Planning method of on-satellite trigger task and ground simulation system
By acquiring the target position and determining the windowing position through a precise positioning payload, and optimizing the spaceborne system in conjunction with a ground simulation system, the problem of inaccurate target positioning during onboard autonomous triggering was solved, thus improving observation efficiency and quality.
Patent Information
- Application Number
- CN202511591444.1
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-11-03
- Publication Date
- 2026-01-02
AI Technical Summary
In existing technologies, the target positioning is inaccurate during the on-board autonomous triggering process, which makes it impossible for the target to stably enter the center of the field of view, affecting the observation efficiency and quality. In addition, there are execution errors and environmental disturbances during attitude maneuvers, which cause the target to deviate from the center of the field of view, resulting in observation risks.
The target position is obtained by precise positioning payload, the window position of mission payload is determined, and the window size is determined according to the confidence level. The spaceborne system is verified and optimized by ground simulation system to ensure that the target accurately enters the center of the field of view for observation.
It effectively narrows the imaging range, improves the quality and efficiency of target observation, ensures the signal-to-noise ratio and spatial resolution of observation data, and avoids observation interruptions.
Smart Images
Figure CN121247092A_ABST
Abstract
Description
Technical Field
[0001] This invention relates to the field of aerospace technology, and in particular to a planning method for on-board triggered missions and a ground simulation system. Background Technology
[0002] As spacecraft on-orbit missions evolve towards higher autonomy and faster response speeds, onboard autonomous target acquisition and observation has become a core technology supporting dynamic missions. It can eliminate reliance on ground-based telemetry and control links, triggering attitude maneuvers through local onboard decisions to quickly bring targets into the field of view of benchmark observation payloads, such as optical telescopes (VT), to meet the timeliness requirements of observations in emergency missions or long-distance scenarios.
[0003] Currently, the onboard autonomous triggering system uses onboard coarse detection modules, such as wide-angle infrared detectors, low-resolution optical cameras, or target prediction modules based on orbital dynamics models, to initially identify targets and output coarse positioning information. After this coarse positioning information is transmitted to the onboard attitude control subsystem, the system generates and executes attitude maneuver commands to drive the satellite to adjust its attitude angle, ultimately guiding the target from outside the initial field of view of the VT to within the field of view, completing the basic steps of target entry pupil.
[0004] However, the target localization step in the aforementioned autonomous triggering process cannot achieve precise position control of the target within the VT field of view. On the one hand, the hardware accuracy of the coarse detection module is limited, with an angular resolution typically only 0.1° to 1°, and it is susceptible to environmental interference such as stray light and cosmic radiation, resulting in a large measurement deviation of the target's initial position. On the other hand, there are execution errors during attitude maneuvers, and the momentum wheel, thruster, and other actuators of the satellite attitude control subsystem have output delays. At the same time, disturbances such as solar radiation pressure and atmospheric drag in the on-orbit environment further affect the accuracy of attitude adjustment, ultimately causing the target to not stably land in the center of the VT field of view when it enters the field of view.
[0005] Target deviation from the center of the VT field of view can pose observation risks and limit mission effectiveness. Specifically, high-resolution VTs generally exhibit aberrations and sensitivity degradation at the edges of the field of view. When the target is in the edge region, the signal-to-noise ratio and spatial resolution of the imaging data will significantly decrease, failing to meet the accuracy requirements for subsequent target feature extraction and geometric parameter measurement. Furthermore, if the target deviation is too large, even minor attitude jitter during subsequent satellite attitude stabilization may cause the target to briefly leave the field of view, resulting in observation interruption. Simultaneously, the VT's fine tracking function requires an initial position within the field of view as a reference; a large initial deviation will exceed the adjustment range of the fine tracking mechanism or cause tracking response lag, ultimately leading to the loss of continuous target observation capability. Summary of the Invention
[0006] To improve observation efficiency, the first aspect of this invention provides a method for planning on-board triggered missions, comprising: After the mission is triggered on the satellite, the target's position is obtained through a precision positioning payload; Based on the current target orientation and the target position provided by the precise positioning payload, determine the windowing position of the mission payload; and Based on the window opening position, the observation task is performed.
[0007] Furthermore, the planning method also includes: Determine the credibility of the observed target; and The window size is determined based on the aforementioned confidence level.
[0008] Furthermore, based on the current target orientation and the target position provided by the precise positioning payload, determining the windowing position of the mission payload includes: Calculate the vector direction of the target in the fine positioning load coordinate system; Based on the target's vector direction in the precise positioning load coordinate system, determine the target's vector direction in the mission load coordinate system; Based on the target's vector direction in the mission payload coordinate system, determine the pixel coordinate position of the target in the mission payload detector; and The window position is determined based on the pixel coordinates.
[0009] Furthermore, the windowing position is calculated using a spaceborne system, and the planning method further includes: Before satellite launch, based on test data, the window position is calculated and compared using both the ground simulation system and the onboard system. If they match, the verification is successful; otherwise, the onboard system is updated and optimized.
[0010] Furthermore, the constraints are calculated using a ground simulation system, and the planning method further includes: After the mission is triggered on-board, the target's position information in the precise positioning payload coordinate system is transmitted; and The window opening position is calculated and uploaded based on the location information using a ground simulation system.
[0011] A second aspect of the present invention provides a ground simulation system, comprising: The input module is used to acquire the target's position information in the precise positioning load coordinate system; and A calculation module, communicatively connected to the input module, is used to calculate the window position based on the information using the method described above.
[0012] Furthermore, the system also includes a display module for displaying the information received by the input module and the calculation results of the calculation module.
[0013] This invention provides a planning method for on-board triggered missions. It acquires the current target position using a precision positioning payload and determines the windowing position of the target VT (Vibration Target) based on the target's orientation. Furthermore, the window size can be determined based on confidence level. This windowing process effectively reduces the payload's imaging range, improving target observation quality and efficiency. Attached Figure Description
[0014] 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.
[0015] Figure 1 A flowchart illustrating a method for planning an on-board triggered mission according to an embodiment of the present invention is shown. Figure 2 A schematic diagram showing the input module and display module of a ground simulation system according to an embodiment of the present invention is provided; and Figure 3 The diagram shows a flowchart illustrating a method for calculating constraints according to an embodiment of the present invention. Detailed Implementation
[0016] 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 in order to provide a comprehensive understanding of embodiments of the invention.
[0017] 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.
[0018] 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.
[0019] After being autonomously triggered, the satellite maneuvers to bring the target into the field of view of the reference payload, such as VT. However, this target positioning provides a coarse positioning, and the target may be located at a position that is far from the center of VT's field of view. Therefore, it is necessary to perform windowing processing on the target based on the position information provided by another fine positioning payload, such as the Soft X-ray Telescope (MXT).
[0020] Figure 1 This diagram illustrates a flowchart of a satellite-triggered mission planning method according to an embodiment of the present invention. Figure 1 As shown, a method for planning on-board triggered missions includes: First, in step 101, precise positioning information is acquired. After the mission is triggered on the satellite, the target's position is obtained through a precise positioning payload. In one embodiment of the present invention, the target's position information is determined using a precise positioning payload such as an MXT. Next, in step 102, the windowing position is determined. Based on the current target orientation and the target position provided by the precision positioning payload, the windowing position of the mission payload is determined. In one embodiment of the present invention, the target position provided by the precision positioning payload is first used to calculate the vector direction of the target in the precision positioning payload coordinate system: , in, , These are the azimuth coordinates of the target in the coordinate system of the precision positioning payload detector; Then, based on the target's vector direction in the precision positioning load coordinate system, the target's vector direction in the mission load coordinate system is determined: , in, The installation matrix of the precision positioning payload detector in the satellite body coordinate system, and This is the installation matrix of the mission payload detector in the satellite body coordinate system; Next, based on the target's vector direction in the mission payload coordinate system, the pixel coordinate position of the target in the mission payload detector is determined, including: First, determining the azimuth angle of the target in the mission payload detector according to the target's vector direction in the mission payload coordinate system: ; Next, based on the azimuth angle, calculate the rectangular coordinates of the target in the tangential plane of the mission payload detector: , , in, These are the azimuth angles corresponding to the center of the field of view of the mission payload detector, respectively, with default values of 0; and Finally, based on the Cartesian coordinates, the pixel coordinates of the target in the mission payload detector are calculated: , , Where a(0), a(1), a(2), a(3), a(4), a(5) are the horizontal axis calibration coefficients, with default values of 1024.0, 1.28205, 0.0, 0.0, 0.0, 0.0; b(0), b(1), b(2), b(3), b(4), b(5) are the vertical axis calibration coefficients, with default values of 1024.0, 1.28205, 0.0, 0.0, 0.0, 0.0; and Finally, the window position is determined based on the pixel coordinates: , , in, The window size is defined as follows: In one embodiment of the invention, the window size is determined based on the confidence level of the observed target. In one embodiment of the invention, if the confidence level is 0, the window size is 770; if the confidence level is 1, the window size is 770; if the confidence level is 2, the window size is 460; if the confidence level is 3, the window size is 300; and if the confidence level is 4, the window size is 230. Finally, in step 103, the observation task is executed. The observation task is performed based on the stated window position.
[0021] In one embodiment of the present invention, the window position is calculated using a spaceborne system. Therefore, the spaceborne system needs to be verified during the ground testing phase before satellite launch. Specifically, before satellite launch, based on test data and / or historical data, the window position is calculated and compared using both a ground simulation system and the spaceborne system. If they match, the verification is successful; otherwise, the spaceborne system is updated and optimized.
[0022] In one embodiment of the present invention, the constraints are calculated by a ground simulation system. Specifically, after the mission is triggered on the satellite, the position information of the target in the precise positioning payload coordinate system is sent down, and then the window position is calculated and uploaded based on the position information by the ground simulation system.
[0023] In one embodiment of the present invention, the ground simulation system includes an input module, a calculation module, and a display module. The input module is used to acquire the target's position information in a precise positioning load coordinate system. The calculation module is communicatively connected to the input module and is used to calculate the window position based on the position information using the method described above. The display module is used to display the information received by the input module and the calculation results from the calculation module. In one embodiment of the present invention, as... Figure 2 As shown, the input module includes an input box that can receive relevant information manually entered by the user. In one embodiment of the invention, the input module can also receive and display information transmitted from a satellite.
[0024] In one embodiment of the present invention, the input module and the display module are software interfaces, such as... Figure 2 As shown, after selecting the GRB-MXT mode from the drop-down menu, the user can input the target's azimuth coordinates (theta and phi) in the MXT, confidence level selection, and other information on the interface. Then, click the Calculate button, wait for the calculation results, and finally display the target's azimuth coordinates, position coordinates, and window position in the VT.
[0025] In one embodiment of the present invention, the ground simulation system can also be used for the calculation and simulation of constraints. Specifically, it calculates constraints based on satellite orbit, pointing, and mission information. The mission information includes the position of the observation target within the onboard payload, such as the ECLAirs observer's field of view. The constraints include attitude constraints such as the quaternion of the observation target, the angles between the satellite's +X-axis, +Y-axis, -Y-axis, and -Z-axis and the line connecting the Sun and Earth under the target's pointing direction, and the angle between the satellite's +X-axis and the line connecting the Moon and the satellite. During the ground testing phase before satellite launch, the simulation results can be compared with the calculation results of the onboard system to verify the performance of the onboard system. After the observation mission is triggered, simulation can be performed based on the position and other information of the observation target sent from the satellite, the constraints can be calculated, and compared with preset conditions. If the preset conditions are met, the mission is executed; otherwise, the mission is canceled. The preset conditions include, for example, thresholds for the angles between the satellite's +X-axis, +Y-axis, -Y-axis, and -Z-axis and the line connecting the Sun and Earth, and the angle between the satellite's +X-axis and the line connecting the Moon and the satellite. If the calculated angles between the satellite's +X axis, +Y axis, -Y axis, and -Z axis and the line connecting the Sun and Earth, as well as the angle between the satellite's +X axis and the line connecting the Moon and the satellite, are within the threshold range, then the mission is considered successful and executed; otherwise, the mission is considered unsuccessful and the current mission is canceled.
[0026] In one embodiment of the present invention, such as Figure 3 As shown, the calculation of the constraints includes: First, in step 301, the vector direction of the observed target is calculated. Based on the azimuth coordinates of the observed target in the detector coordinate system, the vector direction of the observed target in the detector coordinate system is determined and transformed to the J2000 coordinate system. In one embodiment of the present invention, the vector direction of the observed target in the detector coordinate system... The following formula is used for calculation: , in, , These are the azimuth coordinates of the observed target in the detector coordinate system, which can be transformed to the J2000 coordinate system using the mounting matrix and attitude matrix: , in The current satellite points to the quaternion. The corresponding attitude matrix, This is the installation matrix of the detector in the satellite's body coordinate system; Next, in step 302, the solar vector is calculated. The solar vector at the mission start time is calculated. In one embodiment of the invention, the solar vector is calculated according to the following formula: , in, The correction matrix is calculated as follows: , in, , ,in The Julian century number, expressed in centuries, is calculated from the satellite's current onboard time and the accumulated seconds relative to 00:00:00 on January 1, 2020. ; For the true ecliptic longitude of the sun, ,in The angle of the sun's near point. ;as well as The inclination of the sun's orbit. ; Next, in step 303, the attitude matrix is calculated. Based on the solar vector, the attitude matrix at the mission start time is calculated, wherein the attitude matrix aligns with the satellite's +X axis pointing towards the target, and the solar vector lies within the satellite's XOZ plane. The attitude matrix can be used as a coordinate transformation matrix to transform the solar vector to the local coordinate system, etc. In one embodiment of the invention, the attitude matrix... Based on the vector direction of the observed target in the J2000 coordinate system and the solar vector Sure: ; Next, in step 304, the quaternion of the observed target is calculated. The quaternion of the observed target is calculated based on the attitude matrix. In one embodiment of the invention, the attitude matrix is first represented as a square matrix: ; Then calculate and compare the following four values: , , ,as well as Then, based on the maximum value among the four values, the quaternion is determined. ,include: If the maximum value is ,but: ; If the maximum value is ,but: ; If the maximum value is ,but: ;as well as If the maximum value is ,but: ;as well as Finally, the quaternions are normalized. If the value is less than 0, then a negative value is taken; and Finally, in step 305, angular constraints are calculated. Based on the solar vector, angular constraints are calculated, including the angles between the satellite's +X-axis, +Y-axis, -Y-axis, and -Z-axis and the line connecting the Sun and Earth, and the angle between the satellite's +X-axis and the line connecting the Moon and the satellite. In one embodiment of the invention, calculating the angles between the satellite's +X-axis, +Y-axis, -Y-axis, and -Z-axis and the line connecting the Sun and Earth based on the solar vector includes: Calculate the solar vector in this system. : ;as well as Calculate the angles between the satellite's +X axis and the line connecting the Sun and Earth, the +Y axis and the line connecting the Sun and Earth, the -Y axis and the line connecting the Sun and Earth, and the -Z axis and the line connecting the Sun and Earth, respectively: .
[0027] In one embodiment of the present invention, the angle between the satellite's +X axis and the line connecting the lunar satellite is calculated according to the following steps: First, calculate the Moon-Earth vector in the inertial frame. ,in: ,in , ;as well as ,in , , , ,in: , , , in, , ; Next, based on the aforementioned inertial frame Moon-Earth vector and the satellite's position vector relative to Earth Determine the lunar satellite vector in the inertial frame. : ; Next, based on the lunar satellite vector in the inertial frame... and attitude matrix Determine the lunar satellite vector within this system. : ;as well as Finally, based on the lunar satellite vector in this system, the vector... Determine the angle between the satellite's + X-axis and the line connecting the lunar satellite. .
[0028] Therefore, based on the satellite orbit, the position of the observed target in the field of view of the onboard payload, the current direction of the satellite, and the mission start time, the quaternion of the observed target, the angle between the satellite +X axis and the line connecting the Sun and Earth under the target's direction, the angle between the satellite +Y axis and the line connecting the Sun and Earth under the target's direction, the angle between the satellite -Y axis and the line connecting the Sun and Earth under the target's direction, the angle between the satellite -Z axis and the line connecting the Sun and Earth under the target's direction, and the angle between the satellite +X axis and the line connecting the Moon and the Moon under the target's direction can be calculated. These values can then be compared with thresholds to determine whether the mission is feasible.
[0029] 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 planning method for on-board triggered missions, characterized in that, include: After the mission is triggered on the satellite, the target's position is obtained through a precision positioning payload; Based on the current target orientation and the target position provided by the precise positioning payload, determine the window position of the mission payload; as well as Based on the window opening position, the observation task is performed.
2. The planning method as described in claim 1, characterized in that, Based on the current target orientation and the target position provided by the precise positioning payload, determining the windowing position of the mission payload includes: Calculate the vector direction of the target in the fine positioning load coordinate system; Based on the target's vector direction in the precise positioning load coordinate system, determine the target's vector direction in the mission load coordinate system; Based on the target's vector direction in the mission payload coordinate system, determine the pixel coordinate position of the target in the mission payload detector; and The window position is determined based on the pixel coordinates.
3. The planning method as described in claim 2, characterized in that, The vector direction of the target in the precise positioning load coordinate system is calculated according to the following formula: ， in, , These are the azimuth coordinates of the target in the coordinate system of the precision positioning payload detector.
4. The planning method as described in claim 2, characterized in that, The vector direction of the target in the mission payload coordinate system is calculated using the following formula: ， in, The installation matrix of the precision positioning payload detector in the satellite body coordinate system, and This is the installation matrix of the mission payload detector in the satellite's body coordinate system.
5. The planning method as described in claim 2, characterized in that, Determining the pixel coordinate position of the target in the mission payload detector based on the target's vector direction in the mission payload coordinate system includes: Based on the target's vector direction in the mission payload coordinate system, determine the target's azimuth angle within the mission payload detector: ； Based on the azimuth angle, calculate the rectangular coordinates of the target in the tangential plane of the mission payload detector: ， ， in, These are the azimuth angles corresponding to the center of the field of view of the mission payload detector, respectively, with default values of 0; and Based on the rectangular coordinates, calculate the pixel coordinate position of the target in the mission payload detector: ， ， Where a(0), a(1), a(2), a(3), a(4), a(5) are the horizontal axis calibration coefficients, with default values of 1024.0, 1.28205, 0.0, 0.0, 0.0, 0.0; b(0), b(1), b(2), b(3), b(4), b(5) are the vertical axis calibration coefficients, with default values of 1024.0, 1.28205, 0.0, 0.0, 0.0, 0.
0.
6. The planning method as described in claim 2, characterized in that, The window opening position is calculated according to the following formula: ， ， in, This refers to the window size.
7. The planning method as described in claim 6, characterized in that, Also includes: Determine the credibility of the observed target; as well as The window size is determined based on the confidence level. If the confidence level is 0, the window size is 770. If the confidence level is 1, the window size is 770. If the confidence level is 2, the window size is 460. If the confidence level is 3, the window size is 300. If the confidence level is 4, the window size is 230.
8. The planning method as described in claim 1, characterized in that, The window location is calculated using a spaceborne system, and the planning method further includes: Before satellite launch, based on test data, the window position is calculated and compared using both the ground simulation system and the onboard system. If they match, the verification is successful; otherwise, the onboard system is updated and optimized.
9. The planning method as described in claim 1, characterized in that, The constraints are calculated using a ground simulation system, and the planning method further includes: After the mission is triggered on-board, the target's position information in the precise positioning payload coordinate system is transmitted; and The window opening position is calculated and uploaded based on the location information using a ground simulation system.
10. A ground simulation system, characterized in that, include: The input module is configured to acquire the target's position information in the fine positioning load coordinate system; A calculation module, communicatively connected to the input module, and configured to calculate the window position based on the location information using the method described in any one of claims 1 to 9; as well as The display module is configured to display the information received by the input module and the calculation results of the calculation module.