Autonomous mission planning method for fast maneuvering rendezvous imaging of spatially non-planar target

Abstract

The application discloses a self-task planning method for quick maneuvering rendezvous imaging of a spatial space target, and comprises the following steps: step one, multi-pulse control sequence planning: according to task requirements and considering determination of an orbit requirement, a multi-pulse control sequence is given, and the first control is implemented autonomously on the satellite; step two, closed-loop control sequence planning: after the i-th control is finished, the remaining N-i control sequences are given in combination with the orbit of the satellite after control and the orbit of the target, and the i+1-th control is implemented; wherein, i is greater than or equal to 1 and less than or equal to N-2, and the total control times are N; step three, single-pulse transfer planning: after N-1 control is finished, the last control velocity increment is calculated by using a Lambert transfer method in combination with the orbit of the satellite after control and the orbit of the target, and the N-th control is implemented; and step four, imaging task planning. The closed-loop control planning method is given by considering the influence of control errors and orbit errors, the method can ensure that the planning result of each time is basically consistent with the first planning, and the final rendezvous precision requirement is met.

Description

TECHNICAL FIELD

[0001] The patent relates to the field of autonomous rendezvous of space non-planar targets, in particular to the rendezvous imaging task of space non-planar targets, through multiple on-board autonomous task planning and corresponding control implementation, to ensure that the autonomous rendezvous imaging meeting the task requirements is realized under the condition of having orbit error and control error. BACKGROUND

[0002] When designing the rendezvous control of space non-planar targets, fuel optimization is considered, and the general analysis method is to solve the minimum single-pulse velocity increment meeting the rendezvous constraints. Due to the limitation of the longest single jet time of the spacecraft, the single-pulse velocity increment needs to be decomposed into multiple pulses for control.

[0003] During the task implementation process, considering the safety of the spacecraft orbit, orbit determination needs to be performed again after control, and due to the influence of control error and orbit error, if the orbit control strategy at the beginning of the task planning is continued to be executed, it is very likely that the rendezvous imaging requirements cannot be met, so it is necessary to re-plan after each control. In addition, in the actual task implementation process, considering the influence of the measurement and control arc segment, the control time of multiple planning should be ensured to be consistent with the first planning as much as possible. SUMMARY

[0004] For the maneuvering rendezvous imaging task of space non-planar targets, an autonomous task planning method for fast maneuvering rendezvous imaging of space non-planar targets is proposed, considering the influence of control error and orbit error, a closed-loop control planning method is given, which can ensure that each planning result is basically consistent with the first planning, and meet the final rendezvous precision requirements.

[0005] The technical scheme of the present application is: an autonomous task planning method for fast maneuvering rendezvous imaging of space non-planar targets, the specific steps are as follows:

[0006] Step 1, multi-pulse control sequence planning: according to the task requirements, considering the orbit determination requirements, a multi-pulse control sequence is given, and the first control is implemented autonomously on board;

[0007] Step 2, closed-loop control sequence planning: after the i-th control is completed, the remaining N-i control sequence is given combined with the post-control satellite orbit and the target orbit, and the i+1 control is implemented; wherein i≥1 and i≤N-2, the total control times is N times;

[0008] Step 3, single-pulse transfer planning: after N-1 controls are completed, the last control velocity increment is calculated combined with the post-control satellite orbit and the target orbit using the Lambert transfer method, and the Nth control is implemented;

[0009] Step 4, Imaging Task Planning: After the Nth control cycle, imaging task planning is performed, and the payload imaging parameters are given for intersection imaging.

[0010] Furthermore, the multi-pulse control sequence planning described in step one is specifically implemented as follows:

[0011] Step 11: Solve for the minimum velocity increment based on the Lambert equation;

[0012] Lambert's theorem can be used to analyze orbital transfers between any two points in space. Based on Lambert's equations, this paper analyzes and solves for the minimum velocity increment required in a spacecraft rendezvous and imaging mission against an anomalous target. The velocity increment is closely related to the semi-major axis of the transfer orbit. For spacecraft with unrestricted transfer time, once the maneuvering point and rendezvous point are determined, countless transfer orbits will pass through these two points. For spacecraft, due to fuel constraints, the key focus is on determining which transfer orbit's semi-major axis minimizes the velocity increment passing through the two points—this is the minimum velocity increment transfer problem. To solve for the minimum velocity increment, we need to find the derivative of the velocity increment with respect to the semi-major axis of the transfer orbit. Based on the derivative information, the bisection method is used to solve for the minimum velocity increment.

[0013] (111) Set the transfer semi-major axis Search limit and search lower limit ;like Then let the initial Otherwise, let the initial ;like If the current transition interval of the semi-major axis has no solution, then let the initial... ;in, For the semi-major axis of the parking track, To find the minimum value of the semi-major axis of the ellipse;

[0014] (112) For any ,make ,calculate , , ,in Iterate from 0, that is , ;in, For the transfer speed increment, For the first Group transfer semi-major axis search lower limit, For the first upper limit of the semi-major axis search for group transfer For the first Group transition semi-major axis median;

[0015] (113) If , let , ; if , let , , turn (112) continues iteration;

[0016] (114) if or , stop calculation, take , the minimum velocity increment , that is ;

[0017] Step 12, according to the determination of the track requirements and satellite control ability to carry out multi-pulse velocity increment decomposition;

[0018] In the actual task implementation process, due to the existence of control error and on-board extrapolation model error, a control is easy to make the terminal error accumulate too large, generally need to be divided into multiple control to reduce the terminal error; and the actual thrust is not pulse thrust, when the transfer velocity increment is larger, the control time is longer than the longest jet time limit of single control of the thruster, also need to be divided into multiple control to complete, therefore, the multi-pulse velocity increment decomposition problem under complex perturbation needs to be considered.

[0019] Multi-pulse velocity increment decomposition method is: the minimum fuel transfer velocity increment under ideal conditions is divided into controls, each control is carried out along the direction of , so the velocity increment of single control satisfies ;

[0020] Closed orbit is defined as the orbit controlled at a certain point and finally returned to this point; let the last control in controls be the intersection control, the first controls be closed control, according to the time equation, the following relationship can be obtained:

[0021]

[0022] Among them, and are the mean anomaly of the parking satellite and the target satellite at the initial time; is the mean anomaly of the parking satellite at the maneuver point; is the mean anomaly of the target satellite at the intersection point; and are the mean anomaly of points A and B on the transfer orbit; , , and These are the semi-major axes of the parking track, the target track, the closed track after the i-th control, and the transfer track, respectively. This refers to the number of orbits the parking satellite completes in its parking orbit. For the number of orbits the satellite makes in its i-th closed orbit, The number of orbits the target satellite makes in the target orbit;

[0023] The above formula can be rewritten as:

[0024]

[0025]

[0026] The semi-major axis of a closed orbit can be represented as:

[0027]

[0028] Assuming the track control interval is at least one track, then , and The value is , , Solve using a numerical scheme, according to The sequence is set to loop, and the above formula is used to solve for the semi-major axis of the closed orbit. ;

[0029] When the semi-major axis of the closed track Once determined, the speed increment required for each control operation can be calculated. The calculation steps are as follows:

[0030] (121) In the LVLH coordinate system, the velocity increment of the maneuvering point A can be expressed as: Then the speed before control is: The speed after control is: ;

[0031] in, The velocity increment of maneuver point A, The X-axis component of the velocity increment at the maneuvering point. The Y-axis component of the velocity increment at the maneuvering point. To control forward speed, The X-axis component of the control velocity increment, The Y-axis component of the control velocity increment, To control the speed, To control the X-axis component of the velocity increment, The Y-axis component of the control velocity increment;

[0032] (122) When Once confirmed, the first The magnitude of the speed after secondary control is Let the vector expression be: , Two variables, one equation There are infinitely many solutions. This paper uses the proportional method to solve it;

[0033] in, For the first The speed modulus after secondary control Let A be the radius vector. For the semi-major axis of the closed orbit, The gravitational constant of Earth, For the first Speed ​​increment after secondary control For the first The X-axis component of the velocity increment after secondary control. No. The Y-axis component of the velocity increment after secondary control;

[0034] (123) Setting The vector direction and If the vector directions are consistent, then A vector can be represented as: Then from The conditions that need to be met are:

[0035]

[0036] in, To control forward speed, To control the speed, For the speed increment of the i-th control, Let X be the speed increment of the i-th control. Let Y be the Y-axis component of the speed increment for the i-th control. For coefficients;

[0037] (124) Solve That is, to solve for the current i-th The vector of secondary control, where , According to the above iterable velocity increment vector required each time Finally, the multi-pulse velocity increment is obtained. As shown in the following formula:

[0038]

[0039] in, Let X be the X-axis component of the pre-control velocity for the i-th time. Let Y be the Y-axis component of the pre-control velocity for the i-th time. X-axis component of the i-th controlled velocity, Y-axis component of the i-th controlled velocity, Controlled velocity increment of the i-th, Multi-pulse velocity increment, Control instants of the i-th to the K-th, Controlled velocity increments of the i-th to the K-th.

[0040] Further, after the last control implementation is completed, the multi-pulse velocity increment calculated according to step one and the post-control satellite orbit and target orbit are used to perform the closed-loop control sequence planning described in step two, and the specific method is as follows:

[0041] Step 21, extrapolate the satellite post-control orbit and the target orbit to the vicinity of the control time of this planning, update the pre-control orbit semi-major axis and pre-control velocity of this planning, and calculate the remaining controlled velocity increment according to the closed orbit semi-major axis ;

[0042] Step 22, calculate the time error of the spacecraft after applying the remaining multi-pulse velocity increment to the time of arrival at the rendezvous point , and the time of arrival of the space debris at the rendezvous point , then the rendezvous time difference between the spacecraft and the space debris is ;

[0043] Step 23, if the rendezvous time error is small but does not meet the rendezvous time requirement, the first controlled velocity increment of the multi-pulse control sequence is modified. If and , is the rendezvous time difference between the spacecraft and the space debris, is the rendezvous time difference threshold, is the rendezvous time difference iteration exit condition, then according to the rendezvous time difference the semi-major axis of the first control that needs to be adjusted is calculated as follows:

[0044]

[0045]

[0046] wherein, is the semi-major axis of the first control after adjustment, is the semi-major axis of the first control without adjustment, is the correction amount of the semi-major axis of the first control, is the time difference of the rendezvous point per orbit, is the rendezvous time difference, is the number of revolutions of the satellite in the first closed orbit semi-major axis; is the derivative of the intersection period with respect to the semi-major axis, is the Earth non-spherical perturbation J2 term, is the Earth gravitational constant, is the Earth equatorial radius, is the orbit inclination.

[0047] Further, the single pulse transfer planning of step three, after the completion of the N-1 control implementation, the last control planning is carried out, and the Lambert single pulse transfer is used for control sequence planning, and the specific method is as follows:

[0048] Step 31, according to the satellite orbit after control and the target orbit, the satellite orbit and the target orbit of the maneuver point and the intersection point are found;

[0049] Step 32, the transfer velocity increment is solved by using the Lambert equation;

[0050] Step 33, the velocity increment is substituted into the high-precision orbit extrapolation model to calculate the intersection imaging, if the imaging requirement is met, the velocity increment is output as the control result of the last task planning; otherwise, the intersection parameters are updated, and the solving is returned to step 31 until the intersection imaging requirement is met.

[0051] Further, the imaging task planning of step four, after the end of the last control, the imaging task planning is carried out, and the specific method is as follows:

[0052] Step 41, the satellite orbit after control and the target orbit are extrapolated to the vicinity of the orbit intersection point;

[0053] Step 42, according to the satellite orbit and the target orbit, the relative distance, azimuth, elevation, sun angle and other imaging parameters are calculated, and the specific calculation method is as follows:

[0054] a) Relative distance

[0055] Position of satellite in J2000 system Position of target in J2000 system Relative position vector of satellite and target Relative distance ;

[0056] b) Azimuth and elevation

[0057] The azimuth and elevation generally refer to the azimuth and elevation of the target relative to the satellite VVLH coordinate system. In order to obtain the azimuth and elevation in the VVLH system, the relative position vector obtained by the above calculation needs to be transformed into the VVLH system. The vector is converted to the satellite VVLH coordinate system, and the formula is as follows (wherein is the conversion matrix from J2000 system to VVLH coordinate system)

[0058]

[0059]

[0060]

[0061]

[0062] wherein, is the target position in the satellite VVLH coordinate system, is the conversion matrix from J2000 system to VVLH coordinate system, is the X-axis component of the target position in the satellite VVLH coordinate system, is the Y-axis component of the target position in the satellite VVLH coordinate system, is the target azimuth process quantity, is the target azimuth, is the target elevation angle;

[0063] b) Sun angle

[0064] The sun angle is defined as the angle between the target pointing satellite vector and the target pointing sun vector. In the following formula, is the sun vector in the inertial system.

[0065]

[0066] wherein, is the sun angle, is the sun vector in the inertial system, is the position of the target relative to the satellite in the J2000 system;

[0067] Step 43, screening the imaging parameters meeting the load constraints.

[0068] The beneficial effects of the present application are: a kind of fast maneuvering rendezvous imaging self-task planning method for space non-planar target is proposed, the influence of control error and orbit error is considered, and closed-loop control planning method is given, which can ensure that each planning result is basically consistent with the first planning, and meet the final rendezvous precision requirement. DETAILED DESCRIPTION

[0069] In order to make the physical law and efficacy described in the patent easy to understand, the following will be further described in combination with specific embodiments.

[0070] I. Task scenario design

[0071] The patent is directed to the maneuvering rendezvous imaging task of spatially different surface targets, and the on-board autonomous task planning calculation is carried out. The task scene is designed as follows.

[0072] (1) The initial orbits of the spacecraft and the target are as shown in Table 1.

[0073] Table 1

[0074]

[0075] (2) The spacecraft load working constraint conditions are as shown in Table 2.

[0076] Table 2

[0077]

[0078] (3) Rendezvous imaging requirements

[0079] The closest imaging distance of the spacecraft to the target is less than 30 km, and the safety distance is 5 km.

[0080] II. Multi-pulse control sequence planning

[0081] According to step one, the fast maneuvering multi-pulse control sequence planning of the spacecraft to the different surface target is carried out, and the results are shown in Table 3.

[0082] Table 3

[0083]

[0084] According to the first task planning result, the first control-V1 is implemented at the intersection point of the two stars, and the spacecraft orbit semi-major axis is adjusted to 21.063 km, and the eccentricity is changed by 0.002336.

[0085] III. Closed-loop control sequence planning

[0086] After the first control, the orbit parameters of the spacecraft and the target change, and when the closed-loop control sequence planning is carried out, the orbits of the spacecraft and the target are as shown in Table 4.

[0087] Table 4

[0088]

[0089] According to step two, the closed-loop control sequence planning is carried out on the above table orbits, and the results are shown in Table 5.

[0090] Table 5

[0091]

[0092] According to the control result of the second time planning, the second time orbit control-V2 is implemented at the two-star intersection point, the orbit semi-major axis of the satellite is reduced by 6.001 km, and the eccentricity changes by 0.000656.

[0093] Four, single-pulse transfer planning

[0094] After the second control, the orbit parameters of the spacecraft and the target change, and when the single-pulse transfer planning is performed, the orbit of the spacecraft and the target is as shown in Table 6.

[0095] Table 6

[0096]

[0097] According to the single-pulse transfer planning of the above table orbit according to step three, the result is as shown in Table 7.

[0098] Table 7

[0099]

[0100] According to the control result of the third time planning, the third time orbit control-V3 is implemented at the two-star intersection point, the orbit semi-major axis of the satellite is reduced by 1.813 km, and the eccentricity changes by 0.00021.

[0101] Five, imaging task planning

[0102] After the third control (the last control), the out-of-plane rendezvous imaging task planning calculation is started, and the orbit parameters of the spacecraft and the target are as shown in Table 8.

[0103] Table 8

[0104]

[0105] According to the imaging task planning of the orbit after the third control according to step four, the imaging parameters are as shown in Table 9.

[0106] Table 9

[0107] .

[0108] The above only describes the preferred embodiments of the present application, and it should be noted that for ordinary skilled in the art, without departing from the principles of the present application, a number of improvements and refinements can be made, and these improvements and refinements should be considered as the protection scope of the present application.

Claims

1. An autonomous mission planning method for fast maneuvering rendezvous imaging of spatially non-planar targets, characterized in that, The specific steps are as follows: Step one, multi-pulse control sequence planning: according to the task requirements, considering the determination of the orbit requirements, a multi-pulse control sequence is given, and the first control is implemented autonomously on the satellite; Step two, closed-loop control sequence planning: after the i-th control is completed, the remaining N-i control sequence is given in combination with the post-control satellite orbit and the target orbit, and the i+1-th control is implemented; wherein, i≥1 and i≤N-2, and the total control times is N times; After the last control is implemented, the multi-pulse velocity increment calculated in step one, the post-control satellite orbit and the target orbit are used to perform the closed-loop control sequence planning in step two, and the specific method is as follows: Step 21, extrapolate the satellite post-control orbit and target orbit to the vicinity of the last planned control opportunity, update the pre-control orbit semi-major axis and pre-control velocity of the current planning, and calculate the remaining control velocity increment ΔV according to the closed orbit semi-major axis Mi ; Step 22, calculate the intersection time error after applying the remaining multi-pulse velocity increment Mi The time of the later-arriving intersection point is t s The time of the space debris arriving at the intersection point is t t The intersection time difference between the spacecraft and the space debris is Δt r = t t -t s ; Step 23, if the intersection time error is small but does not meet the intersection time requirement, then the first control velocity increment of the multi-pulse control sequence is modified; if Δt r ≤ Δt set and Δt r ≤ Δt Iter , Δt r is the intersection time difference of the spacecraft and the space debris, Δt set is the intersection time difference threshold, Δt Iter is the intersection time difference iteration exit condition, then the semi-major axis a r of the first control which needs to be adjusted is obtained according to the intersection time difference Δt 1_co , and is calculated as follows: a 1_co = a1+ Δa1 wherein a 1_co is the adjusted semi-major axis of the first control, a1 is the unadjusted semi-major axis of the first control, Δa1 is the correction amount of the semi-major axis of the first control, ΔT1 is the time difference of the intersection point per orbit, Δt r is the time difference of intersection, Q1 is the number of orbits of the satellite in the first closed orbit semi-major axis; is the derivative of the intersection period to the semi-major axis, J2 is the J2 term of the earth's non-spherical perturbation, μ is the earth's gravitational constant, R e is the earth's equatorial radius, i s is the orbit inclination; Step three, single-pulse transfer planning: after N-1 controls are completed, the last control velocity increment is calculated using the Lambert transfer method in combination with the post-control satellite orbit and the target orbit, and the N-th control is implemented; Step four, imaging task planning: after the N-th control is completed, the imaging task planning is performed, and the load imaging parameters are given for the encounter imaging.

2. The autonomous task planning method for fast maneuvering imaging of spatially-skewed targets according to claim 1, characterized in that, The multi-pulse control sequence planning in step one has the following specific method: Step 11, minimum velocity increment solving based on Lambert equation; (111) set transfer semi-major axis a tr upper search limit a s and lower search limit a x ; if 0.5a P <a m , then set initial a x0 =a m , otherwise set initial a x0 =0.5a P ; if 1.5a P <a m , then current transfer semi-major axis interval has no solution, otherwise set initial a s0 =1.5a P ; wherein a P is the semi-major axis of the parking orbit, a m is the minimum value of the semi-major axis of the transfer ellipse; (112) for any a xi , a si , let a zi = (a xi + a si ) / 2, compute where i iterates from 0, i.e., a x0 , a s0 ; where Δv A is the transfer velocity increment, a xi is the lower bound of the search for the semimajor axis of the ith transfer, a si is the upper bound of the search for the semimajor axis of the ith transfer, and a zi is the median of the semimajor axis of the ith transfer. (113)if let a x(i+1) = a xi , a s(i+1) = a zi ; if let a x(i+1) = a zi , a s(i+1) = a si , go to (112) for iteration; (114) if or |a si -a xi | < 0, stop calculation, take a tr = a zi minimum speed increment Δv min i.e. Step 12, multi-pulse velocity increment decomposition based on determination of the orbit requirements and satellite control capability; The multi-pulse velocity increment decomposition method is: the most fuel-efficient transfer velocity increment Δv min is controlled K times, each control is in the direction of Δv min , it can be known that the velocity increment Δv i satisfies The closed orbit is defined as an orbit that can be controlled at a certain point and finally returned to the point; assuming that the last control in K controls is the encounter control, and the first K-1 controls are closed controls, according to the time equation, the following relationship can be obtained: wherein M P0 and M T0 are the mean anomaly of the parking satellite and the target satellite at the initial moment; M PA is the mean anomaly of the parking satellite at the maneuver point; M TB is the mean anomaly of the target satellite at the rendezvous point; M trA and M trB are the mean anomaly of points A and B on the transfer orbit; a P , a T , a i , and a tr are the semi-major axes of the parking orbit, the target orbit, the closed orbit after the ith control, and the transfer orbit, respectively; N is the number of revolutions of the parking satellite on the parking orbit, N i is the number of revolutions of the parking satellite on the ith closed orbit, and Z is the number of revolutions of the target satellite on the target orbit; The above formula can be rewritten as: The closed orbit semi-major axis can be expressed as: Assuming that the orbit control interval is at least 1 orbit, N, N i and Z are values of N = 1, 2, 3..., N i = 1, 2, 3..., Z = 1, 2, 3... ; solve by numerical scheme, set the loop in the order of Z, N i , N, and solve the closed orbit semi-major axis a i ; When the closed orbit semi-major axis a i After the determination, the velocity increment required for each control can be solved, and the solving steps are as follows: (121)In LVLH coordinate system, the velocity increment of the maneuver point A can be expressed as: Δv = [Δv X , Δv Y , 0], the pre-control velocity is: v PA = [v PAX , Δv PAY , 0], and the post-control velocity is: v trA = [v trAX , Δv trAY , 0]; where Δv is the velocity increment at the maneuver point A, Δv X is the X-axis component of the velocity increment at the maneuver point, Δv Y is the Y-axis component of the velocity increment at the maneuver point, v PA is the pre-control velocity, v PAX is the X-axis component of the pre-control velocity increment, Δv PAY is the Y-axis component of the pre-control velocity increment, v trA is the post-control velocity, v trAX is the X-axis component of the post-control velocity increment, Δv trAY is the Y-axis component of the post-control velocity increment; (122) when a i The modulus of the velocity after the i-th control is determined as Let the vector expression be: v i = [v ix , v iy , 0], Solve by the equal proportion method; wherein v i is the modulus of the velocity after the i-th control, r A is the vector radius of the maneuver point A, a i is the semi-major axis of the closed orbit, μ is the Earth's gravitational constant, v i is the velocity increment after the i-th control, v ix is the X-axis component of the velocity increment after the i-th control, v iy is the Y-axis component of the velocity increment after the i-th control; (123) Set Δv i The vector direction of Δv i The vector can be expressed as: Δv i = [kΔv x , kΔv y , 0], then the condition that needs to be met from v i-1 → v i is: where v i-1 is the pre-control speed, v i is the post-control speed, Δv i is the speed increment of the i-th control, Δv x is the X-axis component of the speed increment of the i-th control, Δv y is the Y-axis component of the speed increment of the i-th control, k is a coefficient; (124) Solve k, i.e. solve the vector for the current i-th control, where v 0x = v PAx , v 0y = v PAy , v Kx = v trAx , v Ky = v trAy , according to the above, iteratively each time the required velocity increment vector Δv i , finally resulting in a multi-pulse velocity increment ΔV M as follows: wherein v 0x is the X-axis component of the pre-control speed of the i-th time, v 0y is the Y-axis component of the pre-control speed of the i-th time, v Kx is the X-axis component of the post-control speed of the i-th time, v Ky is the Y-axis component of the post-control speed of the i-th time, Δv i is the control speed increment of the i-th time, ΔV M is the multi-pulse speed increment, t1~t K are the control times from the i-th to the K-th time, Δv1~Δv K are the control speed increments from the i-th to the K-th time.

3. The autonomous mission planning method for fast maneuvering imaging of spatially-skewed targets according to claim 1, wherein, After the N-1th control is implemented, the last control planning is performed, and the Lambert single-pulse transfer is used for control sequence planning, and the specific method is as follows: Step 31, satellite orbit and target orbit of the maneuver point and encounter point are found according to the post-control satellite orbit and the target orbit; Step 32, transfer velocity increment solving using Lambert equation; Step 33, the velocity increment is substituted into the high-precision orbit extrapolation model for encounter imaging calculation, if the imaging requirements are met, the velocity increment is output as the control result of the last task planning; otherwise, the encounter parameters are updated, and step 31 is returned to solve until the encounter imaging requirements are met.

4. The autonomous task planning method for fast maneuvering imaging of spatially-skewed targets according to claim 1, characterized in that, The imaging task planning in step four has the following specific method after the last control is completed: Step 41, the post-control satellite orbit and the target orbit are extrapolated to the vicinity of the orbit encounter point; Step 42, the relative distance, azimuth, elevation, and sun angle imaging parameters are calculated according to the satellite orbit and the target orbit, and the specific calculation method is as follows: a) relative distance Position of satellite in J2000 system R s = [R sx , R sy , R sz ], Position of target in J2000 system R t = [R tx , R ty , R tz ], Relative position vector of satellite and target ΔR st = R t - R s , Relative distance ΔR = |R t - R s |; b) azimuth and elevation where ΔR st_VVLH is the target position in the satellite VVLH coordinate system, is the conversion matrix from the J2000 system to the VVLH coordinate system, ΔR st_VVLHx is the X-axis component of the target position in the satellite VVLH coordinate system, ΔR st_VVLHy is the Y-axis component of the target position in the satellite VVLH coordinate system, θ_temp is the target azimuth process variable, and θ is the target azimuth, is the target elevation angle. b) sun angle where a is the solar angle, S i is the solar vector in the inertial system, AR st_J2000 is the position of the target relative to the satellite in the J2000 system; Step 43, the imaging parameters meeting the load constraints are selected.

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