Satellite on-board orbit transfer method and system

Abstract

This invention provides a satellite onboard orbit conversion method and system. The method includes: calculating the corrected Julian day at the current time and calculating the Greenwich Mean Sidereal Time Angle based on the corrected Julian day; converting the Greenwich Mean Sidereal Time Angle into radians; calculating the Earth rotation matrix from the J2000 system to the WGS84 system based on the converted radians; pre-calculating the Earth nutation and precession correction matrices as modifiable parameters for satellite commands and updating them periodically; calculating the conversion matrix from the J2000 system to the WGS84 system based on the Earth nutation and precession correction matrices and the Earth rotation matrix; and calculating the satellite position and velocity in the WGS84 system based on the conversion matrix, thereby realizing the satellite onboard orbit conversion. This invention addresses the need for converting EP satellite onboard orbit information from an inertial frame to the WGS84 coordinate system, meets the requirements for short message stand-alone operation, and can provide a reference for on-orbit conversion of other satellites.

Description

Technical Field

[0001] This invention relates to the field of satellite communication technology, specifically to a satellite onboard orbit conversion method and system, and also to a corresponding satellite-based computer and computer-readable storage medium. Background Technology

[0002] The Einstein Probe (EP) satellite is an astronomical science exploration satellite for time-domain astronomy and high-energy astrophysics, used for all-sky soft X-ray transient source detection and variable source monitoring. It carries a large field-of-view (3600 square degrees) soft X-ray (0.5~4keV) telescope (WXT) and two depth follow-up X-ray (0.3-10keV) telescopes with a field of view of approximately 38 arcminutes. Its purpose is to conduct deep, large field-of-view time-domain surveys to discover and detect various known and unknown sudden transient / burst objects and events in the universe, and to issue alerts to guide onboard equipment and other space and ground-based telescopes at home and abroad to conduct follow-up observations.

[0003] To address the rapid downlink of alert information and rapid uplink of target of opportunity (ToO) information required by the EP satellite, the satellite is equipped with a short message communication unit (SRM) to rapidly transmit and upload information between the satellite and ground via the BeiDou-3 MEO global short message link. This enables rapid satellite-to-ground communication and guides participation in astronomical observations of sudden transient / outburst celestial bodies and events globally. When communicating with the BeiDou satellite, the EP satellite's SRM requires Doppler frequency offset compensation. According to the SRM design, spectrum compensation requires obtaining the satellite's real-time WGS84 orbital information (including position and velocity). The EP satellite has a GNSS receiver, which can provide real-time WGS84 position and velocity under normal operating conditions. However, due to the satellite's operational status, the GNSS receiver experiences positioning anomalies in certain areas. To address this and ensure the reliability of short message uplink and downlink communication, the satellite needs to convert the injected orbital information from the inertial coordinate system to the WGS84 coordinate system in real time. Therefore, there is an urgent need in this field to design a conversion technology that meets the requirements of real-time satellite-borne computing.

[0004] Currently, no descriptions or reports of technologies similar to this invention have been found, and no similar information has been collected domestically or internationally. Summary of the Invention

[0005] To address the aforementioned shortcomings in the prior art, this invention provides a satellite onboard orbit conversion method and system, along with a corresponding satellite-based computer and computer-readable storage medium.

[0006] According to a first aspect of the present invention, a satellite onboard orbit conversion method is provided, comprising: Calculate the corrected Julian day at the current time, and calculate the Greenwich Mean sidereal hour angle based on the corrected Julian day; Convert the Greenwich Mean Sidereal Time Angle into radians; Calculate the Earth's rotation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system based on the transformed radians; The pre-calculated Earth nutation and precession correction matrices are used as parameters that can be modified by satellite commands and are updated regularly. Based on the Earth's nutation and precession correction matrices and the Earth's rotation matrix, calculate the transformation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system; Based on the transformation matrix, the satellite position and velocity in the WGS84 coordinate system are calculated to achieve satellite orbit transformation.

[0007] Preferably, the calculation of the corrected Julian day at the current time includes: Let the current time be t. Based on the current time t, calculate the Julian diary entry as JD. Further calculate the corrected Julian day T as: The JD of Julian's Diary is calculated based on the current time t as follows: In the formula, Y, M, and D represent the year, month, and day (including the decimal part of the day) of the current time t. This indicates that the integer part of the content is taken, and the decimal places are omitted; The sidereal hour angle calculated based on the modified Julian day is: In the formula, The Greenwich Mean Sidereal Time Angle is 0.5.

[0008] Preferably, the conversion of the Greenwich Mean Sidereal Time angle into radians is as follows: In the formula, GMST is the converted radians, and GMST0 is the Greenwich Mean Sidereal Time. Pi is the mathematical constant of a circle.

[0009] Preferably, the step of calculating the Earth's rotation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system based on the transformed radians includes: In the formula, This is the Earth's rotation matrix.

[0010] Preferably, the step of calculating the transformation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system based on the Earth's nutation and precession correction matrix and the Earth's rotation matrix includes: In the formula, This is the transformation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system. This is the Earth's rotation matrix. This is the correction matrix for Earth's nutation and precession.

[0011] Preferably, calculating the satellite position and velocity in the WGS84 coordinate system based on the transformation matrix includes: In the formula, and These represent the satellite's position and velocity in the WGS84 coordinate system. and These represent the satellite's position and velocity in the J2000 inertial coordinate system. This is the transformation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system. This is a constant value, representing the rotational angular velocity matrix of the WGS84 coordinate system relative to the J2000 coordinate system.

[0012] According to a second aspect of the present invention, a satellite onboard orbit transfer system is provided, comprising: The Modified Julian Day Calculation Module is used to calculate the Modified Julian Day at the current time. A Greenwich Mean Sidereal Time Calculation Module, which calculates the Greenwich Mean Sidereal Time based on the Modified Julian Day; Angle-to-radian conversion module, which is used to convert Greenwich Mean Sidereal Time angles into radians; The Earth rotation matrix calculation module is used to calculate the Earth rotation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system based on the transformed radians. The correction matrix module is used to pre-calculate the Earth's nutation and precession correction matrices as parameters that can be modified by satellite commands, and to update them periodically. The coordinate system transformation matrix module is used to correct for Earth's nutation and precession matrices. Given the Earth's rotation matrix, calculate the transformation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system; The WGS84 position and velocity calculation module is used to calculate the position and velocity of satellites in the WGS84 coordinate system based on the transformation matrix.

[0013] According to a third aspect of the present invention, a spaceborne computer is provided, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor, when executing the computer program, can be used to perform the method described in any one of the above inventions.

[0014] According to a fourth aspect of the present invention, a computer-readable storage medium is provided having a computer program stored thereon, which, when executed by a processor, can be used to perform the method described in any one of the preceding inventions.

[0015] By adopting the above technical solution, the present invention has at least one of the following beneficial effects compared with the prior art: The satellite orbit conversion method and system provided by this invention belong to a satellite parameter conversion algorithm under the J2000 inertial coordinate system. By automatically converting the J2000 inertial coordinate system and the WGS84 coordinate system, the WGS84 position and velocity of the satellite can be provided to the short message communication device in real time on the orbit, so as to provide accurate satellite parameters and thus realize real-time satellite-to-ground communication of scientific exploration satellites.

[0016] The satellite orbit conversion method and system provided by this invention have been tested using simulation data. The calculation results of the conversion algorithm are compared with those of STK software. The results show that the algorithm can meet the requirements of the single-machine short message service for scientific exploration satellites.

[0017] The satellite orbit conversion method and system provided by this invention have designed a simplified and optimized algorithm for real-time on-orbit calculation, which reduces computing resources while ensuring the accuracy of the calculation results meets the requirements. This algorithm is a real-time on-orbit calculation algorithm for satellites, which has low requirements for computing resources and meets the accuracy requirements.

[0018] The satellite orbit conversion method and system provided by this invention are applicable to real-time on-orbit calculations of satellites. They have low requirements for on-board resources, high calculation accuracy, and strong applicability, and have good application and market prospects. Attached Figure Description

[0019] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings: Figure 1 This is a flowchart illustrating the satellite orbit conversion method in a preferred embodiment of the present invention.

[0020] Figure 2 This is a comparison chart of the satellite's position calculation results in the WGS84 coordinate system and the STK calculation results in a simulation example of the present invention. The blue result represents the calculation result of the transformation algorithm, and the red result represents the calculation result of STK.

[0021] Figure 3 This is a comparison chart of the satellite velocity calculation results of this algorithm and the STK calculation results in the WGS84 coordinate system in a simulation example of the present invention. In the chart, blue represents the calculation results of the transformation algorithm, and red represents the calculation results of STK.

[0022] Figure 4 This is an error diagram showing the satellite's position calculation results in the WGS84 coordinate system and the STK position simulation results in a simulation example of this invention.

[0023] Figure 5 This is an error diagram showing the velocity calculation results of the satellite in the WGS84 coordinate system and the STK velocity simulation results of the algorithm in a simulation example of the present invention. Detailed Implementation

[0024] The embodiments of the present invention are described in detail below: These embodiments are implemented based on the technical solution of the present invention, and provide detailed implementation methods and specific operation processes. It should be noted that those skilled in the art can make several modifications and improvements without departing from the concept of the present invention, and these all fall within the protection scope of the present invention.

[0025] When the EP satellite's short message communication unit communicates with the BeiDou satellite, it needs to compensate for Doppler frequency offset. According to the design of the short message communication unit, spectrum compensation requires obtaining the satellite's real-time WGS84 orbit information (including position and velocity). Under normal GNSS receiver operation, the EP satellite can provide real-time WGS84 position and velocity. However, depending on the on-orbit satellite's operation, the GNSS receiver experiences positioning anomalies in certain areas. To address this and ensure the reliability of short message uplink and downlink communication, the satellite needs to convert the injected orbit information from the inertial coordinate system to the WGS84 coordinate system in real time. However, existing technologies lack a conversion technique that meets the requirements of real-time satellite-borne computing.

[0026] To address the aforementioned issues, one embodiment of the present invention provides a satellite onboard orbit conversion method. This method, which addresses the requirement of converting EP satellite onboard orbit information from an inertial frame to a WGS84 coordinate system, presents an orbit conversion algorithm. This algorithm can reduce the onboard computing resource requirements, meet the single-machine usage requirements of short message service, and provide a reference for on-orbit conversion of other satellites.

[0027] Specifically, such as Figure 1 As shown, the satellite orbit conversion method provided in this embodiment may include: S1, Calculate the corrected Julian day T at the current time; in some preferred embodiments, this step may further include: Let the current time be t. Based on the current time t, calculate the Julian diary entry as JD. Further calculate the corrected Julian day T as: (1) Furthermore, based on the current time t, the JD of Julian's diary is calculated as follows: M and D represent the year, month, and day (including the decimal part of the day) of the current time t, respectively. This indicates that the integer part of the content is taken, and the decimal places are omitted; S2, calculate the Greenwich Mean Sidereal Time Angle GMST0 (in degrees) based on the Modified Julian Day. In some preferred embodiments, this step may further include: The Greenwich Mean Sidereal Time (GMST0) in degrees is calculated as follows: (2) S3, convert the Greenwich Mean Sidereal Time (GMST0) into radians (rad). In some preferred embodiments, this step may further include: Converting the Greenwich Mean Sidereal Time (GMST0) to radians (rad) gives: (3) In the formula, Pi is the mathematical constant of a circle.

[0028] S4. Calculate the Earth's rotation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system based on the transformed radians. In some preferred embodiments, this step may further include: Earth's rotation matrix for: (4) S5, pre-calculate the Earth's nutation and precession correction matrices as parameters that can be modified by satellite commands, and update them periodically. In some preferred embodiments, this step may further include: Pre-calculate the Earth's nutation and precession correction matrices, denoted as... Because the formulas for calculating Earth's nutation and precession are extremely complex, and changes in this matrix have little impact on the calculation results in the short term, this matrix is ​​stored in the satellite computer as a parameter that can be modified by satellite commands and is updated regularly to ensure accuracy.

[0029] S6, based on the Earth's nutation and precession correction matrix and Earth's rotation matrix Calculate the transformation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system. In some preferred embodiments, this step may further include... Transformation matrix for: (5) S7, based on the transformation matrix Calculate satellite positions in the WGS84 coordinate system and speed This enables the satellite's onboard orbit transfer. In some preferred embodiments, this step may further include: Location and speed They are respectively: (6) (7) In the formula, and These represent the satellite's position and velocity in the J2000 inertial coordinate system. This is a constant value, representing the rotational angular velocity matrix of the WGS84 coordinate system relative to the J2000 coordinate system.

[0030] Based on the same inventive concept, one embodiment of the present invention also provides a satellite onboard orbit conversion system.

[0031] Specifically, the satellite onboard orbit conversion system provided in this embodiment may include: The Modified Julian Day Calculation Module is used to calculate the Modified Julian Day at the current time. The Greenwich Mean Sidereal Time Calculation Module calculates the Greenwich Mean Sidereal Time based on the Modified Julian Day. Angle-to-radian conversion module, which is used to convert Greenwich Mean Sidereal Time angles into radians; The Earth rotation matrix calculation module is used to calculate the Earth rotation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system based on the transformed radians. The correction matrix module is used to pre-calculate the Earth's nutation and precession correction matrices as parameters that can be modified by satellite commands, and to update them periodically. The coordinate system transformation matrix module is used to correct for Earth's nutation and precession matrices. Given the Earth's rotation matrix, calculate the transformation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system; The WGS84 position and velocity calculation module is used to calculate the position and velocity of satellites in the WGS84 coordinate system based on the transformation matrix.

[0032] The implementation of each functional module constituting the system provided in this embodiment is further described in detail below.

[0033] In this embodiment, the satellite's position and velocity in the J2000 inertial coordinate system are denoted as follows: , Calculate the position and velocity in the WGS84 coordinate system, denoted as follows: , The specific implementation methods of each functional module are shown in the simulation example below.

[0034] The revised Julian day calculation module, assuming the current time is t, calculates the Julian diary entry as JD based on the current time, and further calculates the revised Julian day T: (1) The Greenwich Mean Sidereal Time Calculation Module calculates the Greenwich Mean Sidereal Time GMST0 in degrees. (2) Angle-to-radian conversion module converts Greenwich Mean Sidereal Time (GMST0) into radians, unit: rad. (3) The Earth rotation matrix calculation module calculates the Earth rotation matrix from J2000 to WGS84: (4) The correction matrix module calculates the correction matrices for Earth's nutation and precession, denoted as... Because the formulas for calculating Earth's precession and nutation are extremely complex, and changes in this matrix in the short term have little impact on the calculation results, this matrix is ​​stored in the satellite computer as a parameter that can be modified by satellite commands and is updated regularly to ensure accuracy.

[0035] The coordinate system transformation matrix module calculates the transformation matrix from J2000 to WGS84 coordinate system: (5) The WGS84 position and velocity calculation module calculates satellite position and velocity in the WGS84 coordinate system. (6) (7) Among them, the rotational angular velocity matrix of WGS84 relative to the J2000 series is constant. .

[0036] It should be noted that the steps in the method provided by the present invention can be implemented using the corresponding components in the system. Those skilled in the art can refer to the technical solution of the system to implement the steps of the method, and can also refer to the technical solution of the method to implement the composition of the system. That is, the embodiments in the system and the embodiments in the method can be understood as preferred examples of each other, which will not be elaborated here.

[0037] An embodiment of the present invention also provides a space service computer, including a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, it can be used to perform any of the methods described in the above embodiments of the present invention.

[0038] Optionally, the memory is used to store programs; the memory may include volatile memory, such as random-access memory (RAM), such as static random-access memory (SRAM), double data rate synchronous dynamic random-access memory (DDR SDRAM), etc.; the memory may also include non-volatile memory, such as flash memory. The memory is used to store computer programs (such as application programs and functional modules that implement the above methods), computer instructions, etc., and the aforementioned computer programs and computer instructions can be partitioned and stored in one or more memories. Furthermore, the aforementioned computer programs, computer instructions, data, etc., can be accessed by the processor.

[0039] A processor is used to execute computer programs stored in memory to implement the various steps of the methods or various modules of the systems involved in the above embodiments. For details, please refer to the relevant descriptions in the preceding method and system embodiments.

[0040] The processor and memory can be separate structures or integrated structures. When the processor and memory are separate structures, they can be coupled together via a bus.

[0041] An embodiment of the present invention also provides a computer-readable storage medium having a computer program stored thereon, which, when executed by a processor, can be used to perform the methods described above in the present invention.

[0042] Computer-readable media include computer storage media and communication media, wherein communication media include any medium that facilitates the transfer of computer programs from one place to another. Storage media can be any available medium accessible to a general-purpose or special-purpose computer. An exemplary storage medium is coupled to a processor, enabling the processor to read information from and write information to the storage medium. Of course, the storage medium can also be a component of the processor. The processor and storage medium can reside in an ASIC. Alternatively, the ASIC can reside in a user device. Of course, the processor and storage medium can also exist as discrete components in a communication device.

[0043] The effectiveness of the technical solution provided by the above embodiments of the present invention will be further explained in detail below with reference to a specific simulation example.

[0044] In this specific simulation example, the orbital start time of a certain satellite is set to 07:00:00 on 2024 / 1 / 16. The position and velocity of the satellite in the J2000 inertial frame at this moment are shown in Table 1.

[0045] Table 1 The orbital transformation algorithm described above was used to simulate the position and velocity transformation results of the satellite over a one-day flight period, and the results were compared with those obtained using STK simulation. Figures 2-5 As shown. Wherein: Figure 2 The figure shows the change of the satellite's position in the XYZ direction in the WGS84 coordinate system over time. The blue line represents the simulation result calculated by this transformation algorithm, and the red line represents the STK simulation result. Since the two positions are extremely close, it is difficult to distinguish them in this figure. Therefore, the two lines shown in the figure almost overlap.

[0046] Figure 3 The graph shows the satellite's velocity in the XYZ directions over time in the WGS84 coordinate system. The blue line represents the simulation result calculated by this transformation algorithm, while the red line represents the STK simulation result. Since the velocities of the two are extremely close, it is difficult to distinguish them in this graph. Therefore, the two lines shown in the graph almost overlap.

[0047] Figure 4 To determine the position error of the satellite in the WGS84 coordinate system using this transformation algorithm and STK simulation, Figure 5The proposed conversion algorithm and STK simulation were used to measure the velocity error of the satellite in the WGS84 coordinate system. Simulation results show that the average position error of the satellite calculated by the proposed conversion algorithm in the WGS84 coordinate system is 245.75m, with a maximum error of 272.12m; the average velocity error is 0.25m / s, with a maximum error of 0.28m / s. These results meet the requirements of short message communication equipment spectrum compensation, which requires a position error within 500m and a velocity error within 0.5m / s. The simulation results demonstrate the effectiveness of the proposed conversion algorithm.

[0048] The satellite orbit conversion method and system provided in the above embodiments of the present invention address the requirement of converting EP satellite orbit information from an inertial frame to a WGS84 coordinate system. An orbit conversion algorithm is implemented, applicable to satellite-based computer calculations. The algorithm was verified using simulation data and compared with calculation results from STK software. The results show that the algorithm meets the requirements for short message stand-alone operation and can provide a reference for on-orbit conversion of other satellites.

[0049] Any matters not covered in the above embodiments of the present invention are well-known in the art.

[0050] The specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various modifications or variations within the scope of the claims, which do not affect the essence of the present invention.

Claims

1. A satellite onboard orbit conversion method, characterized in that, include: Calculate the corrected Julian day at the current time, and calculate the Greenwich Mean sidereal hour angle based on the corrected Julian day; Convert the Greenwich Mean Sidereal Time Angle into radians; Calculate the Earth's rotation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system based on the transformed radians; The pre-calculated Earth nutation and precession correction matrices are used as parameters that can be modified by satellite commands and are updated regularly. Based on the Earth's nutation and precession correction matrices and the Earth's rotation matrix, calculate the transformation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system; Based on the transformation matrix, the satellite position and velocity in the WGS84 coordinate system are calculated to achieve satellite orbit transformation.

2. The satellite orbit conversion method according to claim 1, characterized in that, The calculation of the corrected Julian day at the current time includes: Let the current time be t. Based on the current time t, calculate the Julian diary entry as JD. Further calculate the corrected Julian day T as: The JD of Julian's Diary is calculated based on the current time t as follows: In the formula, Y, M, and D represent the year, month, and day of the current time t, respectively, including the decimal part of the day. Indicates taking the integer part; The sidereal hour angle calculated based on the modified Julian day is: In the formula, The Greenwich Mean Sidereal Time Angle is 0.

5.

3. The satellite orbit conversion method according to claim 1, characterized in that, The conversion of Greenwich Mean Sidereal Time Angle into radians is as follows: In the formula, GMST is the converted radians, and GMST0 is the Greenwich Mean Sidereal Time. Pi is the mathematical constant of a circle.

4. The satellite orbit conversion method according to claim 1, characterized in that, The calculation of the Earth's rotation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system based on the transformed radians includes: In the formula, This is the Earth's rotation matrix.

5. The satellite orbit conversion method according to claim 1, characterized in that, The calculation of the transformation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system based on the Earth's nutation and precession correction matrices and the Earth's rotation matrix includes: In the formula, This is the transformation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system. This is the Earth's rotation matrix. This is the correction matrix for Earth's nutation and precession.

6. The satellite onboard orbit conversion method according to claim 1, characterized in that, The step of calculating the satellite position and velocity in the WGS84 coordinate system based on the transformation matrix includes: In the formula, and These represent the satellite's position and velocity in the WGS84 coordinate system. and These represent the satellite's position and velocity in the J2000 inertial coordinate system. This is the transformation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system. This is a constant value, representing the rotational angular velocity matrix of the WGS84 coordinate system relative to the J2000 coordinate system.

7. A satellite-borne orbit conversion system, characterized in that, include: The Modified Julian Day Calculation Module is used to calculate the Modified Julian Day at the current time. A Greenwich Mean Sidereal Time Calculation Module, which calculates the Greenwich Mean Sidereal Time based on the Modified Julian Day; Angle-to-radian conversion module, which is used to convert Greenwich Mean Sidereal Time angles into radians; The Earth rotation matrix calculation module is used to calculate the Earth rotation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system based on the transformed radians. The correction matrix module is used to pre-calculate the Earth's nutation and precession correction matrices as parameters that can be modified by satellite commands, and to update them periodically. The coordinate system transformation matrix module is used to correct for Earth's nutation and precession matrices. Given the Earth's rotation matrix, calculate the transformation matrix from the J2000 inertial coordinate system to the WGS84 coordinate system; The WGS84 position and velocity calculation module is used to calculate the position and velocity of satellites in the WGS84 coordinate system based on the transformation matrix.

8. A spacefaring computer, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that, When the processor executes the computer program, it can be used to perform the method of any one of claims 1-6.

9. A computer-readable storage medium having a computer program stored thereon, characterized in that, When executed by a processor, the computer program can be used to perform the method of any one of claims 1-6.

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