CN113985905A - Non-unwinding attitude control method considering actuator saturation and faults - Google Patents

Non-unwinding attitude control method considering actuator saturation and faults Download PDF

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CN113985905A
CN113985905A CN202111149081.8A CN202111149081A CN113985905A CN 113985905 A CN113985905 A CN 113985905A CN 202111149081 A CN202111149081 A CN 202111149081A CN 113985905 A CN113985905 A CN 113985905A
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actuator
attitude
spacecraft
unwinding
representing
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王增
刘长杰
王畅
陶玙
王春阳
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Xian Technological University
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    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
    • G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
    • G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
    • G05D1/0816—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability
    • G05D1/0833—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability using limited authority control

Abstract

The invention provides a non-unwinding attitude control method considering actuator saturation and faults, which solves the problems of unwinding and out-of-control risks of a spacecraft in the prior art, and reduces the energy consumption of attitude control and the out-of-control risks of attitude control. The invention comprises the following steps: step 1, establishing an attitude control model of a spacecraft;
Figure DDA0003286507560000011
Figure DDA0003286507560000013
wherein (q)v,q4)∈R3The x R represents the orientation of the spacecraft body described by the unit quaternion relative to an inertial coordinate system, and the x R satisfy the requirement
Figure DDA0003286507560000012
And 2, designing a non-unwinding attitude controller considering actuator saturation and faults.

Description

Non-unwinding attitude control method considering actuator saturation and faults
The technical field is as follows:
the invention belongs to the technical field of advanced control, and relates to a non-unwinding attitude control method considering actuator saturation and faults.
Background art:
a spacecraft described by a quaternion method belongs to a multi-input multi-output nonlinear equation system, and has two balance points, namely (0,0,0 +/-1)T. When the initial state of the spacecraft approaches the balance point (0,0,0, -1)TIn time, the traditional attitude control technology can cause the spacecraft to converge to the balance point (0,0,0, -1) in a longer pathTCausing waste of excess energy consumption, i.e. the problem of unwinding of the spacecraft. The existing attitude control technology does not consider the unwinding problem of the spacecraft, causes energy consumption waste of attitude control of the spacecraft, does not consider the problems of actuator saturation and faults, and causes out-of-control risk increase.
The invention content is as follows:
the invention aims to provide a non-unwinding attitude control method considering actuator saturation and faults, which solves the problems of unwinding and out-of-control risks of a spacecraft in the prior art, and provides a non-unwinding attitude control method considering actuator saturation and faults, so that the energy consumption of attitude control is reduced, and the out-of-control risk of attitude control is reduced.
In order to achieve the purpose, the invention adopts the technical scheme that:
a non-unwinding attitude control method considering actuator saturation and faults is characterized in that: the method comprises the following steps:
step 1, based on a unit quaternion method, considering external interference torque, and establishing an attitude control model of the rigid-body spacecraft:
Figure BDA0003286507540000021
Figure BDA0003286507540000022
wherein (q)v,q4)∈R3The x R represents the orientation of the spacecraft body described by the unit quaternion relative to an inertial coordinate system, and the x R satisfy the requirement
Figure BDA0003286507540000023
ω∈R3Representing the angular velocity of the spacecraft, I3∈R3×3Represents an identity matrix, J ∈ R3×3Representing a positively-defined symmetric inertial matrix, u ∈ R3And d ∈ R3Representing the control moment and the external disturbance, Γ ═ diag (γ), respectively1(t),γ2(t),γ3(t))∈R3×3Characterizing actuator faults, each diagonal element satisfying γ0≤γi(t) is less than or equal to 1, (i is 1,2,3) characterizes the working efficiency of the actuator, gamma0Is a normal number, γi(t) ═ 1 indicates no actuator failure, 0<γi(t)<1 characterise the ith actuator for partial loss of efficiency, (×)×∈R3×3The representative antisymmetric matrix is expressed in the form:
Figure BDA0003286507540000024
step 2, designing an attitude controller without unwinding problems considering actuator saturation and faults: the attitude controller is as follows:
ui＝u1i+u2i (5)
Figure BDA0003286507540000025
Figure BDA0003286507540000031
Figure BDA0003286507540000032
S＝ωe+hγ2ev (9)
where sign (x) represents a sign function, k1Where α, β, δ are all normal numbers, H ∈ H { -1,1} is an auxiliary variable, H is a constant, and H is a constant+Continuous and hopping sets are defined as
C＝{x∈S3×R3×H:he4＞-η} (11)
D＝{x∈S3×R3×H:he4≤-η} (12)
Wherein x is { q ═ qe,ωeH, η ∈ (0,1) represents the delay gap;
applications of
Figure BDA0003286507540000036
The designed control method can ensure that the output of the actuator meets the requirements
Figure BDA0003286507540000033
By adjusting the controller parameters k, k1Beta ensures that the actuator output is less than the maximum output of the actuator | umaxAnd selecting proper parameters to resist external interference and compensate the partial loss of the actuator on the premise of meeting the requirement of system stability.
The spacecraft attitude tracking control model in the step 1is based on a unit quaternion method, and the kinematics and dynamics model is as follows:
Figure BDA0003286507540000034
Figure BDA0003286507540000035
ωe＝ω-Cωd (3)
wherein ω ∈ R3、ωd∈R3And ωe∈R3Representing the actual angular velocity of the spacecraft, the desired angular velocity and the tracking error of the angular velocity,
Figure BDA0003286507540000041
for attitude tracking errors, where the vector ev＝[e1,e2,e3]∈R3Scalar e4E R, is the relative pose between the actual pose and the desired pose, and a corresponding rotation matrix C e R3×3Is defined as
Figure BDA0003286507540000042
Wherein the rotation matrix satisfies C1 and
Figure BDA0003286507540000043
I3∈R3×3represents an identity matrix, J ∈ R3×3Representing a positively-defined symmetric inertial matrix, Γ ═ diag (γ)1(t),γ2(t),γ3(t))∈R3×3Characterizing actuator faults, each diagonal element satisfying γ0≤γi(t) is less than or equal to 1, (i is 1,2,3) characterizes the working efficiency of the actuator, gamma0Is a normal number, γi(t) ═ 1 indicates no actuator failure, 0<γi(t)<1 represents the partial loss of the ith actuator efficiency, and u belongs to R3And d ∈ R3Respectively, representing the control moment and the external disturbance, (×)×∈R3×3The representative antisymmetric matrix can be expressed in the form:
Figure BDA0003286507540000044
in step 1, the desired pose is described by:
Figure BDA0003286507540000045
wherein, ω isd∈R3Representing the desired angular velocity of the spacecraft, (q)v,q4)∈R3Xr represents the desired attitude of the spacecraft described by the unit quaternion.
Compared with the prior art, the invention has the advantages and effects that:
1. compared with the existing attitude control technology, the method can avoid the problem of attitude control unwinding of the spacecraft described by the quaternion, can greatly reduce the energy consumption of attitude control, simultaneously considers the saturation and the fault of the actuator, ensures that the attitude control torque is preset below the maximum value of the actuator in advance, and reduces the out-of-control risk of the attitude control.
2. According to the invention, by adopting a hybrid system and an advanced control method of passive fault-tolerant control, considering the problems of saturation and fault of the actuator, and designing the unwinding-free attitude controller with limited control torque, the unwinding problem of attitude control can be effectively avoided, the energy consumption of attitude control is reduced, and the risk of attitude out-of-control caused by the saturation and fault of the actuator can be reduced.
Description of the drawings:
FIG. 1is an attitude control map showing unwinding problems, wherein a represents attitude tracking error convergence, b represents angular velocity tracking error convergence, c represents control torque, and d represents an enlarged view of attitude tracking error convergence showing unwinding problems;
FIG. 2 is an attitude control map without unwinding problems, where a represents attitude tracking error convergence, b represents angular velocity tracking error convergence, c represents control torque, and d represents an enlarged view of attitude tracking error convergence without unwinding problems;
FIG. 3 is a graph comparing attitude control energy consumption with and without unwinding problems, a.
FIG. 4 is a schematic flow diagram of the method of the present invention.
The specific implementation mode is as follows:
in order to make the objects, technical solutions and advantages of the present invention more apparent, the present invention is described in further detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative of the invention and are not intended to limit the invention.
Example (b):
in conjunction with the method flow diagram shown in fig. 4, the proposed unwinding-attitude-free control method considering actuator saturation and failure includes the following steps:
step 1, establishing an attitude control model of a spacecraft;
step 2, designing a non-unwinding attitude controller considering actuator saturation and faults;
the specific process of the step 1is as follows:
based on a unit quaternion method, considering external interference torque, establishing an attitude control model of the rigid-body space vehicle:
Figure BDA0003286507540000061
Figure BDA0003286507540000062
wherein (q)v,q4)∈R3The x R represents the orientation of the spacecraft body described by the unit quaternion relative to an inertial coordinate system, and the x R satisfy the requirement
Figure BDA0003286507540000063
ω∈R3Representing the angular velocity of the spacecraft, I3∈R3×3Represents an identity matrix, J ∈ R3×3Representing a positively-defined symmetric inertial matrix, u ∈ R3And d ∈ R3Representing the control moment and the external disturbance, Γ ═ diag (γ), respectively1(t),γ2(t),γ3(t))∈R3×3Characterizing actuator faults, each diagonal element satisfying γ0≤γi(t) is less than or equal to 1, (i is 1,2,3) characterizes the working efficiency of the actuator, gamma0Is a normal number, γi(t) ═ 1 indicates no actuator failure, 0<γi(t)<1 characterisation of the ith actuator for partial loss of efficiency, (. C)×∈R3×3The representative antisymmetric matrix is expressed in the form:
Figure BDA0003286507540000064
the desired pose is described by:
Figure BDA0003286507540000071
the unit quaternion described space vehicle kinematics and dynamics model has no singular point and can describe the 360-degree space vehicle attitude. However, the spacecraft described by the quaternion has two balance points, and the existing control technology does not consider the unwinding problem of attitude control, so that the energy consumption of the attitude control is wasted, and simultaneously does not consider the execution saturation and faults, so that the risk of out-of-control attitude control is increased.
Specifically, the spacecraft attitude tracking control model in the step 1is based on a unit quaternion method, and the kinematics and dynamics model thereof is
Figure BDA0003286507540000072
Figure BDA0003286507540000073
ωe＝ω-Cωd (3)
Wherein ω ∈ R3、ωd∈R3And ωe∈R3Representing the actual angular velocity of the spacecraft, the desired angular velocity and the tracking error of the angular velocity,
Figure BDA0003286507540000074
for attitude tracking errors, where the vector ev＝[e1,e2,e3]∈R3Scalar e4E R, is the relative pose between the actual pose and the desired pose. The corresponding rotation matrix C ∈ R3×3Is defined as
Figure BDA0003286507540000075
Wherein the rotation matrix satisfies C1 and
Figure BDA0003286507540000076
I3∈R3×3represents an identity matrix, J ∈ R3×3Representing a positively-defined symmetric inertial matrix, Γ ═ diag (γ)1(t),γ2(t),γ3(t))∈R3×3Characterizing actuator faults, each diagonal element satisfying γ0≤γi(t) is less than or equal to 1, (i is 1,2,3) characterizes the working efficiency of the actuator, gamma0Is a normal number, and is apparently gammai(t) ═ 1 indicates no actuator failure, 0<γi(t)<1 represents the partial loss of the ith actuator efficiency, and u belongs to R3And d ∈ R3Respectively, representing the control moment and the external disturbance, (×)×∈R3×3The representative antisymmetric matrix can be expressed in the following form
Figure BDA0003286507540000081
The specific process of the step 2 is as follows:
the attitude controller considering the saturation and the fault of the actuator in the step 2 comprises the following steps:
ui＝u1i+u2i (5)
Figure BDA0003286507540000082
Figure BDA0003286507540000083
Figure BDA0003286507540000084
S＝ωe+hγ2ev (9)
where sign (·) represents a sign function, k1Where α, β, δ are all normal numbers, H ∈ H { -1,1} is an auxiliary variable, H is a constant, and H is a constant+Continuous and hopping sets are defined as
C＝{x∈S3×R3×H:he4＞-η} (11)
D＝{x∈S3×R3×H:he4≤-η} (12)
Wherein x is { q ═ qe,ωeH, η ∈ (0,1) represents the delay gap.
Applications of
Figure BDA0003286507540000086
The designed control technology can ensure that the output of the actuator meets the requirements
Figure BDA0003286507540000085
By adjusting the controller parameters k, k1Beta ensures that the actuator output is less than the maximum output of the actuator | umaxAnd selecting proper parameters to resist external interference and compensate the partial loss of the actuator on the premise of meeting the requirement of system stability.
In order to ensure the stability of the system, the stability analysis needs to be carried out on the designed attitude controller.
Using Γ u ═ u1-(I-Γ)u1+Γu2Equation (2) can be written as
Figure BDA0003286507540000091
Defining the Lyapunov candidate function as:
Figure BDA0003286507540000092
by derivation of the above formula, the result is obtained
Figure BDA0003286507540000093
By substituting formula (16) for formula (18), the compound is obtained
Figure BDA0003286507540000094
Wherein the content of the first and second substances,
Figure BDA0003286507540000095
and is
Figure BDA0003286507540000096
Utilizing formulae (6) - (8), and
Figure BDA0003286507540000097
can obtain the product
Figure BDA0003286507540000098
Lambda is less than or equal to | inverse | by applying | |maxIs obtained by
Figure BDA0003286507540000099
Wherein the content of the first and second substances,
Figure BDA00032865075400000910
when x is equal to D, V1Jump occurs to obtain
V1(x+)-V1(x)≤4he4≤0 (22)
The system is stable, and the system state finally converges asymptotically to the equilibrium point.
FIGS. 1-3 show that the present invention is applicable to an environment in which an external disturbance torque and an actuator malfunction occurThe attitude control method provided by the invention can ensure that the spacecraft described by the quaternion can quickly converge to a balance point (0,0,0 +/-1)TI.e. initial attitude q (0) ═ 0.3, -0.2, -0.3,0.8832]TNear the balance point (0,0,0, -1)TIn time, the spacecraft will converge to the equilibrium point (0,0,0, -1) with the shortest pathT(as shown by d in FIG. 2), the 50 second mean moment 0.42NM is simulated, instead of rotating a large half-turn, to converge to the equilibrium point (0,0,0,1) with a longer pathT(as shown in d in figure 1), the simulation of the 50-second average moment is 2.27NM, so that the energy consumption of attitude control of the spacecraft is greatly reduced (as shown in the energy consumption comparison of figure 3), namely, the problem of unwinding of the spacecraft is solved.
FIG. 2c shows the presence of an external disturbance di0.01isin (t/100) and actuator failure gammaiAnd in the case of i being 1,2 and 3, the required maximum torque is 2.5NM, and the actuator can achieve stable attitude control with low energy consumption on the premise of not exceeding the maximum value of 10Nm of the actuator.
Experimental example:
in order to verify the effectiveness of the attitude stabilization controller which is designed by the patent and takes the saturation and fault of the actuator into consideration, the attitude control is carried out on the space aircraft, the effectiveness of the unwinding problem and the fault of the actuator is verified, and whether the energy consumption of the controller is reduced or not is verified. The section mainly carries out validity verification through numerical simulation, and explains the validity of a specific implementation mode and the proposed control algorithm. Assuming that the nominal inertia matrix of the rigid-body spacecraft is J ═ 201.20.9; 1.2171.4, respectively; 0.91.415]kg×m2. The initial desired attitude and angular velocity are set to q, respectivelyd(0)＝[0,0,0,1]TAnd ωd(t)＝0.05[sin(πt/100),sin(2πt/100),sin(3πt/100)]rad/s. The initial attitude and angular velocity are set to q (0) — [0.3, -0.2, -0.3,0.8832, respectively]TAnd ω (0) ═ 0.06, -0.04,0.05]Trad/s。
Suppose the maximum output torque of the actuator is umax＝[10,10,10]TNxm, actuator failure as a time-varying function gammai0.75+0.05isin (t), i 1,2,3, external interference di0.01isin (t/100). Then, by selectingSelecting a suitable control gain so that the output value of the controller is less than the maximum value of the actuator
Figure BDA0003286507540000111
The gain of the controller is selected to be k 3, k13, 0.01, 0.5, and 2, wherein the controller's energy consumption is defined as
Figure BDA0003286507540000112
The simulation adopts continuous saturation function to replace discontinuous sign function
Figure BDA0003286507540000113
Wherein the content of the first and second substances,
Figure BDA0003286507540000114
the system precision is ensured by adjusting the parameter to be an arbitrarily small normal number.
The above embodiments are merely illustrative of the principles and effects of the present invention, and it will be apparent to those skilled in the art that various changes and modifications can be made without departing from the inventive concept of the present invention, and the scope of the present invention is defined by the appended claims.

Claims (3)

1. A non-unwinding attitude control method considering actuator saturation and faults is characterized in that: the method comprises the following steps:
step 1, based on a unit quaternion method, considering external interference torque, and establishing an attitude control model of the rigid-body spacecraft:
Figure FDA0003286507530000011
Figure FDA0003286507530000012
wherein (q)v,q4)∈R3The x R represents the orientation of the spacecraft body described by the unit quaternion relative to an inertial coordinate system, and the x R satisfy the requirement
Figure FDA0003286507530000013
ω∈R3Representing the angular velocity of the spacecraft, I3∈R3×3Represents an identity matrix, J ∈ R3×3Representing a positively-defined symmetric inertial matrix, u ∈ R3And d ∈ R3Representing the control moment and the external disturbance, Γ ═ diag (γ), respectively1(t),γ2(t),γ3(t))∈R3×3Characterizing actuator faults, each diagonal element satisfying γ0≤γi(t) is less than or equal to 1, (i is 1,2,3) characterizes the working efficiency of the actuator, gamma0Is a normal number, γi(t) ═ 1 indicates no actuator failure, 0<γi(t)<1 characterisation of the ith actuator for partial loss of efficiency, (. C)×∈R3×3The representative antisymmetric matrix is expressed in the form:
Figure FDA0003286507530000014
step 2, designing an attitude controller without unwinding problems considering actuator saturation and faults: the attitude controller is as follows:
ui＝u1i+u2i (5)
Figure FDA0003286507530000021
Figure FDA0003286507530000022
Figure FDA0003286507530000023
S＝ωe+hγ2ev (9)
where sign (·) represents a sign function, k1A, β, δ are all normal numbers, H ∈ H { -1,1} is an auxiliary variable, H { -1,1} is a constant, H { -1,1 { -H } H { -H } H { -H } is a { -H } is a { -H } is a { -H } H { -H {+Continuous and hopping sets are defined as
C＝{x∈S3×R3×H:he4＞-η} (11)
D＝{x∈S3×R3×H:he4≤-η} (12)
Wherein x is { q ═ qe,ωeH, η ∈ (0,1) represents the delay gap;
applying C1,
Figure FDA0003286507530000024
|ei|≤1,h2as 1, a control method has been devised to ensure that the output of the actuator satisfies the requirements
Figure FDA0003286507530000025
By adjusting the controller parameters k, k1Beta ensures that the actuator output is less than the maximum output of the actuator | umaxAnd selecting proper parameters to resist external interference and compensate the partial loss of the actuator on the premise of meeting the requirement of system stability.
2. The unwinding-free attitude control method taking into account actuator saturation and failure according to claim 1, characterized in that:
the spacecraft attitude tracking control model in the step 1is based on a unit quaternion method, and the kinematics and dynamics model is as follows:
Figure FDA0003286507530000031
Figure FDA0003286507530000032
ωe＝ω-Cωd (3)
wherein ω ∈ R3、ωd∈R3And ωe∈R3Representing the actual angular velocity of the spacecraft, the desired angular velocity and the tracking error of the angular velocity,
Figure FDA0003286507530000033
for attitude tracking errors, where the vector ev＝[e1,e2,e3]∈R3Scalar e4E R, is the relative pose between the actual pose and the desired pose, and a corresponding rotation matrix C e R3×3Is defined as
Figure FDA0003286507530000034
Wherein the rotation matrix satisfies C1 and
Figure FDA0003286507530000035
I3∈R3×3represents an identity matrix, J ∈ R3×3Representing a positively-defined symmetric inertial matrix, Γ ═ diag (γ)1(t),γ2(t),γ3(t))∈R3×3Characterizing actuator faults, each diagonal element satisfying γ0≤γi(t) is less than or equal to 1, (i is 1,2,3) characterizes the working efficiency of the actuator, gamma0Is a normal number, γi(t) ═ 1 indicates no actuator failure, 0<γi(t)<1 represents the partial loss of the ith actuator efficiency, and u belongs to R3And d ∈ R3Respectively, representing the control moment and the external disturbance, (×)×∈R3×3The representative antisymmetric matrix can be expressed in the form:
Figure FDA0003286507530000036
3. the unwinding-free attitude control method taking into account actuator saturation and failure according to claim 1, characterized in that:
in step 1, the desired pose is described by:
Figure FDA0003286507530000037
wherein, ω isd∈R3Representing the desired angular velocity of the spacecraft, (q)v,q4)∈R3Xr represents the desired attitude of the spacecraft described by the unit quaternion.
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Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN107168357A (en) * 2017-06-30 2017-09-15 北京航空航天大学 It is a kind of to consider posture restraint and the spacecraft attitude maneuver control method of anti-unwinding
CN108646556A (en) * 2018-05-08 2018-10-12 中国人民解放军战略支援部队航天工程大学 Input saturation spacecraft is without unwinding Attitude tracking control method
CN109143846A (en) * 2018-09-25 2019-01-04 浙江工业大学 A kind of rigid aircraft adaptive neural network tracking and controlling method considering actuator constraints problem

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN107168357A (en) * 2017-06-30 2017-09-15 北京航空航天大学 It is a kind of to consider posture restraint and the spacecraft attitude maneuver control method of anti-unwinding
CN108646556A (en) * 2018-05-08 2018-10-12 中国人民解放军战略支援部队航天工程大学 Input saturation spacecraft is without unwinding Attitude tracking control method
CN109143846A (en) * 2018-09-25 2019-01-04 浙江工业大学 A kind of rigid aircraft adaptive neural network tracking and controlling method considering actuator constraints problem

Non-Patent Citations (3)

* Cited by examiner, † Cited by third party
Title
夏钰泉: "考虑驱动器饱和影响的空间飞行器输出反馈控制", CNKI博士学位论文全文库, no. 2021, 15 May 2021 (2021-05-15), pages 48 - 55 *
崔祜涛 等: "考虑有界干扰和输入饱和的航天器姿态抗退绕机动控制", 中国科学:技术科学, vol. 42, no. 9, 20 September 2012 (2012-09-20), pages 1004 - 1015 *
黄兵: "刚体航天器鲁棒抗退绕控制", CNKI博士学位论文全文库, no. 2021, 15 April 2021 (2021-04-15), pages 19 - 21 *

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