Abstract: There is disclosed in one example an inner loop controller for an aircraft flight computer, including: a stall protection circuit to compute, for an attitude angle ?, an attitude limit ?max as a function of a flight path angle (?) and an angle of attack limit (?max); a transfer function circuit to convert ? to an attitude rate {dot over (?)}, wherein {dot over (?)} is a time derivative of ?; and a load protector circuit to compute a limit on {dot over (?)} ({dot over (?)}max) as a function of a load factor limit (Nz,max) and a true airspeed (v).

Abstract: A system can include a flight controller for an aircraft that includes an electric motor that drives blades with a variable pitch, where the flight controller receives a command to change a flight characteristic of the aircraft and creates a torque command and a revolutions per minute (RPM) command. The system can also include a propulsion assembly, where the propulsion assembly creates a current command based at least in part on the torque command and the RPM command, creates a blade pitch command based at least in part on the torque command and the RPM command, communicates the current command to the electric motor to change a mechanical output of the electric motor, and communicates the blade pitch command to blade actuators to control the pitch of the blades. The current command and the blade pitch command cause the blades of the aircraft to rotate at a predetermined RPM.

Abstract: Embodiments are directed to systems and methods for utilizing a two-degree-of-freedom model-following control law within an architecture of nested loop functions. This allows for the separation of command and feedback path requirements and enables the restrictions of the inner loops to be applied to the outer loop feedbacks. This method of restriction allows a better allocation of coordinated authority between the loops.

Abstract: One embodiment is an aircraft including a fuselage; a wing connected to the fuselage; first and second booms connected to the wing on opposite sides of the fuselage; first and second forward propulsion systems attached to forward ends of the first and second booms; first and second aft propulsion systems fixedly attached proximate aft ends of the first and second booms; and first and second wing-mounted propulsion systems connected to outboard ends of wings; wherein the first and second wing-mounted propulsion systems are tiltable between a first position when the aircraft is in a hover mode and a second position when the aircraft is in a cruise mode.

Abstract: There is disclosed in one example an inner loop controller for an aircraft flight computer, including: a stall protection circuit to compute, for an attitude angle ?, an attitude limit ?max as a function of a flight path angle (?) and an angle of attack limit (?max); a transfer function circuit to convert ? to an attitude rate {dot over (?)}, wherein {dot over (?)} is a time derivative of ?; and a load protector circuit to compute a limit on {dot over (?)} ({dot over (?)}max) as a function of a load factor limit (Nz,max) and a true airspeed (v).

Abstract: A method for determining an aircraft angle-of-attack for aircraft stall protection includes providing an output signal from an angle-of-attack sensor and determining an initial angle-of-attack signal based on the output signal. The initial angle-of-attack signal is compensated to provide a pseudo angle-of-attack signal, and the pseudo angle-of-attack signal is mapped to a true angle-of-attack signal based on flight test data. The true angle-of-attack signal is compensated based on roll rate and sideslip or estimated sideslip to provide a compensated angle-of-attack. A complementary filter is applied that complements the compensated angle-of-attack signal with a higher frequency inertial angle-of-attack rate signal, calculated from aircraft inertial data, to provide an angle-of-attack complementary filter output. An angle-of-attack threshold for aircraft stall protection is determined based on one or more compensation parameters.

Abstract: An interconnected flight controller for an aircraft includes a mechanical linkage connecting a pilot interface with a copilot interface. When an input is provided to either of the pilot or copilot interfaces, coordinated motion is provided between them of a proportional magnitude and direction. A mechanical-disconnect element within the mechanical linkage is adapted to actuate mechanical decoupling between the pilot interface and the copilot interface. One or more sensors is coupled to the mechanical linkage to sense inputs and communicate the inputs to a fly-by-wire flight controller. An autopilot servo is coupled to the mechanical linkage for providing autopilot control or feedback and a force-feedback subsystem is connected to the mechanical linkage to simulate and apply an opposing force of aircraft control surfaces to the pilot interface and the copilot interface.

Abstract: A method for monitoring an oscillatory signal from an oscillating device includes filtering the oscillatory signal to within a desired frequency band to provide a filtered signal and extracting an amplitude from the filtered signal. The method further includes switching control of the oscillating device when the amplitude exceeds a predetermined amplitude requirement for a predetermined duration. An oscillatory signal monitor includes a first controller and a second controller each configured to independently control an oscillating device. An oscillatory signal based on a position of the oscillating device is filtered to a desired frequency band, and an amplitude is extracted from the filtered signal. A switch is provided for switching control of the oscillatory device from the first controller to the second controller when the amplitude exceeds a predetermined amplitude requirement for a predetermined duration.

Abstract: An autopilot nonlinear compensation method includes providing an autopilot command for executing an aircraft maneuver, determining a desired aerodynamic moment of the aircraft based on the autopilot command, providing a measured pilot interface position, determining a total aerodynamic moment of the aircraft based on the measured pilot interface position and the autopilot command in combination with the desired aerodynamic moment, determining a ratio of the desired aerodynamic moment to the total aerodynamic moment, and adjusting the autopilot command with a corrective command based on the ratio. The method may be used to stabilize autopilot control of an aircraft following nonlinear deployment of a control surface.

Abstract: An automatic yaw enhancement method for an aircraft having at least one propeller includes providing to a flight controller a pilot command from a pilot interface and avionic data for an airspeed, an angle of attack, and a thrust. A P-factor compensation is determined based on one or more of the airspeed, the angle of attack, and the thrust. A command to a trim device is determined based on a P-factor compensation. When a rudder bias persists, the command to the trim device is repeatedly updated until a rudder force input is nullified. The methods provide automatic pilot assistance for controlling yaw during asymmetric flight conditions and automatic turn coordination while allowing intentional side-slip for facilitating crosswind landings.

Abstract: An actuator hardover monitor for a control surface includes an actuator sensor for detecting an actuator position, a command model of an expected position of the actuator based on an input command, and a monitor to determine whether a difference between the actuator position and the expected position exceeds a threshold for a predetermined duration. A method of preventing a hardover event for a control surface includes commanding an actuator valve to a commanded position, determining continuously when the commanded position, or an actuator valve position, or a control-surface position, or a modeled actuator valve position exceeds a predetermined limit to provide an exceedance. The method may further include filtering a signal of the exceedance based on a time constant to provide a filtered exceedance, and switching to a backup control-surface actuator when the filtered exceedance exceeds the predetermined limit for a predetermined duration.

Abstract: An autopilot nonlinear compensation method includes providing an autopilot command for executing an aircraft maneuver, determining a desired aerodynamic moment of the aircraft based on the autopilot command, providing a measured pilot interface position, determining a total aerodynamic moment of the aircraft based on the measured pilot interface position and the autopilot command in combination with the desired aerodynamic moment, determining a ratio of the desired aerodynamic moment to the total aerodynamic moment, and adjusting the autopilot command with a corrective command based on the ratio. The method may be used to stabilize autopilot control of an aircraft following nonlinear deployment of a control surface.

Abstract: An interconnected flight controller for an aircraft includes a mechanical linkage connecting a pilot interface with a copilot interface. When an input is provided to either of the pilot or copilot interfaces, coordinated motion is provided between them of a proportional magnitude and direction. A mechanical-disconnect element within the mechanical linkage is adapted to actuate mechanical decoupling between the pilot interface and the copilot interface. One or more sensors is coupled to the mechanical linkage to sense inputs and communicate the inputs to a fly-by-wire flight controller. An autopilot servo is coupled to the mechanical linkage for providing autopilot control or feedback and a force-feedback subsystem is connected to the mechanical linkage to simulate and apply an opposing force of aircraft control surfaces to the pilot interface and the copilot interface.

Abstract: Embodiments of a method to determine airspeed for aircraft include determining critical air data parameters without the use of pitot-static systems. Airspeed may be determined by iteratively repeating the method until converging on a stable airspeed value that differs from a previous airspeed value by less than a predetermined threshold. Airspeed may be determined by modeling aircraft lift and repeatedly updating dynamic pressure to converge on an airspeed based on the balance between aircraft lift and weight. Airspeed may be determined based on predetermined relationships between a horizontal control surface position and dynamic pressure. A voting logic method validates or invalidates airspeeds from dissimilar sources, including airspeeds determined using the methods described herein and conventional pitot-static systems.

Abstract: In an embodiment, a method for aircraft weight estimation is provided that includes determining a weight signal based on a dynamic pressure signal, a calibrated angle of attack signal, a lift coefficient signal, a load factor signal, and a wing surface area. In another embodiment, a method to estimate aircraft weight is provided that includes determining a weight based on historical flight data relating horizontal control surface position to dynamic pressure. In another embodiment, a system for continuously estimating aircraft weight during flight is provided that includes a pitot-static subsystem, an angle of attack indicator, an accelerometer, a controller configured to provide a weight signal, and a signal filter for filtering the weight signal to determine a stable aircraft weight.

Abstract: An actuator hardover monitor for a control surface includes an actuator sensor for detecting an actuator position, a command model of an expected position of the actuator based on an input command, and a monitor to determine whether a difference between the actuator position and the expected position exceeds a threshold for a predetermined duration. A method of preventing a hardover event for a control surface includes commanding an actuator valve to a commanded position, determining continuously when the commanded position, or an actuator valve position, or a control-surface position, or a modeled actuator valve position exceeds a predetermined limit to provide an exceedance. The method may further include filtering a signal of the exceedance based on a time constant to provide a filtered exceedance, and switching to a backup control-surface actuator when the filtered exceedance exceeds the predetermined limit for a predetermined duration.

Abstract: In an embodiment, a method for aircraft weight estimation is provided that includes determining a weight signal based on a dynamic pressure signal, a calibrated angle of attack signal, a lift coefficient signal, a load factor signal, and a wing surface area. In another embodiment, a method to estimate aircraft weight is provided that includes determining a weight based on historical flight data relating horizontal control surface position to dynamic pressure. In another embodiment, a system for continuously estimating aircraft weight during flight is provided that includes a pitot-static subsystem, an angle of attack indicator, an accelerometer, a controller configured to provide a weight signal, and a signal filter for filtering the weight signal to determine a stable aircraft weight.

Abstract: Systems and methods are described for automatically counteracting undesirable aircraft movement caused by malfunctioning fly-by-wire aircraft control surfaces, hardover events, control surface disconnection, and other control surface failure events. The systems and methods include control law algorithms for reacting to such events to counteract the undesired aircraft movement. An expected roll rate is generated based on control input and compared to the actual roll rate of the aircraft.

Abstract: Systems and methods are described for automatically counteracting undesirable aircraft movement caused by malfunctioning fly-by-wire aircraft control surfaces, hardover events, control surface disconnection, and other control surface failure events. The systems and methods include control law algorithms for reacting to such events to counteract the undesired aircraft movement. An expected roll rate is generated based on control input and compared to the actual roll rate of the aircraft.

Abstract: A system and method for a controlling an aircraft with flight control surfaces that are controlled both manually and by a computing device is disclosed. The present invention improves overall flight control operation by reducing the mechanical flight control surface components while providing sufficient back-up control capability in the event of either a mechanical or power-related failure. Through the present invention, natural feedback is provided to the operator from the mechanical flight control surface which operates independent of computer-aided flight control surfaces. Further, through the present invention, force input signals received from the pilot are filtered to improve the operation of the computer-aided flight control surfaces.