A turbofan engine defining a circumferential direction is provided. The turbofan engine includes: a turbomachine, the turbomachine defining a pylon attachment location along the circumferential direction; an unducted rotor assembly drivingly coupled to the turbomachine, the unducted rotor assembly including a plurality of unducted rotor blades; and a plurality of outlet guide vanes positioned downstream of the plurality of unducted rotor blades, the plurality of outlet guide vanes including N.sub.OGV number of outlet guide vanes, the plurality of outlet guide vanes including a first outlet guide vane and a second outlet guide vane adjacent the first outlet guide vane, the first and second outlet guide vanes defining a gap spacing greater than 360 degrees divided by N.sub.OGV, and the pylon attachment location positioned outside of between the first and second outlet guide vanes along the circumferential direction.

A rear part for a turbojet engine includes a variable-geometry nozzle in which a convergent flap is provided with a lever supporting a bearing roller arranged axially between an upstream bearing wall and a downstream bearing wall, which are rigidly secured to a movable part capable of moving axially on command with respect to an upstream stator structure in such a way that the bearing roller is free to move with respect to the upstream and downstream bearing walls in a radial direction. During a downstream movement of the movable part, the upstream bearing wall pushes the lever to pivot a downstream end of the convergent flap towards the longitudinal axis. Operation of the drive mechanism for moving the movable part can thus be optimized while limiting the size and the mass of the mechanism for controlling the flaps.

An exemplary gas turbine engine includes a turbine section and a fan mechanically connected to the turbine section such that rotation of the turbine drives rotation of the fan. The fan includes a hub, a plurality of blade bodies extending radially outward from the hub to a first partial shroud, and a plurality of blade tips extending radially outward from the partial shroud.

A turbine stator vane assembly comprises a plurality of circumferentially arranged struts and a plurality of circumferentially arranged stator vanes. Each stator vane has a leading edge, a trailing edge and a chord length. Each strut has a leading edge, a trailing edge, a chord length, a pressure surface and a suction surface. The chord length of the struts is greater than the chord length of the stator vanes. The stator vanes are arranged in groups between adjacent pairs of struts and each group of stator vanes comprises a plurality of stator vanes. The circumferential spacing between adjacent struts is substantially the same and each stator vane in a group of stator vanes has a different chord length to the other stator vanes in that group of stator vanes.

An aircraft engine including a core casing extending circumferentially about an axis and defining a core flow passage, a nacelle located radially outward from and around the casing, and a bypass passage defined between the nacelle and the casing. A mixer includes a peripheral wall having a leading edge and a trailing edge, the leading edge attached to the casing. A first mixer portion and a second mixer portion are circumferentially spaced apart and respectively extend away from the leading edge from a first location to a second location. The first and second mixer portions respectively have a first and second stiffness, the first stiffness greater than the second stiffness and greater than a stiffness of an upstream portion of the wall extending from the first mixer portion at the first location toward the leading edge.

An airfoil includes an airfoil body having a pressure side and a suction side that each radially extend between a first radial boundary end and a second radial boundary end and each axially extend between a leading edge and a trailing edge. The airfoil body defines a plurality of first skin core passages disposed proximate the suction side and radially extend from the first radial boundary end towards the second radial boundary end, and a plurality of second skin core passages that radially extend toward the second radial boundary end and circumferentially extend towards the suction side proximate the second radial boundary end.

A turbine engine, a turbojet, or a turboprop engine, including an annular combustion chamber defined by an inner shell and an outer shell, a turbine distributor arranged downstream of the combustion chamber, with a downstream end of the outer shell and/or of the inner shell of the chamber including a radial rim arranged opposite a radial rim of an upstream end of the distributor, a sealing mechanism including at least one strip extending between the rims to provide a seal between the combustion chamber and the distributor. The sealing strip extends axially and circumferentially between the rims and bears sealingly against free ends of the rims.

A turbofan engine defining a circumferential direction is provided. The turbofan engine includes: a turbomachine, the turbomachine defining a pylon attachment location along the circumferential direction; an unducted rotor assembly drivingly coupled to the turbomachine, the unducted rotor assembly including a plurality of unducted rotor blades; and a plurality of outlet guide vanes positioned downstream of the plurality of unducted rotor blades, the plurality of outlet guide vanes including N.sub.OGV number of outlet guide vanes, the plurality of outlet guide vanes including a first outlet guide vane and a second outlet guide vane adjacent the first outlet guide vane, the first and second outlet guide vanes defining a gap spacing greater than 360 degrees divided by N.sub.OGV, and the pylon attachment location positioned outside of between the first and second outlet guide vanes along the circumferential direction.

A control ring for a stage of variable-pitch vanes for a turbine engine includes at least one member for bearing on a casing and means for fixing the member. The member includes a bushing with an axial bore for the passage of the member or an element for supporting the member. A through-slot opens into the bore and allows for substantial radial deformation of the bushing. The member further include an outer thread for screwing the bushing into a complementary thread of a hole in the body, thereby deforming the bushing. The member or support can be mounted and moved inside the bore, to a second position, in which the bushing is radially constrained and is tightly mounted on the member or support, which is thus immobilized in relation to the bushing.

The present application provides a gas turbine engine. The gas turbine engine may include a compressor discharge casing, a number of combustors configured in an annular array, and a number of aft mounting assemblies. An aft mounting assembly mounts a combustor to an inner diameter of the compressor discharge casing.