CN114070403B - Feedforward tracking control method and system for inter-satellite laser communication system - Google Patents
Feedforward tracking control method and system for inter-satellite laser communication system Download PDFInfo
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Abstract
The invention provides a feedforward tracking control method and a feedforward tracking control system for an inter-satellite laser communication system. The method comprises the following steps: acquiring double-star orbit information, and acquiring light spot centroid position information according to real-time feedback information of a detection camera; if the light spot center of mass is in the fine tracking window, calculating the tracking quantity of the fine aiming mechanism according to the center of mass miss distance, and if the light spot center of mass is outside the fine tracking window, calculating the tracking quantity of the coarse aiming mechanism according to the center of mass miss distance; calculating a tracking error caused by relative motion of the satellite under the condition of considering system time delay according to the response control bandwidth of the tracking system and the relative motion speed between the satellites to obtain a feedforward tracking compensation quantity; and if the feedforward tracking compensation quantity is smaller than the control precision of the coarse aiming mechanism, adding the feedforward compensation quantity and the tracking quantity of the fine aiming mechanism to obtain the tracking angle control quantity. The method effectively improves the tracking control precision and improves the steady-state tracking capability of the link between the laser satellites.
Description
Technical Field
The invention relates to the field of optical communication, in particular to a feedforward tracking control method and a feedforward tracking control system for an inter-satellite laser communication system.
Background
The inter-satellite laser communication system has the advantages of high communication speed, large communication capacity, strong anti-interference capability, good confidentiality, light and small terminal and the like, and is one of important development trends of future space networks. At present, the laser communication test among the satellites is carried out successively in China, the United states, europe and the like. In China, satellite laser communication tests are carried out on ocean No. 2 satellites at the earliest time in 2011. Thereafter, the satellite laser communication experiment is successfully carried out in China on the practice of the No. thirteen satellite and the Beidou system. In 2020, the aerospace science and technology group eight hospitals and the aerospace science and technology group cloud engineering successively performed low-orbit satellite laser communication tests. Because the inter-satellite laser communication distance is long, the relative movement speed between the satellites is high, and the divergence angle of signal beams is small, the accuracy and the frequency response characteristic of tracking control directly influence the stability of a link between the laser satellites, which puts high requirements on the tracking technology of an inter-satellite laser communication system.
The traditional tracking method only calculates the tracking control quantity according to the spot centroid position fed back by the camera, and has the problems of high steady-state tracking difficulty of an inter-satellite laser communication link with high-speed relative motion due to tracking response lag, difficulty in long-time stable communication of the inter-satellite laser link and the like, and has adverse effect on high-speed networking of a space laser network system.
Disclosure of Invention
In order to overcome the defects in the prior art, the invention aims to provide a feedforward tracking control method and a feedforward tracking control system for an inter-satellite laser communication system.
In order to achieve the above object, the present invention provides a feedforward tracking control method for an inter-satellite laser communication system, comprising the following steps:
acquiring double-star orbit information, and acquiring light spot centroid position information according to real-time feedback information of a detection camera;
judging whether the light spot mass center of each satellite in the double satellites is in the respective precise tracking window, and if the light spot mass center of any satellite is in the precise tracking window of the satellite, calculating the precise sighting mechanism tracking quantity of the satellite according to the mass center miss quantity; if the light spot center of mass of any satellite is outside the fine tracking window of the satellite, calculating the tracking quantity of the coarse aiming mechanism of the satellite according to the center of mass miss quantity;
calculating a tracking error caused by relative motion of the satellites under the condition of considering system time delay according to the response control bandwidth of the tracking system and the relative motion speed between the satellites to obtain respective feedforward tracking compensation quantities of the double satellites;
judging the magnitude relation between the respective feedforward tracking compensation quantity of the double satellites and the control precision of the coarse aiming mechanism, if the feedforward tracking compensation quantity is greater than the control precision of the coarse aiming mechanism, adding the feedforward compensation quantity of the satellite and the corresponding tracking quantity of the coarse aiming mechanism to obtain the tracking angle control quantity of the satellite, and controlling the coarse aiming mechanism to perform closed-loop tracking control on the satellite according to the tracking angle control quantity; and if the feedforward tracking compensation quantity is smaller than the control precision of the coarse aiming mechanism, adding the feedforward compensation quantity and the tracking quantity of the fine aiming mechanism corresponding to the feedforward compensation quantity to obtain a tracking angle control quantity, and controlling the fine aiming mechanism to perform closed-loop tracking control on the satellite according to the tracking angle control quantity.
The method effectively improves the tracking control precision and improves the steady-state tracking capability of the link between the laser satellites.
In the preferable scheme of the feedforward tracking control method of the inter-satellite laser communication system, the calculation formula of the tracking quantity of the coarse aiming mechanism is as follows
The tracking quantity calculation formula of the fine aiming mechanism is as follows
Wherein, theta Tracking_Az ，θ Tracking_El Respectively tracking the azimuth and the pitch angle; x and y are respectively the miss distance of the spot centroid in the azimuth axis and the pitch axis of the camera; alpha (alpha) ("alpha") coarse_x ，α coarse_y Multiplying coefficient of miss amount and tracking amount of the coarse aiming mechanism; alpha is alpha refined_x ，α refined_y Multiplying coefficient of miss amount and tracking amount of the fine aiming mechanism; beta is a coarse_x ，β coarse_y Adding coefficients of the miss distance and the tracking distance of the coarse aiming mechanism; beta is a beta refined_x ，β refined_y And adding coefficients of the miss distance and the tracking distance of the fine aiming mechanism.
According to the optimal scheme of the feedforward tracking control method of the inter-satellite laser communication system, the calculation steps of the feedforward tracking compensation amount are as follows:
the method comprises the steps of extrapolating data in real time based on double-star orbit information data, calculating double-star position vectors of a near focus coordinate system at the time delay caused by tracking delay response, converting the double-star position vectors from the near focus coordinate system to a geocentric equatorial coordinate system, subtracting the position vectors under the geocentric equatorial coordinate system to obtain feedforward tracking vectors, converting the feedforward tracking vectors from the geocentric equatorial coordinate system to a satellite orbit coordinate system, and calculating to obtain respective feedforward tracking azimuth/pitching angle compensation quantities of double stars.
According to the optimal scheme of the feedforward tracking control method of the inter-satellite laser communication system, the double-satellite position vector solving step of the near focus coordinate system at the time delay caused by tracking lag response comprises the following steps:
mean angle of approach of two stars at time t:
M A (t)＝E A (t)-e A sin(E A (t))，E A (t) is the angle of approach of satellite A at time t, e A Is the orbital eccentricity of satellite a; m B (t)＝E B (t)-e B sin(E B (t))，E B (t) is the angle of approach of satellite B at time t, e B Is the orbital eccentricity of satellite B;
and (3) calculating the position vectors of the double stars at the current moment t under the near focus coordinate system according to the formula:
position vector r of satellite A at current moment t in near-focus coordinate system P A (t)＝(r A (t)·cos(f A (t)))X p A +(r A (t)·sin(f A (t)))Y p A ；
Position vector r of satellite B under near focus coordinate system at current moment t P B (t)＝(r B (t)·cos(f B (t)))X p B +(r B (t)·sin(f B (t)))Y p B ；
Wherein r is A (t) is the modulus of the satellite A position vector at time t, r B (t) is the modulus of the satellite B position vector at time t, and the expression is as follows:
r A (t)＝a A (1-e A cosE A (t))，r B (t)＝a B (1-e B cosE B (t))；
wherein f is A (t) true anomaly of satellite A at time t，f B (t) is the true anomaly of satellite B at time t, and is expressed as follows:
the position vector of the satellite in the near-focus coordinate system at time t is simplified as follows:
Position vectors of the double stars in the equatorial coordinate system of the geocentric at the moment t:
position vector of satellite A under geocentric equatorial coordinate system at time t
Position vector of satellite B in geocentric equatorial coordinate system at time t
Wherein, the transformation matrix of the near focus coordinate system and the geocentric equatorial coordinate system of the satellite A
Transformation matrix of near-focus coordinate system and geocentric equatorial coordinate system of satellite B
Ω A Is the rising intersection right ascension and omega of the satellite A A Orbital perigee angle, i, for satellite A A Is the orbital inclination of the satellite A, omega B Is the rising intersection right ascension and omega of the satellite B B Orbital perigee angle, i, for satellite B B Is the orbital inclination of satellite B, then X p A 、Y p A Transformation to the geocentric equatorial coordinate system is denoted X p lA 、Y p lA The expression is as follows:
X p B 、Y p B transformation to the equatorial frame of the earth's center is denoted X p lB 、Y p lB The expression is as follows:
feedforward tracking vector v of satellite A to satellite B under geocentric equatorial coordinate system at time t I A The computational expression of (a) is:
respectively are position vectors of the satellite A and the satellite B under the geocentric equatorial coordinate system at the moment t;
feedforward tracking vector v of satellite B to satellite A under geocentric equatorial coordinate system at time t I B The calculation expression of (a) is:
the position vectors of the satellite A and the satellite B in the equatorial coordinate system of the geocentric at the moment t are respectively.
Transforming the feedforward tracking vector of the satellite A from the earth center equatorial coordinate system to the satellite orbit coordinate system to obtain the feedforward tracking vector of the satellite A in the orbit coordinate system
Wherein, the first and the second end of the pipe are connected with each other,
，f A (t) is the true anomaly of satellite A at time t;
according to the feedforward tracking vector v of the satellite A under the orbit coordinate system O A And calculating the feedforward tracking pitch angle compensation quantity of the satellite A to the satellite B under the satellite orbit coordinate system at the time t
Feedforward tracking azimuth angle compensation quantity
transforming the feedforward tracking vector of the satellite B into a satellite orbit coordinate system from the geocentric equator coordinate system to obtain the feedforward tracking vector of the satellite B in the orbit coordinate system
Wherein the content of the first and second substances,
，f B (t) is the true anomaly of satellite B at time t;
according to the feedforward tracking vector v of the satellite B in the orbit coordinate system O B And calculating the feedforward tracking pitching angle compensation quantity of the satellite B to the satellite A under the satellite orbit coordinate system at the time t
Feedforward tracking azimuth angle compensation quantity
The invention further provides a feedforward tracking control system of the inter-satellite laser communication system, which comprises a control unit and a storage unit, wherein the control unit is in communication connection with the storage unit, the storage unit is used for storing at least one executable instruction, and the executable instruction enables the processing unit to execute the operation corresponding to the high-precision aiming pointing method of the inter-satellite laser communication system.
The beneficial effects of the invention are: compared with the existing inter-satellite tracking technology, the method can be applied to low-orbit, medium-orbit and high-orbit satellite laser communication links, integrates factors such as satellite relative motion speed, tracking control time delay, facula mass center miss distance and the like, sets feed-forward tracking compensation quantity, effectively improves tracking control precision, and further improves steady-state tracking capability of the laser inter-satellite links.
Additional aspects and advantages of the invention will be set forth in part in the description which follows and, in part, will be obvious from the description, or may be learned by practice of the invention.
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The above and/or additional aspects and advantages of the present invention will become apparent and readily appreciated from the following description of the embodiments, taken in conjunction with the accompanying drawings of which:
FIG. 1 is a flow chart of a feedforward tracking control method of an inter-satellite laser communication system;
FIG. 2 is a schematic view of a camera tracking window;
FIG. 3 is a schematic diagram of coordinate system conversion;
fig. 4 is a diagram showing simulation results of the inter-satellite feedforward tracking control amount.
Detailed Description
Reference will now be made in detail to embodiments of the present invention, examples of which are illustrated in the accompanying drawings, wherein like or similar reference numerals refer to the same or similar elements or elements having the same or similar function throughout. The embodiments described below with reference to the accompanying drawings are illustrative only for the purpose of explaining the present invention and are not to be construed as limiting the present invention.
In the description of the present invention, unless otherwise specified and limited, it is to be noted that the terms "mounted," "connected," and "connected" are to be interpreted broadly, and may be, for example, a mechanical connection or an electrical connection, a communication between two elements, a direct connection, or an indirect connection via an intermediate medium, and specific meanings of the terms may be understood by those skilled in the art according to specific situations.
As shown in fig. 1, the present invention provides a feedforward tracking control method for an inter-satellite laser communication system, and as shown in fig. 1, the feedforward tracking control method for the inter-satellite laser communication system specifically includes the following steps:
firstly, acquiring double-satellite orbit information according to a satellite-borne GPS or GNSS, and acquiring light spot centroid position information according to real-time feedback information of a detection camera.
In this embodiment, the two satellites are respectively referred to as a satellite a and a satellite B, specifically, the two-satellite orbit information acquired by the satellite-borne GPS and the GNSS includes six double-satellite orbits, including 12 parameters in total, of the two-satellite orbit semimajor axis a, the eccentricity e, the near location angle ω, the ascent point right ascension Ω, the inclination angle i, and the time T of passing through the near location, and the two-satellite orbit information is subjected to high-frequency real-time extrapolation. The detection camera is preferably an InGaAs camera, the coarse aiming mechanism is preferably an aerospace high-precision torque motor, and the fine aiming mechanism is preferably a piezoelectric deflection mirror (PZT).
And then judging whether the light spot centroid of each satellite in the double satellites is in a fine tracking window, wherein the fine tracking window is preset in the system, as shown in fig. 2, the upper surface of the fig. 2 is provided with a window, the fine tracking window and a coarse tracking window are divided, light can be detected by the terminal as long as the light enters the window, and the window in which the light is in can be judged, if the light spot centroid of any satellite is in the fine tracking window of the satellite, the tracking quantity of the fine aiming mechanism of the satellite is calculated according to the centroid miss-target quantity, and if the light spot centroid of any satellite is outside the fine tracking window of the satellite, the tracking quantity of the coarse aiming mechanism of the satellite is calculated according to the centroid miss-target quantity. When the center of mass of the light spot is detected, the center of mass miss distance can be obtained.
The calculation formula of the tracking quantity of the coarse aiming mechanism is as follows
The calculation formula of the tracking quantity of the precision aiming mechanism is as follows
Wherein, theta Tracking_Az ，θ Tracking_El Respectively tracking an azimuth angle and a pitch angle; x and y are respectively the miss distance of the light spot centroid in the azimuth axis and the miss distance of the pitching axis in the camera; alpha is alpha coarse_x ，α coarse_y Multiplying coefficient of miss amount and tracking amount of the coarse aiming mechanism; alpha is alpha refined_x ，α refined_y Multiplying coefficient of miss amount and tracking amount of the fine aiming mechanism; beta is a beta coarse_x ，β coarse_y Adding coefficients of the miss distance and the tracking distance of the coarse aiming mechanism; beta is a refined_x ，β refined_y And adding coefficients of the miss distance and the tracking distance of the fine aiming mechanism.
If the tracking quantity of the fine and coarse aiming mechanisms of the satellite A in the double satellites is calculated, substituting the tracking quantity of the spot centroid of the satellite A in the azimuth axis and the tracking quantity of the pitch axis into the camera to calculate; if the tracking quantity of the fine and coarse aiming mechanisms of the satellite B in the double-star is calculated, the tracking quantity is substituted into the miss distance of the spot centroid of the satellite B in the azimuth axis and the miss distance of the pitch axis in the camera for calculation.
And calculating the tracking error caused by the relative motion of the satellite under the condition of considering the system time delay according to the response control bandwidth of the tracking system and the relative motion speed between the satellites to obtain respective feedforward tracking compensation quantities of the double satellites. The response control bandwidth and the relative motion speed between the satellites can be directly obtained from the tracking system.
The step of calculating the feedforward tracking compensation amount mainly comprises the following steps: and extrapolating data in real time based on the double-star orbit information data, calculating double-star position vectors of a near focus coordinate system at the time delay caused by tracking delay response, converting the double-star position vectors into a geocentric equator coordinate system from the near focus coordinate system, subtracting the position vectors under the double-star geocentric equator coordinate system to obtain feed-forward tracking vectors, converting the feed-forward tracking vectors into a satellite orbit coordinate system from the geocentric equator coordinate system, and calculating to obtain respective feed-forward tracking azimuth/pitch angle compensation quantities of the double stars. The coordinate relationships are shown in fig. 3.
Specifically, the time delay t is the position vector r of two satellites under the near-focus coordinate system p A ，r p B The solving steps are as follows:
the average near point angle M of two satellite time t can be iteratively solved according to the Keplerian third law A ，M B The expression is as follows:
wherein μ =398601.2 is an attraction constant, a A Is the orbital semi-major axis, T, of satellite A A Is the time of the past location of satellite A, a B Is the orbital semi-major axis, T, of satellite B B Is the time of the past location of satellite B.
According to the Kepler equation, the approximate point angle E of the two satellites is calculated by a Newton iteration method A ，E B The expression is as follows:
M A (t)＝E A (t)-e A sin(E A (t))，M B (t)＝E B (t)-e B sin(E B (t))，e A is the orbital eccentricity of the satellite A, e B Is the orbital eccentricity of satellite B.
Calculating the position vectors r of the two satellites at the current time t under the near-focus coordinate system according to the formula p A ，r p B 。
r P A (t)＝(r A (t)·cos(f A (t)))X p A +(r A (t)·sin(f A (t)))Y p A ，
r P B (t)＝(r B (t)·cos(f B (t)))X p B +(r B (t)·sin(f B (t)))Y p B ；
Wherein r is A ，r B The expression for the modulus of the two satellite position vectors is as follows:
r A (t)＝a A (1-e A cosE A (t))，r B (t)＝a B (1-e B cosE B (t))；
wherein, f A ，f B The true anomaly for two satellites can be calculated from the partial anomaly, and the expression is as follows:
and finally, the position vectors r of the two satellites at the time t under the near-focus coordinate system p A ，r p B The simplification can be as follows:
position vector of satellite A in double stars under near focus coordinate system at moment t
Position vector of satellite B
Wherein, X p A ，Y p A Is the unit vector, X, of satellite A in its near focus coordinate system p B ，Y p B Is the unit vector, X, of satellite B in its near focus coordinate system p A ＝[1 0 0] T ，Y p A ＝[0 1 0] T ，X p B ＝[1 0 0] T ，Y p B ＝[0 1 0] T 。
The expression of the position vector of the satellite a in the two-star at time t under the geocentric equatorial coordinate system is as follows:
the expression of the position vector of the satellite B in the isocenter equatorial coordinate system at time t in the two stars is as follows:
the expression of the transformation matrix of the near-focus coordinate system and the geocentric equatorial coordinate system of the satellite A is as follows:
The expression of the transformation matrix of the near-focus coordinate system and the geocentric equatorial coordinate system of the satellite B is as follows:
X p B 、Y p B transformation to the geocentric equatorial coordinate system is denoted X p lB 、Y p lB The expression is as follows:
feedforward tracking vector of satellite A to satellite B under earth center equatorial coordinate system at time t
Is calculated as
The position vectors of the satellite A and the satellite B in the equatorial coordinate system of the earth center at the moment t respectively, and similarly, the feedforward tracking vector of the satellite B to the satellite A
To avoid repeating the description, the feedforward tracking vector of satellite A to satellite B is used
The process of the present invention is described in detail for the purpose of example.
The feedforward tracking vector of the satellite A is converted into a satellite orbit coordinate system from a geocentric equator coordinate system, and the feedforward tracking vector of the satellite A under the orbit coordinate system can be obtainedFeed forward tracking vector v O A The calculation expression is
Wherein, the first and the second end of the pipe are connected with each other,
according to the feedforward tracking vector v of the satellite A under the orbit coordinate system O A And calculating the feedforward tracking pitch angle compensation quantity of the satellite A to the satellite B under the satellite orbit coordinate system at the time t
And feedforward tracking azimuth angle compensation quantity
Respectively mean Vo A (t) the first, second and third terms of the vector, vo A The (t) vector is a matrix of 3 x 1 size.
Judging the magnitude relation between feedforward tracking compensation quantity (feedforward tracking pitch angle compensation quantity and feedforward tracking azimuth angle compensation quantity) of the satellite A to the satellite B at the time delay (t time) caused by considering tracking lag response and the control precision of the coarse aiming mechanism, if the feedforward tracking compensation quantity is greater than the control precision of the coarse aiming mechanism, adding the feedforward compensation quantity and the calculated tracking quantity of the coarse aiming mechanism to obtain the tracking angle control quantity (feedforward tracking control quantity of a terminal azimuth axis and a pitch axis) of the satellite A at the time delay t, and controlling the coarse aiming mechanism to perform closed-loop tracking control on the satellite A according to the tracking angle control quantity; and if the feedforward tracking compensation quantity is smaller than the control precision of the coarse aiming mechanism, adding the feedforward compensation quantity and the tracking quantity of the fine aiming mechanism to obtain the tracking angle control quantity of the satellite A at the time delay t, and controlling the fine aiming mechanism to perform closed-loop tracking control on the satellite A according to the tracking angle control quantity.
It should be noted that, when the magnitude relation between the feedforward tracking compensation amount (feedforward tracking pitch angle compensation amount, feedforward tracking position angle compensation amount) and the control accuracy of the coarse aiming mechanism is determined, if one of the feedforward tracking pitch angle compensation amount and the feedforward tracking position angle compensation amount is greater than the control accuracy of the coarse aiming mechanism, that is, the feedforward tracking compensation amount is greater than the control accuracy of the coarse aiming mechanism, and both the feedforward tracking pitch angle compensation amount and the feedforward tracking position angle compensation amount are less than the control accuracy of the coarse aiming mechanism, the feedforward tracking compensation amount is less than the control accuracy of the coarse aiming mechanism.
Similarly, a feedforward tracking vector v of the satellite B at the time t in the orbit coordinate system can be obtained O B The feedforward tracking pitch angle compensation amount theta of the satellite B to the satellite A El B Feedforward tracking azimuth angle compensation amount theta Az B 。
Judging the magnitude relation between feedforward tracking compensation quantity (feedforward tracking pitch angle compensation quantity and feedforward tracking azimuth angle compensation quantity) of the satellite B to the satellite A at a time delay (time t) caused by considering tracking lag response and the control precision of the coarse aiming mechanism, if the feedforward tracking compensation quantity is greater than the control precision of the coarse aiming mechanism, adding the feedforward compensation quantity and the calculated tracking quantity of the coarse aiming mechanism to obtain the tracking angle control quantity (feedforward tracking control quantity of a terminal azimuth axis and a pitch axis) of the satellite B at the time delay t, and controlling the coarse aiming mechanism to perform closed-loop tracking control on the satellite B according to the tracking angle control quantity; and if the feedforward tracking compensation quantity is smaller than the control precision of the coarse aiming mechanism, adding the feedforward compensation quantity and the tracking quantity of the fine aiming mechanism to obtain the tracking angle control quantity of the satellite B at the time delay t, and controlling the fine aiming mechanism to perform closed-loop tracking control on the satellite B according to the tracking angle control quantity.
The feedforward tracking control quantity of the terminal azimuth axis and the terminal pitch axis of the satellite A at the time delay t is
The feedforward tracking control quantity of the terminal azimuth axis and the terminal pitch axis of the time delay time t satellite B is
According to the method, feedforward tracking compensation quantity simulation is carried out on the different-orbit satellite, wherein the height of the orbit of the satellite is about 1100km, the inclination angle of the orbit is 86.5 degrees, the phase difference of the satellite is 7.5 degrees, and the right ascension of the orbit of the satellite is 0 degree and 15 degrees respectively. As shown in fig. 4, the amount of feed forward tracking compensation for the terminal azimuth axis and the pitch axis.
The invention also provides an embodiment of the feedforward tracking control system of the inter-satellite laser communication system, which comprises a control unit and a storage unit, wherein the control unit is in communication connection with the storage unit, and the storage unit is used for storing at least one executable instruction, and the executable instruction enables the processing unit to execute the operation corresponding to the high-precision aiming pointing method of the inter-satellite laser communication system.
In the description herein, references to the description of the term "one embodiment," "some embodiments," "an example," "a specific example," or "some examples," etc., mean that a particular feature, structure, material, or characteristic described in connection with the embodiment or example is included in at least one embodiment or example of the invention. In this specification, the schematic representations of the terms used above do not necessarily refer to the same embodiment or example. Furthermore, the particular features, structures, materials, or characteristics described may be combined in any suitable manner in any one or more embodiments or examples.
While embodiments of the present invention have been shown and described, it will be understood by those of ordinary skill in the art that: various changes, modifications, substitutions and alterations can be made to the embodiments without departing from the principles and spirit of the invention, the scope of which is defined by the claims and their equivalents.
Claims (8)
1. A feedforward tracking control method for an inter-satellite laser communication system is characterized by comprising the following steps:
acquiring double-star orbit information, and acquiring light spot centroid position information according to real-time feedback information of a detection camera;
judging whether the light spot mass center of each satellite in the double satellites is in the respective precise tracking window, and if the light spot mass center of any satellite is in the precise tracking window of the satellite, calculating the precise sighting mechanism tracking quantity of the satellite according to the mass center miss quantity; if the light spot centroid of any satellite is outside the fine tracking window of the satellite, calculating the tracking quantity of the coarse aiming mechanism of the satellite according to the centroid miss quantity;
according to the response control bandwidth of the tracking system and the relative movement speed between the satellites, calculating the tracking error caused by the relative movement of the satellites under the condition of considering the system time delay to obtain respective feedforward tracking compensation quantities of the double satellites;
judging the magnitude relation between the respective feedforward tracking compensation quantity of the double satellites and the control precision of the coarse aiming mechanism, if the feedforward tracking compensation quantity is greater than the control precision of the coarse aiming mechanism, adding the feedforward tracking compensation quantity of the satellite and the corresponding tracking quantity of the coarse aiming mechanism to obtain the tracking angle control quantity of the satellite, and controlling the coarse aiming mechanism to perform closed-loop tracking control on the satellite according to the tracking angle control quantity; and if the feedforward tracking compensation quantity is smaller than the control precision of the coarse aiming mechanism, adding the feedforward tracking compensation quantity and the tracking quantity of the fine aiming mechanism corresponding to the feedforward tracking compensation quantity to obtain a tracking angle control quantity, and controlling the fine aiming mechanism to perform closed-loop tracking control on the satellite according to the tracking angle control quantity.
2. The feedforward tracking control method for the inter-satellite laser communication system according to claim 1, wherein the calculation formula of the tracking amount of the coarse aiming mechanism is
The tracking quantity calculation formula of the fine aiming mechanism is as follows
Wherein, theta Tracking_Az ，θ Tracking_El Respectively tracking the azimuth and the pitch angle; x and y are respectively the miss distance of the spot centroid in the azimuth axis and the pitch axis of the camera; alpha (alpha) ("alpha") coarse_x ，α coarse_y For tracking miss distance and coarse aiming mechanismA magnitude multiplication coefficient; alpha is alpha refined_x ，α refined_y The multiplication coefficient of the miss distance and the tracking distance of the fine aiming mechanism is obtained; beta is a coarse_x ，β coarse_y Adding coefficients of the miss distance and the tracking distance of the coarse aiming mechanism; beta is a refined_x ，β refined_y And adding coefficients of the miss distance and the tracking distance of the fine aiming mechanism.
3. A feedforward tracking control method for an inter-satellite laser communication system according to claim 1, wherein the calculation of the feedforward tracking compensation amount includes:
the method comprises the steps of extrapolating data in real time based on double-star orbit information data, calculating double-star position vectors of a near focus coordinate system at the time delay caused by tracking delay response, converting the double-star position vectors from the near focus coordinate system to a geocentric equatorial coordinate system, subtracting the position vectors under the geocentric equatorial coordinate system to obtain feedforward tracking vectors, converting the feedforward tracking vectors from the geocentric equatorial coordinate system to a satellite orbit coordinate system, and calculating to obtain respective feedforward tracking azimuth and pitching angle compensation quantities of double stars.
4. A feedforward tracking control method of an inter-satellite laser communication system according to claim 3, wherein the two-star position vector solving step of the near-focus coordinate system at the time delay caused by the tracking lag response is:
mean anomaly of two stars at time t:
M A (t)＝E A (t)-e A sin(E A (t))，E A (t) is time tAngle of approach of star A, e A Is the orbital eccentricity of satellite a; m B (t)＝E B (t)-e B sin(E B (t))，E B (t) is the angle of approach of satellite B at time t, e B Is the orbital eccentricity of the satellite B;
and (3) calculating the position vectors of the double stars at the current moment t under the near focus coordinate system according to the formula:
position vector r of satellite A at current moment t in near-focus coordinate system P A (t)＝(r A (t)·cos(f A (t)))X p A +(r A (t)·sin(f A (t)))Y p A ；
Position vector r of satellite B under near focus coordinate system at current moment t P B (t)＝(r B (t)·cos(f B (t)))X p B +(r B (t)·sin(f B (t)))Y p B ；
Wherein r is A (t) is the modulus of the satellite A position vector at time t, r B (t) is the modulus of the satellite B position vector at time t, and the expression is as follows:
r A (t)＝a A (1-e A cosE A (t))，r B (t)＝a B (1-e B cosE B (t))；
wherein f is A (t) true anomaly of satellite A at time t, f B (t) is the true anomaly of satellite B at time t, and is expressed as follows:
the position vector of the satellite in the near-focus coordinate system at time t is simplified as follows:
X A p 、Y A p are unit vectors, X, of the satellite A in a near focus coordinate system B p 、Y B p Are unit vectors, X, of the satellite B in a near focus coordinate system A p ＝X B p ＝[1 0 0] T ，Y A p ＝Y B p ＝[0 1 0] T 。
5. A feed-forward tracking control method for an inter-satellite laser communication system as claimed in claim 3, wherein the position vectors of the double stars in the geocentric equatorial coordinate system at time t are:
position vector of satellite A under geocentric equatorial coordinate system at time t
Position vector of satellite B in geocentric equatorial coordinate system at time t
Wherein, the transformation matrix of the near focus coordinate system and the geocentric equatorial coordinate system of the satellite A
Transformation matrix of near-focus coordinate system and geocentric equatorial coordinate system of satellite B
Ω A The right ascension point of satellite A、ω A Orbital perigee angle, i, for satellite A A Is the orbital inclination of the satellite A, a A Is the orbital semi-major axis, Ω, of satellite A B Is the rising intersection right ascension and omega of the satellite B B Orbital perigee angle, i, for satellite B B Is the orbital inclination of the satellite B, a B Is the orbital semi-major axis of satellite B, then X p A 、Y p A Transformation to the geocentric equatorial coordinate system is denoted X p lA 、Y p lA The expression is as follows:
X p B 、Y p B transformation to the geocentric equatorial coordinate system is denoted X p lB 、Y p lB The expression is as follows:
6. the feedforward tracking control method for the inter-satellite laser communication system according to claim 3, wherein the feedforward tracking vector v of the satellite A to the satellite B under the geocentric equatorial coordinate system at the time t is the vector I A The computational expression of (a) is:
respectively are position vectors of the satellite A and the satellite B under the geocentric equatorial coordinate system at the moment t;
feedforward tracking vector v of satellite B to satellite A under geocentric equatorial coordinate system at time t I B The computational expression of (a) is:
the position vectors of the satellite A and the satellite B in the equatorial coordinate system of the geocentric at the time t are respectively.
7. A feedforward tracking control method of an inter-satellite laser communication system according to claim 3, wherein the feedforward tracking vector of the satellite A is transformed from the earth center equatorial coordinate system to the satellite orbit coordinate system to obtain the feedforward tracking vector v of the satellite A in the orbit coordinate system O A ，
Wherein, the first and the second end of the pipe are connected with each other,
f A (t) is the true anomaly, Ω, of satellite A at time t A Is the rising crossing right ascension, omega, of satellite A A Is the orbital perigee angle, v, of satellite A I A (t) is a feedforward tracking vector of the satellite A to the satellite B under the geocentric equatorial coordinate system at the moment t;
according to the feedforward tracking vector v of the satellite A under the orbit coordinate system O A And calculating the feedforward tracking pitching angle compensation quantity of the satellite A to the satellite B under the satellite orbit coordinate system at the time t
Feedforward tracking azimuth angle compensation quantity
transforming the feedforward tracking vector of the satellite B from the earth center equatorial coordinate system to the satellite orbit coordinate system to obtain the feedforward tracking vector of the satellite B in the orbit coordinate system
Wherein the content of the first and second substances,
f B (t) is the true anomaly, Ω, of satellite B at time t B Is the rising cross-point right ascension, omega, of satellite B B Is the orbital perigee angle, v, of satellite B I B (t) is a feedforward tracking vector of the satellite B to the satellite A under the geocentric equatorial coordinate system at the moment t;
according to the feedforward tracking vector v of the satellite B in the orbit coordinate system O B And calculating the feedforward tracking pitch angle compensation quantity of the satellite B to the satellite A under the satellite orbit coordinate system at the time t
Feedforward tracking azimuth angle compensation quantity
8. A feedforward tracking control system of an inter-satellite laser communication system, comprising a processing unit and a storage unit, wherein the processing unit is communicatively connected with the storage unit, and the storage unit is used for storing at least one executable instruction, and the executable instruction causes the processing unit to execute an operation corresponding to the feedforward tracking control method of the inter-satellite laser communication system according to any one of claims 1 to 7.
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| CN113568133A (en) * | 2021-07-20 | 2021-10-29 | 西安工业大学 | Advanced aiming and precise tracking two-in-one system and light beam pointing method |
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