1658results about "Aircraft indicators" patented technology

Systems and methods are disclosed herein for influencing mission planning, vehicle maintenance, and identifying an aerial vehicle that provide a greatest probability of mission success. In some examples, a dynamic reliability model of dynamic reliability models can be updated based on at least component reliability data indicative of a reliability of a component of the aerial vehicle of aerial vehicles. Each dynamic reliability model can characterize a reliability of one of the aerial vehicles. Each dynamic reliability model can be executed to compute an indication of vehicle reliability for each aerial vehicle. A mission of success probability for each aerial vehicle can be computed based on a respective indication of vehicle reliability. A given aerial vehicle of the aerial vehicles can be identified for implementing the mission based on an evaluation of the mission of success probability for each aerial vehicle of the aerial vehicles.

A method is disclosed here. The method includes receiving an input from an aircraft, determining a towing status of the aircraft based on the input, and sending a message to a transmitter to transmit the towing status of the aircraft. Also disclosed herein is a system including an aircraft including a nose landing gear, a first controller, and a transmitter, wherein the first controller is configured to determine a towing status of the aircraft based on a number of sensor signals received from the aircraft, and the transmitter is configured to transmit the towing status and a tow truck configured to tow the aircraft, the tow truck including a second controller and a receiver, wherein the receiver is configured to receive the towing status from the transmitter, and the second controller is configured to provide an indication of the towing status of the aircraft.

The invention discloses an aircraft control surface oscillation fault monitoring method based on multi-scale wavelet transformation, and the method comprises the steps: carrying out the calculation through a PID (Proportion Integration Differentiation) control law through an autopilot according to an input flight path command, and obtaining a control surface control instruction deltadepth; a control surface deflection prediction value delta est is generated based on the nonlinear model and the control surface control instruction, the control surface deflection prediction value and an actual deflection value delta mea collected by the sensor module are subtracted to generate a residual signal Rn, and the residual signal is decomposed into a plurality of sub-band signals through wavelet decomposition; wavelet energy calculation is carried out on the sub-bands possibly containing the control surface oscillation fault signals, whether a fault occurs or not is judged according to the magnitude of an energy comparison threshold value, and the judgment threshold value is determined through Monte Carlo simulation. According to the method, instruction signals are separated by generating residual errors, wavelet transform processing is carried out on residual error signals, interference of noise and irrelevant frequency signals on monitoring performance is reduced, meanwhile, small-amplitude oscillation signals can be effectively monitored according to sub-band wavelet energy serving as a fault criterion, and the monitoring coverage rate of oscillation faults is remarkably increased.

The invention belongs to the field of aviation equipment, and discloses an electric pump control method and system based on a protection mechanism, and the control method comprises the steps: collecting an onboard state signal, a flight control system state signal and the hydraulic pressure of each sub hydraulic system in a hydraulic system in real time; according to the onboard state signal, an electric pump closing instruction and a hydraulic pressure logic judgment instruction are output through first electric pump control logic, and an electric pump connected to the sub hydraulic system is closed according to the electric pump closing instruction; a flight control system state signal and hydraulic pressure are obtained according to the hydraulic pressure logic judgment instruction, and an electric pump connected to each hydraulic sub-system is automatically controlled through second electric pump control logic and electric pump protection logic. By designing the automatic electric pump control logic and the electric pump protection logic, the electric pump damage caused by excessive transient current due to simultaneous starting of all the electric pumps can be avoided, and the flight safety is improved.

The invention belongs to the field of electromechanical control, and discloses a redundancy control method and system for an electric pump and a transmission system.The redundancy control method comprises the steps that onboard state signals, flight control system state signals and first hydraulic pressure and second hydraulic pressure of each hydraulic sub-system in a hydraulic system are collected in real time; an electric pump closing instruction and a hydraulic pressure logic judgment instruction are output through an on-board state threshold value and first electric pump control logic according to the on-board state signal, and the electric pump is closed according to the electric pump closing instruction; and a flight control system state signal, the first hydraulic pressure and the second hydraulic pressure are obtained, and each electric pump is automatically controlled through a low-pressure threshold value and a pressure safety threshold value by adopting second electric pump control logic, a hydraulic pipeline protection mechanism and an electric pump protection mechanism. According to the invention, the starting of each electric pump can be automatically controlled, a hydraulic passage can be protected, and the flight safety is ensured.

Collective control using vertical velocity is provided. A system can include a data processing system comprising one or more processors, coupled with memory. The data processing system can receive, from a sensor, a velocity of an aerial vehicle along a vertical axis of a body of the aerial vehicle. The data processing system can generate a command for collective control of a pitch of blades of a rotor of the aerial vehicle based on a comparison of the velocity with a reference. The data processing system can control, via an actuator, the pitch of the blades of the rotor in accordance with the command.

An aircraft includes a flight deck including a flight deck door configured to be moved between an open position and a closed position, a first monument, second monument, and a secondary barrier door moveably secured between the first monument and the second monument. The secondary barrier door is configured to fold into a stowed position in which a path to the flight deck is open, and unfold into a deployed position in which the path to the flight deck is closed.

A method for selecting an altitude changing phase route for an aircraft is presented. The method comprises receiving sequences of multivariate flight data from at least one prior flight, receiving a set of flight parameters for the aircraft including at least a total altitude change and a takeoff weight, and receiving a set of candidate altitude changing phase routes having candidate step profiles. For each candidate altitude changing phase route, a sequence of fuel burn quantities is predicted for the respective candidate step profile based on at least the sequences of multivariate flight data and the set of flight parameters. The fuel burn quantities are summed over the candidate altitude changing phase route to obtain an estimated fuel burn. A preferred candidate altitude changing phase route having a lowest estimated fuel burn is indicated.

The invention belongs to the field of airborne electromechanics, and discloses a hydraulic mechanism heterogeneous control method and system.The method comprises the steps that hydraulic systems are divided into two groups, and on-board state signals, flight control system state signals and first hydraulic pressure and second hydraulic pressure of sub hydraulic systems are collected; giving an on-board state threshold value, a low-pressure threshold value, a hydraulic safety threshold value and a sub-hydraulic system relation table; starting and stopping automatic control is conducted on the electric pumps in the first group through the established first control logic architecture according to the data; and the electric pumps in the second group are automatically controlled to start and stop by adopting the constructed second control logic architecture through the data. The system comprises a grouping module, a control logic construction module and two control units, wherein each control unit comprises an acquisition module, a data scheduling module, a staggered starting module, an anti-overvoltage protection module and an anti-error working module. All the electric pumps are controlled through two heterogeneous logic architectures, and the control precision and safety are improved.

Embodiments relate generally to systems and VTOL aerial vehicles. An example central server system comprises: at least one central server system processor; and central server system memory storing central server system program instructions accessible by the at least one central server system processor, and configured to cause the at least one central server system processor to: receive vehicle registration data associated with a manned VTOL aerial vehicle; and determine that the vehicle registration data is associated with a manned VTOL aerial vehicle that is authorized to communicate with the central server system. Some embodiments include VTOL aerial vehicles in communication or communicable with a central server system.

The invention relates to the technical field of aircraft power supply systems, and discloses an intelligent power distribution and fault diagnosis method for an aircraft power supply system, which comprises the following steps: acquiring a real-time current signal and load working condition information of an aircraft power supply bus, a double-domain composite entropy feature is constructed by extracting a time domain entropy transition index representing the energy domain stability of a current signal and a phase synchronization index representing the time structural domain stability; according to the method, dual inspection of the energy domain and the time structure domain is carried out on the current signals, and cooperative judgment is carried out in combination with the load working condition, so that sensitive capture of real fault forebodes is realized; and meanwhile, tolerance is dynamically shown on normal loads in the airplane operation process, the restriction between safety and usability in the prior art is effectively avoided, and predictive health management is made possible.

A door assembly configured to be mounted in an opening of a cabin area of an aircraft. The door assembly includes a door configured to be movable between a deployed position across the opening and a stowed position out of the opening. A lock is mounted to the door and movable between a locked position and an unlocked position. An engagement assembly is mounted to the door and movable between an engaged position and a disengaged position. The engagement assembly includes a plunger, a lock mechanism, and an elongated member that extends between the lock and the lock mechanism. In the engaged position, the lock mechanism and the elongated member position the plunger in an extended position outward from the door. In the disengaged position, the lock mechanism and the elongated member position the plunger in a retracted position relative to the door.

The present disclosure relates generally to controlling an aircraft to avoid overheating components. In one embodiment, a method is disclosed, comprising: receiving first sensor data indicating at least one attribute of high voltage wiring of the aircraft; receiving a state of operation of the aircraft; determining a proximity of a temperature of the high voltage wiring to a temperature limit based on the first sensor data and the state of operation of the aircraft; and controlling at least one electric propulsion unit of the aircraft to avoid exceeding the temperature limit based on the determined proximity.

The invention provides a safe flight agent system for executing complex language instructions in a high dynamic environment. The system comprises a perception module and a security module based on a visual language model, a planning module based on a language large model and an unmanned aerial vehicle control interface. In a system operation stage, an intelligent agent receives a command from a user, a planning module splits a task, calls a corresponding task memory and a code generator to generate a corresponding executable code, calls a perceived interface to obtain perceptual information related to the task, and finally, after the perceptual information is checked by a security check module, the perceptual information is sent to a server. And executing the corresponding code to complete the instruction of the user. The system achieves safe and efficient execution of tasks in a highly dynamic environment by combining a high-performance low-frame-rate large model and a high-frame-rate traditional method in each module and using a security detection module involving privacy, potential risk and collision.

The invention relates to a short-range unmanned aerial vehicle anti-electromagnetic interference device and method based on deep learning. The device comprises an electromagnetic pulse induction module, a power supply module, an intelligent electromagnetic pulse shielding module, a signal recovery and enhancement module, an intelligent decision and control module and a deep learning and model updating module. The device integrates a plurality of intelligent modules, can sense and analyze electromagnetic pulses in real time, automatically adjust electromagnetic shielding strength, enhance signal recovery capability, and ensure stable operation of equipment in a complex electromagnetic environment; automatic decision making and online prediction capability based on deep learning are realized, and an electromagnetic interference event can be early warned and a protection strategy can be optimized; a high-sensitivity sensor is adopted, missing detection of electromagnetic pulse signals is avoided, meanwhile, a false alarm correction mechanism is established in a deep learning system, output of multiple detection models is comprehensively judged by introducing an integrated learning technology, and if judgment results of the multiple models are consistent, the detection result is more reliable.

A deployable attendant seat assembly that is configured to deploy in a vehicle cabin from a stowed position as an attendant seat into a deployed position configuring the deployable attendant seat assembly into a lockable barrier to prevent passengers from moving along an aisle in a vehicle cabin into a vehicle area forward of the lockable barrier.

The invention discloses an aircraft nozzle protection device and an aircraft, and relates to the technical field of aircraft nozzle protection.The aircraft nozzle protection device comprises an annular shell and a cover plate, the annular shell is arranged outside an aircraft engine in a sleeving mode, and the cover plate is arranged on the outer side of the annular shell; the first side of the cover plate is rotationally connected with the annular shell, the second side of the cover plate is connected with the annular shell and the engine jet pipe in an unlocking mode, when the aircraft is ejected out of the barrel, the engine jet pipe is static, the second side of the cover plate is locked with the annular shell, and when the aircraft is ejected out of the barrel, the engine jet pipe swings, the second side of the cover plate is unlocked from the annular shell. The first side of the cover plate rotates around the annular shell by the preset angle and then is separated from the annular shell, the cover plate bears huge airflow impact force generated by fuel gas when the aircraft is ejected out of the barrel, the situation that an engine jet pipe is damaged by high-temperature and high-pressure fuel gas is avoided, an ejection bottom plate in the launching barrel does not need to be used, and space is saved; and the mass and cost of the launch canister and the aircraft are reduced.

The present application discloses an aircraft configured for vertical take-off and landing and horizontal flight. The aircraft includes (i) a first nacelle fuel tank configured to hold a first fuel supply for a first engine and to selectively provide fuel to the first engine during both (a) a vertical take-off and landing, and (b) horizontal flight; (ii) a fuselage fuel tank, and (iii) a control system configured to redistribute fuel between the first nacelle fuel tank and the fuselage fuel tank.

The invention relates to the technical field of multi-rotor aircrafts, in particular to a blade protection cover detection method, device and equipment and a medium. The method comprises the steps that if it is determined that a multi-rotor aircraft is in a stable hovering state based on the detected moving speed, the detected accelerated speed and the detected angular speed; if yes, acquiring the detected climbing speed and the blade rotating speed of the multi-rotor aircraft; the nominal weight of the multi-rotor aircraft is obtained, and the theoretical climbing acceleration is obtained according to the blade rotating speed and the nominal weight; obtaining a theoretical climbing speed and a disturbance climbing acceleration according to the detected climbing speed and the theoretical climbing acceleration; and determining whether the model of the multi-rotor aircraft is matched with the model of a blade protection cover mounted on the multi-rotor aircraft according to the theoretical climbing speed and the disturbance climbing acceleration. According to the scheme, whether the model of the multi-rotor aircraft is matched with the model of the blade protection cover installed on the multi-rotor aircraft or not can be determined, an operator can be conveniently reminded to replace the unmatched blade protection cover in time, and the user experience is improved.

The invention relates to a front wheel turning abnormal deflection warning method, which comprises the following steps of: (1) finishing power-on self-test of a front wheel turning control box, and starting turning / anti-swing control; (2) a front wheel turning control box judges whether the system is in a turning state or not, if not, the system enters an anti-swing state, front wheel abnormal deviation judgment is not carried out, and an electromagnetic valve in a combined hydraulic valve is powered off in the anti-swing state; if yes, front wheel abnormal deflection and over-limit logic judgment and disposal are carried out; (3) the front wheel turning control box reads the instruction sensor signal ZL, the feedback sensor signal FK and the inertial navigation ground speed DS, and the front wheel turning control box calculates the deflection angle of the front wheel at the current speed, namely the feedback sensor theoretical position signal FKL; (4) the front wheel turning control box outputs control current according to a servo valve control current algorithm according to the difference value between the feedback sensor theoretical position signal FKL and the feedback sensor signal FK; (5) front wheel abnormal deflection judgment and treatment; and (6) front wheel overrun judgment and treatment.

The invention discloses an arresting protection device of a rear propeller type aircraft and the aircraft, and relates to the field of aircrafts. The arresting hook is hinged below the tail part of the aircraft fuselage; the telescopic limiting and buffering mechanism is mounted on the fuselage and located between the arresting hook and the propeller, and the telescopic limiting and buffering mechanism has a retraction state and an extension locking state; the trigger mechanism is arranged on the telescopic limiting buffer mechanism. Therefore, by increasing the first angle a of the arresting hook swinging upwards for the first time, the impact force of the arresting hook on the fuselage is effectively reduced, and lightweight design of the telescopic limiting buffer mechanism and the fuselage bearing area is facilitated; the telescopic limiting buffer mechanism can be switched among different states to limit the upward swinging angle of the arresting hook, so that the arresting cable is always located below the lowest rotating plane of the propeller, collision with the propeller is avoided, safety of the propeller and the aircraft body is guaranteed, and safety of the aircraft with the rear propeller in the recovery process is improved.

The invention discloses a liquid-phase sweating cooling thermal protection structure based on atomization jet flow, and relates to the field of heat reduction and resistance reduction of active flow control of aerospace high-speed aircrafts, the liquid-phase sweating cooling thermal protection structure comprises a coolant cabin, and cooling spray holes are formed in the cabin wall of the coolant cabin; an atomization vibrating body is fixedly connected to the position, located at the cooling spray hole, of the cabin wall of the cooling agent cabin, the atomization vibrating body can make contact with a cooling agent in the cooling agent cabin, and a driver used for enabling the atomization vibrating body to vibrate is further included. The liquid-phase sweating cooling device works on the basis of the ultrasonic atomization principle, through the effect of the piezoelectric ceramic atomization piece, the liquid coolant generates atomization jet flow and enters a boundary layer outside the aircraft, use of a liquid pumping driving device is avoided, and the light weight degree of the liquid-phase sweating cooling device is improved.

The invention discloses an on-board personnel physiological stress early warning method and system, and relates to the field of aviation intelligent health monitoring, and the method comprises the steps: carrying out the ground testing of a tested on-board personnel before on-board, and constructing a personalized physiological baseline threshold library; in the flight process of the aircraft, physiological data of personnel on the aircraft are collected in real time through wearable equipment, and aircraft navigation parameters are synchronously obtained and transmitted to a ground station; the ground station evaluates the stress degree of the personnel on the aircraft by combining a locally deployed physiological large model with a stored personalized physiological baseline threshold library; and when the stress degree exceeds a safety threshold value, a hierarchical regulation and control mechanism is triggered. According to the method, the physiological stability of the low-altitude aircraft personnel in the low-altitude complex airflow environment can be improved, the man-machine work efficiency level of the low-altitude manned aircraft is enhanced, and the stress state of the low-altitude aircraft personnel can be quantitatively evaluated.

Methods and systems for detecting a torque transmission failure in an aircraft power plant are provided. The system comprises a hydrostatic transmission for transmitting torque between a first shaft and a second shaft using hydraulic fluid, an a failure detector including a sensor configured to detect a pressure change in the hydraulic fluid. The failure detector is configured to generate an output indicative of the torque transmission failure in response to the pressure change having crossed a threshold.

The invention relates to a concave angle asymmetric double-layer negative Poisson's ratio superstructure and application thereof. A concave angle asymmetric double-layer negative Poisson's ratio superstructure comprises bearing pressing plates symmetrically arranged at the upper end and the lower end and a plurality of unit cell elements arranged between the two pressing plates, and is characterized in that the unit cell elements are arranged at equal intervals in the transverse direction and the vertical direction respectively to form an m * n array, each unit cell unit comprises a double-layer concave corner body, transverse connecting rods on the two sides and a vertical connecting rod on the upper portion, the lower portion of each double-layer concave corner body is an outer flat plate cell wall, and the ends of the transverse connecting rods between every two transversely adjacent unit cell units are connected into a whole. And the end parts of the vertical connecting rods between the vertically adjacent unit cell elements are connected with the outer flat plate cell walls into a whole. The deformation mode of the superstructure is influenced by the impact speed, when the speed is increased, the deformation mode transits from overall deformation to local deformation, different local deformation zones are generated, the superstructure has different horizontal contraction processes at different impact speeds, the negative Poisson's ratio effect is reduced along with the increase of the speed, and the local deformation zones are reduced. The superstructure can be applied to anti-impact protection parts of automobiles, airplanes and unmanned aerial vehicles and anti-impact protection parts of wearable protection devices.

A device for monitoring a maximum temperature reached during landing by a brake of landing gear of an aircraft. The monitoring device uses a prediction model to estimate a maximum temperature reached by the brake during landing. An error between the estimated maximum temperature and the measured maximum temperature is determined, and if the error is greater than a first predefined threshold, when the measured maximum temperature is greater than the estimated maximum temperature, or a second predefined threshold, when the measured maximum temperature is less than the estimated maximum temperature, and if a total number of errors is greater than a third predefined threshold, then a warning message is generated for attention of the flight crew and / or ground crew. It is thus possible to anticipate maintenance actions on landing-gear components.

A computer implemented method in a communications network for determining location information about an actual location of a drone comprises obtaining (302) a reported location of the drone at a first time point and obtaining (304) a measurement of radio conditions between the drone and a node in the telecommunications network, at the first time point. The method then comprises predicting (306) radio conditions at one or more locations related to the reported location of the drone, and determining 308 the location information about the actual location of the drone based on the measured radio conditions and the predicted radio conditions.

A barrier door system to control movement of persons through an opening in an interior space of an aircraft. The barrier door system having a frame with outer members that extend around a gap. A barrier door has a first panel and a second panel that are pivotally connected together with the barrier door pivotally connected to the frame. The barrier door is movable between a stowed position and a deployed position. The stowed position includes the first panel and the second panel folded together in an overlapping configuration and positioned in the gap in the frame and away from the opening. The deployed position includes the first panel and the second panel positioned away from the gap and extending across the opening.