Autonomous Orbit Maintenance Method and System for Low-Earth Orbit Inter-Satellite Laser Communication Constellations

Abstract

This invention provides an autonomous orbit-keeping method and system for low-Earth orbit inter-satellite laser communication constellations, comprising: generating upper and lower limits for satellite phase-keeping thresholds and periodically updating the thresholds based on the annual changes in laser link performance; recursively calculating the satellite's mean orbit elements at the current moment based on the autonomous positioning results of the onboard GNSS navigation receiver, and calculating the mean latitude argument deviation relative to the nominal orbit; autonomously calculating and generating an autonomous orbit control strategy onboard using a limit loop phase-keeping method; after generating the orbit-keeping strategy, transmitting it to the ground control center, and reviewing the orbit-keeping scheme autonomously generated onboard according to the integrated space-ground autonomous orbit control decision-making process; if the review is successful, the ground issues a permission command for orbit control implementation, and the satellite executes the orbit control operation according to the strategy; if the review fails, the ground re-enters the orbit-keeping strategy. This invention significantly reduces the complexity of ground system constellation management.

Description

Technical Field

[0001] This invention relates to the field of satellite system design, and more specifically, to a method and system for autonomous orbit maintenance of a low-Earth orbit inter-satellite laser communication constellation. Background Technology

[0002] Orbit maintenance is crucial for ensuring constellation configuration and system performance. With the large-scale construction of low-Earth orbit (LEO) constellations in recent years, the pressure on ground-based constellation management has increased dramatically, necessitating improvements in autonomous orbit maintenance capabilities. Furthermore, to meet the demands of large-volume, high-data-rate inter-satellite communication and multi-satellite collaborative applications, LEO constellations often employ inter-satellite laser communication links to achieve efficient interconnection. However, the instantaneous beam angle of inter-satellite laser communication terminals is small, making them prone to link loss during drastic satellite attitude changes, and re-establishing and re-acquiring links after an interruption is time-consuming. Therefore, for LEO constellations using inter-satellite laser communication links, ensuring uninterrupted inter-satellite link establishment is essential during autonomous orbit maintenance. Current research on constellation orbit maintenance has given limited consideration to the requirement for continuous inter-satellite laser communication link establishment.

[0003] One existing method for satellite orbit maintenance control based on two-line roots (publication number CN102591343A) overcomes the error caused by using unstable atmospheric density as the calculation input. It provides a method for orbit prediction and control of satellites requiring ground trajectory maintenance based on publicly available two-line root data, improving the reliability and accuracy of prediction and control calculations. However, this method lacks onboard autonomous implementation capability. Another method for autonomous satellite orbit maintenance based on online gain estimation (publication number CN107554820A) addresses thrust changes caused by satellite mass variations, orbit modeling errors, and jet pressure variations. It uses an online-updable orbit control gain to autonomously calculate the jet duration for orbit control, reducing ground workload. A third method for onboard autonomous orbit maintenance control (publication number CN106542119A) addresses the orbit maintenance problem of ultra-low orbit satellites by proposing a thruster switching maintenance strategy based on average orbital semi-major axis feedback, achieving autonomous orbital altitude maintenance under different atmospheric conditions. Both of these methods focus on the generation and execution of autonomous strategies on the satellite, without considering the autonomous decision-making process on the satellite, and are still some distance from practical application. Summary of the Invention

[0004] To address the shortcomings of existing technologies, the purpose of this invention is to provide an autonomous orbit maintenance method and system for low-Earth orbit inter-satellite laser communication constellations.

[0005] An autonomous orbit maintenance method for a low-Earth orbit inter-satellite laser communication constellation provided by the present invention includes:

[0006] Step S1: Generate the upper and lower limits of the satellite phase preservation threshold, and update the thresholds periodically according to the annual changes in laser link performance;

[0007] Step S2: Based on the autonomous positioning results of the onboard GNSS navigation receiver, recursively calculate the satellite's horizontal orbit elements at the current moment, and calculate the horizontal latitude argument deviation relative to the nominal orbit;

[0008] Step S3: When the satellite's latitude argument deviation and phase holding threshold meet the preset conditions, the satellite adopts the limit loop phase holding method to autonomously calculate and generate an autonomous orbit control strategy.

[0009] Step S4: After generating the orbit maintenance strategy, it is transmitted to the ground control center. According to the space-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit maintenance plan generated autonomously on the satellite. If the review is successful, the ground sends an instruction to allow orbit control to be implemented, and the satellite executes the orbit control operation according to the strategy. If the review is unsuccessful, the ground sends the orbit maintenance strategy.

[0010] Preferably, in step S1:

[0011] The satellite phase preservation threshold calculation process is as follows: the maximum phase difference between two linked satellites is calculated based on the low-Earth orbit constellation altitude, the longest link establishment distance constraint of the inter-satellite laser communication link, and the minimum link establishment altitude constraint above the Earth's surface. Using the absolute phase preservation mode, the upper and lower limits of the satellite phase preservation threshold are set. Based on the on-orbit laser link telemetry data, the communication performance is analyzed, and the longest link establishment distance and the minimum link establishment altitude constraint of the inter-satellite laser communication link are updated. The phase preservation threshold is updated once every preset time period.

[0012] Generate satellite phase preservation threshold Δu lim Low Earth orbit constellation orbital altitude H sat The longest connection establishment distance constraint for inter-satellite laser communication links in a constellation is L. max The lowest chain height above the ground is H. min The maximum phase difference between the two linked satellites is u. max It is calculated by the following formula:

[0013] u res1 =2sin -1 (L max / 2 / (H sat +R E ))

[0014] u res2 =2cos -1 ((H min +R E ) / (H sat +R E ))

[0015] u max =min(u res1 ,u res2 )

[0016] In the formula, R E U is the Earth's radius. res1 u is the maximum inter-satellite phase difference constrained by the inter-satellite laser communication distance. res2 The maximum inter-satellite phase difference constrained by the minimum link establishment altitude for inter-satellite communication;

[0017] The maximum nominal phase difference between two linked satellites is u nom The absolute phase-preserving mode is adopted, and the satellite phase-preserving threshold is set to the upper limit Δu. lim =(u max -u nom ) / 2, with the lower limit set to -Δu lim ;

[0018] Based on the analysis of communication performance using on-orbit laser link telemetry data, the longest link establishment distance L... max The allowable height value H for shortening and building chains min The phase hold threshold Δu is raised and updated every preset time interval. lim This ensures that satellites within the constellation remain connected.

[0019] Preferably, in step S2:

[0020] Latitude Aspect Deviation The calculation process is as follows: Based on the real-time position and velocity of the satellite at the time of resolution by the GNSS navigation receiver, the mean orbit elements corresponding to the current time are obtained through numerical iteration based on the Lagrange planetary equations; the mean latitude argument at the current time is calculated and compared with the nominal mean orbit elements to obtain the satellite mean latitude argument deviation relative to the nominal orbit at the current time.

[0021] Given the real-time satellite position r0 and velocity v0 at time t0 calculated by the GNSS navigation receiver, determine the oscillological elements [ae iΩωM] at time t0 using Kepler's equations. T Where a is the semi-major axis of the auscultatory orbit, e is the eccentricity of the auscultatory orbit, i is the inclination of the auscultatory orbit, Ω is the right ascension of the ascending node of the auscultatory orbit, ω is the argument of the perigee of the auscultatory orbit, and M is the mean perigee of the auscultatory orbit; the current time t is obtained through numerical iteration. CT Corresponding horizontal orbital elements in, For the semi-major axis of the horizontal track, For the eccentricity of the horizontal track, For the inclination angle of the horizontal track, The right ascension of the ascending node of the horizontal orbit. The perigee angle of the horizontal orbit. The iterative formula for the aperitone angle of the horizontal orbit is shown below:

[0022]

[0023] In the formula, and The first-order variation of the orbital elements with time caused by the perturbation is obtained according to the Lagrange planetary equations. The initial value for the iterative calculation of the horizontal orbital elements is [ae iΩωM+n(t CT -t0)] T where n is the angular velocity of the satellite's orbital motion;

[0024] Let the current time t be... CT Argument of latitude is Compare with the nominal orbital square root number to calculate the current time t. CT Satellite mean latitude angle deviation relative to nominal orbit

[0025] Preferably, in step S3:

[0026] The satellite employs a limit cycle phase holding method, using orbit holding control to offset the semi-major axis, forming a drift control loop. Given the current time, the orbit control implementation time, and the nominal orbital elements, the orbit control implementation time t is calculated according to the principle of the limit cycle phase holding method. f The required velocity increment δV can be calculated and the satellite can autonomously generate an orbit maintenance strategy. The satellite's autonomous orbit maintenance is set to a quasi-prohibited state by default, and is triggered only when the ground control center issues a permission command.

[0027] Onboard autonomous orbit-keeping strategy: When the satellite's mean latitude angle deviation... At that time, the satellite employs the extreme loop phase-keeping method to autonomously calculate and generate an autonomous orbit control strategy. f A velocity increment δV is applied continuously and updated every preset time interval;

[0028] The satellite employs a limit loop phase-keeping method, using orbit-keeping control to offset the semi-major axis, forming a drift control loop. Initially, a positive offset of the semi-major axis is applied, causing the phase to shift to the left. As the semi-major axis decays, it gradually falls below its nominal value, initiating a rightward phase shift. When the satellite reaches the right boundary, orbit-keeping control δa is implemented to maintain the phase deviation within -Δu. lim With Δu lim between;

[0029] make

[0030] k1 = n CT -n N

[0031]

[0032] In the formula, μ is the Earth's gravitational constant. Let t be the current time. CT Orbital angular rate, a CT For t CT The semi-major axis of the horizontal orbit at any given moment; For the nominal orbital angular rate, a N k1 represents the semi-major axis of the nominal orbit; k2 is the difference in angular rate between the current orbit and the nominal orbit; and k2 is the coefficient of variation of the orbital angular rate considering the influence of atmospheric drag. This represents the first-order rate of change of the semi-major axis caused by atmospheric drag.

[0033] The phase holding period is then

[0034]

[0035] The target value for the semi-major axis deviation of phase preservation is

[0036]

[0037] When the satellite reaches boundary C, the actual semi-major axis deviation is

[0038]

[0039] In the formula: Δa N =a CT -a N t represents the deviation between the current actual semi-major axis and the nominal semi-major axis. f It is the moment when the satellite reaches boundary C or the moment when orbit control is implemented;

[0040] When the satellite reaches boundary C, the change in the semi-major axis of the satellite's orbital maneuver is:

[0041] δa=Δa T -Δa f

[0042] For near-Earth circular orbits, the corresponding velocity increment requirement is:

[0043]

[0044] The satellite is in a state of default prohibition for maintaining its autonomous orbit. It is in a state of prohibition by default and will be triggered only when a permission command is issued from the ground.

[0045] Preferably, in step S4:

[0046] Space-Ground Integrated Autonomous Orbit Control Decision and Execution: After the satellite autonomously generates its orbit maintenance strategy, it transmits it to the ground control center at a pre-set time. According to the space-ground integrated autonomous orbit control decision process, the ground control center reviews the orbit maintenance plan generated by the satellite. If the review is successful, the ground issues a permission instruction to implement orbit control, and the satellite executes the orbit control operation according to the strategy. If the review is unsuccessful, the ground sends the orbit maintenance strategy back to the satellite.

[0047] The aforementioned space-ground integrated autonomous orbit control decision-making process is as follows:

[0048] Step S4.1: The ground control center sends an instruction to allow the satellite to autonomously maintain its orbit. Once the phase deviation condition is met, the satellite autonomously generates an orbit-maintaining strategy.

[0049] Step S4.2: The autonomous orbit maintenance strategy is sent to the satellite integrated electronic subsystem for review. If approved, proceed to the next step; otherwise, cancel the current autonomous orbit control.

[0050] Step S4.3: The autonomous orbit control strategy is telemetryally transmitted to the ground, and the ground control center reviews it. If the ground approves the implementation, the next step is executed; otherwise, the autonomous orbit control is canceled and the orbit holding strategy is uploaded from the ground.

[0051] Step S4.4: Implement the autonomous orbit holding strategy. Before the orbit control thruster ignites, the orbit control subsystem sends a state establishment request to the propulsion subsystem at a preset time point to establish the propulsion state.

[0052] Step S4.5: The attitude and orbit control subsystem controls the orbit control thruster to ignite;

[0053] Step S4.6: After the orbit control ignition is completed, the attitude and orbit control subsystem sends a request to shut down the propulsion subsystem, and the integrated electronic subsystem shuts down the propulsion subsystem;

[0054] Step S4.7: After the track control is completed, the accuracy of the track control is calibrated using GNSS navigation data;

[0055] By installing thrusters in both directions of satellite flight, the satellite has the ability to control its ascent and descent orbits in both forward and inverted directions within the orbital coordinate system.

[0056] An autonomous orbit-keeping system for a low-Earth orbit inter-satellite laser communication constellation, according to the present invention, comprises:

[0057] Module M1: Generates the upper and lower limits of the satellite phase-preservation threshold and updates the thresholds periodically based on the annual changes in laser link performance;

[0058] Module M2: Based on the autonomous positioning results of the onboard GNSS navigation receiver, recursively calculate the satellite's horizontal orbit elements at the current moment, and calculate the horizontal latitude angle deviation relative to the nominal orbit;

[0059] Module M3: When the satellite's latitude angle deviation and phase holding threshold meet the preset conditions, the satellite adopts the limit loop phase holding method to autonomously calculate and generate an autonomous orbit control strategy.

[0060] Module M4: After generating the orbit maintenance strategy, it is transmitted to the ground control center. According to the space-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit maintenance plan generated autonomously on the satellite. If the review is successful, the ground sends a command to allow orbit control to be implemented, and the satellite executes the orbit control operation according to the strategy. If the review is unsuccessful, the ground will re-inject the orbit maintenance strategy.

[0061] Preferably, in module M1:

[0062] The satellite phase preservation threshold calculation process is as follows: the maximum phase difference between two linked satellites is calculated based on the low-Earth orbit constellation altitude, the longest link establishment distance constraint of the inter-satellite laser communication link, and the minimum link establishment altitude constraint above the Earth's surface. Using the absolute phase preservation mode, the upper and lower limits of the satellite phase preservation threshold are set. Based on the on-orbit laser link telemetry data, the communication performance is analyzed, and the longest link establishment distance and the minimum link establishment altitude constraint of the inter-satellite laser communication link are updated. The phase preservation threshold is updated once every preset time period.

[0063] Generate satellite phase preservation threshold Δu lim Low Earth orbit constellation orbital altitude H sat The longest connection establishment distance constraint for inter-satellite laser communication links in a constellation is L. max The lowest chain height above the ground is H. min The maximum phase difference between the two linked satellites is u. max It is calculated by the following formula:

[0064] u res1 =2sin -1 (L max / 2 / (H sat +R E ))

[0065] u res2 =2cos -1 ((H min +R E ) / (H sat +R E ))

[0066] u max =min(u res1 ,u res2 )

[0067] In the formula, R E U is the Earth's radius. res1 u is the maximum inter-satellite phase difference constrained by the inter-satellite laser communication distance.res2 The maximum inter-satellite phase difference constrained by the minimum link establishment altitude for inter-satellite communication;

[0068] The maximum nominal phase difference between two linked satellites is u nom The absolute phase-preserving mode is adopted, and the satellite phase-preserving threshold is set to the upper limit Δu. lim =(u max -u nom ) / 2, with the lower limit set to -Δu lim ;

[0069] Based on the analysis of communication performance using on-orbit laser link telemetry data, the longest link establishment distance L... max The allowable height value H for shortening and building chains min The phase hold threshold Δu is raised and updated every preset time interval. lim This ensures that satellites within the constellation remain connected.

[0070] Preferably, in module M2:

[0071] Latitude Aspect Deviation The calculation process is as follows: Based on the real-time position and velocity of the satellite at the time of resolution by the GNSS navigation receiver, the mean orbit elements corresponding to the current time are obtained through numerical iteration based on the Lagrange planetary equations; the mean latitude argument at the current time is calculated and compared with the nominal mean orbit elements to obtain the satellite mean latitude argument deviation relative to the nominal orbit at the current time.

[0072] Given the real-time satellite position r0 and velocity v0 at time t0 calculated by the GNSS navigation receiver, determine the oscillological elements [ae iΩωM] at time t0 using Kepler's equations. T Where a is the semi-major axis of the auscultatory orbit, e is the eccentricity of the auscultatory orbit, i is the inclination of the auscultatory orbit, Ω is the right ascension of the ascending node of the auscultatory orbit, ω is the argument of the perigee of the auscultatory orbit, and M is the mean perigee of the auscultatory orbit; the current time t is obtained through numerical iteration. CT Corresponding horizontal orbital elements in, For the semi-major axis of the horizontal track, For the eccentricity of the horizontal track, For the inclination angle of the horizontal track, The right ascension of the ascending node of the horizontal orbit. The perigee angle of the horizontal orbit. The iterative formula for the aperitone angle of the horizontal orbit is shown below:

[0073]

[0074] In the formula, and The first-order variation of the orbital elements with time caused by the perturbation is obtained according to the Lagrange planetary equations. The initial value for the iterative calculation of the horizontal orbital elements is [ae iΩωM+n(t CT -t0)] T where n is the angular velocity of the satellite's orbital motion;

[0075] Let the current time t be... CT Argument of latitude is Compare with the nominal orbital square root number to calculate the current time t. CT Satellite mean latitude angle deviation relative to nominal orbit

[0076] Preferably, in module M3:

[0077] The satellite employs a limit cycle phase holding method, using orbit holding control to offset the semi-major axis, forming a drift control loop. Given the current time, the orbit control implementation time, and the nominal orbital elements, the orbit control implementation time t is calculated according to the principle of the limit cycle phase holding method. f The required velocity increment δV can be calculated and the satellite can autonomously generate an orbit maintenance strategy. The satellite's autonomous orbit maintenance is set to a quasi-prohibited state by default, and is triggered only when the ground control center issues a permission command.

[0078] Onboard autonomous orbit-keeping strategy: When the satellite's mean latitude angle deviation... At that time, the satellite employs the extreme loop phase-keeping method to autonomously calculate and generate an autonomous orbit control strategy. f A velocity increment δV is applied continuously and updated every preset time interval;

[0079] The satellite employs a limit loop phase-keeping method, using orbit-keeping control to offset the semi-major axis, forming a drift control loop. Initially, a positive offset of the semi-major axis is applied, causing the phase to shift to the left. As the semi-major axis decays, it gradually falls below its nominal value, initiating a rightward phase shift. When the satellite reaches the right boundary, orbit-keeping control δa is implemented to maintain the phase deviation within -Δu. lim With Δu lim between;

[0080] make

[0081] k1 = n CT -n N

[0082]

[0083] In the formula, μ is the Earth's gravitational constant. Let t be the current time. CT Orbital angular rate, a CT For t CTThe semi-major axis of the horizontal orbit at any given moment; For the nominal orbital angular rate, a N k1 represents the semi-major axis of the nominal orbit; k2 is the difference in angular rate between the current orbit and the nominal orbit; and k2 is the coefficient of variation of the orbital angular rate considering the influence of atmospheric drag. This represents the first-order rate of change of the semi-major axis caused by atmospheric drag.

[0084] The phase holding period is then

[0085]

[0086] The target value for the semi-major axis deviation of phase preservation is

[0087]

[0088] When the satellite reaches boundary C, the actual semi-major axis deviation is

[0089]

[0090] In the formula: Δa N =a CT -a N t represents the deviation between the current actual semi-major axis and the nominal semi-major axis. f It is the moment when the satellite reaches boundary C or the moment when orbit control is implemented;

[0091] When the satellite reaches boundary C, the change in the semi-major axis of the satellite's orbital maneuver is:

[0092] δa=Δa T -Δa f

[0093] For near-Earth circular orbits, the corresponding velocity increment requirement is:

[0094]

[0095] The satellite is in a state of default prohibition for maintaining its autonomous orbit. It is in a state of prohibition by default and will be triggered only when a permission command is issued from the ground.

[0096] Preferably, in module M4:

[0097] Space-Ground Integrated Autonomous Orbit Control Decision and Execution: After the satellite autonomously generates its orbit maintenance strategy, it transmits it to the ground control center at a pre-set time. According to the space-ground integrated autonomous orbit control decision process, the ground control center reviews the orbit maintenance plan generated by the satellite. If the review is successful, the ground issues a permission instruction to implement orbit control, and the satellite executes the orbit control operation according to the strategy. If the review is unsuccessful, the ground sends the orbit maintenance strategy back to the satellite.

[0098] The aforementioned space-ground integrated autonomous orbit control decision-making process is as follows:

[0099] Module M4.1: The ground control center sends instructions to allow the satellite to autonomously maintain its orbit. Once the phase deviation condition is met, the satellite autonomously generates an orbit-maintaining strategy.

[0100] Module M4.2: The autonomous orbit maintenance strategy is sent to the satellite integrated electronic subsystem for review. If approved, the next step is executed; otherwise, the current autonomous orbit control is canceled.

[0101] Module M4.3: The autonomous orbit control strategy is telemetryally transmitted to the ground, and the ground control center reviews it. If the ground approves the implementation, the next step is executed; otherwise, the autonomous orbit control is canceled and the orbit holding strategy is uploaded from the ground.

[0102] Module M4.4: Implemented according to the autonomous orbit holding strategy, the orbit control subsystem sends a state establishment request to the propulsion subsystem at a preset time point before the orbit control thruster ignites, thus establishing the propulsion state;

[0103] Module M4.5: The attitude and orbit control subsystem controls the orbit control thruster to ignite;

[0104] Module M4.6: After the track control ignition is completed, the attitude and track control subsystem sends a request to shut down the propulsion subsystem, and the integrated electronic subsystem shuts down the propulsion subsystem;

[0105] Module M4.7: After track control is completed, the accuracy of the track control is calibrated using GNSS navigation data;

[0106] By installing thrusters in both directions of satellite flight, the satellite has the ability to control its ascent and descent orbits in both forward and inverted directions within the orbital coordinate system.

[0107] Compared with the prior art, the present invention has the following beneficial effects:

[0108] This invention achieves full-chain autonomous on-board implementation of strategy generation, orbit control decision-making, and mission execution while ensuring the permanent establishment of inter-satellite laser communication links. This significantly reduces the complexity of constellation management on the ground system and has on-orbit application value. Attached Figure Description

[0109] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings:

[0110] Figure 1 A schematic diagram corresponding to the autonomous orbit maintenance method for low-Earth orbit inter-satellite laser communication constellations provided in an embodiment of the present invention;

[0111] Figure 2 A schematic diagram of the limiting loop phase preservation method provided in an embodiment of the present invention;

[0112] Figure 3This is a schematic diagram of the integrated space-ground autonomous orbit control decision-making and execution process provided in an embodiment of the present invention. Detailed Implementation

[0113] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the protection scope of the present invention.

[0114] Example 1:

[0115] This invention discloses an autonomous orbit maintenance method for low-Earth orbit inter-satellite laser communication constellations. The method includes the following steps: Step A, generating and periodically updating the constellation phase maintenance threshold: To meet the constant maintenance requirements of inter-satellite laser links in low-Earth orbit constellations, an upper limit Δu of the satellite phase maintenance threshold is generated. lim and lower limit -Δu lim The threshold is updated periodically based on the annual changes in laser link performance. Step B: Calculation of satellite horizontal orbit elements deviation based on GNSS navigation receiver: Based on the autonomous positioning results of the onboard GNSS navigation receiver, the satellite horizontal orbit elements at the current moment are recursively calculated, and the horizontal latitude argument deviation relative to the nominal orbit is calculated. Step C, Onboard Autonomous Generation of Orbit Maintenance Strategy: When the satellite's mean latitude angle deviation... At that time, the satellite employs the extreme loop phase-keeping method to autonomously calculate and generate an autonomous orbit control strategy (at t). f A velocity increment δV is applied continuously and updated every hour. Step D, Space-Ground Integrated Autonomous Orbit Control Decision and Execution: After the satellite autonomously generates its orbit maintenance strategy, it is transmitted to the ground control center 24 hours in advance. According to the space-ground integrated autonomous orbit control decision process, the ground control center reviews the orbit maintenance plan generated by the satellite. If the review is successful, the ground issues a permission instruction for orbit control implementation, and the satellite executes the orbit control operation according to the strategy. If the review fails, the ground re-injects the orbit maintenance strategy.

[0116] According to the present invention, an autonomous orbit maintenance method for a low-Earth orbit inter-satellite laser communication constellation is provided, such as... Figures 1-3 As shown, it includes:

[0117] Step S1: Generate the upper and lower limits of the satellite phase preservation threshold, and update the thresholds periodically according to the annual changes in laser link performance;

[0118] Specifically, in step S1:

[0119] The satellite phase preservation threshold calculation process is as follows: the maximum phase difference between two linked satellites is calculated based on the low-Earth orbit constellation altitude, the longest link establishment distance constraint of the inter-satellite laser communication link, and the minimum link establishment altitude constraint above the Earth's surface. Using the absolute phase preservation mode, the upper and lower limits of the satellite phase preservation threshold are set. Based on the on-orbit laser link telemetry data, the communication performance is analyzed, and the longest link establishment distance and the minimum link establishment altitude constraint of the inter-satellite laser communication link are updated. The phase preservation threshold is updated once every preset time period.

[0120] Generate satellite phase preservation threshold Δu lim Low Earth orbit constellation orbital altitude H sat The longest connection establishment distance constraint for inter-satellite laser communication links in a constellation is L. max The lowest chain height above the ground is H. min The maximum phase difference between the two linked satellites is u. max It is calculated by the following formula:

[0121] u res1 =2sin -1 (L max / 2 / (H sat +R E ))

[0122] u res2 =2cos -1 ((H min +R E ) / (H sat +R E ))

[0123] u max =min(u res1 ,u res2 )

[0124] In the formula, R E U is the Earth's radius. res1 u is the maximum inter-satellite phase difference constrained by the inter-satellite laser communication distance. res2 The maximum inter-satellite phase difference constrained by the minimum link establishment altitude for inter-satellite communication;

[0125] The maximum nominal phase difference between two linked satellites is u nom The absolute phase-preserving mode is adopted, and the satellite phase-preserving threshold is set to the upper limit Δu. lim =(u max -u nom ) / 2, with the lower limit set to -Δu lim ;

[0126] Based on the analysis of communication performance using on-orbit laser link telemetry data, the longest link establishment distance L... max The allowable height value H for shortening and building chainsmin The phase hold threshold Δu is raised and updated every preset time interval. lim This ensures that satellites within the constellation remain connected.

[0127] Step S2: Based on the autonomous positioning results of the onboard GNSS navigation receiver, recursively calculate the satellite's horizontal orbit elements at the current moment, and calculate the horizontal latitude argument deviation relative to the nominal orbit;

[0128] Specifically, in step S2:

[0129] Latitude Aspect Deviation The calculation process is as follows: Based on the real-time position and velocity of the satellite at the time of resolution by the GNSS navigation receiver, the mean orbit elements corresponding to the current time are obtained through numerical iteration based on the Lagrange planetary equations; the mean latitude argument at the current time is calculated and compared with the nominal mean orbit elements to obtain the satellite mean latitude argument deviation relative to the nominal orbit at the current time.

[0130] Given the real-time satellite position r0 and velocity v0 at time t0 calculated by the GNSS navigation receiver, determine the oscillological elements [ae iΩωM] at time t0 using Kepler's equations. T Where a is the semi-major axis of the auscultatory orbit, e is the eccentricity of the auscultatory orbit, i is the inclination of the auscultatory orbit, Ω is the right ascension of the ascending node of the auscultatory orbit, ω is the argument of the perigee of the auscultatory orbit, and M is the mean perigee of the auscultatory orbit; the current time t is obtained through numerical iteration. CT Corresponding horizontal orbital elements in, For the semi-major axis of the horizontal track, For the eccentricity of the horizontal track, For the inclination angle of the horizontal track, The right ascension of the ascending node of the horizontal orbit. The perigee angle of the horizontal orbit. The iterative formula for the aperitone angle of the horizontal orbit is shown below:

[0131]

[0132] In the formula, and The first-order variation of the orbital elements with time caused by the perturbation is obtained according to the Lagrange planetary equations. The initial value for the iterative calculation of the horizontal orbital elements is [ae iΩωM+n(t CT -t0)] T where n is the angular velocity of the satellite's orbital motion;

[0133] Let the current time t be... CT Argument of latitude is Compare with the nominal orbital square root number to calculate the current time t. CTSatellite mean latitude angle deviation relative to nominal orbit

[0134] Step S3: When the satellite's latitude argument deviation and phase holding threshold meet the preset conditions, the satellite adopts the limit loop phase holding method to autonomously calculate and generate an autonomous orbit control strategy.

[0135] Specifically, in step S3:

[0136] The satellite employs a limit cycle phase holding method, using orbit holding control to offset the semi-major axis, forming a drift control loop. Given the current time, the orbit control implementation time, and the nominal orbital elements, the orbit control implementation time t is calculated according to the principle of the limit cycle phase holding method. f The required velocity increment δV can be calculated and the satellite can autonomously generate an orbit maintenance strategy. The satellite's autonomous orbit maintenance is set to a quasi-prohibited state by default, and is triggered only when the ground control center issues a permission command.

[0137] Onboard autonomous orbit-keeping strategy: When the satellite's mean latitude angle deviation... At that time, the satellite employs the extreme loop phase-keeping method to autonomously calculate and generate an autonomous orbit control strategy. f A velocity increment δV is applied continuously and updated every preset time interval;

[0138] The satellite employs a limit loop phase-keeping method, using orbit-keeping control to offset the semi-major axis, forming a drift control loop. Initially, a positive offset of the semi-major axis is applied, causing the phase to shift to the left. As the semi-major axis decays, it gradually falls below its nominal value, initiating a rightward phase shift. When the satellite reaches the right boundary, orbit-keeping control δa is implemented to maintain the phase deviation within -Δu. lim With Δu lim between;

[0139] make

[0140] k1 = n CT -n N

[0141]

[0142] In the formula, μ is the Earth's gravitational constant. Let t be the current time. CT Orbital angular rate, a CT For t CT The semi-major axis of the horizontal orbit at any given moment; For the nominal orbital angular rate, a N k1 represents the semi-major axis of the nominal orbit; k2 is the difference in angular rate between the current orbit and the nominal orbit; and k2 is the coefficient of variation of the orbital angular rate considering the influence of atmospheric drag. This represents the first-order rate of change of the semi-major axis caused by atmospheric drag.

[0143] The phase holding period is then

[0144]

[0145] The target value for the semi-major axis deviation of phase preservation is

[0146]

[0147] When the satellite reaches boundary C, the actual semi-major axis deviation is

[0148]

[0149] In the formula: Δa N =a CT -a N t represents the deviation between the current actual semi-major axis and the nominal semi-major axis. f It is the moment when the satellite reaches boundary C or the moment when orbit control is implemented;

[0150] When the satellite reaches boundary C, the change in the semi-major axis of the satellite's orbital maneuver is:

[0151] δa=Δa T -Δa f

[0152] For near-Earth circular orbits, the corresponding velocity increment requirement is:

[0153]

[0154] The satellite is in a state of default prohibition for maintaining its autonomous orbit. It is in a state of prohibition by default and will be triggered only when a permission command is issued from the ground.

[0155] Step S4: After generating the orbit maintenance strategy, it is transmitted to the ground control center. According to the space-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit maintenance plan generated autonomously on the satellite. If the review is successful, the ground sends an instruction to allow orbit control to be implemented, and the satellite executes the orbit control operation according to the strategy. If the review is unsuccessful, the ground sends the orbit maintenance strategy.

[0156] Specifically, in step S4:

[0157] Space-Ground Integrated Autonomous Orbit Control Decision and Execution: After the satellite autonomously generates its orbit maintenance strategy, it transmits it to the ground control center at a pre-set time. According to the space-ground integrated autonomous orbit control decision process, the ground control center reviews the orbit maintenance plan generated by the satellite. If the review is successful, the ground issues a permission instruction to implement orbit control, and the satellite executes the orbit control operation according to the strategy. If the review is unsuccessful, the ground sends the orbit maintenance strategy back to the satellite.

[0158] The aforementioned space-ground integrated autonomous orbit control decision-making process is as follows:

[0159] Step S4.1: The ground control center sends an instruction to allow the satellite to autonomously maintain its orbit. Once the phase deviation condition is met, the satellite autonomously generates an orbit-maintaining strategy.

[0160] Step S4.2: The autonomous orbit maintenance strategy is sent to the satellite integrated electronic subsystem for review. If approved, proceed to the next step; otherwise, cancel the current autonomous orbit control.

[0161] Step S4.3: The autonomous orbit control strategy is telemetryally transmitted to the ground, and the ground control center reviews it. If the ground approves the implementation, the next step is executed; otherwise, the autonomous orbit control is canceled and the orbit holding strategy is uploaded from the ground.

[0162] Step S4.4: Implement the autonomous orbit holding strategy. Before the orbit control thruster ignites, the orbit control subsystem sends a state establishment request to the propulsion subsystem at a preset time point to establish the propulsion state.

[0163] Step S4.5: The attitude and orbit control subsystem controls the orbit control thruster to ignite;

[0164] Step S4.6: After the orbit control ignition is completed, the attitude and orbit control subsystem sends a request to shut down the propulsion subsystem, and the integrated electronic subsystem shuts down the propulsion subsystem;

[0165] Step S4.7: After the track control is completed, the accuracy of the track control is calibrated using GNSS navigation data;

[0166] By installing thrusters in both directions of satellite flight, the satellite has the ability to control its ascent and descent orbits in both forward and inverted directions within the orbital coordinate system.

[0167] Example 2:

[0168] Example 2 is a preferred embodiment of Example 1, and is used to illustrate the present invention in more detail.

[0169] The present invention also provides an autonomous orbit maintenance system for a low-Earth orbit inter-satellite laser communication constellation. The autonomous orbit maintenance system for a low-Earth orbit inter-satellite laser communication constellation can be implemented by executing the process steps of the autonomous orbit maintenance method for a low-Earth orbit inter-satellite laser communication constellation. That is, those skilled in the art can understand the autonomous orbit maintenance method for a low-Earth orbit inter-satellite laser communication constellation as a preferred embodiment of the autonomous orbit maintenance system for a low-Earth orbit inter-satellite laser communication constellation.

[0170] An autonomous orbit-keeping system for a low-Earth orbit inter-satellite laser communication constellation, according to the present invention, comprises:

[0171] Module M1: Generates the upper and lower limits of the satellite phase-preservation threshold and updates the thresholds periodically based on the annual changes in laser link performance;

[0172] Specifically, in module M1:

[0173] The satellite phase preservation threshold calculation process is as follows: the maximum phase difference between two linked satellites is calculated based on the low-Earth orbit constellation altitude, the longest link establishment distance constraint of the inter-satellite laser communication link, and the minimum link establishment altitude constraint above the Earth's surface. Using the absolute phase preservation mode, the upper and lower limits of the satellite phase preservation threshold are set. Based on the on-orbit laser link telemetry data, the communication performance is analyzed, and the longest link establishment distance and the minimum link establishment altitude constraint of the inter-satellite laser communication link are updated. The phase preservation threshold is updated once every preset time period.

[0174] Generate satellite phase preservation threshold Δu lim Low Earth orbit constellation orbital altitude H sat The longest connection establishment distance constraint for inter-satellite laser communication links in a constellation is L. max The lowest chain height above the ground is H. min The maximum phase difference between the two linked satellites is u. max It is calculated by the following formula:

[0175] u res1 =2sin -1 (L max / 2 / (H sat +R E ))

[0176] u res2 =2cos -1 ((H min +R E ) / (H sat +R E ))

[0177] u max =min(u res1 ,u res2 )

[0178] In the formula, R E U is the Earth's radius. res1 u is the maximum inter-satellite phase difference constrained by the inter-satellite laser communication distance. res2 The maximum inter-satellite phase difference constrained by the minimum link establishment altitude for inter-satellite communication;

[0179] The maximum nominal phase difference between two linked satellites is u nom The absolute phase-preserving mode is adopted, and the satellite phase-preserving threshold is set to the upper limit Δu. lim =(u max -u nom ) / 2, with the lower limit set to -Δu lim ;

[0180] Based on the analysis of communication performance using on-orbit laser link telemetry data, the longest link establishment distance L... max The allowable height value H for shortening and building chains min The phase hold threshold Δu is raised and updated every preset time interval. lim This ensures that satellites within the constellation remain connected.

[0181] Module M2: Based on the autonomous positioning results of the onboard GNSS navigation receiver, recursively calculate the satellite's horizontal orbit elements at the current moment, and calculate the horizontal latitude angle deviation relative to the nominal orbit;

[0182] Specifically, in module M2:

[0183] Latitude Aspect Deviation The calculation process is as follows: Based on the real-time position and velocity of the satellite at the time of resolution by the GNSS navigation receiver, the mean orbit elements corresponding to the current time are obtained through numerical iteration based on the Lagrange planetary equations; the mean latitude argument at the current time is calculated and compared with the nominal mean orbit elements to obtain the satellite mean latitude argument deviation relative to the nominal orbit at the current time.

[0184] Given the real-time satellite position r0 and velocity v0 at time t0 calculated by the GNSS navigation receiver, determine the oscillological elements [ae iΩωM] at time t0 using Kepler's equations. T Where a is the semi-major axis of the auscultatory orbit, e is the eccentricity of the auscultatory orbit, i is the inclination of the auscultatory orbit, Ω is the right ascension of the ascending node of the auscultatory orbit, ω is the argument of the perigee of the auscultatory orbit, and M is the mean perigee of the auscultatory orbit; the current time t is obtained through numerical iteration. CT Corresponding horizontal orbital elements in, For the semi-major axis of the horizontal track, For the eccentricity of the horizontal track, For the inclination angle of the horizontal track, The right ascension of the ascending node of the horizontal orbit. The perigee angle of the horizontal orbit. The iterative formula for the aperitone angle of the horizontal orbit is shown below:

[0185]

[0186] In the formula, and The first-order variation of the orbital elements with time caused by the perturbation is obtained according to the Lagrange planetary equations. The initial value for the iterative calculation of the horizontal orbital elements is [ae iΩωM+n(t CT -t0)] T where n is the angular velocity of the satellite's orbital motion;

[0187] Let the current time t be... CT Argument of latitude is Compare with the nominal orbital square root number to calculate the current time t. CT Satellite mean latitude angle deviation relative to nominal orbit

[0188] Module M3: When the satellite's latitude angle deviation and phase holding threshold meet the preset conditions, the satellite adopts the limit loop phase holding method to autonomously calculate and generate an autonomous orbit control strategy.

[0189] Specifically, in module M3:

[0190] The satellite employs a limit cycle phase holding method, using orbit holding control to offset the semi-major axis, forming a drift control loop. Given the current time, the orbit control implementation time, and the nominal orbital elements, the orbit control implementation time t is calculated according to the principle of the limit cycle phase holding method. f The required velocity increment δV can be calculated and the satellite can autonomously generate an orbit maintenance strategy. The satellite's autonomous orbit maintenance is set to a quasi-prohibited state by default, and is triggered only when the ground control center issues a permission command.

[0191] Onboard autonomous orbit-keeping strategy: When the satellite's mean latitude angle deviation... At that time, the satellite employs the extreme loop phase-keeping method to autonomously calculate and generate an autonomous orbit control strategy. f A velocity increment δV is applied continuously and updated every preset time interval;

[0192] The satellite employs a limit loop phase-keeping method, using orbit-keeping control to offset the semi-major axis, forming a drift control loop. Initially, a positive offset of the semi-major axis is applied, causing the phase to shift to the left. As the semi-major axis decays, it gradually falls below its nominal value, initiating a rightward phase shift. When the satellite reaches the right boundary, orbit-keeping control δa is implemented to maintain the phase deviation within -Δu. lim With Δu lim between;

[0193] make

[0194] k1 = n CT -n N

[0195]

[0196] In the formula, μ is the Earth's gravitational constant. Let t be the current time. CT Orbital angular rate, a CT For t CT The semi-major axis of the horizontal orbit at any given moment; For the nominal orbital angular rate, a Nk1 represents the semi-major axis of the nominal orbit; k2 is the difference in angular rate between the current orbit and the nominal orbit; and k2 is the coefficient of variation of the orbital angular rate considering the influence of atmospheric drag. This represents the first-order rate of change of the semi-major axis caused by atmospheric drag.

[0197] The phase holding period is then

[0198]

[0199] The target value for the semi-major axis deviation of phase preservation is

[0200]

[0201] When the satellite reaches boundary C, the actual semi-major axis deviation is

[0202]

[0203] In the formula: Δa N =a CT -a N t represents the deviation between the current actual semi-major axis and the nominal semi-major axis. f It is the moment when the satellite reaches boundary C or the moment when orbit control is implemented;

[0204] When the satellite reaches boundary C, the change in the semi-major axis of the satellite's orbital maneuver is:

[0205] δa=Δa T -Δa f

[0206] For near-Earth circular orbits, the corresponding velocity increment requirement is:

[0207]

[0208] The satellite is in a state of default prohibition for maintaining its autonomous orbit. It is in a state of prohibition by default and will be triggered only when a permission command is issued from the ground.

[0209] Module M4: After generating the orbit maintenance strategy, it is transmitted to the ground control center. According to the space-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit maintenance plan generated autonomously on the satellite. If the review is successful, the ground sends a command to allow orbit control to be implemented, and the satellite executes the orbit control operation according to the strategy. If the review is unsuccessful, the ground will re-inject the orbit maintenance strategy.

[0210] Specifically, in module M4:

[0211] Space-Ground Integrated Autonomous Orbit Control Decision and Execution: After the satellite autonomously generates its orbit maintenance strategy, it transmits it to the ground control center at a pre-set time. According to the space-ground integrated autonomous orbit control decision process, the ground control center reviews the orbit maintenance plan generated by the satellite. If the review is successful, the ground issues a permission instruction to implement orbit control, and the satellite executes the orbit control operation according to the strategy. If the review is unsuccessful, the ground sends the orbit maintenance strategy back to the satellite.

[0212] The aforementioned space-ground integrated autonomous orbit control decision-making process is as follows:

[0213] Module M4.1: The ground control center sends instructions to allow the satellite to autonomously maintain its orbit. Once the phase deviation condition is met, the satellite autonomously generates an orbit-maintaining strategy.

[0214] Module M4.2: The autonomous orbit maintenance strategy is sent to the satellite integrated electronic subsystem for review. If approved, the next step is executed; otherwise, the current autonomous orbit control is canceled.

[0215] Module M4.3: The autonomous orbit control strategy is telemetryally transmitted to the ground, and the ground control center reviews it. If the ground approves the implementation, the next step is executed; otherwise, the autonomous orbit control is canceled and the orbit holding strategy is uploaded from the ground.

[0216] Module M4.4: Implemented according to the autonomous orbit holding strategy, the orbit control subsystem sends a state establishment request to the propulsion subsystem at a preset time point before the orbit control thruster ignites, thus establishing the propulsion state;

[0217] Module M4.5: The attitude and orbit control subsystem controls the orbit control thruster to ignite;

[0218] Module M4.6: After the track control ignition is completed, the attitude and track control subsystem sends a request to shut down the propulsion subsystem, and the integrated electronic subsystem shuts down the propulsion subsystem;

[0219] Module M4.7: After track control is completed, the accuracy of the track control is calibrated using GNSS navigation data;

[0220] By installing thrusters in both directions of satellite flight, the satellite has the ability to control its ascent and descent orbits in both forward and inverted directions within the orbital coordinate system.

[0221] Example 3:

[0222] Example 3 is a preferred example of Example 1, and is used to illustrate the present invention in more detail.

[0223] The objective of this invention is achieved through the following technical solution: an autonomous orbit maintenance method for a low-Earth orbit inter-satellite laser communication constellation, the method comprising the following steps: Step A, generating and periodically updating a constellation phase maintenance threshold: To meet the constant maintenance requirements of inter-satellite laser links in low-Earth orbit constellations, a satellite phase maintenance threshold Δu is generated. limThe threshold is updated periodically based on the annual changes in laser link performance; Step B, calculate the satellite horizontal orbit element deviation based on the GNSS navigation receiver: based on the autonomous positioning results of the onboard GNSS navigation receiver, the satellite horizontal orbit element at the current moment is recursively calculated, and the horizontal latitude argument deviation relative to the nominal orbit is calculated. Step C, Onboard Autonomous Generation of Orbit Maintenance Strategy: When the satellite's mean latitude angle deviation... At that time, the satellite employs the extreme loop phase-keeping method to autonomously calculate and generate an autonomous orbit control strategy (t). f Apply velocity increment δV continuously, updating every hour; Step D, Space-Ground Integrated Autonomous Orbit Control Decision and Execution: After the satellite autonomously generates the orbit maintenance strategy, it is transmitted to the ground control center 24 hours in advance. According to the space-ground integrated autonomous orbit control decision process, the ground control center reviews the orbit maintenance plan autonomously generated by the satellite. If the review is successful, the ground issues a permission instruction for orbit control implementation, and the satellite executes the orbit control operation according to the strategy. If the review fails, the ground injects the orbit maintenance strategy.

[0224] Combination Figure 1 As shown, the method includes the following steps:

[0225] Step A: Generate and periodically update the constellation phase preservation threshold: To meet the constant maintenance requirements of inter-satellite laser links in low-Earth orbit constellations, generate the satellite phase preservation threshold Δu. lim The thresholds are updated regularly based on the annual changes in laser link performance.

[0226] Low Earth Orbit Constellation Orbital Altitude H sat The longest connection establishment distance constraint for inter-satellite laser communication links in a constellation is L. max The lowest chain height above the ground is H. min The maximum phase difference between the two linked satellites is u. max It is calculated using the following formula.

[0227] u res1 =2sin -1 (L max / 2 / (H sat +R E ))

[0228] u res2 =2cos -1 ((H min +R E ) / (H sat +R E ))

[0229] u max =min(u res1 ,u res2 )

[0230] In the formula, R E U is the Earth's radius. res1 u is the maximum inter-satellite phase difference constrained by the inter-satellite laser communication distance. res2 This is the maximum inter-satellite phase difference constrained by the minimum link establishment altitude for inter-satellite communication.

[0231] The maximum nominal phase difference between two linked satellites is u nom Considering the use of absolute phase-preserving mode, the satellite phase-preserving threshold is set to an upper limit of Δu. lim =(u max -u nom ) / 2, with the lower limit set to -Δu lim .

[0232] Over time, the performance of the onboard laser link will decline year by year. Based on the analysis of communication performance using telemetry data of the on-orbit laser link, the longest link establishment distance L... max It will shorten the allowable height value H of the chain. min It will rise, and the phase preservation threshold Δu needs to be updated annually. lim This is to ensure that the satellites within the constellation remain connected.

[0233] Step B, Calculate satellite horizontal orbit element deviation based on GNSS navigation receiver: Based on the autonomous positioning results of the onboard GNSS navigation receiver, recursively calculate the satellite horizontal orbit elements at the current moment, and calculate the horizontal latitude argument deviation relative to the nominal orbit.

[0234] Given the real-time satellite position r0 and velocity v0 at time t0 calculated by the GNSS navigation receiver, determine the oscillological elements [ae iΩωM] at time t0 using Kepler's equations. T (a is the semi-major axis of the close orbit, e is the eccentricity of the close orbit, i is the inclination of the close orbit, Ω is the right ascension of the ascending node of the close orbit, ω is the argument of the perigee of the close orbit, and M is the mean perigee of the close orbit.) The current time t is then calculated through numerical iteration. CT Corresponding horizontal orbital elements ( For the semi-major axis of the horizontal track, For the eccentricity of the horizontal track, For the inclination angle of the horizontal track, The right ascension of the ascending node of the horizontal orbit. The perigee angle of the horizontal orbit. (For the horizontal orbital near-point angle), the iterative formula is shown below.

[0235]

[0236] In the formula, and The first-order variation of the orbital elements with time caused by the perturbation is obtained according to the Lagrange planetary equations. The initial value for the iterative calculation of the horizontal orbital elements is [ae iΩωM+n(t CT -t0)] T (n is the angular velocity of the satellite's orbital motion).

[0237] Record the current t CT The angle of the mean latitude at any given time is Compare the current t with the nominal orbital square root number to obtain the current t. CT The deviation of the satellite's mean latitude angle from its nominal orbit at any given time

[0238] Step C, Onboard Autonomous Generation of Orbit Maintenance Strategy: When the satellite's mean latitude angle deviation... At that time, the satellite employs the extreme loop phase-keeping method to autonomously calculate and generate an autonomous orbit control strategy (t). f A velocity increment δV is applied at all times and updated every hour.

[0239] The satellite employs a limit loop phase-keeping method, using orbit-keeping control to offset the semi-major axis, forming a drift control loop, such as... Figure 2 As shown in the diagram, a positive semi-major axis offset is first applied (point A), and the phase begins to shift to the left (moving from A to B). As the semi-major axis decays, it gradually falls below the nominal value, and the phase begins to shift to the right (moving from B to C). When the satellite reaches the right boundary (point C), orbital maintenance control δa is implemented (moving from C to A), thereby maintaining the phase deviation at -Δu. lim With Δu lim between.

[0240] make

[0241] k1 = n CT -n N

[0242]

[0243] In the formula, μ is the Earth's gravitational constant. Let t be the current time. CT Orbital angular rate, a CT For t CT The semi-major axis of the horizontal orbit at any given moment; For the nominal orbital angular rate, a N k1 represents the semi-major axis of the nominal orbit; k2 is the difference in angular rate between the current orbit and the nominal orbit; and k2 is the coefficient of variation of the orbital angular rate considering the influence of atmospheric drag. This represents the first-order rate of change of the semi-major axis caused by atmospheric drag.

[0244] The phase holding period is then

[0245]

[0246] The target value for the semi-major axis deviation of phase preservation is

[0247]

[0248] When the satellite reaches boundary C, the actual semi-major axis deviation is

[0249]

[0250] In the formula: Δa N =a CT -a N t represents the deviation between the current actual semi-major axis and the nominal semi-major axis. f It refers to the moment when the satellite reaches boundary C or the moment when orbit control is implemented.

[0251] When the satellite reaches boundary C, the change in the semi-major axis of the satellite's orbital maneuver is:

[0252] δa=Δa T -Δa f

[0253] For near-Earth circular orbits, the corresponding velocity increment requirement is:

[0254]

[0255] The satellite is in a state of default prohibition for maintaining its autonomous orbit. It is in a state of prohibition by default and will be triggered only when a permission command is issued from the ground.

[0256] Step D, Space-Ground Integrated Autonomous Orbit Control Decision and Execution: After the satellite autonomously generates the orbit maintenance strategy, it is transmitted to the ground control center 24 hours in advance. According to the space-ground integrated autonomous orbit control decision process, the ground control center reviews the orbit maintenance plan generated by the satellite. If the review is successful, the ground issues a permission instruction to implement orbit control, and the satellite executes the orbit control operation according to the strategy. If the review fails, the ground sends the orbit maintenance strategy.

[0257] like Figure 3 As shown, the specific decision-making process for integrated space-ground autonomous orbit control is as follows:

[0258] A. The ground control center sends an instruction to allow the satellite to autonomously maintain its orbit. Once the phase deviation condition is met, the satellite autonomously generates an orbit-maintaining strategy.

[0259] B. The autonomous orbit maintenance strategy is sent to the satellite integrated electronic subsystem for review. If approved, proceed to the next step; otherwise, cancel the current autonomous orbit control.

[0260] C. The autonomous orbit control strategy is telemetryally transmitted to the ground, and the ground control center reviews it. If the ground approves the implementation, the next step is executed; otherwise, the autonomous orbit control is canceled and the orbit maintenance strategy is uploaded from the ground.

[0261] D. Implementing the autonomous orbit maintenance strategy, the attitude and orbit control subsystem sends a state-establishment request to the propulsion subsystem 30 minutes before the ignition of the orbit control thruster to establish the propulsion state;

[0262] E. The attitude and orbit control subsystem controls the orbit control thruster to ignite;

[0263] F. After the orbit control ignition is completed, the attitude and orbit control subsystem sends a request to shut down the propulsion subsystem, and the integrated electronic subsystem shuts down the propulsion subsystem.

[0264] G. After the track control is completed, the accuracy of the track control is determined by the navigation data.

[0265] To avoid link establishment interruptions caused by attitude maneuvers, thrusters are installed in both the forward and reverse directions of satellite flight. It also has the capability to control the ascent and descent of the orbital system in both forward and inverted flight modes. Therefore, attitude maneuvers are not required when switching from the normal Earth-oriented attitude to the orbital control mode, thus preventing the interruption of the inter-satellite laser communication link caused by attitude changes.

[0266] Those skilled in the art will understand that, besides implementing the system and its various devices, modules, and units provided by this invention in the form of purely computer-readable program code, the same functions can be achieved entirely through logical programming of the method steps, making the system and its various devices, modules, and units of this invention function in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers. Therefore, the system and its various devices, modules, and units provided by this invention can be considered as a hardware component, and the devices, modules, and units included therein for implementing various functions can also be considered as structures within the hardware component; alternatively, the devices, modules, and units for implementing various functions can be considered as both software modules implementing the method and structures within the hardware component.

[0267] Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. A method for autonomous orbit maintenance of a low-Earth orbit inter-satellite laser communication constellation, characterized in that, include: Step S1: Generate the upper and lower limits of the satellite phase preservation threshold, and update the thresholds periodically according to the annual changes in laser link performance; Step S2: Based on the autonomous positioning results of the onboard GNSS navigation receiver, recursively calculate the satellite's horizontal orbit elements at the current moment, and calculate the horizontal latitude argument deviation relative to the nominal orbit; Step S3: When the satellite's latitude argument deviation and phase holding threshold meet the preset conditions, the satellite adopts the limit loop phase holding method to autonomously calculate and generate an autonomous orbit control strategy. Step S4: After generating the orbit maintenance strategy, it is transmitted to the ground control center. According to the space-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit maintenance plan generated autonomously on the satellite. If the review is successful, the ground sends an instruction to allow orbit control to be implemented, and the satellite executes the orbit control operation according to the strategy. If the review is unsuccessful, the ground sends the orbit maintenance strategy. In step S3: The satellite employs a limit cycle phase holding method, using orbit holding control to offset the semi-major axis, forming a drift control loop. Given the current time, the orbit control implementation time, and the nominal orbital elements, the orbit control implementation time t is calculated according to the principle of the limit cycle phase holding method. f The required velocity increment δV is calculated, and the satellite autonomously generates an orbit maintenance strategy based on the calculation results. The satellite's autonomous orbit maintenance is set to a quasi-prohibited state by default, and is triggered only when the ground control center issues a permission command. Onboard autonomous orbit-keeping strategy: When the satellite's mean latitude angle deviation... At that time, the satellite employs the extreme loop phase-keeping method to autonomously calculate and generate an autonomous orbit control strategy. f A velocity increment δV is applied continuously and updated every preset time interval; The satellite employs a limit loop phase-keeping method, using orbit-keeping control to offset the semi-major axis, forming a drift control loop. Initially, a positive offset of the semi-major axis is applied, causing the phase to shift to the left. As the semi-major axis decays, it gradually falls below its nominal value, initiating a rightward phase shift. When the satellite reaches the right boundary, orbit-keeping control δa is implemented to maintain the phase deviation within -Δu. lim With Δu lim between; The maximum nominal phase difference between two linked satellites is u nom The absolute phase-preserving mode is adopted, and the satellite phase-preserving threshold is set to the upper limit of Δu. lim =(u max -u nom ) / 2, with the lower limit set to -Δu lim .

2. The autonomous orbit maintenance method for low-Earth orbit inter-satellite laser communication constellations according to claim 1, characterized in that, In step S1: The satellite phase preservation threshold calculation process is as follows: the maximum phase difference between two linked satellites is calculated based on the low-Earth orbit constellation altitude, the longest link establishment distance constraint of the inter-satellite laser communication link, and the minimum link establishment altitude constraint above the Earth's surface. Using the absolute phase preservation mode, the upper and lower limits of the satellite phase preservation threshold are set. Based on the on-orbit laser link telemetry data, the communication performance is analyzed, and the longest link establishment distance and the minimum link establishment altitude constraint of the inter-satellite laser communication link are updated. The phase preservation threshold is updated once every preset time period. Generate satellite phase preservation threshold Δu lim Low Earth orbit constellation orbital altitude H sat The longest connection establishment distance constraint for inter-satellite laser communication links in a constellation is L. max The lowest chain height above the ground is H. min The maximum phase difference between the two linked satellites is u. max It is calculated by the following formula: u res1 *2sin -1 (L max / 2 / (H sat +R E )) u res2 =2cos -1 ((H min +R E ) / (H sat +R E )) in max =min(in res1 ,in res2 ) In the formula, R E U is the Earth's radius. res1 u is the maximum inter-satellite phase difference constrained by the inter-satellite laser communication distance. res2 The maximum inter-satellite phase difference constrained by the minimum link establishment altitude for inter-satellite communication; The maximum nominal phase difference between two linked satellites is u nom The absolute phase-preserving mode is adopted, and the satellite phase-preserving threshold is set to the upper limit of Δu. lim =(u max -u nom ) / 2, with the lower limit set to -Δu lim ; Based on the analysis of communication performance using on-orbit laser link telemetry data, the longest link establishment distance L... max The allowable height H for shortening and building chains min The phase hold threshold Δu is raised and updated every preset time interval. lim This ensures that satellites within the constellation remain connected.

3. The autonomous orbit maintenance method for low-Earth orbit inter-satellite laser communication constellations according to claim 1, characterized in that, In step S2: Latitude Aspect Deviation The calculation process is as follows: Based on the real-time position and velocity of the satellite at the time of solution by the GNSS navigation receiver, the horizontal orbital elements corresponding to the current time are obtained through numerical iteration based on the Lagrange planetary equations. Calculate the mean latitude argument at the current moment and compare it with the nominal mean orbit elements to determine the satellite mean latitude argument deviation relative to the nominal orbit at the current moment. Given the real-time satellite position r0 and velocity v0 at time t0 calculated by the GNSS navigation receiver, determine the oscillological elements [ae iΩωM] at time t0 using Kepler's equations. T Where a is the semi-major axis of the auscultatory orbit, e is the eccentricity of the auscultatory orbit, i is the inclination of the auscultatory orbit, Ω is the right ascension of the ascending node of the auscultatory orbit, ω is the argument of the perigee of the auscultatory orbit, and M is the mean perigee of the auscultatory orbit; the current time t is obtained through numerical iteration. CT The corresponding number of horizontal orbital elements [ in, For the semi-major axis of the horizontal track, For the eccentricity of the horizontal track, For the inclination angle of the horizontal track, The right ascension of the ascending node of the horizontal orbit. The perigee angle of the horizontal orbit. The iterative formula for the aperitone angle of the horizontal orbit is shown below: In the formula, and The first-order variation of the orbital elements with time caused by the perturbation is obtained according to the Lagrange planetary equations. The initial value for the iterative calculation of the horizontal orbital elements is [ae iΩωM+n(t CT -t0)] T where n is the angular velocity of the satellite's orbital motion; Let the current time t be... CT Argument of latitude is Compare with the nominal orbital square root number to calculate the current time t. CT Satellite mean latitude angle deviation relative to nominal orbit 4. The autonomous orbit maintenance method for low-Earth orbit inter-satellite laser communication constellations according to claim 1, characterized in that, make k1=n CT -n N In the formula, μ is the Earth's gravitational constant. Let t be the current time. CT Orbital angular rate, a CT For t CT The semi-major axis of the horizontal orbit at any given moment; For the nominal orbital angular rate, a N k1 represents the semi-major axis of the nominal orbit; k2 is the difference in angular rate between the current orbit and the nominal orbit; and k2 is the coefficient of variation of the orbital angular rate considering the influence of atmospheric drag. This represents the first-order rate of change of the semi-major axis caused by atmospheric drag. The phase holding period is then... The target value for the semi-major axis deviation of phase preservation is When the satellite reaches boundary C, the actual semi-major axis deviation is In the formula: Δa N =a CT -a N t represents the deviation between the current actual semi-major axis and the nominal semi-major axis. f It is the moment when the satellite reaches boundary C or the moment when orbit control is implemented; When the satellite reaches boundary C, the change in the semi-major axis of the satellite's orbital maneuver is: δa=Δa T -Δa f For near-Earth circular orbits, the corresponding velocity increment requirement is: The satellite is in a state of default prohibition for maintaining its autonomous orbit. It is in a state of prohibition by default and will be triggered only when a permission command is issued from the ground.

5. The autonomous orbit maintenance method for low-Earth orbit inter-satellite laser communication constellations according to claim 1, characterized in that, In step S4: Space-Ground Integrated Autonomous Orbit Control Decision and Execution: After the satellite autonomously generates its orbit maintenance strategy, it transmits it to the ground control center at a pre-set time. According to the space-ground integrated autonomous orbit control decision process, the ground control center reviews the orbit maintenance plan generated by the satellite. If the review is successful, the ground issues a permission instruction to implement orbit control, and the satellite executes the orbit control operation according to the strategy. If the review is unsuccessful, the ground sends the orbit maintenance strategy back to the satellite. The aforementioned space-ground integrated autonomous orbit control decision-making process is as follows: Step S4.1: The ground control center sends an instruction to allow the satellite to autonomously maintain its orbit. Once the phase deviation condition is met, the satellite autonomously generates an orbit-maintaining strategy. Step S4.2: The autonomous orbit maintenance strategy is sent to the satellite integrated electronic subsystem for review. If approved, proceed to the next step; otherwise, cancel the current autonomous orbit control. Step S4.3: The autonomous orbit control strategy is telemetryally transmitted to the ground, and the ground control center reviews it. If the ground approves the implementation, the next step is executed; otherwise, the autonomous orbit control is canceled and the orbit holding strategy is uploaded from the ground. Step S4.4: Implement the autonomous orbit holding strategy. Before the orbit control thruster ignites, the orbit control subsystem sends a state establishment request to the propulsion subsystem at a preset time point to establish the propulsion state. Step S4.5: The attitude and orbit control subsystem controls the orbit control thruster to ignite; Step S4.6: After the orbit control ignition is completed, the attitude and orbit control subsystem sends a request to shut down the propulsion subsystem, and the integrated electronic subsystem shuts down the propulsion subsystem; Step S4.7: After the track control is completed, the accuracy of the track control is calibrated using GNSS navigation data; By installing thrusters in both directions of satellite flight, the satellite has the ability to control its ascent and descent orbits in both forward and inverted directions within the orbital coordinate system.

6. An autonomous orbit-keeping system for a low-Earth orbit inter-satellite laser communication constellation, characterized in that, include: Module M1: Generates the upper and lower limits of the satellite phase-preservation threshold and updates the thresholds periodically based on the annual changes in laser link performance; Module M2: Based on the autonomous positioning results of the onboard GNSS navigation receiver, recursively calculate the satellite's horizontal orbit elements at the current moment, and calculate the horizontal latitude angle deviation relative to the nominal orbit; Module M3: When the satellite's latitude angle deviation and phase holding threshold meet the preset conditions, the satellite adopts the limit loop phase holding method to autonomously calculate and generate an autonomous orbit control strategy. Module M4: After generating the orbit maintenance strategy, it is transmitted to the ground control center. According to the space-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit maintenance plan generated autonomously on the satellite. If the review is successful, the ground sends an instruction to allow orbit control to be implemented, and the satellite executes the orbit control operation according to the strategy. If the review is unsuccessful, the ground sends the orbit maintenance strategy. In module M3: The satellite employs a limit cycle phase holding method, using orbit holding control to offset the semi-major axis, forming a drift control loop. Given the current time, the orbit control implementation time, and the nominal orbital elements, the orbit control implementation time t is calculated according to the principle of the limit cycle phase holding method. f The required velocity increment δV is calculated, and the satellite autonomously generates an orbit maintenance strategy based on the calculation results. The satellite's autonomous orbit maintenance is set to a quasi-prohibited state by default, and is triggered only when the ground control center issues a permission command. Onboard autonomous orbit-keeping strategy: When the satellite's mean latitude angle deviation... At that time, the satellite employs the extreme loop phase-keeping method to autonomously calculate and generate an autonomous orbit control strategy. f A velocity increment δV is applied continuously and updated every preset time interval; The satellite employs a limit loop phase-keeping method, using orbit-keeping control to offset the semi-major axis, forming a drift control loop. Initially, a positive offset of the semi-major axis is applied, causing the phase to shift to the left. As the semi-major axis decays, it gradually falls below its nominal value, initiating a rightward phase shift. When the satellite reaches the right boundary, orbit-keeping control δa is implemented to maintain the phase deviation within -Δu. lim With Δu lim between; The maximum nominal phase difference between two linked satellites is u nom The absolute phase-preserving mode is adopted, and the satellite phase-preserving threshold is set to the upper limit of Δu. lim =(u max -u nom ) / 2, with the lower limit set to -Δu lim .

7. The autonomous orbit maintenance system for a low-Earth orbit inter-satellite laser communication constellation according to claim 6, characterized in that, In module M1: The satellite phase preservation threshold calculation process is as follows: the maximum phase difference between two linked satellites is calculated based on the low-Earth orbit constellation altitude, the longest link establishment distance constraint of the inter-satellite laser communication link, and the minimum link establishment altitude constraint above the Earth's surface. Using the absolute phase preservation mode, the upper and lower limits of the satellite phase preservation threshold are set. Based on the on-orbit laser link telemetry data, the communication performance is analyzed, and the longest link establishment distance and the minimum link establishment altitude constraint of the inter-satellite laser communication link are updated. The phase preservation threshold is updated once every preset time period. Generate satellite phase preservation threshold Δu lim Low Earth orbit constellation orbital altitude H sat The longest connection establishment distance constraint for inter-satellite laser communication links in a constellation is L. max The lowest chain height above the ground is H. min The maximum phase difference between the two linked satellites is u. max It is calculated by the following formula: u res1 *2sin -1 (L max / 2 / (H sat +R E )) u res2 =2cos -1 ((H min +R E ) / (H sat +R E )) in max =min(in res1 ,in res2 ) In the formula, R E u is the Earth's radius. res1 u is the maximum inter-satellite phase difference constrained by the inter-satellite laser communication distance. res2 The maximum inter-satellite phase difference constrained by the minimum link establishment altitude for inter-satellite communication; Based on the analysis of communication performance using on-orbit laser link telemetry data, the longest link establishment distance L... max The allowable height H for shortening and building chains min The phase hold threshold Δu is raised and updated every preset time interval. lim This ensures that satellites within the constellation remain connected.

8. The autonomous orbit maintenance system for a low-Earth orbit inter-satellite laser communication constellation according to claim 6, characterized in that, In module M2: Latitude Aspect Deviation The calculation process is as follows: Based on the real-time position and velocity of the satellite at the time of solution by the GNSS navigation receiver, the horizontal orbital elements corresponding to the current time are obtained through numerical iteration based on the Lagrange planetary equations. Calculate the mean latitude argument at the current moment and compare it with the nominal mean orbit elements to determine the satellite mean latitude argument deviation relative to the nominal orbit at the current moment. Given the real-time satellite position r0 and velocity v0 at time t0 calculated by the GNSS navigation receiver, determine the oscillological elements [ae iΩωM] at time t0 using Kepler's equations. T Where a is the semi-major axis of the auscultatory orbit, e is the eccentricity of the auscultatory orbit, i is the inclination of the auscultatory orbit, Ω is the right ascension of the ascending node of the auscultatory orbit, ω is the argument of the perigee of the auscultatory orbit, and M is the mean perigee of the auscultatory orbit; the current time t is obtained through numerical iteration. CT The corresponding number of horizontal orbital elements [ in, For the semi-major axis of the horizontal track, For the eccentricity of the horizontal track, For the inclination angle of the horizontal track, The right ascension of the ascending node of the horizontal orbit. The perigee angle of the horizontal orbit. The iterative formula for the aperitone angle of the horizontal orbit is shown below: In the formula, and The first-order variation of the orbital elements with time caused by the perturbation is obtained according to the Lagrange planetary equations. The initial value for the iterative calculation of the horizontal orbital elements is [ae iΩωM+n(t CT -t0)] T where n is the angular velocity of the satellite's orbital motion; Let the current time t be... CT Argument of latitude is Compare with the nominal orbital square root number to calculate the current time t. CT Satellite mean latitude angle deviation relative to nominal orbit 9. The autonomous orbit maintenance system for a low-Earth orbit inter-satellite laser communication constellation according to claim 6, characterized in that, make k1=n CT -n N In the formula, μ is the Earth's gravitational constant. Let t be the current time. CT Orbital angular rate, a CT For t CT The semi-major axis of the horizontal orbit at any given moment; For the nominal orbital angular rate, a N k1 represents the semi-major axis of the nominal orbit; k2 is the difference in angular rate between the current orbit and the nominal orbit; and k2 is the coefficient of variation of the orbital angular rate considering the influence of atmospheric drag. This represents the first-order rate of change of the semi-major axis caused by atmospheric drag. The phase holding period is then... The target value for the semi-major axis deviation of phase preservation is When the satellite reaches boundary C, the actual semi-major axis deviation is In the formula: Δa N =a CT -a N t represents the deviation between the current actual semi-major axis and the nominal semi-major axis. f It is the moment when the satellite reaches boundary C or the moment when orbit control is implemented; When the satellite reaches boundary C, the change in the semi-major axis of the satellite's orbital maneuver is: δa=Δa T -Δa f For near-Earth circular orbits, the corresponding velocity increment requirement is: The satellite is in a state of default prohibition for maintaining its autonomous orbit. It is in a state of prohibition by default and will be triggered only when a permission command is issued from the ground.

10. The autonomous orbit maintenance system for a low-Earth orbit inter-satellite laser communication constellation according to claim 6, characterized in that, In module M4: Space-Ground Integrated Autonomous Orbit Control Decision and Execution: After the satellite autonomously generates its orbit maintenance strategy, it transmits it to the ground control center at a pre-set time. According to the space-ground integrated autonomous orbit control decision process, the ground control center reviews the orbit maintenance plan generated by the satellite. If the review is successful, the ground issues a permission instruction to implement orbit control, and the satellite executes the orbit control operation according to the strategy. If the review is unsuccessful, the ground sends the orbit maintenance strategy back to the satellite. The aforementioned space-ground integrated autonomous orbit control decision-making process is as follows: Module M4.1: The ground control center sends instructions to allow the satellite to autonomously maintain its orbit. Once the phase deviation condition is met, the satellite autonomously generates an orbit-maintaining strategy. Module M4.2: The autonomous orbit maintenance strategy is sent to the satellite integrated electronic subsystem for review. If approved, the next step is executed; otherwise, the current autonomous orbit control is canceled. Module M4.3: The autonomous orbit control strategy is telemetryally transmitted to the ground, and the ground control center reviews it. If the ground approves the implementation, the next step is executed; otherwise, the autonomous orbit control is canceled and the orbit holding strategy is uploaded from the ground. Module M4.4: Implemented according to the autonomous orbit holding strategy, the orbit control subsystem sends a state establishment request to the propulsion subsystem at a preset time point before the orbit control thruster ignites, thus establishing the propulsion state; Module M4.5: The attitude and orbit control subsystem controls the orbit control thruster to ignite; Module M4.6: After the track control ignition is completed, the attitude and track control subsystem sends a request to shut down the propulsion subsystem, and the integrated electronic subsystem shuts down the propulsion subsystem; Module M4.7: After track control is completed, the accuracy of the track control is calibrated using GNSS navigation data; By installing thrusters in both directions of satellite flight, the satellite has the ability to control its ascent and descent orbits in both forward and inverted directions within the orbital coordinate system.

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