Semi-major axis control method and device of spacecraft and electronic equipment

Abstract

The invention provides a semi-major axis control method and device of a spacecraft and electronic equipment, and relates to the technical field of spacecraft engineering. The semi-major axis control method of the spacecraft comprises the following steps: acquiring control parameters, and judging whether the number of control times in the control parameters reaches a preset number of times or not; wherein the control parameters comprise control times, a semi-major axis target control value, a control type and an initial control moment; when it is judged that the control times do not reach the preset times, a target control strategy and a target control moment are determined based on the control type, the semi-major axis target control value and the initial control moment; and at the target control moment, performing semi-major axis control on the spacecraft according to the target control strategy. According to the invention, a multi-iteration semi-major axis control calculation strategy adaptive to different control targets is provided, the solving efficiency is improved, and the calculation time is saved.

Description

Technical Field

[0001] The present invention relates to the field of spacecraft engineering technology, and in particular to a semi-major axis control method, device and electronic equipment for a spacecraft. Background Art

[0002] Spacecraft maneuvering behavior refers to the deliberate orbit or attitude changes made by a spacecraft in space to achieve specific mission objectives. At present, more than 97% of the maneuvers of spacecraft on orbit are achieved through semi-major axis control within the orbital plane. Semi-major axis control refers to the deliberate adjustment and management of the semi-major axis of the spacecraft's near-circular orbit through a series of means to achieve specific mission objectives (i.e. the above-mentioned spacecraft maneuvering behavior).

[0003] In the related art, the orbit control strategy calculation method is only applicable to one control mode or a control target parameter with a single description. For other types of control modes, the implemented input or algorithm needs to be changed to adapt to the new mode, which is complex in calculation and low in efficiency. Summary of the invention

[0004] In view of this, the embodiments of the present disclosure propose a semi-major axis control method, device and electronic equipment for a spacecraft to solve the problem that the related technology is only applicable to one control mode or has a control target parameter with a single description, and the input or algorithm for implementation of other types of control modes needs to be changed to adapt to the new mode, resulting in complex calculations and low efficiency.

[0005] According to a first aspect of the present disclosure, a method for controlling a semi-major axis of a spacecraft is provided, comprising: obtaining control parameters, and determining whether the number of controls in the control parameters reaches a preset number; wherein the control parameters include the number of controls, a semi-major axis target control value, a control type, and an initial control moment; when it is determined that the number of controls does not reach the preset number, determining a target control strategy and a target control moment based on the control type, the semi-major axis target control value, and the initial control moment; at the target control moment, performing semi-major axis control of the spacecraft with the target control strategy.

[0006] According to a second aspect of the present disclosure, a semi-major axis control device for a spacecraft is provided, comprising: an acquisition module, used to acquire control parameters, and determine whether the number of controls in the control parameters reaches a preset number; wherein the control parameters include the number of controls, a semi-major axis target control value, a control type and an initial control moment; a processing module, used to determine a target control strategy and a target control moment based on the control type, the semi-major axis target control value and the initial control moment when it is determined that the number of controls does not reach the preset number; and a control module, used to perform semi-major axis control of the spacecraft with the target control strategy at the target control moment.

[0007] According to a third aspect of the present disclosure, there is provided an electronic device, comprising: a processor; and a memory storing a program, wherein the program comprises instructions which, when executed by the processor, cause the processor to execute the above-mentioned semi-major axis control method of the spacecraft.

[0008] According to a fourth aspect of the present disclosure, a non-transitory computer-readable storage medium storing computer instructions is provided, wherein the computer instructions are used to enable the computer to execute the semi-major axis control method of the above-mentioned spacecraft.

[0009] The exemplary embodiments of the present disclosure may have some or all of the following beneficial effects:

[0010] In the semi-major axis control method of the spacecraft provided in the example implementation mode of the present disclosure, control parameters are obtained, and it is determined whether the control times in the control parameters have reached the preset times; wherein the control parameters include the control times, the semi-major axis target control value, the control type and the initial control time; when it is determined that the control times have not reached the preset times, the target control strategy and the target control time are determined based on the control type, the semi-major axis target control value and the initial control time; at the target control time, the semi-major axis of the spacecraft is controlled with the target control strategy. When the current control times have not reached the preset times, the embodiment of the present disclosure can implement semi-major axis control for different control targets with different control strategies based on the control parameters, and provides multiple iterative semi-major axis control calculation strategies that are adapted to various control modes and different descriptions of control targets, with a simple implementation process, improved solution efficiency, and saved calculation time. BRIEF DESCRIPTION OF THE DRAWINGS

[0011] Further details, features and advantages of the present disclosure are disclosed in the following description of exemplary embodiments in conjunction with the accompanying drawings, in which:

[0012] Figure 1 A flow chart of a semi-major axis control method of a spacecraft according to an exemplary embodiment of the present disclosure is shown;

[0013] Figure 2 A schematic block diagram of a semi-major axis control device of a spacecraft according to an exemplary embodiment of the present disclosure is shown;

[0014] Figure 3 A structural block diagram of an exemplary electronic device that can be used to implement the embodiments of the present disclosure is shown. DETAILED DESCRIPTION

[0015] Embodiments of the present disclosure will be described in more detail below with reference to the accompanying drawings. Although certain embodiments of the present disclosure are shown in the accompanying drawings, it should be understood that the present disclosure can be implemented in various forms and should not be construed as being limited to the embodiments described herein, which are instead provided for a more thorough and complete understanding of the present disclosure. It should be understood that the drawings and embodiments of the present disclosure are only for exemplary purposes and are not intended to limit the scope of protection of the present disclosure.

[0016] It should be understood that the various steps described in the method embodiments of the present disclosure may be performed in different orders and / or in parallel. In addition, the method embodiments may include additional steps and / or omit the steps shown. The scope of the present disclosure is not limited in this respect.

[0017] The term "including" and its variations used in this document are open inclusions, that is, "including but not limited to". The term "based on" means "based at least in part on". The term "one embodiment" means "at least one embodiment"; the term "another embodiment" means "at least one other embodiment"; the term "some embodiments" means "at least some embodiments". Relevant definitions of other terms will be given in the description below. It should be noted that the concepts of "first", "second", etc. mentioned in this disclosure are only used to distinguish different devices, modules or units, and are not used to limit the order or interdependence of the functions performed by these devices, modules or units.

[0018] It should be noted that the modifications of "one" and "plurality" mentioned in the present disclosure are illustrative rather than restrictive, and those skilled in the art should understand that unless otherwise clearly indicated in the context, it should be understood as "one or more".

[0019] The names of the messages or information exchanged between multiple devices in the embodiments of the present disclosure are only used for illustrative purposes and are not used to limit the scope of these messages or information.

[0020] The embodiments of the present disclosure provide a method, device, electronic device and storage medium for controlling the semi-major axis of a spacecraft. The scheme of the present disclosure is described below with reference to the accompanying drawings.

[0021] The present disclosure first provides a method for controlling the semi-major axis of a spacecraft. Figure 1 As shown, the semi-major axis control method of the spacecraft specifically includes the following steps:

[0022] S110: Acquire control parameters, and determine whether the control times in the control parameters reach a preset number; wherein the control parameters include the control times, the semi-major axis target control value, the control type, and the initial control time;

[0023] S120: When it is determined that the number of control times does not reach the preset number, determining a target control strategy and a target control time based on the control type, the semi-major axis target control value and the initial control time;

[0024] S130: At the target control moment, the spacecraft is controlled along its semi-major axis using the target control strategy.

[0025] In the semi-major axis control method of the spacecraft provided in the example implementation mode of the present disclosure, control parameters are obtained, and it is determined whether the control times in the control parameters have reached the preset times; wherein the control parameters include the control times, the semi-major axis target control value, the control type and the initial control time; when it is determined that the control times have not reached the preset times, the target control strategy and the target control time are determined based on the control type, the semi-major axis target control value and the initial control time; at the target control time, the semi-major axis of the spacecraft is controlled with the target control strategy. When the current control times have not reached the preset times, the embodiment of the present disclosure can implement semi-major axis control for different control targets with different control strategies based on the control parameters, and provides multiple iterative semi-major axis control calculation strategies that are adapted to various control modes and different descriptions of control targets, with a simple implementation process, improved solution efficiency, and saved calculation time.

[0026] Next, in another embodiment, the above steps are described in more detail.

[0027] In step S110, the control parameters are obtained, and it is determined whether the control times in the control parameters reach the preset times; wherein the control parameters include the control times, the semi-major axis target control value, the control type and the initial control time.

[0028] In the embodiment of the present application, the above control parameters are control information for performing semi-major axis control on the spacecraft to achieve maneuvering behavior of the spacecraft.

[0029] In the embodiments of the present application, the above-mentioned spacecraft refers to various types of aircraft that operate in space according to the laws of celestial mechanics and perform specific tasks such as exploring, developing, and utilizing space and celestial bodies. For example, the spacecraft may be a space vehicle such as a satellite, a probe, or a rocket.

[0030] In the embodiment of the present application, the semi-major axis is half the length of the major axis of the orbit in the near-circular orbit of the spacecraft. The semi-major axis intuitively describes the size of the orbit and determines the total energy of the spacecraft in the orbit. Exemplarily, the semi-major axis control can intentionally adjust and manage the semi-major axis of the near-circular orbit of the spacecraft through a series of means to achieve specific mission objectives (i.e., the above-mentioned spacecraft maneuvering behavior).

[0031] In an embodiment of the present application, the above-mentioned spacecraft maneuvering behavior refers to the conscious orbit or attitude change of the spacecraft in space in order to achieve a specific mission goal. Exemplarily, the above-mentioned spacecraft maneuvering behavior may include orbital maneuvers and attitude maneuvers. Among them, the above-mentioned orbital maneuvers refer to changing the orbital parameters of the spacecraft, such as orbital altitude, orbital inclination, eccentricity, etc. Exemplarily, the spacecraft can be transferred from the launch orbit to the working orbit through the above-mentioned orbital maneuvers. The above-mentioned attitude maneuvers refer to changing the direction of the spacecraft, and the attitude of the spacecraft is crucial to the execution of its mission. Exemplarily, a satellite can use the above-mentioned attitude maneuvers to aim its camera at a specific area and maintain a stable attitude to ensure the quality of the captured image. The semi-major axis control method of the spacecraft provided in the embodiment of the present application realizes the above-mentioned spacecraft maneuvering behavior through an iterative semi-major axis control calculation strategy, optimizes the calculation process of the control strategy, and improves the calculation efficiency. In addition, the semi-major axis control method of the spacecraft provided in the embodiment of the present application can also adapt to a variety of control modes and different description methods of control targets, thereby improving the flexibility of the control strategy.

[0032] In the embodiment of the present application, the above-mentioned control parameters include the number of controls, the semi-major axis target control value, the control type and the initial control time.

[0033] In the embodiment of the present application, the above control times are used to indicate the current control number. For example, if the control times id is 2, the current control is the second control among multiple iterative semi-major axis controls.

[0034] In the embodiment of the present application, the above-mentioned semi-major axis target control value is the target value to be achieved by the above-mentioned semi-major axis control. Exemplarily, the above-mentioned semi-major axis target control value can be a semi-major axis target value or a semi-major axis change amount target value. Among them, the above-mentioned semi-major axis target value is the theoretical value of the semi-major axis when the above-mentioned semi-major axis control ends; the above-mentioned semi-major axis change amount target value is the theoretical value of the increase / decrease of the semi-major axis compared with before the control when the above-mentioned semi-major axis control ends.

[0035] Exemplarily, the above control types can be divided into designated time control types and non-designated time control types. Among them, the designated time control type refers to the semi-major axis control of the spacecraft at the designated time; the non-designated time control types include apogee control type, perigee control type, ascending node control type and descending node control type. Specifically, the apogee control type performs semi-major axis control on the spacecraft at the first apogee after the control start time is allowed; the perigee control type performs semi-major axis control on the spacecraft at the first perigee after the control start time is allowed; the ascending node control type performs semi-major axis control on the spacecraft at the first ascending node after the control start time is allowed; the descending node control type performs semi-major axis control on the spacecraft at the first descending node after the control start time is allowed, respectively achieving the predetermined orbital maneuvering goals.

[0036] In the embodiment of the present application, the apogee refers to the point on the near-circular orbit where the spacecraft is farthest from the central celestial body (such as the earth); the perigee is the point where the spacecraft is closest to the central celestial body; the ascending node refers to the intersection of the celestial body with the reference plane when it moves from south to north along the orbit. In the orbit of the celestial body centered on the earth, the earth's equatorial plane is often selected as the reference plane, that is, the intersection of the orbital plane and the earth's equatorial plane when the satellite moves from south to north; the descending node is another intersection of the orbital plane and the equatorial plane, that is, the intersection when the satellite moves from north to south.

[0037] In the embodiment of the present application, the initial control moment is used to determine the start moment of the semi-major axis control. For example, in the specified moment control type, the initial control moment is the specified moment; in the non-specified moment control type, the initial control moment is the moment when the control is allowed to start and reach the target point.

[0038] In the embodiment of the present application, during the above semi-major axis control process, the data structure of a single control parameter may be as follows:

[0039] Struct control input parameters

[0040] {

[0041] id controls the number of times;

[0042] The target semi-major axis controls the target value;

[0043] type controls the type;

[0044] moment allows to control the moment;

[0045] }

[0046] Among them, "id" is the control number, indicating which control this is; "type" is the above control type, which can be filled in with 5 types, namely, 0-moment is the specified moment, 1-moment is the moment when control is allowed to start and find the apogee, 2-moment is the moment when control is allowed to start and find the perigee, 3-moment is the moment when control is allowed to start and find the ascending node, and 4-moment is the moment when control is allowed to start and find the descending node; "target" is the above semi-major axis target control value, which can be filled in with 2 types, namely the above semi-major axis target value and the semi-major axis change target value; "moment" is the above initial control moment, which is an absolute time expressed as year, month, day, hour, minute, second, and millisecond.

[0047] In the embodiment of the present application, the above-mentioned preset number is used to specify the number of iterations of the above-mentioned multiple iterative semi-major axis control calculation strategy. For example, if the preset number is 3, the multiple iterative semi-major axis control performs 3 controls. In this process, after the previous semi-major axis control is completed, the mass of the spacecraft is updated to the original mass minus the fuel mass, and the state of the spacecraft after the control is used as the initial state of the next semi-major axis control calculation.

[0048] In an embodiment of the present application, when the control parameters of a single control are obtained, it is determined whether the number of controls for this control has reached the above-mentioned preset number. If it has reached it, the above-mentioned multiple iterative semi-major axis control processes are stopped; if it has not reached it, the semi-major axis control of the spacecraft is implemented based on the above-mentioned control parameters.

[0049] In step S120, when it is determined that the number of control times does not reach the preset number, the target control strategy and the target control time are determined based on the control type, the semi-major axis target control value and the initial control time.

[0050] In the embodiment of the present application, the above-mentioned target control time is the start time of the semi-major axis control.

[0051] In an embodiment of the present application, when the above-mentioned control type is a specified time control type, the target control time is the above-mentioned initial control time; when the above-mentioned control type is an apogee control type, the target control time is the time when the spacecraft reaches the next apogee after the initial control time; when the above-mentioned control type is a perigee control type, the target control time is the time when the spacecraft reaches the next perigee after the initial control time; when the above-mentioned control type is an ascending node control type, the target control time is the time when the spacecraft reaches the next ascending node after the initial control time; when the above-mentioned control type is a descending node control type, the target control time is the time when the spacecraft reaches the next descending node after the initial control time.

[0052] In the embodiment of the present application, the above-mentioned target control strategy is a strategy for implementing semi-major axis control on the spacecraft. Exemplarily, the target control strategy may include a pulse velocity variation and the above-mentioned target control moment. Specifically, the motion state of the spacecraft after control may be determined based on the above-mentioned pulse velocity variation at the target control moment. Among them, the above-mentioned pulse velocity variation refers to the instantaneous velocity variation caused by the operation such as engine ignition during the orbit control process of the spacecraft. It is a key parameter in aerospace dynamics and directly determines the changes in the spacecraft orbit, such as changes in orbit shape, orbit altitude and orbit plane.

[0053] In an embodiment of the present application, the target control strategy and target control moment are determined based on the control type, the semi-major axis target control value and the initial control moment. Exemplarily, the target control moment can be implemented as follows: the target control moment is determined based on the control type and the initial control moment, and the pre-control motion state and the pre-control semi-major axis of the spacecraft corresponding to the target control moment are obtained; if the semi-major axis target control value is the semi-major axis variation target value, the corresponding pulse velocity variation value is determined based on the semi-major axis variation target value and the semi-major axis control rate, that is, the target control strategy corresponding to the control moment is obtained; if the semi-major axis target control value is the semi-major axis target value, the difference between the semi-major axis target value and the semi-major axis before control is determined, and the corresponding pulse velocity variation value is determined based on the difference and the semi-major axis control rate, that is, the target control strategy corresponding to the control moment is obtained.

[0054] In an embodiment of the present application, the target control moment is determined based on the control type and the initial control moment. Exemplarily, it can be implemented as follows: when the control type is a specified moment control type, the initial control moment is used as the target control moment; when the control type is a non-specified moment control type, the angular velocity corresponding to the spacecraft at the initial control moment is obtained, and the target control moment is determined based on the angular velocity and the initial control moment.

[0055] For example, when the value of "type" in the above control parameters is 0, the first type of control strategy calculation (that is, the above specified moment control type) is entered, and the initial control moment in the control parameters is directly used as the target control moment. In this case, the spacecraft can be extrapolated to the control moment represented by "moment" with high precision and full model orbit, and the motion state of the spacecraft at this time is taken as the motion state before control, and the semi-major axis of the flat root at this time is taken as the semi-major axis of the flat root before control.

[0056] When the value of "type" in the above control parameters is not 0, the second type of control strategy calculation (that is, the above non-specified time control type) is entered. At this time, the above target control time can be determined by the following process: obtain the angular velocity corresponding to the spacecraft at the initial control time, and determine the target control time based on the angular velocity and the initial control time. At this time, the spacecraft can be extrapolated to the allowed control start time represented by "moment" with high precision and full model orbit, and the angular velocity, mean root perigee argument and mean anomaly angle at this moment can be obtained.

[0057] Furthermore, if the above-mentioned non-specified time control type is the apogee control type (that is, the value of "type" is 1), the above-mentioned process of determining the target control time can be implemented as follows: obtain the first angle of the spacecraft to the first apogee after the initial control time, and determine the first time required for the spacecraft to reach the first apogee after the initial control time based on the first angle and angular velocity; determine the target control time based on the initial control time and the first time.

[0058] Specifically, if the value of "type" is 1, since the mean anomaly angle is 180° when the spacecraft is at the apogee, and the mean anomaly angle variation trend of the spacecraft is 0°~360°, considering that the spacecraft rotates uniformly, the angle between the spacecraft and the next apogee (that is, the first angle mentioned above) can be recorded, and the time taken to reach the next apogee (that is, the first duration mentioned above) can be obtained based on the first angle and angular velocity; the first duration is superimposed on the start time of allowed control represented by "moment" to obtain the target control moment of this control; the spacecraft is extrapolated to the above-mentioned target control moment with high precision and full model orbit, and the motion state of the spacecraft at this time is taken as the motion state before control, and the mean root semi-major axis at this time is taken as the mean root semi-major axis before control.

[0059] If the above-mentioned non-specified time control type is the perigee control type (that is, the value of "type" is 2), the above-mentioned process of determining the target control time can be implemented as follows: obtain the second angle of the spacecraft to the first perigee after the initial control time, and determine the second time required for the spacecraft to reach the first perigee after the initial control time based on the second angle and angular velocity; determine the target control time based on the initial control time and the second time.

[0060] Specifically, if the value of "type" is 2, since the mean anomaly angle is 360° when the spacecraft is at perigee, and the mean anomaly angle variation trend of the spacecraft is 0°~360°, considering that the spacecraft rotates uniformly, the angle between the spacecraft and the next perigee (i.e. the second angle mentioned above) can be recorded, and the time taken to reach the next perigee (i.e. the second duration mentioned above) can be obtained based on the second angle and angular velocity; the second duration is superimposed on the start time of allowed control represented by "moment" to obtain the target control moment of this control; the spacecraft is extrapolated to the above-mentioned target control moment with high precision and full model orbit, and the motion state of the spacecraft at this time is taken as the motion state before control, and the mean root semi-major axis at this time is taken as the mean root semi-major axis before control.

[0061] If the above-mentioned non-specified time control type is the ascending node control type (that is, the value of "type" is 3), the above-mentioned process of determining the target control time can be implemented as follows: obtain the third angle of the spacecraft from the first ascending node after the initial control time, and determine the third time required for the spacecraft to reach the first ascending node after the initial control time based on the third angle and angular velocity; determine the target control time based on the initial control time and the third time.

[0062] Specifically, if the value of "type" is 3, since the latitude argument of the spacecraft when it is at the ascending node is 360° (the perigee argument plus the mean anomaly angle), and the latitude argument changing trend of the spacecraft is 0°~360°, considering that the spacecraft rotates uniformly, the angle of the spacecraft from the next ascending node (that is, the third angle mentioned above) can be recorded, and the time taken to reach the next ascending node (that is, the third duration mentioned above) can be obtained based on the third angle and angular velocity; the third duration is superimposed on the start time of allowed control represented by "moment" to obtain the target control time of this control; the spacecraft is extrapolated to the target control time with high precision and full model orbit, and the motion state of the spacecraft at this time is taken as the motion state before control, and the semi-major axis of the root mean square at this time is taken as the semi-major axis of the root mean square before control.

[0063] If the above-mentioned non-specified time control type is the descending node control type (that is, the value of "type" is 4), the above-mentioned process of determining the target control time can be implemented as follows: obtain the fourth angle of the spacecraft to the first descending node after the initial control time, and determine the fourth time required for the spacecraft to reach the first descending node after the initial control time based on the fourth angle and angular velocity; determine the target control time based on the initial control time and the fourth time.

[0064] Specifically, if the value of "type" is 4, since the latitude argument of the spacecraft is 180° when it is at the descending node, and the latitude argument changing trend of the spacecraft is 0°~360°, considering that the spacecraft rotates uniformly, the angle of the spacecraft from the next descending node (that is, the fourth angle mentioned above) can be recorded, and the time taken to reach the next descending node (that is, the fourth time duration mentioned above) can be obtained based on the fourth angle and angular velocity; the fourth time duration is superimposed on the allowed control start time represented by "moment" to obtain the target control time of this control; the spacecraft is extrapolated to the control time with high precision and full model orbit to the control time, and the motion state of the spacecraft at this time is taken as the motion state before control, and the semi-major axis of the flat root at this time is taken as the semi-major axis of the flat root before control.

[0065] In the embodiment of the present application, after determining the target control moment through the above process and obtaining the pre-control motion state and pre-control semi-major axis corresponding to the spacecraft at the target control moment, the target control strategy can be further determined based on the above semi-major axis target control value and the semi-major axis control rate, wherein the semi-major axis control rate is a physical quantity used to describe the relationship between the change of the semi-major axis and the factors causing such change (mainly the pulse velocity change). In aerospace orbital dynamics, the semi-major axis control rate reflects the efficiency and regularity of changing the semi-major axis of the orbit by applying a certain velocity increment or decrement.

[0066] Exemplarily, when the above-mentioned semi-major axis target control value is the target value of the semi-major axis change (that is, the above-mentioned "target" is the target semi-major axis increment / decrement), the above-mentioned process of determining the target control strategy can be implemented as follows: bring in the target semi-major axis increment / decrement, perform iterative calculation using the shooting method according to the semi-major axis control rate, and obtain the corresponding pulse velocity increment / decrement on the premise of meeting the control target error threshold, superimpose the pulse velocity increment on the above-mentioned motion state before control, determine the motion state after control, convert the motion state after control into a classical orbital root representation, obtain the flat root semi-major axis, and make a difference between the flat root semi-major axis and the flat root semi-major axis before control, which is the target semi-major axis increment / decrement that meets the control target, that is, the control target is achieved; record the pulse velocity increment / decrement and control time as the control strategy, and record the motion state after control; use the fuel mass obtained based on the relationship between pulse velocity and fuel consumption as the initial condition for the next control.

[0067] When the above-mentioned semi-major axis target control value is the semi-major axis target value (that is, the above-mentioned "target" is the target semi-major axis), the above-mentioned process of determining the target control strategy can be implemented as follows: subtract the semi-major axis target value from the above-mentioned flat root semi-major axis before control, and bring the obtained difference into the semi-major axis control rate for iterative calculation by the shooting method, and on the premise of meeting the control target error threshold, obtain the corresponding pulse velocity increment / decrement value, superimpose this pulse velocity increment on the motion state before control, and obtain the motion state after control, and convert the motion state after control into the classical orbit root representation to obtain the flat root semi-major axis, which is the target semi-major axis that meets the control target, that is, the control target is achieved; record the pulse velocity increment / decrement value and control time as the control strategy, and record the motion state after control; and use the fuel consumption mass obtained according to the relationship between pulse velocity and fuel consumption as the initial condition for the next control.

[0068] In step S130, at the target control moment, the spacecraft is controlled along the semi-major axis using the target control strategy.

[0069] In an embodiment of the present application, when the current moment is taken as the target control moment, the spacecraft is controlled along its semi-major axis based on the target control strategy determined in the above steps to achieve corresponding spacecraft maneuvering behavior, thereby completing the specified task.

[0070] In the embodiment of the present application, after the semi-major axis control of the current control times is completed, the control parameters can be calculated cyclically. Specifically, it can be implemented as follows: the mass of the spacecraft needs to be updated to the original mass minus the fuel consumption mass, and the state of the spacecraft after control is used as the initial state of the next calculation.

[0071] Below, in a specific embodiment, taking the above-mentioned multiple iterative semi-major axis control calculation strategy as a three-iteration control for a satellite as an example, the semi-major axis control method of the above-mentioned spacecraft is described in detail:

[0072] In the embodiment of the present application, the initial state of the above satellite is as follows:

[0073]

[0074] In addition, the control parameters of the above three iterative controls are:

[0075] First control input parameters

[0076] Second control input parameters

[0077] Third control input parameters

[0078] After obtaining the control parameters of each control number, further, the target control time and target control strategy are determined based on the control parameters, which is specifically implemented as follows:

[0079] S1: The first control is the specified time control, that is, "moment = 2024 / 05 / 23 21:07:01.000" is implemented, and the control target is the control half major axis value "target = 6900000.000", and the calculation results are as follows:

[0080] Satellite motion status before control (the following are all average roots):

[0081]

[0082] Satellite motion status after control:

[0083]

[0084] Within the allowable error range of the target semi-major axis (0.05m), it can be seen from the above that the control target of 6900000.000m was achieved at the specified time.

[0085] S2: The second control is the apogee control, that is, the control is implemented at the first apogee after "monment=2024 / 05 / 23 21:54:18.000", and the control target is the semi-major axis increment value after control is "target=3000.000", and the calculation results are as follows:

[0086] Satellite motion status before control (the following are all average roots):

[0087]

[0088] Satellite motion status after control:

[0089]

[0090] Within the allowable error range of the target semi-major axis (0.05m), it can be seen from the above that the control target of increasing the semi-major axis by 3000.000m was achieved at the specified time.

[0091] S3: The third control is the descending node control, that is, the control is implemented at the first apogee after "moment=2024 / 05 / 23 22:00:09.000", and the control target is the semi-major axis increment value after the control is "target=-5000.000", and the calculation results are as follows:

[0092] Satellite motion status before control (the following are all average roots):

[0093]

[0094] Satellite motion status after control:

[0095]

[0096] Within the allowable error range of the target semi-major axis (0.05 m), it can be seen from the above that the control target of reducing the semi-major axis by 5000.000 m was achieved at the specified time.

[0097] In the above process, the embodiment of the present application calculates the control parameters cyclically, and uses the last post-control state as the next control initial state for a cycle, and can derive the following control strategy that meets the control objectives of the example:

[0098] First control strategy parameters

[0099] Second control strategy parameters

[0100] The third control strategy parameters

[0101] In the semi-major axis control method of the spacecraft provided in the example implementation mode of the present disclosure, control parameters are obtained, and it is determined whether the control times in the control parameters have reached the preset times; wherein the control parameters include the control times, the semi-major axis target control value, the control type and the initial control time; when it is determined that the control times have not reached the preset times, the target control strategy and the target control time are determined based on the control type, the semi-major axis target control value and the initial control time; at the target control time, the semi-major axis of the spacecraft is controlled with the target control strategy. When the current control times have not reached the preset times, the embodiment of the present disclosure can implement semi-major axis control for different control targets with different control strategies based on the control parameters, and provides multiple iterative semi-major axis control calculation strategies that are adapted to various control modes and different descriptions of control targets, with a simple implementation process, improved solution efficiency, and saved calculation time.

[0102] The above mainly introduces the solution provided by the embodiment of the present invention from the perspective of the method. In order to achieve the above functions, it includes hardware structures and / or software modules corresponding to the execution of each function. Those skilled in the art should easily realize that, in combination with the units and algorithm steps of each example described in the embodiments disclosed herein, the present invention can be implemented in the form of hardware or a combination of hardware and computer software. Whether a function is executed in the form of hardware or computer software driving hardware depends on the specific application and design constraints of the technical solution. Professional and technical personnel can use different methods to implement the described functions for each specific application, but such implementation should not be considered to exceed the scope of the present invention.

[0103] The embodiment of the present invention can divide the semi-major axis control device of the spacecraft into functional modules according to the above method example. For example, each functional module can be divided according to each function, or two or more functions can be integrated into one processing module. The above integrated module can be implemented in the form of hardware or in the form of software functional modules. It should be noted that the division of modules in the embodiment of the present invention is schematic and is only a logical functional division. There may be other division methods in actual implementation.

[0104] Accordingly, the embodiment of the present disclosure also provides a semi-major axis control device for a spacecraft. Figure 2 As shown, the semi-major axis control device 200 of the spacecraft may include an acquisition module 210, a processing module 220 and a control module 230, wherein:

[0105] The acquisition module 210 can be used to acquire control parameters and determine whether the control times in the control parameters have reached a preset number; wherein the control parameters include the control times, the semi-major axis target control value, the control type and the initial control time;

[0106] The processing module 220 may be used to determine a target control strategy and a target control time based on the control type, the semi-major axis target control value and the initial control time when it is determined that the control number does not reach the preset number;

[0107] The control module 230 may be used to perform semi-major axis control of the spacecraft using a target control strategy at the target control moment.

[0108] In an embodiment of the present application, the above-mentioned control types include a designated time control type and a non-designated time control type, and the non-designated time control type includes an apogee control type, a perigee control type, an ascending node control type and a descending node control type; among which, the apogee control type performs semi-major axis control on the spacecraft at the first apogee after the initial control time; the perigee control type performs semi-major axis control on the spacecraft at the first perigee after the initial control time; the ascending node control type performs semi-major axis control on the spacecraft at the first ascending node after the initial control time; and the descending node control type performs semi-major axis control on the spacecraft at the first descending node after the initial control time.

[0109] In an embodiment of the present application, the above-mentioned semi-major axis target control value includes a semi-major axis target value and a semi-major axis variation target value; the above-mentioned processing module is specifically used to: determine the target control moment based on the control type and the initial control moment, and obtain the pre-control motion state and pre-control semi-major axis corresponding to the spacecraft at the target control moment; if the semi-major axis target control value is the semi-major axis variation target value, determine the corresponding pulse velocity variation value based on the semi-major axis variation target value and the semi-major axis control rate, and determine the target control strategy based on the pulse velocity variation value, the pre-control motion state and the pre-control semi-major axis; if the semi-major axis target control value is the semi-major axis target value, determine the difference between the semi-major axis target value and the pre-control semi-major axis, determine the corresponding pulse velocity variation value based on the difference and the semi-major axis control rate, and determine the target control strategy based on the pulse velocity variation value, the pre-control motion state and the pre-control semi-major axis.

[0110] In an embodiment of the present application, the above-mentioned processing module is specifically used for: when the control type is a specified time control type, taking the initial control time as the target control time; when the control type is a non-specified time control type, obtaining the angular velocity corresponding to the spacecraft at the initial control time, and determining the target control time based on the angular velocity and the initial control time.

[0111] In an embodiment of the present application, the above-mentioned processing module is specifically used for: when the non-specified time control type is the apogee control type, obtaining the first angle of the spacecraft to the first apogee after the initial control moment, and determining the first time required for the spacecraft to reach the first apogee after the initial control moment based on the first angle and angular velocity; determining the target control moment based on the initial control moment and the first time.

[0112] In an embodiment of the present application, the above-mentioned processing module is specifically used for: when the non-specified time control type is the perigee control type, obtaining the second angle of the spacecraft from the first perigee after the initial control time, and determining the second time required for the spacecraft to reach the first perigee after the initial control time based on the second angle and angular velocity; determining the target control time based on the initial control time and the second time.

[0113] In an embodiment of the present application, the above-mentioned processing module is specifically used for: when the non-specified time control type is the ascending node control type, obtaining the third angle of the spacecraft from the first ascending node after the initial control time, and determining the third time required for the spacecraft to reach the first ascending node after the initial control time based on the third angle and angular velocity; determining the target control time based on the initial control time and the third time.

[0114] In an embodiment of the present application, the above-mentioned processing module is specifically used for: when the non-specified time control type is a descending node control type, obtaining the fourth angle of the spacecraft from the first descending node after the initial control time, and determining the fourth time required for the spacecraft to reach the first descending node after the initial control time based on the fourth angle and angular velocity; determining the target control time based on the initial control time and the fourth time.

[0115] In addition, the specific implementation details of the semi-major axis control device of the above-mentioned spacecraft have been described in detail in the corresponding position of the semi-major axis control method of the spacecraft, so they will not be repeated here.

[0116] Figure 3 Schematic diagram of the structure of an electronic device in the embodiment of the present disclosure. Figure 3 , which shows a structural schematic diagram of an electronic device 300 suitable for implementing the embodiments of the present disclosure. Figure 3 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 disclosure.

[0117] like Figure 3 As shown, the electronic device 300 may include a processing device (e.g., a central processing unit, a graphics processor, etc.) 301, which can perform various appropriate actions and processes according to a program stored in a read-only memory (ROM) 302 or a program loaded from a storage device 308 to a random access memory (RAM) 303 to implement the semi-major axis control method of the spacecraft of the embodiment described in the present disclosure. Various programs and data required for the operation of the electronic device 300 are also stored in the RAM 303. The processing device 301, the ROM 302, and the RAM 303 are connected to each other via a bus 304. An input / output (I / O) interface 305 is also connected to the bus 304.

[0118] Typically, the following devices may be connected to the I / O interface 305: input devices 306 including, for example, a touch screen, a touch pad, a keyboard, a mouse, a camera, a microphone, an accelerometer, a gyroscope, etc.; output devices 307 including, for example, a liquid crystal display (LCD), a speaker, a vibrator, etc.; storage devices 308 including, for example, a magnetic tape, a hard disk, etc.; and communication devices 309. The communication devices 309 may allow the electronic device 300 to communicate with other devices wirelessly or by wire to exchange data. Although Figure 3 The electronic device 300 is shown with various devices, but it should be understood that it is not required to implement or possess all the devices shown. More or fewer devices may be implemented or possessed instead.

[0119] In particular, according to an embodiment of the present disclosure, the process described above with reference to the flowchart can be implemented as a computer software program. For example, an embodiment of the present disclosure includes a computer program product, which includes a computer program carried on a non-transitory computer-readable medium, and the computer program contains program code for executing the method shown in the flowchart, thereby implementing the semi-major axis control method of the spacecraft as described above. In such an embodiment, the computer program can be downloaded and installed from the network through the communication device 309, or installed from the storage device 308, or installed from the ROM 302. When the computer program is executed by the processing device 301, the above-mentioned functions defined in the method of the embodiment of the present disclosure are executed.

[0120] It should be noted that the computer-readable medium disclosed above may be a computer-readable signal medium or a computer-readable storage medium or any combination of the above two. The computer-readable storage medium may be, for example, but not limited to, an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, device or device, or any combination of the above. 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 of the above. In the present disclosure, a computer-readable storage medium may be any tangible medium containing or storing a program that may be used by or in combination with an instruction execution system, device or device. In the present disclosure, a computer-readable signal medium may include a data signal propagated in a baseband or as part of a carrier wave, in which a computer-readable program code is carried. This propagated data signal may take a variety of forms, including but not limited to an electromagnetic signal, an optical signal, or any suitable combination of the above. The computer readable signal medium may also be any computer readable medium other than a computer readable storage medium, which may send, propagate or transmit a program for use by or in conjunction with an instruction execution system, apparatus or device. The program code contained on the computer readable medium may be transmitted using any suitable medium, including but not limited to: wires, optical cables, RF (radio frequency), etc., or any suitable combination of the above.

[0121] In some embodiments, the client and the server may communicate using any currently known or future developed network protocol such as HTTP (Hyper Text Transfer Protocol), and may be interconnected with any form or medium of digital data communication (e.g., a communication network). Examples of communication networks include a local area network ("LAN"), a wide area network ("WAN"), an internet (e.g., the Internet), and a peer-to-peer network (e.g., an ad hoc peer-to-peer network), as well as any currently known or future developed network.

[0122] The computer-readable medium may be included in the electronic device, or may exist independently without being incorporated into the electronic device.

[0123] The computer-readable medium carries one or more programs. When the one or more programs are executed by the electronic device, the electronic device:

[0124] Obtaining control parameters, and determining whether the number of controls in the control parameters reaches a preset number; wherein the control parameters include the number of controls, the semi-major axis target control value, the control type, and the initial control time;

[0125] When it is determined that the number of control times does not reach the preset number, a target control strategy and a target control time are determined based on the control type, the semi-major axis target control value and the initial control time;

[0126] At the target control moment, the spacecraft is controlled along its semi-major axis using the target control strategy.

[0127] Optionally, when the above one or more programs are executed by the electronic device, the electronic device may also execute other steps described in the above embodiments.

[0128] Computer program code for performing the operations of the present disclosure may be written in one or more programming languages ​​or a combination thereof, including, but not limited to, 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).

[0129] The flow chart and block diagram in the accompanying drawings illustrate the possible architecture, function and operation of the system, method and computer program product according to various embodiments of the present disclosure. In this regard, each square 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 implementations as replacements, the functions marked in the square box can also occur in a sequence different from that marked in the accompanying drawings. For example, two square 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 square box in the block diagram and / or flow chart, and the combination of the square boxes in the block diagram and / or flow chart can be implemented with a dedicated hardware-based system that performs the specified function or operation, or can be implemented with a combination of dedicated hardware and computer instructions.

[0130] The units involved in the embodiments described in the present disclosure may be implemented by software or hardware, wherein the name of a unit does not, in some cases, limit the unit itself.

[0131] The functions described above herein may be performed at least in part by one or more hardware logic components. For example, without limitation, exemplary types of hardware logic components that may be used include: field programmable gate arrays (FPGAs), application specific integrated circuits (ASICs), application specific standard products (ASSPs), systems on chips (SOCs), complex programmable logic devices (CPLDs), and the like.

[0132] In the context of the present disclosure, a machine-readable medium may be a tangible medium that may contain or store a program for use by or in conjunction with an instruction execution system, device, or equipment. A machine-readable medium may be a machine-readable signal medium or a machine-readable storage medium. A machine-readable medium may include, but is not limited to, an electronic, magnetic, optical, electromagnetic, infrared, or semiconductor system, device, or equipment, or any suitable combination of the foregoing. A more specific example of a machine-readable storage medium may include an electrical connection based on one or more lines, 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 foregoing.

[0133] The above description is only a preferred embodiment of the present disclosure and an explanation of the technical principles used. Those skilled in the art should understand that the scope of disclosure involved in the present disclosure is not limited to the technical solutions 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 disclosure (but not limited to) by each other to form a technical solution.

[0134] In addition, although each operation is described in a specific order, this should not be understood as requiring these operations to be performed in the specific order shown or in a sequential order. Under certain circumstances, multitasking and parallel processing may be advantageous. Similarly, although some specific implementation details are included in the above discussion, these should not be interpreted as limiting the scope of the present disclosure. Some features described in the context of a separate embodiment can also be implemented in a single embodiment in combination. On the contrary, the various features described in the context of a single embodiment can also be implemented in multiple embodiments individually or in any suitable sub-combination mode.

[0135] Although the subject matter has been described in language specific to structural features and / or methodological logical actions, it should be understood that the subject matter defined in the appended claims is not necessarily limited to the specific features or actions described above. On the contrary, the specific features and actions described above are merely example forms of implementing the claims.

Claims

1. A method for controlling the semi-major axis of a spacecraft, characterized in that: include: Acquire control parameters, and determine whether the control times in the control parameters reach a preset number; wherein the control parameters include the control times, the semi-major axis target control value, the control type, and the initial control time; When it is determined that the control number does not reach the preset number, determining a target control strategy and a target control time based on the control type, the semi-major axis target control value and the initial control time; At the target control moment, the spacecraft is controlled along its semi-major axis using the target control strategy.

2. The semi-major axis control method of a spacecraft according to claim 1, characterized in that: The control type includes a designated time control type and a non-designated time control type, and the non-designated time control type includes an apogee control type, a perigee control type, an ascending node control type and a descending node control type; wherein, the apogee control type performs semi-major axis control on the spacecraft at the first apogee after the initial control time; the perigee control type performs semi-major axis control on the spacecraft at the first perigee after the initial control time; the ascending node control type performs semi-major axis control on the spacecraft at the first ascending node after the initial control time; and the descending node control type performs semi-major axis control on the spacecraft at the first descending node after the initial control time.

3. The semi-major axis control method of a spacecraft according to claim 2, characterized in that: The semi-major axis target control value includes a semi-major axis target value and a semi-major axis variation target value; the target control strategy and target control time are determined based on the control type, the semi-major axis target control value and the initial control time, including: Determine the target control moment based on the control type and the initial control moment, and obtain the pre-control motion state and pre-control semi-major axis of the spacecraft corresponding to the target control moment; If the semi-major axis target control value is the semi-major axis variation target value, a corresponding pulse velocity variation value is determined based on the semi-major axis variation target value and the semi-major axis control rate, and the target control strategy is determined based on the pulse velocity variation value, the pre-control motion state and the pre-control semi-major axis; If the semi-major axis target control value is the semi-major axis target value, determine the difference between the semi-major axis target value and the semi-major axis before control, determine the corresponding pulse velocity change value based on the difference and the semi-major axis control rate, and determine the target control strategy based on the pulse velocity change value, the pre-control motion state and the semi-major axis before control.

4. The semi-major axis control method of a spacecraft according to claim 3, characterized in that: The determining the target control time based on the control type and the initial control time includes: When the control type is the designated time control type, the initial control time is used as the target control time; When the control type is the non-specified time control type, the angular velocity of the spacecraft corresponding to the initial control time is obtained, and the target control time is determined based on the angular velocity and the initial control time.

5. The semi-major axis control method of a spacecraft according to claim 4, characterized in that: When the non-designated time control type is the apogee control type, determining the target control time based on the angular velocity and the initial control time includes: Acquire a first angle of the spacecraft from a first apogee after the initial control moment, and determine a first time required for the spacecraft to reach the first apogee after the initial control moment based on the first angle and the angular velocity; The target control time is determined based on the initial control time and the first duration.

6. The method for controlling the semi-major axis of a spacecraft according to claim 4, characterized in that: When the non-designated time control type is the perigee control type, determining the target control time based on the angular velocity and the initial control time includes: Acquire a second angle of the spacecraft from the first perigee after the initial control moment, and determine a second time required for the spacecraft to reach the first perigee after the initial control moment based on the second angle and the angular velocity; The target control time is determined based on the initial control time and the second duration.

7. The semi-major axis control method of a spacecraft according to claim 4, characterized in that: When the non-designated time control type is the ascending node control type, determining the target control time based on the angular velocity and the initial control time includes: acquiring a third angle of the spacecraft from the first ascending node after the initial control moment, and determining a third time required for the spacecraft to reach the first ascending node after the initial control moment based on the third angle and the angular velocity; The target control time is determined based on the initial control time and the third duration.

8. The method for controlling the semi-major axis of a spacecraft according to claim 4, characterized in that: When the non-designated time control type is the descending node control type, determining the target control time based on the angular velocity and the initial control time includes: Acquire a fourth angle of the spacecraft from the first descending node after the initial control moment, and determine a fourth time required for the spacecraft to reach the first descending node after the initial control moment based on the fourth angle and the angular velocity; The target control time is determined based on the initial control time and the fourth time duration.

9. A semi-major axis control device for a spacecraft, characterized in that: include: An acquisition module, used to acquire control parameters and determine whether the control times in the control parameters have reached a preset number; wherein the control parameters include the control times, the semi-major axis target control value, the control type and the initial control time; a processing module, configured to determine a target control strategy and a target control time based on the control type, the semi-major axis target control value and the initial control time when it is determined that the control number does not reach the preset number; The control module is used to perform semi-major axis control of the spacecraft using the target control strategy at the target control moment.

10. An electronic device comprising: processor; and a memory storing a program, wherein the program comprises instructions for causing the processor to execute the method according to any one of claims 1-8.

Images