CN118145017A - Inclined sailboard satellite long-term sun-facing control method based on three control moment gyroscopes - Google Patents
Inclined sailboard satellite long-term sun-facing control method based on three control moment gyroscopes Download PDFInfo
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Abstract
The invention relates to a long-term sun-facing control method of a diagonal sailboard satellite based on three control moment gyroscopes, which comprises the steps of utilizing an inertial system fixed vector and a solar vector to establish a solar inertial coordinate system and determining a sun-facing orientation posture; three control moment gyroscopes are configured, wherein two control moment gyroscopes are respectively arranged along a Y axis and a Z axis of the satellite body system, and the other control moment gyroscopes are arranged on an angular bisector of an included angle formed by the Z axis and the Y axis of the satellite body system; and calculating a three-axis attitude control command moment according to the determined sun-facing orientation attitude, and designing command rotating speeds of three control moment gyroscopes based on the three-axis attitude control command moment to realize long-term sun-facing control of the inclined sailboard satellite. The solar array can realize the opposite direction of the sailboard installed at any inclined angle, and ensures that the sailboard is subjected to direct sunlight; the simplest configuration of the three control moment gyroscopes is adopted to realize stable control of the daily attitude, so that the problem of energy shortage of the inclined track is solved.
Description
Technical Field
The invention belongs to the technical field of spacecraft control, and relates to a long-term sun-facing control method of a diagonal sailboard satellite based on three control moment gyroscopes.
Background
With the continuous expansion of the aerospace field, many tasks suggest that the use of inclined orbits has more advantages for achieving observations. For satellites operating in inclined orbits, since the solar sailboards are fixed wing mounted, there is no drive mechanism itself, and in order to ensure the whole satellite energy, it is necessary to keep the sailboards oriented for a long time to the sun. Because of the limitations of the sensor mounting layout and the demand for thermal control of satellites, the sailboards of some satellites are not mounted along the main shaft of the body any more, but are mounted obliquely relative to the stars. Therefore, the conventional sun-oriented control method is no longer applicable.
Disclosure of Invention
The invention solves the technical problems that: the method for controlling the sun-facing of the inclined sailboard satellite for a long time based on the three control moment gyroscopes is provided, and the problem of controlling the sun-facing of the inclined sailboard satellite for a long time under the simplest configuration of an actuating mechanism is solved.
The solution of the invention is as follows: a long-term sun-facing control method of a diagonal sailboard satellite based on three control moment gyroscopes comprises the following steps:
establishing a solar inertial coordinate system by using an inertial system fixed vector and a solar vector, and determining a sun orientation posture;
Three control moment gyroscopes are configured, wherein two control moment gyroscopes are respectively arranged along a Y axis and a Z axis of the satellite body system, and the other control moment gyroscopes are arranged on an angular bisector of an included angle formed by the Z axis and the Y axis of the satellite body system;
And calculating three-axis attitude control command moment according to the determined sun-facing orientation attitude, designing command rotating speeds of three control moment gyroscopes based on the three-axis attitude control command moment, and generating three-axis control moment by each control moment gyroscope according to the corresponding command rotating speed to realize long-term sun-facing orientation of the inclined sailboard satellite.
Further, a solar inertial coordinate system is establishedComprising the following steps:
And normalizing;
Let the inertia system fix the vector Then:
And normalizing;
And normalizing;
Wherein, (s xi,syi,szi) is a unit solar vector in an inertial coordinate system.
Further, the determining a relative day orientation pose includes:
s1, calculating a rotation matrix A i←sun of the inertial coordinate system at the current moment relative to the solar inertial coordinate system:
s2, calculating quaternion q i←sun of the inertial coordinate system at the current moment relative to the solar inertial coordinate system:
Order the If |a 23 | > 0.9999, the attitude angle of the inertial coordinate system at the current moment relative to the solar inertial coordinate system is calculated as follows:
Otherwise:
Wherein, The roll angle, θ is the pitch angle, ψ is the yaw angle;
Solving a quaternion q i←sun by the attitude angle:
s3, calculating quaternion of satellite body system relative to solar inertial coordinate system Wherein q b←i is the quaternion of the satellite body system relative to the inertial coordinate system at the current moment;
s4, setting a target quaternion according to the offset installation angle of the sailboard:
If the inclined angle of the sailboard is alpha, the sailboard-X axis sun alignment is required to be realized, and the target quaternion is set as
S5, calculating a deviation quaternion Delq according to a quaternion q b←sun and a target quaternion q aim of the satellite body system relative to a solar inertial coordinate system, and normalizing: Taking the vector part of the deviation quaternion Delq as Delq.v, and calculating a control attitude angle con,i = -2 Delq.v;
S6, setting three-axis target angular velocities omega aim to be 0, setting an output angular velocity reference to be the relative inertial system angular velocity omega bi of the satellite body system, and calculating and controlling the attitude angular velocity omega con,i＝ωbi-ωaim.
Further, a control moment gyro installed along the Y axis of the satellite body system is recorded as a control moment gyro A and is used for Z axis control; the control moment gyro arranged along the Z axis of the satellite body system is marked as a control moment gyro B and is used for Y axis control; the control moment gyroscopes arranged on the horizontal parting lines of the angle clamping angles of the Z axis and the Y axis of the satellite body system are marked as control moment gyroscopes C and are used for X axis control.
Further, the calculating the three-axis command moment includes:
calculating PID control parameters:
Wherein, K P,i is the angle control gain, K D,i is the angular velocity control gain, K I,i is the angle integral control gain, I i is the star main inertia, xi i is the damping ratio, omega ni is the control bandwidth, and K P,i、kD,i is the tuning coefficient;
Calculating a three-axis attitude control command moment T con,i according to the PID control parameter and the sun-oriented attitude:
TPD,i＝KP,i·anglecon,i+KD,i·ωcon,i
Wherein angle con,i is a control attitude angle, ω con,i is a control attitude angular speed, and t c is a control period; for the attitude control integral quantity of the current period, the initial value/>, of the attitude control integral quantity Defaulting to zero.
Further, the design of the command rotational speeds of the three control moment gyroscopes includes:
the corresponding control moment gyro command rotating speed delta calculation method based on the calculated attitude control triaxial command moment T con,i is as follows:
C1＝A1 cosδi-B1 sinδi
Wherein, Delta i is the outer frame position of the three control moment gyroscopes, and h 0 is the nominal angular momentum of the single control moment gyroscope;
according to the actual execution capacity of a single machine, the rotation speed is instructed to three control moment gyroscopes Clipping is performed.
Compared with the prior art, the invention has the beneficial effects that:
(1) In order to realize that the normal line of the inclined sailboard points to the sun, the invention can realize the opposite-sun direction of the sailboard installed at any inclined angle by setting the offset target attitude angle without changing the installation mode of the existing actuating mechanism, and ensures that the sailboard is subjected to direct sun irradiation.
(2) Because large-angle maneuver does not exist in the sun orientation for a long time, the invention can realize stable control of the sun attitude by adopting the simplest configuration of the three control moment gyroscopes, can effectively reduce the whole star power consumption and improve the problem of energy shortage of the inclined orbit.
Drawings
FIG. 1 is a flow chart of a method for long-term sun-to-sun control of a diagonal sailboard satellite according to the invention;
FIG. 2 is a schematic diagram of the relationship between the solar panel coordinate system and the satellite body coordinate system in the present invention;
wherein Xs, ys and Zs represent a solar sailboard coordinate system, and Xb, yb and Zb represent a satellite body coordinate system;
FIG. 3 illustrates the installation of a control moment gyro according to the present invention;
Fig. 4 to 8 are long-term daily stability control mathematical simulation results of a 45-degree oblique sailboard satellite based on three control moment gyroscopes in the embodiment of the present invention, wherein: fig. 4 is a three-axis attitude angle curve, fig. 5 is a three-axis attitude angular velocity curve, fig. 6 is a star angular momentum curve, fig. 7 is a moment gyro frame rotation angle curve, and fig. 8 is a moment gyro frame rotation speed curve.
Detailed Description
The invention is further illustrated in the following figures and examples.
Example 1
The embodiment provides a long-term sun-to-sun control method of an inclined sailboard satellite based on three control moment gyroscopes, as shown in fig. 1, comprising the following steps:
step one, a solar inertial coordinate system is established by utilizing an inertial system fixed vector and a solar vector, and a sun orientation posture is determined. The method specifically comprises the following steps:
110. Establishing a solar inertial coordinate system
And normalizing;
Let the inertia system fix the vector Then:
And normalizing;
And normalizing;
Wherein, (s xi,syi,szi) is a unit solar vector in an inertial coordinate system.
Determining the sun-facing orientation gesture, comprising steps 120-160:
120. Calculating a rotation matrix A i←sun of the inertial coordinate system at the current moment relative to the solar inertial coordinate system:
130. calculating quaternion q i←sun of the inertial coordinate system at the current moment relative to the solar inertial coordinate system:
Order the If |a 23 | > 0.9999, the attitude angle of the inertial coordinate system at the current moment relative to the solar inertial coordinate system is calculated as follows:
Otherwise:
Wherein, The roll angle, θ is the pitch angle, ψ is the yaw angle;
Solving a gesture quaternion q i←sun by the gesture angle:
140. calculating quaternion of satellite body system relative to solar inertial coordinate system Wherein q b←i is the quaternion of the satellite body system relative to the inertial coordinate system at the current moment;
150. setting a target quaternion according to the offset installation angle of the sailboard:
If the oblique angle of the sailboard is α (as shown in fig. 2, α is 45 °) in this embodiment, it is required to realize the opposite sun of the sailboard-X axis, and the target quaternion is set as
160. According to the quaternion q b←sun and the target quaternion q aim of the satellite body system relative to the solar inertial coordinate system, calculating a deviation quaternion Delq and normalizing: Taking the vector part of the deviation quaternion Delq as Delq.v, and calculating a control attitude angle con,i = -2 Delq.v;
170. Setting the three-axis target angular velocities omega aim to be 0, setting the output angular velocity reference to be the relative inertial system angular velocity omega bi of the satellite body system, and calculating the control attitude angular velocity omega con,i＝ωbi-ωaim.
Step two, three control moment gyroscopes are configured from the complexity of system configuration and the simplicity of engineering, wherein two control moment gyroscopes are respectively arranged along the Y axis and the Z axis of the satellite body system, and the other control moment gyroscopes are arranged on an angular bisector of an included angle formed by the Z axis and the Y axis of the satellite body system.
As shown in fig. 3, the control moment gyro a installed along the Y axis of the satellite body system is mainly used for Z axis control, the control moment gyro B installed along the Z axis of the satellite body system is mainly used for Y axis control, and the control moment gyro C installed on the bisector of the angle between the Z axis and the Y axis of the satellite body system is mainly used for X axis control.
And thirdly, calculating a three-axis attitude control command moment according to the sun orientation attitude determined in the first step, and designing command rotating speeds of three control moment gyroscopes based on the three-axis attitude control command moment, wherein each control moment gyroscope generates three-axis control moment according to the corresponding command rotating speed so as to realize long-term sun orientation of the inclined sailboard satellite. The method specifically comprises the following steps:
310. calculating PID control parameters:
Wherein, K P,i is the angle control gain, K D,i is the angular velocity control gain, K I,i is the angle integral control gain, and the three axes are respectively arranged; i i is the star moment of inertia, ζ i is the damping ratio (default 0.707), ω ni is the control bandwidth, k P,i、kD,i is the tuning coefficient, default to 1.
320. Calculating a three-axis attitude control command moment T con,i according to the PID control parameter and the sun-oriented attitude:
TPD,i＝KP,i·anglecon,i+KD,i·ωcon,i
Wherein angle con,i is a control attitude angle, ω con,i is a control attitude angular speed, and t c is a control period; for the attitude control integral quantity of the current period, the initial value/>, of the attitude control integral quantity Defaulting to zero.
330. Obtaining a corresponding control moment gyro command rotating speed based on the calculated attitude control triaxial command moment T con,i The calculation method comprises the following steps:
C1＝A1 cosδi-B1 sinδi
Wherein, Delta i is the position of the outer frame of the three control moment gyroscopes, the unit is the angle, and the angle is converted into radian during calculation. h 0 = 25Nms, which is the nominal angular momentum of the single control moment gyro.
340. According to the actual execution capacity of a single machine, the rotation speed is instructed to three control moment gyroscopesAnd limiting amplitude, and controlling the moment gyro to generate corresponding triaxial control moment according to the instruction rotating speed so as to realize long-term sun control of the inclined sailboard.
The long-term daily stability control mathematical simulation results of the 45-degree inclined sailboard satellite based on the three control moment gyroscopes are shown in fig. 4-8 respectively.
Although the present invention has been described in terms of the preferred embodiments, it is not intended to be limited to the embodiments, and any person skilled in the art can make any possible variations and modifications to the technical solution of the present invention by using the methods and technical matters disclosed above without departing from the spirit and scope of the present invention, so any simple modifications, equivalent variations and modifications to the embodiments described above according to the technical matters of the present invention are within the scope of the technical matters of the present invention.
What is not described in detail in the present specification is a well known technology to those skilled in the art.
Claims (6)
1. The long-term sun-facing control method of the inclined sailboard satellite based on the three control moment gyroscopes is characterized by comprising the following steps of:
establishing a solar inertial coordinate system by using an inertial system fixed vector and a solar vector, and determining a sun orientation posture;
Three control moment gyroscopes are configured, wherein two control moment gyroscopes are respectively arranged along a Y axis and a Z axis of the satellite body system, and the other control moment gyroscopes are arranged on an angular bisector of an included angle formed by the Z axis and the Y axis of the satellite body system;
And calculating three-axis attitude control command moment according to the determined sun-facing orientation attitude, designing command rotating speeds of three control moment gyroscopes based on the three-axis attitude control command moment, and generating three-axis control moment by each control moment gyroscope according to the corresponding command rotating speed to realize long-term sun-facing orientation of the inclined sailboard satellite.
2. The method for long-term sun-to-sun control of a three moment gyro-based tilt-sail satellite of claim 1, wherein a solar inertial coordinate system is establishedComprising the following steps:
And normalizing;
Let the inertia system fix the vector Then:
And normalizing;
And normalizing;
Wherein, (s xi,syi,szi) is a unit solar vector in an inertial coordinate system.
3. The method for long term sun-to-sun control of a three moment gyro-based tilt-array satellite of claim 2, wherein said determining a sun-to-orientation attitude comprises:
s1, calculating a rotation matrix A i←sun of the inertial coordinate system at the current moment relative to the solar inertial coordinate system:
s2, calculating quaternion q i←sun of the inertial coordinate system at the current moment relative to the solar inertial coordinate system:
Order the If |a 23 | > 0.9999, the attitude angle of the inertial coordinate system at the current moment relative to the solar inertial coordinate system is calculated as follows:
θ＝0，/>
Otherwise:
Wherein, The roll angle, θ is the pitch angle, ψ is the yaw angle;
Solving a quaternion q i←sun by the attitude angle:
s3, calculating quaternion of satellite body system relative to solar inertial coordinate system Wherein q b←i is the quaternion of the satellite body system relative to the inertial coordinate system at the current moment;
s4, setting a target quaternion according to the offset installation angle of the sailboard:
If the inclined angle of the sailboard is alpha, the sailboard-X axis sun alignment is required to be realized, and the target quaternion is set as
S5, calculating a deviation quaternion Delq according to a quaternion q b←sun and a target quaternion q aim of the satellite body system relative to a solar inertial coordinate system, and normalizing: Taking the vector part of the deviation quaternion Delq as Delq.v, and calculating a control attitude angle con,i = -2 Delq.v;
S6, setting three-axis target angular velocities omega aim to be 0, setting an output angular velocity reference to be the relative inertial system angular velocity omega bi of the satellite body system, and calculating and controlling the attitude angular velocity omega con,i＝ωbi-ωaim.
4. The method for long-term sun-to-sun control of a three control moment gyro-based tilt-setting sailboard satellite according to claim 3, wherein the control moment gyro installed along the Y-axis of the satellite body system is denoted as control moment gyro a for Z-axis control; the control moment gyro arranged along the Z axis of the satellite body system is marked as a control moment gyro B and is used for Y axis control; the control moment gyroscopes arranged on the horizontal parting lines of the angle clamping angles of the Z axis and the Y axis of the satellite body system are marked as control moment gyroscopes C and are used for X axis control.
5. The method for long-term sun-to-sun control of a three moment gyro-based tilt-sail panel satellite of claim 4, wherein said calculating the attitude control triaxial command moment comprises:
calculating PID control parameters:
Wherein, K P,i is the angle control gain, K D,i is the angular velocity control gain, K I,i is the angle integral control gain, I i is the star main inertia, xi i is the damping ratio, omega ni is the control bandwidth, and K P,i、kD,i is the tuning coefficient;
Calculating a three-axis attitude control command moment T con,i according to the PID control parameter and the sun-oriented attitude:
TPD,i＝KP,i·anglecon,i+KD,i·ωcon,i
Wherein angle con,i is a control attitude angle, ω con,i is a control attitude angular speed, and t c is a control period; for the attitude control integral quantity of the current period, the initial value/>, of the attitude control integral quantity Defaulting to zero.
6. The method for long-term sun-to-sun control of a three moment gyro-based tilt-set sailboard satellite of claim 5, wherein the designing the commanded rotational speed of the three moment gyros comprises:
Obtaining a corresponding control moment gyro command rotating speed based on the calculated attitude control triaxial command moment T con,i The calculation method comprises the following steps:
C1＝A1 cosδi-B1 sinδi
Wherein, Delta i is the outer frame position of the three control moment gyroscopes, and h 0 is the nominal angular momentum of the single control moment gyroscope;
according to the actual execution capacity of a single machine, the rotation speed is instructed to three control moment gyroscopes Clipping is performed.
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