Switching method of relative navigation working modes of spacecraft, electronic equipment and medium

Abstract

The invention discloses a spacecraft relative navigation working mode switching method, an electronic device and a medium, firstly, state information of a measurement single machine is obtained through collection time sequence design, the state information comprises a health state, measurement data and a data validity identifier, and then according to the health state and the data validity identifier, the measurement data of the measurement single machine is obtained. Judging the available state of the measurement data, if the available state of the measurement data is completely available or partially available, determining and switching a relative navigation working mode in combination with task requirements, and performing filtering calculation and giving a convergence mark according to the relative navigation working mode, and if the available state of the measurement data is unavailable, performing orbit extrapolation, and carrying out fault processing on the measurement single machine. The switching method is high in autonomy and good in flexibility, and can effectively improve the self-reliability and robustness of the relative navigation system of the spacecraft under complex space tasks.

Description

Technical Field

[0001] This invention relates to the field of aerospace technology, and in particular to a method, electronic device and medium for switching relative navigation operating modes of a spacecraft. Background Technology

[0002] With the development of complex space missions such as on-orbit servicing, space assembly, and debris removal, spacecraft need to possess highly autonomous close-range relative navigation capabilities. Throughout the mission, including phases such as long-range approach, parking observation, final approach, hovering, and docking / capture, state variables such as relative distance, line-of-sight angle, and relative velocity change over a wide range. Furthermore, environmental interferences such as lighting conditions, target motion status, and the availability of measurement sensors all change dynamically.

[0003] Existing spacecraft relative navigation typically employs a pre-defined, fixed-timeline-based mode-switching strategy, forcibly switching navigation modes according to mission-planned time nodes. This strategy lacks awareness of actual navigation status and performance, suffers from poor flexibility, and cannot adapt to unexpected changes in mission scenarios. Specifically, when the actual flight status deviates from expectations, it may lead to inappropriate switching timing, causing a sharp drop in navigation accuracy or even divergence, jeopardizing mission safety. Furthermore, this strategy heavily relies on ground commands or pre-set fixed logic, making it difficult to cope with rapidly evolving real-time on-orbit missions and exhibiting low autonomy. Summary of the Invention

[0004] To address some or all of the problems in existing technologies and to improve the autonomy, reliability, and robustness of spacecraft relative navigation systems under complex space missions, the first aspect of this invention provides a method for switching spacecraft relative navigation operating modes. This method determines the availability of measurement data based on the status information of the measurement unit and data validity indicators, and determines and switches the relative navigation operating mode based on the availability of the measurement data. The method includes: By designing the timing sequence for data acquisition, the status information of the measurement unit is obtained, including health status, measurement data, and data validity identifier. Based on the health status and data validity identifier, determine the usability of the measurement data: If the measurement data is available in a fully available or partially available state, then based on the task requirements, determine and switch the relative navigation working mode, and perform filtering calculations according to the relative navigation working mode to give a convergence flag; If the measurement data is unavailable, then track extrapolation is performed, and the measurement unit is troubleshooted.

[0005] Furthermore, the status information of the measurement unit is obtained through the onboard computer, including: The onboard computer and all measurement units are time-stamped with the highest precision time reference on the satellite. The onboard computer decomposes the global time series into independent time slices for each measurement unit; Based on the independent time slices obtained from the decomposition, the data acquisition operation of each measurement unit is triggered.

[0006] Furthermore, the onboard computer and each measurement unit synchronize their timestamps using the second pulse signal from the onboard GNSS.

[0007] Furthermore, obtaining the status information of the measurement unit also includes: If a measurement unit cannot achieve timestamp consistency, the continuous observation data with local timestamps within a specified time period prior to the current time will be stored in the temporary data buffer of the onboard computer. The measurement data of that measurement unit will be time-synchronized with the measurement data of other measurement units to achieve spatiotemporal alignment of all measurement data under a unified time reference.

[0008] Further, determining the usability of the measurement data includes: After receiving the measurement data from each measurement unit, the navigation module performs a time matching judgment and checks the consistency of the timestamps of each measurement data. Based on the state estimation and orbital dynamics model of the previous cycle, a one-step state prediction is performed; Based on the prediction results, the predicted value of the measurement data of each measurement unit is calculated, and the difference between the actual received measurement data and the predicted value is compared with a threshold. If the difference exceeds the threshold, the measurement data is unusable; if the difference does not exceed the threshold, the measurement data is usable.

[0009] Furthermore, the measurement data includes angle measurements and range measurements, wherein the angle measurements include the azimuth and elevation angle measurements of the target star obtained by the optical navigation camera, and the range measurements include range data from a laser rangefinder or microwave radar.

[0010] Furthermore, determining the relative navigation operating mode includes: If the angle measurement and range measurement values ​​of any star are available, the relative navigation working mode is angle measurement and range measurement mode; If the angle measurement value of the two satellites is available, but the distance measurement value is not available, then the relative navigation working mode is the angle measurement mode of the two satellites; If the angle measurement value of a single optical camera is available, but the distance measurement value is not available, then the relative navigation working mode is the single-star angle measurement mode.

[0011] Furthermore, switching the relative navigation operating mode includes: Inherit and use the optimal state estimate from the time before the switch; Reset the covariance matrix to the covariance matrix with the initial larger uncertainty corresponding to the current relative navigation working mode.

[0012] Furthermore, the filtering solution includes: Using the state variables, covariance matrix, and system process noise matrix from the previous filtering cycle, combined with the orbital dynamics model, time updates are performed to predict the state variables and covariance matrix at the current moment. Based on the current state variables and the measurement geometry of each measurement unit, we identify the shortcomings of pseudo-measurements and their corresponding observation matrices. Based on the predicted covariance matrix, the observation matrix, and the noise matrix of each measurement unit, the Kalman gain is determined, and the state variables are corrected based on the Kalman gain. The difference between the actual observations and the pseudo-measurements is fused to output the optimal state estimate for the current moment.

[0013] Based on the aforementioned method for switching the relative navigation operating mode of a spacecraft, a second aspect of the present invention provides an electronic device for switching the relative navigation operating mode of a spacecraft, comprising a memory and a processor, wherein the memory is configured to store a computer program that executes the aforementioned method for switching the relative navigation operating mode of a spacecraft when the processor is running.

[0014] A third aspect of the present invention also provides a computer-readable storage medium for switching the relative navigation operating mode of a spacecraft, the medium storing a computer program that, when run on a processor, executes the aforementioned method for switching the relative navigation operating mode of a spacecraft.

[0015] This invention provides a method for switching the relative navigation working mode of a spacecraft. The method determines the availability of measurement data based on the status information of the measurement unit and the data validity identifier, and determines and switches the relative navigation working mode based on the availability of measurement data. It determines the relative navigation working mode based on the actual navigation status and performance, which has good autonomy and high flexibility, and can effectively improve the reliability and robustness of the relative navigation system of the spacecraft under complex space missions. Attached Figure Description

[0016] 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.

[0017] Figure 1 A flowchart illustrating a method for switching relative navigation operating modes of a spacecraft according to an embodiment of the present invention is shown. Figure 2 The diagram shows a flowchart illustrating a method for switching relative navigation operating modes of a spacecraft according to another embodiment of the present invention. Detailed Implementation

[0018] 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 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.

[0019] 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.

[0020] 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.

[0021] In space rendezvous missions, optical observation is a crucial tool. Geometric calculation methods based on optical observation are fundamental for obtaining the target's relative orbit, primarily including three modes: single-satellite angle measurement, dual-satellite angle measurement, and a combination of angle and range measurement. Relative navigation often occurs simultaneously with guidance and control, with navigation results serving as input for guidance calculations. In actual on-orbit rendezvous missions, due to satellite attitude adjustment requirements and the field of view constraints of the measurement unit, measurement data may be unavailable under extreme conditions. To ensure navigation continuity, high-precision orbit extrapolation models are typically configured onboard to enable target positioning when the measurement unit's status is unavailable. Therefore, addressing the issues of poor flexibility and low autonomy in existing spacecraft relative navigation mode switching strategies, this invention provides a method for switching spacecraft relative navigation operating modes. By evaluating multi-source information, including measurement unit status, filtering parameters, and mission requirements, it achieves safe, smooth, and optimal switching of navigation modes, thereby improving the autonomy, reliability, and robustness of spacecraft relative navigation systems under complex space missions.

[0022] The technical solution of the present invention will be further described below with reference to the accompanying drawings of the embodiments.

[0023] Figure 1 This is a flowchart illustrating a method for switching relative navigation operating modes of a spacecraft according to an embodiment of the present invention. Figure 1 As shown, a method for switching relative navigation operating modes of a spacecraft includes: First, in step 101, data acquisition occurs. Through timing design, the status information of the measurement units is obtained, including health status, measurement data, and data validity identifiers. In one embodiment of the invention, the status information of the measurement units is acquired via an onboard computer, and precise timing control and time synchronization of data acquisition from multiple measurement units are achieved through centralized scheduling by the onboard computer. In another embodiment, to achieve strict spatiotemporal alignment of multi-source data, high-precision timestamp synchronization is first performed, synchronizing the onboard computer and each measurement unit with the highest precision time reference on the satellite. In one embodiment, the onboard computer and each measurement unit use the second pulse signal from the onboard GNSS for timestamp synchronization. Simultaneously, to avoid bus conflicts and optimize computing resources, in one embodiment, a time-division acquisition scheduling method is adopted. That is, the global timing is decomposed into independent time slices for each measurement unit by the onboard computer, and then, through hardware trigger signals or software instructions, the data acquisition operation of each measurement unit is triggered strictly according to the plan of the independent time slices, ensuring that the time difference of the measurement data is minimal during navigation filtering. In addition, if a measurement unit cannot achieve timestamp consistency due to hardware delay or trigger asynchrony, the continuous observation data with local timestamps within a specified time period prior to the current time will be stored in the temporary data buffer of the onboard computer, and the measurement data of the measurement unit will be time-stamped with the measurement data of other measurement units to achieve spatiotemporal alignment of all measurement data under a unified time reference. Next, in step 102, the availability status of the measurement data is determined. Based on the health status and data validity identifier, the availability status of the measurement data is judged. If the measurement data is fully available or partially available, the process proceeds to step 131 to determine the relative navigation working mode. If the measurement data is unavailable, the process proceeds to step 141 for orbit extrapolation. In one embodiment of the present invention, determining the availability status of the measurement data includes two parts: time matching and observation matching. First, the measurement information is time-matched, a time consistency flag is given, and the measurement information required for filtering is given. After time matching, observation matching is performed. In one embodiment of the present invention, the navigation module receives data packets from each measurement unit sent by the onboard computer. First, a time matching judgment is performed to check the consistency of the timestamps of each data unit. Then, using the state estimation and orbital dynamics model of the previous cycle, a one-step state prediction is performed. Simultaneously, based on the predicted state, the predicted values ​​of the observation data of each measurement unit are calculated. The actual received measurement values ​​are compared with the predicted values ​​to determine the data validity, i.e., its availability status. In one embodiment of the present invention, the difference between the actual received measurement data and the predicted value is compared with a threshold. If the difference exceeds the threshold, the measurement data is invalid and unusable; if the difference does not exceed the threshold, the measurement data is valid and usable. The measurement values ​​include angle measurements and range measurements. The angle measurements include the azimuth and elevation angle measurements of the target star taken by the optical navigation camera. The range measurements include the range data from a laser rangefinder or microwave radar. The determination of data validity includes: determining whether the difference between the observed and predicted azimuth and elevation angles of the target star taken by the optical navigation camera exceeds a set threshold; if not, the angle measurement data is considered valid; otherwise, it is considered invalid. It also includes determining whether the difference between the observed and predicted range data from the laser rangefinder or microwave radar exceeds a set threshold; if not, the range measurement data is considered valid; otherwise, it is considered invalid. In step 131, the relative navigation working mode is determined. When the measurement data is available in a fully available or partially available state, the relative navigation working mode is determined based on the mission requirements. In one embodiment of the invention, the mission requirements include, for example, ground requirements such as high-precision orbit positioning. In one embodiment of the invention, if both the angle measurement value and the distance measurement value of any satellite are available, the relative navigation working mode is mode A, angle and distance measurement mode; if the angle measurement value of two satellites is available but the distance measurement value is unavailable, the relative navigation working mode is mode B, two-satellite angle measurement mode; if the angle measurement value of a single optical camera is available but the distance measurement value is unavailable, the relative navigation working mode is mode C, single-satellite angle measurement mode. After determining the relative navigation working mode, recursive state information, observation prediction information, filtered measurement information, and the working mode are output. Simultaneously, if the determined relative navigation working mode differs from the previous cycle, step 132 is entered to switch the relative navigation working mode, performing a smooth switch and reset. In step 132, the relative navigation operating mode is switched. To ensure a smooth switch, in one embodiment of the present invention, when switching the relative navigation operating mode, the optimal state estimate value at the previous moment is inherited to maintain state continuity, while the covariance matrix is ​​reset to the covariance matrix with a larger initial uncertainty corresponding to the current mode, so as to accelerate the convergence of the filter under the new observation; Next, in step 133, filtering is performed. Filtering is conducted according to the relative navigation operating mode, and a convergence flag is given. The core of relative navigation calculation is navigation filtering, which fuses observation data and the dynamic model through a filtering algorithm to output the optimal state estimate. In one embodiment of the present invention, the filtering includes: First, state and covariance prediction. Using the state variables, covariance matrix, and system process noise matrix from the previous filtering cycle, combined with the orbital dynamics model, time updates are performed to predict the state variables and covariance matrix at the current moment. Next, the observation matrix is ​​calculated. Based on the current state variables and the measurement geometry of each measurement unit, the pseudo-measurements and their corresponding observation matrices, such as the Jacobian matrix, are calculated. Finally, Kalman gain calculation and state update are performed. Based on the predicted covariance matrix, observation matrix, and noise matrix of each measurement unit, the Kalman gain is calculated. This gain is then used to correct the state variables, fusing the differences between actual and pseudo-measurements to output the optimal state estimate for the current moment, such as position and velocity, and updating the covariance matrix.

[0024] In one embodiment of the present invention, the reliability of the navigation result can be determined based on the navigation filter output value, and an availability flag can be generated. If the current navigation operating mode is mode C, single-satellite angle measurement mode, the navigation result is considered available if both the position correction and velocity correction are within the threshold range; otherwise, it is unavailable. If the current navigation operating mode is mode A or B, i.e., angle and distance measurement mode or dual-satellite angle measurement mode, the convergence and stability of the filtering result are evaluated, and the standard deviation change is continuously monitored over N (N≥10) filtering cycles. If the change in the position and velocity standard deviation is less than the set threshold, it indicates that the filtering has converged and stabilized, and the filtering result of mode A or B is considered available; otherwise, it is unavailable. If the filtering result is unavailable, proceed to step 141, orbit extrapolation. A high-precision orbit extrapolation model is used to predict the current target orbit. In step 141, track extrapolation is performed. When angle measurement and distance measurement are both ineffective, or the filtering results are unavailable, high-precision track extrapolation is performed. The high-precision track extrapolation model is used to predict the current target track and output the current target track. Fault handling is also performed on the measurement unit.

[0025] The switching method described above can also be applied to multiple spacecraft scenarios, such as a primary satellite and a secondary satellite. Both the primary and secondary satellites are equipped with optical observation cameras, microwave radars, and GNSS receivers, and data is transmitted between them via inter-satellite communication. The primary and secondary satellites simultaneously perform angle and range measurements on the target, with the relative navigation algorithm and operating mode switching running on the primary satellite. The operating range of the rangefinders on the primary and secondary satellites can, for example, be within 1 km. In this embodiment, as... Figure 2 As shown, the acquisition of measurement data includes two parts: the primary satellite and the secondary satellite. The optical observation camera and microwave radar of the primary satellite use GNSS second pulse signals for time synchronization. The onboard computer of the primary satellite sends telemetry requests to the optical observation camera and microwave radar every second and performs data validity checks based on Table 1. If the check passes, the measurement data, including the time, is stored in the observation pool; otherwise, it is not updated in the observation pool. The observation pool can store N seconds of data, where the value of N is determined by the communication delay. Similarly, the optical observation camera and microwave radar of the secondary satellite use GNSS second pulse signals for time synchronization. The onboard computer of the secondary satellite sends telemetry requests to the optical observation camera and microwave radar every second and sends the measurement data, including the time, to the primary satellite. The onboard computer of the primary satellite performs data validity checks. If the check passes, the secondary satellite data in the navigation observation pool is updated; otherwise, it is not updated in the observation pool. The onboard computer of the primary satellite receives the measurement data sent by the secondary satellite and performs time matching with the data in the primary satellite's observation pool.

[0026] Table 1 In this embodiment, due to the different time delays of the various measurement data, the onboard computer of the primary satellite needs to perform time matching on the measurement data of the primary and secondary satellites in the observation pool, including data, orbit data, and attitude data, and provide a time consistency flag (1 / 0), as well as the measurement information required for filtering. If the time consistency flag is 1, the matching is successful, and relative navigation filtering is performed; otherwise, the current relative navigation working mode is switched, and navigation data is updated through the orbital dynamics model. The relative navigation filtering calculation uses the extended Kalman filter algorithm, including one-step state prediction, observation matrix calculation, and observation data prediction. The one-step state prediction refers to inputting the optimal estimate of the filter at the previous time step into the orbital dynamics recursion module to predict the state at time k. By recursively applying the formula to time k+1, we obtain the prior estimate. and its prior estimate of covariance The observation matrix calculation includes: calculating pseudo-measurements and their corresponding observation matrices based on the current state variables and the measurement geometry of each measurement unit. Observation data prediction includes a measurement update module that introduces new measurement information to update the prior state, ultimately obtaining a posterior estimate of the state. and its post-test estimated covariance Navigation data updates via the orbital dynamics model employ an on-board orbital dynamics extrapolation algorithm based on HPOP. The HPOP model is a high-precision numerical model for spacecraft orbit calculations. In addition to the Earth's central gravity, the integrator module also considers the effects of perturbations such as the Earth's non-spherical gravity, the gravity of the Sun, Moon, and planets (third bodies), Earth's atmospheric drag, solar radiation pressure perturbations, Earth's tidal perturbations, and relativistic effects. The calculation models for each perturbation and the integrator are shown in Table 2.

[0027] Table 2 After completing the relative navigation filtering solution, the target observations, including the tracked target and the bright target, are first calculated, along with the difference between the observed and predicted values ​​and the pseudo-measurements. It is then determined whether the difference between the observed and predicted values ​​exceeds a set threshold; if it does not, the data is considered valid; otherwise, it is considered invalid. Based on the observation information matching results, the current relative navigation working mode is determined. First, the matching validity of the observation information is checked. If all matches are found, the navigation working mode is determined according to priority, as shown in Table 3.

[0028] Table 3 When the determined navigation mode differs from the previous cycle mode, a smooth switch and reset are performed. The optimal state estimate from the moment before the switch is inherited to maintain state continuity, and the covariance matrix is ​​reset to the covariance matrix with greater initial uncertainty corresponding to the current mode, in order to accelerate filter convergence under new observations. When the determined navigation mode is the same as the previous cycle mode, the settings under this mode are continued, and parameter acquisition and matching are performed again.

[0029] Based on the aforementioned method for switching the relative navigation operating mode of a spacecraft, the present invention also provides an electronic device for switching the relative navigation operating mode of a spacecraft, comprising a memory and a processor, wherein the memory is configured to store a computer program that executes the aforementioned method for switching the relative navigation operating mode of a spacecraft when the processor is running.

[0030] The present invention also provides a computer-readable storage medium for switching the relative navigation operating mode of a spacecraft, which stores a computer program that, when run on a processor, executes the aforementioned method for switching the relative navigation operating mode of a spacecraft.

[0031] 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 method for switching relative navigation operating modes of a spacecraft, characterized in that, include: By designing the timing sequence for data acquisition, the status information of the measurement unit is obtained, including health status, measurement data, and data validity identifier. Based on the health status and data validity identifier, determine the usability of the measurement data: If the measurement data is available in a fully available or partially available state, then based on the task requirements, determine and switch the relative navigation working mode, and perform filtering calculations according to the relative navigation working mode to give a convergence flag; If the measurement data is unavailable, then track extrapolation is performed, and the measurement unit is troubleshooted.

2. The switching method as described in claim 1, characterized in that, The status information of the measurement unit is obtained through the onboard computer, including: The onboard computer and various measurement units are synchronized with the highest precision time reference on the satellite using the second pulse signal from the onboard GNSS. The onboard computer decomposes the global time series into independent time slices for each measurement unit; Based on the independent time slices obtained from the decomposition, the data acquisition operation of each measurement unit is triggered, and the status information of each measurement unit is obtained.

3. The switching method as described in claim 2, characterized in that, Obtaining the status information of the measurement unit also includes: If a measurement unit cannot achieve timestamp consistency, the continuous observation data with local timestamps within a specified time period prior to the current time will be stored in the temporary data buffer of the onboard computer. The measurement data of that measurement unit will be time-synchronized with the measurement data of other measurement units to achieve spatiotemporal alignment of all measurement data under a unified time reference.

4. The switching method as described in claim 1, characterized in that, Determining the availability of the measurement data includes: After receiving the measurement data from each measurement unit, the navigation module performs a time matching judgment and checks the consistency of the timestamps of each measurement data. Based on the state estimation and orbital dynamics model of the previous cycle, a one-step state prediction is performed; Based on the prediction results, the predicted value of the measurement data of each measurement unit is calculated, and the difference between the actual received measurement data and the predicted value is compared with a threshold. If the difference exceeds the threshold, the measurement data is unusable; if the difference does not exceed the threshold, the measurement data is usable.

5. The switching method as described in claim 1, characterized in that, The measurement data includes angle measurements and range measurements. The angle measurements include the azimuth and elevation angle measurements of the target star taken by the optical navigation camera. The range measurements include range data from a laser rangefinder or microwave radar.

6. The switching method as described in claim 5, characterized in that, Determining the relative navigation operating mode includes: If the angle measurement and range measurement values ​​of any star are available, the relative navigation working mode is angle measurement and range measurement mode; If the angle measurement value of the two satellites is available, but the distance measurement value is not available, then the relative navigation working mode is the angle measurement mode of the two satellites; If the angle measurement value of a single optical camera is available, but the distance measurement value is not available, then the relative navigation working mode is the single-star angle measurement mode.

7. The switching method as described in claim 1, characterized in that, Switching to relative navigation working mode includes: Inherit and use the optimal state estimate from the time before the switch; Reset the covariance matrix to the covariance matrix with the initial larger uncertainty corresponding to the current relative navigation working mode.

8. The switching method as described in claim 1, characterized in that, The filtering solution includes: Using the state variables, covariance matrix, and system process noise matrix from the previous filtering cycle, combined with the orbital dynamics model, time updates are performed to predict the state variables and covariance matrix at the current moment. Based on the current state variables and the measurement geometry of each measurement unit, we identify the shortcomings of pseudo-measurements and their corresponding observation matrices. Based on the predicted covariance matrix, the observation matrix, and the noise matrix of each measurement unit, the Kalman gain is determined, and the state variables are corrected based on the Kalman gain. The difference between the actual observations and the pseudo-measurements is fused to output the optimal state estimate for the current moment.

9. An electronic device for switching the relative navigation operating mode of a spacecraft, characterized in that, It includes a memory and a processor, wherein the memory is configured to store a computer program that, when the processor is running, executes the spacecraft relative navigation operating mode switching method as described in any one of claims 1 to 8.

10. A computer-readable storage medium for switching relative navigation operating modes of a spacecraft, characterized in that, The device contains a computer program that, when run on a processor, executes the method for switching the relative navigation operating mode of a spacecraft as described in any one of claims 1 to 8.

Images