Autonomous control method for sun synchronous regression orbit

Abstract

The invention discloses an autonomous control method for a sun synchronous regression orbit, and belongs to the technical field of spacecraft orbit design and autonomous control, and the method comprises the steps: executing an in-plane and out-plane instruction calculation step based on a current instantaneous orbit element vector when determining that a satellite passes through a descending node based on a real-time position vector and a real-time adjacent moment position vector; obtaining in-plane maneuvering final execution time and an out-plane maneuvering instruction; and executing in-plane maneuver based on the in-plane maneuver final execution time, and executing out-plane maneuver when the satellite passes through the ascending node based on the out-plane maneuver instruction. The invention provides an autonomous control method for a sun-synchronous regression orbit, and solves the technical problem that an existing high-precision autonomous orbit control method for the sun-synchronous strict regression orbit is not suitable for the sun-synchronous strict regression orbit at a non-morning-night non-midnight intersection point. And the long-term stability and robustness of the control are enhanced while the high precision of the autonomous control of the sun-synchronous orbit is ensured.

Description

Technical Field

[0001] This invention relates to the field of spacecraft orbit design and autonomous control technology, and in particular to an autonomous control method for a sun-synchronous return orbit. Background Technology

[0002] Currently, for the problem of autonomous and precise control of strictly regressing sun-synchronous orbits, existing technologies mainly rely on high-precision GPS real-time navigation data and orbital analytical perturbation models. These technologies achieve strict regressing orbit control accuracy down to the hundred-meter level through autonomous orbit determination, pipeline error calculation, and orbital maneuver planning. The concept of satellite orbit pipeline control refers to the satellite strictly maintaining its actual trajectory within a circular spatial pipeline centered on a reference trajectory. If the actual trajectory deviates from the pipeline's boundaries, orbit control is required to achieve precise control of the spatial trajectory relative to the reference trajectory.

[0003] However, existing autonomous control technologies for strictly regressing sun-synchronous orbits are mainly geared towards twilight orbits and orbits with local time at noon / midnight nodes. The local time of the descending node (LTDN) of these orbits is fixed and close to 0:00, 6:00, 12:00, or 18:00. By applying velocity pulses when the inclination error exceeds a threshold, the extreme value of the error after control is made equal to the threshold (finding the extreme value by relying on the wave-like change of the error). At the same time, an in-plane control strategy targeting the semi-major axis and the longitude of the descending node is adopted. This leads to a fundamental limitation in the control of strictly regressing sun-synchronous orbits that do not have local time at twilight or midnight nodes: because the solar gravitational perturbation causes the inclination error to show a monotonous changing trend, it undermines the basis of existing out-of-plane control that relies on the wave-like changing trend; at the same time, the increased inclination error significantly changes the drift rate of the descending node longitude through J2 perturbation, causing the mean of the semi-major axis to continuously decay and the control law to shift, causing the existing in-plane strategy to fail due to neglecting coupling. There is an urgent need to develop an orbit control method that can simultaneously address the technical problems of long-term orbit inclination error caused by solar gravitational perturbation leading to control failure in local time sun-synchronous strict regression orbits that are not at dawn / dusk or midnight nodes, as well as the strong in-plane / out-of-plane coupling compensation, so as to achieve precise control of atypical local time orbits at the hundred-meter level. Summary of the Invention

[0004] This invention provides an autonomous control method for a sun-synchronous return orbit, which can solve the technical problem in the prior art where the long-term term of the orbital inclination error caused by solar gravitational perturbation leads to control failure in sun-synchronous strict return orbits that are not at local time (not at dawn / twilight / midnight). It also solves the technical problem that existing high-precision autonomous orbit control methods for sun-synchronous strict return orbits are not applicable to sun-synchronous strict return orbits that are not at local time (not at dawn / twilight / midnight). While ensuring high precision in autonomous control of sun-synchronous orbits, it enhances the long-term stability and robustness of the control.

[0005] This invention provides an autonomous control method for a sun-synchronous return orbit, comprising:

[0006] Real-time position vector and real-time velocity vector are obtained based on real-time navigation data, and the current instantaneous orbital element vector is obtained based on the real-time position vector and real-time velocity vector;

[0007] Obtain the real-time adjacent time position vectors. Based on the real-time position vectors and the real-time adjacent time position vectors, determine when the satellite passes through the descending node. Based on the current instantaneous orbit element vectors, perform in-plane and out-of-plane command calculation steps to obtain the final execution time of the in-plane maneuver and the out-of-plane maneuver commands.

[0008] The in-plane maneuver is executed based on the final execution time of the in-plane maneuver, and the out-of-plane maneuver is executed based on the out-of-plane maneuver command when the satellite passes through the ascending node, so as to achieve autonomous control of the sun-synchronous return orbit;

[0009] The in-plane and out-of-plane command calculation steps include:

[0010] Obtain historical position vectors and historical velocity vectors, and obtain average orbital element errors based on the real-time position vector, real-time velocity vector, historical position vector, and historical velocity vector;

[0011] Obtain the historical node periodic error sequence, and obtain the average semi-major axis decay rate based on the historical node periodic error sequence, the current instantaneous orbital element vector, and the preset atmospheric drag model;

[0012] Based on the average orbital element error, average semi-major axis decay rate, calibrated westward shift limit of descending node longitude, and preset eastward shift limit of descending node longitude, the in-plane maneuver time is predicted, the theoretical execution time of the in-plane maneuver is obtained, and the predicted number of orbits is obtained based on the theoretical execution time of the in-plane maneuver and the predicted number of orbits.

[0013] The orbital inclination error is obtained based on the average orbital element error, and the semi-monthly short-period term is separated based on the orbital inclination error to obtain the semi-monthly orbital inclination error.

[0014] When the inclination error of the semi-lunar orbit exceeds the preset inclination error threshold, the out-of-plane maneuver speed increment of the target is obtained, and the out-of-plane maneuver command is obtained based on the out-of-plane maneuver speed increment of the target.

[0015] This invention provides an autonomous control method for a sun-synchronous return orbit. First, it analyzes navigation data in real time to obtain real-time position and velocity vectors, and uses these to calculate the current instantaneous orbital elements vector. Then, it combines the position vectors of adjacent real-time moments to determine whether the satellite has passed the descending node. The descending node is used as the anchor point for calculating in-plane and out-of-plane maneuver commands. When the satellite passes the descending node, it calculates the out-of-plane maneuver command and the final execution time of the in-plane maneuver based on the current instantaneous orbital elements vector. This allows the satellite to execute the out-of-plane maneuver when it passes the ascending node, ensuring that the satellite calculates and executes the maneuver at the correct latitude argument. The system employs in-plane maneuvers to control sun-synchronous return orbits at local times other than dawn / dusk and midnight, addressing the technical limitation of existing technologies that only support control of sun-synchronous return orbits with descending node local times of 0:00, 6:00, 12:00, or 18:00. During calculations at the descending node, the system first calculates the average orbital element error based on real-time and historical descending node data, quantifying the deviation between the actual and nominal orbits. This transforms the abstract orbital state into a quantifiable error index. Furthermore, the attenuation rate is precisely obtained through instantaneous orbital elements, average element error, and an atmospheric drag model. To ensure accurate calculation of the semi-major axis decay rate, the in-plane maneuver execution time is predicted based on the average orbital element error, average semi-major axis decay rate, and westward shift limit. This allows for advance planning of the in-plane maneuver time window, ensuring that the semi-major axis lift can accurately counteract the eastward shift of the intersection longitude, thus improving the accuracy of autonomous orbit control. Simultaneously, the semi-monthly short-period term is separated from the orbital inclination error to obtain the semi-monthly orbital inclination error. Finally, a threshold judgment mechanism is constructed based on the error threshold. When the orbital inclination error exceeds the preset inclination error threshold, the target out-of-plane maneuver velocity increment is calculated to achieve out-of-plane maneuver control. Precise maintenance of orbital inclination angle transforms the complex orbital inclination angle error prediction and control into threshold-triggered control based on the separation of semi-lunar periodic terms. This reduces computational complexity, ensures that the orbital inclination angle is precisely controlled within the threshold, and is simple, clear, and highly stable. It simplifies the original strategy and enhances robustness. It effectively improves maneuver efficiency and solves the technical problem of control failure caused by the long-term dominance of orbital inclination angle error due to solar gravitational perturbation in strictly regressive local time sun-synchronous orbits that are not at dawn / dusk or midnight nodes. While ensuring high precision in autonomous control of sun-synchronous orbits, it enhances the long-term stability and robustness of control.

[0016] Furthermore, the autonomous control of the sun-synchronous return orbit by performing in-plane maneuvers based on the final execution time of the in-plane maneuvers and by performing out-of-plane maneuvers based on the out-of-plane maneuver commands when the satellite passes the ascending node includes:

[0017] Real-time acquisition of the satellite's instantaneous navigation data; determination based on the instantaneous navigation data when the satellite passes through the ascending node; and determination of the existence of an out-of-plane maneuver command; and execution of an out-of-plane maneuver based on the out-of-plane maneuver command.

[0018] The system acquires real-time instantaneous navigation data from the satellite and obtains instantaneous position and velocity vectors based on the instantaneous navigation data.

[0019] The instantaneous orbital element vector is obtained based on the instantaneous position vector and instantaneous velocity vector;

[0020] Obtain the position vectors of adjacent instantaneous moments;

[0021] When the satellite passes through the ascending node based on the instantaneous position vector and the position vectors of the instantaneous adjacent moments, and when it is determined that there is an out-of-plane maneuver command, an out-of-plane maneuver is executed based on the out-of-plane maneuver command.

[0022] Based on the instantaneous orbital element vector, perform in-plane maneuvering control actions:

[0023] The instantaneous orbital element vector is obtained based on the instantaneous navigation data, and the instantaneous latitude argument is obtained based on the instantaneous orbital element vector;

[0024] Get the current orbital revolution count;

[0025] Based on the current orbital revolutions, instantaneous latitude argument, final execution time of in-plane maneuver, and preset in-plane maneuver constraints, when the preset in-plane maneuver conditions are met, adjacent descending intersection data are obtained;

[0026] The in-plane maneuver velocity increment is obtained based on the adjacent descending intersection point data and the instantaneous orbital element vector;

[0027] In-plane maneuvers are performed based on the in-plane maneuver speed increment.

[0028] Furthermore, prior to the step of performing in-plane maneuvering control actions based on the instantaneous orbital element vector, the method further includes:

[0029] When the satellite does not pass through the ascending node, it is determined based on the instantaneous position vector and the position vector of the instantaneous adjacent time. In this case, the in-plane maneuvering control action is performed directly based on the instantaneous orbital element vector.

[0030] Furthermore, prior to the step of performing in-plane maneuvering control actions based on the instantaneous orbital element vector, the method further includes:

[0031] If the current orbital revolutions, instantaneous latitude argument, final execution time of in-plane maneuver, and preset in-plane maneuver constraints determine that the preset in-plane maneuver conditions are not met, then no in-plane maneuver will be performed.

[0032] In the above scheme, during satellite operation, it is determined whether the satellite has passed through the ascending node by analyzing real-time acquired instantaneous navigation data. If the satellite passes through the ascending node and there are historical out-of-plane maneuver commands, an out-of-plane maneuver is executed, i.e., the satellite's orbit is controlled based on the historical out-of-plane maneuver commands calculated from the previous passing of the descending node. If the satellite passes through the ascending node but there are no historical out-of-plane maneuver commands, or if the satellite has not passed through the ascending node, an in-plane maneuver is directly determined. That is, regardless of whether the satellite has passed through the ascending node or whether an out-of-plane maneuver has been performed, it is necessary to determine whether an in-plane maneuver is required to ensure the continuity and completeness of the orbit control process and avoid the accumulation of orbit parameter coupling deviations after out-of-plane maneuvers. The determination of whether to perform an in-plane maneuver is based on in-plane constraints, ensuring that in-plane maneuvers are only performed when necessary, reducing fuel consumption. Compared to strategies that predict the in-plane maneuver time and velocity increment at the descending node, this invention performs calculations near the maneuver to reduce extrapolation errors, thereby reducing the calculation error of velocity increments and suppressing excessive westward shift of the descending node longitude, enhancing control stability. It ensures that in-plane control is not limited by the ascending node state, enabling continuous monitoring and correction of in-plane parameters throughout the entire orbital cycle, avoiding error accumulation caused by waiting for the ascending node, simplifying the control calculation process when out-of-plane maneuvers are not required, avoiding redundant calculations, and ensuring the real-time performance of in-plane control.

[0033] Furthermore, the out-of-plane maneuver includes: acquiring the target out-of-plane maneuver velocity increment corresponding to the out-of-plane maneuver command, so as to perform out-of-plane maneuver based on the target out-of-plane maneuver velocity increment.

[0034] In the above scheme, firstly, it is determined whether the satellite is currently passing through the ascending node based on real-time instantaneous navigation data and data from adjacent moments. Then, it is determined whether to perform an out-of-plane maneuver by combining the judgment of whether there are historical out-of-plane maneuver commands. Finally, the out-of-plane maneuver velocity increment of the target is obtained based on the out-of-plane maneuver command obtained when the satellite last passed through the descending node, and the out-of-plane maneuver is performed to increase the normal velocity of the satellite by a corresponding value, thereby achieving orbit correction. By judging the ascending node and the out-of-plane maneuver command, it is ensured that the out-of-plane maneuver is executed accurately at the ascending node, maximizing the orbit correction effect and reducing fuel waste.

[0035] Furthermore, the in-plane maneuvering control action based on the instantaneous orbital element vector includes:

[0036] If the current orbital revolutions, instantaneous latitude argument, final execution time of in-plane maneuver, and preset in-plane maneuver constraints determine that the preset in-plane maneuver conditions are not met, then no in-plane maneuver will be performed.

[0037] Further, the step of obtaining the in-plane maneuver velocity increment based on the adjacent descending intersection data and the instantaneous orbital element vector includes:

[0038] Based on the adjacent descending node data, obtain the adjacent average orbital element vector, adjacent descending node longitude error, adjacent node period and adjacent average semi-major axis attenuation rate.

[0039] The eastward limit error value is obtained based on the adjacent average orbital element vector and the adjacent average semi-major axis attenuation rate.

[0040] The average semi-major axis error change value is obtained based on the eastward shift limit error value and the period of adjacent nodes;

[0041] The semi-major axis lift is obtained based on the average semi-major axis error change value, the longitude error of adjacent descending intersection points, and the preset parabolic symmetry axis offset algorithm. Then, the in-plane maneuver speed increment is obtained based on the semi-major axis lift and the instantaneous orbital element vector.

[0042] In the above scheme, when obtaining the in-plane maneuver speed increment, a preset parabolic symmetry axis offset algorithm is used. That is, the parabolic relationship between the longitude error of the descending node and the average semi-major axis error is used. Targeted calculations are performed based on the different degrees of parabolic symmetry axis offset, achieving dynamic correction of the average semi-major axis error. This avoids semi-major axis control failure caused by track inclination error coupling, thus significantly enhancing track control stability. Simultaneously, the dynamic correction of the average semi-major axis error ensures that the calculated speed increment correctly raises the semi-major axis to the appropriate position, avoiding continuous decay of the semi-major axis mean and control rate deviation, thereby solving the failure problem. Compared to existing real-time calculations of the descending node per lap, this invention only calculates in-plane maneuvering when preset in-plane maneuvering conditions are met. This achieves the calculation of in-plane maneuver speed increments when the drift rate change in the neighborhood stabilizes and the time is similar, reducing speed increment calculation errors, significantly enhancing control stability, effectively offsetting the eastward shift error of the descending node longitude, and preventing in-plane parameters from becoming uncontrollable due to coupling effects.

[0043] Further, the acquisition of historical position vectors and historical velocity vectors, and the acquisition of average orbital element errors based on the real-time position vector, real-time velocity vector, historical position vector, and historical velocity vector, includes:

[0044] The crossing time and crossing coordinates are obtained based on the real-time position vector, real-time velocity vector, historical position vector, historical velocity vector, and linear interpolation method.

[0045] Based on the crossing time and crossing coordinates, obtain the longitude of the descending node and the current orbital node period;

[0046] Obtain historical descending intersection data, and based on the historical descending intersection data and a preset trapezoidal integral formula, obtain the average eccentricity vector and the average orbital inclination angle;

[0047] The average orbital element vector is obtained based on the descending node longitude, current orbital node period, average eccentricity vector, and average orbital inclination. The average orbital element error is then obtained based on the average orbital element vector and the nominal orbital element vector.

[0048] In the above scheme, the crossing time and crossing coordinates are obtained by linear interpolation, and then the current orbital node period is obtained. This accurately quantifies the deviation between the actual orbit and the nominal orbit, providing reliable basic data for subsequent attenuation rate calculation, maneuver time prediction, and threshold judgment, thus ensuring the accuracy of control decisions.

[0049] Further, the step of obtaining the historical node periodic error sequence, and obtaining the average semi-major axis decay rate based on the historical node periodic error sequence, the current instantaneous orbital element vector, and a preset atmospheric drag model, includes:

[0050] The initial semi-major axis decay rate is obtained based on the current instantaneous orbital element vector and the preset atmospheric drag model;

[0051] Obtain the historical node periodic error sequence;

[0052] When the historical node periodic error sequence is determined to meet the preset weighted calculation conditions, a weighted least squares method fitting is performed on the historical node periodic error sequence to obtain the node periodic error decay rate.

[0053] A coupled model is constructed based on perturbation theory; the average semi-major axis decay rate is obtained based on the node periodic error decay rate, historical node periodic error sequence, initial semi-major axis decay rate, and the coupled model.

[0054] When the historical node periodic error sequence is determined to be inconsistent with the preset weighted calculation conditions, the initial semi-major axis attenuation rate is directly used as the average semi-major axis attenuation rate.

[0055] In the above scheme, a coupled model is constructed by combining perturbation theory and fitted by weighted least squares method to suppress the error caused by the uncertainty of atmospheric density model, realize dynamic update of semi-major axis decay rate, significantly improve the calculation accuracy and stability of average semi-major axis decay rate, and help to achieve precise long-term orbit control.

[0056] Further, the in-plane maneuver time prediction based on the average orbital element error, average semi-major axis attenuation rate, calibrated descending node longitude westward shift limit, and preset descending node longitude eastward shift limit, to obtain the theoretical execution time and predicted orbital revolutions of the in-plane maneuver, and the final execution time of the in-plane maneuver based on the theoretical execution time and predicted orbital revolutions, includes:

[0057] Obtain the historical longitude error of the descending node, and based on the historical longitude error of the descending node and the average orbital element error, determine when the westward shift limit has been passed, and obtain the measured average semi-major axis error.

[0058] Based on the measured average semi-major axis error, the preset eastward shift limit of the descending node longitude, the average semi-major axis attenuation rate, and the preset orbital dynamics model, the estimated value of the average semi-major axis error is obtained.

[0059] Obtain the historical mean semi-major axis error, historical maneuver time, and historical node period after maneuvering in adjacent planes;

[0060] Based on the historical average semi-major axis error, historical maneuver time, historical node period, average semi-major axis error estimate and average semi-major axis decay rate, the in-plane maneuver theoretical execution time and the predicted orbit number corresponding to the in-plane maneuver theoretical execution time are obtained.

[0061] Based on the in-plane maneuver theory execution time and the predicted orbital number, the average eccentricity error of the satellite during its last passing of the descending node within the predicted orbital number is obtained.

[0062] The in-plane maneuver latitude argument is obtained based on the average eccentricity error, and the in-plane maneuver latitude argument is used as the final execution time of the in-plane maneuver.

[0063] In the above scheme, the westward limit error is dynamically corrected by dynamic correction of the action to predict and obtain a more accurate final execution time of the in-plane maneuver. First, the theoretical execution time and the predicted number of revolutions are calculated. Then, the final execution time of the in-plane maneuver is determined by combining the eccentricity error. The optimal number of orbits and phase of the in-plane maneuver are accurately locked. This ensures that the semi-major axis lifting action can be executed in time before the longitude of the descending node reaches the eastward limit, while the eccentricity error is effectively controlled without the need for additional maneuvers. This avoids the error from exceeding the threshold and maintains the long-term stability of the in-plane parameters.

[0064] This invention provides an autonomous control method for sun-synchronous return orbits, which performs out-of-plane maneuvers at the ascending node and dynamically predicts the timing of in-plane maneuvers at the descending node, thus optimizing orbit control efficiency. The improved out-of-plane control method separates the short-period and long-period terms of orbital inclination error and designs a threshold triggering mechanism to solve the control failure problem caused by the dominance of the long-period term in existing methods. This effectively addresses the problem of the long-period dominance of orbital inclination error caused by solar gravitational perturbations in strictly regressive sun-synchronous orbits at local time (non-dusk, non-midnight nodes), filling a technological gap. Furthermore, it transforms the complex orbital inclination error prediction control into a method based on the semi-lunar period... Threshold-triggered control of the period term reduces computational complexity and enhances robustness. Simultaneously, a parabolic symmetry axis offset compensation strategy is introduced, dynamically correcting the mean offset and rise of the semi-major axis to address the strong coupling problem between in-plane and out-of-plane parameters caused by the monotonic variation of orbital inclination error. This further optimizes the in-plane control method and effectively improves the stability of orbital control. By fusing theoretical models and actual observation data, the uncertainty of the atmospheric model is suppressed, optimizing the long-term calculation accuracy and stability of the key parameter, the mean semi-major axis decay rate. Ultimately, high-precision autonomous control of the strictly regressed orbit is achieved, significantly improving the long-term stability of orbital control. Attached Figure Description

[0065] To more clearly illustrate the technical solution of this application, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.

[0066] Figure 1 This is a schematic diagram of an autonomous control method for a sun-synchronous return orbit provided in this embodiment;

[0067] Figure 2 This is a schematic diagram of an RTN_I coordinate system provided in this embodiment;

[0068] Figure 3 This is a schematic diagram of a process of raising a semi-major shaft to perform in-plane maneuvering, provided in this embodiment.

[0069] Figure 4 This is a schematic diagram of in-plane maneuver time prediction and in-plane maneuver semi-major axis lifting provided in this embodiment;

[0070] Figure 5 This is a schematic diagram illustrating the change of track inclination error over time when the LTDN is 10:00, as provided in this embodiment.

[0071] Figure 6 This embodiment provides a schematic diagram of out-of-plane maneuvering within six months when the LTDN is 10:00.

[0072] Figure 7 This is a schematic diagram of the autonomous control process for a sun-synchronous return orbit provided in this embodiment. Detailed Implementation

[0073] To make the objectives, technical solutions, and advantages of this application clearer, the technical solutions of this application will be clearly and completely described below with reference to the accompanying drawings of the embodiments. Obviously, the described embodiments are only some embodiments of this application, not all embodiments. Based on the embodiments of this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0074] Unless otherwise defined, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this application pertains; the terminology used herein is for the purpose of describing particular embodiments only and is not intended to limit the application; the terms “comprising” and “having”, and any variations thereof, in the specification, claims, and foregoing description of the drawings are intended to cover non-exclusive inclusion.

[0075] In the description of the embodiments of this application, technical terms such as "first" and "second" are used only to distinguish different objects and should not be construed as indicating or implying relative importance or implicitly specifying the number, specific order, or primary and secondary relationship of the indicated technical features. In the description of the embodiments of this application, "multiple" means two or more, unless otherwise explicitly defined.

[0076] In this document, the term "embodiment" means that a particular feature, structure, or characteristic described in connection with an embodiment may be included in at least one embodiment of this application. The appearance of this phrase in various places throughout the specification does not necessarily refer to the same embodiment, nor is it a separate or alternative embodiment mutually exclusive with other embodiments. It will be explicitly and implicitly understood by those skilled in the art that the embodiments described herein can be combined with other embodiments.

[0077] In the description of the embodiments in this application, the term "and / or" is merely a description of the relationship between related objects, indicating that three relationships can exist. For example, A and / or B can represent: A existing alone, A and B existing simultaneously, and B existing alone. Additionally, the character " / " in this document generally indicates that the preceding and following related objects have an "or" relationship.

[0078] In the description of the embodiments of this application, the term "multiple" refers to two or more (including two), similarly, "multiple sets" refers to two or more (including two sets), and "multiple pieces" refers to two or more (including two pieces).

[0079] In the description of the embodiments of this application, unless otherwise expressly specified and limited, technical terms such as "installation," "connection," "joining," and "fixing" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral part; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal communication of two components or the interaction between two components. For those skilled in the art, the specific meaning of the above terms in the embodiments of this application can be understood according to the specific circumstances.

[0080] This embodiment provides an autonomous control method for a sun-synchronous return orbit, such as... Figure 1 As shown, it includes:

[0081] S1. Obtain the real-time position vector and real-time velocity vector based on real-time navigation data, and obtain the current instantaneous orbital element vector based on the real-time position vector and real-time velocity vector;

[0082] S2. Obtain the real-time adjacent time position vector. Based on the real-time position vector and the real-time adjacent time position vector, determine when the satellite passes through the descending node. Based on the current instantaneous orbit element vector, perform the in-plane and out-of-plane command calculation steps to obtain the final execution time of the in-plane maneuver and the out-of-plane maneuver command.

[0083] S3. Perform in-plane maneuvers based on the final execution time of the in-plane maneuvers, and perform out-of-plane maneuvers based on the out-of-plane maneuver commands when the satellite passes the ascending node, thereby achieving autonomous control of the sun-synchronous return orbit;

[0084] The in-plane and out-of-plane command calculation steps include:

[0085] S01. Obtain historical position vector and historical velocity vector, and obtain average orbital element error based on the real-time position vector, real-time velocity vector, historical position vector and historical velocity vector;

[0086] S02. Obtain the historical node periodic error sequence, and obtain the average semi-major axis decay rate based on the historical node periodic error sequence, the current instantaneous orbital element vector, and the preset atmospheric drag model.

[0087] S03. Based on the average orbital element error, average semi-major axis attenuation rate, calibrated westward shift limit of descending node longitude, and preset eastward shift limit of descending node longitude, predict the in-plane maneuver time, obtain the theoretical execution time of the in-plane maneuver and the predicted number of orbits, and obtain the final execution time of the in-plane maneuver based on the theoretical execution time of the in-plane maneuver and the predicted number of orbits.

[0088] S04. Obtain the orbit inclination error based on the average orbit element error, and separate the semi-monthly short-period term based on the orbit inclination error to obtain the semi-monthly orbit inclination error;

[0089] S05. When the inclination error of the semi-lunar orbit exceeds the preset inclination error threshold, obtain the target out-of-plane maneuver speed increment, and obtain the out-of-plane maneuver command based on the target out-of-plane maneuver speed increment.

[0090] This embodiment provides an autonomous control method for a sun-synchronous return orbit. First, it analyzes navigation data in real time to obtain real-time position and velocity vectors, and calculates the current instantaneous orbital element vector. Then, it combines this with the real-time adjacent position vectors from historical data to determine whether the satellite has passed the descending node. The descending node is used as the anchor point for calculating in-plane and out-of-plane maneuver commands. When the satellite passes the descending node, the out-of-plane maneuver command and the final execution time of the in-plane maneuver are calculated. This allows the satellite to execute the out-of-plane maneuver when it passes the ascending node, enabling it to calculate the maneuver magnitude and execute the in-plane maneuver at the correct latitude argument, thus achieving sun synchronization for local times other than dawn / dusk and midnight nodes. The control of the regressive orbit addresses the technical problem of existing technologies that only apply to sun-synchronous regressive orbit control when the local time of the descending node is 0:00, 6:00, 12:00, or 18:00. During calculations at the descending node, the average orbital element error is first calculated based on real-time and historical descending node data to quantify the deviation between the actual and nominal orbits. This transforms the abstract orbital state into a quantifiable error index. The decay rate is then accurately obtained through instantaneous orbital elements, average element error, and an atmospheric drag model, ensuring precise calculation of the semi-major axis decay rate. This allows for the application of the average orbital element error, average semi-major axis decay rate, and westward-moving limit prediction surface. The in-plane maneuver execution time is planned in advance to ensure that the semi-major axis lift can accurately suppress the eastward shift of the intersection longitude, thus improving the accuracy of autonomous orbit control. Simultaneously, the semi-monthly short-period term is separated from the orbit inclination error to obtain the semi-monthly orbit inclination error. Finally, a threshold judgment mechanism is constructed based on the error threshold. When the orbit inclination error exceeds the preset inclination error threshold, the target out-of-plane maneuver velocity increment is calculated to achieve out-of-plane maneuver control, thereby achieving precise maintenance of the orbit inclination. This transforms the complex orbit inclination error prediction control into threshold-triggered control based on the separation of the semi-monthly period term, reducing computational complexity and ensuring that the orbit inclination is precisely controlled within the threshold. Within this scope, it is simple, clear, and highly stable, simplifying the original strategy and enhancing robustness. It enables in-plane and out-of-plane maneuver calculation and prediction at the descending node, and at the ascending node, it executes out-of-plane maneuver commands obtained when the threshold is met, and performs out-of-plane maneuvers in conjunction with the in-plane maneuver execution time when the preset in-plane maneuver conditions are met, thereby improving maneuver efficiency. It solves the technical problem of control failure caused by the long-term dominant orbital inclination error due to solar gravitational perturbation in strictly regressive local time sun-synchronous orbits that are not at dawn / dusk or midnight nodes. While ensuring high precision of autonomous control of sun-synchronous orbits, it enhances the long-term stability and robustness of control.

[0091] In specific implementation, this invention provides an autonomous and precise orbit control method for sun-synchronous return orbits that are not at local time (not at dawn / dusk / midnight). This orbit control method is a high-precision autonomous control method for strictly returning orbits, specifically for sun-synchronous return orbits where the descending node's local time is not close to the exact times 0 / 6 / 12 / 18 (i.e., not at dawn / dusk / midnight). The reference orbit designed in this embodiment for comparison with the actual orbit, and the actual orbit after control in a real-world scenario, both meet the requirement of strict return. This requires the actual trajectory to always remain strictly within a spatial circular tube centered on the reference trajectory, with a small tube radius (less than one kilometer), thus achieving high precision.

[0092] In the specific implementation process, when performing out-of-plane maneuvers, the velocity increment data in the command is extracted for out-of-plane maneuvers: the thrusters are ignited according to the data to generate corresponding thrust, which increases the normal velocity by the corresponding value.

[0093] In the specific implementation process, this embodiment acquires real-time position and velocity vectors based on real-time navigation data each time, and then acquires the current instantaneous orbital element vector based on the real-time position and velocity vectors. This vector needs to be saved for use in the next process. The average orbital element vector is used. It describes the deviation between the reference track and the actual track, and performs a maneuver when the actual deviation between the two exceeds the corresponding control threshold. The vector includes the node period T j Average eccentricity vector Mean orbital inclination and the right ascension longitude λ of the descending node j In this embodiment, the out-of-plane maneuver employs a threshold triggering mechanism to address the technical problem of out-of-plane control strategy failure. Since the original strategy in the prior art cannot be used when no extreme value exists, this embodiment designs a new strategy, namely the threshold triggering mechanism, based on the new changes in the orbital inclination error.

[0094] Optionally, step S3 includes:

[0095] The system acquires real-time instantaneous navigation data from the satellite and obtains instantaneous position and velocity vectors based on the instantaneous navigation data.

[0096] The instantaneous orbital element vector is obtained based on the instantaneous position vector and instantaneous velocity vector;

[0097] Obtain the position vectors of adjacent instantaneous moments;

[0098] When the satellite passes through the ascending node based on the instantaneous position vector and the position vectors of the instantaneous adjacent moments, and when it is determined that there is an out-of-plane maneuver command, an out-of-plane maneuver is executed based on the out-of-plane maneuver command.

[0099] Based on the instantaneous orbital element vector, perform in-plane maneuvering control actions:

[0100] The instantaneous orbital element vector is obtained based on the instantaneous navigation data, and the instantaneous latitude argument is obtained based on the instantaneous orbital element vector;

[0101] Get the current orbital revolution count;

[0102] Based on the current orbital revolutions, instantaneous latitude argument, final execution time of in-plane maneuver, and preset in-plane maneuver constraints, when the preset in-plane maneuver conditions are met, adjacent descending intersection data are obtained;

[0103] The in-plane maneuver velocity increment is obtained based on the adjacent descending intersection point data and the instantaneous orbital element vector;

[0104] In-plane maneuvers are performed based on the in-plane maneuver speed increment.

[0105] In practice, real-time navigation data and instantaneous navigation data are derived from navigation data analyzed in real-time by the onboard GPS receiver. When obtaining the instantaneous orbital element vector based on the instantaneous navigation data, the satellite's current position vector in the geocentric fixed coordinate system is first obtained. and current velocity vector The units are m and m / s, respectively, based on Calculate the instantaneous orbital element vector

[0106] [a,e,i,Ω,ω,M] T These correspond to the semi-major axis, eccentricity, orbital inclination, right ascension of the ascending node, argument of perigee, and mean perigee, respectively. The current orbital number is obtained by summing the data after each orbit. The orbital number is incremented by one each time the descending node is passed, and the current orbital number is updated accordingly.

[0107] In the specific implementation process, to meet the preset in-plane maneuver conditions, three constraints must be met in sequence. The first condition is that the predicted number of orbits N for the in-plane maneuver has already been obtained in the in-plane and out-of-plane command calculation step when the plane passed the descending intersection point last time. pred The second condition is orbital rotation matching, i.e., the current number of orbits N of the satellite. current Equal to the predicted orbital number N pred The third condition is phase angle synchronization, meaning the satellite's instantaneous latitude argument u reaches the target value u. I That is, satisfying |uu I |≤ε u (εu (This is the tolerance threshold).

[0108] Optionally, before performing in-plane maneuvering control actions based on the instantaneous orbital element vector, the method further includes:

[0109] When the satellite passes through the ascending node based on the instantaneous position vector and the position vector of the instantaneous adjacent time, and when it is determined that there is no out-of-plane maneuver command, the in-plane maneuver control action is directly performed based on the instantaneous orbital element vector.

[0110] Optionally, before the step of performing in-plane maneuvering control actions based on the instantaneous orbital element vector, the method further includes:

[0111] When the satellite does not pass through the ascending node, it is determined based on the instantaneous position vector and the position vector of the instantaneous adjacent time. In this case, the in-plane maneuvering control action is performed directly based on the instantaneous orbital element vector.

[0112] In practice, performing an out-of-plane maneuver refers to applying a velocity pulse perpendicular to the orbital plane to the satellite. In this embodiment, the application position of this out-of-plane velocity pulse is preferably controlled at the ascending node to maximize inclination control efficiency. Therefore, when it is determined that the satellite has just passed the ascending node, it is necessary to first determine whether there is an out-of-plane maneuver command to be performed. If so, the out-of-plane maneuver command needs to be transmitted to the propulsion system. Based on the out-of-plane velocity increment value specified in the out-of-plane maneuver command, the propulsion system triggers the thrusters to ignite at the ascending node, generating an instantaneous thrust pulse perpendicular to the orbital plane. This pulse enables the satellite to obtain a corresponding normal velocity increment ΔV. N This enables efficient adjustment of the orbital inclination and completes control of the satellite outside the orbital plane. The orbital maintenance control is decoupled into out-of-plane control and in-plane control. In-plane control (corresponding to in-plane maneuvers) uses only tangential maneuvers to adjust the semi-major axis a and eccentricity e, while out-of-plane control (corresponding to out-of-plane maneuvers) uses normal maneuvers to adjust the orbital inclination i.

[0113] In the specific implementation process, this embodiment defines the satellite's orbital coordinate system (RTN_I coordinate system) based on the geocentric inertial coordinate system (ECI) as follows: Figure 2The RTN_I and RTN_E coordinate systems are shown. In this embodiment, the RTN coordinate system is divided into two categories: one is the RTN_I coordinate system, defined on the geocentric inertial coordinate system. The R-axis is the line connecting the geocenter and the sampling point on the nominal orbit, pointing to the sampling point; the T-axis can form a right angle with the R-axis, pointing to the direction of satellite motion on the geocentric inertial coordinate system ECI; the N-axis is perpendicular to the R and T axes, forming a right-handed rectangular coordinate system. The other is the RTN_E coordinate system, defined on the geocentric geofixed coordinate system ECEF. The difference between the two is that the T-direction of the RTN_I coordinate system is perpendicular to the R-direction and points to the direction of satellite motion on the ECEF coordinate system. The velocity vector difference between the two T-axis comes from the Earth's rotation. Both coordinate systems have their uses. In this embodiment, the orbital tangent plane xoz used for error analysis of the nominal orbit and the actual operating orbit is defined as the RON plane of the RTN_E coordinate system; while the velocity increment component of the pulse maneuver is defined on the RTN_I coordinate system.

[0114] Optionally, the out-of-plane maneuver includes: acquiring the target out-of-plane maneuver velocity increment corresponding to the out-of-plane maneuver command, so as to perform out-of-plane maneuver based on the target out-of-plane maneuver velocity increment.

[0115] In the specific implementation process, when determining whether a satellite has passed through the ascending node, the position vectors at adjacent time points are compared. The Z-axis component is used to achieve this: First, the instantaneous position vector and instantaneous velocity vector at this moment are obtained based on the instantaneous navigation data. Then, the instantaneous position vector of the previous adjacent moment corresponding to the instantaneous position vector is obtained, and the position vectors of the previous moment and this moment are compared. Z-axis component z n and z n+1 Compared with zero respectively, if z satisfies n <0 and z n+1 If the value is greater than 0, the satellite has just crossed the ascending node; otherwise, it has not.

[0116] Optionally, the in-plane maneuvering control action based on the instantaneous orbital element vector includes:

[0117] If the current orbital revolutions, instantaneous latitude argument, final execution time of in-plane maneuver, and preset in-plane maneuver constraints determine that the preset in-plane maneuver conditions are not met, then no in-plane maneuver will be performed.

[0118] Optionally, obtaining the in-plane maneuver velocity increment based on the adjacent descending intersection data and the instantaneous orbital element vector includes:

[0119] Based on the adjacent descending node data, obtain the adjacent average orbital element vector, adjacent descending node longitude error, adjacent node period and adjacent average semi-major axis attenuation rate.

[0120] The eastward limit error value is obtained based on the adjacent average orbital element vector and the adjacent average semi-major axis attenuation rate.

[0121] The average semi-major axis error change value is obtained based on the eastward shift limit error value and the period of adjacent nodes;

[0122] The semi-major axis lift is obtained based on the average semi-major axis error change value, the longitude error of adjacent descending intersection points, and the preset parabolic symmetry axis offset algorithm. Then, the in-plane maneuver speed increment is obtained based on the semi-major axis lift and the instantaneous orbital element vector.

[0123] In the specific implementation process, when obtaining the in-plane maneuvering velocity increment, the tangential velocity increment ΔV is calculated in three stages. T :

[0124] In the first stage, based on the data of adjacent descending nodes of the previous descending node, an orbital error vector is obtained, including the adjacent average orbital element vector, the adjacent descending node longitude error, the adjacent node period, and the adjacent average semi-major axis decay rate. Using the parabolic relationship between the longitude error of the descending node and the mean semi-major axis error, the eastward shift limit error value δλ of the longitude of the descending node can be calculated. I ;

[0125] In the second stage, the longitude error of the descending node is calculated to drift from 0m to δλ. I The absolute value of the change in the mean semi-major axis error during the period The expression is as follows:

[0126]

[0127] In the above formula, k1 and k2 both represent intermediate parameters, and k1 and k2 are the two partial derivatives of the descending intersection longitude error with respect to the semi-major axis and the track inclination angle. and Related; T represents the historical node period;

[0128] In the third stage, a pre-defined parabolic axis of symmetry offset algorithm is used to calculate the velocity increment. Based on the different degrees of parabolic axis of symmetry offset, two methods for calculating the semi-major axis lift Δa are proposed: when the offset is negligible, a simple double lift strategy is adopted. When the symmetry axis is significantly negatively offset, through Compensation offset, forcibly raising the semi-major axis to Nearby, in which the formula This represents the error of the average semi-major axis when the longitude error of the descending node reaches the eastward shift limit; finally, the tangential velocity increment ΔV required for in-plane maneuvering is calculated. T , Where n represents the satellite's orbital angular velocity. In the specific implementation process, after calculating the in-plane maneuver velocity increment, in-plane maneuvers are executed according to the in-plane maneuver velocity increment. The in-plane maneuver planning adopted in this embodiment follows a periodic phase control strategy. Specifically, as the semi-major axis of the orbit decays, the longitude of the descending node gradually shifts eastward to the eastward deviation limit λ. E At this point, in-plane maneuvers are performed to raise the semi-major axis. The semi-major axis increases by an increment Δa, becoming positive. Afterward, the semi-major axis continues to decay, and the longitude of the descending node shifts westward and then eastward to the eastern deflection limit, repeating the cycle. The process of raising the semi-major axis to perform in-plane maneuvers is as follows: Figure 3 As shown.

[0129] Specifically, the in-plane maneuver time prediction diagram is as follows: Figure 4 As shown in (a), the schematic diagram of the in-plane kinematic semi-major shaft lifting is as follows. Figure 4 As shown in (b), where, Figure 4 (a) and Figure 4 (b) The vertical axis represents the semimajor axis error (λ). a The horizontal axis represents the longitude error (δλ) of the descending node, the curve trajectory reflects the evolution of orbital elements, and the arrows indicate the direction of motion. Figure 4 (a) in u I This represents the final execution time before the theoretical execution time determined based on the eccentricity error, (δλ) I ,δa I ) represents the error value at the theoretical execution time, predicting the satellite's arrival at the East limit (λ). E The time interval near the easternmost point (e.g., the time from the eastern margin (East)) provides a basis for selecting the timing of in-plane maneuvers, where (δλ) W ,δa W () indicates the semi-major axis error and longitude error when the westward shift limit is reached; Figure 4 (b) Curves A and B represent two methods for calculating the semi-major axis lift. Curve A represents the maneuvering situation when the axis of symmetry offset is small, while curve B represents the maneuvering situation when the axis of symmetry offset is large, indicating that the semi-major axis error is forcibly lifted to Δa. up The left and right diagrams illustrate the changes in the rise of the semi-major axis error before and after the maneuver. When the satellite finally reaches its eastward shift limit, the semi-major axis error is specifically adjusted, demonstrating the active control effect of in-plane maneuvers on the orbital semi-major axis. This can be used to maintain or adjust the orbital semi-major axis. The eastward margin represents the allowable deviation range in the longitude direction to the east. Compared to the real-time calculation of the descending node point per orbit in existing technologies, this strategy achieves better performance in λ... ENeighborhood calculation has certain advantages, namely, the drift rate changes tend to stabilize in the neighborhood and the time is similar, which reduces the extrapolation error of the descending intersection longitude, reduces the calculation error of the in-plane maneuver speed increment, and thus significantly enhances the stability of the control.

[0130] In the specific implementation process, regardless of whether an in-plane maneuver is performed, it is necessary to determine whether the descending node has been passed. The method for determining this is the same as that for determining whether the ascending node has been passed, as described earlier: if z n >0 and z n+1 If the value is less than 0, the satellite is determined to have just passed the descending node; otherwise, it has not passed the descending node, and the autonomous orbit control process ends. If the satellite has passed the descending node, the in-plane and out-of-plane command calculation steps are performed, which include five strictly sequential sub-steps S01, S02, S03, S04, and S05.

[0131] Optionally, step S01 includes:

[0132] The crossing time and crossing coordinates are obtained based on the real-time position vector, real-time velocity vector, historical position vector, historical velocity vector, and linear interpolation method.

[0133] Based on the crossing time and crossing coordinates, obtain the longitude of the descending node and the current orbital node period;

[0134] Obtain historical descending intersection data, and based on the historical descending intersection data and a preset trapezoidal integral formula, obtain the average eccentricity vector and the average orbital inclination angle;

[0135] The average orbital element vector is obtained based on the descending node longitude, current orbital node period, average eccentricity vector, and average orbital inclination. The average orbital element error is then obtained based on the average orbital element vector and the nominal orbital element vector.

[0136] In the specific implementation process, the average orbital element error is calculated. In order to optimize onboard storage resources, the average orbital element error is adopted. To describe the actual deviation, this embodiment calculates the crossing time t using linear interpolation based on the real-time navigation data, real-time position vector, and position vectors at adjacent times. DNj and crossing coordinates (x) DNj ,y DNj ,0), and then calculate the longitude λ of the descending intersection point. j and the node period T of the j-th track j Next, the average eccentricity vector is calculated using the trapezoidal integral formula based on the node period and the instantaneous orbital element vector. and average orbital inclination Finally, the average orbital element vector is obtained. The average orbital element vector of the actual orbit is known. and the average orbital element vector of the nominal orbit Then, subtracting the two yields the average orbital element error.

[0137] Optionally, step S02 includes:

[0138] The initial semi-major axis decay rate is obtained based on the current instantaneous orbital element vector and the preset atmospheric drag model;

[0139] Obtain the historical node periodic error sequence;

[0140] When the historical node periodic error sequence is determined to meet the preset weighted calculation conditions, a weighted least squares method fitting is performed on the historical node periodic error sequence to obtain the node periodic error decay rate.

[0141] A coupled model is constructed based on perturbation theory; the average semi-major axis decay rate is obtained based on the node periodic error decay rate, historical node periodic error sequence, initial semi-major axis decay rate, and the coupled model.

[0142] When the historical node periodic error sequence is determined to be inconsistent with the preset weighted calculation conditions, the initial semi-major axis attenuation rate is directly used as the average semi-major axis attenuation rate.

[0143] In the specific implementation process, the average track element error is obtained for coupled calculation of track decay rate and accuracy optimization. Firstly, based on track decay dynamics, the decay rate is calculated using the node period error. With average half-major axis attenuation The two satisfy the strict coupling relationship derived from perturbation theory, that is, a coupling model is constructed based on perturbation theory:

[0144]

[0145] Where J2 represents the Earth's oblateness perturbation coefficient, R e This represents the Earth's radius.

[0146] Next, an initial solution is performed using a pre-defined atmospheric drag model based on atmospheric theory. The expression for the pre-defined atmospheric drag model is as follows:

[0147]

[0148] Where ρ represents atmospheric density, C D The value represents the atmospheric drag coefficient, A represents the satellite's frontal area, and m represents the satellite's mass (μ). eThis represents the Earth's gravitational constant; however, due to its accuracy being limited by significant uncertainties in the atmospheric density model, a dual-modal optimization strategy is developed in this embodiment to improve robustness. After calculating the initial value of the average semi-major axis decay rate using an atmospheric theoretical model, this dual-modal optimization strategy acquires historical node period data containing at least 10 orbital cycles. At this point, the preset weighted calculation conditions are met, and the node period error decay rate is obtained by fitting the historical node period error sequence using the weighted least squares method. Then Substituting into the above coupling model, the average semi-major axis attenuation rate is obtained. Based on instantaneous orbital elements and a coupled model, the nodal period error decay rate is converted into the average semi-major axis decay rate. When there are fewer than 10 orbital cycles, the preset weighted calculation conditions are not met, and the initial semi-major axis decay rate is directly used as the average semi-major axis decay rate. Specifically, when there are fewer than 10 historical nodal period errors, no fitting calculation is performed; the average semi-major axis decay rate is calculated using a formula. When there are 10 or more data points, the nodal period error decay rate is fitted, and then the average semi-major axis decay rate is calculated through the coupling relationship, without using a formula. This approach effectively suppresses atmospheric model uncertainties by fusing theoretical modeling with observations, ensuring the stability of long-term orbital control. Furthermore, the nodal period error decay rate is more stable than the nodal period decay rate obtained by fitting the nodal period errors.

[0149] Optionally, step S03 includes:

[0150] Obtain the historical longitude error of the descending node, and based on the historical longitude error of the descending node and the average orbital element error, determine when the westward shift limit has been passed, and obtain the measured average semi-major axis error.

[0151] Based on the measured average semi-major axis error, the preset eastward shift limit of the descending node longitude, the average semi-major axis attenuation rate, and the preset orbital dynamics model, the estimated value of the average semi-major axis error is obtained.

[0152] Obtain the historical mean semi-major axis error, historical maneuver time, and historical node period after maneuvering in adjacent planes;

[0153] Based on the historical average semi-major axis error, historical maneuver time, historical node period, average semi-major axis error estimate and average semi-major axis decay rate, the in-plane maneuver theoretical execution time and the predicted orbit number corresponding to the in-plane maneuver theoretical execution time are obtained.

[0154] Based on the in-plane maneuver theory execution time and the predicted orbital number, the average eccentricity error of the satellite during its last passing of the descending node within the predicted orbital number is obtained.

[0155] The in-plane maneuver latitude argument is obtained based on the average eccentricity error, and the in-plane maneuver latitude argument is used as the final execution time of the in-plane maneuver.

[0156] In practice, if the in-plane prediction is made too early, the longitude of the descending node will shift excessively westward, and may even exceed the longitude control threshold due to continuous westward shift; if the in-plane prediction is made too late, the longitude will exceed the eastward shift limit and thus exceed the threshold.

[0157] In the specific implementation process, to solve the semi-major axis deviation problem caused by tilt angle error coupling, this embodiment combines historical descending node longitude errors to predict and compensate for in-plane maneuvering time. Based on the local time drift characteristics of the descending node, it predicts the number of orbits to be executed for in-plane maneuvering and determines the latitudinal argument u of the in-plane maneuvering. I Among these, the semi-major axis deviation caused by the coupling of tilt angle errors needs to be addressed as a key issue. This occurs when the longitude of the descending node reaches the westward shift limit λ. w At that time, due to tilt angle error Average semi-major axis error generated by J2 perturbation coupling Always negative Instead of the zero value assumed by existing models, this embodiment employs a dynamic correction control strategy to eliminate this influence. When the preset dynamic correction conditions are met based on the average orbital element error and the preset westward shift limit (i.e., after the descending node longitude has passed the westward shift limit), a dynamic correction action is performed: first, the westward shift limit error is calibrated; then, after the descending node longitude has passed the westward shift limit, the average semi-major axis error is updated in real time. The actual measured value, i.e., the mean semi-major axis error measurement value, is shifted westward to the limit (λ). W The error is calibrated. Next, the eastward shift limit error is predicted, which is the calibrated westward shift limit error. Substituting into the dynamic model, the longitude of the descending node reaches the preset eastward shift limit λ. E Mean semi-major axis error prediction Finally, the computer's operating time is based on... Using the average semi-major axis attenuation rate and the target longitude threshold, the internal maneuver theoretical execution time t is calculated. I and the corresponding predicted orbital number N pred The calculation process satisfies the following formula:

[0158]

[0159] in, The error δλ represents the westward shift of the longitude of the descending node. w The mean semi-major axis error change up to the target longitude threshold; the subscript 0 indicates the initial parameters after the last maneuver. The final in-plane maneuver execution time is determined by the in-plane maneuver latitude argument u. IIt is decided that, at the final passage through the descending intersection point before the maneuver, the latitude argument required to compensate for the average eccentricity error at that moment will be used as the final execution time u. I ,

[0160] When calculating the semi-lunar orbital inclination error, if the short-period error is not isolated and the complete inclination error is used directly to determine whether the control threshold is exceeded, the error may still exceed the control threshold after control due to fluctuations in the short-period error. After removing the short-period error from the semi-lunar calculation, the inclination error decreases almost linearly. For example, the change in orbital inclination error with horizontal axis time (days) at the descending node local time LTDN of 10:00 is shown below. Figure 5 As shown, the short-period term of the track inclination error is as follows: Figure 5 As shown in (a) (its vertical axis unit is deg*10) -4 The long-period term of the track inclination error (its vertical axis unit is deg*10). -3 )like Figure 5 As shown in (b), the orbital inclination error periodic term is a combination of the short-period term and the long-period term (its vertical axis magnitude is deg*10). -3 )like Figure 5 As shown in (c), the long-term term of the track inclination error (its vertical axis magnitude is deg) is as follows: Figure 5 As shown in (d), the track inclination error (its vertical axis magnitude is deg) is as follows: Figure 5 As shown in (e), the local time at non-twilight and non-midnight points causes the total error of the orbital inclination to show a fluctuating decreasing trend, with the fluctuation mainly caused by short-period terms. The error is caused by a period of half a month. Subtracting this short-period term will make the track inclination error decrease in a roughly linear manner. In this embodiment, the track inclination error after subtracting the half-month short-period term is used to determine whether it exceeds the preset inclination error threshold. This method is simple, clear, more accurate and stable, and ensures that the track inclination is stably and accurately controlled within the threshold.

[0161] In the specific implementation process, the criterion for judging whether the track inclination error exceeds the control threshold in steps S04 and S05 is to judge the semi-lunar track inclination error at this time. Does it exceed the custom deviation threshold? in That is, the total error of the track inclination angle Subtract short-period terms The triggering condition set in this embodiment is: like Figure 6 The LTDN (Descending Node Local Time) shown is a schematic diagram of out-of-plane maneuvers over the past six months at 10:00. Figure 6As shown by the solid blue line, the local time at points other than dawn / dusk / midnight causes the total error in orbital inclination to show a fluctuating decreasing trend. This fluctuation is mainly due to short-period terms. Caused by a cycle of half a month, after subtracting the short-cycle item, as follows: Figure 6 As shown by the blue dashed line, the orbital inclination error can be reduced in a roughly linear fashion, making this a more stable criterion for judgment. During this process, it is important to pay attention to the deviation threshold of the semi-lunar orbital inclination. The threshold settings need to meet the following requirements: Figure 6 The constraint shown by the red dashed line is: less than the difference between the orbital inclination control threshold and the amplitude of the semi-monthly short-period term, to ensure a certain tolerance margin. Figure 6 The red solid line represents the upper and lower threshold limits of the orbital parameters. In this embodiment, the out-of-plane maneuver control method is simplified from complex predictive control to a threshold triggering mechanism, effectively solving the inclination instability problem of local time sun-synchronous return orbits that are not at dawn / dusk or midnight nodes. Furthermore, the orbital inclination state prediction module is eliminated, reducing algorithm complexity and effectively improving computational efficiency.

[0162] In the specific implementation process, the out-of-plane maneuver speed increment ΔV is calculated. N At that time, the deviation threshold is different from the user-defined threshold. The calculation process for the out-of-plane maneuver speed increment satisfies the following formula: First, the required compensation amount for the track inclination error is calculated. Then converted into the out-of-plane maneuver velocity increment ΔV of the maneuver pulse N Ignoring eccentricity and assuming the velocity direction coincides with the tangential direction, the final ΔV will be... N Saved as out-of-plane maneuver commands so that out-of-plane maneuvers can be performed when the satellite passes the next ascending node.

[0163] In summary, this embodiment provides an autonomous control process for a sun-synchronous return orbit, as follows: Figure 7 As shown, it includes:

[0164] start;

[0165] S11. Read GPS data and determine orbital position: acquire the navigation data of the satellite in real time, calculate the real-time position vector and real-time velocity vector of the current satellite, and use them to calculate the instantaneous orbital element vector;

[0166] S12. Determine whether the satellite has passed through the ascending node. If yes, proceed to step S13; otherwise, proceed to step S15.

[0167] S13. Determine if there is a historical out-of-plane maneuver command. If yes, proceed to step S14; otherwise, proceed to step S15.

[0168] S14. Perform out-of-plane maneuvers: Execute out-of-plane maneuvers based on the historical out-of-plane maneuver commands;

[0169] S15. Determine whether the in-plane maneuver conditions are met: Obtain the current orbit number and instantaneous latitude argument based on the historical orbit number and instantaneous orbital root vector; Determine whether the in-plane maneuver conditions are met based on the current orbit number and instantaneous latitude argument and the preset in-plane maneuver constraints. If yes, proceed to step S16; if no, proceed to step S17.

[0170] S16. Calculate the in-plane maneuver speed increment and execute the in-plane maneuver: acquire adjacent descending intersection data, and obtain the in-plane maneuver speed increment based on the adjacent descending intersection data and the instantaneous orbital element vector; execute the in-plane maneuver based on the in-plane maneuver speed increment.

[0171] S17. Determine whether the satellite has passed through the descending intersection point: Based on the real-time position vector and the position vector of adjacent historical time, determine whether the satellite has passed through the descending intersection point. If yes, execute step S18; otherwise, end the current control process.

[0172] S18. Track element error calculation: Obtain historical position vector and historical velocity vector, and obtain average track element error based on the real-time position vector, real-time velocity vector, historical position vector and historical velocity vector;

[0173] S19. Calculation of orbital decay rate: Obtain the historical node periodic error sequence, and obtain the average semi-major axis decay rate based on the historical node periodic error sequence, the current instantaneous orbital element vector and the preset atmospheric drag model.

[0174] S110. In-plane maneuver time prediction: Based on the average orbital element error, average semi-major axis decay rate, calibrated descending node longitude westward shift limit and preset descending node longitude eastward shift limit, in-plane maneuver time prediction is performed to obtain the theoretical execution time and predicted orbital revolutions of in-plane maneuver. Based on the theoretical execution time and predicted orbital revolutions of in-plane maneuver, the final execution time of in-plane maneuver is obtained.

[0175] S111. Determine whether the track inclination error exceeds the threshold: Obtain the track inclination error based on the average track element error, and separate the semi-monthly short-period term based on the track inclination error to obtain the semi-monthly track inclination error; Determine whether the semi-monthly track inclination error exceeds the preset inclination error threshold. If yes, execute step S112; otherwise, directly end the current control process.

[0176] S112. Calculate the out-of-plane maneuver speed increment and obtain the out-of-plane maneuver command: Obtain the target's out-of-plane maneuver speed increment and obtain the out-of-plane maneuver command based on the target's out-of-plane maneuver speed increment;

[0177] Finish.

[0178] The out-of-plane maneuvering command obtained in step S112 is used for out-of-plane maneuvering control when the satellite passes the ascending node again. After the current control process ends, the next real-time step begins, and the process is repeated. The next real-time step depends on GPS performance, and may involve performing the entire process every second or every two seconds, etc.

[0179] This embodiment provides an autonomous control method for a sun-synchronous return orbit. It performs out-of-plane maneuvers at the ascending node and dynamically predicts the timing of in-plane maneuvers at the descending node. Compared to existing autonomous control methods for strictly return orbits at dawn / dusk or noon / midnight, this embodiment has the following advantages: 1. Existing technologies are only applicable to sun-synchronous return orbits with a descending node local time of 0:00, 6:00, 12:00, or 18:00. This invention, by improving the out-of-plane control method (such as a threshold-triggered control mechanism and semi-lunar period error separation), effectively solves the problem of long-term orbital inclination error dominating due to solar gravitational perturbation in strictly return orbits with local times other than dawn / dusk (such as 10:00), filling a technological gap. 2. Control stability is improved. This embodiment introduces parabolic symmetry axis offset compensation (such as dynamic correction of the mean semi-major axis error) to avoid in-plane control failure caused by coupling of orbital parameters inside and outside the plane, optimizing the in-plane control method. Furthermore, it replaces the complex orbital inclination error prediction control with threshold-triggered control based on the semi-lunar period term, reducing computational complexity and enhancing robustness. 3. This embodiment combines a theoretical model with a least-squares fitting method to dynamically update the semi-major axis attenuation rate, improving the accuracy of atmospheric drag perturbation compensation. It achieves this by separating the short-period, long-period, and long-period terms of the orbital inclination error, designing a threshold triggering mechanism to address the control failure problem caused by the dominance of the long-period term in existing methods. This dynamically corrects the semi-major axis lift, offsetting the influence of the inclination error on the parabolic symmetry axis of the descending node longitude, performing out-of-plane maneuvers at the ascending node, and dynamically predicting the timing of in-plane maneuvers at the descending node, thus optimizing control efficiency.

[0180] The above description represents the preferred embodiments of the present invention. It should be noted that those skilled in the art can make various improvements and modifications without departing from the principles of the present invention, and these improvements and modifications are also considered to be within the scope of protection of the present invention.

Claims

1. An autonomous control method for a sun-synchronous return orbit, characterized in that, include: Real-time position vector and real-time velocity vector are obtained based on real-time navigation data, and the current instantaneous orbital element vector is obtained based on the real-time position vector and real-time velocity vector; Obtain the real-time adjacent time position vectors. Based on the real-time position vectors and the real-time adjacent time position vectors, determine when the satellite passes through the descending node. Based on the current instantaneous orbit element vectors, perform in-plane and out-of-plane command calculation steps to obtain the final execution time of the in-plane maneuver and the out-of-plane maneuver commands. The in-plane maneuver is executed based on the final execution time of the in-plane maneuver, and the out-of-plane maneuver is executed based on the out-of-plane maneuver command when the satellite passes through the ascending node, so as to achieve autonomous control of the sun-synchronous return orbit; The in-plane and out-of-plane command calculation steps include: Obtain historical position vectors and historical velocity vectors, and obtain average orbital element errors based on the real-time position vector, real-time velocity vector, historical position vector, and historical velocity vector; Obtain the historical node periodic error sequence, and obtain the average semi-major axis decay rate based on the historical node periodic error sequence, the current instantaneous orbital element vector, and the preset atmospheric drag model; Based on the average orbital element error, average semi-major axis decay rate, calibrated westward shift limit of descending node longitude, and preset eastward shift limit of descending node longitude, the in-plane maneuver time is predicted, the theoretical execution time of the in-plane maneuver is obtained, and the predicted number of orbits is obtained based on the theoretical execution time of the in-plane maneuver and the predicted number of orbits. The orbital inclination error is obtained based on the average orbital element error, and the semi-monthly short-period term is separated based on the orbital inclination error to obtain the semi-monthly orbital inclination error. When the inclination error of the semi-lunar orbit exceeds the preset inclination error threshold, the out-of-plane maneuver speed increment of the target is obtained, and the out-of-plane maneuver command is obtained based on the out-of-plane maneuver speed increment of the target.

2. The autonomous control method for a sun-synchronous return orbit as described in claim 1, characterized in that, The autonomous control of the sun-synchronous return orbit, achieved by performing in-plane maneuvers based on the final execution time of the in-plane maneuvers and by performing out-of-plane maneuvers based on the out-of-plane maneuver commands when the satellite passes the ascending node, includes: The system acquires real-time instantaneous navigation data from the satellite and obtains instantaneous position and velocity vectors based on the instantaneous navigation data. The instantaneous orbital element vector is obtained based on the instantaneous position vector and instantaneous velocity vector; Obtain the position vectors of adjacent instantaneous moments; When the satellite passes through the ascending node based on the instantaneous position vector and the position vectors of the instantaneous adjacent moments, and when it is determined that there is an out-of-plane maneuver command, an out-of-plane maneuver is executed based on the out-of-plane maneuver command. Based on the instantaneous orbital element vector, perform in-plane maneuvering control actions: The instantaneous orbital element vector is obtained based on the instantaneous navigation data, and the instantaneous latitude argument is obtained based on the instantaneous orbital element vector; Get the current orbital revolution count; Based on the current orbital revolutions, instantaneous latitude argument, final execution time of in-plane maneuver, and preset in-plane maneuver constraints, when the preset in-plane maneuver conditions are met, adjacent descending intersection data are obtained; The in-plane maneuver velocity increment is obtained based on the adjacent descending intersection point data and the instantaneous orbital element vector; In-plane maneuvers are performed based on the in-plane maneuver speed increment.

3. The autonomous control method for a sun-synchronous return orbit as described in claim 2, characterized in that, Before the step of performing in-plane maneuvering control based on the instantaneous orbital element vector, the method further includes: When the satellite passes through the ascending node based on the instantaneous position vector and the position vector of the instantaneous adjacent time, and when it is determined that there is no out-of-plane maneuver command, the in-plane maneuver control action is directly performed based on the instantaneous orbital element vector.

4. The autonomous control method for a sun-synchronous return orbit as described in claim 2, characterized in that, Before the step of performing in-plane maneuvering control based on the instantaneous orbital element vector, the method further includes: When the satellite does not pass through the ascending node, it is determined based on the instantaneous position vector and the position vector of the instantaneous adjacent time. In this case, the in-plane maneuvering control action is performed directly based on the instantaneous orbital element vector.

5. The autonomous control method for a sun-synchronous return orbit as described in claim 2, characterized in that, The out-of-plane maneuver includes: acquiring the target out-of-plane maneuver velocity increment corresponding to the out-of-plane maneuver command, and performing out-of-plane maneuver based on the target out-of-plane maneuver velocity increment.

6. The autonomous control method for a sun-synchronous return orbit as described in claim 2, characterized in that, The in-plane maneuvering control action based on the instantaneous orbital element vector includes: If the current orbital revolutions, instantaneous latitude argument, final execution time of in-plane maneuver, and preset in-plane maneuver constraints determine that the preset in-plane maneuver conditions are not met, then no in-plane maneuver will be performed.

7. The autonomous control method for a sun-synchronous return orbit as described in claim 2, characterized in that, The method of obtaining the in-plane maneuver velocity increment based on the adjacent descending intersection point data and the instantaneous orbital element vector includes: Based on the adjacent descending node data, obtain the adjacent average orbital element vector, adjacent descending node longitude error, adjacent node period and adjacent average semi-major axis attenuation rate. The eastward limit error value is obtained based on the adjacent average orbital element vector and the adjacent average semi-major axis attenuation rate. The average semi-major axis error change value is obtained based on the eastward shift limit error value and the period of adjacent nodes; The semi-major axis lift is obtained based on the average semi-major axis error change value, the longitude error of adjacent descending intersection points, and the preset parabolic symmetry axis offset algorithm. Then, the in-plane maneuver speed increment is obtained based on the semi-major axis lift and the instantaneous orbital element vector.

8. The autonomous control method for a sun-synchronous return orbit as described in claim 1, characterized in that, The process of acquiring historical position vectors and historical velocity vectors, and obtaining average orbital element errors based on the real-time position vector, real-time velocity vector, historical position vector, and historical velocity vector, includes: The crossing time and crossing coordinates are obtained based on the real-time position vector, real-time velocity vector, historical position vector, historical velocity vector, and linear interpolation method. Based on the crossing time and crossing coordinates, obtain the longitude of the descending node and the current orbital node period; Obtain historical descending intersection data, and based on the historical descending intersection data and a preset trapezoidal integral formula, obtain the average eccentricity vector and the average orbital inclination angle; The average orbital element vector is obtained based on the descending node longitude, current orbital node period, average eccentricity vector, and average orbital inclination. The average orbital element error is then obtained based on the average orbital element vector and the nominal orbital element vector.

9. The autonomous control method for a sun-synchronous return orbit as described in claim 1, characterized in that, The step of obtaining the historical node periodic error sequence, and obtaining the average semi-major axis decay rate based on the historical node periodic error sequence, the current instantaneous orbital element vector, and a preset atmospheric drag model, includes: The initial semi-major axis decay rate is obtained based on the current instantaneous orbital element vector and the preset atmospheric drag model; Obtain the historical node periodic error sequence; When the historical node periodic error sequence is determined to meet the preset weighted calculation conditions, a weighted least squares method is performed on the historical node periodic error sequence to obtain the node periodic error decay rate; a coupled model is constructed based on perturbation theory; and the average semi-major axis decay rate is obtained based on the node periodic error decay rate, the historical node periodic error sequence, the initial semi-major axis decay rate, and the coupled model. When the historical node periodic error sequence is determined to be inconsistent with the preset weighted calculation conditions, the initial semi-major axis attenuation rate is directly used as the average semi-major axis attenuation rate.

10. The autonomous control method for a sun-synchronous return orbit as described in claim 1, characterized in that, The method of predicting in-plane maneuver time based on the average orbital element error, average semi-major axis attenuation rate, calibrated westward shift limit of descending node longitude, and preset eastward shift limit of descending node longitude, obtaining the theoretical execution time and predicted orbital number of in-plane maneuvers, and obtaining the final execution time of in-plane maneuvers based on the theoretical execution time and predicted orbital number of in-plane maneuvers, includes: Obtain the historical longitude error of the descending node, and based on the historical longitude error of the descending node and the average orbital element error, determine when the westward shift limit has been passed, and obtain the measured average semi-major axis error. Based on the measured average semi-major axis error, the preset eastward shift limit of the descending node longitude, the average semi-major axis attenuation rate, and the preset orbital dynamics model, the estimated value of the average semi-major axis error is obtained. Obtain the historical mean semi-major axis error, historical maneuver time, and historical node period after maneuvering in adjacent planes; Based on the historical average semi-major axis error, historical maneuver time, historical node period, average semi-major axis error estimate and average semi-major axis decay rate, the in-plane maneuver theoretical execution time and the predicted orbit number corresponding to the in-plane maneuver theoretical execution time are obtained. Based on the in-plane maneuver theory execution time and the predicted orbital number, the average eccentricity error of the satellite during its last passing of the descending node within the predicted orbital number is obtained. The in-plane maneuver latitude argument is obtained based on the average eccentricity error, and the in-plane maneuver latitude argument is used as the final execution time of the in-plane maneuver.

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