Autonomous task planning method for rapid maneuvering rendezvous imaging of space different-plane targets

Abstract

The invention discloses an autonomous task planning method for rapid maneuvering rendezvous imaging of a space different-plane target, and the method comprises the steps: 1, multi-pulse control sequence planning: giving a multi-pulse control sequence according to a task demand and a measurement orbit demand, and autonomously carrying out the first-time control on a satellite; step 2, closed-loop control sequence planning: after the ith control is finished, combining the controlled satellite orbit and the target orbit to give a remaining (N-i)-time control sequence, and implementing the (i + 1)-time control; wherein i is larger than or equal to 1 and smaller than or equal to N-2, and the total control frequency is N; step 3, monopulse transfer planning: after N-1 times of control is finished, calculating the speed increment of the last time of control by using a Lambert transfer method in combination with the controlled satellite orbit and the target orbit, and implementing the Nth time of control; and step 4, imaging task planning. The influence of control errors and orbit errors is considered, a closed-loop control planning method is given, it can be guaranteed that each planning result is basically consistent with the first planning result, and the final intersection precision requirement is met.

Description

Technical Field

[0001] This patent involves the field of autonomous rendezvous of targets in different planes in space, specifically the rendezvous and imaging mission for targets in different planes in space. Through multiple on-board autonomous mission planning and implementation of corresponding control, it is ensured that autonomous rendezvous and imaging that meets mission requirements can be achieved in the presence of orbit determination errors and control errors. Background Art

[0002] When designing the rendezvous control for out-of-plane targets in space, the fuel optimization is taken into consideration. The general analysis method is to first solve the minimum single-pulse velocity increment that meets the rendezvous constraint. However, due to the maximum single jet time of the spacecraft, the single-pulse velocity increment needs to be decomposed into multiple pulses for control.

[0003] During mission execution, to ensure spacecraft orbital safety, orbit re-determination is required after control is complete. Due to control and orbit errors, continuing with the orbit control strategy from the start of mission planning will likely fail to meet rendezvous and imaging requirements. Therefore, re-planning is necessary after each control. Furthermore, during actual mission execution, to account for the impact of the tracking and control arc, the control timing of multiple planning phases should be kept consistent with the initial plan. Summary of the Invention

[0004] Aiming at the implementation ideas of actual engineering projects, an autonomous mission planning method for the rapid maneuverable rendezvous and imaging of space targets with different planes is proposed for the maneuverable rendezvous and imaging of space targets with different planes. Considering the influence of control error and orbit error, a closed-loop control planning method is given. This method can ensure that each planning result is basically consistent with the first planning and meets the final rendezvous accuracy requirements.

[0005] The technical solution of the present invention is: an autonomous mission planning method for rapid maneuvering rendezvous and imaging of targets in different planes of space, the specific steps of which are as follows:

[0006] Step 1: Multi-pulse control sequence planning: Based on the mission requirements and considering the orbit determination requirements, a multi-pulse control sequence is given and the first control is autonomously implemented on board.

[0007] Step 2: Closed-loop control sequence planning: After the i-th control is completed, the remaining Ni control sequences are given by combining the post-control satellite orbit and the target orbit, and the i+1-th control is implemented; where i ≥ 1 and i ≤ N-2, and the total number of controls is N.

[0008] Step 3: Single-pulse transfer planning: After the N-1 control rounds are completed, the Lambert transfer method is used to calculate the final control velocity increment based on the post-control satellite orbit and the target orbit, and the Nth control round is implemented.

[0009] Step 4. Imaging task planning: After the Nth control ends, perform imaging task planning and give the payload imaging parameters for rendezvous imaging.

[0010] Furthermore, the specific method for the multi-pulse control sequence planning described in Step 1 is as follows:

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

[0012] The Lambert theorem can be used to analyze the orbital transfer situation between any two points in space. Based on the Lambert equation, analyze and solve for the minimum velocity increment required in the spacecraft's cross-plane target rendezvous imaging task. The velocity increment is closely related to the semi-major axis of the transfer orbit. For the transfer problem between two points in space where the transfer time is not restricted, when the maneuvering point and the rendezvous point are determined, there will be an infinite number of transfer orbits passing through the two points. For the spacecraft, due to fuel constraints, what kind of semi-major axis of the transfer orbit makes the velocity increment passing through the two points the smallest becomes the focus of attention, that is, the minimum velocity increment transfer problem. To solve for the minimum velocity increment, it is necessary to solve the derivative of the velocity increment with respect to the semi-major axis of the transfer orbit. According to the derivative information, use the bisection method to solve for the minimum velocity increment.

[0013] (111) Set the search upper limit and the search lower limit of the transfer semi-major axis ; if , then set the initial , otherwise set the initial ; if , then there is no solution in the current transfer semi-major axis interval, otherwise set the initial ; where is the semi-major axis of the parking orbit, is the minimum value of the semi-major axis of the transfer ellipse;

[0014] (112) For any , let , calculate , , , where starts iterating from 0, that is , ; where is the transfer velocity increment, is the lower limit of the search for the transfer semi-major axis of the th group, is the upper limit of the search for the transfer semi-major axis of the th group, is the median of the transfer semi-major axis of the th group;

[0015] (113) If , let , ; If , let , , go to (112) and continue iteration;

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

[0017] Step 12: Decompose the multi-pulse velocity increment according to the measured orbit requirements and satellite control capabilities;

[0018] During the actual mission implementation, due to the existence of control errors and on-board extrapolation model errors, a single control is likely to cause the terminal error to accumulate too much. Generally, it is necessary to divide the control into multiple times to reduce the terminal error; and the actual thrust is not a pulse thrust. When the transfer velocity increment is large, the control duration exceeds the maximum jet time limit of a single control of the thruster, and it is also necessary to divide the control into multiple times to complete. Therefore, it is necessary to consider the problem of multi-pulse velocity increment decomposition under complex perturbations.

[0019] The multi-pulse velocity increment decomposition method is as follows: Divide the most fuel-efficient transfer velocity increment in the ideal case into times of control, and each control is carried out along the direction of . Then it can be known that the velocity increment of a single control satisfies ;

[0020] A closed orbit is defined as an orbit that is controlled at a certain point and can finally return to this point; Suppose the last control in times of control is the rendezvous control, and the previous times of control are all closed-loop controls. 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 moment; is the mean anomaly of the parking satellite at the maneuvering point; is the mean anomaly of the target satellite at the rendezvous point; and are the mean anomalies of points A and B on the transfer orbit; , , and are the semi-major axes of the parking orbit, the target orbit, the closed orbit after the i-th control, and the transfer orbit, respectively; is the number of orbits that the parking satellite runs on the parking orbit, is the number of orbits that the parking satellite runs on the i-th closed orbit, is the number of orbits that the target satellite runs on the target orbit;

[0023] The above formula can be rewritten as:

[0024]

[0025]

[0026] The semi-major axis of the closed orbit can be expressed as:

[0027]

[0028] Assume that the orbit control interval is at least 1 orbit, then 、 and take values of , , ; Solve using a numerical scheme, set a loop in the order of to solve the semi-major axis of the closed orbit ;

[0029] When the semi-major axis of the closed orbit is determined, the velocity increment required for each control can be solved. The solution steps are as follows:

[0030] (121) In the LVLH coordinate system, the velocity increment of the maneuver point A can be expressed as: , then the velocity before control is: , and the velocity after control is: ;

[0031] Among them, is the velocity increment of the maneuver point A, is the X-axis component of the velocity increment of the maneuver point, is the Y-axis component of the velocity increment of the maneuver point, is the velocity before control, is the X-axis component of the velocity increment before control, is the Y-axis component of the velocity increment before control, is the velocity after control, is the X-axis component of the velocity increment after control, is the Y-axis component of the velocity increment after control;

[0032] (122) When is determined, the The magnitude of the velocity after the secondary control is , and the vector expression is set as: , , two variables and one equation, has infinitely many solutions. In this paper, the equal ratio method is used for solution;

[0033] Among them, is the magnitude of the velocity after the -th secondary control, is the radius vector of the maneuvering point A, is the semi-major axis of the closed orbit, is the gravitational constant of the earth, is the velocity increment after the -th secondary control, is the X-axis component of the velocity increment after the -th secondary control, The Y-axis component of the velocity increment after the -th secondary control;

[0034] (123) Set The vector direction of to be consistent with the vector direction of , then the vector can be expressed as: , then the condition that

[0035]

[0036] Among them, is the velocity before control, is the velocity after control, is the velocity increment of the -th control, is the X-axis component of the velocity increment of the -th control,

[0037] (124) Solve That is, solve the vector of the current -th control. Among them, , , and the required velocity increment vector for each iteration can be obtained according to the above. Finally, the multi-pulse velocity increment is as follows:

[0038]

[0039] Among them, is the X-axis component of the velocity before the -th control, is the X-axis component of the controlled speed at the i-th time, is the Y-axis component of the controlled speed at the i-th time, is the control speed increment at the i-th time, is the multi-pulse speed increment, are the control times from the i-th time to the K-th time respectively, is the control speed increment from the i-th time to the K-th time.

[0040] Furthermore, after the last control is implemented, according to the multi-pulse speed increment calculated in Step 1 and the controlled satellite orbit and the target orbit, the closed-loop control sequence planning described in the next Step 2 is carried out. The specific method is as follows:

[0041] Step 21: Extrapolate the controlled satellite orbit and the target orbit to the vicinity of the current control timing planned last time, update the semi-major axis of the pre-control orbit and the pre-control speed planned this time, and calculate the remaining control speed increment according to the closed orbit semi-major axis ;

[0042] Step 22: Calculate the rendezvous time error after applying the remaining multi-pulse speed increment. The time for the spacecraft to reach the rendezvous point after applying the multi-pulse speed increment is , and the time for the space debris to reach the rendezvous point is , 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, correct the first control speed increment of the multi-pulse control sequence. 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 obtain the semi-major axis of the first control that needs to be adjusted The calculation is as follows:

[0044]

[0045]

[0046] Among them, is the adjusted semi-major axis of the first control, is the unadjusted semi-major axis of the first control, 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 orbits the satellite runs at the semi-major axis of the first closed orbit; is the derivative of the nodal period with respect to the semi-major axis, is the J2 term of the Earth's non-spherical perturbation, is the Earth's gravitational constant, is the Earth's equatorial radius, is the orbital inclination.

[0047] Furthermore, for the single-pulse transfer planning described in Step 3, after the implementation of the (N - 1)-th control is completed, the last control planning is carried out. The Lambert single-pulse transfer is used for the control sequence planning. The specific method is as follows:

[0048] Step 31: Find the satellite orbits and target orbits of the maneuver point and the rendezvous point according to the post-control satellite orbit and the target orbit;

[0049] Step 32: Use the Lambert equation to solve for the transfer velocity increment;

[0050] Step 33: Substitute the velocity increment into the high-precision orbit extrapolation model for rendezvous imaging calculation. If the imaging requirements are met, output the velocity increment as the control result of the last mission planning; otherwise, update the rendezvous parameters and return to Step 31 for solution until the rendezvous imaging requirements are met.

[0051] Furthermore, for the imaging mission planning described in Step 4, after the last control ends, the imaging mission planning is carried out. The specific method is as follows:

[0052] Step 41: Extrapolate the post-control satellite orbit and the target orbit to the vicinity of the orbit rendezvous point;

[0053] Step 42: Calculate imaging parameters such as the relative distance, azimuth angle, elevation angle, and solar angle according to the satellite orbit and the target orbit. The specific calculation method is as follows:

[0054] a) Relative distance

[0055] The position of the satellite in the J2000 system , the position of the target in the J2000 system , the relative position vector between the satellite and the target , the relative distance ;

[0056] b) Azimuth angle and elevation angle

[0057] The azimuth angle and elevation angle generally refer to the azimuth and elevation of the target relative to the satellite in the VVLH coordinate system. And the relative position vector calculated above is expressed in the J2000 coordinate system. In order to obtain the azimuth and elevation in the VVLH system, it is necessary to The vector is transformed into the satellite VVLH coordinate system, and the formula is as follows (where is the transformation matrix from the J2000 coordinate system to the VVLH coordinate system)

[0058]

[0059]

[0060]

[0061]

[0062] Among them, is the target position in the satellite VVLH coordinate system, is the transformation matrix from the J2000 coordinate system to the 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 process quantity of the target azimuth angle, is the target azimuth angle, is the target elevation angle;

[0063] b) Solar angle

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

[0065]

[0066] Among them, is the solar included angle, is the solar vector in the inertial coordinate system, is the position of the target relative to the satellite in the J2000 coordinate system;

[0067] Step 43: Screen the imaging parameters that meet the payload constraints.

[0068] The beneficial effects of the present invention are: a method for autonomous mission planning for rapid maneuvering rendezvous imaging of space non-coplanar targets is proposed. Considering the influence of control errors and orbit errors, a closed-loop control planning method is given. This method can ensure that each planning result is basically the same as the first planning and meets the requirements of the final rendezvous accuracy. Specific embodiments

[0069] In order to make the physical laws and effects described in this patent easy to understand, the following combines specific embodiments to further elaborate on this patent.

[0070] I. Task scenario design

[0071] This patent implements onboard autonomous mission planning and calculation for the maneuverable rendezvous and imaging mission of out-of-plane targets in space. The mission scenario is designed as follows.

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

[0073] Table 1

[0074]

[0075] (2) Spacecraft payload operating constraints are shown in Table 2.

[0076] Table 2

[0077]

[0078] (3) Intersection imaging requirements

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

[0080] 2. Multi-pulse control sequence planning

[0081] According to step 1, the multi-pulse control sequence planning of the spacecraft's rapid maneuvering of the out-of-plane target is carried out. The results are shown in Table 3. The first planning results of the spacecraft's out-of-plane rendezvous imaging of the target.

[0082] Table 3

[0083]

[0084] According to the results of the first mission planning, the first control - V1 was implemented at the intersection point of the two satellites, achieving a 21.063 km adjustment of the semi-major axis of the spacecraft's orbit and a change of 0.002336 in eccentricity.

[0085] 3. Closed-loop control sequence planning

[0086] After the first control, the orbital parameters of the spacecraft and the target change. During the closed-loop control sequence planning, the orbits of the spacecraft and the target are shown in Table 4.

[0087] Table 4

[0088]

[0089] According to step 2, the closed-loop control sequence planning is performed for the track in the table above. The results are shown in Table 5.

[0090] Table 5

[0091]

[0092] According to the control results obtained from the second plan, the second orbit control - V2 is implemented at the intersection point of the two satellites, achieving a reduction in the semi - major axis of the orbit of this satellite by 6.001 km and a change in eccentricity of 0.000656.

[0093] IV. Single - pulse transfer plan

[0094] After the second control is completed, the orbital parameters of the spacecraft and the target change. When making the single - pulse transfer plan, the orbits of the spacecraft and the target are as shown in Table 6.

[0095] Table 6

[0096]

[0097] According to the single - pulse transfer plan for the orbits in the above table in Step 3, the results are shown in Table 7.

[0098] Table 7

[0099]

[0100] According to the control results obtained from the third plan, the third orbit control - V3 is implemented at the intersection point of the two satellites, achieving a reduction in the semi - major axis of the orbit of this satellite by 1.813 km and a change in eccentricity of 0.00021.

[0101] V. Imaging mission plan

[0102] After the third control (the last control) is completed, the calculation of the imaging mission plan for the out - of - plane rendezvous is started. The orbital parameters of the spacecraft and the target are as shown in Table 8.

[0103] Table 8

[0104]

[0105] According to Step 4, the imaging mission plan is made for the orbit after the third control. The imaging parameters are shown in Table 9.

[0106] Table 9

[0107] 。

[0108] The above - mentioned are only the preferred embodiments of the present invention. It should be pointed out that for those of ordinary skill in the art in this technical field, without departing from the principle of the present invention, several improvements and refinements can still be made, and these improvements and refinements should also be regarded as the protection scope of the present invention.

Claims

1. An autonomous mission planning method for rapid maneuvering rendezvous imaging of spatially non-coplanar targets, characterized in that, The specific steps are as follows: Step 1. Multi-pulse control sequence planning: For the task requirements and considering the orbit determination requirements, a multi-pulse control sequence is given, and the first control is autonomously implemented on the satellite; Step 2. Closed-loop control sequence planning: After the i-th control is completed, combined with the post-control satellite orbit and the target orbit, the remaining N - i control sequences are given, and the (i + 1)-th control is implemented; where i ≥ 1 and i ≤ N - 2, and the total number of controls is N times; Step 3. Single-pulse transfer planning: After N - 1 controls are completed, combined with the post-control satellite orbit and the target orbit, the Lambert transfer method is used to calculate the velocity increment of the last control, and the N-th control is implemented; Step 4. Imaging task planning: After the N-th control is completed, the imaging task planning is carried out, and the payload imaging parameters are given for rendezvous imaging.

2. The autonomous mission planning method for rapid maneuvering rendezvous imaging of space non-coplanar targets according to claim 1, wherein The specific method of the multi-pulse control sequence planning described in Step 1 is as follows: Step 11. Solve for the minimum velocity increment based on the Lambert equation; (111) Set the search upper limit and search lower limit of the transfer semi-major axis ; If , then set the initial , otherwise set the initial ; If , then there is no solution in the current transfer semi-major axis interval, otherwise set the initial ; If , then there is no solution in the current transfer semi-major axis interval, otherwise set the initial ; Among them, is the semi-major axis of the parking orbit, is the minimum value of the semi-major axis of the transfer ellipse; For any , let , calculate , , , where iterates from 0, i.e., , ; where, is the transfer speed increment, is the lower limit of the transfer semi-major axis search for the th group, is the upper limit of the transfer semi-major axis search for the th group, is the median of the transfer semi-major axis for the th group; (113) If , let , ; If , let , , go to (112) and continue iteration; If or , stop the calculation, take , and the minimum speed increment is ; Step 12. Decompose the multi-pulse velocity increment according to the orbit determination requirements and the satellite control capabilities; The multi-pulse velocity increment decomposition method is as follows: the fuel-optimal transfer velocity increment under ideal conditions is divided into multiple times of control, and each control is carried out along the direction of, then it can be known that the velocity increment of a single control satisfies ; A closed orbit is defined as an orbit that is controlled at a certain point and can eventually return to this point; assume the last control in the times of control is rendezvous control, and the previous times of control are all closed-loop controls. According to the time equation, the following relationship can be obtained: Wherein, and are the mean anomaly of the parking satellite and the target satellite at the initial moment; is the mean anomaly of the parking satellite at the maneuvering point; is the mean anomaly of the target satellite at the rendezvous point; and are the mean anomaly of points A and B on the transfer orbit; , , and are the semi-major axes of the parking orbit, the target orbit, the closed orbit after the i-th control, and the transfer orbit, respectively; is the number of orbits of the parking satellite running on the parking orbit, is the number of orbits of the parking satellite running on the i-th closed orbit, is the number of orbits of the target satellite running on the target orbit; The above formula can be rewritten as: The semi-major axis of the closed orbit can be expressed as: Assume that the orbit control interval is at least 1 orbit, then , and take values of , , ; Solve using a numerical scheme, set up a loop in the order of to solve for the semi-major axis of the closed orbit ; When the semi-major axis of the closed orbit is determined, the velocity increment required for each control can be solved. The solution steps are as follows: (121) In the LVLH coordinate system, the velocity increment of the maneuvering point A can be expressed as: Then the velocity before control is: And the velocity after control is: ; Among them, is the velocity increment of the maneuvering point A, is the X-axis component of the velocity increment of the maneuvering point, is the Y-axis component of the velocity increment of the maneuvering point, is the pre-control velocity, is the X-axis component of the pre-control velocity increment, is the Y-axis component of the pre-control velocity increment, is the post-control velocity, is the X-axis component of the post-control velocity increment, is the Y-axis component of the post-control velocity increment; (122) When is determined, the magnitude of the speed after the th control is , and the vector expression is set as: , , and the equal ratio method is used for solution; Among them, is the magnitude of the velocity after the th control, is the radius vector of the maneuvering point A, is the semi-major axis of the closed orbit, is the Earth's gravitational constant, is the th velocity increment after control, is the X-axis component of the th velocity increment after control, the th Y-axis component of the velocity increment after control; (123) Setting The vector direction of is consistent with the vector direction of The vector can be expressed as: , then from The conditions to be satisfied are: Among them, is the pre-control speed, is the post-control speed, is the speed increment of the i-th control, is the X-axis component of the speed increment of the i-th control, is the Y-axis component of the speed increment of the i-th control, is the coefficient; Solve That is, solve for the vector of the current -th control, where , , according to the above-mentioned velocity increment vector required for each iteration , and finally obtain the multi-pulse velocity increment as follows: Among them, is the X-axis component of the speed before control for the i-th time, is the Y-axis component of the speed before control for the i-th time, is the X-axis component of the speed after control for the i-th time, is the Y-axis component of the speed after control for the i-th time, is the control speed increment for the i-th time, is the multi-pulse speed increment, are the control times from the i-th time to the K-th time respectively, is the control speed increment from the i-th time to the K-th time.

3. The autonomous mission planning method for rapid maneuvering rendezvous imaging of spatially non-coplanar targets according to claim 1, characterized in that, After the previous control is completed, according to the multi-pulse velocity increment calculated in Step 1 and the post-control satellite orbit and the target orbit, the closed-loop control sequence planning described in the next Step 2 is carried out. The specific method is as follows: Step 21: Extrapolate the post-control orbit and the target orbit of the satellite to the vicinity of the previously planned control timing for this time, update the semi-major axis of the pre-control orbit and the pre-control velocity planned for this time, and calculate the remaining control velocity increment based on the semi-major axis of the closed orbit ; Step 22, calculate the rendezvous time error after applying the remaining multi-pulse velocity increment. The spacecraft applies the multi-pulse velocity increment and the time to reach the rendezvous point is . The time for the space debris to reach the rendezvous point is . Then the rendezvous time difference between the spacecraft and the space debris is ; Step 23: If the rendezvous time error is small but does not meet the rendezvous time requirement, correct the first control velocity increment of the multi-pulse control sequence. If and , is the rendezvous time difference between the spacecraft and the space debris, is the threshold of the rendezvous time difference, is the iteration exit condition of the rendezvous time difference, then according to the rendezvous time difference obtain the semi-major axis of the first control that needs to be adjusted The calculation is as follows: Among them, 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 intersection point per revolution, is the intersection time difference, is the number of orbits of the satellite running on the semi-major axis of the first closed orbit; is the derivative of the nodal period with respect to the semi-major axis, is the J2 term of the Earth's non-spherical perturbation, is the Earth's gravitational constant, is the Earth's equatorial radius, is the orbital inclination.

4. The autonomous mission planning method for rapid maneuvering rendezvous imaging of spatially skew targets according to claim 1, wherein For the single-pulse transfer planning described in Step 3, after the (N - 1)-th control is completed, the last control planning is carried out, and the Lambert single-pulse transfer is used for the control sequence planning. The specific method is as follows: Step 31. Find the satellite orbits and target orbits of the maneuver point and the rendezvous point according to the post-control satellite orbit and the target orbit; Step 32. Solve for the transfer velocity increment using the Lambert equation; Step 33. Substitute the velocity increment into the high-precision orbit extrapolation model for rendezvous imaging calculation. If the imaging requirements are met, the velocity increment is output as the control result of the last task planning; otherwise, update the rendezvous parameters and return to Step 31 for solution until the rendezvous imaging requirements are met.

5. The autonomous mission planning method for rapid maneuvering rendezvous imaging of spatially non-coplanar targets according to claim 1, wherein For the imaging task planning described in Step 4, after the last control is completed, the imaging task planning is carried out. The specific method is as follows: Step 41. Extrapolate the post-control satellite orbit and the target orbit to the vicinity of the orbit rendezvous point; Step 42. Calculate the imaging parameters such as the relative distance, azimuth angle, elevation angle, and solar angle according to the satellite orbit and the target orbit. The specific calculation methods are as follows: a) Relative distance The position of the satellite in the J2000 coordinate system , the position of the target in the J2000 coordinate system , the relative position vector between the satellite and the target , the relative distance ; b) Azimuth angle and elevation angle wherein, is the target position in the satellite VVLH coordinate system, is the transformation matrix from the J2000 coordinate system to the 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 process quantity of the target azimuth, is the target azimuth, is the target elevation angle; b) Solar angle Among them, is the solar angle, is the solar vector in the inertial system, is the position of the target relative to the satellite in the J2000 system; Step 43. Screen the imaging parameters that meet the payload constraints.

Images