WO2021032201A1 - 一种偏航角的融合方法、装置及飞行器 - Google Patents

一种偏航角的融合方法、装置及飞行器 Download PDF

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Publication number
WO2021032201A1
WO2021032201A1 PCT/CN2020/110592 CN2020110592W WO2021032201A1 WO 2021032201 A1 WO2021032201 A1 WO 2021032201A1 CN 2020110592 W CN2020110592 W CN 2020110592W WO 2021032201 A1 WO2021032201 A1 WO 2021032201A1
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WO
WIPO (PCT)
Prior art keywords
magnetometer
gps
yaw angle
data
angle
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PCT/CN2020/110592
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English (en)
French (fr)
Inventor
张添保
李颖杰
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深圳市道通智能航空技术有限公司
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Application filed by 深圳市道通智能航空技术有限公司 filed Critical 深圳市道通智能航空技术有限公司
Publication of WO2021032201A1 publication Critical patent/WO2021032201A1/zh
Priority to US17/651,835 priority Critical patent/US11669109B2/en

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    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
    • G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
    • G05D1/0816—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
    • G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
    • G01S19/38—Determining a navigation solution using signals transmitted by a satellite radio beacon positioning system
    • G01S19/39—Determining a navigation solution using signals transmitted by a satellite radio beacon positioning system the satellite radio beacon positioning system transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
    • G01S19/396—Determining accuracy or reliability of position or pseudorange measurements
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/10—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration
    • G01C21/12—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning
    • G01C21/16—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation
    • G01C21/165—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation combined with non-inertial navigation instruments
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C1/16—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like specially adapted for mounting power plant
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/20—Instruments for performing navigational calculations
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
    • G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
    • G01S19/38—Determining a navigation solution using signals transmitted by a satellite radio beacon positioning system
    • G01S19/39—Determining a navigation solution using signals transmitted by a satellite radio beacon positioning system the satellite radio beacon positioning system transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
    • G01S19/53—Determining attitude

Definitions

  • This application relates to the technical field of aircraft, and in particular to a method, device and aircraft for yaw angle correction and alignment.
  • Unmanned Aerial Vehicle also known as UAVs
  • UAVs Unmanned Aerial Vehicle
  • the various actions (or attitudes) of the unmanned aerial vehicle are usually realized by controlling the different rotation speeds of multiple driving motors in the power system of the unmanned aerial vehicle.
  • the yaw angle is an important parameter in controlling the flight attitude of the unmanned aerial vehicle. That is, the yaw angle fusion of the unmanned aerial vehicle is particularly important for the attitude control of the unmanned aerial vehicle.
  • the unmanned aerial vehicle cannot fly in the preset direction or trajectory, or it may be linear, and may even become unstable and blow up. Therefore, how to improve the fusion accuracy and convergence speed of the yaw angle of the aircraft is of great significance.
  • the yaw angle of the aircraft (such as UAV, etc.) is usually obtained based on the data collected by the magnetometer, and the yaw angle is corrected based on the data collected by the magnetometer.
  • the yaw angle obtained by this method is easily affected by external factors. Impact, especially when the magnetometer is in a strong magnetic interference environment, the data of the magnetometer may be seriously wrong, resulting in a large deviation in the correction of the yaw angle, and the convergence of the yaw angle is slow.
  • the embodiment of the present invention provides a yaw angle fusion method, device and aircraft, which effectively improves the yaw angle fusion accuracy and convergence speed.
  • the embodiments of the present invention provide the following technical solutions:
  • an embodiment of the present invention provides a yaw angle fusion method, which is applied to an aircraft, and the method includes:
  • GPS data includes GPS location, speed, acceleration information, and GPS speed signal quality
  • IMU data includes IMU acceleration information and IMU angular velocity information
  • a fusion yaw angle is generated.
  • the method further includes:
  • the yaw angle error angle is calculated.
  • calculating the yaw angle of the magnetometer according to the magnetometer data specifically includes:
  • the magnetometer data of the standard magnetic field at the current position of the aircraft is compared to calculate the magnetometer yaw angle.
  • the determining the corrected yaw angle according to the IMU data, the GPS data, and the magnetometer data specifically includes:
  • the GPS acceleration correction weight coefficient is obtained by the product of the GPS speed signal quality and the GPS acceleration signal quality, and the GPS acceleration signal quality is calculated by the GPS acceleration information.
  • the determining the magnetometer alignment deviation angle according to the magnetometer data includes:
  • the determining the magnetometer alignment yaw angle according to the magnetometer yaw angle specifically includes:
  • the magnetometer alignment pulse is input to the enabling module of the aircraft, and when the enabling module detects that the magnetometer alignment pulse has a rising edge, the magnetometer yaw angle is used as the magnetometer pair Quasi-yaw angle.
  • the determining the GPS realignment deviation angle according to the GPS data and the IMU acceleration information includes:
  • the realignment pulse signal is input to the enabling module of the aircraft, and when the enabling module detects that the realignment pulse has a rising edge, the yaw angle deviation angle is taken as the GPS realignment Align the deviation angle.
  • an embodiment of the present invention provides a yaw angle fusion device, which is applied to an aircraft, and the device includes:
  • the acquisition module is used to acquire GPS data, IMU data, and magnetometer data.
  • the GPS data includes GPS position, speed, acceleration information, and GPS speed signal quality
  • the IMU data includes IMU acceleration information and IMU angular velocity information;
  • the corrected yaw angle module is used to determine the corrected yaw angle according to the IMU data, the GPS data and the magnetometer data;
  • the magnetometer alignment deviation angle module is used to determine the magnetometer alignment deviation angle according to the magnetometer data, the GPS data, and the corrected yaw angle;
  • the GPS realignment deviation angle module is used to determine the GPS realignment deviation angle according to the GPS data and the IMU acceleration information
  • the yaw angle fusion module is used to generate a fusion yaw angle based on the corrected yaw angle, magnetometer alignment deviation angle, and GPS realignment deviation angle.
  • the device further includes: a yaw angle error angle module, and the yaw angle error angle module includes:
  • the magnetometer yaw angle unit is used to calculate the magnetometer yaw angle according to the magnetometer data
  • the yaw angle error angle unit is used to calculate the yaw angle error angle according to the magnetometer yaw angle and the fused yaw angle.
  • the magnetometer yaw angle unit is specifically used for:
  • the magnetometer data of the standard magnetic field at the current position of the aircraft is compared to calculate the magnetometer yaw angle.
  • the corrected yaw angle module is specifically used for:
  • the GPS acceleration correction weight coefficient is obtained by the product of the GPS speed signal quality and the GPS acceleration signal quality, and the GPS acceleration signal quality is calculated by the GPS acceleration information.
  • the magnetometer alignment deviation angle module includes:
  • the magnetometer alignment yaw angle unit is used to determine the magnetometer alignment yaw angle according to the magnetometer yaw angle;
  • the magnetometer alignment deviation angle unit is used to determine the magnetometer alignment deviation angle according to the corrected yaw angle and the magnetometer alignment yaw angle.
  • the magnetometer is aligned with the yaw angle unit, specifically for:
  • the magnetometer alignment pulse is input to the enabling module of the aircraft, and when the enabling module detects that the magnetometer alignment pulse has a rising edge, the magnetometer yaw angle is used as the magnetometer pair Quasi-yaw angle.
  • the GPS realignment deviation angle module includes:
  • the yaw angle deviation angle unit is configured to calculate the yaw angle deviation angle according to the GPS acceleration information and the IMU acceleration information;
  • GPS realignment deviation angle unit for:
  • the realignment pulse signal is input to the enabling module of the aircraft, and when the enabling module detects that the realignment pulse has a rising edge signal, it uses the yaw angle deviation angle as the GPS realigns the deviation angle.
  • an embodiment of the present invention provides an aircraft, including:
  • At least one processor and,
  • a memory communicatively connected with the at least one processor; wherein,
  • the memory stores instructions executable by the at least one processor, and the instructions are executed by the at least one processor so that the at least one processor can execute the yaw angle fusion method described above.
  • an embodiment of the present invention also provides a non-volatile computer-readable storage medium, the computer-readable storage medium stores computer-executable instructions, and the computer-executable instructions are used to enable the aircraft to execute the above The fusion method of the yaw angle.
  • the beneficial effect of the embodiment of the present invention is: different from the prior art, the yaw angle fusion method provided by the embodiment of the present invention is applied to an aircraft, and the method includes: acquiring GPS data, IMU data, and magnetic force
  • the GPS data includes: GPS position, speed, acceleration information, and GPS speed signal quality
  • the IMU data includes: IMU acceleration information and IMU angular velocity information
  • the magnetic force Determine the corrected yaw angle based on the gauge data
  • based on the GPS data and the IMU acceleration information Determine the GPS realignment deviation angle
  • the embodiment of the present invention can effectively improve the fusion accuracy and convergence speed of the yaw angle.
  • FIG. 1 is a schematic block diagram of a method for fusion of yaw angles according to an embodiment of the present invention
  • FIG. 2 is a schematic diagram of a yaw angle fusion algorithm provided by an embodiment of the present invention
  • FIG. 3 is an overall schematic diagram of a yaw angle fusion method provided by an embodiment of the present invention.
  • FIG. 4 is a schematic flowchart of a yaw angle fusion method provided by an embodiment of the present invention.
  • FIG. 5 is a schematic diagram of a process for calculating a yaw angle error angle provided by an embodiment of the present invention
  • FIG. 6 is a detailed flowchart of step S401 in FIG. 5;
  • FIG. 7 is a detailed flowchart of step S20 in FIG. 4;
  • FIG. 8 is a detailed flowchart of step S30 in FIG. 4;
  • FIG. 9 is a detailed flowchart of step S31 in FIG. 8;
  • FIG. 10 is a detailed flowchart of step S40 in FIG. 4;
  • FIG. 11 is a detailed flowchart of step S42 in FIG. 10;
  • FIG. 12 is a schematic diagram of a yaw angle fusion device provided by an embodiment of the present invention.
  • Figure 13 is a schematic diagram of the magnetometer alignment deviation angle module in Figure 12;
  • Fig. 14 is a schematic diagram of the GPS realignment deviation angle module in Fig. 12;
  • FIG. 15 is a schematic diagram of the yaw angle error angle module in FIG. 12;
  • 16 is a schematic diagram of the hardware structure of an aircraft according to an embodiment of the present invention.
  • FIG. 17 is a connection block diagram of an aircraft provided by an embodiment of the present invention.
  • Fig. 18 is a schematic diagram of the power system in Fig. 17.
  • the yaw angle fusion method provided by the embodiment of the present invention can be applied to various movable objects driven by motors or motors, including but not limited to aircraft, robots, and the like.
  • the aerial vehicle may include unmanned aerial vehicle (UAV), unmanned aerial vehicle, etc.
  • UAV unmanned aerial vehicle
  • UAV includes the fuselage, controller and power system.
  • the controller is connected with a power system, and the power system is installed on the fuselage to provide flight power for the aircraft.
  • the controller is used to execute the above-mentioned yaw angle fusion method to align and correct the yaw angle of the aircraft, and generate a control command according to the yaw angle of the fused aircraft, and send the control command to the power system
  • the ESC controls the driving motor of the power system through the control command.
  • the controller is used to execute the yaw angle fusion method to align and correct the yaw angle of the aircraft, and send the aligned and corrected yaw angle of the aircraft to the ESC, and the ESC is based on the corrected yaw angle of the aircraft.
  • the flight angle generates a control command, and the drive motor of the power system is controlled by the control command.
  • the fuselage includes a central shell and one or more arms connected with the central shell, and the one or more arms extend radially from the central shell.
  • the connection between the arm and the center housing can be an integral connection or a fixed connection.
  • the power system is installed on the arm.
  • the controller is used to execute the above-mentioned yaw angle fusion method to align and correct the yaw angle of the aircraft, and generate a control command according to the aligned and corrected yaw angle of the aircraft, and send the control command to the electrical system of the power system Adjust so that the ESC controls the drive motor of the power system through the control command.
  • the controller is a device with a certain logic processing capability, such as a control chip, a single-chip microcomputer, a microcontroller unit (Microcontroller Unit, MCU), etc.
  • the power system includes: ESC, drive motor and propeller.
  • the ESC is located in the cavity formed by the arm or the center housing.
  • the ESC is connected to the controller and the drive motor respectively.
  • the ESC is electrically connected to the drive motor, and is used to control the drive motor.
  • the driving motor is installed on the arm, and the rotating shaft of the driving motor is connected to the propeller.
  • the propeller generates a force for moving the UAV under the driving of the driving motor, for example, a lift force or thrust force for moving the UAV.
  • UAV completes each specified speed and action (or posture) by controlling the drive motor through ESC.
  • the full name of ESC is electronic governor, which adjusts the speed of UAV's drive motor according to the control signal.
  • the controller is the executive body that executes the above-mentioned yaw angle fusion method, and the ESC controls the driving motor based on the control command generated by the fusion yaw angle of the aircraft.
  • the principle of the ESC to control the drive motor is roughly as follows: the drive motor is an open-loop control element that converts electrical pulse signals into angular displacement or linear displacement. In the case of non-overload, the speed and stop position of the drive motor only depend on the frequency and pulse number of the pulse signal, and are not affected by the load change.
  • the drive When the drive receives a pulse signal, it drives the drive motor of the power system Rotate a fixed angle in the set direction, and its rotation runs at a fixed angle. Therefore, the ESC can control the angular displacement by controlling the number of pulses, so as to achieve the purpose of accurate positioning; at the same time, the speed and acceleration of the driving motor can be controlled by controlling the pulse frequency, so as to achieve the purpose of speed regulation.
  • the main functions of UAV are aerial photography, real-time image transmission, and detection of high-risk areas.
  • a camera component will be connected to the UAV.
  • the UAV and the camera assembly are connected through a connection structure, such as a vibration damping ball.
  • the camera component is used to obtain the shooting picture during the aerial photography of the UAV.
  • the camera component includes: a pan-tilt and a camera.
  • the pan-tilt is connected to the UAV.
  • the photographing device is mounted on the pan/tilt.
  • the photographing device may be an image acquisition device for collecting images.
  • the photographing device includes but is not limited to a camera, a video camera, a camera, a scanner, a camera phone, etc.
  • the pan/tilt is used to mount the camera to realize the fixation of the camera, or to adjust the posture of the camera at will (for example, change the height, inclination and/or direction of the camera) and to keep the camera stably in the set posture on.
  • the pan/tilt is mainly used to stabilize the camera in a set posture, prevent the camera from shaking and ensure the stability of the camera.
  • PTZ includes: PTZ motor and PTZ base.
  • the gimbal motor is installed on the base of the gimbal.
  • the controller of the aircraft can also control the gimbal motor through the ESC of the power system. Specifically, the controller of the aircraft is connected to the ESC, and the ESC is electrically connected to the gimbal motor.
  • the controller of the aircraft generates the gimbal motor control command. Adjust the PTZ motor control command to control the PTZ motor.
  • the pan/tilt base is connected to the UAV body and is used to fix the camera assembly on the UAV body.
  • the gimbal motors are respectively connected with the gimbal base and the camera.
  • the pan/tilt may be a multi-axis pan/tilt. To adapt to this, there are multiple pan/tilt motors, that is, one pan/tilt motor is provided for each axis. On the one hand, the pan/tilt motor can drive the rotation of the shooting device, so as to meet the adjustment of the horizontal rotation and pitch angle of the shooting shaft.
  • the rotation of the pan/tilt motor cancels the disturbance of the camera in real time, prevents the camera from shaking, and ensures the stability of the shooting picture.
  • the camera is mounted on the pan/tilt, and an inertial measurement unit (IMU) is provided on the camera.
  • the inertial measurement unit is a device for measuring the three-axis attitude angle (or angular velocity) and acceleration of an object.
  • a three-axis gyroscope and three-directional accelerometers are installed in an IMU to measure the angular velocity and acceleration of the object in three-dimensional space, and to calculate the posture of the object.
  • the IMU should be installed at the center of gravity of the UAV.
  • the yaw angle of the UAV is an important parameter in controlling the attitude of the UAV, and it is necessary to control the drive motor based on the yaw angle of the UAV.
  • the yaw angle fusion of the aircraft is a very difficult and important issue.
  • the yaw angle of the aircraft is related to the heading of the aircraft and affects the performance of the aerial drone in the following aspects:
  • Inaccurate yaw angle will affect the heading and cause the problem of oblique flight.
  • the user sends the aircraft forward (backward) flight instructions.
  • the aircraft will not fly in the nose (tail) direction, but fly in an oblique direction, in the lateral direction ( (Left/Right) will have a velocity component.
  • Inaccurate yaw angle will affect the aerial photography performance of orbiting.
  • the plane is flying around, if the gimbal is aimed at the center of the circle to shoot, if the yaw angle is not accurately integrated, or there is a long-term drift, the subject will not be in the center of the image/video, but there will be a certain angle left or Right.
  • the UAV's yaw angle is usually obtained based on the data collected by the magnetometer, but the yaw angle obtained by this method is easily affected by external factors, especially when the magnetometer is in strong magnetic interference. In the environment, the data of the magnetometer may be seriously wrong, resulting in a large deviation in the estimation of the yaw angle, and the estimation accuracy of the yaw angle of the aircraft is low.
  • the main purpose of the embodiments of the present invention is to provide a yaw angle fusion method, device, and aircraft, which can align and correct the yaw angle of the aircraft based on GPS data, IMU data, and magnetometer data , The fusion generates the yaw angle, which effectively improves the fusion accuracy and convergence speed of the yaw angle, thereby improving the safety and stability of the aircraft.
  • the embodiment of the present invention uses GPS data, IMU data, and magnetometer data to fuse the corrected yaw angle, magnetometer alignment deviation angle, and GPS realignment deviation angle.
  • the fusion method can avoid interference from external factors and also That is, in the environment of weak GPS signal and strong magnetic interference, the yaw angle can be quickly converged, and the fusion accuracy can be improved through multiple alignments and corrections, thereby effectively improving the safety and stability of the aircraft.
  • FIG. 1 is a schematic block diagram of a yaw angle fusion method according to an embodiment of the present invention
  • the ground altitude, GPS speed signal quality, GPS acceleration signal quality, yaw angle deviation angle, stationary flag, IMU angular velocity, and magnetic field signal quality are determined respectively .
  • Yaw angle error angle input the ground altitude, GPS speed signal quality, GPS acceleration signal quality, yaw angle deviation angle, stationary flag, IMU angular velocity, magnetic field signal quality, and yaw angle error angle into the yaw angle Fusion algorithm, fusion generates a fusion yaw angle.
  • FIG. 2 is a schematic diagram of a yaw angle fusion algorithm provided by an embodiment of the present invention
  • FIG. 3 is an overall schematic diagram of a yaw angle fusion method provided by an embodiment of the present invention
  • FIG. 4 is a schematic flowchart of a yaw angle fusion method provided by an embodiment of the present invention.
  • the yaw angle fusion method can be executed by various electronic devices with certain logic processing capabilities, such as aircraft, control chips, etc.
  • the aircraft can include unmanned aerial vehicles, unmanned ships, etc.
  • the following electronic equipment takes an aircraft as an example for description.
  • the aircraft is connected with a gimbal.
  • the gimbal includes a gimbal motor and a gimbal base.
  • the gimbal can be a multi-axis gimbal, such as a two-axis gimbal and a three-axis gimbal. Take the following three-axis gimbal as an example Description.
  • the specific structure of the aircraft and the gimbal reference can be made to the above description, and therefore, it will not be repeated here.
  • the method is applied to an aircraft, such as a drone, and the method includes:
  • Step S10 Obtain GPS data, IMU data, and magnetometer data.
  • the GPS data includes GPS position, speed, acceleration information, and GPS speed signal quality
  • the IMU data includes IMU acceleration information and IMU angular velocity information
  • the aircraft is provided with an attitude sensor component
  • the attitude sensor component includes: a global positioning system (Global Positioning System, GPS), an inertial measurement unit (Inertial measurement unit, IMU), a magnetometer, etc.
  • GPS Global Positioning System
  • IMU inertial measurement unit
  • IMU inertial measurement unit
  • the magnetometer is used to obtain magnetometer data
  • the inertial measurement unit includes a gyroscope and an accelerometer
  • the gyroscope is used to obtain IMU angular velocity
  • the accelerometer is used to obtain IMU angular velocity information.
  • the GPS data includes position information (latitude and longitude information), GPS speed information, and GPS acceleration information.
  • the IMU data includes IMU acceleration information and IMU angular velocity information.
  • the magnetometer data includes magnetic field strength information.
  • the latitude and longitude information and GPS speed information are acquired through algorithms such as data processing and filtering.
  • the data processing includes: data preprocessing, such as decompression, coordinate system conversion, unit conversion, etc., and the filtering includes low-pass filtering. And Kalman filter, respectively through low-pass filter and Kalman filter.
  • Step S20 Determine the corrected yaw angle according to the IMU data, the GPS data and the magnetometer data;
  • FIG. 7 is a detailed flowchart of step S20 in FIG. 4;
  • the determining the corrected yaw angle according to the IMU angular velocity information includes:
  • Step S21 Obtain the magnetic field signal quality of the current position of the aircraft
  • the magnetic field signal quality of the current position of the aircraft is obtained by comparing the magnetometer data measured by the magnetometer of the aircraft, that is, the magnetic field intensity. If the magnetometer data measured by the magnetometer of the aircraft differs greatly from the standard magnetic field strength of the aircraft’s current position, it means that the quality of the magnetic field signal is poor. If the magnetometer data measured by the aircraft’s magnetometer differs from the aircraft’s current The magnetic field strength of the standard magnetic field at the position is similar, indicating that the magnetic field signal quality is good. Generally, the quality of the magnetic field signal at the current position of the aircraft is judged by the amplitude and the inclination.
  • Step S22 Generate a GPS angular velocity correction amount according to the yaw angle deviation angle
  • the aircraft is provided with a feedback controller
  • the GPS angular velocity correction amount is calculated by the feedback controller based on a feedback control algorithm on the yaw angle deviation angle
  • the GPS angular velocity correction amount is compared with the yaw angle deviation angle.
  • Step S23 Integrating the GPS acceleration correction weight coefficient and the GPS angular velocity correction amount to generate a GPS weighted correction amount
  • the product is used as the GPS weight correction amount.
  • the GPS acceleration correction weight coefficient is obtained by the product of the GPS speed signal quality and the GPS acceleration signal quality
  • the GPS acceleration signal quality is calculated by the GPS acceleration information
  • the GPS acceleration signal quality is determined by the GPS acceleration after signal quality detection and judgment.
  • the GPS module will give the signal quality, which is already available at the factory.
  • a threshold greater than the threshold indicates the GPS speed signal If the quality is good, if it is less than the threshold, it means that the GPS speed signal quality is poor, and the threshold is set manually.
  • the differential noise can be obtained.
  • the differential quality is judged according to the noise. If the noise is large, it is not good, and the noise is small.
  • the GPS speed signal quality is good and the differential noise is small, the GPS acceleration signal quality is good, otherwise the GPS acceleration signal quality is poor.
  • Step S24 Generate a magnetometer correction amount according to the yaw angle error angle
  • the aircraft is provided with a feedback controller, and the magnetometer correction is calculated by the feedback controller based on a feedback control algorithm for the yaw angle error angle, and the feedback control algorithm is: the magnetic force
  • Step S25 Integrating the magnetic field signal quality and the magnetometer correction amount to generate a magnetometer weighted correction amount
  • the product of the magnetic field signal quality and the magnetometer correction amount is obtained, and the product is used as the magnetometer weighted correction amount.
  • Step S26 fusing the IMU angular velocity information, the GPS weighted correction amount, and the magnetometer weighted correction amount to generate a corrected yaw angular velocity;
  • the IMU angular velocity information is acquired by the angular velocity meter of the inertial measurement unit.
  • the IMU angular velocity information, the GPS weighted correction amount, and the magnetometer weighted correction amount are fused to obtain the IMU angular velocity information,
  • the sum of the GPS weighted correction amount and the magnetometer weighted correction amount is taken as the corrected yaw angular velocity.
  • the corrected yaw angular velocity is the yaw angular velocity after GPS correction and magnetometer correction.
  • Step S27 Integrate the corrected yaw angular velocity to generate a corrected yaw angle.
  • the integral value is taken as the corrected yaw angle.
  • Step S30 Determine the magnetometer alignment deviation angle according to the magnetometer data, the GPS data and the corrected yaw angle
  • FIG. 8 is a detailed flowchart of step S30 in FIG. 4;
  • the determining the alignment deviation angle of the magnetometer according to the magnetometer data includes:
  • Step S31 Determine the alignment of the magnetometer to the yaw angle according to the yaw angle of the magnetometer
  • the aircraft is provided with an enabling module, and alignment is achieved by using the enabling module.
  • the enabling module includes an input terminal and an output terminal. Once the enabling module detects a pulse, the enabling module will The value of the input terminal is sent to the output terminal so that the output terminal outputs a high level, otherwise, the output terminal inputs a low level or is set to zero.
  • the magnetometer alignment yaw angle is the magnitude of the yaw angle measured when the magnetometer is aligned, wherein the pulse is a magnetometer alignment pulse.
  • the enabling module can replace other arithmetic modules, such as a multiplication module, etc., which are all within the protection scope of the embodiments of the present invention.
  • FIG. 9 is a detailed flowchart of step S31 in FIG. 8;
  • the determining the magnetometer alignment yaw angle according to the magnetometer yaw angle includes:
  • Step S311 Obtain the ground altitude and static state position of the aircraft
  • the aircraft is provided with an ultrasonic sensor and/or a TOF sensor (Time of Flight, TOF), the ground height of the aircraft is acquired by the ultrasonic sensor and/or the TOF sensor, and the static state bit represents the The stationary state of the aircraft.
  • TOF Time of Flight
  • the acquisition of the stationary state bit of the aircraft includes:
  • the stationary state of the aircraft is stationary, the stationary state is The value of is 1. If the stationary state of the aircraft is in motion, the value of the stationary state bit is 0. It can be understood that the IMU acceleration and IMU angular velocity data fluctuate very little when stationary, and it is very stable. When the data of the IMU acceleration and IMU angular velocity fluctuate greatly, the stationary state of the aircraft is determined according to this principle.
  • the method specifically includes: transforming the body coordinate system and the ground coordinate system on the IMU data according to the rotation transformation matrix to generate the IMU acceleration in the ground coordinate system and the IMU angular velocity in the ground coordinate system;
  • the IMU acceleration in the coordinate system and the IMU angular velocity in the ground coordinate system determine the stationary state of the aircraft, and generate the stationary flag of the aircraft.
  • Step S312 Perform a logical judgment on the quality of the magnetic field signal and the static flag, and generate a magnetometer initial alignment pulse
  • the judgment logic includes:
  • Step1.i 0, turn to Step2;
  • Step3 Judge whether the pulse signal F changes from 0 to 1 for the first time, if yes, output a pulse to the magnetometer initial alignment pulse, if not output 0, go to Step 4.
  • Step4.i i+1, return to Step2.
  • the magnetic field signal quality threshold can be specifically set according to specific requirements, and both fall within the protection scope of the present invention.
  • Step S313 Perform a logical judgment on the ground height and the quality of the magnetic field signal, and generate a magnetometer realignment pulse;
  • a logical judgment is made on the ground height and the quality of the magnetic field signal, and the judgment logic includes:
  • Step1.i 0, turn to Step2;
  • Step3 Judge whether the pulse signal F changes from 0 to 1 for the first time, if yes, output a pulse to the magnetometer to realign the pulse, if not output 0, go to Step 4.
  • Step4.i i+1, return to Step2.
  • Step S314 Generate a magnetometer alignment pulse according to the magnetometer initial alignment pulse and the magnetometer realignment pulse;
  • the generating the magnetometer alignment pulse according to the magnetometer initial alignment pulse and the magnetometer realignment pulse includes: realigning the magnetometer initial alignment pulse and the magnetometer The pulse is ORed to generate the magnetometer alignment pulse.
  • the magnetic field signal quality threshold and the ground height threshold can be specifically set according to specific requirements, and both fall within the protection scope of the present invention.
  • Step S315 Input the magnetometer alignment pulse to the enabling module of the aircraft, and when the enabling module detects that the magnetometer alignment pulse has a rising edge, it uses the magnetometer yaw angle as the The magnetometer is aligned with the yaw angle.
  • the input terminal of the enabling module even if the enabling terminal is used to receive the magnetometer alignment pulse, input the magnetometer alignment pulse to the enabling module of the aircraft, when the enabling module detects all When the magnetometer alignment pulse signal has a rising edge, the enabling module uses the magnetometer yaw angle as the magnetometer alignment yaw angle.
  • Step S32 Determine the alignment deviation angle of the magnetometer according to the corrected yaw angle and the alignment yaw angle of the magnetometer.
  • the magnetometer alignment deviation angle is fused to generate the magnetometer alignment deviation angle.
  • the Magnetometer alignment deviation angle -corrected yaw angle + magnetometer alignment yaw angle.
  • Step S40 Determine the GPS realignment deviation angle according to the GPS data and the IMU acceleration information
  • the GPS data includes GPS acceleration information
  • the GPS acceleration information is obtained through acceleration estimation of the GPS speed information and position information
  • the IMU acceleration information is obtained through an accelerometer of the inertial measurement unit.
  • FIG. 10 is a detailed flowchart of step S40 in FIG. 4;
  • the determining the GPS realignment deviation angle according to the GPS acceleration information and the IMU acceleration information includes:
  • Step S41 Calculate the yaw angle deviation angle according to the GPS acceleration information and the IMU acceleration information;
  • the IMU acceleration information is the acceleration information in the ground coordinate system obtained after the measurement data of the inertial measurement unit is calibrated by the calibration matrix and coordinate transformation from the body coordinate system to the ground coordinate system. It is understandable that the calibration matrix is calibrated by the user at the place where the user wants to fly. The calibration matrix is different anywhere on the earth. The aircraft reports magnetometer interference and requires the user to calibrate before determining the calibration matrix.
  • Step S42 Determine the GPS realignment deviation angle according to the yaw angle deviation angle.
  • the aircraft is provided with an enabling module.
  • the enabling module includes an input terminal and an output terminal.
  • the input terminal of the enabling module receives a realignment pulse
  • the enabling module The energy module outputs the yaw angle deviation angle as the realignment yaw angle deviation compensation amount, that is, the GPS realignment deviation angle.
  • FIG. 11 is a detailed flowchart of step S42 in FIG. 10;
  • determining the GPS realignment deviation angle according to the yaw angle deviation angle includes:
  • Step S421 Calculate GPS speed signal quality and GPS acceleration signal quality respectively according to the GPS speed information and the GPS acceleration information;
  • the GPS speed signal quality reference value is generated according to the signal quality detection and judgment principle, and the GPS speed signal quality reference value is used as the GPS Speed signal quality, and differentiate the GPS speed information acquired by GPS, and perform filtering processing to generate the GPS acceleration information, and generate the GPS acceleration signal quality reference value according to the signal quality detection and judgment principle, as the GPS acceleration signal quality.
  • the GPS will give the quality of the GPS speed signal.
  • the magnitude of the differential noise can be obtained, and the GPS acceleration can be judged according to the magnitude of the noise. The signal quality, if the noise is large, the GPS acceleration signal quality is poor, and if the noise is small, the GPS acceleration signal quality is good.
  • the GPS speed signal quality and GPS acceleration signal quality are used for auxiliary reference, reflecting the quality of the signal acquired by GPS, and they show a positive correlation. If the quality of the signal acquired by the GPS is good, the GPS speed signal The values of quality and GPS acceleration signal quality are large. Conversely, if the signal quality acquired by the GPS is poor, the values of the GPS speed signal quality and GPS acceleration signal quality are small.
  • Step S422 Integrate the GPS speed signal quality and GPS acceleration signal quality to generate GPS acceleration correction weight coefficient
  • the GPS speed signal quality and the GPS acceleration signal quality are integrated, and the value obtained by the integration is used as the GPS acceleration correction weight coefficient.
  • the GPS speed signal quality and GPS acceleration signal quality are confidence quantification information. If the GPS acceleration correction weight coefficient is larger, the confidence is higher, and vice versa.
  • Step S423 Determine whether the GPS acceleration correction weight coefficient is greater than a preset GPS realignment threshold
  • the aircraft is provided with a realignment flag, and by setting a GPS realignment threshold in advance, the GPS acceleration correction weight coefficient is compared with the GPS realignment threshold. If the GPS acceleration correction weight coefficient is Is greater than the GPS realignment threshold, proceed to step S224: control the realignment flag of the aircraft to output a high level signal, that is, output 1; if the GPS acceleration correction weight coefficient is not greater than the GPS realignment Threshold, control the realignment flag of the aircraft to output a low-level signal, that is, output 0.
  • the realignment threshold in the embodiment of the present invention can be artificially set according to specific requirements.
  • Step S424 Control the realignment flag of the aircraft to output a high level signal
  • Step S425 controlling the realignment flag to output a low-level signal
  • Step S426 If the signal of the realignment flag has a rising edge, generate a realignment pulse signal, and input the realignment pulse signal to the enabling module of the aircraft;
  • the aircraft is provided with an enabling module, which detects whether the realignment flag has a rising edge, if so, generates a realignment pulse signal, and inputs the realignment pulse signal to the enable module.
  • the input terminal of the energy module is provided.
  • Step S427 When the enabling module detects that the realignment pulse has a rising edge, it uses the yaw angle deviation angle as the GPS realignment deviation angle.
  • the enabling module uses the yaw angle deviation angle output as the realignment yaw angle deviation compensation amount, that is, the GPS Realign the deviation angle.
  • the method further includes: calculating the yaw angle error angle, specifically, please refer to FIG. 5, which is a schematic flowchart of calculating the yaw angle error angle provided by an embodiment of the present invention
  • the calculation of the yaw angle error angle includes:
  • Step S401 Calculate the yaw angle of the magnetometer according to the magnetometer data
  • the magnetometer data includes three-axis magnetic field strength, and GPS latitude and longitude information is used to find the standard magnetic field strength, magnetic inclination information, and magnetic declination information of the current position of the aircraft.
  • the magnetometer data transforms the magnetic field intensity measured by the magnetometer through a calibration matrix calibrated in advance to obtain the calibrated magnetometer data, and according to the calibrated magnetometer data, the current of the aircraft is generated The reference value of the magnetic field signal quality of the location.
  • FIG. 6 is a detailed flowchart of step S401 in FIG. 5;
  • the calculation of the magnetometer yaw angle according to the magnetometer data includes:
  • Step S4011 Calibrate the magnetometer data according to a preset calibration matrix to generate calibrated magnetometer data
  • the preset calibration matrix is calibrated by the user at the place to be flown.
  • the calibration matrix is different anywhere on the earth.
  • the aircraft reports magnetometer interference and requires the user to calibrate before determining the calibration matrix.
  • Step S4012 Generate a rotation transformation matrix according to the attitude angle of the aircraft
  • the rotation transformation matrix is used to convert the airframe coordinate system into a ground coordinate system
  • the attitude angle of the aircraft includes: a yaw angle, a pitch angle, and a roll angle
  • the rotation transformation matrix is 3*3 Matrix, which contains the sine and cosine functions of the yaw angle, pitch angle, and roll angle, and select different functions according to the specific situation.
  • the rotation transformation matrix is:
  • ( ⁇ , ⁇ , ⁇ ) is the attitude angle
  • ⁇ is the roll angle in the attitude angle
  • ⁇ is the pitch angle in the attitude angle
  • ⁇ is the yaw angle in the attitude angle
  • Step S4013 Perform the conversion between the body coordinate system and the ground coordinate system according to the rotation transformation matrix, and perform coordinate transformation on the calibrated magnetometer data to generate magnetometer data in the ground coordinate system;
  • the calibrated magnetometer data that is, the magnetic field strength
  • the rotation transformation matrix to generate the magnetic field strength in the ground coordinate system, which is equivalent to performing the body coordinate system on the magnetometer data And the transformation between the ground coordinate system.
  • Step S4014 According to the magnetometer data in the ground coordinate system, compare the magnetometer data of the standard magnetic field at the current position of the aircraft to calculate the magnetometer yaw angle.
  • the current position of the aircraft corresponds to a standard magnetic field
  • the three-axis readings of the magnetometer form a vector
  • the magnetometer data of the standard magnetic field at the current position corresponds to a vector
  • the magnetic force in the ground coordinate system Comparing the gauge data with the magnetometer data of the standard magnetic field at the current position of the aircraft, calculate the vector angle between the two, and use the vector angle as the magnetometer yaw angle.
  • Step S402 Calculate the yaw angle error angle according to the magnetometer yaw angle and the fused yaw angle.
  • Step S50 Generate a fusion yaw angle according to the GPS realignment deviation angle, the corrected yaw angle, and the magnetometer alignment deviation angle.
  • the correction of the yaw angle is embodied in the correction of the yaw angular velocity, including GPS correction and magnetometer correction.
  • the alignment of the yaw angle is embodied in the correction of the yaw angle deviation, and the initial alignment adopts magnetic force.
  • Gauge alignment, realignment includes GPS alignment and magnetometer alignment.
  • the yaw angle error when the yaw angle error is large and the quality of the magnetometer or GPS information is good, it can realize one-step rapid compensation, and the small yaw angle error can ensure convergence through the feedback controller; it solves the harsh magnetic field environment A series of problems caused by the wrong yaw angle, such as the problem of yaw angle drift when flying around for a long time, the yaw angle disorder when the magnetic field is disturbed, and the aircraft oblique flying problem.
  • the method includes: acquiring GPS data, IMU data, and magnetometer data.
  • the GPS data includes: GPS position, speed, acceleration Information and GPS speed signal quality
  • the IMU data includes: IMU acceleration information and IMU angular velocity information; according to the IMU data, the GPS data, and the magnetometer data, determine the corrected yaw angle; according to the magnetometer Data, the GPS data and the corrected yaw angle, determine the magnetometer alignment deviation angle; according to the GPS data and the IMU acceleration information, determine the GPS realignment deviation angle; according to the corrected yaw angle Angle, magnetometer alignment deviation angle and GPS realignment deviation angle to generate a fusion yaw angle.
  • the embodiment of the present invention can effectively improve the fusion accuracy and convergence speed of the yaw angle.
  • FIG. 12 is a schematic diagram of a yaw angle fusion device provided by an embodiment of the present invention.
  • the yaw angle fusion device 120 may be applied to an aircraft, such as an unmanned aerial vehicle.
  • the yaw angle fusion device 120 includes:
  • the acquisition module 121 is configured to acquire GPS data, IMU data, and magnetometer data.
  • the GPS data includes GPS position, speed, acceleration information, and GPS speed signal quality
  • the IMU data includes IMU acceleration information and IMU angular velocity information;
  • the corrected yaw angle module 122 is configured to determine the corrected yaw angle according to the IMU data, the GPS data, and the magnetometer data;
  • the magnetometer alignment deviation angle module 123 is configured to determine the magnetometer alignment deviation angle according to the magnetometer data, the GPS data, and the corrected yaw angle;
  • the GPS realignment deviation angle module 124 is configured to determine the GPS realignment deviation angle according to the GPS data and the IMU acceleration information;
  • the yaw angle fusion module 125 is used to generate a fusion yaw angle according to the GPS realignment deviation angle, the corrected yaw angle, and the magnetometer alignment deviation angle;
  • the yaw angle error angle module 126 is used to calculate the yaw angle error angle
  • FIG. 13 is a schematic diagram of the magnetometer alignment deviation angle module in FIG. 12;
  • the magnetometer alignment deviation angle module 123 includes:
  • the magnetometer alignment yaw angle unit 1231 is used to determine the magnetometer alignment yaw angle according to the magnetometer yaw angle;
  • the magnetometer alignment deviation angle unit 1232 is configured to determine the magnetometer alignment deviation angle according to the corrected yaw angle and the magnetometer alignment yaw angle.
  • the magnetometer is aligned with the yaw angle unit 1231, specifically for:
  • the magnetometer alignment pulse signal has a rising edge
  • the magnetometer yaw angle is used as the magnetometer alignment yaw angle.
  • FIG. 14 is a schematic diagram of the GPS realignment deviation angle module in FIG. 12;
  • the GPS realignment deviation angle module 124 includes:
  • the yaw angle deviation angle unit 1241 is configured to calculate the yaw angle deviation angle according to the GPS acceleration information and the IMU acceleration information;
  • the GPS realignment deviation angle unit 1242 is used to determine the GPS realignment deviation angle according to the yaw angle deviation angle.
  • the GPS realignment deviation angle unit is specifically used for:
  • the signal of the realignment flag has a rising edge, generate a realignment pulse signal, and input the realignment pulse signal to the enabling module of the aircraft;
  • the enabling module When the enabling module receives the realignment pulse signal, it uses the yaw angle deviation angle as the GPS realignment deviation angle.
  • FIG. 15 is a schematic diagram of the yaw angle error angle module in FIG. 12;
  • the yaw angle error angle module 126 includes:
  • the magnetometer yaw angle unit 1261 is used to calculate the magnetometer yaw angle according to the magnetometer data
  • the yaw angle error angle unit 1262 is configured to calculate the yaw angle error angle according to the magnetometer yaw angle and the fused yaw angle.
  • the magnetometer yaw angle unit 1261 is specifically used for:
  • the magnetometer data of the standard magnetic field at the current position of the aircraft is compared to calculate the magnetometer yaw angle.
  • the content of the device embodiment can be quoted from the method embodiment on the premise that the content does not conflict with each other, which will not be repeated here.
  • the device includes: an acquisition module for acquiring GPS data, IMU data, and magnetometer data, the GPS data includes: GPS Position, velocity, acceleration information, and GPS speed signal quality, the IMU data includes: IMU acceleration information and IMU angular velocity information; a corrected yaw angle module, used for according to the IMU data, the GPS data and the magnetometer Data, determine the corrected yaw angle; magnetometer alignment deviation angle module, used to determine the magnetometer alignment deviation angle according to the magnetometer data, the GPS data, and the corrected yaw angle; GPS realignment The quasi deviation angle module is used to determine the GPS realignment deviation angle based on the GPS data and the IMU acceleration information; the yaw angle fusion module is used to determine the deviation angle of the GPS realignment based on the corrected yaw angle and the magnetometer alignment deviation angle And GPS realigns the deviation angle to generate a fusion yaw angle.
  • FIG. 16 is a schematic diagram of the hardware structure of an aircraft according to an embodiment of the present invention.
  • the aircraft may be an unmanned aerial vehicle (UAV), an unmanned aerial vehicle or other electronic equipment.
  • UAV unmanned aerial vehicle
  • UAV unmanned aerial vehicle
  • the aircraft 160 includes one or more processors 161 and a memory 162. Among them, one processor 161 is taken as an example in FIG. 16.
  • the processor 161 and the memory 162 may be connected through a bus or in other ways. In FIG. 16, the connection through a bus is taken as an example.
  • the memory 162 can be used to store non-volatile software programs, non-volatile computer-executable programs and modules, such as a yaw angle fusion in the embodiment of the present invention
  • the unit corresponding to the method (for example, each module or unit described in Figure 12 to Figure 15).
  • the processor 161 executes various functional applications and data processing of the yaw angle fusion method by running the non-volatile software programs, instructions, and modules stored in the memory 162, that is, realizes the yaw angle fusion in the above method embodiment
  • the memory 162 may include a high-speed random access memory, and may also include a non-volatile memory, such as at least one magnetic disk storage device, a flash memory device, or other non-volatile solid-state storage devices.
  • the memory 162 may optionally include memories remotely provided with respect to the processor 161, and these remote memories may be connected to the processor 161 via a network. Examples of the aforementioned networks include but are not limited to the Internet, corporate intranets, local area networks, mobile communication networks, and combinations thereof.
  • the module is stored in the memory 162, and when executed by the one or more processors 161, the yaw angle fusion method in any of the foregoing method embodiments is executed, for example, the above-described FIGS. 4 to 4 are executed.
  • Each step shown in 11; can also realize the function of each module or unit described in Figure 12 to Figure 15.
  • the aircraft 160 further includes a power system 163.
  • the power system 163 is used for the aircraft to provide flight power, and the power system 163 is connected to the processor 161.
  • the power system 163 includes a driving motor 1631 and an ESC 1632, and the ESC 1632 is electrically connected to the driving motor 1631 for controlling the driving motor 1631.
  • the ESC 1632 generates a control command based on the fused yaw angle obtained after the processor 161 executes the yaw angle fusion method described above, and controls the driving motor 1631 through the control command.
  • the aircraft 160 can execute the yaw angle fusion method provided in the first embodiment of the present invention, and has functional modules and beneficial effects corresponding to the execution method.
  • the yaw angle fusion method provided in the first embodiment of the present invention.
  • the embodiment of the present invention provides a computer program product, the computer program product includes a computer program stored on a non-volatile computer-readable storage medium, the computer program includes program instructions, when the program instructions are executed by a computer At this time, the computer is caused to execute the yaw angle fusion method described above. For example, steps S10 to S50 of the method in FIG. 4 described above are executed.
  • An embodiment of the present invention also provides a non-volatile computer storage medium, the computer storage medium stores computer-executable instructions, the computer-executable instructions are executed by one or more processors, for example, a process in FIG. 16
  • the device 161 can make the above-mentioned one or more processors execute the yaw angle fusion method in any of the above method embodiments, for example, execute the yaw angle fusion method in any of the above method embodiments, for example, execute the above description
  • the steps shown in Fig. 4 to Fig. 11; the function of each module or unit described in Fig. 12 to Fig. 15 can also be realized.
  • the device or device embodiments described above are merely illustrative, wherein the unit modules described as separate components may or may not be physically separated, and the components displayed as modular units may or may not be physical units , Which can be located in one place, or can be distributed to multiple network module units. Some or all of the modules may be selected according to actual needs to achieve the objectives of the solutions of the embodiments.
  • each implementation manner can be implemented by means of software plus a general hardware platform, and of course, it can also be implemented by hardware.
  • the above technical solution essentially or the part that contributes to the related technology can be embodied in the form of a software product, and the computer software product can be stored in a computer-readable storage medium, such as ROM/RAM, magnetic disk , CD-ROM, etc., including several instructions until a computer device (which can be a personal computer, server, or network device, etc.) executes the methods described in each embodiment or some parts of the embodiment.

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Abstract

一种偏航角的融合方法、装置及飞行器（160）。偏航角的融合方法，应用于飞行器（160），能够有效提高偏航角的融合精度以及收敛速度。方法包括：获取GPS数据、IMU数据以及磁力计数据，GPS数据包括：GPS位置、速度、加速度信息以及GPS速度信号质量，IMU数据包括：IMU加速度信息以及IMU角速度信息（S10）；根据IMU数据、GPS数据以及磁力计数据，确定修正后偏航角（S20）；根据磁力计数据、GPS数据以及修正后偏航角，确定磁力计对准偏差角（S30）；根据GPS数据以及IMU加速度信息，确定GPS重对准偏差角（S40）；根据修正后偏航角、磁力计对准偏差角以及GPS重对准偏差角，生成融合偏航角（S50）。

Description

一种偏航角的融合方法、装置及飞行器
本申请要求于2019年8月21日提交中国专利局、申请号为201910774173.1、申请名称为“一种偏航角的融合方法、装置及飞行器”的中国专利申请的优先权，其全部内容通过引用结合在本申请中。
技术领域
本申请涉及飞行器技术领域，特别是涉及一种偏航角修正及对准方法、装置及飞行器。
背景技术
飞行器，如无人飞行器(Unmanned Aerial Vehicle，UAV)，也称无人机，以其具有体积小、重量轻、机动灵活、反应快速、无人驾驶、操作要求低等优点，得到了越来越广泛的应用。无人飞行器的各个动作(或姿态)通常是通过控制无人飞行器的动力系统中的多个驱动电机不同转速实现的。其中，偏航角是对无人飞行器的飞行姿态进行控制中的重要参数，也即无人飞行器的偏航角融合对无人飞行器的姿态控制尤其重要，若无人飞行器的偏航角融合误差大，或者融合精度低，轻则无人飞行器无法按照预设的方向或轨迹飞行，重则出现刷锅线性，甚至可能失稳以致炸机。因此，如何提高飞行器的偏航角的融合精度以及收敛速度具有十分重要的意义。
目前，通常是基于磁力计采集的数据得到飞行器(如UAV等)的偏航角，并基于磁力计采集的数据对偏航角进行修正，但采用该方法得到偏航角很容易受到外界因素的影响，尤其是当磁力计处于强磁干扰环境中时，磁力计的数据可能严重错误，导致偏航角的修正出现较大的偏差，并且偏航角的收敛慢。
为了提高UAV的偏航角估算的准确性，目前常见的是利用外置的GPS模组，依靠GPS估算出一个偏航角，用以对基于磁力计的偏航角值进行修正。但GPS信号有时可能不稳定，使得在某些情况下，磁力计估算的偏航角即使出现了偏差却得不到有效修正。也即利用外置的GPS模组修正UAV的偏航角虽然一定程度上可以提高飞行器的偏航角的估算准确度，但效果并不好，特别是当GPS信号较弱，利用外置的GPS模组未能起到有效的航向修正作用。由于缺少有效的对准方案，偏航角收敛速度较慢，且精度不高。
发明内容
本发明实施例提供一种偏航角的融合方法、装置及飞行器，有效提高了偏航角的融合精度以及收敛速度。
为解决上述技术问题，本发明实施例提供以下技术方案：
第一方面，本发明实施例提供一种偏航角的融合方法，应用于飞行器，所述方法包括：
获取GPS数据、IMU数据以及磁力计数据，所述GPS数据包括：GPS位置、速度、加速度信息以及GPS速度信号质量，所述IMU数据包括：IMU加速度信息以及IMU角速度信息；
根据所述IMU数据、所述GPS数据以及所述磁力计数据，确定修正后偏航角；
根据所述磁力计数据、所述GPS数据以及所述修正后偏航角，确定磁力计对准偏差角；
根据所述GPS数据以及所述IMU加速度信息，确定GPS重对准偏差角；
根据所述修正后偏航角、磁力计对准偏差角以及GPS重对准偏差角，生成融合偏航角。
在一些实施例中，所述方法还包括：
根据所述磁力计数据，计算磁力计偏航角；
根据所述磁力计偏航角和所述融合偏航角，计算偏航角误差角。
在一些实施例中，所述根据所述磁力计数据，计算磁力计偏航角，具体包括：
根据预设的校准矩阵对所述磁力计数据进行校准，生成校准后的磁力计数据；
根据所述飞行器的姿态角，生成旋转变换矩阵；
根据所述旋转变换矩阵进行机体坐标系和地面坐标系的转换，对所述校准后的磁力计数据进行坐标变换，生成所述地面坐标系下的磁力计数据；
根据所述地面坐标系下的磁力计数据，对比所述飞行器的当前位置的标准磁场的磁力计数据，计算所述磁力计偏航角。
在一些实施例中，所述根据所述IMU数据、所述GPS数据以及所述磁力计数据，确定修正后偏航角，具体包括：
获取所述飞行器当前位置的磁场信号质量；
根据所述偏航角偏差角，生成GPS角速度修正量；
对GPS加速度修正权重系数以及所述GPS角速度修正量进行求积，生成GPS加权修正量；
根据所述偏航角误差角，生成磁力计修正量；
对所述磁场信号质量以及所述磁力计修正量进行求积，生成磁力计加权修正量；
对所述IMU角速度信息、所述GPS加权修正量以及所述磁力计加权修正量进行融合，生成修正后偏航角角速度；
对所述修正后偏航角角速度进行积分，生成修正后偏航角，
其中，所述GPS加速度修正权重系数由所述GPS速度信号质量和GPS加速度信号质量求积所得，所述GPS加速度信号质量由所述GPS加速度信息计算所得。
在一些实施例中，所述根据所述磁力计数据，确定磁力计对准偏差角，包 括：
根据所述磁力计偏航角，确定磁力计对准偏航角；
根据所述修正后偏航角以及所述磁力计对准偏航角，确定所述磁力计对准偏差角。
在一些实施例中，所述根据所述磁力计偏航角，确定磁力计对准偏航角，具体包括：
获取所述飞行器的对地高度及静止状态位；
对所述磁场信号质量以及所述静止标志位进行逻辑判断，生成磁力计初对准脉冲；
对所述对地高度以及所述磁场信号质量进行逻辑判断，生成磁力计重对准脉冲；
根据所述磁力计初对准脉冲以及所述磁力计重对准脉冲，生成磁力计对准脉冲；
将所述磁力计对准脉冲输入飞行器的使能模块，当所述使能模块检测到所述磁力计对准脉冲存在上升沿时，则将所述磁力计偏航角作为所述磁力计对准偏航角。
在一些实施例中，所述根据所述GPS数据以及所述IMU加速度信息，确定GPS重对准偏差角，包括：
根据所述GPS加速度信息以及所述IMU加速度信息，计算偏航角偏差角；
根据所述GPS加速度信息，计算GPS加速度信号质量；
对所述GPS速度信号质量以及GPS加速度信号质量进行求积，生成GPS加速度修正权重系数；
判断所述GPS加速度修正权重系数是否大于预设GPS重对准阈值，若是，则控制所述飞行器的重对准标志位输出高电平信号，若否，则控制所述重对准标志位输出低电平信号；
若所述重对准标志位的信号存在上升沿，则生成重对准脉冲信号；以及，
将所述重对准脉冲信号输入所述飞行器的使能模块，当所述使能模块检测到所述重对准脉冲存在上升沿时，则将所述偏航角偏差角作为所述GPS重对准偏差角。
第二方面，本发明实施例提供一种偏航角的融合装置，应用于飞行器，所述装置包括：
获取模块，用于获取GPS数据、IMU数据以及磁力计数据，所述GPS数据包括：GPS位置、速度、加速度信息以及GPS速度信号质量，所述IMU数据包括：IMU加速度信息以及IMU角速度信息；
修正后偏航角模块，用于根据所述IMU数据、所述GPS数据以及所述磁力计数据，确定修正后偏航角；
磁力计对准偏差角模块，用于根据所述磁力计数据、所述GPS数据以及所述修正后偏航角，确定磁力计对准偏差角；
GPS重对准偏差角模块，用于根据所述GPS数据以及所述IMU加速度信息，确定GPS重对准偏差角；
偏航角融合模块，用于根据所述修正后偏航角、磁力计对准偏差角以及GPS重对准偏差角，生成融合偏航角。
在一些实施例中，所述装置还包括：偏航角误差角模块，所述偏航角误差角模块包括：
磁力计偏航角单元，用于根据所述磁力计数据，计算磁力计偏航角；
偏航角误差角单元，用于根据所述磁力计偏航角和所述融合偏航角，计算偏航角误差角。
在一些实施例中，所述磁力计偏航角单元，具体用于：
根据预设的校准矩阵对所述磁力计数据进行校准，生成校准后的磁力计数据；
根据所述飞行器的姿态角，生成旋转变换矩阵；
根据所述旋转变换矩阵进行机体坐标系和地面坐标系的转换，对所述校准后的磁力计数据进行坐标变换，生成所述地面坐标系下的磁力计数据；
根据所述地面坐标系下的磁力计数据，对比所述飞行器的当前位置的标准磁场的磁力计数据，计算所述磁力计偏航角。
在一些实施例中，所述修正后偏航角模块，具体用于：
获取所述飞行器当前位置的磁场信号质量；
根据所述偏航角偏差角，生成GPS角速度修正量；
对GPS加速度修正权重系数以及所述GPS角速度修正量进行求积，生成GPS加权修正量；
根据所述偏航角误差角，生成磁力计修正量；
对所述磁场信号质量以及所述磁力计修正量进行求积，生成磁力计加权修正量；
对所述IMU角速度信息、所述GPS加权修正量以及所述磁力计加权修正量进行融合，生成修正后偏航角角速度；
对所述修正后偏航角角速度进行积分，生成修正后偏航角，
其中，所述GPS加速度修正权重系数由所述GPS速度信号质量和GPS加速度信号质量求积所得，所述GPS加速度信号质量由所述GPS加速度信息计算所得。
在一些实施例中，所述磁力计对准偏差角模块，包括：
磁力计对准偏航角单元，用于根据所述磁力计偏航角，确定磁力计对准偏航角；
磁力计对准偏差角单元，用于根据所述修正后偏航角以及所述磁力计对准偏航角，确定所述磁力计对准偏差角。
在一些实施例中，所述磁力计对准偏航角单元，具体用于：
获取所述飞行器的对地高度及静止状态位；
对所述磁场信号质量以及所述静止标志位进行逻辑判断，生成磁力计初对准脉冲；
对所述对地高度以及所述磁场信号质量进行逻辑判断，生成磁力计重对准脉冲；
根据所述磁力计初对准脉冲以及所述磁力计重对准脉冲，生成磁力计对准脉冲；
将所述磁力计对准脉冲输入飞行器的使能模块，当所述使能模块检测到所述磁力计对准脉冲存在上升沿时，则将所述磁力计偏航角作为所述磁力计对准偏航角。
在一些实施例中，所述GPS重对准偏差角模块，包括：
偏航角偏差角单元，用于根据所述GPS加速度信息以及所述IMU加速度信息，计算偏航角偏差角；
GPS重对准偏差角单元，用于：
根据所述GPS速度信息以及所述GPS加速度信息，分别计算GPS速度信号质量以及GPS加速度信号质量；
对所述GPS速度信号质量以及GPS加速度信号质量进行求积，生成GPS加速度修正权重系数；
判断所述GPS加速度修正权重系数是否大于预设GPS重对准阈值，若是，则控制所述飞行器的重对准标志位输出高电平信号，若否，则控制所述重对准标志位输出低电平信号；
若所述重对准标志位的信号存在上升沿，则生成重对准脉冲信号，；以及
将所述重对准脉冲信号输入所述飞行器的使能模块，当所述使能模块检测接收到所述重对准脉冲存在上升沿信号时，则将所述偏航角偏差角作为所述GPS重对准偏差角。
第三方面，本发明实施例提供一种飞行器，包括：
至少一个处理器；以及，
与所述至少一个处理器通信连接的存储器；其中，
所述存储器存储有可被所述至少一个处理器执行的指令，所述指令被所述至少一个处理器执行，以使所述至少一个处理器能够执行如上所述的偏航角的融合方法。
第四方面，本发明实施例还提供了一种非易失性计算机可读存储介质，所述计算机可读存储介质存储有计算机可执行指令，所述计算机可执行指令用于使飞行器能够执行如上所述的偏航角的融合方法。
本发明实施例的有益效果是：区别于现有技术的情况下，本发明实施例提供的一种偏航角的融合方法，应用于飞行器，所述方法包括：获取GPS数据、IMU数据以及磁力计数据，所述GPS数据包括：GPS位置、速度、加速度信息以及GPS速度信号质量，所述IMU数据包括：IMU加速度信息以及IMU角速度信息；根据所述IMU数据、所述GPS数据以及所述磁力计数据，确定修正后偏 航角；根据所述磁力计数据、所述GPS数据以及所述修正后偏航角，确定磁力计对准偏差角；根据所述GPS数据以及所述IMU加速度信息，确定GPS重对准偏差角；根据所述修正后偏航角、磁力计对准偏差角以及GPS重对准偏差角，生成融合偏航角。通过上述方式，本发明实施例能够有效提高偏航角的融合精度以及收敛速度。
附图说明
一个或多个实施例通过与之对应的附图中的图片进行示例性说明，这些示例性说明并不构成对实施例的限定，附图中具有相同参考数字标号的元件表示为类似的元件，除非有特别申明，附图中的图不构成比例限制。
图1是本发明实施例提供的一种偏航角的融合方法的原理框图；
图2是本发明实施例提供的一种偏航角融合算法的示意图；
图3是本发明实施例提供的一种偏航角的融合方法的整体示意图；
图4是本发明实施例提供的一种偏航角的融合方法的流程示意图；
图5是本发明实施例提供的一种计算偏航角误差角的流程示意图；
图6是图5中的步骤S401的细化流程示意图；
图7是图4中的步骤S20的细化流程图；
图8是图4中的步骤S30的细化流程图；
图9是图8中的步骤S31的细化流程图；
图10是图4中的步骤S40的细化流程图；
图11是图10中的步骤S42的细化流程图；
图12是本发明实施例提供的一种偏航角的融合装置的示意图；
图13是图12中的磁力计对准偏差角模块的示意图；
图14是图12中的GPS重对准偏差角模块的示意图；
图15是图12中的偏航角误差角模块的示意图；
图16是本发明实施例提供的一种飞行器的硬件结构示意图；
图17是本发明实施例提供的一种飞行器的连接框图；
图18是图17中的动力系统的示意图。
具体实施方式
为使本发明实施例的目的、技术方案和优点更加清楚，下面将结合本发明实施例中的附图，对本发明实施例中的技术方案进行清楚、完整地描述，显然，所描述的实施例是本发明一部分实施例，而不是全部的实施例。基于本发明中的实施例，本领域普通技术人员在没有做出创造性劳动前提下所获得的所有其他实施例，都属于本发明保护的范围。
此外，下面所描述的本发明各个实施方式中所涉及到的技术特征只要彼此之间未构成冲突就可以相互组合。
本发明实施例提供的偏航角的融合方法可以应用到各种通过电机或马达驱动的可移动物体上，包括但不限于飞行器、机器人等。其中飞行器可包括无人飞行器(unmanned aerial vehicle，UAV)，无人飞船等。现以UAV为例进行说明。UAV包括机身、控制器和动力系统。控制器与动力系统连接，动力系统安装在所述机身上，用于为所述飞行器提供飞行动力。具体的，控制器用于执行上述的偏航角的融合方法以对准及修正飞行器的偏航角，并根据融合后的飞行器的偏航角生成控制指令，并将该控制指令发送给动力系统的电调，电调通过该控制指令控制动力系统的驱动电机。或者，控制器用于执行偏航角的融合方法以对准及修正飞行器的偏航角，并将对准及修正后的飞行器的偏航角发送至电调，电调根据修正后的飞行器的偏航角生成控制指令，并通过该控制指令控制动力系统的驱动电机。
机身包括：中心壳体以及与中心壳体连接的一个或多个机臂，一个或多个机臂呈辐射状从中心壳体延伸出。机臂与中心壳体的连接可以是一体连接或者固定连接。动力系统安装于机臂上。
控制器用于执行上述偏航角的融合方法以对准并修正飞行器的偏航角，并根据对准及修正后的飞行器的偏航角生成控制指令，并将该控制指令发送给动力系统的电调，以便电调通过该控制指令控制动力系统的驱动电机。控制器为具有一定逻辑处理能力的器件，如控制芯片、单片机、微控制单元(Microcontroller Unit，MCU)等。
动力系统包括：电调，驱动电机和螺旋桨。电调位于机臂或中心壳体所形成的空腔内。电调分别与控制器及驱动电机连接。具体的，电调与驱动电机电连接，用于控制所述驱动电机。驱动电机安装在机臂上，驱动电机的转动轴连接螺旋桨。螺旋桨在驱动电机的驱动下产生使得UAV移动的力，例如，使得UAV移动的升力或者推力。
UAV完成各个规定速度、动作(或姿态)是通过电调控制驱动电机以实现的。电调全称电子调速器，根据控制信号调节UAV的驱动电机的转速。其中，控制器为执行上述偏航角的融合方法的执行主体，电调基于融合后的飞行器的偏航角所生成控制指令来控制驱动电机。电调控制驱动电机的原理大致为：驱动电机是将电脉冲信号转变为角位移或线位移的开环控制元件。在非超载的情况下，驱动电机的转速、停止的位置只取决于脉冲信号的频率和脉冲数，而不受负载变化的影响，当驱动器接收到一个脉冲信号，它就驱动动力系统的驱动电机按设定的方向转动一个固定的角度，它的旋转是以固定的角度运行的。因此，电调可以通过控制脉冲个数来控制角位移量，从而达到准确定位的目的；同时可以通过控制脉冲频率来控制驱动电机转动的速度和加速度，从而达到调速的目的。
目前UAV主要功能为航拍、影像实时传输、高危地区探测等。为了实现航拍、影像实时传输、高危地区探测等功能，UAV上会连接有摄像组件。具体的，UAV和摄像组件通过连接结构，如减振球等进行连接。该摄像组件用于在UAV 进行航拍的过程中，获取拍摄画面。
具体的，摄像组件包括：云台及拍摄装置。云台与UAV连接。其中，拍摄装置搭载于所述云台上，拍摄装置可以为图像采集装置，用于采集图像，该拍摄装置包括但不限于：相机、摄影机、摄像头、扫描仪、拍照手机等。云台用于搭载拍摄装置，以实现拍摄装置的固定、或随意调节拍摄装置的姿态(例如，改变拍摄装置的高度、倾角和/或方向)以及使所述拍摄装置稳定保持在设定的姿态上。例如，当UAV进行航拍时，云台主要用于使所述拍摄装置稳定保持在设定的姿态上，防止拍摄装置拍摄画面抖动，保证拍摄画面的稳定。
云台包括：云台电机及云台基座。其中，云台电机安装于云台基座。飞行器的控制器也可通过动力系统的电调来控制云台电机，具体的，飞行器的控制器与电调连接，电调与云台电机电连接，飞行器的控制器生成云台电机控制指令，电调通过云台电机控制指令以控制云台电机。
云台基座与UAV的机身连接，用于将摄像组件固定安装于UAV的机身上。
云台电机分别与云台基座及拍摄装置连接。该云台可以为多轴云台，与之适应的，云台电机为多个，也即每个轴设置有一个云台电机。云台电机一方面可带动拍摄装置的转动，从而满足拍摄转轴的水平旋转和俯仰角度的调节，通过手动远程控制云台电机旋转或利用程序让电机自动旋转，从而达到全方位扫描监控的作用；另一方面，在UAV进行航拍的过程中，通过云台电机的转动实时抵消拍摄装置受到的扰动，防止拍摄装置抖动，保证拍摄画面的稳定。
拍摄装置搭载于云台上，拍摄装置上设置有惯性测量单元(Inertial measurement unit，IMU)，该惯性测量单元用于测量物体三轴姿态角(或角速率)以及加速度的装置。一般的，一个IMU内会装有三轴的陀螺仪和三个方向的加速度计，来测量物体在三维空间中的角速度和加速度，并以此解算出物体的姿态。为了提高可靠性，还可以为每个轴配备更多的传感器。一般而言IMU要安装在UAV的重心上。
在对UAV的姿态进行控制的过程中，UAV的偏航角是对UAV的姿态进行控制中的重要参数，需要基于UAV偏航角，来控制驱动电机。通过飞行器的控制器实时获取UAV的偏航角，为UAV的姿态控制提供必要的姿态信息。也即UAV的偏航角正确估算对UAV的姿态控制尤其重要，若UAV的偏航角估算错误，UAV轻则无法按照预设的方向或轨迹飞行，重则可能失稳以致炸机。
飞行器的偏航角融合问题是一个相当难且重要的问题，飞行器的偏航角关系着飞行器的航向，并影响航拍无人机以下几个方面的性能：
云台的对中性问题。若偏航角不准确，飞机的云台指向的方向和飞机机头所指向的方向不重合，会存在一个夹角，严重影响拍摄和用户体验。
偏航角不准确会影响航向，出现斜飞的问题。飞机在偏航角不准确的情况下，用户给飞机发送向前(后)飞行的指令，飞机并不会沿着机头(尾)方向飞行，而是沿着斜方向飞行，在侧向(左/右)会有一个速度分量。
偏航角不准确会影响环绕飞行的航拍性能。在飞机环绕飞行时，若云台对 准环绕中心进行拍摄，假如偏航角融合不准确，或者存在长时间的漂移，则会导致拍摄对象不在图像/视频中心，而存在一定角度的左偏或右偏。
当偏航角误差较大时，飞机正常飞行中会出现刷锅、炸机等严重问题。
当偏航角误差较大时，会影响控制性能，飞机的稳定性下降。
当飞机的磁力计受到磁场干扰，这时强行起飞时，会出现刷锅、炸机的问题。
在环境复杂、磁场多变、GPS信号不稳定等情况下飞行时，偏航角收敛太慢的问题。
现有的偏航角融合方案中，通常是基于磁力计采集的数据得到UAV的偏航角，但采用该方法得到偏航角很容易受到外界因素的影响，尤其是当磁力计处于强磁干扰环境中时，磁力计的数据可能严重错误，导致偏航角的估算出现较大的偏差，飞行器的偏航角的估算准确度低。
为了提高UAV的偏航角估算的准确性，目前常见的是利用外置的GPS模组，依靠GPS估算出一个偏航角，用以对基于磁力计的偏航角值进行修正。但GPS信号有时可能不稳定，使得在某些情况下，磁力计估算的偏航角即使出现了偏差却得不到有效修正。也即利用外置的GPS模组修正UAV的偏航角虽然一定程度上可以提高飞行器的偏航角的估算准确度，但效果并不好，特别是当GPS信号较弱，利用外置的GPS模组未能起到有效的航向修正作用。由于缺少多次重对准方案，因此偏航角收敛速度较慢，且精度不高。
因此，基于上述问题，本发明实施例主要目的在于提供一种偏航角的融合方法、装置及飞行器，可以基于GPS数据、IMU数据以及磁力计数据，对飞行器的偏航角进行对准及修正，融合生成偏航角，有效提高了偏航角的融合精度以及收敛速度，从而提高飞行器飞行的安全性与稳定性。
本发明实施例通过GPS数据、IMU数据以及磁力计数据，对修正后的偏航角、磁力计对准偏差角以及GPS重对准偏差角进行融合，该融合方法可避免外界因素的干扰，也即在GPS信号弱与强磁干扰环境下，可以进行偏航角的快速收敛，并且通过多次对准及修正能够提高融合精度，从而有效提高飞行器飞行的安全性与稳定性。
下面结合附图，对本发明实施例作进一步阐述。
实施例一
请参阅图1，图1是本发明实施例提供的一种偏航角的融合方法的原理框图；
如图1所示，通过获取GPS数据、IMU数据以及磁力计数据，分别确定对地高度、GPS速度信号质量、GPS加速度信号质量、偏航角偏差角、静止标志位、IMU角速度、磁场信号质量、偏航角误差角，将所述对地高度、GPS速度信号质量、GPS加速度信号质量、偏航角偏差角、静止标志位、IMU角速度、磁场信号质量、偏航角误差角输入偏航角融合算法，融合生成融合偏航角。
请一并参阅图2和图3，图2是本发明实施例提供的一种偏航角融合算法的示意图，图3是本发明实施例提供的一种偏航角的融合方法的整体示意图；
如图2和图3所示，通过获取GPS速度信号质量、GPS加速度信号质量、偏航角偏差角、IMU角速度、偏航角误差角、对地高度、磁场信号质量、静止标志位、磁力计偏航角等数据，并对上述数据进行相应的处理，最终进行融合生成融合偏航角。
请参阅图4，图4是本发明实施例提供的一种偏航角的融合方法的流程示意图。其中，该偏航角的融合方法可由各种具有一定逻辑处理能力的电子设备执行，如飞行器、控制芯片等，该飞行器可以包括无人机、无人船等。以下电子设备以飞行器为例进行说明。其中，飞行器连接有云台，云台包括云台电机及云台基座，其中，云台可以为多轴云台，如两轴云台、三轴云台，以下三轴云台为例进行说明。对于该飞行器及云台的具体结构的描述可以参考上述描述，因此，在此处不作赘述。
如图4所示，所述方法应用于飞行器，比如，无人机，所述方法包括：
步骤S10：获取GPS数据、IMU数据以及磁力计数据，所述GPS数据包括：GPS位置、速度、加速度信息以及GPS速度信号质量，所述IMU数据包括：IMU加速度信息以及IMU角速度信息；
具体的，所述飞行器设置有姿态传感器组件，所述姿态传感器组件包括：全球定位系统(Global Positioning System，GPS)、惯性测量单元(Inertial measurement unit，IMU)、磁力计等，其中，所述GPS用于获取GPS数据，所述惯性测量单元(IMU)用于获取IMU数据，所述磁力计用于获取磁力计数据，所述惯性测量单元包括陀螺仪以及加速度计，所述陀螺仪用于获取IMU角速度，所述加速度计用于获取IMU角速度信息。所述GPS数据包括：位置信息(经纬度信息)、GPS速度信息、GPS加速度信息，所述IMU数据包括：IMU加速度信息以及IMU角速度信息，所述磁力计数据包括：磁场强度信息。
具体的，通过数据处理、滤波等算法获取所述经纬度信息、GPS速度信息，所述数据处理包括：数据预处理，例如：解压缩、变换坐标系、变换单位等，所述滤波包括低通滤波和卡尔曼滤波，分别通过低通滤波器和卡尔曼滤波器完成。
步骤S20：根据所述IMU数据、所述GPS数据以及所述磁力计数据，确定修正后偏航角；
具体的，请再参阅图7，图7是图4中的步骤S20的细化流程图；
如图7所示，所述根据所述IMU角速度信息，确定修正后偏航角，包括：
步骤S21：获取所述飞行器当前位置的磁场信号质量；
具体的，根据所述飞行器当前位置的标准磁场的磁场强度，对比所述飞行器的磁力计测量的磁力计数据，即所述磁场强度，获取所述飞行器当前位置的磁场信号质量，若所述飞行器的磁力计测量的磁力计数据与所述飞行器当前位置的标准磁场的磁场强度相差较大，则说明所述磁场信号质量差，若所述飞行 器的磁力计测量的磁力计数据与所述飞行器当前位置的标准磁场的磁场强度相近，则说明所述磁场信号质量好，一般通过幅值和倾角评判所述飞行器当前位置的磁场信号质量。
步骤S22：根据所述偏航角偏差角，生成GPS角速度修正量；
具体的，所述飞行器设置有反馈控制器，所述GPS角速度修正量通过所述反馈控制器基于反馈控制算法对所述偏航角偏差角计算得出，所述GPS角速度修正量与所述偏航角偏差角负相关，例如：GPS角速度修正量＝-K*偏航角偏差角，其中，所述K为工程师设计值。
步骤S23：对所述GPS加速度修正权重系数以及所述GPS角速度修正量进行求积，生成GPS加权修正量；
具体的，通过对所述GPS加速度修正权重系数以及所述GPS角速度修正量进行求积，将乘积作为所述GPS加权修正量。
如图1所示，其中，所述GPS加速度修正权重系数由所述GPS速度信号质量和GPS加速度信号质量求积所得，所述GPS加速度信号质量由所述GPS加速度信息计算所得，其中，所述GPS加速度信号质量由所述GPS加速度经过信号质量检测与判定之后确定得到，具体的，GPS模块会给出信号质量，出厂就已经有了，通过设置一个门限，大于所述门限则表示GPS速度信号质量好，小于所述门限则表示GPS速度信号质量差，其中，所述门限人为设置。
其中，GPS模块给出GPS速度信号质量之后，而GPS速度信息用来计算微分，可求得微分噪声的大小，根据噪声大小评判微分质量好不好，噪声大就不好，噪声小就好。当GPS速度信号质量好，并且微分的噪声小，则GPS加速度信号质量好，否则GPS加速度信号质量差。
步骤S24：根据所述偏航角误差角，生成磁力计修正量；
具体的，所述飞行器设置有反馈控制器，所述磁力计修正量通过所述反馈控制器基于反馈控制算法对所述偏航角误差角计算得出，所述反馈控制算法为：所述磁力计修正量与所述偏航角误差角负相关，例如：磁力计修正量＝-K*偏航角误差角，其中，所述K为工程师设计值。
步骤S25：对所述磁场信号质量以及所述磁力计修正量进行求积，生成磁力计加权修正量；
具体的，对所述磁场信号质量以及所述磁力计修正量进行求积，将乘积作为所述磁力计加权修正量。
步骤S26：对所述IMU角速度信息、所述GPS加权修正量以及所述磁力计加权修正量进行融合，生成修正后偏航角角速度；
具体的，所述IMU角速度信息由惯性测量单元的角速度计获取，通过对所述IMU角速度信息、所述GPS加权修正量以及所述磁力计加权修正量进行融合，求所述IMU角速度信息、所述GPS加权修正量以及所述磁力计加权修正量的和值，将所述和值作为所述修正后偏航角角速度。其中，所述修正后偏航角角速度为经过GPS修正以及磁力计修正后的偏航角角速度。
步骤S27：对所述修正后偏航角角速度进行积分，生成修正后偏航角。
具体的，通过对所述修正后偏航角角速度进行积分，将积分值作为修正后偏航角。
步骤S30：根据所述磁力计数据、所述GPS数据以及所述修正后偏航角，确定磁力计对准偏差角；
请参阅图8，图8是图4中的步骤S30的细化流程图；
如图8所示，所述根据所述磁力计数据，确定磁力计对准偏差角，包括：
步骤S31：根据所述磁力计偏航角，确定磁力计对准偏航角；
具体的，所述飞行器设置有使能模块，对准采用使能模块来实现，所述使能模块包括输入端和输出端，一旦所述使能模块检测到脉冲，所述使能模块将所述输入端的值送给所述输出端，以使所述输出端输出高电平，否则，所述输出端输入低电平或置零。所述磁力计对准偏航角为所述磁力计对准时所测得的偏航角大小，其中，所述脉冲是磁力计对准脉冲。
可以理解的是，在本发明实施例中，所述使能模块可以替换其他运算模块，例如：乘法运算模块等，均在本发明实施例的保护范围。
具体的，请再参阅图9，图9是图8中的步骤S31的细化流程图；
如图9所示，所述根据所述磁力计偏航角，确定磁力计对准偏航角，包括：
步骤S311：获取所述飞行器的对地高度及静止状态位；
具体的，所述飞行器设置有超声波传感器和/或TOF传感器(Time of Flight，TOF)，所述飞行器的对地高度通过所述超声波传感器和/或TOF传感器获取，所述静止状态位表征所述飞行器的静止状态。
具体的，请复参阅图1，如图1所示，所述获取所述飞行器的静止状态位，包括：
获取所述IMU数据中的IMU加速度以及IMU角速度，通过对所述IMU加速度以及IMU角速度进行静止检测，确定所述飞行器的静止状态，若所述飞行器的静止状态为静止，则所述静止状态位的值为1，若所述飞行器的静止状态为运动，则所述静止状态位的值为0，可以理解的是，静止时所述IMU加速度以及IMU角速度的数据波动特别小，很稳定，运动时所述IMU加速度以及IMU角速度的数据波动变化大，根据此原理确定所述飞行器的静止状态。其中，所述方法具体包括：根据旋转变换矩阵，对所述IMU数据进行机体坐标系和地面坐标系的转换，生成地面坐标系下的IMU加速度以及地面坐标系下的IMU角速度；根据所述地面坐标系下的IMU加速度以及地面坐标系下的IMU角速度，确定所述飞行器的静止状态，生成所述飞行器的静止标志位。
步骤S312：对所述磁场信号质量以及所述静止标志位进行逻辑判断，生成磁力计初对准脉冲；
具体的，所述对所述磁场信号质量以及所述静止标志位进行逻辑判断，其判断逻辑包括：
Step1.i＝0，转Step2；
Step2.当所述静止标志位信号为0时，脉冲信号F＝0，跳转Step4；当静止标志位信号为1时，判断所述磁场信号质量是否大于预设磁场信号质量阈值，在本步骤中，所述磁场信号质量阈值为0.6，若是，所述脉冲信号F输出1、跳转Step3，若否，所述脉冲信号F输出0、跳转Step4。
Step3.判断所述脉冲信号F是否是首次从0变成1，若是，输出一个脉冲给磁力计初对准脉冲，若否输出0，转Step4。
Step4.i＝i+1，返回Step2。
可以理解的是，在对所述磁场信号质量以及所述静止标志位进行逻辑判断时，所述磁场信号质量阈值可以根据具体需求具体设置，均在本发明的保护范围内。
步骤S313：对所述对地高度以及所述磁场信号质量进行逻辑判断，生成磁力计重对准脉冲；
具体的，对所述对地高度以及所述磁场信号质量进行逻辑判断，其判断逻辑包括：
Step1.i＝0，转Step2；
Step2.当所述对地高度小于预设对地高度阈值时，在本步骤中，所述对地高度阈值设置为0.3m，脉冲信号F＝0，跳转Step4；当所述对地高度大于等于所述对地高度阈值时，判断所述磁场信号质量是否大于预设磁场信号质量阈值，在本步骤中，所述磁场信号质量阈值设置为0.6，若是，脉冲信号F输出1，跳转Step3，若否，脉冲信号F输出0，跳转Step4。
Step3.判断所述脉冲信号F是否是首次从0变成1，若是，输出一个脉冲给磁力计重对准脉冲，若否输出0，转Step4。
Step4.i＝i+1，返回Step2。
步骤S314：根据所述磁力计初对准脉冲以及所述磁力计重对准脉冲，生成磁力计对准脉冲；
具体的，所述根据所述磁力计初对准脉冲以及所述磁力计重对准脉冲，生成磁力计对准脉冲，包括：对所述磁力计初对准脉冲以及所述磁力计重对准脉冲进行或运算，生成所述磁力计对准脉冲。
可以理解的是，在对所述对地高度以及所述磁场信号质量进行逻辑判断时，所述磁场信号质量阈值以及所述对地高度阈值可以根据具体需求具体设置，均在本发明的保护范围内。
步骤S315：将所述磁力计对准脉冲输入飞行器的使能模块，当所述使能模块检测到所述磁力计对准脉冲存在上升沿时，则将所述磁力计偏航角作为所述磁力计对准偏航角。
具体的，所述使能模块的输入端，即使能端用于接收所述磁力计对准脉冲，将所述磁力计对准脉冲输入飞行器的使能模块，当所述使能模块检测到所述磁力计对准脉冲的信号存在上升沿时，所述使能模块将所述磁力计偏航角作为所述磁力计对准偏航角。
步骤S32：根据所述修正后偏航角以及所述磁力计对准偏航角，确定所述磁力计对准偏差角。
具体的，请复参阅图2，如图2所示，根据所述修正后偏航角以及所述磁力计对准偏航角，融合生成所述磁力计对准偏差角，具体的，所述磁力计对准偏差角＝-修正后偏航角+磁力计对准偏航角。
步骤S40：根据所述GPS数据以及所述IMU加速度信息，确定GPS重对准偏差角；
具体的，所述GPS数据包括GPS加速度信息，所述GPS加速度信息通过对所述GPS速度信息以及位置信息进行加速度估算获取，所述IMU加速度信息通过所述惯性测量单元的加速度计获取。
请再参阅图10，图10是图4中的步骤S40的细化流程图；
如图10所示，所述根据所述GPS加速度信息以及所述IMU加速度信息，确定GPS重对准偏差角，包括：
步骤S41：根据所述GPS加速度信息以及所述IMU加速度信息，计算偏航角偏差角；
具体的，所述IMU加速度信息为惯性测量单元的测量数据经过校准矩阵进行校准以及机体坐标系到地面坐标系的坐标变换之后，所得到的地面坐标系下的加速度信息。可以理解的是，所述校准矩阵是用户在要飞行的地方校准得到的，校准矩阵在地球上任意地方都不同，飞行器在报磁力计干扰，要求用户校准之后才能确定所述校准矩阵。
步骤S42：根据所述偏航角偏差角，确定GPS重对准偏差角。
具体的，请复参阅图2，所述飞行器设置有使能模块，所述使能模块包括输入端和输出端，当所述使能模块的输入端接收到重对准脉冲时，所述使能模块将所述偏航角偏差角输出，作为重对准偏航角偏差补偿量，即所述GPS重对准偏差角。
具体的，请一并参阅图2和图11，图11是图10中的步骤S42的细化流程图；
如图11所示，所述根据所述偏航角偏差角，确定GPS重对准偏差角，包括：
步骤S421：根据所述GPS速度信息以及所述GPS加速度信息，分别计算GPS速度信号质量以及GPS加速度信号质量；
具体的，根据GPS获取到的GPS位置信息(经纬度信息)、GPS速度信息，根据信号的质量检测与判定原则，生成GPS速度信号质量参考值，将所述GPS速度信号质量参考值作为所述GPS速度信号质量，并且，对GPS获取的GPS速度信息进行求微分，并进行滤波处理，生成所述GPS加速度信息，并根据信号的质量检测与判定原则，生成GPS加速度信号质量参考值，作为所述GPS加速度信号质量。其中，所述GPS会给出所述GPS速度信号质量，通过对所述GPS获取的GPS速度信息进行求微分，可以求得微分噪声的大小，根据所述噪 声的大小，可以评判所述GPS加速度信号质量，若噪声大，则所述GPS加速度信号质量差，若噪声小，则所述GPS加速度信号质量好。
可以理解的是，所述GPS速度信号质量以及GPS加速度信号质量用于辅助参考，反映GPS获取的信号的质量，其呈现正相关，若所述GPS获取的信号质量好，则所述GPS速度信号质量以及GPS加速度信号质量的数值大，反之，若所述GPS获取的信号质量差，则所述GPS速度信号质量以及GPS加速度信号质量的数值小。
步骤S422：对所述GPS速度信号质量以及GPS加速度信号质量进行求积，生成GPS加速度修正权重系数；
具体的，对所述GPS速度信号质量以及GPS加速度信号质量进行求积，将求积得到的值作为所述GPS加速度修正权重系数。其中，所述GPS速度信号质量以及GPS加速度信号质量为置信度量化信息，若所述GPS加速度修正权重系数越大，则置信度越高，反之则越低。
步骤S423：判断所述GPS加速度修正权重系数是否大于预设GPS重对准阈值；
具体的，所述飞行器设置有重对准标志位，通过预先设置GPS重对准阈值，将所述GPS加速度修正权重系数与所述GPS重对准阈值进行对比，若所述GPS加速度修正权重系数大于所述GPS重对准阈值，则进入步骤S224：控制所述飞行器的重对准标志位输出高电平信号，即输出1；若所述GPS加速度修正权重系数不大于所述GPS重对准阈值，则控制所述飞行器的重对准标志位输出低电平信号，即输出0。
可以理解的是，本发明实施例中的重对准阈值可以根据具体需求人为设置。
步骤S424：控制所述飞行器的重对准标志位输出高电平信号；
步骤S425：控制所述重对准标志位输出低电平信号；
步骤S426：若所述重对准标志位的信号存在上升沿，则生成重对准脉冲信号，将所述重对准脉冲信号输入所述飞行器的使能模块；
具体的，所述飞行器设置有使能模块，通过检测所述重对准标志为是否存在上升沿，若存在，则生成重对准脉冲信号，并将所述重对准脉冲信号输入所述使能模块的输入端。
步骤S427：当所述使能模块检测到所述重对准脉冲存在上升沿时，则将所述偏航角偏差角作为所述GPS重对准偏差角。
具体的，当所述使能模块的输入端，即使能端接收到脉冲时，所述使能模块将所述偏航角偏差角输出作为重对准偏航角偏差补偿量，即所述GPS重对准偏差角。
在本发明实施例中，所述方法还包括：计算偏航角误差角，具体的，请参阅图5，图5是本发明实施例提供的一种计算偏航角误差角的流程示意图；
如图5所示，所述计算偏航角误差角，包括：
步骤S401：根据所述磁力计数据，计算磁力计偏航角；
其中，所述磁力计数据，包含三轴磁场强度，GPS经纬度信息用于查找所述飞行器的当前位置的标准磁场强度、磁倾角信息以及磁偏角信息。其中，所述磁力计数据通过事先校准的校准矩阵对所述磁力计测量的磁场强度进行变换，得到校准后的磁力计数据，并根据所述校准后的磁力计数据，生成所述飞行器的当前位置的磁场信号质量参考值。
具体的，请参阅图6，图6是图5中的步骤S401的细化流程示意图；
如图6所示，所述根据所述磁力计数据，计算磁力计偏航角，包括：
步骤S4011：根据预设的校准矩阵对所述磁力计数据进行校准，生成校准后的磁力计数据；
具体的，所述预设的校准矩阵是用户在要飞行的地方校准得到的，校准矩阵在地球上任意地方都不同，飞行器在报磁力计干扰，要求用户校准之后才能确定所述校准矩阵。
步骤S4012：根据所述飞行器的姿态角，生成旋转变换矩阵；
具体的，所述旋转变换矩阵用于将机体坐标系转换为地面坐标系，所述飞行器的姿态角包括：偏航角、俯仰角、翻滚角，其中，所述旋转变换矩阵为3*3的矩阵，其中包含了所述偏航角、俯仰角、翻滚角的正弦余弦函数，并根据具体情况选择不同的函数，一般而言，通过先转动偏航角，再转动俯仰角，最后转动翻滚角，例如：所述旋转变换矩阵为：
Figure PCTCN2020110592-appb-000001
其中，(φ,θ,ψ)为所述姿态角，φ为所述姿态角中的翻滚角，θ为所述姿态角中的俯仰角，ψ为所述姿态角中的偏航角。
步骤S4013：根据所述旋转变换矩阵进行机体坐标系和地面坐标系的转换，对所述校准后的磁力计数据进行坐标变换，生成所述地面坐标系下的磁力计数据；
具体的，将所述校准后的磁力计数据，即所述磁场强度，乘以所述旋转变换矩阵，生成所述地面坐标系下的磁场强度，相当于对所述磁力计数据进行机体坐标系和地面坐标系之间的变换。
步骤S4014：根据所述地面坐标系下的磁力计数据，对比所述飞行器的当前位置的标准磁场的磁力计数据，计算所述磁力计偏航角。
具体的，所述飞行器的当前位置对应一标准磁场，所述磁力计三轴读数组成一个向量，所述当前位置的标准磁场的磁力计数据对应一向量，通过将所述地面坐标系下的磁力计数据，对比所述飞行器的当前位置的标准磁场的磁力计数据，计算两者之间的向量夹角，将所述向量夹角作为所述磁力计偏航角。
步骤S402：根据所述磁力计偏航角和所述融合偏航角，计算偏航角误差角。
具体的，所述飞行器的每一个采样步长都有做一次误差计算，通过反馈回路一直不停歇的无限进行。融合好的融合偏航角要去与磁力计的偏航角作差，生成偏航角误差角，无休止的进行，直到飞机断电。所述融合偏航角也是无休止的在更新，每一个采样时刻对应唯一的融合偏航角，所述偏航角误差角为所述磁力计偏航角和所述融合偏航角的差值。步骤S50：根据所述GPS重对准偏差角、修正后偏航角以及磁力计对准偏差角，生成融合偏航角。
具体的，通过对所述GPS重对准偏差角、修正后偏航角以及磁力计对准偏差角进行融合，生成融合偏航角，其中，所述融合偏航角＝GPS重对准偏差角+修正后偏航角+磁力计对准偏差角。
在本发明实施例中，偏航角的修正体现在偏航角角速度的修正上，包括GPS修正和磁力计修正，偏航角的对准体现在偏航角偏差补正上，初始对准采用磁力计对准，重对准包括GPS对准和磁力计对准。通过采用多次偏航角对准方案，减小了复杂环境下飞行的偏航角收敛慢的问题，同时提高了融合精度；采用多次对准和修正方案，增加了偏航角融合的鲁棒性，当偏航角误差较大、且磁力计或GPS信息质量较好时，能够实现一步快速补偿，并且，小的偏航角误差通过反馈控制器能够保证收敛；解决了恶劣磁场环境下的飞行问题、长时间环绕飞行时偏航角漂移问题、磁场受干扰时偏航角错乱的问题、飞机斜飞问题等因偏航角错误所导致的一系列问题。
在本发明实施例中，通过提供一种偏航角的融合方法，应用于飞行器，所述方法包括：获取GPS数据、IMU数据以及磁力计数据，所述GPS数据包括：GPS位置、速度、加速度信息以及GPS速度信号质量，所述IMU数据包括：IMU加速度信息以及IMU角速度信息；根据所述IMU数据、所述GPS数据以及所述磁力计数据，确定修正后偏航角；根据所述磁力计数据、所述GPS数据以及所述修正后偏航角，确定磁力计对准偏差角；根据所述GPS数据以及所述IMU加速度信息，确定GPS重对准偏差角；根据所述修正后偏航角、磁力计对准偏差角以及GPS重对准偏差角，生成融合偏航角。通过上述方式，本发明实施例能够有效提高偏航角的融合精度以及收敛速度。
实施例二
请参阅图12，图12为本发明实施例提供的一种偏航角的融合装置的示意图，该偏航角的融合装置120可以应用于飞行器，例如无人飞行器。
如图12所示，该偏航角的融合装置120包括：
获取模块121，用于获取GPS数据、IMU数据以及磁力计数据，所述GPS数据包括：GPS位置、速度、加速度信息以及GPS速度信号质量，所述IMU数据包括：IMU加速度信息以及IMU角速度信息；
修正后偏航角模块122，用于根据所述IMU数据、所述GPS数据以及所述磁力计数据，确定修正后偏航角；
磁力计对准偏差角模块123，用于根据所述磁力计数据、所述GPS数据以 及所述修正后偏航角，确定磁力计对准偏差角；
GPS重对准偏差角模块124，用于根据所述GPS数据以及所述IMU加速度信息，确定GPS重对准偏差角；
偏航角融合模块125，用于根据所述GPS重对准偏差角、修正后偏航角以及磁力计对准偏差角，生成融合偏航角；
偏航角误差角模块126，用于计算偏航角误差角；
请再参阅图13，图13是图12中的磁力计对准偏差角模块的示意图；
如图13所示，该磁力计对准偏差角模块123，包括：
磁力计对准偏航角单元1231，用于根据所述磁力计偏航角，确定磁力计对准偏航角；
磁力计对准偏差角单元1232，用于根据所述修正后偏航角以及所述磁力计对准偏航角，确定所述磁力计对准偏差角。
在本发明实施例中，所述磁力计对准偏航角单元1231，具体用于：
获取所述飞行器的对地高度及静止状态位；
对所述磁场信号质量以及所述静止标志位进行逻辑判断，生成磁力计初对准脉冲；
对所述对地高度以及所述磁场信号质量进行逻辑判断，生成磁力计重对准脉冲；
根据所述磁力计初对准脉冲以及所述磁力计重对准脉冲，生成磁力计对准脉冲；
若所述磁力计对准脉冲的信号存在上升沿，则将所述磁力计偏航角作为所述磁力计对准偏航角。
请参阅图14，图14是图12中的GPS重对准偏差角模块的示意图；
如图14所示，该GPS重对准偏差角模块124，包括：
偏航角偏差角单元1241，用于根据所述GPS加速度信息以及所述IMU加速度信息，计算偏航角偏差角；
GPS重对准偏差角单元1242，用于根据所述偏航角偏差角，确定GPS重对准偏差角。
在本发明实施例中，所述GPS重对准偏差角单元，具体用于：
根据所述GPS速度信息以及所述GPS加速度信息，分别计算GPS速度信号质量以及GPS加速度信号质量；
对所述GPS速度信号质量以及GPS加速度信号质量进行求积，生成GPS加速度修正权重系数；
判断所述GPS加速度修正权重系数是否大于预设GPS重对准阈值，若是，则控制所述飞行器的重对准标志位输出高电平信号，若否，则控制所述重对准标志位输出低电平信号；
若所述重对准标志位的信号存在上升沿，则生成重对准脉冲信号，将所述重对准脉冲信号输入所述飞行器的使能模块；
当所述使能模块接收到所述重对准脉冲信号时，将所述偏航角偏差角作为所述GPS重对准偏差角。
请再参阅图15，图15是图12中的偏航角误差角模块的示意图；
如图15所示，该偏航角误差角模块126，包括：
磁力计偏航角单元1261，用于根据所述磁力计数据，计算磁力计偏航角；
偏航角误差角单元1262，用于根据所述磁力计偏航角和所述融合偏航角，计算偏航角误差角。
在本发明实施例中，所述磁力计偏航角单元1261，具体用于：
根据预设的校准矩阵对所述磁力计数据进行校准，生成校准后的磁力计数据；
根据所述飞行器的姿态角，生成旋转变换矩阵；
根据所述旋转变换矩阵进行机体坐标系和地面坐标系的转换，对所述校准后的磁力计数据进行坐标变换，生成所述地面坐标系下的磁力计数据；
根据所述地面坐标系下的磁力计数据，对比所述飞行器的当前位置的标准磁场的磁力计数据，计算所述磁力计偏航角。
由于装置实施例和方法实施例是基于同一构思，在内容不互相冲突的前提下，装置实施例的内容可以引用方法实施例的，在此不赘述。
在本发明实施例中，通过提供一种偏航角的融合装置，应用于飞行器，所述装置包括：获取模块，用于获取GPS数据、IMU数据以及磁力计数据，所述GPS数据包括：GPS位置、速度、加速度信息以及GPS速度信号质量，所述IMU数据包括：IMU加速度信息以及IMU角速度信息；修正后偏航角模块，用于根据所述IMU数据、所述GPS数据以及所述磁力计数据，确定修正后偏航角；磁力计对准偏差角模块，用于根据所述磁力计数据、所述GPS数据以及所述修正后偏航角，确定磁力计对准偏差角；GPS重对准偏差角模块，用于根据所述GPS数据以及所述IMU加速度信息，确定GPS重对准偏差角；偏航角融合模块，用于根据所述修正后偏航角、磁力计对准偏差角以及GPS重对准偏差角，生成融合偏航角。通过上述方式，本发明实施例能够有效提高偏航角的融合精度以及收敛速度。
请参阅图16，图16是本发明实施例提供一种飞行器的硬件结构示意图。其中，该飞行器可以是无人飞行器(unmanned aerial vehicle，UAV)，无人飞船等电子设备。
如图16所示，该飞行器160包括一个或多个处理器161以及存储器162。其中，图16中以一个处理器161为例。
处理器161和存储器162可以通过总线或者其他方式连接，图16中以通过总线连接为例。
存储器162作为一种非易失性计算机可读存储介质，可用于存储非易失性软件程序、非易失性计算机可执行程序以及模块，如本发明实施例中的一种偏航角的融合方法对应的单元(例如，图12至图15所述的各个模块或单元)。 处理器161通过运行存储在存储器162中的非易失性软件程序、指令以及模块，从而执行偏航角的融合方法的各种功能应用以及数据处理，即实现上述方法实施例偏航角的融合方法以及上述装置实施例的各个模块和单元的功能。
存储器162可以包括高速随机存取存储器，还可以包括非易失性存储器，例如至少一个磁盘存储器件、闪存器件、或其他非易失性固态存储器件。在一些实施例中，存储器162可选包括相对于处理器161远程设置的存储器，这些远程存储器可以通过网络连接至处理器161。上述网络的实例包括但不限于互联网、企业内部网、局域网、移动通信网及其组合。
所述模块存储在所述存储器162中，当被所述一个或者多个处理器161执行时，执行上述任意方法实施例中的偏航角的融合方法，例如，执行以上描述的图4至图11所示的各个步骤；也可实现图12至图15所述的各个模块或单元的功能。
请参阅图17和图18，所述飞行器160还包括动力系统163，所述动力系统163用于飞行器提供飞行动力，所述动力系统163与处理器161连接。所述动力系统163包括：驱动电机1631及电调1632，所述电调1632与驱动电机1631电连接，用于控制所述驱动电机1631。具体的，所述电调1632基于处理器161执行上述偏航角的融合方法后得到的融合偏航角，生成控制指令，通过控制指令控制该驱动电机1631。
所述飞行器160可执行本发明实施例一所提供的偏航角的融合方法，具备执行方法相应的功能模块和有益效果。未在飞行器实施例中详尽描述的技术细节，可参见本发明实施例一所提供的偏航角的融合方法。
本发明实施例提供了一种计算机程序产品，所述计算机程序产品包括存储在非易失性计算机可读存储介质上的计算机程序，所述计算机程序包括程序指令，当所述程序指令被计算机执行时，使所述计算机执行如上所述的偏航角的融合方法。例如，执行以上描述的图4中的方法步骤S10至步骤S50。
本发明实施例还提供了一种非易失性计算机存储介质，所述计算机存储介质存储有计算机可执行指令，该计算机可执行指令被一个或多个处理器执行，例如图16中的一个处理器161，可使得上述一个或多个处理器可执行上述任意方法实施例中的偏航角的融合方法，例如，执行上述任意方法实施例中的偏航角的融合方法，例如，执行以上描述的图4至图11所示的各个步骤；也可实现图12至图15所述的各个模块或单元的功能。
以上所描述的装置或设备实施例仅仅是示意性的，其中所述作为分离部件说明的单元模块可以是或者也可以不是物理上分开的，作为模块单元显示的部件可以是或者也可以不是物理单元，即可以位于一个地方，或者也可以分布到多个网络模块单元上。可以根据实际的需要选择其中的部分或者全部模块来实现本实施例方案的目的。
通过以上的实施方式的描述，本领域的技术人员可以清楚地了解到各实施方式可借助软件加通用硬件平台的方式来实现，当然也可以通过硬件。基于这 样的理解，上述技术方案本质上或者说对相关技术做出贡献的部分可以以软件产品的形式体现出来，该计算机软件产品可以存储在计算机可读存储介质中，如ROM/RAM、磁碟、光盘等，包括若干指令用直至得一台计算机设备(可以是个人计算机，服务器，或者网络设备等)执行各个实施例或者实施例的某些部分所述的方法。
最后应说明的是：以上实施例仅用以说明本发明的技术方案，而非对其限制；在本发明的思路下，以上实施例或者不同实施例中的技术特征之间也可以进行组合，步骤可以以任意顺序实现，并存在如上所述的本发明的不同方面的许多其它变化，为了简明，它们没有在细节中提供；尽管参照前述实施例对本发明进行了详细的说明，本领域的普通技术人员应当理解：其依然可以对前述各实施例所记载的技术方案进行修改，或者对其中部分技术特征进行等同替换；而这些修改或者替换，并不使相应技术方案的本质脱离本申请各实施例技术方案的范围。

Claims (15)

  1. 一种偏航角的融合方法，应用于飞行器，其特征在于，所述方法包括：
    获取GPS数据、IMU数据以及磁力计数据，所述GPS数据包括：GPS位置、速度、加速度信息以及GPS速度信号质量，所述IMU数据包括：IMU加速度信息以及IMU角速度信息；
    根据所述IMU数据、所述GPS数据以及所述磁力计数据，确定修正后偏航角；
    根据所述磁力计数据、所述GPS数据以及所述修正后偏航角，确定磁力计对准偏差角；
    根据所述GPS数据以及所述IMU加速度信息，确定GPS重对准偏差角；
    根据所述修正后偏航角、磁力计对准偏差角以及GPS重对准偏差角，生成融合偏航角。
  2. 根据权利要求1所述的方法，其特征在于，所述方法还包括：
    根据所述磁力计数据，计算磁力计偏航角；
    根据所述磁力计偏航角和所述融合偏航角，计算偏航角误差角。
  3. 根据权利要求2所述的方法，其特征在于，所述根据所述磁力计数据，计算磁力计偏航角，具体包括：
    根据预设的校准矩阵对所述磁力计数据进行校准，生成校准后的磁力计数据；
    根据所述飞行器的姿态角，生成旋转变换矩阵；
    根据所述旋转变换矩阵进行机体坐标系和地面坐标系的转换，对所述校准后的磁力计数据进行坐标变换，生成所述地面坐标系下的磁力计数据；
    根据所述地面坐标系下的磁力计数据，对比所述飞行器的当前位置的标准磁场的磁力计数据，计算所述磁力计偏航角。
  4. 根据权利要求3所述的方法，其特征在于，所述根据所述IMU数据、所述GPS数据以及所述磁力计数据，确定修正后偏航角，具体包括：
    获取所述飞行器当前位置的磁场信号质量；
    根据所述偏航角偏差角，生成GPS角速度修正量；
    对GPS加速度修正权重系数以及所述GPS角速度修正量进行求积，生成GPS加权修正量；
    根据所述偏航角误差角，生成磁力计修正量；
    对所述磁场信号质量以及所述磁力计修正量进行求积，生成磁力计加权修正量；
    对所述IMU角速度信息、所述GPS加权修正量以及所述磁力计加权修正量 进行融合，生成修正后偏航角角速度；
    对所述修正后偏航角角速度进行积分，生成修正后偏航角，
    其中，所述GPS加速度修正权重系数由所述GPS速度信号质量和GPS加速度信号质量求积所得，所述GPS加速度信号质量由所述GPS加速度信息计算所得。
  5. 根据权利要求2-4任一项所述的方法，其特征在于，所述根据所述磁力计数据，确定磁力计对准偏差角，包括：
    根据所述磁力计偏航角，确定磁力计对准偏航角；
    根据所述修正后偏航角以及所述磁力计对准偏航角，确定所述磁力计对准偏差角。
  6. 根据权利要求5所述的方法，其特征在于，所述根据所述磁力计偏航角，确定磁力计对准偏航角，具体包括：
    获取所述飞行器的对地高度及静止状态位；
    对所述磁场信号质量以及所述静止标志位进行逻辑判断，生成磁力计初对准脉冲；
    对所述对地高度以及所述磁场信号质量进行逻辑判断，生成磁力计重对准脉冲；
    根据所述磁力计初对准脉冲以及所述磁力计重对准脉冲，生成磁力计对准脉冲；
    将所述磁力计对准脉冲输入飞行器的使能模块，当所述使能模块检测到所述磁力计对准脉冲存在上升沿时，则将所述磁力计偏航角作为所述磁力计对准偏航角。
  7. 根据权利要求1所述的方法，其特征在于，所述根据所述GPS数据以及所述IMU加速度信息，确定GPS重对准偏差角，包括：
    根据所述GPS加速度信息以及所述IMU加速度信息，计算偏航角偏差角；
    根据所述GPS加速度信息，计算GPS加速度信号质量；
    对所述GPS速度信号质量以及GPS加速度信号质量进行求积，生成GPS加速度修正权重系数；
    判断所述GPS加速度修正权重系数是否大于预设GPS重对准阈值，若是，则控制所述飞行器的重对准标志位输出高电平信号，若否，则控制所述重对准标志位输出低电平信号；
    若所述重对准标志位的信号存在上升沿，则生成重对准脉冲信号；以及，
    将所述重对准脉冲信号输入所述飞行器的使能模块，当所述使能模块检测到所述重对准脉冲存在上升沿时，则将所述偏航角偏差角作为所述GPS重对准偏差角。
  8. 一种偏航角的融合装置，应用于飞行器，其特征在于，所述装置包括：
    获取模块，用于获取GPS数据、IMU数据以及磁力计数据，所述GPS数据包括：GPS位置、速度、加速度信息以及GPS速度信号质量，所述IMU数据包括：IMU加速度信息以及IMU角速度信息；
    修正后偏航角模块，用于根据所述IMU数据、所述GPS数据以及所述磁力计数据，确定修正后偏航角；
    磁力计对准偏差角模块，用于根据所述磁力计数据、所述GPS数据以及所述修正后偏航角，确定磁力计对准偏差角；
    GPS重对准偏差角模块，用于根据所述GPS数据以及所述IMU加速度信息，确定GPS重对准偏差角；
    偏航角融合模块，用于根据所述修正后偏航角、磁力计对准偏差角以及GPS重对准偏差角，生成融合偏航角。
  9. 根据权利要求8所述的装置，其特征在于，所述装置还包括：偏航角误差角模块，所述偏航角误差角模块包括：
    磁力计偏航角单元，用于根据所述磁力计数据，计算磁力计偏航角；
    偏航角误差角单元，用于根据所述磁力计偏航角和所述融合偏航角，计算偏航角误差角。
  10. 根据权利要求9所述的装置，其特征在于，所述磁力计偏航角单元，具体用于：
    根据预设的校准矩阵对所述磁力计数据进行校准，生成校准后的磁力计数据；
    根据所述飞行器的姿态角，生成旋转变换矩阵；
    根据所述旋转变换矩阵进行机体坐标系和地面坐标系的转换，对所述校准后的磁力计数据进行坐标变换，生成所述地面坐标系下的磁力计数据；
    根据所述地面坐标系下的磁力计数据，对比所述飞行器的当前位置的标准磁场的磁力计数据，计算所述磁力计偏航角。
  11. 根据权利要求10所述的装置，其特征在于，所述修正后偏航角模块，具体用于：
    获取所述飞行器当前位置的磁场信号质量；
    根据所述偏航角偏差角，生成GPS角速度修正量；
    对GPS加速度修正权重系数以及所述GPS角速度修正量进行求积，生成GPS加权修正量；
    根据所述偏航角误差角，生成磁力计修正量；
    对所述磁场信号质量以及所述磁力计修正量进行求积，生成磁力计加权修 正量；
    对所述IMU角速度信息、所述GPS加权修正量以及所述磁力计加权修正量进行融合，生成修正后偏航角角速度；
    对所述修正后偏航角角速度进行积分，生成修正后偏航角，
    其中，所述GPS加速度修正权重系数由所述GPS速度信号质量和GPS加速度信号质量求积所得，所述GPS加速度信号质量由所述GPS加速度信息计算所得。
  12. 根据权利要求8-11任一项所述的装置，其特征在于，所述磁力计对准偏差角模块，包括：
    磁力计对准偏航角单元，用于根据所述磁力计偏航角，确定磁力计对准偏航角；
    磁力计对准偏差角单元，用于根据所述修正后偏航角以及所述磁力计对准偏航角，确定所述磁力计对准偏差角。
  13. 根据权利要求12所述的装置，其特征在于，所述磁力计对准偏航角单元，具体用于：
    获取所述飞行器的对地高度及静止状态位；
    对所述磁场信号质量以及所述静止标志位进行逻辑判断，生成磁力计初对准脉冲；
    对所述对地高度以及所述磁场信号质量进行逻辑判断，生成磁力计重对准脉冲；
    根据所述磁力计初对准脉冲以及所述磁力计重对准脉冲，生成磁力计对准脉冲；
    将所述磁力计对准脉冲输入飞行器的使能模块，当所述使能模块检测到所述磁力计对准脉冲存在上升沿时，则将所述磁力计偏航角作为所述磁力计对准偏航角。
  14. 根据权利要求8所述的装置，其特征在于，所述GPS重对准偏差角模块，包括：
    偏航角偏差角单元，用于根据所述GPS加速度信息以及所述IMU加速度信息，计算偏航角偏差角；
    GPS重对准偏差角单元，用于：
    根据所述GPS速度信息以及所述GPS加速度信息，分别计算GPS速度信号质量以及GPS加速度信号质量；
    对所述GPS速度信号质量以及GPS加速度信号质量进行求积，生成GPS加速度修正权重系数；
    判断所述GPS加速度修正权重系数是否大于预设GPS重对准阈值，若是， 则控制所述飞行器的重对准标志位输出高电平信号，若否，则控制所述重对准标志位输出低电平信号；
    若所述重对准标志位的信号存在上升沿，则生成重对准脉冲信号，；以及
    将所述重对准脉冲信号输入所述飞行器的使能模块，当所述使能模块检测接收到所述重对准脉冲存在上升沿信号时，则将所述偏航角偏差角作为所述GPS重对准偏差角。
  15. 一种飞行器，其特征在于，包括：
    至少一个处理器；以及，
    与所述至少一个处理器通信连接的存储器；其中，
    所述存储器存储有可被所述至少一个处理器执行的指令，所述指令被所述至少一个处理器执行，以使所述至少一个处理器能够执行权利要求1-7任一项所述的方法。
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