Satellite orbit departure control method, device, equipment and medium based on small thrust

Abstract

The present invention relates to the field of aerospace measurement and control technologies, and provides a satellite deorbiting control method, device, equipment and medium based on a small thrust. It can determine different ignition strategies according to the control efficiency index values of each candidate ignition strategy, and first perform a perigee altitude lift simulation based on the ignition strategy, then perform a circular orbit control simulation, and finally execute an overall orbit altitude lift simulation to continuously lift the satellite to the graveyard orbit. Further, it evaluates the collision risk of the deorbiting process of the target satellite according to the simulation data, and when the collision risk assessment result shows that there is no collision risk between the target satellite and other satellites during the deorbiting process, it generates a deorbiting control strategy for the target satellite based on the different ignition strategies used to assist in performing a safe and efficient deorbiting control task for the satellite.

Description

Technical Field

[0001] The present invention relates to the technical field of aerospace measurement and control, and particularly to a satellite deorbiting control method, device, equipment and medium based on small thrust. Background Art

[0002] The geosynchronous orbit has the advantages of the same speed as the earth's rotation, a wide field of view coverage of the sub-satellite point, and being visible for 24 hours, etc., and has an important strategic position. However, due to the limited high-orbit space resources, geosynchronous orbit satellites need to be deorbited to the graveyard orbit (i.e., 300 kilometers above the geosynchronous orbit) at the end of their service life to avoid affecting the normal operation of other on-orbit satellites.

[0003] In the prior art, chemical propulsion control is mainly adopted to implement the deorbiting control of geosynchronous orbit satellites. The ignition time of such control is relatively large, and the controllable amount that can be applied is relatively large. Generally, 3 - 4 batches of implementation can achieve the deorbiting effect. However, at the end of the satellite's service life, there are often permanent faults such as abnormal thrusters, abnormal attitude control systems, and gas entrapment in the tanks of oxidants and fuels, etc., which make the satellite unable to implement deorbiting control according to the rated thrust, and it is difficult to quickly raise the orbital altitude to the graveyard orbit. Therefore, only long-term orbital control can be implemented through small thrust to slowly raise the orbital altitude. Moreover, according to the orbital characteristics of geosynchronous orbit satellites, there is still a relatively large collision risk with other normally operating satellites during the raising process.

[0004] In view of the above problems, there is an urgent need for a satellite deorbiting control method based on small thrust to reduce the collision risk with other normal on-orbit satellites during the satellite deorbiting control process. Summary of the Invention

[0005] In view of the above, it is necessary to provide a satellite deorbiting control method, device, equipment and medium based on small thrust, aiming to solve the safety risk problems existing in the satellite deorbiting control process.

[0006] A satellite deorbiting control method based on small thrust, the satellite deorbiting control method based on small thrust includes:

[0007] In response to a deorbiting control instruction for a target satellite, determining at least one candidate ignition strategy; wherein, the target satellite is a geosynchronous orbit satellite at the end of its service life;

[0008] Calculating the control efficiency index value of each candidate ignition strategy according to the small thrust orbital dynamics model;

[0009] Selecting a first ignition strategy from the at least one candidate ignition strategy according to the control efficiency index value of each candidate ignition strategy, and performing a perigee altitude elevation simulation on the target satellite based on the first ignition strategy to obtain first simulation data;

[0010] Select a second ignition strategy from the at least one candidate ignition strategy according to the control efficiency index value of each candidate ignition strategy, and perform a circular orbit control simulation on the target satellite based on the second ignition strategy to obtain second simulation data;

[0011] Obtain the configured ignition batch and the reserved duration for orbit determination, select a third ignition strategy from the at least one candidate ignition strategy according to the configured ignition batch and the reserved duration for orbit determination, and perform an overall lift simulation of the orbit height of the target satellite according to the third ignition strategy to obtain third simulation data;

[0012] Perform a collision risk assessment on the deorbiting process of the target satellite based on the first simulation data, the second simulation data, and the third simulation data to obtain a collision risk assessment result;

[0013] When the collision risk assessment result shows that there is no collision risk during the deorbiting process of the target satellite, generate a deorbiting control strategy for the target satellite based on the first ignition strategy, the second ignition strategy, and the third ignition strategy.

[0014] A satellite deorbiting control device based on small thrust, the satellite deorbiting control device based on small thrust includes:

[0015] A determination unit, configured to determine at least one candidate ignition strategy in response to a deorbiting control instruction for a target satellite; wherein, the target satellite is a geostationary orbit satellite at the end of its life;

[0016] A calculation unit, configured to calculate the control efficiency index value of each candidate ignition strategy according to the small thrust orbit dynamics model;

[0017] A simulation unit, configured to select a first ignition strategy from the at least one candidate ignition strategy according to the control efficiency index value of each candidate ignition strategy, and perform a perigee altitude lift simulation on the target satellite based on the first ignition strategy to obtain first simulation data;

[0018] The simulation unit is further configured to select a second ignition strategy from the at least one candidate ignition strategy according to the control efficiency index value of each candidate ignition strategy, and perform a circular orbit control simulation on the target satellite based on the second ignition strategy to obtain second simulation data;

[0019] The simulation unit is further configured to obtain the configured ignition batch and the reserved duration for orbit determination, select a third ignition strategy from the at least one candidate ignition strategy according to the configured ignition batch and the reserved duration for orbit determination, and perform an overall lift simulation of the orbit height of the target satellite according to the third ignition strategy to obtain third simulation data;

[0020] An evaluation unit for performing a collision risk assessment on the orbital decay process of the target satellite based on the first simulation data, the second simulation data, and the third simulation data to obtain a collision risk assessment result;

[0021] A generation unit for generating an orbital decay control strategy for the target satellite based on the first ignition strategy, the second ignition strategy, and the third ignition strategy when the collision risk assessment result shows that there is no collision risk during the orbital decay process of the target satellite.

[0022] A computer device, the computer device comprising:

[0023] A memory storing at least one instruction; and

[0024] A processor for executing the instructions stored in the memory to implement the satellite orbital decay control method based on small thrust.

[0025] A computer-readable storage medium storing at least one instruction, the at least one instruction being executed by a processor in a computer device to implement the satellite orbital decay control method based on small thrust.

[0026] As can be seen from the above technical solutions, the present invention can determine different ignition strategies according to the control efficiency index values of each candidate ignition strategy, and first perform a perigee altitude elevation simulation based on the ignition strategy, then perform a circular orbit control simulation, and finally execute an overall orbital altitude elevation simulation to continuously lift the satellite to the graveyard orbit. Further, a collision risk assessment is performed on the orbital decay process of the target satellite according to the simulation data, and when the collision risk assessment result shows that there is no collision risk with other satellites during the orbital decay process of the target satellite, an orbital decay control strategy for the target satellite is generated based on the different ignition strategies used to assist in performing a safe and efficient orbital decay control task for the satellite. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] Figure 1 is a flowchart of a preferred embodiment of the satellite orbital decay control method based on small thrust of the present invention;

[0028] Figure 2 is a schematic diagram of the first relationship curve and the second relationship curve of the present invention;

[0029] Figure 3 is a functional module diagram of a preferred embodiment of the satellite orbital decay control device based on small thrust of the present invention;

[0030] Figure 4 is a schematic structural diagram of a computer device of a preferred embodiment for implementing the satellite orbital decay control method based on small thrust of the present invention. Detailed implementation manners

[0031] In order to make the objectives, technical solutions and advantages of the present invention clearer, the present invention will be described in detail below with reference to the accompanying drawings and specific embodiments.

[0032] As Figure 1 shown, it is a flowchart of a preferred embodiment of the satellite deorbiting control method based on a small amount of thrust according to the present invention. According to different requirements, the order of steps in this flowchart can be changed, and some steps can be omitted.

[0033] The satellite deorbiting control method based on a small amount of thrust is applied to one or more computer devices. The computer device is a device capable of automatically performing numerical calculations and / or information processing according to pre-set or stored instructions. Its hardware includes, but is not limited to, a microprocessor, an application specific integrated circuit (ASIC), a field-programmable gate array (FPGA), a digital signal processor (DSP), an embedded device, etc.

[0034] The computer device can be any electronic product that can perform human-computer interaction with a user. For example, a personal computer, a tablet computer, a smart phone, a personal digital assistant (PDA), a game console, an Internet Protocol Television (IPTV), a smart wearable device, etc.

[0035] The computer device may further include a network device and / or a user device. Among them, the network device includes, but is not limited to, a single network server, a server group composed of multiple network servers, or a cloud composed of a large number of hosts or network servers based on cloud computing (Cloud Computing).

[0036] The server can be an independent server or a cloud server that provides basic cloud computing services such as cloud services, cloud databases, cloud computing, cloud functions, cloud storage, network services, cloud communications, middleware services, domain name services, security services, content delivery networks (CDNs), and big data and artificial intelligence platforms.

[0037] Among them, Artificial Intelligence (AI) is a theory, method, technology, and application system that uses digital computers or machines controlled by digital computers to simulate, extend, and expand human intelligence, perceive the environment, acquire knowledge, and use knowledge to obtain the best results.

[0038] Artificial intelligence basic technologies generally include technologies such as sensors, dedicated artificial intelligence chips, cloud computing, distributed storage, big data processing technologies, operation / interaction systems, and mechatronics. Artificial intelligence software technologies mainly include several major directions such as computer vision technology, robotics, biometric technology, speech processing technology, natural language processing technology, and machine learning / deep learning.

[0039] The network where the computer device is located includes but is not limited to the Internet, wide area network, metropolitan area network, local area network, Virtual Private Network (VPN), etc.

[0040] S10. In response to the off-orbit control instruction for the target satellite, determine at least one candidate ignition strategy; wherein, the target satellite is a geostationary orbit satellite at the end of its life.

[0041] In this embodiment, the off-orbit control instruction can be triggered according to actual simulation requirements.

[0042] In this embodiment, the determining at least one candidate ignition strategy includes:

[0043] Obtain the working efficiency of the small thruster of the target satellite and the orbital period of the geostationary satellite;

[0044] Calculate the minimum working duration of the small thruster according to the working efficiency of the small thruster;

[0045] Establish a time interval with the minimum working duration as the minimum value and the orbital period as the maximum value;

[0046] Randomly select the at least one candidate ignition strategy from the time interval.

[0047] For example: The working efficiency of the satellite small thruster is relatively low, generally working at least 2 hours. When the minimum working duration of the small thruster is 2 hours and the orbital period of the geostationary satellite is 24 hours, 2, 8, 12, and 16 can be randomly selected as the ignition durations within the range of [2, 24], and the corresponding candidate ignition strategies can be obtained in sequence.

[0048] S11. Calculate the control efficiency index value of each candidate ignition strategy according to the small-thrust orbit dynamics model.

[0049] In this embodiment, the small-thrust orbit dynamics model may include the following formula:

[0050]

[0051] Wherein, represents the change in semi-major axis; n represents the orbital angular velocity; e represents the orbital eccentricity; f represents the true anomaly, represents the velocity increment of the satellite under single-batch ignition; h p represents the perigee altitude; represents the semi-major axis of the satellite's orbit; represents the nominal semi-major axis of the geosynchronous orbit, generally taken as 42,165.78 km; represents the apogee altitude; h' p represents the perigee altitude after continuous ignition control; represents the initial eccentricity; represents the change in eccentricity; represents the apogee altitude after continuous ignition control.

[0052] Then the change in perigee altitude is ; the change in apogee altitude is .

[0053] Wherein, the perigee lift efficiency = change in perigee altitude / (change in perigee altitude + change in apogee altitude); the apogee lift efficiency = change in apogee altitude / (change in perigee altitude + change in apogee altitude).

[0054] Through the above small-thrust orbit dynamics model, the values of each control efficiency index under different candidate ignition strategies can be calculated.

[0055] For example: Taking a certain geosynchronous orbit satellite as an example, it is fixed near 130° east longitude, the nominal semi-major axis of the orbit is 42,165.78 km, and the eccentricity is 0.0009. This satellite is at the end of its life. Due to problems such as thruster anomalies and attitude control system anomalies, the thruster control efficiency is that continuous ignition for 1 hour can lift the orbit by about 720 m. According to the calculation of the small-thrust orbit dynamics model, the values of each control efficiency index under different candidate ignition strategies shown in the following table can be obtained:

[0056] Table 1

[0057]

[0058] S12, select the first ignition strategy from the at least one candidate ignition strategy according to the control efficiency index value of each candidate ignition strategy, and perform a perigee altitude lift simulation on the target satellite based on the first ignition strategy to obtain the first simulation data.

[0059] Generally speaking, the eccentricity of a geostationary satellite is less than 0.001, which is an empirical value in the industry. The altitude range of its orbital operation is about 42 kilometers above and below the synchronous belt (that is, 0.001 times the nominal semi-major axis of the orbit, 42,165.78 kilometers). Therefore, it is first necessary to raise the perigee altitude to more than 42 kilometers above the synchronous orbit belt.

[0060] Considering that during the continuous ignition process of a small thrust, the shorter the ignition duration, the higher the efficiency of raising the perigee altitude, but the overall increase in orbital altitude is small and cannot achieve the goal of quickly raising the perigee; the longer the ignition duration, the greater the overall increase in orbital altitude, but the efficiency of raising the perigee altitude is lower, and the goal cannot be achieved either. Therefore, according to the perigee altitude raising efficiency being at least greater than 80%, and taking into account the orbit determination evaluation in each control, the ignition duration of each batch can be determined.

[0061] Specifically, the steps of selecting the first ignition strategy from the at least one candidate ignition strategy according to the control efficiency index value of each candidate ignition strategy and performing a perigee altitude raising simulation on the target satellite based on the first ignition strategy to obtain the first simulation data include:

[0062] Calculating the initial perigee altitude of the target satellite according to the small thrust orbital dynamics model;

[0063] Obtaining the nominal semi-major axis and the configured eccentricity;

[0064] Calculating the product of the nominal semi-major axis and the configured eccentricity to obtain the target perigee raising altitude;

[0065] Calculating the sum of the target perigee raising altitude and the initial perigee altitude to obtain the control quantity;

[0066] Obtaining the perigee raising altitude corresponding to each candidate ignition strategy from the control efficiency index values of each candidate ignition strategy;

[0067] Calculating the quotient of the control quantity and the perigee raising altitude corresponding to each candidate ignition strategy to obtain the raising time of each candidate ignition strategy;

[0068] Obtaining the perigee raising efficiency corresponding to each candidate ignition strategy from the control efficiency index values of each candidate ignition strategy;

[0069] Selecting the first ignition strategy from the at least one candidate ignition strategy according to the raising time of each candidate ignition strategy and the perigee raising efficiency corresponding to each candidate ignition strategy;

[0070] Obtain the lift time corresponding to the first ignition strategy as the target lift time, obtain the ignition duration corresponding to the first ignition strategy as the first target ignition duration, and obtain the execution period;

[0071] Calculate the quotient of the target lift time and the execution period to obtain the target ignition batches;

[0072] Within each execution period, perform a batch of continuous ignition control simulations on the target satellite according to the first target ignition duration, and stop the simulation until the continuous ignition control simulations of the target ignition batches are completed;

[0073] Obtain the semi-major axis change amount, perigee altitude change amount, and apogee altitude change amount corresponding to the first ignition strategy from the control efficiency index values of each candidate ignition strategy, and obtain the semi-major axis, perigee altitude, and apogee altitude obtained after each batch of continuous ignition control simulations;

[0074] Integrate the obtained data to obtain the first simulation data.

[0075] For example: When it is calculated that the perigee altitude is 38 kilometers below the geosynchronous orbit, the nominal semi-major axis of the orbit is 42165.78 kilometers, and the eccentricity is 0.0009. Multiplying the two can obtain the control target of raising the perigee altitude to 42 kilometers above the geosynchronous orbit, then the control amount is 38 kilometers + 42 kilometers = 80 kilometers.

[0076] Furthermore, as can be seen from Table 1, the 2-hour ignition strategy has the highest efficiency for raising the perigee, but the slowest lifting speed, and it takes about 80 / 2.86, approximately 28 days to complete; similarly, the 12-hour ignition strategy takes 6 days to complete; the 16-hour ignition strategy also takes 6 days, but the total ignition time is longer and the efficiency is lower. Considering both the lifting efficiency and the lifting time, it is necessary to quickly raise the perigee altitude while also taking into account the perigee lifting efficiency. Therefore, the continuous 12-hour ignition strategy is selected as the first ignition strategy.

[0077] Even further, if the execution period is 1 day, that is, one batch of continuous ignition is performed every day, after each ignition orbit control, the simulation data shown in Table 2 below can be obtained:

[0078] Table 2

[0079]

[0080] S13. Select a second ignition strategy from the at least one candidate ignition strategy according to the control efficiency index values of each candidate ignition strategy, and perform a circular orbit control simulation on the target satellite based on the second ignition strategy to obtain second simulation data.

[0081] According to the requirements of eccentricity, circular orbit control needs to be implemented. Orbit control is only implemented near the apogee every day. The main goal is to make the perigee and apogee altitudes the same.

[0082] Specifically, selecting the second ignition strategy from the at least one candidate ignition strategy according to the control efficiency index value of each candidate ignition strategy, and performing circular orbit control simulation on the target satellite based on the second ignition strategy, the obtained second simulation data includes:

[0083] Selecting the candidate ignition strategy with the highest perigee lift efficiency from the at least one candidate ignition strategy as the second ignition strategy;

[0084] Obtaining the ignition duration corresponding to the second ignition strategy as the second target ignition duration;

[0085] Constructing a target ignition area with the apogee as the center and a first preset value as the radius;

[0086] In each execution cycle, performing a batch of continuous ignition control simulations on the target satellite in the target ignition area according to the second target ignition duration until the difference between the perigee altitude and the apogee altitude obtained after the simulation is less than or equal to the configured threshold, and then stopping the next batch of simulations;

[0087] Obtaining the semi-major axis change amount, perigee altitude change amount, and apogee altitude change amount corresponding to the second ignition strategy from the control efficiency index values of each candidate ignition strategy, and obtaining the semi-major axis, perigee altitude, and apogee altitude after each batch of continuous ignition control simulations;

[0088] Integrating the obtained data to obtain the second simulation data.

[0089] Among them, the first preset value and the configured threshold can be configured according to actual orbit control requirements.

[0090] For example: According to the requirements of eccentricity, circular orbit control needs to be implemented to continuously reduce the orbit eccentricity. As can be seen from Table 1, when adopting the ignition strategy of continuous ignition for 2 hours, the perigee lift efficiency is the highest. Therefore, in order to make the orbit circular as soon as possible, the ignition strategy of continuous ignition for 2 hours can be selected as the second ignition strategy.

[0091] Furthermore, orbit control is mainly implemented near the apogee every day. According to the orbit situation after the previous step of control, it is necessary to ignite for 4 days. The specific simulation data can be seen in Table 3 below:

[0092] Table 3

[0093]

[0094] As can be seen from Table 3, through 4 days of continuous ignition control, the perigee altitude can be made approximately equal to the apogee altitude, thereby achieving circularization control of the orbit.

[0095] S14. Obtain the configured ignition batches and the orbit determination reserved duration. Select a third ignition strategy from the at least one candidate ignition strategy according to the configured ignition batches and the orbit determination reserved duration, and perform an overall simulation of raising the orbit altitude of the target satellite according to the third ignition strategy to obtain third simulation data.

[0096] Among them, the configured ignition batches can be 2, that is, only 2 batches of orbit control can be performed per day.

[0097] Among them, the orbit determination reserved duration can be 4 hours, that is, after each ignition control, 4 hours are required for data analysis.

[0098] In this embodiment, the step of selecting a third ignition strategy from the at least one candidate ignition strategy according to the configured ignition batches and the orbit determination reserved duration, and performing an overall simulation of raising the orbit altitude of the target satellite according to the third ignition strategy to obtain third simulation data includes:

[0099] Calculate the product of the configured ignition batches and the orbit determination reserved duration to obtain a first duration;

[0100] Calculate the difference between the execution period and the first duration to obtain a second duration;

[0101] Calculate the quotient of the second duration and the configured ignition batches, and select the third ignition strategy from the at least one candidate ignition strategy according to the calculated quotient;

[0102] Obtain the ignition duration corresponding to the third ignition strategy as the third target ignition duration;

[0103] Construct two ignition regions with the perigee and the apogee as the centers and a second preset value as the radius respectively;

[0104] In each execution period, perform a batch of continuous ignition control simulations on the target satellite in the two ignition regions according to the third target ignition duration until the perigee altitude obtained after the simulation is higher than the preset distance in the synchronous orbit band, and then stop the next batch of simulations;

[0105] Obtain the data obtained by the simulation according to the third ignition strategy as the third simulation data.

[0106] For example: Since there are 2 batches of ignition, with 4 hours reserved each time, a total of 8 hours need to be reserved for the two ignitions. Since the execution cycle is 1 day, then 24 - 8 = 16, that is, there are 16 hours left for 2 batches of ignition, so each batch of ignition can take 8 hours. Therefore, continuous ignition for 8 hours can be selected as the third ignition strategy.

[0107] Furthermore, through the simulations of the previous two ignitions, namely the simulation of raising the perigee altitude of the target satellite based on the first ignition strategy and the simulation of circularizing the orbit control of the target satellite based on the second ignition strategy, the target satellite has reached a position about 57.6 kilometers above the synchronous belt (according to Table 3, 42223.38 kilometers - 42165.78 kilometers is approximately equal to 57.6 kilometers, (57.49 + 57.71) / 2 = 57.6 kilometers), and the collision risk with other service satellites is relatively small. Implement 2 batches of orbit control near the perigee and apogee respectively every day, adopting a continuous 8-hour ignition strategy, so that the perigee altitude of the orbit is higher than the synchronous orbit belt by 300 kilometers (the graveyard orbit is 300 kilometers above the geosynchronous orbit), and the deorbiting control can be completed.

[0108] Specifically, considering the current orbit altitude of the satellite, the orbital altitude difference from the required orbital altitude is 242.4 kilometers (300 kilometers - 57.6 kilometers = 242.4 kilometers). By implementing 2 batches of 8-hour orbit raising control every day, the orbital altitude can be raised by about 11.52 kilometers (5.76 kilometers * 2) every day, and the deorbiting control can be completed in 21 days (242.4 kilometers / 11.52 kilometers).

[0109] S15. Based on the first simulation data, the second simulation data, and the third simulation data, conduct a collision risk assessment on the deorbiting process of the target satellite to obtain a collision risk assessment result.

[0110] After the deorbiting control simulation is completed, a risk assessment is conducted on the entire deorbiting control process. According to the three-step deorbiting control strategy, the corresponding relationship between the sub-satellite point longitude and the perigee altitude passed by the satellite can be calculated. At the same time, the corresponding relationship between the sub-satellite point longitude of other service satellites passed by the satellite and the apogee altitude can be calculated, and by comparing and calculating the two corresponding relationships, the collision risk between the satellite and other service satellites during the deorbiting control process can be obtained.

[0111] The risk assessment during the satellite deorbiting process is analyzed by calculating whether there is an overlap in the orbital altitude between the satellite and other service satellites passing by during the deorbiting process. If so, there is a collision risk; otherwise, there is no collision risk.

[0112] Specifically, the collision risk assessment of the target satellite's deorbiting process based on the first simulation data, the second simulation data, and the third simulation data, and obtaining the collision risk assessment result includes:

[0113] Generating a first relationship curve between the sub-satellite point longitude and the perigee altitude passed by the target satellite during the deorbiting process based on the first simulation data, the second simulation data, and the third simulation data, and a second relationship curve between the sub-satellite point longitude of other service satellites passed by the target satellite during the deorbiting process and the apogee altitude;

[0114] For the same sub-satellite point longitude in the first relationship curve and the second relationship curve, when it is detected that there is a service satellite among the other service satellites whose apogee altitude is greater than or equal to the perigee altitude of the target satellite, the detected service satellite is determined as at least one risk satellite;

[0115] Detecting whether there is a collision risk between each risk satellite and the target satellite;

[0116] When it is not detected that there is a collision risk between a risk satellite and the target satellite, determining that the collision risk assessment result is that there is no collision risk for the target satellite during the deorbiting process; or

[0117] When it is detected that there is a collision risk between a risk satellite and the target satellite, determining that the collision risk assessment result is that there is a collision risk for the target satellite during the deorbiting process;

[0118] Wherein, calculating the difference between the apogee altitude of each risk satellite and the perigee altitude of the target satellite; when it is detected that the difference corresponding to a risk satellite is greater than the configured distance, determining that there is no collision risk between the detected risk satellite and the target satellite; or when it is detected that the difference corresponding to a risk satellite is less than or equal to the configured distance, determining that there is a collision risk between the detected risk satellite and the target satellite.

[0119] For example: Please refer to Figure 2 , Figure 2 is a schematic diagram of the first relationship curve and the second relationship curve of the present invention. As can be seen from Figure 2 it, during the deorbiting control process of the target satellite (the data changes from right to left), at the same sub-satellite point longitude, there are four service satellites whose apogee altitudes are greater than the perigee altitude of the target satellite, that is, the height value of the point on the second relationship curve is greater than the height value of the point on the first curve (for example, for Figure 2At the longitude of the sub-satellite point of 130 degrees east, the apogee altitude of the service satellite is 18 km, which is greater than the perigee altitude of the target satellite (-38 km). Therefore, it is considered that there is a situation where the orbital altitudes coincide with those of the passing service satellites, that is, there is a collision risk between the target satellite and these two service satellites, and further simulation calculations are required. Since 18 - (-38) = 56 km, which is greater than the configured distance of 50 km, it can be determined that there is no collision risk between the target satellite and these two service satellites.

[0120] S16. When the collision risk assessment result shows that there is no collision risk for the target satellite during the deorbiting process, an off-orbit control strategy for the target satellite is generated based on the first ignition strategy, the second ignition strategy, and the third ignition strategy.

[0121] According to the orbital characteristics of geostationary satellites, due to eccentricity, the perigee and apogee altitudes of the satellite are respectively below and above the geosynchronous orbit band, and the satellite crosses the geosynchronous orbit band twice a day. During the deorbiting control process, there is a collision risk with other geostationary satellites. Therefore, during the deorbiting control of a small-thrust satellite, first, it is necessary to quickly raise the perigee of the orbit as soon as possible to reduce the collision risk; second, according to the requirements of eccentricity, it is necessary to implement circular orbit control, and only perform orbit control near the apogee every day. The main goal is to make the perigee and apogee altitudes the same; finally, continuously carry out the overall elevation of the orbital altitude, and perform two batches of orbit control near the perigee and apogee respectively every day, so that the perigee altitude of the orbit is 300 km higher than the geosynchronous orbit band to complete the deorbiting control.

[0122] Specifically, continuing with the previous example, when the collision risk assessment result shows that there is no collision risk for the target satellite during the deorbiting process, the deorbiting control strategy is as follows: first, simulate the elevation of the perigee altitude based on a continuous 12-hour ignition strategy, then simulate the circular orbit control based on a continuous 2-hour ignition strategy, and finally simulate the overall elevation of the orbital altitude based on a continuous 8-hour ignition strategy to continuously raise the satellite to the graveyard orbit.

[0123] In this embodiment, for the deorbiting problem of geostationary orbit satellites at the end of their service life, based on the influence of small thrust and long-term orbit control on the elevation efficiency, a three-step deorbiting control scheme of "first quickly raise the perigee altitude, then circularize the orbit, and finally continuously raise it to the graveyard orbit" is designed, which effectively reduces the collision risk with other normal on-orbit satellites. Considering that the current electric propulsion technology has been widely applied to geostationary orbit satellites, this method can also be extended to the deorbiting control tasks of such satellites.

[0124] In this embodiment, after determining that the collision risk assessment result is that there is a collision risk for the target satellite during the deorbiting process, the method further includes:

[0125] Obtain the delay step size;

[0126] When performing the perigee altitude lift simulation on the target satellite based on the first ignition strategy, perform the delayed ignition simulation according to the delay step size, and sequentially update the first simulation data, the second simulation data, and the third simulation data after each delayed ignition simulation;

[0127] Conduct a collision risk assessment on the deorbiting process of the target satellite based on the updated first simulation data, the second simulation data, and the third simulation data. Until the obtained collision risk assessment result shows that there is no collision risk during the deorbiting process of the target satellite, update the deorbiting control strategy of the target satellite according to the ignition strategy corresponding to the current round of delayed ignition simulation, the total number of rounds of delayed ignition simulation, and the delay step size.

[0128] For example: when the delay step size is 1 day, during the perigee altitude lift simulation of the target satellite based on the first ignition strategy, the ignition can be postponed by one day to widen the approaching situation between the target satellite and the service satellite, and after the postponement, re-conduct the collision risk assessment based on all the updated data. Until there is no collision risk, record the number of delays (i.e., the total number of rounds of delayed ignition simulation) and the delay step size, and update the deorbiting control strategy of the target satellite accordingly, so as to respond and adjust in a timely manner when a collision risk is detected.

[0129] It can be seen from the above technical solutions that the present invention can determine different ignition strategies according to the control efficiency index values of each candidate ignition strategy, first perform the perigee altitude lift simulation based on the ignition strategy, then perform the circular orbit control simulation, and finally execute the overall orbit altitude lift simulation to continuously lift the satellite to the graveyard orbit. Further, conduct a collision risk assessment on the deorbiting process of the target satellite based on the simulation data, and when the collision risk assessment result shows that there is no collision risk between the target satellite and other satellites during the deorbiting process, generate a deorbiting control strategy for the target satellite based on the different ignition strategies used to assist in performing a safe and efficient deorbiting control task for the satellite.

[0130] As Figure 3 shown, it is a functional module diagram of a preferred embodiment of the satellite deorbiting control device based on small thrust of the present invention. The satellite deorbiting control device 11 based on small thrust includes a determination unit 110, a calculation unit 111, a simulation unit 112, an evaluation unit 113, and a generation unit 114. The modules / units referred to in the present invention refer to a series of computer program segments that can be executed by a processor and can complete fixed functions, and are stored in the memory. In this embodiment, the functions of each module / unit will be described in detail in subsequent embodiments.

[0131] Among them, the determination unit 110 is configured to determine at least one candidate ignition strategy in response to a deorbiting control instruction for a target satellite, where the target satellite is a geostationary orbit satellite at the end of its life;

[0132] The calculation unit 111 is configured to calculate the control efficiency index value of each candidate ignition strategy according to the small-thrust orbit dynamics model;

[0133] The simulation unit 112 is configured to select a first ignition strategy from the at least one candidate ignition strategy according to the control efficiency index value of each candidate ignition strategy, and perform a perigee altitude lift simulation on the target satellite based on the first ignition strategy to obtain first simulation data;

[0134] The simulation unit 112 is further configured to select a second ignition strategy from the at least one candidate ignition strategy according to the control efficiency index value of each candidate ignition strategy, and perform a circular orbit control simulation on the target satellite based on the second ignition strategy to obtain second simulation data;

[0135] The simulation unit 112 is further configured to obtain the configured ignition batch and the orbit determination reserved duration, select a third ignition strategy from the at least one candidate ignition strategy according to the configured ignition batch and the orbit determination reserved duration, and perform an overall orbit altitude lift simulation on the target satellite according to the third ignition strategy to obtain third simulation data;

[0136] The evaluation unit 113 is configured to perform a collision risk assessment on the deorbiting process of the target satellite based on the first simulation data, the second simulation data, and the third simulation data to obtain a collision risk assessment result;

[0137] The generation unit 114 is configured to generate a deorbiting control strategy for the target satellite based on the first ignition strategy, the second ignition strategy, and the third ignition strategy when the collision risk assessment result shows that there is no collision risk during the deorbiting process of the target satellite.

[0138] It can be seen from the above technical solutions that the present invention can determine different ignition strategies according to the control efficiency index values of each candidate ignition strategy, and first perform a perigee altitude lift simulation based on the ignition strategy, then perform a circular orbit control simulation, and finally perform an overall orbit altitude lift simulation to continuously lift the satellite to the graveyard orbit. Further, a collision risk assessment is performed on the deorbiting process of the target satellite according to the simulation data, and when the collision risk assessment result shows that there is no collision risk with other satellites during the deorbiting process of the target satellite, a deorbiting control strategy for the target satellite is generated based on the different ignition strategies used to assist in performing a safe and efficient deorbiting control task for the satellite.

[0139] As shown Figure 4 in the figure, it is a schematic structural diagram of a computer device which is a preferred embodiment of the method for controlling the satellite's orbit departure based on a small amount of thrust according to the present invention.

[0140] The computer device 1 may include a memory 12, a processor 13 and a bus (wherein the bus is the arrow in the figure), and may also include a computer program stored in the memory 12 and executable on the processor 13, such as a satellite orbit departure control program based on a small amount of thrust.

[0141] Those skilled in the art can understand that the schematic diagram is only an example of the computer device 1, and does not constitute a limitation on the computer device 1. The computer device 1 may be a bus structure or a star structure, and the computer device 1 may also include more or fewer other hardware or software than shown in the figure, or different component arrangements. For example, the computer device 1 may also include input / output devices, network access devices, etc.

[0142] It should be noted that the computer device 1 is only an example, and other existing or future possible electronic products that can be adapted to the present invention should also be included within the protection scope of the present invention and are included herein by reference.

[0143] Among them, the memory 12 includes at least one type of readable storage medium. The readable storage medium includes flash memory, mobile hard disk, multimedia card, card-type memory (such as SD or DX memory, etc.), magnetic memory, magnetic disk, optical disk, etc. The memory 12 may be an internal storage unit of the computer device 1 in some embodiments, such as the mobile hard disk of the computer device 1. The memory 12 may also be an external storage device of the computer device 1 in other embodiments, such as a plug-in mobile hard disk, a Smart Media Card (SMC), a Secure Digital (SD) card, a Flash Card, etc. equipped on the computer device 1. Further, the memory 12 may include both an internal storage unit and an external storage device of the computer device 1. The memory 12 can be used not only to store application software installed on the computer device 1 and various types of data, such as the code of the satellite orbit departure control program based on a small amount of thrust, but also to temporarily store data that has been output or will be output.

[0144] In some embodiments, the processor 13 may be composed of an integrated circuit. For example, it may be composed of a single packaged integrated circuit, or may be composed of multiple packaged integrated circuits with the same or different functions, including the combination of one or more central processing units (CPUs), microprocessors, digital processing chips, graphics processors, and various control chips, etc. The processor 13 is the control core (Control Unit) of the computer device 1, connecting various components of the entire computer device 1 through various interfaces and circuits. By running or executing programs or modules stored in the memory 12 (such as executing the satellite deorbiting control program based on small thrust, etc.), and by calling the data stored in the memory 12, it performs various functions of the computer device 1 and processes data.

[0145] The processor 13 executes the operating system of the computer device 1 and various installed application programs. The processor 13 executes the application programs to implement the steps in the above-mentioned embodiments of various satellite deorbiting control methods based on small thrust, such as Figure 1 the steps shown.

[0146] Exemplarily, the computer program may be divided into one or more modules / units. The one or more modules / units are stored in the memory 12 and executed by the processor 13 to complete the present invention. The one or more modules / units may be a series of computer-readable instruction segments capable of performing specific functions, and these instruction segments are used to describe the execution process of the computer program in the computer device 1. For example, the computer program may be divided into a determination unit 110, a calculation unit 111, a simulation unit 112, an evaluation unit 113, and a generation unit 114.

[0147] The above-mentioned integrated units implemented in the form of software function modules may be stored in a computer-readable storage medium. The above-mentioned software function modules are stored in a storage medium, including several instructions for causing a computer device (which may be a personal computer, a computer device, or a network device, etc.) or a processor to execute parts of the satellite deorbiting control method based on small thrust described in various embodiments of the present invention.

[0148] If the modules / units integrated in the computer device 1 are implemented in the form of software functional units and sold or used as independent products, they can be stored in a computer-readable storage medium. Based on this understanding, to implement all or part of the processes in the above-described embodiment methods of the present invention, it can also be completed by a computer program instructing relevant hardware devices. 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 various method embodiments can be implemented.

[0149] 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 disk, computer memory, read-only memory (ROM, Read-Only Memory), random access memory, etc.

[0150] Furthermore, the computer-readable storage medium mainly includes a storage program area and a storage data area. Among them, the storage program area can store an operating system, application programs required for at least one function, etc.; the storage data area can store data created according to the use of the blockchain node, etc.

[0151] The blockchain referred to in the present invention is a new application mode of computer technologies such as distributed data storage, peer-to-peer transmission, consensus mechanism, and encryption algorithm. Blockchain, in essence, is a decentralized database, a series of data blocks generated by using cryptographic methods. Each data block contains information about a batch of network transactions, which is used to verify the validity (anti-counterfeiting) of the information and generate the next block. The blockchain can include a blockchain underlying platform, a platform product service layer, an application service layer, etc.

[0152] The bus can be a peripheral component interconnect (PCI) bus or an extended industry standard architecture (EISA) bus, etc. This bus can be divided into an address bus, a data bus, a control bus, etc. For the sake of representation, in Figure 4 only a straight line is used to represent it, but it does not mean that there is only one bus or one type of bus. The bus is set to realize the connection and communication between the memory 12 and at least one processor 13, etc.

[0153] Although not shown, the computer device 1 may further include a power source (such as a battery) for powering each component. Preferably, the power source can be logically connected to the at least one processor 13 through a power management device, so as to implement functions such as charging management, discharging management, and power consumption management through the power management device. The power source may further include any components such as one or more DC or AC power sources, a recharge device, a power failure detection circuit, a power converter or inverter, a power status indicator, etc. The computer device 1 may further include various sensors, a Bluetooth module, a Wi-Fi module, etc., which will not be elaborated here.

[0154] Furthermore, the computer device 1 may further include a network interface. Optionally, the network interface may include a wired interface and / or a wireless interface (such as a WI-FI interface, a Bluetooth interface, etc.), which is generally used to establish a communication connection between the computer device 1 and other computer devices.

[0155] Optionally, the computer device 1 may further include a user interface. The user interface may be a display (Display), an input unit (such as a keyboard (Keyboard)). Optionally, the user interface may also be a standard wired interface or a wireless interface. Optionally, in some embodiments, the display may be an LED display, a liquid crystal display, a touch liquid crystal display, and an OLED (Organic Light-Emitting Diode) toucher, etc. Among them, the display may also be appropriately referred to as a display screen or a display unit, which is used to display the information processed in the computer device 1 and to display a visual user interface.

[0156] It should be understood that the above embodiments are only for illustration purposes and are not limited by this structure in the scope of the patent application.

[0157] Those skilled in the art can understand that Figure 4 the shown structure does not limit the computer device 1, and it may include fewer or more components than shown, or combine certain components, or have different component arrangements.

[0158] In combination with Figure 1 , the memory 12 in the computer device 1 stores multiple instructions to implement a satellite deorbiting control method based on a small thrust, and the processor 13 can execute the multiple instructions to implement:

[0159] In response to a deorbiting control instruction for a target satellite, determining at least one candidate ignition strategy; wherein, the target satellite is a geosynchronous orbit satellite at the end of its life;

[0160] Calculating the control efficiency index value of each candidate ignition strategy according to the small thrust orbit dynamics model;

[0161] Select a first ignition strategy from the at least one candidate ignition strategy according to the control efficiency index value of each candidate ignition strategy, and perform a perigee altitude lift simulation on the target satellite based on the first ignition strategy to obtain first simulation data;

[0162] Select a second ignition strategy from the at least one candidate ignition strategy according to the control efficiency index value of each candidate ignition strategy, and perform a circular orbit control simulation on the target satellite based on the second ignition strategy to obtain second simulation data;

[0163] Obtain the configured ignition batch and the reserved duration for orbit determination, select a third ignition strategy from the at least one candidate ignition strategy according to the configured ignition batch and the reserved duration for orbit determination, and perform an overall orbit altitude lift simulation on the target satellite according to the third ignition strategy to obtain third simulation data;

[0164] Perform a collision risk assessment on the deorbiting process of the target satellite based on the first simulation data, the second simulation data, and the third simulation data to obtain a collision risk assessment result;

[0165] When the collision risk assessment result shows that there is no collision risk during the deorbiting process of the target satellite, generate a deorbiting control strategy for the target satellite based on the first ignition strategy, the second ignition strategy, and the third ignition strategy.

[0166] Specifically, for the specific implementation method of the above instructions by the processor 13, reference can be made to Figure 1 the description of the relevant steps in the corresponding embodiment, which will not be elaborated here.

[0167] It should be noted that all the data involved in this case are legally obtained. The non-company software tools or components appearing in the embodiments of this application are only for illustrative introduction and do not represent actual use.

[0168] In several embodiments provided by the present invention, it should be understood that the disclosed systems, devices, and methods can be implemented in other ways. For example, the device embodiments described above are only illustrative. For example, the division of the modules is only a logical function division, and there can be other division methods in actual implementation.

[0169] The present invention can be used in numerous general-purpose or special-purpose computer system environments or configurations. For example: personal computers, server computers, handheld or portable devices, tablet devices, multi-processor systems, microprocessor-based systems, set-top boxes, programmable consumer electronics, network PCs, minicomputers, mainframe computers, distributed computing environments including any of the above systems or devices, and so on. The present invention can be described in the general context of computer-executable instructions executed by a computer, such as program modules. Generally, program modules include routines, programs, objects, components, data structures, etc. that perform specific tasks or implement specific abstract data types. The present invention can also be practiced in a distributed computing environment where tasks are performed by remote processing devices connected through a communication network. In a distributed computing environment, program modules can be located in local and remote computer storage media including storage devices.

[0170] The modules described as separate components may or may not be physically separated, and the components shown as modules may or may not be physical units, that is, they can be located in one place or distributed to multiple network units. Some or all of the modules can be selected according to actual needs to achieve the purpose of the solution of this embodiment.

[0171] In addition, in each embodiment of the present invention, the functional modules can be integrated in a processing unit, or each unit can exist physically separately, or two or more units can be integrated in one unit. The above integrated units can be implemented in the form of hardware or in the form of a combination of hardware and software functional modules.

[0172] For those skilled in the art, it is obvious that the present invention is not limited to the details of the above-described exemplary embodiments, and without departing from the spirit or basic characteristics of the present invention, the present invention can be implemented in other specific forms.

[0173] Therefore, from any point of view, the embodiments should be regarded as exemplary and non-restrictive. The scope of the present invention is defined by the appended claims rather than the above description. Therefore, all changes falling within the meaning and scope of the equivalent elements of the claims are intended to be encompassed by the present invention. Any reference signs attached to the claims should not be construed as limiting the claims involved.

[0174] In addition, it is obvious that the word "including" does not exclude other units or steps, and the singular does not exclude the plural. The multiple units or devices described in the present invention can also be implemented by one unit or device through software or hardware. Words such as first, second, etc. are used to denote names and do not denote any particular order.

[0175] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and not to limit them. Although the present invention has been described in detail with reference to the preferred embodiments, those of ordinary skill in the art should understand that the technical solutions of the present invention can be modified or equivalently replaced without departing from the spirit and scope of the technical solutions of the present invention.

Claims

1. A satellite deorbiting control method based on a small amount of thrust, characterized in that, The satellite deorbiting control method based on small thrust includes: In response to a deorbiting control instruction for a target satellite, determining at least one candidate ignition strategy; wherein, the target satellite is a geosynchronous orbit satellite at the end of its life; Calculating the control efficiency index value of each candidate ignition strategy according to the small thrust orbit dynamics model; Selecting a first ignition strategy from the at least one candidate ignition strategy according to the control efficiency index value of each candidate ignition strategy, and performing a perigee altitude elevation simulation on the target satellite based on the first ignition strategy to obtain first simulation data; Selecting a second ignition strategy from the at least one candidate ignition strategy according to the control efficiency index value of each candidate ignition strategy, and performing a circular orbit control simulation on the target satellite based on the second ignition strategy to obtain second simulation data; Obtaining the configured ignition batch and the orbit determination reserved duration, selecting a third ignition strategy from the at least one candidate ignition strategy according to the configured ignition batch and the orbit determination reserved duration, and performing an overall orbit altitude elevation simulation on the target satellite according to the third ignition strategy to obtain third simulation data; Performing a collision risk assessment on the deorbiting process of the target satellite based on the first simulation data, the second simulation data, and the third simulation data to obtain a collision risk assessment result; When the collision risk assessment result shows that there is no collision risk during the deorbiting process of the target satellite, generating a deorbiting control strategy for the target satellite based on the first ignition strategy, the second ignition strategy, and the third ignition strategy.

2. The satellite deorbiting control method based on a small amount of thrust according to claim 1, wherein The determining at least one candidate ignition strategy includes: Obtaining the working efficiency of the small thruster of the target satellite and the orbital period of the geostationary satellite; Calculating the minimum working duration of the small thruster according to the working efficiency of the small thruster; Establishing a time interval with the minimum working duration as the minimum value and the orbital period as the maximum value; Randomly selecting the at least one candidate ignition strategy from the time interval.

3. The satellite deorbiting control method based on a small amount of thrust according to claim 1, characterized in that, The selecting a first ignition strategy from the at least one candidate ignition strategy according to the control efficiency index value of each candidate ignition strategy, and performing a perigee altitude elevation simulation on the target satellite based on the first ignition strategy to obtain first simulation data includes: Calculating the initial perigee altitude of the target satellite according to the small thrust orbit dynamics model; Obtaining the nominal semi-major axis and the configured eccentricity; Calculating the product of the nominal semi-major axis and the configured eccentricity to obtain the target perigee elevation height; Calculating the sum of the target perigee elevation height and the initial perigee altitude to obtain the control amount; Obtaining the perigee elevation height corresponding to each candidate ignition strategy from the control efficiency index value of each candidate ignition strategy; Calculating the quotient of the control amount and the perigee elevation height corresponding to each candidate ignition strategy to obtain the elevation time of each candidate ignition strategy; Obtaining the perigee elevation efficiency corresponding to each candidate ignition strategy from the control efficiency index value of each candidate ignition strategy; Select the first ignition strategy from the at least one candidate ignition strategy according to the lift time of each candidate ignition strategy and the perigee lift efficiency corresponding to each candidate ignition strategy; Obtain the lift time corresponding to the first ignition strategy as the target lift time, obtain the ignition duration corresponding to the first ignition strategy as the first target ignition duration, and obtain the execution period; Calculate the quotient of the target lift time and the execution period to obtain the target ignition batch; Within each execution period, perform a batch of continuous ignition control simulations on the target satellite according to the first target ignition duration, and stop the simulation until the continuous ignition control simulations of the target ignition batch are completed; Obtain the semi-major axis change amount, perigee altitude change amount, and apogee altitude change amount corresponding to the first ignition strategy from the control efficiency index values of each candidate ignition strategy, and obtain the semi-major axis, perigee altitude, and apogee altitude obtained after each batch of continuous ignition control simulations; Integrate the obtained data to obtain the first simulation data.

4. The satellite orbit departure control method based on small thrust according to claim 3, characterized in that, The selecting a second ignition strategy from the at least one candidate ignition strategy according to the control efficiency index values of each candidate ignition strategy and performing a circular orbit control simulation on the target satellite based on the second ignition strategy to obtain the second simulation data includes: Select the candidate ignition strategy with the highest perigee lift efficiency from the at least one candidate ignition strategy as the second ignition strategy; Obtain the ignition duration corresponding to the second ignition strategy as the second target ignition duration; Construct a target ignition area with the apogee as the center and a first preset value as the radius; Within each execution period, perform a batch of continuous ignition control simulations on the target satellite in the target ignition area according to the second target ignition duration, and stop the next batch of simulations until the difference between the perigee altitude and the apogee altitude obtained after the simulation is less than or equal to the configured threshold; Obtain the semi-major axis change amount, perigee altitude change amount, and apogee altitude change amount corresponding to the second ignition strategy from the control efficiency index values of each candidate ignition strategy, and obtain the semi-major axis, perigee altitude, and apogee altitude obtained after each batch of continuous ignition control simulations; Integrate the obtained data to obtain the second simulation data.

5. The satellite orbit departure control method based on a small amount of thrust according to claim 3, characterized in that The selecting a third ignition strategy from the at least one candidate ignition strategy according to the configured ignition batch and the orbit determination reserved duration and performing an overall orbit altitude lift simulation on the target satellite according to the third ignition strategy to obtain the third simulation data includes: Calculate the product of the configured ignition batch and the orbit determination reserved duration to obtain the first duration; Calculate the difference between the execution period and the first duration to obtain the second duration; Calculate the quotient of the second duration and the configured ignition batch, and select the third ignition strategy from the at least one candidate ignition strategy according to the calculated quotient; Obtain the ignition duration corresponding to the third ignition strategy as the third target ignition duration; Construct two ignition areas with the perigee and the apogee as the centers and a second preset value as the radius respectively; In each execution cycle, a batch of continuous ignition control simulations are performed on the target satellite in two ignition regions according to the third target ignition duration until the perigee altitude obtained after the simulation is higher than the preset distance in the geosynchronous orbit belt, and then the next batch of simulations is stopped; Obtain the data obtained by simulation according to the third ignition strategy as the third simulation data.

6. The satellite deorbiting control method based on a small amount of thrust according to claim 1, characterized in that The collision risk assessment of the target satellite's deorbiting process based on the first simulation data, the second simulation data, and the third simulation data, and the obtained collision risk assessment results include: Generate a first relationship curve between the sub-satellite point longitude and the perigee altitude passed by the target satellite during the deorbiting process, and a second relationship curve between the sub-satellite point longitude and the apogee altitude of other service satellites passed by the target satellite during the deorbiting process based on the first simulation data, the second simulation data, and the third simulation data; For the same sub-satellite point longitude in the first relationship curve and the second relationship curve, when it is detected that the apogee altitude of a service satellite in the other service satellites is greater than or equal to the perigee altitude of the target satellite, determine the detected service satellite as at least one risk satellite; Detect whether there is a collision risk between each risk satellite and the target satellite; When it is not detected that there is a collision risk between a risk satellite and the target satellite, determine that the collision risk assessment result is that there is no collision risk for the target satellite during the deorbiting process; or When it is detected that there is a collision risk between a risk satellite and the target satellite, determine that the collision risk assessment result is that there is a collision risk for the target satellite during the deorbiting process; Wherein, calculate the difference between the apogee altitude of each risk satellite and the perigee altitude of the target satellite; when it is detected that the difference corresponding to a risk satellite is greater than the configured distance, determine that there is no collision risk between the detected risk satellite and the target satellite; or when it is detected that the difference corresponding to a risk satellite is less than or equal to the configured distance, determine that there is a collision risk between the detected risk satellite and the target satellite.

7. The satellite deorbiting control method based on a small amount of thrust according to claim 6, characterized in that After determining that the collision risk assessment result is that there is a collision risk for the target satellite during the deorbiting process, the method further includes: Obtain a delay step size; When performing a perigee altitude elevation simulation on the target satellite based on the first ignition strategy, perform a delayed ignition simulation according to the delay step size, and sequentially update the first simulation data, the second simulation data, and the third simulation data after each delayed ignition simulation; Perform a collision risk assessment on the deorbiting process of the target satellite according to the updated first simulation data, the second simulation data, and the third simulation data until the obtained collision risk assessment result shows that there is no collision risk for the target satellite during the deorbiting process, and update the deorbiting control strategy of the target satellite according to the ignition strategy corresponding to the current round of delayed ignition simulation, the total number of rounds of delayed ignition simulation, and the delay step size.

8. A satellite de-orbit control device based on a small amount of thrust, characterized in that, The satellite deorbiting control device based on small thrust includes: A determination unit, configured to determine at least one candidate ignition strategy in response to a deorbiting control instruction for a target satellite; wherein, the target satellite is a geosynchronous orbit satellite at the end of its life; A calculation unit, configured to calculate the control efficiency index value of each candidate ignition strategy according to a small-thrust orbit dynamics model; A simulation unit, configured to select a first ignition strategy from the at least one candidate ignition strategy according to the control efficiency index value of each candidate ignition strategy, and perform a perigee altitude elevation simulation on the target satellite based on the first ignition strategy to obtain first simulation data; The simulation unit is further configured to select a second ignition strategy from the at least one candidate ignition strategy according to the control efficiency index value of each candidate ignition strategy, and perform a circular orbit control simulation on the target satellite based on the second ignition strategy to obtain second simulation data; The simulation unit is further configured to obtain a configured ignition batch and an orbit determination reserved duration, select a third ignition strategy from the at least one candidate ignition strategy according to the configured ignition batch and the orbit determination reserved duration, and perform an overall orbit altitude elevation simulation on the target satellite according to the third ignition strategy to obtain third simulation data; An evaluation unit, configured to perform a collision risk assessment on the deorbiting process of the target satellite based on the first simulation data, the second simulation data and the third simulation data to obtain a collision risk assessment result; A generation unit, configured to generate a deorbiting control strategy for the target satellite based on the first ignition strategy, the second ignition strategy and the third ignition strategy when the collision risk assessment result shows that there is no collision risk during the deorbiting process of the target satellite.

9. A computer device, characterized in that, The computer device includes: A memory, storing at least one instruction; and A processor, configured to execute the instructions stored in the memory to implement the small-thrust-based satellite deorbiting control method according to any one of claims 1 to 7.

10. A computer-readable storage medium, characterized in that: At least one instruction is stored in the computer-readable storage medium, and the at least one instruction is executed by a processor in the computer device to implement the small-thrust-based satellite deorbiting control method according to any one of claims 1 to 7.

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