Antenna direction determination method, electronic equipment and medium

Abstract

The invention discloses an antenna direction determination method, electronic equipment and a computer readable medium, and relates to the technical field of satellite communication. The method comprises the following steps: acquiring high-orbit satellite data information from a low-orbit satellite platform carrying the satellite-borne relay terminal and low-orbit satellite data information of a low-orbit satellite carrying the low-orbit satellite platform; predicting the position of the high-orbit satellite in future set time according to the high-orbit satellite data information; determining the position of the low earth orbit satellite in the future set time according to the low earth orbit satellite data information; and determining a pointing angle of an antenna carried on the low-orbit satellite according to the low-orbit satellite data information, the position of the high-orbit satellite in the future set time and the position of the low-orbit satellite in the future set time. The method has the advantages of autonomous prediction and reduction of dependence on ground data.

Description

Technical Field

[0001] The present invention relates to the field of satellite communication technology, and in particular to an antenna pointing determination method, electronic equipment and computer-readable medium. Background Art

[0002] The satellite-borne relay terminal can rely on the relay system to achieve real-time measurement and control and medium- and low-speed data transmission throughout the entire period, and realize high-speed data backhaul through the digital transmission channel.

[0003] In direct communication links between low Earth orbit (LEO) satellites and geostationary orbit (GEO) satellites, the pointing antenna on LEO needs to be adjusted in real time to point toward GEO. However, LEO moves at high speed relative to the ground, while GEO is not completely stationary relative to the ground and exhibits orbital drift. Current antenna pointing calculation methods rely on ground stations updating GEO orbit data every few hours and sending this updated GEO orbit data via an uplink to LEO to maintain stable communication between LEO and GEO.

[0004] This method of frequently updating GEO orbit data not only increases the data transmission burden, but if the ground station fails to update GEO orbit data in a timely manner, LEO will use outdated orbit data, which may lead to inaccurate antenna pointing and communication failure. Summary of the Invention

[0005] The present invention aims to solve one of the technical problems in the related art to a certain extent. To this end, the present invention provides an antenna pointing determination method for a satellite-borne relay terminal, which has the advantages of autonomous prediction and significantly reduced dependence on real-time ground data.

[0006] To achieve the above-mentioned object, as a first aspect of the present invention, a method for determining antenna pointing is provided for use in a satellite-borne relay terminal. The method comprises:

[0007] Acquiring high-orbit satellite data information from a low-orbit satellite platform carrying the onboard relay terminal, and low-orbit satellite data information of a low-orbit satellite carrying the low-orbit satellite platform;

[0008] Predicting the position of the high-orbit satellite at a set time in the future based on the high-orbit satellite data information;

[0009] Determining the position of the low-orbit satellite at the set future time according to the low-orbit satellite data information;

[0010] The pointing angle of the antenna carried on the low-orbit satellite is determined based on the low-orbit satellite data information, the position of the high-orbit satellite at a set time in the future, and the position of the low-orbit satellite at a set time in the future.

[0011] Optionally, the high-orbit satellite data includes the orbital elements of the high-orbit satellite in the geocentric inertial coordinate system and the epoch time corresponding to the orbital elements; the orbital elements include the semi-major axis, eccentricity, orbital inclination, right ascension of the ascending node, argument of perigee, and mean anomaly;

[0012] The predicting the position of the high-orbit satellite at a set future time based on the high-orbit satellite data information includes:

[0013] Predicting the mean anomaly at the future set time based on the mean anomaly corresponding to the epoch, the future set time, formula (1) and formula (2);

[0014] M=M0+n×(t-t0) (1)

[0015]

[0016] Where M0 is the mean anomaly at the epoch time; n is the mean angular velocity; t0 is the epoch time; t is the future set time; M is the mean anomaly at the future set time; μ is the gravitational constant of the central celestial body; a is the semi-major axis;

[0017] Predict the eccentric anomaly at the future setting time based on the mean anomaly, eccentricity, and formula (3);

[0018] M=Ee×sinE (3)

[0019] Where M is the mean anomaly at a future set time, e is the eccentricity, and E is the eccentric anomaly at a future set time;

[0020] The position of the high-orbit satellite in the geocentric inertial coordinates is determined according to the eccentric anomaly angle, semi-major axis, eccentricity, direction cosine matrix, and formula (4) at a future set time; wherein the direction cosine matrix includes a first direction cosine matrix and a second direction cosine matrix, the first direction cosine matrix is ​​determined according to formula (5), and the second direction cosine matrix is ​​determined according to formula (6);

[0021]

[0022] Where a is the semi-major axis; E is the anomaly angle at the future set time; e is the eccentricity; P is the first direction cosine matrix; Q is the second direction cosine matrix; r ECI is the position of the high-orbit satellite in the Earth-centered inertial coordinate system;

[0023]

[0024]

[0025] Where ω is the argument of perigee; Ω is the right ascension of the ascending node; i is the orbital inclination; P is the first direction cosine matrix; and Q is the second direction cosine matrix.

[0026] Optionally, the low-orbit satellite data information includes a coordinate conversion relationship between a geocentric inertial coordinate system and a low-orbit satellite antenna coordinate system; the position of the high-orbit satellite at a set future time and the position of the low-orbit satellite at a set future time are both located in the geocentric inertial coordinate system;

[0027] The determining the pointing angle of the antenna carried on the low-orbit satellite according to the low-orbit satellite data information, the position of the high-orbit satellite at a set time in the future, and the position of the low-orbit satellite at the set time in the future includes:

[0028] Determine a position vector in a geocentric inertial coordinate system based on a position of a high-orbit satellite at a set time in the future and a position of a low-orbit satellite at a set time in the future;

[0029] The pointing angle of the antenna carried on the low-orbit satellite is determined according to the position vector and the coordinate conversion relationship.

[0030] Optionally, the coordinate transformation relationship includes a first coordinate transformation relationship between the geocentric inertial coordinate system and the low-orbit satellite orbit coordinate system, a second coordinate transformation relationship between the low-orbit satellite orbit coordinate system and the low-orbit satellite body coordinate system, and a third coordinate transformation relationship between the low-orbit satellite body coordinate system and the low-orbit satellite antenna coordinate system;

[0031] Determining the pointing angle of the antenna carried on the low-orbit satellite according to the position vector and the coordinate transformation relationship includes:

[0032] Converting the position vector in the geocentric inertial coordinate system to the low-orbit satellite orbit coordinate system according to the first coordinate conversion relationship to obtain the position vector in the low-orbit satellite orbit coordinate system;

[0033] According to the second coordinate transformation relationship, the position vector in the low-orbit satellite orbital coordinate system is transformed into the low-orbit satellite body coordinate system to obtain the position vector in the low-orbit satellite body coordinate system;

[0034] According to the third coordinate transformation relationship, the position vector in the low-orbit satellite body coordinate system is transformed into the low-orbit satellite antenna coordinate system to obtain the position vector in the low-orbit satellite antenna coordinate system;

[0035] The antenna pointing angle is determined according to the position vector in the low-orbit satellite antenna coordinate system.

[0036] Optionally, determining the antenna pointing angle according to the position vector in the low-orbit satellite antenna coordinate system includes:

[0037] Obtaining an azimuth angle and an off-axis angle in the low-orbit satellite antenna coordinate system according to the position vector in the low-orbit satellite antenna coordinate system and trigonometric function transformation;

[0038] Antenna pointing is performed according to the azimuth angle and the off-axis angle.

[0039] Optionally, determining the position vector in the geocentric inertial coordinate system according to the position of the high-orbit satellite at a set time in the future and the position of the low-orbit satellite at a set time in the future includes:

[0040] According to formula (7), the position vector of the high-orbit satellite at the future set time is subtracted from the position vector of the low-orbit satellite at the future set time to obtain the position vector in the geocentric inertial coordinate system;

[0041] dr ECI =r ECI -r0 ECI (7)

[0042] Among them, r ECI Set the position of the high-orbit satellite at a certain time in the future; r0 ECI Set the position of a low-orbit satellite at a certain time in the future; dr ECI is the position vector in the geocentric inertial coordinate system.

[0043] Optionally, the low-orbit satellite data information further includes position and velocity information of the low-orbit satellite in a geocentric inertial coordinate system, and determining the position of the low-orbit satellite at the future set time based on the low-orbit satellite data information includes:

[0044] According to the position and speed of the low-orbit satellite in the geocentric inertial coordinate system at the current time and the relationship between the future time and the current time, the position of the low-orbit satellite at the future time is obtained.

[0045] Optionally, obtaining the position of the low-orbit satellite at a future time based on the position and velocity of the low-orbit satellite in the geocentric inertial coordinate system at the current time and the relationship between the future time and the current time further includes:

[0046] Smoothing of the position and velocity in the geocentric inertial coordinate system.

[0047] As a second aspect of the present invention, an electronic device is provided, comprising:

[0048] one or more processors;

[0049] A memory having one or more computer programs stored thereon, wherein when the one or more computer programs are executed by the one or more processors, the one or more processors implement the antenna pointing determination method provided according to the first aspect of the present invention.

[0050] As a third aspect of the present invention, a computer-readable medium is provided, on which a computer program is stored, wherein when the computer program is executed by a processor, the antenna pointing determination method provided by the first aspect of the present invention is implemented.

[0051] The antenna pointing determination method for a satellite-borne relay terminal provided by the present invention can obtain high-orbit satellite data information and low-orbit satellite data information from a low-orbit satellite platform carrying the satellite-borne relay terminal. First, the position information of the high-orbit satellite at a set time in the future is predicted based on the epoch time and its corresponding orbital element number and the Kepler equation. Second, the real-time position and velocity information of the low-orbit satellite is obtained and the smoothed data is converted to an inertial coordinate system. Subsequently, the position vector between the high-orbit satellite and the low-orbit satellite in the inertial coordinate system is calculated, and based on the coordinate conversion relationship between the low-orbit satellite and the high-orbit satellite and the coordinate conversion relationship between the low-orbit satellite and the antenna on the low-orbit satellite, the position vector is converted into a rotation angle in the antenna coordinate system. Finally, the antenna on the low-orbit satellite is pointed toward the high-orbit satellite based on the rotation angle.

[0052] The antenna pointing determination method provided by the present invention can autonomously predict the antenna pointing angle for future times. Predicting the operating status of high-orbit satellites significantly reduces the need for real-time data updates from ground stations and effectively alleviates data transmission pressure on communication links. Furthermore, by acquiring the operating status of low-orbit satellites in real time and smoothing the data, antenna pointing accuracy is improved. This autonomous antenna pointing angle prediction method also avoids pointing deviations caused by outdated orbital data. It maintains stable communication even in the event of temporary interruptions or delays at the ground station, improving communication success rate and reliability.

[0053] These features and advantages of the present invention will be further disclosed in the following detailed description and accompanying drawings. The preferred embodiments and means of the present invention will be fully illustrated in conjunction with the accompanying drawings, but are not intended to limit the technical solutions of the present invention. Furthermore, although multiple features, elements, and components may be present in each of the following text and accompanying drawings, they may be labeled with different symbols or numbers for convenience, but all represent components with the same or similar structure or function. BRIEF DESCRIPTION OF THE DRAWINGS

[0054] The present invention will be further described below in conjunction with the accompanying drawings:

[0055] Figure 1 A flowchart of an implementation method of an antenna pointing determination method provided by the present invention;

[0056] Figure 2 Flowchart of step S120 of the antenna direction determination method provided by the present invention;

[0057] Figure 3Flowchart of step S140 of the antenna direction determination method provided by the present invention;

[0058] Figure 4 Flowchart of step S142 of the antenna direction determination method provided by the present invention;

[0059] Figure 5 Flowchart of step S142d of the antenna direction determination method provided by the present invention;

[0060] Figure 6 Flowchart of step S141 of the antenna direction determination method provided by the present invention;

[0061] Figure 7 Flowchart of step S130 of the antenna direction determination method provided by the present invention;

[0062] Figure 8 A diagram of a relay system corresponding to the antenna direction determination method provided by the present invention;

[0063] Figure 9 A module diagram of an electronic device provided by the present invention;

[0064] Figure 10 A schematic diagram of a computer-readable medium provided by the present invention.

[0065] Description of Reference Numerals

[0066] Among them, 101 is a processor; 102 is a memory; 103 is an I / O interface; and 104 is a bus. DETAILED DESCRIPTION

[0067] The following describes embodiments of the present invention in detail, examples of which are shown in the accompanying drawings, wherein the same or similar reference numerals throughout represent the same or similar elements or elements having the same or similar functions. The embodiments described in the embodiments are intended to explain the present invention and are not to be construed as limiting the present invention.

[0068] References in this specification to "one embodiment," "an example," or "an example" mean that a particular feature, structure, or characteristic described in connection with the embodiment itself can be included in at least one embodiment disclosed herein. The appearances of the phrase "in one embodiment" in various places in the specification are not necessarily all referring to the same embodiment.

[0069] Current calculations of antenna pointing angles for LEO satellites relative to GEO satellites have significant limitations: They rely heavily on frequently updated GEO satellite orbit data from ground stations. This mechanism not only wastes communication resources but also creates a rigid dependency on the timeliness and reliability of ground station data updates. Any data update delays or transmission interruptions can lead to antenna misalignment and potentially disrupt the communication link.

[0070] In view of this, as a first aspect of the present invention, a method for determining antenna direction is provided for a satellite-borne relay terminal, such as Figure 1 As shown, the method includes:

[0071] In step S110, high-orbit satellite data information and low-orbit satellite data information of the low-orbit satellite carrying the low-orbit satellite platform are acquired from the low-orbit satellite platform carrying the onboard relay terminal;

[0072] In step S120, the position of the high-orbit satellite at a set time in the future is predicted based on the high-orbit satellite data information;

[0073] In step S130, the position of the low-orbit satellite at the future set time is determined according to the low-orbit satellite data information;

[0074] In step S140 , the pointing angle of the antenna carried on the low-orbit satellite is determined based on the low-orbit satellite data information, the position of the high-orbit satellite at a set time in the future, and the position of the low-orbit satellite at a set time in the future.

[0075] The antenna pointing determination method for a satellite-borne relay terminal provided by the present invention can obtain high-orbit satellite data information and low-orbit satellite data information from a low-orbit satellite platform carrying the satellite-borne relay terminal. First, the position information of the high-orbit satellite at a set time in the future is predicted based on the epoch time and its corresponding orbital element number and the Kepler equation. Second, the real-time position and velocity information of the low-orbit satellite is obtained and the smoothed data is converted to an inertial coordinate system. Subsequently, the position vector between the high-orbit satellite and the low-orbit satellite in the inertial coordinate system is calculated, and based on the coordinate conversion relationship between the low-orbit satellite and the high-orbit satellite and the coordinate conversion relationship between the low-orbit satellite and the antenna on the low-orbit satellite, the position vector is converted into a rotation angle in the antenna coordinate system. Finally, the antenna on the low-orbit satellite is pointed toward the high-orbit satellite based on the rotation angle.

[0076] The antenna pointing determination method provided by this invention can autonomously predict antenna pointing angles for future times. By predicting the operational status of high-orbit satellites, the need for real-time data updates from ground stations can be significantly reduced, effectively alleviating data transmission pressure on communication links. This autonomous antenna pointing angle prediction method also avoids pointing deviations caused by outdated orbital data. It can maintain stable communication even in the event of temporary interruptions or delays at the ground station, improving communication success rate and reliability.

[0077] The antenna pointing determination method provided by the present invention can autonomously predict the operating status of a high-orbit satellite, as an optional implementation of step S120, such as Figure 2 As shown, the high-orbit satellite data includes the epoch time of the high-orbit satellite in the geocentric inertial coordinate system and the orbital elements corresponding to the epoch time; the orbital elements include the semi-major axis, eccentricity and mean anomaly;

[0078] The predicting the position of the high-orbit satellite at a set future time based on the high-orbit satellite data information includes:

[0079] In step S121, the mean anomaly at the future set time is predicted based on the mean anomaly corresponding to the epoch time, the future set time, formula (1) and formula (2);

[0080] M=M0+n×(t-t0) (1)

[0081]

[0082] Where M0 is the mean anomaly at the epoch time; n is the mean angular velocity; t0 is the epoch time; t is the future set time; M is the mean anomaly at the future set time; μ is the gravitational constant of the central celestial body; a is the semi-major axis;

[0083] In step S122, the eccentric anomaly at the future set time is predicted based on the mean anomaly, the eccentricity, and formula (3);

[0084] M=Ee×sinE (3)

[0085] Where M is the mean anomaly at a future set time, e is the eccentricity, and E is the eccentric anomaly at a future set time;

[0086] In step S123, the position of the high-orbit satellite in the geocentric inertial coordinates is determined according to the eccentric anomaly angle, semi-major axis, eccentricity, direction cosine matrix, and formula (4) at the future set time; wherein the direction cosine matrix includes a first direction cosine matrix and a second direction cosine matrix, the first direction cosine matrix is ​​determined according to formula (5), and the second direction cosine matrix is ​​determined according to formula (6);

[0087]

[0088] In the above formula (4), a is the semi-major axis; E is the anomaly angle at the future set time; e is the eccentricity; P is the first direction cosine matrix; Q is the second direction cosine matrix; r ECI is the position of the high-orbit satellite in the Earth-centered inertial coordinate system;

[0089]

[0090]

[0091] The above formulas (5)-(6) give the calculation method of the first direction cosine matrix P and the second direction cosine matrix Q, where ω is the argument of perigee; Ω is the right ascension of the ascending node; and i is the orbital inclination.

[0092] The method for completing the high-orbit satellite orbit prediction in the above steps S121-S123 is to calculate the eccentric anomaly angle at the future set time by solving the Kepler equation (3). Since the Kepler equation is a nonlinear equation, the solution process can be completed by analytical method, numerical method, or a combination of analytical method and numerical method. Numerical methods include but are not limited to Newton iteration method, Euler method, Runge-Kutta method, etc.; analytical methods include but are not limited to Taylor series expansion, Bessel function expansion, piecewise analytical approximation, etc. In practical applications, the appropriate solution method can also be selected considering the optimization of initial values, the selection of convergence basis, the actual consumption of computing resources and computing time. The following is a specific solution process for solving nonlinear equations (transcendental equations) using the Newton iteration method. The number of iterations is 3. First, the initial anomaly angle is constructed according to formula (7); secondly, function (8) is constructed according to the Kepler equation; the first-order derivative of function (8) is solved, as shown in formula (9); the first-order derivative is further differentiated to obtain the second-order derivative, as shown in formula (10); and the anomaly angle at a future set time is solved based on the first-order derivative and the second-order derivative, as shown in formula (11). It should be noted that formula (7) is only used to construct the initial anomaly angle. The anomaly angle obtained by solving formulas (8)-(11) will be used as the input of the next Kepler equation (8). The anomaly angle is continued to be solved until the number of iterations meets the preset requirement of 3 iterations, and the value of the anomaly angle after three iterations is output. It should be noted again that in the specific implementation, the number of iterations can be set by comprehensively considering the measurement accuracy and the equipment operation speed.

[0093] E = min([M / (1+e),M+e,M+e×(pi-M) / (1+e)]) (7)

[0094] fE = E – e×sin(E) – M (8)

[0095] fE_d1 = 1 – e×cos(E) (9)

[0096] fE_d2 = e×sin(E) (10)

[0097] E = E - fE / (fE_d1 – fE×fE_d2 / (2×fE_d1)) (11)

[0098] In the above formulas (7)-(11), E is the eccentric anomaly, e is the eccentricity, M is the mean anomaly, fE is a function, fE_d1 is the first-order derivative of the function fE, fE_d2 is the second-order derivative of the function fE, pi is pi, which is 3.14, and min is the function to be minimized.

[0099] The following explains the complete orbital elements: The orbital elements are the six orbital numbers that describe the position of a celestial body or spacecraft at a specific moment along its orbit. Commonly used orbital elements are the semimajor axis a, eccentricity e, inclination i, right ascension of the ascending node Ω, argument of perigee w, and true anomaly v. The semimajor axis and eccentricity describe the shape of the elliptical orbit, the inclination and right ascension of the ascending node describe the plane of the ellipse, the argument of perigee describes the orientation of the ellipse, and the true anomaly describes the position of the celestial body or spacecraft.

[0100] It is also necessary to further explain the high-orbit satellite orbit prediction method. On the TMS570LS3137 onboard simulation platform, by adopting the high-orbit satellite orbit prediction method of the present invention (i.e., the two-extrapolation model) to predict the high-orbit satellite orbit, a group of high-orbit satellite position and velocity in the geocentric inertial coordinate system (J2000 inertial coordinate system) can be output every 1 second, and the maximum position error predicted for 24 hours is 1.248 km.

[0101] The antenna pointing determination method provided by the present invention can be used to determine the pointing angle of the antenna after predicting the position of the high-orbit satellite at a set time in the future. As an implementation of step S140, Figure 3 As shown, the low-orbit satellite data information includes a coordinate conversion relationship between the geocentric inertial coordinate system and the low-orbit satellite antenna coordinate system; the position of the high-orbit satellite at a set time in the future and the position of the low-orbit satellite at a set time in the future are both located in the geocentric inertial coordinate system;

[0102] The determining the pointing angle of the antenna carried on the low-orbit satellite according to the low-orbit satellite data information, the position of the high-orbit satellite at a set time in the future, and the position of the low-orbit satellite at the set time in the future includes:

[0103] In step S141, a position vector in a geocentric inertial coordinate system is determined based on the position of the high-orbit satellite at a set time in the future and the position of the low-orbit satellite at a set time in the future;

[0104] In step S142, the pointing angle of the antenna carried on the low-orbit satellite is determined according to the position vector and the coordinate conversion relationship.

[0105] The coordinate transformation relationship includes a first coordinate transformation relationship between the geocentric inertial coordinate system and the low-orbit satellite orbit coordinate system, a second coordinate transformation relationship between the low-orbit satellite orbit coordinate system and the low-orbit satellite body coordinate system, and a third coordinate transformation relationship between the low-orbit satellite body coordinate system and the low-orbit satellite antenna coordinate system. Figure 4 As shown, determining the pointing angle of the antenna carried on the low-orbit satellite according to the position vector and the coordinate transformation relationship includes:

[0106] In step S142a, according to the first coordinate conversion relationship, the position vector in the geocentric inertial coordinate system is converted to the low-orbit satellite orbit coordinate system to obtain the position vector in the low-orbit satellite orbit coordinate system;

[0107] In step S142b, according to the second coordinate transformation relationship, the position vector in the low-orbit satellite orbit coordinate system is transformed into the low-orbit satellite body coordinate system to obtain the position vector in the low-orbit satellite body coordinate system;

[0108] In step S142c, according to the third coordinate transformation relationship, the position vector in the low-orbit satellite body coordinate system is transformed into the low-orbit satellite antenna coordinate system to obtain the position vector in the low-orbit satellite antenna coordinate system;

[0109] In step S142d, the antenna pointing angle is determined according to the position vector in the low-orbit satellite antenna coordinate system.

[0110] The following describes in detail several coordinate transformations in steps S142a-S142c and the transformation of the position vector in the geocentric inertial coordinate system to the position vector in the low-orbit satellite antenna coordinate system in combination with formulas (12)-(21).

[0111] The position vector in the geocentric inertial coordinate system can be converted to the position vector in the antenna coordinate system according to formula (12):

[0112] R ant =R3×R2×R1×dr ECI (12)

[0113] The operation in formula (12) is a cross multiplication operation, dr ECI is the position vector in the geocentric inertial coordinate system, R antis the position vector in the antenna coordinate system, R1, R2, and R3 correspond to the first coordinate transformation relationship, the second coordinate transformation relationship, and the third coordinate transformation relationship, respectively. The expressions of R1, R2, and R3 are as follows:

[0114]

[0115]

[0116]

[0117] Among them, the various parameters of the expression (13) of the first coordinate transformation relationship R1 are calculated according to formula (16), formula (17), and formula (18), respectively. In formula (16)-formula (18), r and v respectively represent the position vector and velocity vector of the low-orbit satellite in the geocentric inertial coordinate system.

[0118]

[0119]

[0120]

[0121] Among them, the various parameters of the expression (14) of the second coordinate transformation relationship R2 are calculated according to formula (19), formula (20) and formula (21), respectively. Formula (20)-Formula (21) represent the attitude of the low-orbit satellite, θ, ψ represents the pitch angle, roll angle, and yaw angle of the low-orbit satellite attitude, respectively.

[0122]

[0123]

[0124]

[0125] Among them, θ′, ψ′ is the installation Euler angle of the antenna relative to the low-orbit satellite body. Substituting the above Euler angle into formula (19), formula (20), and formula (21) respectively, we can get R y (θ′), R z (χ′).

[0126] After obtaining the position vector in the low-orbit satellite antenna coordinate system, the antenna pointing angle can be determined as an optional implementation of step S142d, such as Figure 5 As shown,

[0127] In step S142d1, the azimuth angle and the off-axis angle in the low-orbit satellite antenna coordinate system are obtained according to the position vector in the low-orbit satellite antenna coordinate system and trigonometric function transformation;

[0128] In step S142d2, antenna pointing is completed according to the azimuth angle and off-axis angle.

[0129] As an optional implementation of step S141, Figure 6 As shown, the method of determining the position vector in the geocentric inertial coordinate system according to the position of the high-orbit satellite at a future set time and the position of the low-orbit satellite at a future set time includes:

[0130] In step S141a, the position vector of the high-orbit satellite at the future set time is subtracted from the position vector of the low-orbit satellite at the future set time according to formula (22) to obtain the position vector in the geocentric inertial coordinate system;

[0131] dr ECI =r ECI -r0 ECI (twenty two)

[0132] Among them, r ECI Set the position of the high-orbit satellite at a certain time in the future; r0 ECI Set the position of a low-orbit satellite at a certain time in the future; dr ECI is the position vector in the geocentric inertial coordinate system.

[0133] The low-orbit satellite data information directly obtained by the onboard relay terminal from the low-orbit satellite platform also includes the position and velocity information of the low-orbit satellite in the geocentric inertial coordinate system; as an optional implementation of step S130, Figure 7 As shown, determining the position of the low-orbit satellite at the future set time based on the low-orbit satellite data information includes:

[0134] In step S131 , the position of the low-orbit satellite at a future time is obtained based on the position and speed of the low-orbit satellite in the orbital coordinate system at the current time and the relationship between the future time and the current time.

[0135] It is important to note that step S131 involves smoothing the position and velocity in the Earth-centered inertial coordinate system during the calculation of the future position of the low-orbit satellite. Smoothing methods include, but are not limited to, third-order spline interpolation, linear interpolation, and averaging to improve antenna pointing accuracy. Using the third-order spline interpolation algorithm for data smoothing reduces the antenna pointing calculation cycle to 100ms, improving pointing accuracy. Figure 8 A relay system diagram corresponding to the antenna pointing determination method provided by the present invention is given.

[0136] The antenna pointing determination method provided by the present invention can autonomously predict the antenna pointing angle in the future. By predicting the operating status of high-orbit satellites, the need for real-time data updates on ground stations can be significantly reduced, and the data transmission pressure on the communication link can be effectively alleviated. In addition, the antenna pointing accuracy is improved by obtaining the operating status of low-orbit satellites in real time and smoothing the data. This method of autonomously predicting the antenna pointing angle can also avoid pointing deviations caused by outdated orbital data. In the event of temporary interruptions or delays at the ground station, stable communication can still be maintained, thereby improving the communication success rate and reliability. It provides data relay, continuous tracking, and measurement and control services between low-orbit spacecraft and between spacecraft and ground stations, achieving 100% coverage of measurement, control, and communication for low-orbit spacecraft.

[0137] As a second aspect of the present invention, there is provided an electronic device, such as Figure 9 As shown, including:

[0138] One or more processors 101;

[0139] The memory 102 stores one or more computer programs. When the one or more computer programs are executed by the one or more processors 101, the one or more processors 101 implement the antenna pointing determination method provided according to the first aspect of the present invention.

[0140] The tool may further include one or more I / O interfaces 103 connected between the processor 101 and the memory 102 and configured to implement information exchange between the processor 101 and the memory 102 .

[0141] Among them, the processor 101 is a device with data processing capabilities, including but not limited to the central processing unit 101 (CPU); the first memory 102 is a device with data storage capabilities, including but not limited to random access memory 102 (RAM, more specifically such as SDRAM, DDR, etc.), read-only memory 102 (ROM), electrically erasable programmable read-only memory 102 (EEPROM), flash memory (FLASH); the I / O interface 103 (read-write interface) is connected between the processor 101 and the memory 102, and can realize information interaction between the processor 101 and the memory 102, including but not limited to the data bus 104 (Bus), etc.

[0142] In some embodiments, the processor 101 , the memory 102 , and the I / O interface 103 are connected to each other via a bus 104 , and further connected to other components of the computing device.

[0143] In addition, as a third aspect of the present invention, a computer readable medium is provided, on which a computer program is stored. Figure 10As shown, when the computer program is executed by a processor, the antenna pointing determination method provided by the first aspect of the present invention is implemented.

[0144] Those skilled in the art will appreciate that all or part of the processes in the above-mentioned embodiment methods can be implemented by instructing the relevant hardware through a computer program. Accordingly, the computer program can be stored in a non-volatile computer-readable storage medium, and when the computer program is executed, it can implement the method of any of the above-mentioned embodiments. Among them, any reference to memory, storage, database or other media used in the embodiments provided in this application may include non-volatile and / or volatile memory. Non-volatile memory may include read-only memory (ROM), programmable ROM (PROM), electrically programmable ROM (EPROM), electrically erasable programmable ROM (EEPROM) or flash memory. Volatile memory may include random access memory (RAM) or external cache memory. By way of illustration and not limitation, RAM is available in many forms, such as static RAM (SRAM), dynamic RAM (DRAM), synchronous DRAM (SDRAM), double data rate SDRAM (DDRSDRAM), enhanced SDRAM (ESDRAM), Synchronous Link DRAM (SLDRAM), Rambus direct RAM (RDRAM), direct memory bus dynamic RAM (DRDRAM), and memory bus dynamic RAM (RDRAM), etc.

[0145] The above are only specific embodiments of the present invention, but the scope of protection of the present invention is not limited thereto. Those skilled in the art should understand that the present invention includes but is not limited to the contents described in the drawings and the above specific embodiments. Any modifications that do not deviate from the functional and structural principles of the present invention are intended to be included within the scope of the claims.

Claims

1. A method for determining antenna pointing, for use in a satellite-borne relay terminal, characterized in that: The method comprises: Acquiring high-orbit satellite data information from a low-orbit satellite platform carrying the onboard relay terminal, and low-orbit satellite data information of a low-orbit satellite carrying the low-orbit satellite platform; Predicting the position of the high-orbit satellite at a set time in the future based on the high-orbit satellite data information; Determining the position of the low-orbit satellite at the set future time according to the low-orbit satellite data information; The pointing angle of the antenna carried on the low-orbit satellite is determined based on the low-orbit satellite data information, the position of the high-orbit satellite at a set time in the future, and the position of the low-orbit satellite at a set time in the future.

2. The antenna direction determination method according to claim 1, characterized in that: The high-orbit satellite data includes the orbital elements of the high-orbit satellite in the geocentric inertial coordinate system and the epoch time corresponding to the orbital elements; the orbital elements include the semi-major axis, eccentricity, orbital inclination, right ascension of the ascending node, argument of perigee, and mean anomaly; The predicting the position of the high-orbit satellite at a set future time based on the high-orbit satellite data information includes: Predicting the mean anomaly at the future set time based on the mean anomaly corresponding to the epoch, the future set time, formula (1) and formula (2); M=M0+n×(t-t0) (1) Where M0 is the mean anomaly at the epoch time; n is the mean angular velocity; t0 is the epoch time; t is the future set time; M is the mean anomaly at the future set time; μ is the gravitational constant of the central celestial body; a is the semi-major axis; Predict the eccentric anomaly at the future setting time based on the mean anomaly, eccentricity, and formula (3); M=Ee×sinE (3) Where M is the mean anomaly at a future set time, e is the eccentricity, and E is the eccentric anomaly at a future set time; The position of the high-orbit satellite in the geocentric inertial coordinates is determined according to the eccentric anomaly angle, semi-major axis, eccentricity, direction cosine matrix, and formula (4) at a future set time; wherein the direction cosine matrix includes a first direction cosine matrix and a second direction cosine matrix, the first direction cosine matrix is ​​determined according to formula (5), and the second direction cosine matrix is ​​determined according to formula (6); Where a is the semi-major axis; E is the anomaly angle at the future set time; e is the eccentricity; P is the first direction cosine matrix; Q is the second direction cosine matrix; r ECI is the position of the high-orbit satellite in the Earth-centered inertial coordinate system; Where ω is the argument of perigee; Ω is the right ascension of the ascending node; i is the orbital inclination; P is the first direction cosine matrix; and Q is the second direction cosine matrix.

3. The method for determining antenna direction according to claim 1, wherein: The low-orbit satellite data information includes a coordinate conversion relationship between a geocentric inertial coordinate system and a low-orbit satellite antenna coordinate system; the position of the high-orbit satellite at a set time in the future and the position of the low-orbit satellite at a set time in the future are both located in the geocentric inertial coordinate system; The determining the pointing angle of the antenna carried on the low-orbit satellite according to the low-orbit satellite data information, the position of the high-orbit satellite at a set time in the future, and the position of the low-orbit satellite at the set time in the future includes: Determine a position vector in a geocentric inertial coordinate system based on a position of a high-orbit satellite at a set time in the future and a position of a low-orbit satellite at a set time in the future; The pointing angle of the antenna carried on the low-orbit satellite is determined according to the position vector and the coordinate conversion relationship.

4. The method for determining antenna direction according to claim 3, wherein: The coordinate transformation relationship includes a first coordinate transformation relationship between the geocentric inertial coordinate system and the low-orbit satellite orbit coordinate system, a second coordinate transformation relationship between the low-orbit satellite orbit coordinate system and the low-orbit satellite body coordinate system, and a third coordinate transformation relationship between the low-orbit satellite body coordinate system and the low-orbit satellite antenna coordinate system; Determining the pointing angle of the antenna carried on the low-orbit satellite according to the position vector and the coordinate transformation relationship includes: Converting the position vector in the geocentric inertial coordinate system to the low-orbit satellite orbit coordinate system according to the first coordinate conversion relationship to obtain the position vector in the low-orbit satellite orbit coordinate system; According to the second coordinate transformation relationship, the position vector in the low-orbit satellite orbital coordinate system is transformed into the low-orbit satellite body coordinate system to obtain the position vector in the low-orbit satellite body coordinate system; According to the third coordinate transformation relationship, the position vector in the low-orbit satellite body coordinate system is transformed into the low-orbit satellite antenna coordinate system to obtain the position vector in the low-orbit satellite antenna coordinate system; The antenna pointing angle is determined according to the position vector in the low-orbit satellite antenna coordinate system.

5. The method for determining antenna direction according to claim 4, wherein: The determining the antenna pointing angle according to the position vector in the low-orbit satellite antenna coordinate system includes: Obtaining an azimuth angle and an off-axis angle in the low-orbit satellite antenna coordinate system according to the position vector in the low-orbit satellite antenna coordinate system and trigonometric function transformation; Antenna pointing is performed according to the azimuth angle and the off-axis angle.

6. The method for determining antenna direction according to claim 3, wherein: Determining the position vector in the geocentric inertial coordinate system according to the position of the high-orbit satellite at a future set time and the position of the low-orbit satellite at a future set time includes: According to formula (7), the position vector of the high-orbit satellite at the future set time is subtracted from the position vector of the low-orbit satellite at the future set time to obtain the position vector in the geocentric inertial coordinate system; dr ECI ＝r ECI -r0 ECI (7) Among them, r ECI Set the position of the high-orbit satellite at a certain time in the future; r0 ECI Set the position of a low-orbit satellite at a certain time in the future; dr ECI is the position vector in the geocentric inertial coordinate system.

7. The method for determining antenna direction according to claim 1, wherein: The low-orbit satellite data further includes position and velocity information of the low-orbit satellite in a geocentric inertial coordinate system, and determining the position of the low-orbit satellite at the future set time based on the low-orbit satellite data information includes: According to the position and speed of the low-orbit satellite in the geocentric inertial coordinate system at the current time and the relationship between the future time and the current time, the position of the low-orbit satellite at the future time is obtained.

8. The antenna pointing calculation method according to claim 7, characterized in that: The method of obtaining the position of the low-orbit satellite at a future time based on the position and velocity of the low-orbit satellite in the geocentric inertial coordinate system at the current time and the relationship between the future time and the current time further includes: Smoothing of the position and velocity in the geocentric inertial coordinate system.

9. An electronic device, characterized in that: include: one or more processors; A memory having one or more computer programs stored thereon, wherein when the one or more computer programs are executed by the one or more processors, the one or more processors implement the antenna pointing determination method according to any one of claims 1 to 8.

10. A computer-readable medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, the antenna direction determination method according to any one of claims 1 to 8 is implemented.

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