A navigation mode transition method, electronic device, and medium

Abstract

This invention discloses a navigation mode transition method, electronic device, and medium. During the operation of the first navigation mode, it monitors the geometric angles between the parent satellite, the service spacecraft, and space debris. When the geometric angles meet the mode switching requirements, it obtains the target orbit position in the J2000 coordinate system through binary-satellite geometric estimation and uses this position as the initial position for the second navigation mode, switching to the second navigation mode. Determining the target orbit position through binary-satellite geometric estimation before switching navigation modes can rapidly improve the accuracy of single-satellite navigation positioning, provide a good initial value for binary-satellite angle measurement filtering, and accelerate the convergence process.

Description

Technical Field

[0001] This invention relates to the field of aerospace technology, and in particular to a navigation mode transition method, electronic device, and medium. Background Technology

[0002] With the increasing frequency of space activities, the number of space debris, such as defunct satellites and rocket final stages, continues to grow, posing a serious threat to spacecraft in orbit. Space debris removal and on-orbit repair have become key technologies for maintaining the safety of the space environment. The core premise of these missions is that the servicing spacecraft can perform high-precision relative navigation and positioning of non-cooperative target debris.

[0003] Currently, on-orbit servicing missions typically employ a collaborative working model between the parent satellite and the servicing spacecraft. A typical relative navigation process includes two phases: single-satellite navigation and dual-satellite angle measurement navigation. In the initial stages of the mission, due to the short baseline distance between the parent satellite and the servicing spacecraft, the spatial geometry is poor, and the observability of dual-satellite angle measurement is weak, making it impossible to effectively utilize dual-satellite angle measurement information. Therefore, in this phase, a single spacecraft, such as the parent satellite, typically uses its own optical cameras and other sensors for single-satellite angle measurement navigation of the target. Once the parent satellite and the servicing spacecraft maneuver to create a sufficiently long baseline, the spatial geometry improves significantly, and observability is enhanced. At this point, the system switches to a dual-satellite angle measurement navigation mode, where both spacecraft simultaneously measure the angle of the target, achieving higher positioning accuracy.

[0004] However, the above process has a key problem: in the first stage, due to the limited navigation accuracy caused by using only single-satellite angle measurement, the target positioning error is relatively large. When switching to the second stage of dual-satellite angle measurement, the positioning result of the previous stage needs to be used as the initial input for the second stage. Since the nonlinear filtering algorithm is sensitive to initial values, the poor target positioning obtained in the first stage will significantly affect the convergence speed and stability of the filter, resulting in a longer convergence time or even divergence, thus failing to achieve rapid collaborative positioning. Summary of the Invention

[0005] To address some or all of the problems in the existing technology, and in order to improve the initial state accuracy of binary satellite navigation during the transition phase of relative navigation mode switching, the first aspect of the present invention provides a navigation mode transition method, comprising:

[0006] During operation in the first navigation mode, the geometric angles between the home satellite, the service spacecraft, and space debris are monitored.

[0007] When the geometric angles meet the mode switching requirements, the position of the target orbit in the J2000 coordinate system is obtained through binary star geometric estimation.

[0008] Use the aforementioned location as the initial location calculated in the second navigation mode, and switch to the second navigation mode.

[0009] Furthermore, the mode switching requirements include:

[0010] The angle between the vector pointing from the parent star to the space debris and the vector pointing from the service spacecraft to the space debris is greater than a set threshold, and the space debris is located within the field of view of the optical cameras of the parent star and the service spacecraft.

[0011] Furthermore, the first navigation mode is a single-star angle measurement navigation mode.

[0012] Furthermore, the first navigation mode includes:

[0013] The parent planet tracks and measures the space debris using its onboard optical camera;

[0014] Using a high-precision orbit extrapolation model, an extended Kalman filter is applied to output the position and velocity of the target orbit.

[0015] Furthermore, the binary star geometric estimation includes:

[0016] At the same time, obtain the GNSS positioning positions of the parent satellite and the service spacecraft in the J2000 coordinate system, as well as the measurement angles relative to the target;

[0017] Based on the GNSS positioning location and the measured angle, calculate the position (x, y, z) of the target orbit in the J2000 coordinate system.

[0018] Furthermore, the velocity in the filtering result of the previous navigation mode is used as the initial velocity calculated in the second navigation mode.

[0019] Furthermore, the second navigation mode is a dual-satellite angle measurement navigation mode.

[0020] Based on the navigation mode transition method described above, a second aspect of the present invention provides an electronic device for controlling navigation mode transition, comprising a memory and a processor, wherein the memory is configured to store a computer program that executes the navigation mode transition method described above when the processor is running.

[0021] A third aspect of the present invention also provides a computer-readable storage medium for controlling navigation mode transitions, which stores a computer program that, when run on a processor, executes the navigation mode transition method as described above.

[0022] The present invention provides a navigation mode transition method that performs dual-satellite geometric estimation within a critical time window when the dual-satellite angle measurement geometric observability condition is met and both dual-satellite angle measurements are effective. This method rapidly improves the accuracy of single-satellite navigation and positioning, provides a good initial value for dual-satellite angle measurement filtering, accelerates the convergence process, and ensures that the cooperative navigation system can quickly enter a high-precision and stable working state, shortens the entire mission cycle, and improves the reliability and efficiency of on-orbit service missions. Attached Figure Description

[0023] 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.

[0024] Figure 1 A flowchart illustrating a navigation mode transition method according to an embodiment of the present invention is shown.

[0025] Figure 2 A schematic diagram illustrating the principle of binary star geometric estimation according to an embodiment of the present invention is shown;

[0026] Figure 3 This diagram illustrates the position error of the target orbit in the J2000 coordinate system under the first navigation mode of an embodiment of the present invention.

[0027] Figure 4 This diagram illustrates the velocity error of the target orbit in the J2000 coordinate system under the first navigation mode of an embodiment of the present invention.

[0028] Figure 5 This diagram illustrates the position error of the target orbit in the J2000 coordinate system under the second navigation mode according to an embodiment of the present invention.

[0029] Figure 6 This diagram illustrates the velocity error of the target orbit in the J2000 coordinate system under the second navigation mode of an embodiment of the present invention. Detailed Implementation

[0030] 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. Furthermore, it should be understood that the embodiments shown in the drawings are illustrative representations and are not necessarily drawn to scale.

[0031] 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.

[0032] 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.

[0033] To rapidly and effectively improve the initial state accuracy of a dual-satellite navigation system during the transition from single-satellite angular measurement navigation to dual-satellite angular measurement navigation, this invention provides a navigation mode transition method. Within a critical time window confirming that the geometric observability conditions of dual-satellite angular measurement are met and that both angular measurements are effective, a "dual-satellite geometric estimation" step is added. This rapidly improves the accuracy of single-satellite navigation positioning, provides a good initial value for dual-satellite angular measurement filtering, and accelerates the convergence process.

[0034] The technical solution of the present invention will be further described below with reference to the accompanying drawings of the embodiments.

[0035] Figure 1 This diagram illustrates a flowchart of a navigation mode transition method according to an embodiment of the present invention. Figure 1 As shown, a navigation mode transition method includes:

[0036] First, in step 101, observability is monitored. During the operation of the first navigation mode, the geometric angles between the parent satellite, the service spacecraft, and the space debris are monitored, and geometric observability is determined based on these geometric angles. In one embodiment of the invention, the first navigation mode refers to a one-way angle measurement navigation mode. In the early stages of the mission, the service spacecraft has just separated from the parent satellite, so the parent satellite and the service spacecraft have not yet established a sufficient baseline, resulting in poor geometric observability. At this time, a single-satellite angle measurement navigation mode is adopted. Specifically, the parent satellite uses its onboard optical camera to track and measure the space debris. Through the angle measurement observations of a single spacecraft, an extended Kalman filter (EKF) is performed using a high-precision orbit extrapolation model to output the target orbital position and velocity. However, due to the limitations of single-satellite observation geometry, the error in the depth direction, i.e., along the line of sight, is usually large. During this process, the navigation system continuously monitors the geometric angles between the parent satellite, the service spacecraft, and the space debris. When the angle between the vector pointing from the parent star to the space debris and the vector pointing from the service spacecraft to the space debris is greater than a set threshold, and the space debris is located within the field of view of the optical cameras of the parent star and the service spacecraft, the navigation system is considered to be able to switch modes.

[0037] Next, in step 102, binary star geometry estimation is performed. When the geometric angles meet the mode switching requirements, the start flag for binary star geometry estimation is set to start, and binary star geometry estimation is performed to obtain the position of the target orbit in the J2000 coordinate system. In one embodiment of the present invention, the positioning positions of the parent satellite and the service spacecraft's GNSS in the J2000 coordinate system, as well as the measured angles relative to the target, are first obtained at the same time. These measured angles include azimuth and elevation angles. Then, based on the GNSS positioning positions and the measured angles, the position (x, y, z) of the target orbit in the J2000 coordinate system is calculated, as follows: Figure 2 As shown:

[0038] ,

[0039] ,

[0040] ,

[0041] in:

[0042] The weighting coefficient is determined based on the different angular measurement accuracies of the parent satellite and the service spacecraft.

[0043] The location of the parent star in the J2000 coordinate system;

[0044] The GNSS positioning position of the service spacecraft in the J2000 coordinate system;

[0045] The distance between the parent planet and the space debris:

[0046] ,

[0047] The distance between the service spacecraft and the space debris:

[0048] ,

[0049] , is the distance between the parent star and space debris M1 projected onto the xoz plane.

[0050] , is the distance between the service spacecraft and space debris M2 projected onto the xoz plane.

[0051] The azimuth angle obtained by measuring the space debris M1 for the parent star. The azimuth angle obtained by measuring the space debris M2 for the service spacecraft. The elevation angle obtained by measuring the space debris M1 for the parent star The pitch angle obtained by measuring the space debris M2 for the service spacecraft. The distance between the space debris M1 and M2:

[0052] ;

[0053] Finally, in step 103, the navigation mode is switched. The position component of the previous single-satellite filtering result is updated to the target orbit position in the J2000 system obtained through binary-satellite geometric estimation in step 102, while the velocity remains unchanged. This position and velocity are used as the initial state variables for the second navigation mode. This initial position depends only on the positioning accuracy of the spacecraft's GNSS and the measurement accuracy of the camera. As long as the accuracy of the single-satellite is high, the position accuracy of the geometric estimation can be guaranteed, thus improving the convergence speed, avoiding result divergence, and providing stable and reliable navigation information for subsequent approximation and rendezvous. In one embodiment of the present invention, the second navigation mode is a binary-satellite angle measurement navigation mode.

[0054] To better illustrate the technical solution of this invention, a relative navigation process for space debris removal is used as an example for detailed explanation. The parent satellite, carrying a service spacecraft, initially operates in a certain orbit to perform a space debris removal mission. After reaching a given position, the service spacecraft detaches from the parent satellite to carry out its mission. The measurement error settings for the parent satellite and the service spacecraft are shown in Table 1.

[0055]

[0056] Table 1

[0057] In the initial stage, due to the close proximity of the service spacecraft to the parent satellite and poor geometric observability, single-satellite navigation was initially performed using the parent satellite. After EKF filtering, the navigation error is as follows: Figure 3 and Figure 4 As shown in the figure, before enabling dual-satellite angle measurement, the navigation position and velocity errors are [-225.547m, -3329.545m, -4499.551m, 5.3496m / s, -1.3849m / s, -1.7590m / s]. After meeting the dual-satellite observability condition and enabling dual-satellite geometric estimation, the navigation positioning error is [-569.04m, 1161.92m, 1484.53m], a significant improvement compared to single-satellite angle measurement accuracy. Finally, after dual-satellite angle measurement filtering, the navigation positioning error is as follows: Figure 5 and Figure 6 As shown in the figure, the navigation accuracy can reach the 30m level after the dual-star angle measurement filter, which is a significant improvement.

[0058] As can be seen, the navigation mode transition method provided by this invention can quickly improve the accuracy of single-satellite navigation and positioning through dual-satellite geometric estimation, provide a good initial value for dual-satellite angle measurement filtering, accelerate the convergence process, and thus ensure that the cooperative navigation system can quickly enter a high-precision and stable working state, shorten the entire mission cycle, and improve the reliability and efficiency of on-orbit service missions.

[0059] Based on the navigation mode transition method described above, the present invention also provides an electronic device for controlling navigation mode transition, comprising a memory and a processor, wherein the memory is configured to store a computer program that executes the navigation mode transition method described above when the processor is running.

[0060] The present invention also provides a computer-readable storage medium for controlling navigation mode transitions, which stores a computer program that, when run on a processor, executes the navigation mode transition method as described above.

[0061] 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 navigation mode transition method, characterized in that, include: During the operation of the first navigation mode, the geometric angles between the parent satellite, the service spacecraft, and space debris are monitored, wherein the first navigation mode is a single-satellite angle measurement navigation mode. When the geometric angle meets the mode switching requirements, the position of the target orbit in the J2000 coordinate system is obtained through binary star geometric estimation. The mode switching requirements include: the angle between the vector pointing from the parent star to the space debris and the vector pointing from the service spacecraft to the space debris is greater than a set threshold, and the space debris is located within the field of view of the optical cameras of the parent star and the service spacecraft. The location is used as the initial location calculated in the second navigation mode, and the system switches to the second navigation mode, which is a dual-star angle measurement navigation mode.

2. The navigation mode transition method as described in claim 1, characterized in that, The first navigation mode includes: The parent planet tracks and measures the space debris using its onboard optical camera; Using a high-precision orbit extrapolation model, an extended Kalman filter is applied to output the position and velocity of the target orbit.

3. The navigation mode transition method as described in claim 1, characterized in that, The binary star geometric estimation includes: At the same time, obtain the GNSS positioning positions of the parent satellite and the service spacecraft in the J2000 coordinate system, as well as the measurement angles relative to the target; Based on the GNSS positioning location and the measured angle, calculate the position (x, y, z) of the target orbit in the J2000 coordinate system.

4. The navigation mode transition method as described in claim 3, characterized in that, The position (x, y, z) of the target orbit in the J2000 coordinate system is calculated using the following formula: ， ， ， in: The weighting coefficient is determined based on the different angular measurement accuracies of the parent satellite and the service spacecraft. The location of the parent star in the J2000 coordinate system; The GNSS positioning position of the service spacecraft in the J2000 coordinate system; The distance between the parent planet and the space debris: ， The distance between the service spacecraft and the space debris: ， , is the distance between the parent star and space debris M1 projected onto the xoz plane. , is the distance between the service spacecraft and space debris M2 projected onto the xoz plane. The azimuth angle obtained by measuring the space debris M1 for the parent star. The azimuth angle obtained by measuring the space debris M2 for the service spacecraft. The elevation angle obtained by measuring the space debris M1 for the parent star The pitch angle obtained by measuring the space debris M2 for the service spacecraft. The distance between the space fragments M1 and M2: 。 5. The navigation mode transition method as described in claim 1, characterized in that, The velocity in the filtering result of the previous first navigation mode is used as the initial velocity calculated in the second navigation mode.

6. An electronic device for controlling navigation mode transitions, characterized in that, It includes a memory and a processor, wherein the memory is configured to store a computer program that executes the navigation mode transition method as described in any one of claims 1 to 5 when the processor is running.

7. A computer-readable storage medium for controlling navigation mode transitions, characterized in that, The system contains a computer program that, when run on a processor, executes the navigation mode transition method as described in any one of claims 1 to 5.

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