Method for avoiding signal interference based on geostationary orbit satellite

Abstract

A satellite signal interference avoidance method based on a geostationary orbit shortens the duration of a task satellite in a detectable or interference space range of a detection satellite through the design of a stationary orbit inclination angle, and on the basis, the design of small orbit eccentricity is introduced for an inclined orbit, so that the satellite signal interference can be avoided. The encounter duration of the task satellite and the detection satellite in one orbit is further greatly reduced, the risk of detection and interference by the stationary orbit satellite is further reduced, and the satellite safety is improved.

Description

Technical Field

[0001] The invention relates to an avoidance method based on geostationary orbit satellite signal interference, and belongs to the field of geosynchronous orbit remote sensing technology design. Background Art

[0002] The orbital altitude of geosynchronous orbit (GEO) satellites is as high as about 36,000 km. A single satellite can cover nearly one-third of the earth's surface, which is much greater than the coverage rate of a single satellite in medium and low orbits. It has the ability to respond quickly to key events and observe small and medium-sized targets in real time, and is particularly suitable for rapid monitoring of disaster factors, disaster scope and risk assessment. However, there are many signal collection satellites in geosynchronous orbit, which detect and interfere with satellites performing observation missions, preventing them from receiving and sending signals, which poses a certain threat to satellite safety. In addition, with the increase in the number of geostationary orbit satellites and the limited orbital resources, satellite communication frequency resources cannot be guaranteed, and satellite communications are subject to mutual interference, which hinders the normal operation of mission satellites.

[0003] In order to avoid the threats posed by signal detection satellites and other interference signals in geostationary orbits, mission satellites usually use complex algorithms and methods to optimize communication signals and eliminate strong interference signals, such as filtering interference signals in the time domain or controlling antennas in the airspace. In addition, they will use maneuvers to move away from detection satellites or interference satellites. However, the interference signal elimination algorithm is complex, the amount of data is large, and the on-board computing power is very high. It is necessary to develop new dimensionality reduction algorithms or improve the performance of the processing board to achieve this. It is difficult to get rid of tracking by maneuvering to avoid interference sources, and it increases the energy consumption of mission satellites and shortens their lifespan. In addition, when the mission satellite and the detection satellite are in the same geostationary orbit, the detection satellite is often in the blind spot of the mission satellite payload's field of view for collecting data against the sun's light, making it impossible for the mission satellite to detect and avoid the detection satellite. Summary of the invention

[0004] The technical problem solved by the present invention is: in view of the risk problem that mission satellites are easily detected and interfered by geostationary orbit satellites in the current prior art, a method for avoiding interference based on geostationary orbit satellite signals is proposed.

[0005] The present invention solves the above technical problems by the following technical solutions:

[0006] A method for avoiding interference based on geostationary orbit satellite signals, comprising:

[0007] Establish the orbital motion equations of detection satellites and mission satellites;

[0008] Calculate the designed orbital inclination of the geosynchronous orbit according to the detectable space range of the detection satellite;

[0009] Taking the designed orbital inclination of the geosynchronous orbit as a constraint condition, presetting a non-zero initial orbital eccentricity, and using the exploration satellite coordinates obtained by solving the orbital motion equations of the exploration satellite and the mission satellite coordinates obtained by solving the orbital motion equations of the mission satellite, the separation distance of the ascending node and the descending node along the geostationary orbit is calculated;

[0010] The distance criterion is formulated using the detectable space range of the detection satellite, and the separation distance of the ascending node and the descending node along the geostationary orbit is judged. According to the judgment result, the eccentricity of the geosynchronous satellite orbit at the current moment is iteratively optimized until the distance criterion is met;

[0011] The eccentricity of the geosynchronous satellite orbit that meets the distance criterion is output as the orbit optimization design value.

[0012] The orbital motion equation of the detection satellite is:

[0013]

[0014] In the formula, for the geostationary orbit, R is the radius, Ω is the longitude of the ascending node, ω is the angular velocity of the earth's rotation, and x, y, and z are the coordinates of the detection satellite;

[0015] The orbital motion equation of the mission satellite is:

[0016]

[0017] Where x', y', z' are the mission satellite coordinates, and α is the orbital inclination.

[0018] The detectable space range of the detection satellite is A, which is determined according to the detection capability of the detection satellite;

[0019] The calculation method of the geosynchronous orbit design orbit inclination α is:

[0020]

[0021] Where α is the designed minimum orbital inclination value. If the actual observation mission requires adjustment of the detectable space range of the detection satellite, the designed orbital inclination α of the geosynchronous orbit is regularly increased.

[0022] The calculation method of the separation distance between the ascending node and the descending node along the geostationary orbit is:

[0023] Determine the six orbital elements of the mission satellite, calculate the satellite vector in the Earth-fixed coordinate system, and convert it to the Earth's inertial coordinate system;

[0024] Calculate the area swept by the mission satellite from ascending node to descending node and calculate the total area of ​​the elliptical orbit;

[0025] The separation distance between the ascending node and the descending node along the geostationary orbit is calculated based on the area swept by the satellite in the mission from the ascending node to the descending node and the total area of ​​the elliptical orbit.

[0026] The satellite vector is calculated in the earth fixed coordinate system The method is:

[0027]

[0028] Calculation of satellite vector in the Earth's inertial coordinate system The method is:

[0029]

[0030] Where α is the designed orbital inclination of the geosynchronous orbit, β is the true anomaly angle, R is the semi-major axis, w is the periapsis angle, Ω is the longitude of the ascending node, e is the preset non-zero initial orbit eccentricity, r is the distance from the satellite to the center of the earth,

[0031] The area S swept by the satellite in the mission from the ascending node to the descending node t The calculation method is:

[0032]

[0033] In the formula, (x 0 ,y 0 ) are the coordinate positions of the ascending node and the descending node in the inertial system;

[0034] The total area of ​​an elliptical orbit is calculated as:

[0035] S e =π(1-e)R 2

[0036] In the formula, S e is the total area of ​​the elliptical orbit.

[0037] The calculation method of the separation distance between the ascending node and the descending node along the geostationary orbit is:

[0038]

[0039] Where L is the separation distance between the ascending node and the descending node along the geostationary orbit, and R is the orbit radius.

[0040] The criteria formulated by the detectable space range of the detection satellite are specifically:

[0041] Compare the detectable space range A of the detection satellite and the separation distance L between the ascending and descending nodes along the geostationary orbit. If L < A, the iterative optimization direction of the eccentricity of the geosynchronous satellite orbit at the current moment is to increase; if L ≥ A, the iterative optimization direction of the eccentricity of the geosynchronous satellite orbit at the current moment is to decrease.

[0042] When performing iterative optimization to increase the eccentricity of the geosynchronous satellite orbit, use the bisection method to determine the range values (e', e), select the maximum value e1, the minimum value e2, and take the median value e3. Update and calculate the separation distance L between the ascending and descending nodes along the geostationary orbit according to the median value e3, and use the criterion |L - A| < ε for judgment. If the criterion condition is satisfied, output the current eccentricity e3 of the geosynchronous satellite orbit as the optimized design value;

[0043] Among them, ε is the distance difference threshold.

[0044] If the calculated separation distance L between the ascending and descending nodes along the geostationary orbit after update still does not satisfy the criterion |L - A| < ε, then re - use the distance criterion for judgment. If the distance criterion is not satisfied, update the eccentricity e3 of the geosynchronous satellite orbit to e2, recalculate the median value of the updated e2 and e1, and use the next median value to update the separation distance L between the ascending and descending nodes along the geostationary orbit until the criterion |L - A| < ε is satisfied;

[0045] Otherwise, update the eccentricity e3 of the geosynchronous satellite orbit to e1, recalculate the median value of the updated e2 and e1, and use the next median value to update the separation distance L between the ascending and descending nodes along the geostationary orbit until the criterion |L - A| < ε is satisfied.

[0046] The advantages of the present invention compared with the prior art are as follows:

[0047] (1) A method for avoiding interference from signals of geostationary orbit satellites provided by the present invention realizes the avoidance of threats to satellite safety and observation tasks caused by interference from detection satellites on the geostationary orbit through orbital inclination design, greatly reducing the duration and probability of the satellite being within the detectable or interference space range of the detection satellite, and ensuring the application stability and robustness of geosynchronous orbit satellites;

[0048] (2) Based on the overall orbital design, the present invention ingeniously solves the problem of communication interference between geostationary orbit satellites from a new dimension other than algorithm and hardware improvement, avoiding the technical problems of optimizing communication signals and eliminating strong interference signals through complex algorithms and means, and solving the problem of the resource capacity of geostationary orbit satellites; at the same time, through orbital design, it can effectively guide the orbital selection and on - orbit use planning of high - value geosynchronous orbit satellites. Brief Description of the Drawings

[0049] Figure 1A schematic diagram of the relative speed of a geosynchronous orbit satellite and a geostationary orbit satellite without eccentricity and with inclination provided by the present invention;

[0050] Figure 2 A schematic diagram of the sub-satellite point trajectory of a geosynchronous orbit without eccentricity and with inclination provided by the present invention;

[0051] Figure 3 A schematic diagram of the relative relationship between the geosynchronous orbit and geostationary orbit satellites with eccentricity and inclination provided by the present invention;

[0052] Figure 4 A schematic diagram of a sub-satellite point trajectory of a geosynchronous orbit with eccentricity and inclination provided by the present invention;

[0053] Figure 5 A flow chart of track avoidance design provided by the present invention. DETAILED DESCRIPTION

[0054] A method for avoiding interference with geostationary orbit satellite signals. By designing the inclination of the geostationary orbit, the time that the mission satellite is in the space detectable or interfered with by the detection satellite is shortened. On this basis, a small orbit eccentricity design is introduced for the inclined orbit to further significantly reduce the encounter time between the mission satellite and the detection satellite in one orbit, thereby reducing the risk of detection and interference by geostationary orbit satellites and improving satellite safety.

[0055] The method for avoiding interference of geostationary orbit satellite signals includes the following steps:

[0056] Establish the orbital motion equations of detection satellites and mission satellites;

[0057] Calculate the designed orbital inclination of the geosynchronous orbit according to the detectable space range of the detection satellite;

[0058] Taking the designed orbital inclination of the geosynchronous orbit as a constraint condition, presetting a non-zero initial orbital eccentricity, and using the exploration satellite coordinates obtained by solving the orbital motion equations of the exploration satellite and the mission satellite coordinates obtained by solving the orbital motion equations of the mission satellite, the separation distance of the ascending node and the descending node along the geostationary orbit is calculated;

[0059] The distance criterion is formulated using the detectable space range of the detection satellite, and the separation distance of the ascending node and the descending node along the geostationary orbit is judged. According to the judgment result, the eccentricity of the geosynchronous satellite orbit at the current moment is iteratively optimized until the distance criterion is met;

[0060] The eccentricity of the geosynchronous satellite orbit that meets the distance criterion is output as the optimized design value.

[0061] The orbital motion equation of the detection satellite is:

[0062]

[0063] In the formula, for the geostationary orbit, R is the radius, Ω is the longitude of the ascending node, ω is the angular velocity of the earth's rotation, and x, y, and z are the coordinates of the detection satellite;

[0064] The orbital motion equation of the mission satellite is:

[0065]

[0066] Where x', y', z' are the mission satellite coordinates.

[0067] The detectable space range of the detection satellite is A, which is determined according to the detection capability of the detection satellite;

[0068] The calculation method of the geosynchronous orbit design orbit inclination α is:

[0069]

[0070] Where α is the designed minimum orbital inclination value. If the actual observation mission requires adjustment of the detectable space range of the detection satellite, the designed orbital inclination α of the geosynchronous orbit is regularly increased.

[0071] The separation distance between the ascending and descending nodes along the geostationary orbit is calculated as:

[0072] Determine the six orbital elements of the mission satellite, calculate the satellite vector in the Earth-fixed coordinate system, and convert it to the Earth's inertial coordinate system;

[0073] Calculate the area swept by the satellite in the mission from ascending node to descending node and calculate the total area of ​​the elliptical orbit;

[0074] The separation distance between the ascending node and the descending node along the geostationary orbit is calculated based on the area swept by the satellite in the mission from the ascending node to the descending node and the total area of ​​the elliptical orbit.

[0075] Calculating satellite vectors in the Earth-fixed coordinate system The method is:

[0076]

[0077] Calculation of satellite vector in the Earth's inertial coordinate system The method is:

[0078]

[0079] Where α is the designed orbital inclination of the geosynchronous orbit, β is the true anomaly angle, R is the semi-major axis, w is the periapsis angle, Ω is the longitude of the ascending node, e is the preset non-zero initial orbit eccentricity, r is the distance from the satellite to the center of the earth,

[0080] The area S swept by the mission satellite from the ascending node to the descending node t The calculation method is as follows:

[0081]

[0082] In the formula, (x 0 , y 0 ) are the coordinate positions of the ascending node and the descending node in the inertial system;

[0083] The calculation method for the total area of the elliptical orbit is as follows:

[0084] S e = π(1 - e)R 2

[0085] In the formula, S e is the total area of the elliptical orbit.

[0086] The calculation method for the separation distance between the ascending node and the descending node along the geostationary orbit is as follows:

[0087]

[0088] In the formula, L is the separation distance between the ascending node and the descending node along the geostationary orbit, and R is the orbital radius.

[0089] The criterion formulated using the detectable space range of the detection satellite is specifically as follows:

[0090] Compare the detectable space range A of the detection satellite and the separation distance L between the ascending node and the descending node along the geostationary orbit. If L < A, the iterative optimization direction of the eccentricity of the geosynchronous satellite orbit at the current moment is to increase; if L ≥ A, the iterative optimization direction of the eccentricity of the geosynchronous satellite orbit at the current moment is to decrease.

[0091] When performing iterative optimization to increase the eccentricity of the geosynchronous satellite orbit, the bisection method is used to determine the range values (e', e), the maximum value e1, the minimum value e2 are selected, the median value e3 is taken, and the separation distance L between the ascending node and the descending node along the geostationary orbit is updated and calculated according to the median value e3. The criterion |L - A| < ε is used for judgment. If the criterion condition is satisfied, the current eccentricity w3 of the geosynchronous satellite orbit is output as the optimized design value;

[0092] Among them, ε is the distance difference threshold.

[0093] If the calculated separation distance L of the ascending node and the descending node along the geostationary orbit after the update still does not satisfy the criterion condition of |LA|<ε, the distance criterion is used again for judgment. If the distance criterion is not satisfied, the eccentricity of the geosynchronous satellite orbit e3 is updated to e2, and the median of the updated median e2 and e1 is recalculated. The separation distance L of the ascending node and the descending node along the geostationary orbit is updated using the next median until the criterion |LA|<ε is satisfied;

[0094] Otherwise, the eccentricity e3 of the geosynchronous satellite orbit is updated to e1, the median of the updated median e2 and e1 is recalculated, and the next median is used to update the separation distance L of the ascending node and the descending node along the geostationary orbit until the criterion |LA|<ε is met.

[0095] The following is further described in conjunction with the accompanying drawings and preferred embodiments of the specification:

[0096] In the current embodiment, the method of using orbit design to avoid geostationary orbit satellite detection and signal interference includes the following specific steps: Figure 5 As shown, including:

[0097] (1) When the satellite orbital parameters are semi-major axis R = 42164 km and the Earth's rotation angular velocity ω = 7.27 × 10 -5 rad / s, the longitude of the ascending node is Ω, the eccentricity is e, the radius of the space range that the detection satellite can detect or interfere with near the mission area is A, and the time for the mission satellite to suspend signal transmission or resist interference is t 0 Based on this, the minimum orbital inclination angle α of the mission satellite geosynchronous orbit design is obtained min If there is no special requirement, α=α min ;

[0098] (2) Under the conditions of the geosynchronous satellite orbit parameters (including the inclination angle α) of step (1), given a non-zero initial orbit eccentricity e, obtain the separation distance L between the ascending node and the descending node along the geostationary orbit;

[0099] (3) Based on L obtained in step (2), combined with the detectable space range A of the detection satellite, the eccentricity e of the geosynchronous satellite orbit is iteratively adjusted and determined to reduce the encounter time between the mission satellite and the detection satellite in one orbit, or to force the detection satellite to maneuver and increase energy consumption.

[0100] The specific calculation process is:

[0101] (1) Satellite orbit radius R = 42164 km, ascending node longitude Ω = 0, eccentricity e = 0, Earth rotation angular velocity ω = 7.27 × 10 -5 rad / s, combined with the satellite's acceptable detection time t 0=5min and the detectable space range of the detection satellite, such as A=100km, the orbital inclination of the mission satellite is calculated as shown below.

[0102]

[0103] Among them, on the equatorial plane with an inclination of 0°, for the geostationary orbit, the radius is R, the longitude of the ascending node is Ω, the angular velocity of the earth's rotation is ω, and the motion parameter equation of the detection satellite in orbit is

[0104]

[0105] If the mission satellite and the exploration satellite will coincide in the equatorial plane, the distance between the mission satellite and the exploration satellite is mainly the distance component perpendicular to the equatorial plane. Under this condition, for each true anomaly, the relative distance between the mission satellite and the exploration satellite is the shortest, which can be used as a critical condition to calculate the orbital inclination. In this case, for a circular geosynchronous orbit with an orbital inclination α, the mission satellite motion parameter equation is

[0106]

[0107] Taking the geostationary orbit detection satellite as the reference system, the coordinates (x', y', z') of the mission satellite are

[0108]

[0109] The corresponding speed is Figure 1 As shown, the size is

[0110]

[0111] Based on the time t that the satellite can accept being detected or interfered 0 , if t=t 0 / 2, the mission satellite is in the equatorial plane, and the mission satellite is in the detection range of the detection satellite for the longest time, so the detection satellite t 0 The distance from the mission satellite is L'.

[0112]

[0113] The sub-satellite points of the detection satellite and mission satellite are as follows: Figure 2 As shown in the figure, the white dot is the subsatellite path of the detection satellite, and also the subsatellite position of the mission satellite ascending node (descending node). Combined with the detectable space range A of the detection satellite, the design orbital inclination α of the geosynchronous orbit is obtained by reverse calculation based on the constraint condition of L'≥A;

[0114] α is the designed minimum orbital inclination value. If the actual observation mission requires the satellite to adjust the observation range, the orbital inclination can be increased based on α.

[0115] (2) The six orbital parameters of the mission satellite are the semi-major axis R, the periapsis angle w, the ascending node longitude Ω, the eccentricity e, the orbital inclination α, and the true anomaly β. In the Earth-fixed coordinate system, the satellite vector is:

[0116]

[0117] Among them, r is the distance from the satellite to the center of the earth,

[0118] For the Earth's inertial coordinate system, the satellite's position vector is:

[0119]

[0120] Substitute the semi-major axis R = 42164 km, pericenter angle w = 60°, ascending node longitude Ω = 0, eccentricity e = 0.0005, and orbital inclination α = 0.1375° to obtain the position vector of the mission satellite:

[0121]

[0122] By z 0 =0, we get β A =-90°, β B =90°, and then obtain the coordinates of the ascending node and the descending node in the inertial system: (x 0 ,y 0 ) A =(36514014.87,-21081994.73)m, (x 0 ,y 0 ) B =(-36514014.87, 78639755.577) m, calculate the area swept from the ascending node to the descending node, which is

[0123]

[0124] The total area of ​​the elliptical orbit is

[0125] S e =π(1-e)R 2 =5.5823×10 15 m 2 (14)

[0126] The elapsed time can be obtained based on Kepler's second law from the ratio of the swept area to the entire ellipse area: Then we get the subsatellite distance between the ascending node and the descending node M = 2π|0.5-StSeRE = 844.39m (RE is the radius of the earth), and the separation distance between the ascending node and the descending node along the geostationary orbit is

[0127] Among them, when taking , corresponding to (x 0 , y 0 ), A , (x 0 , y 0 ), B , its sub-satellite point track is as Figure 4 shown. The ascending node and the descending node are not the same point on the ground, which can avoid two encounters within one orbit with a detection satellite whose sub-satellite point track is a single point. The time experienced can be obtained based on Kepler's second law, from the ratio of the swept area to the area of the entire ellipse, (T E is the Earth's rotation period), and then the separation distance of the ascending node and the descending node along the geostationary orbit is obtained

[0128] (3) Based on step (2), comparing the radius A of the detectable space range of the detection satellite in practice, such as A = 100 km, it can be known that L < A, then the eccentricity needs to be increased. Taking e = 0.005 and repeating the calculation process of step (2), we get:

[0129]

[0130] Then Furthermore, the sub-satellite point distance M of the ascending node and the descending node is obtained as M = 2π|0.5 - StSeRE| = 37444 m, (RE is the Earth's radius), and the separation distance of the ascending node and the descending node along the geostationary orbit is Comparing with the radius A of the detectable space range of the detection satellite, L >> A. Setting the threshold ε of |L - A| not exceeding 3 km, using the bisection method to assign a value to e in the range of 0.0005 to 0.005, and based on the change value of L, iteratively adjusting the design of e. After 5 iterations, when the final eccentricity e = 0.0023, The orbital parameter inclination α and eccentricity e of the mission satellite are designed accordingly.

[0131] Among them, in order to reduce the number of encounters between the mission satellite and the detection satellite within one orbit, under the design conditions of the geosynchronous satellite orbit parameters (including the inclination α) in step (2), given a non-zero initial orbital eccentricity e, the relative position relationship between the mission satellite and the detection satellite is as Figure 3 shown.

[0132] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

[0133] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

Claims

1. A method for avoiding interference with geostationary satellite signals, characterized in that include: Establish the orbital motion equations of detection satellites and mission satellites; Calculate the designed orbital inclination of the geosynchronous orbit according to the detectable space range of the detection satellite; Taking the designed orbital inclination of the geosynchronous orbit as a constraint condition, presetting a non-zero initial orbital eccentricity, and using the exploration satellite coordinates obtained by solving the orbital motion equations of the exploration satellite and the mission satellite coordinates obtained by solving the orbital motion equations of the mission satellite, the separation distance of the ascending node and the descending node along the geostationary orbit is calculated; The distance criterion is formulated using the detectable space range of the detection satellite, and the separation distance of the ascending node and the descending node along the geostationary orbit is judged. According to the judgment result, the eccentricity of the geosynchronous satellite orbit at the current moment is iteratively optimized until the distance criterion is met; The eccentricity of the geosynchronous satellite orbit that meets the distance criterion is output as the orbit optimization design value.

2. The method for avoiding interference of geostationary satellite signals according to claim 1, characterized in that: The orbital motion equation of the detection satellite is: In the formula, for the geostationary orbit, R is the radius, Ω is the longitude of the ascending node, ω is the angular velocity of the earth's rotation, and x, y, and z are the coordinates of the detection satellite; The orbital motion equation of the mission satellite is: Where x', y', z' are the mission satellite coordinates, and α is the orbital inclination.

3. The method for avoiding interference of geostationary satellite signals according to claim 1, characterized in that: The detectable space range of the detection satellite is A, which is determined according to the detection capability of the detection satellite; The calculation method of the geosynchronous orbit design orbit inclination α is: Where α is the designed minimum orbital inclination value. If the actual observation mission requires adjustment of the detectable space range of the detection satellite, the designed orbital inclination α of the geosynchronous orbit is regularly increased.

4. The method for avoiding interference of geostationary satellite signals according to claim 1, characterized in that: The calculation method of the separation distance between the ascending node and the descending node along the geostationary orbit is: Determine the six orbital elements of the mission satellite, calculate the satellite vector in the Earth-fixed coordinate system, and convert it to the Earth's inertial coordinate system; Calculate the area swept by the mission satellite from ascending node to descending node and calculate the total area of ​​the elliptical orbit; The separation distance between the ascending node and the descending node along the geostationary orbit is calculated based on the area swept by the satellite in the mission from the ascending node to the descending node and the total area of ​​the elliptical orbit.

5. The method for avoiding interference of geostationary satellite signals according to claim 4, characterized in that: The satellite vector is calculated in the earth fixed coordinate system The method is: Calculation of satellite vector in the Earth's inertial coordinate system The method is: Where α is the designed orbital inclination of the geosynchronous orbit, β is the true anomaly angle, R is the semi-major axis, w is the periapsis angle, Ω is the longitude of the ascending node, e is the preset non-zero initial orbit eccentricity, r is the distance from the satellite to the center of the earth, 6. The method for avoiding interference of geostationary satellite signals according to claim 4, characterized in that: The area S swept by the satellite in the mission from the ascending node to the descending node t The calculation method is: Where (x0, y0) is the coordinate position of the ascending node and the descending node in the inertial system; The total area of ​​an elliptical orbit is calculated as: Se=π(1-e)R In the formula, S e is the total area of ​​the elliptical orbit.

7. The method for avoiding interference of geostationary satellite signals according to claim 4, characterized in that: The calculation method of the separation distance between the ascending node and the descending node along the geostationary orbit is: Where L is the separation distance between the ascending node and the descending node along the geostationary orbit, and R is the orbit radius.

8. A method for avoiding interference of geostationary orbit satellite signals according to claim 7, characterized in that: The criterion formulated by the detectable space range of the detection satellite is specifically: Compare the detectable space range A of the detection satellite and the separation distance L between the ascending node and the descending node along the geostationary orbit. If L < A, the iterative optimization direction of the geosynchronous satellite orbit eccentricity at the current moment is to increase; if L ≥ A, the iterative optimization direction of the geosynchronous satellite orbit eccentricity at the current moment is to decrease.

9. A method for avoiding interference of geostationary orbit satellite signals according to claim 8, characterized in that: When performing iterative optimization to increase the eccentricity of the geosynchronous satellite orbit, the bisection method is used to determine the range values (e‘, e), the maximum value e1 and the minimum value e2 are selected, and the median value e3 is taken. The separation distance L between the ascending node and the descending node along the geostationary orbit is updated and calculated according to the median value e3, and the criterion |L - A| < ε is used for judgment. If the criterion condition is satisfied, the current geosynchronous satellite orbit eccentricity e3 is output as the optimized design value; where ε is the distance difference threshold.

10. A method for avoiding interference of geostationary orbit satellite signals according to claim 9, characterized in that: If the separation distance L between the updated ascending node and descending node along the geostationary orbit still does not satisfy the criterion |L - A| < ε, the distance criterion is used for judgment again. If the distance criterion is not satisfied, the geosynchronous satellite orbit eccentricity e3 is updated to e2, the median value of the updated e2 and e1 is recalculated, and the separation distance L between the ascending node and the descending node along the geostationary orbit is updated using the next median value until the criterion |L - A| < ε is satisfied; otherwise, the geosynchronous satellite orbit eccentricity e3 is updated to e1, the median value of the updated e2 and e1 is recalculated, and the separation distance L between the ascending node and the descending node along the geostationary orbit is updated using the next median value until the criterion |L - A| < ε is satisfied.

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