CN113968362B - Satellite in-orbit autonomous triaxial rapid maneuvering control method - Google Patents
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Abstract
An autonomous three-axis rapid maneuvering control method for satellite in orbit relates to the technical field of spacecraft attitude determination control, solves the contradiction problem of rapidity and stability when the existing satellite needs three-axis maneuvering, and comprises expected attitude calculation; three processes of three-axis attitude planning and rapid maneuvering control are performed, and after the satellite autonomous three-axis rapid maneuvering control is performed, the satellite is ensured to rapidly acquire data and simultaneously the imaging stability is ensured, so that high-quality image data are acquired for maritime search and rescue, wide area search and rescue after disaster, emergency geographic investigation and other emergency tasks. Therefore, the imaging capability of the low-orbit remote sensing satellite is improved, and the high timeliness of the image data acquired in orbit is ensured.
Description
Technical Field
The invention relates to the technical field of spacecraft attitude determination control, in particular to an on-orbit autonomous triaxial rapid maneuvering control method for satellites.
Background
When the satellite performs emergency tasks such as maritime search and rescue, wide area search and rescue after disaster, emergency geographic investigation and the like, the satellite is required to rapidly perform large-angle attitude maneuver, the moment and the angular momentum of an actuating mechanism of the micro satellite are limited, the angular velocity of the satellite is limited, and the satellite is required to be controlled under the constraint of the satellite to realize the rapidity of the satellite. In the existing research, the rapid maneuvering of the gesture is divided into two directions, one direction is gesture planning, the gesture planning is mainly carried out aiming at the large-angle maneuvering condition that only the side swinging maneuvering exists on the satellite, the rapid side swinging of the satellite is realized, and the method is not applicable when the satellite is subjected to triaxial maneuvering. The second direction is to design novel control methods such as a hierarchical saturated attitude control law based on Euler axis rotation, and the like, control stars to do attitude maneuver around the Euler axis, and do shortest path maneuver, but the contradiction between maneuver rapidity and stability still exists.
The remote sensing satellite runs in orbit for a long time under a sun-oriented triaxial stable mode, and before earth imaging is carried out, the satellite needs to carry out triaxial rapid maneuvering. According to the moment of inertia and angular momentum constraint of the satellite, a three-axis attitude planning-based mode is designed, the rotation axis of the satellite is kept unchanged in the maneuvering process, the satellite is guaranteed to maneuver rapidly, a corresponding control algorithm is designed to achieve maneuver rapidly, and meanwhile the contradiction between rapidity and stability is solved.
Disclosure of Invention
The invention provides an autonomous three-axis fast maneuvering control method for a satellite in order to realize autonomous fast imaging of the satellite in orbit and solve the contradiction problem of the rapidity and the stability when the satellite needs three-axis maneuvering.
An autonomous triaxial rapid maneuvering control method for satellite in orbit is realized by the following steps:
step one, calculating expected postures;
according to the initial attitude and the target attitude of the satellite, calculating an expected quaternion and calculating a rotation axis and a rotation angle corresponding to triaxial maneuver of the satellite; obtaining the desired quaternion q Q Desired rotation angle θ Q And a rotation axis direction e n ；
Step two, planning three-axis gestures;
under the condition of meeting the constraints of the satellite rotational inertia I, the reaction flywheel moment T and the angular momentum H, limiting and constraint setting are carried out on the triaxial angular acceleration and the angular velocity to obtain an angular acceleration limit value alpha LG And angular velocity limit omega LG The method comprises the steps of carrying out a first treatment on the surface of the And the eight-section type attitude planner which ensures that the direction of a rotating shaft is unchanged in the three-axis large-angle maneuvering process of the satellite and simultaneously designs the angular acceleration to be continuous, and the expected rotating angle theta is used Q Angular acceleration limit alpha LG Angular velocity limit omega LG As input, an angular acceleration generating function
The definition is as follows:
wherein Δt is A The rising time of the angular acceleration is limited as a set value, and the value is reasonably selected according to the dynamic performance of the satellite executing mechanism; Δt (delta t) B ，Δt C Is equal to the desired angle theta Q Related to the size of (2);
acquiring a real-time planning angle theta epsilon [0, theta ] through the gesture planner Q ]Planning angular velocity in real time
And real-time planning of angular acceleration +.>
And obtain the planning quaternion q G ＝[cosθ；sinθe n ]Three-axis planning angular velocity of the system>
And triaxial programmed angular acceleration->
Step three, quick maneuvering control;
step three, calculating the deviation angular velocity omega E Sum and deviation quaternion q E ；
The rotational quaternion of the real-time attitude of the satellite relative to the planned attitude, i.e. the deviation quaternion
The quaternion q under the satellite inertial system is a rotation quaternion of the satellite body coordinate system relative to the inertial system; initial quaternion q C A rotational quaternion of an initial attitude of the satellite relative to an inertial frame;
representation of deviation of real-time angular velocity of satellite from planned angular velocity under inertial system, namely deviationDifferential angular velocity
Wherein the track angular velocity omega gui The satellite angular velocity omega is the rotational angular velocity of the satellite body system relative to the inertial system; />
A rotational quaternion that is the initial attitude of the orbital relative to the satellite; r (q) E ) Is q E A corresponding rotation matrix; />
Is->
A corresponding rotation matrix;
step three, the obtained deviation angular velocity omega E Deviation quaternion q E Triaxial programming angular acceleration a G Inputting the satellite into a PD controller to realize the autonomous triaxial rapid maneuvering control of the satellite in orbit;
the PD controller is designed to:
u＝K q α G -K P q E -K d ω E
wherein K is q ,K P ,K d The feedforward control gain matrix, the proportional control increment matrix and the differential control gain matrix are respectively adopted.
The invention has the beneficial effects that:
the invention designs a three-axis attitude planning scheme aiming at the situation of three-axis large-angle maneuver of the satellite, shortens the required maneuver time, improves the maneuver performance of the satellite, designs a rapid maneuver algorithm and ensures the stability while realizing the rapidness of the satellite.
After the satellite in-orbit autonomous three-axis rapid maneuvering control, the satellite is ensured to rapidly acquire data and simultaneously ensure the imaging stability and acquire high-quality image data aiming at emergency tasks such as maritime search and rescue, post-disaster wide area search and rescue, emergency geographic investigation and the like. Therefore, the imaging capability of the low-orbit remote sensing satellite is improved, and the high timeliness of the image data acquired in orbit is ensured.
Drawings
FIG. 1 is a control schematic diagram of an autonomous triaxial maneuver control method for satellite in orbit according to the present invention;
FIG. 2 is a schematic rotation diagram between two coordinate systems;
FIG. 3 is a three-axis pose layout;
FIG. 4 is a graph of the effect of the gesture planner curve; wherein (a) is an angular acceleration generating function
Schematic, (b) is an angular velocity effect graph, and (c) is an angular effect graph;
FIG. 5 is a schematic diagram of satellite attitude conversion;
FIG. 6 is a simulation effect diagram of desired and planned rotation angles;
FIG. 7 is a PD-plan-control angle effect diagram; wherein, (a), (b) and (c) are respectively effect graphs of X-axis angle, Y-axis angle and Z-axis angle;
FIG. 8 is a PD-plan-control angular velocity effect diagram; wherein, (a), (b) and (c) are respectively effect graphs of X-axis angular velocity, Y-axis angular velocity and Z-axis angular velocity;
FIG. 9 is a graph of the effect of offset angle;
fig. 10 is a graph of the effect of offset angular velocity.
Detailed Description
Detailed description of the preferred embodimentthe present embodiment is described with reference to fig. 1 to 5, which are a method for controlling an autonomous triaxial maneuver of a satellite in orbit, and which are defined as follows:
definition of a relative coordinate System
In the present embodiment, the body coordinate system O is used b X b Y b Z b Orbital coordinate system O b X o Y o Z o And inertial system C e X eI Y eI Z eI Three coordinate systems.
(1) Body coordinate system O b X b Y b Z b : origin of coordinates O b Located at the center of mass of the satellite, the three-axis directions are related to the installation of the satellite body, and define X b The axis points to the direction of the sailboard, Z b The axis points to the camera direction, Y b Axis and X b Axis and Z b The axes form a right-hand rectangular coordinate system.
(2) Orbital coordinate system O b X o Y o Z o : the origin of coordinates is the mass center O of the satellite b The Y-axis points to the opposite direction of the angular velocity of the track, Z o The axis pointing to the earth center, X o Axis and Y o Axis and Z o The axes form a right-hand rectangular coordinate system (direction of flight), which is a reference for orientation to the ground.
(3) Inertial system C e X eI Y eI Z eI : the origin of the coordinate system is the earth centroid C e ，X eI The axis points to the flat spring point (1.2000, 1.12) Z eI Axis is directed to the north and south (jd= 2451545.0, 1/2000, 1/12), Y eI Axis and X eI Axis, Z eI The axes form a right-hand rectangular coordinate system, also known as the J2000 earth inertial coordinate system.
In this embodiment, the satellite attitude is described in the form of a quaternion, and the relevant properties are defined as follows:
the description mode of satellite attitude, quaternion represents:
wherein->
q 0 The scale, which is a quaternion, represents the rotation angle phi,
vector part representing rotation axis direction e n ＝[i；j；k]Satisfy i 2 +j 2 +k 2 ＝1。
The four parameters satisfy the constraint equation:
vector product rule:
inverse of quaternion:
quaternion multiplication:
the specific implementation steps of the embodiment are as follows:
step one: calculating expected postures;
and calculating an expected quaternion and a rotation axis and a rotation angle corresponding to triaxial maneuver of the satellite according to the initial attitude and the target attitude of the satellite.
From the definition of the quaternion, the transformation of the pose of the initial coordinate system Oxyz relative to the target coordinate system Ox ' y ' z ' is expressed as
As in fig. 2.
The desired quaternion of the target pose of the satellite relative to the initial pose is
Wherein the initial quaternion q C A rotational quaternion of an initial attitude of the satellite relative to an inertial frame; target quaternion q F A rotation quaternion of a target attitude of the satellite relative to an inertial system;
from the definition of quaternions, q Q The reference part q of (1) Q0 The rotation angle Φ, Φ=2 arccoss (q Q0 ). At the same time by
Can get->
When Φ=0, the corresponding quaternion is q Q ＝[1；0；0；0]The target pose coincides with the initial pose.
Step two: planning three-axis gestures;
the maneuvering process of the satellite is planned in real time according to the performance constraint of the satellite, and a one-dimensional rotation angle is generated through the attitude planner, so that the maneuvering capability of the satellite can be improved.
Under the constraint of satisfying the satellite rotational inertia I, the reaction flywheel moment T and the angular momentum H, in order to realize the initial quaternion q C To the target quaternion q F I.e. the rotation of the whole desired quaternion q Q The rotation axis e is ensured during the maneuvering process n When the gesture planning is carried out, the three-axis angular acceleration and the three-axis angular velocity are required to be limited and restrained.
The input and output of the gesture planner are one-dimensional, the input is a desired rotation angle, an angular acceleration limit value, an angular velocity limit value, and the output is a real-time angle, a real-time angular velocity and a real-time angular acceleration. The three-axis gesture layout is shown in fig. 3.
Angular acceleration limit alpha LG Is calculated as follows:
α LG ＝||α LG || 2 ；
wherein the moment of inertia of the satellite
Reaction flywheel triaxial moment T= [ T ] x ；T y ；T z ]N·m，M max ＝10 20 For a set larger number, the angular acceleration limit α is planned LG ＝[α LGx ；α LGy ；α LGz ]°/s 2 ，[·] min For the minimum value calculation to be performed, I.I 2 Is the modulus of the vector.
Inputting a triaxial angular velocity limit omega Lim ＝[ω Limx ；ω Limy ；ω Limz ]DEG/s; angular velocity limit omega LG Is calculated as follows:
ω LG ＝||ω LG || 2 ；
wherein the angular momentum of the reaction flywheel H= [ H ] x ；H y ；H z ]N.m.s, three-axis planning angular velocity limit omega LG ＝[ω LGx ；ω LGy ；ω LGz ]°/s。
To avoid the abrupt change of angular acceleration, the stable change of flywheel moment is realized, and the rapidity is considered at the same time, so that the desired rotation angle theta Q Angular acceleration limit alpha LG Angular velocity limit omega LG As input, an eight-section gesture planner with continuous angular acceleration is designed, and the angular acceleration generates a function
The definition is as follows:
wherein Δt is A The rise time of the angular acceleration is defined for the set value, and can be reasonably selected according to the dynamic performance of the satellite actuating mechanism. Δt (delta t) B ，Δt C Is equal to the desired angle theta Q Related to the size, Δt B ，Δt C ，
The specific calculation process is as follows:
(1) When the rotation angle is expected
At the time of planning angular acceleration +.>
Up to a maximum of a LG Planning angular velocity +.>
Up to a maximum of ω LG 。/>
(2) When (when)
At the time of planning angular acceleration +.>
Up to a maximum of a LG Planning angular velocity +.>
The maximum value of not reaching omega LG 。Δt B By means of the unitary quadratic equation>
Solving to obtain->
(3) When (when)
At the time of planning angular acceleration +.>
The maximum value of not reaching alpha LG Planning angular velocity +.>
The maximum value of not reaching omega LG . Angle θ of maneuver Q Corresponding time is 4 delta t A 。Δt B ＝0，Δt C ＝0，/>
From the maneuvering angle theta Q =60°, angular acceleration limit α LG ＝0.1161°/s 2 Angular velocity limit omega LG ＝1.5°/s，Δt A =5s as input, generated by the above-mentioned pose planner
Is abbreviated as->
The planned angular velocity and angle are shown in fig. 4.
The total angular acceleration is 8 sections, and the system is divided into a rising section 2 sections, a stable section 4 sections and a falling section 2 sections. Wherein, the rising section 1, the stable section 1 and the descending section 1 of the angular acceleration correspond to the rising section of the angular velocity; the plateau 2 of angular acceleration corresponds to a plateau of angular velocity; an ascending section 2, a stable section 3 and a descending section 2 of the angular acceleration, which correspond to the descending section of the angular velocity; the value of the plateau 4 of angular acceleration is zero, the value of the corresponding angular velocity is also zero, and the angle value reaches the desired angle.
Real-time planning angle θ e [0, θ ] produced by the gesture planner Q ]Angular velocity of
And angular acceleration->
Solving to obtain a planning quaternion q G ＝[cosθ；sinθe n ]Three-axis planning angular velocity of body system>
Triaxial planning angular acceleration +.>
Step three: fast maneuver control
The dynamics and kinematics equations of rigid satellites are described as:
wherein u is control moment, S (&) is an antisymmetric matrix, and ++>
In the maneuvering control of the satellite, a corresponding posture and angular velocity conversion chart under a plurality of coordinate systems is shown in fig. 5; deviation angular velocity omega E Sum and deviation quaternion q E The calculation is as follows:
rotational quaternion of orbital system relative to initial attitude of satellite
Wherein the orbit quaternion q gui Is the relative inertia of the track systemA rotation quaternion of the system;
the rotational quaternion of the real-time attitude of the satellite relative to the planned attitude, i.e. the deviation quaternion
The quaternion q under the satellite inertial system is a rotation quaternion of the satellite body coordinate system relative to the inertial system;
representation of deviation of real-time angular velocity of satellite from planned angular velocity under inertial system, i.e. deviation angular velocity
Wherein the track angular velocity omega gui The satellite angular velocity ω is the rotational angular velocity of the satellite body system relative to the inertial system.
R(q E ) Is q E The corresponding rotation matrix is used to determine the rotation of the rotor,
is->
A corresponding rotation matrix.
To further increase the rapidity of satellite maneuver, a feedforward design is added on the basis of PD control, and the controller is designed to:
u＝K q α G -K P q E -K d ω E
wherein K is q ,K P ,K d Respectively a feedforward control gain matrix, a proportional control increment matrix and a differential control gain matrix, K q ＝K q I,K P ＝K P I,K d ＝K d I,K q ,K P ,K d Is a gain matrix coefficient greater than 0.
A second embodiment is described with reference to fig. 6 to 10, in which the method for controlling satellite autonomous triaxial fast maneuver according to the first embodiment is adoptedThe example is verified, and the result of verification is compared with the traditional PD control scheme without path planning, and the satellite and control parameters of the embodiment are selected as follows: moment of inertia of satellite
Flywheel angular momentum h= [0.01;0.01;0.01]N.m.s; flywheel torque t= [0.003;0.003;0.003]N.m; input angular velocity limit omega Lim ＝[1.2；1.3；1.1]DEG/s; feedforward control gain matrix coefficient K q =0.75; proportional control gain matrix coefficient K p =1.55; differential control gain matrix coefficient K d =1.5; the initial attitude quaternion of the satellite is q C ＝[1；0；0；0]The method comprises the steps of carrying out a first treatment on the surface of the The target is imaged on the ground and is superposed with the track system, and the target posture is q F ＝q gui Initial angular velocity omega C ＝[0；0；0]°/s。
In the PD control scheme, u= -K P1 q ed -K D1 ω ed ，
ω ed ＝ω-R(q E )ω gui ，K P1 ＝K P1 I,K d1 ＝K d1 I, proportional control gain matrix coefficient K p1 =0.12; differential control gain matrix coefficient K d1 ＝0.58。
The desired rotation angle required from the initial pose to the target pose is 100.74 °, and the three-axis pose planning angle curve is shown in fig. 6. The three-axis attitude angle of the PD control scheme, the three-axis attitude angle corresponding to the three-axis attitude planning angle of the present embodiment, and the three-axis attitude angle of the planned feedforward control scheme of the present embodiment (euler angles obtained by rotating quaternion q in ZYX order) are shown in fig. 7, and correspond to the PD control, the desired planning, and the planned control labels, respectively. The corresponding angular velocity is shown in fig. 8. The deviation angle and the deviation angular velocity of the present embodiment are shown in fig. 9 and 10. As can be seen from fig. 7 and 8, the method according to the present embodiment requires a shorter time for rotation by the same angle.
Claims (4)
1. An autonomous triaxial rapid maneuvering control method for satellite in orbit is characterized by comprising the following steps: the method comprises the following steps:
step one, calculating expected postures;
according to the initial attitude and the target attitude of the satellite, calculating an expected quaternion and calculating a rotation axis and a rotation angle corresponding to triaxial maneuver of the satellite; obtaining the desired quaternion q Q Desired rotation angle θ Q And a rotation axis direction e n ；
Step two, planning three-axis gestures;
under the condition of meeting the constraints of the satellite rotational inertia I, the reaction flywheel moment T and the angular momentum H, limiting and constraint setting are carried out on the triaxial angular acceleration and the angular velocity to obtain an angular acceleration limit value alpha LG And angular velocity limit omega LG The method comprises the steps of carrying out a first treatment on the surface of the And the eight-section type attitude planner which ensures that the direction of a rotating shaft is unchanged in the three-axis large-angle maneuvering process of the satellite and simultaneously designs the angular acceleration to be continuous, and the expected rotating angle theta is used Q Angular acceleration limit alpha LG Angular velocity limit omega LG As input, an angular acceleration generating function
The definition is as follows:
wherein Δt is A The rising time of the angular acceleration is limited as a set value, and the value is reasonably selected according to the dynamic performance of the satellite executing mechanism; Δt (delta t) B ，Δt C Is equal to the desired angle theta Q Related to the size of (2);
acquiring a real-time planning angle theta epsilon [0, theta ] through the gesture planner Q ]Planning angular velocity in real time
And real-time planning of angular acceleration +.>
And obtain the planning quaternion q G ＝[cosθ；sinθe n ]Three-axis planning angular velocity of the system>
And triaxial programmed angular acceleration->
Step three, quick maneuvering control;
step three, calculating the deviation angular velocity omega E Sum and deviation quaternion q E ；
The rotational quaternion of the real-time attitude of the satellite relative to the planned attitude, i.e. the deviation quaternion
The quaternion q under the satellite inertial system is a rotation quaternion of the satellite body coordinate system relative to the inertial system; initial quaternion q C A rotational quaternion of an initial attitude of the satellite relative to an inertial frame;
representation of deviation of real-time angular velocity of satellite from planned angular velocity under inertial system, i.e. deviation angular velocity
Wherein the track angular velocity omega gui The satellite angular velocity omega is the rotational angular velocity of the satellite body system relative to the inertial system; />
A rotational quaternion that is the initial attitude of the orbital relative to the satellite; r (q) E ) Is q E A corresponding rotation matrix; />
Is->
A corresponding rotation matrix;
step three, the obtained deviation angular velocity omega E Deviation quaternion q E Triaxial programming angular acceleration a G Inputting the satellite into a PD controller to realize the autonomous triaxial rapid maneuvering control of the satellite in orbit;
the PD controller is designed to:
u＝K q α G -K P q E -K d ω E
wherein K is q ,K P ,K d The feedforward control gain matrix, the proportional control increment matrix and the differential control gain matrix are respectively adopted.
2. The method for autonomous triaxial fast maneuver control of a satellite according to claim 1, characterized in that: the specific process of the first step is as follows:
from the definition of quaternion, the transformation of the pose of the initial coordinate system Oxyz relative to the target coordinate system Ox ' y ' z ' is expressed as
Phi is the rotation angle;
the desired quaternion of the target pose of the satellite relative to the initial pose is
Wherein the initial quaternion q C A rotational quaternion of an initial attitude of the satellite relative to an inertial frame; target quaternion q F A rotation quaternion of a target attitude of the satellite relative to an inertial system;
from the definition of quaternions, q Q The reference part q of (1) Q0 The rotation angle Φ, Φ=2 arccoss (q Q0 ) The method comprises the steps of carrying out a first treatment on the surface of the At the same time by
Obtain the rotation axis +.>
When Φ=0, the corresponding quaternion is q Q ＝[1；0；0；0]The target pose coincides with the initial pose.
3. The method for autonomous triaxial fast maneuver control of a satellite according to claim 1, characterized in that: in the second step, the angular acceleration limit value alpha LG And angular velocity limit omega LG The calculation process of (1) is as follows:
the angular acceleration limit alpha LG Is calculated as follows:
α LG ＝||α LG || 2 ；
in the moment of inertia of the satellite
Reaction flywheel triaxial moment T= [ T ] x ；T y ；T z ]N·m，M max ＝10 20 For a set larger number, the angular acceleration limit α is planned LG ＝[α LGx ；α LGy ；α LGz ]°/s 2 ，[·] min For the minimum value calculation to be performed, I.I 2 Modulo the vector;
inputting a triaxial angular velocity limit omega Lim ＝[ω Limx ；ω Limy ；ω Limz ]DEG/s; angular velocity limit omega LG Is calculated as follows:
ω LG ＝||ω LG || 2 ；
wherein the angular momentum of the reaction flywheel H= [ H ] x ；H y ；H z ]N.m.s, three-axis planning angular velocity limit omega LG ＝[ω LGx ；ω LGy ；ω LGz ]°/s。
4. The method for autonomous triaxial fast maneuver control of a satellite according to claim 1, characterized in that: in the second step, the Δt B ，Δt C ，
The specific calculation process is as follows:
(1) When the rotation angle is expected
At the time of planning angular acceleration +.>
Up to a maximum of a LG Planning angular velocity +.>
Up to a maximum of ω LG ，/>
(2) When (when)
At the time of planning angular acceleration +.>
Up to a maximum of a LG Planning angular velocity +.>
The maximum value of not reaching omega LG ；Δt B By means of the unitary quadratic equation>
Solving to obtain Deltat C ＝0，/>
(3) When (when)
At the time of planning angular acceleration +.>
The maximum value of not reaching alpha LG Planning angular velocity +.>
The maximum value of not reaching omega LG Angle θ of maneuver Q Corresponding time is 4 delta t A ，Δt B ＝0，Δt C ＝0，/>
From the maneuvering angle theta Q =60°, angular acceleration limit α LG ＝0.1161°/s 2 Angular velocity limit omega LG ＝1.5°/s，Δt A =5s as input, generated by the above-mentioned pose planner
Is abbreviated as->
Angular velocity and angle are planned.
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Denomination of invention: A Method for Autonomous Three Axis Rapid Maneuver Control of Satellites in Orbit Granted publication date: 20230707 Pledgee: Jilin Chuncheng Rural Commercial Bank Co.,Ltd. Pledgor: Changguang Satellite Technology Co.,Ltd. Registration number: Y2024220000054 |