CN111483617B - Illumination optimal attitude maneuver path planning method suitable for Mars detection - Google Patents
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Abstract
An illumination optimal attitude maneuver path planning method suitable for Mars detection belongs to the technical field of Mars detection attitude control and comprises the following steps: s1, according to the long-term steady state of the detector, the detector points to the sun along the X axis, the sun-sun reference of the-Z constraint to the ground and the constraint that the solar wing can be driven in one dimension according to the Y axis, the three-dimensional maneuvering Euler axis is decomposed, the maximum time of the sun vector in the maneuvering process is ensured to be in the XOZ plane, the sun tracking with the maximum performance is ensured under the one-dimensional driving support of the solar wing, and the optimal illumination is realized; and S2, planning the attitude maneuver path in a segmented manner according to the projection decomposition scheme in the S1, and combining the maximum capacity of the flywheel to realize rapid attitude maneuver. The invention aims at ensuring that the normal line of the solar wing points to the sun to the maximum extent in the attitude maneuver process, comprehensively considers the process energy consumption, has simple and feasible method, and provides an on-orbit implementation plan of effective energy guarantee for Mars detection.
Description
Technical Field
The invention relates to an illumination optimal attitude maneuver path planning method suitable for Mars detection, and belongs to the technical field of Mars detection attitude control.
Background
China will launch Mars detectors in 2020, and the conventional three-step flow of 'winding', 'falling' and 'patrolling' is realized at one time. The distance between the Mars and the sun is about 1.523AU, the illumination intensity of the sun is reduced due to the fact that the Mars are far away from the sun, and the illumination intensity of the Mars in one period around the sun is about 0.3-0.5 compared with the illumination intensity of the earth. The maximum illuminated area of the solar wing is about 14m due to the constraint of the emission quality and the mechanical size of the Mars detector2The energy balance margin is larger in the ground fire transfer process, and the energy balance is more tense in the surrounding process.
In addition, in the process, the functions of complex attitude maneuver, track control and the like need to be completed on the fire-surrounding track surrounding device, the solar wing cannot be guaranteed to be sunned to the maximum extent, and long-time charging operation is needed to guarantee energy balance after one action, so that an optimal illumination method for the attitude motion process is urgently needed in the Mars surrounding stage to guarantee smooth completion of the Mars detection task.
Disclosure of Invention
The technical problem to be solved by the invention is as follows: the method for planning the maneuver path of the optimal illumination posture suitable for Mars detection overcomes the defects of the prior art, and comprises the following steps: s1, according to the long-term steady state of the detector, the detector points to the sun along the X axis, the sun-sun reference of the-Z constraint to the ground and the constraint that the solar wing can be driven in one dimension according to the Y axis, the three-dimensional maneuvering Euler axis is decomposed, the maximum time of the sun vector in the maneuvering process is ensured to be in the XOZ plane, the sun tracking with the maximum performance is ensured under the one-dimensional driving support of the solar wing, and the optimal illumination is realized; and S2, planning the attitude maneuver path in a segmented manner according to the projection decomposition scheme in the S1, and combining the maximum capacity of the flywheel to realize rapid attitude maneuver. The invention aims at ensuring that the normal line of the solar wing points to the sun to the maximum extent in the attitude maneuver process, comprehensively considers the process energy consumption, has simple and feasible method, and provides an on-orbit implementation plan of effective energy guarantee for Mars detection.
The purpose of the invention is realized by the following technical scheme:
an illumination optimal attitude maneuver path planning method suitable for Mars detection comprises the following steps:
s1, winding the detector around the Euler axis
The primary attitude maneuver is decomposed into two attitude maneuvers; obtaining attitude quaternion, Euler angle and Euler axis of the two attitude maneuvers;
s2, obtaining parameters of the path planning of the two attitude maneuvers by adopting an attitude maneuver path planning algorithm according to the Euler angles and the Euler axes of the two attitude maneuvers in S1;
s3, calculating the path parameters of the single attitude maneuver path planning in each control period by using the parameters of the two attitude maneuver path planning in the S1;
and S4, determining a control torque in the maneuvering process of the detector according to the path parameters of the single-attitude maneuvering path planning in each control period, wherein the control torque is used for controlling the executing structure of the detector to perform attitude maneuvering.
Preferably, in step S4, the method for planning the illumination-optimized attitude maneuver path suitable for Mars detection calculates a closed-loop control error quaternion and a feedforward control moment according to the path parameter of the single-attitude maneuver path planning in each control period; and determining the control torque of the detector in the maneuvering process by utilizing the closed-loop control error quaternion and the feedforward control torque.
Preferably, the control moment in the maneuvering process of the detector is calculated and determined by utilizing a closed-loop control error quaternion and a feedforward control moment and adopting an equal-proportion amplitude limiting control algorithm.
Preferably, in the method for planning the maneuver path of the optimal illumination posture suitable for Mars detection, in step S1, the detector is wound around the euler axis in space
The method for decomposing the primary attitude maneuver into the secondary attitude maneuvers comprises the following steps:
s11, calculating the Euler axis of the detector
A primary maneuver quaternion to perform a primary attitude maneuver; calculating the Euler angle of the primary attitude maneuver by utilizing the corresponding relation between the quaternion of the primary maneuver and the Euler angle of the attitude;
and S12, calculating and decomposing the attitude quaternion, the Euler angle and the Euler axis into two attitude maneuvers by using the Euler angle in S11.
Preferably, in S11, the euler angle of the primary attitude maneuver is calculated according to 123 rotation sequence by using the correspondence between the primary maneuver quaternion and the attitude euler angle.
Preferably, in the lighting optimal attitude maneuver path planning method suitable for Mars detection, the parameters of the two-time attitude maneuver path planning in S2 include the sum of the maximum angular acceleration around the space Euler axis, the maximum angular velocity around the space Euler axis, the maneuver acceleration time, the maneuver uniform speed time and the deceleration time.
Preferably, in the method for planning the illumination-optimized attitude maneuver path suitable for Mars detection, the path parameters for planning the single attitude maneuver path in each control period in S3 include a rotation angle around the Euler axis of the secondary attitude maneuver at the relative maneuver starting time, a rotation angular velocity around the Euler axis of the secondary attitude maneuver at the relative maneuver starting time, and a maneuver quaternion at the relative maneuver starting time.
Preferably, in the method for planning the maneuver path of the optimal illumination posture suitable for Mars detection, the control moment in the maneuver process of the detector is as follows:
Tctrl＝Kpqerrv+Ki∫qerrv+Kdωerr+Tc
in the formula, TctrlTo control the moment, KpAs a proportional coefficient of the controller, KiAs an integral coefficient of the controller, KdAs a differential coefficient of the controller, qerrvFor closed-loop control of error quaternion, omegaerrFor closed-loop control of error angular velocity, TcThe control torque is fed forward to the controller.
Compared with the prior art, the invention has the following beneficial effects:
(1) the method has the characteristic that the solar wing is optimally illuminated in the maneuvering process under the condition that the solar wing is limited to drive, and is based on the sectional maneuvering in the attitude maneuvering process, the first section of maneuvering guarantees energy, and the second section of maneuvering guarantees are rapid;
(2) the method has the capability of quick maneuvering; the maneuvering process adopts path planning, and compared with the traditional controller-based amplitude limiting control and open-loop control means, the method is based on the maximum capacity constraint of an actuating mechanism, and the fastest maneuvering to the target position is realized;
(3) the method has the capability of avoiding the occurrence of angle singularity in the conventional large-angle maneuver; in the maneuvering process, a quaternion vector part is adopted to replace the traditional Euler angle as a control quantity, so that large-angle maneuvering singularity in the control process can be effectively avoided;
(4) the method has the stability in the whole maneuvering process and the stability after the maneuvering is in place quickly, the maneuvering process adopts a closed-loop feedback and feedforward control strategy, a planned path is quickly tracked in the maneuvering process, and the influence on the steady-state process after the maneuvering is in place is reduced through the design of the trapezoidal path deceleration process.
Drawings
Fig. 1 is an exploded schematic view of a posture maneuver rotating shaft;
FIG. 2 is an illustration of the rotation process versus the daily vector after decomposition;
FIG. 3 is a flow chart of the steps of the method of the present invention.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention more apparent, embodiments of the present invention will be described in detail with reference to the accompanying drawings.
An illumination optimal attitude maneuver path planning method suitable for Mars detection comprises the following steps:
s1, winding the detector around the Euler axis
The primary attitude maneuver is decomposed into two attitude maneuvers; obtaining attitude quaternion, Euler angle and Euler axis of the two attitude maneuvers;
s2, obtaining parameters of the path planning of the two attitude maneuvers by adopting an attitude maneuver path planning algorithm according to the Euler angles and the Euler axes of the two attitude maneuvers in S1; parameters of the two-time attitude maneuver path planning comprise the maximum angular acceleration around the space Euler axis, the maximum angular velocity around the space Euler axis, the maneuver acceleration time, the maneuver constant speed time and the deceleration time;
s3, calculating the path parameters of the single attitude maneuver path planning in each control period by using the parameters of the two attitude maneuver path planning in the S1; the path parameters of the single attitude maneuver path planning in each control period comprise a rotation angle around the Euler axis of the single attitude maneuver at the relative maneuver starting time, a rotation angular speed around the Euler axis of the single attitude maneuver at the relative maneuver starting time and a maneuver quaternion at the relative maneuver starting time;
s4, calculating a closed-loop control error quaternion and a feedforward control moment according to the path parameters planned by the single attitude maneuver path in each control period; and calculating and determining the control torque in the detector maneuvering process by using a closed-loop control error quaternion and a feedforward control torque and adopting an equal-proportion amplitude limiting control algorithm, wherein the control torque is used for controlling an execution structure of the detector to perform attitude maneuvering. The control torque in the maneuvering process of the detector is as follows:
Tctrl＝Kpqerrv+Ki∫qerrv+Kdωerr+Tc
in the formula, TctrlTo control the moment, KpAs a proportional coefficient of the controller, KiAs an integral coefficient of the controller, KdAs a differential coefficient of the controller, qerrvFor closed-loop control of error quaternion, omegaerrFor closed-loop control of error angular velocity, TcThe control torque is fed forward to the controller.
S1, the detector is wound around the Euler axis
The method for decomposing the primary attitude maneuver into the secondary attitude maneuvers comprises the following steps:
s11, calculating the Euler axis of the detector
A primary maneuver quaternion to perform a primary attitude maneuver; calculating the Euler angle of the primary attitude maneuver according to 123 rotation sequence by using the corresponding relation between the quaternion of the primary maneuver and the Euler angle of the attitude;
and S12, calculating and decomposing the attitude quaternion, the Euler angle and the Euler axis into two attitude maneuvers by using the Euler angle in S11.
Example (b):
the method is based on the decomposition of one-time attitude maneuver, combines solar wings and constraint on daily reference, and realizes the maximum illumination in the maneuvering process by means of path planning in the two-section maneuvering and maneuvering process and closed-loop feedback + feedforward torque in the maneuvering process, and is fast in maneuvering and stable in process. The method for planning the maneuver path of the illumination optimal posture comprises the following steps:
s1, according to structural layout constraint of the Mars probe: the +/-Y solar wing is arranged on a +/-Y-direction side plate of the detector body and has one-dimensional driving capability around the Y axis of the detector; the ground directional antenna is arranged in the-Z direction of the detector body and has two-dimensional driving capability around the X axis and the Y axis of the detector. Defining a long-term cruising attitude reference of a detector: the detector + X axis points to the sun, and the-Z axis is constrained in the plane formed by the earth, the detector and the sun and is positioned on one side of the earth. Under the attitude reference, the sun wing can be ensured to face the sun and the directional antenna points to the earth under the long-term stable state of the detector based on the one-dimensional driving of the sun wing and the two-dimensional driving capability of the directional antenna. Due to the requirements of track control and scientific detection, the Mars detector needs to perform autonomous attitude maneuver to realize specific target pointing on the basis of the current steady-state sun-facing reference according to the constraint of track control attitude and scientific detection attitude. The conventional attitude maneuver adopts a primary attitude maneuver around a space Euler axis, the shortest maneuvering path can be guaranteed in the process, but the sun wing cannot guarantee the sun tracking and orientation with the longest duration in the maneuvering process.
Based on the one-dimensional drive constraint of the sun wing around the Y axis (the sun wing can only be driven one-dimensionally around the Y axis) and the definition of the steady state (the steady state condition of the detector means that the detector is in the cruise section or the fire surrounding section) of the sun reference, the sun wing is wound around the Euler axis of the space
The primary attitude maneuver carries out the decomposition under the system of the detector, and the decomposition is into a winding vector
And a winding vector
Two attitude maneuvers. Vector
Sum vector
Split into euler axes after two attitude maneuvers, and then go to S2.
S1.1 calculating primary maneuvering target quaternion and corresponding Euler angle
Wherein q is a primary maneuvering quaternion, qtIs a target quaternion, q, in the J2000 inertial systembiIs the current quaternion of the maneuvering time of the J2000 inertial system.
And calculating the Euler angle of the q according to 123 rotation sequences according to the corresponding relation between the attitude quaternion and the attitude Euler angle.
θ＝arcsin[2(qbo0_nqbo2_n+qbo1_nqbo3_n)]
Wherein theta is a pitch angle,
roll angle, psi yaw angle, qbo0_n、qbo1_n、qbo2_n、qbo3_nAre in the form of q components.
S1.2, calculating and decomposing the attitude quaternion into an attitude quaternion after two attitude maneuvers, and an Euler angle and an Euler axis of each attitude maneuver.
In the formula, q1Attitude quaternion, q, for the first attitude maneuver2Is the attitude quaternion for the second attitude maneuver.
Further obtaining:
Φ1＝2arccos(q10)
Φ2＝2arccos(q20)
in the formula, q10、q11、q12、q13Is q1Component form of (a); q. q.s20、q21、q22、q23Is q2In the form of components. Phi1Euler angle, v, for first attitude maneuver1Euler axis for first attitude maneuver, phi2Euler angle, v, for second attitude maneuver2The Euler axis for the second attitude maneuver. As shown in fig. 1 and 2.
And S2, calculating path planning parameters of the two attitude maneuvers by adopting an attitude maneuver path planning algorithm according to the Euler axes and Euler angles of the two attitude maneuvers calculated in the S1, calculating a control target attitude quaternion of each period in the maneuvering process, and realizing the tracking control of the maneuvering path by adopting a feedforward closed-loop quaternion tracking control algorithm.
And S2.1, calculating path planning parameters of the two attitude maneuvers by adopting an attitude maneuver path planning algorithm. The calculation method of the path planning parameters of the first attitude maneuver comprises the following steps:
the path planning parameters for the second attitude maneuver are as follows:
wherein: t isJwThe maximum acting torque of the actuating mechanism reacting with the flywheel; hJwIs the maximum angular momentum of the reaction flywheel; c is an installation matrix of a flywheel configured on the device, and the default is a unit matrix; a ismax1For winding a spatial Euler shaft
Maximum angular acceleration of (a); omegamax1 is a winding space Euler shaft
The maximum angular velocity of; t is tjs1Is the first maneuver acceleration time; t is tys1Is the sum of the first maneuvering uniform speed time and the first deceleration time. a ismax2For winding a spatial Euler shaft
Maximum angular acceleration of (a); omegamax2For winding a spatial Euler shaft
The maximum angular velocity of; t is tjs2Is a second maneuver acceleration time; t is tys2Is the sum of the second maneuvering uniform speed time and the second deceleration time.
S2.2, calculating the path parameters of the path planning of the single attitude maneuver.
The path parameters of the path planning of the first attitude maneuver are as follows:
wherein: t is the timing of the relative maneuver starting time; phitempThe rotation angle of the Euler axis around the attitude maneuver at the relative maneuver starting time; omegatempThe rotation angular velocity around the Euler axis of the attitude maneuver relative to the maneuver start time; q. q.stempThe quaternion is maneuvered relative to the maneuver start time.
The path parameter of the path planning of the second attitude maneuver adopts the calculation method of the path parameter of the path planning of the first attitude maneuver to calculate the a in the formulamax1Is replaced by amax2，tjs1Is replaced by tjs2，tys1Is replaced by tys2，
Is replaced by
And S2.3, calculating a closed-loop control error quaternion and a feedforward control moment.
The method for calculating the closed-loop control error quaternion and the feedforward control moment by the first attitude maneuver comprises the following steps:
ωerr＝C(qerr)ωtemp-ωb
Tc＝I*amax1v1
wherein q iserrIs a closed-loop control error quaternion, wherein the vector part is brought into the controller; omegaerrError angular velocity is closed-loop controlled; t iscFeedforward control moment for the controller; i is an inertia matrix of the detector; qb is a quaternion measured relative to the maneuver start time; omegabThe angular velocity of the detector measured by the gyro sensor.
The method for calculating the closed-loop control error quaternion and the feedforward control moment by the second attitude maneuver refers to the method for calculating the closed-loop control error quaternion and the feedforward control moment by the first attitude maneuver, and uses a in the formulamax1Is replaced by amax2，
Is replaced by
And S2.4, calculating the control torque of the maneuvering process by adopting an equal-proportion amplitude limiting control algorithm.
Tctrl＝Kpqerrv+Ki∫qerrv+Kdωerr+Tc
In the formula, TctrlTo control the moment, KpAs a proportional coefficient of the controller, KiAs an integral coefficient of the controller, KdAs a differential coefficient of the controller, qerrvFor closed-loop control of error quaternion, omegaerrFor closed-loop control of error angular velocity, TcThe control torque is fed forward to the controller.
FIG. 3 is a flowchart illustrating steps of an embodiment of the method of the present invention.
Those skilled in the art will appreciate that those matters not described in detail in the present specification are well known in the art.
Although the present invention has been described with reference to the preferred embodiments, it is not intended to limit the present invention, and those skilled in the art can make variations and modifications of the present invention without departing from the spirit and scope of the present invention by using the methods and technical contents disclosed above.
Claims (5)
1. An illumination optimal attitude maneuver path planning method suitable for Mars detection is characterized by comprising the following steps:
s1, decomposing one attitude maneuver of the detector around the space Euler axis into two attitude maneuvers; obtaining attitude quaternion, Euler angle and Euler axis of the two attitude maneuvers;
s2, obtaining parameters of the path planning of the two attitude maneuvers by adopting an attitude maneuver path planning algorithm according to the Euler angles and the Euler axes of the two attitude maneuvers in S1;
s3, calculating the path parameters of the single attitude maneuver path planning in each control period by using the parameters of the two attitude maneuver path planning in the S2;
s4, calculating a closed-loop control error quaternion and a feedforward control moment according to the path parameters planned by the single attitude maneuver path in each control period; calculating and determining a control moment in the maneuvering process of the detector by using a closed-loop control error quaternion and a feedforward control moment and adopting an equal-proportion amplitude limiting control algorithm, wherein the control moment is used for controlling an execution structure of the detector to perform attitude maneuvering;
parameters of the two-time attitude maneuver path planning in the S2 comprise the sum of maneuver uniform speed time and deceleration time, the maximum angular acceleration around the space Euler axis, the maximum angular velocity around the space Euler axis and maneuver acceleration time.
2. The method for planning the optimal illumination attitude maneuver path suitable for Mars detection according to claim 1, wherein in S1, the method for decomposing the one attitude maneuver of the detector around the spatial Euler axis into two attitude maneuvers comprises:
s11, calculating a primary maneuvering quaternion of the primary attitude maneuvering of the detector around the space Euler axis; calculating the Euler angle of the primary attitude maneuver by utilizing the corresponding relation between the quaternion of the primary maneuver and the Euler angle of the attitude;
and S12, calculating and decomposing the attitude quaternion, the Euler angle and the Euler axis into two attitude maneuvers by using the Euler angle in S11.
3. The method for planning an optimal illumination attitude maneuver path suitable for Mars detection according to claim 2, wherein in S11, the Euler angle of the primary attitude maneuver is calculated according to 123 turns by using the corresponding relationship between the quaternion of the primary maneuver and the Euler angle of the attitude.
4. The method for planning the illumination-optimized attitude maneuver path suitable for Mars detection according to any one of claims 1 to 3, wherein the path parameters for planning the single attitude maneuver path in each control period in S3 comprise a rotation angle around the Euler axis of the secondary attitude maneuver at the relative maneuver start time, a rotation angular velocity around the Euler axis of the secondary attitude maneuver at the relative maneuver start time, and a quaternion maneuvered at the relative maneuver start time.
5. The method for planning the maneuver path of the Mars optimal illumination posture suitable for Mars detection according to any one of claims 1 to 3, wherein the control torque in the maneuver process of the detector is as follows:
Tctrl＝Kpqerrv+Ki∫qerrv+Kdωerr+Tc
in the formula, TctrlTo control the moment, KpAs a proportional coefficient of the controller, KiAs an integral coefficient of the controller, KdAs a differential coefficient of the controller, qerrvFor closed-loop control of error quaternion, omegaerrFor closed-loop control of error angular velocity, TcThe control torque is fed forward to the controller.
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Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH04231295A (en) * | 1990-05-14 | 1992-08-20 | Hughes Aircraft Co | Spaceship earth-direction attitude acquiring method |
| EP0375496B1 (en) * | 1988-12-23 | 1993-09-22 | AEROSPATIALE Société Nationale Industrielle | Attitude determination and control system for a spin-stabilized satellite in an elliptic transfer orbit |
| CN102004491A (en) * | 2010-10-11 | 2011-04-06 | 北京控制工程研究所 | Initial sun capturing method during initial injection stage of satellite |
| CN103019247A (en) * | 2011-09-27 | 2013-04-03 | 上海航天控制工程研究所 | Gyroscope-free independent space attitude maneuver control method of Martian probe |
| CN104062976A (en) * | 2014-06-10 | 2014-09-24 | 北京控制工程研究所 | Fast aircraft attitude maneuvering method based on fact that angular acceleration derivatives are in sine curve |
| EP3379200A2 (en) * | 2017-03-21 | 2018-09-26 | Thales | Method for adjusting the joining trajectory for an aircraft |
| CN110901956A (en) * | 2019-12-10 | 2020-03-24 | 中国人民解放军国防科技大学 | Satellite stable sun-to-day orientation method with earth orientation deviation as constraint |
-
2020
- 2020-04-09 CN CN202010274480.6A patent/CN111483617B/en active Active
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0375496B1 (en) * | 1988-12-23 | 1993-09-22 | AEROSPATIALE Société Nationale Industrielle | Attitude determination and control system for a spin-stabilized satellite in an elliptic transfer orbit |
| JPH04231295A (en) * | 1990-05-14 | 1992-08-20 | Hughes Aircraft Co | Spaceship earth-direction attitude acquiring method |
| CN102004491A (en) * | 2010-10-11 | 2011-04-06 | 北京控制工程研究所 | Initial sun capturing method during initial injection stage of satellite |
| CN103019247A (en) * | 2011-09-27 | 2013-04-03 | 上海航天控制工程研究所 | Gyroscope-free independent space attitude maneuver control method of Martian probe |
| CN104062976A (en) * | 2014-06-10 | 2014-09-24 | 北京控制工程研究所 | Fast aircraft attitude maneuvering method based on fact that angular acceleration derivatives are in sine curve |
| EP3379200A2 (en) * | 2017-03-21 | 2018-09-26 | Thales | Method for adjusting the joining trajectory for an aircraft |
| CN110901956A (en) * | 2019-12-10 | 2020-03-24 | 中国人民解放军国防科技大学 | Satellite stable sun-to-day orientation method with earth orientation deviation as constraint |
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