145 results about "Airspeed" patented technology

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 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.

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.

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.

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.

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 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.

[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 neural network model predicts a maximum altitude for an aircraft based on factors. The neural network model is trained using aircraft performance tables that include specified parameters and variable parameters. The specified parameters include a residual rate of climb (RROC) threshold and a maneuver margin threshold. The variable parameters include an aircraft gross weight, a temperature, a cruise airspeed, and a present altitude. The variable parameters change within the table while the specified parameters do not. The altitude is incremented and the gross aircraft weight is determined based on a climb profile using the energy method to determine fuel consumption. After these changes, it is determined whether the parameters fail criteria for the maximum altitude. If one of the criteria is not met, then the maximum altitude for the table is determined based on the parameters. The tables and maximum altitudes are used to train the neural network model.

The invention discloses a flight control system structure of an electric tilting six-rotor unmanned aerial vehicle in a propeller aircraft mode. The flight control system structure is composed of a layered closed-loop architecture composed of outer ring attitude / height (or climbing rate) control and inner ring attitude angular rate control, and an airspeed single-loop control, trajectory guidance and actuator distribution / constraint management module. The flight control system structure of the electric tilting six-rotor unmanned aerial vehicle in the propeller aircraft mode has the remarkable advantages that under the influence of the control surface efficiency changing along with the air speed / attack angle, external disturbance and coupling, the rolling and pitching channel adjusting time is remarkably shortened, and overshoot and steady-state errors can be neglected; the airspeed is kept without overshoot, and the steady-state error is small; the yaw rate response is fast, and the coupling inhibition effect is good; the system structure calculation load is low, parameter calibration is friendly, and robustness is high; the method is suitable for attitude, airspeed and height / climbing rate cooperative control and route autonomous tracking of the electric tilt six-rotor unmanned aerial vehicle in a propeller aircraft mode.

The invention discloses an airspeed calibration method and device, a storage medium and wind tunnel calibration equipment, and the method comprises the steps: uniformly determining a plurality of real wind speeds of an airspeed sensor in an aircraft according to a target navigational speed range of the aircraft, and determining the measurement airspeed of the airspeed sensor at each real wind speed; calibrating the initial airspeed calculation parameter according to each real wind speed and the measured airspeed corresponding to each real wind speed to obtain a calibrated airspeed calculation parameter corresponding to each real wind speed; and applying the calibrated airspeed calculation parameter to airspeed calculation of the airspeed sensor. The method comprises the following steps: uniformly determining a plurality of real wind speeds according to a target navigational speed range, measuring corresponding measurement airspeed, calibrating initial airspeed calculation parameters to obtain calibrated airspeed calculation parameters under different real wind speeds, and applying the parameters to airspeed calculation of an airspeed sensor. According to the invention, rapid calibration of the airspeed sensor of the unmanned aerial vehicle can be realized without professional personnel, and the difficulty and complexity of calibration operation are reduced.

The invention discloses a method and a system for evaluating the regional residence capability of a long-endurance aircraft, and belongs to the technical field of aerospace. According to the method, based on battery parameters, solar radiation and historical data of a wind field, an airspeed time change rule of the long-endurance aircraft which meets day and night energy circulation and maximally utilizes solar energy is determined through a nested optimization algorithm; combining the airspeed and the wind speed vector to synthesize the ground speed, updating the position by adopting a fixed time step length, and iteratively solving the farthest distance of the residence point capable of going back and forth in the specified time in each direction through a dichotomy method; and finally, a closed circle is defined by the farthest points in all the directions, and quantitative evaluation of the long-endurance aircraft area residence capability is achieved. According to the method, the all-day optimal airspeed change of the long-endurance aircraft is determined through the nested optimization algorithm, and the farthest residence point in each direction is iteratively solved in combination with vector calculation and the dichotomy, so that the problem of lack of an effective evaluation method for the regional residence capability of the long-endurance aircraft in the prior art is solved; and an important basis can be provided for task planning and performance optimization of the long-endurance aircraft.

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

The invention discloses a control method for gliding parachute-opening landing of an unmanned aerial vehicle, belongs to the technical field of safe landing of the unmanned aerial vehicle, and utilizes the unmanned aerial vehicle to enter gliding and adjust various states and parameters of the unmanned aerial vehicle to enable the unmanned aerial vehicle to meet conditions of safe parachute landing. The method comprises the following steps that an aircraft follows a route in an approaching state, and an automatic accelerator is controlled by speed; in the lifting rate zeroing stage, transverse guidance is closed, longitudinal guidance is controlled by the lifting rate, and an automatic accelerator is closed; in the gliding stage, longitudinal guidance is controlled by speed, and airspeed is equal to recovery speed; in the recovery stage, the longitudinal guidance is closed, the pitch angle is equal to the sum of the increment of the climbing angle and the increment of the attack angle of the parachute opening action, and the recovery system is started and parachute opening is performed until the aircraft lands safely to complete landing.

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.

The present application relates to a kind of static pressure correction method and the drag cone system used by it, by being arranged at the position where the influence of drag cone tube static pressure hole is less, inclination sensor can obtain the angle of attack of drag cone tube in flight calibration process, improve the precision of drag cone calibration.The beneficial effects of the present application are that traditional drag cone method has the advantages of wide airspeed range and simple supporting facilities.In addition, it can further eliminate the static pressure measurement error caused by the angle of attack of drag cone tube, and after correction, it can improve the accuracy of reference static pressure measurement, and the installed inclination sensor can not cause significant adverse effects on static pressure measurement.

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 present application relates to the technical field of aircraft control, and discloses a distributed power ground effect aircraft convenient to control and a control method thereof; the present application replaces traditional rudder control with a distributed power system, realizes full-range efficient control from static to high speed, and directly generates multi-dimensional control torque through power difference by adopting independent control of positive and negative rotors in cooperation with differential adjustment, so that not only the failure defect of traditional rudder at low speed stage is completely eliminated, but also precise attitude adjustment and in-place turning are completed at zero airspeed, and the response capability of the aircraft in a complex environment is improved; in combination with air cushion effect optimization design, directional airflow of front and rear rotors forms a stable air cushion in combination with a skid structure, low-speed maneuverability is improved, a body-level turning radius is realized, and the risk of touching water is avoided. The distributed architecture simplifies the mechanical system, relies on pure electric driving, seamlessly integrates short take-off and landing, cruising and landing processes, reduces energy consumption, and enhances adaptability in narrow water areas.

The application relates to a longitudinal double-rotor unmanned aerial vehicle avionics system, which comprises a starting motor, an engine, a generator, an ACDC airborne power supply, a starting battery, a first battery, a second battery, a knob switch, a gear shifting switch, an ECU, a wind cooling fan, a water cooling fan, a control collection box, a flight control, a radio, a gyroscope, a rudder, a temperature sensor, an oil level sensor, a receiver, a lamp panel and an air speed meter; the generator serves as a main power supply of the unmanned aerial vehicle; the first battery and the second battery jointly constitute a standby power supply of the unmanned aerial vehicle; the wind cooling fan control signal, the water cooling fan control signal, the air speed meter signal and other control signals which have little influence on flight and are prone to interference are concentrated in the control collection box and are separately controlled with the flight control; the application determines the minimum avionics system which satisfies normal flight of the longitudinal double-rotor unmanned aerial vehicle, clearly defines the key components of the unmanned aerial vehicle flight, and improves the flight safety of the unmanned aerial vehicle through improvement of the power supply system operation scheme and the separate control method.

An electrically powered aircraft is configured estimate a true angle-of-attack (AoA) for an aircraft. A controller may receive one or more measurements associated with an aircraft, such as vane angle values, inertial measurement unit (IMU) data, and airspeed data. Discrepancies between vane angle values measured by different AoA sensors and an aircraft's true AoA may include fuselage effects, such as flow interference around the nose at different combinations of true AoA and angle-of-sideslip (AoS) causing different local AoA at the different vanes. The controller may generate and utilize a side slip estimate to determine if the difference in vane angle values is due to side slip or a malfunctioning AoA sensor. In some cases, if the difference in the vane angle values is determined to be due to a malfunctioning AoA sensor, then the controller may perform one or more actions (e.g., deactivate a stall barrier protection)

The application belongs to the technical field of radar signal processing, and specifically discloses a method for detecting airspeed targets based on long-time accumulation, which comprises the following steps: firstly, segmenting the echo signals after pulse compression according to the number of pulses, and compensating the distance migration and Doppler spread of the envelope of the echo signals in each segment; secondly, performing in-segment coherent accumulation on the compensated results in each segment; and finally, performing envelope migration and non-coherent accumulation between segments on all the echo signals, so that the energy of the echo signals can be effectively accumulated. The long-time accumulation method of time segmentation, in-segment coherent accumulation and inter-segment non-coherent accumulation can solve the problems of long-time coherent accumulation calculation and echo signal coherent deterioration, improve the detection performance, and is easy to implement in engineering.

The invention relates to the technical field of trajectory planning, in particular to an energy-optimal aircraft static gliding whole-process flight trajectory planning method, which comprises the following steps: acquiring a vertical rising wind speed under a circling radius at any flight height; constructing a flight performance model considering the unpowered gliding performance of the aircraft with circling maneuver; obtaining net energy obtaining power of the aircraft; obtaining an optimal roll angle and an optimal hovering airspeed corresponding to the optimal hovering radius of the aircraft at different flight heights; acquiring a flight path in the upflow; and obtaining a whole-process flight path. According to the invention, an integrated planning framework of deep fusion of multiple engineering and security constraints is used to solve the problem of how to safely complete flight in an actual task of the unmanned aerial vehicle in a real flight scene, and meanwhile, the whole-process trajectory planning of entering, hovering energy obtaining and exiting is accurately completed with low consumption; therefore, a complete, safe, efficient and executable unmanned aerial vehicle static gliding energy obtaining solution is provided.

The invention discloses a fixed-wing unmanned aerial vehicle self-adaptive flight control method based on a variable-pitch propeller and the self-adaptive variable-pitch propeller. The method comprises the steps that multi-sensor data such as the air speed, the height, the vertical speed, the accelerator and the rotating speed are collected in real time; based on a predefined comprehensive scoring model, scores of all flight stages are calculated through piecewise scoring functions of five dimensions of height, air speed, vertical speed, accelerator and time, and a strict conversion logic and hysteresis mechanism is combined to intelligently judge that the aircraft is in a take-off, climbing, cruising or descending stage at present; according to the judged stages and the real-time state, the pitch angle of the variable pitch propeller and the position of an engine throttle are dynamically calculated, control parameters are adjusted according to the stages, and height and rotating speed compensation is introduced; meanwhile, three-level safety protection including warning, active intervention and emergency response is implemented, and thrust and system efficiency are calculated online to optimize energy efficiency. According to the invention, self-adaptive, high-safety and high-efficiency control in the whole flight process is realized.