An electrically powered rotary-wing aircraft with a first predetermined number of thrust producing units and a second predetermined number of batteries. Each one of the first predetermined number of thrust producing units may include a rotor, and an electrical drive unit with at least two electric motors. Each battery of the second predetermined number of batteries is coupled to at most one electric motor of the at least two electric motors of at least one of the first predetermined number of thrust producing units, and each electric motor of the at least one of the first predetermined number of thrust producing units is coupled to at most one of the second predetermined number of batteries.

An electrically powered rotary-wing aircraft with a first predetermined number of thrust producing units and a second predetermined number of batteries. Each one of the first predetermined number of thrust producing units may include a rotor, and an electrical drive unit with at least two electric motors. Each battery of the second predetermined number of batteries is coupled to at most one electric motor of the at least two electric motors of at least one of the first predetermined number of thrust producing units, and each electric motor of the at least one of the first predetermined number of thrust producing units is coupled to at most one of the second predetermined number of batteries.

A helicopter autopilot system includes an inner loop for attitude hold for the flight of the helicopter including a given level of redundancy applied to the inner loop. An outer loop is configured for providing a navigation function with respect to the flight of the helicopter including a different level of redundancy than the inner loop. An actuator provides a braking force on a linkage that serves to stabilize the flight of the helicopter during a power failure. The actuator is electromechanical and receives electrical drive signals to provide automatic flight control of the helicopter without requiring a hydraulic assistance system in the helicopter. The autopilot can operate the helicopter in a failed mode of the hydraulic assistance system. A number of flight modes are described with associated sensor inputs including rate based and true attitude modes.

A helicopter autopilot system includes an inner loop for attitude hold for the flight of the helicopter including a given level of redundancy applied to the inner loop. An outer loop is configured for providing a navigation function with respect to the flight of the helicopter including a different level of redundancy than the inner loop. An actuator provides a braking force on a linkage that serves to stabilize the flight of the helicopter during a power failure. The actuator is electromechanical and receives electrical drive signals to provide automatic flight control of the helicopter without requiring a hydraulic assistance system in the helicopter. The autopilot can operate the helicopter in a failed mode of the hydraulic assistance system. A number of flight modes are described with associated sensor inputs including rate based and true attitude modes.

An anti-torque rotor of a helicopter, having: a supporting body; a drive shaft which rotates about a first axis with respect to the supporting body; a hub connected operatively to drive shaft and angularly fixed with respect to first axis; at least one blade which is connected operatively to hub, is angularly fixed with respect to first axis, and is angularly movable with respect to a second axis to adjust the pitch angle of blade; and an actuator which can be operated to rotate blade about second axis to adjust the pitch angle of blade; actuator has an electric motor which generates torque along the first axis; and a mechanical stage interposed between the electric motor and blade, and designed to convert the torque into rotation of blade about the respective second axes; electric motor is fixed to supporting body.

An anti-torque rotor of a helicopter, having: a supporting body; a drive shaft which rotates about a first axis with respect to the supporting body; a hub connected operatively to drive shaft and angularly fixed with respect to first axis; at least one blade which is connected operatively to hub, is angularly fixed with respect to first axis, and is angularly movable with respect to a second axis to adjust the pitch angle of blade; and an actuator which can be operated to rotate blade about second axis to adjust the pitch angle of blade; actuator has an electric motor which generates torque along the first axis; and a mechanical stage interposed between the electric motor and blade, and designed to convert the torque into rotation of blade about the respective second axes; electric motor is fixed to supporting body.

A method of retrofitting a mechanically controlled aircraft with a fly-by-wire system includes removing a mechanical links between mechanical pilot inputs and actuators operable to drive flight surfaces. Electromechanical actuators are coupled between a plurality of vehicle management computers and the actuators. Each of the electromechanical actuators is operable to receive commands from the vehicle management computers and output a mechanical force to an input linkage of one of the actuators. Electromechanical pilot input modules are coupled to the mechanical pilot inputs. Each of the electromechanical pilot input modules is operable to convert a pilot-driven input force of an instance of the mechanical pilot inputs into an electronic signal indicative of the pilot-driven input force. At least one high performance computer is coupled to at least one of the vehicle management computers. The high performance computer executes one or more high level intelligence algorithms to selectively operate the aircraft autonomously.

A rotor blade pitch control system (15) comprising a rotor blade (19a, 19b, 19c, 19d) rotatable about both a central axis (20) and a pitch axis (24a, 24b, 24c, 24d), a pitch drive rotor (32a, 32b, 32c, 32d) rotatable about the central axis independently of rotation of the rotor blade about the central axis, a pitch follower (40a, 40b, 40c, 40d) rotatable relative to the pitch drive rotor, the pitch drive rotor and the pitch follower having an eccentric axis (33a, 33b, 33c, 33d), a linkage (50a, 50b, 50c, 50d) between the pitch follower and the rotor blade configured such that the pitch follower rotates with rotation of the rotor blade about the central axis, the pitch drive rotor, the pitch follower and the linkage configured such that the pitch drive rotor may be driven to control an angular displacement of the pitch drive rotor relative to the pitch follower about the central axis and thereby control the pitch of the rotor blade about the pitch axis.

A rotor blade pitch control system (15) comprising a rotor blade (19a, 19b, 19c, 19d) rotatable about both a central axis (20) and a pitch axis (24a, 24b, 24c, 24d), a pitch drive rotor (32a, 32b, 32c, 32d) rotatable about the central axis independently of rotation of the rotor blade about the central axis, a pitch follower (40a, 40b, 40c, 40d) rotatable relative to the pitch drive rotor, the pitch drive rotor and the pitch follower having an eccentric axis (33a, 33b, 33c, 33d), a linkage (50a, 50b, 50c, 50d) between the pitch follower and the rotor blade configured such that the pitch follower rotates with rotation of the rotor blade about the central axis, the pitch drive rotor, the pitch follower and the linkage configured such that the pitch drive rotor may be driven to control an angular displacement of the pitch drive rotor relative to the pitch follower about the central axis and thereby control the pitch of the rotor blade about the pitch axis.

Embodiments are directed to a flight control system for a helicopter comprises a pilot interface configured to receive a control input, at least one electronically controlled actuator, and a computing device configured to translate the control input to an actuator command, wherein the computing device is further configured to apply yaw rate limits to the actuator command to avoid loss of tail rotor effectiveness. The yaw rate limits are associated with a vortex ring state (VRS) envelope for a tail rotor of the helicopter. The electronically controlled actuator comprises a tail rotor actuator. The control input is a pedal input.