The present disclosure relates generally to an aircraft with counter-rotating pusher props powered by a gas turbine engine having a power turbine disposed substantially perpendicular to the compressor, combustor and turbine gas generator power core axis, as well as to the aircraft longitudinal axis.

Systems and methods are disclosed for controlling the pitch angle of a propeller and rotor assembly that selectively limit the pitch angle according to a selected mode of operation. The system includes an actuator having forward and aft chambers, an oil transfer bearing (OTB), and a fine stop collar. The fine stop collar including a first passage in fluid communication with the OTB and forward chamber during a ground-based mode of operation, and a second fluid passage being in fluid communication with the OTB and forward chamber during a flight-based mode of operation.

An unducted thrust producing system, includes a rotating element, a stationary element. An inlet may be located forward or aft of the rotating element and the stationary element. An exhaust may be located forward, aft, or between the rotating element and the stationary element.

An aviation turbine engine having at least one unducted rotary propeller having a plurality of blades, each blade including: a blade body made of composite material including fiber reinforcement densified by a matrix, the fiber reinforcement of the blade body presenting three-dimensional weaving, the body extending between a leading edge and a trailing edge, and a protective fitting for protecting the leading edge and made of composite material having fiber reinforcement densified by a matrix, the fitting being adhesively bonded onto the leading edge of the blade body, the fitting being formed from a dry fiber preform injection molded with a densifying resin, and a polyurethane film for providing protection against erosion covering the blade body and the fitting.

A unshrouded vane assembly for an unducted propulsion system includes a plurality of vanes which have non-uniform characteristics configured to generate a desired vane exit swirl angle.

A propulsion assembly for an aircraft, the assembly including a turbojet having at least one unducted propulsive propeller, and an attachment pylon for attaching the turbojet to a structural element of the aircraft, the pylon being positioned on the turbojet upstream from the propeller and having a streamlined profile defined by two opposite side faces extending transversely between a leading edge and a trailing edge. The pylon includes a plurality of blow nozzles situated in the vicinity of its trailing edge and configured to blow air taken from a pressurized portion of the turbojet, the blow nozzles being positioned over at least a fraction of the trailing edge of the pylon that extends longitudinally facing at least a portion of the propeller. A method of reducing the noise generated by a pylon attaching a turbojet to an aircraft is presented.

A propulsion assembly for aircraft, the assembly including a turbojet having at least one unducted propulsion propeller; and an attachment pylon for attaching the turbojet to a structural element of the aircraft, the pylon being positioned on the turbojet upstream from the propeller and having an airfoil extending transversely between a leading edge and a trailing edge, the trailing edge of the airfoil of the pylon includes a cutout extending longitudinally over a fraction of the trailing edge facing at least a portion of the propeller, the cutout being configured to increase locally the distance between the trailing edge and the propeller, the cutout presenting an outline having a curved shape presenting at least two points of inflection.

Disclosed is a system for a gas turbine engine, the gas turbine engine comprising a primary shaft, the system including a rotor shaft; a plurality of components connected to the rotor shaft, including a wound field synchronous main machine (MM) and a permanent magnet generator (PMG); and wherein the PMG, alone or with the MM provide torque to change rotational speed of the rotor shaft, thereby changing rotational speed of the primary shaft.

A turbine engine is provided with a longitudinal rotation axis having at least one shaft with a radial axis, in particular a pitch change system for the blades of a propeller, said shaft traversing a radial passage of a substantially cylindrical case around the longitudinal axis. The turbine engine includes an annular oil guiding device around the radial shaft. The device has first and second annular parts nested in one another and secured to one another by hooping. The first part is secured by hooping to the radial shaft, and the second part is configured so as to be separated from the first part under the action of a force oriented along the radial axis and exerted on the second part by a member of the turbine engine forming a stop during a radial movement of the radial shaft.

Hybrid aircraft propulsors having electrically-driven augmentor fans are disclosed. An example apparatus includes a turbofan having a core engine and a ducted fan to be rotated via the core engine. The ducted fan includes a plurality of ducted fan blades arranged circumferentially around the core engine and circumscribed by a nacelle. The example apparatus further includes an augmentor fan having an augmentor hub ring and a plurality of augmentor fan blades. The augmentor fan blades are arranged circumferentially around the augmentor hub ring and project outwardly relative to an outer surface of the nacelle. The augmentor fan is to rotate separately from the ducted fan. The example apparatus further includes an electrical drive to rotate the augmentor hub ring in response to a supply of electrical energy provided to the electrical drive.