Autonomous orbit keeping method and system for low-orbit inter-satellite laser communication constellation

Abstract

The invention provides an autonomous orbit keeping method and system for a low-orbit inter-satellite laser communication constellation, and the method comprises the steps: generating an upper limit and a lower limit of a satellite phase keeping threshold value, and updating the threshold value regularly according to the year-by-year change condition of the performance of a laser link; according to the self-positioning result of the satellite-borne GNSS navigation receiver, recurring the number of satellite flat orbits at the current moment, and calculating the flat latitude argument deviation relative to a nominal orbit; a limit cycle phase keeping method is adopted on the satellite, and an autonomous orbit control strategy is generated through autonomous calculation; after an orbit keeping strategy is generated, the orbit keeping strategy is downloaded to a ground management and control center, the ground management and control center rechecks an orbit keeping scheme which is autonomously generated on a satellite according to a satellite-ground integrated autonomous orbit control decision process, if the recheck is passed, the ground sends an orbit control implementation permission instruction, the satellite executes orbit control operation according to the strategy, and if the recheck is not passed, the ground sends an orbit control implementation permission instruction; and if so, injecting a track maintaining strategy from the ground. According to the invention, the complexity of ground system constellation management is greatly reduced.

Description

Technical Field

[0001] The present invention relates to the field of satellite system design, and in particular to a method and system for autonomous orbit keeping of a low-orbit inter-satellite laser communication constellation. Background Art

[0002] The orbit maintenance of the constellation is an important guarantee for maintaining the constellation configuration and system application efficiency. With the construction of large-scale low-orbit constellations in recent years, the pressure on constellation management of the ground system has increased dramatically, and there is an urgent need to improve the autonomous orbit maintenance capability of the constellation. In addition, in response to the needs of inter-satellite communications and multi-satellite collaborative applications with large amounts of data and high data rates, low-orbit constellations are often equipped with inter-satellite laser communication links to achieve efficient interconnection and interoperability of the constellation. However, the instantaneous beam angle of the inter-satellite laser communication terminal is small, and it is easy to cause a link break when the satellite's attitude changes dramatically, and it takes a long time to re-establish the link after the interruption. Therefore, for low-orbit constellations that use inter-satellite laser communication links, it is necessary to ensure that the inter-satellite link establishment is not interrupted during the autonomous orbit maintenance process. At present, research on constellation orbit maintenance has given little consideration to the need for permanent link establishment for inter-satellite laser communications.

[0003] There is a satellite orbit maintenance control method based on two-row elements (publication number CN102591343A). This method overcomes the error caused by using unstable atmospheric density as calculation input, and provides a method for orbit prediction and control of satellites with ground trajectory maintenance requirements based on public two-row element data, which improves the reliability and accuracy of prediction and control calculations. However, this method does not have the ability to be implemented autonomously on board. There is also a satellite autonomous orbit maintenance method based on gain online estimation (publication number CN107554820A). This method uses an orbit control gain that can be updated online to autonomously calculate the jet duration of orbit control in response to changes in satellite mass, modeling errors of the orbit model, and changes in jet pressure, thereby reducing the workload on the ground. There is also an on-board autonomous orbit maintenance control method (publication number CN106542119A). This method aims at the orbit maintenance problem of ultra-low orbit satellites and proposes a thruster switch maintenance strategy based on average orbit semi-major axis feedback to achieve autonomous maintenance of orbit altitude under different atmospheric environments. Both methods focus on on-board autonomous strategy generation and autonomous execution design, without considering the on-board autonomous decision-making process, and are still some distance away from practical application. Summary of the invention

[0004] In view of the defects in the prior art, the object of the present invention is to provide a method and system for autonomous orbit keeping of a low-orbit inter-satellite laser communication constellation.

[0005] According to the present invention, a method for autonomous orbit maintenance of a low-orbit inter-satellite laser communication constellation is provided, comprising:

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

[0007] Step S2: According to the autonomous positioning result of the onboard GNSS navigation receiver, the number of satellite horizontal orbit elements at the current moment is recursively calculated, and the horizontal latitude argument deviation relative to the nominal orbit is calculated;

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

[0009] Step S4: After the orbit keeping strategy is generated, it is transmitted to the ground control center. According to the satellite-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit keeping plan autonomously generated on the satellite. If it passes the review, the ground will issue an instruction to allow orbit control to be implemented, and the satellite will perform orbit control operations according to the strategy. If it fails, the ground will inject the orbit keeping strategy.

[0010] Preferably, in step S1:

[0011] The calculation process of the satellite phase hold threshold is as follows: the maximum phase difference between two link-building satellites is calculated according to the orbital height of the low-orbit constellation, the longest link-building distance constraint of the inter-satellite laser communication link, and the minimum link-building height constraint from the ground surface; the absolute phase hold mode is adopted to set the upper and lower limits of the satellite phase hold threshold, and the communication performance is analyzed according to the telemetry data of the on-orbit laser link, and the longest link-building distance of the inter-satellite laser communication link and the minimum link-building height constraint from the ground surface are updated, and the phase hold threshold is updated once every preset time period;

[0012] Generate satellite phase hold threshold Δu lim , LEO constellation orbit height H sat , the longest link establishment distance constraint of the constellation inter-satellite laser communication link is L max , the lowest link height from the ground is H min ; The maximum phase difference between two satellites is u max 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 is the radius of the Earth, u res1 is the maximum inter-satellite phase difference constrained by the inter-satellite laser communication distance, u res2 is the maximum inter-satellite phase difference constrained by the minimum link establishment altitude of inter-satellite communication;

[0017] The maximum nominal phase difference between two satellites in the link is u nom , using the absolute phase hold mode, the satellite phase hold threshold is set to the upper limit set to Δu lim =(u max -u nom ) / 2, the lower limit is set to -Δu lim ;

[0018] According to the telemetry data of the on-orbit laser link, the communication performance is analyzed and the longest link distance L max The allowable height value H for shortening and chain building min The phase hold threshold Δu is updated every preset time period. lim , ensuring that satellites within the constellation are uninterrupted.

[0019] Preferably, in step S2:

[0020] Mean latitude argument deviation The calculation process is as follows: according to the real-time position and speed of the satellite at the time of solution by the GNSS navigation receiver, based on the Lagrange planetary equation, the number of horizontal orbit elements corresponding to the current time is obtained through numerical iteration; the horizontal latitude argument at the current time is calculated and compared with the nominal horizontal orbit element number to obtain the satellite horizontal latitude argument deviation relative to the nominal orbit at the current time

[0021] Given the real-time position r0 and velocity v0 of the satellite corresponding to time t0 solved by the GNSS navigation receiver, the osculating orbit element number at time t0 [ae iΩωM] can be obtained according to the Kepler equation. T , where a is the semi-major axis of the osculating orbit, e is the eccentricity of the osculating orbit, i is the inclination of the osculating orbit, Ω is the right ascension of the ascending node of the osculating orbit, ω is the argument of perigee of the osculating orbit, and M is the mean anomaly of the osculating orbit. The current time t is obtained by numerical iteration. CT The corresponding number of flat orbits in, is the semi-major axis of the flat orbit, is the eccentricity of the flat orbit, is the flat track inclination, is the right ascension of the ascending node of the mean orbit, is the argument of perigee of the horizontal orbit, is the mean anomaly angle of the mean orbit, and the iteration formula is as follows:

[0022]

[0023] In the formula, and is the first-order variation of the orbital elements with time caused by the perturbation, which is obtained according to the Lagrangian planetary equation. The initial value of the iterative calculation of the mean orbital elements is [ae iΩωM+n(t CT -t0)] T , n is the angular velocity of the satellite orbit;

[0024] Note the current time t CT The parallel latitude angle is Compare with the nominal track average root number to find the current time t CT Satellite mean latitude argument deviation relative to nominal orbit

[0025] Preferably, in step S3:

[0026] The limit cycle phase holding method is used on the satellite to bias the semi-major axis through orbit holding control to form a drift control loop. The current time, orbit control implementation time and nominal orbit elements are known. 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 used to generate an orbit keeping strategy autonomously on board according to the calculation results. The trigger of autonomous orbit keeping on board is set to the prohibited state. The on-board is in the prohibited execution state by default and will be triggered after the ground control center issues a permission command.

[0027] Autonomous orbit keeping strategy generated on board: When the satellite's latitude angle deviates When the satellite uses the limit cycle phase holding method, it autonomously calculates and generates autonomous orbit control strategies, t f Apply a velocity increment δV at all times, updating it every preset time;

[0028] The limit cycle phase holding method is used on the satellite. The semi-major axis is biased through orbit holding control to form a drift control loop. The semi-major axis is first positively biased, and the phase begins to shift to the left. As the semi-major axis decays, the semi-major axis gradually becomes smaller than the nominal value, and the phase begins to shift to the right. When the satellite reaches the right boundary, orbit holding control δa is implemented to maintain the phase deviation at -Δu lim With Δu lim between;

[0029] make

[0030] k1=n CT -n N

[0031]

[0032] Where μ is the Earth's gravitational constant, is the current time t CT Orbital angular velocity, a CT t CT The semi-major axis of the mean orbit at the moment; is the nominal orbital angular velocity, a N is the semi-major axis of the nominal orbit; k1 is the angular velocity difference between the current orbit and the nominal orbit; k2 is the orbital angular velocity variation coefficient considering the influence of atmospheric drag, is the first-order rate of change of the semi-major axis caused by atmospheric drag.

[0033] The phase holding period is

[0034]

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

[0036]

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

[0038]

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

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

[0041] δa=Δa T -Δa f

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

[0043]

[0044] The trigger setting for autonomous orbit maintenance on the satellite is in the prohibited state. By default, it is in the prohibited execution state on the satellite and waits for the ground to issue a permission command for triggering.

[0045] Preferably, in step S4:

[0046] Satellite-ground integrated autonomous orbit control decision-making and execution: After the satellite autonomously generates the orbit keeping strategy, it is transmitted to the ground control center at a preset time in advance. According to the satellite-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit keeping plan autonomously generated on the satellite. If it passes the review, the ground will issue an instruction to allow orbit control to be implemented, and the satellite will execute the orbit control operation according to the strategy. If it fails, the ground will fill in the orbit keeping strategy;

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

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

[0049] Step S4.2: The autonomous orbit keeping 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;

[0050] Step S4.3: The autonomous orbit control strategy is telemetered down to the ground, and the ground control center reviews it. After the ground allows it to be implemented, the next step is executed. Otherwise, the current autonomous orbit control is canceled, and the ground will inject the orbit keeping strategy;

[0051] Step S4.4: According to the autonomous orbit keeping strategy, the attitude and orbit control subsystem sends a state establishment request to the propulsion subsystem at a preset time before the ignition of the orbit control thruster 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 orbit control is completed, the accuracy of the orbit control is calibrated using GNSS navigation data;

[0055] Thrusters are installed in the forward and reverse directions of the satellite's flight, and the satellite has the ability to control the ascending and descending orbits in forward and reverse flight in the orbital coordinate system.

[0056] According to the present invention, a low-orbit intersatellite laser communication constellation autonomous orbit keeping system is provided, comprising:

[0057] Module M1: Generates the upper and lower limits of the satellite phase hold threshold and regularly updates the threshold according to the annual changes in laser link performance;

[0058] Module M2: According to the autonomous positioning results of the onboard GNSS navigation receiver, the number of satellite horizontal orbit elements at the current moment is recursively calculated, and the horizontal latitude argument deviation relative to the nominal orbit is calculated;

[0059] Module M3: When the satellite's horizontal latitude argument deviation and phase holding threshold meet the preset conditions, the satellite uses the limit cycle phase holding method to autonomously calculate and generate an autonomous orbit control strategy;

[0060] Module M4: After the orbit keeping strategy is generated, it is transmitted to the ground control center. According to the satellite-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit keeping plan autonomously generated on the satellite. If it passes the review, the ground will issue an instruction to allow orbit control to be implemented, and the satellite will perform orbit control operations according to the strategy. If it fails, the ground will inject the orbit keeping strategy.

[0061] Preferably, in the module M1:

[0062] The calculation process of the satellite phase hold threshold is as follows: the maximum phase difference between two link-building satellites is calculated according to the orbital height of the low-orbit constellation, the longest link-building distance constraint of the inter-satellite laser communication link, and the minimum link-building height constraint from the ground surface; the absolute phase hold mode is adopted to set the upper and lower limits of the satellite phase hold threshold, and the communication performance is analyzed according to the telemetry data of the on-orbit laser link, and the longest link-building distance of the inter-satellite laser communication link and the minimum link-building height constraint from the ground surface are updated, and the phase hold threshold is updated once every preset time period;

[0063] Generate satellite phase hold threshold Δu lim , LEO constellation orbit height H sat , the longest link establishment distance constraint of the constellation inter-satellite laser communication link is L max , the lowest link height from the ground is H min ; The maximum phase difference between two satellites is u max 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 is the radius of the Earth, u res1 is the maximum inter-satellite phase difference constrained by the inter-satellite laser communication distance, ures2 is the maximum inter-satellite phase difference constrained by the minimum link establishment altitude of inter-satellite communication;

[0068] The maximum nominal phase difference between two satellites in the link is u nom , using the absolute phase hold mode, the satellite phase hold threshold is set to the upper limit set to Δu lim =(u max -u nom ) / 2, the lower limit is set to -Δu lim ;

[0069] According to the telemetry data of the on-orbit laser link, the communication performance is analyzed and the longest link distance L max The allowable height value H for shortening and chain building min The phase hold threshold Δu is updated every preset time period. lim , ensuring that satellites within the constellation are uninterrupted.

[0070] Preferably, in the module M2:

[0071] Mean latitude argument deviation The calculation process is as follows: according to the real-time position and speed of the satellite at the time of solution by the GNSS navigation receiver, based on the Lagrange planetary equation, the number of horizontal orbit elements corresponding to the current time is obtained through numerical iteration; the horizontal latitude argument at the current time is calculated and compared with the nominal horizontal orbit element number to obtain the satellite horizontal latitude argument deviation relative to the nominal orbit at the current time

[0072] Given the real-time position r0 and velocity v0 of the satellite corresponding to time t0 solved by the GNSS navigation receiver, the osculating orbit element number at time t0 [ae iΩωM] can be obtained according to the Kepler equation. T , where a is the semi-major axis of the osculating orbit, e is the eccentricity of the osculating orbit, i is the inclination of the osculating orbit, Ω is the right ascension of the ascending node of the osculating orbit, ω is the argument of perigee of the osculating orbit, and M is the mean anomaly of the osculating orbit. The current time t is obtained by numerical iteration. CT The corresponding number of flat orbits in, is the semi-major axis of the flat orbit, is the eccentricity of the flat orbit, is the flat track inclination, is the right ascension of the ascending node of the mean orbit, is the argument of perigee of the horizontal orbit, is the mean anomaly angle of the mean orbit, and the iteration formula is as follows:

[0073]

[0074] In the formula, and is the first-order variation of the orbital elements with time caused by the perturbation, which is obtained according to the Lagrangian planetary equation. The initial value of the iterative calculation of the mean orbital elements is [ae iΩωM+n(t CT -t0)] T , n is the angular velocity of the satellite orbit;

[0075] Note the current time t CT The parallel latitude angle is Compare with the nominal track average root number to find the current time t CT Satellite mean latitude argument deviation relative to nominal orbit

[0076] Preferably, in the module M3:

[0077] The limit cycle phase holding method is used on the satellite to bias the semi-major axis through orbit holding control to form a drift control loop. The current time, orbit control implementation time and nominal orbit elements are known. 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 used to generate an orbit keeping strategy autonomously on board according to the calculation results. The trigger of autonomous orbit keeping on board is set to the prohibited state. The on-board is in the prohibited execution state by default and will be triggered after the ground control center issues a permission command.

[0078] Autonomous orbit keeping strategy generated on board: When the satellite's latitude angle deviates When the satellite uses the limit cycle phase holding method, it autonomously calculates and generates autonomous orbit control strategies, t f Apply a velocity increment δV at all times, updating it every preset time;

[0079] The limit cycle phase holding method is used on the satellite. The semi-major axis is biased through orbit holding control to form a drift control loop. The semi-major axis is first positively biased, and the phase begins to shift to the left. As the semi-major axis decays, the semi-major axis gradually becomes smaller than the nominal value, and the phase begins to shift to the right. When the satellite reaches the right boundary, orbit holding control δa is implemented to maintain the phase deviation at -Δu lim With Δu lim between;

[0080] make

[0081] k1=n CT -n N

[0082]

[0083] Where μ is the Earth's gravitational constant, is the current time t CT Orbital angular velocity, a CT t CTThe semi-major axis of the mean orbit at the moment; is the nominal orbital angular velocity, a N is the semi-major axis of the nominal orbit; k1 is the angular velocity difference between the current orbit and the nominal orbit; k2 is the orbital angular velocity variation coefficient considering the influence of atmospheric drag, is the first-order rate of change of the semi-major axis caused by atmospheric drag.

[0084] The phase holding period is

[0085]

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

[0087]

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

[0089]

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

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

[0092] δa=Δa T -Δa f

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

[0094]

[0095] The trigger setting for autonomous orbit maintenance on the satellite is in the prohibited state. By default, it is in the prohibited execution state on the satellite and waits for the ground to issue a permission command for triggering.

[0096] Preferably, in the module M4:

[0097] Satellite-ground integrated autonomous orbit control decision-making and execution: After the satellite autonomously generates the orbit keeping strategy, it is transmitted to the ground control center at a preset time in advance. According to the satellite-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit keeping plan autonomously generated on the satellite. If it passes the review, the ground will issue an instruction to allow orbit control to be implemented, and the satellite will execute the orbit control operation according to the strategy. If it fails, the ground will fill in the orbit keeping strategy;

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

[0099] Module M4.1: The ground control center issues instructions to allow the satellite to autonomously maintain its orbit. After the phase deviation conditions are met, the satellite autonomously generates an orbit-keeping strategy.

[0100] Module M4.2: The autonomous orbit keeping 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 transmitted to the ground through telemetry. The ground control center will review it. If the ground allows it to be implemented, the next step will be executed. Otherwise, the autonomous orbit control will be canceled and the ground will fill in the orbit keeping strategy.

[0102] Module M4.4: According to the autonomous orbit keeping strategy, the attitude and orbit control subsystem sends a state establishment request to the propulsion subsystem at a preset time before the ignition of the orbit control thruster to establish the propulsion state;

[0103] Module M4.5: The attitude and orbit control subsystem controls the orbit control thrusters to implement ignition;

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

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

[0106] Thrusters are installed in the forward and reverse directions of the satellite's flight, and the satellite has the ability to control the ascending and descending orbits in forward and reverse flight in the orbital coordinate system.

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

[0108] Under the premise of ensuring the permanent establishment of inter-satellite laser communication links, the present invention realizes the autonomous on-board realization of the entire chain including strategy generation, orbit control decision-making and mission execution, greatly reduces the complexity of constellation management of ground systems, and has on-orbit application value. BRIEF DESCRIPTION OF THE DRAWINGS

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

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

[0111] Figure 2 A schematic diagram of a limit cycle phase maintaining method provided by an embodiment of the present invention;

[0112] Figure 3A schematic diagram of the satellite-ground integrated autonomous orbit control decision and execution process provided by an embodiment of the present invention. DETAILED DESCRIPTION

[0113] The present invention is described in detail below in conjunction with specific embodiments. The following embodiments will help those skilled in the art to further understand the present invention, but are not intended to limit the present invention in any form. It should be noted that, for those of ordinary skill in the art, several changes and improvements can also be made without departing from the concept of the present invention. These all belong to the protection scope of the present invention.

[0114] Embodiment 1:

[0115] The present invention discloses a method for autonomous orbit keeping of a low-orbit inter-satellite laser communication constellation, the method comprising the following steps: Step A, generating a constellation phase keeping threshold and updating it regularly: generating a satellite phase keeping threshold upper limit Δu according to the requirement of constant keeping of the inter-satellite laser link of the low-orbit constellation; lim and lower limit -Δu lim , and regularly update the threshold value according to the annual changes in laser link performance. Step B, satellite horizontal orbit element deviation calculation based on GNSS navigation receiver: According to 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 angle deviation relative to the nominal orbit is calculated. Step C: Autonomously generate orbit-keeping strategy on board: When the satellite's horizontal latitude angle deviation When the satellite uses the limit cycle phase holding method, it autonomously calculates and generates an autonomous orbit control strategy (at t f The velocity increment δV is applied at all times and updated every hour. Step D, satellite-ground integrated autonomous orbit control decision and execution: After the satellite generates the orbit keeping strategy autonomously, it is transmitted to the ground control center 24 hours in advance. According to the satellite-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit keeping plan autonomously generated on the satellite. If it passes the review, the ground will issue an instruction to allow orbit control to be implemented, and the satellite will execute the orbit control operation according to the strategy. If it fails, the ground will inject the orbit keeping strategy.

[0116] According to the present invention, a method for autonomous orbit maintenance of a low-orbit intersatellite laser communication constellation is provided. Figure 1-Figure 3 As shown, including:

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

[0118] Specifically, in step S1:

[0119] The calculation process of the satellite phase hold threshold is as follows: the maximum phase difference between two link-building satellites is calculated according to the orbital height of the low-orbit constellation, the longest link-building distance constraint of the inter-satellite laser communication link, and the minimum link-building height constraint from the ground surface; the absolute phase hold mode is adopted to set the upper and lower limits of the satellite phase hold threshold, and the communication performance is analyzed according to the telemetry data of the on-orbit laser link, and the longest link-building distance of the inter-satellite laser communication link and the minimum link-building height constraint from the ground surface are updated, and the phase hold threshold is updated once every preset time period;

[0120] Generate satellite phase hold threshold Δu lim , LEO constellation orbit height H sat , the longest link establishment distance constraint of the constellation inter-satellite laser communication link is L max , the lowest link height from the ground is H min ; The maximum phase difference between two satellites is u max 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 is the radius of the Earth, u res1 is the maximum inter-satellite phase difference constrained by the inter-satellite laser communication distance, u res2 is the maximum inter-satellite phase difference constrained by the minimum link establishment altitude of inter-satellite communication;

[0125] The maximum nominal phase difference between two satellites in the link is u nom , using the absolute phase hold mode, the satellite phase hold threshold is set to the upper limit set to Δu lim =(u max -u nom ) / 2, the lower limit is set to -Δu lim ;

[0126] According to the telemetry data of the on-orbit laser link, the communication performance is analyzed and the longest link distance L max The allowable height value H for shortening and chain buildingmin The phase hold threshold Δu is updated every preset time period. lim , ensuring that satellites within the constellation are uninterrupted.

[0127] Step S2: According to the autonomous positioning result of the onboard GNSS navigation receiver, the number of satellite horizontal orbit elements at the current moment is recursively calculated, and the horizontal latitude argument deviation relative to the nominal orbit is calculated;

[0128] Specifically, in step S2:

[0129] Mean latitude argument deviation The calculation process is as follows: according to the real-time position and speed of the satellite at the time of solution by the GNSS navigation receiver, based on the Lagrange planetary equation, the number of horizontal orbit elements corresponding to the current time is obtained through numerical iteration; the horizontal latitude argument at the current time is calculated and compared with the nominal horizontal orbit element number to obtain the satellite horizontal latitude argument deviation relative to the nominal orbit at the current time

[0130] Given the real-time position r0 and velocity v0 of the satellite corresponding to time t0 solved by the GNSS navigation receiver, the osculating orbit element number at time t0 [ae iΩωM] can be obtained according to the Kepler equation. T , where a is the semi-major axis of the osculating orbit, e is the eccentricity of the osculating orbit, i is the inclination of the osculating orbit, Ω is the right ascension of the ascending node of the osculating orbit, ω is the argument of perigee of the osculating orbit, and M is the mean anomaly of the osculating orbit. The current time t is obtained by numerical iteration. CT The corresponding number of flat orbits in, is the semi-major axis of the flat orbit, is the eccentricity of the flat orbit, is the flat track inclination, is the right ascension of the ascending node of the mean orbit, is the argument of perigee of the horizontal orbit, is the mean anomaly angle of the mean orbit, and the iteration formula is as follows:

[0131]

[0132] In the formula, and is the first-order variation of the orbital elements with time caused by the perturbation, which is obtained according to the Lagrangian planetary equation. The initial value of the iterative calculation of the mean orbital elements is [ae iΩωM+n(t CT -t0)] T , n is the angular velocity of the satellite orbit;

[0133] Note the current time t CT The parallel latitude angle is Compare with the nominal track average root number to find the current time t CTSatellite mean latitude argument deviation relative to nominal orbit

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

[0135] Specifically, in step S3:

[0136] The limit cycle phase holding method is used on the satellite to bias the semi-major axis through orbit holding control to form a drift control loop. The current time, orbit control implementation time and nominal orbit elements are known. 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 used to generate an orbit keeping strategy autonomously on board according to the calculation results. The trigger of autonomous orbit keeping on board is set to the prohibited state. The on-board is in the prohibited execution state by default and will be triggered after the ground control center issues a permission command.

[0137] Autonomous orbit keeping strategy generated on board: When the satellite's latitude angle deviates When the satellite uses the limit cycle phase holding method, it autonomously calculates and generates autonomous orbit control strategies, t f Apply a velocity increment δV at all times, updating it every preset time;

[0138] The limit cycle phase holding method is used on the satellite. The semi-major axis is biased through orbit holding control to form a drift control loop. The semi-major axis is first positively biased, and the phase begins to shift to the left. As the semi-major axis decays, the semi-major axis gradually becomes smaller than the nominal value, and the phase begins to shift to the right. When the satellite reaches the right boundary, orbit holding control δa is implemented to maintain the phase deviation at -Δu lim With Δu lim between;

[0139] make

[0140] k1=n CT -n N

[0141]

[0142] Where μ is the Earth's gravitational constant, is the current time t CT Orbital angular velocity, a CT t CT The semi-major axis of the mean orbit at the moment; is the nominal orbital angular velocity, a N is the semi-major axis of the nominal orbit; k1 is the angular velocity difference between the current orbit and the nominal orbit; k2 is the orbital angular velocity variation coefficient considering the influence of atmospheric drag, is the first-order rate of change of the semi-major axis caused by atmospheric drag.

[0143] The phase holding period is

[0144]

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

[0146]

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

[0148]

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

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

[0151] δa=Δa T -Δa f

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

[0153]

[0154] The trigger setting for autonomous orbit maintenance on the satellite is in the prohibited state. By default, it is in the prohibited execution state on the satellite and waits for the ground to issue a permission command for triggering.

[0155] Step S4: After the orbit keeping strategy is generated, it is transmitted to the ground control center. According to the satellite-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit keeping plan autonomously generated on the satellite. If it passes the review, the ground will issue an instruction to allow orbit control to be implemented, and the satellite will perform orbit control operations according to the strategy. If it fails, the ground will inject the orbit keeping strategy.

[0156] Specifically, in step S4:

[0157] Satellite-ground integrated autonomous orbit control decision-making and execution: After the satellite autonomously generates the orbit keeping strategy, it is transmitted to the ground control center at a preset time in advance. According to the satellite-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit keeping plan autonomously generated on the satellite. If it passes the review, the ground will issue an instruction to allow orbit control to be implemented, and the satellite will execute the orbit control operation according to the strategy. If it fails, the ground will fill in the orbit keeping strategy;

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

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

[0160] Step S4.2: The autonomous orbit keeping 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;

[0161] Step S4.3: The autonomous orbit control strategy is telemetered down to the ground, and the ground control center reviews it. After the ground allows it to be implemented, the next step is executed. Otherwise, the current autonomous orbit control is canceled, and the ground will inject the orbit keeping strategy;

[0162] Step S4.4: According to the autonomous orbit keeping strategy, the attitude and orbit control subsystem sends a state establishment request to the propulsion subsystem at a preset time before the ignition of the orbit control thruster 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 orbit control is completed, the accuracy of the orbit control is calibrated using GNSS navigation data;

[0166] Thrusters are installed in the forward and reverse directions of the satellite's flight, and the satellite has the ability to control the ascending and descending orbits in forward and reverse flight in the orbital coordinate system.

[0167] Embodiment 2:

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

[0169] The present invention also provides a low-orbit inter-satellite laser communication constellation autonomous orbit maintaining system, which can be implemented by executing the process steps of the low-orbit inter-satellite laser communication constellation autonomous orbit maintaining method, that is, those skilled in the art can understand the low-orbit inter-satellite laser communication constellation autonomous orbit maintaining method as a preferred implementation of the low-orbit inter-satellite laser communication constellation autonomous orbit maintaining system.

[0170] According to the present invention, a low-orbit intersatellite laser communication constellation autonomous orbit keeping system is provided, comprising:

[0171] Module M1: Generates the upper and lower limits of the satellite phase hold threshold and regularly updates the threshold according to the annual changes in laser link performance;

[0172] Specifically, in the module M1:

[0173] The calculation process of the satellite phase hold threshold is as follows: the maximum phase difference between two link-building satellites is calculated according to the orbital height of the low-orbit constellation, the longest link-building distance constraint of the inter-satellite laser communication link, and the minimum link-building height constraint from the ground surface; the absolute phase hold mode is adopted to set the upper and lower limits of the satellite phase hold threshold, and the communication performance is analyzed according to the telemetry data of the on-orbit laser link, and the longest link-building distance of the inter-satellite laser communication link and the minimum link-building height constraint from the ground surface are updated, and the phase hold threshold is updated once every preset time period;

[0174] Generate satellite phase hold threshold Δu lim , LEO constellation orbit height H sat , the longest link establishment distance constraint of the constellation inter-satellite laser communication link is L max , the lowest link height from the ground is H min ; The maximum phase difference between two satellites is u max 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 is the radius of the Earth, u res1 is the maximum inter-satellite phase difference constrained by the inter-satellite laser communication distance, u res2 is the maximum inter-satellite phase difference constrained by the minimum link establishment altitude of inter-satellite communication;

[0179] The maximum nominal phase difference between two satellites in the link is u nom , using the absolute phase hold mode, the satellite phase hold threshold is set to the upper limit set to Δu lim =(u max -u nom ) / 2, the lower limit is set to -Δu lim ;

[0180] According to the telemetry data of the on-orbit laser link, the communication performance is analyzed and the longest link distance L max The allowable height value H for shortening and chain building min The phase hold threshold Δu is updated every preset time period. lim , ensuring that satellites within the constellation are uninterrupted.

[0181] Module M2: According to the autonomous positioning results of the onboard GNSS navigation receiver, the number of satellite horizontal orbit elements at the current moment is recursively calculated, and the horizontal latitude argument deviation relative to the nominal orbit is calculated;

[0182] Specifically, in the module M2:

[0183] Mean latitude argument deviation The calculation process is as follows: according to the real-time position and speed of the satellite at the time of solution by the GNSS navigation receiver, based on the Lagrange planetary equation, the number of horizontal orbit elements corresponding to the current time is obtained through numerical iteration; the horizontal latitude argument at the current time is calculated and compared with the nominal horizontal orbit element number to obtain the satellite horizontal latitude argument deviation relative to the nominal orbit at the current time

[0184] Given the real-time position r0 and velocity v0 of the satellite corresponding to time t0 solved by the GNSS navigation receiver, the osculating orbit element number at time t0 [ae iΩωM] can be obtained according to the Kepler equation. T , where a is the semi-major axis of the osculating orbit, e is the eccentricity of the osculating orbit, i is the inclination of the osculating orbit, Ω is the right ascension of the ascending node of the osculating orbit, ω is the argument of perigee of the osculating orbit, and M is the mean anomaly of the osculating orbit. The current time t is obtained by numerical iteration. CT The corresponding number of flat orbits in, is the semi-major axis of the flat orbit, is the eccentricity of the flat orbit, is the flat track inclination, is the right ascension of the ascending node of the mean orbit, is the argument of perigee of the horizontal orbit, is the mean anomaly angle of the mean orbit, and the iteration formula is as follows:

[0185]

[0186] In the formula, and is the first-order variation of the orbital elements with time caused by the perturbation, which is obtained according to the Lagrangian planetary equation. The initial value of the iterative calculation of the mean orbital elements is [ae iΩωM+n(t CT -t0)] T , n is the angular velocity of the satellite orbit;

[0187] Note the current time t CT The parallel latitude angle is Compare with the nominal track average root number to find the current time t CT Satellite mean latitude argument deviation relative to nominal orbit

[0188] Module M3: When the satellite's horizontal latitude argument deviation and phase holding threshold meet the preset conditions, the satellite uses the limit cycle phase holding method to autonomously calculate and generate an autonomous orbit control strategy;

[0189] Specifically, in the module M3:

[0190] The limit cycle phase holding method is used on the satellite to bias the semi-major axis through orbit holding control to form a drift control loop. The current time, orbit control implementation time and nominal orbit elements are known. 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 used to generate an orbit keeping strategy autonomously on board according to the calculation results. The trigger of autonomous orbit keeping on board is set to the prohibited state. The on-board is in the prohibited execution state by default and will be triggered after the ground control center issues a permission command.

[0191] Autonomous orbit keeping strategy generated on board: When the satellite's latitude angle deviates When the satellite uses the limit cycle phase holding method, it autonomously calculates and generates autonomous orbit control strategies, t f Apply a velocity increment δV at all times, updating it every preset time;

[0192] The limit cycle phase holding method is used on the satellite. The semi-major axis is biased through orbit holding control to form a drift control loop. The semi-major axis is first positively biased, and the phase begins to shift to the left. As the semi-major axis decays, the semi-major axis gradually becomes smaller than the nominal value, and the phase begins to shift to the right. When the satellite reaches the right boundary, orbit holding control δa is implemented to maintain the phase deviation at -Δu lim With Δu lim between;

[0193] make

[0194] k1=n CT -n N

[0195]

[0196] Where μ is the Earth's gravitational constant, is the current time t CT Orbital angular velocity, a CT t CT The semi-major axis of the mean orbit at the moment; is the nominal orbital angular velocity, a Nis the semi-major axis of the nominal orbit; k1 is the angular velocity difference between the current orbit and the nominal orbit; k2 is the orbital angular velocity variation coefficient considering the influence of atmospheric drag, is the first-order rate of change of the semi-major axis caused by atmospheric drag.

[0197] The phase holding period is

[0198]

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

[0200]

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

[0202]

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

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

[0205] δa=Δa T -Δa f

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

[0207]

[0208] The trigger setting for autonomous orbit maintenance on the satellite is in the prohibited state. By default, it is in the prohibited execution state on the satellite and waits for the ground to issue a permission command for triggering.

[0209] Module M4: After the orbit keeping strategy is generated, it is transmitted to the ground control center. According to the satellite-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit keeping plan autonomously generated on the satellite. If it passes the review, the ground will issue an instruction to allow orbit control to be implemented, and the satellite will perform orbit control operations according to the strategy. If it fails, the ground will inject the orbit keeping strategy.

[0210] Specifically, in the module M4:

[0211] Satellite-ground integrated autonomous orbit control decision-making and execution: After the satellite autonomously generates the orbit keeping strategy, it is transmitted to the ground control center at a preset time in advance. According to the satellite-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit keeping plan autonomously generated on the satellite. If it passes the review, the ground will issue an instruction to allow orbit control to be implemented, and the satellite will execute the orbit control operation according to the strategy. If it fails, the ground will fill in the orbit keeping strategy;

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

[0213] Module M4.1: The ground control center issues instructions to allow the satellite to autonomously maintain its orbit. After the phase deviation conditions are met, the satellite autonomously generates an orbit-keeping strategy.

[0214] Module M4.2: The autonomous orbit keeping 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 transmitted to the ground through telemetry. The ground control center will review it. If the ground allows it to be implemented, the next step will be executed. Otherwise, the autonomous orbit control will be canceled and the ground will fill in the orbit keeping strategy.

[0216] Module M4.4: According to the autonomous orbit keeping strategy, the attitude and orbit control subsystem sends a state establishment request to the propulsion subsystem at a preset time before the ignition of the orbit control thruster to establish the propulsion state;

[0217] Module M4.5: The attitude and orbit control subsystem controls the orbit control thrusters to implement ignition;

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

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

[0220] Thrusters are installed in the forward and reverse directions of the satellite's flight, and the satellite has the ability to control the ascending and descending orbits in forward and reverse flight in the orbital coordinate system.

[0221] Embodiment 3:

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

[0223] The object of the present invention is achieved through the following technical solutions: A method for autonomous orbit maintenance of a low-orbit inter-satellite laser communication constellation, the method comprising the following steps: Step A, generating a constellation phase maintenance threshold and regularly updating: generating a satellite phase maintenance threshold Δu according to the requirement of constant maintenance of the inter-satellite laser link of the low-orbit constellation lim, and regularly update the threshold value according to the annual changes in laser link performance; Step B, calculate the satellite horizontal orbit element deviation based on the GNSS navigation receiver: According to the autonomous positioning results of the onboard GNSS navigation receiver, recursively deduce the satellite horizontal orbit element at the current moment, and calculate the horizontal latitude angle deviation relative to the nominal orbit Step C: Autonomously generate orbit-keeping strategy on board: When the satellite's horizontal latitude angle deviation When the satellite uses the limit cycle phase-keeping method, it autonomously calculates and generates an autonomous orbit control strategy (t f The velocity increment δV is applied at all times and updated every hour; Step D, satellite-ground integrated autonomous orbit control decision and execution: After the satellite autonomously generates the orbit keeping strategy, it is transmitted to the ground control center 24 hours in advance. According to the satellite-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit keeping plan autonomously generated on the satellite. If it passes the review, the ground will issue an instruction to allow orbit control to be implemented, and the satellite will execute the orbit control operation according to the strategy. If it fails, the ground will inject the orbit keeping strategy.

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

[0225] Step A, generate the constellation phase hold threshold and update it regularly: according to the requirement of constant hold of inter-satellite laser links of low-orbit constellations, generate the satellite phase hold threshold Δu lim , and regularly update the thresholds based on the year-to-year changes in laser link performance.

[0226] Low Earth Orbit Constellation Orbit Height H sat , the longest link establishment distance constraint of the constellation inter-satellite laser communication link is L max , the lowest link height from the ground is H min The maximum phase difference between the two satellites is u max Calculated by 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 is the radius of the Earth, u res1 is the maximum inter-satellite phase difference constrained by the inter-satellite laser communication distance, u res2 It is the maximum inter-satellite phase difference constrained by the minimum link establishment altitude of inter-satellite communication.

[0231] The maximum nominal phase difference between two satellites in the link is u nom , considering the absolute phase hold mode, the satellite phase hold threshold is set to the upper limit set to Δu lim =(u max -u nom ) / 2, the lower limit is set to -Δu lim .

[0232] As time goes by, the performance of the on-board laser link will decline year by year. According to the telemetry data of the on-orbit laser link, the longest link distance L max Will shorten the allowable height H of the chain min The phase holding threshold Δu needs to be updated once a year. lim to ensure that satellites in the constellation are uninterrupted.

[0233] Step B, calculate the satellite horizontal orbit element deviation based on the GNSS navigation receiver: According to 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 angle deviation relative to the nominal orbit is calculated.

[0234] Given the real-time position r0 and velocity v0 of the satellite corresponding to time t0 solved by the GNSS navigation receiver, the number of osculating orbit elements at time t0 [ae iΩωM] can be obtained according to the Kepler equation. T (a is the semi-major axis of the osculating orbit, e is the eccentricity of the osculating orbit, i is the inclination of the osculating orbit, Ω is the right ascension of the ascending node of the osculating orbit, ω is the argument of perigee of the osculating orbit, and M is the mean anomaly angle of the osculating orbit). Then, the current time t is obtained through numerical iteration. CT The corresponding number of flat orbits ( is the semi-major axis of the flat orbit, is the eccentricity of the flat orbit, is the flat orbit inclination, is the right ascension of the ascending node of the mean orbit, is the argument of perigee of the horizontal orbit, is the mean anomaly angle of the mean orbit), and the iterative formula is shown below.

[0235]

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

[0237] Remember the current t CT The argument of mean latitude at this moment is Compare with the nominal track average root number to find the current t CT Satellite mean latitude argument deviation relative to the nominal orbit at any given moment

[0238] Step C: Autonomously generate orbit-keeping strategy on board: When the satellite's horizontal latitude angle deviation When the satellite uses the limit cycle phase-keeping method, it autonomously calculates and generates an autonomous orbit control strategy (t f A velocity increment δV is applied at all times and updated every hour.

[0239] The limit cycle phase holding method is used on the satellite to bias the semi-major axis through orbit holding control to form a drift control loop, such as Figure 2 That is, the semi-major axis is first positively biased (point A), the phase starts to shift to the left (A moves to B), and as the semi-major axis decays, the semi-major axis gradually becomes smaller than the nominal value, and the phase starts to shift to the right (B moves to C). When the satellite reaches the right boundary (point C), orbit keeping control δa (C moves to A) is implemented, so that the phase deviation is maintained at -Δu lim With Δu lim between.

[0240] make

[0241] k1=n CT -n N

[0242]

[0243] Where μ is the Earth's gravitational constant, is the current time t CT Orbital angular velocity, a CT t CT The semi-major axis of the mean orbit at the moment; is the nominal orbital angular velocity, a N is the semi-major axis of the nominal orbit; k1 is the angular velocity difference between the current orbit and the nominal orbit; k2 is the orbital angular velocity variation coefficient considering the influence of atmospheric drag, is the first-order rate of change of the semi-major axis caused by atmospheric drag.

[0245] The phase holding period is

[0246]

[0247] The target value of the semi-major axis deviation for phase preservation is

[0248]

[0249] When the satellite reaches boundary C, the actual semi-major axis deviation is

[0250]

[0251] Where: Δa N =a CT -a N , is the deviation between the current actual semi-major axis and the nominal semi-major axis; t f It is the moment when the satellite reaches boundary C or the moment when orbit control is implemented.

[0252] When the satellite reaches the boundary C, the change in the semi-major axis of the satellite orbit maneuver is

[0253] δa=Δa T -Δa f

[0254] For a near-Earth circular orbit, the corresponding velocity increment requirement is

[0255]

[0256] The trigger setting for autonomous orbit maintenance on the satellite is in the prohibited state. By default, it is in the prohibited execution state on the satellite and waits for the ground to issue a permission command for triggering.

[0257] Step D, satellite-ground integrated autonomous orbit control decision and execution: After the satellite autonomously generates the orbit keeping strategy, it will be transmitted to the ground control center 24 hours in advance. According to the satellite-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit keeping plan autonomously generated on the satellite. If it passes the review, the ground will issue an instruction to allow orbit control to be implemented, and the satellite will execute the orbit control operation according to the strategy. If it fails, the ground will fill in the orbit keeping strategy.

[0258] like Figure 3 As shown in the figure, the satellite-ground integrated autonomous orbit control decision-making process is as follows:

[0259] A. The ground control center issues instructions to allow the satellite to autonomously maintain its orbit. After the phase deviation conditions are met, the satellite autonomously generates an orbit-keeping strategy;

[0260] B. The autonomous orbit keeping 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;

[0261] C. The autonomous orbit control strategy is transmitted to the ground through telemetry, and the ground control center reviews it. After the ground allows it to be implemented, the next step is executed. Otherwise, the autonomous orbit control is canceled and the ground will inject the orbit keeping strategy;

[0262] D. According to the autonomous orbit keeping strategy, 30 minutes before the ignition of the orbit control thruster, the attitude and orbit control subsystem sends a request for the propulsion subsystem to establish the propulsion state;

[0263] E. The attitude and orbit control subsystem controls the orbit control thrusters to ignite;

[0264] F. After the orbit control ignition is completed, the attitude and orbit control subsystem sends a shutdown request to the propulsion subsystem, and the integrated electronic subsystem shuts down the propulsion subsystem.

[0265] G. After the orbit control is completed, the accuracy of the orbit control is calibrated using the navigation data.

[0266] In order to avoid link interruption caused by attitude maneuvers, thrusters are installed in the forward and reverse directions of the satellite's flight. At the same time, the satellite has the ability to control the ascending and descending orbits in both forward and reverse flight. Therefore, no attitude maneuvers are required to switch from the usual ground-directed attitude to the orbit control mode, thereby preventing the interruption of the inter-satellite laser communication link caused by attitude changes.

[0267] Those skilled in the art know that, in addition to realizing the system and its various devices, modules, and units provided by the present invention in a purely computer-readable program code, it is entirely possible to realize the same functions in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers by logically programming the method steps. Therefore, the system and its various devices, modules, and units provided by the present invention can be considered as a hardware component, and the devices, modules, and units included therein for realizing various functions can also be regarded as structures within the hardware component; the devices, modules, and units for realizing various functions can also be regarded as both software modules for realizing the method and structures within the hardware component.

[0268] The above describes the specific embodiments of the present invention. It should be understood that the present invention is not limited to the above specific embodiments, and those skilled in the art can make various changes or modifications within the scope of the claims, which does not affect the essence of the present invention. In the absence of conflict, the embodiments of the present application and the features in the embodiments can be combined with each other arbitrarily.

Claims

1. A method for autonomous orbit maintenance of a low-orbit intersatellite laser communication constellation, characterized in that: include: Step S1: Generate the upper and lower limits of the satellite phase hold threshold, and regularly update the threshold according to the annual changes in laser link performance; Step S2: According to the autonomous positioning result of the onboard GNSS navigation receiver, the number of satellite horizontal orbit elements at the current moment is recursively calculated, and the horizontal latitude argument deviation relative to the nominal orbit is calculated; Step S3: When the satellite's horizontal latitude argument deviation and phase holding threshold meet the preset conditions, the satellite uses the limit cycle phase holding method to autonomously calculate and generate an autonomous orbit control strategy; Step S4: After the orbit keeping strategy is generated, it is transmitted to the ground control center. According to the satellite-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit keeping plan autonomously generated on the satellite. If it passes the review, the ground will issue an instruction to allow orbit control to be implemented, and the satellite will perform orbit control operations according to the strategy. If it fails, the ground will inject the orbit keeping strategy.

2. The autonomous orbit maintaining method for low-orbit intersatellite laser communication constellation according to claim 1 is characterized in that: In step S1: The calculation process of the satellite phase hold threshold is as follows: the maximum phase difference between two link-building satellites is calculated according to the orbital height of the low-orbit constellation, the longest link-building distance constraint of the inter-satellite laser communication link, and the minimum link-building height constraint from the ground surface; the absolute phase hold mode is adopted to set the upper and lower limits of the satellite phase hold threshold, and the communication performance is analyzed according to the telemetry data of the on-orbit laser link, and the longest link-building distance of the inter-satellite laser communication link and the minimum link-building height constraint from the ground surface are updated, and the phase hold threshold is updated once every preset time period; Generate satellite phase hold threshold Δu lim , LEO constellation orbit height H sat , the longest link establishment distance constraint of the constellation inter-satellite laser communication link is L max , the lowest link height from the ground is H min ; The maximum phase difference between two satellites is u max 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 is the radius of the Earth, u res1 is the maximum inter-satellite phase difference constrained by the inter-satellite laser communication distance, u res2 is the maximum inter-satellite phase difference constrained by the minimum link establishment altitude of inter-satellite communication; The maximum nominal phase difference between two satellites in the link is u nom , using the absolute phase hold mode, the satellite phase hold threshold is set to the upper limit set to Δu lim =(u max -u nom ) / 2, the lower limit is set to -Δu lim ; According to the telemetry data of the on-orbit laser link, the communication performance is analyzed and the longest link distance L max The allowable height value H for shortening and chain building min The phase hold threshold Δu is updated every preset time period. lim , ensuring that satellites within the constellation are uninterrupted.

3. The autonomous orbit maintaining method for low-orbit intersatellite laser communication constellation according to claim 1 is characterized in that: In step S2: Mean latitude argument deviation The calculation process is as follows: according to the real-time position and velocity of the satellite at the time of solution by the GNSS navigation receiver, based on the Lagrange planetary equation, the number of horizontal orbit elements corresponding to the current time is obtained through numerical iteration; Calculate the current moment's mean latitude argument and compare it with the nominal mean orbit elements to find the current moment's satellite mean latitude argument deviation relative to the nominal orbit Given the real-time position r0 and velocity v0 of the satellite corresponding to time t0 solved by the GNSS navigation receiver, the number of osculating orbit elements at time t0 [ae iΩωM] can be obtained according to the Kepler equation. T , where a is the semi-major axis of the osculating orbit, e is the eccentricity of the osculating orbit, i is the inclination of the osculating orbit, Ω is the right ascension of the ascending node of the osculating orbit, ω is the argument of perigee of the osculating orbit, and M is the mean anomaly of the osculating orbit. The current time t is obtained by numerical iteration. CT The corresponding number of flat orbits in, is the semi-major axis of the flat orbit, is the eccentricity of the flat orbit, is the flat track inclination, is the right ascension of the ascending node of the mean orbit, is the argument of perigee of the horizontal orbit, is the mean anomaly angle of the mean orbit, and the iteration formula is as follows: In the formula, and is the first-order variation of the orbital elements with time caused by the perturbation, which is obtained according to the Lagrangian planetary equation. The initial value of the iterative calculation of the mean orbital elements is [ae iΩωM+n(t CT -t0)] T , n is the angular velocity of the satellite orbit; Note the current time t CT The parallel latitude angle is Compare with the nominal track average root number to find the current time t CT Satellite mean latitude argument deviation relative to nominal orbit 4. The autonomous orbit maintaining method for low-orbit intersatellite laser communication constellation according to claim 1, characterized in that: In step S3: The limit cycle phase holding method is used on the satellite to bias the semi-major axis through orbit holding control to form a drift control loop. The current time, orbit control implementation time and nominal orbit elements are known. 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 used to generate an autonomous orbit keeping strategy on board according to the calculation result. The trigger of autonomous orbit keeping on board is set to the prohibited state. The on-board is in the prohibited execution state by default and will be triggered after the ground control center issues a permission command. Autonomous orbit keeping strategy generated on board: When the satellite's latitude angle deviates When the satellite uses the limit cycle phase holding method, it autonomously calculates and generates autonomous orbit control strategies, t f Apply a velocity increment δV at all times, updating it every preset time; The limit cycle phase holding method is used on the satellite. The semi-major axis is biased through orbit holding control to form a drift control loop. The semi-major axis is first positively biased, and the phase begins to shift to the left. As the semi-major axis decays, the semi-major axis gradually becomes smaller than the nominal value, and the phase begins to shift to the right. When the satellite reaches the right boundary, orbit holding control δa is implemented to maintain the phase deviation at -Δu lim With Δu lim between; make k1=n CT -n N Where μ is the Earth's gravitational constant, is the current time t CT Orbital angular velocity, a CT t CT The semi-major axis of the mean orbit at the moment; is the nominal orbital angular velocity, a N is the semi-major axis of the nominal orbit; k1 is the angular velocity difference between the current orbit and the nominal orbit; k2 is the orbital angular velocity variation coefficient considering the influence of atmospheric drag, is the first-order rate of change of the semi-major axis caused by atmospheric drag. The phase holding period is The target value of the semi-major axis deviation for phase preservation is When the satellite reaches boundary C, the actual semi-major axis deviation is Where: Δa N =a CT -a N , is the deviation between the current actual semi-major axis and the nominal semi-major axis; t f It is the moment when the satellite reaches the boundary C or the moment when orbit control is implemented; When the satellite reaches the boundary C, the change in the semi-major axis of the satellite orbit maneuver is: δa=Δa T -Δa f For near-Earth circular orbit, the corresponding velocity increment requirement is: The trigger setting for autonomous orbit maintenance on the satellite is in the prohibited state. By default, it is in the prohibited execution state on the satellite and waits for the ground to issue a permission command for triggering.

5. The autonomous orbit maintaining method for low-orbit intersatellite laser communication constellation according to claim 1, characterized in that: In step S4: Satellite-ground integrated autonomous orbit control decision-making and execution: After the satellite autonomously generates the orbit keeping strategy, it is transmitted to the ground control center at a preset time in advance. According to the satellite-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit keeping plan autonomously generated on the satellite. If it passes the review, the ground will issue an instruction to allow orbit control to be implemented, and the satellite will execute the orbit control operation according to the strategy. If it fails, the ground will fill in the orbit keeping strategy; The satellite-ground integrated autonomous orbit control decision-making process is specifically as follows: Step S4.1: The ground control center issues an instruction to allow the satellite to autonomously maintain its orbit. After the phase deviation condition is met, the satellite autonomously generates an orbit-keeping strategy. Step S4.2: The autonomous orbit keeping 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; Step S4.3: The autonomous orbit control strategy is telemetered down to the ground, and the ground control center reviews it. After the ground allows it to be implemented, the next step is executed. Otherwise, the current autonomous orbit control is canceled, and the ground will inject the orbit keeping strategy; Step S4.4: According to the autonomous orbit keeping strategy, the attitude and orbit control subsystem sends a state establishment request to the propulsion subsystem at a preset time before the ignition of the orbit control thruster 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 orbit control is completed, the accuracy of the orbit control is calibrated using GNSS navigation data; Thrusters are installed in the forward and reverse directions of the satellite's flight, and the satellite has the ability to control the ascending and descending orbits in forward and reverse flight in the orbital coordinate system.

6. An autonomous orbit keeping system for low-orbit intersatellite laser communication constellation, characterized in that: include: Module M1: Generates the upper and lower limits of the satellite phase hold threshold and regularly updates the threshold according to the annual changes in laser link performance; Module M2: According to the autonomous positioning results of the onboard GNSS navigation receiver, the number of satellite horizontal orbit elements at the current moment is recursively calculated, and the horizontal latitude argument deviation relative to the nominal orbit is calculated; Module M3: When the satellite's horizontal latitude argument deviation and phase holding threshold meet the preset conditions, the satellite uses the limit cycle phase holding method to autonomously calculate and generate an autonomous orbit control strategy; Module M4: After the orbit keeping strategy is generated, it is transmitted to the ground control center. According to the satellite-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit keeping plan autonomously generated on the satellite. If it passes the review, the ground will issue an instruction to allow orbit control to be implemented, and the satellite will perform orbit control operations according to the strategy. If it fails, the ground will inject the orbit keeping strategy.

7. The autonomous orbit keeping system for low-orbit intersatellite laser communication constellation according to claim 6 is characterized in that: In the module M1: The calculation process of the satellite phase hold threshold is as follows: the maximum phase difference between two link-building satellites is calculated according to the orbital height of the low-orbit constellation, the longest link-building distance constraint of the inter-satellite laser communication link, and the minimum link-building height constraint from the ground surface; the absolute phase hold mode is adopted to set the upper and lower limits of the satellite phase hold threshold, and the communication performance is analyzed according to the telemetry data of the on-orbit laser link, and the longest link-building distance of the inter-satellite laser communication link and the minimum link-building height constraint from the ground surface are updated, and the phase hold threshold is updated once every preset time period; Generate satellite phase hold threshold Δu lim , LEO constellation orbit height H sat , the longest link establishment distance constraint of the constellation inter-satellite laser communication link is L max , the lowest link height from the ground is H min ; The maximum phase difference between two satellites is u max 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 is the radius of the Earth, u res1 is the maximum inter-satellite phase difference constrained by the inter-satellite laser communication distance, u res2 is the maximum inter-satellite phase difference constrained by the minimum link establishment altitude of inter-satellite communication; The maximum nominal phase difference between two satellites in the link is u nom , using the absolute phase hold mode, the satellite phase hold threshold is set to the upper limit set to Δu lim =(u max -u nom ) / 2, the lower limit is set to -Δu lim ; According to the telemetry data of the on-orbit laser link, the communication performance is analyzed and the longest link distance L max The allowable height value H for shortening and chain building min The phase hold threshold Δu is updated every preset time period. lim , ensuring that satellites within the constellation are uninterrupted.

8. The autonomous orbit keeping system for low-orbit intersatellite laser communication constellation according to claim 6 is characterized in that: In the module M2: Mean latitude argument deviation The calculation process is as follows: according to the real-time position and velocity of the satellite at the time of solution by the GNSS navigation receiver, based on the Lagrange planetary equation, the number of horizontal orbit elements corresponding to the current time is obtained through numerical iteration; Calculate the current moment's mean latitude argument and compare it with the nominal mean orbit elements to find the current moment's satellite mean latitude argument deviation relative to the nominal orbit Given the real-time position r0 and velocity v0 of the satellite corresponding to time t0 solved by the GNSS navigation receiver, the number of osculating orbit elements at time t0 [ae iΩωM] can be obtained according to the Kepler equation. T , where a is the semi-major axis of the osculating orbit, e is the eccentricity of the osculating orbit, i is the inclination of the osculating orbit, Ω is the right ascension of the ascending node of the osculating orbit, ω is the argument of perigee of the osculating orbit, and M is the mean anomaly of the osculating orbit. The current time t is obtained by numerical iteration. CT The corresponding number of flat orbits in, is the semi-major axis of the flat orbit, is the eccentricity of the flat orbit, is the flat orbit inclination, is the right ascension of the ascending node of the mean orbit, is the argument of perigee of the horizontal orbit, is the mean anomaly angle of the mean orbit, and the iteration formula is as follows: In the formula, and is the first-order variation of the orbital elements with time caused by the perturbation, which is obtained according to the Lagrangian planetary equation. The initial value of the iterative calculation of the mean orbital elements is [ae iΩωM+n(t CT -t0)] T , n is the angular velocity of the satellite orbit; Note the current time t CT The parallel latitude angle is Compare with the nominal track average root number to find the current time t CT Satellite mean latitude argument deviation relative to nominal orbit 9. The autonomous orbit keeping system for low-orbit intersatellite laser communication constellation according to claim 6 is characterized in that: In the module M3: The limit cycle phase holding method is used on the satellite to bias the semi-major axis through orbit holding control to form a drift control loop. The current time, orbit control implementation time and nominal orbit elements are known. 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 used to generate an autonomous orbit keeping strategy on board according to the calculation result. The trigger of autonomous orbit keeping on board is set to the prohibited state. The on-board is in the prohibited execution state by default and will be triggered after the ground control center issues a permission command. Autonomous orbit keeping strategy generated on board: When the satellite's latitude angle deviates When the satellite uses the limit cycle phase holding method, it autonomously calculates and generates autonomous orbit control strategies, t f Apply a velocity increment δV at all times, updating it every preset time; The limit cycle phase holding method is used on the satellite. The semi-major axis is biased through orbit holding control to form a drift control loop. The semi-major axis is first positively biased, and the phase begins to shift to the left. As the semi-major axis decays, the semi-major axis gradually becomes smaller than the nominal value, and the phase begins to shift to the right. When the satellite reaches the right boundary, orbit holding control δa is implemented to maintain the phase deviation at -Δu lim With Δu lim between; make k1=n CT -n N Where μ is the Earth's gravitational constant, is the current time t CT Orbital angular velocity, a CT t CT The semi-major axis of the mean orbit at the moment; is the nominal orbital angular velocity, a N is the semi-major axis of the nominal orbit; k1 is the angular velocity difference between the current orbit and the nominal orbit; k2 is the orbital angular velocity variation coefficient considering the influence of atmospheric drag, is the first-order rate of change of the semi-major axis caused by atmospheric drag. The phase holding period is The target value of the semi-major axis deviation for phase preservation is When the satellite reaches boundary C, the actual semi-major axis deviation is Where: Δa N =a CT -a N , is the deviation between the current actual semi-major axis and the nominal semi-major axis; t f It is the moment when the satellite reaches the boundary C or the moment when orbit control is implemented; When the satellite reaches the boundary C, the change in the semi-major axis of the satellite orbit maneuver is: δa=Δa T -Δa f For near-Earth circular orbit, the corresponding velocity increment requirement is: The trigger setting for autonomous orbit maintenance on the satellite is in the prohibited state. By default, it is in the prohibited execution state on the satellite and waits for the ground to issue a permission command for triggering.

10. The autonomous orbit keeping system for low-orbit intersatellite laser communication constellation according to claim 6, characterized in that: In the module M4: Satellite-ground integrated autonomous orbit control decision-making and execution: After the satellite autonomously generates the orbit keeping strategy, it is transmitted to the ground control center at a preset time in advance. According to the satellite-ground integrated autonomous orbit control decision-making process, the ground control center reviews the orbit keeping plan autonomously generated on the satellite. If it passes the review, the ground will issue an instruction to allow orbit control to be implemented, and the satellite will execute the orbit control operation according to the strategy. If it fails, the ground will fill in the orbit keeping strategy; The satellite-ground integrated autonomous orbit control decision-making process is specifically as follows: Module M4.1: The ground control center issues instructions to allow the satellite to autonomously maintain its orbit. After the phase deviation conditions are met, the satellite autonomously generates an orbit-keeping strategy. Module M4.2: The autonomous orbit keeping 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 transmitted to the ground through telemetry. The ground control center will review it. If the ground allows it to be implemented, the next step will be executed. Otherwise, the autonomous orbit control will be canceled and the ground will fill in the orbit keeping strategy. Module M4.4: According to the autonomous orbit keeping strategy, the attitude and orbit control subsystem sends a state establishment request to the propulsion subsystem at a preset time before the ignition of the orbit control thruster to establish the propulsion state; Module M4.5: The attitude and orbit control subsystem controls the orbit control thrusters to implement ignition; Module M4.6: After the orbit control ignition is completed, the attitude and orbit control subsystem sends a shutdown request to the propulsion subsystem, and the integrated electronic subsystem shuts down the propulsion subsystem; Module M4.7: After the orbit control is completed, the accuracy of the orbit control is calibrated through GNSS navigation data; Thrusters are installed in the forward and reverse directions of the satellite's flight, and the satellite has the ability to control the ascending and descending orbits in forward and reverse flight in the orbital coordinate system.

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