339 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.

Systems and techniques for wind estimation are presented herein, such as using Uncrewed Aircraft (UA). No on-board flow sensor is required, and the present subject matter does not require estimation of thrust or drag forces. For example, using only GPS (or another indicium of groundspeed) and orientation sensors, wind vectors in the Earth-fixed frame can be estimated during turning maneuvers. The “Wind-Arc” technique described herein was verified using simulation. Also, when applied to experimental flight test data, the technique works and follows both airspeed and wind speed trends.

The invention provides a pilot landing operation analysis method and device, electronic equipment and a storage medium. The pilot landing operation analysis method comprises the steps that external environment parameters and aircraft running state parameters formed in the process that a pilot operates an aircraft to land are obtained; 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, hidden states corresponding to the state transition matrix are obtained, so that pilot landing operation analysis is completed, the hidden states are used for representing all states of an attention distribution mode of a pilot, and the hidden states comprise the state without attention loss, the state of attention loss of a yaw angle and the state of attention loss of a height; and any one or more of states in which attention to airspeed is missing. The landing operation of the pilot is accurately and objectively analyzed, so that the risk is conveniently managed and controlled in a targeted manner, and the flight safety level of the pilot is improved.

The invention provides an unmanned aerial vehicle fault important characteristic parameter optimal selection method and system based on causal reasoning and a medium, and the method comprises the steps: S1, collecting fault data which comprises flight fault data collected by a gyroscope, a GPS, an airspeed meter and the like; s2, performing data processing on the fault data; s3, causal structure learning: performing causal structure learning on the data after data processing through a VAE decoder and a VAE encoder in sequence to obtain a weighted causal DAG representing a causal relationship between the features; and S4, important fault feature selection: after the weighted causal DAG is obtained, node scores are calculated, then the node scores are sorted in a descending order, the top F fault feature variables are selected as feature subsets, and F is a positive integer. The method can effectively screen out the characteristic parameters of the fault of the unmanned aerial vehicle, can be used for fault diagnosis, facilitates the reduction of the fault frequency of the unmanned aerial vehicle, and improves the safety and reliability of the unmanned aerial vehicle.

The invention relates to the field of flight control of aircrafts, and provides a control method and system for high-precision tracking of a landing glide line by an unmanned aerial vehicle, and the unmanned aerial vehicle. The control method comprises the steps that S1, a control structure for controlling the height of the elevator is determined, specifically, a double-ring cascade PID control structure is adopted, a pitch angle ring is a PID control structure, and a height ring is a PID control structure; a lag network is introduced between the pitch angle ring and the height ring; s2, an accelerator channel adopts a control structure that the air speed is controlled by an accelerator; s3, a landing height instruction and an air speed instruction are obtained, and the unmanned aerial vehicle is controlled to land along the designed glide line track through a control structure for controlling the height through an elevator and a control structure for controlling the air speed through an accelerator. The control system comprises a control module for controlling the height of the elevator, an accelerator control module and a landing glide control module. The method has the beneficial effects that the accuracy of tracking the landing glide track height by the unmanned aerial vehicle is effectively improved, and the safety of the unmanned aerial vehicle in the landing process is ensured.

The invention relates to the field of aviation safety, and provides a flight bumping area prediction method and system based on air-ground data fusion, and the method comprises the steps: obtaining a measurement area, and obtaining a barometric altitude through meteorological data; acquiring air detection data of the measurement aircraft so as to calculate a horizontal airspeed component; checking according to the horizontal airspeed component and the ground speed component to obtain a high-altitude wind speed component; generating an airborne data matrix, calculating the distance between a measurement point in the airborne data matrix and a grid point in a measurement area to obtain a latitude and longitude distance, and performing weight accumulation to obtain a grid point wind speed; wind field deformation is calculated, a jolting index is calculated according to the wind field deformation and normalized, a normalized jolting index is obtained, and a comprehensive jolting index is obtained through the normalized jolting index; and obtaining a target jolting index according to the ground measurement jolting index and the comprehensive jolting index to realize jolting prediction. According to the invention, accurate flight bumping prediction can be realized.

An aircraft control system includes a longitudinal control module, an engine torque control module, and an actuator control system. The longitudinal control module is configured to generate a desired torque value and a desired elevator position value for an aircraft based on a desired airspeed value, a desired altitude value, an actual airspeed value, and an actual altitude value. The engine torque control module is configured to generate a desired power lever position value based on the desired torque value and a measured engine torque value that indicates a measured engine torque in the aircraft. The actuator control system is configured to generate a power lever position command and an elevator position command for the aircraft based on the desired power lever position value and the desired elevator position value.

A method is disclosed herein. The method includes receiving, by a processor, an ejection command for ejecting an ejection seat from an aircraft; determining, by the processor, an airspeed and an altitude of the ejection seat; responsive to the altitude of the ejection seat indicating a second mode of operation, determining, by the processor, whether the airspeed of the ejection seat is at or above a first threshold and below a second threshold; and responsive to the airspeed of the ejection seat being at or above the first threshold and below the second threshold: sending, by the processor, a command to deploy a drogue parachute; and, responsive to a first predetermined time period expiring, sending, by the processor, a command to deploy a main parachute and sever the drogue parachute based on a variable drogue severance delay.

A method is disclosed herein. The method includes receiving, by a processor, an ejection command for ejecting an ejection seat from an aircraft; determining, by the processor, an equivalent airspeed and an equivalent altitude of the ejection seat; and, responsive to the airspeed and the altitude indicating a second mode of operation: sending, by the processor, a command to deploy a drogue parachute; determining, by the processor, a variable drogue severance delay; and responsive to the variable drogue severance delay expiring, sending, by the processor, a command to sever the drogue parachute and deploy a main parachute.

The invention discloses a composite wing unmanned aerial vehicle transition process control method based on a fractional order sliding mode, and particularly relates to the field of composite wing unmanned aerial vehicle flight control. A Newton-Euler equation is adopted to carry out integrated modeling on a composite wing unmanned aerial vehicle transition process, and a longitudinal nonlinear mathematical model is deduced; designing a novel fractional order reaching law by combining a double-power reaching law and a fractional order calculus thought on the basis of a traditional exponential power reaching law, and constructing a pitch angle, air speed and four-rotor height fractional order sliding mode controller; and force and torque of the four-rotor system and the fixed-wing system are distributed by adopting an airspeed weight control distribution strategy, so that output of an actuator is solved. The sliding mode variable structure control is applied to the transition control of the composite wing unmanned aerial vehicle, and the proposed fractional order sliding mode control method improves the dynamic quality of the system, reduces the control steady state error, improves the robustness of the flight control system of the composite wing unmanned aerial vehicle, and realizes the stable control of the transition process.

The invention discloses a fixed-wing aircraft large-angle-of-attack landing control method based on a dynamic reference trajectory, and belongs to the technical field of fixed-wing aircraft autonomous landing, and the method comprises the following steps: establishing a landing control strategy based on a kinematic model and an attitude angle conversion model in the fixed-wing aircraft autonomous landing process, designing two-stage annular auxiliary line radiuses of an air flight stage of the fixed-wing aircraft, designing an expected track yaw angle, an expected track inclination angle and an expected roll angle of the fixed-wing aircraft, solving the expected yaw angle and the expected pitch angle, and constructing a control law of an attack angle when the fixed-wing aircraft is about to land; aiming at a flight stage and a taxiing stage of fixed-wing aircraft landing, a proper angular velocity controller and an air speed controller are designed based on a dynamic reference trajectory, so that the pose of the fixed-wing aircraft and a large attack angle when the fixed-wing aircraft is about to land are controlled, the dangerous conditions of collision and stall are avoided, and the safety of the fixed-wing aircraft is improved. And the accuracy and safety of autonomous landing control are improved.

The invention relates to the technical field of aircraft control, in particular to an aircraft airspeed attack angle estimation method based on augmented generalized Kalman filtering, which comprises the following steps: acquiring sensor data and correcting acceleration, calculating initial estimation of an airspeed attack angle, constructing an error matrix and scoring, extracting variation trend adjustment parameters, and calculating a state variation coefficient. And multi-source information is fused to output the combined airspeed angle of attack quantity. According to the method, through acceleration component body coordinate system projection and back projection correction, an airspeed initial estimation value is constructed in combination with a height difference value, a physical matching degree is enhanced, a slope score and trend change extraction are performed to establish a dynamic compensation mechanism, and the adaptive capacity to a non-stationary flight state is improved; the introduction of the pitch angle speed, the control surface deflection and the aerodynamic coefficient bias makes the state evolution more structural, a change coefficient is established through ratio calculation, the coupling of state prediction and maneuvering dynamic is enhanced, the weighted fusion of an observation residual error and a predicted value is realized, and the convergence capability of an estimated value to complex aerodynamic interference is improved.

The invention discloses a combined type car running vehicle pneumatic flight intelligent device and method. According to the device and method, technologies such as multi-rotor wings, especially turning, folding and pneumatic layout of a vehicle door into wings, 5G communication, Beidou and a Beidou enhancement system, a composite new material and the like are integrated to form a multifunctional carrying tool; ict technology detection and 5G remote audio and video interactive consultation are used for monitoring life dynamic information and fatigue state indexes of a pilot in real time, and six symmetrical and obliquely upward rotary wings are used for assisting flight and hybrid braking; a rotary aeronautical instrument cross pod is additionally arranged, and a one-key starting and airline-selecting take-off flight control system is arranged; a mobile phone navigator and a support are arranged in a cockpit, an intelligent mobile phone, a CAN bus and an IMU inertial unit are used for detecting the acceleration and the flight attitude, and an aerocar joystick, an airspeed tube and an empennage structure of the airspeed tube are iteratively upgraded and energized, such as an elevator, a navigation light and a rudder; the system is integrated, and 200-meter short-distance take-off and landing and run-up take-off are carried out; therefore, the hovercar can stably fly on land, sea and air and safely drive.

The invention discloses an airspeed tube with deicing and drainage functions. The airspeed tube is composed of main parts such as a fusiform drainage opening, a waterproof deicing plate, an adapter flange, a protective sleeve and a mounting seat. The front end of the airspeed tube adopts a pointed-cone-shaped streamline design, so that airflow fluctuation and turbulence can be reduced, and the flight energy consumption is reduced. And the gas diversion trenches are arranged around, so that water drops or ice crystals can be prevented from adhering, and the measurement precision is improved. The total pressure port is located at the foremost end of the fusiform drainage port and can make contact with stable airflow, and measurement reliability is guaranteed; the static pressure openings are formed in the radial side edge of the fusiform drainage opening, and redundant design is provided for the multiple static pressure openings. And the waterproof deicing plate is provided with a deicing heating wire and a miniature water outlet and is used for heating, deicing and draining water. The total pressure pipeline and the static pressure pipeline filter melt water through a waterproof breathable film, and the melt water is prevented from entering the rear-end pressure pipeline. The heating wire and the deicing heating wire wound outside the total pressure conduit and the static pressure conduit form a series loop to improve the cooling efficiency. The design of the total pressure pagoda joint and the static pressure pagoda joint with different lengths at the tail saves the structural space and is convenient to install and maintain. While light-weight and small-size development is met, the functions of low-power-consumption intelligent deicing and drainage after deicing are achieved at the same time.

The invention belongs to the field of flight tests of unmanned aerial vehicles, and particularly relates to a calibration method for an airspeed system of a round-trip flight unmanned aerial vehicle with a fixed course angle. Comprising the following steps: step 1, selecting a test point in an aircraft flight envelope; step 2, the aircraft flies back and forth under the condition corresponding to the test point, and flight data of the going path and the returning path in the stable level flight stage are obtained; step 3, respectively selecting a plurality of flight data points of a going path and a returning path from the flight data, acquiring a course angle and a track angle of each flight data point, and according to the course angle and the track angle, judging whether the non-uniformity of the wind field in the flight process meets the requirement; if so, entering step 4; if not, returning to the step 2, and re-executing the test point; 4, acquiring the ground speed of each flight data point, and calculating the wind speed according to the ground speed; 5, calculating the vacuum speed in the flight process according to the wind speed and the ground speed; and 6, calibrating the air speed of the test point according to the vacuum speed.

A system and method for tilt dead reckoning is provided. The system and method allows an autopilot of an unmanned aerial vehicle (UAV) to perform dead reckoning with a hovering vehicle during GNSS signal loss by estimating the position and velocity of the vehicle based on its pitch and roll angles and known vehicle dynamics. The position and velocity are estimated using tables set up by a UAV integration engineer that provide the expected airspeed at given pitch and roll angles in steady state. This allows the UAV to attempt to follow waypoints when GNSS signal is lost without using any additional sensors.

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 a flexible pressure sensor array for measuring the airspeed of an unmanned aerial vehicle and a preparation method of the flexible pressure sensor array. The sensor array comprises a net-shaped structure and a plurality of flexible pressure sensors arranged at nodes of the net-shaped structure. The sensor is of a sandwich structure, a sensitive layer is arranged in the middle, a flexible substrate layer and an electrode layer are symmetrically arranged up and down, the sensitive layer is composite conductive foam with a conductive elastomer filling framework, and the composite conductive foam is obtained by dipping foam in a conductive material solution for multiple times. The sensitive layer framework is obtained by forming through holes in composite conductive foam in a laser array mode and filling conductive elastomers with the through holes. According to the sensor array, heterogeneous network porous foam with good conductivity and multi-layer pressure response characteristics is used as a sensitive layer, and the sensor array has a wide pressure response range; the conductive elastomer framework improves the problem that the foam structure is easy to collapse due to compression deformation, and the service life is prolonged; by utilizing the internet design, the Young modulus of the whole structure is reduced, and the device can be closely attached to the curved surface of the leading edge of the wing in a conformal manner.

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.

A method for determining airspeed of an aircraft that includes determining a rotor model relating a power coefficient of a propeller of the aircraft to an axial inflow velocity through the propeller as a function of a set of rotor operating parameters; determining the set of rotor operating parameters by sampling an electronic control signal associated with an electric motor actuating the propeller; computing the axial inflow velocity through the propeller based on the set of rotor operating parameters using the rotor model; and determining the airspeed based on the axial inflow velocity.

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.

The invention discloses a flight state estimation method and system for a low-cost small propeller aircraft, and belongs to the technical field of navigation, guidance and control, and the method mainly comprises the steps: building parameterized models of aerodynamic coefficients and propeller thrust coefficients, the difference between the air speed and the ground speed and the derivative deviation of the normal force coefficient to the attack angle are used as system state quantities to be estimated, and the axial apparent acceleration measured by an accelerometer is used as a quantity for measurement; calculating a theoretical value of the quantity measurement by using a kinetic model; and estimating the flight state of the aircraft based on extended Kalman filtering recursion. According to the method, an existing navigation system is utilized, and a simple, practical and effective flight state estimation means is provided for the propeller propulsion aircraft on the premise that no hardware needs to be added.

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.

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

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 relates to an air data indication and calibration device (1) and a method for a vertical take-off and landing aircraft, in particular a helicopter (50), for providing information of the airspeed of the vertical take-off and landing aircraft and for providing information of the height of the vertical take-off and landing aircraft. The vertical take-off and landing aircraft comprises pitot tube equipment (51) used for determining stagnation point air pressure at the position where the pitot tube equipment (51) is located; and the static pressure hole equipment (52) is used for determining the static pressure at the position of the static pressure hole equipment (52). The air data indicating device (1) comprises an airspeed and altitude determining module (2) used for determining the airspeed and altitude of the vertical take-off and landing aircraft in real time according to flight data by using a regression device (3) obtained by training an artificial neural network through training data.

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 an automatic landing control method for a large fixed-wing unmanned aerial vehicle. The landing stage of the unmanned aerial vehicle is divided into an approach flat flight stage, an approach glide stage, a first landing leveling stage, a second landing leveling stage and a landing running stage. In the approaching level flight stage, the unmanned aerial vehicle is controlled to keep the target level flight speed, the target approaching level flight height, the lateral offset distance and the track angle aligned and track the runway extension line; in the approaching gliding stage, the unmanned aerial vehicle is controlled to keep the target gliding speed and the target gliding angle; in the first stage of landing leveling, the unmanned aerial vehicle is controlled to keep the target vertical speed and airspeed; in the second stage of landing leveling, the unmanned aerial vehicle is controlled to keep the target vertical speed and the air speed; in the landing and running stage, the speed of the unmanned aerial vehicle is controlled to be reduced to 0, and meanwhile the unmanned aerial vehicle is controlled to keep the target side offset distance and the course angle on the runway center line. Automatic landing of the unmanned aerial vehicle can be accurately controlled, and the robustness and safety of automatic landing of the unmanned aerial vehicle are greatly improved.