Spacecraft orbital element data management method and device

Abstract

The invention discloses a spacecraft orbital element data management method and device, and the method comprises the steps: obtaining an orbital element set, sorting the orbital elements in the orbital element set based on the epoch time from early to late, and obtaining a to-be-managed orbital element set; traversing the orbital elements to be treated in the orbital element set to be treated in sequence, and performing deletion completion processing on the orbital elements to be treated based on a preset deletion completion rule to obtain a first orbital element; performing exception correction processing on the first orbital element based on a preset exception correction rule to obtain a second orbital element; carrying out redundancy deduplication processing on the second orbit elements based on a preset redundancy deduplication rule to obtain a target orbit element, and entering the next traversal; and obtaining target orbital element data until all orbital elements to be treated in the orbital element set to be treated are traversed, and completing orbital element data treatment.

Description

Technical Field

[0001] The present invention relates to the technical field of spacecraft measurement, operation and control, and in particular to a method and device for governing spacecraft orbital element data. Background Art

[0002] The orbital elements of a spacecraft are a set of parameters used to describe the position and motion characteristics of the spacecraft in orbit, mainly divided into six Keplerian elements and two-line elements. The six Keplerian elements include six parameters: semi-major axis of the orbit, eccentricity, inclination of the orbit, right ascension of the ascending node, argument of perigee, and mean anomaly. They are adapted to the High Precision Orbit Propagator (HPOP) model and are mainly used for short-term orbit prediction. The two-line elements (TLE) are a simplified and standardized method for representing spacecraft orbital data and are widely used for quickly publishing and updating the orbital information of Earth-orbiting spacecraft. The two-line elements consist of two lines of text, each containing specific orbital parameters and check codes. These parameters are the comprehensive result of Keplerian orbital elements and some other perturbation factors (such as non-spherical gravitational perturbation of the Earth, gravitational perturbations of the Sun and the Moon, etc.), but are given in a form that is more convenient for dissemination and rapid update. The two-line elements are adapted to the Simplified General Perturbations 4 (SGP4) model and are mainly used for medium- and long-term orbit prediction.

[0003] Due to the high timeliness requirement of observation needs, the operation and control of remote sensing satellites generally adopt a daily plan operation mode, that is, every day, according to the received observation needs, the satellite work plan for the next day is formulated. Considering the accuracy requirements, the six Keplerian elements are used for orbit prediction. The satellite operator receives the orbital elements from the satellite measurement and control center and saves them in the database. Due to orbit change control, network congestion or other reasons, sometimes multiple sets of orbital element data at the same epoch time are received. The business system generally selects the latest set of orbital elements by the storage time for orbit prediction to support subsequent mission planning and simulation deduction. The orbital elements at the same epoch time may be pre-control orbital elements, post-control orbital elements, or there may be errors in the orbital element data due to human error in the file receiving and sending link, or data loss during storage in the database, resulting in missing orbital element data. Therefore, there are various situations such as missing, abnormal, and redundant orbital element data in the database, which brings inconvenience to the application of orbital elements. Whether the business system selects the orbital elements with the latest or earliest storage time may lead to large errors in orbit prediction, thus affecting the actual effect of the satellite photographing ground targets. Summary of the Invention

[0004] The present invention provides a method and device for governing spacecraft orbital element data, so as to correct the orbital elements and improve the accuracy of orbital prediction.

[0005] To solve the above technical problems, an embodiment of the present invention provides a method for governing spacecraft orbital element data, including: Obtain an orbital element set, sort the orbital elements in the orbital element set in ascending order of epoch time and storage time, and obtain an orbital element set to be governed; Traverse the orbital elements to be governed in the orbital element set to be governed in sequence, perform missing value filling processing on the orbital elements to be governed based on a preset missing value filling rule to obtain a first orbital element; perform abnormal correction processing on the first orbital element based on a preset abnormal correction rule to obtain a second orbital element; perform redundancy removal processing on the second orbital element based on a preset redundancy removal rule to obtain a target orbital element, and proceed to the next traversal; Until all the orbital elements to be governed in the orbital element set to be governed are traversed, obtain the target orbital element data, and complete the governance of the orbital element data.

[0006] By precisely sorting the orbital element set according to the epoch time and storage time, and successively performing missing value filling, abnormal correction, and redundancy removal processing on each orbital element, the present invention can systematically clean and improve various orbital element data in the database, enhancing the integrity and consistency of the orbital element data; at the same time, traversing all the data to be governed at once, processing and saving the governance results in order simplifies the call process of the orbital elements in the business system, reduces the risks of human misselection and system errors, and thus provides reliable and accurate orbital reference data for subsequent satellite mission planning and simulation deduction.

[0007] Further, the performing missing value filling processing on the orbital elements to be governed based on a preset missing value filling rule to obtain a first orbital element includes: Determine a reference orbital element based on the epoch time of the orbital element to be governed, perform filling processing on the orbital circle number of the orbital element to be governed based on the reference orbital element to obtain a first filled orbital element; Match the orbit control record based on the epoch time of the orbital element to be governed, perform orbit control filling processing on the first filled orbital element based on the orbit control record to obtain a second filled orbital element; Match the target orbital element based on the epoch time of the orbital element to be governed, perform Kepler filling processing on the second filled orbital element based on the six Kepler parameters of the target orbital element to obtain a first orbital element.

[0008] The present invention performs multi-dimensional missing complementation processing of orbital elements based on epoch time, introducing operations such as complementing the orbit circle number of the reference orbital elements, complementing the orbit control records, and complementing the six Keplerian elements, significantly improving the information completeness of orbital elements in the time series and physical parameter dimensions, thereby enhancing the overall continuity and physical consistency of orbital data.

[0009] Further, determining the reference orbital elements based on the epoch time of the orbital elements to be governed, and performing complementation processing on the orbit circle number of the orbital elements to be governed based on the reference orbital elements to obtain the first complemented orbital elements, includes: Calculating the epoch time difference between the target orbital elements and the orbital elements to be governed, and taking the target orbital elements with the smallest absolute value of the epoch time difference as the reference orbital elements; Obtaining the orbital period and circle change point information of the reference orbital elements, and calculating the circle difference based on the orbital period, circle change point information, and epoch time difference; Using the orbit circle number of the reference orbital elements plus the circle difference as the orbit circle number of the orbital elements to be governed, completing the complementation processing, and obtaining the first complemented orbital elements.

[0010] The present invention can realize the dynamic recovery and automatic calculation of the orbit circle number by introducing the orbital period and circle change point information of the reference orbital elements, and combining the epoch time difference to calculate the circle difference and then complement the orbit circle number, improving the data complementation efficiency and accuracy.

[0011] Further, performing abnormal correction processing on the first orbital elements based on a preset abnormal correction rule to obtain the second orbital elements, includes: Determining whether there are orbit circle number abnormalities and six Keplerian element abnormalities in the first orbital elements according to the reference orbital elements; the reference orbital elements are the governed target orbital elements; When there are orbit circle number abnormalities in the first orbital elements, performing abnormal correction processing on the orbit circle number of the first orbital elements according to the orbit circle number and orbit circle difference of the reference orbital elements to obtain the second orbital elements; When there are six Keplerian element abnormalities in the first orbital elements, performing orbit extrapolation respectively according to the reference orbital elements and the first orbital elements, calculating the position error at the same moment based on the ephemeris data, and performing abnormal judgment processing according to the position error to obtain the second orbital elements.

[0012] The present invention uses the governed target orbital elements as the reference orbital elements to perform correction and marking processing on the orbit circle number abnormalities and six Keplerian element abnormalities respectively, which can not only effectively correct the systematic deviation in the orbital data, but also establish an explicit abnormal identification mechanism for the data with physical abnormalities, thereby improving the credibility and usability of the orbital element data.

[0013] Further, the redundant duplicate removal process for the second orbital elements based on the preset redundant duplicate removal rule to obtain the target orbital elements includes: Screen out all the orbital elements with the same epoch time as the second orbital elements in the orbital element data as the redundant orbital element set; Calculate the similarity between any two orbital elements in the redundant orbital element set. When the similarity is greater than the preset threshold, mark the orbital element with an earlier storage time as redundant; Remove the orbital elements marked as redundant in the redundant orbital element set to obtain the target orbital elements.

[0014] The present invention identifies the redundant data of the orbital elements through the epoch time screening and similarity calculation methods, and performs retention and removal processing based on the sequence of storage times, effectively realizing the redundant duplicate removal within the data set. While improving the uniqueness and storage utilization efficiency of the orbital element data, this solution also avoids the orbital determination conflict problem caused by multi-source data under the same epoch time, further ensuring the consistency and processing performance of the orbital element data.

[0015] In a second aspect, the present invention provides a spacecraft orbital element data governance device, including: a sorting module, a governance module, and a storage module; The sorting module is used to obtain an orbital element set, sort the orbital elements in the orbital element set in ascending order based on the epoch time and storage time, and obtain the orbital elements to be governed; The governance module is used to sequentially traverse the orbital elements to be governed in the orbital elements to be governed set, perform missing complement processing on the orbital elements to be governed based on the preset missing complement rule to obtain the first orbital elements; perform abnormal correction processing on the first orbital elements based on the preset abnormal correction rule to obtain the second orbital elements; perform redundant duplicate removal processing on the second orbital elements based on the preset redundant duplicate removal rule to obtain the target orbital elements, and proceed to the next traversal; The storage module is used to complete the governance of the orbital element data until all the orbital elements to be governed in the orbital elements to be governed set are traversed and the target orbital element data is obtained.

[0016] Further, the governance module is used to perform missing complement processing on the orbital elements to be governed based on the preset missing complement rule to obtain the first orbital elements, including: Determine the reference orbital elements based on the epoch time of the orbital elements to be governed, and perform complement processing on the orbital circle number of the orbital elements to be governed based on the reference orbital elements to obtain the first complemented orbital elements; Epoch time matching orbit control records based on the orbit elements to be rectified, and performing orbit control completion processing on the first completed orbit elements based on the orbit control records to obtain second completed orbit elements; Epoch time matching target orbit elements based on the orbit elements to be rectified, and performing Kepler completion processing on the second completed orbit elements based on the six Kepler parameters of the target orbit elements to obtain first orbit elements.

[0017] Furthermore, the rectification module is used to determine reference orbit elements based on the epoch time of the orbit elements to be rectified, and perform completion processing on the orbit circle number of the orbit elements to be rectified based on the reference orbit elements to obtain first completed orbit elements, including: Calculating the epoch time difference between the target orbit elements and the orbit elements to be rectified, and taking the target orbit elements with the smallest absolute value of the epoch time difference as the reference orbit elements; Obtaining the orbit period and orbit change point information of the reference orbit elements, and calculating the number of orbit differences based on the orbit period, orbit change point information and epoch time difference; Taking the orbit circle number of the reference orbit elements plus the number of orbit differences as the orbit circle number of the orbit elements to be rectified, completing the completion processing, and obtaining first completed orbit elements.

[0018] Furthermore, the rectification module is used to perform anomaly correction processing on the first orbit elements based on a preset anomaly correction rule to obtain second orbit elements, including: Determining whether there are anomalies in the orbit circle number and Kepler six parameters of the first orbit elements according to the reference orbit elements; the reference orbit elements are the rectified target orbit elements; When there is an anomaly in the orbit circle number of the first orbit elements, anomaly correction processing is performed on the orbit circle number of the first orbit elements according to the orbit circle number and orbit difference of the reference orbit elements to obtain second orbit elements; When there are anomalies in the Kepler six parameters of the first orbit elements, orbit extrapolation is performed according to the reference orbit elements and the first orbit elements respectively, the position error at the same moment is calculated based on the ephemeris data, and anomaly judgment processing is performed according to the position error to obtain second orbit elements.

[0019] Furthermore, the rectification module is used to perform redundancy removal processing on the second orbit elements based on a preset redundancy removal rule to obtain target orbit elements, including: Screening out all orbit elements with the same epoch time as the second orbit elements in the orbit element data as the redundant orbit element set; Calculate the similarity between any two sets of orbital elements in the redundant set of orbital elements. When the similarity is greater than the preset threshold, mark the set of orbital elements with an earlier storage time as redundant. Remove the orbital elements marked as redundant from the redundant set of orbital elements to obtain the target orbital elements. Description of the Drawings

[0020] Figure 1 It is a schematic flow chart of a method for managing spacecraft orbital element data provided in an embodiment of the present invention. Figure 2 It is another schematic flow chart of a method for managing spacecraft orbital element data provided in an embodiment of the present invention. Figure 3 It is a schematic flow chart of a method for completing the orbit circle number provided in an embodiment of the present invention. Figure 4 It is a schematic flow chart of a method for completing the orbit control flag provided in an embodiment of the present invention. Figure 5 It is a schematic flow chart of a method for completing the six Kepler parameters provided in an embodiment of the present invention. Figure 6 It is a schematic flow chart of a method for correcting abnormal orbit circle numbers provided in an embodiment of the present invention. Figure 7 It is a schematic flow chart of a method for correcting abnormal six Kepler parameters provided in an embodiment of the present invention. Figure 8 It is a schematic flow chart of a method for redundant deduplication provided in an embodiment of the present invention. Detailed Embodiments

[0021] The following combines the drawings and embodiments to further describe in detail the specific embodiments of the present invention. The following embodiments are used to illustrate the present invention, but are not used to limit the scope of the present invention.

[0022] The terms "first" and "second" in the specification, claims and drawings of this application are used to distinguish different objects, rather than to describe a specific order. In addition, the terms "comprising" and "having" and any variations thereof are intended to cover non-exclusive inclusion. For example, a process, method, system, product or device that includes a series of steps or units is not limited to the listed steps or units, but may optionally further include steps or units not listed, or may optionally further include other steps or units inherent to these processes, methods, products or devices.

[0023] References to "embodiments" in this specification mean that the particular features, structures, or characteristics described in connection with the embodiments can be included in at least one embodiment of the present application. The phrase appears in various places in the specification and does not necessarily refer to the same embodiment, nor is it an independent or alternative embodiment mutually exclusive with other embodiments. Those skilled in the art will explicitly and implicitly understand that the embodiments described herein can be combined with other embodiments.

[0024] Embodiment 1 See Figure 1 , Figure 1 which is a schematic flowchart of a method for governing spacecraft orbital element data provided in an embodiment of the present invention. An embodiment of the present invention provides a method for governing spacecraft orbital element data, including steps 101 to 103, as follows: Step 101: Obtain a set of orbital elements, sort the orbital elements in the set of orbital elements in ascending order based on the epoch time and the storage time, and obtain a set of orbital elements to be governed. Step 102: Sequentially traverse the orbital elements to be governed in the set of orbital elements to be governed, perform missing value filling processing on the orbital elements to be governed based on a preset missing value filling rule to obtain a first set of orbital elements; perform anomaly correction processing on the first set of orbital elements based on a preset anomaly correction rule to obtain a second set of orbital elements; perform redundancy removal processing on the second set of orbital elements based on a preset redundancy removal rule to obtain target orbital elements, and proceed to the next traversal. Step 103: Until all the orbital elements to be governed in the set of orbital elements to be governed have been traversed, obtain target orbital element data and complete the governance of orbital element data.

[0025] In this embodiment, first, all orbital element data is obtained from the database, and these orbital elements are sorted in ascending order according to the epoch time. When the epoch times are the same, they are sorted in ascending order according to the storage time. Then, missing value filling, anomaly correction, and redundancy removal are sequentially performed on the sorted orbital elements, and corresponding marks are added. Each correctly governed orbital element can be used as a reference orbital element for subsequent governance of orbital elements, and then the governed orbital elements are saved to the database.

[0026] In this embodiment, by accurately sorting the orbital element number set according to the epoch time and the storage time, and performing missing completion, anomaly correction and redundant deduplication processing on each orbital element in turn, it is possible to systematically clean and improve various types of orbital element number data in the database, and improve the integrity and consistency of the orbital element number data; at the same time, all the data to be governed are traversed at one time, and the governance results are processed in sequence and saved, which simplifies the calling process of the business system for the orbital element numbers, reduces the risks of human misselection and system errors, and thus provides reliable and accurate orbital reference data for subsequent satellite mission planning and simulation deduction.

[0027] Please refer to Figure 2 , Figure 2 Another flowchart of a method for managing spacecraft orbital element data provided by an embodiment of the present invention.

[0028] In this embodiment, all the orbital element data of spacecraft are first extracted from the database and sorted from early to late according to the epoch time; for records with the same epoch time, they are further sorted from early to late according to the storage time. The system performs missing completion, abnormal correction and redundant deduplication operations on the first orbital element after sorting, and sets it as the benchmark element after the treatment is completed; then, the remaining orbital elements are completed, corrected and deduplicated one by one according to the same rules until all records have completed data treatment, and finally all the treated target orbital elements are uniformly saved in the database.

[0029] In this embodiment, the target orbital element number after governance is the reference orbital element number. The reference orbital element number in each traversal and each governance can be determined according to the current orbital element number to be governed.

[0030] In this embodiment, the step of performing missing completion processing on the number of orbital elements to be managed based on a preset missing completion rule to obtain the first number of orbital elements includes: Determine a reference orbital element based on the epoch time of the orbital element to be governed, and complete the orbital circle number of the orbital element to be governed based on the reference orbital element to obtain a first completed orbital element; Matching orbit control records based on the epoch time of the orbit element to be governed, performing orbit control complement processing on the first complemented orbit element based on the orbit control record, and obtaining a second complemented orbit element; The target orbital element is matched based on the epoch time of the orbital element to be governed, and the second complement orbital element is subjected to Kepler complement processing based on the Kepler six parameters of the target orbital element to obtain the first orbital element.

[0031] In this embodiment, first, according to the epoch time of the orbital element to be governed, the record closest to the epoch time is selected from the orbital element library that has been governed and precision-checked as the reference orbital element, where "closest" means that the time difference is the smallest and the reference record has been marked with an accuracy mark; then, the orbital period and circle change point information of the circle where the reference orbital element is located are extracted, and the circle difference ΔN is calculated in combination with the epoch time difference between the element to be governed and the reference element, and the first completed orbital circle number is obtained by adding ΔN to the original circle number of the reference element, thereby forming the first completed orbital element.

[0032] Next, read the orbit control record, which contains the start and end time of each orbit control operation, and determine its state before or after control by matching the epoch time of the element number to be governed with the record time period range; write the orbit control flag into the first completed orbit element number to obtain the second completed orbit element number.

[0033] Finally, the target orbital element with the same epoch time as the element to be governed and containing the complete Kepler six parameters is retrieved, and its orbital semi-major axis, eccentricity, orbital inclination, right ascension of the ascending node, argument of perigee and mean anomaly are assigned to the second completed orbital element one by one to complete the six-parameter calibration, thereby outputting the first orbital element that is complete and verified.

[0034] In this embodiment, the missing orbital element data includes missing orbital circle number, missing orbital control mark and missing Kepler six parameters. For missing orbital circle number, the correct orbital circle number can be obtained by calculating the orbital circle difference between the orbital element to be governed and the reference orbital element. For missing orbital control mark, by judging the relationship between the orbital element to be governed and the orbital control recording period range, it can be determined whether it is the pre-control element or the post-control element. For missing Kepler six parameters, the six parameters of the orbital elements at the same epoch time are obtained to complete the missing Kepler six parameters.

[0035] Please refer to Figure 3 , Figure 3 A schematic diagram of a track circle number completion process provided in an embodiment of the present invention.

[0036] In this embodiment, determining a reference orbital element based on the epoch time of the orbital element to be governed, and completing the orbital circle number of the orbital element to be governed based on the reference orbital element to obtain a first completed orbital element includes: Calculating the epoch time difference between the target orbital element and the orbital element to be governed, and taking the target orbital element with the smallest absolute value of the epoch time difference as the reference orbital element; Obtaining the orbital period and the circle change point information of the reference orbital element, and calculating the circle difference based on the orbital period, the circle change point information and the epoch time difference; Add the orbit circle number based on the reference orbit elements and the difference in the number of orbits as the orbit circle number of the orbit elements to be corrected, complete the complement process, and obtain the first complement orbit elements.

[0037] In this embodiment, when it is found that the orbit circle number recorded in a certain orbit element is missing, the system will retrieve the reference orbit elements closest to the epoch time of this record, calculate the epoch time difference between the two accordingly, and extract the orbit period information of the orbit where the reference orbit elements are located at the same time; Subsequently, use the reference elements to perform orbit prediction to obtain the ephemeris data for the corresponding time period, and determine the orbit change point closest to the target epoch time in the ephemeris result; Combining the time difference, orbit period, and the position of the orbit change point, the difference in the number of orbits can be calculated, and this difference is added to the orbit circle number of the reference elements to generate the circle number of the record to be complemented. Finally, add a missing complement flag to this record for distinction.

[0038] Please refer to Figure 4 , Figure 4 which is a schematic diagram of a track control flag complement process provided by an embodiment of the present invention.

[0039] In this embodiment, first obtain the complete track control records, and compare the epoch time of the orbit elements to be processed with the start and end times of each track control record; If the epoch time falls within the control time range of a certain track control record, the "before control" mark is given to the orbit elements; If its epoch time is after the end of a certain track control record and does not exceed the preset threshold range, the "after control" mark is given; If neither of them is satisfied, the original state is maintained without marking.

[0040] Please refer to Figure 5 , Figure 5 which is a schematic diagram of a Kepler six-parameter complement process provided by an embodiment of the present invention.

[0041] In this embodiment, when it is detected that the Kepler six-parameters of a certain orbit element record are incomplete, the system will retrieve the auxiliary record with the same epoch time and complete six-parameters; If a complete record that meets the conditions is found, directly use the six-parameters of this record to complement the target record, and add a missing complement flag to this record for subsequent tracking and auditing; If the corresponding complete record cannot be found, the original record parameters are retained and the complement process is terminated.

[0042] In this embodiment, the abnormal conditions of the orbit element data include abnormal orbit circle numbers and abnormal Kepler six-parameters. For abnormal orbit circle numbers, compare the correctly calculated orbit circle number with the orbit circle number of the orbit elements to be corrected to determine whether the orbit circle number is abnormal. For abnormal Kepler six-parameters, calculate the ephemeris data for a future period based on the orbit elements to be corrected and the reference orbit elements, and determine whether the position error between the two ephemeris data at the same moment is greater than the threshold to determine whether the Kepler six-parameters are abnormal.

[0043] In this embodiment, the abnormal correction process of the first orbital elements based on a preset abnormal correction rule to obtain the second orbital elements includes: Determining whether there are abnormal orbital circle numbers and abnormal Kepler six parameters in the first orbital elements according to the reference orbital elements; the reference orbital elements are the target orbital elements after treatment; When there is an abnormal orbital circle number in the first orbital elements, the abnormal orbital circle number of the first orbital elements is corrected according to the orbital circle number and the orbital circle difference of the reference orbital elements to obtain the second orbital elements; When there are abnormal Kepler six parameters in the first orbital elements, orbital extrapolation is respectively performed according to the reference orbital elements and the first orbital elements, the position error at the same moment is calculated based on the ephemeris data, and abnormal judgment processing is performed according to the position error to obtain the second orbital elements.

[0044] In this embodiment, to effectively correct the possible abnormal information in the first orbital elements, the system first makes an abnormal judgment on the current first orbital elements based on a preset abnormal correction rule, using the treated target orbital elements as the reference orbital elements. This judgment process includes two dimensions: one is the logical consistency of the orbital circle number, and the other is the orbital fitting rationality of the Kepler six parameters. If it is judged that there is a difference between the circle number of the current orbital elements and the circle number that should be possessed by the reference orbital elements at the corresponding epoch time, the system calculates the epoch time difference between the two, calculates the orbital circle difference in combination with the orbital period and the circle change point information of the reference orbital elements, and corrects the orbital circle number of the current orbital elements based on this circle difference to obtain the second orbital elements. Further, the system judges whether there are abnormal Kepler six parameters in the current orbital elements, such as extrapolating the ephemeris data of the corresponding time period using the orbital dynamics model and comparing it with the ephemeris data of the reference orbital elements. If the position error at the same moment exceeds the set threshold, it is considered that there is a parameter abnormality, and an abnormal mark is then placed on the current orbital elements. Through the above two-stage abnormal identification and correction mechanism, it is ensured that the output second orbital elements have high data accuracy and orbital dynamics consistency.

[0045] Please refer to Figure 6 , Figure 6 which is a schematic diagram of an abnormal correction process for an orbital circle number provided by an embodiment of the present invention.

[0046] In this embodiment, when correcting the orbit circle number anomaly, the system first determines whether the orbit circle number of the record has been completed; for the record that has not been completed, the reference orbit element closest to its epoch time is retrieved, and the time difference between the two and the orbit period of the circle are calculated accordingly; the reference element is used to carry out orbit prediction to obtain the ephemeris data of the corresponding time period, and the circle change point closest to the target epoch time is extracted therefrom; after calculating the circle difference by combining the epoch time difference, orbit period and circle change point information, the difference is superimposed with the circle number of the reference element to generate a theoretical circle number; the system then compares the theoretical value with the circle number in the current record, and if any inconsistency is found, the theoretical value is used to correct it, and an abnormal correction mark is added to the record; if the original circle number is consistent with the theoretical value, the status quo is retained without marking.

[0047] Please refer to Figure 7 , Figure 7 A schematic diagram of a Kepler six-parameter anomaly correction process provided by an embodiment of the present invention.

[0048] In this embodiment, when the orbital element to be evaluated has not been marked as "post-control", the system will first retrieve the reference orbital element closest to its epoch time, and extrapolate the ephemeris data of the same future time period based on this element and the current element respectively; then, by comparing the spatial position difference between the two sets of ephemeris at the same time, the position error value is calculated and compared with the preset threshold - if the error exceeds the threshold, an abnormal mark is added to the orbital element for subsequent review and processing.

[0049] In this embodiment, the redundancy of the orbital element data includes complete consistency and incomplete consistency of the orbital elements. For complete consistency of the orbital elements, by comparing the values of each parameter one by one, it can be determined whether the orbital elements are redundant. For incomplete consistency of the orbital elements, by calculating the similarity of each parameter and determining whether the similarity is greater than a threshold, it is determined whether the orbital elements are redundant.

[0050] Please refer to Figure 8 , Figure 8 A schematic diagram of a redundant deduplication process provided by an embodiment of the present invention.

[0051] In this embodiment, in the case where there are multiple orbital element records at the same epoch time, all orbital elements at that moment are first retrieved and sorted from early to late according to the storage time; then, the sorted adjacent records are compared two by two in turn - if two records are completely consistent in all parameters, or their difference is found to be greater than a preset threshold after similarity calculation, the earlier stored record is marked as redundant; after traversing all records, the system eliminates the orbital elements marked as redundant and only retains the target record with the best quality and the most representative, to ensure that only unique and non-redundant orbital element data exists in each epoch time in the database.

[0052] In this embodiment, the redundant duplicate removal processing of the second number of orbital elements based on a preset redundant duplicate removal rule to obtain the target number of orbital elements includes: Screen out all orbital elements with the same epoch time as the second number of orbital elements from the orbital element data as the redundant orbital element set; Calculate the similarity between any two orbital elements in the redundant orbital element set. When the similarity is greater than the preset threshold, mark the orbital element with an earlier storage time as redundant; Remove the orbital elements marked as redundant from the redundant orbital element set to obtain the target number of orbital elements.

[0053] An embodiment of the present invention also provides a spacecraft orbital element data governance device, including: a sorting module, a governance module, and a storage module; The sorting module is used to obtain an orbital element set, sort the orbital elements in the orbital element set in ascending order of epoch time and storage time, and obtain an orbital element set to be governed; The governance module is used to sequentially traverse the orbital elements to be governed in the orbital element set to be governed, perform missing value filling processing on the orbital elements to be governed based on a preset missing value filling rule to obtain the first number of orbital elements; perform abnormal correction processing on the first number of orbital elements based on a preset abnormal correction rule to obtain the second number of orbital elements; perform redundant duplicate removal processing on the second number of orbital elements based on a preset redundant duplicate removal rule to obtain the target number of orbital elements, and enter the next traversal; The storage module is used to complete the governance of the orbital element data until all the orbital elements to be governed in the orbital element set to be governed are traversed and the target orbital element data is obtained.

[0054] In this embodiment, the governance module is used to perform missing value filling processing on the orbital elements to be governed based on a preset missing value filling rule to obtain the first number of orbital elements, including: Determine the reference orbital elements based on the epoch time of the orbital elements to be governed, and perform filling processing on the orbital circle number of the orbital elements to be governed based on the reference orbital elements to obtain the first filled orbital elements; Match the orbit control record based on the epoch time of the orbital elements to be governed, and perform orbit control filling processing on the first filled orbital elements based on the orbit control record to obtain the second filled orbital elements; Match the target orbital elements based on the epoch time of the orbital elements to be governed, and perform Kepler filling processing on the second filled orbital elements based on the six Kepler parameters of the target orbital elements to obtain the first number of orbital elements.

[0055] In this embodiment, the governance module is configured to determine the reference orbital elements based on the epoch time of the orbital elements to be governed, and perform a complement processing on the orbital circle number of the orbital elements to be governed based on the reference orbital elements, so as to obtain the first complemented orbital elements, including: Calculate the epoch time difference between the target orbital elements and the orbital elements to be governed, and use the target orbital elements with the smallest absolute value of the epoch time difference as the reference orbital elements; Obtain the orbital period and the information of the orbit-changing point of the reference orbital elements, and calculate the number of circle differences based on the orbital period, the information of the orbit-changing point and the epoch time difference; Use the orbital circle number of the reference orbital elements plus the number of circle differences as the orbital circle number of the orbital elements to be governed, complete the complement processing, and obtain the first complemented orbital elements.

[0056] In this embodiment, the governance module is configured to perform an anomaly correction process on the first orbital elements based on a preset anomaly correction rule to obtain the second orbital elements, including: Determine whether there are anomalies in the orbital circle number and the six Keplerian parameters of the first orbital elements according to the reference orbital elements; the reference orbital elements are the governed target orbital elements; When there is an anomaly in the orbital circle number of the first orbital elements, perform an anomaly correction process on the orbital circle number of the first orbital elements according to the orbital circle number and the number of circle differences of the reference orbital elements to obtain the second orbital elements; When there is an anomaly in the six Keplerian parameters of the first orbital elements, perform orbital extrapolation according to the reference orbital elements and the first orbital elements respectively, calculate the position error at the same moment based on the ephemeris data, and perform an anomaly judgment process according to the position error to obtain the second orbital elements.

[0057] In this embodiment, the governance module is configured to perform a redundancy removal process on the second orbital elements based on a preset redundancy removal rule to obtain the target orbital elements, including: Screen out all the orbital elements with the same epoch time as the second orbital elements in the orbital element data as the redundant orbital element set; Calculate the similarity between any two orbital elements in the redundant orbital element set. When the similarity is greater than a preset threshold, mark the orbital element with an earlier storage time as redundant; Remove the orbital elements marked as redundant from the redundant orbital element set to obtain the target orbital elements.

[0058] In an embodiment of the present invention, a terminal device is further provided, including a processor, a memory, and a computer program stored in the memory and configured to be executed by the processor. When the processor executes the computer program, the above-mentioned spacecraft orbital element data governance method is implemented.

[0059] In an embodiment of the present invention, a computer-readable storage medium is further provided. The computer-readable storage medium includes a stored computer program, wherein when the computer program runs, it controls the device where the computer-readable storage medium is located to execute the above-mentioned spacecraft orbital element data governance method.

[0060] Exemplarily, the computer program can be divided into one or more modules. One or more modules are stored in the memory and executed by the processor to complete the present invention. One or more modules can be a series of computer program instruction segments capable of completing specific functions, and these instruction segments are used to describe the execution process of the computer program in the terminal device.

[0061] The terminal device can be a computing device such as a desktop computer, a notebook, a palm computer, and a cloud server. The terminal device may include, but is not limited to, a processor, a memory, and a display. Those skilled in the art can understand that the above components are only examples of the terminal device and do not constitute a limitation on the terminal device. It may include more or fewer components than those described, or combine certain components, or different components. For example, the terminal device may further include input / output devices, network access devices, a bus, etc.

[0062] The so-called processor may be a central processing unit (CPU), or may also be other general-purpose processors, a digital signal processor (DSP), an application specific integrated circuit (ASIC), a field-programmable gate array (FPGA), or other programmable logic devices, discrete gate or transistor logic devices, discrete hardware components, etc. The general-purpose processor may be a microprocessor or the processor may also be any conventional processor, etc. The processor is the control center of the terminal device and connects various parts of the entire terminal device through various interfaces and lines.

[0063] The memory can be used to store computer programs and / or modules. By running or executing the computer programs and / or modules stored in the memory and invoking the data stored in the memory, the processor can implement various functions of the terminal device. The memory mainly includes a program storage area and a data storage area. Among them, the program storage area can store an operating system, application programs required for at least one function (such as a sound playback function, a text conversion function, etc.); the data storage area can store data created according to the use of the mobile phone (such as audio data, text message data, etc.). In addition, the memory can include high-speed random access memory and can also include non-volatile memory, such as a hard disk, memory, plug-in hard disk, smart media card (SMC), secure digital (SD) card, flash card, at least one magnetic disk storage device, flash memory device, or other volatile solid-state storage devices.

[0064] Among them, when the module for governing the orbital elements data of the spacecraft is implemented in the form of a software functional unit and sold or used as an independent product, it can be stored in a computer-readable storage medium. Based on such an understanding, to implement all or part of the processes in the above-mentioned embodiment methods of the present invention, it can also be completed by instructing relevant hardware through a computer program. The computer program can be stored in a computer-readable storage medium. When the computer program is executed by a processor, the steps of the above-mentioned various method embodiments can be implemented. Among them, the computer program includes computer program code, and the computer program code can be in the form of source code, object code, executable file, or some intermediate form, etc. The computer-readable medium can include: any entity or device capable of carrying the computer program code, recording medium, USB flash drive, mobile hard disk, magnetic disk, optical disc, computer memory, read-only memory (ROM), random access memory (RAM), electrical carrier signal, telecommunication signal, and software distribution medium, etc. Those of ordinary skill in the art can understand and implement it without creative work.

[0065] The specific embodiments described above have further elaborated on the purpose, technical solutions, and beneficial effects of the present invention. It should be understood that the above descriptions are only specific embodiments of the present invention and are not used to limit the protection scope of the present invention. It is particularly pointed out that for those skilled in the art, any modifications, equivalent replacements, improvements, etc. made within the spirit and principles of the present invention shall be included in the protection scope of the present invention.

Claims

1. A method for governing spacecraft orbital element data, characterized in that, Including: Obtain a set of orbital elements, sort the orbital elements in the set of orbital elements based on the ascending order of the epoch time and the warehousing time, and obtain a set of orbital elements to be processed; Traverse the orbital elements to be processed in the set of orbital elements to be processed in sequence, perform missing data filling processing on the orbital elements to be processed based on a preset missing data filling rule, and obtain the first set of orbital elements; Perform abnormal correction processing on the first set of orbital elements based on a preset abnormal correction rule, and obtain the second set of orbital elements; Perform redundant data removal processing on the second set of orbital elements based on a preset redundant data removal rule, obtain the target orbital elements, and enter the next traversal; Until all the orbital elements to be processed in the set of orbital elements to be processed are traversed, obtain the target orbital element data, and complete the governance of the orbital element data.

2. The method for managing spacecraft orbital element data according to claim 1, wherein, The performing missing data filling processing on the orbital elements to be processed based on a preset missing data filling rule and obtaining the first set of orbital elements includes: Determine the reference orbital elements based on the epoch time of the orbital elements to be processed, perform filling processing on the orbit number of the orbital elements to be processed based on the reference orbital elements, and obtain the first filled orbital elements; Match the orbit control record based on the epoch time of the orbital elements to be processed, perform orbit control filling processing on the first filled orbital elements based on the orbit control record, and obtain the second filled orbital elements; Match the target orbital elements based on the epoch time of the orbital elements to be processed, perform Kepler filling processing on the second filled orbital elements based on the six Kepler parameters of the target orbital elements, and obtain the first set of orbital elements.

3. The method for governing spacecraft orbital element data according to claim 2, wherein The determining the reference orbital elements based on the epoch time of the orbital elements to be processed, performing filling processing on the orbit number of the orbital elements to be processed based on the reference orbital elements, and obtaining the first filled orbital elements includes: Calculate the epoch time difference between the target orbital elements and the orbital elements to be processed, and use the target orbital elements with the smallest absolute value of the epoch time difference as the reference orbital elements; Obtain the orbital period and the orbit change point information of the reference orbital elements, and calculate the orbit cycle difference based on the orbital period, the orbit change point information, and the epoch time difference; Use the orbit number of the reference orbital elements plus the orbit cycle difference as the orbit number of the orbital elements to be processed, complete the filling processing, and obtain the first filled orbital elements.

4. The method for governing spacecraft orbital element data according to claim 3, characterized in that, The performing abnormal correction processing on the first set of orbital elements based on a preset abnormal correction rule and obtaining the second set of orbital elements includes: Determine whether there are orbit number anomalies and Kepler six-parameter anomalies in the first set of orbital elements according to the reference orbital elements; the reference orbital elements are the processed target orbital elements; When there is an orbit number anomaly in the first set of orbital elements, perform abnormal correction processing on the orbit number of the first set of orbital elements according to the orbit number and the orbit cycle difference of the reference orbital elements, and obtain the second set of orbital elements; When there is a Kepler six-parameter anomaly in the first set of orbital elements, perform orbit extrapolation on the reference orbital elements and the first set of orbital elements respectively, calculate the position error at the same moment based on the ephemeris data, and perform abnormal judgment processing according to the position error to obtain the second set of orbital elements.

5. The method for governing spacecraft orbital element data according to claim 4, wherein Performing redundancy removal processing on the second orbital elements based on a preset redundancy removal rule to obtain target orbital elements includes: Screening out all orbital elements with the same epoch time as the second orbital elements in the orbital element data as a redundant orbital element set; Calculating the similarity between any two orbital elements in the redundant orbital element set. When the similarity is greater than a preset threshold, the orbital element with an earlier storage time is marked as redundant; Removing the orbital elements marked as redundant from the redundant orbital element set to obtain target orbital elements.

6. A device for managing spacecraft orbital element data, characterized in that, Including: A sorting module, a governance module, and a storage module; The sorting module is used to obtain an orbital element set, sort the orbital elements in the orbital element set in ascending order of epoch time and storage time to obtain an orbital element set to be governed; The governance module is used to sequentially traverse the orbital elements to be governed in the orbital element set to be governed, perform missing data filling processing on the orbital elements to be governed based on a preset missing data filling rule to obtain first orbital elements; Performing abnormal correction processing on the first orbital elements based on a preset abnormal correction rule to obtain second orbital elements; Performing redundancy removal processing on the second orbital elements based on a preset redundancy removal rule to obtain target orbital elements, and entering the next traversal; The storage module is used to obtain target orbital element data until all the orbital elements to be governed in the orbital element set to be governed are traversed, completing the governance of the orbital element data.

7. The spacecraft orbital element data governance device according to claim 6, characterized in that, The governance module is used to perform missing data filling processing on the orbital elements to be governed based on a preset missing data filling rule to obtain first orbital elements, including: Determining a reference orbital element based on the epoch time of the orbital element to be governed, and performing circle number filling processing on the circle number of the orbital element to be governed based on the reference orbital element to obtain a first filled orbital element; Matching orbit control records based on the epoch time of the orbital element to be governed, and performing orbit control filling processing on the first filled orbital element based on the orbit control records to obtain a second filled orbital element; Matching target orbital elements based on the epoch time of the orbital element to be governed, and performing Kepler filling processing on the second filled orbital element based on the six Kepler parameters of the target orbital elements to obtain first orbital elements.

8. The spacecraft orbital element data governance device according to claim 7, characterized in that, The governance module is used to determine a reference orbital element based on the epoch time of the orbital element to be governed, and perform circle number filling processing on the circle number of the orbital element to be governed based on the reference orbital element to obtain a first filled orbital element, including: Calculating the epoch time difference between the target orbital elements and the orbital element to be governed, and using the target orbital element with the smallest absolute value of the epoch time difference as the reference orbital element; Obtaining the orbital period and circle change point information of the reference orbital element, and calculating the circle number difference based on the orbital period, circle change point information, and epoch time difference; Using the circle number of the reference orbital element plus the circle number difference as the circle number of the orbital element to be governed to complete the filling processing and obtain a first filled orbital element.

9. The spacecraft orbital element data governance device according to claim 8, characterized in that, The governance module is used to perform anomaly correction processing on the first orbital elements based on a preset anomaly correction rule to obtain second orbital elements, including: Determining whether there are anomalies in the orbital circle number and Kepler six parameters of the first orbital elements according to the reference orbital elements; the reference orbital elements are the target orbital elements after governance; When there is an anomaly in the orbital circle number of the first orbital elements, anomaly correction processing is performed on the orbital circle number of the first orbital elements according to the orbital circle number and the orbital circle order difference of the reference orbital elements to obtain second orbital elements; When there are anomalies in the Kepler six parameters of the first orbital elements, orbital extrapolation is performed respectively according to the reference orbital elements and the first orbital elements, the position error at the same moment is calculated based on the ephemeris data, and anomaly judgment processing is performed according to the position error to obtain second orbital elements.

10. A spacecraft orbital element data governance device as described in claim 9, characterized in that, The governance module is used to perform redundancy removal processing on the second orbital elements based on a preset redundancy removal rule to obtain target orbital elements, including: Screening out all orbital elements with the same epoch time as the second orbital elements in the orbital element data as the redundant orbital element set; Calculating the similarity between any two orbital elements in the redundant orbital element set, and when the similarity is greater than a preset threshold, marking the orbital element with an earlier storage time as redundant; Removing the orbital elements marked as redundant in the redundant orbital element set to obtain target orbital elements.

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