A hydraulic assembly for an aircraft engine includes a pump with a housing and a rotor arranged inside the housing. The housing has at least one hydraulic line configured for the passage of a fluid; and a hydraulic transfer unit comprising: a hydraulic transfer block attached to the housing and comprising an internal fluid passage channel which communicates with the hydraulic line, and a casing which covers the transfer block and which is guided in rotation about an axis of revolution (Y) and about the transfer block, the casing having a fluid outlet port that communicates with the internal channel.

An assembly including a propeller blade and a system for angularly setting the pitch of the blade for an aircraft turbomachine, the blade having a root extending from an upper end connected to an airfoil of the blade to a free lower end, the root having a bulging segment, which bulging segment is referred to as the bulb, the system for angularly setting the pitch of the blade including a cup which is radially defined by an annular wall and which includes a lower bottom closed by a bottom wall and an upper opening through which the bulb is intended to be axially inserted into the cup, the bottom wall including a protrusion which engages with a cavity having a complementary shape in the free end of the root.

A trunnion-to-disk connection for use on an open fan configuration of a gas turbine engine may include an integral trunnion and blade spar inserted through a trunnion aperture of a fan disk and supported by top bearing and a bottom bearing. A cavity can be provided between a trunnion of the integral trunnion and blade spar and the fan disk, as well as between the top bearing and bottom bearing. Pressurized hydraulic fluid can be supplied to the cavity to urge the integral trunnion and blade spar in a direction to preload the bearings. Prior to pressurization, and prior to installation of the bottom bearing, the trunnion can be inserted into a trunnion aperture of the fan disk such that an end of the trunnion extends past the fan disk to provide sufficient space to insert the bottom bearing from within the open interior of the fan disk.

An assembly for an aircraft turbine engine is provided. The assembly generally includes an un-ducted propeller blade of which the root has a bulbous portion; a system for setting the pitch of the blade about a pitch setting axis, the pitch setting system including a cup configured to be placed inside the outer casing, has a flared open top end for axial insertion of the root and in which the root of the blade can be axially affixed by a locking member; and at least one radial retention prong in a radial space reserved between the wall of the cup and the root.

Disclosed is a variable pitch propeller blade for an aircraft turbine engine. The blade includes an airfoil connected to a root, the root having a body housed in an annular barrel that extends around a pitch axis of the blade, the body including a free end opposite the airfoil; a shoulder towards the airfoil; and a bulb located between the free end and the shoulder. The bulb can have a cross-section, referred to as the middle section, that has a maximum value Sm, and the barrel can be affixed to the body and covers and matches at least a portion of the bulb and the shoulder.

A gas turbine engine with split variable fan exit guide vanes including a fan duct supporting a circumferential pattern of split variable fan exit guide vanes, the split variable fan exit guide vanes comprising an upper section and a lower section, the upper section and the lower section each being adjustable about an axis extending along a span of each of the split variable fan exit guide vanes; an upper actuator in operative communication with the upper section, the upper actuator configured to independently adjust an incidence angle of the upper section responsive to predetermined gas turbine operating conditions; and a lower actuator in operative communication with the lower section, the lower actuator configured to independently adjust an incidence angle of the lower section responsive to the predetermined gas turbine operating conditions.

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.

A variable pitch fan for a gas turbine engine includes a hub, a fan blade supported by the hub, the fan blade including a blade root, and a retention system including a bearing and a retainer, the retainer securing the bearing to the blade root. The hub is rotatably supported by the bearing.

A variable pitch fan for a gas turbine engine includes a hub, a fan blade supported by the hub, the fan blade including a blade root, and a retention system including a bearing and a retainer, the retainer securing the bearing to the blade root. The hub is rotatably supported by the bearing.

A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by

[00001]$\frac{N_{\mathrm{FB}}{D}_{\mathrm{FT}}}{L_{\mathrm{AXIAL}}\left(\frac{R_{\mathrm{TB}}}{N_{\mathrm{FB}}}\right)}.$

N.sub.FB is a number of the fan blades, D.sub.FT is a fan tip diameter of the fan blades, R.sub.TB is a thrust bearing radius of the radial thrust bearings, and L.sub.AXIAL is an axial length from a fan hub tip to the fan bearings.