2561results about "Aircraft indicators" patented technology

A control unit for an aircraft capable of hovering or for a flight simulation system of the aircraft is described; the control unit is programmed to:receive in input a plurality of data associated with equipment and / or kits actually installed or simulated on the aircraft and / or with the operating configuration of the equipment and / or kits; the equipment and / or kits causing a reduction of the actual or simulated forward velocity of the aircraft;store a table defining a correspondence between each datum and a respective value of a first signal associated with a value of maximum forward velocity of the aircraft; anddisplay the lowest of the first signals.

A control unit for an aircraft capable of hovering or for a flight simulation system of the aircraft is described, the control unit is programmed to receive at input at least a first signal associated with real or simulated flight parameters of the aircraft and to provide at output a second signal associated with a forward velocity to never exceed of the aircraft; the control unit is programmed to process a third signal associated with an actual or simulated weight of the aircraft; to associate plurality of tables to respective intervals of values of the third signal; each table associating a plurality of values of the second signal with a respective first signal; and to process the second signal, based on the table associated with the relative interval of the third signal and the respective first signals.

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.

The invention discloses a supersonic aircraft head thermal protection structure, and belongs to the technical field of hot end component cooling, the supersonic aircraft head thermal protection structure comprises a cooling system, a preheating system, a CBC power generation system and a secondary cooler, the cooling system comprises a coolant liquid storage tank, an electromagnetic adjustable pump and a cooling channel arranged in the wall surface of the aircraft head, a cooling medium flowing in a direction opposite to that of external high-temperature gas is arranged in the cooling channel; the preheating system comprises a fuel tank, a fuel adjustable pump and a fuel channel; the secondary cooler is used for transferring heat in the cooling channel into the fuel channel; the CBC power generation system comprises a heater, a primary cooler and a circulating working medium, the heater is used for transferring heat in the cooling channel to the circulating working medium, and the primary cooler is used for transferring heat in the circulating working medium into the fuel channel. Efficient thermal protection of the head of the supersonic aircraft can be achieved, the risk of structural failure caused by high temperature is reduced, and the safety of the aircraft is improved.

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 relates to a hydraulic oil leakage monitoring method and system for an A320 hydraulic system. The method comprises the steps that after oil mass parameter data and auxiliary parameter data of an A320 hydraulic system are obtained, a normal state message is obtained and sent through a preset hydraulic oil leakage normal state message triggering model, meanwhile, whether the current oil mass parameter data exceed a current preset threshold value or not is judged through a preset hydraulic oil leakage alarm message triggering model, and if yes, the normal state message is sent to the A320 hydraulic system; if yes, generating and sending a hydraulic oil leakage alarm message based on the current oil quantity parameter data and the auxiliary parameter data, otherwise, returning to obtain the parameter data, and correcting the preset threshold by the current auxiliary parameter data; receiving and analyzing the current normal message and the hydraulic oil leakage alarm message, and correcting and filtering the current hydraulic oil leakage alarm message by using the normal message to obtain an effective message; and displaying and sending the effective message. Compared with the prior art, the method has the advantages that the real-time performance and the accuracy of leakage detection of the A320 hydraulic system are improved, and the like.

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 discloses a stable coolant supply structure for a sweating cooling system of a hot end component of an aircraft, the hot end component is arranged to be of a porous structure, that is, directional pore channels are densely distributed in the hot end component of the aircraft in the advancing direction; meanwhile, the inner space of the aircraft hot end component is divided into one or more coolant cavities, the front end of each coolant cavity corresponds to a directional pore channel of the corresponding cooling area, and the rear end of each coolant cavity is communicated with a branch pipeline with a control valve. The branch pipelines are gathered and then communicated with an output port of the coolant driving system, and all the control valves are connected with the regulator. The invention further discloses a stable coolant supply method of the sweating cooling system for the hot end part of the aircraft. The invention belongs to the technical field of thermal protection of aerospace aircrafts, and solves the problem that a coolant in a sweating cooling system of a hot end component of an aircraft cannot be stably supplied in a variable incoming flow environment in the prior art.

Apparatuses, systems, and methods are disclosed for establishing visually detectable ground markings, including in daylight, and designating visually detectable unsafe ground traffic areas from safe ground traffic areas proximate to and beneath an aircraft by providing visually detectable ground images formed from high intensity light energy sources located on an aircraft.

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.

A sidewall assembly includes a window and a sidewall portion having an inner-facing side and at least partially surrounding the window. The sidewall assembly further includes a transparent display panel coupled to the sidewall portion covering the inner-facing side of the sidewall portion and the window. The sidewall assembly further includes a display controller operatively coupled with the transparent display panel. The sidewall assembly further includes a window shade movable between an open position to allow light to pass between the window and the transparent display panel and a closed position to obstruct light from passing between the window and the transparent display panel. The sidewall assembly further includes a blackout feature disposed between the window and the transparent display panel and a blackout control device configured to control the blackout feature to selectively obstruct light from passing between the window and the transparent display panel.

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 belongs to the technical field of aircrafts, and particularly relates to a multifunctional photovoltaic panel cleaning unmanned aerial vehicle and a cleaning method thereof. The invention provides a multifunctional photovoltaic panel cleaning unmanned aerial vehicle. The multifunctional photovoltaic panel cleaning unmanned aerial vehicle comprises a vehicle body; the multiple extension frames are arranged in the circumferential direction of the machine body; the spraying device is arranged at the lower end of the extension frame; the protection device is arranged on the outer wall of the spraying device in a lifting and sleeving manner and is suitable for limiting the direction of liquid sprayed by the spraying device; when a rotating disc of the spraying device rotates to spray liquid to the periphery, the protection device moves downwards to be arranged on the outer side of the rotating disc in a surrounding mode so as to limit the spraying coverage range of the rotating disc. And through the arrangement of the protection device, the sprayed liquid can be prevented from splashing to the outside of the photovoltaic panel when the spraying device moves to the edge position of the photovoltaic panel.

A dual purpose aircraft lavatory component that can be a dual purpose aircraft lavatory door that is configured to deploy in a vehicle cabin from an initial position as an aircraft lavatory monument component into a deployed position converting aircraft lavatory monument component into an aircraft cabin lockable barrier to prevent passengers from moving along an aisle in a vehicle cabin into a vehicle area forward of the lockable barrier.