diameter greater than the diameters of the first and second portions and second sections flanking the first section. The first section includes at least one fusible section with at least one through-hole and being configured to break when the value of a torque applied to the first portion exceeds a predetermined threshold value.

Abstract: A reducing assembly includes a planet carrier that is flexible and fixed, borne by a casing which surrounds it by two distinct connections with this casing, with a device for applying a torsional prestress to the planet carrier between these two connections. Prestressing the flexible planet carrier makes it possible to limit its level of mechanical stress when it is in service, in order to simplify its design and dimensioning.

Abstract: An impeller (230) for a planet carrier of a planetary gear speed reducer of a turbomachine, is configured intended to be rotatably secured to the planet carrier and to be rotated about an axis A of the speed reducer. The impeller has an annular shape about the axis and includes lubrication means (43, 45, 238), in particular for lubricating bearings of planet gears of the speed reducer. The lubrication means include an annular cavity (238) situated at the inner periphery of the impeller. The impeller includes an inner peripheral wall (246) closing the cavity (238) in the radial direction, and the impeller includes an annular port (248) that extends around the axis and that opens in the axial direction into the cavity in order to supply it with lubricating oil.

Abstract: The invention relates to a method for regulating an acceleration of a turbomachine, the turbomachine comprising a combustion chamber, a high-pressure turbine located downstream of the combustion chamber and which drives a high-pressure shaft in rotation, and a low-pressure turbine located downstream of the high-pressure turbine and which drives a low-pressure shaft in rotation, characterized in that the method comprises the following steps: Injecting (100) mechanical power onto the high-pressure shaft until a speed of the high-pressure shaft attains a target value, then Extracting (200) mechanical power from the high-pressure shaft so as to maintain the speed at the target value.

Abstract: Dynamic sealing device for an aircraft turbomachine, comprising two rotors (3, 8) which are configured to rotate about the one same axis (A) and different speeds, the said rotors extending at least in part around one another and dynamic sealing means (14) being mounted between the said rotors, characterized in that: —the said sealing means comprise annular rings (38) extending about the said axis, these rings being mounted with the ability to slide on a first (3) of the said rotors, in a radial direction with respect to the said axis, and being able to bear radially via their external periphery against a second (8) of the said rotors which at least partially surrounds the said first rotor, and in that it further comprises: —oil discharge means comprising an annular cavity (44) formed in the said second rotor, the said annular cavity opening radially towards the inside in the vicinity of the said sealing means and being in fluidic communication with oil passage ports (45) formed in the bottom of the cavity and

Abstract: A method for regulating the temperature of the exhaust gases of a turbomachine, the method including regulation of the injection of fuel into a combustion chamber of the turbomachine so that the turbomachine generates a target thrust; regulation of the injection of mechanical power by an electric motor onto a shaft driven in rotation by a turbine, the electric motor being activated when a clearance between a casing and the blades of the turbine exceeds a threshold value.

Abstract: The invention relates to a reducing assembly (26) comprising a planet carrier (27) that is flexible and fixed, borne by a casing (32) which surrounds it by means of two distinct connections with this casing (32), with means (44) for applying a torsional prestress to the planet carrier (27) between these two connections. Prestressing the flexible planet carrier (27) makes it possible to limit its level of mechanical stress when it is in service, in order to simplify its design and dimensioning.

Abstract: A drive shaft of an aircraft turbine engine includes a first portion and a second portion, and connecting means connecting the first and second portions and being configured to transmit a torque from the second portion to the first portion. The connecting means has at least one bellows with a first section having a diameter greater than the diameters of the first and second portions and second sections flanking the first section. The first section includes at least one fusible section with at least one through-hole and being configured to break when the value of a torque applied to the first portion exceeds a predetermined threshold value.

Abstract: A turbomachine includes a single ducted fan including a first shaft rotated by a second shaft via a speed reduction gearset, the second shaft being rotated by a third shaft of a turbine, the first shaft being guided in rotation with respect to a fixed structure via a first bearing and a second bearing placed upstream of the speed reduction gearset. The second shaft is guided in rotation with respect to the first shaft via a rolling bearing placed upstream of the speed reduction gearset, the rolling bearing comprising an outer ring housed in the first shaft, an inner ring connected to the second shaft and rolling elements arranged between the inner and outer rings.

Abstract: Aircraft turbine engine, including at least one first compressor, an annular combustion chamber and at least one first turbine, which define a first flow duct for a primary flow. Between the combustion chamber and the first turbine is a device for discharging at least part of the primary flow.

Abstract: An aircraft turbomachine has a longitudinal axis A and includes a reduction gear, at least one turbine shaft, an upstream end of which is connected to an input shaft of a reduction gear, a fan, one shaft of which is connected to an output shaft of the reduction gear, and dynamic sealing means between said reduction gear input shaft and the fan shaft. The sealing means includes an annular cowling, the axial cross-section of which is substantially U-shaped An outer annular leg of the annular cowling is fastened to the fan shaft, and an inner annular leg supports at least one annular sealing joint that cooperates with the reduction gear input shaft. The cowling is elastically deformable in the radial direction in relation to the axis A.

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 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 gas turbine engine includes 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 retractable axial retainer may include a fluid supply configured for supplying at least one cavity with fluid. In some examples, the retractable axial retainer is mounted movably or slidingly so as to generate movement of the retractable axial retainer between the operation and non-operational positions.

Abstract: Dynamic sealing device for an aircraft turbomachine, comprising two rotors (3, 8) which are configured to rotate about the one same axis (A) and different speeds, the said rotors extending at least in part around one another and dynamic sealing means (14) being mounted between the said rotors, characterized in that:—the said sealing means comprise annular rings (38) extending about the said axis, these rings being mounted with the ability to slide on a first (3) of the said rotors, in a radial direction with respect to the said axis, and being able to bear radially via their external periphery against a second (8) of the said rotors which at least partially surrounds the said first rotor, and in that it further comprises:—oil discharge means comprising an annular cavity (44) formed in the said second rotor, the said annular cavity opening radially towards the inside in the vicinity of the said sealing means and being in fluidic communication with oil passage ports (45) formed in the bottom of the cavity and p

Abstract: An aircraft turbomachine has a longitudinal axis A and includes a reduction gear, at least one turbine shaft, an upstream end of which is connected to an input shaft of a reduction gear, a fan, one shaft of which is connected to an output shaft of the reduction gear, and dynamic sealing means between said reduction gear input shaft and the fan shaft. The sealing means includes an annular cowling, the axial cross-section of which is substantially U-shaped An outer annular leg of the annular cowling is fastened to the fan shaft, and an inner annular leg supports at least one annular sealing joint that cooperates with the reduction gear input shaft. The cowling is elastically deformable in the radial direction in relation to the axis A.

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: 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.