A Human-in-the-Loop Non-Cooperative Target Proximity Tracking Control Method

Abstract

The present invention relates to a human-in-the-loop non-cooperative target approaching tracking control method, belonging to the field of spacecraft control technology. By adding the control input of a ground operator to the automatic control process during the approaching process of a chasing spacecraft in space, the human-in-the-loop approaching rendezvous control of the spacecraft is realized, ensuring that the chasing spacecraft can still implement a safe and reliable rendezvous approaching strategy under the condition that there are deviations in the dynamic and orbital information of the target spacecraft during the approaching process, and ensuring the safety and reliability of the approaching process. It breaks through the operation interaction method based on the image tracking of the target spacecraft and solves the problem of generating a spacecraft maneuvering strategy that integrates the active perception of the operator under the condition that the state of the target spacecraft is unknown.

Description

Technical Field

[0001] The present invention belongs to the technical field of spacecraft control, and particularly relates to a method for approaching and tracking control of non-cooperative targets with human-in-the-loop. Background Technique

[0002] The "Geosynchronous Space Situational Awareness Program" (GSSAP) in the United States uses the relative drift between satellites operating in low Earth synchronous orbit and geosynchronous orbit for orbital patrol, and can detect, catalog, and precisely reconnoiter spacecraft of interest. When necessary, it can also approach geosynchronous orbit targets through orbital maneuvers for close-up inspection to obtain high-definition images of the targets. Currently, the autonomous ability of spacecraft is weak, and it is unable to independently detect and avoid attacks by GSSAP satellites. Therefore, adopting a "human-in-the-loop" scheme for decision-making and operation interaction has advantages such as sufficient prior foundation, short implementation cycle, and fast response speed, and is the preferred solution to solve the aforementioned problems. Space offensive and defensive confrontation is the main battlefield for future major power games. Establishing a space offensive and defensive system with human participation can enhance the flexibility of the system and improve technical generality.

[0003] Facing possible space security confrontation problems, how to determine the flight purpose of enemy high-dynamic maneuvering spacecraft and achieve early defense requires further research. Therefore, there is potential research value in conducting active tracking research on target spacecraft that may be aggressive. The present invention is based on possible space security problems, and in response to the need to improve the game maneuvering ability of our spacecraft, conducts research on "human-in-the-loop" spacecraft game control, breaks through the operation interaction method based on image tracking of target spacecraft, and solves the problem of generating spacecraft maneuvering strategies that integrate the active perception of operators under the condition that the state of the target spacecraft is unknown. Summary of the Invention

[0004] The technical problem to be solved by the present invention is:

[0005] Aiming at the problem of tracking non-cooperative target spacecraft during the security defense process and improving the game maneuvering ability of our spacecraft, the present invention provides a control method for approaching and tracking non-cooperative targets with human-in-the-loop.

[0006] To solve the above technical problems, the technical solution adopted by the present invention is:

[0007] A method for approaching and tracking control of non-cooperative targets with human-in-the-loop, characterized by including the following steps:

[0008] Step 1: Give the position x of the chaser spacecraft in the geocentric coordinate system at the initial moment P , the position x of the target spacecraft e , and give the conversion relationship between the orbital coordinate system and the geocentric coordinate system;

[0009] Step 2: Obtain the positions of the two spacecraft in the orbital coordinate system through the transformation relationship, and give the position x of the virtual target point at the current moment. d = x e , and establish a relative motion dynamics model.

[0010] Step 3: Collect the current joint angles of the hand controller, calculate the current end position through forward kinematics, and calculate the direct control quantity u1 of the chasing spacecraft given by the operator through the hand controller at the current moment.

[0011] Step 4: Calculate the position increment of the chasing spacecraft in the orbital coordinate system, obtain the position increment in the geocentric coordinate system through coordinate transformation, and calculate the automatic control quantity u2 of the chasing spacecraft to automatically approach the target spacecraft.

[0012] Step 5: Finally, obtain the control quantity u of the chasing spacecraft as u = u1 + u2, input the control quantity into the relative motion dynamics model established in Step 2, and solve to obtain the position of the chasing spacecraft under the control input; and within each update cycle, update the position of the virtual target point of the target spacecraft; repeat the process of Steps 3 - 5 to obtain the final position of the chasing spacecraft.

[0013] Further technical solution of the present invention: The transformation relationship between the orbital coordinate system and the geocentric coordinate system in Step 1 is:

[0014]

[0015] where r0 = [r 0x , r 0y , r 0z T is the position coordinate of the origin of the entrained inertial coordinate system in the geocentric equatorial inertial coordinate system, A is a matrix parameter, [x L , y L , z L T is the position of the spacecraft in the orbital coordinate system, and [x r1 , y r1 , z r1 T is the position of the spacecraft in the geocentric coordinate system.

[0016] Further technical solution of the present invention: The relative motion dynamics model in Step 2 is:

[0017]

[0018] where [x, y, z] is the difference between the virtual position of the target spacecraft and the position of the chasing spacecraft in the orbital coordinate system of the chasing spacecraft, u x , u y , u z ​​​For the control input of the chaser spacecraft, ω is the angular velocity of the circular orbit at the initial moment of the chaser spacecraft's period. Respectively represent the second-order derivatives of x, y, and z. Respectively represent the first-order derivatives of x and y.

[0019] A further technical solution of the present invention: Step 3 is specifically as follows:

[0020] Obtain the current end position (x F , y F , z F ) through forward kinematics:

[0021] x F = -sinθ1(L2sinθ3 + L1cosθ2)

[0022] y F = -L2cosθ3 + L1sinθ2 + L3

[0023] z F = L2cosθ1sinθ3 + L1cosθ1cosθ2 - L4

[0024] In the formula, θ1, θ2, and θ3 are the rotation angles of the first three joints respectively, L1 and L2 are the lengths of link 1 and link 2 respectively, L3 is the height distance difference between the handheld rod and the first joint point, and L4 is the horizontal distance difference between the handheld rod and the first joint point;

[0025] Calculate the direct control quantity u1 = [u x , u y , u z given by the operator to the chaser spacecraft through the hand controller: T :

[0026]

[0027] Among them, k dx , k dy , k dz are gain coefficients, and (x zero , y zero , z zero ) is the neutral position point.

[0028] A further technical solution of the present invention: Step 4 is specifically as follows:

[0029] Obtain the position increment Δx of the chaser spacecraft in the orbital coordinate system through calculation:

[0030] Δx = [x dT[k] , y dT[k] , z dT[k] T - [xdT[k-1] , y dT[k-1] , z dT[k-1] T

[0031] Through coordinate transformation, the position increment Δρ of the chasing spacecraft in the geocentric coordinate system can be obtained as Δρ = A T Δx, and the actual control quantity is calculated as u2 = -K P Δρ - K D Δρ′, where K P , K D respectively represent the gain coefficients of the proportional and differential links in PD control, and Δρ′ represents the change rate of the position increment in the geocentric equatorial inertial coordinate system.

[0032] A computer system, characterized by comprising: one or more processors, and a computer-readable storage medium for storing one or more programs, wherein, when the one or more programs are executed by the one or more processors, the one or more processors are caused to implement the above method.

[0033] A computer-readable storage medium, characterized by storing computer-executable instructions which are used to implement the above method when executed.

[0034] The beneficial effects of the present invention are as follows:

[0035] A human-in-the-loop non-cooperative target approaching tracking control method provided by the present invention, under the condition that the imaging conditions of the target spacecraft are updated stage by stage by the optical payload, enables the operator to observe the imaging of the target spacecraft on the ground, and uses the interaction device to generate the orbital maneuver strategy of the chasing spacecraft, so as to realize the active tracking of the target spacecraft by "human in the loop".

[0036] By adding the control input of the ground operator to the automatic control process during the approaching process of the chasing spacecraft in space, the human-in-the-loop spacecraft approaching and rendezvous control is realized, ensuring that the chasing spacecraft can still implement a safe and reliable rendezvous and approaching strategy under the condition that there are deviations in the dynamics and orbital information of the target spacecraft during the approaching process, and ensuring the safety and reliability of the approaching process. It breaks through the operation and interaction method based on the image tracking of the target spacecraft, and solves the problem of generating the spacecraft maneuver strategy by fusing the active perception of the operator under the condition that the state of the target spacecraft is unknown. Description of the Drawings

[0037] The drawings are only for the purpose of illustrating specific embodiments and are not considered to be a limitation of the present invention. Throughout the drawings, the same reference signs denote the same components.

[0038] Figure 1 Schematic diagram of the geocentric coordinate system and the orbital coordinate system;

[0039] ​Figure 2 Schematic diagram of the structure of the hand controller;

[0040] Figure 3 Structure diagram of the spacecraft approach tracking system;

[0041] Figure 4 Positions of the chaser star and the target star on the x-axis in the geocentric coordinate system;

[0042] Figure 5 Positions of the chaser star and the target star on the y-axis in the geocentric coordinate system;

[0043] Figure 6 Positions of the chaser star and the target star on the z-axis in the geocentric coordinate system. Specific implementation manners

[0044] In order to make the objectives, technical solutions and advantages of the present invention clearer and more understandable, the present invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not used to limit the present invention. In addition, the technical features involved in the various embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.

[0045] The embodiment of the present invention provides a human-in-the-loop non-cooperative target approach tracking control method. By adding the control input of the ground operator to the automatic control process during the approach process of the chaser spacecraft in space, the human-in-the-loop operation interaction of the spacecraft approach tracking based on the image tracking of the target spacecraft is realized, ensuring that the chaser spacecraft can still implement a safe and reliable rendezvous approach strategy in the case of deviations in the dynamics and orbital information of the target spacecraft during the approach process, and ensuring the safety and reliability of the approach process. The method includes the following steps:

[0046] Step 1: Give the position x of the chaser spacecraft at the initial moment in the geocentric coordinate system P , the position x of the target spacecraft e , and give the conversion relationship between the orbital coordinate system and the geocentric coordinate system.

[0047] Step 2: Obtain the positions of the two spacecraft in the orbital coordinate system through the conversion relationship, and give the position x of the virtual target point at the current moment d = x e , and establish a relative motion dynamics model.

[0048] Step 3: Collect the current joint angles of the hand controller, calculate the current end position through forward kinematics, and calculate the direct control quantity u1 of the chaser spacecraft given by the operator through the hand controller at the current moment.

[0049] Step 4: Calculate the position increment of the chaser spacecraft in the orbital coordinate system, obtain the position increment in the geocentric coordinate system through coordinate transformation, and calculate the automatic control quantity u2 given by the chaser spacecraft to automatically approach the target spacecraft.

[0050] Step 5: Finally, obtain the control quantity u = u1 + u2 of the chaser spacecraft, and input the control quantity into the relative motion dynamics model established in Step 2. u = [u x , u y , u z T , and solve to obtain the position of the chaser spacecraft under the control input. And within each update cycle, update the position of the virtual target point of the target spacecraft. Repeating the process of Steps 3 - 5, the final position result of the chaser spacecraft can be obtained.

[0051] The above steps are as follows:

[0052] Step 1: Define an orbital coordinate system with the origin at the centroid of the chaser spacecraft, and establish the coordinate transformation relationship between the chaser spacecraft and the target spacecraft in the geocentric equatorial inertial coordinate system and the coordinate system of the chaser spacecraft's orbit (LVLH).

[0053] Establish the spacecraft orbital coordinate system Oxyz (S o ), whose origin is at the centroid O of the chaser spacecraft. The x-axis points along the direction from the earth's center to the chaser spacecraft, the y-axis is perpendicular to the x-axis in the orbital plane and points in the direction of the chaser spacecraft's velocity, and the z-axis forms a right-handed orthogonal coordinate system with the x-axis and y-axis and is parallel to the normal of the orbital plane.

[0054] The definitions and relationships between the geocentric equatorial inertial coordinate system and the spacecraft orbital coordinate system (LVLH) are as Figure 1 shown. In the figure, Ω, ω, and θ are respectively the right ascension of the ascending node, the argument of perigee, and the true anomaly of the chaser spacecraft.

[0055] Coordinate system transformation relationship

[0056] 1. The transformation relationship between the non-rotating inertial coordinate system and the geocentric equatorial inertial coordinate system

[0057] There is only a translational relationship between the non-rotating inertial coordinate system and the geocentric equatorial inertial coordinate system. Therefore, a non-rotating inertial coordinate system is established at a certain moment position of the chaser spacecraft. The origin of the non-rotating inertial coordinate system of the chaser spacecraft is at the centroid of the chaser spacecraft. Assuming this point as the origin, a coordinate system with all axes parallel and in the same direction as the geocentric inertial coordinate system is established and named Ox1y1z1. This coordinate system can be obtained by translating the geocentric inertial coordinate system, and the translation amount is the coordinate value of the origin in this coordinate system. Let r1 and r2 respectively represent the same position point in the geocentric inertial coordinate system O i x i y​i z i (S i ) under and the position coordinates in the entrained inertial coordinate system Ox1y1z1, the conversion relationship between the two can be expressed as

[0058] r2 = r1 - r0 (1)

[0059] In the formula, r0 represents the position coordinates of the origin of the entrained inertial coordinate system in the inertial coordinate system.

[0060] 2. Conversion relationship between the entrained inertial coordinate system and the chasing spacecraft orbit coordinate system

[0061] The orbit coordinate system can be obtained by rotating the entrained inertial coordinate system. The reference coordinate position of the chasing spacecraft in the orbit coordinate system can be obtained through transformation to its position in the geocentric inertial coordinate system.

[0062] First, assume the position (X, Y, Z) in the entrained inertial coordinate system Ox1y1z1 T , through the following rotation transformation:

[0063] Ox1y1z1 rotates by an angle Ω around the z-axis to obtain Ox2y2z2, and the transformation matrix between these two coordinate systems is

[0064]

[0065] Ox2y2z2 rotates by an angle i around the x-axis to obtain Ox3y3z3, and the conversion relationship between these two coordinate systems is

[0066]

[0067] Assume that the perigee of the circular orbit is at the ascending node position, and the initial position of the spacecraft is at the ascending node. Ox3y3z3 rotates by an angle f around the z-axis, where f is the angle corresponding to the position where the chasing spacecraft establishes the coordinate system, to obtain the orbit coordinate system. The conversion relationship between the two coordinate systems is

[0068]

[0069] Combining the above transformations, the direct conversion relationship between the orbit coordinate system Oxyz and the reference coordinate system Ox1y1z1 is

[0070]

[0071] In the formula, X, Y, Z represent the coordinate positions in the entrained inertial coordinate system, [x L , y L , z L represents the position of a point in the spacecraft orbit coordinate system.

[0072] If the components in the orbital coordinate system are known and it is necessary to determine its vector components in the reference coordinate system, then utilize the property that the inverse matrix of the orthogonal transformation between the two coordinate systems is equal to its transpose matrix, i.e., A -1 = A T to obtain

[0073]

[0074] 3. Transformation relationship between the orbital coordinate system of the chasing spacecraft and the geocentric equatorial inertial coordinate system

[0075] According to the above two transformation relationships, the transformation relationship between the orbital coordinate system of the chasing spacecraft and the geocentric equatorial inertial coordinate system can be obtained, and the positions of the two spacecraft in the orbital coordinate system of the chasing spacecraft are updated to the geocentric inertial coordinate system. Assume that the position of a point in the spacecraft orbital coordinate system is [x L , y L , z L T , and assume that the position coordinates of the origin of the attached inertial coordinate system in the geocentric equatorial inertial coordinate system are r0 = [r 0x , r 0y , r 0z T . The position coordinates of this point in the spacecraft orbital coordinate system in the geocentric equatorial inertial coordinate system can be obtained as follows:

[0076]

[0077] Step 2: Establish the relative motion dynamics model of the chasing spacecraft in the orbital coordinate system

[0078] In the geocentric equatorial inertial coordinate system, at the current time t, the position of the chasing satellite is r C , and the position of the target satellite is r T . The positions of the two satellites respectively satisfy the orbital dynamics

[0079]

[0080] where r represents the position vector of the spacecraft in the geocentric equatorial inertial coordinate system, r represents the position scalar of the spacecraft in the geocentric equatorial inertial coordinate system, which is the distance from the spacecraft to the origin. u represents the control input of the spacecraft, expressed as u = [u x , u y , u z T , μ is the gravitational parameter, and for the Earth its value is: μ = 3.986012×10 3 km 3 / s 2 .

[0081] ​​​Assume that the relative position vector of the chaser spacecraft with respect to the target spacecraft is ρ. By subtracting the orbital dynamics equations of the two spacecraft, the relative motion dynamics equation of the target spacecraft with respect to the chaser spacecraft in the geocentric equatorial inertial coordinate system can be obtained as follows:

[0082]

[0083] where, u T represents the control input of the chaser spacecraft, and u C represents the control input of the target spacecraft, and r T represents the position vector of the chaser spacecraft in the geocentric equatorial inertial coordinate system, and r C represents the position vector of the target spacecraft in the geocentric equatorial inertial coordinate system.

[0084] According to the derivative relationship between the geocentric equatorial inertial coordinate system and the spacecraft orbital coordinate system, we have

[0085]

[0086] where, ω represents the orbital angular velocity of the chaser spacecraft. In the spacecraft orbital coordinate system, ω = [0, 0, ω] T .

[0087] In the process of this method, since the target spacecraft is a non-cooperative target, its control quantity is unknown, and its position is updated periodically through sensors and ground information. Therefore, in this scheme, its periodically updated position is set as the virtual target point, and the control target of the chaser spacecraft is to approach the virtual target point in each period to achieve tracking. Therefore, the motion trajectory of the chaser spacecraft in each period can be approximately regarded as a part of a circular orbit. Therefore, assuming that the chaser spacecraft operates on a circular orbit, its angular velocity ω can be obtained as constant. Therefore where, at the beginning of each period, the orbital angular velocity ω needs to be updated, and the relative motion dynamics equation when the orbit of the chaser spacecraft is a circular orbit can be obtained. Equation (9) can be written as:

[0088]

[0089] On this basis, since the actual position coordinate information of the target spacecraft cannot be accurately obtained in the actual process, the virtual coordinate point of the target spacecraft, that is, the virtual target point, is introduced and regarded as the expected position r d of the chaser spacecraft, replacing the actual position of the target spacecraft in the moving coordinate system. The above equation can be written as:

[0090]

[0091] where, ρ is the difference between the current actual positions of the two spacecraft. By simplification, the above equation can be written as:

[0092]

[0093] Among them, [x, y, z] is the difference between the virtual position of the target spacecraft and the position of the chaser spacecraft in the orbital coordinate system of the chaser spacecraft. ω is the angular velocity of the circular orbit at the initial moment of the chaser spacecraft's period. u x 、u y 、u z are the control inputs of the chaser spacecraft.

[0094] Step 3: Determine the calculation and conversion relationship of the hand controller control input

[0095] In the present invention, through human-computer interaction, the operator can, according to the position of the target spacecraft in the visual field of the front camera of the chaser spacecraft, send a control input signal to the control system through the hand controller to remotely operate and control the chaser spacecraft. The specific implementation method is that the operator controls the end position of the hand controller, obtains the control quantity input by the operator to the spacecraft through the force-position mapping relationship, and finally converts the control quantity into the control quantity required to be applied in each direction on the spacecraft body through coordinate conversion, so as to realize the close tracking of the non-cooperative target by the chaser spacecraft with the human in the loop.

[0096] The hand controller of the present invention adopts a Phantom Omni hand controller, as Figure 2 shown. Its first three joints realize three-dimensional space positioning, and the last three joints realize attitude adjustment. The hand-held end of this hand controller has a universal joint, which intersects the rotation axes in three directions at one point to achieve decoupling of position and attitude. Therefore, the position of the end of the hand controller is not affected by the attitude.

[0097] Since the consideration of attitude is not involved in the present invention, only the first three joints of the hand controller are analyzed. According to the structure diagram, the forward kinematic equation of this hand controller is:

[0098] x F =-sinθ1(L2sinθ3+L1cosθ2) (13)

[0099] y F =-L2cosθ3+L1sinθ2+L3 (14)

[0100] z F =L2cosθ1sinθ3+L1cosθ1cosθ2-L4 (15)

[0101] In the above formula, x F , y F , z FThey are the positions of point F on the x-axis, y-axis, and z-axis in the coordinate system respectively. θ1, θ2, and θ3 are the rotation angles of the first three joints respectively. L1 and L2 are the lengths of link 1 and link 2 respectively. L3 is the difference in height between the handheld rod and the first joint point. L4 is the difference in horizontal distance between the handheld rod and the first joint point.

[0102] During the operation, the joint angles in the joint space at the end of the controller are collected in real time. Through forward kinematics, the current end position (x F , y F , z F ) is obtained. Assuming the neutral position point (x zero , y zero , z zero ), the control quantity u1 = [u x , u y , u z given by the operator acting on the spacecraft is obtained by calculating the position difference. T It is:

[0103]

[0104] where k dx , k dy , k dz are gain coefficients.

[0105] Step 4: Establish a control method for the automatic approaching process of the target spacecraft with a fixed time interval update

[0106] To reduce the distance between the two spacecraft and achieve the goal of successful pursuit, the above equation can be regarded as a following problem. The control target is ρ1 → 0. By adopting PD control, the desired control quantity can be obtained as

[0107]

[0108] That is

[0109]

[0110] By substituting it into (12) to solve the differential equation, the desired tracking trajectory [x dT[k] , y dT[k]， z dT[k] of the chasing spacecraft in the orbital coordinate system of the chasing spacecraft can be obtained. T The position increment Δx of the chasing spacecraft in the orbital coordinate system can be obtained through calculation:

[0111] Δx = [x dT[k] , y dT[k] , z dT[k] T - [x dT[k-1] , y​dT[k-1] , z dT[k-1] T (19)

[0112] Based on the above, through coordinate transformation, the position increment Δρ of the chaser spacecraft in the geocentric coordinate system can be obtained as Δρ = A T Δx, and the actual control quantity is calculated as u2 = -K P Δρ - K D Δρ′, where K P , K D respectively represent the gain coefficients of the proportional and derivative links in PD control, and are selected according to the actual situation. Δρ′ represents the change rate of the position increment in the geocentric equatorial inertial coordinate system, that is, its first derivative.

[0113] Step 5: Based on the calculation results of the above two-stage control, give the control input for the human-in-the-loop spacecraft approaching process.

[0114] According to the calculations in Step 3 and Step 4, the control input finally applied to the chaser spacecraft is:

[0115] u = u1 + u2 (20)

[0116] Introduce the control quantity input into the relative motion dynamics model established in Step 2, u = [u x , u y , u z T , solve to obtain the position of the chaser spacecraft under the control input. Repeat the process of Steps 3 - 5 to obtain the final position of the chaser spacecraft, and transmit the position result to the visual system to achieve visual observation by personnel.

[0117] The above is only the specific implementation manner of the present invention, but the protection scope of the present invention is not limited thereto. Any person skilled in the art within the technical scope disclosed by the present invention can easily think of various equivalent modifications or substitutions, and these modifications or substitutions should all be covered within the protection scope of the present invention.​​

Claims

1. A human-in-the-loop non-cooperative target approaching tracking control method, characterized in that Including the following steps: Step 1: Give the position \(x\) of the chaser spacecraft in the geocentric coordinate system at the initial moment P , the position \(x\) of the target spacecraft e , and give the conversion relationship between the orbital coordinate system and the geocentric coordinate system; Step 2: Obtain the positions of the two spacecraft in the orbital coordinate system through the transformation relationship, and give the position x of the virtual target point at the current moment d = x e , and establish a relative motion dynamics model; Step 3: Collect the current joint angles of the manual controller, calculate the current end position through forward kinematics, and calculate the current direct control quantity u1 of the chasing spacecraft given by the operator through the manual controller; Step 4: Calculate the position increment of the chasing spacecraft in the orbital coordinate system, obtain the position increment in the geocentric coordinate system through coordinate transformation, and calculate the automatic control quantity u2 of the chasing spacecraft to automatically approach the target spacecraft; Step 5: Finally, obtain the control quantity u of the chasing spacecraft, u = u1 + u2. Input the control quantity into the relative motion dynamics model established in Step 2 to solve for the position of the chasing spacecraft under the control input; and within each update cycle, update the position of the virtual target point of the target spacecraft; repeat the process of Steps 3 - 5 to obtain the final position of the chasing spacecraft.

2. The method for approaching and tracking control of a non-cooperative target with a human-in-the-loop according to claim 1, wherein The conversion relationship between the orbital coordinate system and the geocentric coordinate system in Step 1 is: Among them, r0 = [r 0x , r 0y , r 0z T is the position coordinate of the origin of the entrained inertial coordinate system in the geocentric equatorial inertial coordinate system, A is a matrix parameter, [x L , y L , z L T is the position of the spacecraft in the orbital coordinate system, [x r1 , y r1 , z r1 T is the position of the spacecraft in the geocentric coordinate system.​​​ 3. The method for human-in-the-loop non-cooperative target approaching tracking control according to claim 2, characterized in that The relative motion dynamics model in Step 2 is: where [x, y, z] is the difference between the virtual position of the target spacecraft and the position of the chaser spacecraft in the orbital coordinate system of the chaser spacecraft, u x , u y , u z are the control inputs of the chaser spacecraft, ω is the angular velocity of the circular orbit at the initial moment of the chaser spacecraft's period, respectively represent the second-order derivatives of x, y, z, respectively represent the first-order derivatives of x, y.

4. The method for approaching and tracking control of a non-cooperative target with a human-in-the-loop according to claim 3, wherein Step 3 is specifically: Obtain the current end - effector position (x F , y F , z F ) through forward kinematics: x F = -sinθ1(L2 sinθ3 + L1 cosθ2) y F = -L2 cosθ3 + L1 sinθ2 + L3 z F = L2 cosθ1 sinθ3 + L1 cosθ1cosθ2 - L4 In the formula, θ1, θ2, and θ3 are the rotation angles of the first three joints respectively, L1 and L2 are the lengths of the first link and the second link respectively, L3 is the height distance difference between the handheld rod and the first joint point, and L4 is the horizontal distance difference between the handheld rod and the first joint point; Calculate the current direct control quantity u1 = [u x , u y , u z given by the operator through the manual controller T : where k dx , k dy , k dz is the gain coefficient, and (x zero , y zero , z zero ) is the neutral position point.

5. The method for close-range tracking control of a non-cooperative target with human-in-the-loop according to claim 4, characterized in that Step 4 is specifically: Calculate the position increment Δx of the chasing spacecraft in the orbital coordinate system through calculation: Δx = [x dT[k] , y dT[k] , z dT[k] T - [x dT[k-1] , y dT[k-1] , z dT[k-1] T ​​ The position increment Δρ = A of the chaser spacecraft in the geocentric coordinate system can be obtained through coordinate transformation T Δx, and the actual control quantity is calculated as u2 = -K P Δρ - K D Δρ', where K P , K D represent the gain coefficients of the proportional and derivative links in PD control respectively, and Δρ′ represents the change rate of the position increment in the geocentric equatorial inertial coordinate system.

6. A computer system, characterized in that Including: One or more processors, a computer-readable storage medium for storing one or more programs, wherein when the one or more programs are executed by the one or more processors, the one or more processors implement the method described in Claim 1.

7. A computer-readable storage medium, characterized in that Stored with computer-executable instructions that are used to implement the method described in Claim 1 when executed.

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