CN115950419A - Combined navigation method, device and system for subminiature unmanned aerial vehicle - Google Patents
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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
Technical Field
The application relates to the field of navigation, in particular to a combined navigation method, device and system for a subminiature unmanned aerial vehicle.
Background
Modern satellite navigation systems have long time intervals between two positioning, require more than 10 minutes of tracking for each positioning, cannot continuously provide unmanned aerial vehicle position information, and therefore are often combined with an inertial navigation system.
The satellite global positioning system is adopted to enable the unmanned aerial vehicle to obtain position and speed information in real time in any area. However, when the unmanned aerial vehicle is in violent maneuvering operation or when the signal-to-noise ratio of the satellite global positioning system is low, the navigation precision is greatly reduced. Under the condition that the precision of the inertial navigation equipment and the precision of the satellite navigation equipment are determined, the positioning precision of the inertial/satellite combined navigation is difficult to improve by the discrete structure.
In view of the above problems, no effective solution has been proposed.
Disclosure of Invention
The embodiment of the application provides a combined navigation method, a device and a system for a subminiature unmanned aerial vehicle, which are used for at least solving the technical problems of discrete inertia and inaccurate satellite navigation.
According to an aspect of an embodiment of the present application, there is provided a combined navigation method for a subminiature unmanned aerial vehicle, including: 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 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.
According to another aspect of the embodiments of the present application, there is also provided a combined navigation device for a subminiature unmanned aerial vehicle, including: the strapdown calculating module is configured to acquire acceleration information and angular velocity information of the unmanned aerial vehicle through a strapdown inertial measurement unit, and perform attitude calculation 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.
According to still another aspect of the embodiments of the present application, there is also provided a combined navigation system for a subminiature unmanned aerial vehicle, including: a strapdown inertial measurement unit configured to acquire acceleration information and angular velocity information of the unmanned aerial vehicle; a satellite differential positioning system configured to transmit satellite differential correction information at a preset frequency; the strapdown inertial measurement unit comprises a navigation computer, and the navigation computer is the combined navigation device for the subminiature unmanned aerial vehicle.
In the embodiment of the application, satellite differential correction information sent by a ground station at a preset frequency is received, and the pose information is corrected by using the satellite differential correction information; and navigating the unmanned aerial vehicle based on the corrected pose information, so that the technical problems of discrete inertia and inaccurate satellite navigation are solved.
Drawings
The accompanying drawings, which are included to provide a further understanding of the application and are incorporated in and constitute a part of this application, illustrate embodiment(s) of the application and together with the description serve to explain the application and not to limit the application. In the drawings:
fig. 1 is a flowchart of a combined navigation method for a subminiature drone according to an embodiment of the present application;
fig. 2A is a flowchart of another integrated navigation method for a subminiature drone according to an embodiment of the present application;
FIG. 2B is a schematic diagram of a geographic coordinate system according to an embodiment of the present application;
FIG. 2C is a schematic diagram of a loosely coupled structure according to an embodiment of the present application;
FIG. 3 is a flow diagram of a Kalman filtering error correction method in accordance with an embodiment of the present application;
FIG. 4 is a schematic diagram of a strapdown inertial measurement unit system composition according to an embodiment of the present application;
FIG. 5 is a circuit diagram of a hardware design of an inertial measurement unit according to an embodiment of the present application;
FIG. 6 is a circuit diagram of a hardware design of a navigation computer according to an embodiment of the present application;
fig. 7 is a circuit diagram of a hardware design of a secondary power supply module according to an embodiment of the present application;
fig. 8 is a circuit diagram of a hardware design of a satellite receiver according to an embodiment of the present application.
Detailed Description
In order to make the technical solutions of the present application better understood by those skilled in the art, the technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the drawings in the embodiments of the present application, and it is obvious that the described embodiments are only some embodiments of the present application, and not all embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present application.
It should be noted that the terms "first," "second," and the like in the description and claims of this application and in the drawings described above are used for distinguishing between similar elements and not necessarily for describing a particular sequential or chronological order. It is to be understood that the data so used is interchangeable under appropriate circumstances such that the embodiments of the application described herein are capable of operation in sequences other than those illustrated or described herein. Furthermore, the terms "comprises," "comprising," and "having," and any variations thereof, are intended to cover a non-exclusive inclusion, such that a process, method, system, article, or apparatus that comprises a list of steps or elements is not necessarily limited to those steps or elements expressly listed, but may include other steps or elements not expressly listed or inherent to such process, method, article, or apparatus.
According to an embodiment of the present application, there is provided a combined navigation method for a subminiature unmanned aerial vehicle, as shown in fig. 1, the method including:
step S102, acceleration information and angular velocity information of the unmanned aerial vehicle are obtained through a strapdown inertial measurement unit, and attitude calculation is carried out based on the acceleration information and the angular velocity information, so that pose information of the unmanned aerial vehicle is obtained.
Firstly, acquiring acceleration information and angular velocity information of the unmanned aerial vehicle through a strapdown inertial measurement unit, and then performing attitude calculation by using a quaternion method based on the angular velocity information to obtain an attitude angle of the carrier. Specifically, the pose information is described by using quaternions, and the relation between a conversion relation from a projectile coordinate system to a geographic coordinate system and the quaternions is determined 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 carrier.
And then, performing strapdown calculation based on the acceleration information to obtain the speed information and the position information of the carrier. Specifically, the change rate of the ground speed in a navigation coordinate system is determined based on the change rate of the ground speed in an inertial coordinate system, a 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 carrier in the navigation coordinate system along the true north direction, the east direction and the local vertical direction based on the navigation equation. Determining a rotation angular rate of the geographic coordinate system relative to the earth-bound coordinate system, for example, based on a rate of change of longitude and latitude of the carrier; determining a local gravity vector of the carrier based on the angular rate of rotation; determining components of the velocity of the vehicle in the navigational coordinate system in true north, east and local vertical directions based on an angular deviation of the local gravity vector direction from the local vertical direction due to the gravity anomaly, the current latitude, the current longitude and the current altitude from the earth's surface, and the local gravity vector.
And step S104, receiving satellite differential correction information sent by the ground station at a preset frequency, and correcting the pose information by using the satellite differential correction information.
First, errors of position, velocity, and attitude are estimated using a kalman filter based on the satellite difference correction information. Specifically, the position, the speed and the attitude of a navigation solution from a navigation satellite system are input to the kalman filter as measurement information as initial estimation values; 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.
Then, the pose information is corrected based on the estimated error.
And S106, navigating the unmanned aerial vehicle based on the corrected pose information.
In the embodiment of the application, satellite differential correction information sent by a ground station at a preset frequency is received, and the pose information is corrected by using the satellite differential correction information; navigating the unmanned aerial vehicle based on the corrected pose information, thereby solving the technical problems of discrete inertia and inaccurate satellite navigation
Example 2
According to an embodiment of the present application, there is also provided another combined navigation method for a subminiature unmanned aerial vehicle, as shown in fig. 2A, the method including:
and step S202, designing a coordinate system.
The navigation coordinate system adopts a north-sky-east coordinate system. The origin of coordinates is taken at the aircraft centroid, ox n Axial north pole, oy n The axis is directed perpendicular to the local horizontal plane to the sky, oz n The axis points to east and ox n Axle, oy n The axis is in the right hand rule, as shown in FIG. 2B.
The origin of coordinates of the carrier coordinate system is at the center of mass, ox, of the aircraft b Axis is positive, oy, along the longitudinal axis of the aircraft, pointing towards the aircraft head b Axial in the aircraft longitudinalInto the symmetry plane, perpendicular to ox b Axial pointing up is positive, oz b Shaft and ox b Axis and oy b The shaft is in right hand rule and is fixedly connected with the carrier.
True heading angle psi available using positive Euler N And a pitch angle
The roll angle γ directly describes the conversion relation from the projectile coordinate system to the geographic coordinate system as follows:
when the pitch angle and the roll angle of the real-time coordinate system and the carrier coordinate system are defined the same using positive euler, the relation between the yaw angle and the true heading angle is as follows:
where a represents an angular deviation.
In the embodiment, the coordinate system is converted by adopting the above mode, so that errors in coordinate conversion can be avoided, and subsequent navigation calculation is more accurate.
And step S204, resolving the attitude.
In the implementation, a quaternion method is adopted, all pose information is described by four elements, and the quaternion q can be expressed as:
q＝a+bi+cj+dk (1)
wherein a, b, c and d are real parts of quaternions, and i, j and k are imaginary units.
Provided with a b-system middle vector v b Which is represented by v in the n system n And then:
in the formula q * = a-bi-cj-dk is the complex conjugate of the quaternion q.
Can be obtained from the above formula
Relationship to quaternion:
solving by adopting quaternion attitude, the following equation needs to be solved:
wherein p = [0, ω = T ]，
Representing a differential form of a quaternion, q representing a quaternion, the component of the angular velocity ω being ω x 、ω y 、ω z . Equation (4) can also be expressed in matrix form:
in the formula
A solution quaternion is computed from the angular increments, whose solution can be expressed as:
wherein q is k Quaternion, t, representing time k k The integration calculates the time instant.
Bring W into
The final solution can be:
in the formula,
t is the system sampling period, and I is the identity matrix of 4x 4.
The attitude angle can thus be obtained by solving the quaternion, the formula being:
θ＝arcsin[-C 31 ]＝arcsin[2(bd-ac)]
wherein the matrix C is formula (3).
In the embodiment, the attitude angle is calculated by adopting the method, so that the calculated attitude angle is more accurate, and accurate basic data is provided for subsequent navigation.
And step S206, strapdown resolving.
The navigation equation can be expressed in the form:
where r represents a position vector, f represents a specific force, and the velocity is obtained by first integration and the position is obtained by second integration.
Wherein v is n Representing the navigation coordinate system velocity. Navigation system working on earth in local geographical coordinate system, ground speed expressed as
Its rate of change relative to the navigational coordinate system can be represented by its rate of change under the inertial system: />
Wherein,
v e surface speed, omega en Representing the angular rate of rotation, omega, of the navigational coordinate system relative to the terrestrial coordinate system ie Representing the angular rate of rotation, g, of the terrestrial coordinate system relative to the inertial coordinate system 1 ＝g-ω ie ×[ω ie ×r]The substitution formula (13) includes:
the navigation equation can be expressed in the form:
the components of the speed along the true north, east and local vertical directions are:
wherein v is N ，v E ，v D Respectively, represents speed in the true north, east and local vertical directions, respectively>
Representing the rotation angle and speed of the terrestrial coordinate system relative to the inertial coordinate system>
Representing the angular rate of rotation, or ratio, of the navigational coordinate system relative to the terrestrial coordinate system>
Representing the ground speed.
fw is the specific force vector measured by a set of 3 accelerometers, resolved into the local geographical reference frame as:
f n ＝[f N f E f D ] T (16)
wherein, L0 represents a group of a compound represented by,
the rotation angular rate, i.e. transfer rate, of the local geographic coordinate system relative to the earth-fixed coordinate system is expressed in terms of a rate of change of longitude and latitude as follows:
make it
Wherein, L represents latitude, and is as follows:
in the formula (19), R 0 Is the radius of the earth; h is the height from the earth's surface.
where Ω represents the earth rotation angular rate, the navigation equation can be expressed in the form of the following components:
in the formula: ξ, η are the first angular deviation and the second angular deviation of the local gravity vector direction relative to the local vertical direction due to gravity anomaly.
The latitude, longitude and altitude from the earth's surface are given by the following equations:
wherein,
represents a differential form of latitude, is selected>
Represents a differentiated version of longitude, is>
Representing a highly differentiated form from the earth's surface.
In this embodiment, when calculating the velocity information, not only the latitude, the longitude, and the height from the earth surface are considered, but also an angular deviation of the local gravity vector direction with respect to the local vertical direction due to gravity anomaly is introduced, so that the calculated velocity information is more accurate.
And step S208, integrated navigation.
The micro strapdown inertial measurement unit and a mobile station satellite receiver (GNSS) form combined navigation, and the GNSS is responsible for resetting accumulated errors of the inertial navigation within a certain time period. With the loose assembly structure, the estimated position, velocity and attitude errors are used to correct the solution of inertial navigation, and the specific architecture is shown in fig. 2C.
The loose assembly structure of the embodiment is a series system, the navigation solution output position and speed from the GNSS are used as measurement information and input into a Kalman filter, and the Kalman filter is used for estimating the error of strapdown solution. The specific flow of the Extended Kalman Filter (EKF) method is shown in fig. 3, and includes the following steps: step S302, a system model and a measurement model are established; step S304, inputting an initial estimation value; step S306, carrying out linearization processing in a prediction stage, and estimating error covariance; step S308, calculating Kalman gain, and re-estimating error covariance based on the Kalman gain.
The main advantages of the loosely coupled integration provided by the present embodiment are simplicity and redundancy, which can be used for any micro inertial measurement unit (i.e. strapdown inertial measurement unit) and GNSS device, and is particularly suitable for improving algorithm application, in a loosely coupled configuration, there exists one independent GNSS available navigation solution, which is outside the combined navigation solution. When the open-loop micro-inertial measurement unit is corrected, an independent micro-inertial measurement unit navigation solution also exists, and the basic parallel navigation solution is supported.
In the error model of the micro inertial measurement unit of the embodiment, the error model is based on a small misalignment angle, and a nonlinear differential equation of the error of the micro inertial measurement unit is derived by expressing the influence of error factors through a small disturbance. And combining the state variables of the system according to the error state equations of the micro inertial measurement unit and the GNSS, wherein the state variables are formed by combining the error state variables of the micro inertial measurement unit and the GNSS and can be expressed by an equation (27).
Wherein,
representing a differential form of a system state variable, F (t) representing a state transition matrix, X (t) Representing the system state vector, B (t) representing the control matrix, u (t) representing the control variable, G (t) representing the system noise driving matrix, and w (t) representing the system noise matrix.
The errors of the vehicle-mounted micro-inertia measurement unit comprise a roll angle deviation delta alpha in the X-axis direction and a pitch angle deviation delta in the Y-axis direction β Deviation of yaw angle δ in Z-axis direction γ Three speed errors delta VN ，δ VE ，δ VD And three position errors δ L, δ L, δ h, δ f x δf y δf z Acceleration deviations (m/s) of X-axis, Y-axis and Z-axis, respectively 2 )，δ ωx δ ωy δ ωz X-axis, Y-axis, Z-axis gyro bias (°/s), respectively, represented by the following equation (28):
simplifying the error model to obtain a state transition matrix F (t) Is (29):
wherein, F S-9×9 Represents the error of the 9X 9-dimensional inertial navigation system, F DCM-6×6 Representing a 6 x 6 dimensional directional cosine matrix, DCM 3×3 Representing a 3 x 3 dimensional direction cosine matrix.
The error relation of the direction cosine matrix from the vehicle-mounted reference frame to the navigation reference frame and the gyroscope is represented by an equation (31).
Where C represents a direction cosine matrix element.
The F matrix has elements represented by equations (32) to (43):
wherein R is the radius of the earth, V N ，V E ，V D Is that the micro inertial measurement unit indicates the north, east and earth speed, f n ，f E ，f D The method is characterized in that the method is used for measuring north, east and earth direction specific force of an MIMU, L is the latitude output by a micro inertial measurement unit, omega is the rotational angular velocity of the earth, the height is negligible relative to the radius of the earth, and simplified processing is omitted.
The measurement model of the combined system is embodied in a matrix H, which relates the measurement model to the filter states. The observation information of this embodiment is the speed, position and attitude of the GNSS system, the attitude includes a yaw angle α and a pitch angle β, and the GNSS cannot provide a roll angle. GNS provides current location information of a vehicleUsing the geographic reference system as the navigation reference system, longitude L G Latitude l G Height h G Are parameters. GNSS provides vehicle speed information using Cartesian velocity [ V ] N V E V D ]As a GNSS velocity measurement.
The GNSS provided position is expressed as the sum of the geographical true and error values, as (44):
wherein L is G Representing latitude under a geographic coordinate system,/ G Representing longitude, h, in a geographic coordinate system G Representing height, L, in a geographic coordinate system N Representing latitude, N, in a navigational coordinate system N Indicating the north position error, N, of the satellite receiver E Indicating an east position error, R, of the satellite receiver N Denotes the transverse radius of curvature, h N And the height and the antenna position error of the Nh satellite receiver are expressed under a navigation coordinate system. δ L, δ h are position error measurements of the GNSS relative to the navigation reference frame as (45):
wherein L is I Representing the latitude, l, in an inertial frame I Representing longitude, l, in an inertial frame G Representing longitude, R, in a geographic coordinate system M Denotes the meridian radius of curvature, R N Denotes the transverse radius of curvature, P GN Representing the north position, P, under a geographical coordinate system GE Representing an east position, P, in a geographic coordinate system GD Representing the downward position in a geographic coordinate system. The speed measurement information of the micro-inertia measurement unit is the sum of the true value of the navigation system and the corresponding speed error, and the GNSS speed measurement information is expressed as the difference between the true value of the N system and the corresponding speed error. M N ，M E ，M D For the component of the GNSS velocity measurement error term on the north-east coordinate axis, the velocity measurement is as shown in equation (46):
wherein, V IN Representing the north velocity, V, in an inertial frame GN Representing the north velocity, V, in a geographic coordinate system IE Representing east velocity, V, in an inertial frame GE Representing east velocity, V, in a geographic coordinate system ID Representing the downward velocity, V, of the inertial frame GD Representing the downward speed, M, in a geographic coordinate system N Representing the speed of the satellite receiver in the north direction error, M E Indicating east error velocity, M, of the satellite receiver D Representing the earth-oriented error velocity, delta, of the satellite receiver VN Representing a component of error, δ, in the north direction VE Representing the error component of the ground-direction velocity, delta VD The ground-direction velocity error component is represented, HV represents the metrology matrix, X (t) represents the system state vector, and Vv (t) represents the metrology noise vector.
In the above formula
The standard deviation of the velocity is represented by equation (47).
Wherein HDOP represents a position accuracy factor,
representing a position standard deviation differential form.
The attitude measurement of the micro-inertia measurement unit is expressed as a true value and a corresponding attitude error under a navigation system, the attitude measurement of the GNSS is expressed as a true value and a corresponding attitude error under an N system, and a course angle and pitch angle measurement equation is (48):
wherein,
represents a measurement equation of the attitude, and>
represents the heading angle under the inertial coordinate system, and>
representing a heading angle in a geographic coordinate system>
Representing a depression elevation angle in an inertial coordinate system>
Representing a depression elevation angle in a geographical coordinate system, and>
represents a posture observation matrix, X (t) represents a system state vector, and->
Representing the measured noise matrix.
The position, velocity and attitude quantities (45) (46) (48) are measured and expressed as shown in equation 49:
wherein Z represents the measurement equation, Z p(t) Representing the position measurement equation, Z v(t) The equation for measuring the speed is expressed,
representing the attitude measurement equation, H V Represents a velocity measurement matrix, H P Represents a position measurement matrix, based on the position of the sensor>
Measurement matrix for representing attitude，v p Indicating errors in the position measurements of the satellite receiver, v v Indicating the error in the measurement of the satellite receiver velocity, X k+1 Represents the system estimation state at the moment of k +1, k represents the current navigation resolving moment, V k+1 Denotes the measurement noise sequence at time k + 1, and H denotes the measurement matrix.
Noise is observed for the GNSS.
In the embodiment, the measurement equation of the combined system is established by adopting the speed, position and attitude errors, and the method is more suitable for the change of a dynamic system than directly adopting the state value.
Step S210, initial alignment.
Reference information measured with a satellite reference station: longitude, latitude, altitude, attitude angle, pitch angle, and heading angle information. And transmitting the related reference information to a strapdown inertial measurement unit, receiving the reference information by the strapdown inertial measurement unit, and finishing transmission alignment under speed and position matching by utilizing the built EKF filter and combining self inertia and satellite data. Thereby achieving the initial alignment at the unmanned aerial vehicle end.
The application solves the technical problems of large volume and high cost of the split inertial and satellite navigation systems in the prior art, and has the advantages of small volume and low cost.
It should be noted that, for simplicity of description, the above-mentioned method embodiments are described as a series of acts or combination of acts, but those skilled in the art will recognize that the present application is not limited by the order of acts described, as some steps may occur in other orders or concurrently depending on the application. Further, those skilled in the art should also appreciate that the embodiments described in the specification are preferred embodiments and that the acts and modules referred to are not necessarily required in this application.
Through the above description of the embodiments, those skilled in the art can clearly understand that the method according to the above embodiments can be implemented by software plus a necessary general hardware platform, and certainly can also be implemented by hardware, but the former is a better implementation mode in many cases. Based on such understanding, the technical solutions of the present application or portions thereof that contribute to the prior art may be embodied in the form of a software product, where the computer software product is stored in a storage medium (such as a ROM/RAM, a magnetic disk, and an optical disk), and includes several instructions for enabling a terminal device (which may be a mobile phone, a computer, a server, or a network device) to execute the method described in the embodiments of the present application.
Example 3
According to the embodiment of the application, the combined navigation system for the subminiature unmanned aerial vehicle comprises a strapdown inertial measurement unit and a satellite differential positioning system.
The strapdown inertial measurement unit, as shown in FIG. 4, includes an inertial measurement component 42, a navigation computer 44, a secondary power module 40, a crystal oscillator 48, and external connectors 46, a mobile station satellite receiver 49.
1) Inertial measurement unit
The inertial measurement unit 42 mainly obtains acceleration and angular velocity information of the carrier, and uploads the data to the navigation computer 44 through the four-wire system SPI interface. The inertial measurement unit 42 houses a 3-axis gyroscope and a 3-axis accelerometer, which are specified in table 1.
TABLE 1
The hardware circuit layout is shown in fig. 5. The chip is connected with the navigation computer 44 through a four-wire system SPI interface, the power supply is 3.3V, and an independent bypass capacitor is arranged for filtering processing.
2) Navigation computer
The navigation computer 44 mainly performs navigation algorithm calculation, in this embodiment, the working dominant frequency of the navigation computer 44 is set to 160MHz, a floating point calculation unit is provided, the self-storage program space is 2MB, and the memory is 800KB, so that the navigation calculation requirement of one period of 5ms can be met. In addition, the navigation computer 44 also has a rich DMA channel for off-chip operations and logic. The specific resource allocation is shown in table 2 below.
| Serial number | Resource name | Association device | Remarks to note |
| 1 | USART0 | External communication | Navigation data output TTL @3.3 |
| 2 | USART1 | External communication | Navigation data |
| 3 | SPI0 | External communication | Four wire system SPI, |
| 4 | SPI1 | Inertial measurement assembly | Four-wire system SPI |
TABLE 2
The specific hardware design is shown in fig. 6. The minimal system of the chip comprises: a starting capacitor, a reset circuit, a clock source and the like.
3) Secondary power supply module
As shown in fig. 7, the secondary power supply module 40 mainly adjusts and matches internal power supply. In this embodiment, two separate chips are provided for power output, one for the mobile station satellite receiver 49 and one for the other digital circuits on the board. In order to improve the power performance and reduce the interference between devices, pi-type filter design is carried out on input and output in the power design process.
4) Crystal oscillator
The crystal oscillator 48 in the implementation has the characteristics of good temperature characteristic, high reliability and the like, and the stability of the crystal oscillator is less than or equal to 20ppm.
5) To external connector
In this embodiment, pins are attached to the surface of the external connector 46 at a pitch of 1.0mm, and gold is plated, so that contact resistance is reduced on the premise of ensuring stability.
The satellite differential positioning system comprises 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 line, a reference station satellite navigation system, a reference station satellite receiver, a reference station satellite receiving antenna, a reference station satellite receiving feeder line, a reference station satellite antenna connector, a secondary power supply module, an interface conversion module and a data fusion unit.
1) Mobile station satellite navigation system
Specific indices of the mobile station satellite navigation system are shown in table 3.
TABLE 3
The specific hardware design circuit is shown in fig. 8.
2) Reference station satellite navigation system
The satellite navigation system of the reference station has the functions of positioning and attitude measurement at the same time, and is matched with the high-precision inertial measurement unit to realize high-precision reference pose acquisition. The main technical criteria are shown in table 4 below.
TABLE 4
The satellite receiver is used for receiving high-precision satellite positioning data and realizing high-precision aviation-level positioning precision. And the effective data fusion is realized by combining the high-precision inertial measurement unit and the data fusion unit, and the final position and attitude initial measurement is finished.
Example 4
According to an embodiment of the present application, there is also provided a combined navigation device for a subminiature unmanned aerial vehicle, including: 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 revising.
Optionally, the specific examples in this embodiment may refer to the examples described in embodiment 1 and embodiment 2, and this embodiment is not described herein again.
Example 5
Embodiments of the present application also provide a storage medium. The storage medium is configured to store program codes for executing the methods in embodiments 1 and 2 above.
Optionally, in this embodiment, the storage medium may include, but is not limited to: a U-disk, a Read-Only Memory (ROM), a Random Access Memory (RAM), a removable hard disk, a magnetic or optical disk, and other various media capable of storing program codes.
The above-mentioned serial numbers of the embodiments of the present application are merely for description and do not represent the merits of the embodiments.
The integrated unit in the above embodiments, if implemented in the form of a software functional unit and sold or used as a separate product, may be stored in the above computer-readable storage medium. Based on such understanding, the technical solutions of the present application, which are essential or part of the technical solutions contributing to the prior art, or all or part of the technical solutions, may be embodied in the form of a software product, which is stored in a storage medium and includes several instructions for causing one or more computer devices (which may be personal computers, servers, network devices, or the like) to execute all or part of the steps of the methods described in the embodiments of the present application.
In the above embodiments of the present application, the descriptions of the respective embodiments have respective emphasis, and for parts that are not described in detail in a certain embodiment, reference may be made to related descriptions of other embodiments.
In the several embodiments provided in the present application, it should be understood that the disclosed client may be implemented in other manners. The above-described embodiments of the apparatus are merely illustrative, and for example, the division of the units is only one type of division of logical functions, and there may be other divisions when actually implemented, for example, a plurality of units or components may be combined or may be integrated into another system, or some features may be omitted, or not executed. In addition, the shown or discussed mutual coupling or direct coupling or communication connection may be an indirect coupling or communication connection through some interfaces, units or modules, and may be in an electrical or other form.
The units described as separate parts may or may not be physically separate, and parts displayed as units may or may not be physical units, may be located in one place, or may be distributed on a plurality of network units. Some or all of the units can be selected according to actual needs to achieve the purpose of the solution of the embodiment.
In addition, functional units in the embodiments of the present application may be integrated into one processing unit, or each unit may exist alone physically, or two or more units are integrated into one unit. The integrated unit may be implemented in the form of hardware, or may also be implemented in the form of a software functional unit.
The foregoing is only a preferred embodiment of the present application and it should be noted that those skilled in the art can make several improvements and modifications without departing from the principle of the present application, and these improvements and modifications should also be considered as the protection scope 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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