The present invention thus proposes an inlet cone for an aircraft turbine engine, comprising a frustoconical body and a tip made from elastically deformable material fixed to an end of smaller diameter of said body, the tip comprising a top configured to be situated on an axis of rotation of the cone and a fastening base for attachment on said end of said body. Said base extends in a connecting plane P. Said connecting plane P is inclined relative to said axis of rotation. Said base has a generally circular or oval shape. According to the invention, said connecting plane P is inclined relative to a transverse plane T perpendicular to said axis of rotation.

A triple-flow turbomachine for an aircraft, including a power transmission module including a torque input connected to a turbine shaft, a first torque output of a gearbox connected to a main shaft for rotatably driving a main fan propeller, and a second torque output of a planet gear connected to a secondary shaft for rotatably driving a secondary fan propeller. The planet gear is independent of the gearbox and arranged downstream of the gearbox.

An assembly adapted for use in a gas turbine engine includes a carrier and a blade track segment. The carrier extends at least partway about an axis. The blade track segment is supported by the carrier radially relative to the axis to define a portion of a gas path of the assembly.

The invention concerns a flow straightener unit (1) for a fan module of a turbomachine, the straightener unit (1) comprising a plurality of blades (2) distributed about an axis of rotation, each blade (2) is made of a composite material and comprises an aerofoil (21) and a root (22) intended to be assembled on a hub (4) of the turbomachine. The unit (1) comprises a centring and attachment plate (3) of the blade (2) on the turbomachine intended to be attached to the hub (4) at a determined azimuthal position and to the root (22) of the blade (2), the plate (3) is designed to be screwed to the hub (4) by screws (51a) that are longitudinal with respect to the axis of rotation of the unit (1) and screwed to the root (22) of the blade (2) by screws (52) that are radial with respect to the axis of rotation of the unit (1).

A mechanical reduction gear for turbomachine, in particular for aircraft. The reduction gear includes a sun gear having an axis of rotation, a ring gear which extends around the sun gear, planet gears which are meshed with the sun gear and the ring gear and which each includes a first toothing and a second toothing each including two series of teeth located on either side of a median plane, each of the planet gears being centred and guided in rotation by bearings, including an upstream bearing interposed axially between the series of upstream teeth of the second toothing and the plane, and a downstream bearing interposed axially between the series of downstream teeth of the second toothing and the plane.

A cast turbine nozzle includes an airfoil having a body including a suction side, a pressure side opposing the suction side, a leading edge spanning between the pressure side and the suction side, a trailing edge opposing the leading edge and spanning between the pressure side and the suction side, and a cooling cavity defined by an inner surface of the body. The nozzle also includes at least one endwall connected with the airfoil along the suction side, the pressure side, the trailing edge and the leading edge, and a plurality of heat transfer protrusions extending inwardly from the inner surface within the body, the plurality of heat transfer protrusions extending from the leading edge along the suction side and along the pressure side in a radially staggered columnar pattern. The inner surface includes a planar surface extending between adjacent heat transfer protrusions.

A shroud having an annular member extending along an axis from a first end to a second end. The annular member has an inside surface radially inward from an outside surface that diverges from the first end towards the second end relative to the axis. The shroud further includes a flange that extends radially outward from the outside surface and a recess, each located at the second end. The recess extends radially outward from the inside surface and circumferentially about the axis. An axial extent of the recess at least partially coincides with an axial extent of the flange.

A method of operating a combustor in a turbomachine is provided. The method includes a step of injecting a first central flow of liquid fuel from each outer liquid fuel cartridge in the first group of outer liquid fuel cartridges into a combustion zone at a first central angle. A first group of outer liquid fuel cartridges includes a first pilot tip. A second group of outer liquid fuel cartridges includes a second pilot tip. The method further includes a step of injecting a second central flow of liquid fuel from each outer liquid fuel cartridge in the second group of outer liquid fuel cartridges into the combustion zone at a second central angle. The first central angle is different than the second central angle.

A gas turbine engine includes a platform that has a gas path side, a non-gas path side, a first mate face, and a second mate face. The second mate face has a beveled edge sloping towards the first mate face. The gas turbine engine also includes a coverplate that includes a first bend, a flat portion substantially parallel to the first mate face and a first wing substantially parallel to the second mate face.

A fiber-reinforced component for use in a gas turbine engine includes a fiber sleeve forming a cooling channel and a plurality of fiber plies enclosing the fiber sleeve, with the plurality of fiber plies forming first and second walls separated by the fiber sleeve. The fiber-reinforced component further includes a matrix material between fibers of the fiber sleeve and the plurality of fiber plies.