231 results about "Airspeed" patented technology

The invention relates to the technical field of flight control, in particular to a fixed-wing unmanned aerial vehicle attitude correction and precise flight control method fused with visual navigation, which comprises the following steps: extracting ORB feature points and constructing a dynamic feature point cloud; simultaneously acquiring angular velocity and acceleration data of the inertial measurement unit; performing tight coupling fusion processing on the dynamic feature point cloud and the inertial measurement data to generate real-time pose data; inputting the real-time pose data into an attitude correction parameter calculation module, and generating an attitude correction parameter through a dynamic weight distribution algorithm in combination with a preset flight path planning model and GPS positioning data; and a control surface deflection instruction and an accelerator control instruction are generated through the multi-constraint optimization model, and unmanned aerial vehicle flight attitude closed-loop correction is completed. According to the method, aerodynamic parameters such as the air speed, the wind field lateral disturbance and the attack angle are introduced to serve as coupling input, the control surface deflection amount and limiting conditions are dynamically adjusted, and self-adaptive modeling and control over the nonlinear flight characteristics are achieved.

The invention relates to the technical field of unmanned aerial vehicles, in particular to an intelligent flight path planning and energy management system and method for a long-endurance fixed-wing unmanned aerial vehicle. Comprising an environment sensing unit; the flight path planning unit is used for planning a flight path meeting task requirements, safety requirements and energy constraints based on the flight environment information of the unmanned aerial vehicle acquired by the environment sensing unit and pre-stored performance parameters of the unmanned aerial vehicle; and the energy management unit realizes energy dynamic management and optimization according to the real-time energy state of the unmanned aerial vehicle, the flight task and the planning result of the flight path planning unit. The flight path planning unit can call working condition energy consumption data such as navigational speed and height output by the energy management unit in real time, and dynamically adjust the weight of the path to avoid high-energy-consumption flight segments; the energy management unit synchronously and intelligently adjusts a main / standby battery charging and discharging strategy to optimize energy distribution according to a task time sequence (such as waypoint priority and track curvature) of a track.

An ejection seat includes an ejection seat sequencing system. The sequencing system enables time delays to various components of the system, such the drogue parachute, the main parachute, and harness release cartridge. A sequence controller receives data for a parameter, such as airspeed, from at least two sensors. The sequence controller selects a sequence timing delay mode to implement the time delays for the components of the sequencing system. The sequence controller also accesses a lookup table in a database accessible by the sequence controller. The sequence controller uses a delay parameter of the lookup table to modify a time delay within the sequence timing delay mode, as configured by a user.

The invention relates to the technical field of aircraft electric heating deicing, in particular to a pulse heating aircraft deicing method and system.The method comprises the following steps that the ice layer thickness, the ice layer type, the airspeed and the attack angle of each aircraft surface area are collected, the inherent pneumatic sensitivity coefficient of each area is called, the aerodynamic penalty index of each area is established, and the aerodynamic penalty index of each area is calculated; and according to the values of the aerodynamic penalty indexes of the areas, performing descending order arrangement to generate a deicing area priority sequence. Through multi-parameter joint evaluation of the ice layer thickness, the ice layer type, the airspeed and the incidence angle, the heating resource distribution sequence is optimized, energy waste is avoided, in the local heating execution stage, temperature change slope real-time monitoring is adopted to replace fixed time control, the actual thermal response behavior of the ice layer can be dynamically responded, and the real-time thermal response performance of the ice layer is improved. The initial phase change process of the ice crystal is identified through a phase change starting slope threshold value, and then a medium-power pulse control mode is switched, so that a steady-state regulation heating path is formed, and the phase change control continuity is strengthened.

The invention provides an unmanned aerial vehicle draught fan wake flow observation method and system based on multi-source fusion positioning. The method comprises the steps that basic sensing data used for draught fan wake flow observation and local coordinate system parameters of a multi-source sensor are acquired; constructing a unified coordinate system based on the basic sensing data of the unmanned aerial vehicle; performing zero offset compensation on the original reading of the airspeed sensor in a static state to obtain a compensated airspeed value; based on the basic sensing data of the unmanned aerial vehicle and the compensated airspeed value, self-adaptive track control is carried out, boundary approximation processing is carried out on the fan wake flow area, and a wake flow loss field of the fan wake flow of the unmanned aerial vehicle is obtained; and calculating a wind speed difference according to the wake flow loss field. According to the method, the unmanned aerial vehicle carries various sensors, so that the key parameters of the wake flow field of the fan are accurately obtained, and reliable data support is provided for power prediction, operation optimization, layout design and the like of a wind power plant.

An air data indicating and calibrating arrangement and method for a vertical take-off and landing aircraft, in particular a helicopter, for providing information on an air speed of the vertical take-off and landing aircraft and for providing information on an altitude of the vertical take-off and landing aircraft are described. The vertical take-off and landing aircraft includes a pitot tube device for determining a stagnation air pressure at a position of the pitot tube device and a static port device for determining a static air pressure at a position of the static port device. The air data indicating arrangement includes an air speed and altitude determination module to determine in real-time the air speed and the altitude of the vertical take-off and landing aircraft from flight data by using a regressor obtained by training an artificial neural network with training data.

A vertical takeoff and landing (VTOL) aircraft includes a maximum takeoff weight (MTOW) of about 1700-kg, a cruising speed of about 80-knots equivalent airspeed (EAS) to about 100-knots EAS, a nominal range of about 300-km, a nominal flight time of about 4-hours, and at least one pair of rotors configured to provide vertical thrust. Each rotor of the at least one pair includes an about 2-meter propeller diameter, a disk loading of about 67-kg / m2 to about 110-kg / m2, and an operating speed of about 1800 rotations-per-minute (RPM) to about 2700-RPM.

The invention belongs to the technical field of atmosphere data detection, and particularly discloses a helicopter distributed three-axis atmosphere data measuring device which comprises a left three-axis atmosphere measuring assembly and a right three-axis atmosphere measuring assembly which are installed on the two sides of a nose and are the same in structure. Each airspeed tube consists of a rotary airspeed tube and an electronic component; the rotary airspeed tube and the electronic component are electrically connected through the fuselage skin opening; the rotating airspeed tubes are mounted outside skins on two sides of the machine head and used for sensing total pressure, static pressure, local attack angle and local sideslip angle of the synthetic gas flow in real time; the electronic component is arranged in the machine head skin and used for collecting the total pressure, the static pressure, the local attack angle, the local sideslip angle and the atmospheric temperature of the synthetic gas flow, conducting parameter calculation and outputting atmospheric parameters. According to the invention, the problem that a centralized atmosphere data system, a distributed atmosphere data system and a three-axis atmosphere data system have different restrictions on the use of a helicopter at present is solved.

The invention discloses a helicopter flight control performance determination method and device, equipment, a medium and a product, and relates to the field of helicopter flight control performance evaluation.The method comprises the steps that flight data of helicopter tail rotor fault flight are collected; calculating each index value by using an index model according to the flight data; the index model comprises a process index model and a transient index model; the process index model comprises a tracking index model, an airspeed keeping index model and a safety boundary index model; the transient index model comprises a landing state index model, a landing position index model and a controller delay index model; determining a comprehensive evaluation value of the flight control performance according to each index value and the weight after weight changing; according to the method and the device, the accuracy of helicopter flight control performance evaluation under the condition of the tail rotor fault is improved.

The invention discloses a cross-medium aircraft, an airspeed tube drainage method of the cross-medium aircraft and a product of the airspeed tube drainage method, and is applied to the technical field of aircrafts. The cross-medium aircraft comprises an airspeed tube and an airspeed tube drainage system, and the airspeed tube drainage system comprises a detection module used for detecting the relative humidity in a first tube or the water pressure in the first tube in the airspeed tube when the operation working condition of the cross-medium aircraft is the water surface working condition or the air working condition; and the drainage module is used for conveying gas into the airspeed tube to discharge accumulated water in the airspeed tube under the condition that the relative humidity in the first tube is greater than or equal to a first preset relative humidity threshold value or the water pressure in the first tube is greater than or equal to a first preset water pressure threshold value. According to the method, gas can be conveyed into the airspeed tube through the drainage module, accumulated water is discharged through the gas, it is ensured that measurement of fluid pressure by the airspeed tube is not interfered by the accumulated water, and the problem that fluid pressure measurement of a cross-medium aircraft is inaccurate due to the accumulated water in the airspeed tube is effectively solved.

The invention relates to the technical field of unmanned aerial vehicle control, in particular to a transition control method of a tailstock type unmanned aerial vehicle, which comprises the following steps of: establishing an inertia, airframe and airflow coordinate system, and constructing a longitudinal dynamic model based on the coordinate system; calculating a velocity trajectory by adopting a polynomial and deriving, solving coefficients by combining initial and conversion completion conditions, and generating velocity, acceleration and jerk trajectories; related variables are calculated according to the kinetic model and the speed trajectory; an expected airspeed and a pitch angle are tracked through an airspeed controller and a pitch angle controller respectively, and an accelerator control quantity and a pitch control quantity are output; and when the conversion time exceeds a specified value, judging whether mode conversion is completed or not according to whether the actual airspeed reaches the cruise airspeed, and if not, continuously tracking parameters. According to the technical scheme, the speed trajectory is smooth and controllable, it is ensured that the unmanned aerial vehicle with any takeoff weight safely completes mode conversion, and the problem that stable and reliable flight mode conversion cannot be achieved for the tailstock type unmanned aerial vehicle with any takeoff weight is solved.

Systems and methods include a controller configured to actuate a rotor assembly of an aircraft vehicle. The controller is configured to transmit an instruction to rotate the rotor assembly to an initial angle with respect to a vertical axis. The controller is further configured to receive one or more measurements associated with an airspeed of the aircraft vehicle. The controller is further configured to determine if the airspeed of the aircraft vehicle is greater than or equal to an airspeed threshold value. In response to a determination that the airspeed of the aircraft vehicle is not greater than or equal to the airspeed threshold value, controller is further configured to: determine a subsequent angle for rotating the first rotor assembly and to transmit a subsequent instruction to rotate the first rotor assembly to the subsequent angle.

This disclosure relates to flight control of electric aircraft and other vehicles. A computer-implemented method for estimating available range of an aircraft in flight is disclosed, comprising: receiving electrical information of one or more batteries measured using a first sensor; estimating aircraft-level energy based on electrical information of the one or more batteries; receiving one or more of an altitude of the aircraft or a current airspeed of the aircraft measured using a second sensor; estimating a steady-state force based on the one or more of the altitude of the aircraft or the current airspeed of the aircraft; estimating one or more of a vertical landing range or a horizontal landing range based on the one or more of the estimated aircraft-level energy or the estimated steady-state force; and displaying the one or more of the estimated vertical landing range or the estimated horizontal landing range on a display.

The invention discloses an aircraft control response data fitting method and system based on a neural network and a computer program product, and belongs to the field of flight simulation, and the method comprises the steps: obtaining historical flight parameter data, constructing a neural network model, and training the neural network model based on the historical flight parameter data; setting an initial environment state and an aircraft state of the aircraft, and determining an airspeed, an attack angle and a sideslip angle of the aircraft according to the environment state and the aircraft state; acquiring real-time state data and manipulation data of the aircraft; constructing real-time input data, and predicting through the trained neural network model to obtain acceleration, angular acceleration and engine speed change rate; and updating the aircraft state and the engine rotating speed according to the predicted acceleration, the angular acceleration and the engine rotating speed change rate to realize fitting of aircraft control response data. According to the method, the complexity can be reduced, the real-time performance, the accuracy and the dynamic adaptability of aircraft control response data fitting are improved, and the response characteristics of the aircraft in different control scenes are effectively simulated.

A final approach and takeoff (FATO) surface for a VTOL aircraft redirects downwash as outwash and may capture outwash from the VTOL aircraft to generate electricity. The FATO surface system may include a central solid plateau configured to capture downwash from the VTOL aircraft to enhance a ground effect for optimal stabilization during takeoff and landing. A floor grate surrounding the plateau may allow outwash to pass through to a subsurface where the outwash is channeled to exhausts or at least one turbine for generating electricity from the outwash. Redirecting the downwash to outwash reduces undesirable effects such as noise and airspeed associated with the VTOL aircraft. The electricity may be used to recharge the VTOL, electric vehicles parked at a vertiport of the FATO surface system, may be used by a building at the vertiport, or may be delivered to a local power grid.

An aircraft system having a LIDAR system, one or more sensors, and a control unit. The control unit includes processing circuitry configured to calculate during flight a pressure altitude, calibrated airspeed, Mach number, equivalent airspeed, static temperature, static pressure, and dynamic pressure of an aircraft based on a combination of air data measurements from the LIDAR system and the one or more sensors.

The application provides a fixed-wing aircraft airspeed calibration flight test method based on a standard machine. The method comprises the following steps: determining the weight of the fixed-wing aircraft as a first weight, setting the aircraft to a specified take-off configuration, and taking off according to a take-off procedure; the standard machine flies to a specified airspace according to a plan; at an initial specified flight height, a tow cone is released to a specified length, and the tow cone can obtain an accurate static pressure value in the current area without being affected by a spoiler; the flap and slot wing configuration of the fixed-wing aircraft is set to a first flap and slot wing configuration; the aircraft is trimmed at a given speed; the speed in the above step is set to other speeds within the speed envelope of the first flap and slot wing configuration, and the steps are repeated; the standard machine obtains airspeed errors under the first flap and slot wing configuration and different speeds; and the airspeed errors under the first flap and slot wing configuration and different speeds are fitted to obtain an airspeed correction curve under the first weight, the first flap and slot wing configuration. More accurate airspeed errors can be obtained.

A method for operating an aircraft includes determining (902A, 902B, 904, 905) a shaft horse power requirement of a multi-engine power plant, comparing (906) the shaft horse power requirement to a threshold value, and activating (908) an economy mode of operation or inhibiting (910) the economy mode of operation based upon the results of the comparison step (906). The threshold value may be determined from a power rating for an engine of the multi-engine power plant, and may be adjusted and / or delayed, depending upon whether the current state is economy mode active or economy inhibited. The shaft horse power requirement can be calculated from a control input, such as collective position, from a collection of flight parameters, including airspeed, aircraft gross weight, altitude, and the like, or from a measurement of actual shaft horse power production by the engines

The present invention discloses a multi-rotor airspeed solution method, device and computer-readable storage medium, wherein the method comprises: establishing an airspeed solution mathematical model; defining a function f(V) according to the airspeed solution mathematical model in each solution cycle, and combining the real-time measured θ inclination angle and a x Acceleration, using the Newton iterative algorithm to solve the root of the equation f(V) = 0 in real time, as the real-time airspeed V. The present invention implements an airspeed calculation solution based on dynamic principles, fully utilizing dynamic principles and combining the unique control method of multi-rotor. Without adding hardware sensors, known attitude measurements and acceleration measurements are used as algorithm inputs, and the software algorithm calculates airspeed data with the required accuracy. This simplifies the hardware structure, reduces the risk of electrical failure, and saves manufacturing, debugging, and maintenance costs.

The application discloses a UAV cluster integrated control system based on formation flight, and relates to the field of flight control technology.The system comprises a flight acquisition module, an atmosphere acquisition module, a cluster control module and a display module.The flight acquisition module acquires data and stabilizes the output of posterior coordinates, ground speed and attitude angle through drift compensation Kalman filtering.The atmosphere acquisition module relies on a database of aerodynamic characteristics, and solves airspeed, barometric altitude and angle of attack through a model of atmospheric solution that is fused with model-specific aerodynamic correction coefficients and adaptive weighting.The cluster control module establishes an airflow disturbance model to quantify airflow disturbance intensity and generate disturbance coefficients of UAVs and adjacent UAVs, and dynamically outputs adjustment instructions with the minimum formation deviation and the stable attitude of a single machine as the target.The display module is used for 3D visualization monitoring of flight formation.The system significantly improves the stability and adaptability of formation flight, and is suitable for UAV cluster cooperative operation in complex scenarios.

The application discloses a kind of subsonic aircraft airspeed measurement method and system, it is related to aerospace field, the method comprises: obtaining the dynamic pressure voltage value of aircraft external gas measured by silicon piezoresistive pressure difference sensor and the temperature value of silicon piezoresistive pressure difference sensor measured by temperature sensor under different standard temperature points different standard dynamic pressure pressure value;Characteristic coefficient value of each subset is calculated;According to the characteristic coefficient value of subset and the dynamic pressure voltage value and temperature value in subset, the uncorrected dynamic pressure corresponding to each subset is calculated;Uncorrected dynamic pressure is corrected and calibrated, and the calibrated dynamic pressure corresponding to each subset is obtained;Set range dynamic pressure is calculated, and set range dynamic pressure is obtained by total pressure value minus static pressure value;Target dynamic pressure is calculated according to calibrated dynamic pressure and set range dynamic pressure;According to target dynamic pressure, the airspeed corresponding to each target dynamic pressure is determined, and the high-precision measurement of subsonic aircraft airspeed is realized.

The invention relates to the technical field of unmanned aerial vehicle parts, in particular to an unmanned aerial vehicle waterproof airspeed tube convenient to disassemble and assemble, which comprises an air inlet piece, a pressure separation piece, an air inlet pipe and an airspeed tube. The air inlet pipe and the airspeed tube are both mounted in the pressure separation piece and are not directly communicated with each other; the air inlet piece is connected with the pressure separation piece, and external air enters the airspeed tube through the air inlet pipe to be measured. An air inlet cavity and a water baffle are arranged in the air inlet piece, and primary drainage is achieved through a first drainage assembly; the pressure separation part comprises a horizontal pressure transmission part and an inclined height difference flow guide part and is provided with a second drainage assembly, and the second drainage assembly comprises a water collection cavity and a second drainage port and is installed between the air inlet pipe and the airspeed pipe to achieve air-water separation and secondary drainage. The airspeed tube air inlet hole is higher than the air inlet tube exhaust hole, and accumulated water is further prevented from flowing backwards. The device is reasonable in structure, has good waterproof and drainage capacity, can improve the airspeed measurement precision, and is suitable for various unmanned aerial vehicle platforms with fixed wings, vertical take-off and landing and the like.

This application relates to the field of eyepoint data compensation technology, providing a method and system for compensating eyepoint data in flight simulators, solving the problems of poor real-time performance and low accuracy of eyepoint data in flight simulations. The method includes: establishing a shared memory region between the simulation logic processing module and the eyepoint data sending module of the flight simulator; this region contains updated state variables in the form of frame IDs and a first-in-first-out shared eyepoint data area; after completing a frame attitude calculation, the simulation logic processing module updates the frame ID; the eyepoint data sending module obtains the latest eyepoint data through the position corresponding to the latest frame ID; then, combined with the instructor's station settings, airspeed, and ground speed data, the aircraft attitude mode is determined to be stationary, stable, or maneuvering; finally, linear interpolation is used for the stable mode, and a pre-trained lightweight LSTM is used for the maneuvering mode to predictively compensate the latest eyepoint data, obtaining the compensated data. This application improves the real-time performance and accuracy of eyepoint data in flight simulations.

This invention discloses a control system and launch method for an airborne folding-wing unmanned aerial vehicle (UAV). Before the launch of the sub-unit, the mother aircraft sends a voltage signal to the sub-unit, which is continuously monitored by the sub-unit's flight control system. Upon reaching the launch point, the mother aircraft is powered off, the unlocking device is activated, and the flight control system initiates control after the voltage monitoring disappears for a period of time. At this moment, the V-tail controls pitch and participates in roll control. After a certain period, the flight control system unlocks the locking servo, causing the wings to begin unfolding until they are fully extended. At this time, the ailerons participate in control, entering the unpowered dive-pull control phase. After a certain period, the wing control is locked. Subsequently, the flight control system determines that the attitude and airspeed are stable. Once stable, the engine is started, and the engine speed is determined. If the start fails, the start operation is repeated three times. After the engine starts, the flight path is executed normally, the ground station launch process is exited, and normal flight control logic is entered. This invention enables full-process control of an airborne folding-wing UAV.

[Object] To detect airspeed and an airflow direction with respect to an airframe of a motorized aircraft with high accuracy without increasing the cost and weight and rapidly control attitudes of an electric propulsion system and the airframe in accordance with fluctuations of the airspeed and airflow direction.[Solving Means] An electric propulsion system control device includes: a first airspeed measurement unit that is mounted on an airframe of an aircraft and includes a first propulsion system parameter detector that detects a propulsion system parameter, the propulsion system parameter being a parameter of an electric propulsion system, the electric propulsion system being driven by an electric motor and rotating about a rotation axis as a center, and a first airspeed calculator that calculates first airspeed on a basis of the propulsion system parameter, the first airspeed being airspeed with respect to a first direction that is a direction of the rotation axis; a second airspeed measurement unit that is mounted on the airframe and measures second airspeed, the second airspeed being airspeed with respect to a second direction different from the first direction; and an airflow calculator that calculates airspeed and airflow direction with respect to the airframe on a basis of the first direction and the first airspeed and the second direction and the second airspeed.

The application provides a pilot landing operation analysis method and device, electronic equipment and a storage medium, comprising: obtaining external environment parameters and aircraft operation state parameters formed in the process of the pilot operating the aircraft to land; calling a neural network model and inputting the external environment parameters and the aircraft operation state parameters into the neural network model to obtain a state transition matrix at each time point; based on the state transition matrix, obtaining a hidden state corresponding to the state transition matrix to complete the pilot landing operation analysis, wherein the hidden state is used to represent each state of the attention allocation mode of the pilot, and the hidden state includes any one or several of a state without attention loss, a state of missing attention to a yaw angle, a state of missing attention to an altitude, and a state of missing attention to an airspeed. The pilot landing operation is accurately and objectively analyzed, so as to facilitate risk control and improve the flight safety level of the pilot.

An airspeed instrument that employs one or more circulation chambers to receive and circulate incoming airflow to separate moisture from the air that can be then expelled via a drainage port. The exemplary airspeed instrument apparatus can be implemented as a small instrument appropriately sized for small UASs and aircraft. The separation chamber has an angled bottom surface that leads to a sampling port positioned at the top of the chamber, which prevents moisture from entering the sampling port into the measuring electronics. In some embodiments, the exemplary airspeed instrument apparatus is configured with an integrated static port sensor.

The invention relates to the technical field of aircraft power control, and discloses an artificial intelligence optimized vertical fixed wing power control method and system, and the method comprises the steps: obtaining the original airspeed data of an airspeed sensor, the ground speed data of a global positioning system and the acceleration data of an inertial measurement unit, and generating the ground speed of an aircraft based on the data fusion. Under the low-wind-speed condition, the ground speed of the aircraft is used as the reference airspeed, the difference between the original airspeed data of the airspeed sensor and the reference airspeed is continuously compared, and the sliding average value of the difference is calculated. When the sliding average value continuously exceeds the preset tolerance range, the method can identify that the airspeed sensor has continuous calibration drift. And then, according to the identified continuous calibration drift, performing real-time compensation on original airspeed data of an airspeed sensor to obtain calibrated airspeed data, and inputting the calibrated airspeed data into flight mode switching judgment logic to trigger an action instruction of a tilting mechanism.

A sensor unit located on an air vehicle enables the measurement of air speed, engine torque and lateral acceleration of the air vehicle. A flight control computer enables a first adjustment command to be generated using the air speed and engine torque data measured by the sensor unit. A controller run in the flight control computer enables a second adjustment command to be generated using the lateral acceleration data. A gain coefficient of the controller is determined by the flight control computer based on the air speed measured by the sensor unit. The flight control computer collects the first adjustment command and the second adjustment command is generated by the controller so that it generates a final adjustment command.

The invention provides an automatic balancing device and method for the braking torque of an airplane in the high-speed taxiing stage. The device comprises a braking pressure sensor, an anti-skid control unit, a servo valve, a speed sensor, an instruction sensor and a central controller. The signal output end of the brake pressure sensor is respectively connected with the left brake pressure signal input end and the right brake pressure signal input end of the central controller; the signal output end of the speed sensor is connected with the airplane wheel rotating speed signal input end of the antiskid control unit; the output end of the instruction sensor is connected with the brake instruction input end of the central controller; the collaborative pressure relief instruction output end of the central controller is connected with the collaborative pressure relief instruction input end of the anti-skid control unit; an anti-skid current output end of the anti-skid control unit is connected with a control current input end of the servo valve; the pressure output end of the servo valve is connected with a brake pipeline of a corresponding airplane wheel; the input end of the instruction sensor is connected with a brake pedal; and the airspeed signal input end of the central controller is connected with an airspeed system.