Disclosed herein is a fan module with variable pitch blades for a propulsion unit, the fan module comprising a rotor carrying the blades and comprising an inner shaft and an outer casing defining between them a space, a control device for controlling the pitch of the blades, and a feathering device for feathering the blades. The rotation of the rotor is guided by a bearing, and the module further comprises means for recovering and guiding a liquid lubricant of the bearing, with the recovery and guidance means being configured to recover and guide said lubricant from an axial upstream end of the casing, axially from upstream to downstream and radially from the inside to the outside, under the centrifugal effect.

An unducted fan for an aircraft turbine engine, including a propeller having a hub and an annular array of blades extending substantially radially outwards from the hub, is provided. The hub includes an annular array of cowls, each of which is mounted between the radially inner ends of the blades and suitable for being fixed to an upstream collar and to a downstream collar of the hub. At least one cowl includes, at a circumferential end, a joint edge which is contiguous with a complementary joint edge of an adjacent cowl. The at least one cowl and at least one of the upstream and downstream collars are configured so that the cowl can be mounted and dismounted by a movement including a pivoting about an initial pivot axis circumferentially separated from the at least one joint edge.

A gas turbine engine may have a fan section that includes a variable pitch mechanism for changing the pitch of fan blades. The gas turbine engine may include a fan hub that is rotatable about an engine central longitudinal axis of the gas turbine engine and a blade receiver that is for holding a fan blade and that is rotatably coupled to the fan hub and may be rotatable about a radial axis of the gas turbine engine. The gas turbine engine may include a spring element disposed between the fan hub and the blade receiver to bias the blade receiver radially outward and an annular retaining ring may be disposed around the radial axis and between the fan hub and the blade receiver to limit radially outward movement of the blade receiver to prevent the blade receiver from decoupling and disengaging from the fan hub.

A gas turbine engine may have a blade receiver for holding a fan blade, wherein the blade receiver is rotatable about a radial axis of the gas turbine engine. The gas turbine engine may also include a variable pitch mechanism comprising an actuation arm. The gas turbine engine may also include a splined index ring disposed between the actuation arm of the variable pitch mechanism and the blade receiver. The splined index ring may include an outer spline and an inner spline.

A variable pitch fan for a gas turbine engine includes a fan blade defining a pitch axis and attached at a radially inner end to a trunnion mechanism. The variable pitch fan also includes a disk having a disk segment with the trunnion mechanism extending at least partially through the disk segment. A key is positioned at least partially in a key slot defined in a base of the trunnion mechanism, and further positioned adjacent to a support member of the disk segment. The key defines a first contact line between the key and the key slot and a second contact line between the key and the support member. The first and second contact lines respectively define a first contact reference line in a second contact reference line. The first and second contact reference lines define an angle with the pitch axis of the fan blade greater than zero degrees and less than ninety degrees.

The invention relates to a device comprising a fixed actuator for a control system for the orientation of fan blades of a turbomachine, comprising: at least one actuator body housed inside an annular casing of the turbomachine forming a bearing support, the annular casing being centred on a longitudinal axis of the turbomachine an inclined relative thereto, the actuator body having an internal volume in which a piston is able to slide, separating the internal volume into two chambers isolated from one another; and a plurality of actuator rods passing through the annular casing forming a bearing support and each having one end fixed to the piston and an opposite end intended to be connected to an internal ring of a movement transfer bearing in order to drive it in translation, the internal volume of the actuator body being closed downstream by a chamber bottom which forms a single part with the annular casing forming a bearing support.

The present invention relates to a vane (7) of a turbine engine, comprising: a structure which has an aerodynamic profile (20); a spar (21) comprising a vane root portion (24), a blade portion (25) arranged inside the structure which has an aerodynamic profile (20), and a stilt portion (26), the blade portion (25) comprising a body (27) connected to the vane root portion (24) and two arms extending radially from the body (27); and a structural reinforcement (31) extending from the vane root portion (24) to the body (27) of the blade portion (25), the structural reinforcement (31) being rigidly connected to the vane root portion (24) and configured to form a force path separate from the stilt portion (26) in the event of failure of the spar (21) within the stilt portion (26).

A gas turbine engine includes a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order; a fan defining a fan axis and comprising a plurality of fan blades rotatable about the fan axis; and a pitch change mechanism operable with the plurality of fan blades to control a pitch of the plurality of fan blades. The pitch change mechanism imparts a torque to each fan blade of the plurality of fan blades via first and second forces imparted in opposite directions to each fan blade of the plurality of fan blades.

A fan for a gas turbine engine is provided. The fan includes a plurality of fan blades, a disk, and a trunnion mechanism for attaching the fan blades to the disk. The disk can be formed of a plurality of individual disk segments, with the trunnion mechanism attaching one of the plurality of fan blades to a respective disk segment. A retention member is also provided. The retention member includes a means for catching a portion of the trunnion mechanism should a primary attachment system of the trunnion mechanism fail.

A gas turbine engine is provided. The gas turbine engine includes a core and a variable pitch fan arranged in flow communication with the core. The variable pitch fan has a disk and at least nine fan blades coupled to the disk for rotation together with the disk. The gas turbine engine further includes a rotatable nacelle covering the disk such that the engine has a fan hub radius ratio of between about 0.1 and about 0.4.