24results about "Control initiation means" patented technology

In some embodiments, an electric aircraft may include a plurality of flight components including a plurality of control surfaces; a plurality of lift propulsors; at least a thrust propulsor; and a plurality of electric motors configured to power the plurality of propulsors; a flight controller, communicatively connected to a pilot input and flight components. In some embodiments, a flight controller may be configured to receive control datum from a pilot input; and generate an output datum as a function of the control datum.

A vertical take-off and landing aircraft includes: a fuselage and 2N tilting rotors. Both sides of the fuselage are provided with a wing symmetrically, a tail of the fuselage is provided with an empennage, and ruddervators are provided at the empennage. 2N tilting rotors are symmetrically provided at both sides of the fuselage, and a part of the 2N tilting rotors are provided on the empennage. N is a natural number greater than or equal to 2; in a vertical take-off and landing state, projections of propellers of the 2N tilting rotors on a horizontal plane are centrally symmetrical about point B, point B and a center of gravity point G of the vertical take-off and landing aircraft are both provided in a symmetry plane of the fuselage.

Describes a micro spoiler for enhancing the effectiveness of a lateral control surface of an aircraft wing. An example aircraft includes a wing, a lateral control surface, and a micro spoiler. The lateral control surface is movably coupled to the wing. The lateral control surface is movable between a neutral position, a first upward deflection position, and a second upward deflection position that extends beyond the first upward deflection position. The micro spoiler is located on or in front of the lateral control surface. The micro spoiler is movable between a retracted position and a deployed position. The micro spoiler is configured to move from the retracted position to the deployed position based on the lateral control surface moving from the neutral position to the first upward deflection position or moving toward the first upward deflection position.

Embodiments relate to an aircraft control router for an aircraft. The aircraft control router may include a command processing module, sensor validation module, aircraft state estimation module, and control laws module. The command processing module may be configured to generate aircraft trajectory values based on received aircraft control inputs. The sensor validation module may be configured to validate sensor signals generated by sensors of the aircraft. The aircraft state estimation module may be configured to determine an estimated aircraft state of the aircraft based on the validated sensor signals. The control laws module may be configured to generate actuator commands for actuators of the aircraft to adjust control surfaces of the aircraft, where the generated actuator commands are based on aircraft trajectory values, validated sensor signals, and an estimated aircraft state. The aircraft control router may transmit the generated actuator commands to actuators of the aircraft.

In some embodiments, an electric aircraft may include a plurality of flight components including a plurality of control surfaces; a plurality of lift propulsors; at least a thrust propulsor; and a plurality of electric motors configured to power the plurality of propulsors; a flight controller, communicatively connected to a pilot input and flight components. In some embodiments, a flight controller may be configured to receive control datum from a pilot input; and generate an output datum as a function of the control datum.

A system and a method are disclosed for a vehicle control and interface system configured to facilitate control of different vehicles through universal mechanisms. The vehicle control and interface system can be integrated with different types of vehicles (e.g., rotorcraft, fixed-wing aircraft, motor vehicles, watercraft, etc.) in order to facilitate operation of the different vehicles using universal vehicle control inputs. In particular, the vehicle control and interface system converts universal vehicle control inputs describing a requested trajectory of a vehicle received from one or more universal vehicle control interfaces into commands for specific actuators of the vehicle configured to adjust a current trajectory of the vehicle to the requested trajectory. In order to convert the universal vehicle control inputs to actuator commands the vehicle control and interface system processes the universal vehicle control inputs using a universal vehicle control router.

The subject matter of this specification can be embodied in, among other things, a handlebar-shaped housing (102) having an elongate central body (110) having a first hand grip (120a) at a first axial end (112a), a first flexible paddle (130a) affixed to the first hand grip, a first deflection sensor (136a) configured to identify a first amount of deflection of the first flexible paddle, a second hand grip (120b) at a second axial end (112b), a second flexible paddle (130b) affixed to the second hand grip, a second deflection sensor (136b) configured to identify a second amount of deflection of the second flexible paddle, and circuitry configured to identify one or more of a pitch, a roll, and a yaw of the elongate central body, and a controller configured to receive an orientation sensor signals, deflection signals from the deflection sensors, and provide a control signal based on one or more of the sensor signals.

A method and system for processing parameter values from a redundant sensor (34) configured to sense a parameter used in the control of an aircraft engine (10) is provided. The method includes: a) receiving a plurality of parameter values from a redundant sensor (34) by sensing the same parameter at the same time; b) identifying mismatched parameter values; c) producing a predicted parameter value using an artificial intelligence (AI) model having a database of parameter values representative of the sensed parameter; d) providing the predicted parameter value to a control unit (36); and e) operating the control unit (36) to select a first parameter value or a second parameter value using the predicted parameter for use in the control of the aircraft engine (10).

A method of controlling a powered lift aircraft comprises receiving, from a pilot input device, an input indicative of one of a powered lift enabled mode or a powered lift disabled mode and controlling the at least one powered lift element to operate the aircraft in a selected one of the powered lift enabled mode or the powered lift disabled mode based on the received input, wherein in the powered lift enabled mode, the at least one powered lift element is controlled based on a state of the aircraft and, wherein in the powered lift disabled mode, the at least one powered lift element is controlled to disable powered lift.

The present application provides a vertical take-off and landing aircraft and a control method thereof. The vertical take-off and landing aircraft includes: a fuselage and 2N tilting rotors. Both sides of the fuselage are provided with a wing symmetrically, a tail of the fuselage is provided with an empennage, and ruddervators are provided at the empennage. 2N tilting rotors are symmetrically provided at both sides of the fuselage, and a part of the 2N tilting rotors are provided on the empennage. N is a natural number greater than or equal to 2; in a vertical take-off and landing state, projections of propellers of the 2N tilting rotors on a horizontal plane are centrally symmetrical about point B, point B and a center of gravity point G of the vertical take-off and landing aircraft are both provided in a symmetry plane of the fuselage, and point B is provided at a side of point G close to the empennage. During the transition of a vertical take-off and landing aircraft from a vertical take-off and landing state to a cruising state, both point G and point B is configured to move along the symmetry plane, and point B is always provided at the side of point G close to the empennage. This configuration enables the vertical take-off and landing aircraft to have relatively good pitch trimming capability.

A drive for a harmonic gear assembly, comprising an input shaft having a shaft axis (A) and an epicyclic carrier unit removably mounted around, and coaxial with the input shaft, the epicyclic carrier unit comprising a carrier having an interior circumference designed to match the outer circumference of the shaft and further comprising a first roller assembly mounted at a first location around the circumference of the carrier unit and comprising a first roller shaft extending substantially parallel to the shaft axis and a first roller rotatably mounted around the first rotor shaft, and a second roller assembly mounted at a second location around the circumference of the carrier unit and comprising a second roller shaft extending substantially parallel to the shaft axis and a second roller rotatably mounted around the second rotor shaft, the first and second rollers extending radially outwards from the carrier to define a major axis between the radially outer surfaces of the first and second rollers, and wherein the outer surface of the carrier unit at locations other than the first and second locations defines a minor axis.

A vertical take-off and landing (VTOL) aircraft (10) comprises a fuselage (24) first and second forward wings (20, 22) and first and second rearward wings (30, 32), each wing having a fixed leading edge (25, 35) and a trailing control surface (50) which is pivotal about a generally horizontal axis. Electric rotors (60) are mounted to the wings (20, 22, 30, 32), the electric rotors (60) being pivotal with the trailing control surface (50) between a first position in which each rotor (60) has a generally vertical axis of rotation, and a second position in which each rotor (60) has a generally horizontal axis of rotation; wherein at least one of the wings (20, 22, 30, 32) has a first and a second electric rotor (60) which are each mounted having non-parallel axes of rotation so that the thrust lines of the first and second electric rotors are different.

The present disclosure relates to a battery pack diagnosis control apparatus, a battery system including the same, and a method thereof. An exemplary embodiment of the present disclosure provides a battery system including: at least one battery pack; a flight control apparatus configured to transmit at least one of a battery discharge mode command or a motor discharge command; a battery control apparatus configured to transmit a relay control command for a relay based on receiving the battery discharge mode command, and transmit a diagnosis command based on receiving the motor discharge command; and at least one battery pack control apparatus configured to control the relay based on receiving the relay control command, and diagnose a failure of the at least one battery pack by measuring a current of the at least one battery pack based on receiving the diagnosis command.

The invention relates to a redundant control device (10) for a mobile or stationary system, in particular for an aircraft, wherein - the redundant control device (10) has a first control module (16a) and a second control module (16b) which is preferably identical in construction to the first control module (16a) for the purpose of redundancy, - the control modules (16a, 16b) are arranged together in an interior space (14) of the control device (10), - each of the control modules (16a, 16b) comprises a carrier element (18a, 18b) with a plurality of components (20a, 20b) arranged on the respective carrier element (18a, 18b) and - the carrier elements (20a, 20b) with the components (20a, 20b) are coated with a sealing layer (22a, 22b) in order to prevent the components (20a, 20b) from becoming detached from the respective carrier element (18a, 18b).

The present invention relates to a control system for an aircraft comprising a computing unit and a subunit, in particular a valve unit, which includes a control unit, wherein the computing unit is connected to the control unit only via one or more single-channel signal channels.

A system for and method of indicating an angle of attack of an aircraft is provided. The system utilizes measurements from existing aircraft sensors so that a dedicated angle of attack sensor is not required. The system enables and benefits from a unique method that includes performing standard flight operations during a calibration flight. The system does not need any aerodynamic model of the aircraft because it utilizes information from the calibration flight to develop an estimated aerodynamic model of the specific airplane. The system utilizes the aerodynamic model and other information to estimate an angle of attack of the aircraft during flight. Finally, the system utilizes additional information to indicate the estimated angle of attack of the aircraft relative to an optimum angle of attack for the aircraft and / or a critical angle of attack of the aircraft. The system further enables recalibration, so as to accommodate aircraft aerodynamic modifications.

The utility model belongs to the field of aviation aircraft design, and belongs to a hybrid flight control connection control link. Comprising a remote controller, a receiver, a flight controller A, a flight controller B and a wireless coach module, the receiver is wirelessly connected with the remote controller; the receiver is respectively connected with the input ends of the flight controller A and the flight controller B through a bus; the output ends of the flight controller A and the flight controller B are connected with each steering engine through a wireless coach module; the receiver is connected with the control end of the wireless coach module through a single-channel signal line. The receiver is used for receiving a signal transmitted by the remote controller and transmitting the transmitted signal in three paths which are respectively supplied to the flight control A, the flight control B and the wireless coach module; the flight control A and the flight control B are used for receiving the bus signal, making a control signal for the steering engine according to respective flight control programs, and sending the control signal to the wireless coach module; and the wireless coach module is used for judging and selectively transmitting one path of control signals of the flight control A or the flight control B to the steering engine through the single-channel PWM signal of the receiver.

A system and a method are disclosed for a vehicle control and interface system configured to facilitate control of different vehicles through universal mechanisms. The vehicle control and interface system can be integrated with different types of vehicles (e.g., rotorcraft, fixed-wing aircraft, motor vehicles, watercraft, etc.) in order to facilitate operation of the different vehicles using universal vehicle control inputs. In particular, the vehicle control and interface system converts universal vehicle control inputs describing a requested trajectory of a vehicle received from one or more universal vehicle control interfaces into commands for specific actuators of the vehicle configured to adjust a current trajectory of the vehicle to the requested trajectory. In order to convert the universal vehicle control inputs to actuator commands the vehicle control and interface system processes the universal vehicle control inputs using a universal vehicle control router.

This application provides a vertical take-off and landing aircraft and a control method thereof. [Solution] The vertical take-off and landing aircraft comprises an airframe and 2N tiltrotors. Main wings are provided on both sides of the airframe, a tail fin is provided at the rear of the airframe, and a rudder is provided on the tail fin. The 2N tiltrotors are mounted symmetrically on both sides of the airframe, and some of the 2N tiltrotors are located at the tail fin. Here, N is a natural number greater than or equal to 2. In the vertical take-off and landing state, the projection of the propellers of the 2N tiltrotors in the horizontal plane is center-symmetric with respect to point B, and both point B and the center of gravity G of the vertical take-off and landing aircraft are located within the plane of symmetry of the airframe, with point B located on the side of point G closer to the tail fin. In the process of the vertical take-off and landing aircraft transitioning from the vertical take-off and landing state to the cruising state, both point G and point B move toward the nose side along the plane of symmetry, and point B is always located on the side of point G closer to the tail fin. This layout improves the trim-pitch capability of the vertical take-off and landing aircraft.

A method of controlling a powered lift aircraft includes receiving an input from a pilot input device indicative of one of a powered lift enabled mode or a powered lift disabled mode; and controlling at least one powered lift element based on the received input to cause the aircraft to operate in a selected one of the powered lift enabled mode, in which the at least one powered lift element is controlled based on a state of the aircraft, or the powered lift disabled mode, in which the at least one powered lift element is controlled based on a state of the aircraft, to cause the aircraft to operate in the selected one of the powered lift enabled mode and the powered lift disabled mode. And wherein in the powered lift deactivation mode, the at least one powered lift element is controlled to deactivate the powered lift.

A method for controlling a powered lift aircraft includes receiving an input from a pilot input device indicating either a powered lift enabled mode or a powered lift disabled mode, and controlling at least one powered lift element to operate the aircraft in the selected powered lift enabled mode or powered lift disabled mode based on the received input, wherein in the powered lift enabled mode, at least one powered lift element is controlled based on the aircraft's state, and in the powered lift disabled mode, at least one powered lift element is controlled to disable the powered lift.

A method of controlling a powered lift aircraft comprises receiving, from at least one pilot input device, an input indicative of a landing of the powered lift aircraft, controlling, using at least one control circuit, at least one aircraft powered lift element to operate based on the received input. When the input is indicative of a vertical landing, controlling comprises allowing the at least one aircraft powered lift element to provide powered lift to the aircraft based on a sensed airspeed of the aircraft and when the input is indicative of a non-vertical landing, controlling comprises preventing the at least one aircraft powered lift element from providing powered lift to the aircraft.