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: The invention proposes a geared turbomachine comprising a fan (16), a low pressure shaft (28) which drives the fan (16) rotationally around the main axis (A) of the turbomachine (10), a speed reducer (40) interposed between the fan (16) and the low pressure shaft (28), and an alternator (52) driven notably by the low pressure shaft (28), characterised in that the alternator (52) comprises an inductor (54) and an armature (56) of which one of the inductor (54) and the armature (56) is integral with the fan (16) and the other of the inductor (54) and the armature (56) is integral with the low pressure shaft (28).

Abstract: A turbomachine includes a fan shaft driven by a turbine shaft via a device for reducing a speed of rotation. The turbomachine includes an uncoupling device interposed between the reduction device and the turbine shaft. The uncoupling device is configured to uncouple the reduction device and the turbine shaft in response to the exceeding of a determined resistant torque exerted by the reduction device on the turbine shaft.

Abstract: A turbine engine comprising a fan comprising a first shaft driven in rotation via a reduction gear, the first shaft being guided in rotation with respect to a support of a fixed structure via an upstream bearing and a downstream bearing, the downstream bearing guiding in rotation the first shaft with respect to a downstream branch of the support, the first shaft and the support together defining an enclosure, the reduction gear comprising a ring gear fixed to the structure via a shell, the turbine engine comprising lubricator composed of at least one pipe for conveying a liquid lubricant, wherein the pipe is located in the enclosure, the pipe passing through an orifice made in the shell and an opening made in the downstream branch.

Abstract: A turbomachine having a fan shaft supported by a first bearing positioned downstream of the fan, the first bearing including an outer ring attached to an annular support secured to the stator. The turbomachine includes at least one gas feed duct leading into an enclosure positioned against the disc, the gas feed duct being adapted to be fed at a second end with pressurized gas taken from an airstream of a high-pressure compressor of the turbomachine, so that the gas applies an axial force towards upstream on the disc during some operating phases of the turbomachine. The turbomachine further includes a ferrule imperviously sealing the enclosure.

Abstract: A turbomachine (10) comprising a fan (12) shaft driven by a turbine shaft (16) by intermediary of a rotational speed reduction device (20), said turbomachine comprising decoupling means (28) interposed between the reduction device (20) and the fan shaft (18), which decoupling means are able to decouple the reduction device (20) and the fan shaft (18), characterized in that the reduction device (20) is coupled to the fan shaft via the intermediary of a coupling (17) having trapezoidal teeth of the curvic type which comprises said decoupling means (28), said decoupling means (28) being configured to decouple the reduction device (20) and the fan shaft (18) in response to a predetermined resistive torque (CD), referred to as the decoupling torque, being exceeded, which torque acts between the fan shaft (18) and the reduction device (20).

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 propulsive assembly including a turbine engine configured to be fastened to an aircraft by means of a suspension, having certain casing portions suitable for being decoupled so as to attenuate certain bending modes of the turbine engine in the event of an incident, the propulsive assembly comprising at least first and second casing portions (33a, 22) extending axially one after another, one of said casing portions being cantilevered out relative to said suspension, wherein a first casing portion (33a) possesses a first fastener portion (41) connected rigidly to the body (43) of the first casing portion (33a) and a second fastener portion (42) connected more flexibly to the body (43) of the first casing portion (33a), and wherein the second casing portion (22) is fastened to the second fastener portion (42) of the first casing portion (33a) in a manner that is permanent, and to the first fastener portion (41) of the first casing portion (33a) in a manner that is releasable.

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 propulsive assembly including a turbine engine configured to be fastened to an aircraft by means of a suspension, having certain casing portions suitable for being decoupled so as to attenuate certain bending modes of the turbine engine in the event of an incident, the propulsive assembly comprising at least first and second casing portions (33a, 22) extending axially one after another, one of said casing portions being cantilevered out relative to said suspension, wherein a first casing portion (33a) possesses a first fastener portion (41) connected rigidly to the body (43) of the first casing portion (33a) and a second fastener portion (42) connected more flexibly to the body (43) of the first casing portion (33a), and wherein the second casing portion (22) is fastened to the second fastener portion (42) of the first casing portion (33a) in a manner that is permanent, and to the first fastener portion (41) of the first casing portion (33a) in a manner that is releasable.

Abstract: The invention relates to a turbine engine comprising a fixed structure (7), a fan (1) rotor (3), having an axis (XX) of rotation, and emergency braking means (12, 13) of the rotor (3), in particular in the event of the loss of a blade of the fan (1), characterized in that said emergency braking means (12, 13) comprise a first (18) and a second (22) member supported by the fixed structure (7), said members (18, 22) respectively being configured so as to interact frictionally with a first (20) element which is complementary to the rotor (3), by forming a stop of the rotor (3) towards the rear along the axis (XX), and a second (24) element which is complementary to the rotor (3), by forming a stop of the rotor (3) towards the front along the axis (XX), when the emergency braking means (12, 13) are active.

Abstract: A turbojet engine includes a fan shaft driven by a turbine shaft via a device for reducing the speed of rotation. The engine includes an uncoupling device interposed between the reduction device and the turbine shaft. The uncoupling device is configured to uncouple the reduction device and the turbine shaft in response to the exceeding of a determined resistant torque exerted by the reduction device on the turbine shaft.

Abstract: The invention relates to a turbine engine comprising a fixed structure (7), a fan (1) rotor (3), having an axis (XX) of rotation, and emergency braking means (12, 13) of the rotor (3), in particular in the event of the loss of a blade of the fan (1), characterised in that said emergency braking means (12, 13) comprise a first (18) and a second (22) member supported by the fixed structure (7), said members (18, 22) respectively being configured so as to interact frictionally with a first (20) element which is complementary to the rotor (3), by forming a stop of the rotor (3) towards the rear along the axis (XX), and a second (24) element which is complementary to the rotor (3), by forming a stop of the rotor (3) towards the front along the axis (XX), when the emergency braking means (12, 13) are active.