Constellation phase keeping method, constellation phase keeping device and electronic equipment

Abstract

The invention provides a constellation phase keeping method and device and electronic equipment, and the method comprises the steps: obtaining the mean root semi-major axis of each satellite in a target constellation in a target epoch, and selecting a satellite with the maximum mean root semi-major axis as a reference satellite; and constellation phase keeping is carried out based on the reference satellite. According to the scheme, only a relatively small period difference needs to be adjusted, the overall orbit adjustment amplitude and frequency can be reduced, and then fuel consumption and the complexity of orbit control operation are reduced; the influence of atmospheric resistance on the reference satellite is relatively small, so that the condition that other satellites frequently perform orbit control due to frequent change of the reference satellite orbit can be reduced; constellation configuration confusion caused by instability of the reference satellite itself is reduced in the face of various perturbation force and environment changes; the operation period of the reference satellite is long, the orbit is relatively stable, and too complex orbit change conditions do not need to be considered when an orbit control strategy is formulated, so that an orbit control algorithm and operation are easier to implement.

Description

Technical Field

[0001] The present application relates to the technical field of constellation configuration maintenance, and in particular to a constellation phase maintenance method, device and electronic equipment. Background Art

[0002] Maintaining the configuration of a low-orbit satellite constellation is fundamental to ensuring the proper functioning of its services. Due to orbital deviations, the relative positions of satellites in the constellation experience varying perturbations, leading to a diverging constellation configuration. This impacts the constellation's coverage characteristics and service operations. Therefore, continuous configuration maintenance is essential after constellation deployment. Maintaining the constellation configuration is one of the technical challenges that must and will continue to be faced in the development of low-orbit satellite internet.

[0003] Constellation configuration is reflected in the semi-major axis, eccentricity, inclination, right ascension of the ascending node, and phase. The inclination and right ascension of the ascending node drift by the same amount for each orbital plane, and their impact on the constellation configuration while in orbit is essentially negligible. For near-circular orbits, the long-term drift of eccentricity is approximately zero, making its impact negligible. Configuration changes for low-Earth orbit constellations are primarily due to phase changes caused by changes in the semi-major axis due to atmospheric drag. Therefore, maintaining constellation configuration primarily involves phase maintenance in engineering.

[0004] Constellation phase hold can be divided into two methods: absolute phase hold and relative phase hold. Absolute phase hold controls the satellite at the nominal position in a certain coordinate system, usually requiring compensation for orbital semi-major axis and phase drift. Relative phase hold only maintains the relative position relationship between all satellites, controlling the inter-satellite phase difference within the maximum allowable deviation range.

[0005] Existing constellation phase-keeping methods usually use the satellite with the smallest cumulative phase deviation relative to each satellite as the reference satellite. From the perspective of the entire satellite life cycle, the total fuel consumption of this method is relatively high. Currently, there is an urgent need for a constellation phase-keeping method with low total fuel efficiency. Summary of the Invention

[0006] This specification provides a constellation phase maintenance method, device, and electronic device to overcome the drawbacks of existing constellation phase maintenance methods.

[0007] A first aspect of the present specification provides a constellation phase maintenance method, comprising: obtaining the mean root semi-major axis of each satellite in a target constellation at a target epoch, and selecting the satellite with the largest mean root semi-major axis as a reference satellite; and performing constellation phase maintenance based on the reference satellite.

[0008] In some embodiments, constellation phase maintenance is performed based on the reference satellite, including: calculating a virtual phase difference of each satellite in the target constellation relative to the reference satellite, where the virtual phase difference is the difference between an actual phase difference and a nominal phase difference; and maintaining the constellation configuration when the maximum value of the virtual phase difference of each satellite is greater than or equal to a phase control threshold.

[0009] In some embodiments, the method further includes: if the maximum value of the virtual phase difference of each satellite is less than a phase control threshold, not maintaining the constellation configuration.

[0010] In some embodiments, maintaining the constellation configuration includes controlling the target satellite movement in the following manner: determining a target value of the semi-major axis change of the target satellite; calculating the orbital angular velocity of the reference satellite; calculating a velocity increment based on the orbital angular velocity of the reference satellite; calculating a control moment based on the velocity increment; calculating control parameters of the thruster based on the velocity increment and various parameters of the thruster; and controlling the thruster movement based on the control parameters.

[0011] In some embodiments, determining a target value of the semi-major axis variation of a target satellite includes: obtaining a semi-major axis control error of the target satellite; calculating a target value of the semi-major axis variation of the target satellite, wherein the target value of the semi-major axis variation is the difference between the semi-major axis of a reference satellite and the semi-major axis of the target satellite, and then summing the difference with the semi-major axis control error.

[0012] In some embodiments, after calculating the target value of the semi-major axis variation of the target satellite, the method further includes: correcting the target value of the semi-major axis variation according to the semi-major axis variation rate of the target satellite.

[0013] In some embodiments, the target value of the semi-major axis variation is corrected based on the semi-major axis variation rate of the target satellite, including: when the semi-major axis variation rate of the target satellite is less than the semi-major axis variation rate of the reference satellite, calculating a correction amount based on the semi-major axis and semi-major axis variation rate of the reference satellite, the semi-major axis variation rate of the target satellite, the phase control threshold, the virtual phase difference of the target satellite, and the orbital angular velocity of the reference satellite; and using the correction amount to correct the target value of the semi-major axis variation of the target satellite.

[0014] In some embodiments, the correction amount is calculated as follows: Among them, Δa represents the correction amount, a k is the square root semi-major axis of the reference satellite, is the mean root mean square semi-major axis rate of change of the target satellite, is the root mean square semi-major axis change rate of the reference satellite, x is the phase difference threshold, η is the control coefficient, Δλ i is the virtual phase difference of the target satellite, and n is the orbital angular velocity of the reference satellite.

[0015] In some embodiments, the semi-major axis change rate of the target satellite is calculated in the following manner: the square root of the semi-major axis of the target satellite at the time of the current orbit determination is subtracted from the square root of the semi-major axis at the time of the last orbit determination to obtain the semi-major axis change amount; the time interval between the last orbit determination and the current orbit determination is determined; the semi-major axis change amount is divided by the time interval to obtain the semi-major axis change rate of the target satellite.

[0016] In some embodiments, the velocity increment is calculated based on the orbital angular velocity of the reference satellite, including using the following formula to calculate the velocity increment: ΔV = n / 2*a c , where ΔV = represents the velocity increment, n is the orbital angular velocity of the reference satellite, and a c is the target value of the semi-major axis variation.

[0017] In some embodiments, the control moment is calculated based on the speed increment, including: when the speed increment is greater than 0, setting the control moment at the apogee, and Tc=(π-M) / n+T now , where Tc is the control time, M is the mean anomaly of the target satellite at the current time, n is the orbital angular velocity of the reference satellite, and T now is the current time; and / or, when the velocity increment is less than 0, the control time is set at the perigee of the orbit, then the phase difference between the current and perigee is 2π-M, then the control time Tc=(2π-M) / n+T now , where n is the orbital angular velocity of the reference satellite, T now For the current moment.

[0018] In some embodiments, the method further includes: calculating the virtual phase difference of each satellite in the target constellation relative to the reference satellite, wherein the virtual phase difference is the difference between the actual phase difference and the nominal value of the phase difference; calculating the virtual semi-major axis of each satellite in the target constellation relative to the reference satellite, wherein the virtual semi-major axis is the difference between the mean square root semi-major axis of the satellite and the mean square root semi-major axis of the reference satellite; and drawing a constellation virtual phase drift situation diagram of the target constellation at the target epoch based on the virtual phase difference and virtual semi-major axis of each satellite relative to the reference satellite; wherein the constellation virtual phase drift situation diagram includes a horizontal coordinate representing the phase difference and a vertical coordinate identifying the mean square root semi-major axis.

[0019] A second aspect of this specification provides a constellation phase holding device, including: an acquisition unit, used to obtain the mean root semi-major axis of each satellite in the target constellation at the target epoch, and select the satellite with the largest mean root semi-major axis as the reference satellite; a processing unit, used to perform constellation phase holding based on the reference satellite.

[0020] In some embodiments, the processing unit includes: a first calculation subunit, used to calculate the virtual phase difference of each satellite in the target constellation relative to the reference satellite, where the virtual phase difference is the difference between the actual phase difference and the nominal value of the phase difference; a first determination subunit, used to maintain the constellation configuration when the maximum value of the virtual phase difference of each satellite is greater than or equal to the phase control threshold.

[0021] In some embodiments, the apparatus further includes: a second determining subunit, configured to not maintain the constellation configuration when the maximum value of the virtual phase difference of each satellite is less than a phase control threshold.

[0022] In some embodiments, the first determination subunit includes: a third determination subunit, used to determine the target value of the semi-major axis change of the target satellite; a second calculation subunit, used to calculate the orbital angular velocity of the reference satellite; a third calculation subunit, used to calculate the velocity increment based on the orbital angular velocity of the reference satellite; a fourth calculation subunit, used to calculate the control moment based on the velocity increment; a fifth calculation subunit, used to calculate the control parameters of the thruster based on the velocity increment and various parameters of the thruster; and a control subunit, used to control the thruster action based on the control parameters.

[0023] In some embodiments, the third determination subunit includes: a first acquisition subunit, used to obtain the semi-major axis control error of the target satellite; a sixth calculation subunit, used to calculate the target value of the semi-major axis change of the target satellite, the target value of the semi-major axis change is the difference between the semi-major axis of the reference satellite and the semi-major axis of the target satellite, and then the difference is summed with the semi-major axis control error.

[0024] In some embodiments, the third determining subunit further includes: a correction subunit, configured to correct the target value of the semi-major axis variation according to the semi-major axis variation rate of the target satellite.

[0025] In some embodiments, the correction subunit includes: a seventh calculation subunit, used to calculate the correction amount based on the semi-major axis and semi-major axis change rate of the reference satellite, the semi-major axis change rate of the target satellite, the phase control threshold, the virtual phase difference of the target satellite, and the orbital angular velocity of the reference satellite when the semi-major axis change rate of the target satellite is less than the semi-major axis change rate of the reference satellite; and a correction subunit, used to use the correction amount to correct the target value of the semi-major axis change amount of the target satellite.

[0026] In some embodiments, the correction amount is calculated as follows: Among them, Δa represents the correction amount, a k is the square root semi-major axis of the reference satellite, is the mean root mean square semi-major axis rate of change of the target satellite, is the root mean square semi-major axis change rate of the reference satellite, x is the phase difference threshold, η is the control coefficient, Δλ i is the virtual phase difference of the target satellite, and n is the orbital angular velocity of the reference satellite.

[0027] In some embodiments, the semi-major axis change rate of the target satellite is calculated in the following manner: the square root of the semi-major axis of the target satellite at the time of the current orbit determination is subtracted from the square root of the semi-major axis at the time of the last orbit determination to obtain the semi-major axis change amount; the time interval between the last orbit determination and the current orbit determination is determined; the semi-major axis change amount is divided by the time interval to obtain the semi-major axis change rate of the target satellite.

[0028] In some embodiments, the third calculation subunit calculates the velocity increment using the following formula: ΔV=n / 2*a c , where ΔV = represents the velocity increment, n is the orbital angular velocity of the reference satellite, and a c is the target value of the semi-major axis variation.

[0029] In some embodiments, the fourth calculation subunit includes: a fourth determination subunit, configured to set the control time at the apogee when the speed increment is greater than 0, and Tc=(π-M) / n+T now , where Tc is the control time, M is the mean anomaly of the target satellite at the current time, n is the orbital angular velocity of the reference satellite, and T now is the current moment; and / or, a fifth determining subunit for setting the control moment at the perigee of the orbit when the velocity increment is less than 0, then the phase difference between the current and perigee is 2π-M, then the control moment Tc=(2π-M) / n+T now , where n is the orbital angular velocity of the reference satellite, T now For the current moment.

[0030] In some embodiments, the device also includes: an eighth calculation subunit, used to calculate the virtual phase difference of each satellite in the target constellation relative to the reference satellite, wherein the virtual phase difference is the difference between the actual phase difference and the nominal value of the phase difference; a ninth calculation subunit, used to calculate the virtual semi-major axis of each satellite in the target constellation relative to the reference satellite, wherein the virtual semi-major axis is the difference between the mean square root semi-major axis of the satellite and the mean square root semi-major axis of the reference satellite; a drawing unit, used to draw a constellation virtual phase drift situation diagram of the target constellation at the target epoch based on the virtual phase difference and virtual semi-major axis of each satellite relative to the reference satellite; wherein the constellation virtual phase drift situation diagram includes a horizontal coordinate representing the phase difference and a vertical coordinate identifying the mean square root semi-major axis.

[0031] A third aspect of this specification provides an electronic device, comprising: a memory and a processor, wherein the processor and the memory are communicatively connected to each other, the memory stores computer instructions, and the processor implements the constellation phase maintenance method described in any one of the first aspects by executing the computer instructions.

[0032] A fourth aspect of this specification provides a computer storage medium, wherein the computer storage medium stores computer program instructions, and when the computer program instructions are executed, the steps of the constellation phase maintenance method described in any one of the first aspects are implemented.

[0033] A fifth aspect of this specification provides a computer program product, comprising a computer program, which, when executed by a processor, implements the steps of the constellation phase maintenance method described in any one of the first aspects.

[0034] Compared to the prior art, the constellation phase hold method, apparatus, and electronic device provided in this specification select the satellite with the largest mean root semi-major axis in the constellation as the reference satellite and perform phase hold based on this reference satellite, which has the following advantages:

[0035] 1. Reduce the need for overall orbital adjustments and lower fuel consumption: According to Kepler's third law, the square of the satellite's operating period is proportional to the cube of its semi-major axis, that is, the larger the semi-major axis, the longer the satellite's operating period. Using the satellite with the largest square root semi-major axis as the reference satellite means selecting a satellite with a relatively long operating period. When other satellites maintain a relative phase relationship with it, due to their relatively short operating periods, only relatively small period differences need to be adjusted during orbital adjustments. Compared to selecting a satellite with a smaller semi-major axis (shorter operating period) as the reference satellite, the overall orbital adjustment amplitude and frequency can be reduced, thereby reducing fuel consumption and the complexity of orbit control operations.

[0036] 2. Reduce the impact of atmospheric drag: Generally speaking, orbital altitude is related to the semi-major axis. Satellites with larger semi-major axes typically have higher orbital altitudes. The higher the orbital altitude, the lower the atmospheric density, and the less atmospheric drag the satellite experiences. The satellite with the largest square root semi-major axis is selected as the reference satellite. This satellite is less affected by atmospheric drag and its orbital changes are more stable. This results in smaller orbital changes for the reference satellite during relative phase hold, making it easier for other satellites to maintain their relative relationship with it. This reduces the need for frequent orbital control for other satellites due to frequent changes in the reference satellite's orbit, and facilitates the stable maintenance of the entire constellation configuration.

[0037] 3. Enhance constellation system stability: The satellite with the largest semi-major axis (PSA) has a more representative and stable orbital profile. Using it as a reference satellite, the other satellites adjust their relative phases around it, helping to maintain a relatively stable constellation configuration. During constellation operation, facing various perturbations and environmental changes, this relatively stable reference helps maintain overall constellation functionality and performance, reducing disruptions to the constellation configuration caused by inherent instability of the reference satellite, and ultimately improving constellation system reliability and stability.

[0038] 4. Simplified Orbit Control Strategy: Because the satellite with the largest RPSA has a longer operating period and a relatively stable orbit, orbit control strategies can be simplified. For other satellites, adjustments only need to be made to the relative phase and period differences with the reference satellite, without having to consider complex orbital variations. This simplifies orbit control algorithms and operations, reducing the difficulty and cost of constellation operations management. BRIEF DESCRIPTION OF THE DRAWINGS

[0039] In order to more clearly illustrate the implementation methods of the present application or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the implementation methods or the description of the prior art. Obviously, the drawings described below are only some implementation methods recorded in this application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative labor.

[0040] Figure 1 A flowchart of the constellation phase maintenance method provided in this specification;

[0041] Figure 2 Another flowchart of the constellation phase maintenance method provided in this specification;

[0042] Figure 3 This is another flowchart of the constellation phase maintenance method provided in this specification;

[0043] Figure 4 This is another flowchart of the constellation phase maintenance method provided in this specification;

[0044] Figure 5 A schematic diagram of a process for maintaining constellation configuration provided in this manual;

[0045] Figure 6 Another flowchart for maintaining constellation configuration provided for this manual;

[0046] Figure 7 A schematic diagram of another process for maintaining constellation configuration provided in this specification;

[0047] Figure 8 This is a structural block diagram of the constellation phase holding device provided in this specification;

[0048] Figure 9 This is a schematic diagram of the structure of the electronic device provided in this manual. DETAILED DESCRIPTION

[0049] In order to enable those skilled in the art to better understand the technical solutions in this application, the following will clearly and completely describe the technical solutions in the embodiments of this application in conjunction with the drawings in the embodiments of this application. Obviously, the described embodiments are only part of the embodiments of this application, not all of the embodiments. Based on the embodiments of this application, all other embodiments obtained by ordinary technicians in this field without making creative efforts should fall within the scope of protection of this application.

[0050] Prior to proposing the constellation phase hold method provided in this specification, the inventors conducted the following research on a relative phase hold method: selecting a satellite in the constellation as a reference satellite to minimize the cumulative value of the actual phase differences of all other satellites in the constellation relative to the reference satellite, thereby achieving the lowest orbit control frequency. The specific steps are as follows: 1. Calculating the actual phase differences of the remaining satellites relative to the reference satellite using each satellite in the constellation as the reference satellite; 2. Calculating the cumulative value of the actual phase differences of the remaining satellites when satellite i is used as the reference satellite; 3. Selecting the satellite with the smallest cumulative value and the least critical threshold as reference satellite k, and using orbit extrapolation and curve fitting based on the phase difference variation pattern of the measured orbit data to determine the satellites that need to be phase held and estimate the phase hold time; 4. When the phase difference of satellite i relative to reference satellite k approaches the threshold, the orbit altitude is raised or lowered based on the phase angle drift direction and the reference satellite as a standard to implement orbit control on satellite i; 5. Calculating the orbit control strategy to align the semi-major axis of the satellite with the reference satellite k, and changing the phase drift direction between satellites by changing the orbit semi-major axis to achieve relative phase adjustment. The above i and k represent different satellites.

[0051] Through the above research, it is found that the relative phase maintenance method has the following disadvantages: 1. The selection of the reference satellite leads to long-term fuel waste. Although the fuel consumption is minimized in the current single control, the total fuel consumption is not minimized from the perspective of the satellite's entire life. 2. The selection process of the reference satellite is relatively cumbersome, with a large amount of calculations, a long time required for calculation, and a large amount of computer memory. When the number of satellites in the constellation is n, the time complexity of the algorithm is O(n 2 ), the space complexity is O(n 2); 3. There is still room for optimization in the orbit control strategy. Taking the benchmark satellite as the standard, each orbit control will raise or lower the orbit semi-major axis, and the control direction will be accelerated or decelerated. From the perspective of long-term operation and control, this increases the operational complexity and error risk of the operation and control personnel.

[0052] To overcome the drawbacks of the aforementioned constellation phase-holding method, this specification provides a constellation phase-holding method, which belongs to the relative phase-holding method. This method can be used for phase-holding in low-orbit mega-constellations. The characteristics of low-orbit mega-constellations include: 1. A large number of satellites (hundreds, thousands, or even tens of thousands), which can cover a wider area and provide more stable and powerful service capabilities. However, the large number of satellites also leads to a more complex configuration; 2. Compared with high-orbit satellites, low-orbit satellites are closer to the Earth. The signal transmission delay is small and the data transmission rate is high, which can achieve more accurate Earth observation and higher-resolution image acquisition, providing more detailed data for remote sensing, geographic information systems and other fields.

[0053] like Figure 1 As shown, the constellation phase maintaining method provided in this specification includes the following S10 and S20.

[0054] S10: Obtain the mean root semi-major axis of each satellite in the target constellation at the target epoch, and select the satellite with the largest mean root semi-major axis as the reference satellite.

[0055] An epoch is a specific point in time used in astronomy and satellite orbit calculations as a time reference for describing a satellite's position and status. Satellite orbital parameters vary at different times, so a specific point in time is necessary to determine a satellite's position and status. The term "target epoch" used in various sections of this manual refers to the same point in time.

[0056] like Figure 2 As shown, S10 may include the following S11 to S13.

[0057] S11: The latest orbit determination results of each satellite in the constellation may be obtained first, and the latest orbit determination results may be extrapolated to the same epoch T.

[0058] Satellite orbit determination refers to the determination of the satellite's orbital parameters in space, including the orbital shape (such as the semi-major axis and eccentricity), the orientation of the orbital plane (such as the inclination and right ascension of the ascending node), and the satellite's position in orbit (such as the argument of perigee and true anomaly). These orbital parameters can be calculated through various observational methods (such as radar and laser ranging observations from ground-based observation stations, as well as data from sensors carried by the satellite itself), combined with orbital mechanics and mathematical models, to obtain the satellite's orbit determination results.

[0059] Orbital elements are a set of parameters that describe a satellite's orbit. Common orbital elements include the semi-major axis, eccentricity, inclination, right ascension of the ascending node, argument of perigee, and true anomaly. These parameters can be used to determine a satellite's position in orbit, as well as its shape, orientation, and other information.

[0060] Extrapolation is to calculate the orbital parameters of the satellite at epoch T based on the laws of satellite motion (such as Kepler's law and orbital dynamics models that take into account the influence of various perturbations) using known orbit determination results and time information.

[0061] S12: Convert the instantaneous root of each satellite in the constellation into a square root, and obtain the square root semi-major axis a0~a of each satellite n 0~n represent the n+1 satellites in the constellation.

[0062] Instantaneous orbital roots (instantaneous orbital elements), or instantaneous orbital parameters, reflect the satellite's orbital state at a given moment. Because satellites are subject to various perturbations during their orbit (such as the Earth's non-spherical gravity, atmospheric drag, and solar radiation pressure), their instantaneous orbital elements change over time. This means that the satellite's instantaneous orbital elements vary at different times.

[0063] The flat root is the average orbital element. To more concisely and intuitively describe a satellite's orbital characteristics, after accounting for the long-term and long-period effects of perturbations, the instantaneous orbital elements are averaged using mathematical methods. The resulting relatively stable orbital element is the average orbital element, or flat root. The flat root eliminates high-frequency variations caused by perturbations to a certain extent, better reflecting the average state and long-term characteristics of the satellite's orbit.

[0064] Converting instantaneous roots to flat roots involves processing instantaneous orbital elements using mathematical models and algorithms. For example, when considering the Earth's non-spherical gravitational perturbations, complex perturbation theory and computational methods are required to correct and average the various parameters in the instantaneous orbital elements to obtain flat roots. This process aims to remove short-term variations caused by perturbations and highlight the secular and long-period characteristics of the orbit.

[0065] S13: Select the satellite with the largest square root semi-major axis in the constellation as the reference satellite k. k is the identifier of the reference satellite.

[0066] S20: Perform constellation phase maintenance based on the reference satellite.

[0067] This application selects the satellite with the largest semi-major axis in the constellation as the reference satellite, and performs phase maintenance based on this reference satellite, which has the following advantages:

[0068] 1. Reduce the need for overall orbital adjustments: According to Kepler's third law, the square of the satellite's operating period is proportional to the cube of its semi-major axis, that is, the larger the semi-major axis, the longer the satellite's operating period. Using the satellite with the largest square root semi-major axis as the reference satellite means selecting a satellite with a relatively long operating period. When other satellites maintain a relative phase relationship with it, due to their relatively short operating periods, only relatively small period differences need to be adjusted during orbital adjustments. Compared to selecting a satellite with a smaller semi-major axis (shorter operating period) as the reference satellite, the overall orbital adjustment amplitude and frequency can be reduced, thereby reducing fuel consumption and the complexity of orbit control operations.

[0069] 2. Reduce the impact of atmospheric drag: Generally speaking, orbital altitude is related to the semi-major axis. Satellites with larger semi-major axes typically have higher orbital altitudes. The higher the orbital altitude, the lower the atmospheric density, and the less atmospheric drag the satellite experiences. The satellite with the largest square root semi-major axis is selected as the reference satellite. This satellite is less affected by atmospheric drag and its orbital changes are more stable. This results in smaller orbital changes for the reference satellite during relative phase hold, making it easier for other satellites to maintain their relative relationship with it. This reduces the need for frequent orbital control for other satellites due to frequent changes in the reference satellite's orbit, and facilitates the stable maintenance of the entire constellation configuration.

[0070] 3. Enhance constellation system stability: The satellite with the largest semi-major axis (PSA) has a more representative and stable orbital profile. Using it as a reference satellite, the other satellites adjust their relative phases around it, helping to maintain a relatively stable constellation configuration. During constellation operation, facing various perturbations and environmental changes, this relatively stable reference helps maintain overall constellation functionality and performance, reducing disruptions to the constellation configuration caused by inherent instability of the reference satellite, and ultimately improving constellation system reliability and stability.

[0071] 4. Simplified Orbit Control Strategy: Because the satellite with the largest RPSA has a longer operating period and a relatively stable orbit, orbit control strategies can be simplified. For other satellites, adjustments only need to be made to the relative phase and period differences with the reference satellite, without having to consider complex orbital variations. This simplifies orbit control algorithms and operations, reducing the difficulty and cost of constellation operations management.

[0072] In some embodiments, as Figure 3 As shown, S20 includes the following S21 and S22.

[0073] S21: Calculate the virtual phase difference of each satellite in the target constellation relative to the reference satellite, where the virtual phase difference is the difference between the actual phase difference and the nominal phase difference value.

[0074] In satellite orbits, phase angles describe the relative positions of satellites in their orbits. Imagine a satellite's orbit as a circular motion, with different satellite positions on this circle corresponding to different phase angles. When multiple satellites are in operation, the phase angle relationship between them is critical to the proper operation and functionality of the constellation. For example, in a communications constellation, a specific phase relationship facilitates signal relay transmission and the proper division of coverage areas.

[0075] Phase difference, or phase angle deviation, is a phenomenon affecting satellites as they orbit in space. These factors, such as the Earth's non-spherical gravity, atmospheric drag, solar radiation pressure, and the gravitational pull of other celestial bodies, can cause deviations between the satellites' actual orbits and their ideal orbits, leading to changes in the relative phase angles between the satellites.

[0076] The "actual phase difference" mentioned above refers to the phase difference of each satellite relative to the reference satellite. For each satellite in the constellation, the change in phase angle between it and the reference satellite is calculated. Through precise orbital calculations and measurements, the orbital parameters of each satellite at different times are obtained, and its phase angle relative to the reference satellite is determined. The phase angles at different times are then compared to calculate the difference in phase angles, which is the actual phase difference.

[0077] For example, assuming that at a certain moment the phase angle of satellite A relative to the reference satellite is θ1, and after a period of time, the phase angle of satellite A changes to θ2, then the actual phase difference of satellite A relative to the reference satellite is |θ2-θ1|.

[0078] The nominal phase difference value refers to the ideal phase difference that each satellite should maintain relative to a reference satellite, pre-determined during the constellation design phase based on the constellation's functional requirements (such as communication coverage and observation missions). This nominal phase difference value is a theoretical standard used to guide the normal operation and functional implementation of the constellation.

[0079] The virtual phase difference is the difference between the actual phase difference and the nominal phase difference. It reflects the deviation between the actual operating state of each satellite and its ideal design state. For example, if a satellite's actual phase difference relative to a reference satellite is 45°, and the nominal design phase difference is 30°, then the virtual phase difference of that satellite relative to the nominal phase determined by the reference satellite's phase is 45° - 30° = 15°.

[0080] S22: When the maximum value of the virtual phase difference of each satellite is greater than or equal to the phase control threshold, maintain the constellation configuration.

[0081] The phase control threshold can be predetermined. For example, a phase difference threshold can be determined. When the virtual phase difference of a satellite exceeds the phase difference threshold, it indicates that the phase of the satellite needs to be adjusted. If the virtual phase difference of a satellite is less than the phase difference threshold, no phase adjustment is required.

[0082] In some cases, for some satellites, it may be necessary to adjust their phase even if the phase difference does not reach the phase difference threshold. To this end, based on the phase difference threshold, the satellite's control coefficient η (η∈(0,1]) can be further determined. The control coefficient is a manually set parameter, and its value depends on factors such as the specific design requirements of the constellation, the performance of the satellite, and the expectation of constellation configuration stability. The control coefficient can be different for each satellite.

[0083] The phase control threshold is a standard value used to determine whether the satellite's current phase state exceeds the permitted range. If no control coefficient is set, the phase control threshold is the phase difference threshold. If both the phase difference threshold and the control coefficient are set, the phase control threshold can be the product of the phase difference threshold and the control coefficient. Specifically, the phase difference threshold is x*η, where x represents the phase difference threshold, η represents the control coefficient, and * represents the product operation.

[0084] Maintaining the constellation configuration involves using various technical means and control strategies to maintain the predetermined relative positions and overall layout of satellites within the constellation during its operation, ensuring the constellation can continuously and stably achieve its intended functions and performance indicators. This involves activating the satellites' position control actuators (such as thrusters) to change their phases.

[0085] In some embodiments, the virtual semi-major axis of each satellite in the target constellation relative to the reference satellite can also be calculated. The virtual semi-major axis is the difference between the plane root semi-major axis of the satellite and the plane root semi-major axis of the reference satellite. Then, a virtual phase drift situation map of the target constellation can be drawn based on the virtual semi-major axis and the above-mentioned virtual phase difference. Specifically, a virtual drift reference coordinate system of the target constellation can be constructed based on the constellation design configuration and the nominal value of the phase difference. The abscissa of this virtual drift reference coordinate system is the satellite motion trajectory, which is represented by the phase difference. The positive value on the abscissa indicates the direction of satellite motion; the ordinate is the plane root semi-major axis. The positive value of the ordinate points to the sky. The origin of the virtual drift reference coordinate system represents the reference constellation. By marking the virtual phase difference and virtual semi-major axis of each satellite in the virtual drift reference coordinate system, a virtual phase drift situation map of the target constellation at the target epoch can be obtained. Figure 4 A schematic diagram of the virtual phase drift situation diagram.

[0086] In some embodiments, as Figure 5As shown, the above S23 can maintain the constellation configuration in the manner described in the following S41 to S46.

[0087] S41: Determine the target value of the semi-major axis variation of the target satellite.

[0088] The target value of the change in the semi-major axis of the target satellite refers to how much the semi-major axis of the target satellite needs to change in order to maintain the phase relationship between the target satellite and the reference satellite to meet the control requirements.

[0089] The target change in the semi-major axis of the target satellite can be determined in the operation of maintaining the constellation configuration according to the following:

[0090] 1. Constellation Design Requirements: During the constellation design phase, the satellite orbital parameters are clearly defined, including the relative positions, spacing, and coverage of the satellites. These design requirements determine the ideal orbital state for each satellite. For example, in a communications satellite constellation, to achieve seamless global coverage and avoid signal interference, satellites must maintain a specific relative position. The semi-major axis is a key parameter that influences the altitude and period of the satellite's orbit. Based on the constellation's coverage and communication requirements, the ideal value of each satellite's semi-major axis can be calculated. By comparing the target satellite's current semi-major axis with the ideal value, the target value for the change in the semi-major axis can be determined.

[0091] 2. Relative position and phase relationship: The relative position and phase relationship between satellites in a constellation is crucial to maintaining the constellation configuration. If there is a deviation in the relative position or phase between the target satellite and the reference satellite (or other reference satellite), it is necessary to correct it by adjusting the semi-major axis. For example, in satellite formation flight, satellites need to maintain a specific relative distance and phase difference. When it is found that the phase of the target satellite is ahead or behind the satellite, the semi-major axis change required to restore the correct phase relationship can be calculated based on the principle of orbital dynamics. Generally speaking, a slight change in the semi-major axis will cause a change in the satellite's orbital period, thereby affecting the satellite's speed and position in orbit, and then adjusting the relative phase between satellites.

[0092] 3. Influence of orbital perturbations: In space, satellites are affected by a variety of orbital perturbations, such as the unevenness of the Earth's gravitational field, solar radiation pressure, atmospheric drag (for low-orbit satellites), etc. These perturbations will cause the satellite's orbital parameters to gradually change, including the semi-major axis. In order to offset the impact of orbital perturbations and maintain the constellation configuration, it is necessary to determine the adjustment amount of the target satellite's semi-major axis based on the analysis and prediction of the orbital perturbations. For example, for low-orbit satellites, atmospheric drag will gradually reduce the satellite's orbital altitude and the semi-major axis. By modeling and calculating atmospheric drag, the change in the semi-major axis within a certain period of time can be predicted, and then the target value of the semi-major axis change can be determined based on this predicted value, so that orbital adjustments can be made in a timely manner.

[0093] 4. Mission Requirements: Different satellite missions have different requirements for constellation configuration. For example, in Earth observation missions, to achieve continuous or high-resolution observation of a specific area, it may be necessary to adjust the satellite's orbital parameters, including its semi-major axis. Based on the specific requirements of the mission, such as the scope of the observation area, observation frequency, and resolution requirements, the appropriate value of the target satellite's semi-major axis can be calculated, and the target value of the semi-major axis change can be determined.

[0094] 5. Mathematical Modeling and Simulation Analysis: Utilizing mathematical models of orbital dynamics and celestial mechanics, we accurately model and simulate the satellite's orbital motion. By inputting various parameters, such as the satellite's initial orbital parameters, perturbation factors, and control strategies, we simulate the satellite's orbital changes under different circumstances. Based on the simulation results, we can evaluate the impact of different semi-major axis changes on the constellation configuration and determine the optimal target semi-major axis change. Furthermore, we can calibrate and optimize the mathematical model based on actual monitoring data to improve the accuracy of the target value.

[0095] Determining the target value of the semi-major axis change of the target satellite is a complex process, which requires comprehensive consideration of multiple factors such as constellation design, relative position relationship, orbital perturbation, mission requirements, etc., and the use of mathematical models and simulation analysis tools for precise calculation and optimization.

[0096] In some cases, the impact of the above 3-5 aspects on the “target change of the target satellite semi-major axis” can be expressed by the control error. Figure 6 As shown, S41 can determine the target value of the semi-major axis variation of the target satellite by the following methods described in S411 and S412.

[0097] S411: Obtain the semi-major axis control error of the target satellite.

[0098] S412: Calculate the target value of the semi-major axis variation of the target satellite, where the target value of the semi-major axis variation is the difference between the semi-major axis of the reference satellite and the semi-major axis of the target satellite, and then sum the difference with the semi-major axis control error.

[0099] For example, for target satellite i, its current square root semi-major axis is a i , the square root semi-major axis of the reference satellite is a k , the semi-major axis control error is a m , then the target value a of the semi-major axis variation is c =a k -a i +a mak-ai is the negative of the virtual semi-major axis of the i-th satellite. This target value is used to compensate for the virtual semi-major axis. For example, if the RPSI of the i-th satellite is 7 and the RPSI of the reference satellite is 10, the virtual semi-major axis of the i-th satellite is -3. Therefore, ignoring control errors, the target value for the change in the semi-major axis is 3.

[0100] In some cases, after determining the target value of the semi-major axis variation in the above manner, it is still necessary to calibrate the target value of the semi-major axis variation. Figure 6 As shown, after S412, the process further includes S413: correcting the target value of the semi-major axis variation according to the semi-major axis variation rate of the target satellite. That is, correction can be performed according to the "semi-major axis variation rate" of the target satellite.

[0101] The semi-major axis change rate of the target satellite can be calculated in the following manner: subtract the square root semi-major axis of the target satellite at the time of the previous orbit determination from the square root semi-major axis of the target satellite at the time of the current orbit determination to obtain the semi-major axis change amount; determine the time interval between the last orbit determination and the current orbit determination; divide the semi-major axis change amount by the time interval to obtain the semi-major axis change rate of the target satellite.

[0102] In some embodiments, if the semi-major axis change rates of the satellites in the constellation are the same or have a small difference, that is, (in, represents the mean root mean square semi-major axis change rate of each satellite in the target constellation relative to the reference satellite), then the target value of the above-mentioned semi-major axis change amount does not need to be corrected.

[0103] In some embodiments, as Figure 7 As shown, S413 may specifically include the following S4131 and S4132.

[0104] S4131: When the semi-major axis change rate of the target satellite is smaller than the semi-major axis change rate of the reference satellite, the correction amount is calculated based on the semi-major axis and semi-major axis change rate of the reference satellite, the semi-major axis change rate of the target satellite, the phase control threshold, the virtual phase difference of the target satellite, and the orbital angular velocity of the reference satellite.

[0105] In some embodiments, the correction amount is calculated as follows: Among them, Δa represents the correction amount, a k is the square root semi-major axis of the reference satellite, is the mean root mean square semi-major axis rate of change of the target satellite, is the root mean square semi-major axis change rate of the reference satellite, x is the phase difference threshold, η is the control coefficient, Δλ i is the virtual phase difference of the target satellite, and n is the orbital angular velocity of the reference satellite.

[0106] The derivation process of the above formula is: Simplified satellite phase drift rate (unit rad / s), the simplified phase shift acceleration is (unit rad / s 2 ). The ratio of 1 unit rad / s to 1 unit rad / day is 1 / 86400, and the phase deviation Δu b =η*x+Δλ i , based on the satellite motion formula Combining the above variable relationships, we can get:

[0107] S4132: Correct the target value of the semi-major axis variation of the target satellite using the correction amount.

[0108] That is, the target value a of the semi-major axis variation of the target satellite c =a k -a i +a m +Δa.

[0109] S42: Calculate the orbital angular velocity of the reference satellite.

[0110] The orbital angular velocity of a reference satellite is the angle per unit time that the reference satellite rotates in its orbit around a central celestial body (such as Earth). Based on Kepler's laws and knowledge of circular motion, the satellite's orbital angular velocity is related to parameters such as the orbital radius (often approximated by the semi-major axis in elliptical orbits).

[0111] For an elliptical orbit, the satellite's velocity varies on the orbit, but the average orbital angular velocity can be approximately calculated using the semi-major axis using the formula: Among them, ω is the orbital angular velocity, μ is the gravitational constant of the central celestial body, and a is the orbital radius of the satellite.

[0112] S43: Calculate the velocity increment according to the orbital angular velocity of the reference satellite.

[0113] Velocity increment refers to the change in satellite velocity over a certain period of time.

[0114] In some embodiments, S43 may calculate the speed increment using the following formula:

[0115] ΔV=n / 2*a c , where ΔV is the velocity increment, n is the orbital angular velocity of the reference satellite, and a c is the target value of the semi-major axis variation.

[0116] S44: Calculate the control time according to the speed increment.

[0117] In some embodiments, when the velocity increment is greater than 0 (i.e., the angular velocity is in an accelerating state), the control time is set at the apogee of the orbit. Assuming that the mean anomaly angle of the target satellite's orbital element at the current moment is M, when M<π, the phase difference between the current and apogee is π-M, then the control time Tc = (π-M) / n+T now , where n is the orbital angular velocity of the reference satellite, T now is the current time; when M>π, the phase difference between the current and the apogee is 3π-M, then the control time Tc=(3π-M) / n+T now The apogee in this paragraph is the apogee closest to the current position of the target satellite. The control time can be postponed to the next apogee according to actual needs.

[0118] If the reference satellite's orbital velocity increment is greater than 0, the satellite's initial angular velocity at apogee is relatively low. This increase in angular velocity after control causes the satellite to rotate a greater angle in the same amount of time during subsequent orbital motion than before control. Consequently, the satellite's instantaneous phase gradually advances after control. This means that the satellite reaches a specific position in orbit earlier than without control, resulting in a change in instantaneous phase.

[0119] Because the proposed scheme uses the satellite with the largest semi-major axis as the reference satellite, meaning it is at the highest altitude, controlling other satellites to maintain the constellation configuration typically requires raising their orbits. At apogee, a satellite's velocity is lowest, resulting in the lowest kinetic energy and the highest potential energy. If orbital altitude increases are desired, controlling at apogee takes advantage of its lower velocity, consuming less energy to achieve orbital changes.

[0120] Because the satellite's motion is relatively stable at apogee, external interference is relatively small, and control operations at apogee are based on a relatively stable initial state. When the velocity increment is greater than 0, angular velocity changes are more easily controlled within a relatively stable range compared to control at perigee (where velocity is high and motion changes are rapid). Large fluctuations in angular velocity due to external interference or control errors during the control process are less likely to occur, resulting in relatively small angular velocity oscillations.

[0121] During apogee control, the primary change is in the satellite's velocity (thus affecting its angular velocity), while the orbital radius changes relatively slowly near apogee. After control, the orbital radius does not experience drastic changes due to the relatively stable energy changes of the satellite (because kinetic energy is low at apogee, energy changes during control are relatively small), resulting in minimal oscillations in the orbital radius parameter. However, during perigee control, due to the high perigee velocity and the large energy changes during control, this can cause a significant change in the orbital radius, leading to more pronounced oscillations in the orbital radius parameter.

[0122] In some embodiments, when the velocity increment is less than 0 (i.e., the angular velocity is in a deceleration state), the control time is set at the perigee of the orbit. The phase difference between the current and perigee is 2π-M, and the control time Tc = (2π-M) / n+T now , where n is the orbital angular velocity of the reference satellite, T now The perigee in this section is the closest perigee to the current position of the target satellite. The control time can be postponed to the next perigee according to actual needs.

[0123] Satellites at perigee have the highest linear velocity. Given a given radius, the greater the linear velocity, the greater the angular velocity. When angular velocity needs to be reduced (i.e., the velocity delta is less than 0), maneuvers at perigee, applying braking or deceleration forces, can leverage this greater linear velocity to achieve a significant change in linear velocity in a short period of time, thereby rapidly reducing the angular velocity of other satellites and maintaining the constellation configuration.

[0124] A satellite's kinetic energy is greatest at perigee. According to the principle of energy conversion, deceleration control at perigee converts this high kinetic energy into potential energy, thereby reducing angular velocity. Compared to control at apogee (where kinetic energy is lower), perigee allows for more efficient utilization of the satellite's kinetic energy. From an energy efficiency perspective, this makes it possible to achieve a more efficient reduction in angular velocity while consuming less energy.

[0125] It should be noted that due to the influence of the perturbation force, the altitude of the satellite will usually gradually decrease, that is, the altitude of the satellite needs to be increased; since the satellite with the largest plane root semi-major axis (that is, the highest) is used as the reference satellite in this scheme, the satellite is generally accelerated when maintaining the constellation configuration, so the control time is generally at the apogee of the orbit.

[0126] S45: Calculating control parameters of the thruster according to the speed increment and various parameters of the thruster.

[0127] Various parameters of the thruster, such as thrust, specific impulse, installation angle, whether to use electric thruster or chemical thruster, etc.

[0128] Thruster control parameters, such as thrust size, thruster direction, propellant flow, ignition time, ignition duration, and flameout time.

[0129] S46: Controlling the thruster action according to the control parameters.

[0130] Control the thruster action, that is, control the corresponding thrusters to ignite according to the control parameters to adjust the satellite's position and attitude.

[0131] In some embodiments, when the maximum value of the virtual phase difference between satellites is less than the phase control threshold, it may be determined that the constellation configuration is not maintained.

[0132] Compared with the existing technology, the above-mentioned constellation phase maintenance method has the following advantages: 1. From the perspective of the entire life of the satellite and the long-term operation of the constellation, the total fuel consumption is minimized. 2. The calculation process of the reference satellite is simplified. The selection of the reference satellite is carried out through the semi-major axis of the orbital root, which can simplify the amount of calculation and reduce the time required for calculation and the computer memory occupied. When the number of satellites in the constellation is n, the time complexity of the algorithm is O(n) and the space complexity is O(n). 3. The orbit control strategy is simplified. This scheme takes into account the long-term perturbations of near-Earth satellites. The control direction only has acceleration control and no deceleration control. The simpler strategy makes it less likely for operation and control personnel to make mistakes in long-term satellite management. 4. This scheme takes into account the different semi-major axis change rates of each satellite during the on-orbit operation of the constellation, and calculates the correction amount of configuration maintenance control based on this, thereby optimizing the control strategy.

[0133] This specification provides a constellation phase holding device that can be used to implement any of the above constellation phase holding methods. Figure 8 As shown, the device includes an acquisition unit 10 and a processing unit 20.

[0134] The acquisition unit 10 is used to obtain the mean root semi-major axis of each satellite in the target constellation at the target epoch, and select the satellite with the largest mean root semi-major axis as the reference satellite.

[0135] The processing unit 20 is configured to perform constellation phase maintenance based on the reference satellite.

[0136] In some embodiments, the processing unit includes a first calculation subunit, a second calculation subunit, and a first determination subunit.

[0137] The first calculation subunit is used to calculate the virtual phase difference of each satellite in the target constellation relative to the reference satellite, where the virtual phase difference is the difference between the actual phase difference and the nominal phase difference.

[0138] The second calculation subunit is used to calculate the virtual semi-major axis of each satellite in the target constellation relative to the reference satellite, where the virtual semi-major axis is the difference between the square root semi-major axis of the satellite and the square root semi-major axis of the reference satellite.

[0139] The first determining subunit is configured to maintain the constellation configuration when the maximum value of the virtual phase difference of each satellite is greater than or equal to the phase control threshold.

[0140] In some embodiments, the apparatus further comprises a second determining subunit, configured to not maintain the constellation configuration when the maximum value of the virtual phase difference of each satellite is less than a phase control threshold.

[0141] In some embodiments, the first determining subunit includes a third determining subunit, a second calculating subunit, a third calculating subunit, a fourth calculating subunit, a fifth calculating subunit, and a control subunit.

[0142] The third determining subunit is used to determine a target value of the semi-major axis variation of the target satellite.

[0143] The second calculation subunit is used to calculate the orbital angular velocity of the reference satellite.

[0144] The third calculation subunit is used to calculate the velocity increment according to the orbital angular velocity of the reference satellite.

[0145] The fourth calculation subunit is configured to calculate the control moment according to the speed increment.

[0146] The fifth calculation subunit is used to calculate the control parameters of the thruster according to the speed increment and various parameters of the thruster.

[0147] A control subunit is used to control the thruster action according to the control parameters.

[0148] In some embodiments, the third determining subunit includes a first acquiring subunit and a sixth calculating subunit.

[0149] The first acquisition subunit is used to acquire the semi-major axis control error of the target satellite.

[0150] The sixth calculation subunit is used to calculate the target value of the semi-major axis variation of the target satellite. The target value of the semi-major axis variation is the difference between the semi-major axis of the reference satellite and the semi-major axis of the target satellite, and then the difference is summed with the semi-major axis control error.

[0151] In some embodiments, the third determining subunit further includes a correcting subunit for correcting the target value of the semi-major axis variation according to the semi-major axis variation rate of the target satellite.

[0152] In some embodiments, the syndrome subunit includes a seventh calculation subunit and a correction subunit.

[0153] The seventh calculation subunit is used to calculate the correction amount based on the semi-major axis and semi-major axis change rate of the reference satellite, the semi-major axis change rate of the target satellite, the phase control threshold, the virtual phase difference of the target satellite, and the orbital angular velocity of the reference satellite when the semi-major axis change rate of the target satellite is less than the semi-major axis change rate of the reference satellite.

[0154] The correction subunit is used to correct the target value of the semi-major axis variation of the target satellite using the correction amount.

[0155] In some embodiments, the correction amount is calculated as follows: Among them, Δa represents the correction amount, a k is the square root semi-major axis of the reference satellite, is the mean root mean square semi-major axis rate of change of the target satellite, is the root mean square semi-major axis change rate of the reference satellite, x is the phase difference threshold, η is the control coefficient, Δλ i is the virtual phase difference of the target satellite, and n is the orbital angular velocity of the reference satellite.

[0156] In some embodiments, the semi-major axis change rate of the target satellite is calculated in the following manner: the square root of the semi-major axis of the target satellite at the time of the current orbit determination is subtracted from the square root of the semi-major axis at the time of the last orbit determination to obtain the semi-major axis change amount; the time interval between the last orbit determination and the current orbit determination is determined; the semi-major axis change amount is divided by the time interval to obtain the semi-major axis change rate of the target satellite.

[0157] In some embodiments, the third calculation subunit calculates the velocity increment using the following formula: ΔV=n / 2*a c , where ΔV = represents the velocity increment, n is the orbital angular velocity of the reference satellite, and a c is the target value of the semi-major axis variation.

[0158] In some embodiments, the fourth calculation subunit includes a fourth determination subunit and / or a fifth determination subunit.

[0159] The fourth determining subunit is used to set the control time at the apogee when the speed increment is greater than 0, and Tc=(π-M) / n+T now , where Tc is the control time, M is the mean anomaly of the target satellite at the current time, n is the orbital angular velocity of the reference satellite, and T now For the current moment.

[0160] The fifth determining subunit is used to set the control time at the perigee of the orbit when the velocity increment is less than 0. The phase difference between the current and perigee is 2π-M, so the control time Tc=(2π-M) / n+T now , where n is the orbital angular velocity of the reference satellite, T now For the current moment.

[0161] In some embodiments, the apparatus further includes an eighth computing subunit, a ninth computing subunit, and a rendering unit.

[0162] The eighth calculation subunit is used to calculate the virtual phase difference of each satellite in the target constellation relative to the reference satellite, where the virtual phase difference is the difference between the actual phase difference and the nominal phase difference value.

[0163] The ninth calculation subunit is used to calculate the virtual semi-major axis of each satellite in the target constellation relative to the reference satellite, where the virtual semi-major axis is the difference between the square root semi-major axis of the satellite and the square root semi-major axis of the reference satellite.

[0164] The drawing unit is used to draw a virtual phase drift situation diagram of the target constellation at the target epoch based on the virtual phase difference and virtual semi-major axis of each satellite relative to the reference satellite;

[0165] The constellation virtual phase drift situation diagram includes a horizontal coordinate indicating the phase difference and a vertical coordinate indicating the square root semi-major axis.

[0166] The description and functions of the above-mentioned devices can be understood by referring to the contents of the corresponding method part and will not be repeated here.

[0167] The embodiment of the present invention further provides an electronic device, which can be the user client running the browser, or a server. Figure 9 As shown, the electronic device may include a processor 901 and a memory 902, wherein the processor 901 and the memory 902 may be connected via a bus or other means. Figure 9 The bus connection is taken as an example.

[0168] The processor 901 may be a central processing unit (CPU). The processor 901 may also be other general-purpose processors, digital signal processors (DSP), application-specific integrated circuits (ASIC), field-programmable gate arrays (FPGA), or other programmable logic devices, discrete gate or transistor logic devices, discrete hardware components, or a combination of the above chips.

[0169] The memory 902 is a non-transitory computer-readable storage medium that can be used to store non-transitory software programs, non-transitory computer executable programs, and modules, such as the program instructions / modules corresponding to the constellation phase maintenance method in the embodiment of the present invention (e.g. Figure 8 The processor 901 executes various functional applications and data processing of the processor by running the non-transitory software programs, instructions, and modules stored in the memory 902, that is, implementing the constellation phase maintenance method in the above method embodiment.

[0170] The memory 902 may include a program storage area and a data storage area, wherein the program storage area may store an operating system and application programs required for at least one function; the data storage area may store data created by the processor 901, etc. In addition, the memory 902 may include a high-speed random access memory, and may also include a non-volatile memory, such as at least one disk storage device, a flash memory device, or other non-volatile solid-state storage device. In some embodiments, the memory 902 may optionally include a memory remotely located relative to the processor 901, and these remote memories may be connected to the processor 901 via a network. Examples of the above-mentioned network include, but are not limited to, the Internet, an intranet, a local area network, a mobile communication network, and combinations thereof.

[0171] The one or more modules are stored in the memory 902 and when executed by the processor 901, perform the following steps: Figure 1 The constellation phase maintenance method in the illustrated embodiment.

[0172] The specific details of the above electronic device can be understood by referring to the corresponding descriptions and effects in the method embodiment, and will not be repeated here.

[0173] This specification also provides a computer storage medium, which stores computer program instructions. When the computer program instructions are executed, the steps of the above-mentioned constellation phase maintenance method are implemented.

[0174] This specification also provides a computer program product, comprising a computer program, which implements the steps of the above constellation phase maintenance method when executed by a processor.

[0175] Those skilled in the art will appreciate that all or part of the processes in the above-mentioned embodiments can be implemented by instructing the relevant hardware through a computer program. The program can be stored in a computer-readable storage medium. When the program is executed, it can include the processes of the embodiments of the above-mentioned methods. The storage medium can be a magnetic disk, an optical disk, a read-only memory (ROM), a random access memory (RAM), a flash memory, a hard disk drive (HDD), or a solid-state drive (SSD). The storage medium can also include a combination of the above-mentioned types of memory.

[0176] The various embodiments in this specification are described in a progressive manner, and the same or similar parts between the various embodiments can be referred to each other. Each embodiment focuses on the differences from other embodiments.

[0177] The systems, devices, modules or units described in the above embodiments may be implemented by computer chips or entities, or by products with certain functions.

[0178] For the convenience of description, the above devices are described as being divided into various units according to their functions. Of course, when implementing this application, the functions of each unit can be implemented in the same or multiple software and / or hardware.

[0179] Through the description of the above embodiments, it can be seen that those skilled in the art can clearly understand that the present application can be implemented by means of software plus the necessary general hardware platform. Based on this understanding, the technical solution of the present application, or the part that contributes to the prior art, can be embodied in the form of a software product. The computer software product can be stored in a storage medium, such as ROM / RAM, a magnetic disk, an optical disk, etc., and includes a number of instructions for enabling a computer device (which can be a personal computer, a server, or a network device, etc.) to execute certain parts of the methods of each embodiment of the present application.

[0180] The present application can be used in a wide variety of general-purpose or specialized computer system environments or configurations, such as personal computers, server computers, handheld or portable devices, tablet devices, multiprocessor systems, microprocessor-based systems, set-top boxes, programmable consumer electronics, network PCs, minicomputers, mainframe computers, and distributed computing environments that include any of the above.

[0181] The present application may be described in the general context of computer-executable instructions executed by a computer, such as program modules. Generally, program modules include routines, programs, objects, components, data structures, etc. that perform specific tasks or implement specific abstract data types. The present application may also be practiced in distributed computing environments where tasks are performed by remote processing devices connected through a communications network. In a distributed computing environment, program modules may be located in local and remote computer storage media, including storage devices.

[0182] Although the present application has been described through embodiments, those skilled in the art will appreciate that there are many modifications and variations to the present application without departing from the spirit of the present application. It is intended that the appended claims include these modifications and variations without departing from the spirit of the present application.

Claims

1. A constellation phase maintenance method, characterized in that: include: Obtain the mean root semi-major axis of each satellite in the target constellation at the target epoch, and select the satellite with the largest mean root semi-major axis as the reference satellite; Constellation phase maintenance is performed based on the reference satellite.

2. The method according to claim 1, characterized in that Performing constellation phase maintenance based on the reference satellite includes: Calculating a virtual phase difference of each satellite in the target constellation relative to the reference satellite, where the virtual phase difference is the difference between the actual phase difference and the nominal phase difference; When the maximum value of the virtual phase difference of each satellite is greater than or equal to the phase control threshold, the constellation configuration is maintained.

3. The method according to claim 1, characterized in that The method further comprises: When the maximum value of the virtual phase difference of each satellite is less than the phase control threshold, the constellation configuration is not maintained.

4. The method according to claim 1, wherein Maintaining the constellation configuration includes controlling the target satellite motions by: Determine a target value of the semi-major axis variation of the target satellite; Calculate the orbital angular velocity of the reference satellite; Calculate the velocity increment based on the orbital angular velocity of the reference satellite; Calculating a control moment according to the speed increment; Calculating control parameters of the thruster according to the speed increment and various parameters of the thruster; The thruster action is controlled according to the control parameters.

5. The method according to claim 4, characterized in that Determine the target value of the semi-major axis variation of the target satellite, including: Obtain the semi-major axis control error of the target satellite; The target value of the semi-major axis variation of the target satellite is calculated, where the target value of the semi-major axis variation is the difference between the semi-major axis of the reference satellite and the semi-major axis of the target satellite, and the difference is then summed with the semi-major axis control error.

6. The method according to claim 5, characterized in that After calculating the target value of the semi-major axis variation of the target satellite, it also includes: The target value of the semi-major axis variation is corrected according to the semi-major axis variation rate of the target satellite.

7. The method according to claim 6, characterized in that Correcting the target value of the semi-major axis variation according to the semi-major axis variation rate of the target satellite includes: When the semi-major axis change rate of the target satellite is less than the semi-major axis change rate of the reference satellite, a correction amount is calculated based on the semi-major axis and semi-major axis change rate of the reference satellite, the semi-major axis change rate of the target satellite, the phase control threshold, the virtual phase difference of the target satellite, and the orbital angular velocity of the reference satellite. The correction amount is used to correct the target value of the semi-major axis variation of the target satellite.

8. The method according to claim 7, characterized in that The calculation formula for the correction amount is: Among them, Δa represents the correction amount, a k is the square root semi-major axis of the reference satellite, is the mean root mean square semi-major axis rate of change of the target satellite, is the root mean square semi-major axis change rate of the reference satellite, x is the phase difference threshold, η is the control coefficient, Δλ i is the virtual phase difference of the target satellite, and n is the orbital angular velocity of the reference satellite.

9. The method according to claim 6, characterized in that The semi-major axis change rate of the target satellite is calculated as follows: The change of the semi-major axis is obtained by subtracting the square root semi-major axis of the target satellite at the time of previous orbit determination from the square root semi-major axis of the target satellite at the time of current orbit determination. Determine the time interval between the last orbit determination and the current orbit determination; The semi-major axis change amount is divided by the time interval to obtain the semi-major axis change rate of the target satellite.

10. The method according to claim 4, characterized in that The velocity increment is calculated based on the orbital angular velocity of the reference satellite, including calculating the velocity increment using the following formula: ΔV=n / 2*a c Where, ΔV = represents the velocity increment, n is the orbital angular velocity of the reference satellite, and a c is the target value of the semi-major axis variation.

11. The method according to claim 4, characterized in that Calculating the control moment according to the speed increment includes: When the speed increment is greater than 0, set the control time at the far point, and Tc=(π-M) / n+T now , where Tc is the control time, M is the mean anomaly of the target satellite at the current time, n is the orbital angular velocity of the reference satellite, and T now For the current moment; and / or, When the velocity increment is less than 0, the control time is set at the perigee of the orbit. The phase difference between the current and perigee is 2π-M, so the control time Tc=(2π-M) / n+T now , where n is the orbital angular velocity of the reference satellite, T now For the current moment.

12. The method according to claim 1, characterized in that The method further comprises: Calculating a virtual phase difference of each satellite in the target constellation relative to the reference satellite, where the virtual phase difference is the difference between the actual phase difference and the nominal phase difference; Calculating a virtual semi-major axis of each satellite in the target constellation relative to the reference satellite, where the virtual semi-major axis is the difference between the square root semi-major axis of the satellite and the square root semi-major axis of the reference satellite; A virtual phase drift situation diagram of the target constellation at the target epoch is drawn based on the virtual phase difference and virtual semi-major axis of each satellite relative to the reference satellite; The constellation virtual phase drift situation diagram includes a horizontal coordinate indicating the phase difference and a vertical coordinate indicating the square root semi-major axis.

13. A constellation phase holding device, characterized in that: include: an acquisition unit, configured to acquire the mean root semi-major axis of each satellite in the target constellation at the target epoch, and select the satellite with the largest mean root semi-major axis as the reference satellite; A processing unit is configured to perform constellation phase maintenance based on the reference satellite.

14. An electronic device, characterized in that: include: A memory and a processor, wherein the processor and the memory are communicatively connected to each other, the memory stores computer instructions, and the processor implements the constellation phase maintenance method according to any one of claims 1 to 12 by executing the computer instructions.

15. A computer storage medium, characterized in that The computer storage medium stores computer program instructions, and when the computer program instructions are executed, the steps of the constellation phase maintenance method according to any one of claims 1 to 12 are implemented.

16. A computer program product, characterized in that The invention comprises a computer program, which, when executed by a processor, implements the steps of the constellation phase maintenance method according to any one of claims 1 to 12.

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