Inter-satellite pointing on-orbit calculation method and system for ephemeris data time delay

Abstract

The invention provides an inter-satellite pointing on-orbit calculation method and system for ephemeris data delay, and the method comprises the steps: a data solving step: solving ephemeris data and inter-satellite relative positioning data through employing a numerical integration algorithm, and obtaining the position and speed of a satellite at a current on-satellite moment; and a coordinate system conversion step: performing coordinate system conversion on the obtained position and speed at the on-satellite moment to obtain an inter-satellite pointing vector. The ephemeris data and the inter-satellite relative positioning data have high positioning precision, even if the orbit attitude data has time delay, the inter-satellite pointing vector with high precision can be obtained through numerical integration recursion and coordinate system conversion, the in-orbit calculation steps are simple and convenient, and the method is suitable for large-scale popularization and application. The risk of inter-satellite link communication interruption is effectively reduced; the method can be used for inter-satellite pointing calculation of formation satellites, the time delay problem of ephemeris data is solved through numerical integration recursion, and the method is of great significance to improvement of continuous stability of work of inter-satellite communication links.

Description

Technical Field

[0001] The present invention relates to the field of aerospace technology, and in particular to an on-orbit calculation method and system for inter-satellite pointing of ephemeris data delay. Background Art

[0002] As the complexity of satellite missions increases, more and more formation flying satellites are put into use. In order to achieve inter-satellite collaboration, it is necessary to establish inter-satellite communication links between satellites. Accurately grasping the azimuth relationship of other satellites relative to the inter-satellite communication antenna of this satellite has become the key to whether the inter-satellite communication links can operate continuously and stably.

[0003] The circling formation satellites have put forward higher requirements on the full-time and omnidirectionality of the intersatellite communication links. Due to the limitations of satellite layout, a single antenna is bound to have an interference area, making it difficult to form an omnidirectional pattern. Multiple antennas need to be arranged so that their fields of view can be spliced ​​to form an omnidirectional intersatellite communication antenna array group, leaving a certain overlapping range between the fields of view. By switching and selecting different intersatellite communication antennas, the circling formation satellites can establish a continuous and stable intersatellite communication link at any time.

[0004] In order to accurately select the corresponding intersatellite communication antenna, it is necessary to obtain the relative position information of other satellites relative to the local satellite. The traditional method of calculating intersatellite pointing by using the satellite orbit parameters recorded on the ground has poor satellite orbit position accuracy after a long period of relative parameter recording. If the orbit recursion accuracy is to be improved, the complexity of the on-board calculation software must be increased, which will affect the reliability of the satellite's stable operation:

[0005] Patent document CN114741647A discloses a vector calculation method for establishing a link between satellites in different orbits, but it requires the position vector of the other satellite in the inertial system to be known, which results in a large calculation error in terms of navigation accuracy and inter-satellite formation distance.

[0006] Patent document CN113472421B discloses a low-orbit networked satellite inter-satellite beam pointing method based on Lagrangian interpolation, and introduces an inter-satellite beam pointing method that requires reading the orbital parameters of two satellites and interpolating the calculations. When the time is long relative to the orbital injection time, the accuracy is poor and the dependence on ground injection is too strong.

[0007] Patent document CN116252968A discloses a relative pointing tracking control method without tracking and aiming conditions, and introduces a relative pointing calculation and fitting extrapolation method. However, the orbit recursion accuracy will directly affect the accuracy of the pointing calculation, which is not suitable for the high-precision scenario involved in the present invention.

[0008] Patent document CN117254847A discloses a method and device for rapid satellite search of a low-orbit satellite terminal, and introduces a method for calculating antenna pointing information between a satellite terminal and a satellite, but it also requires the use of satellite ephemeris information for recursion, which is not suitable for the high-precision scenario involved in the present invention.

[0009] Patent document CN114002710A discloses an autonomous prediction method for the on-board orbital position of a low-orbit satellite with small eccentricity, and introduces an on-board orbit recursion method. However, the long-term recursion accuracy is still lower than the satellite navigation positioning accuracy, and the high-precision data on the satellite is not effectively utilized.

[0010] With the development of satellite-borne navigation and positioning systems, the accuracy of satellite positioning obtained through navigation has improved. However, for formation satellite missions, the use of differential technology to position formation satellites between satellites has higher accuracy. It is necessary to make good use of this information to bring new breakthroughs in solving the on-orbit calculation problem of inter-satellite pointing. Summary of the invention

[0011] In view of the defects in the prior art, the object of the present invention is to provide an on-orbit calculation method and system for inter-satellite pointing of ephemeris data delay.

[0012] According to the present invention, a method for calculating inter-satellite pointing on-orbit of ephemeris data delay is provided, comprising:

[0013] Data solving steps: Use numerical integration algorithm to solve the ephemeris data and inter-satellite relative positioning data to obtain the position and velocity of the satellite at the current time on the satellite;

[0014] Coordinate system conversion step: convert the position and velocity obtained on the satellite to obtain the inter-satellite pointing vector.

[0015] Preferably, the data solving step includes:

[0016] Step S1.0: Based on the latest ephemeris data on the satellite, the position and velocity vector of the satellite in the inertial system at the time of the ephemeris data are calculated;

[0017] Step S1.1: Recursively extrapolate the position and velocity obtained from the ephemeris data through a numerical integration algorithm to solve the position and velocity of the satellite in the inertial system at the current time on the satellite.

[0018] Preferably, the coordinate system conversion step includes:

[0019] Step S2.1: According to the current on-board time, the approximate conversion matrix between the inertial system and the earth-fixed system is calculated, and the inter-satellite relative position vector in the earth-fixed system output by the inter-satellite navigation system is converted to the inertial system;

[0020] Step S2.2: According to the position and velocity vector of the local satellite in the inertial system, the transformation matrix between the orbital system and the inertial system is calculated, and the inter-satellite relative position vector is transformed from the inertial system to the orbital system;

[0021] Step S2.3: Calculate the intersatellite pointing vector in the local system of the local star.

[0022] Preferably, step S2.3 includes linearly estimating the attitude angle and attitude angular velocity at the time of the ephemeris data to obtain the attitude angle at the current time on the satellite, calculating the transformation matrix between the orbital system and the present system according to the corresponding Euler angle conversion sequence, compensating for the deviation of the installation position of the intersatellite communication antenna relative to the satellite center of mass, and obtaining the intersatellite pointing vector under the present system.

[0023] Preferably, the solved data is updated on the data bus in the form of an ephemeris table; in the data solving step, the orbital attitude data of the ephemeris has no time delay relative to the current time of the satellite.

[0024] According to the present invention, an inter-satellite pointing on-orbit calculation system for ephemeris data delay is provided, comprising:

[0025] Data solution module: Use numerical integration algorithm to solve the ephemeris data and inter-satellite relative positioning data to obtain the position and velocity of the satellite at the current time on the satellite;

[0026] Coordinate system conversion module: convert the position and velocity obtained on the satellite to obtain the inter-satellite pointing vector.

[0027] Preferably, the data solving module includes:

[0028] Module M1.0: Based on the latest ephemeris data on the satellite, the position and velocity vector of the satellite in the inertial system at the time of the ephemeris data are calculated;

[0029] Module M1.1: The position and velocity obtained from the ephemeris data are recursively deduced through a numerical integration algorithm to solve the position and velocity of the satellite in the inertial system at the current time on the satellite.

[0030] Preferably, the coordinate system conversion module includes:

[0031] Module M2.1: According to the current on-board time, the approximate conversion matrix between the inertial system and the earth-fixed system is calculated, and the inter-satellite relative position vector in the earth-fixed system output by the inter-satellite navigation system is converted to the inertial system;

[0032] Module M2.2: Based on the position and velocity vector of the satellite in the inertial system, calculate the transformation matrix between the orbital system and the inertial system, and transform the inter-satellite relative position vector from the inertial system to the orbital system;

[0033] Module M2.3: Calculate the intersatellite pointing vector in the local system of the local star.

[0034] Preferably, the module M2.3 includes linearly estimating the attitude angle and attitude angular velocity at the time of the ephemeris data to obtain the attitude angle at the current time on the satellite, calculating the transformation matrix between the orbital system and the present system according to the corresponding Euler angle conversion sequence, compensating for the deviation of the installation position of the intersatellite communication antenna relative to the satellite center of mass, and obtaining the intersatellite pointing vector under the present system.

[0035] Preferably, the solved data is updated on the data bus in the form of an ephemeris table; in the data solving module, the orbital attitude data of the ephemeris has no time delay relative to the current time of the satellite.

[0036] Compared with the prior art, the present invention has the following beneficial effects:

[0037] 1. The method proposed in the present invention makes full use of the processed navigation and positioning data, which is generally updated on the data bus in the form of ephemeris. The overlapping field of view of the inter-satellite communication antenna ensures that high-precision inter-satellite relative position information can be used as input, and the inter-satellite pointing vector under the local satellite system is obtained through coordinate transformation.

[0038] 2. The present invention takes into account the delay in updating ephemeris data and obtains the position and attitude information at the current satellite time through numerical integration, thereby ensuring the accuracy of the inter-satellite pointing algorithm. The on-orbit calculation steps are simple, effectively reducing the risk of inter-satellite link communication interruption.

[0039] 3. The present invention can be used for inter-satellite pointing calculation of formation satellites, and solves the delay problem of ephemeris data through numerical integration recursion, which is of great significance to improving the continuous stability of inter-satellite communication links.

[0040] Other beneficial effects of the present invention will be explained in the specific implementation manner through the introduction of specific technical features and technical solutions. Through the introduction of these technical features and technical solutions, those skilled in the art should be able to understand the beneficial technical effects brought about by the technical features and technical solutions. BRIEF DESCRIPTION OF THE DRAWINGS

[0041] Other features, objects and advantages of the present invention will become more apparent from the detailed description of non-limiting embodiments made with reference to the following drawings:

[0042] Figure 1 The figure is a flow chart of the method of the present invention. DETAILED DESCRIPTION

[0043] The present invention is described in detail below in conjunction with specific embodiments. The following embodiments will help those skilled in the art to further understand the present invention, but are not intended to limit the present invention in any form. It should be noted that, for those of ordinary skill in the art, several changes and improvements can also be made without departing from the concept of the present invention. These all belong to the protection scope of the present invention.

[0044] Reference Figure 1 As shown, a method for calculating inter-satellite pointing on-orbit of ephemeris data delay includes:

[0045] Step 1: Calculate the position and velocity of the satellite in the inertial system at the time of the ephemeris data;

[0046] The position and velocity vector of the satellite in the inertial system are calculated by calculating the orbital elements in the ephemeris.

[0047] Step 2: Calculate the position and velocity of the star in the inertial system at the current moment;

[0048] Since there is a time delay in updating the ephemeris data relative to the current on-board time, it is necessary to recursively extrapolate the position and velocity obtained from the ephemeris data using a numerical integration algorithm to solve the position and velocity of the satellite at the current on-board time.

[0049] Step 3: Calculate the relative position between satellites in the inertial system;

[0050] The relative positions of other satellites obtained by inter-satellite measurements are generally expressed in the Earth-fixed system. According to the current on-board time, the approximate conversion matrix between the inertial system and the Earth-fixed system can be calculated using a simplified formula to convert the inter-satellite relative positions in the Earth-fixed system to the inertial system.

[0051] Step 4: Calculate the relative position between stars in the local star orbit system;

[0052] According to the current position and velocity vector of the local star, the conversion matrix between the inertial system and the orbital system is calculated, and the relative position between the stars in the inertial system is converted to the local star's orbital system.

[0053] Step 5: Calculate the intersatellite pointing vector of the local star and local system;

[0054] The attitude angle and attitude angular velocity at the time of ephemeris data are linearly estimated to obtain the attitude angle at the current time on the satellite. The transformation matrix between the orbital system and this system is calculated according to the corresponding Euler angle conversion sequence. After compensating for the deviation of the installation position of the intersatellite communication antenna relative to the satellite center of mass, the intersatellite pointing vector under this system is obtained.

[0055] The method proposed in the present invention makes full use of the processed navigation and positioning data, which is generally updated on the data bus in the form of an ephemeris table, and the overlapping fields of view of inter-satellite communication antennas ensure that high-precision inter-satellite relative position information can be used as input, and the inter-satellite pointing vector under the local satellite system is obtained through coordinate conversion; the present invention also takes into account the update delay of the ephemeris data, and obtains the position and attitude information at the current on-satellite time through numerical integration, thereby ensuring the accuracy of the inter-satellite pointing algorithm, and the on-orbit calculation steps are simple, which effectively reduces the risk of interruption of inter-satellite link communication.

[0056] The above is a basic embodiment of the present invention. The technical solution of the present invention is further described below through a preferred embodiment.

[0057] Example 1

[0058] Reference Figure 1 As shown, a method for calculating inter-satellite pointing on-orbit of ephemeris data delay includes:

[0059] 1. Calculate the position and velocity of the satellite in the inertial system at the time of the ephemeris data;

[0060] The satellite's orbital elements are generally continuously updated in the satellite's ephemeris, but relative to the current on-board time, the data in the ephemeris is the result of a certain time in the past, and there is a certain time delay. These orbital elements can be used to calculate the satellite's position and speed at the time corresponding to the data.

[0061] The orbital elements in the ephemeris table can be in various forms. The algorithm for converting various forms of orbital elements into position and velocity in the inertial system is relatively mature. Here, the ephemeris data time t z The orbital instantaneous root (orbital semi-major axis a) of the J2000.0 inertial system s , eccentricity e s , orbital inclination i s , right ascension of ascending node Ω s 、Arg of perigee ω s 、True anomaly f s ) as an example to calculate the position in the inertial system and speed The formula is as follows:

[0062]

[0063]

[0064]

[0065]

[0066]

[0067]

[0068] Where μ is the Earth's gravitational constant, h s is the angular momentum, r p 、v p are the position and velocity vectors in the near-focus coordinate system, Q is the transformation matrix from the near-focus coordinate system to the inertial system, and the rotation matrix function R with α as the independent variable X (α), R Z (α) The expressions are:

[0069]

[0070]

[0071] 2. Calculate the position and velocity of the star in the inertial system at the current moment

[0072] Since there is a time delay between the update of ephemeris data and the current on-board time (i.e., the ephemeris data time t z Less than the current on-board time t x ), the position obtained from the ephemeris data needs to be integrated by numerical integration algorithm and speed Perform recursion to solve the position of the star at the current time on the star and speed

[0073] The algorithm is as follows:

[0074] Integration step h = t x -t z The initial value of the integral is

[0075]

[0076] The following numerical calculation formula is used

[0077]

[0078] Among them, the differential equation system f(p) is a 6×1 column vector, p is the term in brackets from k1 to k4, let The simplified approximate expression of f(p) is:

[0079]

[0080] Where R e is the equatorial radius of the Earth, and J2 is the second-order harmonic coefficient of the Earth's non-spherical gravitational perturbation term. The output result is:

[0081]

[0082]

[0083] 3. Calculate the relative position between satellites in the inertial system:

[0084] The relative positions of other stars obtained by intersatellite measurements Generally speaking, in the ground fixed system, according to the current satellite time t x , the simplified formula is used to calculate the approximate transformation matrix between the inertial system and the earth-fixed system, and the relative position between satellites in the earth-fixed system is Convert to relative position between stars in inertial system

[0085] Let M ECF2ECI is the transformation matrix from the Earth-fixed system (WGS84 coordinate system) to the J2000.0 inertial system, then:

[0086]

[0087] Among them, M ECI2ECF is the transformation matrix from the J2000.0 inertial system to the WGS84 coordinate system, and its expression is:

[0088] M ECI2ECF =EP·ER·NR·PR;

[0089] Among them, EP is the polar motion matrix, which is simplified to a 3×3 unit matrix; ER is the Earth rotation matrix, NR is the nutation matrix, and PR is the precession matrix. They are calculated according to the following formulas:

[0090]

[0091] ER=R Z (S G );

[0092] NR=R X (-(ε+Δε))R Z (-Δψ)R X (ε);

[0093] PR=R Z (-Z A )R Y (θ A )R Z (-ζ A );

[0094] Rotation matrix function R Y for

[0095]

[0096] The calculations are as follows:

[0097] 1.ζA ,θ A , Z A

[0098] The current time t on the satellite in the form of integrated seconds relative to the epoch J2000.0 (12:00 on January 1, 2000) x Convert to the Julian century form of the corresponding dynamic time TDT:

[0099]

[0100] Among them, t m is the accumulated number of leap seconds, the precession constant ζ A ,θ A , Z A The calculation of is as follows:

[0101]

[0102] (2) Δε, ε, Δψ

[0103] The nutation of the ecliptic longitude Δψ, the obliquity ε and the nutation Δε are:

[0104]

[0105] Where Ω m =(125.044555555556-1934.136185×T TDT )×d2r.

[0106] (3)S G

[0107] Greenwich true sidereal time S G for:

[0108]

[0109] in For Greenwich mean sidereal time, the right ascension nutation Δμ = Δψcosε.

[0110] calculate The current time t is required x The corresponding Julian century number for UT1

[0111]

[0112]

[0113] 4. Calculate the relative position between stars in the orbital system of the local star

[0114] According to the current position velocity vector of the star Calculate the transformation matrix M between the inertial system and the orbital system orb2ECI , the relative position between the stars in the inertial system Convert to relative inter-satellite position in the local orbital system

[0115] Coordinate transformation matrix M from satellite orbit system to J2000.0 inertial system orb2ECI for:

[0116]

[0117] in, Both are 3×1 column vectors. ||X a || represents X a Model.

[0118] Relative position in satellite orbit coordinate system for:

[0119]

[0120] 5. Calculate the intersatellite pointing vector of the local system

[0121] Since the satellite's attitude angular velocity is low in the wheel control state, the recursive time is short and can be linearly estimated. The attitude angle (roll angle) at the time of the ephemeris data Pitch angle θ0, yaw angle ψ0) and attitude angular velocity (roll angular velocity Pitch angular velocity Yaw angular velocity ), and the attitude angle (roll angle) at the current time on the satellite is obtained by linear estimation Pitch angle θ, yaw angle ψ):

[0122]

[0123] Calculate the transformation matrix M from the satellite orbit coordinate system to the satellite body coordinate system according to the corresponding Euler angle conversion sequence orb2body Here we take 3-1-2 as an example

[0124]

[0125] The installation position of the intersatellite communication antenna in the satellite body coordinate system is R az , then the intersatellite pointing vector is

[0126]

[0127] The present invention also provides an on-orbit calculation system for inter-satellite pointing of ephemeris data delay. The on-orbit calculation system for inter-satellite pointing of ephemeris data delay can be implemented by executing the process steps of the on-orbit calculation method for inter-satellite pointing of ephemeris data delay, that is, those skilled in the art can understand the on-orbit calculation method for inter-satellite pointing of ephemeris data delay as a preferred implementation of the on-orbit calculation system for inter-satellite pointing of ephemeris data delay.

[0128] Specifically, an on-orbit calculation system for inter-satellite pointing of ephemeris data delay includes:

[0129] Data solution module: Use numerical integration algorithm to solve the ephemeris data and inter-satellite relative positioning data to obtain the position and velocity of the satellite at the current time on the satellite;

[0130] Coordinate system conversion module: convert the position and velocity obtained on the satellite to obtain the inter-satellite pointing vector.

[0131] The data solving module comprises:

[0132] Module M1.0: Based on the latest ephemeris data on the satellite, the position and velocity vector of the satellite in the inertial system at the time of the ephemeris data are calculated;

[0133] Module M1.1: The position and velocity obtained from the ephemeris data are recursively deduced through a numerical integration algorithm to solve the position and velocity of the satellite in the inertial system at the current time on the satellite.

[0134] The coordinate system conversion module comprises:

[0135] Module M2.1: According to the current on-board time, the approximate conversion matrix between the inertial system and the earth-fixed system is calculated, and the inter-satellite relative position vector in the earth-fixed system output by the inter-satellite navigation system is converted to the inertial system;

[0136] Module M2.2: Based on the position and velocity vector of the satellite in the inertial system, calculate the transformation matrix between the orbital system and the inertial system, and transform the inter-satellite relative position vector from the inertial system to the orbital system;

[0137] Module M2.3: Calculate the intersatellite pointing vector in the local system of the local star.

[0138] The module M2.3 includes linearizing the attitude angle and attitude angular velocity at the time of the ephemeris data to obtain the attitude angle at the current time on the satellite, calculating the transformation matrix between the orbital system and this system according to the corresponding Euler angle conversion sequence, compensating for the deviation of the installation position of the intersatellite communication antenna relative to the satellite center of mass, and obtaining the intersatellite pointing vector under this system.

[0139] The solved data is updated on the data bus in the form of an ephemeris table; in the data solving module, the orbit attitude data of the ephemeris has no time delay relative to the current time of the satellite.

[0140] Those skilled in the art know that, in addition to realizing the system and its various devices, modules, and units provided by the present invention in a purely computer-readable program code, it is entirely possible to realize the same functions in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers by logically programming the method steps. Therefore, the system and its various devices, modules, and units provided by the present invention can be considered as a hardware component, and the devices, modules, and units included therein for realizing various functions can also be regarded as structures within the hardware component; the devices, modules, and units for realizing various functions can also be regarded as both software modules for realizing the method and structures within the hardware component.

[0141] In the description of the present application, it should be understood that the terms "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside", etc., indicating orientations or positional relationships, are based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present application and simplifying the description, and do not indicate or imply that the referred device or element must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as a limitation on the present application.

[0142] The above describes the specific embodiments of the present invention. It should be understood that the present invention is not limited to the above specific embodiments, and those skilled in the art can make various changes or modifications within the scope of the claims, which does not affect the essence of the present invention. In the absence of conflict, the embodiments of the present application and the features in the embodiments can be combined with each other arbitrarily.

Claims

1. A method for calculating inter-satellite pointing on-orbit based on ephemeris data delay, characterized in that: include: Data solving steps: Use numerical integration algorithm to solve the ephemeris data and inter-satellite relative positioning data to obtain the position and velocity of the satellite at the current time on the satellite; Coordinate system conversion step: convert the position and velocity obtained on the satellite to obtain the inter-satellite pointing vector.

2. The on-orbit calculation method for inter-satellite pointing of ephemeris data delay according to claim 1, characterized in that: The data solving step comprises: Step S1.0: Based on the latest ephemeris data on the satellite, the position and velocity vector of the satellite in the inertial system at the time of the ephemeris data are calculated; Step S1.1: Recursively extrapolate the position and velocity obtained from the ephemeris data through a numerical integration algorithm to solve the position and velocity of the satellite in the inertial system at the current time on the satellite.

3. The on-orbit calculation method for inter-satellite pointing of ephemeris data delay according to claim 2, characterized in that: The coordinate system conversion step comprises: Step S2.1: According to the current on-board time, the approximate conversion matrix between the inertial system and the earth-fixed system is calculated, and the inter-satellite relative position vector in the earth-fixed system output by the inter-satellite navigation system is converted to the inertial system; Step S2.2: According to the position and velocity vector of the local satellite in the inertial system, the transformation matrix between the orbital system and the inertial system is calculated, and the inter-satellite relative position vector is transformed from the inertial system to the orbital system; Step S2.3: Calculate the intersatellite pointing vector in the local system of the local star.

4. The on-orbit calculation method for inter-satellite pointing of ephemeris data delay according to claim 3, characterized in that: The step S2.3 includes linearly estimating the attitude angle and attitude angular velocity at the time of the ephemeris data to obtain the attitude angle at the current time on the satellite, calculating the transformation matrix between the orbital system and the current system according to the corresponding Euler angle conversion sequence, compensating for the deviation of the installation position of the intersatellite communication antenna relative to the satellite center of mass, and obtaining the intersatellite pointing vector under the current system.

5. The on-orbit calculation method for inter-satellite pointing of ephemeris data delay according to claim 1, characterized in that: The solved data is updated on the data bus in the form of an ephemeris table; in the data solving step, the orbit attitude data of the ephemeris has no time delay relative to the current time of the satellite.

6. An on-orbit calculation system for inter-satellite pointing based on ephemeris data delay, characterized in that: include: Data solution module: Use numerical integration algorithm to solve the ephemeris data and inter-satellite relative positioning data to obtain the position and velocity of the satellite at the current time on the satellite; Coordinate system conversion module: convert the position and velocity obtained on the satellite to obtain the inter-satellite pointing vector.

7. The on-orbit calculation system for inter-satellite pointing of ephemeris data delay according to claim 6, characterized in that: The data solving module comprises: Module M1.0: Based on the latest ephemeris data on the satellite, the position and velocity vector of the satellite in the inertial system at the time of the ephemeris data are calculated; Module M1.1: The position and velocity obtained from the ephemeris data are recursively deduced through a numerical integration algorithm to solve the position and velocity of the satellite in the inertial system at the current time on the satellite.

8. The on-orbit calculation system for inter-satellite pointing of ephemeris data delay according to claim 7, characterized in that: The coordinate system conversion module comprises: Module M2.1: According to the current on-board time, the approximate conversion matrix between the inertial system and the earth-fixed system is calculated, and the inter-satellite relative position vector in the earth-fixed system output by the inter-satellite navigation system is converted to the inertial system; Module M2.2: Based on the position and velocity vector of the satellite in the inertial system, calculate the transformation matrix between the orbital system and the inertial system, and transform the inter-satellite relative position vector from the inertial system to the orbital system; Module M2.3: Calculate the intersatellite pointing vector in the local system of the local star.

9. The on-orbit calculation system for inter-satellite pointing of ephemeris data delay according to claim 8, characterized in that: The module M2.3 includes linearizing the attitude angle and attitude angular velocity at the time of the ephemeris data to obtain the attitude angle at the current time on the satellite, calculating the transformation matrix between the orbital system and this system according to the corresponding Euler angle conversion sequence, compensating for the deviation of the installation position of the intersatellite communication antenna relative to the satellite center of mass, and obtaining the intersatellite pointing vector under this system.

10. The on-orbit calculation system for inter-satellite pointing of ephemeris data delay according to claim 6, characterized in that: The solved data is updated on the data bus in the form of an ephemeris table; in the data solving module, the orbit attitude data of the ephemeris has no time delay relative to the current time of the satellite.

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