CN106407619A - Linear-feedback global stabilization method for controlling limited spacecraft rendezvous control system - Google Patents
Linear-feedback global stabilization method for controlling limited spacecraft rendezvous control system Download PDFInfo
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Abstract
The invention provides a linear-feedback global stabilization method for controlling a limited spacecraft rendezvous control system and relates to a controller design method of the spacecraft rendezvous control system. A thruster arranged on a chasing spacecraft can only provide finite thrust, so that the system must face the problem of control limitation. The control quality is reduced or even the system instability is generated if the problem is ignored. Aiming at the defect of the prior art, a linear state feedback-based global stabilization control law is provided and an optimal selection scheme of control law parameters is given, so that the condition that a closed-loop system has a fastest convergence speed is ensured. The linear-feedback global stabilization method has the advantages that: firstly, the provided control law is linear and is convenient to design and implement; secondly, the global asymptotic stability of the closed-loop system is ensured; and finally the optimal parameters are given to ensure that the closed-loop system has the fastest convergence speed. The linear-feedback global stabilization method is used for the field of spacecraft rendezvous control.
Description
Technical field
The controller design method of Spacecraft Rendezvous control system of the present invention.
Background technology
Spacecraft Rendezvous docking is to realize spacecraft maintenance, intercept, enter a port, and many spacecrafts assemble on a large scale, multi satellite network
Networkization cooperates, and the basis of the related space mission such as spacefarer's rescue.The success of Spacecraft Rendezvous docking directly affects above-mentioned
The realization of task.Controller design for Spacecraft Rendezvous system has important engineering significance.
Limited thrust can only be provided due to being arranged on the carry-on thruster of pursuit, thus limited control can only be produced
Acceleration processed.If not considering this problem during controller design, thruster needs the acceleration producing to be possible to more than thrust
The peak acceleration that device itself can provide.Now, actual closed loop system will not rule desirably be operated, and this is not
But the Control platform of actual control system can be reduced, result even in the unstability of closed loop system, cause catastrophic consequence.
The controller design problem of Spacecraft Rendezvous docking system under Control constraints situation, has been obtained for domestic and international in a large number
The research of scientist, achieves some good results.Control law using these existing design of control method typically has two kinds
Form：One is Nonlinear control law, its shortcoming be control law realize more complicated it is difficult to debugging；Two is Linear Control rule, its
Shortcoming is the local stability (partial linear control law can achieve half Stabilization) that can only ensure closed loop system.By the end of mesh
Before, the Linear Control rule realizing Stabilization there is no related ends to report.
Content of the invention
The present invention is the design problem in order to solve the Spacecraft Rendezvous control system under Control constraints situation, and propose
A kind of linear feedback Stabilization method of Control constraints Spacecraft Rendezvous control system.
A kind of linear feedback Stabilization method of Control constraints Spacecraft Rendezvous control system is realized according to the following steps：
Step one：Introduce passive space vehicle orbital coordinate system o-xyz, its initial point o is located at the barycenter of passive space vehicle, x-axis
The direction flown along pursuit spacecraft along the direction of circular orbit radius, y-axis, z-axis is pointed to outside orbit plane and x-axis and y-axis
Constitute right-handed coordinate system；The Spacecraft Rendezvous system linearity model in x-y plane based on the description of C-W equation, as x and y
When thruster on direction of principal axis all works, set up the linear Feedback Control rule realizing Stabilization：
U=FX (1)
WhereinState vector, wherein x andPosition and the speed in x-axis direction respectively, y andPoint
It is not position and the speed in y-axis direction；U=[u1,u2]TIt is input vector, wherein u1And u2It is thruster respectively in x and y-axis side
The normalization acceleration producing upwards；N is the orbit angular velocity of passive space vehicle；δ1And δ2It is thruster respectively in x and y-axis side
The peak acceleration that can produce upwards；Parameter k1,k2,k3,k4,k5Meet following condition：
Wherein
Step 2：To the control law (1) obtaining in step one, set up and ensure that closed loop system has the bar of rapid convergence speed
Part；
Step 3：The Spacecraft Rendezvous system linearity model on z-axis direction based on the description of C-W equation, sets up real
The linear Feedback Control of existing Stabilization；
WhereinIt is state vector, z,It is position and the speed in z-axis direction respectively；u3It is thruster in z-axis side
The normalization acceleration producing upwards；δ3It is the peak acceleration that thruster produces in z-axis direction；f1It is any nonnegative constant；f2
It is any normal number.
Invention effect：
The present invention is directed to the excellent of the control method of Spacecraft Rendezvous control system proposition under Control constraints situation
Point has three aspects.First, the control law being proposed is linear, conveniently designs and realizes.Secondly, the overall situation of closed loop system is gradually
Nearly stability is guaranteed.Finally, the parameter giving optimum has convergence rate the fastest to ensure closed loop system.By right
Typical Spacecraft Rendezvous control system is emulated, and compared with the conventional method relatively, result shows proposed control method
Intersection task can be quickly completed, control performance is far superior to existing control method.Additionally, in simulation study, due to set
The linear controller of meter directly acts on real Spacecraft Rendezvous nonlinear model, so simulation result also shows this
The control program of bright proposition has stronger robustness to linearized stability.
Brief description
Fig. 1 position signalling curve chart under the effect of the present invention carried control law by Spacecraft Rendezvous system；
Fig. 2 rate signal curve chart under the effect of the present invention carried control law by Spacecraft Rendezvous system；
Fig. 3 control signal curve under the effect of the present invention carried control law by Spacecraft Rendezvous system；
Fig. 4 is position signalling curve chart under different control laws act on for the Spacecraft Rendezvous system；
Fig. 5 is rate signal curve chart under different control laws act on for the Spacecraft Rendezvous system.
Specific embodiment
Specific embodiment one：A kind of linear feedback Stabilization method bag of Control constraints Spacecraft Rendezvous control system
Include following steps：
Step one：Introduce passive space vehicle orbital coordinate system o-xyz, its initial point o is located at the barycenter of passive space vehicle, x-axis
The direction flown along pursuit spacecraft along the direction of circular orbit radius, y-axis, z-axis is pointed to outside orbit plane and x-axis and y-axis
Constitute right-handed coordinate system；The Spacecraft Rendezvous system linearity model in x-y plane based on the description of C-W equation, as x and y
When thruster on direction of principal axis all works, set up the linear Feedback Control rule realizing Stabilization：
U=FX (1)
WhereinState vector, wherein x andPosition and the speed in x-axis direction respectively, y and
It is position and the speed in y-axis direction respectively；U=[u1,u2]TIt is normalization input vector, wherein u1And u2It is thruster respectively in x
With on y-axis direction produce normalization acceleration,N is the orbit angular velocity of passive space vehicle；
δ1And δ2It is the peak acceleration that can produce on x with y-axis direction for the thruster respectively；Parameter k1,k2,k3,k4,k5Meet such as
Lower condition：
Wherein
Step 2：To the control law (1) obtaining in step one, set up and ensure that closed loop system has the bar of rapid convergence speed
Part：
Step 3：The Spacecraft Rendezvous system linearity model on z-axis direction based on the description of C-W equation, sets up real
The linear Feedback Control of existing Stabilization；
WhereinIt is state vector, z,It is position and the speed in z-axis direction respectively；u3It is thruster in z-axis side
The normalization acceleration producing upwards；δ3It is the peak acceleration that thruster produces in z-axis direction；f1It is any nonnegative constant；f2
It is any normal number.
Specific embodiment two：Present embodiment from unlike specific embodiment one：Set up real in described step one
Now the detailed process of the linear Feedback Control rule of Stabilization is：
Step is one by one：The foundation of system model；
Using linear C-W equation as Spacecraft Rendezvous control system mathematical model, its concrete form is：
Wherein ax,ay,azIt is mounted in respectively and pursue what spaceborne thruster produced on x, tri- coordinate axess of y, z
Acceleration,For saturation function；
σδ(·):R → [- δ, δ] is the saturation function of standard, and concrete form is
For the purpose of simplifying the description, σ (x) is made to represent σ hereinafter1(x).
Spacecraft Rendezvous process description is state vectorInitial value φ (t from non-zero0) turn
Change to state φ (tfThe process of)=0, t0Represent intersection task start time, tfRepresent intersection task and complete the moment；By formula
(4) it is full decoupled for obtaining Spacecraft Rendezvous control system controlling with control in the z-axis direction in x-y plane, because
This separately can be designed to them.
For the Spacecraft Rendezvous control system in x-y plane, its equation is：
Choose state vector X respectively and normalization input vector u is as follows：
Then formula (5) is written as the form of state space：
Wherein A, B are constant matricess, and form is as follows：
Step one two：The conversion of closed-loop model；
The closed loop system constituting for formula (1) and formula (6), introduces linear transformation χ=TX, wherein matrix T-shaped formula such as
Under：
Due to：
Therefore T is a nonsingular square matrix, then, under here conversion, intersection model (6) is rewritten as：
Wherein：
Control law (1) is written as form：
Closed loop system is：
Step one three：The inspection of closed loop system global stability；
Select positive semidefinite matrix P0：
Define following positive definite diagonal matrix：
Choose following Lyapunov function：
Wherein fiRepresent F0The i-th row, and if only if for V (χ)=0 when below three formulas are set up simultaneously：
χTP0χ=0, ρ1χTf1 Tf1χ=0, ρ1χTf2 Tf2χ=0 (16)
Due to k2>0, so ρ1>0,ρ2>0；Observation matrix：
The Principal Minor Sequences at different levels of calculating matrix (17) are as follows：
P is understood by above formula0+F0 TF0>0, meet (18) formula χ be only 0, so V (χ) is positive definite integral form, from but close
Suitable Lyapunov function.
Lyapunov function V (x) along the time-derivative of closed loop system (12) is：
Inequality σT(u)T0(u- σ (u)) >=0, T in formula0For any positive semidefinite matrix, make：
Then following formula is set up：
Calculate：
Due to k4Meet formula (2), matrix S above0It is positive definite, composite type (21), can be obtained by formula (19)：
From LaSalle invariant set principle, state vector can converge to set Ξ={ x:F0X=0 } in, in set Ξ
Closed loop system (12) is changed into：
By
Matrix is to (A0,F0) it is all observable to any μ >=0, in set Ξ, unique element is 0, therefore, closed loop system
(12) it is globally asymptotically stable.
Other steps and parameter are identical with specific embodiment one.
Specific embodiment three：Unlike one of present embodiment and specific embodiment one to two：Described step 2
Middle foundation ensures that the detailed process that closed loop system has the condition of rapid convergence speed is：
Because closed loop system (1) and (6) (or (12)) are asymptotically stable in the larges, therefore after finite time, system can work
Make in linear zone, become linear system.Now the convergence rate of system depends on the position of the set of poles λ (A+BF) of system：
λ (A+BF)=λ (nA0+nB0F0)=n λ (A0+B0F0) (26)
Wherein λ (A0+B0F0) unrelated with angular velocity n, make closed loop system have convergence rate the fastest, feedback oscillator F0It is
The optimal solution of following extreme-value problems：
λ(A0+B0F0) related to μ；Even if in the case that μ gives, above-mentioned extreme-value problem is also one and is difficult to the non-of solution
Linear non-convex optimization problem.Here only provide its locally optimal solution.
X is identical with the lifting force device configuration of y direction, then μ=1；Interval as follows：
By linear search, a locally optimal solution trying to achieve formula (29) is：
(k1,k2,k3,k4,k5)=(0.61,2.97,2.17,5.7138,1.21) (29)
The characteristic value collection of the linearized system of closed loop system is：
nλ(A0+B0F0)=n { -1.5015 ± 0.3125i, -1.5079 ± 0.1372i } (30)
One of other steps and parameter and specific embodiment one to two are identical.
Specific embodiment four：Unlike one of present embodiment and specific embodiment one to three：Described step 3
The detailed process that the linear Feedback Control of Stabilization is realized in middle foundation is：
Step 3 one：Set up Spacecraft Rendezvous system kinetics equation in the z-axis direction；
The Spacecraft Rendezvous control system (4) being described from C-W equation, x-y plane internal dynamicss equation is dynamic with z-axis
Mechanical equation is decoupling；Therefore, understand that kinetics equation is Spacecraft Rendezvous system in the z-axis direction by formula (4)：
Take state vector Z and the dominant vector u of system3Form as follows：
Then formula (31) uses following system description：
The form of Φ, Ψ is as follows：
Step 3 two：
Introduce following state transformations：
Then formula (33) is changed into following form：
Concrete form as follows：
The form of controller (3) is changed into simultaneously：
u3=H0ζ,H0=[- f1-f2] (38)
Closed loop system is write as：
Step 3 three：The inspection of closed loop system global stability；
Choose following positive definite matrix：
Q0Meet equationChoose following Lyapunov function again：
WhereinIt can be seen that W (ζ) is positive definite, it is suitable Lyapunov function；Make Π0=1, then：
Lyapunov function W (ζ) along the derivative of closed loop system (39) is：
Because matrix is to (Φ0,H0) for arbitrary f1≥0,f2>0 be observable, by LaSalle invariant set principle
Understand that closed loop system (39) is globally asymptotically stable.
One of other steps and parameter and specific embodiment one to three are identical.
Embodiment one：
The orbit altitude of hypothesis passive space vehicle is 500km, and orbit radius is R=6.8781 × 106m.If
δ1=δ2=8 × 10-3m/s2,δ3=6 × 10-3m/s2
I.e. μ=1.For the controller (3) on z-axis direction, take f1=f2=5, now the set of poles of closed loop system is λ
(Φ0+Ψ0H0)={ -2, -3 }.For the controller (1) in x-y plane, parameter therein is chosen as (29)；
It is directed to the primary nonlinear equation of Spacecraft Rendezvous control system
Emulated, here
If initial condition is
The condition responsive curve of closed loop system and controlling curve are separately recorded on Fig. 1, Fig. 2 and Fig. 3.
The intersection deadline that can be seen that closed loop system from Fig. 1 and Fig. 2 is tf1=8000s ≈ 1.4Tp, T herepIt is
The orbital period of passive space vehicle；In addition, from simulation result it is also seen that in the presence of this control law, holding in initial time
Row device is all saturation.This explanation closed loop system presents the nonlinear characteristic of essence.Additionally, for used by controller design
Nonlinear model (44) used by inearized model (4) and emulation is visibly different, so simulation result also illustrates the present invention
The control program being proposed has stronger robustness to linearized stability.
Finally, by control law proposed by the invention and document [B.Zhou, Q.Wang, Z.Lin and G.Duan, Gain
scheduled control of linear systems subject to actuator saturation with
application to spacecraft rendezvous,IEEE Transactions on Control Systems
Technology, Vol.22, No.5, pp.2031 2038,2014] in propose linear gain scheduling controlling rule and document
[B.Zhou,N.G.Cui and G.Duan,Circular orbital rendezvous with actuator
saturation and delay:A parametric Lyapunov equation approach,IET Control
Theory＆Applications, Vol.6, No.9, pp.1281-1287,2012] the middle linear half Stabilization control law proposing
Compare, simulation result is as shown in Figure 4 and Figure 5.The control effect that can be seen that control law proposed by the present invention from this two figure is remote
It is much better than two kinds of control methods of report in above-mentioned document.
Claims (4)
1. a kind of linear feedback Stabilization method of Control constraints Spacecraft Rendezvous control system is it is characterised in that described control
The linear feedback Stabilization method of Spacecraft Rendezvous control system limited by system comprises the following steps：
Step one：Introduce passive space vehicle orbital coordinate system o-xyz, its initial point o be located at passive space vehicle barycenter, x-axis along
The direction of circular orbit radius, the direction that y-axis is flown along pursuit spacecraft, z-axis is pointed to outside orbit plane and x-axis and y-axis composition
Right-handed coordinate system；The Spacecraft Rendezvous system linearity model in x-y plane based on the description of C-W equation, when x and y-axis side
When thruster upwards all works, set up the linear Feedback Control rule realizing Stabilization：
U=FX (1)
WhereinState vector, wherein x andPosition and the speed in x-axis direction respectively, y andIt is y respectively
Axial position and speed；U=[u1,u2]TIt is normalization input vector, wherein u1And u2It is thruster respectively in x and y-axis side
The normalization acceleration producing upwards,N is the orbit angular velocity of passive space vehicle；δ1And δ2Point
It is not the peak acceleration that can produce on x with y-axis direction for the thruster；Parameter k1,k2,k3,k4,k5Meet following condition：
Wherein
Step 2：To the control law (1) obtaining in step one, set up and ensure that closed loop system has the condition of rapid convergence speed；
Step 3：The Spacecraft Rendezvous system linearity model on z-axis direction based on the description of C-W equation, sets up and realizes entirely
Office's quelling linear Feedback Control；
WhereinIt is state vector, z,It is position and the speed in z-axis direction respectively；u3Be thruster in the z-axis direction
The normalization acceleration producing；δ3It is the peak acceleration that thruster produces in z-axis direction；f1It is any nonnegative constant；f2It is to appoint
Meaning normal number.
2. the linear feedback Stabilization side of a kind of Control constraints Spacecraft Rendezvous control system according to claim 1
Method is it is characterised in that the detailed process setting up the linear Feedback Control rule realizing Stabilization in described step one is：
Step is one by one：The foundation of system model；
Using linear C-W equation as Spacecraft Rendezvous control system mathematical model, its concrete form is：
Wherein ax,ay,azIt is mounted in respectively and pursue the acceleration that spaceborne thruster produces on x, tri- coordinate axess of y, z
Degree,For saturation function；
Spacecraft Rendezvous process description is state vectorInitial value φ (t from non-zero0) be transformed into
State φ (tfThe process of)=0, t0Represent intersection task start time, tfRepresent intersection task and complete the moment；Obtained by formula (4)
It is full decoupled to Spacecraft Rendezvous control system controlling with control in the z-axis direction in x-y plane；
For the Spacecraft Rendezvous control system in x-y plane, its equation is：
Choose state vector X respectively and normalization input vector u is as follows：
Then formula (5) is written as the form of state space：
Wherein A, B are constant matricess, and form is as follows：
Step one two：The conversion of closed-loop model；
The closed loop system constituting for formula (1) and formula (6), introduces linear transformation χ=TX, and wherein matrix T-shaped formula is as follows：
Due to：
Therefore T is a nonsingular square matrix, then, under here conversion, intersection model (6) is rewritten as：
Wherein：
Control law (1) is written as form：
Closed loop system is：
Step one three：The inspection of closed loop system global stability；
Select positive semidefinite matrix P0：
Define following positive definite diagonal matrix：
Choose following Lyapunov function：
Wherein fiRepresent F0The i-th row, V (χ)=0 and if only if below three formulas set up simultaneously：
Due to k2>0, so ρ1>0,ρ2>0；Observation matrix：
The Principal Minor Sequences at different levels of calculating matrix (17) are as follows：
From above formulaThe χ meeting (18) formula is only 0, so V (χ) is positive definite integral form；
Lyapunov function V (x) along the time-derivative of closed loop system (12) is：
Inequality σT(u)T0(u- σ (u)) >=0, T in formula0For any positive semidefinite matrix, make：
Then following formula is set up：
Calculate：
Due to k4Meet formula (2), matrix S above0It is positive definite, composite type (21), can be obtained by formula (19)：
From LaSalle invariant set principle, state vector can converge to set Ξ={ x:F0X=0 } in；Closed loop in set Ξ
System (12) is changed into：
By
Matrix is to (A0,F0) it is all observable to any μ >=0, in set Ξ, unique element is 0；Therefore, closed loop system (12)
It is globally asymptotically stable.
3. the linear feedback Stabilization side of a kind of Control constraints Spacecraft Rendezvous control system according to claim 2
It is characterised in that setting up in described step 2, method ensures that the detailed process that closed loop system has the condition of rapid convergence speed is：
Because closed loop system (1) and (6) are asymptotically stable in the larges, now the convergence rate of system depends on the set of poles of system
The position of λ (A+BF)：
λ (A+BF)=λ (nA0+nB0F0)=n λ (A0+B0F0) (26)
Wherein λ (A0+B0F0) unrelated with angular velocity n, make closed loop system have convergence rate the fastest, feedback oscillator F0It is following pole
The optimal solution of value problem：
λ(A0+B0F0) related to μ；
X is identical with the lifting force device configuration of y direction, then μ=1；Interval as follows：
By linear search, the locally optimal solution trying to achieve formula (29) is：
(k1,k2,k3,k4,k5)=(0.61,2.97,2.17,5.7138,1.21) (29)
The characteristic value collection of the linearized system of closed loop system is：
nλ(A0+B0F0)=n { -1.5015 ± 0.3125i, -1.5079 ± 0.1372i } (30).
4. the linear feedback Stabilization side of a kind of Control constraints Spacecraft Rendezvous control system according to claim 3
Method is it is characterised in that the detailed process setting up the linear Feedback Control realizing Stabilization in described step 3 is：
Step 3 one：Set up Spacecraft Rendezvous system kinetics equation in the z-axis direction；
The Spacecraft Rendezvous control system (4) being described from C-W equation, x-y plane internal dynamicss equation and kinetics in z-axis
Equation is decoupling；Therefore, understand that kinetics equation is Spacecraft Rendezvous system in the z-axis direction by formula (4)：
Take state vector Z and the dominant vector u of system3Form as follows：
Then formula (31) uses following system description：
The form of Φ, Ψ is as follows：
Step 3 two：
Introduce following state transformations：
Then formula (33) is changed into following form：
Concrete form as follows：
The form of controller (3) is changed into simultaneously：
u3=H0ζ,H0=[- f1-f2] (38)
Closed loop system is write as：
Step 3 three：The inspection of closed loop system global stability；
Choose following positive definite matrix：
Q0Meet equationChoose following Lyapunov function again：
WhereinIt can be seen that W (ζ) is positive definite, it is suitable Lyapunov function；Make Π0=1, then：
Lyapunov function W (ζ) along the derivative of closed loop system (39) is：
Because matrix is to (Φ0,H0) for arbitrary f1≥0,f2>0 be observable, from LaSalle invariant set principle
Closed loop system (39) is globally asymptotically stable.
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Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN108388134A (en) * | 2018-03-21 | 2018-08-10 | 哈尔滨工业大学 | A kind of linear feedback attitude control method of Control constraints axial symmetry spacecraft |
| CN110262225A (en) * | 2018-08-24 | 2019-09-20 | 杭州电子科技大学 | The switch controller design method of controlled space device orbital rendezvous system |
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| CN112346472A (en) * | 2020-11-20 | 2021-02-09 | 杭州电子科技大学 | Design method of cost-guaranteed robust gain scheduling controller of spacecraft rendezvous system |
| CN112882390A (en) * | 2021-01-19 | 2021-06-01 | 杭州电子科技大学 | Finite time output feedback control method for spacecraft rendezvous system |
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Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104076818A (en) * | 2014-07-02 | 2014-10-01 | 哈尔滨工业大学 | Space rendezvous system gain scheduling control method with linearization errors taken into consideration |
| CN104881036A (en) * | 2015-06-19 | 2015-09-02 | 哈尔滨工业大学 | Control-limited minisatellite three-axis magnetic moment attitude control method based on algebraic Lyapunov equation |
| CN105005651A (en) * | 2015-07-09 | 2015-10-28 | 北京航空航天大学 | Gradient separate zone optimization design method for spacecraft pulse rendezvous trajectory |
-
2016
- 2016-11-16 CN CN201611009932.8A patent/CN106407619B/en active Active
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104076818A (en) * | 2014-07-02 | 2014-10-01 | 哈尔滨工业大学 | Space rendezvous system gain scheduling control method with linearization errors taken into consideration |
| CN104881036A (en) * | 2015-06-19 | 2015-09-02 | 哈尔滨工业大学 | Control-limited minisatellite three-axis magnetic moment attitude control method based on algebraic Lyapunov equation |
| CN105005651A (en) * | 2015-07-09 | 2015-10-28 | 北京航空航天大学 | Gradient separate zone optimization design method for spacecraft pulse rendezvous trajectory |
Non-Patent Citations (2)
| Title |
|---|
| BIN ZHOU: "Global stabilization of periodic linear systems by bounded controls with applications to spacecraft magnetic attitude control", 《AUTOMATICA》 * |
| 王茜等: "输入饱和系统的离散增益调度控制及其在在轨交会中的应用", 《自动化学报》 * |
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| CN108388134B (en) * | 2018-03-21 | 2020-10-02 | 哈尔滨工业大学 | Linear feedback attitude control method for controlling limited axisymmetric spacecraft |
| CN110262225A (en) * | 2018-08-24 | 2019-09-20 | 杭州电子科技大学 | The switch controller design method of controlled space device orbital rendezvous system |
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