A fly-by-wire system for a rotorcraft includes a computing device having control laws. The control laws are operable to engage a level-and-climb command in response to a switch of a pilot control assembly being selected. The level-and-climb command establishes a roll-neutral (“wings level”) attitude with the rotorcraft increasing altitude. The switch may be disposed on a collective control of the pilot control assembly (e.g., a button on a grip of the collective control). Selection of the switch may correspond to a button depress. The level-and-climb command may include a roll command and a collective pitch command. One or more control laws may be further operable to increase or decrease forward airspeed in response to pilot engagement of the level-and-climb command. The level-and-climb command may correspond to a go-around maneuver, an abort maneuver, or an extreme-attitude-recovery maneuver to be performed by the rotorcraft.

A vertical landing and take-off aircraft VTOL transitions from a vertical takeoff state to a cruise state where the vertical takeoff state uses propellers to generate lift and the cruise state uses wings to generate lift. The aircraft has an M-wing configuration with propellers located on the wingtip nacelles, wing booms, and tail boom. The wing boom and/or the tail boom can include boom control effectors. Hinged control surfaces on the wings, tail boom, and tail tilt during takeoff and landing to yaw the vehicle. The boom control effectors, cruise propellers, stacked propellers, and control surfaces can have different positions during different modes of operation in order to control aircraft movement and mitigate noise generated by the aircraft.

A method for optimizing the operation of at least one first propeller and of at least one second propeller of a hybrid helicopter. The method comprises the following step during a control phase: deflection, with an autopilot system, of at least one aerodynamic stabilizer member into a setpoint position having, with respect to a reference position, a target deflection angle that is a function of a setpoint deflection angle, the setpoint deflection angle being calculated by the autopilot system in order to compensate for a torque exerted by the lift rotor at zero sideslip.

Embodiments of the present invention are a flight control method and device, and an unmanned aerial vehicle. The method comprises firstly acquiring the current flight velocity of the unmanned aerial vehicle, then obtaining the current optimum inclination angle corresponding to the unmanned aerial vehicle according to the current flight velocity, and further adjusting the flight state of the unmanned aerial vehicle according to the current optimum inclination angle. The method can relieve the restrictions on the flight freedom of unmanned aerial vehicles and make the user experience rapid flight pleasure.

A method of operating an under-actuated actuator system including a plurality of actuators (3), preferably for operating a multiactuator aerial vehicle (1), wherein the actuators (3) are individual propulsion units of the multiactuator aerial vehicle (1), each actuator having a maximum physical capacity u.sup.max, the method including: controlling the actuators (3) by with an actual control input u∈.sup.k computed from an allocation equation u=D.sup.−1u.sub.p, wherein D.sup.−1 is an inverse allocation matrix and u.sub.p∈.sup.m is a pseudo control input defined by a system dynamics equation m(x){umlaut over (x)}+c(x,{dot over (x)})+g(x)+G(x)u.sub.p=f.sub.ext, wherein x∈n is an n-dimensional configuration vector of the system, m(x)∈.sup.n×n is a state dependent generalized moment of inertia, c(x,{dot over (x)})∈.sup.n are state dependent Coriolis force, g(x)∈.sup.n are gravitational forces and f.sub.ext∈.sup.n are external forces and torques, and G(x)∈.sup.n×m is a control input matrix which contains the information of under-actuation. The system is under-actuated if Rank

A system for flight control configured for use in an electric aircraft includes an inertial measurement unit (IMU) and configured to detect an aircraft angle and an aircraft angle rate. The system includes a flight controller including an outer loop controller configured to receive the input datum from the sensor, receive the aircraft angle from the IMU, and generate a rate setpoint as a function of the input datum. The system includes an inner loop controller configured to receive the aircraft angle rate, receive the rate setpoint from the outer loop controller, and generate a moment datum as a function of the rate setpoint. The system includes a mixer configured to receive the moment datum, perform a torque allocation as a function of the moment datum, and generate a motor command datum as a function of the torque allocation.

The present disclosure provides methods and systems for providing a lateral center of gravity of an aircraft on an aircraft display. A fuel distribution in the aircraft fuel tanks is determined. A lateral center of gravity of the aircraft is determined based on the fuel distribution. The lateral center of gravity is sent to the aircraft display. The present disclosure further provides an aircraft display for displaying the lateral center of gravity of an aircraft.

The present invention includes an a plurality of first variable speed motors mounted on a tail boom of the helicopter; one or more fixed pitch blades attached to each of the plurality of first variable speed motors; and wherein a speed of one or more of the plurality of first variable speed motors is varied to provide an anti-torque thrust.

A method includes generating, in a normal mode, an aircraft power setpoint and an aircraft pitch setpoint based on a desired airspeed setpoint and a desired altitude setpoint. The method includes transitioning from the normal mode to an underpower mode when the aircraft is unable to maintain the desired airspeed setpoint. The method includes setting, in the underpower mode, the aircraft power setpoint to a full power setting, generating, in the underpower mode, the aircraft pitch setpoint based on the desired altitude setpoint, and transitioning from the underpower mode to an underspeed mode when the aircraft airspeed is less than an airspeed threshold value while the aircraft power setpoint is set to the full power setting. The method includes maintaining, in the underspeed mode, the aircraft power setpoint at the full power setting and generating, in the underspeed mode, the aircraft pitch setpoint based on the desired airspeed setpoint.

An aeronautical vehicle that rights itself from an inverted state to an upright state has a self-righting frame assembly has a protrusion extending upwardly from a central vertical axis. The protrusion provides an initial instability to begin a self-righting process when the aeronautical vehicle is inverted on a surface. A propulsion system, such as rotor driven by a motor can be mounted in a central void of the self-righting frame assembly and oriented to provide a lifting force. A power supply is mounted in the central void of the self-righting frame assembly and operationally connected to the at least one rotor for rotatably powering the rotor. An electronics assembly is also mounted in the central void of the self-righting frame for receiving remote control commands and is communicatively interconnected to the power supply for remotely controlling the aeronautical vehicle to take off, to fly, and to land on a surface.