An autonomous path planning method for close-range observation missions of small spacecraft

Abstract

The present invention provides an autonomous path planning method for a microsatellite in a close-range observation mission, which is characterized by including: Step 1, setting the close distance and observation duration relative to the observation target according to the observation ability and safety requirements of the microsatellite; Step 2, setting the distribution spherical surface of the close-range observation points according to the close distance, modeling the distribution spherical surface by using the spherical grid method, and encoding the grid by using the polar coordinate method; Step 9, fitting the angles swept by adjacent grids on the spherical surface by using a polynomial, so that the microsatellite smoothly transitions within each grid, and connecting the trajectories between adjacent grids to obtain a relatively smooth transfer path during the close-range observation process of the microsatellite. The present invention combines the advantages of continuous optimization methods and discrete search methods, solves the problems that traditional trajectory planning algorithms are complex and cannot be autonomously planned and implemented on spacecraft, and can meet the autonomous path planning requirements of microsatellite close-range detection missions.

Description

Technical Field

[0001] The present invention relates to the field of close-range detection of space targets, and particularly to an autonomous path planning method in the close-range observation mission of a microsatellite. Background Art

[0002] Performing close-range detection of space targets is an important means to obtain information such as the on-orbit state and morphological characteristics of space targets. Compared with general trajectory planning, the trajectory planning of the close-range observation mission of a spacecraft needs to consider more influencing factors and constraints. It not only needs to consider the safety of the overall mission and the rapidity of the planning algorithm, but also further emphasizes on-orbit autonomous operation.

[0003] The existing methods for close-range relative motion path planning of spacecraft mainly fall into two categories: continuous optimization and discrete search. The trajectory generated by the continuous optimization method is smooth, but its numerical solution depends on the initial value guess, the algorithm is complex, it is difficult to solve on the spacecraft, and convergence cannot be ensured. The discrete search method can theoretically obtain a globally optimal solution, but it is necessary to reasonably discretize the search space, and the generated trajectory is not smooth. In order to ensure the feasibility and safety of the close-range observation mission of a microsatellite and at the same time meet the on-orbit autonomous planning requirements, it is necessary to combine the background of different missions and study the design and optimization of the mission trajectory according to the characteristics and requirements of specific missions. Summary of the Invention

[0004] The technical problem solved by the present invention is: aiming at the autonomous planning requirements in the close-range observation mission of a microsatellite, in order to solve the problems such as complex calculation and difficult on-orbit solution of traditional trajectory planning, the purpose of the present invention is to propose an autonomous path planning method based on artificial potential function and polynomial fitting. Using the present invention improves the safety and autonomy of the close-range observation mission of a microsatellite and meets the on-orbit autonomous planning observation path requirements of a microsatellite.

[0005] The technical solution of the present invention is:

[0006] An autonomous path planning method in the close-range observation mission of a microsatellite, characterized by comprising:

[0007] Step 1, set the close distance and observation duration relative to the observation target according to the observation ability and safety requirements of the microsatellite;

[0008] Step 2, set the distribution spherical surface of the close-range observation points according to the close distance, model the distribution spherical surface by using the spherical grid method, and encode the grid by using the polar coordinate method;

[0009] Step 3, calculate the position of the sun at the current moment, and combine the initial distribution spherical surface of the close-range observation points obtained in Step 2 to judge the sunlit observation azimuth;

[0010] Step 4: Calculate the lunar position at the current moment, and combine it with the initial distribution spherical surface of the close observation points obtained in Step 2 to determine the azimuth for avoiding lunar interference light;

[0011] Step 5: According to the absolute position of the target at the current moment, and combine it with the initial distribution spherical surface of the close observation points obtained in Step 2 to determine the azimuth for avoiding earth - atmosphere interference light;

[0012] Step 6: Apply the theory of artificial potential field method to generate attraction in the direction of the forward - light observation azimuth and repulsion in the direction of the reverse - light avoidance azimuth, and calculate the resultant potential force received by the current grid;

[0013] Step 7: If it is the first calculation, traverse all grids to find the optimal observation position; otherwise, traverse the grids adjacent to the current position to find the optimal observation position;

[0014] Step 8: Update the observation time, and repeat Steps 3 - 7 to determine new grids that meet the observation requirements until the observation ends;

[0015] Step 9: Use polynomial fitting for the angles swept by adjacent grids on the spherical surface to enable smooth transition of the micro - spacecraft within each grid, and connect the trajectories between adjacent grids to obtain a relatively smooth transfer path for the micro - spacecraft during the close - range observation process.

[0016] Furthermore, in Step 1, the close - range observation task of the micro - spacecraft is subject to safety constraints, and the close - range observation trajectory should always be outside the target flight restricted area; at the same time, it is subject to the observation ability constraints of the on - board equipment of the micro - spacecraft, and the close - range detection distance needs to be within the optimal observation distance range of the on - board equipment. Therefore, it is necessary to reasonably evaluate and determine the close - range distance of the micro - spacecraft. In addition, design the observation duration as required in combination with the mission objectives.

[0017] Furthermore, in Step 2, taking into account both safety and observation requirements, the motion trajectory of the micro - spacecraft for close - range observation is constrained on a spherical surface at a fixed distance from the target center. It is a trajectory composed of multiple arcs spliced on the fixed spherical surface. The spherical grid method is used to perform grid - based modeling on this spherical surface, and the polar coordinate method is used for encoding.

[0018] Furthermore, in Step 3, an analytical formula is used to calculate the solar position vector in the J2000 inertial system at the current moment, and the transformation matrix from the J2000 inertial system to the target VVLH orbital system is calculated through the absolute position and velocity vectors of the target. Then, the forward - light observation azimuth in the target orbital system is calculated through vector subtraction and coordinate transformation. The analytical formula for calculating the relative solar position vector is:

[0019] a s = 1.00000102

[0020] e s= 0.01670862 - 0.00004204t - 0.00000124t 2

[0021] i s = 23.439291° - 0.01300417°t - 0.00000016°t 2

[0022] Ω s = 0°

[0023] ω s = 282.937347° + 0.32256206°t - 0.00015757°t 2

[0024] M s = 357.5291° + 0.98556200804°d - 0.0007734°t 2

[0025] f s = MTof(e s ,M s )

[0026] r s = CoeToRV(a s ,e s ,i s ,Ω s ,ω s ,f s ) (1)

[0027] In the formula, a s is the average distance from the sun to the earth, with the unit of AU, e s is the eccentricity, i s is the orbital inclination, Ω s is the right ascension of the ascending node, ω s is the argument of periapsis, M s and f s are the mean anomaly and the true anomaly respectively, t and d are the Julian centuries and days counted from J2000, r s is the solar position vector in the J2000 coordinate system, MTof() is the function of converting the mean anomaly to the true anomaly, and CoeToRV() is the function of converting orbital elements to position and velocity;

[0028] The formula for calculating the transformation matrix from the J2000 coordinate system to the target VVLH orbital system is:

[0029]

[0030] In the formula, r taris the position vector of the target in the geocentric J2000 inertial system, v tar is the velocity vector of the target in the geocentric J2000 inertial system;

[0031] Thus, the sunlit observation azimuth in the target VVLH orbital system is represented by a unit vector as:

[0032]

[0033] Furthermore, in step 4, an analytical formula is used to calculate the lunar position vector in the J2000 inertial system at the current moment, and the transformation matrix from the J2000 inertial system to the target VVLH orbital system is calculated through the absolute position and velocity vectors of the target. Then, the lunar interference light avoidance azimuth in the target orbital system is obtained through vector subtraction and coordinate transformation. The analytical formula for calculating the lunar position vector in the geocentric J2000 inertial system is:

[0034] λ ecliptic = 218.32° + 481267.883°t + 6.29°sin(134.9° + 477198.85°t)

[0035] - 1.27°sin(259.2° - 413335.38°t) + 0.66°sin(235.7° + 890534.23°t)

[0036] + 0.21°sin(269.9° + 954397.70°t) - 0.19°sin(357.5° + 35999.05°t)

[0037] - 0.11°sin(186.6° + 966404.05°t)

[0038]

[0039]

[0040] ε = 23.439291° - 0.0130042°t

[0041]

[0042] In the formula, λ ecliptic is the lunar longitude, is the lunar latitude, is the half angular separation of the Earth relative to the lunar center, ε is the obliquity of the ecliptic, a e is the equatorial radius of the Earth, r m is the position vector of the moon in the geocentric J2000 inertial system, t is the Julian century number counted from J2000.0,

[0043] Therefore, the lunar interference light avoidance azimuth in the target VVLH orbital system is represented by a unit vector as follows:

[0044]

[0045] Furthermore, in step 5, the terrestrial interference light avoidance azimuth is represented by a unit vector as follows:

[0046] V e = TR×r tar (6).

[0047] Furthermore, in step 6, using the artificial potential field method theory, a gravitational force is generated in the forward-light observation azimuth and a repulsive force is generated in the backlight avoidance azimuth, and the resultant potential force received by the current grid is calculated. The specific process is as follows: Convert the polar coordinate encoding of the grid into a unit vector V, and then calculate the angles between this vector and the forward-light observation vector V s , the lunar interference light avoidance vector V m and the terrestrial interference light avoidance vector V e in turn:

[0048]

[0049]

[0050]

[0051] The attractive potential field function of the forward-light observation vector is represented using the arctangent function:

[0052]

[0053] Therefore, the attractive force function generated by the forward-light observation vector is:

[0054]

[0055] In the formula, the attractive force F s is a bounded value. When the angle between the unit vector represented by the grid and the forward-light observation vector is large, all attractive forces approach the threshold k t ,

[0056] A smooth step function is introduced to represent the repulsive force functions of the lunar interference light and the terrestrial interference light avoidance vectors:

[0057]

[0058]

[0059] In the formula, k m , k e are repulsive force control parameters, and α0, β0 are critical repulsive angles.

[0060] Thus, the overall potential field resultant force on the current grid is as follows:

[0061] F = F s + F m + F e (14).

[0062] Furthermore, in step 7, the optimal observation position refers to the grid with the smallest angle between the potential field resultant force.

[0063] Furthermore, in step 8, the microspacecraft approaching observation mission also has requirements for the observation duration. In order to ensure that the observation conditions of the microspacecraft are always good during the observation period, the observation azimuth will be continuously adjusted, thereby discretizing the observation time and calculating the observation azimuths at multiple observation time points.

[0064] Furthermore, in step 9, the grids corresponding to the observation azimuths at different time points are determined according to steps 3 to 8, and then the angles swept by adjacent grids on the sphere are polynomially fitted.

[0065] Compared with the prior art, the technical beneficial effects of the present invention are as follows:

[0066] An autonomous path planning method in a microspacecraft approaching observation mission proposed by the present invention first determines good observation positions at key observation time points through the artificial potential field method, and then uses polynomials to fit the trajectories of adjacent observation points, and further obtains a relatively smooth transfer path during the approaching observation process. The method proposed by the present invention combines the advantages of continuous optimization methods and discrete search methods, solves the problems that traditional trajectory planning algorithms are complex and cannot be autonomously planned and implemented on spacecraft, and can meet the autonomous path planning requirements of microspacecraft approaching detection missions. BRIEF DESCRIPTION OF THE DRAWINGS

[0067] Figure 1 is a flowchart of the autonomous path planning method in the microspacecraft approaching observation mission of the present invention;

[0068] Figure 2 is a schematic diagram of the design principle of the approaching detection distance of the microspacecraft;

[0069] Figure 3 is a schematic diagram of the spherical grid method modeling;

[0070] Figure 4 is a schematic diagram of polar coordinate parameters;

[0071] Figure 5 is a schematic diagram of the forward light observation potential field, the lunar light avoidance potential field, and the earth-air light avoidance potential field. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0072] The following will be combined with Figures 1 to 5 to further describe in detail an autonomous path planning method in the close - proximity observation mission of a small spacecraft according to the present invention.

[0073] Figure 1 is a flowchart of the autonomous path planning method in the close - proximity observation mission of the small spacecraft according to the present invention. As Figure 1 shown, the present invention provides an autonomous path planning method for the close - proximity observation mission of a small spacecraft, and the method includes nine steps.

[0074] Step 1: According to the observation ability and safety requirements of the small spacecraft, set the close - proximity distance relative to the observation target, and set the observation duration as required.

[0075] The close - proximity observation mission of the small spacecraft is subject to safety constraints, and the close - proximity observation trajectory should always be outside the target flight no - fly zone; at the same time, it is subject to the observation ability constraints of the on - board equipment of the small spacecraft, and the close - proximity detection distance needs to be within the optimal observation distance range of the on - board equipment. Therefore, it is necessary to reasonably evaluate and determine the close - proximity distance of the small spacecraft. In addition, the observation duration can be designed as required in combination with the mission objectives.

[0076] Figure 2 is a schematic diagram of the design principle of the close - proximity detection distance of the small spacecraft. In Step 1, set the close - proximity distance relative to the observation target as Figure 2 shown, L max and L min are respectively the inner and outer envelope spheres of the motion area of the small spacecraft determined by the observation ability and the target safety requirements. To balance the observation effect and safety, select R=(L min +L max ) / 2 as the close - proximity distance. Step 2: According to the close - proximity distance, set the distribution sphere of the close - proximity observation points, model the distribution sphere using the spherical grid method, and encode the grid using the polar coordinate method.

[0077] Taking into account safety and observation requirements, the motion trajectory of the small spacecraft's close - proximity observation is constrained on a sphere at a fixed distance from the target center. For the trajectory composed of multiple circular arc segments distributed on the fixed sphere, use the spherical grid method to grid - model the sphere and use the polar coordinate method for encoding. Figure 3 is a schematic diagram of the spherical grid method for modeling. In Step 2, model the distribution sphere using the spherical grid method. As Figure 3 shown, divide the spherical longitude and latitude vertically and cross - wise into multiple grids. Figure 4 is a schematic diagram of polar coordinate parameters. Then, use the polar coordinates Figure 4 shown in to encode the grid, and the unit vector corresponding to the grid can be obtained as

[0078] Step 3: Calculate the solar position at the current moment, and combine it with the initial distribution spherical surface of the approaching observation points obtained in Step 2 to determine the sunlit observation azimuth.

[0079] Figure 5 It is a schematic diagram of the sunlit observation potential field, the lunar light avoidance potential field, and the earth atmosphere light avoidance potential field. As Figure 5 shown, use the analytical formula to calculate the solar position vector in the J2000 inertial system at the current moment, and calculate the transformation matrix from the J2000 inertial system to the target VVLH orbital system through the absolute position and velocity vector of the target. Then, the sunlit observation azimuth in the target orbital system can be calculated through vector subtraction and coordinate transformation. The analytical formula for calculating the relative solar position vector is:

[0080] a s = 1.00000102

[0081] e s = 0.01670862 - 0.00004204t - 0.00000124t 2

[0082] i s = 23.439291° - 0.01300417°t - 0.00000016°t 2

[0083] Ω s = 0°

[0084] ω s = 282.937347° + 0.32256206°t - 0.00015757°t 2

[0085] M s = 357.5291° + 0.98556200804°d - 0.0007734°t 2

[0086] f s = MTof(e s , M s )

[0087] r s = CoeToRV(a s , e s , i s , Ω s , ω s , f s ) (1)

[0088] In the formula, a s is the average distance from the sun to the earth, with the unit of AU, e sis the eccentricity, i s is the orbital inclination, Ω s is the right ascension of the ascending node, ω s is the argument of perigee, M s and f s are the mean anomaly and the true anomaly respectively, t and d are the Julian centuries and days counted from J2000, r s is the solar position vector in the J2000 coordinate system, MTof() is the function for converting the mean anomaly to the true anomaly, and CoeToRV() is the function for converting orbital elements to position and velocity.

[0089] The formula for calculating the transformation matrix from the J2000 coordinate system to the target VVLH orbital system is:

[0090]

[0091] In the formula, r tar is the position vector of the target in the geocentric J2000 inertial system, v tar is the velocity vector of the target in the geocentric J2000 inertial system. Thus, the sunward observation azimuth in the target VVLH orbital system can be represented by a unit vector as:

[0092]

[0093] Step 4: Calculate the lunar position at the current moment, and combine it with the initial distribution spherical surface of the close observation points obtained in Step 2 to determine the lunar interference light avoidance azimuth;

[0094] Figure 5 is a schematic diagram of the sunward observation potential field, the lunar light avoidance potential field, and the earth - atmosphere light avoidance potential field. As Figure 5 shown, use the analytical formula to calculate the lunar position vector in the J2000 inertial system at the current moment, and calculate the transformation matrix from the J2000 inertial system to the target VVLH orbital system through the absolute position and velocity vectors of the target. Then, the lunar interference light avoidance azimuth in the target orbital system can be obtained through vector subtraction and coordinate transformation. The analytical formula for calculating the lunar position vector in the geocentric J2000 inertial system is:

[0095] λ ecliptic = 218.32° + 481267.883°t + 6.29°sin(134.9° + 477198.85°t)

[0096] - 1.27°sin(259.2° - 413335.38°t) + 0.66°sin(235.7° + 890534.23°t)

[0097] +0.21°sin(269.9° + 954397.70°t) - 0.19°sin(357.5° + 35999.05°t)

[0098] -0.11°sin(186.6° + 966404.05°t)

[0099]

[0100]

[0101] ε = 23.439291° - 0.0130042°t

[0102]

[0103] In the formula, λ ecliptic is the lunar ecliptic longitude, is the lunar ecliptic latitude, is the half angular separation of the Earth relative to the lunar center, ε is the obliquity of the ecliptic, a e is the equatorial radius of the Earth, r m is the position vector of the Moon in the geocentric J2000 inertial system, and t is the Julian century number counted from J2000.0. Then, the azimuth for avoiding lunar interference light in the target VVLH orbital system can be represented by a unit vector as follows:

[0104]

[0105] Step 5: According to the target absolute position at the current moment and combined with the initial distribution spherical surface of the approaching observation points obtained in Step 2, judge the azimuth for avoiding Earth - atmosphere interference light;

[0106] Figure 5 Schematic diagrams of the potential field for observing with the light, the potential field for avoiding lunar light, and the potential field for avoiding Earth - atmosphere light are shown. As Figure 5 shown, the azimuth for avoiding Earth - atmosphere interference light can be represented by a unit vector as follows:

[0107] V e = TR×r tar (6)

[0108] Step 6: Apply the theory of the artificial potential field method to generate gravitational force in the azimuth for observing with the light and repulsive force in the azimuth for avoiding the light, and calculate the resultant potential force on the current grid.

[0109] In step 6, combined with the initial position of the microsatellite, the artificial potential field method theory is used to generate gravitational force in the forward-light observation direction and repulsive force in the backlight avoidance direction, calculate the resultant potential force received by the current grid, and thus determine the grids that meet the requirements of the next observation among the grids adjacent to the current position. The specific process is as follows: convert the polar coordinate encoding of the grid into a unit vector V, and then calculate the angles between this vector and the forward-light observation vector V s , the lunar interference light avoidance vector V m , and the earth-air interference light avoidance vector V e in turn:

[0110]

[0111]

[0112]

[0113] Use the arctangent function to represent the attractive potential field function of the forward-light observation vector:

[0114]

[0115] Therefore, the attractive force function generated by the forward-light observation vector is:

[0116]

[0117] In the formula, the attractive force F s is a bounded value. When the angle between the unit vector represented by the grid and the forward-light observation vector is large, all attractive forces approach the threshold k t .

[0118] Introduce a smooth step function to represent the repulsive force functions of the lunar interference light and the earth-air interference light avoidance vector:

[0119]

[0120]

[0121] In the formula, k m , k e are repulsive force control parameters, and α0, β0 are critical repulsive angles.

[0122] Therefore, the resultant potential force received by the current grid is:

[0123] F = F s + F m + F e (14)

[0124] Step 7: If it is the first calculation, traverse all grids to find the optimal observation position; otherwise, traverse the grids adjacent to the current position to find the optimal observation position. The optimal observation position refers to the grid with the smallest angle between the potential field resultant force.

[0125] Step 8: Update the observation time, and repeat Steps 3 - 7 to determine new grids that meet the observation requirements until the observation ends. In Step 8, since the microspacecraft's close - approach observation mission also has requirements for the observation duration, in order to ensure that the observation conditions of the microspacecraft are always good during the observation period, it is necessary to continuously adjust the observation azimuth. Therefore, it is necessary to discretize the observation time and calculate the observation azimuths at multiple observation time points.

[0126] Step 9: Fit the angles swept by adjacent grids on the sphere using a polynomial to enable the microspacecraft to smoothly transition within each grid. Connect the trajectories between adjacent grids, and a relatively smooth transfer path during the microspacecraft's close - approach observation process can be obtained. In Step 9, based on Steps 3 - 8, determine the grids corresponding to the observation azimuths at different time points, and then fit the angles swept by adjacent grids on the sphere using a polynomial.

[0127] The following introduces the application process of the present invention in combination with specific embodiments:

[0128] Taking the close - approach observation of China's space station as an example, the present invention will be further described in detail.

[0129] Step 1: According to the capabilities of the microspacecraft's observation camera and in combination with the design of the safety no - fly zone of the space station, set the observation distance to 200 m. Considering that low - orbit spacecraft will enter and exit the earth's shadow, set the observation duration to 30 min. Step 2: Divide the 200 - m sphere surface into several grids at 5° intervals and represent them in polar coordinates. Step 3: Calculate the sun - facing observation azimuth at the current moment and determine the corresponding grid coordinates. Step 4: Calculate the lunar - light avoidance azimuth at the current moment and determine the corresponding grid coordinates. Step 5: Calculate the earth - atmosphere - light avoidance azimuth at the current moment and determine the corresponding grid coordinates. Step 6: Apply the artificial potential field method theory to generate attraction for the sun - facing observation azimuth and repulsion for the back - light avoidance azimuth, and calculate the potential field resultant force. Step 7: For the first calculation, traverse all grids to find the optimal observation position; for subsequent calculations, traverse the grids adjacent to the current position to find the optimal observation position. Step 8: Advance the observation time by 5 min. If it does not exceed the total observation duration, repeat Steps 3 - 7. Step 9: Use a cubic polynomial to fit adjacent grids, and a relatively smooth motion trajectory can be obtained.

[0130] The above embodiments are only used to illustrate the content of the present invention. In addition to the above - mentioned implementation manners, the present invention has other implementation manners. All technical solutions formed by equivalent replacement or equivalent deformation fall within the protection scope of the present invention.

Claims

1. An autonomous path planning method in the close observation mission of a small spacecraft, characterized in that Including: Step 1: Set the approaching distance and observation duration relative to the observation target according to the observation capabilities and safety requirements of the microsatellite. Step 2: Set the distribution sphere of the approaching observation points according to the approaching distance, model the distribution sphere using the spherical grid method, and encode the grid using the polar coordinate method. Step 3: Calculate the solar position at the current moment, and combine it with the initial distribution sphere of the approaching observation points obtained in Step 2 to determine the sunlit observation azimuth. Step 4: Calculate the lunar position at the current moment, and combine it with the initial distribution sphere of the approaching observation points obtained in Step 2 to determine the azimuth for avoiding lunar interference light. Step 5: According to the absolute position of the target at the current moment, and combine it with the initial distribution sphere of the approaching observation points obtained in Step 2 to determine the azimuth for avoiding earth-air interference light. Step 6: Apply the theory of artificial potential field method to generate attraction in the sunlit observation azimuth and repulsion in the backlight avoidance azimuth, and calculate the resultant potential force received by the current grid. Step 7: If it is the first calculation, traverse all grids to find the optimal observation position; otherwise, traverse the adjacent grids of the current position to find the optimal observation position. Step 8: Update the observation time, and repeat Steps 3 - 7 to determine a new grid that meets the observation requirements until the observation ends. Step 9: Fit the angles swept by adjacent grids on the sphere using polynomials to enable the microsatellite to smoothly transition within each grid, and connect the trajectories between adjacent grids to obtain a relatively smooth transfer path during the approaching observation process of the microsatellite.

2. The autonomous path planning method in the close-range observation mission of a small spacecraft according to claim 1, wherein, In Step 1, the approaching observation task of the microsatellite is constrained by safety, and the approaching observation trajectory should always be outside the target flight restricted area; at the same time, it is constrained by the observation capabilities of the microsatellite's on-board equipment, and the approaching detection distance needs to be within the optimal observation distance range of the on-board equipment. Therefore, it is necessary to reasonably evaluate and determine the approaching distance of the microsatellite. In addition, design the observation duration as required in combination with the mission objectives.

3. The autonomous path planning method in the close-range observation mission of a small spacecraft according to claim 1, wherein In Step 2, taking into account safety and observation requirements, the movement trajectory of the microsatellite's approaching observation is constrained on a sphere at a fixed distance from the target center. It is a trajectory composed of multiple arcs spliced on a fixed sphere. The spherical grid method is used to grid-model this sphere, and the polar coordinate method is used for encoding.

4. The autonomous path planning method in the close-range observation mission of a small spacecraft according to claim 1, characterized in that, In Step 3, an analytical formula is used to calculate the solar position vector in the J2000 inertial system at the current moment, and the transformation matrix from the J2000 inertial system to the target VVLH orbital system is calculated through the absolute position and velocity vector of the target. Then, the sunlit observation azimuth in the target orbital system is calculated through vector subtraction and coordinate transformation. The analytical formula for calculating the relative solar position vector is as follows: a s ＝1.00000102 e s = 0.01670862 - 0.00004204t - 0.00000124t 2 i s = 23.439291° - 0.01300417°t - 0.00000016°t 2 Ω s = 0° ω s = 282.937347° + 0.32256206°t - 0.00015757°t 2 M s = 357.5291° + 0.98556200804°d - 0.0007734°t 2 f s = MTof(e s , M s ) r s = CoeToRV(a s , e s , i s , Ω s , ω s , f s ) (1) where a s is the average distance from the Sun to the Earth, in AU, e s is the eccentricity, i s is the inclination of the orbit, Ω s is the right ascension of the ascending node, ω s is the argument of perigee, M s and f s are the mean anomaly and the true anomaly respectively, t and d are the number of Julian centuries and days counted from J2000, r s is the solar position vector in the J2000 coordinate system, MTof() is the function for converting the mean anomaly to the true anomaly, and CoeToRV() is the function for converting orbital elements to position and velocity; The formula for calculating the transformation matrix from the J2000 coordinate system to the target VVLH orbital system is: where r tar is the position vector of the target in the geocentric J2000 inertial frame, and v tar is the velocity vector of the target in the geocentric J2000 inertial frame; Thus, the sunlit observation azimuth in the target VVLH orbital system is expressed as a unit vector:

5. The autonomous path planning method in the close observation mission of a small spacecraft according to claim 1, wherein In step 4, the position vector of the moon in the J2000 inertial system at the current moment is calculated using an analytical formula, and the transformation matrix from the J2000 inertial system to the target VVLH orbital system is calculated through the absolute position and velocity vector of the target. Then, the azimuth for avoiding lunar interference light in the target orbital system is calculated through vector subtraction and coordinate transformation. The analytical formula for calculating the position vector of the moon in the geocentric J2000 inertial system is as follows: λ ecliptic = 218.32° + 481267.883°t + 6.29° sin(134.9° + 477198.85°t) - 1.27° sin(259.2° - 413335.38°t) + 0.66° sin(235.7° + 890534.23°t) + 0.21° sin(269.9° + 954397.70°t) - 0.19° sin(357.5° + 35999.05°t) - 0.11° sin(186.6° + 966404.05°t) ε = 23.439291° - 0.0130042°t where λ ecliptic is the lunar longitude, is the lunar latitude, is the semi-aperture angle of the Earth relative to the lunar center, ε is the obliquity of the ecliptic, a e is the equatorial radius of the Earth, r m is the position vector of the Moon in the geocentric J2000 inertial frame, and t is the number of Julian centuries since J2000.0, Therefore, the azimuth for avoiding lunar interference light in the target VVLH orbital system is represented by a unit vector as:

6. The autonomous path planning method in the close-range observation mission of a small spacecraft according to claim 1, wherein In step 5, the azimuth for avoiding earth-air interference light is represented by a unit vector as: V e = TR × r tar (6).

7. The autonomous path planning method in the close-range observation mission of a small spacecraft according to claim 1, characterized in that In the said step 6, the artificial potential field method theory is used to generate an attractive force in the front-light observation direction and a repulsive force in the back-light avoidance direction, and calculate the resultant potential force received by the current grid. The specific process is as follows: Convert the polar coordinate encoding of the grid into a unit vector V, and then calculate the angles between this vector and the front-light observation vector V s , the lunar interference light avoidance vector V m , and the earth-air interference light avoidance vector V e in turn: The attractive potential field function of the forward-light observation vector is represented using the arctangent function: Therefore, the attractive function generated by the forward-light observation vector is: where the attractive force F s is a bounded value. When the included angle between the unit vector represented by the grid and the frontlight observation vector is large, all attractive forces approach the threshold value k t , A smooth step function is introduced to represent the repulsive force function of the lunar interference light and the earth-air interference light avoidance vectors: where k m , k e are repulsive force control parameters, α0 and β0 are critical repulsive angles Therefore, the total resultant potential force received by the current grid is: F = F s + F m + F e (14).

8. The autonomous path planning method in the close observation mission of a small spacecraft according to claim 1, characterized in that In step 7, the optimal observation position refers to the grid with the smallest angle with the resultant potential force.

9. The autonomous path planning method in the close observation mission of a small spacecraft according to claim 1, wherein In step 8, the micro-spacecraft approaching observation mission also has requirements for the observation duration. To ensure that the observation conditions of the micro-spacecraft are always good during the observation period, the observation azimuth is continuously adjusted, thereby discretizing the observation time and calculating the observation azimuths at multiple observation time points.

10. The autonomous path planning method in the close-range observation mission of a small spacecraft according to claim 9, wherein In step 9, the grids corresponding to the observation azimuths at different time points are determined according to steps 3 to 8, and then the angles swept by adjacent grids on the sphere are polynomially fitted.

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