Abstract: A controller area network (CAN) distribution system for an aircraft includes flight control computers (FCCS) for operation of the aircraft; a first bus bundle is in communication with the FCCs, and directs communication from the FCCs to a left front rotor assembly and a right rear rotor assembly; a second bus bundle in communication with the FCCs, the second bus bundle directs communication from the FCCs to a right front rotor assembly and a left rear rotor assembly; and a third bus bundle in communication with the FCCs, the third bus bundle directs communication from the FCCs to a left wing tip rotor assembly and a right wing tip rotor assembly; the first bus bundle, the second bus bundle, and the third bus bundle are to balance bandwidth, optimize total bus length and width, and provide adequate control after a failure of any two busses.

Abstract: A hybrid flight control system includes an operator input device having a force sensor and/or a position sensor. The operator input is connected to a first portion of the control surface via a mechanical linkage. The system includes an electro-mechanical actuator and an electro-mechanical actuator control system in data communication with the electro-mechanical actuator, as well as a flight control computing system in data communication with the operator input device and the electro-mechanical actuator control system. Upon receiving an input from an operator, the operator input device sends a signal to the flight control computing system. Upon receiving a signal from the operator input device, the flight control computing system sends a signal to the electro-mechanical actuator control system. Upon receiving a signal from the electro-mechanical actuator control system, the electro-mechanical actuator moves a second portion of the control surface according to a predetermined set of conditions.

Abstract: A flight control system includes a first and second operator input means configured to receive a mechanical input from an operator and are connected to an artificial feel system and at least one sensor. The system includes a flight control computer (FCC) in data communication with the sensor, wherein the FCC receives a signal from the sensor in response to an input by the operator. The flight control system has first and second servo controllers in data communication with first and second servos, respectively, and the FCC. The servos are connected to control surfaces that, when moved, cause the aircraft to change attitude. The flight control system also includes a third and a fourth servo controller, the third and fourth servo controllers being in communication with the FCC and a dual-lane servo operably connected to a third control surface that, when moved, causes the aircraft to change attitude.

Abstract: Embodiments are directed to systems and methods for controlling a tiltrotor aircraft using a flight control system. A flight control computer is configured to control aircraft effectors in response to inputs from inceptors. The flight control computer stores instructions for controlling aircraft effectors. The instructions cause the flight control computer to perform the steps of converting a signal representing longitudinal motion of a first inceptor into a fore-and-aft translational rate command for the tiltrotor aircraft; converting a signal representing lateral motion of the first inceptor into a side-to-side translational rate command for the tiltrotor aircraft; converting a signal representing longitudinal motion of a second inceptor into a height rate command for the tiltrotor aircraft; and converting a signal representing lateral motion of the second inceptor to a yaw rate command for the tiltrotor aircraft.

Abstract: An aircraft emergency descent method includes setting a pre-set maximum collective blade pitch and a pre-set altitude as part of a failure procedure; monitoring rotor assemblies through an aircraft control system and a failure detection module; determining when a rotor assembly has failed; and activating the failure procedure. The failure procedure includes commanding a maximum torque to a motor of each rotor assembly such that the rotational velocity of functioning rotors increases; detecting the increase in rotational velocity; adjusting either motor torque or a collective blade pitch to regulate rotational velocity; monitoring altitude of the aircraft; and upon determining when the aircraft reaches the pre-set altitude, adjusting the collective blade pitch to the pre-set maximum collective blade pitch via the at least one governor such that momentum is conserved, causing a descent rate of the aircraft to decrease as the aircraft approaches a ground surface.

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: A landing gear system for an aerial device includes a mounting structure and a main landing gear. The mounting structure is configured to be attached relative to a fuselage of the aerial device. The main landing gear includes a pair of aft swept struts and wheels mounted on each of the aft swept struts. The pair of aft swept struts is attached relative to the mounting structure and extends downwardly and rearwardly relative thereto.

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: 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.