A method and electronic equipment for joint control of low-Earth orbit satellite formation

Abstract

The present application provides a method and electronic equipment for joint control of a satellite formation, wherein the satellite formation includes a master satellite and a slave satellite, and the method includes: calculating the nominal formation configuration parameters of the master satellite and the master satellite orbit configuration maintenance error according to the semi-major axis attenuation rate of the master satellite orbit; using the master satellite orbit configuration maintenance error to calculate the ignition latitude argument and pulse size of the master satellite; calculating the nominal formation configuration parameters of the slave satellite at the ignition moment of the master satellite according to the formation relative semi-major axis attenuation rate and the ignition latitude argument of the master satellite; calculating the slave satellite formation configuration error according to the current slave satellite nominal formation configuration parameters, and using the slave satellite formation configuration error to calculate the ignition latitude argument and pulse size of the slave satellite.

Description

Technical Field

[0001] The present application relates to the field of satellite formation and orbit control, and specifically to a method and electronic equipment for joint control of a low-Earth orbit satellite formation. Background Art

[0002] Satellite formation flying has been a hot topic that has attracted widespread attention at home and abroad in recent years. Formation flying is a new satellite networking method, which means that several satellites are operating in close orbits and forming a specific shape. While maintaining this shape, each satellite flies around the earth with the same orbital period. Compared with a single satellite, satellite formation flying can complete more dimensional payload tasks. Summary of the invention

[0003] The present application aims to provide a method and electronic equipment for joint control of a low-Earth orbit satellite formation, which simultaneously completes the joint control of the satellite formation's in-plane and out-of-plane orbit formations through three pulses of a single satellite, significantly reducing the satellite control time compared to traditional methods.

[0004] According to one aspect of the present application, a method for joint control of a satellite formation is proposed, wherein the satellite formation includes a master satellite and a slave satellite, and the method includes:

[0005] The nominal formation configuration parameters of the primary star and the configuration maintenance error of the primary star orbit are calculated according to the semi-major axis decay rate of the primary star orbit;

[0006] Utilizing the primary satellite orbit configuration maintenance error, the ignition latitude argument and pulse size of the primary satellite are calculated;

[0007] Calculating the nominal formation configuration parameters of the slave satellite at the time of ignition of the master satellite according to the relative semi-major axis decay rate of the formation and the ignition latitude argument of the master satellite;

[0008] According to the nominal formation configuration parameters of the current slave satellite, the slave satellite formation configuration error is calculated, and the ignition latitude argument and pulse size of the slave satellite are calculated using the slave satellite formation configuration error.

[0009] According to some embodiments, the method includes: controlling the pulses of the master satellite and the slave satellite to include a first pulse, a second pulse, and a third pulse.

[0010] According to some embodiments, the method includes: the primary satellite and the secondary satellite have consistent ignition latitude arguments at the first pulse and the second pulse.

[0011] According to some embodiments, the method includes: calculating the nominal formation configuration parameters of the slave satellite at the time of ignition of the master satellite according to the relative semi-major axis attenuation rate of the formation and the ignition latitude argument of the master satellite, including:

[0012] Calculate the nominal relative eccentricity vector and the nominal relative semi-major axis from the star.

[0013] According to some embodiments, the method includes: calculating the nominal formation configuration parameters of the slave satellite at the time of ignition of the master satellite according to the relative semi-major axis attenuation rate of the formation and the ignition latitude argument of the master satellite, and further includes:

[0014] The actual nominal relative eccentricity vector of the slave satellite relative to the master satellite under the ignition latitude argument is calculated so that the error between the actual nominal relative eccentricity vector and the expected nominal relative eccentricity vector is minimized.

[0015] According to some embodiments, the method includes: calculating the nominal formation configuration parameters of the slave satellite at the time of ignition of the master satellite according to the relative semi-major axis attenuation rate of the formation and the ignition latitude argument of the master satellite, and further includes:

[0016] Calculate the change of the relative semi-major axis within the control period according to the relative semi-major axis attenuation rate of the master and slave satellites and the along-track separation introduced by the first pulse and the second pulse;

[0017] The nominal relative semi-major axis is calculated according to the calculated variation of the relative semi-major axis.

[0018] According to some embodiments, the method further comprises:

[0019] According to the calculated pulse sizes and ignition latitude angles of the master and slave satellites, the start and end times of continuous thrust ignition at each pulse are calculated using the propulsion system parameters.

[0020] According to some embodiments, the method further comprises:

[0021] The deviation of the in-plane formation configuration parameters is calculated based on the in-plane configuration parameters of the current orbit.

[0022] According to another aspect of the present application, there is provided an electronic device, including:

[0023] 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, any one of the above methods is implemented.

[0024] According to another aspect of the present application, a computer program product is provided, comprising a computer program or instructions, which implement any one of the above methods when executed by a processor.

[0025] According to the exemplary embodiment of the present application, by using the absolute position navigation result of the master satellite and the relative position navigation result of the formation, the optimal pulse size and application time of the master and slave satellites for orbit formation maintenance are autonomously calculated and generated. And by using the pulse size and application time, the continuous low-thrust orbit formation configuration joint control of the low-Earth orbit satellite formation is realized.

[0026] It should be understood that the foregoing general description and the following detailed description are exemplary only and are not restrictive of the present application. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] In order to more clearly illustrate the technical solutions in the embodiments of the present application, the following briefly introduces the drawings required for describing the embodiments.

[0028] Figure 1 A flow chart showing a method for joint control of a low-Earth orbit satellite formation according to an exemplary embodiment of the present application.

[0029] Figure 2 A schematic diagram showing satellite positions and reference orbits according to an exemplary embodiment of the present application is shown.

[0030] Figure 3 A schematic diagram of low-Earth orbit satellite formation configuration maintenance control according to an exemplary embodiment of the present application is shown.

[0031] Figure 4 A block diagram of an electronic device according to an exemplary embodiment is shown. DETAILED DESCRIPTION

[0032] Example embodiments will now be described more fully with reference to the accompanying drawings. However, example embodiments can be implemented in many forms and should not be construed as limited to the embodiments set forth herein; rather, these embodiments are provided so that this application will be comprehensive and complete and will fully convey the concepts of the example embodiments to those skilled in the art. The same reference numerals in the figures represent the same or similar parts, and thus their repeated description will be omitted.

[0033] In addition, described feature, structure or characteristic can be combined in one or more embodiments in any suitable manner. In the following description, many specific details are provided to provide a full understanding of the embodiments of the present application. However, those skilled in the art will appreciate that the technical scheme of the present application can be put into practice without one or more of the specific details, or other methods, components, devices, steps, etc. can be adopted. In other cases, known methods, devices, realizations or operations are not shown or described in detail to avoid blurring the various aspects of the application.

[0034] The block diagrams shown in the accompanying drawings are merely functional entities and do not necessarily correspond to physically independent entities. That is, these functional entities may be implemented in software form, or in one or more hardware modules or integrated circuits, or in different networks and / or processor devices and / or microcontroller devices.

[0035] The flowcharts shown in the accompanying drawings are only exemplary and do not necessarily include all the contents and operations / steps, nor must they be executed in the order described. For example, some operations / steps can be decomposed, and some operations / steps can be combined or partially combined, so the actual execution order may change according to actual conditions.

[0036] It should be understood that although the terms first, second, third, etc. may be used herein to describe various components, these components should not be limited by these terms. These terms are used to distinguish one component from another component. Therefore, the first component discussed below can be referred to as the second component without departing from the teachings of the concepts of the present application. As used herein, the term "and / or" includes any one of the associated listed items and all combinations of one or more.

[0037] Those skilled in the art will appreciate that the drawings are merely schematic diagrams of example embodiments, and the modules or processes in the drawings are not necessarily necessary for implementing the present application, and therefore cannot be used to limit the scope of protection of the present application.

[0038] The orbit control of the satellite formation is used to correct the absolute orbit deviation caused by the perturbation force, and the formation control is used to eliminate the alienation of the formation configuration caused by the difference in the surface mass ratio of the main and auxiliary satellites. During the orbit control process, the satellite has a large maneuvering range, while the distance between the formation satellites is very small. There is a great risk of collision when controlling some satellites alone, so the formation satellites need to maneuver at the same time.

[0039] The existing technology adopts two schemes. Scheme 1: Control the master and slave satellites in sequence. The master satellite calculates the control amount and then maneuvers. After the master satellite reaches the target position, the slave satellite calculates the control amount and then maneuvers. Scheme 2: After the master satellite calculates the control amount, it synchronizes the control amount and control time to the slave satellite. The master and slave satellite adopt the same control amount and control time.

[0040] The inventors found that both of the above two solutions have the following disadvantages.

[0041] The configuration cannot be maintained during the control process, resulting in safety issues. The distance between the master and slave satellites is very small, and the destruction of the configuration may cause the risk of collision between the formation satellites. Controlling the master and slave satellites one after another or taking the same control amount for the master and slave satellites will destroy the original formation configuration, leading to possible collision risks.

[0042] The control process is complex and time-consuming. To control the master and slave satellites separately, the master satellite needs to maneuver first, and then the slave satellites after the master satellite is in place, which requires more orbits for control; after the master and slave satellites take the same control amount, the configuration is difficult to maintain, and formation control needs to be performed again, which also consumes more orbits.

[0043] In addition to solving the above shortcomings of the existing technology, in order to save energy and simplify control, orbit control and formation control are carried out simultaneously to minimize the number of controls and control time. According to the specific constraints of the actual formation satellite project, an orbit formation joint control strategy for satellite formation is designed.

[0044] The present application proposes a method for joint control of a satellite formation, which enables orbit control and formation control to be performed simultaneously, reducing the number of controls and control time. In addition, the formation satellites need to maneuver at the same time to avoid the risk of collision.

[0045] Satellite formation refers to the process in which the master and slave satellites simultaneously orbit the earth in a circle, and the slave satellites maintain a specific configuration relative to the master satellite. Orbital control refers to the overall control of the satellite formation, including orbit raising; formation control refers to the operation of maintaining the relative position relationship between the master and slave satellites, and only controls the slave satellites.

[0046] In this application, the master satellite and the slave satellite performed three pulse controls respectively. The first two times were in-plane control, and the control center time (i.e., latitude angle u) of the master and slave satellites were consistent. The third time was out-of-plane control, and the master and slave satellites adopted their own values ​​respectively.

[0047] This application solves the problem of maintaining safety and configuration accuracy during the orbital control of satellite formations. It provides a method for joint orbital formation control of low-Earth orbit satellite formations that operates autonomously on board. It uses the absolute position navigation results of the master satellite and the relative position navigation results of the formation to autonomously calculate and generate the optimal pulse size and application time of the master and slave satellites for orbital formation maintenance. And using the pulse size and application time, it realizes the joint control of the continuous low-thrust orbital formation configuration of the low-Earth orbit satellite formation.

[0048] The exemplary embodiments of the present application are described below with reference to the accompanying drawings.

[0049] Figure 1 A flow chart showing a method for joint control of a low-Earth orbit satellite formation according to an exemplary embodiment of the present application.

[0050] According to some embodiments, the formation configuration is described using a relative E / I vector, and the in-plane formation configuration parameter deviation is calculated based on the current in-plane configuration parameters.

[0051] The relative E / I vector is defined as follows:

[0052]

[0053] In the formula, a k ,e k ,i k ,ω k ,Ω k ,u k , k = 1, 2, the meaning of each variable is: the semi-major axis a of the formation satellite orbit k , eccentricity e k , orbital inclination i k 、Arg of perigee ω k , right ascension of ascending node Ω k and latitude argument u k The subscripts k = 1 and 2 represent the reference orbit and the current orbit respectively. Δa is the relative semi-major axis, (Δe x Δe y ) is the relative eccentricity vector, (Δi x Δi y ) is the relative inclination vector, and Δu is the relative latitude argument.

[0054] In the following text, variables with subscript m refer to the master star relative to the virtual target master star, variables with subscript c refer to the slave star relative to the virtual target master star, and variables with subscript r refer to the slave star relative to the master star.

[0055]

[0056] See also Figure 1 In S101, the primary star nominal formation configuration parameters and the primary star orbit configuration maintenance error are calculated according to the primary star orbit semi-major axis decay rate.

[0057] The nominal formation configuration parameters of the primary star are estimated according to the decay rate of the primary star orbit semi-major axis, and the primary star orbit configuration maintenance error is calculated. Furthermore, the simplified Gaussian equation is used to calculate the primary star control ignition latitude angle with the goal of minimizing fuel consumption. and tangential pulse size The superscript k=1, 2, 3 represent the first pulse, the second pulse and the third pulse respectively.

[0058] Here the decay rate of the primary star's orbit semi-major axis is expressed as k m It indicates that it is the semi-major axis decay rate of the primary star relative to the target orbit.

[0059] The goal of the primary satellite orbit control is to always operate in an orbit with a radius of r centered on the reference orbit. Figure 2 Schematic diagram of satellite positions and reference orbits shown.

[0060] The orbital deviation of the satellite from the center of the strictly regressive orbit includes the deviation within the orbital plane and the deviation outside the orbital plane. The deviation within the plane can be expressed as the deviation of the semi-major axis of the satellite orbit from the strictly regressive orbit Δam and the eccentricity deviation Δe m :

[0061]

[0062] Where k m is the decay rate of the semi-major axis of the primary star's orbit.

[0063] The orbital deviation can be expressed as the inclination deviation Δi between the satellite orbit and the strict regression orbit. m :

[0064]

[0065] Furthermore, the orbital configuration maintenance error of the primary star is calculated:

[0066]

[0067] In S103, the primary satellite orbit configuration maintenance error is used to calculate the primary satellite ignition latitude argument and pulse size.

[0068] For two spacecraft flying closely together in a near-circular orbit, the HCW equation can be used:

[0069]

[0070] The solution to the equation can be expressed in terms of orbital elements:

[0071]

[0072] Furthermore, the relative orbital elements are expressed as a function of the pulse:

[0073]

[0074] Where n is the average orbital angular velocity of the primary star, t M is the maneuvering time, t is the current time, and u is the latitude angle at time tM. Δv=(Δv R ,Δv T ,Δv N ) is the vector representation of the velocity change caused by the main star performing the pulse maneuver in the virtual target star RTN (radial, Radial; tangential, transverse, normal; Normal) coordinate system.

[0075] From formula (7), it can be seen that the optimal fuel configuration control ignition direction of the in-plane control (including the first pulse and the second pulse) is selected in the direction of the vector change in the configuration plane, thereby determining the in-plane configuration control ignition latitude argument:

[0076]

[0077] and the pulse size (speed increment)

[0078]

[0079] Where v is the average satellite speed.

[0080] The optimal configuration control ignition direction of the fuel for out-of-plane control (third pulse) is selected in the direction of the vector change outside the configuration plane, thereby determining the out-of-plane configuration control ignition latitude argument:

[0081]

[0082] And the pulse size (speed increment):

[0083]

[0084] In S105, the nominal formation configuration parameters of the slave satellite at the time of ignition of the master satellite are calculated according to the relative semi-major axis attenuation rate of the formation and the argument of the ignition latitude of the master satellite.

[0085] According to the formation relative semi-major axis attenuation rate and the main satellite ignition latitude argument, the nominal in-plane formation configuration parameters of the slave satellite at the time of main satellite ignition are calculated. The relative semi-major axis attenuation rate of the formation here is the relative semi-major axis attenuation rate of the master and slave satellites, which represents the semi-major axis attenuation rate k of the slave satellite relative to the master satellite. r .

[0086] In order to ensure the formation mission, the formation configuration needs to be maintained with high precision by calculating the nominal relative eccentricity vector Overshoot control is performed on e, and the nominal relative semi-major axis is calculated To control the radial offset u.

[0087] The expected nominal relative eccentricity vector is (real deviation, predicted):

[0088]

[0089] Where R represents the basic rotation matrix in the relative eccentric plane, is the relative eccentricity vector angle attenuation rate, control period T man is m times the orbital period T, that is, T man =mT.

[0090] To avoid the risk of collision, the master satellite and the slave satellite maneuver simultaneously during the in-plane control process. For this purpose, the slave satellite uses the value calculated by the master satellite for the in-plane control ignition latitude angle:

[0091]

[0092] By adopting the same ignition latitude angle, the master satellite and the slave satellite can maneuver at the same time, effectively avoiding the risk of collision between formation satellites.

[0093] Calculate the actual nominal relative eccentricity of the slave star relative to the master star at this latitude argument Minimize the error between the relative eccentricity vector and the expected nominal value:

[0094]

[0095] Furthermore, the control amount of the slave star relative to the master star is calculated:

[0096]

[0097] Through the two-pulse in-plane formation maintenance strategy, the relative semi-major axis Δa is introduced r The bias is used to separate the track Δu r Therefore, it is necessary to calculate the relative semi-major axis offset according to the change of the along-track separation during the control cycle, that is, the nominal relative semi-major axis

[0098] Relative orbital angle Δu within the control period r The total change is:

[0099]

[0100] The separation along the track introduced by two pulses and They are:

[0101]

[0102] In the formula, k r is the relative semi-major axis attenuation rate of the master and slave satellites. According to the control mechanism, the change of the relative semi-major axis in one control cycle caused by two maneuvers can be calculated:

[0103]

[0104] The relative semi-major axis after the first pulse can be calculated from (16)(17)(18): Relative semi-major axis after the second pulse and the nominal relative semi-major axis

[0105]

[0106] In S107, the slave satellite formation configuration error is calculated according to the current slave satellite nominal formation configuration parameters, and the slave satellite control ignition latitude argument and pulse size are calculated using the slave satellite formation configuration error.

[0107] According to some embodiments, the slave satellite formation configuration holding error is calculated based on the current slave satellite nominal formation configuration parameters:

[0108]

[0109] Calculate the pulse size from the configuration control in the star orbit plane:

[0110]

[0111] The optimal configuration of the fuel controlled out-of-plane ignition direction is selected in the direction of the vector change outside the configuration plane, thereby determining the latitude angle of the out-of-plane configuration controlled ignition:

[0112]

[0113] And the pulse size (speed increment):

[0114]

[0115] After obtaining the optimal pulse size and ignition latitude angle according to the above calculations, the propulsion system parameters are used to calculate the start and end times of continuous thrust ignition at each optimal pulse, thereby achieving continuous low-thrust configuration control of the formation with optimal satellite fuel consumption.

[0116] The start and end time of each pulse ignition is:

[0117]

[0118] In the formula, t UTC is the current time, is the current latitude argument of the controlled satellite; and is the optimal implementation time and pulse size calculated in steps 2 to 4; M s and F Th are the mass of the controlled star and the thrust of the propeller respectively; k=1,2,3 represent three pulses respectively, s=m,c represent the master star and the slave star respectively.

[0119] Formula (9) in step S103 and formula (21) in this step can give similar results by changing their forms, and their essence is the same.

[0120] The control results obtained by the autonomous orbit formation joint control strategy for the low-Earth orbit satellite formation proposed in this application are all given explicitly, with the characteristics of simple form and small calculation amount, and are easy to implement on the satellite and operate autonomously.

[0121] On the basis of proposing the optimal solution of pulse control, according to actual engineering constraints such as the actual thrust size of the thrusters, a continuous small thrust configuration maintenance control method for the formation is given to achieve the goals of simultaneously satisfying the control accuracy of the formation configuration maintenance and the optimal fuel consumption.

[0122] By performing three pulse controls on each satellite, the satellite formation's in-plane and out-of-plane orbital formation joint control is completed at the same time, which significantly reduces the satellite control time compared to traditional methods and reduces the impact on satellite mapping or remote sensing services. In addition, by reducing the frequency of satellite on-orbit control, the control process is simplified and the risk of failure is reduced.

[0123] During the control process, the relative control of the master and slave satellites is carried out using high-precision relative navigation results, avoiding the influence of absolute navigation errors and ensuring the control accuracy of the relative positions of the satellite formation during orbit control.

[0124] Figure 3 A schematic diagram of low-Earth orbit satellite formation configuration maintenance control according to an exemplary embodiment of the present application is shown.

[0125] According to some embodiments, a low-Earth orbit satellite formation configuration maintenance control method includes the following steps.

[0126] The formation configuration is described by relative E / I vector, and the deviation of the in-plane formation configuration parameters is calculated according to the current in-plane configuration parameters; then the nominal formation configuration parameters are calculated according to the formation configuration parameter attenuation rate estimation; and then the formation configuration maintenance error is calculated according to the current formation configuration and the nominal formation configuration parameters. The current formation configuration maintenance error is used to minimize fuel consumption, and the simplified Gaussian equation is used to calculate the configuration control ignition latitude angle and pulse size; finally, after the optimal pulse size and implementation time are calculated, the propulsion system parameters are used to calculate the continuous thrust ignition start and end time at each optimal pulse, so as to achieve the formation continuous low thrust configuration control with optimal satellite fuel consumption.

[0127] In this embodiment, the semi-major axis of the master satellite a = 6907633m, the orbital angular velocity n = 0.0011rad / s, and the target relative orbital element of the slave satellite and the master satellite in the satellite formation is Δα r_t =(0m,3.51e-6,5.8e-5,2.46e-05,-1.13e-05,2.67e-05)

[0128] Step 1: According to the actual main star orbit elements and the target main star orbit elements, calculate the relative orbit element deviation between the actual main star orbit at the current moment and the virtual target main star orbit as Δα m=(-35.8m,7.26e-6,5.52e-6,1.11e-6,-1.33e-07,3.71e-07), according to the actual slave star orbit element, the actual master star orbit element and the target master star relative orbit element, the relative orbit element deviation between the actual slave star orbit and the actual master star orbit at the current moment is calculated as Δα r =(-1.56m,5.25e-6,5.96e-5,2.12e-5,1.12e-05,2.35e-05)

[0129] Step 2: In this embodiment, the relative semi-major axis decay rate of the primary star is k m =-30m / day, the control period is m=30, and the state vector of the main star formation is δΔα m =(65.6m,-7.26e-6,-5.52e-6,-1.11e-6,1.33e-07,-3.71e-07), so we can get the three pulse times And three pulse sizes

[0130]

[0131] Step 3: In this embodiment, the relative semi-major axis attenuation rate is k m =-1.5m / day, the relative eccentricity vector angle decay rate is The control period m = 15, so from formula (9) we get: Further, the nominal relative eccentricity Δe can be obtained man =(5.70e-5,-2.08e-5). From (17), we can get the nominal relative semi-major axis

[0132] Step 4: Calculate the slave satellite configuration error δΔα based on the current slave satellite nominal formation configuration parameters c =(68.4m,-1.74e-6,-6.87e-06,-1.46e-06,-0.22e-06,2.82e-06), and then we can get the third pulse time And three pulse sizes

[0133]

[0134] Step 5: During formation flight, the current on-board time t UTC =35980s, latitude angle u=5, primary star mass M m =311.53kg, auxiliary star mass M c =261.81kg, thrust F Th =0.015N. Therefore, the start and end time of the three pulse ignition of the main satellite can be calculated using formula (18): And the start and end time of the three pulse ignition from the star

[0135]

[0136] It should be clearly understood that the present application describes how to form and use specific examples, but the present application is not limited to any details of these examples. On the contrary, based on the teaching of the content disclosed in the present application, these principles can be applied to many other embodiments.

[0137] Those skilled in the art will appreciate that all or part of the steps to implement the above embodiments are implemented as a computer program executed by a CPU. When the computer program is executed by the CPU, the program for executing the above functions defined by the above method provided in the present application may be stored in a computer-readable storage medium, which may be a read-only memory, a disk or an optical disk, etc.

[0138] In addition, it should be noted that the above figures are only schematic illustrations of the processes included in the method according to the exemplary embodiments of the present application, and are not intended to be limiting. It is easy to understand that the processes shown in the above figures do not indicate or limit the time sequence of these processes. In addition, it is also easy to understand that these processes can be performed synchronously or asynchronously, for example, in multiple modules.

[0139] Through the description of the example embodiments, it will be easy for those skilled in the art to understand that the method for joint control of a low-Earth orbit satellite formation according to the embodiments of the present application has at least one or more of the following advantages.

[0140] According to the example embodiment, the satellite formation's in-plane and out-of-plane orbital formation joint control is completed simultaneously through three pulses, which significantly reduces the satellite control time compared to the traditional method and reduces the impact on satellite mapping or remote sensing services. By ensuring that the primary and auxiliary satellites maneuver at the same time, the risk of collision between formation satellites is effectively avoided. In addition, by reducing the frequency of satellite on-orbit control, the control process is simplified and the risk of failure is reduced.

[0141] According to the example embodiment, during the control process, the relative control of the master and slave satellites is performed using relative orbital elements calculated from high-precision relative navigation results, thereby avoiding the influence of absolute navigation errors and ensuring the control accuracy of the relative positions of the satellite formation during the orbital control process.

[0142] According to the example embodiments, the control results obtained by the proposed autonomous orbit formation joint control strategy for the low-Earth orbit satellite formation are all given explicitly, have the characteristics of simple form and small calculation amount, and are easy to implement on the satellite and operate autonomously.

[0143] According to the example embodiment, on the basis of proposing the optimal solution of pulse control, a continuous small thrust configuration maintenance control method for the formation is given according to actual engineering constraints such as the actual thrust size of the thruster, so as to achieve the goals of simultaneously satisfying the control accuracy of the formation configuration maintenance and the optimal fuel consumption.

[0144] Figure 4 A block diagram of an electronic device according to an exemplary embodiment is shown.

[0145] Refer to the following Figure 4 The electronic device 200 according to this embodiment of the present application is described. Figure 4 The electronic device 200 shown is merely an example and should not bring any limitation to the functions and scope of use of the embodiments of the present application.

[0146] like Figure 4 As shown, the electronic device 200 is in the form of a general computing device. The components of the electronic device 200 may include, but are not limited to: at least one processing unit 210, at least one storage unit 220, a bus 230 connecting different system components (including the storage unit 220 and the processing unit 210), a display unit 240, etc.

[0147] The storage unit stores program codes, and the program codes can be executed by the processing unit 210, so that the processing unit 210 executes the methods described in this specification according to various exemplary embodiments of the present application.

[0148] The storage unit 220 may include a readable medium in the form of a volatile storage unit, such as a random access memory unit (RAM) 2201 and / or a cache memory unit 2202 , and may further include a read-only memory unit (ROM) 2203 .

[0149] The storage unit 220 may also include a program / utility 2204 having a set (at least one) of program modules 2205, such program modules 2205 including but not limited to: an operating system, one or more application programs, other program modules, and program data, each of which or some combination may include an implementation of a network environment.

[0150] Bus 230 may represent one or more of several types of bus structures, including a memory unit bus or memory unit controller, a peripheral bus, an accelerated graphics port, a processing unit, or a local bus using any of a variety of bus architectures.

[0151] The electronic device 200 may also communicate with one or more external devices 300 (e.g., keyboards, pointing devices, Bluetooth devices, etc.), may also communicate with one or more devices that enable a user to interact with the electronic device 200, and / or communicate with any device that enables the electronic device 200 to communicate with one or more other computing devices (e.g., routers, modems, etc.). Such communication may be performed via an input / output (I / O) interface 250. Furthermore, the electronic device 200 may also communicate with one or more networks (e.g., a local area network (LAN), a wide area network (WAN), and / or a public network, such as the Internet) via a network adapter 260. The network adapter 260 may communicate with other modules of the electronic device 200 via the bus 230. It should be understood that, although not shown in the figure, other hardware and / or software modules may be used in conjunction with the electronic device 200, including but not limited to: microcode, device drivers, redundant processing units, external disk drive arrays, RAID systems, tape drives, and data backup storage systems.

[0152] Through the description of the above implementation methods, it is easy for those skilled in the art to understand that the example implementation methods described here can be implemented by software or by combining software with necessary hardware. The technical solution according to the implementation method of the present application can be embodied in the form of a software product, which can be stored in a non-volatile storage medium (which can be a CD-ROM, a USB flash drive, a mobile hard disk, etc.) or on a network, including several instructions to enable a computing device (which can be a personal computer, a server, or a network device, etc.) to execute the above method according to the implementation method of the present application.

[0153] The software product may use any combination of one or more readable media. The readable medium may be a readable signal medium or a readable storage medium. The readable storage medium may be, for example, but not limited to, a system, device or device of electricity, magnetism, light, electromagnetic, infrared, or semiconductor, or any combination thereof. More specific examples (non-exhaustive list) of readable storage media include: an electrical connection with one or more wires, a portable disk, a hard disk, a random access memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM or flash memory), an optical fiber, a portable compact disk read-only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination thereof.

[0154] Computer readable storage media may include data signals propagated in baseband or as part of a carrier wave, wherein readable program codes are carried. Such propagated data signals may take various forms, including but not limited to electromagnetic signals, optical signals, or any suitable combination thereof. The readable storage medium may also be any readable medium other than a readable storage medium, which may send, propagate, or transmit a program for use by or in conjunction with an instruction execution system, device, or device. The program codes contained on the readable storage medium may be transmitted using any suitable medium, including but not limited to wireless, wired, optical cable, RF, etc., or any suitable combination thereof.

[0155] Program code for performing the operations of the present application may be written in any combination of one or more programming languages, including object-oriented programming languages ​​such as Java, C++, etc., and conventional procedural programming languages ​​such as "C" or similar programming languages. The program code may be executed entirely on the user computing device, partially on the user device, as a separate software package, partially on the user computing device and partially on a remote computing device, or entirely on a remote computing device or server. In the case of a remote computing device, the remote computing device may be connected to the user computing device through any type of network, including a local area network (LAN) or a wide area network (WAN), or may be connected to an external computing device (e.g., via the Internet using an Internet service provider).

[0156] Those skilled in the art will appreciate that the above modules can be distributed in the device according to the description of the embodiment, or can be changed accordingly and only used in one or more devices different from the embodiment. The modules of the above embodiments can be combined into one module, or further divided into multiple sub-modules.

[0157] The exemplary embodiments of the present application are specifically shown and described above. It should be understood that the present application is not limited to the detailed structures, configurations or implementations described herein; on the contrary, the present application is intended to cover various modifications and equivalent configurations included in the spirit and scope of the appended claims.

Claims

1. A method for joint control of a satellite formation, wherein the satellite formation comprises a master satellite and a slave satellite, characterized in that: The method comprises: The nominal formation configuration parameters of the primary star and the configuration maintenance error of the primary star orbit are calculated according to the semi-major axis decay rate of the primary star orbit; Utilizing the primary satellite orbit configuration maintenance error, the ignition latitude argument and pulse size of the primary satellite are calculated; Among them, the optimal configuration control ignition direction of the fuel controlled in the plane is selected in the direction of the vector change in the configuration plane, and the latitude angle of the ignition controlled in the plane configuration is determined: And the pulse size: Where v is the average satellite speed; The optimal configuration of the fuel controlled out-of-plane ignition direction is selected in the direction of the vector change outside the configuration plane, and the latitude angle of the out-of-plane configuration controlled ignition is determined: And the pulse size: Calculating the nominal formation configuration parameters of the slave satellite at the time of ignition of the master satellite according to the relative semi-major axis decay rate of the formation and the ignition latitude argument of the master satellite; Calculate the slave satellite formation configuration error according to the current slave satellite nominal formation configuration parameters, and use the slave satellite formation configuration error to calculate the ignition latitude argument and pulse size of the slave satellite; Calculate the pulse size from the configuration control in the star orbit plane: The optimal configuration of the fuel controlled out-of-plane ignition direction is selected in the direction of the vector change outside the configuration plane, thereby determining the latitude angle of the out-of-plane configuration controlled ignition: And the pulse size:

2. The method according to claim 1, characterized in that The pulses for controlling the master star and the slave star include a first pulse, a second pulse and a third pulse.

3. The method according to claim 2, characterized in that The master satellite and the slave satellite have the same ignition latitude arguments during the first pulse and the second pulse.

4. The method according to claim 3, characterized in that The calculation of the nominal formation configuration parameters of the slave satellite at the time of the master satellite ignition according to the relative semi-major axis attenuation rate of the formation and the ignition latitude argument of the master satellite includes: Calculate the nominal relative eccentricity vector and the nominal relative semi-major axis from the star.

5. The method according to claim 4, characterized in that The method of calculating the nominal formation configuration parameters of the slave satellite at the time of ignition of the master satellite according to the relative semi-major axis attenuation rate of the formation and the ignition latitude argument of the master satellite also includes: The actual nominal relative eccentricity vector of the slave satellite relative to the master satellite under the ignition latitude argument is calculated so that the error between the actual nominal relative eccentricity vector and the expected nominal relative eccentricity vector is minimized.

6. The method according to claim 5, characterized in that The method of calculating the nominal formation configuration parameters of the slave satellite at the time of ignition of the master satellite according to the relative semi-major axis attenuation rate of the formation and the ignition latitude argument of the master satellite also includes: Calculate the change of the relative semi-major axis within the control period according to the relative semi-major axis attenuation rate of the master and slave satellites and the along-track separation introduced by the first pulse and the second pulse; The nominal relative semi-major axis is calculated according to the calculated variation of the relative semi-major axis.

7. The method according to claim 1, characterized in that Also includes: According to the calculated pulse sizes and ignition latitude angles of the master and slave satellites, the start and end times of continuous thrust ignition at each pulse are calculated using the propulsion system parameters.

8. The method according to claim 1, characterized in that Also includes: The deviation of the in-plane formation configuration parameters is calculated based on the in-plane configuration parameters of the current orbit.

9. An electronic device, characterized in that: include: A memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor implements the method according to any one of claims 1 to 8 when executing the computer program.

10. A computer program product comprising a computer program or instructions, characterized in that When the computer program or instruction is executed by a processor, the method according to any one of claims 1 to 8 is implemented.

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