An off-orbit control method, system, electronic device and storage medium for a satellite

Abstract

The invention discloses a deorbit control method, system, electronic device and storage medium of a satellite, and relates to the field of satellite control technology. The method comprises: determining the optimal passive deorbit initial orbit height of the satellite; when the active deorbit orbit height is not zero, calculating the time plan for the sun; according to the active deorbit orbit height, the preset ignition strategy and the time plan for the sun, using the active deorbit scheme, transferring the satellite from the orbit when the autonomous deorbit is started to the orbit of the optimal passive deorbit initial orbit height, and then using the passive deorbit scheme to control the satellite to deorbit. The present invention adopts a combination of the active deorbit scheme and the passive deorbit scheme to control the satellite to deorbit, which can make full use of the atmospheric drag effect, reduce the consumption of active deorbit propellant, reduce the passive deorbit time, and has strong operability in engineering implementation.

Description

Background Art

[0002] The booming development of low-earth orbit (LEO) satellite communication has become one of the most remarkable competition focuses in the field of science and technology in the world today. Advanced LEO 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 400 km to 1500 km to provide broadband Internet access services to the ground and realizes global coverage through networking of multiple satellites.

[0003] According to the regulations of the Inter-Agency Space Debris Coordination Committee (IADC), a spacecraft must deorbit within 25 years after failure to reduce the number of space debris, which is the "25-year rule" for satellites. The disposal methods after the satellite's life ends vary according to the different orbits of the satellite. Generally, there are two methods. One is for satellites operating in low orbits. The satellite can be made to reduce its orbital altitude and return to the Earth. During the process of entering the Earth, the satellite will experience intense friction with the Earth's atmosphere, generating high temperatures and burning out completely. The other is for satellites operating in geostationary orbits. Since the cost of re-entering and returning to the Earth is huge, the satellite is usually pushed into a higher orbit, so that it can stay away from other normal satellites and orbit the Earth for hundreds of years.

[0004] LEO communication satellites operate in inclined or polar orbits and are often equipped with a two-axis and two-wing solar array drive assembly (SADA) to align the normal of the solar panels towards the sun, ensuring stable energy during the operation of the entire satellite. Different satellite constellation construction plans and requirements for satellite deorbiting are different. Some are fully implemented in accordance with the 25-year regulations of the IADC, while others hope to complete deorbiting in a shorter time, such as 7 years or 10 years.

[0005] Generally, satellite orbit control and deorbiting tasks are implemented using ground command control methods, which require a large amount of ground mission planning resources and measurement, operation, and control resources. In a large-scale Internet constellation, the expected method for satellite deorbiting is to achieve on-board autonomous deorbiting, but there is less research and implementation in this area currently. At the same time, there is even less implementation of autonomous deorbiting engineering that fully utilizes atmospheric drag and solar panel control strategies. Summary of the Invention

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

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

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

[0009] When the active deorbiting orbit altitude is non-zero, calculate the total cumulative ignition duration for the satellite to transfer from the orbit at the start of autonomous deorbiting to the orbit with the optimal passive deorbiting initial orbit altitude. Based on the total cumulative ignition duration, calculate the ignition duration within each target operating cycle of the satellite within a preset duration, where the active deorbiting orbit altitude is the altitude difference between the optimal passive deorbiting initial orbit altitude and the altitude of the orbit of the satellite at the start of autonomous deorbiting;

[0010] Based on the ignition duration within each target operating cycle of the satellite within a preset duration, and under the constraint of energy balance, calculate the duration of continuous sun-pointing of the satellite's solar panel drive device within each target operating cycle to obtain the sun-pointing time plan;

[0011] According to the active deorbiting orbit altitude, the preset ignition strategy, and the sun-pointing time plan, use the active deorbiting plan to transfer the satellite from the orbit at the start of autonomous deorbiting to the orbit with the optimal passive deorbiting initial orbit altitude, and then use the passive deorbiting plan to control the satellite's deorbiting.

[0012] The beneficial effects of an off-orbit control method for a satellite provided by the present invention are as follows:

[0013] When the altitude difference between the optimal passive deorbiting initial orbit altitude and the altitude of the orbit of the satellite at the start of autonomous deorbiting, i.e., the active deorbiting orbit altitude, is non-zero, adopt a combination of the active deorbiting plan and the passive deorbiting plan to control the satellite's deorbiting, which can make full use of the atmospheric drag effect, reduce the consumption of active deorbiting propellant, reduce the passive deorbiting duration, is applicable to low-orbit communication satellites carrying large-area solar panels, and has strong engineering operability.

[0014] Based on the above solution, an off-orbit control method for a satellite of the present invention can be further improved as follows.

[0015] Further, it further includes:

[0016] When the active deorbiting orbit altitude is zero, use the passive deorbiting plan to control the satellite's deorbiting.

[0017] Further, calculating the total cumulative ignition duration for the satellite to transfer from the orbit at the start of autonomous deorbiting to the orbit with the optimal passive deorbiting initial orbit altitude includes:

[0018] Based on the preset ignition strategy, the optimal passive deorbiting initial orbit altitude, and the altitude of the orbit of the satellite at the start of autonomous deorbiting, and using the Hohmann transfer principle, calculate the total cumulative ignition duration for the satellite to transfer from the orbit at the start of autonomous deorbiting to the orbit with the optimal passive deorbiting initial orbit altitude.

[0019] Further, determining the optimal passive deorbiting initial orbit altitude of the satellite includes:

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

[0021] 2) Second, the present invention also provides an orbital decay control system for a satellite, and the specific technical solution is as follows:

[0022] It includes an optimal initial orbital altitude determination module for passive orbital decay, a total cumulative ignition duration calculation module, an ignition duration calculation module for each target orbit revolution, a sun-time planning acquisition module, and an orbital decay control module;

[0023] The optimal initial orbital altitude determination module for passive orbital decay is used to: determine the optimal initial orbital altitude for the satellite's passive orbital decay;

[0024] The total cumulative ignition duration calculation module is used to: when the active orbital decay altitude is not zero, calculate the total cumulative ignition duration for the satellite to transfer from the orbit at the start of autonomous orbital decay to the orbit with the optimal initial orbital altitude for passive orbital decay, where the active orbital decay altitude is the altitude difference between the optimal initial orbital altitude for passive orbital decay and the altitude of the orbit at the start of the satellite's autonomous orbital decay;

[0025] The ignition duration calculation module for each target orbit revolution is used to: based on the total cumulative ignition duration, calculate the ignition duration within each target orbit revolution of the satellite within a preset duration;

[0026] The sun-time planning acquisition module is used to: based on the ignition duration within each target orbit revolution of the satellite within a preset duration, and under the constraint of energy balance, calculate the continuous sun-time of the satellite's solar panel drive device within each target orbit revolution to obtain the sun-time planning;

[0027] The orbital decay control module is used to: according to the active orbital decay altitude, the preset ignition strategy, and the sun-time planning, use the active orbital decay plan to transfer the satellite from the orbit at the start of autonomous orbital decay to the orbit with the optimal initial orbital altitude for passive orbital decay, and then use the passive orbital decay plan to control the satellite's orbital decay.

[0028] Based on the above solution, the orbital decay control system for a satellite of the present invention can be further improved as follows.

[0029] Furthermore, the orbital decay control module is also used to:

[0030] when the active orbital decay altitude is zero, use the passive orbital decay plan to control the satellite's orbital decay.

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

[0032] Based on a preset ignition strategy, the optimal initial orbit altitude for passive deorbiting, and the altitude of the satellite's orbit when it starts autonomous deorbiting, and using the Hohmann transfer principle, calculate the total cumulative ignition duration for the satellite to transfer from the orbit at the start of autonomous deorbiting to the orbit with the optimal initial orbit altitude for passive deorbiting.

[0033] Furthermore, the optimal initial orbit altitude determination module is specifically configured to: based on the structure, mass characteristic parameters of the satellite, and the deorbiting duration constraint, use STK software to determine the optimal initial orbit altitude for passive deorbiting of the satellite.

[0034] 3) Thirdly, the present invention also provides an electronic device. The electronic device includes a processor, the processor is coupled to a memory, and at least one computer program is stored in the memory. The at least one computer program is loaded and executed by the processor so that the electronic device implements any one of the above satellite deorbiting control methods.

[0035] 4) Fourthly, the present invention also provides a computer-readable storage medium. A computer program is stored on the computer-readable storage medium, and when the computer program is executed by a processor, it implements any one of the above satellite deorbiting control methods.

[0036] It should be noted that for the beneficial effects obtained by the technical solutions and corresponding possible implementation manners of the second to fourth aspects of the present invention, reference may be made to the technical effects of the first aspect and its corresponding possible implementation manners above, and details are not described herein again. 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 description in the embodiments of the present invention:

[0038] Figure 1 It is a schematic flowchart of a satellite deorbiting control method according to an embodiment of the present invention;

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

[0040] Figure 3 It is a schematic structural diagram of a satellite's solar panel drive device;

[0041] Figure 4 It is an ignition schematic diagram for the first target operation cycle;

[0042] Figure 5 It is an ignition schematic diagram for non-first target operation cycles;

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

[0044] Figure 7Schematic diagram of the structure of an electronic device according to an embodiment of the present invention. Detailed implementation manners

[0045] The principles and features of the present invention will be described below. The examples given are only for explaining the present invention and are not intended 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 technical problems will be described in detail below with specific embodiments. These 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 with reference to the accompanying drawings.

[0047] As Figure 1 shown, a deorbiting control method for a satellite according to an embodiment of the present invention includes the following steps:

[0048] S1. Determine the optimal initial orbit altitude for passive deorbiting of the satellite. Specifically:

[0049] Based on the structure, mass characteristic parameters, and deorbiting duration constraint of the satellite, use STK software to determine the optimal initial orbit altitude for passive deorbiting of the satellite, including the following steps:

[0050] S10. Preset the initial value of the initial orbit altitude for passive deorbiting as M, with the unit of km.

[0051] S11. Combine M, the structure, and mass characteristic parameters of the satellite, and use the satellite life analysis function of STK software to calculate. Calculate the passive deorbiting duration t of the satellite, and determine whether t is less than or equal to the deorbiting duration 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 initial orbit altitude for passive deorbiting as the optimal initial orbit altitude for passive deorbiting.

[0052] Among them, the deorbiting duration constraint is: the duration difference between the mission-required deorbiting duration of the satellite and the preset duration. Among them, the preset duration can be 1 year or can be set according to the actual situation.

[0053] S12. Let m′ = M - ΔH × N, M = 1, and use the calculated m′ as M. Execute S11 again. When the first judgment result is no, let N = N + 1, and use the calculated m′ as M. Execute S11 again until the first judgment result is yes. Determine the current m′ as the optimal initial orbit altitude for passive deorbiting of the satellite, denoted as m′′. Among them, ΔH is the iteration step size, with the unit of km.

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

[0055] S2. When the active deorbiting orbital altitude is non - zero, calculate the total cumulative ignition duration for the satellite to transfer from the orbit at the start of autonomous deorbiting to the orbit with the optimal passive deorbiting initial orbital altitude. And based on the total cumulative ignition duration, calculate the ignition duration within each target operating cycle of the satellite within a preset duration. Here, the active deorbiting orbital altitude is the altitude difference between the optimal passive deorbiting initial orbital altitude and the altitude of the orbit when the satellite starts autonomous deorbiting.

[0056] Among them, based on the optimal passive deorbiting initial orbital altitude calculated according to S1, the active deorbiting orbital altitude can be calculated as: H 0 -m″, where H 0 is the orbital altitude when the satellite receives the ground command and starts autonomous deorbiting, and m″ is the optimal passive deorbiting initial orbital altitude.

[0057] In S2, the specific implementation process of calculating the total cumulative ignition duration for the satellite to transfer from the orbit at the start of autonomous deorbiting to the orbit with the optimal passive deorbiting initial orbital altitude is as follows:

[0058] Based on the preset ignition strategy, the optimal passive deorbiting initial orbital altitude, and the altitude of the orbit when the satellite starts autonomous deorbiting, and using the Hohmann transfer principle, calculate the total cumulative ignition duration for the satellite to transfer from the orbit at the start of autonomous deorbiting to the orbit with the optimal passive deorbiting initial orbital altitude.

[0059] As Figure 2 shown in the Hohmann transfer schematic diagram, orbit A and orbit B are different orbits for the satellite's operation. Take the orbit when the satellite starts autonomous deorbiting as orbit A, and the orbit with the optimal passive deorbiting initial orbital altitude as orbit B. The altitude of orbit A is r A , and the altitude of orbit B is r B , r B =m″. Denote v A and v B as the magnitudes of the velocities required for the satellite to orbit along orbit A and orbit B respectively, and their values are: where μ = 3.9860044118×10 5 km 3 / s 2 , which is the geocentric gravitational constant. The calculation methods of v 1 and v 2 are as follows: Thus, the two velocity increments of the satellite during the Hohmann transfer are: Δv 1 =v 1 -v A , Δv 2 =v B -v 2, therefore, the total velocity increment -ΔV required for the satellite to maneuver from orbit B to orbit A is: -ΔV = k * (-Δv 1 -Δv 2 ), where k is the efficiency conversion coefficient between the engine pulse thrust model and the engine continuous low thrust model. The value range of k is 1.2 to 1.5, and k can be specifically determined by conducting multiple simulations using the STK orbit maneuver module.

[0060] Furthermore, through the formula: T all = m × ΔV / F, the total cumulative ignition duration T all is calculated. The unit is s; m is the mass of the satellite when the deorbiting mission starts, with the unit kg; F is the thrust magnitude of the satellite's electric propulsion thruster.

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

[0062] Calculate the number of orbit revolutions of the satellite within the preset duration as N all0 , and take the even number N all0 nearest to N all . Then there are N all target orbit revolutions. The ignition duration within each target orbit revolution of the satellite 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 deorbiting duration constraint. Taking a low-earth orbit communication satellite as an example, a low-earth orbit communication satellite usually operates in an orbit range with an altitude of 400 Km to 1500 Km, and one year can be reserved as the preset duration.

[0064] In S3, based on the ignition duration within each target orbit revolution of the satellite within the preset duration, and under the constraint of energy balance, calculate the duration for the satellite's solar panel drive device to face the sun within each target orbit revolution, and obtain the sun-facing time plan, which specifically includes the following steps:

[0065] According to the energy consumption on the satellite and the power generation efficiency when the solar panel has a 5-degree error facing the sun, the duration T SADA for the satellite to keep the solar panel facing the sun within each target orbit revolution can be calculated.

[0066] S4. According to the active deorbiting orbital altitude, the preset ignition strategy, and the sun-pointing time plan, use the active deorbiting plan to transfer the satellite from the orbit at the start of autonomous deorbiting to the orbit with the optimal passive deorbiting initial orbital altitude, and then use the passive deorbiting plan to control the satellite's deorbiting.

[0067] Among them, as Figure 3 shown, the satellite's solar panel drive device includes three rotating mechanisms arranged in sequence. The rotation axis of the rotating mechanism in the middle is perpendicular to the rotation axes of the rotating mechanisms on both sides. The rotating mechanisms on both sides are respectively connected to a solar panel, and a sun sensor is arranged on each solar panel. Specifically:

[0068] According to the deployment state of the satellite's solar panels, a rotating mechanism is configured for each solar panel. The satellite includes two solar panels, denoted as the first solar panel and the second solar panel respectively. The rotating mechanism configured for the first solar panel is denoted as B1, and the rotating mechanism configured for the second solar panel is denoted as B2. A rotating mechanism is arranged 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 part remaining after the satellite excludes the solar panels and the structures connecting the solar panels. Taking the centroid of the satellite body as the origin, a satellite centroid body coordinate system O-X-Y-Z is established. When both solar panels are deployed, the normal of the solar cell on the solar panel is consistent with the negative direction of the Z-axis of the satellite centroid body coordinate system. Define the position of B1 at this time as the zero position of B1, define the position of B2 at this time as the zero position of B2, and define the position of A at this time as the zero position of A. When B1, B2, and A are all in their 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 centroid 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 centroid 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 centroid body coordinate system.

[0069] Stepper motors are used inside B1, B2, and A. B1 is used to drive the first solar panel to rotate around the rotation axis of B1 through the stepper motor. B2 is used to drive the second solar panel to rotate around the rotation axis of B2 through the stepper motor. A is used to drive the first solar panel and the second solar panel to rotate around the rotation axis of A simultaneously. In this process, the right-hand rule is used to determine the rotation directions of the first solar panel and the second solar panel.

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

[0071] B1, B2, and A all support the speed - priority motion mode, position - priority motion mode, and position - holding mode. The speed - priority motion mode is: setting the angular velocity of the motion mechanism to move according to the set angular velocity magnitude and direction; the position - priority motion mode is: setting both the motion position and motion speed of the motion mechanism, 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 judged by the rotating mechanism internally according to the principle of the shortest path; the position - holding mode is that the driving mechanism keeps the current position unchanged.

[0072] Among them, A can rotate continuously by 360°, and B1 and B2 can move within the set motion range, and the motion range parameters are set comprehensively according to the Beta angle and star body occlusion. Mechanical limit mechanisms are designed outside the comprehensively determined motion range of B1 and B2.

[0073] A sun sensor is arranged on one side of the normal positive direction of the battery cells of the first solar panel and the second solar panel respectively. The sun sensor is in a pyramid configuration. The sun sensor arranged on the first solar panel is denoted as the first sun sensor, and the sun sensor arranged on the second solar panel is denoted as the second sun sensor. The layout positions are as Figure 3 shown. The field - of - view ranges of the first sun sensor and the second sun sensor are: - 90° to + 90°. The first sun sensor and the second sun sensor collect solar vector information in the solar sensor measurement coordinate system set by themselves, or the solar sensor measurement coordinate system can be reset, and an electric propulsion thruster is configured in the negative X - axis direction of the satellite centroid body coordinate system.

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

[0075] Among them, the active de - orbiting scheme can be an active de - orbiting scheme with cumulative even - numbered orbits, and the preset ignition strategy is an ignition strategy with symmetric ignition positions for adjacent orbits, as Figure 4 and Figure 5 shown.

[0076] Among them, the active de - orbiting scheme with cumulative even - numbered orbits is specifically as follows:

[0077] Denote the M'th target operating orbit as Q M′ , where M' is a positive integer and M' is not greater than N all . When the satellite receives the autonomous de - orbiting task start instruction sent from the ground, it first performs a yaw 180° attitude maneuver, and subsequent de - orbiting tasks are implemented in this attitude state. Specifically:

[0078] 1) As shown in Figure 4 , for the first target operation cycle Q 1 , within the first attitude adjustment duration T1, control A, B1, and B2 to all return to zero, ignite the electric propulsion thruster, and the ignition lasts for T Q . After that, the electric propulsion thruster shuts off. According to the orbit determination data of the satellite's current orbit, calculate the orbital period of the satellite on the current orbit, and perform a pitch - 90° attitude maneuver on the satellite's attitude. After the second attitude adjustment duration T2, the attitude adjustment is completed. Among them, T1 and T2 are estimated through mathematical simulation during the satellite scheme design, or determined through on - orbit tests before the start of the de - orbit mission.

[0079] 2) As shown in Figure 5 , for the 2nd to the Nth all target operation cycles, within the attitude adjustment duration T1, control A, B1, and B2 to all return to zero and complete the attitude maneuver. The adjusted attitude is: pitch angle is 0°, yaw angle is 180°, roll angle is 0°. The electric propulsion thruster is ignited, and the ignition lasts for T Q . After that, the electric propulsion thruster shuts off. According to the orbit determination data of the satellite's current orbit, calculate the orbital period of the satellite on the current orbit, and perform a pitch - 90° attitude maneuver on the satellite's attitude. The attitude adjustment is completed after T2 time. As shown in Figure 5 , after the attitude adjustment is completed after T2 time, after another T Sat_M′ + T Sat_M′ / 2 - T Q - T2 - T1 time, start the Q M′+1 attitude adjustment. Among them, record the orbital period corresponding to the M'th target operation cycle as T Sat_M′ .

[0080] After each attitude adjustment is completed, according to the orbit determination data of the satellite's current orbit, calculate the angular velocity and Beta angle of the current orbit. Record the angular velocity of the orbit corresponding to the M'th target operation cycle as Orbit M′ , and record the Beta angle of the orbit corresponding to the M'th target operation cycle as Beta M′ . The unit of angular velocity is deg / s, and the unit of Beta angle is deg.

[0081] According to the orbit determination data of the current orbit and the solar vector parameters, determine whether the satellite is in the earth's shadow area to obtain the second judgment result.

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

[0083] If the second judgment result is negative, control the combined movement of A, B1, and B2 to drive the solar panel to track the sun. Set the movement angular velocity of A within the M'-th target operation cycle as: 100×Orbit M′ ; The movement angular positions of B1 and B2 in the M'-th target operation cycle are: Ang_Beta M′ and -Ang_Beta M′ . If the angle information measured by the sun sensor is less than 5 degrees within the time duration T sun , and T sun < T SADA , then set the movement angular velocity of A as: Orbit M′ , until the attitude adjustment moment before the next ignition of the propulsion system, control A, B1, and B2 to return to the zero position.

[0084] After the ignition of the last target operation cycle ends, calculate the current orbit altitude based on the orbit parameters at the previous moment. If the orbit at the start of autonomous deorbiting is lower than the optimal passive deorbiting initial orbit altitude, the active deorbiting process ends and the passive deorbiting plan starts to be executed. Otherwise, continue to execute the active deorbiting plan for cumulative even cycles.

[0085] Among them, the passive deorbiting plan is specifically as follows:

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

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

[0088] The passive deorbiting plan can make full use of the atmospheric drag effect 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 further includes:

[0091] S5. When the active deorbiting orbit altitude is zero, use the passive deorbiting plan 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 can adjust the execution order of S1, S2, etc. according to the actual situation, and this is also within the protection scope of the present invention. It can be understood that in some embodiments, it may include some or all of the above embodiments.

[0093] Such as Figure 6As shown in the figure, an orbit-removal control system 200 of a satellite according to an embodiment of the present invention includes an optimal passive orbit-removal initial orbit altitude determination module 201, an accumulated ignition total duration calculation module 202, an operation cycle ignition duration calculation module 203, a sun-facing time plan acquisition module 204, and a control orbit-removal module 205;

[0094] The optimal passive orbit-removal initial orbit altitude determination module 201 is configured to: determine the optimal passive orbit-removal initial orbit altitude of the satellite;

[0095] The accumulated ignition total duration calculation module 202 is configured to: when the active orbit-removal orbit altitude is not zero, calculate the accumulated ignition total duration of the satellite from the orbit at the start of autonomous orbit-removal to the orbit of the optimal passive orbit-removal initial orbit altitude, where the active orbit-removal orbit altitude is the altitude difference between the optimal passive orbit-removal initial orbit altitude and the altitude of the orbit of the satellite at the start of autonomous orbit-removal;

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

[0097] The sun-facing time plan acquisition module 204 is configured to: based on the ignition duration within each target operation cycle of the satellite within a preset duration, and under the constraint of energy balance, calculate the continuous sun-facing duration of the satellite's solar panel drive device within each target operation cycle to obtain a sun-facing time plan;

[0098] The control orbit-removal module 205 is configured to: according to the active orbit-removal orbit altitude, a preset ignition strategy, and the sun-facing time plan, use an active orbit-removal plan to transfer the satellite from the orbit at the start of autonomous orbit-removal to the orbit of the optimal passive orbit-removal initial orbit altitude, and then use a passive orbit-removal plan to control the satellite to remove the orbit.

[0099] Optionally, in the above technical solution, the control orbit-removal module 205 is further configured to:

[0100] When the active orbit-removal orbit altitude is zero, use a passive orbit-removal plan to control the satellite to remove the orbit.

[0101] Optionally, in the above technical solution, the accumulated ignition total duration calculation module 202 is specifically configured to:

[0102] Based on a preset ignition strategy, the optimal passive orbit-removal initial orbit altitude, and the altitude of the orbit of the satellite at the start of autonomous orbit-removal, and using the Hohmann transfer principle, calculate the accumulated ignition total duration of the satellite from the orbit at the start of autonomous orbit-removal to the orbit of the optimal passive orbit-removal initial orbit altitude.

[0103] Optionally, in the above technical solution, the optimal passive deorbiting initial orbit altitude determination module 201 is specifically configured to: determine the optimal passive deorbiting initial orbit altitude of the satellite by using the STK software based on the structure, mass characteristic parameters of the satellite, and the deorbiting duration constraint.

[0104] It should be noted that the beneficial effects of the satellite deorbiting control method provided in the above embodiment are the same as those of the satellite deorbiting control system 200 provided in the above, and will not be elaborated here. In addition, when the system provided in the above embodiment realizes its functions, only the division of the above function modules is used as an example. In actual applications, the above functions can be allocated to different function modules according to needs, that is, the system can be divided into different function modules according to the actual situation 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 can be seen in the method embodiment, which will not be elaborated here.

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

[0106] In some embodiments, the satellite deorbiting control system of the present invention can be implemented in a combination of software and hardware. As an example, the satellite deorbiting 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 deorbiting 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] Among them, the modules described in the embodiments of the present invention can be implemented by software or by hardware. Among them, the name of the module does not constitute a limitation to 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 on the memory and executable on the processor. When the processor executes the computer program, the off-orbit control method of any one of the above satellites 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 off-orbit control method of the satellite shown in any embodiment of the present invention by calling the computer program.

[0109] In an alternative embodiment, an electronic device is provided, as Figure 7 shown, Figure 7 the electronic device 4000 shown includes: a processor 4001 and a memory 4003. Among them, the processor 4001 and the memory 4003 are connected, such as connected through a bus 4002. Optionally, the electronic device 4000 may further include a transceiver 4004, and the transceiver 4004 may be used for data interaction between this electronic device and other electronic devices, such as sending and / or receiving data, etc. It should be noted that in practical applications, the transceiver 4004 is not limited to one, and the structure of the electronic device 4000 does not constitute a limitation to the embodiment of the present invention.

[0110] The 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 can implement or execute various exemplary logic blocks, modules, and circuits described in connection with the disclosure of the present invention. The processor 4001 may also be a combination for implementing computing functions, such as a combination including one or more microprocessors, a combination of a DSP and a microprocessor, etc.

[0111] The bus 4002 may include a path for transmitting information between the above components. The bus 4002 may be a PCI (Peripheral Component Interconnect) bus or an EISA (Extended Industry Standard Architecture) bus, etc. The bus 4002 may be divided into an address bus, a data bus, a control bus, etc. For ease of representation, Figure 7The bus 4002 is represented only by a thick line in the figure, but it 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. It can also be an EEPROM (Electrically Erasable Programmable Read Only Memory), a CD-ROM (Compact Disc Read Only Memory), or other optical disc storage, optical disc storage (including compact discs, laser discs, optical discs, digital versatile discs, Blu-ray discs, etc.), magnetic disk storage media, or other magnetic storage devices, 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 thereto.

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

[0114] Among them, the electronic device can also be a terminal device. The terminal device can be any terminal device that can install an application and access a web page through the application, including at least one of a smart phone, a tablet computer, a notebook computer, a desktop computer, a smart speaker, a smart watch, a smart TV, and a smart vehicle-mounted device.

[0115] It should be noted that Figure 7 The electronic device shown is only an example and should not impose any limitations on the functions and usage scope of the embodiments of the present invention.

[0116] A computer-readable storage medium according to an embodiment of the present invention has a computer program stored thereon. When the computer program is executed by a processor, it implements any one of the above satellite deorbiting control methods.

[0117] Optionally, the computer-readable storage medium can be a Read-Only Memory (ROM), a Random Access Memory (RAM), a Compact Disc Read-Only Memory (CD-ROM), magnetic tape, floppy disk, and optical data storage device, etc.

[0118] In an exemplary embodiment, a computer program product or a computer program is further provided. The computer program product or the computer program includes computer instructions, and the computer instructions are 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 executes any one of the above satellite deorbiting control methods.

[0119] Computer program code for performing the operations of the present invention may be written in one or more programming languages or combinations thereof. The above programming languages include object-oriented programming languages such as Java, Smalltalk, C++, and also include conventional procedural programming languages such as the "C" language or similar programming languages. The program code may be executed entirely on the user's computer, partially on the user's computer, executed as a stand-alone software package, partially on the user's computer and partially on a remote computer, or entirely on a remote computer or server. In the case of 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 (for example, by using an Internet service provider to connect through the Internet).

[0120] It should be understood that the flowcharts and block diagrams in the accompanying drawings illustrate the possible architectures, functions, and operations of methods and computer program products according to various embodiments of the present invention. In this regard, each block in the flowchart or block diagram may represent a module, a program segment, or a part of code, and the module, program segment, or part of code contains one or more executable instructions for implementing the specified logical function. It should also be noted that in some alternative implementations, the functions marked in the blocks may occur in a different order than marked in the accompanying drawings. For example, two consecutive blocks shown may actually be executed substantially in parallel, and they may sometimes be executed in the reverse order, depending on the functions involved. It should also be noted that each block in the block diagram and / or flowchart, and the combination of blocks in the block diagram and / or flowchart, may be implemented by a dedicated hardware-based system for performing the specified functions or operations, or may be implemented by a combination of dedicated hardware and computer instructions.

[0121] The computer-readable storage medium provided by the embodiments of the present invention may be, but is not limited to, a system, device or component of electricity, magnetism, light, electromagnetic, infrared ray, or semiconductor, or any combination of the above. More specific examples of the computer-readable storage medium 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 of the above. In the present invention, the computer-readable storage medium may be any tangible medium that contains or stores a program, and the program may be used by or in combination with an instruction execution system, device or component.

[0122] The above computer-readable storage medium carries one or more programs, and when the above one or more programs are executed by the electronic device, the electronic device is caused to execute the method shown in the above embodiments.

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

[0124] It should be noted that the terms "first", "second", etc. in the description and claims of the present application are used to distinguish similar objects, and represent a limitation on a specific order or sequence. Under appropriate circumstances, the use order of similar objects may be interchanged so that the embodiments of the present application described herein can be implemented in an order other than the illustrated or described order.

[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, that is: it can be completely hardware, can also be completely software (including firmware, resident software, microcode, etc.), and can also be in the form of 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 media 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 should not be construed as limiting the present invention. Those of ordinary skill in the art can make changes, modifications, substitutions, and variations to 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 orbit 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 orbit 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 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; 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 total ignition 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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