CN115950419A - Combined navigation method, device and system for subminiature unmanned aerial vehicle - Google Patents

Abstract

The application discloses a combined navigation method, a device and a system for a subminiature unmanned aerial vehicle. Wherein, the method comprises the following steps: acquiring acceleration information and angular velocity information of the unmanned aerial vehicle through an inertial measurement unit, and performing attitude calculation based on the acceleration information and the angular velocity information to obtain pose information of the unmanned aerial vehicle; receiving satellite differential correction information sent by a ground station at a preset frequency, and correcting the pose information by using the satellite differential correction information; navigating the unmanned aerial vehicle based on the corrected pose information. The method and the device solve the technical problems of discrete inertia and inaccurate satellite navigation in the related technology.

Description

Combined navigation method, device and system for ultra-small unmanned aerial vehicle

Technical Field

The present application relates to the field of navigation, and more specifically, to a combined navigation method, device and system for ultra-small unmanned aerial vehicles.

Background Art

The time interval between two positionings of modern satellite navigation systems is long, and each positioning requires more than 10 minutes of tracking. It cannot provide drone location information continuously, so it is often combined with an inertial navigation system.

The use of satellite global positioning system can enable drones to obtain real-time position and speed information in any area. However, when drones make violent maneuvers or when the signal-to-noise ratio of the satellite global positioning system is low, the navigation accuracy will be greatly reduced. When the accuracy of inertial navigation equipment and satellite navigation equipment is determined, it is difficult to improve the positioning accuracy of inertial/satellite combined navigation with a discrete structure.

To address the above-mentioned problems, no effective solution has been proposed yet.

Summary of the invention

The embodiments of the present application provide a combined navigation method, device and system for an ultra-small unmanned aerial vehicle, so as to at least solve the technical problem of inaccurate discrete inertial and satellite navigation.

According to one aspect of an embodiment of the present application, a combined navigation method for an ultra-small unmanned aerial vehicle is provided, comprising: acquiring acceleration information and angular velocity information of the unmanned aerial vehicle through a strapdown inertial measurement unit, and performing attitude calculation based on the acceleration and angular velocity information to obtain the position and posture information of the unmanned aerial vehicle; receiving satellite differential correction information sent by a ground station at a preset frequency, and using the satellite differential correction information to correct the position and posture information; and navigating the unmanned aerial vehicle based on the corrected position and posture information.

According to another aspect of an embodiment of the present application, a combined navigation device for an ultra-small unmanned aerial vehicle is also provided, including: a strapdown solution module, configured to obtain acceleration information and angular velocity information of the unmanned aerial vehicle through a strapdown inertial measurement unit, and perform attitude solution based on the acceleration and angular velocity information to obtain the posture information of the unmanned aerial vehicle; a correction module, configured to receive satellite differential correction information sent by a ground station at a preset frequency, and use the satellite differential correction information to correct the posture information; a navigation module, configured to navigate the unmanned aerial vehicle based on the corrected posture information.

According to another aspect of the embodiment of the present application, a combined navigation system for an ultra-small unmanned aerial vehicle is also provided, including: a strapdown inertial measurement unit, configured to obtain acceleration information and angular velocity information of the unmanned aerial vehicle; a satellite differential positioning system, configured to send satellite differential correction information at a preset frequency; the strapdown inertial measurement unit includes a navigation computer, and the navigation computer is the combined navigation device for the ultra-small unmanned aerial vehicle as described above.

In an embodiment of the present application, satellite differential correction information sent by a ground station at a preset frequency is received, and the satellite differential correction information is used to correct the posture information; based on the corrected posture information, the UAV is navigated, thereby solving the technical problem of inaccurate discrete inertial and satellite navigation.

BRIEF DESCRIPTION OF THE DRAWINGS

The drawings described herein are used to provide a further understanding of the present application and constitute a part of the present application. The illustrative embodiments of the present application and their descriptions are used to explain the present application and do not constitute an improper limitation on the present application. In the drawings:

FIG1 is a flow chart of a combined navigation method for an ultra-small UAV according to an embodiment of the present application;

FIG2A is a flow chart of another combined navigation method for ultra-small UAV according to an embodiment of the present application;

FIG2B is a schematic diagram of a geographic coordinate system according to an embodiment of the present application;

FIG2C is a schematic diagram of a loosely coupled structure according to an embodiment of the present application;

FIG3 is a flow chart of a Kalman filter error correction method according to an embodiment of the present application;

4 is a schematic diagram of the composition of a strapdown inertial measurement unit system according to an embodiment of the present application;

5 is a hardware design circuit diagram of an inertial measurement assembly according to an embodiment of the present application;

6 is a hardware design circuit diagram of a navigation computer according to an embodiment of the present application;

7 is a hardware design circuit diagram of a secondary power supply module according to an embodiment of the present application;

FIG8 is a hardware design circuit diagram of a satellite receiver according to an embodiment of the present application.

DETAILED DESCRIPTION

In order to enable those skilled in the art to better understand the solution of the present application, the technical solution in the embodiments of the present application will be clearly and completely described below in conjunction with the drawings in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without creative work should fall within the scope of protection of the present application.

It should be noted that the terms "first", "second", etc. in the specification and claims of the present application and the above-mentioned drawings are used to distinguish similar objects, and are not necessarily used to describe a specific order or sequence. It should be understood that the data used in this way can be interchangeable where appropriate, so that the embodiments of the present application described herein can be implemented in an order other than those illustrated or described herein. In addition, the terms "including" and "having" and any of their variations are intended to cover non-exclusive inclusions, for example, a process, method, system, product or device comprising a series of steps or units is not necessarily limited to those steps or units clearly listed, but may include other steps or units that are not clearly listed or inherent to these processes, methods, products or devices.

According to an embodiment of the present application, a combined navigation method for an ultra-small UAV is provided, as shown in FIG1 , the method comprising:

Step S102 obtains acceleration information and angular velocity information of the UAV through a strapdown inertial measurement unit, and performs attitude calculation based on the acceleration and angular velocity information to obtain the position and posture information of the UAV.

First, after obtaining the acceleration information and angular velocity information of the UAV through the strapdown inertial measurement unit, the attitude is solved using the quaternion method based on the angular velocity information to obtain the attitude angle of the carrier. Specifically, the posture information is described using quaternions, and the relationship between the conversion relationship from the missile body coordinate system to the geographic coordinate system and the quaternion is determined based on the description; the quaternion describing the posture information is solved based on the determined relationship between the conversion relationship and the quaternion using the angular velocity increment; and the attitude is solved based on the solved quaternion to obtain the attitude angle of the carrier.

Then, based on the acceleration information, a strapdown solution is performed to obtain the speed information and position information of the carrier. Specifically, based on the rate of change of the ground speed in the inertial coordinate system, the specific force vector in the acceleration information, and the angular velocity of the earth's rotation, the rate of change of the ground speed in the navigation coordinate system is determined; based on the rate of change of the ground speed in the navigation coordinate system, a navigation equation is determined, and based on the navigation equation, the components of the speed of the carrier in the true north, east and local vertical directions in the navigation coordinate system are determined. For example, based on the rate of change of the longitude and latitude of the carrier, the angular rate of rotation of the geographic coordinate system relative to the earth's fixed coordinate system is determined; based on the angular rate of rotation, the local gravity vector of the carrier is determined; based on the angular deviation of the local gravity vector direction relative to the local vertical direction caused by gravity anomaly, the current latitude, the current longitude and the current height from the earth's surface, and the local gravity vector, the components of the speed of the carrier in the navigation coordinate system along the true north, east and local vertical directions are determined.

Step S104, receiving satellite differential correction information sent by the ground station at a preset frequency, and using the satellite differential correction information to correct the posture information.

First, based on the satellite differential correction information, the Kalman filter is used to estimate the errors of position, velocity and attitude. Specifically, the position, velocity and attitude of the navigation solution from the navigation satellite system are input into the Kalman filter as measurement information as initial estimation values; in the prediction stage, the Kalman filter linearizes the initial estimation values and determines the error covariance based on the initial estimation values after linearization; determines the Kalman gain based on the error covariance, and redetermines the error covariance based on the Kalman gain; estimates the errors of the position, velocity and attitude of the strapdown solution based on the redetermined error covariance.

Then, the pose information is corrected based on the estimated error.

Step S106: Navigate the UAV based on the corrected position information.

In the embodiment of the present application, satellite differential correction information sent by a ground station at a preset frequency is received, and the satellite differential correction information is used to correct the posture information; based on the corrected posture information, the drone is navigated, thereby solving the technical problem of inaccurate discrete inertial and satellite navigation.

Example 2

According to an embodiment of the present application, another combined navigation method for an ultra-small UAV is provided, as shown in FIG2A , the method comprising:

Step S202: coordinate system design.

The navigation coordinate system adopts the north-sky-east coordinate system. The origin of the coordinate is taken at the center of mass of the aircraft, the ox _n axis points to the North Pole, the oy _n axis is perpendicular to the local horizontal plane and points to the sky, and the oz _n axis points to the east in the right-hand rule with the ox _n axis and the oy _n axis, as shown in Figure 2B.

The origin of the carrier coordinate system is at the center of mass of the aircraft, the ox _b axis is along the longitudinal axis of the aircraft, pointing to the head of the aircraft is positive, the oy _b axis is in the longitudinal symmetry plane of the aircraft, perpendicular to the ox _b axis and pointing upwards is positive, the oz _b axis, the ox _b axis and the oy _b axis are in the right-hand rule, and are fixed to the carrier.

滚转角γ来直接描述弹体坐标系到地理坐标系的转换关系式为：Using positive Euler, the true heading angle ψ _N and pitch angle

The roll angle γ is used to directly describe the transformation relationship from the missile body coordinate system to the geographic coordinate system:

When using positive Euler, the pitch angle and roll angle of the geographic coordinate system and the carrier coordinate system are defined the same, and the relationship between the yaw angle and the true heading angle is as follows:

Wherein, A represents the angle deviation.

In this embodiment, the coordinate system is converted in the above manner, which can avoid errors in coordinate conversion, thereby making subsequent navigation solutions more accurate.

Step S204: posture calculation.

This implementation uses the quaternion method, using four elements to describe all posture information. The quaternion q can be expressed as:

q＝a+bi+cj+dk (1)

Among them, a, b, c, d are the real parts of the quaternion, and i, j, k are the imaginary units.

Suppose there is a vector v ^b in the b system, which is represented by v ⁿ in the n system, then:

Wherein q ^* =a-bi-cj-dk is the complex conjugate of the quaternion q.

与四元数之间的关系：From the above formula, we can get

Relationship with quaternions:

Using quaternion attitude solution, you need to solve the following equation:

表示四元数的微分形式，q表示四元数，角速度ω的分量为ω_x、ω_y、ω_z。式(4)也可用矩阵形式表示：where p = [0, ω ^T ],

represents the differential form of the quaternion, q represents the quaternion, and the components of the angular velocity ω are ω _x , ω _y , ω _z . Equation (4) can also be expressed in matrix form:

In the formula

The quaternion solution is calculated by the angle increment, and its solution can be expressed as:

Where q _k represents the quaternion at time k, and t _k represents the integral calculation time.

Substitute W and let

The final solution is:

T为系统采样周期，I为4x4的单位矩阵。In the formula,

T is the system sampling period, and I is the 4x4 identity matrix.

Therefore, the attitude angle can be obtained by solving the quaternion. The formula is as follows:

θ＝arcsin[-C ₃₁ ]＝arcsin[2(bd-ac)]

The matrix C in the formula is formula (3).

In this embodiment, the above-mentioned method is used to solve the attitude angle, so that the solved attitude angle can be more accurate, thereby providing accurate basic data for subsequent navigation.

Step S206, strapdown solution.

The navigation equation can be expressed as follows:

Among them, r represents the position vector, f represents the specific force, the velocity is obtained by the first integration, and the position is obtained by the second integration.

它相对于导航坐标系的变化率可通过其在惯性系下的变化率表示：Where v _n represents the navigation coordinate system speed. For a navigation system working in the local geographic coordinate system on the earth, the ground speed is expressed as

Its rate of change relative to the navigation coordinate system can be expressed by its rate of change in the inertial system:

v_e表地速，ω_en表示导航坐标系相对于地球坐标系的转动角速率，ω_ie表示地球坐标系相对于惯性坐标系的转动角速率，g₁＝g-ω_ie×[ω_ie×r]代入式(13)则有：in,

v _e represents the ground speed, ω _en represents the rotational angular velocity of the navigation coordinate system relative to the earth coordinate system, ω _ie represents the rotational angular velocity of the earth coordinate system relative to the inertial coordinate system, g ₁ =g-ω _ie ×[ω _ie ×r] Substituting into formula (13), we have:

The navigation equation can be expressed as follows:

其中，v_N，v_E，v_D分别表示沿真北、东向和当地垂线方向的速度，

表示地球坐标系相对于惯性坐标系的转动角速率、

表示导航坐标系相对于地球坐标系的转动角速率、

表示地速。The components of velocity along true north, east and local vertical are:

Where, _vN , _vE , and _vD represent the velocities along the true north, east, and local vertical directions, respectively.

represents the angular velocity of the Earth coordinate system relative to the inertial coordinate system,

Indicates the angular rate of rotation of the navigation coordinate system relative to the earth coordinate system,

Indicates ground speed.

fw is a set of specific force vectors measured by three accelerometers, decomposed into the local georeferenced coordinate system as:

f ⁿ = [f _N f _E f _D ] ^T (16)

是当地地理坐标系中地球的自转角速度：

is the angular velocity of the Earth's rotation in the local geographic coordinate system:

表示当地地理坐标系相对于地球固连坐标系的转动角速率，即转移速率，其值可以用经度和纬度的变化率表示如下：Among them, L0 means,

It represents the angular rate of rotation of the local geographic coordinate system relative to the earth's fixed coordinate system, that is, the transfer rate. Its value can be expressed by the rate of change of longitude and latitude as follows:

其中，L表示纬度，得：make

Where L represents latitude, we get:

In formula (19), _R0 is the radius of the earth; h is the height from the earth's surface.

是当地重力矢量，它由地球的质量引力(g)和地球转动产生的向心加速度(ω_ie×ω_ie×R)组成。因此，可以写成：

is the local gravity vector, which is composed of the mass gravitational force of the earth (g) and the centripetal acceleration caused by the rotation of the earth (ω _ie ×ω _ie ×R). Therefore, it can be written as:

Among them, Ω represents the angular rate of the earth's rotation, and the navigation equation can be expressed in the following component form:

Where: ξ, η are the first angle deviation and second angle deviation of the local gravity vector direction relative to the local vertical direction caused by gravity anomaly.

Latitude, longitude, and altitude above the Earth's surface are given by the following formulas:

表示纬度微分形式，

表示经度微分形式，

表示距地球表面高度微分形式。in,

represents the latitude differential form,

represents the differential form of longitude,

Represents the differential form of the height above the earth's surface.

In this embodiment, when calculating the speed information, not only the latitude, longitude and altitude from the earth's surface are taken into account, but also the angular deviation of the local gravity vector direction relative to the local vertical direction caused by gravity anomaly is introduced, so that the calculated speed information is more accurate.

Step S208, combined navigation.

The integrated navigation is composed of a micro strapdown inertial measurement unit and a mobile station satellite receiver (GNSS). The GNSS is responsible for resetting the accumulated error of the inertial navigation within a certain time period. The loose combination structure is adopted, and the estimated position, velocity and attitude errors are used to correct the inertial navigation solution. The specific architecture is shown in Figure 2C.

The loose combination structure of this embodiment is a series system, and the position and velocity output from the navigation solution of GNSS are input into the Kalman filter as measurement information, and the Kalman filter is used to estimate the error of the strapdown solution. The specific process of the extended Kalman filter (EKF) method is shown in Figure 3, including the following steps: step S302, establishing a system model and a measurement model; step S304, inputting an initial estimate; step S306, performing linearization processing in the prediction stage to estimate the error covariance; step S308, calculating the Kalman gain, and re-estimating the error covariance based on the Kalman gain.

The main advantages of the loosely coupled integration provided by this embodiment are simplicity and redundancy, which can be used for any micro-inertial measurement unit (i.e., strapdown inertial measurement unit) and GNSS equipment, and is particularly suitable for improved algorithm applications. In the loosely coupled structure, there is an independent GNSS available navigation solution, which is outside the combined navigation solution. When the open-loop micro-inertial measurement unit correction is executed, there is also an independent micro-inertial measurement unit navigation solution, which supports basic parallel navigation solutions.

In the error model of the micro-inertial measurement unit in this embodiment, an error model based on a small misalignment angle is used to represent the influence of the error factor by a small disturbance, and a nonlinear differential equation of the error of the micro-inertial measurement unit is derived. According to the error state equations of the micro-inertial measurement unit and the GNSS, the state variables of the combined system are formed by combining the error state variables of the micro-inertial measurement unit and the GNSS, which can be expressed by formula (27).

表示系统状态变量微分形式，F(t)表示状态转移矩阵，X_(t)表示系统状态向量，B(t)表示控制矩阵，u(t)表示控制变量，G(t)表示系统噪声驱动矩阵，w(t)表示系统噪声矩阵。in,

represents the differential form of the system state variable, F(t) represents the state transfer matrix, X _(t) represents the system state vector, B(t) represents the control matrix, u(t) represents the control variable, G(t) represents the system noise driving matrix, and w(t) represents the system noise matrix.

The errors of the vehicle-mounted micro-inertial measurement unit include the roll angle deviation δα in the X-axis direction, the pitch angle deviation δ _β in the Y-axis direction, and the yaw angle deviation δ _γ in the Z-axis direction. The three velocity errors δ _VN , δ _VE , δ _VD and the three position errors δL, δl, δh, δf _x δf _y δf _z are the acceleration deviations of the X-axis, Y-axis and Z-axis (m/s ² ), and δ _ωx δ _ωy δ _ωz are the gyro deviations of the X-axis, Y-axis and Z-axis (°/s), respectively, which are expressed by the following formula (28):

The state transfer matrix F _(t) is obtained by simplifying the error model as (29):

Among them, _FS-9×9 represents the 9×9 dimensional inertial navigation system error, _FDCM-6×6 represents the 6×6 dimensional direction cosine matrix, and DCM _3×3 represents the 3×3 dimensional direction cosine matrix.

The attitude angle error is expressed by equation (31) through the direction cosine matrix from the vehicle reference system to the navigation reference system and the gyro error.

Where C represents the direction cosine matrix element.

The elements of the F matrix are expressed as follows:

Where R is the radius of the Earth, V _N , _VE , and V _D are the north, east, and ground velocities indicated by the micro-inertial measurement unit, f _n , f _E , and f _D are the north, east, and ground specific force measurements of the MIMU, L is the latitude output by the micro-inertial measurement unit, Ω is the angular velocity of the Earth's rotation, and the height is negligible relative to the radius of the Earth, so the height is omitted for simplification.

The measurement model of the combined system is embodied in the matrix H, which links the measurement model with the filtering state. The observation information of this embodiment is the speed, position and attitude of the GNSS system. The attitude includes the yaw angle α and the pitch angle β. GNSS cannot provide the roll angle. GNS provides the current position information of the vehicle, with the geographic reference system as the navigation reference system, and the longitude _LG , latitude _lG , and altitude _hG as parameters. GNSS provides vehicle _speed information, using the Cartesian speed [ _{VNVED ]} as the GNSS speed measurement.

The position provided by GNSS is expressed as the sum of the true value and the error in the geographic context, as expressed in (44):

Wherein, L _G represents the latitude in the geographic coordinate system, l _G represents the longitude in the geographic coordinate system, h _G represents the altitude in the geographic coordinate system, L _N represents the latitude in the navigation coordinate system, N _N represents the north position error of the satellite receiving meter, N _E represents the east position error of the satellite receiving meter, R _N represents the lateral curvature radius, h _N represents the altitude in the navigation coordinate system, and Nh represents the celestial position error of the satellite receiving meter. δL, δl, δh are the position error measurements of GNSS relative to the navigation reference system, as expressed in (45):

Wherein, _LI represents the latitude in the inertial coordinate system, _IL represents the longitude in the inertial coordinate system, _LG represents the longitude in the geographic coordinate system, _RM represents the radius of curvature of the meridian plane, _RN represents the radius of transverse curvature, _PGN represents the north position in the geographic coordinate system, _PGE represents the east position in the geographic coordinate system, and _PGD represents the ground position in the geographic coordinate system. The velocity measurement information of the micro-inertial measurement unit is the sum of the true value of the navigation system and the corresponding velocity error, and the GNSS velocity measurement information is expressed as the difference between the true value in the N system and the corresponding velocity measurement error. _MN , _ME , _MD are the components of the GNSS velocity measurement error term on the north-east coordinate axis. The velocity measurement is shown in formula (46):

Wherein, V _IN represents the north velocity in the inertial coordinate system, V _GN represents the north velocity in the geographic coordinate system, V _IE represents the east velocity in the inertial coordinate system, V _GE represents the east velocity in the geographic coordinate system, V _ID represents the ground velocity in the inertial coordinate system, V _GD represents the ground velocity in the geographic coordinate system, _MN represents the north error velocity of the satellite receiver, _ME represents the east error velocity of the satellite receiver, _MD represents the ground error velocity of the satellite receiver, δ _VN represents the north velocity error component, δ _VE represents the ground velocity error component, δ _VD represents the ground velocity error component, HV represents the measurement matrix, X(t) represents the system state vector, and Vv(t) represents the measurement noise vector.

速度的标准差由式(47)表示。In the above formula

The standard deviation of the velocity is expressed by equation (47).

表示位置标准差微分形式。Among them, HDOP represents the position precision factor,

Represents the differential form of the location standard deviation.

The attitude measurement of the micro-inertial measurement unit is expressed as the true value and the corresponding attitude error in the navigation system, and the attitude measurement of the GNSS is expressed as the true value and the corresponding attitude error in the N system. The heading angle and pitch angle measurement equations are (48):

表示姿态量测方程，

表示惯性坐标系下航向角，

表示地理坐标系下航向角，

表示惯性坐标系下俯仰角，

表示地理坐标系下俯仰角，

表示姿态观测矩阵，X(t)表示系统状态向量，

表示量测噪声矩阵。in,

represents the attitude measurement equation,

represents the heading angle in the inertial coordinate system,

Indicates the heading angle in the geographic coordinate system.

represents the pitch angle in the inertial coordinate system,

Indicates the elevation angle in the geographic coordinate system.

represents the attitude observation matrix, X(t) represents the system state vector,

represents the measurement noise matrix.

The position, velocity and attitude measurements (45)(46)(48) are combined and expressed as shown in formula 49:

表示姿态量测方程，H_V表示速度量测矩阵，H_P表示位置量测矩阵，

表示姿态量测矩阵，v_p表示卫星接收计位置测量误差，v_v表示卫星接收计速度测量误差，X_k+1表示k+1时刻系统估计状态，k表示当前导航解算时刻，V_k+1表示k+1时刻量测噪声序列，H表示量测矩阵。

为GNSS观测噪声。Wherein, z represents the measurement equation, Z _p(t) represents the position measurement equation, and Z _v(t) represents the velocity measurement equation.

represents the attitude measurement equation, H _V represents the velocity measurement matrix, _HP represents the position measurement matrix,

represents the attitude measurement matrix, v _p represents the satellite receiver position measurement error, v _v represents the satellite receiver velocity measurement error, X _k+1 represents the system estimated state at time k+1, k represents the current navigation solution time, V _k+1 represents the measurement noise sequence at time k+1, and H represents the measurement matrix.

is the GNSS observation noise.

In this embodiment, the measurement equation of the combined system is established by using speed, position and posture errors, which is more suitable for changes in the dynamic system than directly using state values.

Step S210: initial alignment.

The reference information measured by the satellite reference station: longitude, latitude, altitude, attitude angle, pitch angle and heading angle information. The relevant reference information is passed to the strapdown inertial measurement unit. The strapdown inertial measurement unit receives the reference information and uses the established EKF filter, combined with its own inertia and satellite data, to complete the transfer alignment under speed and position matching. This achieves the initial alignment of the drone end.

The present application solves the technical problems of large size and high cost of discrete inertial and satellite navigation systems in the prior art, and has the beneficial effects of small size and low cost.

It should be noted that, for the above-mentioned method embodiments, for the sake of simplicity, they are all expressed as a series of action combinations, but those skilled in the art should be aware that the present application is not limited by the order of the actions described, because according to the present application, certain steps can be performed in other orders or simultaneously. Secondly, those skilled in the art should also be aware that the embodiments described in the specification are all preferred embodiments, and the actions and modules involved are not necessarily required by the present application.

Through the description of the above implementation methods, those skilled in the art can clearly understand that the method according to the above embodiment can be implemented by means of software plus a necessary general hardware platform, and of course by hardware, but in many cases the former is a better implementation method. Based on such an understanding, the technical solution of the present application, or the part that contributes to the prior art, can be embodied in the form of a software product, which is stored in a storage medium (such as ROM/RAM, magnetic disk, optical disk), and includes a number of instructions for a terminal device (which can be a mobile phone, computer, server, or network device, etc.) to execute the methods described in each embodiment of the present application.

Example 3

According to an embodiment of the present application, a combined navigation system for an ultra-small unmanned aerial vehicle is provided, comprising a strapdown inertial measurement unit and a satellite differential positioning system.

As shown in FIG. 4 , the strapdown inertial measurement unit includes an inertial measurement assembly 42 , a navigation computer 44 , a secondary power supply module 40 , a crystal oscillator 48 , an external connector 46 , and a mobile station satellite receiver 49 .

1) Inertial Measurement Unit

The inertial measurement unit 42 mainly acquires the acceleration and angular velocity information of the carrier, and uploads the data to the navigation computer 44 through a four-wire SPI interface. The inertial measurement unit 42 has a built-in 3-axis gyroscope and a 3-axis accelerometer, and its specific indicators are shown in Table 1.

Table 1

The hardware circuit design is shown in Figure 5. The chip is connected to the navigation computer 44 via a four-wire SPI interface, the power supply is 3.3V, and is equipped with an independent bypass capacitor for filtering.

2) Navigation Computer

The navigation computer 44 is mainly used for solving the navigation algorithm. In this embodiment, the operating main frequency of the navigation computer 44 is set to 160MHz, equipped with a floating-point calculation unit, 2MB of its own program storage space, and 800KB of memory, which can meet the navigation solution requirement of a cycle of 5ms. In addition, the navigation computer 44 also has abundant DMA channels for off-chip chip operations and logic operations. The specific resource allocation is shown in Table 2 below.

Serial number Resource Name Related equipment Remark 1 USART0 External Communications Navigation data output TTL@3.3V 2 USART1 External Communications Navigation data output RS-422@3.3V 3 SPI0 External Communications Four-wire SPI, speed 25MHz 4 SPI1 Inertial Measurement Unit Four-wire SPI

Table 2

The specific hardware design is shown in Figure 6. The minimum system of the chip includes: startup capacitor, reset circuit and clock source, etc.

3) Secondary power module

The secondary power supply module 40 is shown in FIG7 , and is mainly used to adjust and match the internal power supply. In this embodiment, two independent chips are placed for power output, one for providing power to the mobile station satellite receiver 49, and the other for providing power to other digital circuits on the board. In order to improve the power supply performance and reduce interference between devices, a π-type filter design is performed on both the input and output during the power supply design process.

4) Crystal oscillator

In this embodiment, the crystal oscillator 48 has the characteristics of good temperature characteristics and high reliability, and its stability is ≤20ppm.

5) External connector

In this embodiment, the external connector 46 has surface mount pins with a pitch of 1.0 mm and is gold plated, so that the contact impedance is reduced while ensuring stability.

The satellite differential positioning system includes a satellite differential positioning system, a mobile station satellite navigation system, a mobile station satellite machine, a mobile station satellite receiving antenna, a mobile station satellite receiving feeder, a base station satellite navigation system, a base station satellite receiver, a base station satellite receiving antenna, a base station satellite receiving antenna, a base station satellite receiving feeder, a base station satellite receiving feeder, a base station satellite antenna connector, a base station satellite antenna connector, a secondary power supply module, an interface conversion module and a data fusion unit.

1) Mobile station satellite navigation system

The specific indicators of the mobile station satellite navigation system are shown in Table 3.

Table 3

The specific hardware design circuit is shown in Figure 8.

2) Base station satellite navigation system

The base station satellite navigation system has both positioning and attitude measurement functions, and cooperates with the high-precision inertial measurement unit to achieve high-precision base attitude acquisition. The main technical indicators are shown in Table 4.

Table 4

The satellite receiver is used to receive high-precision satellite positioning data to achieve high-precision aviation-grade positioning accuracy. Combined with the high-precision inertial measurement unit and data fusion unit, effective data fusion is achieved to complete the final initial measurement of position and attitude.

Example 4

According to an embodiment of the present application, a combined navigation device for an ultra-small unmanned aerial vehicle is also provided, including: a strapdown solution module, configured to obtain acceleration information and angular velocity information of the unmanned aerial vehicle through an inertial measurement unit, and perform attitude solution based on the acceleration and angular velocity information to obtain the posture information of the unmanned aerial vehicle; a correction module, configured to receive satellite differential correction information sent by a ground station at a preset frequency, and use the satellite differential correction information to correct the posture information; a navigation module, configured to navigate the unmanned aerial vehicle based on the corrected posture information.

Optionally, the specific examples in this embodiment may refer to the examples described in the above-mentioned Embodiment 1 and Embodiment 2, and this embodiment will not be described in detail here.

Example 5

The embodiment of the present application further provides a storage medium, which is configured to store program codes for executing the methods in the above embodiments 1 and 2.

Optionally, in this embodiment, the above-mentioned storage medium may include but is not limited to: a USB flash drive, a read-only memory (ROM), a random access memory (RAM), a mobile hard disk, a magnetic disk or an optical disk, and other media that can store program codes.

The serial numbers of the above-mentioned embodiments of the present application are for description only and do not represent the advantages or disadvantages of the embodiments.

If the integrated units in the above embodiments are implemented in the form of software functional units and sold or used as independent products, they can be stored in the above computer-readable storage medium. Based on this understanding, the technical solution of the present application, or the part that contributes to the prior art, or all or part of the technical solution can be embodied in the form of a software product, which is stored in a storage medium and includes several instructions for enabling one or more computer devices (which may be personal computers, servers, or network devices, etc.) to execute all or part of the steps of the methods described in the various embodiments of the present application.

In the above embodiments of the present application, the description of each embodiment has its own emphasis. For parts that are not described in detail in a certain embodiment, please refer to the relevant descriptions of other embodiments.

In the several embodiments provided in the present application, it should be understood that the disclosed client can be implemented in other ways. Among them, the device embodiments described above are only schematic. For example, the division of the units is only a logical function division. There may be other division methods in actual implementation. For example, multiple units or components can be combined or integrated into another system, or some features can be ignored or not executed. Another point is that the mutual coupling or direct coupling or communication connection shown or discussed can be through some interfaces, indirect coupling or communication connection of units or modules, which can be electrical or other forms.

The units described as separate components may or may not be physically separated, and the components shown as units may or may not be physical units, that is, they may be located in one place or distributed on multiple network units. Some or all of the units may be selected according to actual needs to achieve the purpose of the solution of this embodiment.

In addition, each functional unit in each embodiment of the present application may be integrated into one processing unit, or each unit may exist physically separately, or two or more units may be integrated into one unit. The above-mentioned integrated unit may be implemented in the form of hardware or in the form of software functional units.

The above is only a preferred implementation of the present application. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the principles of the present application. These improvements and modifications should also be regarded as the scope of protection of the present application.

Claims (10)

1. An integrated navigation method for a subminiature unmanned aerial vehicle, comprising:

acquiring acceleration information and angular velocity information of the unmanned aerial vehicle through an inertial measurement unit, and performing attitude calculation based on the acceleration information and the angular velocity information to obtain pose information of the unmanned aerial vehicle;

receiving satellite differential correction information sent by a ground station at a preset frequency, and correcting the pose information by using the satellite differential correction information;

navigating the drone based on the revised pose information.

2. The method of claim 1, wherein performing an attitude solution based on the acceleration information and angular velocity information comprises:

carrying out attitude calculation by utilizing a quaternion method based on the angular velocity information to obtain an attitude angle of the unmanned aerial vehicle;

and performing strapdown calculation based on the acceleration information to obtain the speed information and the position information of the unmanned aerial vehicle.

3. The method of claim 2, wherein performing attitude calculation by using a quaternion method based on the angular velocity information to obtain an attitude angle of the drone comprises:

describing the pose information by using quaternions, and determining a relation between a conversion relation from a projectile coordinate system to a geographic coordinate system and the quaternions based on the description;

solving the quaternion describing the pose information based on the determined relationship between the conversion relation and the quaternion using an angular velocity increment;

and carrying out attitude calculation based on the solved quaternion to obtain an attitude angle of the unmanned aerial vehicle.

4. The method according to claim 2, wherein performing strapdown solution based on the acceleration information to obtain the velocity information of the drone comprises:

determining the change rate of the ground speed in a navigation coordinate system based on the change rate of the ground speed in an inertial coordinate system, the specific force vector in the acceleration information and the rotation angular velocity of the earth;

and determining a navigation equation based on the change rate of the ground speed in a navigation coordinate system, and determining components of the speed of the unmanned aerial vehicle along the true north direction, the east direction and the local vertical direction in the navigation coordinate system based on the navigation equation.

5. The method of claim 4, wherein determining the components of the velocity of the drone along true north, east, and local vertical directions in the navigational coordinate system based on the navigational equations comprises:

determining the rotation angular rate of a geographic coordinate system relative to an earth fixed connection coordinate system based on the change rate of the longitude and the latitude of the unmanned aerial vehicle;

determining a local gravity vector of the drone based on the angular rate of rotation;

determining components of the velocity of the drone in the navigation coordinate system along true north, east and local vertical directions based on an angular deviation of a local gravity vector direction relative to a local vertical direction due to a gravity anomaly, a current latitude, a current longitude and a current altitude from the earth's surface, and the local gravity vector.

6. The method according to claim 1, wherein correcting the pose information using the satellite differential correction information comprises:

estimating errors of position, velocity and attitude with a kalman filter based on the satellite difference correction information;

correcting the pose information based on the estimated error.

7. The method of claim 1, wherein estimating errors for position, velocity, and attitude using a kalman filter comprises:

inputting the position, the speed and the attitude of a navigation solution from a navigation satellite system into the Kalman filter as measurement information to serve as an initial estimation value;

in a prediction stage, the Kalman filter carries out linearization processing on the initial estimation value and determines an error covariance based on the initial estimation value after linearization processing;

determining a Kalman gain based on the error covariance, and re-determining the error covariance based on the Kalman gain;

estimating errors of the position, velocity, and attitude of the strapdown solution based on the re-determined error covariance.

8. An integrated navigation device for a subminiature unmanned aerial vehicle, comprising:

the strapdown resolving module is configured to acquire acceleration information and angular velocity information of the unmanned aerial vehicle through an inertial measurement unit, and perform attitude resolving based on the acceleration information and the angular velocity information to obtain pose information of the unmanned aerial vehicle;

the correction module is configured to receive satellite differential correction information sent by a ground station at a preset frequency, and correct the pose information by using the satellite differential correction information;

a navigation module configured to navigate the drone based on the pose information after the correction.

9. An integrated navigation system for a subminiature unmanned aerial vehicle, comprising:

a strapdown inertial measurement unit configured to acquire acceleration information and angular velocity information of the drone;

a satellite differential positioning system configured to transmit satellite differential correction information at a preset frequency;

wherein the strapdown inertial measurement unit includes a navigation computer, the navigation computer being the integrated navigation device for a subminiature unmanned aerial vehicle according to claim 8.

10. A computer-readable storage medium, on which a program is stored, which, when executed, causes a computer to carry out the method according to any one of claims 1 to 7.

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