A set of commands for each of a plurality of actuators to alter an aircraft's state responsive to one or more inputs is produced. The set of commands is provided to fewer than all actuators comprising the plurality of actuators.

A method for an aircraft may include determining a transition from an airplane mode to a helicopter mode for a propulsion system of the aircraft, wherein the propulsion system comprises a plurality of rotor blades; determining whether a descent rate condition is satisfied, wherein the descent rate condition is associated with a maximum allowable descent rate for the aircraft; and controlling a collective pitch angle for the plurality of rotor blades based on the descent rate condition.

A technique for controlling an unmanned aerial vehicle (UAV) includes monitoring a sensed airspeed of the UAV, obtaining a commanded speed for the UAV, wherein the commanded speed representing a command to fly the UAV at a given speed relative to an airmass or to Earth, and when the commanded speed is greater than the sensed airspeed, using the commanded speed in lieu of the sensed airspeed to inform flight control decisions of the UAV.

An example method of limiting an aircraft to a pitch axis envelope includes determining aircraft state limits associated with multiple pitch axis variables of an aircraft, determining predicted aircraft states, comparing the predicted aircraft states to the aircraft state limits to produce aircraft state errors, translating the aircraft state errors into a set of positive and negative limit elevator commands, selecting a highest priority positive limit elevator command, selecting a highest priority negative limit elevator command, limiting a primary pitch axis control law elevator command of the aircraft to a value that is less than or equal to the highest priority positive limit elevator command and greater than or equal to the highest priority negative limit elevator command, and controlling the aircraft according to the primary pitch axis control law elevator command limited to the value.

A flight control computer of an aircraft is likely to operate in a so-called incidence protection mode in which it is configured to compute the deflection orders of an elevator as a function of incidence angle values supplied by a set of incidence probes, so as to keep the incidence angle of the aircraft within a range of acceptable incidence angle values. The flight control computer is configured to, when only one incidence probe is operational: compute a first estimated incidence angle value of the aircraft, by a first estimator and a second estimated incidence angle value of the aircraft, by a second estimator unlike the first estimator; and keep the incidence protection mode activated as long as the incidence angle value supplied by the single operational incidence probe is consistent with at least one out of the first estimated incidence angle value and the second estimated incidence angle value.

In one embodiment, a method is performed by a control computer. The method includes receiving a time series of control inputs in relation to a control axis of an aircraft, where the control computer causes actuation in response to each control input in the time series as the control input is received. The method also includes determining aircraft oscillation over a sample period corresponding to the time series. The method also includes evaluating information related to the determined aircraft oscillation using engagement settings associated with a control filter. The method also includes engaging the control filter responsive to the information satisfying the engagement settings, where the engaged control filter systematically attenuates future control inputs in relation to the control axis prior to actuation responsive thereto.

Flight control systems, flight control laws, and aircraft are provided. An flight control system includes an input configured to receive a pitch rate command, a processor operative to receive the pitch angle command, to calculate a pitch angle saturation limit, to compare the sum of the pitch rate command, the scaled pitch rate, and the scaled pitch angle to the pitch angle saturation limit, to convert the pitch rate command system to the pitch angle command system in response to the sum exceeding the pitch angle saturation limit value to limit the pilot pitch-up pitch rate command, and to couple the pitch rate command to an aircraft control surface for the failure case of one of control surface, and the aircraft control surface configured to adjust an aircraft control surface setting in response to the pitch rate command and/or pitch angle command to protect an aircraft from being in stall condition.

A set of commands for each of a plurality of actuators to alter an aircraft's state responsive to one or more inputs is produced. The set of commands is provided to fewer than all actuators comprising the plurality of actuators.

Aircraft and methods and systems for controlling performance of an aircraft. The methods and systems allow the aircraft to meet a performance requirement using a set of aircraft flight data and actuators connected to control surfaces. Performance data for primary and secondary actuators are obtained to select between a primary control law for controlling the primary control surface, a secondary control law for controlling the secondary control surface, and a blended control law that controls the primary and secondary control surfaces together. If the primary control surface cannot meet the aircraft performance requirement using the primary control law, the blended control law is implemented if the primary and secondary control surfaces can be used together to meet the performance requirement; otherwise the secondary control surface is used, using the secondary control law, to meet the aircraft performance requirement.

In an example, method of controlling thrust produced by a plurality of engines on an aircraft for assisting with nose-down recovery of the aircraft is described. The method includes selecting a maximum value of an aircraft parameter, measuring a value of the aircraft parameter while the aircraft is in flight, based on a comparison of the maximum value and the measured value determining that the measured value exceeds the maximum value, and reducing a thrust produced by each of the engines of the plurality of engines to bring the measured value of the aircraft parameter below the maximum value of the aircraft parameter.