Multi-information fusion attitude determination method and device for high dynamic load

Abstract

The present invention discloses a multi-information fusion attitude determination method and device for high dynamic payloads, belonging to the field of satellite attitude control. The method includes: when the satellite is in a star-sensitive invisible sky area, the attitude extrapolation of the gyro is used to perform attitude determination of the payload platform; when the satellite enters a star-sensitive visible sky area, the attitude quaternion of the gyro is reset by using the star-sensitive measurement attitude, and the attitude extrapolation of the gyro is performed based on the reset attitude quaternion, so as to perform fusion attitude determination by using the attitude extrapolation of the gyro in combination with the star-sensitive correction, and the first double star-sensitive group is used for fusion attitude determination in the first half of the arc section of the visible sky area, and the second double star-sensitive group is used for fusion attitude determination in the second half of the arc section of the visible sky area; the two star sensors in each double star-sensitive group are installed as an integrated whole, and the projections of the two star sensors in each double star-sensitive group on the YOZ plane overlap. The present invention can ensure the star-sensitive measurement accuracy of high dynamic payloads and meet the attitude determination accuracy requirements under the high-speed motion characteristics of the on-orbit payload platform of the optical satellite.

Description

Technical Field

[0001] The present invention relates to the technical field of satellite attitude control, and in particular to a multi-information fusion attitude determination method and device for high dynamic load. Background Art

[0002] Optical satellites are composed of two platforms: a high-speed rotating payload platform and a satellite platform. The three-axis attitude determination of the payload platform directly depends on the accuracy of the high dynamic star sensor of the payload platform and the gyro joint attitude determination of the payload platform to provide real-time attitude data for the payload platform.

[0003] The payload platform of traditional satellites does not rotate, while the new optical satellite payload platform rotates at high speed. When the payload platform rotates at high speed around the roll axis, the gyro measurement accuracy will drop significantly. Therefore, when the satellite is in orbit within the star-sensitive visible sky area, it needs to rely on the star-sensitive measurement accuracy of the payload platform. However, due to the motion characteristics of the high-speed rotation of the payload platform, only the star-sensitive measurement sensors within a certain rotation angle range are available during the rotation process. Therefore, it is necessary to coordinate the configuration and installation of star sensors for the high-speed rotating payload of this optical satellite. Since the payload platform of traditional satellites does not rotate and the gyro measurement accuracy of the payload platform is relatively high, the fusion attitude determination method of the payload platform of traditional satellites is not suitable for optical satellites with high-speed rotating payloads at all, and cannot meet the attitude determination accuracy requirements of the optical satellite's on-orbit payload platform under the high-speed motion characteristics.

[0004] Therefore, there is an urgent need to provide a multi-information fusion posture determination method and device for high dynamic loads. Summary of the invention

[0005] In order to solve the problem that there is currently no method for determining the attitude of an optical satellite under high-speed rotation of an on-orbit payload platform, an embodiment of the present invention provides a multi-information fusion attitude determination method and device for a high-dynamic payload.

[0006] On the one hand, a multi-information fusion attitude determination method for a high dynamic load is provided, the method comprising:

[0007] When the satellite is in the sky area where the star sensor is not visible, the gyro attitude extrapolation is used to determine the attitude of the payload platform;

[0008] When the satellite enters the star-sensing visible sky area, the attitude quaternion of the gyro is reset by using the star-sensing attitude measurement, and the gyro attitude is extrapolated based on the reset attitude quaternion, so as to use the gyro attitude extrapolation and star-sensing correction to perform fusion attitude determination, and the first double star-sensing group is used for fusion attitude determination in the front half arc section of the visible sky area, and the second double star-sensing group is used for fusion attitude determination in the back half arc section of the visible sky area; wherein, the two star sensors in each double star-sensing group are installed as an integrated unit, and the projections of the two star sensors in each double star-sensing group on the YOZ plane coincide.

[0009] On the other hand, a multi-information fusion posture determination device for high dynamic load is provided, which is used to implement the steps described in any method embodiment of the specification, and the device includes:

[0010] The first attitude determination unit is used to determine the attitude of the payload platform by using the attitude extrapolation of the gyroscope when the satellite is in the sky area where the star sensor is not visible;

[0011] The second attitude determination unit is used for resetting the attitude quaternion of the gyroscope by using the attitude measured by the star sensor when the satellite enters the visible sky area of ​​the star sensor, and performing gyroscope attitude extrapolation based on the reset attitude quaternion, so as to perform fusion attitude determination by using the attitude extrapolation of the gyroscope in combination with the star sensor correction, and using the first double star sensor group for fusion attitude determination in the front half arc section of the visible sky area, and using the second double star sensor group for fusion attitude determination in the back half arc section of the visible sky area; wherein the two star sensors in each double star sensor group are installed as an integrated unit, and the projections of the two star sensors in each double star sensor group on the YOZ plane overlap.

[0012] On the other hand, a computer device is provided, which includes a memory and a processor, wherein the memory is used to store a computer program, and the processor is used to execute the computer program stored in the memory to implement the steps of the above-mentioned method.

[0013] On the other hand, a computer-readable storage medium is provided, wherein a computer program is stored in the storage medium, and when the computer program is executed by a processor, the steps of the above-mentioned method are implemented.

[0014] On the other hand, a computer program product is provided, comprising a computer program, wherein the computer program implements the steps of the above method when executed by a processor.

[0015] The technical solution provided by the present invention can at least bring the following beneficial effects:

[0016] Two groups of double star sensors are set up to perform star sensor filtering on the front and back arc segments of the visible sky area, respectively, so as to cooperate with the gyro attitude extrapolation to complete the joint attitude determination of the high dynamic payload. In order to ensure that the optical axis attitude determination of the two star sensors is used in the front and back arc segments of the payload imaging, it is necessary to ensure that the two star sensors in each double star sensor group are effective at the same time in the front and back arc segments. Therefore, the installation positions of the two star sensors in each double star sensor group are set to meet the projection overlap in the YOZ plane, so as to ensure the star sensor measurement accuracy of the high dynamic payload and meet the attitude determination accuracy requirements under the high-speed motion characteristics of the optical satellite on-orbit payload platform. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0018] Figure 1 This is a flow chart of a multi-information fusion posture determination method for a high dynamic load provided by an embodiment of the present invention;

[0019] Figure 2 This is a structural diagram of a multi-information fusion posture determination device for a high dynamic load provided by an embodiment of the present invention;

[0020] Figure 3 It is a hardware architecture diagram of a computer device provided by an embodiment of the present invention. DETAILED DESCRIPTION

[0021] In order to make the purpose, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, rather than all the embodiments. Based on the embodiments in the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0022] The specific implementation of the above concept is described below.

[0023] Please refer to Figure 1 , an embodiment of the present invention provides a multi-information fusion posture determination method for a high dynamic load, the method comprising:

[0024] Step 100: The satellite is in a sky area that is not visible to the star sensor, and the attitude of the payload platform is determined by using the gyro attitude extrapolation;

[0025] Step 102: When the satellite enters the visible sky area of ​​the star sensor, the attitude quaternion of the gyro is reset by using the star sensor to measure the attitude, and the gyro attitude is extrapolated based on the reset attitude quaternion, so as to use the gyro attitude extrapolation combined with the star sensor correction to perform fusion attitude determination, and the first double star sensor group is used for fusion attitude determination in the first half of the visible sky area, and the second double star sensor group is used for fusion attitude determination in the second half of the visible sky area; wherein, the two star sensors in each double star sensor group are installed as an integrated unit, and the projections of the two star sensors in each double star sensor group on the YOZ plane overlap.

[0026] In the embodiment of the present invention, two groups of dual star-sensor groups are set to perform star-sensor filtering on the front and rear arc segments of the visible sky area of ​​the star-sensor, respectively, so as to cooperate with the gyro attitude extrapolation to complete the joint attitude determination of the high dynamic payload. In order to ensure that the optical axis attitude determination of the two star-sensors is used for both the front and rear arc segments of the payload imaging, it is necessary to ensure that the two star-sensors in each dual star-sensor group are effective at the same time during the front and rear arc segments. Therefore, the installation positions of the two star-sensors in each dual star-sensor group are set to meet the projection overlap in the YOZ plane, so as to ensure the star-sensor measurement accuracy of the high dynamic payload and meet the attitude determination accuracy requirements under the high-speed motion characteristics of the optical satellite on-orbit payload platform. In addition, the quaternion is reset when entering the visible sky area from the invisible sky area of ​​the star-sensor, which can ensure that the gyro constant drift estimation does not introduce the gyro extrapolation model deviation in the non-visible area, thereby ensuring the attitude determination accuracy.

[0027] Described below Figure 1 How the various steps are performed.

[0028] For step 100:

[0029] In this step, as the payload platform rotates around the rotation axis, the star-sensing combination enters the visible sky area and the invisible sky area in turn. The star-sensing combination uses the gyro information to perform attitude extrapolation in the invisible sky area, and only uses the gyro to perform satellite attitude determination. However, when the payload platform rotates at high speed, the gyro measurement accuracy will be significantly reduced. In order to improve the payload attitude determination accuracy in the invisible sky area, the gyro error calibration can be performed in the first several cycles of the satellite's on-orbit motion to correct the error of the gyro attitude extrapolation result.

[0030] Specifically, the gyro error calibration is performed during the first several cycles of the satellite's on-orbit motion, which may include:

[0031] The load is pronated and the attitude maneuver sequence is completed to calibrate the installation error, initial constant deviation and scale factor error of the non-coplanar mounted gyro assembly;

[0032] After the payload is spun up, the noise characteristics of the gyro error model are amplified five times to determine the filtering parameters of the first and second double star-sensing groups. The gyro constant deviation is calibrated in real time on orbit, and the gyro attitude extrapolation result is corrected for installation error, constant deviation and scale factor error.

[0033] In this embodiment, in order to improve the filter convergence speed and the estimation accuracy of the constant deviation, the noise characteristics of the gyro error model can be amplified by 5 times to design the filter parameters of the first double star sensitive group and the second double star sensitive group, so that the double star sensitive group can be used to perform on-orbit real-time calibration of the gyro constant deviation to improve the accuracy of on-orbit calibration of the constant deviation. By correcting the installation error, constant deviation and scale factor error of the gyro attitude extrapolation result, the attitude determination accuracy of the payload platform in the invisible sky area and the visible sky area can be improved.

[0034] Regarding step 102:

[0035] In the traditional payload attitude determination method, since the payload platform does not rotate, the payload platform only needs to be installed with a set of star sensors, and the camera imaging task is performed by adjusting the satellite platform attitude. For the optical satellite of the present invention, the payload platform rotates at high speed around the roll axis during on-orbit operation, and only the star-sensing measurement sensors within a certain rotation angle range are available during the rotation process. The camera imaging task needs to be performed in the visible sky area of ​​the star-sensing. In each on-orbit motion cycle of the satellite, the camera imaging area occupies about 1 / 3 of the area of ​​the entire sky area, and the effective sky area of ​​a set of dual star-sensing groups occupies about 1 / 6 of the area of ​​the entire sky area. Therefore, at least two sets of dual star-sensing groups need to be set up to complete the first half arc segment filtering and the second half arc segment filtering of the camera imaging area (i.e., the visible sky area) respectively.

[0036] When the payload platform rotates at high speed around the rotation axis, the projections of the two star sensors in the double star sensor group overlap in the YOZ plane, which can ensure the maximum effective sky area of ​​the double star sensor. Therefore, the AB star sensors in the first double star sensor group are installed as an integrated unit, and the CD star sensor in the second double star sensor group is installed as an integrated unit, and the projections of the AB star sensors need to overlap in the YOZ plane, and the projections of the CD star sensors need to overlap in the YOZ plane.

[0037] In some embodiments, when the satellite is operating in a sun-synchronous orbit and the camera is installed on the payload platform body axis - Y axis, and the payload platform rotates at a high speed around the body rotation axis X axis, the projection of the two star sensors in the first double star sensitive group on the YOZ plane has an angle of 75 degrees with the Z axis and an angle of 15 degrees with the Y axis, and the projection of the two star sensors in the second double star sensitive group on the YOZ plane has an angle of 25 degrees with the Z axis and an angle of 65 degrees with the Y axis.

[0038] In the visible sky area of ​​star sensors, when both the first double star sensor group and the second double star sensor group are in good condition, the attitude extrapolation of the first double star sensor group + gyroscope is used for fusion attitude determination in the first half of the arc, and the attitude extrapolation of the second double star sensor group + gyroscope is used for fusion attitude determination in the second half of the arc. However, the first double star sensor group and the second double star sensor group may fail, so it is necessary to design a fusion attitude determination method under star sensor failure to ensure the accuracy of payload attitude determination.

[0039] In some embodiments, when the satellite is within the visible sky area of ​​the star sensor, the payload platform performs fusion attitude determination in the following manner:

[0040] For each on-orbit motion cycle, the following are executed:

[0041] Determine whether the first double star sensor group and the second double star sensor group are faulty;

[0042] When only a single star in the first double star-sensor group fails, the attitude extrapolation of the other single star-sensor in the first double star-sensor group that is not faulty is used to perform fusion attitude determination in the first half of the arc of the visible sky area, and the attitude extrapolation of the second double star-sensor group is used to perform fusion attitude determination normally in the second half of the arc;

[0043] When only the double star in the first double star group fails, the second double star group filtering result of the second half arc of the previous motion cycle and the gyro attitude extrapolation of 2 / 3 of the sky area are used to perform attitude determination in the first half arc of the current motion cycle, and the second double star group filtering and gyro attitude extrapolation are used normally in the second half arc for fusion attitude determination;

[0044] When only the second double star sensor group has a fault, the first double star sensor group filtering and the gyro's attitude extrapolation are used normally for fusion attitude determination in the first half of the visible sky segment, and the first double star sensor group filtering results in the first half of the current motion cycle and the gyro's attitude extrapolation are used for attitude determination in the second half of the arc; among them, the faults of the second double star sensor group include single-star faults and double-star faults.

[0045] In this embodiment, in each movement cycle, the first double star-sensitivity group is used in the first half arc segment of the visible sky area, and the second double star-sensitivity group is used in the second half arc segment. If a single star failure occurs in the first double star-sensitivity group in the first half arc segment, the filtering result of the second double star-sensitivity group in the second half arc segment of the previous movement cycle is also used to make the gyro attitude extrapolate 2 / 3 of the sky area. This 2 / 3 of the sky area has no star-sensitivity correction, which will inevitably introduce a large gyro cumulative attitude determination error. The attitude determination accuracy is not as high as the accuracy of the fusion attitude determination using another single star-sensitivity filtering without failure in the first double star-sensitivity group + gyro attitude extrapolation. Therefore, the first double star-sensitivity group is divided into two cases: single star-sensitivity failure and double star-sensitivity failure.

[0046] When a single star of the second double star-sensitivity group used in the second half of the arc fails, the attitude determination accuracy of the second half of the arc is directly performed by using the filtering result of the first double star-sensitivity group in the first half of the current motion cycle + the attitude extrapolation of the gyro, which is higher than the attitude extrapolation accuracy of another single star-sensitivity group that is not faulty in the second double star-sensitivity group + the gyro. Therefore, regardless of whether the second double star-sensitivity group has a single star-sensitivity failure or a double star-sensitivity failure, the attitude determination of the second half of the arc is performed by using the filtering result of the first double star-sensitivity group in the first half of the current motion cycle + the attitude extrapolation of the gyro. In summary, the fusion attitude determination method of this embodiment can improve the load attitude determination accuracy under star-sensitivity failure.

[0047] In order to further improve the accuracy of star sensor filtering, the embodiment of the present invention further provides a method for calibrating the installation deviations of the first double star sensor group and the second double star sensor group.

[0048] Specifically, the star-sensing installation deviation calibration method for the first several cycles of the satellite's on-orbit motion includes:

[0049] In the first several cycles, in the overlapping visible area of ​​the first double star-sensitivity group and the second double star-sensitivity group, the first double star-sensitivity group and the second double star-sensitivity group are calibrated on-orbit to obtain the installation deviation of the first double star-sensitivity group and the second double star-sensitivity group;

[0050] Carry out on-orbit calibration of the installation deviation of the payload platform star sensor and the satellite platform star sensor relative to the rotation axis;

[0051] The installation deviation of the star sensor of the load platform obtained by mutual calibration and the installation deviation of the star sensor relative to the rotation axis are used to correct the installation deviation of the filtering result of the corresponding star sensor.

[0052] In this embodiment, the first double star sensor group and the second double star sensor group can be calibrated on-orbit in the overlapping visible area. The overlapping visible area of ​​the two star sensor groups is generally greater than 10 degrees. In addition, the installation deviation relative to the rotation axis can be calibrated. The installation deviation can be jointly determined by two methods to improve the filtering accuracy of the star sensor.

[0053] It should be noted that the payload platform only completes the attitude determination of the payload platform star sensor. The result of attitude determination does not introduce the control closed loop, and only has observation capability. Under the mechanical bearing connection condition, the satellite platform momentum wheel controls the pointing of the entire satellite. The result of the satellite platform star sensor attitude determination will introduce the control closed loop, which has the ability to affect the motion characteristics. Therefore, the installation deviation of the satellite platform star sensor relative to the rotation axis is calibrated. After correcting the installation deviation of the satellite platform star sensor, the installation deviation of the payload platform star sensor relative to the rotation axis is calibrated again. This can make the installation deviation of the payload platform star sensor relative to the rotation axis more accurate, so as to further improve the filtering accuracy of the star sensor.

[0054] In some embodiments, the on-orbit calibration of the relative rotation axis installation deviation of the payload platform star-sensing installation and the satellite platform star-sensing installation includes:

[0055] The constant terms of the attitude of the payload platform relative to the dynamic coordinate system and the constant terms of the attitude of the payload platform relative to the static coordinate system are obtained from the satellite, so as to use these two constant terms to calibrate the installation deviation of the satellite platform and the star-sensor relative rotation axis.

[0056] In this embodiment, two sets of variable telemetry information are configured on the satellite, which are used to calculate two attitude determination results of the star sensor, namely, the attitude of the payload platform relative to the dynamic coordinate system and the attitude of the payload platform relative to the static coordinate system, wherein the attitude of the payload platform relative to the dynamic coordinate system includes a constant term (payload platform star sensor installation deviation) and a sine term (satellite platform star sensor installation deviation); the attitude of the payload platform relative to the static coordinate system includes a constant term (satellite platform star sensor installation deviation) + a sine term (payload platform star sensor installation deviation). The constant term in the attitude of the payload platform relative to the dynamic coordinate system and the constant term in the attitude of the payload platform relative to the static coordinate system can be taken to calibrate the installation deviation of the satellite platform and the payload platform star sensor relative to the rotation axis to correct the installation deviation.

[0057] Please refer to Figure 2 The embodiment of the present invention provides a multi-information fusion posture determination device for high dynamic load, which is used to implement the steps of any method embodiment in the specification, and the device includes:

[0058] The first attitude determination unit 201 is used for determining the attitude of the payload platform by using the attitude extrapolation of the gyroscope when the satellite is in the sky area where the star sensor is not visible;

[0059] The second attitude determination unit 202 is used for resetting the attitude quaternion of the gyro by using the attitude measured by the star sensor when the satellite enters the visible sky area of ​​the star sensor, and performing gyro attitude extrapolation based on the reset attitude quaternion, so as to perform fusion attitude determination by using the attitude extrapolation of the gyro in combination with the star sensor correction, and using the first double star sensor group for fusion attitude determination in the front half of the visible sky area, and using the second double star sensor group for fusion attitude determination in the back half of the visible sky area; wherein the two star sensors in each double star sensor group are installed as an integrated unit, and the projections of the two star sensors in each double star sensor group on the YOZ plane overlap.

[0060] In one embodiment of the present invention, when the satellite is operating in a sun-synchronous orbit and the camera is installed on the Y-axis of the payload platform body, and the payload platform rotates at a high speed around the X-axis of the body rotation, the projection of the two star sensors in the first double star sensitive group in the second attitude determination unit 202 on the YOZ plane has an angle of 75 degrees with the Z axis and an angle of 15 degrees with the Y axis, and the projection of the two star sensors in the second double star sensitive group on the YOZ plane has an angle of 25 degrees with the Z axis and an angle of 65 degrees with the Y axis.

[0061] In one embodiment of the present invention, the satellite in the second attitude determination unit 202 is within the visible sky area of ​​the star sensor, and the payload platform performs fusion attitude determination in the following manner:

[0062] For each on-orbit motion cycle, the following are executed:

[0063] Determine whether the first double star sensor group and the second double star sensor group are faulty;

[0064] When only a single star in the first double star-sensor group fails, the attitude extrapolation of the other single star-sensor in the first double star-sensor group that is not faulty is used to perform fusion attitude determination in the first half of the arc of the visible sky area, and the attitude extrapolation of the second double star-sensor group is used to perform fusion attitude determination normally in the second half of the arc;

[0065] When only the double star in the first double star group fails, the second double star group filtering result of the second half arc of the previous motion cycle and the gyro attitude extrapolation of 2 / 3 of the sky area are used to perform attitude determination in the first half arc of the current motion cycle, and the second double star group filtering and gyro attitude extrapolation are used normally in the second half arc for fusion attitude determination;

[0066] When only the second double star sensor group has a fault, the first double star sensor group filtering and the gyro's attitude extrapolation are used normally for fusion attitude determination in the first half of the visible sky segment, and the first double star sensor group filtering results in the first half of the current motion cycle and the gyro's attitude extrapolation are used for attitude determination in the second half of the arc; among them, the faults of the second double star sensor group include single-star faults and double-star faults.

[0067] In one embodiment of the present invention, a calibration unit 203 is further included. During the first several cycles of the satellite's on-orbit motion, the calibration unit 203 is used to:

[0068] The load is pronated and the attitude maneuver sequence is completed to calibrate the installation error, initial constant deviation and scale factor error of the non-coplanar mounted gyro assembly;

[0069] After the payload is spun up, the noise characteristics of the gyro error model are amplified five times to determine the filtering parameters of the first and second double star-sensing groups. The gyro constant deviation is calibrated in real time on orbit, and the gyro attitude extrapolation result is corrected for installation error, constant deviation and scale factor error.

[0070] In one embodiment of the present invention, the calibration unit 203 is further used for:

[0071] In the first several cycles, in the overlapping visible area of ​​the first double star-sensitivity group and the second double star-sensitivity group, the first double star-sensitivity group and the second double star-sensitivity group are calibrated on-orbit to obtain the installation deviation of the first double star-sensitivity group and the second double star-sensitivity group;

[0072] Carry out on-orbit calibration of the installation deviation of the payload platform star sensor and the satellite platform star sensor relative to the rotation axis;

[0073] The installation deviation of the star sensor of the load platform obtained by mutual calibration and the installation deviation of the star sensor relative to the rotation axis are used to correct the installation deviation of the filtering result of the corresponding star sensor.

[0074] In one embodiment of the present invention, when performing on-orbit calibration of the relative rotation axis installation deviation of the payload platform star-sensing installation and the satellite platform star-sensing installation, the calibration unit 203 is used to:

[0075] The constant terms of the attitude of the payload platform relative to the dynamic coordinate system and the constant terms of the attitude of the payload platform relative to the static coordinate system are obtained from the satellite, so as to use these two constant terms to calibrate the installation deviation of the satellite platform and the star-sensor relative rotation axis.

[0076] It should be noted that the multi-information fusion posture determination device for high dynamic loads provided in the above embodiment is only illustrated by the division of the above functional units. In practical applications, the above functions can be assigned to different functional units as needed, that is, the internal structure of the device can be divided into different functional units to complete all or part of the functions described above. In addition, the above device embodiment and the method embodiment belong to the same concept, and the specific implementation process is detailed in the method embodiment, which will not be repeated here.

[0077] The embodiment of the present application also provides a computer device, please refer to Figure 3 The computer device includes a processor and a memory, wherein the memory stores at least one instruction, at least one program, code set or instruction set, and the at least one instruction, at least one program, code set or instruction set is loaded and executed by the processor to implement the multi-information fusion posture determination method for high dynamic load provided by the above-mentioned method embodiments.

[0078] An embodiment of the present application also provides a computer-readable storage medium, on which is stored at least one instruction, at least one program, code set or instruction set, and the at least one instruction, at least one program, code set or instruction set is loaded and executed by a processor to implement the multi-information fusion posture determination method for high dynamic loads provided by the above-mentioned method embodiments.

[0079] An embodiment of the present application also provides a computer program product, which includes a computer program. A processor of a computer device reads the computer program from a computer-readable storage medium, and the processor executes the computer program, so that the computer device executes any of the high dynamic load multi-information fusion posture determination methods in the above-mentioned embodiments.

[0080] For the convenience of description, the above system or device is described by dividing it into various modules or units according to its functions. Of course, when implementing the present application, the functions of each unit can be implemented in the same or multiple software and / or hardware.

[0081] It can be known from the description of the above implementation methods that those skilled in the art can clearly understand that the present application can be implemented by means of software plus a necessary general hardware platform. Based on such an understanding, the technical solution of the present application can be essentially or partly embodied in the form of a software product that contributes to the prior art. The computer software product can be stored in a storage medium such as ROM / RAM, a magnetic disk, an optical disk, etc., and includes several instructions for enabling a computer device (which can be a personal computer, a server, or a network device, etc.) to execute the methods of each embodiment of the present application or some parts of the embodiments.

[0082] Finally, it should be noted that, in this article, relational terms such as first, second, third and fourth are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device. In the absence of further restrictions, the elements defined by the sentence "comprise a ..." do not exclude the presence of other identical elements in the process, method, article or device including the elements.

[0083] The above are only preferred implementations of the present application. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the principles of the present application. These improvements and modifications should also be regarded as the scope of protection of the present application.

Claims

1. A multi-information fusion attitude determination method for high dynamic loads, characterized in that: The method comprises: When the satellite is in the sky area where the star sensor is not visible, the gyro attitude extrapolation is used to determine the attitude of the payload platform; When the satellite enters the star-sensing visible sky area, the attitude quaternion of the gyro is reset by using the star-sensing attitude measurement, and the gyro attitude is extrapolated based on the reset attitude quaternion, so as to use the gyro attitude extrapolation combined with the star-sensing correction to perform fusion attitude determination, and the first double star-sensing group is used for fusion attitude determination in the first half of the arc section of the visible sky area, and the second double star-sensing group is used for fusion attitude determination in the second half of the arc section of the visible sky area; wherein, the two star sensors in each double star-sensing group are installed as an integrated whole, and the projections of the two star sensors in each double star-sensing group on the YOZ plane coincide; When the satellite is in the visible sky area, the payload platform performs fusion attitude determination in the following ways: For each on-orbit motion cycle, the following are executed: Determining whether the first dual star-sensing group and the second dual star-sensing group are faulty; When only a single star in the first double star-sensitivity group fails, another single star-sensitivity in the first double star-sensitivity group that is not at fault is used in the first half of the arc segment of the visible sky region to perform fusion attitude determination by filtering and gyro attitude extrapolation, and the second double star-sensitivity group filtering and gyro attitude extrapolation are normally used in the second half of the arc segment to perform fusion attitude determination; When only the double star in the first double star sensitive group fails, the filtering result of the second double star sensitive group in the second half of the previous motion cycle and the gyro attitude extrapolation of 2 / 3 of the sky area are used to perform attitude determination in the first half of the current motion cycle, and the second double star sensitive group filtering and gyro attitude extrapolation are used normally in the second half of the arc for fusion attitude determination; When only the second double star sensor group has a fault, the first double star sensor group filtering and the gyro attitude extrapolation are used normally in the first half of the visible sky area for fusion attitude determination, and the first double star sensor group filtering result in the first half of the current motion cycle and the gyro attitude extrapolation are used for attitude determination in the second half; wherein, the faults in the second double star sensor group include single star faults and double star faults.

2. The method according to claim 1, characterized in that When the satellite is operating in a sun-synchronous orbit, and the camera is installed on the Y-axis of the payload platform body, and the payload platform rotates at a high speed around the X-axis of the body rotation, the projections of the two star sensors in the first double star sensitive group on the YOZ plane have an angle of 75 degrees with the Z-axis and an angle of 15 degrees with the Y-axis, and the projections of the two star sensors in the second double star sensitive group on the YOZ plane have an angle of 25 degrees with the Z-axis and an angle of 65 degrees with the Y-axis.

3. The method according to claim 1, characterized in that The first several cycles of the satellite's orbital motion also include: The load is pronated and the attitude maneuver sequence is completed to calibrate the installation error, initial constant deviation and scale factor error of the non-coplanar mounted gyro assembly; After the payload starts to spin, the noise characteristics of the gyro error model are amplified 5 times to determine the filtering parameters of the first double star sensor group and the second double star sensor group, the gyro constant deviation is calibrated in real time on orbit, and the installation error, the constant deviation and the scale factor error are corrected for the gyro attitude extrapolation result.

4. The method according to claim 1, characterized in that The first several cycles of the satellite's orbital motion also include: In the first several cycles, in the overlapping visible area of ​​the first double star-sensitive group and the second double star-sensitive group, the first double star-sensitive group and the second double star-sensitive group are calibrated on-orbit to obtain the installation deviation of the first double star-sensitive group and the second double star-sensitive group; Carry out on-orbit calibration of the installation deviation of the payload platform star sensor and the satellite platform star sensor relative to the rotation axis; The installation deviation of the star sensor of the load platform obtained by mutual calibration and the installation deviation of the star sensor relative to the rotation axis are used to correct the installation deviation of the filtering result of the corresponding star sensor.

5. The method according to claim 4, characterized in that The on-orbit calibration of the relative rotation axis installation deviation of the payload platform star-sensing installation and the satellite platform star-sensing installation includes: The constant terms of the attitude of the payload platform relative to the dynamic coordinate system and the constant terms of the attitude of the payload platform relative to the static coordinate system are obtained from the satellite, so as to use these two constant terms to calibrate the installation deviation of the satellite platform and the star-sensor relative rotation axis.

6. A multi-information fusion posture determination device for high dynamic load, used to implement the steps of any of the methods described in claims 1-5 above, characterized in that: The device comprises: The first attitude determination unit is used to determine the attitude of the payload platform by using the attitude extrapolation of the gyroscope when the satellite is in the sky area where the star sensor is not visible; The second attitude determination unit is used for resetting the attitude quaternion of the gyroscope by using the attitude measured by the star sensor when the satellite enters the visible sky area of ​​the star sensor, and performing gyroscope attitude extrapolation based on the reset attitude quaternion, so as to perform fusion attitude determination by using the attitude extrapolation of the gyroscope in combination with the star sensor correction, and using the first double star sensor group for fusion attitude determination in the front half arc section of the visible sky area, and using the second double star sensor group for fusion attitude determination in the back half arc section of the visible sky area; wherein the two star sensors in each double star sensor group are installed as an integrated unit, and the projections of the two star sensors in each double star sensor group on the YOZ plane overlap.

7. A computer device, characterized in that: The computer device includes a memory and a processor, the memory is used to store a computer program, and the processor is used to execute the computer program stored in the memory to implement the steps of any of the methods described in claims 1-5.

8. A computer-readable storage medium, characterized in that: The storage medium stores a computer program, and when the computer program is executed by a processor, the steps of the method described in any one of claims 1 to 5 are implemented.

9. A computer program product, characterized in that The method comprises a computer program, which, when executed by a processor, implements the steps of the method according to any one of claims 1 to 5.

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