CN112684697A - Split type satellite in-orbit two-cabin rotational inertia identification method and system - Google Patents
Split type satellite in-orbit two-cabin rotational inertia identification method and system Download PDFInfo
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Abstract
The invention provides an identification method and system for the rotational inertia of two in-orbit compartments of a split type satellite, which are used for measuring the rotational inertia of the two compartments of the satellite around the respective mass centers after the two compartments of the satellite are separated, and comprise the following steps: powering off a repeated locking mechanism between two satellite cabins, electrifying a magnetic floating actuator between the two cabins, generating a magnetic control couple with known size, and driving the two cabins to rotate; the control system keeps open loop and uncontrolled, collects the angular velocity information of the two cabins and the three axes, fits the angular velocity curve, and differentiates to obtain the curve of the three axes angular acceleration; and magnetic control couple information output by combining the magnetic suspension actuator and three-axis angular acceleration resolving to obtain two-cabin rotational inertia. According to the method, through the design of the on-orbit identification method of the rotational inertia of the two cabins, the accurate identification of the rotational inertia of the two cabins is realized, dynamic data support is provided for the on-orbit attitude control of the satellite, and the technical problem that accurate identification is difficult due to the influence of gravity and an expansion part on the ground is solved.
Description
Technical Field
The invention relates to a composite control technology of a load cabin of a satellite platform with ultrahigh pointing accuracy and ultrahigh stability (double-super), in particular to a novel method for identifying the rotational inertia of two in-orbit cabins of a split satellite.
Background
The requirements of the advanced spacecraft on the attitude pointing accuracy and the stability in the future are two orders of magnitude higher than those of the advanced spacecraft in the prior art. The traditional design of fixedly connecting a load and a platform is adopted, the dynamic characteristics of the two are deeply coupled, so that the load double super indexes are difficult to realize, and although certain effects are achieved by adopting methods such as active and passive micro vibration suppression and the like, the double super indexes are difficult to realize due to the defect of the limited fixedly connected design.
The 'double-super' satellite platform breaks through the traditional fixed connection design, adopts a non-contact, high-precision and time-delay-free displacement sensor to realize the separation of a load (cabin) only provided with a quiet component and a platform (cabin) provided with a movable component, and thoroughly eliminates the micro-vibration influence. The traditional control logic mainly based on a satellite platform is changed, and a brand new method of 'load cabin driving, platform cabin driven and two cabin relative positions cooperative decoupling control' is adopted for the first time, so that the double super-precision of the load cabin can be realized.
The accurate identification of the rotational inertia positions of the two cabins is the premise of realizing high-precision attitude control of the load cabin and collaborative quick attitude maneuver feedforward control of the two cabins. The ground is limited by the influence of gravity and unfolding components and is difficult to accurately identify, so that accurate two-cabin rotational inertia information needs to be acquired by an on-orbit calibration method.
Patent document CN106249749A discloses a master-slave non-contact dual-super satellite platform variable-centroid and variable-inertia attitude control system, which does not solve the technical problem of accurate identification of two-compartment rotational inertia position.
Patent document CN102620886A discloses a two-step on-orbit identification combined spacecraft rotational inertia estimation method, which combines an EKF algorithm and a least square algorithm to realize optimal estimation of spacecraft rotational inertia, but the research object is an integrated satellite and does not relate to the research of two-cabin rotational inertia.
In the document [1] [2] [3] [4], a method for identifying the quality characteristics of the on-orbit spacecraft based on different algorithms is proposed, but all research objects of the method are all integrated satellites, and the method is not related to and is not suitable for identifying the quality characteristics of the split satellites in a separated state.
[1]Bergmann EV,Walker BK,Levy D R.Mass property estimation for control of asymmetrical satellites[J].Journal of Guidance,Control and Dynamics,1987,10(2):483-492.
[2]Bergmann E V,Dzielski J.Spacecraft mass property identification with torque-generating control[J].Journal of Guidance,Control,and Dynamics,1990,13(2):99-103.
[3]Wilson E,Lages C,Mah R.On-line,gyro-based,mass-property identification for thruster-controlled spacecraft using recursive least squares[C]//Proceedings of the 45th Midwest Symposium on Circuits and Systems.Moffett Field,California,Ames Research Center，Aug.4-7,2002
[4]Tanygin S,Williams T.Mass property estimation using coasting maneuvers[J].Journal of Guidance,Control,and Dynamics,1997,20(4):625-632。
Disclosure of Invention
Aiming at the requirement of high-precision attitude control of a load cabin of a two-cabin non-contact type 'double-super' satellite platform, the invention aims to provide a novel method for identifying the rotational inertia of two in-orbit cabins of a split type satellite and a novel system for identifying the rotational inertia of two in-orbit cabins of the split type satellite.
According to an aspect of the present invention, there is provided a method for identifying an in-orbit two-compartment rotational inertia of a split satellite, where a schematic diagram of identifying the in-orbit two-compartment rotational inertia is shown in fig. 1, and the method includes:
a driving step: powering off a repeated locking mechanism between two satellite cabins, electrifying a magnetic floating actuator between the two cabins, generating a magnetic control couple with known size, and driving the two cabins to rotate;
the collection step comprises: the control system keeps open loop and uncontrolled, collects the angular velocity information of the two cabins and the three axes, fits the angular velocity curve, and differentiates to obtain the curve of the three axes angular acceleration;
two-cabin rotational inertia resolving step: and magnetic control couple information output by combining the magnetic suspension actuator and three-axis angular acceleration resolving to obtain two-cabin rotational inertia.
Preferably, the driving step adopts a plurality of magnetic suspension actuators arranged between two cabins and is used for finishing the three-axis attitude control of the load cabin and the control of the relative mass center position between the two cabins. Taking 2 of the platform, outputting a magnetic control couple (outputting a directional couple every time), generating equal torque to act on the platform cabin and the load cabin, and recording the torque as TLO。
Preferably, let the platform cabin angular velocity be ωbAngular velocity of load compartment is ωpSetting the two-cabin rotational inertia matrix as JbAnd JpThen there is
Angular velocity information is acquired sensitively through the gyros configured in the two cabins, and angular acceleration is acquired through linear fitting. The platform cabin fitting angular velocity is recorded as omegab_NHAnd the fitting angular velocity of the load compartment is recorded as omegap_NH，
Is the angular acceleration of the platform cabin,
the angular acceleration of the load compartment is represented by a linear fitting formula
Wherein k isbAngular acceleration of the platform cabin, k, obtained for fittingpFor the fitted angular acceleration of the load compartment, bbAnd bpRespectively is a constant coefficient, and t is the action time of the magnetic control couple.
Preferably, the fitted angular acceleration of the load compartment is set as
The angular acceleration of the platform cabin obtained by fitting is
Then there is
Thereby obtaining the identification result of the rotational inertia of the two cabins as
According to another aspect of the present invention, there is also provided a novel split type satellite in-orbit two-compartment rotational inertia identification system, including:
a driving module: powering off a repeated locking mechanism between two satellite cabins, electrifying a magnetic floating actuator between the two cabins, generating a magnetic control couple with known size, and driving the two cabins to rotate;
an acquisition module: the control system keeps open loop and uncontrolled, collects the angular velocity information of the two cabins and the three axes, fits the angular velocity curve, and differentiates to obtain the curve of the three axes angular acceleration;
two cabin inertia of rotation solve the module: and magnetic control couple information output by combining the magnetic suspension actuator and three-axis angular acceleration resolving to obtain two-cabin rotational inertia.
Preferably, the driving module adopts a plurality of magnetic suspension actuators arranged between two cabins and is used for finishing three-axis attitude control of the load cabin and relative mass center position control between the two cabins. Taking 2 of the two sets of magnetic control force couples (one output at a time)Directional couple) to generate equal moments acting on the platform compartment and the load compartment, denoted TLO。
Preferably, let the platform cabin angular velocity be ωbAngular velocity of load compartment is ωpSetting the two-cabin rotational inertia matrix as JbAnd JpThen there is
Angular velocity information is sensitively obtained through the gyros arranged in the two cabins, and angular acceleration is obtained through linear fitting. The platform cabin fitting angular velocity is recorded as omegab_NHAnd the fitting angular velocity of the load compartment is recorded as omegap_NH，
Is the angular acceleration of the platform cabin,
the angular acceleration of the load compartment is represented by a linear fitting formula
Wherein k isbAngular acceleration of the platform cabin, k, obtained for fittingpFor the fitted angular acceleration of the load compartment, bbAnd bpRespectively is a constant coefficient, and t is the action time of the magnetic control couple.
Preferably, the fitted angular acceleration of the load compartment is set as
The angular acceleration of the platform cabin obtained by fitting is
Thereby obtaining the identification result of the rotational inertia of the two cabins as
Compared with the prior art, the invention has the following beneficial effects:
1. the accurate identification of the rotational inertia of the two cabins is realized through the design of the on-orbit identification method of the rotational inertia of the two cabins;
2. providing dynamic data support for the attitude control of the satellite in orbit;
3. the technical problem that accurate identification is difficult due to the influence of gravity and an unfolding component on the ground is solved.
Drawings
Other features, objects and advantages of the invention will become more apparent upon reading of the detailed description of non-limiting embodiments with reference to the following drawings:
FIG. 1 is a schematic diagram of the principle of identifying the on-orbit two-compartment rotational inertia of a novel split satellite;
FIG. 2 is a schematic diagram of an in-orbit two-compartment rotational inertia identification method for a novel split satellite;
FIG. 3 is a schematic view of load compartment angular velocity measurement and fitting;
fig. 4 is a schematic diagram of the measurement and fitting of the pod angular velocity.
Detailed Description
The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the invention, but are not intended to limit the invention in any way. It should be noted that, for those skilled in the art, variations and modifications can be made without departing from the concept of the present invention, such as variations in the magnitude or direction of the thrust of the magnetic levitation actuator, variations in the installation position and orientation of the magnetic levitation actuator, variations in the method of identifying angular velocity, etc. All falling within the scope of the present invention.
As shown in fig. 1, the present invention provides a novel method for identifying the in-orbit two-compartment rotational inertia of a split satellite. More specifically, the method comprises the following steps:
a driving step: powering off a repeated locking mechanism between two satellite cabins, electrifying a magnetic floating actuator between the two cabins, generating a magnetic control couple with known size, and driving the two cabins to rotate;
the collection step comprises: the control system keeps open loop and uncontrolled, collects the angular velocity information of the two cabins and the three axes, fits the angular velocity curve, and differentiates to obtain the curve of the three axes angular acceleration;
two-cabin rotational inertia resolving step: and magnetic control couple information output by combining the magnetic suspension actuator and three-axis angular acceleration resolving to obtain two-cabin rotational inertia.
And in the driving step, a plurality of magnetic suspension actuators are arranged between two cabins and are used for finishing the three-axis attitude control of the load cabin and the control of the relative mass center position between the two cabins. Taking 2 of the platform, outputting a magnetic control couple (outputting a directional couple every time), generating equal torque to act on the platform cabin and the load cabin, and recording the torque as TLO。
Let the angular velocity of the platform cabin be omegabAngular velocity of load compartment is ωpSetting the two-cabin rotational inertia matrix as JbAnd JpThen there is
Angular velocity information is acquired sensitively through the gyros configured in the two cabins, and angular acceleration is acquired through linear fitting. The platform cabin fitting angular velocity is recorded as omegab_NHAnd the fitting angular velocity of the load compartment is recorded as omegap_NH，
Is the angular acceleration of the platform cabin,
the angular acceleration of the load compartment is represented by a linear fitting formula
Wherein k isbAngular acceleration of the platform cabin, k, obtained for fittingpFor the fitted angular acceleration of the load compartment, bbAnd bpRespectively is a constant coefficient, and t is the action time of the magnetic control couple.
The angular acceleration of the load compartment obtained by fitting is set as
The angular acceleration of the platform cabin obtained by fitting is
Then there is
Thereby obtaining the identification result of the rotational inertia of the two cabins as
In the present embodiment, the parameter setting rule is as shown in fig. 2. The output couple of the magnetic suspension actuator between the two given cabins is 0.01Nm, namely
TLO＝0.01Nm
The system is in an open-loop uncontrolled state, under the action of the output couple of the magnetic suspension actuator, the angular velocities of the two cabins and the fitting result thereof are shown in figures 3 and 4, and the angular accelerations of the two cabins around the center of mass are obtained according to the slope of the fitting curve
The results of the identification of the moment of inertia about the two compartments are shown in the table below,
it can be seen that the maximum deviation from the set theoretical value is 3.5%, and the accuracy depends on the measurement accuracy of the triaxial angular velocity.
The invention provides a novel split type satellite in-orbit two-cabin rotational inertia identification system, which comprises:
a driving module: powering off a repeated locking mechanism between two satellite cabins, electrifying a magnetic floating actuator between the two cabins, generating a magnetic control couple with known size, and driving the two cabins to rotate;
an acquisition module: the control system keeps open loop and uncontrolled, collects the angular velocity information of the two cabins and the three axes, fits the angular velocity curve, and differentiates to obtain the curve of the three axes angular acceleration;
two cabin inertia of rotation solve the module: and magnetic control couple information output by combining the magnetic suspension actuator and three-axis angular acceleration resolving to obtain two-cabin rotational inertia.
Those skilled in the art will appreciate that, in addition to implementing the system and its various devices, modules, units provided by the present invention as pure computer readable program code, the system and its various devices, modules, units provided by the present invention can be fully implemented by logically programming method steps in the form of logic gates, switches, application specific integrated circuits, programmable logic controllers, embedded microcontrollers and the like. Therefore, the system and various devices, modules and units thereof provided by the invention can be regarded as a hardware component, and the devices, modules and units included in the system for realizing various functions can also be regarded as structures in the hardware component; means, modules, units for performing the various functions may also be regarded as structures within both software modules and hardware components for performing the method.
The foregoing description of specific embodiments of the present invention has been presented. It is to be understood that the present invention is not limited to the specific embodiments described above, and that various changes or modifications may be made by one skilled in the art within the scope of the appended claims without departing from the spirit of the invention. The embodiments and features of the embodiments of the present application may be combined with each other arbitrarily without conflict.
Claims (8)
1. A split type satellite in-orbit two-cabin rotational inertia identification method is characterized by comprising the following steps:
a driving step: powering off a repeated locking mechanism between two satellite cabins, electrifying a magnetic floating actuator between the two cabins, generating a magnetic control couple with known size, and driving the two cabins to rotate;
the collection step comprises: the control system keeps open loop and uncontrolled, collects the angular velocity information of the two cabins and the three axes, fits the angular velocity curve, and differentiates to obtain the curve of the three axes angular acceleration;
two-cabin rotational inertia resolving step: and according to the triangular shaft acceleration curve obtained in the acquisition step, combining magnetic control couple information output by the magnetic suspension actuator and three-shaft angular acceleration resolving to obtain two-cabin rotational inertia.
2. The method for identifying the in-orbit two-compartment rotational inertia of the split-type satellite according to claim 1, wherein in the driving step, a plurality of magnetic levitation actuators are arranged between the two compartments for completing the three-axis attitude control of the load compartment and the relative centroid position control between the two compartments, 2 of the magnetic control couples are output, and the magnetic control couples in one direction are output each time to generate equal torque to act on the platform compartment and the load compartment, which is marked as TLO。
3. The split-type satellite in-orbit two-compartment rotational inertia identification method according to claim 2,
let the angular velocity of the platform cabin be omegabAngular velocity of load compartment is ωpSetting the two-cabin rotational inertia matrix as JbAnd JpThen there is
Angular velocity information is sensitively obtained through a gyroscope configured in two cabins, angular acceleration is obtained through linear fitting, and the fitting angular velocity of the platform cabin is recorded asωb_NHAnd the fitting angular velocity of the load compartment is recorded as omegap_NH，
Is the angular acceleration of the platform cabin,
the angular acceleration of the load compartment is represented by a linear fitting formula
Wherein k isbI.e. the angular acceleration of the platform cabin, k, obtained by fittingpI.e. the fitted angular acceleration of the load compartment, bbAnd bpRespectively is a constant coefficient, and t is the action time of the magnetic control couple.
4. The method for identifying the in-orbit two-compartment rotational inertia of a split-type satellite according to claim 3, wherein the angular acceleration of the load compartment obtained by fitting is set as
The angular acceleration of the platform cabin obtained by fitting is
Then there is
Thereby obtaining the identification result of the rotational inertia of the two cabins as
5. The utility model provides a two cabin inertia identification system of split type satellite in orbit which characterized in that includes the module:
a driving module: powering off a repeated locking mechanism between two satellite cabins, electrifying a magnetic floating actuator between the two cabins, generating a magnetic control couple with known size, and driving the two cabins to rotate;
an acquisition module: the control system keeps open loop and uncontrolled, collects the angular velocity information of the two cabins and the three axes, fits the angular velocity curve, and differentiates to obtain the curve of the three axes angular acceleration;
two cabin inertia of rotation solve the module: and according to the triangular shaft acceleration curve obtained by the acquisition module, magnetic control couple information output by the magnetic suspension actuator and three-shaft angular acceleration are combined to obtain two-cabin rotational inertia.
6. The split-type satellite on-orbit two-cabin rotational inertia identification system according to claim 5, wherein in the driving module, a plurality of magnetic levitation actuators are arranged between the two cabins and used for completing three-axis attitude control of the load cabin and relative centroid position control between the two cabins, 2 of the magnetic control couples are output, and the magnetic control couples in one direction are output each time to generate equal torque to act on the platform cabin and the load cabin and are marked as TLO。
7. The split satellite in-orbit two-compartment rotational inertia identification system of claim 6,
let the angular velocity of the platform cabin be omegabAngular velocity of load compartment is ωpSetting the two-cabin rotational inertia matrix as JbAnd JpThen there is
Angular velocity information is sensitively obtained through a gyroscope configured in two cabins, angular acceleration is obtained through linear fitting, and the fitting angular velocity of the platform cabin is recorded as omegab_NHAnd the fitting angular velocity of the load compartment is recorded as omegap_NH，
Is the angular acceleration of the platform cabin,
the angular acceleration of the load compartment is represented by a linear fitting formula
Wherein k isbI.e. the angular acceleration of the platform cabin, k, obtained by fittingpI.e. the fitted angular acceleration of the load compartment, bbAnd bpRespectively is a constant coefficient, and t is the action time of the magnetic control couple.
8. The split-type satellite in-orbit two-compartment rotational inertia identification system according to claim 7, wherein the angular acceleration of the load compartment obtained by fitting is set as
The angular acceleration of the platform cabin obtained by fitting is
Then there is
Thereby obtaining the identification result of the rotational inertia of the two cabins as
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Cited By (4)
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| CN113609579A (en) * | 2021-07-21 | 2021-11-05 | 上海机电工程研究所 | Explosive load identification method and system for initiating explosive actuator |
| CN114408220A (en) * | 2022-01-25 | 2022-04-29 | 上海卫星工程研究所 | On-track calibration method and system for force arm of magnetic suspension actuator |
| CN114919774A (en) * | 2022-05-20 | 2022-08-19 | 南京航空航天大学 | On-orbit calibration method for Lorentz force actuator of non-contact load undisturbed satellite platform |
| CN113609579B (en) * | 2021-07-21 | 2024-05-10 | 上海机电工程研究所 | Method and system for identifying explosive load of initiating explosive device |
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Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
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| CN113609579A (en) * | 2021-07-21 | 2021-11-05 | 上海机电工程研究所 | Explosive load identification method and system for initiating explosive actuator |
| CN113609579B (en) * | 2021-07-21 | 2024-05-10 | 上海机电工程研究所 | Method and system for identifying explosive load of initiating explosive device |
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| CN114919774A (en) * | 2022-05-20 | 2022-08-19 | 南京航空航天大学 | On-orbit calibration method for Lorentz force actuator of non-contact load undisturbed satellite platform |
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