Low earth orbit satellite formation configuration maintenance method, device, equipment and medium

Abstract

This application provides a method, apparatus, device, and medium for maintaining the configuration of low-Earth orbit satellite formations, which can solve problems such as frequent control, high fuel consumption, and shortened on-orbit life caused by ignoring the differences in aerodynamic drag between satellites in related technologies. The method includes: acquiring the difference in the semi-major axis decay rate between a reference satellite and a companion satellite in the formation; predicting the relative drift trajectory of the companion satellite relative to the reference satellite in the velocity direction based on the difference in the semi-major axis decay rate and the initial semi-major axis difference; determining an orbital control strategy based on the relative drift trajectory and a preset configuration maintenance threshold; executing the orbital control strategy to adjust the orbital altitude of the companion satellite when the drift reaches or exceeds the configuration maintenance threshold; resetting the semi-major axis difference after adjusting the orbital altitude of the companion satellite, and returning to the step of acquiring the difference in the semi-major axis decay rate.

Description

Technical Field

[0001] This application relates to the field of spacecraft orbit control technology, and in particular to a method, apparatus, equipment and medium for maintaining low-Earth orbit satellite formation configuration. Background Technology

[0002] When low-Earth orbit satellite formations are in orbit, aerodynamic drag is a core environmental perturbation factor affecting the stability of the formation configuration (the lower the orbital altitude, the more significant the impact of aerodynamic drag). Even if the satellites in the formation have the same design configuration, the aerodynamic drag experienced by each satellite will still have inherent differences due to factors such as manufacturing tolerances, slight attitude deviations, surface-to-mass ratio, and actual deviations in mass. If the satellites in the formation have different configurations, the differences in aerodynamic drag will be further amplified.

[0003] Traditional formation maintenance schemes often assume that the satellites in the formation are subject to the same disturbances and use conventional orbit control strategies to correct the configuration. They do not specifically consider the differences in aerodynamic drag of individual satellites, which can easily lead to over-control of some satellites and under-control of others. This results in frequent maintenance control, affects the normal on-orbit missions of the satellites, generates unnecessary fuel consumption, and shortens the on-orbit service life of the formation satellites.

[0004] Therefore, there is an urgent need for a low-Earth orbit satellite formation maintenance scheme that can utilize differences in aerodynamic drag to reduce control frequency and fuel consumption. Summary of the Invention

[0005] This application provides a method, apparatus, device, and medium for maintaining the low-Earth orbit satellite formation configuration, which can solve the problems in related technologies such as frequent control, high fuel consumption, and shortened on-orbit life caused by ignoring the differences in aerodynamic drag between satellites.

[0006] To achieve the above objectives, this application adopts the following technical solution: Firstly, a method for maintaining the low-Earth orbit satellite formation configuration considering differences in aerodynamic drag is provided, including: Obtain the difference in the semi-major axis decay rate between the reference satellite and the accompanying satellite in the formation. ;in, , The semi-major axis attenuation rate of the accompanying satellite, The semi-major axis attenuation rate of the reference satellite; Based on the difference in the semi-major axis attenuation rate of the orbit And the initial semi-major axis difference Δa0, to predict the relative drift trajectory of the companion satellite relative to the reference satellite in the velocity direction, the relative drift trajectory including the change of drift amount over time; where Δa0 = a c -a s a c For the semi-major axis of the companion satellite, a sThe semi-major axis of the reference satellite; Based on the relative drift trajectory and the preset configuration maintenance threshold ΔL max Determine the track control strategy; When the drift amount reaches or exceeds the configuration maintenance threshold ΔL max At that time, an orbit control strategy is executed to adjust the orbital altitude of the companion satellite; After adjusting the orbital altitude of the companion satellite, reset the semi-major axis difference Δa0 and return to the step of obtaining the difference in the orbital semi-major axis attenuation rate.

[0007] Secondly, a low-Earth orbit satellite formation maintenance device considering aerodynamic drag differences is provided, comprising: The acquisition module is used to obtain the difference in the semi-major axis decay rate of the orbit between the reference satellite and the accompanying satellite in the formation. ;in, , The semi-major axis attenuation rate of the accompanying satellite, The semi-major axis attenuation rate of the reference satellite; The prediction module is used to predict the difference in the semi-major axis decay rate of the orbit. And the initial semi-major axis difference Δa0, to predict the relative drift trajectory of the companion satellite relative to the reference satellite in the velocity direction, the relative drift trajectory including the change of drift amount over time; where Δa0 = a c -a s a c For the semi-major axis of the companion satellite, a s The semi-major axis of the reference satellite; The determination module is used to maintain a threshold ΔL based on the relative drift trajectory and a preset configuration. max Determine the track control strategy; An execution module is configured to respond when the drift amount reaches or exceeds the configuration maintenance threshold ΔL. max At that time, an orbit control strategy is executed to adjust the orbital altitude of the companion satellite; The reset module is used to reset the semi-major axis difference Δa0 after adjusting the orbital altitude of the companion satellite, and return to the step of obtaining the difference in the orbital semi-major axis attenuation rate.

[0008] Thirdly, an electronic device is provided, comprising: One or more processors; Storage device for storing one or more programs; When the one or more programs are executed by the one or more processors, the one or more processors perform the method as described in the first aspect.

[0009] Fourthly, a computer program product is provided, comprising a computer program or instructions that, when executed on a computer, cause the computer to perform the method described in the first aspect.

[0010] The beneficial effects of this invention are: This application provides a method for maintaining the configuration of low-Earth orbit satellite formations. It utilizes the difference in attenuation rate to achieve automatic reversal of the drift direction, avoiding the shortcomings of traditional schemes that require active control every time the boundary is reached, and significantly reducing the control frequency. Since only a single-direction (ascending or descending) control needs to be performed, there is no longer a need to alternate between the two control methods, and fuel consumption can be reduced by about half. The reduction in control frequency and fuel consumption directly extends the effective on-orbit service time of the formation satellites, solving the problems of frequent control, high fuel consumption, and shortened on-orbit life caused by ignoring the differences in aerodynamic drag between satellites in related technologies. Attached Figure Description

[0011] Figure 1 The diagram schematically illustrates the steps of a method for maintaining the low-Earth orbit satellite formation configuration.

[0012] Figure 2 This diagram illustrates a conventional formation satellite velocity and directional control law in the prior art.

[0013] Figure 3 A schematic diagram of a velocity-direction control law based on differences in aerodynamic characteristics is shown.

[0014] Figure 4 A block diagram of a low-Earth orbit satellite formation maintenance device is shown schematically.

[0015] Figure 5 A block diagram of an electronic device is shown schematically.

[0016] Figure 6 A block diagram of a computer-readable medium is shown schematically. Detailed Implementation

[0017] Exemplary embodiments will now be described more fully with reference to the accompanying drawings. However, these exemplary embodiments can be implemented in many forms and should not be construed as limited to the embodiments set forth herein; rather, they are provided so that this application will be thorough and complete, and will fully convey the concept of the exemplary embodiments to those skilled in the art. The same reference numerals in the drawings denote the same or similar parts, and therefore repeated descriptions of them will be omitted.

[0018] Furthermore, the described features, structures, or characteristics can be combined in any suitable manner in one or more embodiments. Numerous specific details are provided in the following description to give a thorough understanding of embodiments of this application. However, those skilled in the art will recognize that the technical solutions of this application can be practiced without one or more of the specific details, or other methods, components, apparatuses, steps, etc., can be employed. In other instances, well-known methods, apparatuses, implementations, or operations are not shown or described in detail to avoid obscuring various aspects of this application.

[0019] 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 can be implemented in software, in one or more hardware modules or integrated circuits, or in different network and / or processor devices and / or microcontroller devices.

[0020] The flowcharts shown in the accompanying drawings are merely illustrative and do not necessarily include all content and operations / steps, nor do they necessarily have to be performed in the described order. For example, some operations / steps can be broken down, while others can be combined or partially combined; therefore, the actual execution order may change depending on the specific circumstances.

[0021] 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. Therefore, the first component discussed below may be referred to as the second component without departing from the teachings of this application. As used herein, the term "and / or" includes all combinations of any and one or more of the listed applications.

[0022] Those skilled in the art will understand that the accompanying drawings are merely schematic diagrams of exemplary embodiments, and the modules or processes in the drawings are not necessarily essential for implementing this application, and therefore cannot be used to limit the scope of protection of this application.

[0023] To address the problems of frequent control, high fuel consumption, and shortened on-orbit lifespan caused by neglecting the differences in aerodynamic drag between satellites in related technologies, this invention proposes a low-Earth orbit satellite formation maintenance scheme that considers differences in aerodynamic drag. By identifying the aerodynamic characteristics of the formation satellites during the on-orbit testing phase, quantifying the impact of aerodynamic drag differences on relative motion, and adjusting the relative motion trajectory and orbit control strategy of individual satellites accordingly, the scheme aims to reduce control frequency and fuel consumption, extend the maintenance period, improve on-orbit lifespan, and enhance the overall on-orbit performance of the formation.

[0024] To facilitate understanding, let's first explain the conventional formation flight control schemes in existing technologies. Relative orbital dynamics is the theoretical basis of formation flight control. In low Earth orbit (LEO) satellite formation control, the CW equations are often used. Assuming a near-circular orbit, with a certain satellite as the reference satellite, in the LVLH coordinate system, the x-axis is the vector direction from the geocenter to the satellite's center of mass, the z-axis points to the orbital plane normal, and the y-axis conforms to the right-hand rule. In a near-circular orbit, this direction is basically consistent with the velocity direction. The in-plane relative motion of the satellite with respect to the reference satellite can be described as (when the distance between the satellite and the reference satellite is close):

[0025] Where the reference satellite is s, the companion satellite is c, n is the angular velocity of the satellite, x and y are the positions of the companion satellite relative to the reference satellite, t is time, τ is the perigee of the companion satellite, and ω is the angular velocity of the satellite. c The perigee argument of the accompanying satellite, ΔΩ=Ω c -Ω s That is, the difference in right ascension of the ascending nodes of the companion satellite and the reference satellite; B=ae c Where a is the semi-major axis of the orbit, e c For the eccentricity of the accompanying satellite, i s Let μ be the inclination angle of the reference satellite, and Δμ(τ) = μ c (τ)-μ s (τ) represents the latitudinal argument difference between the accompanying satellite and the reference satellite at perigee, Δa=a c -a s This represents the difference in the semi-major axis of the orbit between the companion satellite and the reference satellite.

[0026] Based on the above formula, under this assumption, the x-direction exhibits only constant and periodic motion, while the y-direction exhibits constant, periodic, and linear drift motion. The y-direction represents velocity, meaning it is unstable and will gradually drift over time, disrupting the configuration. Therefore, maintaining the configuration of a formation satellite primarily involves keeping the relative motion in the y-direction within a certain range.

[0027] As can be seen from the formula, the relative motion in the y-direction in the CW equation is 1.5nΔat, which is the drift velocity of the satellite relative to the reference satellite. This drift velocity is caused by the difference in the semi-major axes of the two satellites. If the aerodynamic characteristics of the two satellites are identical, their decay rates are essentially the same and can be considered constant. However, in reality, even satellites with the same configuration will have slight differences, leading to variations in relative velocity. Typically, low-Earth orbit satellites are mainly affected by atmospheric drag, causing their altitude to continuously decrease. According to the formula for atmospheric drag:

[0028] In the formula: S is the equivalent cross-sectional area for atmospheric drag, and C... DV is the drag coefficient, m is the satellite mass, ρ is the atmospheric density at the satellite's location, and V=vv a The velocity vectors of the satellite relative to the atmosphere are v and v0. a These are the velocity vectors of the satellite and the atmosphere relative to the Earth's center of mass, respectively. Since the satellites in the formation are relatively close together, their atmospheric density can be considered uniform. Furthermore, during formation, the orbital elements of the satellites are essentially the same, and their velocity vectors relative to the atmosphere can also be considered uniform. Therefore, the difference in attenuation rate is mainly related to the equivalent cross-sectional area and mass of the satellite relative to atmospheric drag; the ratio of these two is usually called the surface-to-mass ratio. The surface-to-mass ratio is a relatively stable value; therefore, the difference in attenuation rate among the satellites in the formation can usually be considered a constant. Furthermore, the change in relative velocity can be considered a linear change.

[0029] In the control of satellite formations, due to thruster errors and orbit determination errors, it is difficult to control the orbital altitude of the satellites in the formation to be consistent. In other words, there is a linear drift in the relative motion of the satellites in the velocity direction. In engineering, the linear drift in the velocity direction is usually handled by offsetting the satellites. That is, by deliberately adjusting the satellites to a higher or lower orbital altitude than the reference satellite, the satellites can drift back and forth within a certain range.

[0030] Taking a 3.8m altitude difference between the two stars as an example, assuming the two stars have the same decay rate, the relative motion amplitude in the velocity direction is 1km, and the required relative motion threshold in the velocity direction is 1.5km, meaning the maximum linear drift is 1km, based on this value, using the relative motion formula, the estimated maintenance period is 2 days. A schematic diagram of the maintenance effect is shown below. Figure 2 As shown, each time the control boundary is reached, the track height needs to be raised or lowered to adjust the drift direction to the opposite direction, thereby achieving control within a certain range.

[0031] Existing technical solutions do not take into account the differences in aerodynamic characteristics of satellites in orbit. In close formation, this may lead to frequent control and wasted fuel. Taking the example above, if the attenuation rate of a satellite is higher than that of the reference satellite, when controlling it separately according to the altitude deviation of ±3.8m during formation holding, there will be cases of over-control and under-control. At an offset of +3.8m, the holding period will be longer, but at an offset of -3.8m, the linear drift speed will be faster than expected, resulting in a shorter holding period. This causes the control period to be inconsistent in length and results in some fuel waste.

[0032] Atmospheric drag continuously reduces satellite orbital altitude, a significant factor affecting satellite lifespan. In formation maintenance, conventional control schemes employ either ascent or descent control at different boundaries to adjust drift direction, inevitably involving descent control. Although the amount of each control operation is small, over long periods of formation maintenance, this accumulates into substantial fuel waste, thus shortening the satellite formation's lifespan. This application addresses the above situation, considering the instability in formation configuration caused by differences in satellite attenuation rates. It proposes an orbit control strategy based on the attenuation rate differences of the formation satellites, reducing control frequency, extending the maintenance period, and thereby improving satellite utilization.

[0033] According to a first specific embodiment of the present invention, such as Figure 1 As shown, the present invention provides a method for maintaining the low-Earth orbit satellite formation configuration considering differences in aerodynamic drag, comprising: Step S11: Obtain the difference in the semi-major axis decay rate between the reference satellite and the accompanying satellite in the formation. .

[0034] in, , The semi-major axis attenuation rate of the accompanying satellite, The semi-major axis attenuation rate of the reference satellite; Specifically, step S11 may include the following steps: S111, during the set period of stable flight of the two satellites in formation, performs linear fitting or segmented calculation on the semi-major axis data of the companion satellite and the reference satellite based on the orbit determination results under uncontrolled conditions, in order to obtain the semi-major axis attenuation rate of each satellite.

[0035] S112, the difference in the semi-major axis attenuation rate of the orbit is calculated based on the obtained semi-major axis attenuation rate of each satellite.

[0036] For example, during the initial deployment, a baseline and companion satellite are determined to complete the initial formation of the two satellites. Then, the two satellites are initiated to maintain stable flight for 7 days. Based on the orbit determination results under uncontrolled conditions, the altitude change rate of a single satellite is obtained by linear fitting or piecewise calculation of the semi-major axis, which is then converted into an altitude decay rate, and the difference in decay rate is calculated.

[0037] Step S12, based on the difference in the semi-major axis attenuation rate of the orbit And the initial semi-major axis difference Δa0, to predict the relative drift trajectory of the companion satellite relative to the reference satellite in the velocity direction, the relative drift trajectory including the change of drift amount over time; where Δa0 = a c -a s a c For the semi-major axis of the companion satellite, a s The semi-major axis of the reference satellite; Furthermore, predicting the relative drift trajectory of the companion satellite relative to the reference satellite in the velocity direction includes establishing a model of how the drift amount ΔL(t) changes with time t: ; Where, when Δa0 and When the signs are opposite, the drift amount ΔL(t) has an extreme value, and the time tpeak when the extreme value occurs is: ; The magnitude of the extreme value ΔLpeak is: ; Where ΔL0 is the initial drift, which can be set to ΔL0=0, n0 is the orbital angular velocity, and a is the semi-major axis of the orbit.

[0038] Step S13: Based on the relative drift trajectory and the preset configuration maintenance threshold ΔL max Determine the track control strategy; Specifically, the orbit control strategy includes: when When the drift is greater than 0, monitor whether the companion satellite is behind the reference satellite. When the drift reaches or is about to reach the configuration maintenance threshold ΔL, max At that time, the accompanying satellite was subjected to orbit reduction control to drift forward; when When the drift is less than 0, monitor whether the companion satellite is in front of the reference satellite. When the drift reaches or is about to reach the configuration maintenance threshold ΔL, max At that time, orbital ascent control was performed on the accompanying satellite to cause it to drift backward.

[0039] It should be noted that the configuration maintenance threshold ΔL max The configuration is set according to the mission requirements of the satellite formation. That is, ΔL. max The requirements for satellite formation configuration depend on the range that the satellites need to maintain during their missions, and there are no fixed requirements.

[0040] Furthermore, step S13 may include the following steps: S131, based on the relative motion dynamics equation corrected by introducing the difference in the semi-major axis decay rate of the orbit, the relationship between the drift amount and time is obtained, and the extreme value of the drift amount and the time when the extreme value occurs are calculated.

[0041] S132, based on the extreme value and the configuration maintenance threshold ΔL max Calculate the required altitude difference between the two satellites after orbit control, so that the extreme value of the drift falls within the configuration maintenance threshold ΔL. max Within.

[0042] The altitude difference between the two satellites after orbit control satisfies the following relationship: ; Right now ; Where, ΔL max To maintain the configuration threshold, n is the angular velocity of the satellite's motion. In this formula, Δa0 is also the altitude difference between the two satellites after orbit control. The difference in attenuation rate between the companion satellite and the reference satellite.

[0043] Furthermore, the control amount can be calculated based on the height difference, so that the relative drift amount can be naturally maintained at the configuration maintenance threshold ΔL without active control. max Within. The control quantity (tangential velocity pulse) Δv = n × Δa0. If the current altitude of the companion satellite is lower than the target altitude, then orbital ascent control (positive tangential pulse) is executed to raise it to the target altitude.

[0044] Step S14, when the drift amount reaches or exceeds the configuration maintenance threshold ΔL max At that time, an orbit control strategy is executed to adjust the orbital altitude of the companion satellite.

[0045] Step S15: After adjusting the orbital altitude of the companion satellite, reset the semi-major axis difference Δa0, and return to the step of obtaining the orbital semi-major axis attenuation rate difference.

[0046] The following example illustrates the velocity-directional control law based on the differences in aerodynamic characteristics between two low-Earth orbit close-packed satellites (reference satellite and companion satellite). Figure 2 As shown.

[0047] First, in the initial deployment phase: determine the base satellite and companion satellite, and complete the initial formation of the two satellites. Set the configuration preservation threshold ΔL required by the mission. max = 1000m (that is, the relative drift in the velocity direction must not exceed ±1000m).

[0048] Next, the difference in semi-major axis attenuation rate is obtained. The two satellites were initiated and stably tracked each other for 7 days without control. During this period, the daily half-major axis sequence of the two satellites was obtained through ground-based orbit determination. Linear fitting was performed on the half-major axis time series of the reference satellite to obtain its decay rate. =-5 m / day (the negative sign indicates an altitude decrease); linear fitting was performed on the semi-major axis time series of the accompanying star to obtain... =-6 m / day. Therefore, the difference in attenuation rate... =(-6)-(-5) =-1m / day. That is, the attenuation rate of the companion satellite is greater than that of the reference satellite (attenuation is faster).

[0049] Then, a drift prediction model is established and a control strategy is determined: the model is improved based on the CW equation, and the drift amount changes with time as follows: ; Assume the initial altitude difference between the companion satellite and the reference satellite is Δa0 = +3.8m (the companion satellite is higher than the reference satellite), and the initial drift is ΔL0 = 0. The orbital angular velocity is n0 ≈ 0.0011 rad / s, and the semi-major axis is a ≈ 6878 km.

[0050] Since Δa0>0 and =-1m / day < 0, the two have opposite signs, which meets the condition for the occurrence of an extreme value: Time of extreme value occurrence: tpeak=-Δa0 / =-3.8 / -1=3.8 days; Extreme value size:

[0051] The calculation result is exactly equal to the task threshold of 1000m, indicating that the initial height difference of 3.8m is appropriate.

[0052] according to If the value is negative, according to the strategy of this invention, it is necessary to monitor whether the companion satellite is in front of the reference satellite (because when the attenuation rate difference is negative, the companion satellite will gradually drift forward). During the drift process, the companion satellite gradually drifts forward (in the Y direction) from its initial position, and the relative drift amount increases. When the drift amount approaches 1000m (for example, around 3.8 days), the altitude difference between the companion satellite and the reference satellite drops to zero, at which point the drift direction automatically changes from forward to backward. Therefore, no active control is required throughout the entire cycle, greatly reducing the number of control operations.

[0053] Next, the stage of determining and executing the control quantity: If active control is required in the next cycle due to environmental changes or initial deviation adjustments, the required semi-major axis deviation is calculated backward from the task threshold. ; The control quantity (tangential velocity pulse) Δv = n·Δa0 ≈ 0.0011 × 3.8 ≈ 0.00418 km / s = 4.18 m / s. If the current altitude of the companion satellite is lower than the target altitude, then an orbital ascent control (positive tangential pulse) is executed to raise it to the deviation value from the target.

[0054] Finally, in the parameter iteration and update phase: after completing one control cycle (approximately 7.6 days), the latest orbit determination data is acquired again, and the current parameters are recalculated. If the change exceeds the preset threshold, the control strategy parameters are adjusted, and the above steps are repeated to form a closed-loop maintenance.

[0055] Figure 3Assuming the companion satellite and the reference satellite are at an altitude of 3.8m, and the configuration maintenance threshold is set to 1000m, if the attenuation rate difference is -1m / d, then exactly at the drift boundary, the companion satellite and the reference satellite attenuate to the same altitude, and the drift direction changes, drifting in opposite directions. No additional control is needed to adjust the drift direction, reducing the number of control operations and extending the maintenance period. According to the theoretical formula, tpeak=-Δa0 / Calculations show that the extreme value occurs on day 3.8. Considering symmetry, the maintenance period is approximately 7.6 days. After exceeding the limit, the control value is 7.6m. During the maintenance process, only elevation or descent control is needed, without the need for alternating control. Furthermore, if the attenuation rate of the accompanying satellite is higher than that of the reference satellite, only elevation control is required, without descent control. This can further save fuel for low-Earth orbit satellites. Simulation analysis shows that, under the condition of consistent initial altitude difference, fuel consumption can be reduced by about half.

[0056] Right now, Figure 2 The existing technical solution requires control (alternating between raising and lowering the orbit) approximately every two days. Figure 3 The proposed solution only requires one control operation before exceeding the limit (ideally, no control is even needed in this embodiment), extending the maintenance period to 7.6 days and avoiding orbit reduction control (because...). (If the value is negative, only an orbital ascent is required), fuel consumption is estimated to be reduced by about 50%.

[0057] Compared with the prior art, this application has the following beneficial effects: 1. Significantly reduced control frequency: By utilizing the difference in attenuation rate, the drift direction is automatically reversed, avoiding the drawback of traditional solutions that require active control every time the boundary is reached. Taking typical values ​​(initial height difference 3.8m, attenuation rate difference -1m / day, threshold 1000m) as an example, the maintenance period of the present invention can reach about 7.6 days, while the traditional solution is only about 2 days.

[0058] 2. Significantly reduced fuel consumption: Since only single-direction control (ascending or descending) needs to be performed, eliminating the need for alternating between two control methods, fuel consumption can be reduced by approximately half. Furthermore, the allocation of control parameters can be optimized using extreme value formulas to avoid overcontrol.

[0059] 3. Extend satellite on-orbit lifespan: The reduction in control frequency and fuel consumption directly extends the effective on-orbit service time of the formation satellites.

[0060] 4. High adaptability: By updating the decay rate difference parameter after each control cycle, it can adapt to long-term changes in the atmospheric environment, and the close flight of the formation satellites ensures that the difference parameter is relatively stable.

[0061] 5. Simplified engineering implementation: The control strategy only needs to be based on... The algorithm for determining whether to ascend or descend an orbit is simple and easy to implement on satellite or on the ground.

[0062] 6. Strong industrial applicability: This application can be applied to various low-Earth orbit satellite formation missions, especially Earth observation, synthetic aperture radar interferometry, and formation flight demonstration and verification missions with strict requirements for configuration maintenance accuracy and fuel consumption. Because it takes into account the actual differences in aerodynamic drag in orbit, this application has better engineering applicability and operability, and can effectively extend the on-orbit service life of formation satellites.

[0063] According to a second specific embodiment of the present invention, such as Figure 4 As shown, the present invention provides a low-Earth orbit satellite formation maintenance device 100 that takes into account differences in aerodynamic drag, comprising: Module 110 is used to acquire the difference in the semi-major axis decay rate of the orbit between the reference satellite and the companion satellite in the formation. ;in, , The semi-major axis attenuation rate of the accompanying satellite, The semi-major axis attenuation rate of the reference satellite; Prediction module 120 is used to predict the difference in the semi-major axis decay rate of the orbit. And the initial semi-major axis difference Δa0, to predict the relative drift trajectory of the companion satellite relative to the reference satellite in the velocity direction, the relative drift trajectory including the change of drift amount over time; where Δa0 = a c -a s a c For the semi-major axis of the companion satellite, a s The semi-major axis of the reference satellite; The determining module 130 is used to maintain a threshold ΔL based on the relative drift trajectory and a preset configuration. max Determine the track control strategy; Execution module 140 is configured to, when the drift amount reaches or exceeds the configuration maintenance threshold ΔL max At that time, an orbit control strategy is executed to adjust the orbital altitude of the companion satellite; The reset module 150 is used to reset the semi-major axis difference Δa0 after adjusting the orbital altitude of the companion satellite, and return to the step of obtaining the difference in the orbital semi-major axis attenuation rate.

[0064] The following reference Figure 5 To describe an electronic device 200 according to this embodiment of the present application. Figure 5 The electronic device 200 shown is merely an example and should not impose any limitations on the functionality and scope of use of the embodiments of this application.

[0065] like Figure 5 As shown, the electronic device 200 is presented in the form of a general-purpose 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 storage unit 220 and processing unit 210), a display unit 240, etc.

[0066] The storage unit stores program code that can be executed by the processing unit 210, causing the processing unit 210 to perform the steps described in this specification according to various exemplary embodiments of this application. For example, the processing unit 210 can perform actions such as... Figure 1 The steps are shown in the figure.

[0067] 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 storage unit 2202, and may further include a read-only memory unit (ROM) 2203.

[0068] The storage unit 220 may also include a program / utility 2204 having a set (at least one) program module 2205, such program module 2205 including but not limited to: an operating system, one or more application programs, other program modules and program data, each or some combination of these examples may include an implementation of a network environment.

[0069] Bus 230 can represent one or more of several types of bus structures, including a memory cell bus or memory cell controller, a peripheral bus, a graphics acceleration port, a processing unit, or a local bus using any of the various bus structures.

[0070] Electronic device 200 can also communicate with one or more external devices 200' (e.g., keyboard, pointing device, Bluetooth device, etc.), enabling users to communicate with devices that interact with electronic device 200, and / or any device that allows electronic device 200 to communicate with one or more other computing devices (e.g., router, modem, etc.). This communication can be performed via input / output (I / O) interface 250. Furthermore, electronic device 200 can also communicate with one or more networks (e.g., local area network (LAN), wide area network (WAN), and / or public networks, such as the Internet) via network adapter 260. Network adapter 260 can communicate with other modules of electronic device 200 via bus 230. It should be understood that, although not shown in the figures, other hardware and / or software modules can be used in conjunction with 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.

[0071] Through the above description of the embodiments, those skilled in the art will readily understand that the exemplary embodiments described herein can be implemented by software or by combining software with necessary hardware.

[0072] Therefore, according to a fourth specific embodiment of the present invention, the present invention provides a computer-readable medium. For example... Figure 6 As shown, the technical solution according to the embodiments of the present invention can be embodied in the form of a software product. The software product can be stored in a non-volatile storage medium (such as a CD-ROM, USB flash drive, mobile hard drive, etc.) or on a network, and includes several instructions to cause a computing device (such as a personal computer, server, or network device, etc.) to execute the above-described method according to the embodiments of the present invention.

[0073] The software product may employ any combination of one or more readable media. A readable medium may be a readable signal medium or a readable storage medium. A readable storage medium may be, for example, but not limited to, an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or any combination thereof. More specific examples of readable storage media (a non-exhaustive list) include: an electrical connection having one or more wires, a portable disk, a hard disk, random access memory (RAM), read-only memory (ROM), erasable programmable read-only memory (EPROM or flash memory), optical fiber, portable compact disk read-only memory (CD-ROM), optical storage devices, magnetic storage devices, or any suitable combination thereof.

[0074] The computer-readable storage medium may include data signals propagated in baseband or as part of a carrier wave, carrying readable program code. 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, capable of transmitting, propagating, or transmitting programs for use by or in connection with an instruction execution system, apparatus, or device. The program code contained on the readable storage medium may be transmitted using any suitable medium, including but not limited to wireless, wired, optical fiber, RF, etc., or any suitable combination thereof.

[0075] Program code for performing the operations of this invention can be written in any combination of one or more programming languages, including object-oriented programming languages ​​such as Java and C++, and conventional procedural programming languages ​​such as C or similar languages. The program code can execute entirely on the user's computing device, partially on the user's device, as a standalone software package, partially on the user's computing device and partially on a remote computing device, or entirely on a remote computing device or server. In cases involving remote computing devices, the remote computing device can be connected to the user's computing device via any type of network, including a local area network (LAN) or a wide area network (WAN), or it can be connected to an external computing device (e.g., via the Internet using an Internet service provider).

[0076] The aforementioned computer-readable medium carries one or more programs, which, when executed by a device, cause the computer-readable medium to perform the functions of the first embodiment.

[0077] Those skilled in the art will understand that the above modules can be distributed in the device as described in the embodiments, or they can be modified accordingly and placed in one or more devices that are unique to this embodiment. The modules in the above embodiments can be combined into one module, or they can be further divided into multiple sub-modules.

[0078] Through the description of the above embodiments, those skilled in the art will readily understand that the exemplary embodiments described herein can be implemented by software or by combining software with necessary hardware. Therefore, the technical solutions of the embodiments of the present invention can be embodied in the form of a software product, which can be stored in a non-volatile storage medium (such as a CD-ROM, USB flash drive, portable hard drive, etc.) or on a network, including several instructions to cause a computing device (such as a personal computer, server, mobile terminal, or network device, etc.) to execute the methods according to the embodiments of the present invention.

[0079] The above are merely preferred embodiments of the present invention and are not intended to limit the present invention. Various modifications and variations can be made to the present invention by those skilled in the art. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A method for maintaining the low-Earth orbit satellite formation configuration considering differences in aerodynamic drag, characterized in that, The method includes: Obtain the difference in the semi-major axis decay rate between the reference satellite and the accompanying satellite in the formation. ;in, , The semi-major axis attenuation rate of the accompanying satellite, The semi-major axis attenuation rate of the reference satellite; Based on the difference in the semi-major axis attenuation rate of the orbit And the initial semi-major axis difference Δa0, to predict the relative drift trajectory of the companion satellite relative to the reference satellite in the velocity direction, the relative drift trajectory including the change of drift amount over time; where Δa0 = a c -a s a c For the semi-major axis of the companion satellite, a s The semi-major axis of the reference satellite; Based on the relative drift trajectory and the preset configuration maintenance threshold ΔL max Determine the track control strategy; When the drift amount reaches or exceeds the configuration maintenance threshold ΔL max At that time, an orbit control strategy is executed to adjust the orbital altitude of the companion satellite; After adjusting the orbital altitude of the companion satellite, reset the semi-major axis difference Δa0 and return to the step of obtaining the difference in the orbital semi-major axis attenuation rate.

2. The method according to claim 1, characterized in that, The relative drift trajectory and the preset configuration maintenance threshold ΔL are used as the basis for this. max Determine the orbital control strategy, including: when When the drift is greater than 0, monitor whether the companion satellite is behind the reference satellite. When the drift reaches or is about to reach the configuration maintenance threshold ΔL, max At that time, the accompanying satellite was subjected to orbit reduction control to drift it forward; when When the drift is less than 0, monitor whether the companion satellite is in front of the reference satellite. When the drift reaches or is about to reach the configuration maintenance threshold ΔL, max At that time, orbital ascent control was performed on the accompanying satellite to cause it to drift backward.

3. The method according to claim 2, characterized in that, The acquisition of the difference in the semi-major axis decay rate between the reference satellite and the accompanying satellite in the formation includes: During the set period of stable flight of the two satellites, based on the orbit determination results under uncontrolled conditions, linear fitting or piecewise calculation is performed on the semi-major axis data of the companion satellite and the reference satellite to obtain the semi-major axis attenuation rate of each satellite. The difference in the semi-major axis attenuation rate of the orbit is calculated based on the obtained semi-major axis attenuation rate of each satellite.

4. The method according to claim 3, characterized in that, The relative drift trajectory and the preset configuration maintenance threshold ΔL are used as the basis for this. max Determine the orbital control strategy, including: Based on the relative motion dynamics equations corrected by introducing the difference in the semi-major axis decay rate of the orbit, the relationship between the drift amount and time is obtained, and the extreme values ​​of the drift amount and the time when the extreme values ​​occur are calculated. Based on the extreme value and the configuration maintenance threshold ΔL max Calculate the required altitude difference between the two satellites after orbit control, so that the extreme value of the drift falls within the configuration maintenance threshold ΔL. max Within.

5. The method according to claim 4, characterized in that, The prediction of the relative drift trajectory of the companion satellite relative to the reference satellite in the velocity direction includes: Establish a model for the drift ΔL(t) as a function of time t: ； Where, when Δa0 and When the signs are opposite, the drift amount ΔL(t) has an extreme value, and the time tpeak when the extreme value occurs is: ； The magnitude of the extreme value ΔLpeak is: ； Where ΔL0 is the initial drift, n0 is the orbital angular velocity, and a is the semi-major axis of the orbit.

6. The method according to claim 5, characterized in that, The altitude difference between the two satellites after orbit control satisfies the following relationship: ； Where, ΔL max To maintain the configuration threshold, n is the angular velocity of the satellite's motion. In this formula, Δa0 is also the altitude difference between the two satellites after orbit control. The difference in attenuation rate between the companion satellite and the reference satellite.

7. The method according to claim 1, characterized in that, The configuration maintenance threshold ΔL max The configuration is set according to the mission requirements of the satellite formation.

8. A low-Earth orbit satellite formation maintenance device considering aerodynamic drag differences, characterized in that, The device includes: The acquisition module is used to obtain the difference in the semi-major axis decay rate of the orbit between the reference satellite and the accompanying satellite in the formation. ;in, , The semi-major axis attenuation rate of the accompanying satellite, The semi-major axis attenuation rate of the reference satellite; The prediction module is used to predict the difference in the semi-major axis decay rate of the orbit. And the initial semi-major axis difference Δa0, to predict the relative drift trajectory of the companion satellite relative to the reference satellite in the velocity direction, the relative drift trajectory including the change of drift amount over time; where Δa0 = a c -a s a c For the semi-major axis of the companion satellite, a s The semi-major axis of the reference satellite; The determination module is used to maintain a threshold ΔL based on the relative drift trajectory and a preset configuration. max Determine the track control strategy; An execution module is configured to respond when the drift amount reaches or exceeds the configuration maintenance threshold ΔL. max At that time, an orbit control strategy is executed to adjust the orbital altitude of the companion satellite; The reset module is used to reset the semi-major axis difference Δa0 after adjusting the orbital altitude of the companion satellite, and return to the step of obtaining the difference in the orbital semi-major axis attenuation rate.

9. An electronic device, characterized in that, include: One or more processors; Storage device for storing one or more programs; When the one or more programs are executed by the one or more processors, the one or more processors implement the method as described in any one of claims 1-7.

10. A computer-readable storage medium, characterized in that, The computer-readable storage medium includes a computer program or instructions that, when executed on a computer, cause the computer to perform the method as described in any one of claims 1-7.

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