CN110764119A

CN110764119A - Satellite antenna autonomous measurement and control method and system independent of satellite navigation information

Info

Publication number: CN110764119A
Application number: CN201911081026.2A
Authority: CN
Inventors: 伍宗伟; 沈晓卫; 袁丁; 张峰干; 金伟; 朱丰超; 姚敏立; 贾维敏
Original assignee: Rocket Force University of Engineering of PLA
Current assignee: Rocket Force University of Engineering of PLA
Priority date: 2019-11-07
Filing date: 2019-11-07
Publication date: 2020-02-07
Anticipated expiration: 2039-11-07
Also published as: CN110764119B

Abstract

The invention discloses a satellite antenna autonomous measurement and control method and a satellite antenna autonomous measurement and control system independent of satellite navigation information. The method comprises the steps of firstly, obtaining initial values of a pitch angle and a roll angle according to a measurement value of an accelerometer, starting satellite finding, determining an initial course angle through back-stepping calculation after a target satellite is loaded, and then entering a stable tracking state; under the stable tracking state of the satellite antenna, carrying out attitude angle fusion estimation according to the measurement values of the accelerometer and the gyroscope, and determining the pitch angle and roll angle estimation values of the azimuth turntable; correcting and estimating the course angle of the azimuth turntable by adopting the course angle error obtained by tracking; and finally, adjusting the three-dimensional attitude angle of the azimuth turntable according to the three attitude angle estimated values, keeping the sight aiming between the satellite antenna and the target satellite, and realizing the autonomous measurement and control of the satellite antenna. The method can realize autonomous measurement and control under the condition of no satellite navigation information assistance by adopting a simple and practical algorithm, and has important practical significance under the conditions of limited satellite navigation information use and military use.

Description

Satellite antenna autonomous measurement and control method and system independent of satellite navigation information

Technical Field

The invention relates to the technical field of mobile satellite communication, in particular to a satellite antenna autonomous measurement and control method and system independent of satellite navigation information.

Background

With the development of economic globalization and informatization, people are urgently required to transmit or receive broadband and large-capacity multimedia information such as voice, data, images, videos and the like in real time at any place and any time in a journey so as to master the time bureau of the change of the moment more quickly and accurately. Satellite communication is the only communication system that can simultaneously provide different service requirements in different environments.

Since the bandwidth allocated to the satellite mobile service by ITU (International Telecommunication Union) is low, it is difficult to meet the requirement of broadband communication service. Satellite fixed service based "mobile communication" communication systems offer the possibility of this ideal implementation. The satellite communication system in motion refers to a system in which a mobile carrier (such as an automobile, a train, an airplane, a ship and the like) provided with a satellite antenna can establish a communication link with a stationary satellite (i.e. a synchronous orbit satellite) and can keep the communication link unblocked in the process of rapid movement of the carrier so as to realize real-time communication. In the process, the sight aiming between the antenna and the satellite is always kept to be the task needed to be solved by the measurement and control system.

The measurement and control system keeps the satellite pointing, the current course, pitching and rolling three-dimensional attitude angles of the carrier need to be sensed by means of the inertial device, the instruction angle for keeping the pointing inertial space stable is obtained through the coordinate system transformation, and the driving motor keeps the pointing, so that the real-time estimation of the three-dimensional attitude angles of the carrier becomes the key task of the measurement and control system. Currently, a Global Navigation Satellite System (GNSS) and an Inertial Navigation System (INS) are used to perform combined Navigation to obtain an accurate attitude angle estimation. However, once the satellite navigation system is interfered or loses the satellite navigation information signal, the device error of the inertial navigation system becomes larger and larger along with the time, and the attitude estimation can be invalid. In an urban environment, a canyon mountain region and a military application environment, signals of a satellite navigation system are very easily shielded or interfered, so that a large estimation error occurs in an attitude angle of a measurement and control system, and pointing deviation of an antenna is brought.

Disclosure of Invention

The invention aims to provide a satellite antenna autonomous measurement and control method and a satellite antenna autonomous measurement and control system which do not depend on satellite navigation information, so that the satellite antenna can still be stably tracked under the condition of no satellite navigation information assistance in satellite communication, and autonomous measurement and control are further realized.

In order to achieve the purpose, the invention provides the following scheme:

a satellite antenna independent measurement and control method independent of satellite navigation information is based on a micro-mechanical inertia measurement unit; the micro-mechanical inertia measurement unit is arranged on an azimuth turntable of the satellite antenna; the micro-mechanical inertial measurement unit internally comprises a gyroscope and an accelerometer which are orthogonally arranged in three axes; the satellite antenna autonomous measurement and control method comprises the following steps:

acquiring accelerometer measurements acquired by the accelerometer;

determining a pitch angle initial value and a roll angle initial value of the azimuth turntable according to the accelerometer measured value;

starting satellite searching according to the pitch angle initial value and the roll angle initial value, and determining an initial course angle by performing backward calculation on the upper target satellite;

adjusting the satellite antenna to be aligned with a target satellite according to the initial course angle, and entering a stable tracking state;

acquiring a gyro measurement value acquired by the gyro in a stable tracking state of the satellite antenna;

performing attitude angle fusion estimation according to the accelerometer measurement value and the gyroscope measurement value, and determining a pitch angle estimation value and a roll angle estimation value of the azimuth turntable;

acquiring a course angle error by adopting a satellite signal maximum tracking method;

correcting and estimating the course angle of the azimuth turntable by adopting the course angle error to obtain a course angle estimated value;

and adjusting the three-dimensional attitude angle of the azimuth turntable according to the pitch angle estimated value, the roll angle estimated value and the course angle estimated value, keeping the sight aiming between the satellite antenna and the target satellite, and realizing the autonomous measurement and control of the satellite antenna.

Optionally, determining the pitch angle initial value and the roll angle initial value of the azimuth turntable according to the accelerometer measurement value specifically includes:

using a formula based on said accelerometer measurements

Determining an initial pitch angle value of the azimuth turntable;

using a formula

Determining the initial value of the roll angle of the azimuth turntable; wherein a is_x,a_y,a_zAccelerometer measurements collected for three orthogonally placed accelerometers, respectively; g is a gravity acceleration value; let P₀＝P_a,R₀＝R_a，P₀As an initial value of pitch angle, R₀Is the initial value of the roll angle.

Optionally, the starting of satellite finding according to the pitch angle initial value and the roll angle initial value and the determining of the initial course angle by performing back-stepping calculation after the target satellite is located include:

let P be P₀，R＝R₀，H＝H₀When it is 0, the formula is adopted

Calculating to obtain the pitch angle and the polarization angle β gamma of the satellite antenna on the turntable, wherein

A transformation matrix from a geographic coordinate system to an antenna pointing coordinate system is represented, A, E, V represents the azimuth angle, the pitch angle and the polarization angle of the fixed satellite earth station aimed at the target satellite, respectively;

a transformation matrix representing a transformation from an antenna turntable coordinate system to an antenna pointing coordinate system;representing a transformation matrix from a geographic coordinate system to a turntable coordinate system, and H, P, R representing a heading angle, a pitch angle and a roll angle of the azimuth turntable respectively;

adjusting the satellite antenna to the pitch angle and the polarization angle according to the β and gamma, and rotating an azimuth turntable to start azimuth scanning;

after scanning for one week, judging whether the signal intensity of a beacon machine received by the satellite beacon is greater than a satellite capturing threshold value X_TObtaining a first judgment result;

if the first judgment result is that the received signal intensity of the beacon machine is less than or equal to the satellite capturing threshold value X_TAnd then adjust the pitch angle to β + (-1)ⁿ2n, returning to the step of adjusting the satellite antenna to the pitch angle and the polarization angle according to the β and gamma, and rotating an azimuth turntable to start azimuth scanning, wherein n is the number of rotation turns;

if the first judgment result is that the received signal intensity of the beacon machine is greater than the satellite capturing threshold value X_TThen according to β gamma indicated by the encoder at the maximum of the satellite signal strength, the formula is adopted

Calculating course angle H of the current azimuth turntable by reverse deduction_{Master and slave}；

The course angle H_{Master and slave}And converting the value range to the real value range of the course angle to obtain the initial course angle.

Optionally, the attitude angle fusion estimation is performed according to the accelerometer measurement value and the gyro measurement value, and the pitch angle estimation value and the roll angle estimation value of the azimuth turntable are determined, which specifically includes:

according to the accelerometer measurement value and the gyro measurement value, adopting a formulaDetermining the regulation parameter mu for judging the mobility of carrier, where Δ P is the difference between integral pitch angle of gyro and pitch angle measured by accelerometer, and Δ R is the difference between integral roll angle of gyro and roll angle measured by accelerometer, ξ₁Is a first threshold value, ξ₂Is a second threshold value;

judging an adjusting parameter mu according to the accelerometer measured value, the gyroscope measured value and the carrier maneuver by adopting a formula

Determining a pitch angle estimated value and a roll angle estimated value of the azimuth turntable; wherein P is_a,k-1The initial value of the pitch angle of the azimuth turntable at the kth-1 time point is obtained; r_a,k-1The initial value of the roll angle of the azimuth turntable at the k-1 time point; p_k-1The estimated value of the pitch angle of the azimuth turntable at the kth-1 time point is obtained; r_k-1The roll angle estimated value of the azimuth turntable at the k-1 time point; k is a radical of_pFor the first fusion weighting parameter, k_iA second fusion weighting parameter; δ t represents a calculation interval time; omega_x，k-1、ω_y，k-1、ω_z，k-1Respectively representing the gyro measurement values acquired by three orthogonally-installed gyros at the k-1 time point; p_kFor the k time point the orientation dial(ii) a pitch angle estimate; r_kAnd the roll angle estimated value of the azimuth turntable at the kth time point.

Optionally, the correcting and estimating the course angle of the azimuth turntable by using the course angle error to obtain a course angle estimated value specifically includes:

according to the course angle error, adopting a course angle correction algorithm

Correcting and estimating the course angle of the azimuth turntable to obtain a course angle estimated value; where Δ θ_a，k-1Tracking deviation angle for azimuth at time point k-1 β_k-1The pitch angle of the satellite antenna at the kth-1 time point; Δ H_k-1The error is the course angle error of the kth-1 time point; k is a radical of_p，HProportional coefficient, k, corrected for proportional integral_i，HAn integration coefficient corrected for proportional integral; h_k-1The estimated value of the course angle of the azimuth turntable at the kth-1 time point; h_kAnd the estimated value of the course angle of the azimuth turntable at the kth time point.

A satellite antenna autonomous survey and control system independent of satellite navigation information, the system comprising:

the accelerometer measured value acquisition module is used for acquiring accelerometer measured values acquired by an accelerometer;

the pitch angle and roll angle initial value determining module is used for determining a pitch angle initial value and a roll angle initial value of the azimuth turntable according to the accelerometer measured value;

the initial course angle determining module is used for starting satellite searching according to the pitch angle initial value and the roll angle initial value and determining an initial course angle through backstepping calculation after a target satellite is loaded;

the stable tracking module is used for adjusting the satellite antenna to be aligned with the target satellite according to the initial course angle and entering a stable tracking state;

the gyroscope measured value acquisition module is used for acquiring a gyroscope measured value acquired by the gyroscope in a stable tracking state of the satellite antenna;

the attitude angle fusion estimation module is used for performing attitude angle fusion estimation according to the accelerometer measurement value and the gyroscope measurement value and determining a pitch angle estimation value and a roll angle estimation value of the azimuth turntable;

the course angle error acquisition module is used for acquiring a course angle error by adopting a satellite signal maximum tracking method;

the course angle error correction module is used for correcting and estimating the course angle of the azimuth turntable by adopting the course angle error to obtain a course angle estimation value;

and the autonomous measurement and control module is used for adjusting the three-dimensional attitude angle of the azimuth turntable according to the pitch angle estimated value, the roll angle estimated value and the course angle estimated value, keeping sight aiming between the satellite antenna and the target satellite and realizing autonomous measurement and control of the satellite antenna.

Optionally, the pitch angle and roll angle initial value determining module specifically includes:

a pitch angle initial value determining unit for adopting a formula according to the accelerometer measurement value

Determining an initial pitch angle value of the azimuth turntable;

a roll angle initial value determining unit for employing a formula

Optionally, the initial heading angle determining module specifically includes:

a pitch angle and polarization angle calculation unit for setting P to P₀，R＝R₀，H＝H₀When it is 0, the formula is adopted

A transformation matrix from a geographic coordinate system to an antenna pointing coordinate system is represented, A, E, V represents the azimuth angle, the pitch angle and the polarization angle of the fixed satellite earth station aimed at the target satellite, respectively;a transformation matrix representing a transformation from an antenna turntable coordinate system to an antenna pointing coordinate system;representing a transformation matrix from a geographic coordinate system to a turntable coordinate system, and H, P, R representing a heading angle, a pitch angle and a roll angle of the azimuth turntable respectively;

the azimuth scanning unit is used for adjusting the satellite antenna to the pitch angle and the polarization angle according to the β and gamma, and rotating an azimuth turntable to start azimuth scanning;

a beacon machine signal intensity judging unit for judging whether the beacon machine signal intensity received by the satellite beacon is greater than the satellite capturing threshold value X or not after scanning for one circle_TObtaining a first judgment result;

a cyclic scanning unit for determining whether the received beacon signal strength is less than or equal to a satellite capture threshold X_TAnd then adjust the pitch angle to β + (-1)ⁿ2n, returning to the step of adjusting the satellite antenna to the pitch angle and the polarization angle according to the β and gamma, and rotating an azimuth turntable to start azimuth scanning, wherein n is the number of rotation turns;

a course angle calculation unit for calculating the course angle of the received beacon if the first judgment result is that the received signal intensity of the beacon is greater than the satellite capture threshold X_TThen according to the encoder indication at the maximum of the satellite signal strengthβ, gamma, using the formula

An initial course angle determining unit for determining the course angle H_{Master and slave}And converting the value range to the real value range of the course angle to obtain the initial course angle.

Optionally, the attitude angle fusion estimation module specifically includes:

a carrier maneuvering judgment and adjustment parameter calculation unit for adopting a formula according to the accelerometer measurement value and the gyroscope measurement valueDetermining the regulation parameter mu for judging the mobility of carrier, where Δ P is the difference between integral pitch angle of gyro and pitch angle measured by accelerometer, and Δ R is the difference between integral roll angle of gyro and roll angle measured by accelerometer, ξ₁Is a first threshold value, ξ₂Is a second threshold value;

a pitching and rolling angle estimation value determining unit for judging the adjusting parameter mu according to the accelerometer measurement value, the gyroscope measurement value and the carrier maneuver by adopting a formula

Determining a pitch angle estimated value and a roll angle estimated value of the azimuth turntable; wherein P is_a,k-1The initial value of the pitch angle of the azimuth turntable at the kth-1 time point is obtained; r_a,k-1The initial value of the roll angle of the azimuth turntable at the k-1 time point; p_k-1The estimated value of the pitch angle of the azimuth turntable at the kth-1 time point is obtained; r_k-1The roll angle estimated value of the azimuth turntable at the k-1 time point; k is a radical of_pFor the first fusion weighting parameter, k_iA second fusion weighting parameter; δ t represents a calculation interval time; omega_x，k-1、ω_y，k-1、ω_z，k-1Individual watchShowing the gyro measurement values acquired by three orthogonally mounted gyros at the k-1 st time point; p_kThe estimated value of the pitch angle of the azimuth turntable at the kth time point is obtained; r_kAnd the roll angle estimated value of the azimuth turntable at the kth time point.

Optionally, the heading angle error correction module specifically includes:

a course angle error correction unit for adopting course angle correction algorithm according to the course angle error

According to the specific embodiment provided by the invention, the invention discloses the following technical effects:

the invention provides a satellite antenna autonomous measurement and control method and a system independent of satellite navigation information, wherein the method comprises the steps of firstly determining a pitch angle initial value and a roll angle initial value of an azimuth turntable according to an accelerometer measured value; starting satellite searching according to the pitch angle initial value and the roll angle initial value, and determining an initial course angle by performing backward calculation on the upper target satellite; adjusting the satellite antenna to be aligned with a target satellite according to the initial course angle, and entering a stable tracking state; under the stable tracking state of the satellite antenna, carrying out attitude angle fusion estimation according to the measurement value of the accelerometer and the measurement value of the gyroscope, and determining a pitch angle estimation value and a roll angle estimation value of the azimuth turntable; correcting and estimating the course angle of the azimuth turntable by adopting the course angle error obtained by tracking to obtain a course angle estimated value; and finally, adjusting the three-dimensional attitude angle of the azimuth turntable according to the pitch angle estimated value, the roll angle estimated value and the course angle estimated value, keeping the sight aiming between the satellite antenna and the target satellite, and realizing the autonomous measurement and control of the satellite antenna. The method can completely separate the assistance of navigation positioning information, independently and autonomously work, can realize autonomous measurement and control under the condition of no satellite navigation information assistance by adopting a simple and practical algorithm, and has important practical significance under the conditions of limited satellite navigation information use and military use.

Drawings

In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings needed to be used in the embodiments will be briefly described below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and it is obvious for those skilled in the art to obtain other drawings without inventive exercise.

FIG. 1 is a schematic view of an installation of a satellite azimuth turntable and an inertial measurement unit provided by the present invention;

FIG. 2 is a flowchart of a method for initial alignment and initial assignment of attitude angles for satellites according to the present invention;

FIG. 3 is a schematic diagram of a gyroscope and accelerometer fusion estimation attitude angle provided by the present invention;

FIG. 4 is a schematic diagram of azimuthal sinusoidal scan maximum tracking provided by the present invention;

FIG. 5 is a flowchart of a method for satellite occlusion recovery and course angle reassignment according to the present invention.

Detailed Description

The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the 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 invention.

The invention provides a satellite antenna autonomous measurement and control method and a satellite antenna autonomous measurement and control system independent of satellite navigation information, aiming at the problem that the conventional measurement and control system must depend on the satellite navigation information.

In order to make the aforementioned objects, features and advantages of the present invention comprehensible, embodiments accompanied with figures are described in further detail below.

The invention aims to solve the technical problem that when no satellite navigation information exists, errors accumulated by an independent inertia measurement unit along with time cannot be eliminated, and a new data fusion structure and a new data fusion method are required to complete the estimation of the three-dimensional attitude angle. Except the inertial measurement unit, the only means which can be selected by the current satellite communication-in-motion system is a satellite beacon receiver, under the condition of aligning a satellite, a signal maximum value tracking method is adopted to estimate the pointing deviation of the satellite in real time, and the key of the invention for correcting the attitude angle and correcting the gyro error is that the invention can realize autonomous measurement and control.

The sensor on which the method of the invention is based comprises: an Inertial Measurement Unit (IMU) and a satellite beacon receiver. The two are placed on the azimuth turntable of the satellite antenna, the adoption of the placing mode is beneficial to realizing the pointing stability of the inertial space by utilizing the space velocity measurement characteristic of the gyroscope, and the transformation of the coordinate system is simple.

The method comprises three aspects:

firstly, initial attitude angle assignment is carried out, gyroscope integration can be carried out to obtain uninterrupted attitude angle estimation only if the initial assignment is completed by the three-axis angles of azimuth, elevation and roll, and under the condition of satellite navigation information assistance, a real attitude angle can be quickly obtained by utilizing a combined navigation algorithm and a ground speed and course angle output by satellite navigation information, but under the condition of no satellite navigation information assistance, the initial assignment of the attitude angle becomes a practical problem;

secondly, the problem of stable tracking of pointing of a satellite antenna behind a satellite is solved, the attitude angle obtained by integrating a gyroscope is influenced by zero offset of the gyroscope, the long-time accuracy cannot be guaranteed, particularly, the accumulated error of a low-cost MEMS (Micro-Electro-Mechanical System, Micro Electro Mechanical System) gyroscope in one hour can reach 30 degrees or even more, and how to keep the accuracy of attitude estimation by using the received signal strength of a satellite beacon receiver is the core problem of the invention;

and finally, the problems of satellite reacquisition and attitude reassignment after the satellite signal is shielded are solved, and once the satellite signal is shielded, the problem that the gyro accumulated error cannot be corrected in an auxiliary manner due to the fact that the satellite navigation information is not assisted exists, at the moment, the attitude angle has estimation error, and how to realign the target satellite and the attitude angle reassignment after the satellite signal is not shielded exists, so that the problem that the satellite link recovery is considered by the autonomous measurement and control system is solved.

The invention provides a satellite antenna autonomous measurement and control method independent of satellite navigation information, which specifically comprises the following steps:

1. attitude angle initialization in initial satellite finding

The invention adopts a micro-Mechanical Inertial Measurement Unit (MIMU) which is arranged on an azimuth turntable of an antenna and is arranged at a position close to an azimuth rotating shaft as much as possible, thereby reducing the influence of the rotation of the azimuth turntable on inertial measurement. Compared with the mode of being installed on the vehicle body, the IMU signal wire adopting the installation mode can directly enter the main controller without passing through the azimuth slip ring, and the attitude of the antenna can be directly measured, so that the coordinate system conversion operation is reduced. In addition, under the condition of stable azimuth, the azimuth turntable is in an inertial space stable state, the output value of the azimuth gyro is almost kept near a zero value, and the measurement range and the precision requirement of the azimuth gyro can be properly reduced. The inside of the MIMU comprises a gyroscope and an accelerometer which are orthogonally arranged in three axes, and sensitive axes respectively point to the right front, right side and right top of an antenna turntable (namely an azimuth turntable). As shown in FIG. 1, the measured values are all represented in the turntable coordinate system b (three-dimensional coordinate x)_b、y_b、z_b). The other two coordinate systems are respectively a local geographic coordinate system n (origin O of the coordinate system)_nThree, 2Dimensional coordinate x_n、y_n、z_n) And an antenna beam pointing coordinate system s (three-dimensional coordinate x)_s、y_s、z_s) The definitions are shown in FIG. 1.

The basic principle of the satellite communication in motion can be illustrated by the following formula,

thereinRepresenting a transformation matrix from a geographical coordinate system to an antenna pointing coordinate system,

representing a transformation matrix from the antenna turntable coordinate system to the antenna pointing coordinate system,

representing a transformation matrix from the geographical coordinate system to the turntable coordinate system. Wherein:

for example, in the case of a seven-satellite in western ann, the satellite longitude is 105.5 ° east longitude, and according to the latitude and longitude of the target satellite, the east longitude is 109.0 ° north latitude 34.3 °, and the satellite angles are calculated statically as azimuth angle a 186.1 °, pitch angle E50.0 °, and polarization angle V-5 °, the satellite antenna fixed on the ground needs to be adjusted according to the three angles A, E, V to align the antenna for communication, but since the satellite antenna is mounted on a moving carrier, the carrier moves to cause three attitude angles H, P, R (i.e., heading angle, pitch angle, roll angle) of an azimuth turntable (i.e., an azimuth turntable), the orientation of the antenna is affected, and therefore the orientation of the satellite antenna relative to the turntable must be adjusted in real time (i.e., the antenna is adjusted relative to azimuth angle α, pitch angle β, and polarization angle e., a y-direction of the antenna relative to the turntable), so that the inertial orientation of the antenna is stable in space, and therefore, if the orientation of the satellite antenna is stabilized by a steering angle γ -orientation formula, the orientation of the antenna is calculated as if the orientation of the antenna is stabilized on the ground, the same as if the orientation of the antenna, the orientation of the satellite antenna is stabilized orientation of the antenna, the orientation of the satellite is calculated by the orientation of the satellite antenna, the satellite, the orientation of the satellite, the satellite is stabilized orientation of the satellite, the satellite is stabilized orientation of the satellite, the satellite is calculated by the satellite

The attitude angle H, P, R of the azimuth turntable can be integrated by a 3-axis gyro mounted on the turntable, and the precise knowledge of the initial value becomes the first important issue that must be addressed. The method is characterized in that the pitch angle P and the roll angle R of the turntable are obtained through an accelerometer gravity measurement value placed on the turntable, and the course angle H is obtained through driving the azimuth turntable to scan for finding a satellite by 360 degrees and then calculating through backstepping.

Under the condition that the carrier is static, the pitch angle and the roll angle of the rotary table can be calculated by the accelerometer through gravity measurement:

a in the above formula_x,a_y,a_zFor three orthogonally placed accelerometer measurements, g is the gravitational acceleration value. Let P₀＝P_a,R₀＝R_aObtaining the initial values P of the pitch angle P and the roll angle R₀,R₀Then, assume the initial value of course angle H at this time₀If the initial values of H, P, R are all given, calculating and adjusting the pitch angle β and the polarization angle γ of the antenna on the turntable according to the formula (2), then driving the azimuth turntable to scan a circle to search for a target satellite, and reversely deducing the initial values of the heading angles, the specific method is (as shown in fig. 2):

1) giving initial values P to the attitude angles P and R of the turntable according to the accelerometer measurement value of the current position of the antenna system₀,R₀The course angle is directly set to H₀The pitch angle and the polarization angle α, γ of the antenna on the turntable at this time need to be adjusted are calculated by equation (1), and the heading angle H is unknown, so even if the antenna is adjusted to the calculated target angle, the antenna still cannot be aligned with the satellite;

2) rotating the azimuth turntable to start to find the star;

3) after rotating the azimuth for one circle, if the satellite cannot be locked according to the signal of the satellite beacon receiver (the standard for locking the satellite is that the received signal strength of the satellite is greater than the satellite acquisition threshold X_TThe threshold value X_TTypically empirically based on the characteristics of the beacon receiver), the pitch angle is adjusted to β + (-1)ⁿAnd 2n, continuing to rotate the azimuth finder. The significance of adjusting the pitch angle is to reduce the influence of pitch angle error and improve the reliability of azimuth scanning satellite finding. Where n is the number of rotations, n > - θ/2, and θ is the pitch beam width. Considering that the common satellite communication-in-motion antennas are small-caliber antennas, theta is generally 3-5 degrees.

4) If the received signal strength of the beacon machine is larger than the satellite capturing threshold value X_TIf yes, the satellite is captured, the course angle H of the current azimuth turntable is calculated through formula (2) in a reverse-deducing mode according to β and gamma indicated by an encoder at the maximum value of the satellite signal intensity, and the target satellite capturing and the initial course angle obtaining under the condition of no initial course angle are completed, otherwise, the step 2 is returned.

The initial value of P, R is determined through the above steps, namely step 1), by the accelerometer measurement value, the heading angle H at this time is assumed to be 0, then the azimuth turntable scans circularly to search for the satellite, after the target satellite is found, the true H is calculated through reverse estimation, and the autonomous satellite searching and heading angle initial value assigning under the condition without GNSS assistance are realized. The course angle reverse-deducing calculation method in the step 4) comprises the following steps:

from

The first two data sinHcosP and cosHcosP of the second row are only taken out from the expression (A), so that the backward calculation of H can be realized, as shown in the following:

note that the derived heading angle H in equation (6)_{Master and slave}Is (-90,90) in order to be within a range of [0,360 ] from the true value of the heading angle H]Correspondingly, an angle conversion is also required, as shown in table 1 below:

TABLE 1 course Angle conversion mapping Table

C₂₁	C₂₂	H true value
			+	+	H_{Master and slave}
+	–	180°+H_{Master and slave}
			–	+	360°+H_{Master and slave}
–	–	180°+H_{Master and slave}

For the communication in motion, after the target satellite is aligned, the satellite enters a stable tracking state, the attitude angle of the rotary table can be uninterruptedly calculated by utilizing the gyro integral, the command angle α of the control motor is calculated by the formula (2), and the gamma keeps the opposite star state.

2. Attitude angle fusion estimation in stable tracking

Under the condition of stable tracking of the antenna, attitude estimation mainly depends on gyro integral acquisition, the short-time precision of a gyro is higher, the estimation of an attitude angle in limited time is more accurate, and the estimation algorithm is as follows:

omega in the formula_x、ω_y、ω_zRespectively representing three gyro measurements mounted orthogonally inside the IMU, each representing an estimate of the attitude angle R, P, H, the estimation of attitude is achieved by the above integration algorithm. But the gyroscope contains a slowly time-varying zero offset, namely:

ω＝ω_t+b+η_ω(8)

where ω is a gyro measurement, ω_tFor true angular velocity, b is zero offset error, which typically varies relatively slowly, η_ωTo measure noise.

In order to remove the influence of zero offset and gyro noise, other error correction means must be considered, and auxiliary measurement means available for the system are accelerometer and beacon receiver signal strength information.

The pitch angle and the roll angle can be obtained through accelerometer measurement, as shown in formula (3), but the measurement value is inevitably influenced by carrier maneuvering acceleration and noise such as vibration, the short-time precision is poor, but the long-time precision is good, so the attitude angle estimation with high precision can be obtained through fusion estimation of the attitude estimation values of the gyroscope and the accelerometer. The general fusion estimation method adopts a Kalman filter frame, but the calculated amount is large, in order to facilitate embedded use, the invention provides a common PI (proportional integral) control frame, the attitude angle error adopts a direct proportional weighting mode to fuse the attitude angle information of a gyroscope and an accelerometer, the zero offset error in the gyroscope data is regarded as the first-order static error of a control system, an integral link is adopted for correction, the method is simple and practical, signal processing is carried out from the angle of the control system, and the realized schematic block diagram is shown in figure 3. The implementation algorithm is as follows:

wherein R is_a、P_aFor the attitude angle obtained by the accelerometer, see equation (3), t is the integration time.

Under the condition of short calculation interval time, the above result is subjected to approximate discretization to obtain an algorithm formula of the final estimated pitch angle P and roll angle R:

wherein the index k denotes the variable value at the kth point in time for an iteration over time, e.g. R_kAnd P_kRespectively representing roll angle and pitch angle at the k-th time point, R_k-1And P_k-1Respectively represent the roll angle and the pitch angle, omega, at the k-1 th time point_x，k-1、ω_y，k-1、ω_z，k-1Three gyro measurements are shown at time points k-1, respectively. Wherein the attitude angle (roll angle and pitch angle) R is obtained by an accelerometer_a,k-1、P_a,k-1Is obtained by the formula (3), δ t represents the calculation interval time, δ ω_x,k-1And δ ω_y,k-1And respectively representing the error correction values of the x-axis gyroscope and the y-axis gyroscope at the k-1 time point. k is a radical of_p、k_iFusion weighting parameters corresponding to the accelerometer and the gyroscope respectively, and mu is a carrier maneuver judgment and adjustment parameter. Typically, after the accelerometer and gyroscope are selected, the weighting parameter k is_p、k_iCan be selected by field adjustment. Mu is another important parameter which needs to be adjusted to reduce attitude angle measurement errors of the accelerometer due to disturbance of vehicle maneuvering acceleration.

It is known that the pitch angle and the roll angle of an accelerometer can be calculated by measuring the gravity vector under a static condition, but the accelerometer is easily influenced by the maneuvering acceleration of a carrier after being installed on the carrier, and large errors are brought to the measurement of the attitude angle of the accelerometer. Therefore, the determination of the mobility of the support is required to set the parameter μ, and the present invention employs the following strategy to adjust the parameter,

where μ is the rule of judgment, a_x,a_y,a_zIs three-axis withThe measurement of the speedometer, g is the acceleration of gravity, which can be approximated to be 9.8m/s2, Δ R, Δ P is the difference between the gyro integral attitude angle and the accelerometer measurement attitude angle, ξ₁，ξ₂Two thresholds.

First judgment formulaMeaning that if the square root of the sum of squares of the three-axis accelerometer measurements is greater than the threshold ξ compared to the acceleration of gravity g₁Then the disturbance of the motoric acceleration is considered to be present, the principle being that the measurement of the accelerometer should contain only the acceleration of gravity g if the carrier is free of disturbance of the motoric acceleration. Second judgment rule

The meaning of (1) is that if the square root value of the error value sum of squares obtained by integrating the accelerometer and the gyroscope is larger than the threshold value ξ₂It indicates that a large interference may be experienced. If both the two judgment rules are smaller than the threshold value, the value mu is equal to 0, the interference of the inorganic dynamic acceleration is considered, if a certain judgment formula is larger than the threshold value, the value mu is equal to 1, and the interference of the acceleration is considered.

The estimated values of the pitch angle and the roll angle are obtained through the fusion algorithm of the gyroscope and the accelerometer

However, the course angle is not corrected in an auxiliary way by any measuring means, and the error of the course angle is acquired by tracking the maximum value of the satellite signal so as to correct the error of the course angle.

There are many ways for tracking the maximum value of the satellite signal, such as step tracking, cone scanning, single pulse tracking, etc., in the present invention, a way of half cone scanning is used to realize satellite tracking, i.e. the azimuth is scanned separately, and a sine wave scanning way is used to realize the method, and the implementation schematic diagram is shown in fig. 4. In fig. 4, δ θ represents an angle error value, RF_powerIndicating the beacon received signal strength. For example, in position 1, it is apparent that the antenna is pointing to the right, when this is the caseWhen scanning the positive half cycle, the beacon receives a signal stronger than that of the negative half cycle, and in order to avoid the error caused by the interference of the noise signal, the signal intensity integration mode in one scanning cycle is adopted to obtain the current angle error, and in order to obtain the direction of the error signal, an operation similar to signal inversion is required, that is, the negative half cycle signal is inverted into a dotted line form in the figure as shown in fig. 4, that is, the deviation direction can be obtained in the integration result of one cycle. Received signal strength RF within the satellite beamwidth_powerThe relationship to the pointing deviation angle δ θ is generally approximated as a quadratic curve, namely:

RF_power＝(-a·δθ²+c)+n (12)

the parameter δ θ is an azimuth dimension pointing deviation angle, a and c are parameters of a quadratic curve and an antenna with a certain caliber, the two parameters can be obtained through a design index and an actual measurement result of the antenna, n is measurement noise, the measurement noise has a large influence on the judgment of an error angle, and particularly for satellite communication, the interference and noise of a satellite channel and a satellite beacon machine are usually large, so if a step tracking method is adopted, the influence of the noise can cause tracking failure.

In order to remove the influence of the noise, a method of integrating within one period is adopted to improve the tracking accuracy. Assuming the satellite beam azimuth deviation angle as delta theta_aThe azimuth scanning angle is a sine signal theta_scan＝θ_s·sinω_st, then the received signal strength is:

RF_power＝-a·(Δθ_a-θ_s·sinω_st)²+c+n (13)

the signal is multiplied by the original scan signal, at which time, for the negative half cycle signal, the inversion is achieved,

wherein I is an integral value, RF_powerFor received signal strength, theta_sFor sweeping the amplitude, omega, of the sinusoidal signal_sT is one scanning period for the scanning signal period. Substituting the formula (13) into the formula (14), and calculating to obtain:

in this integration (15), the result of the first part is obviously 0, leaving only the second part, namely:

thus, the resulting deviation angle is:

the azimuth tracking deviation angle Delta theta is obtained_aThe following task is then to correct the error of the heading angle H by this deviation. Before this, an error transfer equation must be obtained, and through a secant compensation, the azimuth dimension tracking error of the antenna beam coordinate system is firstly converted into an azimuth turntable control error of the antenna, namely:

where Δ α is the antenna azimuth turntable control error angle, which can be considered to be caused entirely by the error in heading angle H, and β is the pitch angle of the antenna at that time, which can be accurately measured by a pitch encoder, the following equation (1) holds:

wherein:

for simplicity, only the matrix is considered

The second line first element sin α cos β, in conjunction with equation (5), has:

since P and R are considered to have been accurately estimated by the accelerometer, we differentiate the above equation (20) considering only the H error to yield:

[C₂₁C₂₂C₂₃]representing coordinate transformation matrices

The second line of (2) as shown in equation (5).

Substituting equation (21) to obtain the result:

dα＝dH (23)

therefore, if the error obtained from the tracking result is an azimuth error, the azimuth tracking error angle Δ α is a heading angle error Δ H:

like pitch angle and roll angle estimation, the course angle H is corrected and estimated by adopting the idea of PI correction, and a course angle correction algorithm is obtained by referring to a formula (10):

in which the parameters define a reference equation (10) and the index k denotes the value of the variable at the kth point in time for an iteration over time, e.g.Δθ_a，k-1Indicating the satellite beam bearing deviation angle at the k-1 time point, β_k-1Represents the antenna pitch angle at the k-1 time point, the only difference being the heading angle error Δ H at the k-1 time point_k-1Sinusoidal scan error Δ θ from the azimuthal dimension_a。k_p，H、k_i，HProportional, integral fusion weighting parameter for course angle correction, H_k-1Denotes the heading angle, H, at the k-1 st time point_kIndicating the heading angle at the kth time point.

In conclusion, under the condition that the satellite antenna normally tracks the satellite, the pitch angle and the roll angle are adopted to be corrected and estimated by using the accelerometer, the heading angle is estimated autonomously by adopting a method of fusion and estimation of tracking information of a satellite beacon machine, a real-time α and gamma value is calculated according to the three attitude angles of the pitch angle, the roll angle and the heading angle estimated in real time, and then the driving motor adjusts the corresponding command angle to realize communication in motion.

3. Satellite occlusion recovery and course angle reassignment

Under the condition that the satellite signals are shielded, the pitch angle P and the roll angle R can still be corrected in an auxiliary mode by means of the accelerometer. The course angle H cannot be corrected by the tracking signal and can only be maintained by means of gyro integration. After the satellite signal is recovered, the satellite signal needs to be detected quickly and recovered to a stable tracking state, and meanwhile, the heading angle H needs to be re-assigned to obtain a correct three-dimensional attitude angle again so as to perform stable tracking on the satellite.

The method adopted by the invention is that the signal output of the satellite beacon receiver is utilized, when the signal is detected to be continuously smaller than a certain threshold value, the signal shielding mode is considered to be entered, at the moment, the azimuth angle of the satellite antenna starts to scan left and right, and the scanning window gradually becomes larger along with the time so as to accord with the objective rule that the integral error of the gyroscope gradually increases along with the time. When the satellite signal is detected, the satellite is considered to be acquired again, and the course angle at the moment can be calculated through a formula (4), so that the satellite re-acquisition and the course angle re-assignment are completed. The implementation method is shown in fig. 5, and comprises the following specific steps:

1) determining that a beacon receiver detects a signal R_LAnd a signal threshold L_trackThe size of (d); if the beacon receiver detects the signal R_LGreater than a signal threshold L_trackAdjusting the satellite antenna to perform stable tracking according to the real-time estimated three-dimensional attitude angle; if the beacon receiver detects the signal R_LLess than or equal to the signal threshold L_trackIf the vehicle enters the shielding mode, the pitch angle and the roll angle are still corrected by the accelerometer, but the heading angle is only subjected to gyro integral estimation; starting to time the shielding time t;

2) judging the shielding time T and the shielding time threshold T_gyroThe size of (d); if the occlusion time T is less than or equal to the occlusion time threshold T_gyroCorrecting a pitch angle and a roll angle by adopting an accelerometer, and only carrying out gyro integral estimation on a course angle; if the occlusion time t>Occlusion time threshold T_gyroI.e. the signal blocking time exceeds the time for which gyro integration can ensure that the beam deviation is less than the beam width, and the azimuth window scanning must be started considering that the satellite is lost at this time. For example, if the beam width of a 0.6m aperture antenna is 3.6 °, the null error of a micro-mechanical gyro is 0.02 °/s, and the gyro offset is deviated by more than half the beam width (the beam width is symmetric about the beam center), it is considered that it is difficult to acquire a satellite again at this time, and T is a value_gyro3.6 °/2/0.02 °/s is 90s, i.e. when the time t is>At 90s, the azimuth window scan must be started, assuming that the satellite has been lost at this time.

3) The method comprises the following steps of (1) scanning an azimuth window, wherein a scanned window theta is dynamically adjusted along with time, and the adjustment algorithm is as follows:

θ＝θ₀+(Δω_z) t, θ therein₀For the initial scanning window,. DELTA.omega_zThe amount of the gyro drift hedging increase can be set to 1.5 times of the gyro drift. During the scanning process, the level R of the beacon receiver is continuously detected_LWhether or not it is greater than the capture threshold L_acq。

4) When receiving signal level R_LAbove capture threshold L_acqConsidering that the satellite signal is currently received, the satellite is reacquired, and then the scanning is stopped in the azimuth,and calculating course angle assignment by backstepping according to a formula (4) and entering a stable tracking state.

Compared with the prior art, the autonomous measurement and control method of the satellite antenna provided by the invention has the following advantages:

1. the method can ensure that the satellite antenna completely breaks away from the assistance of navigation positioning information and independently works, which is an important function that the satellite communication in China lacks at present, and is especially important in the environment that the satellite navigation information is interfered or is not allowed to be used.

2. At present, the satellite communication-in-motion in China can perform initial satellite finding only after a satellite navigation and inertial navigation combined algorithm completes navigation solution and attitude angle convergence, and the method of the invention obtains initial values of a pitching angle and a rolling angle of a rotary table by using an accelerometer under the condition of only a gyroscope and the accelerometer, rotates an azimuth rotary table to perform initial satellite finding, captures the satellite and completes initial assignment of a course angle. And a dynamic pitch angle adjusting strategy is adopted in the satellite searching process, so that the satellite searching reliability is enhanced.

3. In the normal stable tracking of satellite communication-in-motion, the method not only utilizes the gravity measurement value of the accelerometer to assist in correcting the errors of the pitch angle and the roll angle, but also provides a sinusoidal scanning method for resolving the azimuth tracking error through the received signal of the beacon machine. Because the error correction link in the fusion estimation mainly comprises an attitude angle and a gyro zero offset, and the gyro zero offset is a slow time-varying approximate constant error, a PI correction technology is provided to eliminate the attitude estimation error and the gyro zero offset by taking the PID control theory as reference. Aiming at the problem that the maneuvering acceleration of the carrier can interfere with gravity measurement to influence attitude angle estimation, a carrier maneuvering judgment rule is designed to adjust PI fusion parameters, and the reliability and accuracy of estimation are improved.

4. In the method, under the condition that the satellite signal is shielded, the pitch angle and the roll angle are still corrected by the aid of acceleration, but the course angle error is not corrected by the aid of the heading angle error. Therefore, a timing azimuth scanning and scanning window self-adaptive algorithm is provided for re-acquisition of satellite signals and re-assignment of course angles. The timing azimuth scanning utilizes the short-time stability of the gyroscope, and the scanning window self-adaptive algorithm fully considers the angle drift which possibly occurs for a long time, and the scanning window is continuously increased along with the accumulation of time, so that the reliability of recapture is improved.

Based on the satellite antenna autonomous measurement and control method provided by the invention, the invention also provides a satellite antenna autonomous measurement and control system independent of satellite navigation information, and the system comprises:

The pitch angle and roll angle initial value determining module specifically comprises:

a pitch angle initial value determining unit for adopting a formula according to the accelerometer measurement valueDetermining an initial pitch angle value of the azimuth turntable;

a roll angle initial value determining unit for employing a formula

The initial course angle determining module specifically comprises:

a transformation matrix representing a transformation from an antenna turntable coordinate system to an antenna pointing coordinate system;

representing a transformation matrix from a geographic coordinate system to a turntable coordinate system, and H, P, R representing a heading angle, a pitch angle and a roll angle of the azimuth turntable respectively;

a course angle calculation unit for calculating the course angle of the received beacon if the first judgment result is that the received signal intensity of the beacon is greater than the satellite capture threshold X_TThen according to β gamma indicated by the encoder at the maximum of the satellite signal strength, the formula is adopted

The attitude angle fusion estimation module specifically comprises:

a carrier maneuvering judgment and adjustment parameter calculation unit for adopting a formula according to the accelerometer measurement value and the gyroscope measurement value

Determining a carrier maneuvering judgment adjusting parameter mu; wherein Δ P is the integral pitch angle of the gyroscope and the pitch angle measured by the accelerometerThe difference between the integral roll angle of the gyro and the roll angle measured by the accelerometer, and a delta R ξ₁Is a first threshold value, ξ₂Is a second threshold value;

Determining a pitch angle estimated value and a roll angle estimated value of the azimuth turntable; wherein P is_a,k-1The initial value of the pitch angle of the azimuth turntable at the kth-1 time point is obtained; r_a,k-1The initial value of the roll angle of the azimuth turntable at the k-1 time point; p_k-1The estimated value of the pitch angle of the azimuth turntable at the kth-1 time point is obtained; p_k-1The roll angle estimated value of the azimuth turntable at the k-1 time point; k is a radical of_pFor the first fusion weighting parameter, k_iA second fusion weighting parameter; δ t represents a calculation interval time; omega_x，k-1、ω_y，k-1、ω_z，k-1Respectively representing the gyro measurement values acquired by three orthogonally-installed gyros at the k-1 time point; p_kThe estimated value of the pitch angle of the azimuth turntable at the kth time point is obtained; r_kAnd the roll angle estimated value of the azimuth turntable at the kth time point.

The course angle error correction module specifically comprises:

Correcting and estimating the course angle of the azimuth turntable to obtain a course angle estimated value; where Δ θ_a，k-1Tracking deviation angle for azimuth at time point k-1 β_k-1The pitch angle of the satellite antenna at the kth-1 time point; Δ H_k-1At the k-1 time pointCourse angle error; k is a radical of_p，HProportional coefficient, k, corrected for proportional integral_i，HAn integration coefficient corrected for proportional integral; h_k-1The estimated value of the course angle of the azimuth turntable at the kth-1 time point; h_kAnd the estimated value of the course angle of the azimuth turntable at the kth time point.

In conclusion, the autonomous measurement and control method and system for the satellite antenna are comprehensive in consideration, simple in steps, reasonable in design and convenient to implement. Aiming at three key working states of satellite communication-in-motion, a simple and practical algorithm is designed one by one to realize an autonomous measurement and control method under the condition of no satellite navigation information assistance, and the method has important practical significance under the conditions of limited satellite navigation information use and military use.

The embodiments in the present description are described in a progressive manner, each embodiment focuses on differences from other embodiments, and the same and similar parts among the embodiments are referred to each other. For the system disclosed by the embodiment, the description is relatively simple because the system corresponds to the method disclosed by the embodiment, and the relevant points can be referred to the method part for description.

The principles and embodiments of the present invention have been described herein using specific examples, which are provided only to help understand the method and the core concept of the present invention; meanwhile, for a person skilled in the art, according to the idea of the present invention, the specific embodiments and the application range may be changed. In view of the above, the present disclosure should not be construed as limiting the invention.

Claims

1. A satellite antenna independent measurement and control method independent of satellite navigation information is characterized in that the satellite antenna independent measurement and control method is based on a micro-mechanical inertia measurement unit; the micro-mechanical inertia measurement unit is arranged on an azimuth turntable of the satellite antenna; the micro-mechanical inertial measurement unit internally comprises a gyroscope and an accelerometer which are orthogonally arranged in three axes; the satellite antenna autonomous measurement and control method comprises the following steps:

acquiring accelerometer measurements acquired by the accelerometer;

2. The satellite antenna autonomous measurement and control method according to claim 1, wherein the determining of the initial pitch angle value and the initial roll angle value of the azimuth turntable according to the accelerometer measurement value specifically includes:

using a formula based on said accelerometer measurements

Determining an initial pitch angle value of the azimuth turntable;

using a formula

Determining the roll angle of the azimuth circleAn initial value; wherein a is_x,a_y,a_zAccelerometer measurements collected for three orthogonally placed accelerometers, respectively; g is a gravity acceleration value; let P₀＝P_a,R₀＝R_a，P₀As an initial value of pitch angle, R₀Is the initial value of the roll angle.

3. The autonomous measurement and control method of a satellite antenna according to claim 2, wherein the starting of satellite finding according to the initial pitch angle value and the initial roll angle value and the determination of the initial heading angle by back-stepping calculation after the satellite is mounted specifically include:

let P be R₀，R＝R₀，H＝H₀When it is 0, the formula is adopted

after scanning for one week, judging whether the signal intensity of the beacon machine received by the satellite beacon is greater than a satellite capturing threshold valueX_TObtaining a first judgment result;

4. The satellite antenna autonomous measurement and control method according to claim 3, wherein the attitude angle fusion estimation is performed according to the accelerometer measurement value and the gyro measurement value to determine a pitch angle estimation value and a roll angle estimation value of the azimuth turntable, and specifically includes:

according to the accelerometer measurement value and the gyro measurement value, adopting a formula

Determining the regulation parameter mu for judging the mobility of carrier, where Δ P is the difference between integral pitch angle of gyro and pitch angle measured by accelerometer, and Δ R is the difference between integral roll angle of gyro and roll angle measured by accelerometer, ξ₁Is a first threshold value, ξ₂Is a second threshold value;

Determining a pitch angle estimated value and a roll angle estimated value of the azimuth turntable; wherein P is_a,k-1The initial value of the pitch angle of the azimuth turntable at the kth-1 time point is obtained; r_a,k-1The initial value of the roll angle of the azimuth turntable at the k-1 time point; p_k-1The estimated value of the pitch angle of the azimuth turntable at the kth-1 time point is obtained; r_k-1The roll angle estimated value of the azimuth turntable at the k-1 time point; k is a radical of_pFor the first fusion weighting parameter, k_iA second fusion weighting parameter; δ t represents a calculation interval time; omega_x，k-1、ω_y，k-1、ω_z，k-1Respectively representing the gyro measurement values acquired by three orthogonally-installed gyros at the k-1 time point; p_kThe estimated value of the pitch angle of the azimuth turntable at the kth time point is obtained; r_kAnd the roll angle estimated value of the azimuth turntable at the kth time point.

5. The autonomous measurement and control method for the satellite antenna according to claim 4, wherein the correcting and estimating of the course angle of the azimuth turntable by using the course angle error to obtain a course angle estimated value specifically comprises:

Correcting and estimating the course angle of the azimuth turntable to obtain a course angle estimated value; where Δ θ_a，k-1Tracking deviation angle for azimuth at time point k-1 β_k-1The pitch angle of the satellite antenna at the kth-1 time point; Δ H_k-1The error is the course angle error of the kth-1 time point; k is a radical of_p，HProportional coefficient, k, corrected for proportional integral_i，HAn integration coefficient corrected for proportional integral; h_k-1The estimated value of the course angle of the azimuth turntable at the kth-1 time point; h_kNavigation of the azimuth turntable at the kth time pointAnd (4) estimating the angle.

6. An autonomous measurement and control system for a satellite antenna independent of satellite navigation information, the system comprising:

7. The satellite antenna autonomous measurement and control system according to claim 6, wherein the module for determining the initial values of the pitch angle and the roll angle specifically comprises:

Determining an initial pitch angle value of the azimuth turntable;

a roll angle initial value determining unit for employing a formulaDetermining the initial value of the roll angle of the azimuth turntable; wherein a is_x,a_y,a_zAccelerometer measurements collected for three orthogonally placed accelerometers, respectively; g is a gravity acceleration value; let P₀＝P_a,R₀＝R_a，P₀As an initial value of pitch angle, R₀Is the initial value of the roll angle.

8. The satellite antenna autonomous measurement and control system according to claim 7, wherein the initial heading angle determining module specifically comprises:

indicating the orientation of the antenna from the antenna turntable coordinate system to the antennaA transformation matrix of the object system;representing a transformation matrix from a geographic coordinate system to a turntable coordinate system, and H, P, R representing a heading angle, a pitch angle and a roll angle of the azimuth turntable respectively;

9. The satellite antenna autonomous measurement and control system according to claim 8, wherein the attitude angle fusion estimation module specifically comprises:

a carrier maneuvering judgment and adjustment parameter calculation unit for measuring the adjustment parameter according to the accelerometerThe magnitude and the gyro measurement value are calculated by formulaDetermining the regulation parameter mu for judging the mobility of carrier, where Δ P is the difference between integral pitch angle of gyro and pitch angle measured by accelerometer, and Δ R is the difference between integral roll angle of gyro and roll angle measured by accelerometer, ξ₁Is a first threshold value, ξ₂Is a second threshold value;

10. The satellite antenna autonomous measurement and control system according to claim 9, wherein the heading angle error correction module specifically comprises: