Abstract: A turbofan engine has a primary duct, in which there flows a primary flow at a primary pressure and a secondary duct, which radially surrounds the primary duct and in which there flows a secondary flow at a secondary pressure. The primary duct includes at least one compressor configured to compress the primary flow, a turbine driving the compressor in rotation and a combustion chamber designed to receive, at an inlet, the primary air flow compressed by the compressor. The turbomachine further includes a cooling circuit extending between the compressor and the turbine. The cooling circuit has an air flow rate regulating device arranged upstream of the turbine and having at least one valve that is configured to move between an open position and a closed position, located between the compressor and the combustion chamber in the primary duct and the pressure (PS) in the secondary duct.

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: A shaft assembly (22, 32) for an aircraft turbine engine (1), comprising a first outer shaft (32) and a second inner shaft (22), the first outer shaft being intended to be engaged axially on the second shaft and comprising inner longitudinal splines (34) for coupling with outer longitudinal splines (24) of the second shaft, characterised in that, when the first and second shafts are in the coupling position, the inner and outer splines are engaged in one another and cooperate with each other in an axial coupling area (Z), the inner and outer teeth being situated outside this coupling area.

Abstract: A turbofan engine has a primary duct , in which there flows a primary flow at a primary pressure and a secondary duct, which radially surrounds the primary duct and in which there flows a secondary flow at a secondary pressure. The primary duct includes at least one compressor configured to compress the primary flow , a turbine driving the compressor in rotation and a combustion chamber designed to receive, at an inlet, the primary air flow compressed by the compressor The turbomachine further includes a cooling circuit extending between the compressor and the turbine.The cooling circuit has an air flow rate regulating device arranged upstream of the turbine and having at least one valve that is configured to move between an open position and a closed position, located between the compressor and the combustion chamber in the primary duct and the pressure (PS) in the secondary duct.

Abstract: An aircraft turbomachine with a reduction gear has a first shaft and a second shaft having one same axis of rotation, the second shaft being rotationally driven by the first shaft via the reduction gear, the first shaft having elastically deformable means having bellows section(s) and being connected to the reduction gear by a connecting system likewise having elastically deformable means involving a hairpin or bellows section(s).

Abstract: A turbine engine including a fan shaft driven by a turbine shaft with a rotational speed reduction device. A decoupling element is interposed between the reduction device and the fan shaft which operates to decouple the reduction device and the fan shaft. The reduction device is coupled to the fan shaft by a coupling having trapezoidal teeth of the curvic type which are used during the said decoupling operation. The decoupling element is configured to decouple the reduction device and the fan shaft in response to a predetermined resistive torque, referred to as the decoupling torque, which acts between the fan shaft and the reduction device.

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: 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: 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 vane (1) for a turbomachine rotor having a first blade (2) and a second blade (3). A platform (4) forms an angular wall sector in an axial direction (X), with the two blades (2, 3) connected to the platform (4). Two individual supports (12, 13), extend from the platform (4), and bear an attachment member (16, 17) at an end opposite the platform (4). The attachment member extends parallel to the axial direction (X) and has at least one surface (S1, S2) facing the platform (4) with a constant cross-sectional profile perpendicular to the axial direction (X) to form a bearing surface for retaining the vane (1) in a radial direction (R) oriented from the platform to the blades (2, 3).

Abstract: The invention concerns a power transmission system of a turbomachine, comprising: a speed reducer (12) comprising a sun gear (15) rotationally secured to a power shaft (5) with a longitudinal axis, an outer ring gear (18) rotating a rotor shaft along the longitudinal axis, and a planet carrier (17), and a device (40) for recovering oil ejected by centrifugal effect and comprising an annular gutter (41) for recovering the ejected oil, the gutter being attached to a fixed annular housing (26) and having a recovery chamber (42) and a first wall portion (43) disposed at least partially facing oil ejection means (30) of the speed reducer for directing the oil to the recovery chamber. According to the invention, the recovery chamber is provided with an inlet opening (45) directed radially outwardly and defined in a plane radially lower than a plane where an outlet port (33) of the ejection means is defined.

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 shaft assembly (22, 32) for an aircraft turbine engine (1), comprising a first outer shaft (32) and a second inner shaft (22), the first outer shaft being intended to be engaged axially on the second shaft and comprising inner longitudinal splines (34) for coupling with outer longitudinal splines (24) of the second shaft, characterised in that, when the first and second shafts are in the coupling position, the inner and outer splines are engaged in one another and cooperate with each other in an axial coupling area (Z), the inner and outer teeth being situated outside this coupling area.

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 with a reduction gear has a first shaft and a second shaft having one same axis of rotation, the second shaft being rotationally driven by the first shaft via the reduction gear, the first shaft having elastically deformable means having bellows section(s) and being connected to the reduction gear by a connecting system likewise having elastically deformable means involving a hairpin or bellows section(s).