CN115352656A - Satellite attitude control method, system and equipment for replacing fault flywheel by magnetic torquer - Google Patents
Satellite attitude control method, system and equipment for replacing fault flywheel by magnetic torquer Download PDFInfo
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Abstract
The invention relates to the technical field of satellite attitude control, in particular to a satellite attitude control method, a satellite attitude control system and satellite attitude control equipment for replacing a fault flywheel by a magnetic torquer. The method comprises the following steps: calculating the target output torque of each non-fault flywheel based on the installation matrixes corresponding to all the flywheels of the satellite; calculating the target output magnetic moment of each magnetic torquer of the satellite according to the expected moment of the fault flywheel; and controlling each non-fault flywheel to output corresponding target output torque, and controlling each magnetic torquer to output corresponding target output magnetic moment so as to control the attitude of the satellite. When any flywheel on the satellite breaks down, the magnetic torquer arranged in the existing three-orthogonal mode on the satellite is adopted to complete the control function of the broken flywheel, so that the configuration of a plurality of redundant flywheels is avoided, the development cost of the satellite is greatly reduced, and the weight of the satellite is reduced.
Description
Technical Field
The invention relates to the technical field of satellite attitude control, in particular to a satellite attitude control method, a satellite attitude control system and satellite attitude control equipment for replacing a fault flywheel by a magnetic torquer.
Background
For satellites such as minisatellites and micro-nano satellites, flywheels are important executive components in attitude control systems of the satellites. In order to ensure the reliability and meet the long service life requirement, most of the existing attitude control systems adopt a redundant backup mode, namely four or more flywheels are adopted, when a certain flywheel breaks down, the redundant flywheel replaces the flywheel, so that the control effect similar to that of the original system is realized, and in order to ensure the ideal flywheel configuration and control effect, a plurality of backup flywheels in different installation directions are required to realize the backup function of any shaft so as to meet the requirement of attitude control.
However, the traditional method of backing up a failed flywheel by adopting a mode of configuring more flywheels increases the weight of the whole satellite control system of the satellite and also increases the development cost of the whole satellite.
Disclosure of Invention
The invention aims to solve the technical problem of the prior art and provides a satellite attitude control method, a satellite attitude control system and satellite attitude control equipment for replacing a fault flywheel by a magnetic torquer.
The invention relates to a satellite attitude control method for replacing a fault flywheel by a magnetic torquer, which adopts the technical scheme as follows:
calculating the target output torque of each non-fault flywheel based on the installation matrixes corresponding to all the flywheels of the satellite;
calculating the target output magnetic moment of each magnetic torquer of the satellite according to the expected moment of the fault flywheel, wherein one magnetic torquer is arranged on each of three coordinate axes in a body coordinate system of the satellite;
and controlling each non-fault flywheel to output a corresponding target output torque, and controlling each magnetic torquer to output a corresponding target output magnetic moment so as to control the attitude of the satellite.
The satellite attitude control method using the magnetic torquer to replace the fault flywheel has the following beneficial effects:
when any flywheel on the satellite breaks down, the magnetic torquer arranged in the existing three-orthogonal mode on the satellite is adopted to complete the control function of the broken flywheel, so that the configuration of a plurality of redundant flywheels is avoided, the development cost of the satellite is greatly reduced, and the weight of the satellite is reduced.
On the basis of the scheme, the satellite attitude control method for replacing the fault flywheel by the magnetic torquer can be further improved as follows.
Further, the process of calculating the target output torque for each non-faulty flywheel includes:
deleting elements related to the fault flywheel from the installation matrix to obtain a corrected installation matrix;
and calculating the target output torque of each non-fault flywheel according to the corrected mounting matrix and the expected torque of each non-fault flywheel.
Further, the method also comprises the following steps:
and obtaining the expected torque of the fault flywheel according to the installation matrix and the expected torque of each non-fault flywheel.
Further, the calculating the target output magnetic moment of each magnetic torquer of the satellite according to the expected moment of the fault flywheel comprises:
calculating the target output magnetic moment of each magnetic torquer corresponding to the fault flywheel according to a first formula, wherein the first formula is as follows:
wherein,
a measurement value of a geomagnetic intensity vector representing an orbital position of a satellite in a body coordinate system of the satellite,
represents the expected torque of the faulty flywheel,
and MT comprises 3 elements and corresponds to the target output magnetic moment of the magnetic torquer on each coordinate axis in the body coordinate system of the satellite.
The invention relates to a technical scheme of a satellite attitude control system using a magnetic torquer to replace a fault flywheel, which comprises the following steps:
the first computing module is to: calculating the target output torque of each non-fault flywheel based on the installation matrixes corresponding to all the flywheels of the satellite;
the second calculation module is configured to: calculating the target output magnetic moment of each magnetic torquer of the satellite according to the expected moment of the fault flywheel, wherein one magnetic torquer is arranged on each of three coordinate axes in a body coordinate system of the satellite;
the control module is used for: and controlling each non-fault flywheel to output a corresponding target output torque, and controlling each magnetic torquer to output a corresponding target output magnetic moment so as to control the attitude of the satellite.
The satellite attitude control system using the magnetic torquer to replace the fault flywheel has the following beneficial effects:
when any flywheel on the satellite breaks down, the magnetic torquer arranged in the existing three-orthogonal mode on the satellite is adopted to complete the control function of the broken flywheel, so that the configuration of a plurality of redundant flywheels is avoided, the development cost of the satellite is greatly reduced, and the weight of the satellite is reduced.
On the basis of the scheme, the satellite attitude control system using the magnetic torquer to replace the fault flywheel can be further improved as follows.
Further, the first computing module is specifically configured to:
deleting elements associated with the fault flywheel from the installation matrix to obtain a corrected installation matrix;
and calculating the target output torque of each non-fault flywheel according to the corrected mounting matrix and the expected torque of each non-fault flywheel.
Further, a third computing module is included, the third computing module being configured to:
and obtaining the expected torque of the fault flywheel according to the installation matrix and the expected torque of each non-fault flywheel.
Further, the second calculation module is specifically configured to:
calculating the target output magnetic moment of each magnetic torquer corresponding to the fault flywheel according to a first formula, wherein the first formula is as follows:
wherein,
a measurement value of a geomagnetic intensity vector representing an orbital position of a satellite in a body coordinate system of the satellite,
represents the expected torque of the faulty flywheel,
and MT comprises 3 elements and corresponds to the target output magnetic moment of the magnetic torquer on each coordinate axis in the body coordinate system of the satellite.
The storage medium of the present invention stores instructions, and when the instructions are read by a computer, the computer is caused to execute any one of the above-described satellite attitude control methods using a magnetic moment device to replace a faulty flywheel.
An apparatus of the present invention includes a processor and the storage medium described above, the processor executing instructions in the storage medium.
Drawings
FIG. 1 is a schematic flowchart of a satellite attitude control method using a magnetic torquer to replace a failed flywheel according to an embodiment of the present invention;
fig. 2 is a schematic structural diagram of a satellite attitude control system using a magnetic torquer to replace a failed flywheel according to an embodiment of the present invention.
Detailed Description
As shown in fig. 1, a satellite attitude control method using a magnetic torquer to replace a failed flywheel according to an embodiment of the present invention includes the following steps:
s1, calculating target output torque of each non-fault flywheel based on mounting matrixes corresponding to all flywheels of a satellite;
the satellite is provided with 3 flywheels for explanation:
the mounting matrix Mn corresponding to all flywheels is:
wherein, a 11 Cosine value, a, representing the angle between the output shaft of the first flywheel and the x-axis of the body coordinate system of the satellite 12 Cosine value, a, representing the angle between the output axis of the first flywheel and the y-axis of the body coordinate system of the satellite 13 Cosine value representing the angle between the output shaft of the first flywheel and the z-axis of the body coordinate system of the satellite, a 21 Cosine value, a, representing the angle between the output shaft of the second flywheel and the x-axis of the body coordinate system of the satellite 22 Cosine value, a, representing the angle between the output shaft of the second flywheel and the y-axis of the body coordinate system of the satellite 23 Cosine value, a, representing the angle between the output shaft of the second flywheel and the z-axis of the body coordinate system of the satellite 31 Cosine value, a, representing the angle between the output shaft of the third flywheel and the x-axis of the body coordinate system of the satellite 32 Cosine value, a, representing the angle between the output shaft of the third flywheel and the y-axis of the body coordinate system of the satellite 33 Showing the output shaft of a third flywheel andcosine values of the included angles between the z-axes of the body coordinate system of the satellite.
Wherein, the body coordinate system of the satellite refers to: and establishing a body coordinate system of the satellite by taking the center of mass of the satellite as an origin and three inertia main shafts of the satellite as an x axis, a y axis and a z axis.
And S2, calculating the target output magnetic moment of each magnetic torquer of the satellite according to the expected moment of the fault flywheel, wherein one magnetic torquer is arranged on each of three coordinate axes in a body coordinate system of the satellite, and the three magnetic torquers are installed on the satellite.
And S3, controlling each non-fault flywheel to output a corresponding target output torque, and controlling each magnetic torquer to output a corresponding target output magnetic torque so as to control the attitude of the satellite.
When any flywheel on the satellite breaks down, the magnetic torquer arranged in the existing three-orthogonal mode on the satellite is adopted to complete the control function of the broken flywheel, so that the configuration of a plurality of redundant flywheels is avoided, the development cost of the satellite is greatly reduced, and the weight of the satellite is reduced.
Optionally, in the foregoing technical solution, in S1, the step of calculating the target output torque of each non-faulty flywheel includes:
s10, deleting elements related to the fault flywheel from the installation matrix to obtain a corrected installation matrix, specifically:
1) For example, the first flywheel is used as a failed flywheel, the second flywheel and the third flywheel are used as non-failed flywheels, and in the installation matrix, the element associated with the first flywheel is the cosine value a of the included angle between the output shaft of the first flywheel and the x axis of the body coordinate system of the satellite 11 All elements of a row and all elements of a column, a 11 All elements of the row include: a is 21 And a 31 ，a 11 All elements listed include: a is 12 And a 13 Obtaining a corrected mounting matrix of
2) E.g. secondThe flywheels are used as fault flywheels, the first flywheel and the third flywheel are used as non-fault flywheels, in the installation matrix, the element related to the second flywheel is the cosine value a of the included angle between the output shaft of the second flywheel and the x axis of the body coordinate system of the satellite 22 All elements of a row and all elements of a column, a 22 All elements of a row include: a is 12 And a 32 ，a 11 All elements listed include: a is 21 And a 23 Obtaining a corrected mounting matrix of
3) For example, the third flywheel is used as a failed flywheel, the first flywheel and the second flywheel are used as non-failed flywheels, and in the installation matrix, the element associated with the third flywheel is a cosine value a of an included angle between an output shaft of the third flywheel and an x axis of a body coordinate system of the satellite 33 All elements of a row and all elements of a column, a 33 All elements of the row include: a is 13 And a 23 ，a 33 All elements listed include: a is a 31 And a 32 Obtaining a corrected mounting matrix of
And S11, calculating the target output torque of each non-fault flywheel according to the corrected installation matrix and the expected torque of each non-fault flywheel. Specifically, the method comprises the following steps:
1) For example, the first flywheel being a failed flywheel, the second flywheel and the third flywheel being non-failed flywheels, the resulting modified mounting matrix
For illustration, specifically:
(1) the calculation process of the expected torque of the second flywheel and the expected torque of the third flywheel is as follows:
wherein,
E 1 the coefficient matrix (Tm) corresponding to the first flywheel being a failed flywheel 2 ) ' indicates the desired moment of the second flywheel, (Tm) 3 ) ' represents a desired torque of a third flywheel, tcx represents a control torque required for the x-axis of the body coordinate system of the satellite, tcy represents a control torque required for the y-axis of the body coordinate system of the satellite, tcz represents a control torque required for the z-axis of the body coordinate system of the satellite, and Tcx, tcy, and Tcz are preset by a person;
(2) the calculation process of the target output torque of the second flywheel and the target output torque of the third flywheel is as follows:
(Twm 2 ) ' indicates the target output torque of the second flywheel, (Twm) 3 ) ' represents a target output torque of the third flywheel;
2) For example, the second flywheel acts as a failed flywheel, the first flywheel and the third flywheel act as non-failed flywheels, and the resulting modified mounting matrix
For illustration, specifically:
(1) the calculation process of the expected torque of the second flywheel and the expected torque of the third flywheel is as follows:
wherein,
E 2 the coefficient matrix (Tm) corresponding to the second flywheel being a failed flywheel 1 ) "indicates the desired moment of the first flywheel, (Tm) 3 ) "represents the desired moment of the third flywheel, tcx represents the control moment required for the x-axis of the satellite's body coordinate system, and Tcy represents the satellite's bodyThe control moment required by the y axis of the coordinate system is Tcz which represents the control moment required by the z axis of the body coordinate system of the satellite, and Tcx, tcy and Tcz are preset by people;
(2) the calculation process of the target output torque of the second flywheel and the target output torque of the third flywheel is as follows:
(Twm 1 ) "indicates the target output torque of the first flywheel, (Twm) 3 ) ' represents a target output torque of the third flywheel;
3) For example, the third flywheel acts as a failed flywheel, the first flywheel and the second flywheel act as non-failed flywheels, and the resulting modified mounting matrix
For illustration, specifically:
(1) the expected torque of the first flywheel and the expected torque of the second flywheel are calculated by the following steps:
(2) the calculation process of the target output torque of the first flywheel and the target output torque of the second flywheel is as follows:
(Twm 1 ) "' indicates the target output torque of the first flywheel, (Twm) 2 ) And represents the target output torque of the second flywheel.
Optionally, in the above technical solution, the method further includes:
and obtaining the expected torque of the fault flywheel according to the mounting matrix and the expected torque of each non-fault flywheel. Specifically, the method comprises the following steps:
1) For example, when the first flywheel is a failed flywheel, the expected torque of the first flywheel, i.e., the failed flywheel, is calculated using the following formula
2) For example, when the second flywheel is a failed flywheel, the expected torque of the second flywheel, i.e., the failed flywheel, is calculated using the following formula
3) For example, when the third flywheel is a failed flywheel, the expected torque of the third flywheel, i.e., the failed flywheel, is calculated using the following formula
Optionally, in the foregoing technical solution, in S2, calculating a target output magnetic moment of each magnetic torquer of the satellite according to the expected moment of the faulty flywheel, includes:
s20, calculating the reason according to the first formulaThe target output magnetic moment of each magnetic torquer corresponding to the barrier flywheel is as follows:
wherein,
the measurement value of the geomagnetic intensity vector representing the orbit position of the satellite in the body coordinate system of the satellite is a vector and is provided by a satellite-borne magnetometer on the satellite,
indicating the expected torque of the failed flywheel,
the MT includes 3 elements, and respectively corresponds to target output magnetic moments of the magnetic torquer on three coordinate axes in the body coordinate system of the satellite, specifically:
1) For example, when the first flywheel is a failed flywheel, at this time,
the target output magnetic moment of each magnetotorquer of the satellite is calculated using the following formula, at which time the first formula becomes:
therefore, 3 elements in MT and MT are calculated and respectively correspond to target output magnetic moments of the magnetic torquer on three coordinate axes in a body coordinate system of the satellite;
2) For example, when the second flywheel is a failed flywheel, at this time,
the target output magnetic moment of each magnetotorquer of the satellite is calculated using the following formula, at which time the first formula becomes:
therefore, 3 elements in MT and MT are calculated and respectively correspond to target output magnetic moments of the magnetic torquer on three coordinate axes in a body coordinate system of the satellite;
3) When, for example, the third flywheel is a faulty flywheel, at this time,
the target output magnetic moment of each magnetotorquer of the satellite is calculated using the following formula, at which time the first formula becomes:
therefore, 3 elements in MT and MT are calculated and respectively correspond to the target output magnetic moment of the magnetic torquer on three coordinate axes in the body coordinate system of the satellite.
In the above embodiments, although the steps are numbered as S1, S2, etc., but only the specific embodiments are given in the present application, and a person skilled in the art may adjust the execution sequence of S1, S2, etc. according to the actual situation, which is also within the protection scope of the present invention, it is understood that some embodiments may include some or all of the above embodiments.
As shown in fig. 2, a satellite attitude control system 200 using a magnetic torquer to replace a failed flywheel according to an embodiment of the present invention includes a first computing module 210, a second computing module 220, and a control module 230;
the first calculation module 210 is configured to: calculating the target output torque of each non-fault flywheel based on the mounting matrixes corresponding to all the flywheels of the satellite;
the second calculation module 220 is configured to: calculating the target output magnetic moment of each magnetic torquer of the satellite according to the expected moment of the fault flywheel, wherein one magnetic torquer is arranged on each of three coordinate axes in a body coordinate system of the satellite;
the control module 230 is configured to: and controlling each non-fault flywheel to output corresponding target output torque, and controlling each magnetic torquer to output corresponding target output magnetic moment so as to control the attitude of the satellite.
When any flywheel on the satellite breaks down, the magnetic torquer arranged in the existing three-orthogonal mode on the satellite is adopted to complete the control function of the broken flywheel, so that the configuration of a plurality of redundant flywheels is avoided, the development cost of the satellite is greatly reduced, and the weight of the satellite is reduced.
Optionally, in the above technical solution, the first calculating module 210 is specifically configured to:
deleting elements related to the fault flywheel from the installation matrix to obtain a corrected installation matrix;
and calculating the target output torque of each non-fault flywheel according to the corrected installation matrix and the expected torque of each non-fault flywheel.
Optionally, in the above technical solution, the computing apparatus further includes a third computing module, and the third computing module is configured to:
and obtaining the expected torque of the fault flywheel according to the mounting matrix and the expected torque of each non-fault flywheel.
Optionally, in the foregoing technical solution, the second calculating module 220 is specifically configured to:
calculating the target output magnetic moment of each magnetic torquer corresponding to the fault flywheel according to a first formula, wherein the first formula is as follows:
wherein,
a measurement value of a geomagnetic intensity vector indicating an orbital position of a satellite in a body coordinate system of the satellite,
indicating the expected torque of the failed flywheel,
and a module for representing a vector of the geomagnetic intensity, wherein the MT comprises 3 elements and respectively corresponds to the target output magnetic moment of the magnetic torquer on each coordinate axis in the body coordinate system of the satellite.
The above steps for realizing the corresponding functions of the parameters and the unit modules in the satellite attitude control system 200 using a magnetic torquer to replace a failed flywheel of the present invention can refer to the parameters and the steps in the above embodiment of the satellite attitude control method using a magnetic torquer to replace a failed flywheel, which are not described herein again.
The storage medium of an embodiment of the present invention stores instructions, and when the instructions are read by a computer, the computer is caused to execute a satellite attitude control method in which a magnetic torquer is used to replace a faulty flywheel in any of the above embodiments.
An apparatus according to an embodiment of the present invention includes a processor and the storage medium described above, and the processor executes instructions in the storage medium. The equipment is electronic equipment, and specifically can select computers, mobile phones and the like.
As will be appreciated by one skilled in the art, the present invention may be embodied as a system, method or computer program product.
Accordingly, the present disclosure may be embodied in the form of: may be embodied entirely in hardware, entirely in software (including firmware, resident software, micro-code, etc.) or in a combination of hardware and software, and may be referred to herein generally as a "circuit," module "or" system. Furthermore, in some embodiments, the invention may also be embodied in the form of a computer program product in one or more computer-readable media having computer-readable program code embodied in the medium.
Any combination of one or more computer-readable media may be employed. The computer readable medium may be a computer readable signal medium or a computer readable storage medium. A computer readable storage medium may be, for example, but not limited to, an electronic, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or any combination of the foregoing. More specific examples (a non-exhaustive list) of the computer-readable storage medium include an electrical connection having one or more wires, a portable computer diskette, a hard disk, a Random Access Memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM or flash memory), an optical fiber, a portable compact disc read-only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination of the foregoing. In the context of this document, a computer readable storage medium may be any tangible medium that can contain, or store a program for use by or in connection with an instruction execution system, apparatus, or device.
Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention, and that variations, modifications, substitutions and alterations can be made to the above embodiments by those of ordinary skill in the art within the scope of the present invention.
Claims (10)
1. A satellite attitude control method for replacing a failed flywheel by a magnetic torquer is characterized by comprising the following steps:
calculating the target output torque of each non-fault flywheel based on the mounting matrixes corresponding to all the flywheels of the satellite;
calculating the target output magnetic moment of each magnetic torquer of the satellite according to the expected moment of the fault flywheel, wherein one magnetic torquer is arranged on each of three coordinate axes in a body coordinate system of the satellite;
and controlling each non-fault flywheel to output corresponding target output torque, and controlling each magnetic torquer to output corresponding target output magnetic moment so as to control the attitude of the satellite.
2. The satellite attitude control method for replacing a failed flywheel with a magnetic torquer as claimed in claim 1, wherein the process of calculating the target output torque of each non-failed flywheel comprises:
deleting elements associated with the fault flywheel from the installation matrix to obtain a corrected installation matrix;
and calculating the target output torque of each non-fault flywheel according to the corrected mounting matrix and the expected torque of each non-fault flywheel.
3. The satellite attitude control method for replacing a failed flywheel with a magnetic torquer according to claim 2, further comprising:
and obtaining the expected torque of the fault flywheel according to the installation matrix and the expected torque of each non-fault flywheel.
4. The method as claimed in claim 3, wherein the calculating the target output magnetic moment of each magnetic torquer of the satellite according to the expected moment of the failed flywheel comprises:
calculating the target output magnetic moment of each magnetic torquer corresponding to the fault flywheel according to a first formula, wherein the first formula is as follows:
wherein,
a measurement value of a geomagnetic intensity vector representing an orbital position of a satellite in a body coordinate system of the satellite,
represents the expected torque of the faulty flywheel,
and MT comprises 3 elements and corresponds to the target output magnetic moment of the magnetic torquer on each coordinate axis in the body coordinate system of the satellite.
5. A satellite attitude control system using a magnetic torquer to replace a fault flywheel is characterized by comprising a first calculation module, a second calculation module and a control module;
the first computing module is to: calculating the target output torque of each non-fault flywheel based on the installation matrixes corresponding to all the flywheels of the satellite;
the second calculation module is configured to: calculating the target output magnetic moment of each magnetic torquer of the satellite according to the expected moment of the fault flywheel, wherein one magnetic torquer is arranged on each of three coordinate axes in a body coordinate system of the satellite;
the control module is used for: and controlling each non-fault flywheel to output corresponding target output torque, and controlling each magnetic torquer to output corresponding target output magnetic moment so as to control the attitude of the satellite.
6. The satellite attitude control system using a magnetic torquer to replace a failed flywheel according to claim 5, wherein the first computing module is specifically configured to:
deleting elements associated with the fault flywheel from the installation matrix to obtain a corrected installation matrix;
and calculating the target output torque of each non-fault flywheel according to the corrected installation matrix and the expected torque of each non-fault flywheel.
7. The satellite attitude control system using a magnetic torquer to replace a failed flywheel according to claim 6, further comprising a third computing module, wherein the third computing module is configured to:
and obtaining the expected torque of the fault flywheel according to the installation matrix and the expected torque of each non-fault flywheel.
8. The satellite attitude control system using a magnetic torquer to replace a failed flywheel according to claim 7, wherein the second computing module is specifically configured to:
calculating the target output magnetic moment of each magnetic torquer corresponding to the fault flywheel according to a first formula, wherein the first formula is as follows:
wherein,
geomagnetic intensity vector representing orbital position of satellite in body of satelliteThe measured values in the coordinate system are,
represents the expected torque of the faulty flywheel,
and MT comprises 3 elements and corresponds to the target output magnetic moment of the magnetic torquer on each coordinate axis in the body coordinate system of the satellite.
9. A storage medium having stored therein instructions which, when read by a computer, cause the computer to carry out a method of satellite attitude control using a magnetic torquer in place of a faulty flywheel as claimed in any one of claims 1 to 4.
10. An apparatus comprising a processor and the storage medium of claim 9, the processor executing instructions in the storage medium.
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| CN116331524A (en) * | 2023-05-30 | 2023-06-27 | 北京钧天航宇技术有限公司 | Method and device for determining installation position of satellite magnetic torquer |
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| CN116331524A (en) * | 2023-05-30 | 2023-06-27 | 北京钧天航宇技术有限公司 | Method and device for determining installation position of satellite magnetic torquer |
| CN116331524B (en) * | 2023-05-30 | 2023-07-21 | 北京钧天航宇技术有限公司 | Method and device for determining installation position of satellite magnetic torquer |
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