CN117647990A - Spacecraft attitude active self-stabilization control method of orbit virtual field - Google Patents
Spacecraft attitude active self-stabilization control method of orbit virtual field Download PDFInfo
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Abstract
The invention discloses a spacecraft attitude active self-stabilization control method of an orbit virtual field, which is implemented according to the following steps: extrapolating the orbit position and speed of the spacecraft by using the GNSS data of the current spacecraft; calculating space disturbance moment T of the spacecraft according to the extrapolated orbit position and speed of the spacecraft and the correlation constant of the spacecraft; judging whether the space disturbance moment T reaches a threshold value or not, and determining that the spacecraft needs to be subjected to active attitude adjustment; calculating a spacecraft attitude adjustment strategy through the space disturbance moment T; and controlling the spacecraft to adjust the posture at a specific moment according to the posture of the spacecraft. The spacecraft attitude active self-stabilization control method realizes the spacecraft control of attitude orbit control when no measuring mechanism or a measuring mechanism is not credible; for the ultra-low cost micro spacecraft, the attitude control of the micro spacecraft can be realized while the compression cost is realized.
Description
Technical Field
The invention belongs to the technical field of spacecraft attitude control, and particularly relates to a spacecraft attitude active self-stabilization control method of an orbit virtual field.
Background
The current spacecraft attitude orbit control design is based on real-time measurement data monitoring of star-sensitive, hypersensitive, gyroscopic and other on-board measuring instruments, the change of the attitude of the spacecraft is found, and then the quantity to be corrected is output to an executing mechanism such as a magnetic torquer, a thruster and the like through calculation of a spaceborne computer. The accuracy dependence of the method on the measured value of the measuring instrument is strong, if the on-board measuring device is interfered by a magnetic field, single particles and the like, the measurement is deviated, the posture and the position of the spacecraft are easily deviated, and the load application is affected.
Disclosure of Invention
The invention aims to provide an active self-stabilizing control method for the attitude of a spacecraft in an orbit virtual field, which solves the problem that the existing control method has strong dependence on the measured value of a measuring instrument, so that the attitude and the position of the spacecraft are easy to deviate.
The technical scheme adopted by the invention is that the spacecraft attitude active self-stabilization control method of the orbit virtual field is implemented according to the following steps:
step 1, extrapolating the orbit position and speed of a spacecraft by using GNSS data of the current spacecraft;
step 2, calculating space disturbance moment T of the spacecraft according to the extrapolated orbit position and speed of the spacecraft and the correlation constant of the spacecraft;
step 3, judging whether the space interference moment T reaches a threshold value or not, and determining that the spacecraft needs to be subjected to active attitude adjustment;
step 4, calculating a spacecraft attitude adjustment strategy through the space disturbance moment T;
and 5, controlling the spacecraft to adjust the posture at a specific moment according to the posture of the spacecraft.
The specific process of the step 1 is as follows:
the current spacecraft GNSS receiver obtains accurate UTC time service information t and current GNSS orbit position and speed of the spacecraft from the GNSS system, then carries out orbit extrapolation, and the extrapolation time length is deltat, thereby obtaining the position and speed between the outer space of the estimated spacecraft.
The space disturbance moment in the step 2 comprises solar pressure moment, gravity gradient moment, geomagnetic moment and aerodynamic moment.
The specific process of the step 2 is as follows:
calculating sun light pressure moment T s ：
Moment T generated by solar pressure in spacecraft body coordinates s ＝(Τ sx Τ sy Τ sz ) The unified representation is:
wherein K is t K is the tangential light pressure coefficient n The normal light pressure coefficient is that sigma is the normal included angle between sunlight and a sailboard, eta s Is the angular distance between the zenith of the spacecraft and the sun, and ζ is the azimuth angle of deflection of the rotating shaft of the sailboard, ω 0 Is the angular velocity of spacecraft movement, r x 、r z The deviation amounts of the solar sailboard rotating shaft in the x and z axis directions of the spacecraft body system are respectively r y1 、r y2 The distances of the centers of the single sailboards along the y-axis direction of the spacecraft body are respectively;
calculating gravity gradient moment T g ：
Pair of earth center gravitational fieldThe gravity gradient moment generated by the spacecraft is not only related to the gesture, but also closely related to the quality characteristic of the spacecraft, and under the condition that the spacecraft deviates from the orbit system by a small amount to have zero gesture, the gravity gradient moment T g ＝(T gx T gy T gz ) Expressed as:
wherein,is the inertia tensor of the spacecraft, and is obtained by ground measurement, wherein the inertia tensor is phi, theta, phi and phi>Respectively yaw, roll and pitch angles of the spacecraft;
calculating the geomagnetic moment T m ：
Determining the geomagnetic induction intensity B of the position of the spacecraft according to the extrapolated spacecraft orbit position;
the spacecraft magnetic moment interacts with the earth's magnetic field and generates a moment, expressed as:
Τ m ＝Μ m ×Β (3)
wherein, M m Is the magnetic moment of the spacecraft, and
wherein b e For the total intensity of the earth magnetic moment, θ m Is the included angle between the radial direction of the spacecraft and the geomagnetic equator, e r Is the unit vector of the spacecraft position vector, z m Is a magnetic dipole vector unit vector;
calculating aerodynamic moment T a ：
Determining a velocity vector v of a windward side of a spacecraft relative to the atmosphere according to an extrapolated spacecraft orbit velocity s ；
For a spacecraft of a certain altitude range, the aerodynamic moment T a Is mainly a dryDisturbing moment, assuming that the incident molecule loses its full energy in the collision, then:
Τ a ＝ρ s ×F s (5)
wherein ρ is s Is the position vector of the pressure center of the windward side of the spacecraft relative to the mass center of the spacecraft, and
where ρ is the atmospheric density, S is the frontal area of the spacecraft, n is the unit vector perpendicular to the frontal area, and
v s ＝v o -w e ×r s (7)
wherein v is o Is the orbit speed of the spacecraft, r s Is the geocentric position vector of the spacecraft, which is determined according to the extrapolated spacecraft orbit position, w e An angular velocity vector that is the rotation of the earth;
sun light pressure moment T s Gravity gradient moment T g T of geomagnetic moment m And aerodynamic moment T a The sum is the space disturbance moment of the spacecraft, expressed as:
Τ＝Τ s +Τ g +Τ m +Τ a 。
the specific process of the step 3 is as follows:
setting threshold |T max Judging whether the space interference moment T reaches the threshold value T max I, if I T I is greater than or equal to I T max I, needing posture adjustment, and entering step 4; otherwise, the posture adjustment is not needed, and the step 1 is returned.
The specific process of the step 4 is as follows:
converting the space interference torque T under the orbit coordinate system into the space interference torque under the spacecraft body system;
under the condition that the spacecraft does not have a momentum control device, the attitude stability control is carried out by completely relying on small thrust, and the angular momentum H of the spacecraft system s Expressed as:
H s ＝Ιω (8)
wherein I represents an inertia tensor of the spacecraft in an orbit coordinate system, and w represents an angular velocity vector of the spacecraft;
the spacecraft attitude dynamics equation is:
taking a main inertia axis of the spacecraft as a spacecraft body coordinate system, wherein the inertia array is a diagonal array, and I=diag (I x I y I z ) The method comprises the steps of carrying out a first treatment on the surface of the M represents a moment vector stressed by the spacecraft under the spacecraft body coordinate system;
attitude of spacecraft in orbital coordinate system is determined by roll angleThe pitch angle θ and the yaw angle ψ are represented, and in the three-axis stability control, the spacecraft attitude angles are small, and the attitude matrix is represented as:
the rotational speed of the spacecraft orbit coordinate system in space is (0-omega) 0 0) The rotational speed of the spacecraft is expressed in the spacecraft body coordinate system as:
bringing the expression of ω into a spacecraft attitude dynamics equation to obtain an attitude stability dynamics equation based on the micro thrust as follows:
wherein c= (C x C y C z ) The moment is controlled for a small amount on the satellite; taking a small control moment C on the satellite as a space trunk under the spacecraft body systemDisturbance moment.
The specific process of the step 5 is as follows: and (3) driving the posture adjustment of the spacecraft small thruster according to the spacecraft posture adjustment strategy obtained in the step (4).
The beneficial effects of the invention are as follows:
according to the spacecraft attitude active self-stabilization control method of the orbit virtual field, solar pressure moment, gravity gradient moment, geomagnetic moment and aerodynamic moment of a spacecraft on a subsequent orbit are calculated respectively according to orbit information such as the spacecraft on-board GNSS data, when the predicted space disturbance combined moment reaches a preset maximum threshold, the spacecraft starting attitude is active and stable, and after the moment under an orbit coordinate system is converted into a spacecraft body coordinate system, the space disturbance moment is actively balanced by an executing mechanism. The method realizes spacecraft control of attitude and orbit control when no measuring mechanism or unreliable measuring mechanism exists; for the ultra-low cost micro spacecraft, the attitude control of the micro spacecraft can be realized while the compression cost is realized.
Drawings
FIG. 1 is a flow chart of a spacecraft attitude initiative self-stabilization control method of an orbit virtual field of the invention;
FIG. 2 is a schematic diagram of the control result of the control method of the present invention.
Detailed Description
The invention will be described in detail below with reference to the drawings and the detailed description.
Example 1
The active self-stabilizing control method for the spacecraft attitude of the orbit virtual field in the embodiment is implemented as shown in fig. 1 specifically according to the following steps: extrapolating the orbit position and speed of the spacecraft by using the GNSS data of the current spacecraft; calculating space disturbance moment T of the spacecraft according to the extrapolated orbit position and speed of the spacecraft and the correlation constant of the spacecraft; judging whether the space disturbance moment T reaches a threshold value or not, and determining that the spacecraft needs to be subjected to active attitude adjustment; calculating a spacecraft attitude adjustment strategy through the space disturbance moment T; the spacecraft attitude adjustment strategy drives the attitude adjustment of the spacecraft small thruster. According to the position information of the extrapolated spacecraft, the space interference resultant moment borne by the subsequent orbit position of the spacecraft in orbit is calculated, when the predicted space interference resultant moment reaches a maximum threshold value, the value under the orbit coordinate system is converted into the spacecraft body coordinate system, and the control quantity of each executing mechanism is calculated, so that the active self-stabilizing control of the spacecraft attitude is realized. Meanwhile, the ground evaluates the self-stabilization control of the spacecraft by monitoring the orbit change of the spacecraft, and the spacecraft dynamically updates the virtual field by combining the orbit measurement result, the thrust calibration and the orbit deviation condition fed back by the ground, wherein the result is shown in figure 2. Prepare for the next self-stabilization control.
Example 2
In this embodiment, the specific process of extrapolating the orbit position and speed of the spacecraft by using the GNSS data of the current spacecraft is as follows: the current spacecraft GNSS receiver obtains accurate UTC time service information t and current GNSS orbit position and speed of the spacecraft from the GNSS system, then carries out orbit extrapolation, and the extrapolation time length is deltat, thereby obtaining the position and speed between the outer space of the estimated spacecraft.
The process is mainly based on the measured data (position velocity) of the GNSS receiver, and the numerical integration is performed after the relevant mechanical model is added, so that the position velocity of a period of time in the future is obtained. Therefore, the accuracy of the process is mainly affected by two aspects, namely, the accurate determination of the current measurement data is determined by the performance of the GNSS receiver; and secondly, the accuracy of the mechanical model is extrapolated by the orbit, the more complex the model is, the higher the accuracy is, and the longer the extrapolation time is.
Example 3
On the basis of embodiment 1, the spatial disturbance moment in this embodiment includes solar pressure moment, gravity gradient moment, geomagnetic moment and aerodynamic moment, and the specific process is as follows:
calculating sun light pressure moment T s ：
Moment T generated by solar pressure in spacecraft body coordinates s ＝(Τ sx Τ sy Τ sz ) The unified representation is:
wherein K is t K is the tangential light pressure coefficient n The normal light pressure coefficient is that sigma is the normal included angle between sunlight and a sailboard, eta s Is the angular distance between the zenith of the spacecraft and the sun, and ζ is the azimuth angle of deflection of the rotating shaft of the sailboard, ω 0 Is the angular velocity of spacecraft movement, r x 、r z The deviation amounts of the solar sailboard rotating shaft in the x and z axis directions of the spacecraft body system are respectively r y1 、r y2 The distances of the centers of the single sailboards along the y-axis direction of the spacecraft body are respectively;
calculating gravity gradient moment T g ：
The gravity gradient moment generated by the earth center gravitational field on the spacecraft is not only related to the gesture, but also closely related to the quality characteristic of the spacecraft, and under the condition that the spacecraft deviates from the orbit system by a small amount by zero gesture, the gravity gradient moment T g ＝(T gx T gy T gz ) Expressed as:
wherein,is the inertia tensor of the spacecraft, and is obtained by ground measurement, wherein the inertia tensor is phi, theta, phi and phi>Respectively yaw, roll and pitch angles of the spacecraft;
calculating the geomagnetic moment T m ：
Determining the geomagnetic induction intensity B of the position of the spacecraft according to the extrapolated spacecraft orbit position;
the spacecraft magnetic moment interacts with the earth's magnetic field and generates a moment, expressed as:
Τ m ＝Μ m ×Β (3)
wherein, M m Is the magnetic moment of the spacecraft, and
wherein b e For the total intensity of the earth magnetic moment, θ m Is the included angle between the radial direction of the spacecraft and the geomagnetic equator, e r Is the unit vector of the spacecraft position vector, z m Is a magnetic dipole vector unit vector;
calculating aerodynamic moment T a ：
Determining a velocity vector v of a windward side of a spacecraft relative to the atmosphere according to an extrapolated spacecraft orbit velocity s ；
For a spacecraft of a certain altitude range, the aerodynamic moment T a Is the dominant disturbing moment, assuming that the incident molecule loses all of its energy in the collision:
Τ a ＝ρ s ×F s (5)
wherein ρ is s Is the position vector of the pressure center of the windward side of the spacecraft relative to the mass center of the spacecraft, and
where ρ is the atmospheric density, S is the frontal area of the spacecraft, n is the unit vector perpendicular to the frontal area, and
v s ＝v o -w e ×r s (7)
wherein v is o Is the orbit speed of the spacecraft, r s Is the geocentric position vector of the spacecraft, which is determined according to the extrapolated spacecraft orbit position, w e An angular velocity vector that is the rotation of the earth;
sun light pressure moment T s Gravity gradient moment T g T of geomagnetic moment m And aerodynamic moment T a The sum is the space disturbance moment of the spacecraft, expressed as:
Τ＝Τ s +Τ g +Τ m +Τ a 。
when the predicted space disturbance moment reaches a preset maximum threshold, the satellite is startedThe specific process for judging whether the space interference moment T reaches the threshold value is as follows: setting threshold |T max Judging whether the space interference moment T reaches the threshold value T max I, if I T I is greater than or equal to I T max I, needing posture adjustment; otherwise, no attitude adjustment is required. The setting of the threshold value is mainly determined by different requirements of the on-orbit spacecraft on the attitude stability.
In this embodiment, the spatial disturbance moments are vectors and are time functions, and are all changed in real time according to the position and speed of the on-orbit spacecraft.
Example 4
Based on embodiment 1, in this embodiment, a spacecraft attitude adjustment strategy is calculated through a space disturbance moment t, and the specific process is as follows:
converting the space interference torque T under the orbit coordinate system into the space interference torque under the spacecraft body system; after the moment under the orbit coordinate system is converted into the satellite body coordinate system, the space interference moment is actively balanced by using an executing mechanism.
Under the condition that the spacecraft does not have a momentum control device, the attitude stability control is carried out by completely relying on small thrust, and the angular momentum H of the spacecraft system s Expressed as:
H s ＝Ιω (8)
wherein I represents an inertia tensor of the spacecraft in an orbit coordinate system, and w represents an angular velocity vector of the spacecraft;
the spacecraft attitude dynamics equation is:
taking a main inertia axis of the spacecraft as a spacecraft body coordinate system, wherein the inertia array is a diagonal array, and I=diag (I x I y I z ) The method comprises the steps of carrying out a first treatment on the surface of the M represents a moment vector stressed by the spacecraft under the spacecraft body coordinate system;
attitude of spacecraft in orbital coordinate system is determined by roll angleThe pitch angle θ and the yaw angle ψ are represented, and in the three-axis stability control, the spacecraft attitude angles are small, and the attitude matrix is represented as:
the rotational speed of the spacecraft orbit coordinate system in space is (0-omega) 0 0) The rotational speed of the spacecraft is expressed in the spacecraft body coordinate system as:
bringing the expression of ω into a spacecraft attitude dynamics equation to obtain an attitude stability dynamics equation based on the micro thrust as follows:
wherein c= (C x C y C z ) The moment is controlled for a small amount on the satellite; the small control moment C on the satellite is taken as the space disturbance moment under the spacecraft body system.
The precision of the attitude measurement element on the in-orbit spacecraft is easily constrained by the performance and cost of the spacecraft platform, and the process in the control method does not depend on the attitude measurement element on the satellite, but calculates to obtain the attitude disturbance moment according to the position speed of the in-orbit spacecraft and the high-precision data model of the ground.
By means of the mode, the spacecraft attitude active self-stabilization control method of the orbit virtual field is used for respectively calculating the solar pressure moment, the gravity gradient moment, the geomagnetic moment and the aerodynamic moment of the spacecraft on the subsequent orbit according to orbit information such as the satellite-borne GNSS data of the spacecraft, when the predicted space disturbance combined moment reaches a preset maximum threshold, the spacecraft starting attitude is actively stabilized, the moment under the orbit coordinate system is converted into the spacecraft body coordinate system, and then the executing mechanism is used for actively balancing the space disturbance moment. The method realizes spacecraft control of attitude and orbit control when no measuring mechanism or unreliable measuring mechanism exists; for the ultra-low cost micro spacecraft, the attitude control of the micro spacecraft can be realized while the compression cost is realized.
Claims (7)
1. The spacecraft attitude active self-stabilization control method of the orbit virtual field is characterized by comprising the following steps of:
step 1, extrapolating the orbit position and speed of a spacecraft by using GNSS data of the current spacecraft;
step 2, calculating space disturbance moment T of the spacecraft according to the extrapolated orbit position and speed of the spacecraft and the correlation constant of the spacecraft;
step 3, judging whether the space interference moment T reaches a threshold value or not, and determining that the spacecraft needs to be subjected to active attitude adjustment;
step 4, calculating a spacecraft attitude adjustment strategy through the space disturbance moment T;
and 5, controlling the spacecraft to adjust the posture at a specific moment according to the posture of the spacecraft.
2. The method for actively self-stabilizing the attitude of a spacecraft in an orbit virtual field according to claim 1, wherein the specific process of the step 1 is as follows:
the current spacecraft GNSS receiver obtains accurate UTC time service information t and current GNSS orbit position and speed of the spacecraft from the GNSS system, then carries out orbit extrapolation, and the extrapolation time length is deltat, thereby obtaining the position and speed between the outer space of the estimated spacecraft.
3. The method for actively self-stabilizing the attitude of a spacecraft in an orbit virtual field according to claim 1, wherein the spatial disturbance moment in step 2 comprises a solar pressure moment, a gravity gradient moment, a geomagnetic moment and an aerodynamic moment.
4. The method for actively self-stabilizing the attitude of a spacecraft in an orbit virtual field according to claim 3, wherein the specific process of the step 2 is as follows:
calculating sun light pressure moment T s ：
Moment T generated by solar pressure in spacecraft body coordinates s ＝(Τ sx Τ sy Τ sz ) The unified representation is:
wherein K is t K is the tangential light pressure coefficient n The normal light pressure coefficient is that sigma is the normal included angle between sunlight and a sailboard, eta s Is the angular distance between the zenith of the spacecraft and the sun, and ζ is the azimuth angle of deflection of the rotating shaft of the sailboard, ω 0 Is the angular velocity of spacecraft movement, r x 、r z The deviation amounts of the solar sailboard rotating shaft in the x and z axis directions of the spacecraft body system are respectively r y1 、r y2 The distances of the centers of the single sailboards along the y-axis direction of the spacecraft body are respectively;
calculating gravity gradient moment T g ：
The gravity gradient moment generated by the earth center gravitational field on the spacecraft is not only related to the gesture, but also closely related to the quality characteristic of the spacecraft, and under the condition that the spacecraft deviates from the orbit system by a small amount by zero gesture, the gravity gradient moment T g ＝(T gx T gy T gz ) Expressed as:
wherein,is the inertia tensor of the spacecraft, and is obtained by ground measurement, wherein the inertia tensor is phi, theta, phi and phi>Respectively yaw, roll and pitch angles of the spacecraft;
calculating the geomagnetic moment T m ：
Determining the geomagnetic induction intensity B of the position of the spacecraft according to the extrapolated spacecraft orbit position;
the spacecraft magnetic moment interacts with the earth's magnetic field and generates a moment, expressed as:
Τ m ＝Μ m ×Β (3)
wherein, M m Is the magnetic moment of the spacecraft, and
wherein b e For the total intensity of the earth magnetic moment, θ m Is the included angle between the radial direction of the spacecraft and the geomagnetic equator, e r Is the unit vector of the spacecraft position vector, z m Is a magnetic dipole vector unit vector;
calculating aerodynamic moment T a ：
Determining a velocity vector v of a windward side of a spacecraft relative to the atmosphere according to an extrapolated spacecraft orbit velocity s ；
For a spacecraft of a certain altitude range, the aerodynamic moment T a Is the dominant disturbing moment, assuming that the incident molecule loses all of its energy in the collision:
Τ a ＝ρ s ×F s (5)
wherein ρ is s Is the position vector of the pressure center of the windward side of the spacecraft relative to the mass center of the spacecraft, and
where ρ is the atmospheric density, S is the frontal area of the spacecraft, n is the unit vector perpendicular to the frontal area, and
v s ＝v o -w e ×r s (7)
wherein v is o Is the orbit speed of the spacecraft, r s Is the geocentric position vector of the spacecraft, and the geocentric position vector of the spacecraft is according to the outsideDetermination of orbit position of pushing spacecraft, w e An angular velocity vector that is the rotation of the earth;
sun light pressure moment T s Gravity gradient moment T g T of geomagnetic moment m And aerodynamic moment T a The sum is the space disturbance moment of the spacecraft, expressed as:
Τ＝Τ s +Τ g +Τ m +Τ a 。
5. the spacecraft attitude initiative self-stabilization control method of an orbit virtual field according to claim 1, wherein the specific process of the step 3 is as follows:
setting threshold |T max Judging whether the space interference moment T reaches the threshold value T max I, if I T I is greater than or equal to T max I, needing posture adjustment, and entering step 4; otherwise, the posture adjustment is not needed, and the step 1 is returned.
6. The method for actively self-stabilizing the attitude of a spacecraft in an orbit virtual field according to claim 1, wherein the specific process of the step 4 is as follows:
converting the space interference torque T under the orbit coordinate system into the space interference torque under the spacecraft body system;
under the condition that the spacecraft does not have a momentum control device, the attitude stability control is carried out by completely relying on small thrust, and the angular momentum H of the spacecraft system s Expressed as:
H s ＝Ιω (8)
wherein I represents an inertia tensor of the spacecraft in an orbit coordinate system, and w represents an angular velocity vector of the spacecraft;
the spacecraft attitude dynamics equation is:
taking a main inertia axis of the spacecraft as a spacecraft body coordinate system, wherein the inertia array is a diagonal array, and I=diag (I x I y I z ) The method comprises the steps of carrying out a first treatment on the surface of the M represents a moment vector stressed by the spacecraft under the spacecraft body coordinate system;
attitude of spacecraft in orbital coordinate system is determined by roll angleThe pitch angle θ and the yaw angle ψ are represented, and in the three-axis stability control, the spacecraft attitude angles are small, and the attitude matrix is represented as:
the rotational speed of the spacecraft orbit coordinate system in space is (0-omega) 0 0) The rotational speed of the spacecraft is expressed in the spacecraft body coordinate system as:
bringing the expression of ω into a spacecraft attitude dynamics equation to obtain an attitude stability dynamics equation based on the micro thrust as follows:
wherein c= (C x C y C z ) The moment is controlled for a small amount on the satellite; the small control moment C on the satellite is taken as the space disturbance moment under the spacecraft body system.
7. The method for actively self-stabilizing the attitude of a spacecraft in an orbit virtual field according to claim 1, wherein the specific process of the step 5 is as follows: and (3) driving the posture adjustment of the spacecraft small thruster according to the spacecraft posture adjustment strategy obtained in the step (4).
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