Satellite deorbit control method and system, electronic equipment and storage medium

Abstract

The invention discloses a satellite deorbit control method and system, electronic equipment and a storage medium, and relates to the technical field of satellite control, and the method comprises the steps: determining the optimal passive deorbit initial orbit height of a satellite; when the height of the active deorbiting orbit is not zero, calculating to obtain a daily time plan; according to the active deorbit orbit height, the preset ignition strategy and the daily time plan, the active deorbit scheme is used for transferring the satellite from the orbit when autonomous deorbit is started to the orbit of the optimal passive deorbit initial orbit height, and then the passive deorbit scheme is used for controlling the satellite to deorbit. According to the method, the active deorbit scheme and the passive deorbit scheme are combined, the satellite deorbit is controlled, the atmospheric resistance effect can be fully utilized, consumption of active deorbit propellants is reduced, the passive deorbit duration is shortened, and the engineering implementation operability is high.

Description

Background Art

[0002] The booming development of low-orbit satellite communications has become one of the most eye-catching competitive focuses in the world's science and technology field. Advanced low-orbit satellite communication technology will not only achieve global real-time connection, but also provide strong support for applications in many fields. The Internet satellite constellation currently under construction uses a group of satellites operating at an orbital altitude of 400km to 1500km to provide broadband Internet access services to the ground, and achieves global coverage through a network of multiple satellites.

[0003] According to the Interstellar Debris Coordination Committee (IADC), spacecraft must be deorbited within 25 years after failure to reduce the amount of space debris. This is the "25-year rule" for satellites. The treatment of satellites after their lifespan varies depending on the satellite's orbit. There are usually two methods. One is that for satellites in low orbit, the satellite can be lowered and returned to Earth. During the process of entering the Earth, the satellite will rub against the Earth's atmosphere violently, resulting in high temperatures and burning out. The other is that for satellites in geostationary orbit, due to the huge cost of re-entering the Earth, the satellite is usually pushed into a higher orbit, so that it can be away from other normal satellites and orbit the Earth for hundreds of years.

[0004] Low-orbit communication satellites operate in inclined or polar orbits, and are often equipped with dual-axis dual-wing SADA to achieve normal solar sails to the sun, ensuring energy stability during the operation of the entire satellite. Different satellite constellation construction plans and satellite deorbit requirements are different. Some are fully implemented in accordance with the 25-year regulations of IADC, while others hope to complete deorbit in a shorter time, such as 7 or 10 years.

[0005] Usually, satellite orbit control and deorbit missions are implemented using ground command control, which requires large ground mission planning resources and consumes large amounts of measurement, operation and control resources. In large-scale Internet constellations, autonomous deorbiting of satellites on board is the desired method for constellation operation and control. Currently, there is little research and implementation in this area, and even fewer autonomous deorbiting projects that take full advantage of atmospheric damping and solar panel control strategies are being implemented. Summary of the invention

[0006] The technical problem to be solved by the present invention is to address the deficiencies of the prior art and specifically provide a satellite deorbit control method, system, electronic device and storage medium, as follows:

[0007] 1) In a first aspect, the present invention provides a satellite deorbit control method, and the specific technical solution is as follows:

[0008] Determine the optimal initial orbital altitude of the satellite for passive deorbiting;

[0009] When the active deorbit orbit altitude is not zero, the cumulative total ignition time of the satellite transferred from the orbit when the autonomous deorbit was initiated to the orbit of the optimal passive deorbit initial orbit altitude is calculated, and the ignition time of the satellite in each target operation circle within the preset time is calculated based on the cumulative total ignition time, wherein the active deorbit orbit altitude is the height difference between the optimal passive deorbit initial orbit altitude and the altitude of the orbit when the satellite initiated the autonomous deorbit;

[0010] Based on the ignition duration of the satellite in each target operation cycle within the preset duration, and under the constraint of energy balance, the duration of the satellite's solar sail drive device facing the sun in each target operation cycle is calculated to obtain the time plan for facing the sun;

[0011] According to the active deorbit orbital height, preset ignition strategy and solar time planning, the active deorbit plan is used to transfer the satellite from the orbit when the autonomous deorbit is initiated to the orbit with the optimal passive deorbit initial orbital height, and then the passive deorbit plan is used to control the satellite to deorbit.

[0012] The beneficial effects of a satellite deorbit control method provided by the present invention are as follows:

[0013] When the height difference between the optimal passive deorbiting initial orbital altitude and the altitude of the orbit when the satellite initiates autonomous deorbiting, that is, the active deorbiting orbital altitude is not zero, a combination of active deorbiting scheme and passive deorbiting scheme is adopted to control the satellite deorbiting. This can make full use of the atmospheric drag, reduce the consumption of active deorbiting propellant, and shorten the passive deorbiting time. It is suitable for low-orbit communication satellites carrying large-area solar panels, and the project implementation is highly operational.

[0014] Based on the above scheme, the satellite deorbit control method of the present invention can also be improved as follows.

[0015] Furthermore, it also includes:

[0016] When the active deorbit orbit altitude is zero, the passive deorbit scheme is used to control the satellite to deorbit.

[0017] Furthermore, the total accumulated ignition time of the satellite from the orbit at the time of starting the autonomous deorbit to the orbit at the optimal passive deorbit initial orbit altitude is calculated, including:

[0018] Based on the preset ignition strategy, the optimal passive deorbit initial orbital altitude and the altitude of the satellite's orbit when it starts autonomous deorbit, and using the Hohmann orbit change principle, the total cumulative ignition time of the satellite's transfer from the orbit when it starts autonomous deorbit to the orbit at the optimal passive deorbit initial orbital altitude is calculated.

[0019] Further, determining the optimal passive deorbit initial orbital height of the satellite includes:

[0020] Based on the satellite's structure, mass characteristic parameters and deorbit time constraints, the STK software is used to determine the optimal initial orbital altitude of the satellite for passive deorbiting.

[0021] 2) In a second aspect, the present invention further provides a satellite deorbit control system, and the specific technical solution is as follows:

[0022] It includes the optimal passive deorbit initial orbit height determination module, the cumulative ignition total duration calculation module, the operation circle ignition duration calculation module, the sun time planning acquisition module and the deorbit control module;

[0023] The optimal passive deorbit initial orbit height determination module is used to: determine the optimal passive deorbit initial orbit height of the satellite;

[0024] The cumulative ignition total duration calculation module is used to: when the active deorbit orbit altitude is not zero, calculate the cumulative ignition total duration of the satellite transferred from the orbit when the autonomous deorbit is initiated to the orbit of the optimal passive deorbit initial orbit altitude, wherein the active deorbit orbit altitude is the height difference between the optimal passive deorbit initial orbit altitude and the altitude of the orbit when the satellite initiates the autonomous deorbit;

[0025] The operation circle ignition duration calculation module is used to: calculate the ignition duration of each target operation circle of the satellite within a preset duration according to the accumulated total ignition duration;

[0026] The solar time planning acquisition module is used to: based on the ignition duration of the satellite in each target operation cycle within the preset duration, and under the constraint of energy balance, calculate the duration of the solar sail driving device of the satellite facing the sun in each target operation cycle, and obtain the solar time planning;

[0027] The deorbit control module is used to: according to the active deorbit orbital height, preset ignition strategy and solar time planning, use the active deorbit plan to transfer the satellite from the orbit when the autonomous deorbit is started to the orbit with the optimal passive deorbit initial orbital height, and then use the passive deorbit plan to control the satellite deorbit.

[0028] Based on the above scheme, the satellite deorbit control system of the present invention can also be improved as follows.

[0029] Furthermore, the control deorbit module is also used for:

[0030] When the active deorbit orbit altitude is zero, the passive deorbit scheme is used to control the satellite to deorbit.

[0031] Furthermore, the cumulative total ignition duration calculation module is specifically used for:

[0032] Based on the preset ignition strategy, the optimal passive deorbit initial orbital altitude and the altitude of the satellite's orbit when it starts autonomous deorbit, and using the Hohmann orbit change principle, the total cumulative ignition time of the satellite's transfer from the orbit when it starts autonomous deorbit to the orbit at the optimal passive deorbit initial orbital altitude is calculated.

[0033] Furthermore, the optimal passive deorbit initial orbital height determination module is specifically used to determine the optimal passive deorbit initial orbital height of the satellite using STK software based on the satellite's structure, mass characteristic parameters and deorbit time constraints.

[0034] 3) In a third aspect, the present invention further provides an electronic device, comprising a processor, the processor being coupled to a memory, the memory storing at least one computer program, the at least one computer program being loaded and executed by the processor, so that the electronic device implements any of the above-mentioned satellite deorbit control methods.

[0035] 4) In a fourth aspect, the present invention further provides a computer-readable storage medium having a computer program stored thereon, and when the computer program is executed by a processor, the computer program implements any of the above-mentioned satellite deorbit control methods.

[0036] It should be noted that the beneficial effects achieved by the technical solutions of the second to fourth aspects of the present invention and the corresponding possible implementation methods can be found in the above-mentioned technical effects of the first aspect and its corresponding possible implementation methods, and will not be repeated here. BRIEF DESCRIPTION OF THE DRAWINGS

[0037] In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the following briefly introduces the drawings required for describing the embodiments of the present invention:

[0038] Figure 1 A schematic flow chart of a satellite deorbit control method according to an embodiment of the present invention;

[0039] Figure 2 This is a schematic diagram of the Hohmann transfer;

[0040] Figure 3 It is a schematic diagram of the structure of the solar sail panel driving device of the satellite;

[0041] Figure 4 Schematic diagram of firing for the first target run;

[0042] Figure 5 Schematic diagram of firing for non-first target runs;

[0043] Figure 6 A schematic structural diagram of a satellite deorbit control system according to an embodiment of the present invention;

[0044] Figure 7The figure is a schematic diagram of the structure of an electronic device according to an embodiment of the present invention. DETAILED DESCRIPTION

[0045] The principles and features of the present invention are described below. The examples given are only used to explain the present invention and are not used to limit the scope of the present invention.

[0046] The technical solution of the present invention and how the technical solution of the present invention solves the above-mentioned technical problems are described in detail below with specific embodiments. The following specific embodiments can be combined with each other, and the same or similar concepts or processes may not be repeated in some embodiments. The embodiments of the present invention will be described below in conjunction with the accompanying drawings.

[0047] like Figure 1 As shown, a satellite deorbit control method according to an embodiment of the present invention comprises the following steps:

[0048] S1. Determine the optimal initial orbital height of the satellite for passive deorbiting, specifically:

[0049] Based on the satellite's structure, mass characteristic parameters and deorbit time constraints, the optimal passive deorbit initial orbital height of the satellite is determined using STK software, including the following steps:

[0050] S10. Preset the initial value of the initial orbital altitude of the passive deorbiting to be M, in km.

[0051] S11. Combine M, the structure and quality characteristic parameters of the satellite, and use the satellite life analysis function of the STK software to calculate the passive deorbit time t of the satellite, and determine whether t is less than or equal to the deorbit time constraint to obtain a first judgment result. When the first judgment result is no, continue to execute S12. When the first judgment result is yes, use the initial value M of the passive deorbit initial orbital height as the optimal passive deorbit initial orbital height.

[0052] The deorbit duration constraint is: the difference between the deorbit duration required by the satellite's mission and the preset duration. The preset duration can be 1 year or can be set according to actual conditions.

[0053] S12, let m′=M-ΔH×N, M=1, use the calculated m′ as M, and execute S11 again. When the first judgment result is no, let N=N+1, and use the calculated m′ as M, and execute S11 again until the first judgment result is yes, and determine the current n′ as the optimal passive deorbit initial orbital height of the satellite, recorded as m′′, where ΔH is the iteration step size, the unit is km.

[0054] Among them, M can be 450km, ΔH can be 5km, and both M and ΔH can be set according to actual conditions.

[0055] S2. When the active deorbit orbit altitude is not zero, calculate the cumulative total ignition time of the satellite from the orbit when the autonomous deorbit is initiated to the orbit of the optimal passive deorbit initial orbit altitude, and calculate the ignition time of the satellite in each target operation circle within the preset time according to the cumulative total ignition time, wherein the active deorbit orbit altitude is the height difference between the optimal passive deorbit initial orbit altitude and the altitude of the orbit when the satellite initiates the autonomous deorbit.

[0056] Among them, according to the optimal passive deorbit initial orbital height calculated by S1, the active deorbit orbital height can be calculated as: H0-m″, where H0 is the orbital height when the satellite receives the ground command and starts autonomous deorbit, and m″ is the optimal passive deorbit initial orbital height.

[0057] In S2, the total accumulated ignition time of the satellite from the orbit at the time of autonomous deorbiting to the orbit at the optimal passive deorbiting initial orbital altitude is calculated. The specific implementation process is as follows:

[0058] Based on the preset ignition strategy, the optimal passive deorbit initial orbital altitude and the altitude of the satellite's orbit when it starts autonomous deorbit, and using the Hohmann orbit change principle, the total cumulative ignition time of the satellite's transfer from the orbit when it starts autonomous deorbit to the orbit at the optimal passive deorbit initial orbital altitude is calculated.

[0059] like Figure 2 The schematic diagram of Hohmann transfer is shown in FIG. 1 , where orbit A and orbit B are different orbits of the satellite. The orbit of the satellite when it starts autonomous deorbiting is taken as orbit A, and the orbit of the optimal passive deorbiting initial orbit height is taken as orbit B. The height of orbit A is r A , the height of track B is r B , r B =m″, record v A and v B They are the speeds required for the satellite to run along orbit A and orbit B, respectively, and their values ​​are: where μ = 3.9860044118 × 10 5 km 3 / s 2 , is the gravitational constant. The calculation methods of v1 and v2 are as follows: Therefore, the two velocity increments of the satellite during the Hohmann transfer process are: Δv1 = v1-v A , Δv2=v B-v2, therefore, the total speed increment -ΔV required for the satellite to maneuver from orbit B to orbit A is: -ΔV=k*(-Δv1-Δv2), where k is the efficiency conversion coefficient between the engine pulse thrust model and the engine continuous low thrust model, and the value range of k is 1.2 to 1.5. The specific value of k can be determined by performing multiple simulations using the STK orbital maneuvering module.

[0060] Further through the formula: T all =m×ΔV / F to calculate the total ignition time T all , unit is s; m is: the mass of the satellite when the deorbit mission is initiated, unit is kg; F is the thrust of the satellite's electric propulsion thruster.

[0061] In S2, the ignition duration of the satellite in each target operation cycle within the preset duration is calculated based on the accumulated total ignition duration. The specific implementation process is as follows:

[0062] Calculate the number of orbits of the satellite within the preset time as N all0 , take and N all0 The nearest even number N all , then there are N all target operation cycles, the ignition duration of the satellite in each target operation cycle is: T Q =T all / N all It should be noted that if N all0 is an even number, then N all0 =N all .

[0063] The preset duration is the preset duration in the deorbit duration constraint. Taking a low-orbit communication satellite as an example, a low-orbit communication satellite usually operates in an orbit range of 400 km to 1500 km in altitude, and one year can be reserved as the preset duration.

[0064] S3, based on the ignition duration of the satellite in each target operation cycle within the preset duration, and under the constraint of energy balance, calculate the duration of the satellite's solar sail driving device facing the sun in each target operation cycle, and obtain the time plan for facing the sun, which specifically includes the following steps:

[0065] Based on the energy consumption on the satellite and the power generation efficiency of the solar panels when the error of the solar panels is 5 degrees, it can be calculated that the time T that the satellite needs the solar panels to remain in the state of facing the sun in each target operation circle is SADA .

[0066] S4. Based on the active deorbit orbital height, preset ignition strategy and solar time planning, the active deorbit scheme is used to transfer the satellite from the orbit when the autonomous deorbit is initiated to the orbit with the optimal passive deorbit initial orbital height, and then the passive deorbit scheme is used to control the satellite to deorbit.

[0067] Among them, Figure 3 As shown, the satellite's solar sail driving device includes three rotating mechanisms arranged in sequence, the rotating axis of the rotating mechanism in the middle is perpendicular to the rotating axes of the rotating mechanisms on both sides, the rotating mechanisms on both sides are respectively connected to a solar sail, and each solar sail is provided with a sun sensor, specifically:

[0068] According to the unfolding state of the satellite's solar sail panels, a rotation mechanism is configured for each solar sail panel. The satellite includes two solar sail panels, which are respectively denoted as the first solar sail panel and the second solar sail panel. The rotation mechanism configured for the first solar sail panel is denoted as B1, and the rotation mechanism configured for the second solar sail panel is denoted as B2. A rotation mechanism is set between B1 and B2, denoted as A. B1, A and B2 are arranged in sequence. Figure 2 The cube in is used to represent the satellite body. The satellite body is: the remaining part of the satellite except the solar panels and the structure connecting the solar panels. The satellite body coordinate system is established with the center of mass of the satellite body as the origin: OXYZ. When both solar panels are deployed, the normal of the solar panel battery is consistent with the negative direction of the Z axis of the satellite body coordinate system. The position of B1 at this time is defined as the zero position of B1, the position of B2 at this time is defined as the zero position of B2, and the position of A at this time is defined as the zero position of A. When B1, B2 and A are all in the corresponding zero positions, the positive direction of the rotation axis of B1 is consistent with the negative direction of the X axis of the satellite body coordinate system, the positive direction of the rotation axis of B2 is consistent with the positive direction of the X axis of the satellite body coordinate system, and the positive direction of the rotation axis of A is consistent with the positive direction of the Y axis of the satellite body coordinate system.

[0069] B1, B2 and A all use stepper motors inside. B1 is used to drive the first solar sail panel to rotate around the rotation axis of B1 through the stepper motor. B2 is used to drive the second solar sail panel to rotate around the rotation axis of B2 through the stepper motor. A is used to drive the first solar sail panel and the second solar sail panel to rotate around the rotation axis of A at the same time through the stepper motor. In this process, the right-hand rule is used to determine the rotation direction of the first solar sail panel and the second solar sail panel.

[0070] Among them, B1, B2 and A return status information through telemetry, and the status information includes angle measurement information and whether the status of the rotating mechanism is normal, etc. Moreover, B1, B2 and A can receive instructions to realize functions such as "moving at a set speed", "moving at a set angle position" and "maintaining a state after moving to a set position".

[0071] B1, B2 and A all support speed priority motion mode, position priority motion mode and position hold mode. The speed priority motion mode is: set the motion mechanism's angular velocity to move according to the set angular velocity magnitude and direction; the position priority motion mode is: set the motion position and motion speed of the motion mechanism at the same time, and the rotating mechanism moves to the set position at the set speed, where the set motion speed only sets the speed magnitude, and the motion direction is determined by the rotating mechanism according to the shortest path principle; the position hold mode is for the driving mechanism to keep its current position.

[0072] Among them, A can rotate 360° continuously, B1 and B2 can move within the set range of motion, and the range of motion parameter setting is determined comprehensively based on the Beta angle and the occlusion of the star. Mechanical limit mechanisms are designed for B1 and B2 outside the comprehensively determined range of motion.

[0073] A solar sensor is arranged on each side of the positive direction of the normal line of the first solar panel and the second solar panel. The solar sensor is in a pyramid configuration. The solar sensor arranged on the first solar panel is recorded as the first solar sensor, and the solar sensor arranged on the second solar panel is recorded as the second solar sensor. The layout positions are as follows: Figure 3 The field of view of the first solar sensor and the second solar sensor is -90° to +90°. The first solar sensor and the second solar sensor collect solar vector information in the solar sensor measurement coordinate system set by themselves. The solar sensor measurement coordinate system can also be reset, and the electric propulsion thruster is configured in the negative direction of the X-axis of the satellite mass center body coordinate system.

[0074] When the sun enters the field of view of the first solar sensor and the second solar sensor, solar vector information can be obtained. If the sun does not enter the field of view of the first solar sensor and the second solar sensor, the obtained solar vector information is a zero vector.

[0075] Among them, the active de-orbit scheme can be an active de-orbit scheme with cumulative even-numbered laps, and the preset ignition strategy is an ignition strategy with symmetrical ignition positions in adjacent laps, such as Figure 4 and Figure 5 shown.

[0076] Among them, the active de-orbit plan for cumulative even-numbered laps is as follows:

[0077] The M′th target running circle is recorded as Q M′ , M' is a positive integer, and M' is not greater than N all When the satellite receives the autonomous deorbit mission start command sent by the ground, it first performs a 180° yaw attitude maneuver, and then implements the subsequent deorbit mission in this attitude state. Specifically:

[0078] 1) If Figure 4As shown, for the first target operation circle Q1, within the first attitude adjustment time T1, control A, B1, and B2 are all reset to zero, and the electric propulsion thruster is ignited, and the ignition lasts for T Q After that, the electric propulsion thruster is turned off, and the orbital period of the satellite in the current orbit is calculated according to the orbit determination data of the satellite's current orbit, and the satellite's attitude is pitched -90° attitude maneuvered, and the attitude adjustment is completed after the second attitude adjustment time T2. Among them, T1 and T2 are estimated by mathematical simulation during satellite program design, or determined by on-orbit tests before the deorbit mission is launched.

[0079] 2) If Figure 5 As shown, for the second target, the number of cycles is run to the Nth all If the target operation circle is completed, then within the attitude adjustment time T1, control A, B1, and B2 are all reset to zero and the attitude maneuver is completed. The adjusted attitude is: pitch angle is 0°, yaw angle is 180°, roll angle is 0°, and the electric propulsion thruster is ignited for T Q After that, the electric propulsion thruster is turned off, and the orbital period of the satellite in the current orbit is calculated according to the orbit determination data of the satellite's current orbit, and the satellite's attitude is pitched -90°, and the attitude adjustment is completed after T2 time. Figure 5 As shown, after T2 time, the posture adjustment is completed, and then after T Sat_M′ +T Sat_M′ / 2-T Q -After T2-T1 time, start Q M′+1 Attitude adjustment. The orbital period corresponding to the M′th target running circle is recorded as T Sat_M′ .

[0080] After each attitude adjustment is completed, the angular velocity and Beta angle of the current orbit are calculated according to the orbit determination data of the satellite's current orbit, and the angular velocity of the orbit corresponding to the M'th target running circle is recorded as Orbit M′ , the Beta angle of the orbit corresponding to the M′th target running circle is recorded as Beta M′ , the unit of angular velocity is deg / s, and the unit of Beta angle is deg.

[0081] It is determined whether the satellite is in the earth's shadow area according to the orbit determination data of the current orbit and the solar vector parameter to obtain a second determination result.

[0082] If the second judgment result is yes, A, B1, and B2 are all controlled to return to zero, and the state remains unchanged after returning to zero.

[0083] If the second judgment result is no, then control A, B1, and B2 to move together, drive the solar sail to track the sun, and set the angular velocity of A in the M'th target running circle to: 100×Orbit M′ ; The angular position of B1 and B2 in the M′th target running cycle is: Ang_Beta M′ and-Ang_Beta M′ , if the angle information obtained by Taimin measurement is within the time T sun The inner and outer angles are less than 5 degrees, and T sun <T SADA , then set the angular velocity of A to: Orbit M′ , until the attitude adjustment moment before the next ignition of the propulsion system, control A, B1, and B2 all return to zero position.

[0084] After the ignition of the last target operation circle is completed, the current orbital altitude is calculated based on the orbital parameters at the previous moment. If the orbit when the autonomous deorbit is initiated is lower than the optimal passive deorbit initial orbital altitude, the active deorbit process ends and the passive deorbit plan is implemented. Otherwise, the active deorbit plan with accumulated even circles continues to be implemented.

[0085] Among them, the passive deorbit scheme is as follows:

[0086] 1) Control the satellite to perform a -90° pitch attitude maneuver, and during the maneuver, keep the current positions of A, B1, and B2 unchanged.

[0087] 2) When the satellite completes the -90° pitch attitude maneuver, controls A, B1, and B2 are all returned to zero and remain unchanged.

[0088] The passive deorbiting scheme can make full use of the atmospheric drag and reduce the consumption of active deorbiting propellant.

[0089] Among them, controlling A, B1, and B2 to return to zero means: controlling A, B1, and B2 to move to their respective zero positions.

[0090] Optionally, in the above technical solution, it also includes:

[0091] S5. When the active deorbit orbit height is zero, the passive deorbit scheme is used to control the satellite to deorbit.

[0092] In the above embodiments, although the steps are numbered S1, S2, etc., these are only specific embodiments given by the present invention. Those skilled in the art may adjust the execution order of S1, S2, etc. according to actual conditions, which is also within the protection scope of the present invention. It can be understood that in some embodiments, some or all of the above embodiments may be included.

[0093] like Figure 6As shown, a satellite deorbit control system 200 according to an embodiment of the present invention includes an optimal passive deorbit initial orbit height determination module 201, a cumulative ignition total duration calculation module 202, an operation circle ignition duration calculation module 203, a daily time planning acquisition module 204 and a deorbit control module 205;

[0094] The optimal passive deorbit initial orbit height determination module 201 is used to: determine the optimal passive deorbit initial orbit height of the satellite;

[0095] The cumulative ignition total duration calculation module 202 is used to: when the active deorbit orbit altitude is not zero, calculate the cumulative ignition total duration of the satellite transferred from the orbit when the autonomous deorbit is initiated to the orbit of the optimal passive deorbit initial orbit altitude, wherein the active deorbit orbit altitude is the height difference between the optimal passive deorbit initial orbit altitude and the altitude of the orbit when the satellite initiates the autonomous deorbit;

[0096] The operation cycle ignition duration calculation module 203 is used to: calculate the ignition duration of each target operation cycle of the satellite within a preset duration according to the accumulated total ignition duration;

[0097] The solar time planning acquisition module 204 is used to: calculate the duration of the solar sail driving device of the satellite facing the sun in each target operation cycle within a preset duration, based on the ignition duration of the satellite in each target operation cycle within a preset duration, and under the constraint of energy balance, to obtain the solar time planning;

[0098] The deorbit control module 205 is used to: according to the active deorbit orbital height, the preset ignition strategy and the sun time planning, use the active deorbit scheme to transfer the satellite from the orbit when the autonomous deorbit is started to the orbit of the optimal passive deorbit initial orbital height, and then use the passive deorbit scheme to control the satellite to deorbit.

[0099] Optionally, in the above technical solution, the control derailment module 205 is further used for:

[0100] When the active deorbit orbit altitude is zero, the passive deorbit scheme is used to control the satellite to deorbit.

[0101] Optionally, in the above technical solution, the cumulative total ignition duration calculation module 202 is specifically used for:

[0102] Based on the preset ignition strategy, the optimal passive deorbit initial orbital altitude and the altitude of the satellite's orbit when it starts autonomous deorbit, and using the Hohmann orbit change principle, the total cumulative ignition time of the satellite's transfer from the orbit when it starts autonomous deorbit to the orbit at the optimal passive deorbit initial orbital altitude is calculated.

[0103] Optionally, in the above technical solution, the optimal passive deorbit initial orbital height determination module 201 is specifically used to determine the optimal passive deorbit initial orbital height of the satellite using STK software based on the structure, mass characteristic parameters and deorbit time constraints of the satellite.

[0104] It should be noted that the beneficial effects of the satellite deorbit control method provided in the above embodiment are the same as the beneficial effects of the satellite deorbit control system 200, which will not be described in detail here. In addition, when the system provided in the above embodiment realizes its functions, only the division of the above functional modules is used as an example. In practical applications, the above functions can be assigned to different functional modules as needed, that is, the system can be divided into different functional modules according to actual conditions to complete all or part of the functions described above. In addition, the system and method embodiments provided in the above embodiment belong to the same concept, and the specific implementation process is detailed in the method embodiment, which will not be described in detail here.

[0105] Among them, the satellite deorbit control system of the present invention can be a computer program (including program code) running in a computer device. For example, the satellite deorbit control system of the present invention is an application software that can be used to execute the corresponding steps in the satellite deorbit control method of the present invention.

[0106] In some embodiments, the satellite deorbit control system of the present invention can be implemented by a combination of software and hardware. As an example, the satellite deorbit control system of the present invention can be a processor in the form of a hardware decoding processor, which is programmed to execute the satellite deorbit control method of the present invention. For example, the processor in the form of a hardware decoding processor can adopt one or more application specific integrated circuits (ASICs), DSPs, programmable logic devices (PLDs), complex programmable logic devices (CPLDs), field programmable gate arrays (FPGAs) or other electronic components.

[0107] The modules involved in the embodiments of the present invention may be implemented in software or hardware. The name of a module does not limit the module itself in some cases.

[0108] An electronic device according to an embodiment of the present invention includes a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, any of the above-mentioned satellite deorbit control methods is implemented. That is to say, an electronic device according to an embodiment of the present invention may include but is not limited to: a processor and a memory; the memory is used to store the computer program; the processor is used to execute the satellite deorbit control method shown in any embodiment of the present invention by calling the computer program.

[0109] In an alternative embodiment, an electronic device is provided, such as Figure 7 As shown, Figure 7 The electronic device 4000 shown includes: a processor 4001 and a memory 4003. The processor 4001 and the memory 4003 are connected, such as through a bus 4002. Optionally, the electronic device 4000 may also include a transceiver 4004, which may be used for data interaction between the electronic device and other electronic devices, such as data transmission and / or data reception. It should be noted that in actual applications, the transceiver 4004 is not limited to one, and the structure of the electronic device 4000 does not constitute a limitation on the embodiments of the present invention.

[0110] Processor 4001 may be a CPU (Central Processing Unit), a general-purpose processor, a DSP (Digital Signal Processor), an ASIC (Application Specific Integrated Circuit), an FPGA (Field Programmable Gate Array) or other programmable logic devices, transistor logic devices, hardware components or any combination thereof. It may implement or execute various exemplary logic blocks, modules and circuits described in conjunction with the disclosure of the present invention. Processor 4001 may also be a combination that implements computing functions, such as a combination of one or more microprocessors, a combination of a DSP and a microprocessor, etc.

[0111] The bus 4002 may include a path to transmit information between the above components. The bus 4002 may be a PCI (Peripheral Component Interconnect) bus or an EISA (Extended Industry Standard Architecture) bus. The bus 4002 may be divided into an address bus, a data bus, a control bus, etc. For ease of representation, Figure 7In the figure, only one thick line is used to represent the bus 4002, but this does not mean that there is only one bus or one type of bus.

[0112] The memory 4003 can be a ROM (Read Only Memory) or other types of static storage devices that can store static information and instructions, a RAM (Random Access Memory) or other types of dynamic storage devices that can store information and instructions, or an EEPROM (Electrically Erasable Programmable Read Only Memory), a CD-ROM (Compact Disc Read Only Memory) or other optical disk storage, optical disk storage (including compressed optical disk, laser disk, optical disk, digital versatile disk, Blu-ray disk, etc.), a magnetic disk storage medium or other magnetic storage device, or any other medium that can be used to carry or store the desired program code in the form of instructions or data structures and can be accessed by a computer, but is not limited to this.

[0113] The memory 4003 is used to store the application code (computer program) for executing the solution of the present invention, and the execution is controlled by the processor 4001. The processor 4001 is used to execute the application code stored in the memory 4003 to implement the content shown in the above method embodiment.

[0114] Among them, the electronic device can also be a terminal device, and the terminal device can be any terminal device that can install applications and access web pages through applications, including at least one of a smart phone, a tablet computer, a laptop computer, a desktop computer, a smart speaker, a smart watch, a smart TV, and a smart car device.

[0115] It should be noted that Figure 7 The electronic device shown is only an example and should not bring any limitation to the functions and scope of use of the embodiments of the present invention.

[0116] A computer-readable storage medium according to an embodiment of the present invention stores a computer program, and when the computer program is executed by a processor, any of the above-mentioned satellite deorbit control methods is implemented.

[0117] Optionally, the computer-readable storage medium may be a read-only memory (ROM), a random access memory (RAM), a compact disc (CD-ROM), a magnetic tape, a floppy disk, an optical data storage device, and the like.

[0118] In an exemplary embodiment, a computer program product or a computer program is also provided, the computer program product or the computer program comprising computer instructions, the computer instructions being stored in a computer-readable storage medium. A processor of an electronic device reads the computer instructions from the computer-readable storage medium, and the processor executes the computer instructions, so that the electronic device performs any of the above-mentioned satellite deorbit control methods.

[0119] Computer program code for performing the operations of the present invention may be written in one or more programming languages, or a combination thereof, including object-oriented programming languages, such as Java, Smalltalk, C++, and conventional procedural programming languages, such as "C" or similar programming languages. The program code may be executed entirely on the user's computer, partially on the user's computer, as a separate software package, partially on the user's computer and partially on a remote computer, or entirely on a remote computer or server. In cases involving a remote computer, the remote computer may be connected to the user's computer through any type of network, including a local area network (LAN) or a wide area network (WAN), or may be connected to an external computer (e.g., through the Internet using an Internet service provider).

[0120] It should be understood that the flow charts and block diagrams in the accompanying drawings illustrate the possible implementation architecture, functions and operations of the method and computer program product according to various embodiments of the present invention. In this regard, each box in the flow chart or block diagram can represent a module, a program segment, or a part of a code, and the module, the program segment, or a part of the code contains one or more executable instructions for realizing the specified logical function. It should also be noted that in some alternative implementations, the functions marked in the box can also occur in a sequence different from that marked in the accompanying drawings. For example, two boxes represented in succession can actually be executed substantially in parallel, and they can sometimes be executed in the opposite order, depending on the functions involved. It should also be noted that each box in the block diagram and / or flow chart, and the combination of the boxes in the block diagram and / or flow chart can be implemented with a dedicated hardware-based system that performs a specified function or operation, or can be implemented with a combination of dedicated hardware and computer instructions.

[0121] The computer-readable storage medium provided by the embodiment of the present invention may be, but is not limited to, an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, device or component, or any combination thereof. More specific examples of computer-readable storage media may include, but are not limited to: an electrical connection with one or more wires, a portable computer disk, a hard disk, a random access memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM or flash memory), an optical fiber, a portable compact disk read-only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination thereof. In the present invention, a computer-readable storage medium may be any tangible medium containing or storing a program that may be used by or in conjunction with an instruction execution system, device, or component.

[0122] The computer-readable storage medium carries one or more programs. When the one or more programs are executed by the electronic device, the electronic device executes the method shown in the above embodiment.

[0123] The above description is only a preferred embodiment of the present invention and an explanation of the technical principles used. Those skilled in the art should understand that the disclosure scope involved in the present invention is not limited to the technical solution formed by a specific combination of the above technical features, but should also cover other technical solutions formed by any combination of the above technical features or their equivalent features without departing from the above disclosed concept. For example, the above features are replaced with the technical features with similar functions disclosed in the present invention (but not limited to) by each other.

[0124] It should be noted that the terms "first", "second", etc. in the specification and claims of the present application are used to distinguish similar objects and represent the definition of a specific order or sequence. The order of use of similar objects can be interchanged where appropriate, so that the embodiments of the present application described herein can be implemented in an order other than the order shown or described.

[0125] Those skilled in the art know that the present invention can be implemented as a system, method or computer program product. Therefore, the present invention can be specifically implemented in the following forms, namely: it can be complete hardware, it can be complete software (including firmware, resident software, microcode, etc.), or it can be a combination of hardware and software, which is generally referred to as "circuit", "module" or "system" herein. In addition, in some embodiments, the present invention can also be implemented in the form of a computer program product in one or more computer-readable media, and the computer-readable medium contains computer-readable program code.

[0126] Although the embodiments of the present invention have been shown and described above, it is to be understood that the above embodiments are exemplary and are not to be construed as limitations of the present invention. A person skilled in the art may change, modify, replace and vary the above embodiments within the scope of the present invention.

Claims

1. A satellite deorbit control method, characterized in that: include: Determine the optimal initial orbital altitude of the satellite for passive deorbiting; When the active deorbit orbital altitude is not zero, the cumulative total ignition duration of the satellite transferred from the orbit when the autonomous deorbit is initiated to the orbit of the optimal passive deorbit initial orbital altitude is calculated, and according to the cumulative total ignition duration, the ignition duration of the satellite in each target operation circle within the preset duration is calculated, wherein the active deorbit orbital altitude is the height difference between the optimal passive deorbit initial orbital altitude and the altitude of the orbit when the satellite initiates the autonomous deorbit; Based on the ignition duration of the satellite in each target operation cycle within a preset duration, and under the constraint of energy balance, the duration of the solar sail driving device of the satellite facing the sun in each target operation cycle is calculated to obtain a time plan for facing the sun; According to the active deorbit orbital height, the preset ignition strategy and the solar time planning, the active deorbit scheme is used to transfer the satellite from the orbit when the autonomous deorbit is started to the orbit of the optimal passive deorbit initial orbital height, and then the passive deorbit scheme is used to control the satellite to deorbit.

2. A satellite deorbit control method according to claim 1, characterized in that: Also includes: When the active deorbit orbit height is zero, the satellite is controlled to deorbit using a passive deorbit scheme.

3. The satellite deorbit control method according to claim 1, characterized in that: Calculating the total accumulated ignition time of the satellite from the orbit at the time of starting the autonomous deorbit to the orbit of the optimal passive deorbit initial orbit altitude, including: Based on the preset ignition strategy, the optimal passive deorbit initial orbital altitude and the altitude of the satellite's orbit when autonomous deorbit is initiated, and using the Hohmann orbit change principle, the total cumulative ignition duration of the satellite's transfer from the orbit when autonomous deorbit is initiated to the orbit at the optimal passive deorbit initial orbital altitude is calculated.

4. The satellite deorbit control method according to claim 1, characterized in that: Determine the optimal initial orbital height of the satellite for passive deorbiting, including: Based on the structure, mass characteristic parameters and deorbiting time constraints of the satellite, the optimal passive deorbiting initial orbital altitude of the satellite is determined using STK software.

5. A satellite deorbit control system, characterized in that: It includes the optimal passive deorbit initial orbit height determination module, the cumulative ignition total duration calculation module, the operation circle ignition duration calculation module, the sun time planning acquisition module and the deorbit control module; The optimal passive deorbit initial orbit height determination module is used to: determine the optimal passive deorbit initial orbit height of the satellite; The cumulative ignition total duration calculation module is used to: when the active deorbit orbit altitude is not zero, calculate the cumulative ignition total duration of the satellite transferred from the orbit when the autonomous deorbit is initiated to the orbit of the optimal passive deorbit initial orbit altitude, wherein the active deorbit orbit altitude is the height difference between the optimal passive deorbit initial orbit altitude and the altitude of the orbit when the satellite initiates the autonomous deorbit; The operating circle ignition duration calculation module is used to calculate the ignition duration of each target operating circle of the satellite within a preset duration according to the accumulated total ignition duration; The solar time planning acquisition module is used to calculate the duration of the solar panel driving device of the satellite facing the sun in each target operation cycle within a preset duration and under the constraint of energy balance, so as to obtain the solar time planning; The deorbit control module is used to: according to the active deorbit orbital height, the preset ignition strategy and the sun time planning, use the active deorbit scheme to transfer the satellite from the orbit when the autonomous deorbit is started to the orbit of the optimal passive deorbit initial orbital height, and then use the passive deorbit scheme to control the satellite to deorbit.

6. A satellite deorbit control system according to claim 5, characterized in that: The control deorbit module is also used for: When the active deorbit orbit height is zero, the satellite is controlled to deorbit using a passive deorbit scheme.

7. A satellite deorbit control system according to claim 5, characterized in that: The cumulative ignition total duration calculation module is specifically used for: Based on the preset ignition strategy, the optimal passive deorbit initial orbital altitude and the altitude of the satellite's orbit when autonomous deorbit is initiated, and using the Hohmann orbit change principle, the total cumulative ignition duration of the satellite's transfer from the orbit when autonomous deorbit is initiated to the orbit at the optimal passive deorbit initial orbital altitude is calculated.

8. A satellite deorbit control system according to claim 5, characterized in that: The optimal passive deorbit initial orbit height determination module is specifically used to determine the optimal passive deorbit initial orbit height of the satellite using STK software based on the structure, mass characteristic parameters and deorbit time constraints of the satellite.

9. An electronic device, characterized in that: The invention comprises a memory, a processor and a computer program stored in the memory and executable on the processor, wherein when the processor executes the computer program, a satellite deorbit control method as claimed in any one of claims 1 to 4 is implemented.

10. A computer-readable storage medium, characterized in that: The computer-readable storage medium stores a computer program, and when the computer program is executed by a processor, the deorbit control method for a satellite as described in any one of claims 1 to 4 is implemented.

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