CN113641182A - High-precision aiming and pointing method and system for inter-satellite laser communication system - Google Patents

High-precision aiming and pointing method and system for inter-satellite laser communication system Download PDF

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CN113641182A
CN113641182A CN202110850342.2A CN202110850342A CN113641182A CN 113641182 A CN113641182 A CN 113641182A CN 202110850342 A CN202110850342 A CN 202110850342A CN 113641182 A CN113641182 A CN 113641182A
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梅博
杨中华
李梦男
朱林玉
郭茂
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Dongfanghong Satellite Mobile Communication Co Ltd
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    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
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    • G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
    • G—PHYSICS
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    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/10—Simultaneous control of position or course in three dimensions
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    • H04—ELECTRIC COMMUNICATION TECHNIQUE
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    • H04—ELECTRIC COMMUNICATION TECHNIQUE
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Abstract

The invention provides a high-precision aiming pointing method and system for an inter-satellite laser communication system. Acquiring double-star orbit information to obtain position vectors of double stars under a near-focus coordinate system, converting the position vectors under the near-focus coordinate system into an equatorial coordinate system, and performing vector subtraction on the position vectors of the double stars under the equatorial coordinate system to obtain an inter-star aiming vector; converting the inter-satellite aiming vector from an equatorial coordinate system to a satellite orbit coordinate system, calculating to obtain a satellite orbit aiming angle, and calculating aiming angle compensation quantity; converting the inter-satellite aiming vector from a satellite orbit coordinate system to a satellite platform coordinate system and then to a terminal coordinate system; and calculating an inter-satellite initial aiming angle, and adding the inter-satellite initial aiming angle and the aiming angle compensation quantity to obtain an inter-satellite aiming pointing angle. The method effectively improves the aiming pointing precision and promotes the quick link establishment/re-link efficiency of the inter-satellite laser communication system.

Description

High-precision aiming and pointing method and system for inter-satellite laser communication system
Technical Field
The invention relates to a high-precision aiming pointing method and a high-precision aiming pointing system for an inter-satellite laser communication system, and belongs to the technical field of aiming, capturing and tracking of satellite laser communication terminals.
Background
The inter-satellite laser communication system has the advantages of high communication speed, large communication capacity, strong anti-interference capability, good confidentiality, terminal miniaturization and the like, and is one of important development trends of future space networks. At present, the laser communication test among satellites is carried out successively in China, America and Europe. In China, satellite laser communication tests are carried out on ocean No. 2 satellites at the earliest possible year of 2012. In 8 months in 2020, Starlink in the United states has completed the first in-orbit test of inter-satellite laser communication, and in the same year, the eight institute 802 of China aerospace science and technology group has also successfully developed the low-orbit satellite laser communication test. Because the inter-satellite laser communication has the characteristics of long communication distance, small signal beam divergence angle and the like, the method puts high requirements on the aiming pointing technology of the inter-satellite laser communication system.
The traditional inter-satellite aiming method has the problems of low aiming pointing precision, large uncertain domain range generated by aiming angle error, long link building time, high light spot tracking difficulty and the like, and has adverse effect on the rapid link building/re-linking capability of an inter-satellite laser communication system. In order to adapt to the development trend of low power consumption and light size of the current satellite laser communication terminal and the overall design requirement of rapid chain establishment/complex chain of the inter-satellite laser network, and better meet the large-scale construction and networking requirements of China on low-orbit constellations in the future, the research on the high-precision aiming pointing technology of the inter-satellite laser communication system is more urgent.
Disclosure of Invention
In order to overcome the defects in the prior art, the invention aims to provide a high-precision aiming pointing method and system for an inter-satellite laser communication system.
In order to achieve the above object, the present invention provides a high-precision aiming pointing method for an inter-satellite laser communication system, comprising the following steps:
acquiring double-satellite orbit information through a satellite-borne GPS and a GNSS, and calculating the orbit information in real time to obtain position vectors of double satellites under a near focus coordinate system;
converting the position vectors of the double stars under the near-focus coordinate system into the equatorial coordinate system to obtain the position vectors of the double stars under the equatorial coordinate system;
carrying out vector subtraction on position vectors of the double stars under an equatorial coordinate system to obtain respective inter-star aiming vectors of the double stars;
converting the inter-satellite aiming vectors of the double stars from an equatorial coordinate system to a satellite orbit coordinate system, and calculating to obtain respective satellite orbit aiming angles of the double stars;
calculating the delay time generated by optical transmission delay and satellite-borne computer information processing delay caused by relative motion between satellites, and calculating the satellite orbit aiming angle deviation caused by the delay time to obtain respective aiming angle compensation quantities of the two satellites;
converting the inter-satellite aiming vectors of the double satellites from a satellite orbit coordinate system to a satellite platform coordinate system according to attitude information measured by a satellite platform, and then converting the inter-satellite aiming vectors of the double satellites from the satellite platform coordinate system to a terminal coordinate system;
and calculating the initial aiming angle between the two stars according to the inter-star aiming vector of the two stars under the terminal coordinate system, and adding the initial aiming angle between the two stars and the aiming angle compensation amount to obtain the inter-star aiming pointing angle of the two stars.
The method comprehensively considers factors such as satellite orbit information, satellite platform attitude measurement information, the installation position of the laser communication terminal on the satellite platform, optical transmission delay and information processing delay caused by relative motion between satellites and the like, effectively improves aiming pointing accuracy, and improves the rapid link establishment/re-link efficiency of the inter-satellite laser communication system.
The preferred scheme of the high-precision aiming pointing method is as follows: mean satellite near-point angle of two stars at pointing time t
Figure BDA0003182216470000031
Wherein μ ═ 398601.2 is an attraction constant, a is an orbit semi-major axis of the satellite, and T is the time of the satellite passing by;
m (t), e (t), esin (e (t)), e (t), an angle of approach of the satellite at time t, and e, an orbital eccentricity of the satellite;
calculating the position vector r of the satellite at the current time t under the near-focus coordinate system according to the formulaP(t)＝(r(t)·cos(f(t)))Xp+(r(t)·sin(f(t)))Yp；
Where r (t) is a module of the satellite position vector at time t, and the expression is as follows:
r(t)＝a(1-ecosE(t))；
wherein f (t) is the true near point angle of the satellite at the time t, and the expression is as follows:
Figure BDA0003182216470000032
the position vector of the satellite in the near-focus coordinate system at the final pointing time t is simplified as follows:
Figure BDA0003182216470000033
Xp、Ypare unit vectors, X, of the satellite in a near focus coordinate systemp＝[1 0 0]T，Yp＝[0 1 0]T。
The preferred scheme of the high-precision aiming pointing method is as follows: position vector of satellite in equatorial coordinate system at time t
Figure BDA0003182216470000034
Wherein, the transformation matrix of the near focus coordinate system and the equator coordinate system of the satellite
Figure BDA0003182216470000035
Omega is the ascension point of the satellite, omega is the orbital perigee angle of the satellite, and i is the orbital inclination angle of the satellite, then Xp，YpTransformation to the geocentric equatorial coordinate system is denoted Xp l，Yp lThe expression is asThe following:
Figure BDA0003182216470000041
the preferred scheme of the high-precision aiming pointing method is as follows: respective intersatellite aiming vectors v of double stars under an equatorial coordinate system at the moment tI(t)＝r”I(t)-r'I(t)，r'I(t) is the position vector of the body satellite in the double star at the time of t under the equatorial coordinate system, r "IAnd (t) is a position vector of the opposite satellite in the double stars under the equatorial coordinate system at the time t.
The preferred scheme of the high-precision aiming pointing method is as follows: respective aiming vectors of the two stars are converted into a satellite orbit coordinate system from an equator coordinate system to obtain the aiming vectors v of each satellite in the two stars under the orbit coordinate systemO(t)＝RI-O(t)vI(t) wherein,
Figure BDA0003182216470000042
f (t) is the true near point angle of the body satellite at the time t;
according to aiming vector v of the body satellite under the orbit coordinate systemOAnd calculating the aiming pitch angle of the body satellite pointing to the opposite satellite under the satellite orbit coordinate system at the moment t
Figure BDA0003182216470000043
Aiming azimuth
Figure BDA0003182216470000044
Angular distance to two satellites
Figure BDA0003182216470000045
vO (1)(t)、vO (2)(t)、vO (3)(t) denotes the first, second and third terms of the vector at time t, vo (t), respectively, which is a matrix of size 3 x 1.
The heightThe preferred scheme of the precision aiming pointing method is as follows: lag time when two-satellite-center body satellite points to the opposite satellite
Figure BDA0003182216470000051
Where c is the speed of light, tdataThe signal data response and processing time of the body satellite are shown, wherein rho (t) refers to the slant distance between the body satellite and the two satellites at the pointing moment t;
recalculating the pitching aiming angles theta of the two satellites under the satellite orbit coordinate systems of the two satellites after considering the lag timeE(t +. DELTA.t), azimuth aiming angle thetaA(t +. DELTA.t) and the pitching and azimuth aiming angles theta of the two satellites at the time t, which are calculated when the lag time is not considered, under the satellite orbit coordinate systemE(t)，θAAnd (t) subtracting to obtain aiming angle compensation quantity of each of the two satellites at the time t.
The preferred scheme of the high-precision aiming pointing method is as follows: according to the attitude information measured by the satellite platform, the inter-satellite aiming vector is converted from the satellite orbit coordinate system to the satellite platform coordinate system,
the aiming vector of each satellite in the double-star under the orbit coordinate system is converted into an expression v under the satellite platform coordinate systemS(t)＝RO-S(t)vO(t)，vO(t) aiming vectors of the satellites under the orbit coordinate system at the aiming moment;
Figure BDA0003182216470000052
wherein, in the step (A),
Figure BDA0003182216470000053
θG0roll angle, theta, transmitted for the respective satelliteF0Pitch angle, theta, transmitted for the respective satelliteP0The yaw angle transmitted for the respective satellite,
Figure BDA0003182216470000054
rolling angular velocity transmitted for each satellite,
Figure BDA0003182216470000055
Pitch rate transmitted for each satellite,
Figure BDA0003182216470000056
Yaw rate transmitted for the respective satellite.
The preferred scheme of the high-precision aiming pointing method is as follows: aiming vector v under terminal coordinate systemL(t)＝RS-LvS(t)；
If the terminal coordinate system is coincident with the x-axis of the satellite platform coordinate system, the matrix is converted
Figure BDA0003182216470000061
Alpha is an angle which needs to rotate around the x axis when the satellite platform coordinate system is overlapped with the terminal coordinate system;
if the terminal coordinate system is coincident with the y-axis of the satellite platform coordinate system, the matrix is converted
Figure BDA0003182216470000062
Alpha is an angle which needs to rotate around the y axis when the satellite platform coordinate system is overlapped with the terminal coordinate system;
if the terminal coordinate system is coincident with the z-axis of the satellite platform coordinate system, the matrix is converted
Figure BDA0003182216470000063
Alpha is an angle which needs to rotate around the z axis when the satellite platform coordinate system is overlapped with the terminal coordinate system.
The preferred scheme of the high-precision aiming pointing method is as follows: according to aiming vector v under terminal coordinate systemLObtaining the initial pitching and azimuth aiming angle theta of the satellite at the time tEl(t)，θAz(t), the expression is as follows:
Figure BDA0003182216470000064
Figure BDA0003182216470000065
wherein v isL (1)、vL (2)、vL (3)Are each vLA first term, a second term, a third term of the vector;
aiming pointing angle between stars is
Figure BDA0003182216470000066
Figure BDA0003182216470000067
Aiming at the pitch angle between the stars at the time t,
Figure BDA0003182216470000068
for inter-satellite aiming pointing azimuth angle at time t, Δ θEl(t) the amount of aiming pitch angle compensation at time t, Δ θAzAnd (t) is the aiming pointing azimuth angle compensation amount at the time t.
The application also provides a high-precision aiming pointing system for 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.
The invention has the beneficial effects that: the invention is applied to low-orbit, middle-orbit and high-orbit satellite laser communication links, utilizes the characteristics of large pointing range and high pointing accuracy of a terminal aiming mechanism (motor), and completes the conversion and correction of aiming pointing angles among a geocentric equatorial coordinate system, a near-focus coordinate system, a satellite orbit coordinate system, a satellite platform coordinate system and a laser terminal coordinate system by comprehensively considering factors such as satellite orbit information, satellite platform attitude measurement information, the installation position of a laser communication terminal on a satellite platform, optical transmission delay and information processing delay caused by relative motion among satellites and the like, thereby effectively improving the pointing accuracy of the intersatellite laser aiming.
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 process flow of the present invention;
FIG. 2 is a schematic diagram of coordinate system transformation;
FIG. 3 is a diagram illustrating the deviation of the aiming angle caused by time delay;
FIG. 4(a) is a diagram showing simulation results of initial aiming angles of adjacent satellites in the same orbit;
FIG. 4(b) is a diagram showing the simulation result of the initial aiming angle of the different-orbit adjacent satellite;
FIG. 5(a) is a diagram showing the simulation result of the pointing angle compensation of the adjacent satellite in the same orbit;
fig. 5(b) is a diagram showing a simulation result of the amount of the off-orbit adjacent satellite aiming angle compensation.
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 high-precision aiming pointing method for an inter-satellite laser communication system, which completes the calculation of a high-precision aiming pointing angle of the inter-satellite laser communication system by acquiring the dual-satellite orbit information of a GPS and a GNSS, the attitude information measured by a satellite platform, considering the installation position of a laser communication terminal on the satellite platform, the optical transmission delay and the information processing delay caused by the relative motion between satellites, and relating to the conversion and correction of an aiming vector between a geocentric equatorial coordinate system, a near-focus coordinate system, a satellite orbit coordinate system, a satellite platform coordinate system, and a laser terminal coordinate system, as shown in fig. 2.
As shown in fig. 1, the method specifically comprises:
step 1, calculating the orbit information in real time through the double-satellite orbit information acquired by the satellite-borne GPS and the GNSS, and solving the position vector of the double satellites under a near focus coordinate system.
In this embodiment, the two satellites are respectively referred to as a satellite a and a satellite B, and the two-satellite orbit information acquired by the satellite-borne GPS and the GNSS includes six dual-satellite orbits, including 12 parameters of a semi-major axis a of the dual-satellite orbit, an eccentricity e, an angle of approach ω, a right ascension Ω of a rising intersection, an inclination i, and a time T of passing through the approach. And further extrapolating and updating the six orbits at the two-star pointing time t according to the six orbits obtained in real time. In the present application, when a satellite a in a double star points to a satellite B at a pointing time t, the satellite a is referred to as a body satellite, and the satellite B is referred to as an opposite satellite; when the satellite B points to the satellite a, the satellite B is referred to as a body satellite, and the satellite a is referred to as an opposite satellite.
When the double stars aim at the pointing direction, the position vector r of the double stars under the near focus coordinate system at the pointing time tp A，rp BThe solving steps are as follows:
the average near point angle M of the pointing moments t of the two satellites can be iteratively solved according to the Kepler's third lawA，MBThe expression is as follows:
Figure BDA0003182216470000091
where μ is 398601.2 as an attractive force constant, aAIs the orbital semi-major axis, T, of satellite AAIs the time of the past location of satellite A, aBIs the orbital semi-major axis, T, of the satellite BBIs 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 methodA，EBThe expression is: mA(t)＝EA(t)-eAsin(EA(t)),MB(t)＝EB(t)-eBsin(EB(t))，eAIs the orbital eccentricity of the satellite A, eBIs 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 formulap A，rp B。
rP A(t)＝(rA(t)·cos(fA(t)))Xp A+(rA(t)·sin(fA(t)))Yp A，
rP B(t)＝(rB(t)·cos(fB(t)))Xp B+(rB(t)·sin(fB(t)))Yp B；
Wherein r isA，rBThe expression, modulo the two satellite position vectors, is as follows:
rA(t)＝aA(1-eAcosEA(t))，rB(t)＝aB(1-eBcosEB(t))；
wherein f isA，fBThe true near point angle for two satellites can be obtained from the approximate near point angle, and the expression is as follows:
Figure BDA0003182216470000093
Figure BDA0003182216470000094
and finally, the position vectors r of the two satellites at the time t in the near-focus coordinate systemp A，rp BThe simplification can be as follows:
when pointing toPosition vector of satellite A in double stars under near focus coordinate system at moment t
Figure BDA0003182216470000101
Position vector of satellite B in two satellites
Figure BDA0003182216470000102
Wherein the content of the first and second substances,
Figure BDA0003182216470000103
is a unit vector, X, of the satellite A in a near-focus coordinate systemp B，Yp BIs a unit vector, X, of satellite B in a near focus coordinate systemp A＝[1 0 0]T，Yp A＝[0 1 0]T，Xp B＝[1 0 0]T，Yp B＝[0 1 0]T。
And 2, converting the position vectors of the double stars under the near-focus coordinate system into the equatorial coordinate system to obtain the position vectors of the double stars under the equatorial coordinate system.
Specifically, the position vector of the satellite a in the equatorial coordinate system in the two stars at time t:
Figure BDA0003182216470000104
the position vector of the satellite B in the double star at the moment t under the equatorial coordinate system is as follows:
Figure BDA0003182216470000105
wherein, the transformation matrix of the near focus coordinate system and the equator coordinate system of the satellite A
Figure BDA0003182216470000106
ΩAIs the rising intersection right ascension and omega of the satellite AAIs the orbital perigee angle, i, of satellite AAIs the orbital inclination of satellite a.
Rotation of near-focus coordinate system and equatorial coordinate system of satellite BChange matrix
Figure BDA0003182216470000107
ΩBIs the rising intersection right ascension and omega of the satellite BBOrbital perigee angle, i, for satellite BBOrbital inclination of satellite B, then Xp A、Yp ATransformation to the geocentric equatorial coordinate system is denoted Xp lA、Yp lAThe expression is as follows:
Figure BDA0003182216470000108
Xp B、Yp Btransformation to the geocentric equatorial coordinate system is denoted Xp lB、Yp lBThe expression is as follows:
Figure BDA0003182216470000111
and 3, carrying out vector subtraction on the position vectors of the double stars in the equatorial coordinate system to obtain an inter-star aiming vector.
Specifically, the laser aiming vector of the satellite A pointing to the satellite B under the equatorial coordinate system at the moment t is calculated
Figure BDA0003182216470000112
The calculation expression is:
Figure BDA0003182216470000113
respectively, the position vectors of the satellite A and the satellite B in the equatorial coordinate system at the time t, and the aiming vector of the satellite B pointing to the satellite A
Figure BDA0003182216470000114
And 4, converting the inter-satellite aiming vector from an equatorial coordinate system to a satellite orbit coordinate system, and calculating to obtain the satellite orbit aiming angle.
To avoid repetition, a satellite A is aimed at a satellite BQuasi-vector
Figure BDA0003182216470000115
The process of the present invention is described in detail for the purpose of example.
Specifically, aiming vectors from the satellite A to the satellite B in the orbit coordinate system are obtained after aiming vectors from the satellite A to the satellite B in the double-star are converted into the satellite orbit coordinate system from the equatorial coordinate system
Figure BDA0003182216470000116
Wherein the content of the first and second substances,
Figure BDA0003182216470000117
，fA(t) is the true anomaly of satellite A at time t;
according to aiming vector v of satellite A in its orbit coordinate systemO AAnd solving the aiming pitch angle of the satellite A pointing to the satellite B under the satellite orbit coordinate system at the moment t
Figure BDA0003182216470000119
Aiming azimuth
Figure BDA00031822164700001110
Angular distance to two satellites
Figure BDA0003182216470000121
Figure BDA0003182216470000122
Respectively mean VoA(t) the first, second and third terms of the vector, VoAThe (t) vector is a matrix of 3 x 1 size.
Similarly, the aiming pitch angle of the satellite B pointing to the satellite A is obtained
Figure BDA0003182216470000123
Aiming azimuthCorner
Figure BDA0003182216470000124
And the slant distance rho between the two satellitesB(t)。
And 5, considering the optical transmission delay caused by the relative motion between the satellites and the information processing delay of the satellite-borne computer system, calculating the delay time generated by the delay, and calculating the aiming angle deviation of the satellite orbit caused by the delay time according to the steps 1 to 4, wherein the aiming angle deviation is shown in figure 3, and the aiming angle compensation quantity is obtained.
Specifically, the lag time when satellite a points to satellite B in the two satellites:
Figure BDA0003182216470000125
where c is the speed of light, tA dataSignal data response and processing time, p, for satellite AA(t) is the slant range between satellite A and two satellites at pointing time t; lag time when satellite B points at satellite a in two stars:
Figure BDA0003182216470000126
where c is the speed of light, tB dataSignal data response and processing time, p, for satellite BB(t) is the slant range from satellite B to both satellites at pointing time t. In fact, the change amount of the inter-satellite distance in the lag time is very small, and within the controllable error range, the rho can be considered asA(t)≈ρA(t+ΔtA)，ρB(t)≈ρB(t+ΔtB) Then, the calculation expression of the hysteresis time Δ t can be simplified as: lag time when satellite A points to satellite B in two satellites
Figure BDA0003182216470000127
Lag time when satellite B points to satellite A in two satellites
Figure BDA0003182216470000128
Recalculating the satellite A and the satellite B in the satellite orbit coordinate system after the lag time (t + delta t moment) according to the steps 1-4Lower elevation aiming angle thetaE A(t+△t)、θE B(t +. DELTA.t), azimuth aiming angle thetaA A(t+△t)、θA B(t +. DELTA.t) and the pitching aiming angles theta of the satellites A and B under the satellite orbit coordinate system at the time t obtained by calculation in the step 4E A(t)、θE B(t), azimuth aiming angle θA A(t)、θA B(t) subtracting to obtain the compensation quantity delta theta of the pitching aiming angle when the satellite A points to the satellite BEl AAzimuth aiming angle compensation quantity delta thetaAz AThe specific expression is as follows:
Figure BDA0003182216470000131
pitching aiming angle compensation quantity delta theta when satellite B points to satellite AEl BAzimuth aiming angle compensation quantity delta thetaAz BThe specific expression is as follows:
Figure BDA0003182216470000132
and 6, taking the deviation between the satellite orbit coordinate system and the satellite platform coordinate system into consideration, and converting the aiming vector from the satellite orbit coordinate system to the satellite platform coordinate system according to the attitude information measured by the satellite platform.
Specifically, aiming vectors of the satellite A in the orbital coordinate system of the two satellites are converted into an expression v in the coordinate system of the satellite platformS A(t)＝RO-S A(t)vO A(t)，
Figure BDA0003182216470000133
The aiming vector of the satellite B in the two satellites under the orbital coordinate system is converted into an expression v under the satellite platform coordinate systemS B(t)＝RO-S B(t)vO B(t)，
Figure BDA0003182216470000134
Wherein the content of the first and second substances,
Figure BDA0003182216470000135
the roll angle transmitted for satellite a star,
Figure BDA0003182216470000136
the pitch angle transmitted for satellite a satellite,
Figure BDA0003182216470000137
the yaw angle transmitted for satellite a satellite,
Figure BDA0003182216470000138
rolling angular velocity transmitted for satellite A satellite,
Figure BDA0003182216470000139
Pitch rate transmitted for satellite a satellite,
Figure BDA00031822164700001310
A yaw rate sent for satellite a house traffic;
Figure BDA0003182216470000141
the roll angle transmitted for satellite B satellite traffic,
Figure BDA0003182216470000142
the pitch angle transmitted for satellite B satellite traffic,
Figure BDA0003182216470000143
the yaw angle transmitted for satellite B satellite,
Figure BDA0003182216470000144
rolling angular velocity transmitted for satellite B,
Figure BDA0003182216470000145
Pitch rate transmitted for satellite B,
Figure BDA0003182216470000146
The yaw rate transmitted for satellite B satellite.
And 7, converting the inter-satellite aiming vector from the satellite platform coordinate system to the terminal coordinate system by considering the installation position of the terminal on the satellite platform.
Aiming vector v under terminal coordinate systemL A,BThe expression of (a) is as follows:
vL A,B(t)＝RS-L A,BvS A,B(t)
in the formula, a conversion matrix RS-L A,BThe installation angle of the terminal on the satellite platform is determined, and in general, the terminal is installed in such a manner that a certain axis coincides with the satellite platform.
If the terminal coordinate system coincides with the x axis of the satellite platform coordinate system, the two coordinate systems can coincide only by rotating the satellite platform coordinate system by an angle around the x axis, and if the rotation angle is alpha, the conversion matrix expression is as follows:
Figure BDA0003182216470000149
if the y axis of the terminal coordinate system is coincident with the y axis of the satellite platform coordinate system, and the rotation angle is alpha, the conversion matrix expression is as follows:
Figure BDA00031822164700001410
if the terminal coordinate system is coincident with the z axis of the satellite platform coordinate system and the rotation angle is alpha, the conversion matrix expression is as follows:
Figure BDA00031822164700001411
aiming vector v of satellite A under terminal coordinate systemL A(t)＝RS-L AvS A(t); in formula (I) a conversion matrix RS-L ABy the terminalIn determining the installation angle on the satellite platform, the terminal is generally installed so that one axis coincides with the satellite platform.
If the terminal coordinate system is coincident with the x-axis of the satellite platform coordinate system, the matrix is converted
Figure BDA0003182216470000151
Alpha is an angle which needs to rotate around the x axis when the satellite platform coordinate system is overlapped with the terminal coordinate system;
if the terminal coordinate system is coincident with the y-axis of the satellite platform coordinate system, the matrix is converted
Figure BDA0003182216470000152
Alpha is an angle which needs to rotate around the y axis when the satellite platform coordinate system is overlapped with the terminal coordinate system;
if the terminal coordinate system is coincident with the z-axis of the satellite platform coordinate system, the matrix is converted
Figure BDA0003182216470000153
Alpha is an angle which needs to rotate around the z axis when the satellite platform coordinate system is coincided with the terminal coordinate system;
aiming vector v of satellite B under terminal coordinate systemL B(t)＝RS-L BvS B(t); in formula (I) a conversion matrix RS-L BThe installation angle of the terminal on the satellite platform is determined, and in general, the terminal is installed in such a manner that a certain axis coincides with the satellite platform.
If the terminal coordinate system is coincident with the x-axis of the satellite platform coordinate system, the matrix is converted
Figure BDA0003182216470000154
Alpha is an angle which needs to rotate around the x axis when the satellite platform coordinate system is overlapped with the terminal coordinate system;
if the terminal coordinate system is coincident with the y-axis of the satellite platform coordinate system, the matrix is converted
Figure BDA0003182216470000155
Alpha is an angle which needs to rotate around the y axis when the satellite platform coordinate system is overlapped with the terminal coordinate system;
if the terminal coordinate system is coincident with the z-axis of the satellite platform coordinate system, the matrix is converted
Figure BDA0003182216470000161
Alpha is an angle which needs to rotate around the z axis when the satellite platform coordinate system is overlapped with the terminal coordinate system.
And 8, calculating an initial aiming angle between the satellites according to the aiming vector between the satellites under the terminal coordinate system, and adding the initial aiming angle to the aiming angle compensation quantity obtained in the step 5 to finally obtain the high-precision aiming pointing angle between the satellites.
In particular, according to the aiming vector v under the terminal coordinate systemL AThe initial pitching and azimuth aiming angle theta of the satellite A at the time t can be obtainedEl A(t)，θAz A(t), the expression is as follows:
Figure BDA0003182216470000162
Figure BDA0003182216470000163
Figure BDA0003182216470000164
are respectively as
Figure BDA0003182216470000165
The first term, the second term, and the third term of the vector.
According to aiming vector v under terminal coordinate systemL BThe initial pitching and azimuth aiming angle theta of the satellite B at the time t can be obtainedEl B(t)，θAz B(t), the expression is as follows:
Figure BDA0003182216470000166
Figure BDA0003182216470000167
Figure BDA0003182216470000168
are respectively as
Figure BDA0003182216470000169
The first term, the second term, and the third term of the vector.
Adding the initial aiming angles of the satellite A and the satellite B at the time t with the aiming angle compensation quantity obtained in the step 5 respectively to finally obtain a high-precision inter-satellite biaxial aiming pointing angle when the satellite A aims at the satellite B
Figure BDA00031822164700001610
Figure BDA00031822164700001611
For inter-satellite aiming pointing pitch angle of satellite a at time t,
Figure BDA00031822164700001612
aiming pointing azimuth angle between satellites of the satellite A at the time t;
the high-precision inter-satellite biaxial aiming pointing angle when the satellite B aims at the satellite A is
Figure BDA0003182216470000171
Figure BDA0003182216470000172
For inter-satellite aiming pointing pitch angle for satellite B at time t,
Figure BDA0003182216470000173
pointing azimuth angle for inter-satellite aiming of satellite B at time t.
The application also provides a high-precision aiming pointing system for 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.
Aiming pointing simulation is performed between adjacent satellites in the same orbit/different orbits according to the method. The height of the satellite orbit is 1200km, the orbit eccentricity is 0.0012, the orbit inclination angle is 86.5 degrees, the phase difference of the in-orbit satellite is 30 degrees, the phase difference of the out-of-orbit satellite is 7.5 degrees, the ascension of the orbit of the in-orbit satellite is 0 degree, and the ascension of the orbit of the out-of-orbit satellite is 0 degree and 15.18 degrees respectively.
As shown in FIG. 4(a), the satellite is operated for about 109 minutes for one turn, and the in-orbit satellite is initially aimed at the pointing azimuth angle θAzAlways keeping 0 degree, and the deviation is not more than 2 e-15; pitch angle thetaElThe range is-15.13 ° -14.95 °. As shown in FIG. 4(b), the off-orbit satellite initially aims at the azimuth angle θAzRanging from-62.18 DEG to 62.21 DEG; pitch angle thetaElThe range is-8.71 ° -4.10 °.
As shown in FIG. 5(a), the azimuth angle compensation amount Δ θ of the in-orbit satelliteAzAlways keeping 0 degree, the deviation is not more than 2.8 e-15; delta thetaElThe pitch angle compensation amount ranges from-3.75 e-06 degrees to 2.62e-06 degrees. As shown in FIG. 5(b), the off-orbit satellite azimuth angle compensation amount Δ θAzRanging from-9.32 e-04 ° -9.30 e-04 °; compensation quantity delta theta of pitching angleElRanging from-5.75 e-05 deg. -5.76 e-05 deg..
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 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 (10)

1. A high-precision aiming and pointing method for an inter-satellite laser communication system is characterized by comprising the following steps:
acquiring double-satellite orbit information through a satellite-borne GPS and a GNSS, and calculating the orbit information in real time to obtain position vectors of double satellites under a near focus coordinate system;
converting the position vectors of the double stars under the near-focus coordinate system into the equatorial coordinate system to obtain the position vectors of the double stars under the equatorial coordinate system;
carrying out vector subtraction on position vectors of the double stars under an equatorial coordinate system to obtain respective inter-star aiming vectors of the double stars;
converting the inter-satellite aiming vectors of the double stars from an equatorial coordinate system to a satellite orbit coordinate system, and calculating to obtain respective satellite orbit aiming angles of the double stars;
calculating the delay time generated by optical transmission delay and satellite-borne computer information processing delay caused by relative motion between satellites, and calculating the satellite orbit aiming angle deviation caused by the delay time to obtain respective aiming angle compensation quantities of the two satellites;
converting the inter-satellite aiming vectors of the double satellites from a satellite orbit coordinate system to a satellite platform coordinate system according to attitude information measured by a satellite platform, and then converting the inter-satellite aiming vectors of the double satellites from the satellite platform coordinate system to a terminal coordinate system;
and calculating the initial aiming angle between the two stars according to the inter-star aiming vector of the two stars under the terminal coordinate system, and adding the initial aiming angle between the two stars and the aiming angle compensation amount to obtain the inter-star aiming pointing angle of the two stars.
2. The method of claim 1, wherein the satellite mean-near-point angle of the two stars at pointing time t is
Figure FDA0003182216460000021
Wherein μ ═ 398601.2 is an attraction constant, a is an orbit semi-major axis of the satellite, and T is the time of the satellite passing by;
m (t), e (t), esin (e (t)), e (t), an angle of approach of the satellite at time t, and e, an orbital eccentricity of the satellite;
calculating the position vector r of the satellite at the current time t under the near-focus coordinate system according to the formulaP(t)＝(r(t)·cos(f(t)))Xp+(r(t)·sin(f(t)))Yp；
Where r (t) is a modulus of the satellite position vector at time t, the expression is as follows:
r(t)＝a(1-ecosE(t))；
wherein f (t) is the true near point angle of the satellite at the time t, and the expression is as follows:
Figure FDA0003182216460000022
the position vector of the satellite in the near-focus coordinate system at the final pointing time t is simplified as follows:
Figure FDA0003182216460000023
Xp、Ypare unit vectors, X, of the satellite in a near focus coordinate systemp＝[1 0 0]T，Yp＝[0 1 0]T。
3. The method of claim 2, wherein the position vector of the satellite in the equatorial coordinate system at time t is pointed
Figure FDA0003182216460000024
Wherein, the transformation matrix of the near focus coordinate system and the equator coordinate system of the satellite
Figure FDA0003182216460000025
Omega is the ascension point of the satellite, omega is the orbital perigee angle of the satellite, and i is the orbital inclination angle of the satellite, then Xp，YpTransformation to the geocentric equatorial coordinate system is denoted Xp l，Yp lThe expression is as follows:
Figure FDA0003182216460000031
4. the method of claim 1, wherein the inter-satellite aiming vectors v of the two satellites are respectively pointed under the equatorial coordinate system at the time tI(t)＝r″I(t)-r′I(t)，r′I(t) is the position vector of the body satellite in the double star at the time of t in the equatorial coordinate system, r ″IAnd (t) is a position vector of the opposite satellite in the double stars under the equatorial coordinate system at the time t.
5. The method of claim 1, wherein the aiming vectors of the satellites in the two satellites are transformed from the equatorial coordinate system to the orbital coordinate system of the satellite, and the aiming vectors v of the satellites in the two satellites in the orbital coordinate system are obtainedO(t)＝RI-O(t)vI(t) wherein,
Figure FDA0003182216460000032
f (t) is the true near point angle of the body satellite at the time t;
according to aiming vector v of the body satellite under the orbit coordinate systemOFinding the orbit of the satellite at time tAiming pitch angle of body satellite pointing to opposite satellite under road coordinate system
Figure FDA0003182216460000033
Aiming azimuth
Figure FDA0003182216460000034
Angular distance to two satellites
Figure FDA0003182216460000035
vO (1)(t)、vO (2)(t)、vO (3)(t) denotes the first, second and third terms of the vector at time t, vo (t), respectively, which is a matrix of size 3 x 1.
6. The high-precision aiming and pointing method for the inter-satellite laser communication system as claimed in claim 1, wherein the dead time of the two-satellite body satellite pointing to the opposite satellite
Figure FDA0003182216460000036
Where c is the speed of light, tdataThe signal data response and processing time of the body satellite are shown, wherein rho (t) refers to the slant distance between the body satellite and the opposite satellite at the pointing moment t;
recalculating the pitching aiming angles theta of the two satellites under the satellite orbit coordinate systems of the two satellites after considering the lag timeE(t +. DELTA.t), azimuth aiming angle thetaA(t +. DELTA.t) and the pitching and azimuth aiming angles theta of the two satellites at the time t, which are calculated when the lag time is not considered, under the satellite orbit coordinate systemE(t)，θAAnd (t) subtracting to obtain aiming angle compensation quantity of each of the two satellites at the time t.
7. The high precision aiming pointing method for inter-satellite laser communication system as claimed in claim 1, wherein the inter-satellite aiming vector is converted from the satellite orbit coordinate system to the satellite platform coordinate system based on the attitude information measured by the satellite platform,
the aiming vector of each satellite in the double-star under the orbit coordinate system is converted into an expression v under the satellite platform coordinate systemS(t)＝RO-S(t)vO(t)，vO(t) aiming vectors of the satellites at the aiming time t in the orbit coordinate system;
Figure FDA0003182216460000041
wherein, in the step (A),
Figure FDA0003182216460000042
θG0roll angle, theta, transmitted for the respective satelliteF0Pitch angle, theta, transmitted for the respective satelliteP0The yaw angle transmitted for the respective satellite,
Figure FDA0003182216460000043
rolling angular velocity transmitted for each satellite,
Figure FDA0003182216460000044
Pitch rate transmitted for each satellite,
Figure FDA0003182216460000045
Yaw rate transmitted for the respective satellite.
8. The method of claim 1, wherein the aiming vector v is a terminal coordinate systemL(t)＝RS-LvS(t)；
If the terminal coordinate system is coincident with the x-axis of the satellite platform coordinate system, the matrix is converted
Figure FDA0003182216460000051
Alpha is a coordinate system of the satellite platform and a terminal seatThe angle of the satellite platform coordinate system which needs to rotate around the x axis when the coordinate systems are overlapped;
if the terminal coordinate system is coincident with the y-axis of the satellite platform coordinate system, the matrix is converted
Figure FDA0003182216460000052
Alpha is an angle which needs to rotate around the y axis when the satellite platform coordinate system is overlapped with the terminal coordinate system;
if the terminal coordinate system is coincident with the z-axis of the satellite platform coordinate system, the matrix is converted
Figure FDA0003182216460000053
Alpha is an angle which needs to rotate around the z axis when the satellite platform coordinate system is overlapped with the terminal coordinate system.
9. The method of claim 1, wherein aiming pointing is performed according to aiming vector v in terminal coordinate systemLObtaining the initial pitching and azimuth aiming angle theta of the satellite at the time tEl(t)，θAz(t), the expression is as follows:
Figure FDA0003182216460000054
Figure FDA0003182216460000055
wherein v isL (1)、vL (2)、vL (3)Are each vLA first term, a second term, a third term of the vector;
aiming pointing angle between stars is
Figure FDA0003182216460000056
Figure FDA0003182216460000057
Aiming at the pitch angle between the stars at the time t,
Figure FDA0003182216460000058
for inter-satellite aiming pointing azimuth angle at time t, Δ θEl(t) the amount of aiming pitch angle compensation at time t, Δ θAzAnd (t) is the aiming pointing azimuth angle compensation amount at the time t.
10. A high-precision aiming pointing system for an inter-satellite laser communication system, which comprises a control unit and a storage unit, wherein the control unit is connected with the storage unit in a communication way, and the storage unit is used for storing at least one executable instruction which causes the processing unit to execute the corresponding operation of the high-precision aiming pointing method of the inter-satellite laser communication system according to any one of claims 1 to 9.
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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114070403A (en) * 2021-11-15 2022-02-18 东方红卫星移动通信有限公司 Feedforward tracking control method and system for inter-satellite laser communication system
CN115833919A (en) * 2023-02-14 2023-03-21 鹏城实验室 Iterative learning-based inter-satellite laser link rapid networking method and related equipment
CN116659543A (en) * 2023-06-21 2023-08-29 中国人民解放军61540部队 Satellite position and attitude estimation method and device based on remote sensing satellite orbit number

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101719792A (en) * 2009-11-27 2010-06-02 哈尔滨工业大学 Platform for simulating link satellite photo-communication terminal-to-terminal relative sighting angle movement
CN101873169A (en) * 2010-05-14 2010-10-27 北京大学 Receiving system in wireless optical communication and signal receiving method thereof
CN102142899A (en) * 2011-01-06 2011-08-03 中国科学院上海光学精密机械研究所 Satellite laser communication composite tracking and aiming vibration suppression device and control method
CN102207614A (en) * 2011-03-16 2011-10-05 中国科学院上海技术物理研究所 Deep space optical communication tracking and aiming system and method
CN111669224A (en) * 2020-06-02 2020-09-15 武汉光谷航天三江激光产业技术研究院有限公司 Inter-satellite laser communication aiming deviation on-orbit measurement and correction method

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101719792A (en) * 2009-11-27 2010-06-02 哈尔滨工业大学 Platform for simulating link satellite photo-communication terminal-to-terminal relative sighting angle movement
CN101873169A (en) * 2010-05-14 2010-10-27 北京大学 Receiving system in wireless optical communication and signal receiving method thereof
CN102142899A (en) * 2011-01-06 2011-08-03 中国科学院上海光学精密机械研究所 Satellite laser communication composite tracking and aiming vibration suppression device and control method
CN102207614A (en) * 2011-03-16 2011-10-05 中国科学院上海技术物理研究所 Deep space optical communication tracking and aiming system and method
CN111669224A (en) * 2020-06-02 2020-09-15 武汉光谷航天三江激光产业技术研究院有限公司 Inter-satellite laser communication aiming deviation on-orbit measurement and correction method

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
陶坤宇;凌志华;舒锐;周彦平;: "一种卫星间跟瞄方向的控制策略与实现", 宇航学报 *

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114070403A (en) * 2021-11-15 2022-02-18 东方红卫星移动通信有限公司 Feedforward tracking control method and system for inter-satellite laser communication system
CN114070403B (en) * 2021-11-15 2022-12-02 东方红卫星移动通信有限公司 Feedforward tracking control method and system for inter-satellite laser communication system
CN115833919A (en) * 2023-02-14 2023-03-21 鹏城实验室 Iterative learning-based inter-satellite laser link rapid networking method and related equipment
CN116659543A (en) * 2023-06-21 2023-08-29 中国人民解放军61540部队 Satellite position and attitude estimation method and device based on remote sensing satellite orbit number
CN116659543B (en) * 2023-06-21 2024-05-07 中国人民解放军61540部队 Satellite position and attitude estimation method and device based on remote sensing satellite orbit number

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