Method for controlling the descending node local time of a remote sensing satellite based on orbit altitude adjustment

Abstract

The application aims to provide a remote sensing satellite descending node local time control method based on orbit height adjustment, which has simple control and high control precision. The method comprises the following steps: S1, acquiring orbit elements of a remote sensing satellite, wherein the orbit elements are instantaneous elements or two-line elements; S2, calculating a current descending node local time of the satellite based on the orbit elements; S3, calculating a drift deviation of the current descending node local time from a nominal descending node local time △T ; S4, calculating a current orbit plane rotation rate based on a J2 perturbation parameter of the earth; S5, calculating a difference between the current orbit plane rotation rate and an earth revolution rate; S6, designing an expected orbit plane rotation rate based on the difference calculated in the step S5; S7, calculating an expected orbit height according to the expected orbit plane rotation rate; and S8, adjusting the orbit height of the satellite to the expected orbit height. The application is applied to the technical field of remote sensing satellite measurement and control.

Description

Technical Field

[0001] This invention relates to the field of remote sensing satellite telemetry, tracking, and command (TT&C) technology, specifically to a method for controlling the local time of the descending node of a remote sensing satellite based on orbital altitude adjustment. This method is applicable to remote sensing satellite TT&C scenarios related to the manufacturing of spacecraft such as launch vehicles and sounding rockets, as well as aerospace-related equipment such as launch vehicle ground equipment and spacecraft ground equipment. It provides technical support for the high-quality development of the aerospace industry. Based on statistical results of the drift process of the satellite's descending node local time, the satellite's orbital altitude is adjusted to achieve control of the local time of the descending node. Background Technology

[0002] With the development of microsatellite technology, low-Earth orbit sun-synchronous orbit remote sensing satellite constellations have become the best choice for global data acquisition and observation due to their advantages such as flexible deployment, convenient launch, wide coverage area, and high cost performance. Many remote sensing satellite constellation systems have been developed and operated both domestically and internationally, including China's Gaofen series satellite constellation, the United States' Landsat series satellite constellation, the Jilin-1 satellite constellation operated by Changguang Company, and the Zhuhai-1 satellite constellation operated by Zhuhai Aerospace Microelectronics Co., Ltd.

[0003] Remote sensing satellites mostly employ a sun-synchronous orbit design. This design ensures that the local solar time at the nadir point remains constant each time the satellite passes over the same region. Due to the Earth's revolution around the sun, even with the same local solar time, ground illumination conditions vary significantly across seasons. However, over a period of time, illumination conditions can be considered roughly the same, which is beneficial for remote sensing data observation and multi-temporal comparative analysis. The descending node local time is a crucial parameter in sun-synchronous orbit design. For example, the designed local times for the descending node of the Zhuhai-1 02 and 03 remote sensing satellites are 11:00 and 13:00, respectively.

[0004] Due to orbital perturbations and orbital insertion deviations, the local time at the descending node of remote sensing satellites can drift, affecting satellite observation results. Therefore, descending node local time control is necessary. Currently, the descending node local time of remote sensing satellites is mostly controlled by ensuring accuracy at launch and then adjusting the orbital inclination after launch. For microsatellites, while adjusting the orbital inclination after launch can be done in one step, it is technically complex, consumes a lot of fuel, and significantly shortens the satellite's lifespan.

[0005] In contrast, while the local time control of the descending node of remote sensing satellites based on orbital altitude adjustment requires gradual correction, it consumes less fuel and is simpler and easier to implement.

[0006] With the rapid development of the manufacturing industry of spacecraft such as launch vehicles and sounding rockets and aerospace-related equipment, the deployment scale of low-Earth orbit sun-synchronous orbit remote sensing satellite constellations continues to expand, which puts forward higher requirements for the stability, long life and telemetry, tracking and control efficiency of satellites in orbit. Precise control of the local time of the descending node has become one of the key technologies to ensure the collaborative work of aerospace equipment and improve the effectiveness of satellite applications.

[0007] Currently, there is no research on controlling the local time of the descending node of remote sensing satellites based on orbital altitude adjustment. This invention is designed to solve this problem. Summary of the Invention

[0008] To address the shortcomings of existing technologies, this invention proposes a simple and highly accurate method for controlling the local time of the descending node of remote sensing satellites based on orbital altitude adjustment.

[0009] The technical solution adopted in this invention is a method for controlling the local time of the descending node of a remote sensing satellite based on orbital altitude adjustment. This method includes the following steps:

[0010] S1. Obtain the orbital elements of the remote sensing satellite, wherein the orbital elements are instantaneous elements or two rows of elements;

[0011] S2. Calculate the current descending node local time of the satellite based on the orbital elements, including: determining the UTC time and longitude of the satellite when it descends through the equator based on the orbital dynamics model, converting the UTC time into seconds, and obtaining the descending node local time in seconds by combining the conversion relationship between longitude and local time, and normalizing the local time to make it within the range of 0 to 86400 seconds.

[0012] S3. Calculate the drift deviation between the current local time of the descending node and the nominal local time of the descending node. △T Set a safety value for the difference. T safe Determine drift deviation △T Is the absolute value greater than the safety value of the difference? T safe If yes, proceed to step S4 to perform the local time control flow for the descending intersection; otherwise, return to step S1.

[0013] S4. Calculate the current orbital plane rotation rate based on the Earth's J2 perturbation parameters;

[0014] S5. Calculate the difference between the current orbital plane rotation rate and the Earth's revolution rate;

[0015] S6. Design the desired orbital surface rotation rate based on the difference calculated in step S5.

[0016] S7. Calculate the desired track height based on the desired track surface rotation rate;

[0017] S8. Adjust the satellite's orbital altitude to the desired orbital altitude.

[0018] Furthermore, in step S1, the instantaneous roots are obtained through the satellite telemetry and control agency, and the two rows of roots are obtained from NASA's public website.

[0019] Furthermore, step S2 specifically includes:

[0020] S21. Based on the obtained satellite orbit elements, use STK software or satellite telemetry and control software to calculate the descending node time (UTC) and corresponding longitude (LON) when the satellite descends through the Earth's equator, where the hours, minutes, and seconds of the descending node time (UTC) are converted to seconds.

[0021] S22. Convert the hour HH, minute MM, and second SS of the descending intersection time UTC to the second count SEC, then we have:

[0022] SEC = HH × 3600 + MM × 60 + SS,

[0023] Using the relationship between longitude (LON) and local time, the formula for calculating the local time (LOCALTIME) at the descending node is as follows:

[0024] LOCALTIME = SEC + LON × 4 × 60;

[0025] S23. Normalize the local time to ensure its value is within the range of 0 to 86400 seconds:

[0026] When LOCALTIME < 0, LOCALTIME can be obtained by the following formula.

[0027] LOCALTIME+86400;

[0028] When LOCALTIME > 86400, LOCALTIME can be obtained by the following formula.

[0029] LOCALTIME-86400.

[0030] Furthermore, in step S2, the local time of the current descending node of the satellite is calculated using an algorithm based on the orbital dynamics model. When the satellite orbital elements are instantaneous elements, the orbital dynamics model is J2 or J4; when the satellite orbital elements are two rows of elements, the orbital dynamics model SGP4 is used for calculation.

[0031] Furthermore, the specific steps of step S3 are as follows:

[0032] S31. Calculate the difference between the current local time at the descending node and the nominal local time at the descending node. △T The calculation formula is as follows:

[0033] △T = LOCALTIME - T des ,

[0034] in T des The local time at the nominal descending node is measured in seconds.

[0035] S32, Set the difference safety value T safe When drift deviation △T The absolute value is greater than the difference safety value T safe When the drift deviation occurs, local time control at the descending intersection point is performed, proceeding to step S4; △T The absolute value is less than the safety value of the difference. T safe If no local time control of the descending intersection is performed, return to step 1.

[0036] Furthermore, the specific steps of step S4 are as follows:

[0037] S41. Calculate the current orbital plane rotation rate: Using the Earth's oblateness perturbation parameter J2, the formula for calculating the current orbital plane rotation rate under the perturbation of the Earth's oblateness parameter J2 is as follows:

[0038] ,

[0039] In the formula, μ is the gravitational constant. R e The average equatorial radius of the Earth. a For the semi-major axis of the satellite orbit, e For orbital eccentricity, i For the track inclination angle, J 2 = 0.001082, where,

[0040] a = R e +H,

[0041] H represents the satellite's orbital altitude.

[0042] Further, in step S5, the difference Δ between the current orbital plane rotation rate and the Earth's revolution rate is calculated using the following formula. oh :

[0043] ,

[0044] If the difference in the rotational speed of the track surface is Δ ohA value greater than 0 indicates that the orbital plane rotates faster than the Earth's revolution around the Sun, and the local time at the descending node shifts forward. If the difference is Δ... oh A value less than 0 indicates that the orbital plane rotates slower than the Earth's revolution around the Sun, and the local time of the descending node is shifted backward.

[0045] Further, step S6 specifically involves: when there is a difference in the orbital plane rotation rate, causing the satellite's local time at the descending node to drift for a certain period of time, in order to drift back, the desired orbital plane rotation rate ωq is designed as follows:

[0046] ωq=0.9856-△ oh .

[0047] Further, step S7 specifically involves: calculating the desired orbital height H based on the desired orbital surface rotation rate ωq. The desired orbital height H and the desired orbital surface rotation rate are calculated using the following formula:

[0048] ,

[0049] The semi-major axis 'a' of the satellite orbit is obtained using Newton's iteration method. The desired orbital altitude 'H' is then calculated as follows:

[0050] H=a- R e ,

[0051] in, R e It is the Earth's average equatorial radius.

[0052] Furthermore, the gravitational constant μ = 3.986006e+5 km 3 / S 2 The Earth's average equatorial radius Re = 6378.14 km; the nominal descending node local time is the target descending node local time set during the satellite design phase; the remote sensing satellite is a sun-synchronous orbit satellite with a preset fixed orbital inclination.

[0053] The beneficial effects of this invention are as follows:

[0054] This invention calculates the orbital plane rotation rate using Earth's J2 perturbation parameters and corrects the local time drift at the descending node by fine-tuning the orbital altitude, instead of employing the traditional orbital inclination adjustment method. In principle, orbital inclination adjustment requires the satellite to overcome Earth's gravity and make significant changes to its orbital attitude, consuming enormous amounts of fuel. In contrast, orbital altitude adjustment only requires small ascents and descents via thrusters, resulting in far lower energy consumption. This lightweight adjustment method based on perturbation principles significantly reduces the satellite's fuel consumption during on-orbit operation, solving the core problem of limited fuel reserves for microsatellites. This, in turn, significantly extends the satellite's on-orbit service life, allowing it to continuously and stably conduct observation missions and indirectly enhancing its overall utilization value.

[0055] This invention constructs a complete closed-loop process from obtaining orbital elements to adjusting orbital altitude. In principle, it relies on mature orbital dynamics models (J2, J4, SGP4 models) and standardized calculation tools (STK software, Newton's iteration method, etc.), eliminating the need to develop complex attitude control algorithms. In terms of steps, standardized operations such as calculating the local time drift deviation at the descending node, designing the desired orbital plane rotation rate, and solving for the desired orbital altitude achieve modularization and streamlining of the control process. Compared to the complex attitude calibration and multi-system coordination involved in traditional orbital inclination adjustment, this method does not require complex hardware upgrades or sophisticated attitude control equipment. Ordinary satellite telemetry and control personnel can complete the operation using conventional telemetry and control software, lowering the technical threshold and operational difficulty, and facilitating its widespread application in various microsatellite constellations.

[0056] This invention ensures control accuracy through multiple precise calculations: first, the descending node UTC time is converted into seconds, and then the accurate descending node local time is obtained by combining the conversion relationship between longitude and local time, and normalization processing is used to ensure data validity; then, a difference safety value T is set. safe To avoid orbital instability caused by frequent adjustments, and to accurately compensate for local time drift by designing the desired orbital plane rotation rate based on the Earth's orbital speed (0.9856° / day), this refined calculation logic can stably control the local time at the descending node within a preset safe range, effectively avoiding changes in ground illumination conditions caused by local time drift. For sun-synchronous orbit remote sensing satellites, stable illumination conditions are crucial for ensuring the consistency and comparability of multi-temporal remote sensing data. This ensures that the satellite acquires high-quality data in scenarios such as weather forecasting, environmental monitoring, and surface change analysis, providing reliable support for subsequent data processing and applications.

[0057] The control method of this invention is naturally adapted to the operation and maintenance needs of satellite constellations. This method is applicable to both instantaneous roots obtained from satellite telemetry, tracking, and command (TT&C) agencies and two rows of roots obtained from NASA's public website, adapting to orbital parameters from different sources. Furthermore, its standardized control process enables automated execution, meeting the batch management needs of multiple satellites within the constellation. Compared to the limitations of traditional methods that struggle to adapt to multi-satellite collaborative control, this method eliminates the need for individual control schemes for each satellite. It allows for batch adjustments based on the unified nominal descending node local time of the satellites, significantly improving the operational efficiency of the satellite constellation and facilitating efficient and collaborative global observation.

[0058] The control method of this invention can form a technical synergy with aerospace-related equipment such as launch vehicle ground equipment and spacecraft ground equipment, adapt to the lightweight and automated needs of the spacecraft manufacturing industry for satellite telemetry, tracking, and control, and help the technological upgrading and efficient development of the satellite and application industry. Attached Figure Description

[0059] Figure 1 This is a simplified flowchart of the method of the present invention. Detailed Implementation

[0060] like Figure 1 As shown, this invention provides a method for controlling the local time of the descending node of a remote sensing satellite based on orbital altitude adjustment. The method includes the following steps:

[0061] S1. Obtain the orbital elements of the remote sensing satellite, wherein the orbital elements are instantaneous elements or two rows of elements;

[0062] S2. Calculate the current descending node local time of the satellite based on the orbital elements, including: determining the UTC time and longitude of the satellite when it descends through the equator based on the orbital dynamics model, converting the UTC time into seconds, and obtaining the descending node local time in seconds by combining the conversion relationship between longitude and local time, and normalizing the local time to make it within the range of 0 to 86400 seconds.

[0063] S3. Calculate the drift deviation between the current local time of the descending node and the nominal local time of the descending node. △T Set a safety value for the difference. T safe Determine drift deviation △T Is the absolute value greater than the safety value of the difference? T safe If yes, proceed to step S4 to perform the local time control flow for the descending intersection; otherwise, return to step S1. In this invention, the difference safety value... T safe The allowable range is ±1800s;

[0064] S4. Calculate the current orbital plane rotation rate based on the Earth's J2 perturbation parameters;

[0065] S5. Calculate the difference between the current orbital plane rotation rate and the Earth's revolution rate;

[0066] S6. Design the desired orbital surface rotation rate based on the difference calculated in step S5.

[0067] S7. Calculate the desired track height based on the desired track surface rotation rate;

[0068] S8. Adjust the satellite's orbital altitude to the desired orbital altitude.

[0069] The specific steps are as follows:

[0070] S1: Obtain satellite orbital elements

[0071] Satellite orbital elements can be either instantaneous satellite elements or two-line satellite elements. Instantaneous elements can be obtained from satellite tracking and control departments, while two-line satellite elements can be obtained from NASA's public website. The instantaneous satellite elements and two-line satellite elements can also be converted.

[0072] S2: Calculate the local time of the satellite's current descending node.

[0073] 1) Based on the obtained satellite orbital elements, use STK software or satellite telemetry and control software to calculate the longitude (LON) corresponding to the UTC time (YYYY-MM-DD HH:MM:SS) when the satellite descends through the Earth's equator. Alternatively, an algorithm can be designed and implemented based on the orbital dynamics model. If the satellite orbital elements are instantaneous elements, use the orbital dynamics model J2 or J4; if the satellite orbital elements are two rows of elements, use the orbital dynamics model SGP4 for calculation.

[0074] 2) Convert the descending node UTC time to seconds (SEC), then SEC = HH * 3600 + MM * 60 + SS. The relationship between longitude (LON) and local time is: a 1-degree difference in longitude corresponds to a 4-minute difference in local time; therefore, the formula for calculating the local time of the descending node is as follows:

[0075] LOCALTIME = SEC + LON * 4 * 60

[0076] Considering the local time range of the descending node is 0-24 hours, corresponding to a second count of 0-86400, then we have

[0077] When LOCALTIME < 0, LOCALTIME can be obtained by the following formula.

[0078] LOCALTIME+86400;

[0079] When LOCALTIME > 86400, LOCALTIME can be obtained by the following formula.

[0080] LOCALTIME-86400.

[0081] S3: Determine whether to perform descending node local time control.

[0082] 1) Calculate the difference between the current local time of the descending node and the nominal local time of the descending node. △T The calculation formula is as follows:

[0083] △T = LOCALTIME - T des，

[0084] in T desThe local time of the nominal descending node is also measured in seconds, such as 39600 for 11:00 AM.

[0085] 2) To avoid frequent local time control at the descending intersection, a safety value is designed. T safe It can generally be set to 30 minutes, i.e., 1800, when the difference... △T When the absolute value is greater than the safety value, perform local time control at the descending intersection point and proceed to step S4. When the difference... △T When the absolute value is less than the safety value, no descending intersection local time control is performed, and the process returns to step S1 and repeats.

[0086] S4: Calculate the current orbital surface rotation rate

[0087] Because the Earth is a non-uniform sphere, its mass distribution and density are also non-uniform, resulting in a non-uniform dynamic effect on the satellite's gravity. The calculation of the current orbital plane rotation rate utilizes the Earth's oblateness perturbation parameter J2. Under the J2 perturbation, the formula for calculating the current orbital plane rotation rate is:

[0088] ,

[0089] In the formula: gravitational constant μ = 3.986006e + 5 km 3 / S 2 Re is the Earth's mean equatorial radius, Re = 6378.14 km; a is the semi-major axis of the satellite orbit, a = Re + H, where H is the satellite orbital altitude; e is the orbital eccentricity; i is the orbital inclination. J 2 = 0.001082.

[0090] S5: Calculate the difference in rotational speed of the track surface

[0091] Since the Earth's angular velocity around the Sun is 0.9856° / day, sun synchronization requires the satellite's orbital plane to precess around its central axis in the same direction as the Earth's revolution, also by 0.9856° / day. Therefore, the difference in orbital plane rotation rate is Δ. oh for

[0092] ,

[0093] The difference in the orbital plane rotation rate will cause a shift in the local time at the descending node of the satellite. If the difference is greater than 0, it means that the orbital plane rotates faster than the Earth's revolution around the Sun, and the local time at the descending node will shift forward. If the difference is less than 0, it means that the orbital plane rotates slower than the Earth's revolution around the Sun, and the local time at the descending node will shift backward.

[0094] S6: Desired rotation rate of the orbital surface

[0095] The current discrepancy in the orbital plane rotation rate has caused the satellite's descending node local time to drift for a certain period of time. To enable it to drift back, the desired orbital plane rotation rate is designed as follows:

[0096] ωq=0.9856-△ Oh.

[0097] S7: Calculate the desired orbital altitude

[0098] The desired orbit height is calculated based on the desired orbital surface rotation rate. The desired orbit height is related to the desired orbital surface rotation rate by the following formula:

[0099] ,

[0100] The semi-major axis *a* is obtained by using Newton's iteration method or other numerical methods. The desired orbital height is then calculated as follows:

[0101] H = a - Re.

[0102] Here, Newton's iteration method is used for calculation and solution. When using Newton's iteration method, the iteration termination condition is that the satellite's semi-major axis obtained after two iterations meets the semi-major axis accuracy requirement epsilon, which can be set to the satellite positioning accuracy, such as 0.01km. That is, |an+1-an| <epsilon。

[0103] S8: Perform orbital altitude adjustment: Adjust the satellite's orbital altitude to the desired orbital altitude.

[0104] The present invention will now be further illustrated with specific examples.

[0105] S1: Obtain the two rows of satellite roots (such as TLE data corresponding to the NORAD number) or instantaneous roots from NASA's public platform or telemetry and control department.

[0106] S2: Use the SGP4 model to calculate the descending node UTC time (e.g., 2023-10-01 10:30:00) and longitude (e.g., 120.5°). Convert the time to seconds (SEC = 10 × 3600 + 30 × 60 = 37800), and calculate the local time (LOCALTIME) = 37800 + 120.5 × 4 × 60 = 66690 seconds. Normalization: If the result exceeds 86400 seconds, subtract 86400.

[0107] S3: Nominal Local Time (T) des =11:00 (39600 seconds), the difference ΔT = 66690 - 39600 = 27090 seconds > the safety value of 1800 seconds, and the system enters control.

[0108] S4: Assume the satellite's orbital altitude H = 500 km, semi-major axis a = 6378.14 + 500 = 6878.14 km, eccentricity e ≈ 0, and inclination i = 97.4°. Substitute these values ​​into the formula to calculate the rotation rate. .

[0109] S5: Calculate the difference Δω= -0.9856, if Δω=0.5° / day>0, it indicates that the local time has drifted forward.

[0110] S6: Design ωq=0.9856-0.5=0.4856° / day to achieve reverse rate compensation.

[0111] S7: Solving the ωq formula using Newton's iteration method, we obtain the expected semi-major axis a≈6900km and the expected height H=6900-6378.14≈521.86km.

[0112] S8: Control the thrusters to increase the altitude from 500km to 521.86km, and after adjustment, verify that the local time drift has been reduced to a safe range.

[0113] This invention achieves precise control through highly fine-tuning, and the entire process can be automated by standard measurement and control software.

[0114] Finally, it should be emphasized that the above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, the present invention can have various changes and modifications. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A method for controlling the local time of the descending node of a remote sensing satellite based on orbital altitude adjustment, characterized in that, The method includes the following steps: S1. Obtain the orbital elements of the remote sensing satellite, wherein the orbital elements are instantaneous elements or two rows of elements; S2. Calculate the current descending node local time of the satellite based on the orbital elements, including: determining the UTC time and longitude of the satellite when it descends through the equator based on the orbital dynamics model, converting the UTC time into seconds, and obtaining the descending node local time in seconds by combining the conversion relationship between longitude and local time, and normalizing the local time to make it within the range of 0 to 86400 seconds. S3. Calculate the drift deviation ΔT between the current descending node local time and the nominal descending node local time, and set a safety value T for the difference. safe Determine whether the absolute value of the drift deviation ΔT is greater than the safety value T of the difference. safe If yes, proceed to step S4 to perform the local time control flow for the descending intersection; otherwise, return to step S1. S4. Calculate the current orbital plane rotation rate based on the Earth's J2 perturbation parameters; S5. Calculate the difference between the current orbital plane rotation rate and the Earth's revolution rate; S6. Based on the difference calculated in step S5, design the desired orbital surface rotation rate. S7. Calculate the desired track height based on the desired track surface rotation rate; S8. Adjust the satellite orbital altitude to the desired orbital altitude.

2. The method according to claim 1, characterized in that, In step S1, the instantaneous roots are obtained through the satellite telemetry and control agency, and the two rows of roots are obtained from NASA's public website.

3. The method according to claim 1, characterized in that, Step S2 is as follows: S21. Based on the obtained satellite orbit elements, use STK software or satellite telemetry and control software to calculate the descending node time (UTC) and corresponding longitude (LON) when the satellite descends through the Earth's equator, where the hours, minutes, and seconds of the descending node time (UTC) are converted to seconds. S22. Convert the hour HH, minute MM, and second SS of the descending intersection time UTC to the second count SEC, then we have: SEC = HH × 3600 + MM × 60 + SS, Using the relationship between longitude (LON) and local time, the formula for calculating the local time (LOCALTIME) at the descending node is as follows: LOCALTIME = SEC + LON × 4 × 60; S23. Normalize the local time to ensure its value is within the range of 0 to 86400 seconds: When LOCALTIME < 0, LOCALTIME is obtained by the following formula. LOCALTIME+86400; When LOCALTIME > 86400, LOCALTIME is obtained by the following formula. LOCALTIME-86400.

4. The method according to claim 3, characterized in that, In step S2, the local time of the current descending node of the satellite is calculated. The algorithm is implemented based on the orbital dynamics model. When the satellite orbital elements are instantaneous elements, the orbital dynamics model is J2 or J4; when the satellite orbital elements are two rows of elements, the orbital dynamics model SGP4 is used for calculation.

5. The method according to claim 1, characterized in that, The specific steps of step S3 are as follows: S31. Calculate the difference ΔT between the current local time of the descending node and the nominal local time of the descending node. The calculation formula is as follows: △T=LOCALTIME-T des , Where T des The local time at the nominal descending node is measured in seconds. S32, Set the difference safety value T safe When the absolute value of the drift deviation ΔT is greater than the safety value T of the difference safe When the local time of the descending intersection is controlled, proceed to step S4; when the absolute value of the drift deviation ΔT is less than the safety value T of the difference. safe If no local time control of the descending intersection is performed, return to step 1.

6. The method according to claim 1, characterized in that, The specific steps of step S4 are as follows: S41. Calculate the current orbital plane rotation rate: Using the Earth's oblateness perturbation parameter J2, the formula for calculating the current orbital plane rotation rate under the perturbation of the Earth's oblateness parameter J2 is as follows: ， In the formula, μ is the gravitational constant, and R e Let be the Earth's mean equatorial radius, a be the semi-major axis of the satellite orbit, e be the orbital eccentricity, i be the orbital inclination, and J2 = 0.001082. a=R e +H, H represents the satellite's orbital altitude.

7. The method according to claim 6, characterized in that, In step S5, the difference Δω between the current orbital plane rotation rate and the Earth's revolution rate is calculated using the following formula: ， If the difference in the orbital plane rotation rate Δω is greater than 0, it means that the orbital plane rotates faster than the Earth revolves around the Sun, and the local time of the descending node shifts forward. If the difference Δω is less than 0, it means that the orbital plane rotates slower than the Earth revolves around the Sun, and the local time of the descending node shifts backward.

8. The method according to claim 7, characterized in that, Step S6 specifically involves: When there is a difference in the orbital plane rotation rate, causing the satellite's local time at the descending node to drift for a certain period of time, in order to drift back, the desired orbital plane rotation rate ωq is designed as follows: ωq=0.9856-△ω.

9. The method according to claim 8, characterized in that, Step S7 specifically involves: calculating the desired orbital height H based on the desired orbital surface rotation rate ωq. The desired orbital height H and the desired orbital surface rotation rate are calculated using the following formula: ， The semi-major axis 'a' of the satellite orbit is obtained using Newton's iteration method. The desired orbital altitude 'H' is then calculated as follows: H=a-R e , Among them, R e It is the Earth's average equatorial radius.

10. The method according to claim 9, characterized in that, The gravitational constant μ = 3.986006e+5 km 3 / S 2 The Earth's average equatorial radius Re = 6378.14 km; the nominal descending node local time is the target descending node local time set during the satellite design phase; the remote sensing satellite is a sun-synchronous orbit satellite with a preset fixed orbital inclination.

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