CN114756039A - Multi-body coupling attitude control method and system based on zero force control - Google Patents
Multi-body coupling attitude control method and system based on zero force control Download PDFInfo
- Publication number
- CN114756039A CN114756039A CN202210384578.6A CN202210384578A CN114756039A CN 114756039 A CN114756039 A CN 114756039A CN 202210384578 A CN202210384578 A CN 202210384578A CN 114756039 A CN114756039 A CN 114756039A
- Authority
- CN
- China
- Prior art keywords
- rotating mechanism
- control
- torque
- attitude
- satellite platform
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
Images
Classifications
-
- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
- G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
- G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
Abstract
The invention discloses a multi-body coupling attitude control method and a system based on zero force control, which are used for acquiring disturbance torque between a satellite platform and a rotating mechanism, which is generated due to the movement of the rotating mechanism; and calculating a compensation torque according to the disturbance torque and the expected zero torque, and controlling the corresponding rotating mechanism to act according to the compensation torque so as to counteract the influence of the movement of the rotating mechanism on the control of the satellite platform. The rotating mechanism between the satellite platform and the effective load is used as an intermediate medium, the compensation of the satellite platform is realized by controlling the action of the rotating mechanism according to the compensation moment, the influence of the rotation of the load on the control of the satellite platform is reduced, the attitude control precision and the stability of the satellite platform are improved, the decoupling of the control of the satellite platform and the rotation of the load is realized, and the control requirement of the satellite platform when a plurality of loads are provided at the same time and point in different spaces can be met.
Description
Technical Field
The invention relates to the technical field of satellite-borne load application, in particular to a multi-body coupling attitude control method and system based on zero-force control.
Background
The space vehicle is restrained by the flight orbit, so that the task requirement of remote sensing on the ground target at any time cannot be met. In order to improve the response speed to the ground target and shorten the revisit period, the spacecraft is often required to have the capability of adjusting the remote sensing load attitude so as to improve the remote sensing range of the load.
The space load tends to have higher resolution, so that high requirements are put on the pointing accuracy and stability of the control. The existing main design scheme can not solve the problems that an aircraft platform carries various loads with different requirements on space direction and the load inertia is basically equivalent to the platform mass.
The current design scheme mainly aims at the following situations:
1. the inertia of the aircraft platform is far greater than that of a load rotating part (mostly seen in a space station), and the inertia of the load is far less than the mass of the platform, so that even if a plurality of turntable mechanisms are maneuvered simultaneously, the stability of the platform is hardly influenced, which is obviously not suitable for the requirement that the satellite platform is equivalent to the mass of the load;
2. although the load has high requirements on the pointing accuracy and stability of control, the requirement on the aircraft platform is not high, so that the spatial pointing and stability of the load are preferentially ensured during design, which is obviously not suitable for multi-load platforms because the load always has requirements on control, and the control of the loads is established on the basis of platform dynamics, on one hand, the platform is unstable and cannot simultaneously meet multiple loads and ensure the control to carry out tasks, on the other hand, the multiple loads are simultaneously controlled, and the influence on the platform can also cause the control divergence of the platform due to the mutual coupling effect of the multiple turntable mechanism controls, thereby causing the influence on the platform safety;
3. The platform load integration scheme can only meet the requirements of a single-load aircraft or the conditions that the direction of the load to the space pointing is the same or the mutual relation is determined, and cannot meet the requirement of simultaneously providing a plurality of loads with different space pointing.
Disclosure of Invention
Aiming at the defects of the prior art, the invention provides a multi-body coupling attitude control method and system based on zero force control, which aim to solve the problem that the existing satellite platform has a plurality of rotary table mechanisms and carries a large inertia load at the same time, and the high precision and stability of the platform and the load attitude control cannot be ensured.
In a first aspect, the invention provides a multi-body coupling attitude control method based on zero force control, which comprises the following steps:
acquiring disturbance torque between a satellite platform and a rotating mechanism, which is generated by the movement of the rotating mechanism;
and calculating a compensation torque according to the disturbance torque and the expected zero torque, and controlling the corresponding rotating mechanism to act according to the compensation torque so as to counteract the influence of the movement of the rotating mechanism on the control of the satellite platform.
Further, the specific calculation formula of the compensation torque is as follows:
wherein, Te＝To-Tr，ToFor desired zero moment, TrIs disturbance torque, kp1Is the proportionality coefficient, k, of the PI controller I1Is the integral coefficient of the PI controller, t0Is the starting time of the movement of the rotating mechanism, T is the current time, TcIs the compensation torque of the rotating mechanism.
Further, the specific calculation formula of the compensation torque is as follows:
wherein, Te＝To-Tr，ToFor desired zero moment, TrAs disturbance torque, TqFor a feed-forward moment, kp1Is the proportionality coefficient, k, of the PI controllerI1Is the integral coefficient of the PI controller, t0Is the starting time of the movement of the rotating mechanism, T is the current time, TcIs the compensation torque of the rotating mechanism.
Further, the feed-forward torque TqThe method is obtained by calculation according to the relative movement between the load and the corresponding rotating mechanism, and the specific calculation formula is as follows:
wherein, IqIs the inertia of the rotating mechanism and is,
is the angular acceleration of the rotating mechanism.
In a second aspect, the invention provides a multi-body coupling attitude control system based on zero-force control, which comprises a load control system, a satellite platform control system and a turntable control system arranged between the load control system and the satellite platform control system;
the turntable control system is used for acquiring a disturbance torque between the satellite platform and the rotating mechanism, which is generated by the movement of the rotating mechanism; and the compensation torque is calculated according to the disturbance torque and the expected zero torque, and the corresponding rotating mechanism is controlled to act according to the compensation torque so as to counteract the influence of the movement of the rotating mechanism on the control of the satellite platform.
Further, the turntable control system comprises a turntable controller, a torque sensor and a rotating mechanism;
the moment sensor is used for acquiring disturbance moment between the satellite platform and the rotating mechanism, which is generated by the movement of the rotating mechanism;
and the turntable controller calculates a compensation torque according to the disturbance torque and the expected zero torque, and controls the corresponding rotating mechanism to act according to the compensation torque.
Further, the satellite platform control system comprises a first attitude sensor, a Kalman filter, a platform controller and a platform executing mechanism;
the first attitude sensor acquires on-orbit real-time attitude information of the satellite platform, the Kalman filter filters the on-orbit real-time attitude information of the satellite platform, and the platform controller outputs a control instruction according to the filtered on-orbit real-time attitude information and target attitude information of the satellite platform to control the action of the platform executing mechanism so as to realize attitude control of the satellite platform.
Further, the first attitude sensor comprises a star sensor and a first three-axis fiber optic gyroscope.
Further, the load control system comprises a second attitude sensor, a load controller and a load executing mechanism;
The second attitude sensor acquires real-time attitude information of the payload, and the load controller outputs a control instruction according to the real-time attitude information of the payload and target attitude information thereof to control the action of the load executing mechanism, so that the directional control of the payload is realized.
Further, the load executing mechanism is a momentum wheel, the moment of the momentum wheel is distributed by adopting rotation speed control, and the specific formula is as follows:
TT＝-Cw·hw＝-Cw·Iw·Ωw
wherein, TTBeing the moment of the momentum wheel, CwIs a mounting matrix of the momentum wheel, hwIs the angular momentum of the momentum wheel, IwIs the inertia matrix of the momentum wheel, omegawThe rotational speed of the momentum wheel.
Further, the second attitude sensor includes a photoelectric encoder and a second triaxial fiber optic gyroscope.
The invention has the beneficial effects that:
the invention provides a multi-body coupling attitude control method and system based on zero force control, which take a rotating mechanism between a satellite platform and an effective load as an intermediate medium, calculate a compensation torque through the output torque of the satellite platform and the disturbance torque of the rotating mechanism, control the corresponding rotating mechanism to act according to the compensation torque to realize the compensation of the satellite platform, reduce the influence of load rotation on the satellite platform control, improve the attitude control precision and stability of the satellite platform, realize the decoupling of the satellite platform control and the load rotation, and can meet the control requirement of the satellite platform when a plurality of loads are directed in different spaces;
The rotating mechanism is used as an intermediate medium between the effective load and the satellite platform, a follow-up strategy is adopted, the transmission torque of the effective load and the satellite platform main body at the joint is zero, namely zero torque transmission, the control precision and the control stability of the effective load and the satellite platform are ensured, and the design difficulty of carrying a large-inertia load by a multi-body platform is simplified.
Drawings
In order to more clearly illustrate the technical solution of the present invention, the drawings required to be used in the description of the embodiments are briefly introduced below, it is obvious that the drawings in the following description are only one embodiment of the present invention, and it is obvious for those skilled in the art to obtain other drawings without creative efforts.
Fig. 1 is a block diagram of a control strategy of a turntable control system according to an embodiment of the present invention;
FIG. 2 is a block diagram of a control strategy of a satellite platform control system according to an embodiment of the present invention;
FIG. 3 is a block diagram of a control strategy for a load control system in an embodiment of the present invention;
fig. 4 shows a connection relationship between a load, a turntable, and a satellite body according to an embodiment of the present invention.
Detailed Description
The technical solutions in the present invention are clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be obtained by a person skilled in the art without inventive efforts based on the embodiments of the present invention, shall fall within the scope of protection of the present invention.
The technical solution of the present application will be described in detail below with specific examples. The following several specific embodiments may be combined with each other, and details of the same or similar concepts or processes may not be repeated in some embodiments.
The aircraft is oriented to the ground, the orientation of the load is realized through a rotating mechanism, the aircraft (such as a small satellite) is small in size and light in weight and is easily influenced by the rotation of the load, the rotating mechanism (a space turntable, a scanning mirror, a solar sailboard and the like) of the on-satellite load has larger mass relative to the small satellite, the rotation of the load can cause larger rotation inertia fluctuation of the small satellite, the operation stability and the attitude control of the satellite are seriously influenced, the operation of the satellite body can be seriously interfered by the existence of disturbance moment of the rotating mechanism and vibration, the deviation of the satellite attitude is caused, and the precision of a rotation detection sensor is reduced; the system bandwidth of hardware of a satellite platform control system and a turntable control system is limited, the influence is eliminated by a rotating mechanism and the satellite platform in a self-adaptive mode, and the effect is poor. Based on the method, the compensation torque is calculated according to the output torque of the satellite platform and the disturbance torque of the rotating mechanism, the corresponding rotating mechanism (such as a rotary table) is controlled to act according to the compensation torque to counteract the influence of the movement of the rotating mechanism on the satellite platform, and the decoupling of the control of the satellite platform and the rotation of the load is realized.
The embodiment of the invention provides a multi-body coupling attitude control method based on zero force control, wherein a multi-body coupling attitude control system comprises a load control system, a satellite platform control system and a rotary table control system arranged between the load control system and the satellite platform control system. The specific control method comprises the following steps:
step 1: and acquiring disturbance torque between the satellite platform and the rotating mechanism, which is generated by the movement of the rotating mechanism.
And 2, step: and (2) calculating a compensation moment according to the disturbance moment and the expected zero moment in the step (1), and controlling the corresponding rotating mechanism to act according to the compensation moment so as to counteract the influence of the movement of the rotating mechanism on the control of the satellite platform.
In one embodiment of the present invention, as shown in fig. 1, a turntable control system includes a turntable controller, a torque sensor, and a rotation mechanism; the moment sensor collects disturbance moment generated between the satellite platform and the rotating mechanism due to movement of the rotating mechanism and sends the disturbance moment to the rotary table controller, the rotary table controller calculates compensation moment according to the disturbance moment and expected zero moment, and the specific calculation formula is as follows:
wherein, Te＝To-Tr，ToAt the desired zero moment, i.e. 0, TrAs disturbance torque, kp1Is the proportionality coefficient, k, of the PI controller I1Being PI controllersIntegral coefficient, t0Is the starting time of the movement of the rotating mechanism, T is the current time, TcIs the compensation moment of the rotating mechanism.
If only the disturbance torque T is relied onrIn the feedback control, the control response speed is slow and the tracking error is large due to the hysteresis effect of the turntable controller, so in another embodiment of the present invention, the specific calculation formula of the compensation torque is as follows:
wherein, TqFor feed-forward torque, feed-forward torque TqThe method is obtained by calculation according to the relative movement between the load and the corresponding rotating mechanism, and the specific calculation formula is as follows:
wherein, IqIs the inertia of the corresponding rotating mechanism,
is the angular acceleration of the corresponding rotating mechanism. When the motion planning of load pointing is performed based on the known current pointing direction of the satellite platform (namely, the space attitude reference established by the satellite platform) and the expected pointing direction of the task target load, the angular speed and the angular acceleration of the rotating mechanism can be obtained, and the angular acceleration of the rotating mechanism is used as the input of control feedforward.
The feedforward part is added in the calculation of the compensation moment, so that the response speed of a control system can be greatly improved, the tracking error of the rotating mechanism in the follow-up process is reduced, and the control stability is improved.
Turntable controller according to TcControl instructions are output to control the rotating mechanism to act, so that the torque generated by the rotating mechanism is TcThe method and the device have the advantages that the influence of the movement of the rotating mechanism on the ground orientation attitude of the satellite platform is counteracted, and the decoupling of the control of the satellite platform and the rotation of the load is realized. To pairAnd in the plurality of rotating mechanisms, each rotating mechanism is controlled according to the corresponding compensation torque, so that the control requirement of the satellite platform when a plurality of loads are provided at different space directions can be met. In the present embodiment, the movement of the rotating mechanism is driven by a brushless motor.
In one embodiment of the present invention, as shown in fig. 2, a satellite platform control system includes a first attitude sensor, a kalman filter, a platform controller, and a platform actuator; the first attitude sensor acquires the in-orbit real-time attitude information of the satellite platform, the Kalman filter carries out filtering estimation on the in-orbit real-time attitude information of the satellite platform, and the platform controller carries out filtering estimation on the in-orbit real-time attitude information of the satellite platform according to the in-orbit real-time attitude information of the satellite platform after filtering estimation
And target attitude information thereof
θref_sAnd outputting a control instruction to control the platform executing mechanism to act so as to realize the attitude control of the satellite platform.
In an embodiment of the present invention, the first attitude sensor includes a star sensor and a first triaxial fiber optic gyroscope, the star sensor and the first triaxial fiber optic gyroscope are fixedly mounted on a platform of the satellite platform, the first triaxial fiber optic gyroscope is sensitive to an angular velocity of the satellite platform, and the star sensor is sensitive to an azimuth of the satellite platform to provide measurement information for the first triaxial fiber optic gyroscope. The Kalman filter filters and estimates the attitude information actually measured by the first triaxial fiber gyroscope and the star sensor, and the platform controller filters and estimates the attitude information according to the actually measured attitude information 
And target attitude information thereof
θref_sObtaining the moment T of the platform actuatorsAccording to the torque T of the platform actuatorsSending out control instruction to make platform executeThe mechanism generates a torque T under the control instructionsThe concrete formula is as follows:
Ts＝-kp2·sgn(q0)·I·qv-kd2·I·ω+ω×H (4)
wherein, TsMoment information, k, output for satellite platformsp2Is the proportionality coefficient, k, of the PD controllerd2Sgn () is a sign function, q, for the differential coefficient of the PD controller0Is an error quaternion scalar section; q. q.s0By target attitude
θref_sWith the actual attitude
The calculation is obtained, and the calculation mode is the prior art; q. q.svAs part of the error quaternion vector, qvBy target attitude
θref_sWith the actual attitude
The calculation is obtained, and the calculation mode is the prior art; i is a star inertia matrix (representing star mass characteristics), omega is a star angular velocity, omega is detected and obtained by an angular velocity measuring sensor, and H is star angular momentum. In FIG. 2, the reaction wheel is the platform actuator, i.e., torque TsThe actuator of (1).
In one embodiment of the present invention, as shown in fig. 3, the load control system includes a second attitude sensor, a load controller, and a load actuator; the second attitude sensor collects real-time attitude information of the payload, and the load controller outputs a control instruction to control the action of the load executing mechanism according to the real-time attitude information of the payload and target attitude information of the payload, so that the directional control of the payload is realized.
In one embodiment of the present invention, the second attitude sensor includes a photoelectric encoder and a second attitude sensorThe two-three-axis optical fiber gyroscope comprises a photoelectric encoder and a second three-axis optical fiber gyroscope which are fixedly arranged on a load, wherein the second three-axis optical fiber gyroscope is sensitive to the rotation angle rate of the load and the relative rotation angle information of the load sensitive by the photoelectric encoder. The load controller is based on the measured attitude information
And target attitude information thereof
θref_TObtaining the torque T of the load actuatorTAccording to the torque T of the load actuatorTSending a control command to enable the load executing mechanism to generate a torque T under the control commandT. In this embodiment, the load executing mechanism is a momentum wheel, the moment of the momentum wheel is realized by controlling the rotating speed of the momentum wheel, and the specific distribution formula is as follows:
TT＝-Cw·hw＝-Cw·Iw·Ωw (5)
wherein, TTMoment of momentum wheel, CwIs a mounting matrix of the momentum wheel, hwIs the angular momentum of the momentum wheel, IwIs the inertia matrix of the momentum wheel, omegawThe rotational speed of the momentum wheel.
The satellite platform control system establishes three-axis attitude reference coordinates to provide a dynamic basis for control of each rotating mechanism, and when the load needs to adjust the current spatial orientation, the load control system calculates a control instruction of the momentum wheel based on the three-axis attitude reference coordinates established by the satellite platform to realize high-precision orientation control of the effective load; according to the disturbance torque acquired by the torque sensor, the turntable controller calculates the compensation torque, and controls the rotating mechanism according to the compensation torque, so that zero torque transmission between the satellite platform and the load is realized.
The embodiment of the invention also provides a multi-body coupling attitude control system based on zero-force control, which comprises a load control system, a satellite platform control system and a rotary table control system arranged between the load control system and the satellite platform control system. As shown in fig. 4, the satellite platform is connected to a rotating mechanism via a connection support structure, and the rotating mechanism establishes a control connection between the satellite platform and the payload.
In one embodiment of the present invention, as shown in fig. 1, a turntable control system includes a turntable controller, a torque sensor, and a rotating mechanism; the moment sensor collects disturbance moment generated between the satellite platform and the rotating mechanism due to movement of the rotating mechanism and sends the disturbance moment to the rotary table controller, and the rotary table controller calculates compensation moment according to the disturbance moment and expected zero moment, wherein a specific formula is shown as a formula (1) or (2). Turntable controller according to TcControl instructions are output to control the rotating mechanism to act, so that the torque generated by the rotating mechanism is TcThe method and the device have the advantages that the influence of the movement of the rotating mechanism on the ground orientation attitude of the satellite platform is counteracted, and the decoupling of the control of the satellite platform and the rotation of the load is realized.
In one embodiment of the present invention, as shown in fig. 2, the satellite platform control system includes a star sensor, a first triaxial fiber optic gyroscope, a kalman filter, a platform controller, and a platform actuator; the star sensor and the first triaxial fiber optic gyroscope are fixedly arranged on a platform surface of the satellite platform, the first triaxial fiber optic gyroscope is sensitive to the angular speed of the satellite platform, and the star sensor is sensitive to the azimuth of the satellite platform, so that measurement information is provided for the first triaxial fiber optic gyroscope; the Kalman filter filters and estimates the attitude information actually measured by the first triaxial fiber gyroscope and the star sensor, and the platform controller filters and estimates the attitude information according to the actually measured attitude information 
And its target attitude information
θref_sObtaining the moment T of the platform actuatorsAccording to the torque T of the platform actuatorsSending a control instruction to enable the platform actuating mechanism to generate a torque T under the control instructionsThe specific formula is shown as formula (4), and the real-time control of the satellite platform attitude is realized.
In one embodiment of the present invention, as shown in fig. 3, the load control system includes an optical-electrical encoder, a second three-axis fiber optic gyroscope, a load controller, and a load actuator; the photoelectric encoder and the second three-axis optical fiber gyroscope are fixedly arranged on the load, the second three-axis optical fiber gyroscope is sensitive to the rotation angular rate of the load and the relative rotation angle information of the load sensitive by the photoelectric encoder; the load controller is based on the measured attitude information
And its target attitude information
θref_TObtaining the torque T of the load actuatorTAccording to the torque T of the load actuatorTSending a control command to enable the load executing mechanism to generate a torque T under the control commandTAnd realizing the directional control of the payload.
While preferred embodiments of the present invention have been described, additional variations and modifications in those embodiments may occur to those skilled in the art once they learn of the basic inventive concepts. Therefore, it is intended that the appended claims be interpreted as including preferred embodiments and all such alterations and modifications as fall within the scope of the invention. It will be apparent to those skilled in the art that various changes and modifications may be made in the present invention without departing from the spirit and scope of the invention. Thus, if such modifications and variations of the present invention fall within the scope of the claims of the present invention and their equivalents, the present invention is also intended to include such modifications and variations.
Claims (10)
1. A multi-body coupling attitude control method based on zero force control is characterized by comprising the following steps:
acquiring disturbance torque between a satellite platform and a rotating mechanism, which is generated by the movement of the rotating mechanism;
and calculating a compensation torque according to the disturbance torque and the expected zero torque, and controlling the corresponding rotating mechanism to act according to the compensation torque so as to counteract the influence of the movement of the rotating mechanism on the control of the satellite platform.
2. The multi-body coupling attitude control method based on zero-force control as claimed in claim 1, wherein the specific calculation formula of the compensation moment is:
wherein, Te＝To-Tr，ToTo expect zero moment, TrAs disturbance torque, kp1Is the proportionality coefficient of PI controller, kI1Is the integral coefficient of the PI controller, t0Is the starting time of the movement of the rotating mechanism, T is the current time, TcIs the compensation torque of the rotating mechanism.
3. The multi-body coupling attitude control method based on zero-force control as claimed in claim 1, wherein the specific calculation formula of the compensation moment is:
wherein, Te＝To-Tr，ToFor desired zero moment, TrAs disturbance torque, TqFor a feed-forward moment, kp1Is the proportionality coefficient, k, of the PI controllerI1Is the integral coefficient of the PI controller, t 0Is the starting time of the movement of the rotating mechanism, T is the current time, TcIs the compensation moment of the rotating mechanism.
4. The multi-body coupling attitude control method based on zero-force control as claimed in claim 3, characterized in that the feedforward torque TqThe method is obtained by calculation according to the relative movement between the load and the corresponding rotating mechanism, and the specific calculation formula is as follows:
wherein, IqIs the inertia of the rotating mechanism and is,
is the angular acceleration of the rotating mechanism.
5. A multi-body coupling attitude control system based on zero force control comprises a load control system, a satellite platform control system and a rotary table control system arranged between the load control system and the satellite platform control system; the method is characterized in that:
the turntable control system is used for acquiring a disturbance torque between the satellite platform and the rotating mechanism, which is generated by the movement of the rotating mechanism; and the compensation torque is calculated according to the disturbance torque and the expected zero torque, and the corresponding rotating mechanism is controlled to act according to the compensation torque so as to counteract the influence of the movement of the rotating mechanism on the control of the satellite platform.
6. The zero-force-control-based multi-body-coupled attitude control system of claim 5, wherein the turntable control system comprises a turntable controller, a torque sensor, and a rotation mechanism;
The moment sensor is used for acquiring disturbance moment between the satellite platform and the rotating mechanism, which is generated by the movement of the rotating mechanism;
and the turntable controller calculates a compensation torque according to the disturbance torque and the expected zero torque, and controls the corresponding rotating mechanism to act according to the compensation torque.
7. The zero-force-control-based multi-body-coupled attitude control system according to claim 6, wherein the specific calculation formula of the compensation moment is as follows:
wherein, Te＝To-Tr，ToFor desired zero moment, TrIs disturbance torque, kp1Is the proportionality coefficient, k, of the PI controllerI1Is the integral coefficient of the PI controller, t0Is the starting time of the movement of the rotating mechanism, T is the current time, TcIs the compensation torque of the rotating mechanism.
8. The zero-force-control-based multi-body-coupled attitude control system according to claim 6, wherein the specific calculation formula of the compensation moment is as follows:
wherein, Te＝To-Tr，ToFor desired zero moment, TrAs disturbance torque, TqFor a feed-forward moment, kp1Is the proportionality coefficient, k, of the PI controllerI1Is the integral coefficient of the PI controller, t0Is the starting time of the movement of the rotating mechanism, T is the current time, TcIs the compensation torque of the rotating mechanism.
9. The zero-force-control-based multi-body-coupled attitude control system according to any one of claims 5 to 8, wherein the satellite platform control system comprises a first attitude sensor, a Kalman filter, a platform controller and a platform actuator;
The first attitude sensor acquires on-orbit real-time attitude information of the satellite platform, the Kalman filter filters the on-orbit real-time attitude information of the satellite platform, and the platform controller outputs a control instruction according to the filtered on-orbit real-time attitude information and target attitude information of the satellite platform to control the action of the platform executing mechanism so as to realize attitude control of the satellite platform.
10. The zero-force-control-based multi-body-coupled attitude control system according to any one of claims 5 to 8, wherein the load control system comprises a second attitude sensor, a load controller and a load actuator;
the second attitude sensor acquires real-time attitude information of the payload, and the load controller outputs a control instruction according to the real-time attitude information of the payload and target attitude information thereof to control the action of the load executing mechanism so as to realize the directional control of the payload;
preferably, the load executing mechanism is a momentum wheel, the moment of the momentum wheel is distributed by adopting rotation speed control, and the specific formula is as follows:
TT＝-Cw·hw＝-Cw·Iw·Ωw
wherein, TTMoment of momentum wheel, CwIs a mounting matrix of the momentum wheel, hwIs the angular momentum of the momentum wheel, IwIs the inertia matrix of the momentum wheel, omega wIs the rotational speed of the momentum wheel.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210384578.6A CN114756039A (en) | 2022-04-13 | 2022-04-13 | Multi-body coupling attitude control method and system based on zero force control |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210384578.6A CN114756039A (en) | 2022-04-13 | 2022-04-13 | Multi-body coupling attitude control method and system based on zero force control |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN114756039A true CN114756039A (en) | 2022-07-15 |
Family
ID=82330821
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202210384578.6A Pending CN114756039A (en) | 2022-04-13 | 2022-04-13 | Multi-body coupling attitude control method and system based on zero force control |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN114756039A (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116442240A (en) * | 2023-05-26 | 2023-07-18 | 中山大学 | Robot zero-force control method and device based on high-pass filtering decoupling |
-
2022
- 2022-04-13 CN CN202210384578.6A patent/CN114756039A/en active Pending
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116442240A (en) * | 2023-05-26 | 2023-07-18 | 中山大学 | Robot zero-force control method and device based on high-pass filtering decoupling |
| CN116442240B (en) * | 2023-05-26 | 2023-11-14 | 中山大学 | Robot zero-force control method and device based on high-pass filtering decoupling |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| Kim et al. | Automatic mass balancing of air-bearing-based three-axis rotational spacecraft simulator | |
| JP2635821B2 (en) | Three-axis stabilizing satellite pointing at the earth and method for capturing the attached sun and earth | |
| CN111099045B (en) | Full physical simulation method for double super satellite dynamics and control air floatation platform | |
| CN110329550B (en) | Attitude control for agile satellite applications | |
| US8620496B2 (en) | Systems and method of controlling a spacecraft using attitude sensors | |
| EP1977972B1 (en) | Active control of soft hybrid isolation systems | |
| US20080315039A1 (en) | System and methods for space vehicle torque balancing | |
| US20220155800A1 (en) | Method and apparatus for yaw fusion and aircraft | |
| CN109677637B (en) | Space non-cooperative target tracking and pointing method based on optical angle measuring camera | |
| JP4550347B2 (en) | System and method for controlling the attitude of a spacecraft | |
| CN110361829A (en) | A kind of telescope Pointing Calibration method and telescope | |
| US20050242241A1 (en) | Method and system for steering a momentum control system | |
| CN111880410A (en) | Four-rotor unmanned aerial vehicle fault-tolerant control method for motor faults | |
| CN114756039A (en) | Multi-body coupling attitude control method and system based on zero force control | |
| US5875676A (en) | Non colocated rate sensing for control moment gyroscopes | |
| CN108427429B (en) | Spacecraft visual axis maneuvering control method considering dynamic pointing constraint | |
| CN110750110B (en) | Pointing control device and control method for space laser communication | |
| CN110955255A (en) | High-precision orbit control attitude maintaining method, system and medium based on CMG | |
| CN113761664B (en) | Flywheel rotation speed optimization method during remote sensing satellite imaging | |
| CA1325666C (en) | Nutation sensor and nutation control system for a dual-spin stabilized satellite | |
| CN111891402B (en) | Mars detection ground antenna pointing recovery method based on autonomous maneuvering | |
| Reijneveld et al. | Attitude control system of the Delfi-n3Xt satellite | |
| EP4267465A1 (en) | Systems and methods for payload stabilization | |
| JPH11129997A (en) | Attitude control device of artificial satellite | |
| EP1134640B1 (en) | Attitude control system for a spacecraft |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination |