A spacecraft autonomous approach guidance control method, electronic equipment and medium

Abstract

This invention discloses an autonomous approach guidance and control method, electronic equipment, and medium for spacecraft. It determines the approach control sequence and dwell time based on the spacecraft's relative orbit to the target, nominal approach position, attitude maneuver allowance, and nominal guidance transfer time. Then, it corrects the guidance transfer time and adjusts the approach control sequence accordingly. Next, it determines the dwell control sequence based on the spacecraft's relative orbit to the target, upper and lower limits of the nominal dwell time, and lateral acceleration. Finally, it determines the lateral control strategy, normal control strategy, and radial control strategy based on the spacecraft's relative orbit to the target and attitude maneuver allowance. This method combines fixed-time CW guidance with independent lateral, normal, and radial three-axis guidance, achieving high-precision approach control within a given time range.

Description

Technical Field

[0001] This invention relates to the field of aerospace technology, and in particular to an autonomous approach guidance and control method, electronic equipment, and medium for spacecraft. Background Technology

[0002] With the surge in the number of spacecraft in orbit, the space debris problem has become increasingly serious. Close-range observation, on-orbit status assessment, or removal of specific hazardous debris has become a critical task in maintaining the safety of the space environment. A prerequisite for achieving these tasks is that spacecraft can autonomously and reliably maneuver to a predetermined location near the target debris.

[0003] This process places extremely stringent and interconnected requirements on the approach guidance and control system. First, mission planning often requires the service spacecraft to arrive at the target point at a predetermined time to meet constraints such as lighting conditions and multi-spacecraft coordination. Second, to ensure the safety of subsequent operations, the relative position and velocity at the terminal stage require very high control precision. However, the two core indicators, "arrival at a given time" and "high-precision terminal state," present design challenges at the dynamics and control levels. In engineering implementation, this manifests as follows: open-loop or nominal trajectory guidance methods designed to ensure time determinism are often extremely sensitive to initial deviations, model uncertainties, and space environment perturbations, making it difficult to guarantee terminal accuracy; while closed-loop feedback control methods used to improve accuracy and robustness often face difficulties in controlling the convergence time. Summary of the Invention

[0004] To address some or all of the problems in existing technologies, and in order to combine the planning capability of arrival at a given time with high-precision robustness in complex space environments to achieve rapid and high-precision approach control of space debris, the first aspect of this invention provides a spacecraft autonomous approach guidance and control method that combines fixed-time CW guidance with three-axis independent guidance along the transverse normal path to achieve high-precision approach control within a given time range, including:

[0005] The approach control sequence and dwell time are determined based on the spacecraft's relative orbit to the target, the nominal approach position, the attitude maneuver allowance, and the nominal guidance transfer time.

[0006] The guidance transfer time is corrected, and the approach control sequence is modified based on the current relative orbit of the spacecraft to the target, the nominal approach position, the attitude maneuver allowance, and the corrected guidance transfer time.

[0007] The dwell control sequence is determined based on the current relative orbit of the spacecraft to the target, the upper and lower limits of the nominal dwell time search, and the lateral acceleration.

[0008] The lateral control strategy, normal control strategy, and radial control strategy are adjusted based on the current relative orbit of the spacecraft to the target and the time allotted for attitude maneuvers.

[0009] Furthermore, determining the approach control sequence and the dwell time includes:

[0010] Based on the current relative orbit of the spacecraft with respect to the target, and after determining the attitude maneuver reserve time based on the CW equation, the relative orbit of the spacecraft with respect to the target is determined.

[0011] Based on the attitude maneuver time, the spacecraft's relative orbit to the target, the reference value of the relative position of the dwell point, and the nominal guidance transfer time are used to determine the two-point transfer velocity increment.

[0012] Based on the CW equation, finite thrust extrapolation is performed to obtain the relative orbit after the control ends. Based on the relative orbit after the control ends, extrapolation is performed to the dwell time to obtain the relative orbit of the spacecraft relative to the target at the dwell time.

[0013] If the dwell error is less than the minimum relative position of the dwell point target, then the minimum relative position of the dwell point target is updated to the value of the dwell error. The relative orbit of the spacecraft to the target at the dwell time is used as the guidance planning dwell point output, and the two-point transfer velocity increment is used as the control sequence. If the magnitude of the dwell error is greater than the guidance approach control accuracy, then the reference value of the relative position of the dwell point is updated, the two-point transfer velocity increment is determined, and the iteration is performed until the number of iterations reaches the upper limit, or the magnitude of the dwell error is less than or equal to the guidance approach control accuracy.

[0014] Furthermore, the correction of guidance transfer time includes subtracting the current time and attitude preparation reserve time from the dwell time to obtain the correction of guidance transfer time.

[0015] Furthermore, the method for modifying the approach control sequence is the same as the method for determining the approach control sequence.

[0016] Furthermore, the stay control sequence is determined to include:

[0017] The possible dwell time period is determined based on the upper and lower limits of the nominal dwell time search;

[0018] Relative orbit calculations are performed for all spacecraft and target orbits within the possible dwell time period;

[0019] The point with the smallest radial position is taken as the dwell point, and its corresponding relative velocity is obtained.

[0020] Furthermore, the lateral control strategy includes lateral control speed increments and control sequences, and modifying the lateral control strategy includes:

[0021] Based on the current relative orbit of the spacecraft to the target, the relative orbit of the spacecraft to the target after the attitude maneuver allowance time is obtained by extrapolation using the CW equation;

[0022] Based on the spacecraft's relative orbit to the target after the attitude maneuver allowance time, the lateral control velocity increment is determined, and the control sequence is determined based on the lateral control velocity increment.

[0023] Furthermore, the modification of the normal control strategy includes:

[0024] Based on the current relative orbit of the spacecraft to the target, the relative orbit of the spacecraft to the target after the attitude maneuver allowance time is obtained by extrapolation using the CW equation;

[0025] Determine the direction of the normal acceleration based on the normal component of the spacecraft's current orbit relative to the target;

[0026] The nominal acceleration duration is determined based on the normal acceleration, normal relative position, and normal velocity.

[0027] Based on the upper and lower limits of the nominal acceleration time, determine the minimum normal position after acceleration and deceleration control, and its corresponding acceleration time, deceleration time, and relative orbit.

[0028] Based on the acceleration time, deceleration time, and normal acceleration, determine the normal acceleration control speed increment, the normal deceleration control speed increment, and the control sequence.

[0029] Furthermore, the modification of the radial control strategy includes:

[0030] Based on the current relative orbit of the spacecraft to the target, the relative orbit of the spacecraft to the target after the attitude maneuver allowance time is obtained by extrapolation using the CW equation;

[0031] Determine the direction of radial acceleration based on the radial component of the spacecraft's current orbit relative to the target;

[0032] The nominal acceleration duration is determined based on the radial acceleration, radial relative position, and radial velocity.

[0033] Based on the upper and lower limits of the nominal acceleration time, determine the minimum radial position after acceleration and deceleration control, and its corresponding acceleration time, deceleration time, and relative track.

[0034] Based on the acceleration time, deceleration time, and radial acceleration, determine the radial acceleration control speed increment, the radial deceleration control speed increment, and the control sequence.

[0035] Based on the spacecraft autonomous approach guidance and control method described above, a second aspect of the present invention provides an electronic device for spacecraft autonomous approach guidance and control, comprising a memory and a processor, wherein the memory is configured to store a computer program that executes the spacecraft autonomous approach guidance and control method described above when the processor is running.

[0036] A third aspect of the present invention also provides a computer-readable storage medium for autonomous approach guidance and control of a spacecraft, which stores a computer program that, when run on a processor, executes the autonomous approach guidance and control method for a spacecraft as described above.

[0037] The present invention provides an autonomous approach guidance and control method for spacecraft that comprehensively considers time constraints, precision control and engineering feasibility. It can combine the planning of arrival at a given time with high precision robustness in complex space environments to achieve rapid and high-precision approach control of space debris. Attached Figure Description

[0038] To further illustrate the above and other advantages and features of the various embodiments of the present invention, a more specific description of the various embodiments of the present invention will be presented with reference to the accompanying drawings. It is to be understood that these drawings depict only typical embodiments of the invention and are therefore not intended to limit its scope. In the drawings, identical or corresponding parts will be indicated by identical or similar reference numerals for clarity.

[0039] Figure 1 This diagram illustrates a flow chart of an autonomous approach guidance and control method for a spacecraft according to an embodiment of the present invention.

[0040] Figure 2 This diagram illustrates the changes in the relative position of the target during the autonomous approach and rendezvous guidance planning process.

[0041] Figure 3 This diagram illustrates the change in the relative velocity of the target during the autonomous approach and rendezvous guidance planning process. Detailed Implementation

[0042] In the following description, the invention is described with reference to various embodiments. However, those skilled in the art will recognize that the embodiments may be practiced without one or more specific details or in conjunction with other alternatives and / or additional methods or components. In other instances, well-known structures, materials, or operations are not shown or described in detail so as not to obscure the inventive points of the invention. Similarly, for illustrative purposes, specific numbers and configurations are set forth to provide a comprehensive understanding of embodiments of the invention. However, the invention is not limited to these specific details.

[0043] In this specification, references to "an embodiment" or "this embodiment" mean that a particular feature, structure, or characteristic described in connection with that embodiment is included in at least one embodiment of the invention. The phrase "in one embodiment" appearing throughout this specification does not necessarily refer to the same embodiment in all instances.

[0044] It should be noted that the embodiments of the present invention describe the method steps in a specific order; however, this is only for illustrating the specific embodiment and not for limiting the order of the steps. On the contrary, in different embodiments of the present invention, the order of the steps can be adjusted according to actual needs.

[0045] Existing approach guidance and control systems, in order to ensure time determinism, are often extremely sensitive to initial deviations, model uncertainties, and space environment perturbations, making it difficult to guarantee terminal accuracy. Closed-loop feedback control methods, used to improve accuracy and robustness, often suffer from difficulties in controlling the convergence time. To address this issue, this invention proposes an autonomous approach guidance and control method for spacecraft. This method comprehensively considers time constraints, accuracy control, and engineering feasibility, combining the planning of arrival at a given time with high-precision robustness in complex space environments to achieve rapid and high-precision approach control of space debris. Specifically, it combines fixed-time CW guidance with three-axis independent guidance along the transverse normal path to achieve high-precision approach control within a given time range.

[0046] The technical solution of the present invention will be further described below with reference to the accompanying drawings of the embodiments.

[0047] Figure 1 This diagram illustrates a flow chart of an autonomous approach guidance and control method for a spacecraft according to an embodiment of the present invention. Figure 1 As shown, a spacecraft autonomous approach guidance and control method includes:

[0048] First, in step 101, approach planning guidance is performed. This is based on the spacecraft's current relative orbit to the target. Approaching the nominal location and attitude maneuver reserve time Nominal guidance transfer time Determine the approach control sequence and length of stay The aforementioned dwell time This is a global variable, calculated during the initial approach mission planning, and remains unchanged thereafter. In one embodiment of the invention, the approach control sequence is determined. and length of stay include:

[0049] Based on the current relative orbit of the spacecraft with respect to the target Determined based on CW equation After a certain period of time, the spacecraft's relative orbit to the target ;

[0050] Reference value of the relative position of the main point Set to near nominal position Set the minimum relative position of the garrison point target. The iteration count is initialized to 1, and the iteration entry flag Per is set to 1 to begin iteration. Each iteration includes: determining the initial relative orbital state at time t0. Relative position of the ends The approach control pulse velocity increment is calculated using the CW equation. And further based on the nominal guidance transfer time Determine the velocity increment of the two-point transfer. The dwell time is the current time plus the transfer time, i.e. By using the CW equations for finite thrust extrapolation, the relative orbit after the control phase ends is obtained. Based on this orbit, extrapolation to the dwell time yields the spacecraft's relative orbit to the target at the dwell time. Determine the dwell error If the dwell error is less than the minimum value of the relative position of the dwell point target Then, the minimum value of the relative position of the dwelling point target is updated to the value of the dwelling error. And store the spacecraft's relative orbit to the target during its stay as As the output of the guidance planning dwell point, the two-point transfer velocity increment serves as the control sequence. Confirm whether the number of iterations has reached the upper limit. If not, determine whether the magnitude of the dwell error is greater than the guidance and approach control accuracy. If it is greater than, then update the relative position reference value of the dwell point. The two-point transfer velocity increment is redefined, and the next iteration begins. The iteration number j is incremented by 1. If the magnitude of the dwell error is less than or equal to the guidance approach control accuracy, the iteration ends, the iteration entry flag Per is set to 0, and the iteration convergence flag is set to 1. When the iteration number j is greater than the maximum iteration number, the iteration ends, and the iteration entry flag Per is set to 0.

[0051] Next, in step 102, the planned guidance is corrected. The guidance transfer time is adjusted based on the current relative orbit of the spacecraft with respect to the target. Approaching the nominal location and attitude maneuver reserve time Corrected guidance transfer time Correct the approach control sequence In one embodiment of the invention, the guidance transfer time is corrected. Equal to dwell time Subtract the current time and attitude maneuver reserve time In one embodiment of the invention, the method for modifying the approach control sequence is the same as the method for determining the approach control sequence.

[0052] Next, in step 103, the spacecraft will be guided to stay in place. This will be based on the spacecraft's current relative orbit to the target. Nominal dwell time search upper and lower limits and lateral acceleration Determine the dwell time control sequence. In one embodiment of the invention, the possible dwell time is determined based on the upper and lower limits of the nominal dwell time search:

[0053] ,

[0054] Among them, if At the current target relative orbit time Previously, The search time range is ,like Then the current target relative orbit will be extrapolated using the CW equation. get Relative orbit at time Let the minimum radial position be... Regarding the possible length of stay The relative orbits of all satellites and the target orbit within the range are calculated, and the point with the smallest radial position is found as the dwell point. And obtain the relative velocity at that moment, that is, the dwell velocity increment is The dwell time control period is ,in .

[0055] Next, in step 104, lateral correction guidance is performed. This is based on the current relative orbit of the spacecraft with respect to the target. and attitude maneuver reserve time The lateral control strategy is modified. In one embodiment of the invention, this is based on the current relative orbit of the spacecraft with respect to the target. The result is obtained by extrapolation from the CW equation. The spacecraft's relative orbit to the target after the specified time:

[0056] ;

[0057] Then based on the relative orbit Determine the lateral control speed increment And determine the control sequence based on the lateral control speed increment. .

[0058] Next, in step 105, normal correction guidance is performed. This is based on the current relative orbit of the spacecraft with respect to the target. and attitude maneuver reserve time Correcting the normal control strategy. In one embodiment of the invention, firstly based on the current spacecraft's relative orbit to the target. The result is obtained by extrapolation from the CW equation. The spacecraft's relative orbit to the target after the specified time:

[0059] ;

[0060] Then, based on the relative orbit Normal component in VVLH coordinate system Determine the normal acceleration The direction. Specifically, if the normal component If it is less than zero, then the normal acceleration The direction is positive, that is, the normal acceleration of acceleration control. Otherwise, normal acceleration The direction is negative, that is, the normal acceleration of acceleration control. Then, based on the aforementioned normal acceleration... Relative position of normal direction Normal velocity Determine the nominal acceleration time The acceleration time that minimizes the normal position after acceleration and deceleration control is calculated based on the upper and lower limits of the acceleration duration. For each moment within the acceleration period, the displacement after the acceleration and deceleration phases are calculated, and the minimum normal position and the corresponding acceleration time are identified. Deceleration time And the relative orbit. In one embodiment of the present invention, the calculation of the relative orbit after acceleration and deceleration includes: firstly, calculating the acceleration using the CW equation. Relative orbit of time Take the normal velocity relative to the orbit. If the normal velocity is less than 0 at this point, then the direction of the second acceleration is positive, which is the normal acceleration for deceleration control. Otherwise, the direction of the second acceleration is negative, that is, the normal acceleration of the deceleration control. The deceleration time is , to thrust and deceleration time Calculating deceleration using the CW equation Relative orbit of time Finally, based on the acceleration time... Deceleration time And normal acceleration correction normal guidance. Specifically, the control timing is the current target time relative to the orbit plus the attitude maneuver allowance time. Normal acceleration control speed increment Normal deceleration control speed increment and control sequences are Once acceleration control ends, deceleration control begins.

[0061] Finally, in step 106, radial correction guidance is performed. This is based on the current relative orbit of the spacecraft with respect to the target. and attitude maneuver reserve time A modified radial control strategy is employed. In one embodiment of the invention, this is first based on the current relative orbit of the spacecraft with respect to the target. The result is obtained by extrapolation from the CW equation. The spacecraft's relative orbit to the target after the specified time:

[0062] ;

[0063] Then, based on the relative orbit radial components in the VVLH coordinate system Determine radial acceleration The direction. Specifically, if the radial component If less than zero, then radial acceleration The direction is positive, that is, the radial acceleration of acceleration control. Otherwise, radial acceleration The direction is negative, that is, the normal acceleration of acceleration control. Then, based on the radial acceleration... Radial relative position radial velocity Determine the nominal acceleration time The acceleration time that minimizes the normal position after acceleration and deceleration control is calculated based on the upper and lower limits of the acceleration duration. For each moment within the acceleration period, the displacement after the acceleration and deceleration phases are calculated, and the minimum radial position and the corresponding acceleration time are identified. Deceleration time And the relative orbit. In one embodiment of the present invention, the calculation of the relative orbit after acceleration and deceleration includes: firstly, calculating the acceleration using the CW equation. Relative orbit of time Take the radial velocity relative to the orbit. If the radial velocity is less than 0 at this point, then the direction of the second acceleration is positive, which is the radial acceleration for deceleration control. Otherwise, the direction of the second acceleration is negative, which is the radial acceleration of the deceleration control. The deceleration time is , to thrust and deceleration time Calculating deceleration using the CW equation Relative orbit of time Finally, based on the acceleration time... Deceleration time And radial acceleration correction radial guidance. Specifically, the control timing is the current target time relative to the orbit plus the attitude maneuver allowance. Radial acceleration control speed increment Radial deceleration control speed increment and control sequences are Once acceleration control ends, deceleration control begins.

[0064] To verify the performance of the spacecraft autonomous approach guidance and control method described above, the spacecraft's initial position was set to 5 km behind the space debris, and it needed to reach a position 400 m behind the target within 10 to 12 minutes. The specific mission parameters are shown in Table 1.

[0065]

[0066] Table 1

[0067] The spacecraft autonomous approach guidance and control method described above is then used to sequentially perform approach planning guidance, correction planning guidance, dwell planning guidance, lateral correction guidance, normal correction guidance, and radial correction guidance. The dwell point positions planned for each stage are shown in Table 2.

[0068]

[0069] Table 2

[0070] The control sequence results are shown in Table 3:

[0071]

[0072] Table 3

[0073] The changes in the target's relative position and velocity during the control process are as follows: Figure 2 , Figure 3As shown, the final results of the spacecraft's autonomous approach and rendezvous guidance planning mission are as follows: target relative to the orbit radial position Rx: 4.257m, lateral position Ry: 374.9704m, normal position Rz: -0.32607m, radial velocity Vx: 0.035933m / s, lateral velocity Vy: 0.091399m / s, and normal velocity Vx: 0.00067097m / s. These results meet the mission requirements and enable rapid and high-precision approach to space debris.

[0074] Based on the spacecraft autonomous approach guidance and control method described above, the present invention also provides an electronic device for spacecraft autonomous approach guidance and control, which includes a memory and a processor, wherein the memory is configured to store a computer program, and the computer program executes the spacecraft autonomous approach guidance and control method described above when the processor is running.

[0075] The present invention also provides a computer-readable storage medium for autonomous approach guidance and control of spacecraft, which stores a computer program that, when run on a processor, executes the autonomous approach guidance and control method of spacecraft as described above.

[0076] Although various embodiments of the invention have been described above, it should be understood that they are presented by way of example only and not as limitations. It will be apparent to those skilled in the art that various combinations, modifications, and alterations can be made without departing from the spirit and scope of the invention. Therefore, the breadth and scope of the invention disclosed herein should not be limited by the exemplary embodiments disclosed above, but should be defined solely by the appended claims and their equivalents.

Claims

1. A spacecraft autonomous approach guidance and control method, characterized in that, include: Based on the current relative orbit of the spacecraft to the target Determined based on CW equation After a certain period of time, the spacecraft's relative orbit to the target ,in Allow time for attitude maneuvering; Based on the relative orbit Reference values ​​for the relative positions of the outposts Nominal guidance transfer time Determine the approach control pulse velocity increment This allows us to determine the velocity increments at the two points. ; Based on the CW equations, finite thrust extrapolation is performed to obtain the relative orbit after the control period ends. This relative orbit after the control period is then extrapolated to the dwell time. The relative orbit of the spacecraft to the target at the time of stay is obtained. The dwell time is equal to the current time plus the nominal guidance transfer time; Determine the dwelling error If the dwell error is less than the minimum value of the relative position of the dwell point target Then, the minimum value of the relative position of the dwelling point target is updated to the value of the dwelling error. The spacecraft's relative orbit to the target at the dwell time is used as the guidance planning dwell point output, and the two-point transfer velocity increment is used as the approach control sequence. If the magnitude of the dwell error is greater than the guidance and approach control accuracy Then update the relative position reference value of the dwell point. The process involves iterating and redetermining the two-point transfer velocity increment until the number of iterations reaches the upper limit, or the magnitude of the dwell error is less than or equal to the guidance approach control accuracy. Correct the guidance transfer time and adjust it according to the current relative orbit of the spacecraft with respect to the target. Approaching the nominal location and attitude maneuver reserve time Corrected guidance transfer time Correct the approach control sequence ; Search upper and lower limits based on nominal dwell time Determine the possible duration of stay: ， like At the current target relative orbit time Previously, The search time range is ,like Then the current target relative to the orbit Extrapolation using CW equations get Relative orbit at time ; Let the minimum radial position be For possible stay periods Relative orbit calculations were performed on all spacecraft and target orbits within the area. The point with the smallest radial position As a point of residence, and the relative velocity at that moment. As a residence velocity increment, based on lateral acceleration Determine the stay control time as ,in , For lateral speed, For normal velocity, Radial velocity; Based on the current relative orbit of the spacecraft to the target and attitude maneuver reserve time Revise the lateral control strategy, normal control strategy, and radial control strategy.

2. The spacecraft autonomous approach guidance and control method as described in claim 1, characterized in that, Correcting the guidance transfer time includes: adjusting the dwell time. Subtract the current time and attitude maneuver reserve time The corrected guidance transfer time is obtained. .

3. The spacecraft autonomous approach guidance and control method as described in claim 1, characterized in that, The method for modifying the approach control sequence is the same as the method for determining the approach control sequence.

4. The spacecraft autonomous approach guidance and control method as described in claim 1, characterized in that, The lateral control strategy includes lateral control speed increments. and control sequences, the modification of the lateral control strategy includes: Based on the current relative orbit of the spacecraft to the target The result is obtained by extrapolation of the CW equation. The spacecraft's relative orbit to the target after the specified time: ； Based on the relative orbit Determine the lateral control speed increment And determine the control sequence based on the lateral control speed increment. .

5. The spacecraft autonomous approach guidance and control method as described in claim 1, characterized in that, The modified normal control strategy includes: Based on the current relative orbit of the spacecraft to the target The result is obtained by extrapolation of the CW equation. The spacecraft's relative orbit to the target after the specified time: ； According to the relative orbit Normal component in VVLH coordinate system Determine the normal acceleration The direction; According to the normal acceleration Relative position of normal direction Normal velocity Determine the nominal acceleration time ; Based on the upper and lower limits of the nominal acceleration time, determine the minimum normal position after acceleration / deceleration control and its corresponding acceleration time. Deceleration time and relative orbits; Based on the acceleration time Deceleration time and normal acceleration determination normal acceleration control speed increment Normal deceleration control speed increment and control sequences: 。 6. The spacecraft autonomous approach guidance and control method as described in claim 1, characterized in that, The modified radial control strategy includes: Based on the current relative orbit of the spacecraft to the target The result is obtained by extrapolation of the CW equation. The spacecraft's relative orbit to the target after the specified time: ； According to the relative orbit radial components in the VVLH coordinate system Determine radial acceleration The direction; According to the radial acceleration Radial relative position radial velocity Determine the nominal acceleration time ; Based on the upper and lower limits of the nominal acceleration time, determine the minimum radial position after acceleration / deceleration control and its corresponding acceleration time. Deceleration time and relative orbits; Based on the acceleration time Deceleration time and radial acceleration determines the radial acceleration control speed increment Radial deceleration control speed increment and control sequences: 。 7. An electronic device for autonomous approach guidance and control of spacecraft, characterized in that, It includes a memory and a processor, wherein the memory is configured to store a computer program that executes the spacecraft autonomous approach guidance and control method as described in any one of claims 1 to 6 when the processor is running.

8. A computer-readable storage medium for autonomous approach guidance and control of spacecraft, characterized in that, The device contains a computer program that, when run on a processor, executes the spacecraft autonomous approach guidance and control method as described in any one of claims 1 to 6.

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