CN103241390A - Device and method for controlling flight attitude of micro/nano satellite - Google Patents

Device and method for controlling flight attitude of micro/nano satellite Download PDF

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Publication number
CN103241390A
CN103241390A CN2013102112562A CN201310211256A CN103241390A CN 103241390 A CN103241390 A CN 103241390A CN 2013102112562 A CN2013102112562 A CN 2013102112562A CN 201310211256 A CN201310211256 A CN 201310211256A CN 103241390 A CN103241390 A CN 103241390A
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controller
fuzzy
power density
micro
flight attitude
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CN103241390B (en
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尤政
郑伦贵
张高飞
王梦赑
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Tsinghua University
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Tsinghua University
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Abstract

The invention provides a device and a method for controlling a flight attitude of a micro/nano satellite. The device comprises a fuzzy adaptive PI (Proportional-Integral) controller, a driver, a high power density motor and a flywheel body, wherein the fuzzy adaptive PI controller is used for sending an adjustable duty ratio pulse signal to the driver; the driver is connected with the fuzzy adaptive PI controller, and used for converting the adjustable duty ratio pulse signal into an angular momentum control signal and sending the angular momentum control signal to the high power density motor; the high power density motor is connected with the driver, and used for controlling the flywheel body according to the angular momentum control signal; and the flywheel body is connected with the high power density motor, and used for outputting variable angular momentum under the driving of the high power density motor to control the flight attitude of the micro/nano satellite. The device for controlling the flight attitude of the micro/nano satellite has the advantages that the device can adaptively absorb various disturbing moments during an on-orbit period, responds at superspeed, is high in accuracy, has a zero tracking error, and can adaptively resist external disturbance.

Description

Micro-nano satellite flight attitude control setup and method
Technical field
The invention belongs to satellite attitude control technical field, be specifically related to a kind of micro-nano satellite flight attitude control setup and method.
Background technology
Along with hi-tech developments such as micrometer/nanometers, the research of 10 kilograms of following nanometer satellite technologies becomes one of focus of INSAT international satellite's technical study, various countries carry out the technical study of micro-nano satellite energetically, it can be had at aspects such as military affairs, communication, geoexploration, environment and disaster monitoring, Meteorological Services, scientific experiment, surveies of deep space more specifically use.Direct vital function is implemented and finished to attitude control as one of satellite gordian technique to aerial mission, reaction wheel/the momentum wheel of nanometer satellite appearance control actr is installed in three principal moments direction of principal axis of satellite body system of axes, by adjusting the flywheel rotating speed, adopt the momentum exchange mode to come absorbing environmental moment to the interference of satellite body, keep the satellite three axis stabilization, simultaneously can make satellite do attitude maneuver around pitching, rolling and three axles of driftage, to satisfy the needs of the work of satellite capacity weight and flight test.Nanometer satellite own vol is little, inertia is little, in light weight, be subjected to many unknown disturbance moment loadings during flight, long-time accumulation can make attitude deviation predetermined ideal state, influence is operation normally.Require little momentum wheel theoretical model convergence real system model, along with flywheel temperature, rotation speed change, nonlinearities change also takes place in friction moment thereupon, and big ratio of inertias design these factors to receive the dynamic property of satellite attitude adjustment and accurately control claim.
Summary of the invention
The present invention one of is intended to solve the problems of the technologies described above at least to a certain extent or provides a kind of useful commerce to select at least.For this reason, first purpose of the present invention is to propose a kind of micro-nano satellite flight attitude control setup, and second purpose of the present invention is to propose a kind of micro-nano satellite flight attitude control method.
According to micro-nano satellite flight attitude control setup of the present invention, comprise: fuzzy self-adaption PI controller, actuator, high power density motor and flywheel body, wherein: described fuzzy self-adaption PI controller is used for sending adjustable duty cycle pulse signal to described actuator; Described actuator links to each other with described fuzzy self-adaption PI controller, is used for described adjustable duty cycle pulse signal is converted to the moment of momentum control signal, and is sent to described high power density motor; Described high power density motor links to each other with described actuator, is used for controlling described flywheel body according to described moment of momentum control signal; Described flywheel body links to each other with described high power density motor, is used for exporting variable angular momentum and controls described micro-nano satellite flight attitude under the driving of described high power density motor.
Preferably, described actuator also is used for gathering torque voltage signal and the protection voltage signal of described high power density motor, and is sent to described fuzzy self-adaption PI controller.
Preferably, described fuzzy self-adaption PI controller comprises: speed measuring module, comparator, self adaptation rotational speed governor, treater, adaptive electro machine controller, comparison filter and current controller, wherein: described speed measuring module links to each other with described high power density motor, be used for gathering the tach signal of described high power density motor, and be sent to described comparator; Described comparator links to each other with described speed measuring module, is used for described tach signal and given controlling quantity are compared, and obtains speed error, and is sent to described self adaptation rotational speed governor; Described self adaptation rotational speed governor links to each other with described comparator, is used for described speed error is converted to spin rate control quantity, and is sent to described treater; Described adaptive electro machine controller links to each other with described actuator with described high power density motor respectively, be used for to gather the motor torque characteristic electric current of described high power density motor, and is converted to the torque current controlling quantity and is sent to described treater; Described treater links to each other with described adaptive electro machine controller with described self adaptation rotational speed governor, is used for described spin rate control quantity and described torque current controlling quantity are converted to the voltage signal controlling quantity and are sent to described relatively filter; Described relatively filter links to each other with described treater with described actuator, is used for described torque voltage signal, protection voltage signal and described voltage signal controlling quantity are compared, and obtains the comparative voltage signal, and is sent to described current controller; Described current controller links to each other with described relatively filter with described adaptive electro machine controller, is used for described comparative voltage signal is converted to the comparison current signal, and is sent to described adaptive electro machine controller; Described adaptive electro machine controller is converted to described adjustable duty cycle pulse signal with described relatively current signal, and is sent to described actuator.
Micro-nano satellite flight attitude control setup of the present invention has possessed the advantages such as various distrubing moments, hyper-speed response, pinpoint accuracy, zero tracking error and self adaptation opposing external disturbance during self adaptation is absorbed in rail.
According to micro-nano satellite flight attitude control method of the present invention, be applied to the described micro-nano satellite flight attitude of first purpose of the present invention control setup, may further comprise the steps: A: the math modeling of setting up described micro-nano satellite flight attitude control setup; B: the controlling quantity that designs the fuzzy self-adaption PI controller of described micro-nano satellite flight attitude control setup.
Preferably, steps A further comprises:
A1: the every phase phase voltage of described high power density motor equals winding resistance pressure drop and winding induced potential sum, do reasonable assumption after, winding A, B, C three phasevoltage can be expressed as:
U A = Ri A + d dt ( L A i A + M AB i B + M AC i C + Ψ pm )
U B = Ri B + d dt ( L B i B + M BA i A + M BC i C + Ψ pm )
U C = Ri C + d dt ( L C i C + M CA i A + M CB i B + Ψ pm )
Wherein: U A, U BAnd U CBe phase voltage, i A, i BAnd i CBe phase current, L A, L BAnd L CBe self-induction, M AB, M BA, M AC, M CA, M BCAnd M CBBe mutual inductance, Ψ PmBe winding permanent magnetism magnetic linkage, when rotor rotated, described winding permanent magnetism magnetic linkage magnetic flux changed with angle;
A2: when described rotor position angle was a, described winding permanent magnetism magnetic linkage was:
Ψ pm = N ∫ - π 2 + a π 2 + a B ( θ ) Sdθ
N is number of turns of winding, and B (θ) is the close distribution of described rotor permanent magnet radial air gap magnetic, and S is that described winding surrounds area on the diameter of stator bore surface;
A3: described high power density motor counter potential is:
e A = d dt NS ∫ - π 2 + θ π 2 + θ B ( x ) Sdx = NSw [ B ( π 2 + θ ) - B ( - π 2 + θ ) ] ;
A4: described high power density motor adopts the Y type to connect winding current: i A+ i B+ i C=0;
A5: described high power density motor line voltage equation is:
U AB U BC U CA = R - R 0 0 R - R - R 0 R i A i B i C + L - M M - L 0 0 L - M M - L M - L 0 L - M d dt i A i B i C + e A - e B e B - e C e C - e A ;
A6: described high power density motor is operated under 120 ° of conducting mode of operation, obtains voltage equation and is:
U AB = 2 Ri + 2 ( L - M ) di dt + 2 e A
Torque equation is: T e=K Ti
K wherein TBe described high power density motor torque factor, phase current when i is stable state;
A7: the described high power density motor equation of motion is: T e - T L = J dΩ dt + B v Ω
T in the formula LBe load torque, J is described rotor moment of inertia, B vBe viscid friction coefficient.
Preferably, step B further comprises:
B1: micro-nano satellite flight attitude control setup adopts electric current loop, the two closed loop controls of der Geschwindigkeitkreis, control system choose tach signal and given controlling quantity difference be speed error E, the speed error rate of change is EC,
The fuzzy set of described speed error E correspondence is:
A=[NB?NM?NS?ZO?PS?PM?PB]
The fuzzy set of described speed error rate of change EC correspondence is:
B=[NB?NM?NS?ZO?PS?PM?PB]
Fuzz variable NB, NM, NS, ZO, PS, PM and PB represent respectively negative big, negative in, negative little, zero, just little, just neutralize honest,
The fuzzy subset
C={NB?NM?NS?ZO?PS?PM?PB};
B2: described fuzzy self-adaption PI controller is based on the PI controller, in conjunction with a kind of controller that fuzzy adaptive controller constitutes, and revises intermediate-frequency bandwidth h and the minimum resonance peak M of described control system in real time by adjusting the PI parameter equivalent Min, the pass between described fuzzy self-adaption PI controller output and the input fuzzy set is:
After the described fuzz variable of input, output and the ternary fuzzy relation of importing For:
R ~ = ∪ i [ ( A ~ × B ~ ) T 1 × C ~ i ]
Wherein, Be the fuzzy subset on the domain of speed error E Be the fuzzy subset on the domain of speed error rate of change EC Be described fuzzy self-adaption PI controller output, i.e. Δ K pWith Δ K iFuzzy subset on the domain Subscript i=1,2, wherein By fuzzy relation matrix The nm vector that constitutes, n and m are respectively the domain element number;
B3: given micro-nano satellite flight attitude control setup rotary speed instruction, obtain described speed error and speed error rate of change by feedback element after, can try to achieve described fuzzy self-adaption PI controller Δ K pWith Δ K iCorresponding output fuzzy set,
Wherein, " ο " is the cartesian product computing;
B4: obtain the precisely controlled amount of de-fuzzy as a result to described by fuzzy reasoning, obtain accurate controlling quantity by gravity model appoach,
y = Σ L = 1 M y ‾ [ μ B L ( y ‾ L ) ] Σ L = 1 M [ μ B L ( y ‾ L ) ] ;
B5: the controlling quantity that finally is applied on the described micro-nano satellite flight attitude control setup is:
K p=K p+ΔK p,K i=K i+ΔK i
Micro-nano satellite flight attitude control method of the present invention has possessed the advantages such as various distrubing moments, hyper-speed response, pinpoint accuracy, zero tracking error and self adaptation opposing external disturbance during self adaptation is absorbed in rail.
Additional aspect of the present invention and advantage part in the following description provide, and part will become obviously from the following description, or recognize by practice of the present invention.
Description of drawings
Above-mentioned and/or additional aspect of the present invention and advantage are from obviously and easily understanding becoming the description of embodiment in conjunction with following accompanying drawing, wherein:
Fig. 1 is the constructional drawing of the micro-nano satellite flight attitude control setup of the embodiment of the invention;
Fig. 2 is the optimum flywheel main body structure of the micro-nano satellite flight attitude control setup figure of the embodiment of the invention;
Fig. 3 is the constructional drawing of the micro-nano satellite flight attitude control setup fuzzy self-adaption PI controller of the embodiment of the invention;
Fig. 4 is the diagram of circuit of the micro-nano satellite flight attitude control method of the embodiment of the invention;
Fig. 5 is the schematic diagram of the micro-nano satellite flight attitude control method of the embodiment of the invention;
Fig. 6 is the fuzzy self-adaption PI controller figure of the micro-nano satellite flight attitude control method of the embodiment of the invention;
Fig. 7 is the response curve of the micro-nano satellite flight attitude control method fuzzy self-adaption PI controller of the embodiment of the invention;
Fig. 8 is the control surface chart of the micro-nano satellite flight attitude control method fuzzy self-adaption PI controller of the embodiment of the invention;
Fig. 9 is the analogous diagram of the micro-nano satellite flight attitude control method of the embodiment of the invention;
Figure 10 is the response curve of flywheel body of the micro-nano satellite flight attitude control method of the embodiment of the invention;
Figure 11 is the steady state error figure of flywheel body of the micro-nano satellite flight attitude control method of the embodiment of the invention;
Figure 12 is the Disturbance Rejection design sketch of flywheel body of the micro-nano satellite flight attitude control method of the embodiment of the invention.
The specific embodiment
Describe embodiments of the invention below in detail, the example of described embodiment is shown in the drawings, and wherein identical or similar label is represented identical or similar elements or the element with identical or similar functions from start to finish.Be exemplary below by the embodiment that is described with reference to the drawings, be intended to for explaining the present invention, and can not be interpreted as limitation of the present invention.
In description of the invention, it will be appreciated that, term " " center "; " vertically "; " laterally "; " length "; " width "; " thickness ", " on ", D score, " preceding ", " back ", " left side ", " right side ", " vertically ", " level ", " top ", " end " " interior ", " outward ", " cw ", close the orientation of indications such as " conter clockwises " or position is based on orientation shown in the drawings or position relation, only be that the present invention for convenience of description and simplification are described, rather than device or the element of indication or hint indication must have specific orientation, with specific orientation structure and operation, therefore can not be interpreted as limitation of the present invention.
In addition, term " first ", " second " only are used for describing purpose, and can not be interpreted as indication or hint relative importance or the implicit quantity that indicates indicated technical characterictic.Thus, one or more these features can be expressed or impliedly be comprised to the feature that is limited with " first ", " second ".In description of the invention, the implication of " a plurality of " is two or more, unless clear and definite concrete restriction is arranged in addition.
In the present invention, unless clear and definite regulation and restriction are arranged in addition, broad understanding should be done in terms such as term " installation ", " linking to each other ", " connection ", " fixing ", for example, can be captive joint, also can be to removably connect, or connect integratedly; Can be mechanical connection, also can be to be electrically connected; Can be directly to link to each other, also can link to each other indirectly by intermediary, can be the connection of two element internals.For the ordinary skill in the art, can understand above-mentioned term concrete implication in the present invention as the case may be.
In the present invention, unless clear and definite regulation and restriction are arranged in addition, first feature second feature it " on " or D score can comprise that first and second features directly contact, can comprise that also first and second features are not directly contacts but by the contact of the additional features between them.And, first feature second feature " on ", " top " and " above " comprise first feature directly over second feature and oblique upper, or only represent that the first characteristic level height is higher than second feature.First feature second feature " under ", " below " and " below " comprise first feature under second feature and tiltedly, or only represent that the first characteristic level height is less than second feature.
As shown in Figure 1, the constructional drawing for the micro-nano satellite flight attitude control setup of the embodiment of the invention comprises fuzzy self-adaption PI controller 100, actuator 200, high power density motor 300 and flywheel body 400, wherein:
Fuzzy self-adaption PI controller 100 is used for sending adjustable duty cycle pulse signal to actuator 200; actuator 200 links to each other with fuzzy self-adaption PI controller 100; be used for adjustable duty cycle pulse signal is converted to the moment of momentum control signal; and be sent to high power density motor 300; actuator 200 also is used for gathering torque voltage signal and the protection voltage signal of high power density motor 300, and is sent to fuzzy self-adaption PI controller 100.
Changed by gravity gradient torque, solar wind radiation moment, the asymmetric many disturbance factors such as thermal radiation moment that cause of celestial body each several part temperature according to micro-nano satellite when the orbital flight, and the nonlinearities change of uncertain sudden distrubing moment and running temperature and friction moment, fuzzy self-adaption PI controller 100 and actuator 200 absorb distrubing moment, with the output of fuzzy self-adaption mode operating angle momentum control signal, real-time stabilization celestial body attitude.
High power density motor 300 links to each other with actuator 200, be used for according to moment of momentum control signal control flywheel body 400, in reduced volume and quality, keep enough power, flywheel body 400 links to each other with high power density motor 300, is used for output variable angular momentum control micro-nano satellite flight attitude under the driving of high power density motor 300.The various distrubing moments of flywheel body 400 during being absorbed in rail adaptively under the drive of high power density motor 300.
As shown in Figure 2, be the constructional drawing of the optimum flywheel body 400 of the micro-nano satellite flight attitude control setup of the embodiment of the invention.
For satisfying micro-nano satellite totally to the requirement of configuration, size, quality and the power consumption of flywheel body 400, reduce the volume of flywheel body 400 as far as possible, improve the inertia/mass ratio of flywheel body 400 simultaneously, parameters such as the configuration of flywheel body 400, size, quality are optimized design.Under above constraint condition, flywheel body 400 has adopted tray type structure as shown in Figure 2, has guaranteed quality that the intensity of flywheel body 400 makes flywheel body 400 away from rotating shaft, thereby has improved inertia/mass ratio, realizes maximum rotation inertia with the little quality of small size.Wherein h represents the height of flywheel body 400, and t represents the thickness on flywheel body 400 chassis, and R and r represent external diameter and the internal diameter of flywheel body 400 respectively.
As shown in Figure 3, constructional drawing for the micro-nano satellite flight attitude control setup fuzzy self-adaption PI controller 100 of the embodiment of the invention comprises: speed measuring module 110, comparator 120, self adaptation rotational speed governor 130, treater 140, adaptive electro machine controller 150, relatively filter 160 and current controller 170.
Speed measuring module 110 links to each other with high power density motor 300, is used for gathering the tach signal of high power density motor 300, and is sent to comparator 120.
Comparator 120 links to each other with speed measuring module 110, is used for tach signal and given controlling quantity are compared, and obtains speed error, and is sent to self adaptation rotational speed governor 130.
Self adaptation rotational speed governor 130 links to each other with comparator 120, is used for speed error is converted to spin rate control quantity, and is sent to treater 140.
Adaptive electro machine controller 150 links to each other with actuator 200 with high power density motor 300 respectively, be used for to gather the motor torque characteristic electric current of high power density motor 300, and is converted to the torque current controlling quantity and is sent to treater 140.
Treater 140 links to each other with adaptive electro machine controller 150 with self adaptation rotational speed governor 130, is used for spin rate control quantity and torque current controlling quantity are converted to the voltage signal controlling quantity and are sent to comparison filter 160.
Relatively filter 160 links to each other with treater 140 with actuator 200, is used for torque voltage signal, protection voltage signal and voltage signal controlling quantity are compared, and obtains the comparative voltage signal, and is sent to current controller 170.
Current controller 170 links to each other with comparison filter 160 with adaptive electro machine controller 150, is used for the comparative voltage signal is converted to the comparison current signal, and is sent to adaptive electric machine controller 150.
Adaptive electro machine controller 150 will compare current signal and be converted to adjustable duty cycle pulse signal, and be sent to actuator 200.
Micro-nano satellite flight attitude control setup according to the embodiment of the invention, can make micro-nano satellite break away from fast and stable attitude after the rocket injection, make micro-nano satellite at the attitude maneuver that is subjected to keep the stable of attitude under the multiple external interference moment environment during the follow-up flight and makes necessity, the embodiment of the invention is according to micro-nano satellite celestial body attitude and suffered external disturbance moment, drive optimum flywheel body by high power density motor, output angle momentum control signal control micro-nano satellite flight attitude has possessed the various distrubing moments during self adaptation is absorbed in rail, the hyper-speed response, pinpoint accuracy, characteristics such as zero tracking error and self adaptation opposing external disturbance.
As shown in Figure 4, be the diagram of circuit of the micro-nano satellite flight attitude control method of the embodiment of the invention, and in conjunction with the schematic diagram of the micro-nano satellite flight attitude control method of the embodiment of the invention shown in Figure 5, may further comprise the steps:
A: the math modeling of setting up micro-nano satellite flight attitude control setup.
Steps A further comprises:
Micro-nano satellite flight attitude control setup is made up of fuzzy self-adaption PI controller, actuator, high power density motor and flywheel body, wherein, and when high power density motor is preferably the FaulHarber-2036B permanent-magnet brushless DC electric machine:
A1: the every phase phase voltage of high power density motor equals winding resistance pressure drop and winding induced potential sum, do reasonable assumption after, winding A, B, C three phasevoltage can be expressed as:
U A = Ri A + d dt ( L A i A + M AB i B + M AC i C + Ψ pm )
U B = Ri B + d dt ( L B i B + M BA i A + M BC i C + Ψ pm )
U C = Ri C + d dt ( L C i C + M CA i A + M CB i B + Ψ pm )
Wherein: U A, U BAnd U CBe phase voltage, i A, i BAnd i CBe phase current, L A, L BAnd L CBe self-induction, M AB, M BA, M AC, M CA, M BCAnd M CBBe mutual inductance, Ψ PmBe winding permanent magnetism magnetic linkage, when rotor rotated, winding permanent magnetism magnetic linkage magnetic flux changed with angle.
A2: when rotor position angle was a, winding permanent magnetism magnetic linkage was:
Ψ pm = N ∫ - π 2 + a π 2 + a B ( θ ) Sdθ
N is number of turns of winding, and B (θ) is the close distribution of rotor permanent magnet radial air gap magnetic, and S is that winding surrounds area on the diameter of stator bore surface.
A3: the high power density motor counter potential is:
e A = d dt NS ∫ - π 2 + θ π 2 + θ B ( x ) Sdx = NSw [ B ( π 2 + θ ) - B ( - π 2 + θ ) ] .
A4: high power density motor FaulHarber-2036B adopts the Y type to connect winding current: i A+ i B+ i C=0.
A5: high power density motor line voltage equation is:
U AB U BC U CA = R - R 0 0 R - R - R 0 R i A i B i C + L - M M - L 0 0 L - M M - L M - L 0 L - M d dt i A i B i C + e A - e B e B - e C e C - e A .
A6: high power density motor is operated under 120 ° of conducting mode of operation, obtains voltage equation and is:
U AB = 2 Ri + 2 ( L - M ) di dt + 2 e A
Torque equation is: T e=K Ti
K wherein TBe the high power density motor torque factor, phase current when i is stable state.
A7: the high power density motor equation of motion is:
T in the formula LBe load torque, J is rotor moment of inertia, B vBe viscid friction coefficient.
As the above analysis, as load torque T LChange or along with rotation speed change and temperature traverse viscid friction coefficient B vAlso present nonlinearities change.
Further, UC1625 motor fuzzy self-adaption PI controller and driver model can equivalence be first order inertial loop by analysis:
G ( s ) = K Ts + 1
B: the controlling quantity of the fuzzy self-adaption PI controller of design micro-nano satellite flight attitude control setup.
Micro-nano satellite flight attitude control setup adopts electric current loop, the two closed loop controls of der Geschwindigkeitkreis.For the employed permanent-magnet brushless DC electric machine of the high power density motor of the embodiment of the invention, this controlled object is the model of non-linear a, multivariate, close coupling.Work under perturbation action, traditional PI controller difficulty reaches ideal effect, the adaptive controller of dynamically-adjusting parameter when the characteristics of micro-nano satellite flight attitude control setup self and working environment need to have determined to adopt to change according to working environment.
Fuzzy self-adaption PI controller is based on the PI controller, in conjunction with a kind of controller of fuzzy adaptive controller formation.Revise intermediate-frequency bandwidth h and the minimum resonance peak M of momentum wheel system in real time by adjusting the PI parameter equivalent Min, improve big ratio of inertias system dynamic characteristic.Be applicable to highly non-linear, parameter with the operation point change is big, cross-coupled is serious, environmental factor disturb strong, math modeling is changeable or uncertain controling environment.Fuzzy self-adaption PI controller as shown in Figure 6.
B1: micro-nano satellite flight attitude control setup adopts electric current loop, the two closed loop controls of der Geschwindigkeitkreis, control system choose tach signal and given controlling quantity difference be speed error E, the speed error rate of change is EC, the domain of speed error E is [6000,6000], the quantizing factor K of speed error E E=0.001083, the domain of speed error rate of change EC is [40000,40000], the quantizing factor K of speed error rate of change EC EC=0.0001625,
The fuzzy set of speed error E correspondence is:
A=[NB?NM?NS?ZO?PS?PM?PB]
The fuzzy set of speed error rate of change EC correspondence is:
B=[NB?NM?NS?ZO?PS?PM?PB]
Fuzz variable NB, NM, NS, ZO, PS, PM and PB represent respectively negative big, negative in, negative little, zero, just little, just neutralize honest,
Fuzzy self-adaption PI controller output Δ K pWith Δ K iDomain on the fuzzy set is
{-6?-5?-4?-3?-2?-1?0?1?2?3?4?5?6}
The fuzzy subset
C={NB?NM?NS?ZO?PS?PM?PB}
As shown in Figure 7, being the response curve of the micro-nano satellite flight attitude control method fuzzy self-adaption PI controller of the embodiment of the invention, is to improve dynamic performance and steady-state behaviour, and response curve is at OA section proportionality coefficient K pShould be first big after small, integral coefficient K iBig behind the Ying Xianxiao, in order to export very fast convergence stable state.AB section proportionality coefficient K pShould increase gradually, error is eliminated as early as possible, integral coefficient K iShould reduce slowly to prevent that integration is saturated.BC section proportionality coefficient K pShould be less and reduce gradually with error, integral coefficient should increase.CD section proportionality coefficient K pIncrease integral coefficient K gradually iReduce gradually.DE section proportionality coefficient K pShould reduce gradually, integral coefficient should increase gradually.Last K pAnd K iTend towards stability.
B2: a kind of controller that fuzzy self-adaption PI controller is based on the PI controller, constitutes in conjunction with fuzzy adaptive controller, by adjusting intermediate-frequency bandwidth h and the minimum resonance peak M of the real-time Correction and Control of PI parameter equivalent system Min, the pass between the output of fuzzy self-adaption PI controller and the input fuzzy set is:
After the input fuzz variable, output and the ternary fuzzy relation of importing For:
R ~ = ∪ i [ ( A ~ × B ~ ) T 1 × C ~ i ]
Wherein, Be the fuzzy subset on the domain of speed error E Be the fuzzy subset on the domain of speed error rate of change EC Be the output of fuzzy self-adaption PI controller, i.e. Δ K pWith Δ K iFuzzy subset on the domain Subscript i=1,2, wherein By fuzzy relation matrix The nm vector that constitutes, n and m are respectively the domain element number.
B3: given micro-nano satellite flight attitude control setup rotary speed instruction, behind feedback element acquisition speed error and speed error rate of change, can try to achieve fuzzy self-adaption PI controller Δ K pWith Δ K iCorresponding output fuzzy set,
Wherein, " ο " is the cartesian product computing, as shown in Figure 8, is the control surface chart of the micro-nano satellite flight attitude control method fuzzy self-adaption PI controller of the embodiment of the invention.
B4: to obtained the precisely controlled amount of de-fuzzy as a result by fuzzy reasoning, obtain accurate controlling quantity by gravity model appoach,
y = Σ L = 1 M y ‾ [ μ B L ( y ‾ L ) ] Σ L = 1 M [ μ B L ( y ‾ L ) ]
B5: the controlling quantity that finally is applied on the micro-nano satellite flight attitude control setup is:
K p=K p+ΔK p,K i=K i+ΔK i
As shown in Figure 9, be the analogous diagram of the micro-nano satellite flight attitude control method of the embodiment of the invention, certain is received satellite fine motion amount wheel system, parameter is as follows: high power density motor rotor moment of inertia J=1.95gcm 2, Rotary Inertia of Flywheel J '=1.34 * 10 -4Kgm 2, high power density motor phase resistance R=3.4 Ω, high power density motor self-induction L A=L B=L C=148uH, FaulHarber-2036B velocity constant k n=1506rpm/V, EMF(Electromotive Force, electro-motive force) constant k E=0.664mV/rpm, high power density motor torque factor k T=6.34mNm/A, mechanical time constant τ m=16ms, namely high power density motor rotating speed under rated voltage and zero load reaches 63% needed time of rated speed of rotation, actuator equivalent time constant τ=0.0148ms.
Be the analogous diagram of the micro-nano satellite flight attitude control method of the embodiment of the invention as shown in Figure 9, NS-2 flywheel body operating range wherein is 0~6000rpm, when the given rotating speed controlling quantity is 6024rpm constantly from t=5s, flywheel body response curve as shown in figure 10,1. curve is flywheel body step response curve under the no controlled reset, curve is 2. for adopting the flywheel body step response curve of traditional PI controller, 3. for adopting the flywheel body step response curve of fuzzy self-adaption PI controller, the flywheel body adopts fuzzy self-adaption PI controller response time littler than weak point and the overshoot of traditional PI controller to curve.As shown in figure 11, the rotating speed of locating at 7 seconds is 6024.1rpm, and the steady state error of this moment is 0.00166%, reaches the designing requirement of steady state error 1/6000rpm, and response time is about 0.2s.Apply a disturbing signal constantly at 7.1s, the disturbed appearance fluctuation of flywheel body rotating speed, fuzzy self-adaption PI controller rapid Adjustment System when disturbance occurring, with with respect to traditional PI controller faster speed, disturbed output amplitude is tending towards stable state, its simulated effect as shown in figure 12, curve is flywheel body response curve when adopting the traditional PI controller to be subjected to disturbance 1., curve is flywheel body response curve when adopting fuzzy self-adaption PI controller to be subjected to disturbance 2..
Micro-nano satellite flight attitude control method according to the embodiment of the invention, fuzzy self-adaption PI controller can make flywheel body super fast response, and response time obtains extra small steady state error simultaneously smaller or equal to 0.2s, steady state error is 0.00166%, satisfies designing requirements such as high anti-interference high precision.
Micro-nano satellite flight attitude control method according to the embodiment of the invention, can make micro-nano satellite break away from fast and stable attitude after the rocket injection, make micro-nano satellite at the attitude maneuver that is subjected to keep the stable of attitude under the multiple external interference moment environment during the follow-up flight and makes necessity, the embodiment of the invention is according to micro-nano satellite celestial body attitude and suffered external disturbance moment, drive optimum flywheel body by high power density motor, output angle momentum control signal control micro-nano satellite flight attitude has possessed the various distrubing moments during self adaptation is absorbed in rail, the hyper-speed response, pinpoint accuracy, characteristics such as zero tracking error and self adaptation opposing external disturbance.
Describe and to be understood that in the diagram of circuit or in this any process of otherwise describing or method, expression comprises module, fragment or the part of code of the executable instruction of the step that one or more is used to realize specific logical function or process, and the scope of preferred implementation of the present invention comprises other realization, wherein can be not according to order shown or that discuss, comprise according to related function by the mode of basic while or by opposite order, carry out function, this should be understood by the embodiments of the invention person of ordinary skill in the field.
In the description of this specification sheets, concrete feature, structure, material or characteristics that the description of reference term " embodiment ", " some embodiment ", " example ", " concrete example " or " some examples " etc. means in conjunction with this embodiment or example description are contained at least one embodiment of the present invention or the example.In this manual, the schematic statement to above-mentioned term not necessarily refers to identical embodiment or example.And concrete feature, structure, material or the characteristics of description can be with the suitable manner combination in any one or more embodiment or example.
Although illustrated and described embodiments of the invention above, be understandable that, above-described embodiment is exemplary, can not be interpreted as limitation of the present invention, those of ordinary skill in the art can change above-described embodiment under the situation that does not break away from principle of the present invention and aim within the scope of the invention, modification, replacement and modification.

Claims (6)

1. a micro-nano satellite flight attitude control setup is characterized in that, comprising: fuzzy self-adaption PI controller, actuator, high power density motor and flywheel body, wherein:
Described fuzzy self-adaption PI controller is used for sending adjustable duty cycle pulse signal to described actuator;
Described actuator links to each other with described fuzzy self-adaption PI controller, is used for described adjustable duty cycle pulse signal is converted to the moment of momentum control signal, and is sent to described high power density motor;
Described high power density motor links to each other with described actuator, is used for controlling described flywheel body according to described moment of momentum control signal;
Described flywheel body links to each other with described high power density motor, is used for exporting variable angular momentum and controls described micro-nano satellite flight attitude under the driving of described high power density motor.
2. micro-nano satellite flight attitude control setup as claimed in claim 1 is characterized in that, described actuator also is used for gathering torque voltage signal and the protection voltage signal of described high power density motor, and is sent to described fuzzy self-adaption PI controller.
3. as claim 1 and 2 described micro-nano satellite flight attitude control setups, it is characterized in that, described fuzzy self-adaption PI controller comprises: speed measuring module, comparator, self adaptation rotational speed governor, treater, adaptive electro machine controller, comparison filter and current controller, wherein:
Described speed measuring module links to each other with described high power density motor, is used for gathering the tach signal of described high power density motor, and is sent to described comparator;
Described comparator links to each other with described speed measuring module, is used for described tach signal and given controlling quantity are compared, and obtains speed error, and is sent to described self adaptation rotational speed governor;
Described self adaptation rotational speed governor links to each other with described comparator, is used for described speed error is converted to spin rate control quantity, and is sent to described treater;
Described adaptive electro machine controller links to each other with described actuator with described high power density motor respectively, be used for to gather the motor torque characteristic electric current of described high power density motor, and is converted to the torque current controlling quantity and is sent to described treater;
Described treater links to each other with described adaptive electro machine controller with described self adaptation rotational speed governor, is used for described spin rate control quantity and described torque current controlling quantity are converted to the voltage signal controlling quantity and are sent to described relatively filter;
Described relatively filter links to each other with described treater with described actuator, is used for described torque voltage signal, protection voltage signal and described voltage signal controlling quantity are compared, and obtains the comparative voltage signal, and is sent to described current controller;
Described current controller links to each other with described relatively filter with described adaptive electro machine controller, is used for described comparative voltage signal is converted to the comparison current signal, and is sent to described adaptive electro machine controller;
Described adaptive electro machine controller is converted to described adjustable duty cycle pulse signal with described relatively current signal, and is sent to described actuator.
4. a micro-nano satellite flight attitude control method is characterized in that, is applied to the micro-nano satellite flight attitude control setup described in the claim 1-3, may further comprise the steps:
A: the math modeling of setting up described micro-nano satellite flight attitude control setup;
B: the controlling quantity that designs the fuzzy self-adaption PI controller of described micro-nano satellite flight attitude control setup.
5. micro-nano satellite flight attitude control method as claimed in claim 4 is characterized in that steps A further comprises:
A1: the every phase phase voltage of described high power density motor equals winding resistance pressure drop and winding induced potential sum, do reasonable assumption after, winding A, B, C three phasevoltage can be expressed as:
U A = Ri A + d dt ( L A i A + M AB i B + M AC i C + Ψ pm )
U B = Ri B + d dt ( L B i B + M BA i A + M BC i C + Ψ pm )
U C = Ri C + d dt ( L C i C + M CA i A + M CB i B + Ψ pm )
Wherein: U A, U BAnd U CBe phase voltage, i A, i BAnd i CBe phase current, L A, L BAnd L CBe self-induction, M AB, M BA, M AC, M CA, M BCAnd M CBBe mutual inductance, Ψ PmBe winding permanent magnetism magnetic linkage, when rotor rotated, described winding permanent magnetism magnetic linkage magnetic flux changed with angle;
A2: when described rotor position angle was a, described winding permanent magnetism magnetic linkage was:
Ψ pm = N ∫ - π 2 + a π 2 + a B ( θ ) Sdθ
N is number of turns of winding, and B (θ) is the close distribution of described rotor permanent magnet radial air gap magnetic, and S is that described winding surrounds area on the diameter of stator bore surface;
A3: described high power density motor counter potential is:
e A = d dt NS ∫ - π 2 + θ π 2 + θ B ( x ) Sdx = NSw [ B ( π 2 + θ ) - B ( - π 2 + θ ) ] ;
A4: described high power density motor adopts the Y type to connect winding current: i A+ i B+ i C=0;
A5: described high power density motor line voltage equation is:
U AB U BC U CA = R - R 0 0 R - R - R 0 R i A i B i C + L - M M - L 0 0 L - M M - L M - L 0 L - M d dt i A i B i C + e A - e B e B - e C e C - e A ;
A6: described high power density motor is operated under 120 ° of conducting mode of operation, obtains voltage equation and is:
U AB = 2 Ri + 2 ( L - M ) di dt + 2 e A
Torque equation is: T e=K Ti
K wherein TBe described high power density motor torque factor, phase current when i is stable state;
A7: the described high power density motor equation of motion is:
T in the formula LBe load torque, J is described rotor moment of inertia, B vBe viscid friction coefficient.
6. as claim 4 and 5 described micro-nano satellite flight attitude control methods, it is characterized in that step B further comprises:
B1: micro-nano satellite flight attitude control setup adopts electric current loop, the two closed loop controls of der Geschwindigkeitkreis, control system choose tach signal and given controlling quantity difference be speed error E, the speed error rate of change is EC,
The fuzzy set of described speed error E correspondence is:
A=[NB?NM?NS?ZO?PS?PM?PB]
The fuzzy set of described speed error rate of change EC correspondence is:
B=[NB?NM?NS?ZO?PS?PM?PB]
Fuzz variable NB, NM, NS, ZO, PS, PM and PB represent respectively negative big, negative in, negative little, zero, just little, just neutralize honest,
The fuzzy subset
C={NB?NM?NS?ZO?PS?PM?PB};
B2: described fuzzy self-adaption PI controller is based on the PI controller, in conjunction with a kind of controller that fuzzy adaptive controller constitutes, and revises intermediate-frequency bandwidth h and the minimum resonance peak M of described control system in real time by adjusting the PI parameter equivalent Min, the pass between described fuzzy self-adaption PI controller output and the input fuzzy set is:
After the described fuzz variable of input, output and the ternary fuzzy relation of importing For:
R ~ = ∪ i [ ( A ~ × B ~ ) T 1 × C ~ i ]
Wherein, Be the fuzzy subset on the domain of speed error E Be the fuzzy subset on the domain of speed error rate of change EC Be described fuzzy self-adaption PI controller output, i.e. Δ K pWith Δ K iFuzzy subset on the domain Subscript i=1,2, wherein By fuzzy relation matrix The nm vector that constitutes, n and m are respectively the domain element number;
B3: given micro-nano satellite flight attitude control setup rotary speed instruction, obtain described speed error and speed error rate of change by feedback element after, can try to achieve described fuzzy self-adaption PI controller Δ K pWith Δ K iCorresponding output fuzzy set,
Wherein, " ο " is the cartesian product computing;
B4: obtain the precisely controlled amount of de-fuzzy as a result to described by fuzzy reasoning, obtain accurate controlling quantity by gravity model appoach,
y = Σ L = 1 M y ‾ [ μ B L ( y ‾ L ) ] Σ L = 1 M [ μ B L ( y ‾ L ) ] ;
B5: the controlling quantity that finally is applied on the described micro-nano satellite flight attitude control setup is:
K p=K p+ΔK p,K i=K i+ΔK i
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Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN104443433A (en) * 2014-11-06 2015-03-25 北京控制工程研究所 Variable-period control method of satellite emergency system
CN105048896A (en) * 2015-07-08 2015-11-11 河南科技大学 Brushless DC motor direct torque adaptive fuzzy control method
CN106251763A (en) * 2016-08-25 2016-12-21 中国人民解放军国防科学技术大学 A kind of flywheel demo system and demenstration method thereof
CN106872102A (en) * 2016-12-28 2017-06-20 中国科学院长春光学精密机械与物理研究所 The telescope shafting parameter identification method and device of DC motor Driver
CN107628273A (en) * 2017-09-27 2018-01-26 上海航天控制技术研究所 A kind of satellite attitude control method based on variable controlling cycle
CN108681310A (en) * 2018-05-14 2018-10-19 西安交通大学 A kind of controller accelerating start and stop towards mechanical main shaft height
CN109178345A (en) * 2018-09-29 2019-01-11 北京控制工程研究所 A kind of holder direction and celestial body posture cooperative control method for aerial tracking of maneuvering target
CN109823571A (en) * 2019-01-23 2019-05-31 清华大学 A kind of multistage attitude control method of remote sensing micro-nano satellite
CN110209190A (en) * 2019-03-01 2019-09-06 苏州纳飞卫星动力科技有限公司 A kind of method of the unbiased flight control of satellite nominal track

Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101800505A (en) * 2010-03-12 2010-08-11 北京航空航天大学 Method for controlling rotary speed of magnetically suspended flywheel

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101800505A (en) * 2010-03-12 2010-08-11 北京航空航天大学 Method for controlling rotary speed of magnetically suspended flywheel

Non-Patent Citations (3)

* Cited by examiner, † Cited by third party
Title
杨宁等: "高精度飞轮控制系统方案分析研究", 《航天控制》, vol. 22, no. 03, 30 June 2004 (2004-06-30) *
林伟杰: "模糊自适应PI控制永磁同步电机交流伺服系统", 《中小型电机》, vol. 32, no. 03, 31 December 2005 (2005-12-31), pages 11 - 1 *
王蜀泉: "基于模糊控制的卫星姿态控制方法研究", 《中国优秀硕士学位论文全文数据库工程科技Ⅱ辑》, no. 07, 15 November 2005 (2005-11-15) *

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CN107628273B (en) * 2017-09-27 2019-06-04 上海航天控制技术研究所 A kind of satellite attitude control method based on the variable control period
CN108681310A (en) * 2018-05-14 2018-10-19 西安交通大学 A kind of controller accelerating start and stop towards mechanical main shaft height
CN109178345A (en) * 2018-09-29 2019-01-11 北京控制工程研究所 A kind of holder direction and celestial body posture cooperative control method for aerial tracking of maneuvering target
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CN109823571B (en) * 2019-01-23 2020-07-03 清华大学 Multi-stage attitude control method for remote sensing micro-nano satellite
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