A GEO satellite intelligent orbit change detection method based on TLE ephemeris

Abstract

The present invention discloses a method for detecting an intelligent orbit change of a GEO satellite based on TLE ephemeris, comprising the following steps: Step 1, determining a start threshold for satellite orbit change; Step 2, determining an end threshold for satellite orbit change; Step 3, intelligent detection of a GEO satellite orbit change. The beneficial effects of the present invention are as follows: In the past, orbit change detection for GEO satellites mostly adopted a manual interpretation method, and the process from satellite data collation and inversion to orbit change analysis was complicated and the accuracy was difficult to guarantee. The method disclosed in the present invention can automatically perform satellite ephemeris processing and orbit change detection analysis based on a large amount of TLE format ephemeris data, quickly obtain the orbit change situation of the GEO satellite, and greatly improve work efficiency and analysis accuracy.

Description

Technical Field

[0001] The present invention belongs to the field of aerospace technology, and in particular relates to an intelligent orbit change detection and analysis method for GEO (Geosynchronous Earth Orbit) satellites in this field. Background Art

[0002] Satellite orbital slots, particularly those for GEO satellites, are limited. With the advancement and development of aerospace technology, more and more GEO satellites are being launched into space, orbiting along pre-set orbits to provide services to users on Earth and in space. Given the limited availability of orbital slots, the more GEO satellites are launched, the tighter the orbital slots become, and the more difficult it becomes to declare, coordinate, and utilize them. Therefore, it is crucial to understand the dynamic changes in the number of GEO satellites in orbit.

[0003] The TLE (Two Line Elements) format ephemeris regularly published by the North American Aerospace Defense Command records information such as each satellite's number, satellite name, launch date, satellite orbit inclination, and period. The SGP4 / SDP4 orbit model can be used to invert the satellite's orbital information at the corresponding observation time. The obtained satellite orbital information can be used to specifically describe the changes in the satellite's orbit.

[0004] "Software Implementation of Calculating Satellite Orbit Evolution Trajectory Using Public Ephemeris" discloses a software for calculating satellite orbit evolution trajectory using public ephemeris. From the perspective of software implementation, the article gives examples of using public TLE format ephemeris for parameter analysis, and then analyzes the normal working satellite orbit control process, satellite decommissioning process, satellite network data activation process of in-orbit satellites, and the implementation of the co-orbit protocol. However, the article only gives the software simulation results and does not provide a description of the relevant methods. Summary of the Invention

[0005] The technical problem to be solved by the present invention is to provide an intelligent orbit change detection method for GEO satellites based on the public TLE format ephemeris.

[0006] The present invention adopts the following technical solutions:

[0007] A method for detecting an intelligent orbit change of a GEO satellite based on TLE ephemeris is improved in that it comprises the following steps:

[0008] Step 1: Determine the satellite orbit change start threshold:

[0009] Step 11: Get all TLE format ephemeris of 50 GEO satellites, keep only the latest ephemeris of each satellite every day, and arrange them in ascending order of date;

[0010] Step 12: Process the ephemeris data using the SGP4 or SDP4 orbit model to obtain the satellite longitude, latitude, and altitude position information of the satellite at 0:00 every day;

[0011] Step 13, obtain the longitude value of each satellite at 0 o'clock on the 100th day when the orbit is not changed;

[0012] Step 14, remove outliers;

[0013] Step 15, calculate the orbital position difference delta of each satellite for two consecutive days;

[0014] Step 16: Take the largest delta among the orbital position differences of each satellite and calculate the average of the corresponding deltas of 50 satellites as the orbit change start threshold, denoted as δ. When the orbital position difference of a satellite is greater than δ for two consecutive days, the satellite begins to change its orbit.

[0015] Step 2: Determine the satellite orbit change end threshold:

[0016] After a GEO satellite undergoes an orbit change, it needs to go through a stabilization period. During this period, the orbit position fluctuation may exceed a threshold value δ. The stabilization period is set to 20 consecutive days when the orbit position difference is less than δ, which is used as the threshold condition for determining the end of the orbit change.

[0017] Step 3: GEO satellite orbit change intelligent detection:

[0018] Step 31, construct an ephemeris array, denoted as TLE_ARR, obtain all ephemeris of the specified satellite, keep only the latest ephemeris of the satellite every day, sort them by date from smallest to largest, and obtain the ephemeris array;

[0019] Step 32, construct a date array, denoted as DATE_ARR, with an interval of 1 day, starting from 30 days after the satellite launch date and continuing to the maximum date with an ephemeris;

[0020] Step 33: Construct an orbit array, denoted as ORBIT_LONG_ARR, traverse the date array DATE_ARR, find the ephemeris data for the corresponding date in the ephemeris array, and use the SGP4 or SDP4 algorithm model to calculate the satellite orbit position parameters at 0:00 every day. After removing outliers, an array is obtained, in which each element is the orbit position of the satellite at 0:00 every day. If there is no ephemeris for a day, it is skipped and the element is retained in the array.

[0021] Step 34: Construct an orbit difference array, denoted as LONG_delta_ARR. Since the first element of the ORBIT_LONG_ARR array is the starting position, the change for this day is recorded as null, so the first element of the LONG_delta_ARR array is recorded as null. Starting from the second element of the ORBIT_LONG_ARR array, compare it with the previous element, calculate the orbit difference between the two, and record them in the orbit difference array LONG_delta_ARR in sequence. And so on. Since there is no ephemeris data for some dates, the orbit position of the corresponding date index position in the ORBIT_LONG_ARR array is null and cannot be compared with other elements, so the corresponding position in the LONG_delta_ARR array is also recorded as null.

[0022] Step 35: Obtain four arrays through the above four steps, with the same length and the same index, and the information recorded is the same date;

[0023] Step 36: Construct the result array, denoted as RESULT_ARR[], and traverse the LONG_delta_ARR array, taking 20 elements at a time as a group. Set the empty elements in the LONG_delta_ARR array to 0. If an element greater than δ appears in the group, it is considered that the satellite has started orbit change. Extract the index of the element in the LONG_delta_ARR array and retrieve the corresponding date from the DATA_ARR array. Record this as the date when the satellite starts orbit change.

[0024] If the values ​​of all elements from the i-th element to the last element in the group are less than δ, then take the i-th element as the first element and take 20 consecutive elements as a group to traverse again, and so on, until there is no value greater than δ in 20 consecutive elements, then it is considered that the satellite has completed the orbit change, and use the index of the first element of the group in the array LONG_delta_ARR to get the corresponding date in the array DATA_ARR, which is recorded as the date when the satellite completes the orbit change;

[0025] Repeat the above steps until the traversal is completed, and an array of orbit change records is obtained. Each element records the satellite's start date and orbital position, end date and orbital position, historical stop orbital position, and number of orbit changes.

[0026] It is not ruled out that the last element in the orbit change result array does not have the orbit change completion date and orbit position, indicating that the satellite is still in orbit change according to the logic of this orbit change analysis method.

[0027] The beneficial effects of the present invention are:

[0028] In the past, orbit change detection for GEO satellites mostly relied on manual interpretation, a complex process involving satellite data collation and inversion, and orbit change analysis, with limited accuracy. The method disclosed in this paper, however, can automatically process satellite ephemeris and perform orbit change detection and analysis based on a large amount of TLE-formatted ephemeris data, rapidly determining the orbit change status of GEO satellites and significantly improving work efficiency and analysis accuracy. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] Figure 1 It is a schematic flow diagram of the method of the present invention;

[0030] Figure 2 This is a schematic diagram of the continuous ephemeris of GEO satellites;

[0031] Figure 3 is a schematic diagram of outliers;

[0032] Figure 4 This is a schematic diagram of track change. DETAILED DESCRIPTION

[0033] In order to make the purpose, technical solutions and advantages of the present invention more clearly understood, the present invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not intended to limit the present invention.

[0034] Example 1: This example discloses a GEO satellite intelligent orbit change detection method based on TLE ephemeris, which can realize automatic identification and extraction of orbit change information based on traditional orbit evolution analysis. Figure 1 As shown, the following steps are included:

[0035] Step 1: Determine the satellite orbit change start threshold:

[0036] Step 11: Get all TLE format ephemeris of 50 GEO satellites, keep only the latest ephemeris of each satellite every day, and arrange them in ascending order of date; Figure 2 This is a schematic diagram of the continuous ephemeris of GEO satellites.

[0037] Step 12: Process the ephemeris data using the SGP4 or SDP4 orbit model to obtain the satellite longitude, latitude, and altitude position information of the satellite at 0:00 every day, with the longitude being the main consideration here.

[0038] Step 13, obtain the longitude value of each satellite at 0 o'clock on the 100th day when the orbit is not changed;

[0039] Step 14, remove outliers; Figure 3 It is a schematic diagram of outliers (dashed box).

[0040] Step 15, calculate the orbital position difference delta of each satellite for two consecutive days;

[0041] Step 16: Take the largest delta among the orbital position differences of each satellite and calculate the average of the corresponding deltas of 50 satellites as the orbit change start threshold, denoted as δ. When the orbital position difference of a satellite is greater than δ for two consecutive days, the satellite begins to change its orbit. In this embodiment, δ = 0.5°.

[0042] Step 2: Determine the satellite orbit change end threshold:

[0043] After a GEO satellite undergoes an orbit change, it must go through a stabilization period. During this period, the orbit fluctuation may exceed a threshold of 0.5°. A stabilization period of 20 consecutive days with an orbital position difference of less than 0.5° is used as the threshold condition for determining the end of the orbit change. Figure 4 It is a schematic diagram of track change (dashed box).

[0044] Step 3: GEO satellite orbit change intelligent detection:

[0045] Step 31, construct an ephemeris array, denoted as TLE_ARR, obtain all ephemeris of the EXPRESS-AM3 (satellite number: 28707) satellite, retain only the latest ephemeris of the satellite every day, sort them by date from smallest to largest, and obtain an ordered array of ephemeris data;

[0046] Step 32, construct a date array, denoted as DATE_ARR, with an interval of 1 day, starting from 30 days after the satellite launch date and continuing to the maximum date with an ephemeris;

[0047] Step 33: Construct an orbit array, denoted as ORBIT_LONG_ARR. Traverse the date array DATE_ARR and find the ephemeris data for the corresponding date in the ephemeris array. Use the SGP4 or SDP4 algorithm model to calculate the satellite orbit position (i.e., longitude) at 0:00 each day. After removing outliers, an array is obtained, in which each element is the orbit position of the satellite at 0:00 each day. If there is no ephemeris for a day, it is skipped and the element is retained in the array.

[0048] Step 34: Construct an orbit difference array, denoted as LONG_delta_ARR. Since the first element of the ORBIT_LONG_ARR array is the starting position, the change for this day is recorded as null, so the first element of the LONG_delta_ARR array is recorded as null. Starting from the second element of the ORBIT_LONG_ARR array, compare it with the previous element, calculate the orbit difference between the two, and record them in the orbit difference array LONG_delta_ARR in sequence. And so on. Since there is no ephemeris data for some dates, the orbit position of the corresponding date index position in the ORBIT_LONG_ARR array is null and cannot be compared with other elements, so the corresponding position in the LONG_delta_ARR array is also recorded as null.

[0049] Step 35: Obtain four arrays through the above four steps, with the same length and the same index, and the information recorded is the same date;

[0050] Step 36: Construct the result array, denoted as RESULT_ARR[], and traverse the LONG_delta_ARR array, taking 20 elements (that is, 20 days) as a group at a time. Set the empty elements in the LONG_delta_ARR array as 0. If an element with an angle greater than 0.5° appears in the group, it is considered that the satellite has started orbit change. Extract the index of the element in the LONG_delta_ARR array and get the corresponding date from the DATA_ARR array, which is recorded as the date when the satellite starts orbit change.

[0051] If the values ​​of all elements in the group from the i-th element to the last element are less than 0.5°, then the i-th element is taken as the first element, and 20 consecutive elements are taken as a group to traverse again. This process is repeated until no value of the 20 consecutive elements is greater than 0.5°. In this case, the satellite is considered to have completed the orbit change. The index of the first element of the group in the array LONG_delta_ARR is used to retrieve the corresponding date from the array DATA_ARR and record it as the date when the satellite orbit change ends.

[0052] Repeat the above steps until the traversal is completed, and an array of orbit change records is obtained. Each element records the satellite's start date and orbital position, end date and orbital position, historical stop orbital position, and number of orbit changes.

[0053] Orbit change results: EXPRESS-AM3 (satellite number: 28707) has changed its orbit three times. Details are as follows:

[0054]

Claims

1. A GEO satellite intelligent orbit change detection method based on TLE ephemeris, characterized in that: The steps include: Step 1: Determine the satellite orbit change start threshold: Step 11: Get all TLE format ephemeris of 50 GEO satellites, keep only the latest ephemeris of each satellite every day, and arrange them in ascending order of date; Step 12: Process the ephemeris data using the SGP4 or SDP4 orbit model to obtain the satellite longitude, latitude, and altitude position information of the satellite at 0:00 every day; Step 13, obtain the longitude value of each satellite at 0 o'clock on the 100th day when the orbit is not changed; Step 14, remove outliers; Step 15, calculate the orbital position difference delta of each satellite for two consecutive days; Step 16: Take the largest delta among the orbital position differences of each satellite and calculate the average of the corresponding deltas of 50 satellites as the orbit change start threshold, denoted as δ. When the orbital position difference of a satellite is greater than δ for two consecutive days, the satellite begins to change its orbit. Step 2: Determine the satellite orbit change end threshold: After a GEO satellite undergoes an orbit change, it needs to go through a stabilization period. During this period, the orbit position fluctuation may exceed a threshold value δ. The stabilization period is set to 20 consecutive days when the orbit position difference is less than δ, which is used as the threshold condition for determining the end of the orbit change. Step 3: GEO satellite orbit change intelligent detection: Step 31, construct an ephemeris array, denoted as TLE_ARR, obtain all ephemeris of the specified satellite, keep only the latest ephemeris of the satellite every day, sort them by date from smallest to largest, and obtain the ephemeris array; Step 32, construct a date array, denoted as DATE_ARR, with an interval of 1 day, starting from 30 days after the satellite launch date and continuing to the maximum date with an ephemeris; Step 33: Construct an orbit array, denoted as ORBIT_LONG_ARR, traverse the date array DATE_ARR, find the ephemeris data for the corresponding date in the ephemeris array, and use the SGP4 or SDP4 algorithm model to calculate the satellite orbit position parameters at 0:00 every day. After removing outliers, an array is obtained, in which each element is the orbit position of the satellite at 0:00 every day. If there is no ephemeris for a day, it is skipped and the element is retained in the array. Step 34: Construct an orbit difference array, denoted as LONG_delta_ARR. Since the first element of the ORBIT_LONG_ARR array is the starting position, the change for this day is recorded as null, so the first element of the LONG_delta_ARR array is recorded as null. Starting from the second element of the ORBIT_LONG_ARR array, compare it with the previous element, calculate the orbit difference between the two, and record them in the orbit difference array LONG_delta_ARR in sequence. And so on. Since there is no ephemeris data for some dates, the orbit position of the corresponding date index position in the ORBIT_LONG_ARR array is null and cannot be compared with other elements, so the corresponding position in the LONG_delta_ARR array is also recorded as null. Step 35: Obtain four arrays through the above four steps, with the same length and the same index, and the information recorded is the same date; Step 36: Construct the result array, denoted as RESULT_ARR[], and traverse the LONG_delta_ARR array, taking 20 elements at a time as a group. Set the empty elements in the LONG_delta_ARR array to 0. If an element greater than δ appears in the group, it is considered that the satellite has started orbit change. Extract the index of the element in the LONG_delta_ARR array and retrieve the corresponding date from the DATA_ARR array. Record this as the date when the satellite starts orbit change. If the values ​​of all elements from the i-th element to the last element in the group are less than δ, then take the i-th element as the first element and take 20 consecutive elements as a group to traverse again, and so on, until there is no value greater than δ in 20 consecutive elements, then it is considered that the satellite has completed the orbit change, and use the index of the first element of the group in the array LONG_delta_ARR to get the corresponding date in the array DATA_ARR, which is recorded as the date when the satellite completes the orbit change; Repeat the above steps until the traversal is completed, and an array of orbit change records is obtained. Each element records the satellite's start date and orbital position, end date and orbital position, historical stop orbital position, and number of orbit changes. It is not ruled out that the last element in the orbit change result array does not have the orbit change completion date and orbit position, indicating that the satellite is still in orbit change according to the logic of this orbit change analysis method.

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