Satellite orbit adjustment time orbit determination filter autonomous initialization method, system, medium and device

Abstract

The application provides a satellite orbit adjustment time orbit determination filter autonomous initialization method, system, medium and equipment, comprising the following steps: step 1, adjacent acceleration difference out-of-range event identification in a unit time is performed; step 2, burr identification and disposal of the out-of-range event are performed; and step 3, thruster ignition or thruster shutdown event identification and disposal are performed. The application realizes the function of autonomously initializing the filter at the thruster ignition time and the thruster shutdown time, and ensures that the satellite orbit adjustment time navigation receiver can quickly output high-precision positioning data and orbit determination data.

Description

Technical Field

[0001] This invention relates to the field of satellite orbit technology, specifically to a method, system, medium, and device for autonomous initialization of orbit determination filters during satellite orbit adjustment. Background Technology

[0002] The mechanical factors influencing satellite orbital perturbations include: the inhomogeneity of Earth's gravitational field, atmospheric drag, celestial gravity, solar radiation pressure, and the geomagnetic field. Therefore, satellite orbits require periodic adjustments. Taking a certain type of 400km orbital satellite as an example, its semi-major axis decays by approximately 600m per day, requiring two autonomous orbit control maneuvers daily and one ground-based adjustment every two weeks. With the development of electric thruster technology, satellites widely employ autonomous orbit adjustment technology to reduce the frequency of ground-based intervention.

[0003] Spaceborne navigation receivers can process positioning and orbit determination data in real time, replacing ground-based telemetry and control systems for orbit determination and track marking. To improve orbit determination and positioning accuracy, Kalman filters are commonly used, suitable for processing orbit determination and positioning data during normal orbit operation. However, due to the trade-off between the accuracy and convergence time of the orbit determination filter, the filtered positioning and orbit determination data cannot promptly reflect changes in the orbit at the time of thruster ignition, resulting in the output positioning and orbit determination data failing to meet the load's position accuracy requirements.

[0004] Based on the aforementioned technological status, the orbit determination filter needs to be initialized during satellite orbit adjustment to clear accumulated prior parameters and enable the filter to converge quickly again, thus ensuring the accuracy of the positioning and orbit determination data output by the navigation receiver. Currently, orbit determination filter initialization mainly includes two methods: autonomous initialization based on orbit adjustment telemetry and passive initialization based on ground commands. Both methods require external input parameters.

[0005] Patent document CN101435863A discloses a method for real-time precise orbit determination of navigation satellites. This method divides the satellite orbit into continuous, equally long short arc segments based on observation time, performs sliding processing according to these short arc segments, and establishes observation equations for the initial state and mechanical model parameters of the satellite within each arc segment, thereby improving orbit determination accuracy. However, this invention does not analyze the orbit determination accuracy during satellite orbit adjustment, resulting in discrepancies with actual application in the model.

[0006] Patent document CN102305630A discloses a method for autonomous orbit determination of SAR satellites based on extended Kalman filtering. This method establishes satellite motion equations based on orbital dynamics, observation equations using the distance between the SAR and ground markers and the Doppler frequency shift between the SAR and ground markers as observation quantities, and a recursive equation for extended Kalman filtering to obtain satellite state information. However, this invention is used to achieve autonomous satellite orbit determination, without considering the orbit determination accuracy analysis during satellite orbit adjustment.

[0007] Patent document CN10200221A discloses a method for calibrating an accelerometer using orbit determination data. This method corrects the pulse equivalent during track control, ensuring control accuracy. In contrast, this invention uses acceleration differences to initialize the orbit determination filter, enabling the navigation receiver's orbit determination filter to reconverge and rapidly output high-precision positioning and orbit determination data.

[0008] Patent document CN103424116A discloses a precise orbit determination method for geostationary satellites adapted to orbital maneuvers. This method uses computer dynamic acceleration, observation data, and observation residual sequences to obtain the system deviation of the ground control station. After subtracting the system deviation, the observation data is used to re-improve the orbit. In contrast, this invention does not use ground station data. The navigation receiver autonomously determines the satellite's orbit and promptly corrects the navigation receiver's filter error after orbit adjustment, enabling rapid output of high-precision positioning and orbit determination data.

[0009] Patent document CN1959430A discloses a precise positioning system and its implementation method for medium and low Earth orbit satellites. It utilizes a novel precise orbit determination method combining the Galileo positioning system, satellite laser ranging technology, and satellite communication technology, but does not address orbit determination methods after satellite orbit adjustment. This invention primarily aims to reduce orbit determination errors and filtering convergence time after satellite orbit adjustment, thereby ensuring the timeliness of positioning and orbit determination data output. Summary of the Invention

[0010] To address the shortcomings of existing technologies, the purpose of this invention is to provide a method, system, medium, and device for autonomous initialization of orbit determination filters during satellite orbit adjustment.

[0011] The satellite orbit determination filter autonomous initialization method provided by the present invention includes:

[0012] Step 1: Identify events where the difference between adjacent accelerations exceeds the range within a unit of time;

[0013] Step 1.1: Continuously calculate the difference in acceleration between adjacent single points after the navigation receiver has determined the position and orbit, and divide it by the adjacent time interval to obtain the difference in acceleration between adjacent points per unit time in real time;

[0014] Step 1.2: Set the threshold for exceeding the range of adjacent acceleration differences per unit time;

[0015] Step 1.3: In real time, determine whether the difference in acceleration between adjacent points under positioning and orbit determination conditions per unit time is greater than the out-of-range threshold;

[0016] Step 1.4: When the difference between adjacent accelerations per unit time is greater than the out-of-range threshold, glitch identification and handling are performed; when the difference between adjacent accelerations is not greater than the out-of-range threshold, thruster ignition or thruster shutdown events are identified.

[0017] Step 2: Identify and handle glitch events that exceed the permitted scope;

[0018] Step 2.1: Set a filtering threshold when the difference between adjacent accelerations per unit time exceeds the out-of-range threshold;

[0019] Step 2.2: If the time difference between two adjacent acceleration differences exceeding the range within a unit time is less than the filtering threshold, it is judged as a glitch;

[0020] Step 2.3: When the acceleration difference exceeding the range event between adjacent units of time is a glitch, initialize the acceleration difference exceeding the range time within the unit of time and set the acceleration difference exceeding the range time within the unit of time to a fixed reference value.

[0021] Step 2.4: When the difference between adjacent accelerations per unit time is a non-glitch jump, update the acceleration difference when it exceeds the range;

[0022] Step 3: Identify and handle thrust ignition or thrust shutdown events;

[0023] Step 3.1: When the difference between adjacent accelerations per unit time is not greater than the out-of-range threshold, calculate the out-of-range duration by subtracting the current time from the out-of-range time.

[0024] Step 3.2: Set the time difference threshold corresponding to the duration of exceeding the range;

[0025] Step 3.3: When there is a jump in the adjacent acceleration difference within a unit time, the subsequent acceleration difference maintains a stable change state after the jump and the duration is within the time difference threshold. The satellite is identified as having a thruster ignition event or a thruster shutdown event during orbit adjustment, and the orbit determination filter is initialized autonomously. At the same time, the acceleration difference exceeding the range within a unit time is initialized autonomously.

[0026] Preferably, in step 1.3, the acceleration difference per unit time when the satellite's minimum thruster ignites is used as the out-of-range threshold.

[0027] Preferably, in step 2.1, half of the shortest ignition time of the satellite thruster is used as the filtering threshold for glitch events.

[0028] Preferably, in step 3.2, half of the shortest ignition time of the satellite thruster is used as the lower limit of the time difference threshold, and the difference between the satellite launch time and the fixed reference value is used as the upper limit of the time difference threshold.

[0029] The satellite orbit adjustment and orbit determination filter autonomous initialization system provided by the present invention includes:

[0030] Module M1: Identifies events where the difference between adjacent accelerations exceeds the range within a unit of time;

[0031] Module M1.1: Continuously calculates the difference between the positioning accelerations of adjacent single points after the navigation receiver has determined the positioning and orbit, and divides it by the adjacent time interval to obtain the difference in adjacent accelerations per unit time in real time;

[0032] Module M1.2: Sets the out-of-range threshold for the difference between adjacent accelerations per unit time;

[0033] Module M1.3: Real-time determination of whether the difference in acceleration between adjacent points under positioning and orbit determination conditions per unit time exceeds the out-of-range threshold;

[0034] Module M1.4: When the difference between adjacent accelerations per unit time is greater than the out-of-range threshold, glitch identification and handling are performed; when the difference between adjacent accelerations is not greater than the out-of-range threshold, thruster ignition or thruster shutdown events are identified.

[0035] Module M2: Performs glitch identification and handling for out-of-range events;

[0036] Module M2.1: Sets the filtering threshold when the difference between adjacent accelerations per unit time exceeds the out-of-range threshold;

[0037] Module M2.2: Events where the time difference between two adjacent acceleration differences exceeding the range within a unit time period is less than the filtering threshold are judged as glitches;

[0038] Module M2.3: When the acceleration difference exceeding the range event between adjacent units of time is a glitch, initialize the acceleration difference exceeding the range time within the unit of time and set the acceleration difference exceeding the range time within the unit of time to a fixed reference value;

[0039] Module M2.4: When the difference between adjacent accelerations per unit time is a non-glitch jump, update the acceleration difference when it exceeds the range;

[0040] Module M3: Recognizes and handles thrust ignition or thrust shutdown events;

[0041] Module M3.1: When the difference between adjacent accelerations within a unit of time is not greater than the out-of-range threshold, the out-of-range duration is obtained by subtracting the current time from the out-of-range time.

[0042] Module M3.2: Sets the time difference threshold corresponding to the duration of exceeding the range;

[0043] Module M3.3: When there is a jump in the adjacent acceleration difference within a unit time, the subsequent acceleration difference maintains a stable change state after the jump and the duration is within the time difference threshold. It identifies the satellite as a thruster ignition event or thruster shutdown event during orbit adjustment, and autonomously initializes the orbit determination filter. At the same time, it autonomously initializes the acceleration difference when it exceeds the range within a unit time.

[0044] Preferably, in module M1.3, the acceleration difference per unit time when the satellite's minimum thruster ignites is used as the out-of-range threshold.

[0045] Preferably, in module M2.1, half of the shortest ignition time of the satellite thruster is used as the filtering threshold for glitch events.

[0046] Preferably, in module M3.2, half of the shortest ignition time of the satellite thruster is used as the lower limit of the time difference threshold, and the difference between the satellite launch time and a fixed reference value is used as the upper limit of the time difference threshold.

[0047] According to the computer-readable storage medium storing a computer program provided by the present invention, when the computer program is executed by a processor, it implements the steps of the autonomous initialization method for the orbit determination filter during satellite orbit adjustment.

[0048] The electronic device provided by the present invention includes a memory, a processor, and a computer program stored in the memory and executable on the processor. When the computer program is executed by the processor, it implements the steps of the autonomous initialization method for the orbit determination filter during satellite orbit adjustment.

[0049] Compared with the prior art, the present invention has the following beneficial effects:

[0050] (1) This invention solves the problem of long reconvergence time of the orbit determination filter of the navigation receiver, ensuring that the navigation receiver can output high-precision positioning data and orbit determination data in a timely and autonomous manner, so as to meet the position accuracy requirements of the payload imaging.

[0051] (2) This invention does not require external input parameters. It uses real-time positioning acceleration data to autonomously identify whether the satellite has experienced an orbit adjustment event and whether the acceleration jump is a glitch signal, so as to achieve autonomous initialization of the orbit determination filter when the thruster is ignited and shut down. Attached Figure Description

[0052] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings:

[0053] Figure 1 Characteristic diagram of the jump change of adjacent acceleration difference per unit time 10 minutes after thruster ignition;

[0054] Figure 2 This is a feature map of the jump in the difference between adjacent accelerations per unit time caused by data jumps;

[0055] Figure 3 This is a characteristic diagram of the jump in adjacent acceleration difference per unit time caused by data jumps during thruster ignition.

[0056] Figure 4 This is a flowchart of the autonomous initialization process of the orbit determination filter during satellite orbit adjustment. Detailed Implementation

[0057] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the protection scope of the present invention.

[0058] Example 1

[0059] After powering on, the navigation receiver initializes for out-of-range times, begins receiving navigation signals, and performs positioning and orbit determination. Following positioning and orbit determination, it autonomously identifies satellite orbit adjustments according to the method of this invention.

[0060] like Figure 4 This invention proposes an autonomous initialization method for orbit determination filters during satellite orbit adjustment. The method specifically includes three events: identification of adjacent acceleration difference exceeding the range within a unit time, identification and handling of glitch events exceeding the range, and identification and handling of thruster ignition or thruster shutdown events.

[0061] 1. Identification of events where the difference between adjacent accelerations exceeds the range per unit time.

[0062] To avoid abrupt changes in acceleration differences caused by discontinuous data from the navigation receiver, the difference in acceleration is divided by the time interval to obtain the difference between adjacent accelerations per unit time. In the WGS84 coordinate system, the difference between adjacent accelerations per unit time fluctuates within a very small range (less than 0.0001 m / s²). During thruster ignition or shutdown, the change in acceleration difference generated by the satellite is related to the thrust magnitude and the satellite's weight, but the change in acceleration difference is much greater than 0.0001 m / s². Therefore, the change in acceleration difference can be used to identify thruster ignition or shutdown events. (Appendix) Figure 1 The characteristic diagram of the jump in adjacent acceleration difference within a unit time of 10 minutes after the ignition of a 10N thruster can clearly identify the moment when the thruster ignites or shuts down.

[0063] This invention sets an out-of-range threshold for the difference between adjacent accelerations per unit time. The acceleration difference per unit time during the ignition of the satellite's minimum thrust is used as the out-of-range threshold, which can be modified according to the on-orbit conditions. For example, for a satellite weighing 3850 kg with a minimum thrust of 10 N, the resultant acceleration difference during thrust ignition or de-ignition is approximately 0.0025 m / s². Therefore, this out-of-range threshold can be set to 0.001 m / s².

[0064] When the difference between adjacent accelerations per unit time exceeds the out-of-range threshold, the process enters the glitch identification and handling stage; when the difference between adjacent accelerations does not exceed the out-of-range threshold, the process enters the thruster ignition or thruster shutdown event identification stage.

[0065] 2. Burr identification and handling of out-of-range events

[0066] When the acceleration data output by the navigation receiver changes abruptly, out-of-range events can also occur. Such glitches need to be removed to prevent incorrect initialization of the orbit determination filter.

[0067] The method proposed in this invention uses the difference between adjacent accelerations as the criterion for out-of-range events. A single acceleration data jump will result in two out-of-range events, and the interval between the two out-of-range events is short, generally lasting for the acceleration data update cycle (e.g., 1 second). Meanwhile, to avoid frequent autonomous orbit adjustments, satellites are designed with a minimum thruster ignition time requirement. Taking a satellite in a 400km orbit as an example, the minimum thruster ignition time is 30 seconds. Therefore, this invention uses the duration of the out-of-range event as a filtering threshold for glitch events. Half of the minimum thruster ignition time can be used as the filtering threshold for glitch events, and this threshold can be modified according to the on-orbit situation. When an out-of-range event of the acceleration difference between adjacent units of time is a glitch, the out-of-range time of the acceleration difference within a unit of time is initialized, generally set to a fixed reference value.

[0068] Glitches caused by data transitions fall into two categories: glitch events caused by data transitions during normal orbit and glitch events caused by data transitions during thruster ignition. A glitch event during normal orbit corresponds to two acceleration difference transitions, as shown in the attached figure. Figure 2 As shown, because the duration of the event is less than the filter threshold, this event will not initialize the orbit determination filter. Glitches during thrust ignition are shown in the attached figure. Figure 3As shown, at the start of thrust ignition, the first jump time is used as the starting time for glitch detection. Since the difference between the data jump time and the ignition start time is less than the filter threshold, the first and second jumps are identified as glitches, and the out-of-range time is initialized. The third jump time is used again as the starting time for glitch detection. When the duration exceeds the filter threshold, the orbit determination filter is initialized for the first time, and the out-of-range time is also initialized. At the start of ignition shutdown, the starting time for glitch detection is used again. When the duration exceeds the filter threshold, the orbit determination filter is initialized for the second time, and the out-of-range time is also initialized.

[0069] 3. Identification and handling of thruster ignition or thruster shutdown events

[0070] At the moments of thruster ignition and extinguishing, the difference in acceleration between adjacent units of time changes normally. The time difference between two such changes is the out-of-range duration, which is also the satellite thruster ignition duration. This invention sets a time difference threshold for the out-of-range duration: half of the shortest ignition duration of the satellite thruster is used as the lower limit of the time difference threshold, and the difference between the satellite launch time and a fixed reference value is used as the upper limit of the time difference threshold. The time difference threshold can be modified according to the on-orbit situation. If the minimum ignition duration of the satellite thruster is 30 seconds, the lower limit of the time difference threshold can be set to 15 seconds, and the upper limit can be set to the difference between the satellite launch time and the initial out-of-range time.

[0071] When the thruster ignition time is 10 minutes, the thruster ignition start moment corresponds to the first data jump occurrence time. If the difference between this moment and the fixed reference value is greater than the upper limit of the time difference threshold, the out-of-range time is updated to the data jump occurrence time, and no operation is performed on the orbit determination filter. If no second jump occurs within 15 seconds (i.e., the duration is greater than the lower limit of the time difference threshold but less than the upper limit), this data jump is identified as a thruster ignition or thruster shutdown event, and the orbit determination filter and the out-of-range time are initialized. After the orbit determination filter is initialized, the navigation receiver uses the latest single-point positioning data for reorientation, ensuring that the positioning and orbit determination data output after thruster ignition meets the position accuracy requirements of payload imaging.

[0072] The thruster shutdown time corresponds to the second data jump occurrence time. If the difference between this time and the fixed reference value is greater than the upper limit of the time difference threshold, the out-of-range time is updated to the data jump occurrence time, and no operation is performed on the orbit determination filter. If no third jump occurs within 15 seconds (i.e., the duration is greater than the lower limit of the time difference threshold but less than the upper limit), this data jump is identified as a thruster ignition or shutdown event, and the orbit determination filter and the out-of-range time are initialized. After the orbit determination filter is initialized, the navigation receiver uses the latest single-point positioning data for reorientation, ensuring that the positioning and orbit determination data output after thruster shutdown meets the position accuracy requirements of payload imaging.

[0073] Example 2

[0074] The present invention also provides an autonomous initialization system for orbit determination filters during satellite orbit adjustment. The autonomous initialization system for orbit determination filters during satellite orbit adjustment can be implemented by executing the process steps of the autonomous initialization method for orbit determination filters during satellite orbit adjustment. That is, those skilled in the art can understand the autonomous initialization method for orbit determination filters during satellite orbit adjustment as a preferred embodiment of the autonomous initialization system for orbit determination filters during satellite orbit adjustment.

[0075] The satellite orbit adjustment and orbit determination filter autonomous initialization system provided by the present invention includes:

[0076] Module M1: Identifies events where the difference between adjacent accelerations exceeds the range within a unit of time;

[0077] Module M1.1: Continuously calculates the difference between the positioning accelerations of adjacent single points after the navigation receiver has determined the positioning and orbit, and divides it by the adjacent time interval to obtain the difference in adjacent accelerations per unit time in real time;

[0078] Module M1.2: Sets the out-of-range threshold for the difference between adjacent accelerations per unit time;

[0079] Module M1.3: Real-time determination of whether the difference in acceleration between adjacent points under positioning and orbit determination conditions per unit time exceeds the out-of-range threshold;

[0080] Module M1.4: When the difference between adjacent accelerations per unit time is greater than the out-of-range threshold, glitch identification and handling are performed; when the difference between adjacent accelerations is not greater than the out-of-range threshold, thruster ignition or thruster shutdown events are identified.

[0081] Module M2: Performs glitch identification and handling for out-of-range events;

[0082] Module M2.1: Sets the filtering threshold when the difference between adjacent accelerations per unit time exceeds the out-of-range threshold;

[0083] Module M2.2: Events where the time difference between two adjacent acceleration differences exceeding the range within a unit time period is less than the filtering threshold are judged as glitches;

[0084] Module M2.3: When the acceleration difference exceeding the range event between adjacent units of time is a glitch, initialize the acceleration difference exceeding the range time within the unit of time and set the acceleration difference exceeding the range time within the unit of time to a fixed reference value;

[0085] Module M2.4: When the difference between adjacent accelerations per unit time is a non-glitch jump, update the acceleration difference when it exceeds the range;

[0086] Module M3: Recognizes and handles thrust ignition or thrust shutdown events;

[0087] Module M3.1: When the difference between adjacent accelerations within a unit of time is not greater than the out-of-range threshold, the out-of-range duration is obtained by subtracting the current time from the out-of-range time.

[0088] Module M3.2: Sets the time difference threshold corresponding to the duration of exceeding the range;

[0089] Module M3.3: When there is a jump in the adjacent acceleration difference within a unit time, the subsequent acceleration difference maintains a stable change state after the jump and the duration is within the time difference threshold. It identifies the satellite as a thruster ignition event or thruster shutdown event during orbit adjustment, and autonomously initializes the orbit determination filter. At the same time, it autonomously initializes the acceleration difference when it exceeds the range within a unit time.

[0090] In module M1.3, the acceleration difference per unit time when the satellite's minimum thruster ignites is used as the out-of-range threshold.

[0091] In module M2.1, half of the shortest ignition time of the satellite thruster is used as the filtering threshold for glitch events.

[0092] In module M3.2, half of the shortest ignition time of the satellite thruster is used as the lower limit of the time difference threshold, and the difference between the satellite launch time and the fixed reference value is used as the upper limit of the time difference threshold.

[0093] Those skilled in the art will understand that, in addition to implementing the system, apparatus, and their modules provided by this invention in purely computer-readable program code, the same program can be implemented in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers by logically programming the method steps. Therefore, the system, apparatus, and their modules provided by this invention can be considered a hardware component, and the modules included therein for implementing various programs can also be considered structures within the hardware component; alternatively, modules for implementing various functions can be considered both software programs implementing the method and structures within the hardware component.

[0094] Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. A method for autonomous initialization of a station keeping filter during orbit adjustment of a satellite, characterized in that, The method comprises the following steps: Step 1: identifying the adjacent acceleration difference out-of-range event in unit time; Step 1.1: continuously subtracting the adjacent single-point positioning accelerations after the positioning and orbiting of the navigation receiver and dividing by the adjacent time interval to obtain the adjacent acceleration difference in unit time in real time; Step 1.2: setting the out-of-range threshold of the adjacent acceleration difference in unit time; Step 1.3: judging whether the adjacent acceleration difference in unit time under the positioning and orbiting condition is greater than the out-of-range threshold in real time; Step 1.4: identifying the burr and disposing when the adjacent acceleration difference in unit time is greater than the out-of-range threshold, and identifying the thruster ignition or thruster extinction event when the adjacent acceleration difference is not greater than the out-of-range threshold; Step 2: identifying and disposing the burr of the out-of-range event; Step 2.1: setting the filtering threshold when the adjacent acceleration difference in unit time is greater than the out-of-range threshold; Step 2.2: judging that the burr is the time difference between two adjacent acceleration difference out-of-range events in unit time being less than the filtering threshold; Step 2.3: initializing the acceleration difference out-of-range time in unit time when the adjacent acceleration difference out-of-range event in unit time is the burr, and setting the acceleration difference out-of-range time in unit time as a fixed reference value; Step 2.4: updating the acceleration difference out-of-range time when the adjacent acceleration difference in unit time is a non-burr jump; Step 3: identifying and disposing the thruster ignition or thruster extinction event; Step 3.1: obtaining the out-of-range duration by subtracting the current time from the out-of-range time when the adjacent acceleration difference in unit time is not greater than the out-of-range threshold; Step 3.2: setting the time difference threshold corresponding to the out-of-range duration; Step 3.3: identifying the thruster ignition event or thruster extinction event of the orbiting satellite when the time duration of the subsequent acceleration difference after the jump is within the time difference threshold range, and autonomously initializing the orbiting filter and the acceleration difference out-of-range time in unit time.

2. The autonomous initialization method for a satellite orbit determination filter during orbit maneuvering according to claim 1, characterized in that, In the step 1.3, the acceleration difference value in unit time when the satellite minimum thruster is ignited is taken as the out-of-range threshold.

3. The autonomous initialization method for a satellite orbit determination filter during orbit maneuvering according to claim 1, wherein, In the step 2.1, 1 / 2 of the shortest thruster ignition duration of the satellite is taken as the filtering threshold of the burr event.

4. The autonomous initialization method for a satellite orbit determination filter during orbit maneuvering according to claim 1, wherein, In the step 3.2, 1 / 2 of the shortest thruster ignition duration of the satellite is taken as the lower limit of the time difference threshold, and the difference between the satellite launch time and the fixed reference value is taken as the upper limit of the time difference threshold.

5. A system for autonomous initialization of a station keeping filter during orbit adjustment of a satellite, characterized in that, The method comprises the following steps: Module M1: identifying the adjacent acceleration difference out-of-range event in unit time; Module M1.1: continuously subtracting the adjacent single-point positioning accelerations after the positioning and orbiting of the navigation receiver and dividing by the adjacent time interval to obtain the adjacent acceleration difference in unit time in real time; Module M1.2: setting the out-of-range threshold of the adjacent acceleration difference in unit time; Module M1.3: judging whether the adjacent acceleration difference in unit time under the positioning and orbiting condition is greater than the out-of-range threshold in real time; Module M1.4: identifying the burr and disposing when the adjacent acceleration difference in unit time is greater than the out-of-range threshold, and identifying the thruster ignition or thruster extinction event when the adjacent acceleration difference is not greater than the out-of-range threshold. Module M1.4: Spur identification and disposal when the adjacent acceleration difference in unit time is greater than the out-of-range threshold, thruster ignition or thruster extinction event identification when the adjacent acceleration difference in unit time is not greater than the out-of-range threshold; Module M2: Spur identification and disposal of out-of-range events; Module M2.1: Set a filter threshold when the adjacent acceleration difference in unit time is greater than the out-of-range threshold; Module M2.2: The time difference between two adjacent acceleration difference in unit time out-of-range events is less than the filter threshold, and is determined to be a spur; Module M2.3: When the adjacent acceleration difference in unit time is an out-of-range event, initialize the acceleration difference in unit time, and set the acceleration difference in unit time to a fixed reference value; Module M2.4: When the adjacent acceleration difference in unit time is a non-spur jump, update the acceleration difference out-of-range time; Module M3: Thruster ignition or thruster extinction event identification and disposal; Module M3.1: When the adjacent acceleration difference in unit time is not greater than the out-of-range threshold, the out-of-range duration is obtained by subtracting the current time from the out-of-range time; Module M3.2: Set a time difference threshold corresponding to the out-of-range duration; Module M3.3: When the adjacent acceleration difference in unit time jumps once, the subsequent acceleration difference remains in a stable state after the jump, and the duration is within the time difference threshold range to identify the thruster ignition event or thruster extinction event of the satellite in orbit transfer, and automatically initialize the orbit transfer filter, and initialize the acceleration difference in unit time.

6. The autonomous initialization of a satellite orbit determination filter during orbit maneuvering system of claim 5 wherein, In the module M1.3, the acceleration difference in unit time when the satellite minimum thruster is ignited is used as the out-of-range threshold.

7. The autonomous initialization system for a station-keeping filter during orbit maneuvering of a satellite of claim 5, wherein, In the module M2.1, half of the shortest thruster ignition duration of the satellite is used as the filter threshold of the spur event.

8. The autonomous initialization system for a station-keeping filter during orbit maneuvering of a satellite of claim 5, wherein, In the module M3.2, half of the shortest thruster ignition duration of the satellite is used as the lower limit of the time difference threshold, and the difference between the satellite launch time and the fixed reference value is used as the upper limit of the time difference threshold.

9. A computer readable storage medium storing a computer program, characterized in that, The computer program is executed by the processor to realize the steps of the satellite orbit transfer filter automatic initialization method in any one of claims 1 to 4.

10. An electronic device comprising a memory, a processor, and a computer program stored on the memory and executable on the processor, characterized in that, The computer program is executed by the processor to realize the steps of the satellite orbit transfer filter automatic initialization method in any one of claims 1 to 4.

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