Multi-load observation task planning method

Abstract

The invention provides a multi-load observation task planning method, which comprises the following steps of: after any load on a satellite triggers an observation task, converting the position of an observation target into a position in a reference load view field, then calculating a constraint condition according to the position of the observation target in the reference load view field, comparing the constraint condition with a preset condition, and if the preset condition is met, determining that the observation target is a target. If not, the task is cancelled. By setting the reference load and calculating the constraint condition based on the reference load, it can be ensured that all loads in the observation task can observe the target, and then the task success rate and the observation accuracy are ensured.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aerospace technology, in particular to a multi-load observation task planning method. BACKGROUND

[0002] In the on-orbit task of a spacecraft, multi-load collaborative observation has become one of the core modes to improve task efficiency and data integrity. Such a task usually triggers a target signal by a certain specific load on the satellite first, and then needs to schedule other associated loads to observe the target synchronously to obtain multi-dimensional and multi-parameter target information to support subsequent data processing and task decision-making.

[0003] However, in actual engineering applications, it is difficult to achieve complete coaxiality between multiple collaborative loads on the satellite due to the influence of the structural design of the spacecraft, load installation process and complex on-orbit environment. Specifically, the generation of the visual axis deviation is mainly caused by three factors: first, the mechanical positioning error existing in the load installation process, which causes the optical axes of the loads to be unable to be strictly aligned; second, the temperature alternation, space micro-vibration and other environmental factors during the on-orbit operation of the spacecraft, which will cause the slight deformation of the load bracket or optical structure, further aggravating the visual axis deviation; third, the inherent deviation of the optical system of part of the loads will also cause the actual observation visual axis to deviate from the theoretical design visual axis.

[0004] The current multi-load collaborative observation scheme mostly assumes that the visual axes of the loads are in a coaxial state, or only uses a simplified deviation compensation strategy, without fully considering the dynamic and complex visual axis deviation. This directly leads to the fact that in the actual observation process, part of the loads cannot effectively capture the target signal due to the visual axis deviation exceeding the observation field of view range, and further causes problems such as missing observation data and mismatching of multi-load data. For example, in a space target tracking task, after the target is triggered by the guide load, the imaging load may miss the target due to the visual axis deviation, resulting in the inability to obtain a clear image of the target; in an earth observation task, the visual axis deviation of the spectral load and the imaging load will cause the spatial position of the spectral data and the image data to be mismatched, greatly reducing the application value of the data. SUMMARY

[0005] In order to ensure that each load can observe the target in a multi-load task, the present application provides a multi-load observation task planning method, which comprises: After any load on the satellite triggers an observation task, the position of the observation target is converted into a position in the field of view of a reference load; and The constraint condition is calculated according to the position of the observation target in the field of view of the reference load, and compared with the preset condition, if the preset condition is met, the task is executed, otherwise the task is cancelled.

[0006] Further, the reference load is the load with the smallest field of view among the loads required for the observation task.

[0007] Further, the position of the observation target is converted into a position in the reference load field of view through a bias matrix between the reference load.

[0008] Further, the constraint condition comprises: a quaternion of the observation target, a target pointing down, an included angle between a +X axis, a +Y axis, a -Y axis, a -Z axis of the satellite and a sun-earth line, and an included angle between the +X axis of the satellite and a moon-satellite line.

[0009] Further, the calculation of the constraint condition based on the position of the observation target in the reference load field of view comprises: determining a vector direction of the observation target in a reference load coordinate system based on the position of the observation target in the reference load field of view, and converting into a J2000 coordinate system; calculating a sun vector at a task start time, and determining an attitude matrix of the satellite based on the sun vector and the vector direction of the observation target in the J2000 coordinate system, wherein the attitude matrix is an attitude matrix when the +X axis of the satellite points to the target and the sun vector is in the XOZ plane of the satellite; calculating a quaternion of the observation target based on the attitude matrix; calculating the included angle between the +X axis, the +Y axis, the -Y axis, the -Z axis of the satellite and the sun-earth line based on the sun vector; and calculating the included angle between the +X axis of the satellite and the moon-satellite line.

[0010] Further, the constraint condition is calculated by a satellite-borne system, and the planning method further comprises: before the satellite is launched, the constraint condition is calculated by a ground simulation system and a satellite-borne system respectively based on test data, and compared, if consistent, the verification is passed, if not consistent, the satellite-borne system is updated and optimized.

[0011] Further, the constraint condition is calculated by a ground simulation system, and the planning method further comprises: after the task is triggered on the satellite, the position information of the observation target is issued; the constraint condition is calculated based on the position information by the ground simulation system, and compared with a preset condition, if satisfied, the task is executed, otherwise, the observation task is cancelled.

[0012] The planning method for a multi-load observation task provided by the application determines a load with the smallest field of view as a measurement reference, calibrates and calculates a bias matrix of other loads and the reference load, and after the task is triggered, judges whether the constraint condition meets the requirements based on the reference load field of view, so as to ensure that all loads can see the target, and ensure the task efficiency and the accuracy of the observation result. BRIEF DESCRIPTION OF DRAWINGS

[0013] To further clarify the above and other advantages and features of the present embodiments, a more particular description of embodiments of the application will be rendered by reference to specific embodiments thereof which are illustrated in the appended drawings. It is appreciated that these drawings depict only typical embodiments of the application and are therefore not to be considered limiting of its scope. The drawings incorporate graphics for purposes of clarity. In the drawings:

[0014] Figure 1 A flow chart illustrating a method for planning a multi-load observation task according to an embodiment of the present application; and Figure 2 A flow chart illustrating a method for calculating constraints according to an embodiment of the present application. DETAILED DESCRIPTION

[0015] In the following description, the present application is described with reference to various embodiments thereof. However, it will be appreciated that the various embodiments can be practiced without one or more of the specific details set forth herein or with other methods or components. In other instances, well-known structures or operations are not shown or described in detail in order to avoid obscuring the inventive aspect of the present application. Similarly, specific numbers and configurations are set forth in this description in order to provide a thorough understanding of the embodiments of the present application.

[0016] In this specification, reference to "one embodiment" or "the embodiment" means that a particular feature, structure, or characteristic described is included in at least one embodiment of the present application. The appearance of the phrase "in one embodiment" in various places in the specification are not necessarily all referring to the same embodiment.

[0017] It is to be appreciated that embodiments of the present application are described in a particular order of steps, however, this is not meant to limit the present application to that particular order of steps. Rather, the steps of the present application can be performed in any order, as the steps of the present application are not interdependent.

[0018] In order to ensure that all loads can see the observation target in a multi-load simultaneous observation task, the present application provides a method for planning a multi-load observation task, which determines the load with the smallest field of view as a measurement reference, calibrates and calculates the bias matrix of other loads relative to the reference load, and calculates the target position based on the reference load.

[0019] The technical solutions of the present application will be further described below with reference to the accompanying drawings of the embodiments.

[0020] Figure 1 A flow chart illustrating a method for planning a multi-load observation task according to an embodiment of the present application. As shown in FIG. 1, the method comprises the following steps: Figure 1As shown, a multi-load observation task planning method comprises: First, in step 101, a reference load is determined. After a certain load on the satellite triggers a target position, it is determined whether the target needs to be observed by multiple loads simultaneously. If so, the load with the smallest field of view among the required multiple loads is taken as the reference load; Next, in step 102, a bias matrix is calculated. The bias matrix of the other loads in the multiple loads, especially the load triggering the task and the reference load, is calibrated and calculated. In an embodiment of the present application, the bias matrix refers to the boresight bias matrix of two loads; Next, in step 103, the target position is converted. Based on the bias matrix between the load triggering the task and the reference load, the position of the observation target is converted to the position in the field of view of the reference load; Next, in step 104, it is determined whether the constraint condition meets the requirements. The constraint condition is calculated according to the position of the observation target in the field of view of the reference load, and compared with the preset condition. If the preset condition is met, the task is executed in step 105, otherwise the task is cancelled in step 106. Since the field of view of the reference load is the smallest, as long as the target can meet the observation condition of the reference load, i.e. can be observed by the reference load, it can be observed by other loads, thereby meeting the requirement of multi-load observation.

[0021] In an embodiment of the present application, the constraint condition is calculated by the on-board system. In order to ensure the accuracy of the calculation result, in an embodiment of the present application, before the satellite is launched, the constraint condition is calculated by the ground simulation system and the on-board system based on test data or historical data respectively, and the calculation result of the on-board system is compared with the calculation result of the ground simulation system. If they are consistent, the verification is passed, otherwise the on-board system is updated and optimized.

[0022] In an embodiment of the present application, the constraint condition is calculated by the ground simulation system. Specifically, after the task is triggered on the satellite, the satellite sends down the position information of the observation target, and then the ground simulation system receives the position information, calculates the constraint condition based on the position information, i.e. the orbit, execution, etc. of the target, and compares it with the preset condition. If it meets the requirement, the task is executed, otherwise the observation task is cancelled.

[0023] In one embodiment of the present application, the constraint conditions are calculated based on satellite orbit, pointing and mission information. The mission information includes the position of the observation target in the field of view of the reference payload, etc. The constraint conditions include the attitude constraints such as the observation target quaternion, the target pointing direction, the angles between the +X axis, +Y axis, -Y axis, -Z axis of the satellite and the line connecting the sun and the earth, and the angle between the +X axis of the satellite and the line connecting the satellite and the moon, etc. The preset conditions include the threshold values of the angles between the +X axis, +Y axis, -Y axis, -Z axis of the satellite and the line connecting the sun and the earth, and the angle between the +X axis of the satellite and the line connecting the satellite and the moon, etc. When the calculated angles between the +X axis, +Y axis, -Y axis, -Z axis of the satellite and the line connecting the sun and the earth, and the angle between the +X axis of the satellite and the line connecting the satellite and the moon are within the threshold values, the mission is considered to be successful, and the mission is executed. Otherwise, the mission is considered to be unsuccessful, and the current mission is cancelled.

[0024] Figure 2 A flowchart of a method for calculating constraint conditions according to one embodiment of the present application is shown. As shown in FIG. 2, the calculation of the constraint conditions includes: Figure 2 First, in step 201, the vector direction of the observation target is calculated. Based on the position of the observation target in the field of view of the reference payload, the vector direction of the observation target in the reference payload coordinate system is determined and converted to the J2000 coordinate system. In one embodiment of the present application, the vector direction of the observation target in the reference payload coordinate system is calculated according to the following formula: wherein are the azimuth coordinates of the observation target in the reference payload coordinate system, which can be converted to the J2000 coordinate system through the installation matrix and the attitude matrix: wherein is the current satellite pointing quaternion corresponding attitude matrix, is the installation matrix of the reference payload in the satellite body coordinate system; Next, in step 202, the sun vector is calculated. The sun vector at the start time of the mission is calculated. In one embodiment of the present application, the sun vector is calculated according to the following formula: wherein is the correction matrix, which is calculated as follows: wherein wherein ​​​​​​​​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 203, 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 204, 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 present 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 less than 0, take the negative value; and Finally, at step 205, the angle constraints are calculated. Based on the sun vector, the angle constraints are calculated, including the angles between the +X axis, +Y axis, -Y axis, -Z axis of the satellite and the geodesic, and the angle between the +X axis of the satellite and the selenodesic. In one embodiment of the present application, the angles between the +X axis, +Y axis, -Y axis, -Z axis of the satellite and the geodesic are calculated based on the sun vector, including: The sun vector is calculated in the body frame : ; and The angles between the +X axis, +Y axis, -Y axis, -Z axis of the satellite and the geodesic are calculated respectively: .

[0025] In one embodiment of the present application, the angle between the +X axis of the satellite and the selenodesic is calculated according to the following steps: First, the inertial system selenodesic vector is calculated , wherein: , wherein , ; and , wherein , , , , wherein: , , , wherein, , ; Next, based on the inertial system selenodesic vector and the position vector of the satellite relative to the Earth , the vector of the selenodesic in the inertial system is determined : ; Next, based on the vector of the selenodesic in the inertial system and the attitude matrix , the vector of the selenodesic in the body frame is determined : ; and Finally, based on the vector of the selenodesic in the body frame determining the angle between the satellite +X axis and the moon- satellite connecting line .

[0026] Thus, based on the satellite orbit, the position of the observation target in the reference load field of view, the current satellite pointing, and the mission start time, the quaternion of the observation target, the angle between the satellite +X axis and the sun- satellite connecting line in the target pointing direction, the angle between the satellite +Y axis and the sun- satellite connecting line in the target pointing direction, the angle between the satellite -Y axis and the sun- satellite connecting line in the target pointing direction, the angle between the satellite -Z axis and the sun- satellite connecting line in the target pointing direction, and the angle between the satellite +X axis and the moon- satellite connecting line in the target pointing direction can be calculated and compared with the threshold value to determine whether the mission is feasible.

[0027] In addition, the application further provides a ground simulation system, which comprises an input module, a calculation module and a display module, wherein the input module is used for receiving the satellite orbit, pointing and mission information input by the user in the ground test stage, and is used for receiving the observation target position information sent by the satellite after the satellite is launched, the calculation module is communicatively connected with the input module, is used for calculating the constraint condition based on the mission information, satellite orbit, pointing and the like by the method as described above, and comparing with the preset condition, and the display module is used for displaying the input data, the received information and the calculation result of the calculation module. In an embodiment of the application, the input module and the display module are software interfaces, the user can input the UTC time, the current satellite position, the quaternion, the target right ascension and the target declination and the like in the corresponding input boxes of the interface, then click the calculation button, wait for the calculation result, and finally the calculation result is displayed at the constraint angle and the target quaternion.

[0028] Although the embodiments of the application are described above, it should be understood that they are presented only as examples, and not as limitations. It is obvious to those skilled in the related art that various combinations, modifications and changes can be made without departing from the spirit and scope of the application. Therefore, the width and scope of the application disclosed herein should not be limited by the above disclosed exemplary embodiments, but should be defined only according to the appended claims and their equivalent replacements.

Claims

1. A planning method for multi-payload observation tasks, characterized in that, include: After any onboard payload triggers an observation mission, the position of the observed target is converted into the position in the field of view of the reference payload; as well as The constraints are calculated based on the position of the observed target in the field of view of the reference payload, and compared with the preset conditions. If the preset conditions are met, the task is executed; otherwise, the task is canceled.

2. The planning method as described in claim 1, characterized in that, The reference load is the load with the smallest field of view among the loads required for the observation mission.

3. The planning method as described in claim 1, characterized in that, Based on the offset matrix between the triggering task's payload and the reference payload, the position of the observed target is converted into its position within the reference payload's field of view.

4. The planning method as described in claim 1, characterized in that, The constraints include: the quaternion of the observed target, 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.

5. The planning method as described in claim 1, characterized in that, The constraints calculated based on the position of the observed target in the field of view of the reference load include: Based on the position of the observed target in the reference load field of view, determine the vector direction of the observed target in the reference load coordinate system, and transform it to the J2000 coordinate system; Calculate the solar vector at the mission start time, and determine the satellite's attitude matrix based on the solar vector and the vector direction of the observed target in the J2000 coordinate system, wherein the attitude matrix is ​​the attitude matrix when the satellite's +X axis points to the target and the solar vector is located in the satellite's XOZ plane. The quaternion of the observed target is calculated based on the attitude matrix; Based on the solar vector, calculate the angles between the satellite's +X axis, +Y axis, -Y axis, and -Z axis and the line connecting the Sun and Earth; and Calculate the angle between the satellite's +X axis and the line connecting the moon and the satellite.

6. The planning method as described in claim 5, characterized in that, The attitude matrix of the satellite is determined according to the following formula: ， in, ,in The current satellite points to the quaternion. The corresponding attitude matrix, The installation matrix of the detector in the satellite's body coordinate system, and The vector direction of the observed target in the reference load coordinate system is: ， in, , These are the azimuth coordinates of the observed target in the reference load coordinate system; and The solar vector is calculated using the following formula: ， in, The correction matrix is ​​calculated as follows: ， in, , ,in The Julian century number is as of 12:00 on January 1, 2000, in centuries; For the true ecliptic longitude of the sun; and This represents the inclination of the sun's orbit.

7. The planning method as described in claim 6, characterized in that, Calculating the quaternion of the observed target based on the attitude matrix includes: The attitude matrix is ​​represented in square matrix form: ； Calculate and compare the following four values: , , ,as well as ; Determine the quaternion based on the maximum value among the four values. : 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 Normalize the quaternion if If the value is less than 0, then a negative value is taken.

8. The planning method as described in claim 5, characterized in that, Based on the vectors in the aforementioned solar vector system, calculate the angles between the satellite's +X axis, +Y axis, -Y axis, and -Z axis and the line connecting the Sun and Earth; and / or The angle between the satellite's +X axis and the line connecting the moon and satellite is determined based on the lunar satellite vector within this system.

9. The planning method as described in claim 1, characterized in that, The constraints are calculated using a spaceborne system, and the planning method further includes: Before satellite launch, the constraints are calculated using both the ground simulation system and the onboard system based on test data, and then compared. If the two are consistent, the verification is successful; otherwise, the onboard system is updated and optimized.

10. 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 the satellite, the location information of the observed target is transmitted; and The ground simulation system calculates the constraints based on the location information and compares them with preset conditions. If the conditions are met, the task is executed; otherwise, the observation task is canceled.

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