Systems and methods are described for governing the speed of a propeller on a propeller-based engine in an aircraft. The method comprises obtaining a synthesized or estimated blade angle for the propeller of the engine, determining one or more gain for a controller of the propeller based on the synthesized or estimated blade angle and one or more engine or aircraft parameter, determining a difference between a reference propeller speed and an actual propeller speed, applying the one or more gain to the difference via the controller in order to generate a command signal for controlling the propeller, and governing the propeller of the engine using the command signal.

A method of controlling an electric propulsion system of an aircraft. The propulsion system comprising an electric motor configured to drive a variable pitch propulsor. The method comprises determining a commanded thrust setting; determining one or more flight parameters; determining either a corresponding rotor governor speed and motor torque set-point, or a corresponding motor speed and rotor pitch angle set point, which provides the commanded thrust setting at the determined flight parameter having a maximum propulsion system efficiency; and controlling the rotor governor and electric motor in accordance with the determined respective set-points.

A system includes a thermal engine operatively connected to drive a propeller. An electric motor is operatively connected to the thermal engine to drive the propeller together with the thermal engine. An external input system is configured to accept input and output a thrust command. A protection function module is configured to enforce limits on the thermal engine, electric motor, and propeller. A low select module is operatively connected to receive input from the external input system and form the protection function module and to output the lower of input from the protection function module and external input system to the thermal engine, the electric motor, and the propeller.

Freezing of an electrical component of a VTOL rotor is prevented. The aircraft 100 includes a fuselage 12, a VTOL rotor 20 including one or more blades 23 that is supported on the boom 18 to be spaced apart from the fuselage for generating thrust in a vertical direction during take-off and landing, a motor 21 that is stored in the boom and is configured to cause the one or more blades to rotate, and an inverter 22 for controlling the motor, a detection unit 80 configured to detect a temperature of at least one apparatus of the motor or the inverter, and a control unit 99 configured to run the at least one apparatus based on a thrust request for a plurality of VTOL rotors and a detection result of the temperature.

A method and control system for an aircraft engine comprising a gas turbine driving a fan propeller with a mechanical gear-train and a dedicated pitch change mechanism for the fan propeller includes a fuel flow signal input; a pitch change mechanism signal input; a controlled plant for relating a pitch change mechanism pitch angle (BetaP) and a fuel flow (Wf) to at least two controlled outputs and a set of constraints. A decoupling control decoupling the controlled plant and/or the constraints into two separate single-input single-output (SISO) control loops for the first and second controlled outputs and a decoupling control decoupling the constraints from the decoupled controlled outputs and the constraints from one another provide gas turbine and fan propeller coordinate control while coordinately controlling constraints and outputs. A feedforward control can compensate the load change effect on engine speed and fan propeller rotor speed control.

A control device (1) of a flying object includes a generator (11), a driving source (12), a battery (13), an electric motor (3), a rotor blade (4), a battery status determination part (5), a variable pitch mechanism (6), and a pitch change control part (7). The electric motor (3) is driven by electric power supplied from at least one of the generator (11) and the battery (13). The rotor blade (4) is driven by the electric motor (3). The battery status determination part (5) determines a state of charge of the battery (13). The variable pitch mechanism (6) changes a pitch of the rotor blade (4). The pitch change control part (7) determines whether the pitch of the rotor blade (4) is changed based on a charging rate of the battery determined by the battery status determination part (5).

To reduce the overall mass of an engine assembly for aircraft, this assembly comprises a part of the fuselage of an aircraft, a turbomachine comprising an unfaired propeller, together with an annular row of unfaired after-guide vanes located aft of the propeller and rotationally fixed in relation to a longitudinal axis of the turbomachine, and a mounting pylon. At least part of the leading edge of the pylon is incorporated within the annular row between two after-guide vanes.

One aspect is a flight control system for a rotary wing aircraft that includes flight control computer configured to interface with a main rotor system, a translational thrust system, and an engine control system. The flight control computer includes processing circuitry configured to execute control logic. The control logic includes a primary flight control configured to produce flight control commands for the main rotor system and the translational thrust system. Main rotor engine anticipation logic is configured to produce a rotor power demand associated with the main rotor system. Propulsor loads engine anticipation logic is configured to produce an auxiliary propulsor power demand associated with the translational thrust system. The auxiliary propulsor power is combined with the rotor power demand to produce a total power demand anticipation signal for the engine control system.

An engine control device having a calculator for calculating a pitch setpoint for at least one propeller of the engine, the calculator taking account at least of a flight speed.

A pitch control system configured to vary a pitch angle of at least one of a plurality of propeller blades of a propeller system is provided including a switch movable between a neutral position and a plurality of non-neutral positions. Movement of the switch to a first non-neutral position generates a command to move the propeller blades in a first direction. Movement of the switch to a second non-neutral position generates a command to move the propeller blades in a second direction. Movement of the switch to a third non-neutral position generates a command to move the propeller blades to a zero thrust position.