77results about "Actuated automatically" patented technology

Operation control unit controls flight driving mechanism such that drone performs an operation by which detection values of sensors included in sensors included in drone match a target value. Detection unit detects, when a plurality of drones perform an operation by which detection values of the sensors included in sensors included in the plurality of drones match a target value, a difference in an operation result of each of drones. Clustering unit classifies a plurality of drones, including drone in which clustering unit is included, into at least one cluster based on differences in operation results detected by detection unit. Correction unit corrects the detection values of the sensors based on the operation results of drones included in a maximum cluster of the classified clusters.

The present invention relates to an electric flight control system (5) for controlling an aircraft (1).The flight control system (5) comprises a plurality of functional subsets (95), each having at least one dissimilar primary computer (35) and at least one dissimilar secondary computer (36), at least one functional subset (95) being a simplex / duplex subset (96) comprising at least three identical simplex primary computers (41, 42, 43) and at least one duplex secondary computer (44, 45), each duplex secondary computer (44, 45) of a simplex / duplex subset (96) comprising an independent and synchronized control computing channel (46) and a monitoring computing channel (47), each simplex primary computer (41, 42, 43) of a simplex / duplex subset (96) having a single computing channel, and in that, barring failure, the three primary computers (41, 42, 43) of a subset simplex / duplex (96) form two virtual pseudo-duplex computers. Abbreviation figure: figure 1.

A system comprising an unmanned aerial vehicle (UAV) (100) configured to transition (520) from a terminal homing mode (510) to a target search mode (530), responsive to an uplink signal (451) and / or an autonomous determination of scene change.

The present disclosure relates to a an aircraft assistance system measuring a wind speed and direction in a vicinity of an aerodrome (10) and calculating a wind speed and direction in a path (112), along which the aircraft (100) approaches or departs, and predict wind data representing a wind speed and direction in the path (112) for a predetermined future time period. The present disclosure further relates to an aircraft that receives the predicted wind data transmitted by the aircraft assistance system and determines the predicted wind speed and direction along the path (112), and displays a visualisation about the predicted wind speed and direction or provides an autopilot system adapting a flight path accordingly.

The invention relates to an aerodynamic profile body with an outer flow surface for flow through a fluid, wherein the aerodynamic profile body comprises:a first surface layer and a second surface layer, which form at least part of the outer flow surface and are designed to be flexible in an end section of the aerodynamic profile body,a connection arrangement that flexibly connects the first surface layer to the second surface layer of the profile body in the end section, andan actuator device with at least one actuator provided in an interior of the aerodynamic profile body which interacts with at least one of the surface layers in the flexible end section in such a way that the cross-section of the aerodynamic profile body in the flexible end section can be changed by the at least one actuator,characterized in thatthe outer flow surface of at least one of the surface layers in the flexible end section has, at least in sections, a surface profile that extends parallel to the direction of flow.

A method of controlling an aircraft includes receiving a pilot input command and determining which one of a plurality of aircraft flight parameters to hold constant. The aircraft flight parameters include an altitude of the aircraft, a vertical speed of the aircraft, and a flight path angle of the aircraft. The determination of which one of a plurality of aircraft flight parameters to hold constant is based on both a speed of the aircraft and the determination of whether the pilot input command is commanding an ascent or descent of the aircraft. The method of controlling ascent of an aircraft includes controlling the aircraft based on the pilot input command while holding the determined aircraft flight parameter constant.

The invention relates to a force application device for an aircraft control stick, in which the computer is configured: a) for an angle (ANG1) smaller than or equal to a prescribed virtual stop threshold (S1, S2), to control the motor so that it applies a resistive force (EFF1) to the lever, according to the angle (ANG1), and to deactivate the brake; b) for an angle (ANG1) which is increasing beyond the prescribed threshold (S1, S2), to control the motor and activate the brake so that they apply a resistive force (EFF2) to the lever that is higher than the value (V1); c for an angle (ANG1) which is decreasing while being higher than the prescribed virtual stop threshold (S1, S2), to control the motor so that it applies a return force (EFF3) to the lever according to the angle (ANG1) and to deactivate the brake.

The application discloses a distributed large eVTOL electric governor control method, relates to the technical field of aircraft control systems, and discloses a distributed large eVTOL electric governor control method, which realizes rapid detection and isolation of faults, maintains system stability through accurate dynamic thrust compensation, can improve flight safety, reduce fault propagation risk, and ensure that the system can quickly recover to stability when hardware faults occur.

Aspects of the present disclosure generally relate to systems and methods for flight control of aircrafts driven by electric propulsion systems and in other types of vehicles. In some embodiments, a computer-implemented method for command prioritization in an aircraft is disclosed. The method comprises receiving a pilot command, analyzing the pilot command to determine characteristics associated with the pilot command, wherein the characteristics to airspeed and climb of an aircraft, assigning weights to characteristics associated with the pilot command based on constraint data, determining priority of execution between airspeed and climb based on the weights assigned to the characteristics associated with the pilot command, calculating a correction factor to be applied to the characteristics associated with the pilot command based on determined priority and generating at least one actuator command to control the aircraft based on determined priority of execution.

An unmanned aerial vehicle (UAV) launch tube (100) that comprises at least one inner layer of prepreg substrate (370) disposed about a right parallelepiped aperture (305), at least one outer layer of prepreg substrate (380) disposed about the right parallelepiped aperture (305), and one or more structural panels (341-344) disposed between the at least one inner layer of prepreg substrate (340) and the at least one outer layer of prepreg substrate (380). An unmanned aerial vehicle (UAV) launch tube (100) that comprises a tethered sabot (700,740) configured to engage a UAV within a launcher volume defined by an inner wall, the tethered sabot (700,740) dimensioned to provide a pressure seal at the inner wall and tethered to the inner wall, and wherein the tethered sabot (700,740) is hollow having an open end oriented toward a high pressure volume and a tether (740) attached within a hollow (910) of the sabot (700) and attached to the inner wall retaining the high pressure volume or attach to the inner base wall (1013). A system comprising a communication node (1500-1505) and a launcher (1520) comprising an unmanned aerial vehicle (UAV) in a pre-launch state configured to receive and respond to command inputs from the communication node (1500-1505).

Systems, methods, and machine-readable media can facilitate the control of aircraft. Actuators can be controlled based on the following: A first control signal can be output using a first command module to control the actuator. The first command module may include a first electronic configuration. Commands and / or system response signals generated by the first command module can be monitored using a monitoring module. The monitoring module may include a third electronic configuration different from the first electronic configuration. When the first command module is deactivated or malfunctions, a second control signal can be output using a second command module to control the actuator. The second command module may include a second electronic configuration different from the first and third electronic configurations. When the first command module is deactivated or malfunctions, the actuator can be controlled using the second control signal.

Method of monitoring motion of an aircraft (100) during flight. The method includes determining sensed motion of the aircraft (100) about an axis during the flight and determining predictive motion of the aircraft (100) about the axis during the flight (based on a digital model (200)). An error is determined based on the sensed motion and the predictive motion. An effector (110) on the aircraft (100) is determined that controls the motion of the aircraft about the axis. The error is converted into an amount of necessary movement of the effector (110) to correct the error.

The present invention is a powered electric lift augmentation systems (ELAS's) to add drag while minimally reducing or enhancing the lift augmentation of an airplane. At least one ELAS is attached to an airfoil, and each ELAS comprises: a nacelle; and electric duct fan (EDF) housed within the nacelle; an optional pylon perpendicularly connected beneath the nacelle; an optional slat; and one or more drag elements, pivotably connected through a range of angles to airflow. Each drag element comprises a hinged plate, or a pair of opposing plates, connected to a pilot's control element able to rotate the plates to create a range of drag levels or to stow away. The drag elements are connected to the airfoil leading edge, the nacelle, pylon, or slat; and may further comprises holes. As ELAS assembly may comprise one or more ELAS systems attached to an airfoil (aerodynamic lifting element—ALE).

A grouped flying object system (A) in which it possible for a plurality of flying objects (B) to fly while connected to a single supported object (C), said system (A) comprising: connection mechanisms (D) that connect the flying objects (B) to the supported object (C); and a grouped flight control unit that causes the plurality of flying objects (B) to fly in a coordinated manner.

A quantitative flight test data gathering system is configured to, during flight test operation of an aircraft, obtain, from one or more flight data sensors, from a flight control system, or both, one or more feedback signals indicative of a state of the aircraft. The quantitative flight test data gathering system is configured to generate, based at least in part on the feedback signals, one or more signals associated with a flight test maneuver. The quantitative flight test data gathering system is configured to provide the one or more signals as input to the flight control system to cause one or more control laws of the flight control system to generate control signals to cause the aircraft to operate according to the flight test maneuver.

This aircraft flight trim system comprises a friction module that has variable friction and is coupled on one side to a drive shaft and on the other side to an output shaft via a kinematic chain.The aircraft flight trim system has a control circuit which controls the friction module on the basis of an angular displacement value of the output shaft provided by a position sensor and on the basis of a force feedback control signal provided by a flight control computer.

The invention relates to a rotorcraft (1) with at least one main rotor (1a) and at least one engine (1g) for powering the at least one main rotor (1a) in a normal flight mode, comprising an autorotation training mode control system (7) that is activatable for switching rotorcraft operation between the normal flight mode and an autorotation training mode configured to enable training of autorotation, wherein the autorotation training mode control system (7) comprises at least one main control element that is manually operable for activating the autorotation training mode control system (7) and switching the rotorcraft operation from the normal flight mode to the autorotation training mode to engage the autorotation training mode.

The present invention provides an aircraft capable of reliably transmitting traction force to a work device and appropriately towing the said work device. [Solution] The flying device comprises an airframe and a plurality of rotors attached to the airframe that can change the altitude of the airframe, the airframe being connected to a mobile work device via a plurality of rope members and capable of towing the work device, the plurality of rotors increasing the altitude of the airframe as the attitude of the airframe and / or the attitude of the work device towed by the plurality of rope members becomes unstable. The plurality of rotors also increase the altitude of the airframe as the difference in the tensions acting on the plurality of rope members increases.

A quantitative flight test data gathering system is configured to, during flight test operation of an aircraft, obtain, from one or more flight data sensors, from a flight control system, or both, one or more feedback signals indicative of a state of the aircraft. The quantitative flight test data gathering system is configured to generate, based at least in part on the feedback signals, one or more signals associated with a flight test maneuver. The quantitative flight test data gathering system is configured to provide the one or more signals as input to the flight control system to cause one or more control laws of the flight control system to generate control signals to cause the aircraft to operate according to the flight test maneuver.

The present invention provides a system and method for a radar / optical signal integrated gust detection system. [Solution] In a particular embodiment, the system includes an optical system mounted on an aircraft and configured to provide optical wind measurements from a volume of space along the direction of travel of the aircraft. Furthermore, the system includes a radar system mounted on an aircraft and configured to provide radar wind measurements from a volume of space along the direction of travel of the aircraft. The system also includes a processor that receives optical wind measurements from the optical system and radar wind measurements from the radar system, the processor configured to characterize gusts in the volume of space by fusing the optical wind measurements with radar wind measurements, the fusing of optical wind measurements with radar wind measurements based on meteorological conditions in the volume of space.

A control device that comprises a processor and a memory that is connected to or housed in the processor. The processor: obtains the vertical position of a marker detected by an optical sensor mounted on a flight vehicle, for a marker that can change position in the vertical direction by using a displacement mechanism; and controls the flight vehicle on the basis of the vertical position of the marker, so as to maintain or change the vertical position of the flight vehicle.

To provide an aircraft capable of landing in a stable posture. [Solution] The flying device comprises an airframe, skids attached to the underside of the airframe, a plurality of rotors attached to the airframe that can change the altitude of the airframe, and a drive unit attached to the airframe that can change the relative position of the work device to the airframe by winding and unwinding a string member connected to the work device. The plurality of rotors are configured such that when the work device is positioned above the lower end of the skids by the drive unit, contact with the ground surface of the skids is permitted as the altitude of the airframe is changed, and when the work device is positioned below the lower end of the skids by the drive unit, contact with the ground surface of the skids is restricted as the altitude of the airframe is changed.

Aspects of the present disclosure relate to flight control of electric aircraft and other vehicles. In one embodiment, an aircraft is disclosed which comprises a fuselage, one or more flight control computers configured to provide control signals, four sets of electrically driven propellers, the first set being positioned in front of the second set and both on one side of the fuselage, the third set being positioned in front of the fourth set and both on the other side of the fuselage, and a plurality of electric buses, the one or more flight control computers configured to provide control signals to at least one of the first set of propellers, at least one of the fourth set of propellers, and at least one of the third set of propellers and at least one of the second set of propellers via one of the plurality of electric buses.

An electrical flight control system for controlling an aircraft. The flight control system comprises a plurality of functional subassemblies each having at least one primary computer and at least one dissimilar secondary computer, at least one functional subassembly being a simplex / duplex subassembly comprising at least three identical simplex primary computers and at least one duplex secondary computer, each duplex secondary computer of a simplex / duplex subassembly comprising an independent and synchronized control calculation channel and a monitoring calculation channel, each simplex primary computer of a simplex / duplex subassembly having a single calculation channel, and wherein, except in the event of failure, the three primary computers of a simplex / duplex subassembly form two virtual pseudo-duplex computers.

To provide a flight control system capable of performing safe manual operation of a flying body without needing a high level operation technique.SOLUTION: A flight control system includes an operation device and an unmanned aircraft. The unmanned aircraft includes: a propulsion mechanism containing a plurality of rotary vanes; a position estimation unit which estimates a position of the unmanned aircraft; an attitude estimation unit which estimates an attitude of the unmanned aircraft; a speed estimation unit which estimates a moving speed of the unmanned aircraft; and a flight control unit which controls the propulsion mechanism on the basis of information containing the current position, attitude, and speed of the unmanned aircraft. The flight control unit controls the propulsion mechanism so that the flight is performed along a virtual flight line previously set on map data on the basis of an operation signal transmitted from the operation device according to operation by an operator along one axis direction of an operation input unit.SELECTED DRAWING: Figure 1

An aircraft operating method comprises receiving inputs from flight controls; responding to the inputs from the flight controls by providing control surface commands to control flight; and responding to alternative operation of the flight controls to select and perform an aircraft maintenance operation other than during flight.

A fly-by-wire system that lets a single-engine aircraft mimic multi-engine behavior for real-flight pilot training, especially asymmetric thrust effects that create yaw / roll. A cockpit central thrust control with two independent thrust levers generates an overall power command plus a dissymmetry signal; a calculator applying a dedicated control law converts these into corrected setpoints for the rudder, left / right ailerons, and propeller parameters to reproduce multi-engine handling. Instructors may trigger a simulated engine failure and then reset back to normal single-engine operation. Optional wingtip electric thrusters powered by a battery may increase asymmetric effects while maintaining safety margin, and the system may display / record simulated multi-engine data for debriefing-reducing training cost and CO2 versus operating real twins

The invention relates to an aircraft wing comprising a wing structure and a spoiler. The spoiler is movable between a stowed configuration and a deployed configuration in a "pop-up" manner. The spoiler comprises a hinged top flap movable between a first position, in which the hinged top flap is constrained by an actuator, and a second position, in which the hinged top flap is unconstrained by the actuator. When the hinged top flap is in the second position, airflow across the top surface of the aircraft wing acts to pull the spoiler to the deployed position. The invention also relates to a method of deploying a stowed spoiler in an aircraft wing and a method of retracting a deployed spoiler in an aircraft wing.

The application provides an aircraft landing assisting method, system, aircraft, medium and program product. The method comprises: in response to a spoiler abnormality signal, obtaining a set of operation parameters of a current flight phase; based on the set of operation parameters and a preset aircraft performance database, determining slat aerodynamic performance data and flap aerodynamic performance data; according to the slat aerodynamic performance data and the flap aerodynamic performance data, determining a cooperative deflection combination of the slat and the flap in the current flight phase, the cooperative deflection combination comprising a cooperative deflection angle and a cooperative deflection timing; based on the cooperative deflection combination, generating a landing assisting instruction for the spoiler abnormality, the landing assisting instruction being used to control the slat and the flap to adjust the cooperative deflection angle at the cooperative deflection timing. The method of the application utilizes the resistance generated by the flap and the slat to alleviate the influence caused by the spoiler abnormality.