CN111177951A - Spacecraft reconfigurability evaluation method - Google Patents
Spacecraft reconfigurability evaluation method Download PDFInfo
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- CN111177951A CN111177951A CN202010071106.6A CN202010071106A CN111177951A CN 111177951 A CN111177951 A CN 111177951A CN 202010071106 A CN202010071106 A CN 202010071106A CN 111177951 A CN111177951 A CN 111177951A
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Abstract
A spacecraft reconfigurability evaluation method belongs to the general technical field of spacecrafts, and comprehensively considers the time constraint, the performance degradation degree, the all-direction reconfiguration capability distribution and the reliability degradation influence. Firstly, providing an evaluation index form based on an integral quadratic function; then, through detailed design of a weight matrix in the indexes, the reconstruction capability distribution condition of the system and the reliability reduction influence caused by faults are introduced into the evaluation indexes; finally, the effectiveness and the correctness of the evaluation method provided by the patent are verified through a simulation example. The method comprehensively considers the problems of limited energy, limited time, interference influence and other limitation constraints, is closer to the actual engineering background, and can enable designers to know the real reconstruction capability of the system under the constraint of actual performance.
Description
Technical Field
The invention relates to a spacecraft reconfigurability evaluation method, and belongs to the general technical field of spacecrafts.
Background
For the problem of spacecraft reconstruction, whether the system is subjected to reconfigurability optimization at the initial design stage or the system is subjected to online performance evaluation at the fault stage, a quantitative index capable of describing the size of the system reconfiguration capacity is required to be used as a basis for design and evaluation, so that the problems of whether redundancy allocation is reasonable, whether reconstruction measures are effective, whether a fault system can be reconfigured and the like are quantitatively evaluated.
Because the actuator is in the running state for a long time, the actuator is easy to wear and age, and faults are frequent, the research on the system reconfigurability under the condition of actuator faults is of great significance. For actuator faults, the existing method mostly starts from the perspective of residual energy controllability and measures the reconfigurability of the system. However, such methods have the following disadvantages:
(1) there is a lack of consideration of the impact of practical factors on reconfigurability, such as constraints and interference. It is well known that real physical systems tend to have a number of practical constraints, such as limited energy, time, and control inputs. For example, spacecraft control systems, which are limited in energy availability by windsurfing power generation and propellant carrying capacity; certain specific tasks (such as orbit change, landing and the like) need to be completed in a specific time window and are subject to corresponding time constraints; the control input torque is bounded due to the saturation of the flywheel speed. In addition, the spacecraft which runs on the orbit inevitably receives the effects of various interference moments in space, such as environmental interference moments including gravity gradient moment, solar radiation moment, pneumatic moment, geomagnetic moment and the like, and non-environmental interference moments including rotating moment of a movable part, friction moment inside a flywheel, driving moment of a solar cell array, coupling moment of a flexible structure and the like. Therefore, to describe the actual reconfiguration capability of a control system, in addition to the consideration of the remaining controllability, the actual problems of safe time, operating conditions, resource allocation and the like need to be considered. Otherwise, the failure system may lose the reconfigurability in a practical sense because of the excessive reconfiguration cost and the long time required even though the failure system is still controllable.
(2) It is difficult to reflect the rationality of the system reconstruction potential distribution and direct comparison of different systems is not possible. The traditional reconfigurability evaluation method mostly takes the minimum absolute value of the residual control capability of the system as an index to measure reconfigurability. However, since the expected reconstruction capability is different between different systems and different directions of the same system, and the minimum reconstruction capability is not necessarily the weakest, the method described above is difficult to reflect the rationality of the reconstruction potential distribution of the system, and cannot directly compare different systems. Based on the evaluation indexes, the optimization design effect of the system is limited, and the situation of partial redundant directions may occur, so that the quality and the cost are improved and the waste of available resources is caused.
(3) The influence of the reliability reduction on the system reconfigurability is not introduced into the evaluation index. Most of the existing researches separate the reconfigurability from the reliability for discussion, however, the reconfigurability and the reliability can be influenced by the tool belonging to the quality characteristic of the improved system. When a certain actuator fails, the reliability of the actuator is reduced, and at the moment, if the actuator undertakes an overweight control task, the diffusion of the failure is accelerated, and the actual reconfigurability of the system is weakened. Therefore, even if the failure actuator has sufficient residual controllability without considering the influence of reliability, the entire reconstruction process cannot be smoothly completed with high reliability.
Disclosure of Invention
The technical problem to be solved by the invention is as follows: the method overcomes the defects of the prior art, provides a spacecraft reconfigurability evaluation method, and comprehensively considers the time constraint, the performance degradation degree, the all-direction reconfiguration capability distribution and the reliability degradation influence. Firstly, providing an evaluation index form based on an integral quadratic function; then, through detailed design of a weight matrix in the indexes, the reconstruction capability distribution condition of the system and the reliability reduction influence caused by faults are introduced into the evaluation indexes; finally, the effectiveness and the correctness of the evaluation method provided by the patent are verified through a simulation example.
The purpose of the invention is realized by the following technical scheme:
a spacecraft reconfigurability evaluation method is characterized by comprising the following steps:
s1, establishing a spacecraft system fault model under the conditions of actuator fault, environmental interference and time limitation; the time-limited condition means that the fault occurrence time is unknown and the task completion time is determined;
s2, establishing a spacecraft reconfigurability evaluation model according to the system control precision requirement, the health state of an executing mechanism and an allowable performance threshold;
s3, utilizing H according to the spacecraft system fault model in S12The control method comprises the steps of solving the maximum value (the minimum value of J) of an evaluation parameter rho in a spacecraft reconfigurability evaluation model;
s4, based on the spacecraft reconfigurability evaluation model in S3, solving a numerical solution of indexes by using a fine integration method;
s5, if the numerical value of the reconfigurability quantization index is solved to be 0, judging that the spacecraft is not reconfigurable, and entering the step (6); if the numerical solution of the quantitative index is not 0, judging that the spacecraft is reconfigurable, outputting the quantitative reconfigurable size, and entering the step (6);
and S6, ending.
Preferably, when the spacecraft is determined to be reconfigurable, the failed actuator in S1 is switched to a backup actuator, or the failed actuator in S1 is switched from the current control mode to another control mode.
Preferably, the spacecraft system fault model described in S1 is:
wherein
In the formula Ix,Iy,IzRespectively, the three-axis moment of inertia of the satellite; x is formed by Rn、u∈Rm、
State vector, control input vector and external interference vector of the nominal system respectively; a is an element of Rn×n,Bu∈Rn×m,
Respectively a state matrix, a control input matrix and an external interference matrix; t is tfTime of occurrence of a failure, tmissetting the angle of control mechanism as α beta, setting the moment distribution matrix as phi (α beta), setting A as diag { theta } as the configuration of control mechanism1,θ2,...,θmIs the actuator failure factor matrix, θi∈[0,1],i＝1,2,...,m，θiSmaller means lower residual efficiency of the respective actuator; i is3×3Is an identity matrix; 03×3Is a zero matrix.
Preferably, the spacecraft reconfigurability evaluation model in S2 is:
wherein rho is a quantitative evaluation index, kappa is an allowable performance threshold value when the spacecraft is reconfigurable, Q and R are symmetric matrixes, and Q-Q is satisfiedT≥0，R＝RT＞0；x∈Rn、u∈RmRespectively a state vector and a control input vector of a nominal system; t is tmisIs the task completion time.
Compared with the prior art, the invention has the following beneficial effects:
(1) the reconfigurability evaluation index proposed by the existing research does not consider various practical performance constraints of spaceflight, and the obtained evaluation result cannot reflect the real reconfiguration capacity of the system, so that the reconfigurability evaluation index has no strong practical significance. Compared with the traditional method, the method comprehensively considers the problems of limited energy, limited time, interference influence and other limitation constraints, is closer to the actual engineering background, and can enable designers to know the real reconstruction capability of the system under the actual performance constraint;
(2) the existing research mainly measures the reconfigurability of the system according to the absolute size of the residual control capacity after the system fails, the reconfigurability of the system is equivalent to the residual control capacity, specific task requirements and task completion degree which can be maintained by the system are not considered, and therefore an obtained evaluation result can only show whether the system is still controllable or not and cannot reflect whether the system has enough residual capacity to complete a target task under certain constraints or not. Compared with the traditional method, the method carries out coordinate scaling on the state space according to the requirements in all directions, and the obtained new coordinates reflect the degree of satisfaction of the system to the control capability requirements in all directions; in addition, the method also performs normalization processing on the evaluation indexes based on the performance threshold value, so that systems with different magnitudes can also perform reconfigurable comparison, and the method is more universal;
(3) the existing research does not consider the weakening problem of the actual reconfigurability of the system caused by the reliability reduction of the fault part, so that the confidence coefficient of an evaluation result is reduced, and the originally determined reconfigurable system is possibly deteriorated due to the fault and completely fails before the reconfiguration task is completed, so that the original reconfigurability is lost. Compared with the traditional method, the method disclosed by the invention has the advantages that the weight is distributed to the working cost of each part according to the real-time reliability of each actuator, so that the reliability reduction influence caused by the fault is introduced into the evaluation index, the obtained result is more consistent with the service life characteristic of the part, and the confidence coefficient is higher.
Drawings
FIG. 1 is a surface showing the variation of reconfigurability with failure factors.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention more apparent, embodiments of the present invention will be described in detail with reference to the accompanying drawings.
Example 1:
a spacecraft reconfigurability evaluation method comprises the following steps:
(1) establishing a spacecraft system fault model under the conditions of actuator fault, environmental interference and time limitation
Wherein, Ix,Iy,IzRespectively, the three-axis moment of inertia of the satellite; x is formed by Rn、u∈Rm、
State vectors, control input vectors and external disturbances of the nominal system, respectively; a is an element of Rn×n,Bu∈Rn×m,
Respectively a state matrix, a control input matrix and an external interference matrix; t is tfTime of occurrence of a failure, tmissetting the angle of control mechanism as α beta, setting the moment distribution matrix as phi (α beta), setting A as diag { theta } as the configuration of control mechanism1,θ2,...,θmIs the actuator failure factor matrix, θi∈[0,1],i＝1,2,...,m，θiSmaller is indicative of lower residual efficiency of the respective actuator. I is3×3Is an identity matrix; 03×3Is a zero matrix.
Get Bf＝BuΛ, which is the equivalent control matrix after failure of the actuator, BfIs a parameter related to the installation of the actuator
And a matrix function of the degree of failure Λ
(2) Establishing a spacecraft reconfigurability evaluation model according to the system control precision requirement, the health state of an actuating mechanism and an allowable performance threshold;
wherein rho is a quantitative evaluation index,
kappa is an allowable performance threshold value when the spacecraft is reconfigurable, Q and R are symmetric matrixes, and the condition that Q is equal to QT≥0，R＝RT＞0。
And determining a weight coefficient in the reconfigurability quantization index according to the task requirement of the spacecraft and the health state of the current control mechanism.
After the fault occurs, considering the reduction of the system performance, the requirement on the control precision can be properly reduced within an allowable range, and according to the precision requirements of the system in different directions, a weight matrix Q of a state item in an index is determined:
wherein a is more than 0, the proportional coefficient for balancing the control precision and the energy consumption is epsiloni(i ═ 1, 2.. times, n) is the minimum accuracy requirement that the system needs to achieve in the ith direction, εmin＝min{ε1,ε2,...,εn}. If the minimum accuracy requirement in the i direction is properly reduced under the action of (6), the state deviation weight in the direction is correspondingly reduced.
The failure probability of the actuator increases with decreasing failure factor, resulting in reduced reliability. In order to prolong the service life of the actuators as much as possible and ensure that the system can complete the reconstruction task with high reliability, a weight matrix R is designed according to the reliability of each actuator:
R＝Wu TWu(7)
wherein λ isi(i ═ 1, 2.. times, m) is the failure probability distribution of the actuator i and
λi0is the nominal failure rate, k, of the actuator i in normal operation>0 is a scaling factor, λ, related to the actuator parameter and its loadmin＝min{λ i1, (i) ═ 1, 2. Under the action of the matrix shown in the formula (8), the output cost of the actuator which is lower in reliability is higher, and theta is higheriIs the failure factor of the ith actuator.
(3) Utilizing H according to the spacecraft system fault model in S12Solving the maximum value (the minimum value of J) of the spacecraft reconfigurability evaluation model evaluation parameter rho in the step (2);
get
Then there is
CTC＝Q,DTD＝R,DTC＝0,CTD＝0 (10)
On this basis, equation (4) can be converted into:
wherein the content of the first and second substances,
y(t)＝Cx(t)+Du(t) (12)
based on the above processing, the integral quadratic form in equation (4) is converted into L of the signal y shown in equation (11)2Norm squared form. Based on this, utilize H2Control methodThe method establishes mathematical function relations among indexes, system structure parameters, specified task time and external interference:
wherein W (t) is a positive definite symmetric time-varying matrix satisfying the following matrix differential equation:
(4) based on the mathematical expression of the reconfigurability quantization index obtained in the step (3), parameters A and B brought into the spacecraftf,BdThe index weights Q and R are used for solving a numerical solution of the reconfigurability quantization index based on a fine integral method to realize quantitative evaluation of the reconfigurability of the spacecraft;
(5) if the quantitative analysis result of the reconfigurability is 0, judging that the spacecraft under the current design is not reconfigurable, and entering the step (6); if the quantitative analysis result of the reconfigurability is not 0, judging that the spacecraft under the current design is reconfigurable, outputting the quantified reconfigurability, and entering the step (6);
(6) and (6) ending.
Example 2:
by applying the method of embodiment 1, in order to verify the effectiveness of the method provided by the patent of the present invention, a linear model of a typical satellite attitude control system is used as a specific embodiment for explanation, the satellite is a symmetric quad-tilt wheel control system, and relevant parameters are shown in table 1.
TABLE 1 satellite and orbital parameters thereof
(1) Establishing a spacecraft system model;
(2) designing quantitative evaluation indexes of the reconfigurability of the spacecraft:
ρ＝max{1-minJ,0}
wherein the content of the first and second substances,
(3) utilizing H according to the spacecraft system fault model in S12Solving the maximum value (J minimum value) of the spacecraft reconfigurability evaluation model evaluation parameter rho in the step (2);
wherein W (t) is a positive definite symmetric time-varying matrix satisfying the following matrix differential equation:
(4) solving a numerical solution of the reconfigurability quantization index based on a fine integral method to realize quantitative evaluation of the reconfigurability of the spacecraft;
(5) judging the reconfigurability of the spacecraft under the current design, and outputting the quantified reconfigurability;
taking 1 and 2 rounds of failures as objects, and setting a failure factor alphai(i is 1,2) is respectively taken from 0 to 1, and the system reconfigurable degree rho is obtained according to the failure factor alphaiThe varying curves of (a) are shown in fig. 1. The observation shows that: the system reconfiguration degree is related to the fault degree, and the more serious the fault is, the more serious the fault canthe smaller the reconstruction degree, the asymmetry of the curved surface shows that the influence of the flywheel 2 fault on the system is more serious than the flywheel 1 fault, when the alpha is2And when the value is less than or equal to 0.32, the system reconfigurable degree is zero even if the flywheel 1 has no fault. Therefore, to improve the safety quality characteristics of the system, the reliability of the flywheel 2 should be improved.
(6) And (6) ending.
In summary, the above embodiments verify the feasibility and effectiveness of the spacecraft reconfigurability analysis method under the influence of the interference provided by the present invention.
Example 3:
a computer-readable storage medium, on which a computer program is stored which, when executed by a processor, carries out the steps of the method of embodiment 1.
Those skilled in the art will appreciate that those matters not described in detail in the present specification are well known in the art.
Although the present invention has been described with reference to the preferred embodiments, it is not intended to limit the present invention, and those skilled in the art can make variations and modifications of the present invention without departing from the spirit and scope of the present invention by using the methods and technical contents disclosed above.
Claims (7)
1. A spacecraft reconfigurability evaluation method is characterized by comprising the following steps:
s1, establishing a spacecraft system fault model under the conditions of actuator fault, environmental interference and time limitation;
s2, establishing a spacecraft reconfigurability evaluation model;
s3, utilizing H according to the spacecraft system fault model in S12Solving the minimum value of an evaluation parameter J in a spacecraft reconfigurability evaluation model;
s4, based on the spacecraft reconfigurability evaluation model in S3, solving a numerical solution of indexes by using a fine integration method;
s5, if the numerical value of the reconfigurability quantization index is 0, judging that the spacecraft is not reconfigurable, and entering the step S6; if the numerical solution of the quantitative index is not 0, judging that the spacecraft is reconfigurable, outputting the quantitative reconfigurable size, and entering step S6;
and S6, ending.
2. A spacecraft reconfigurability evaluation method according to claim 1, wherein when it is determined that the spacecraft is reconfigurable, the failed actuator in S1 is switched to a backup actuator, or the failed actuator in S1 is switched from a current control mode to another control mode.
3. A spacecraft reconfigurability evaluation method according to claim 1, wherein the spacecraft system failure model in S1 is:
wherein
In the formula Ix,Iy,IzRespectively, the three-axis moment of inertia of the satellite; x is formed by Rn、u∈Rm、
State vector, control input vector and external interference vector of the nominal system respectively; a is an element of Rn×n,Bu∈Rn×m,
Respectively a state matrix, a control input matrix and an external interference matrix; t is tfTime of occurrence of a failure, tmissetting the angle of control mechanism as α beta, setting the moment distribution matrix as phi (α beta), setting A as diag { theta } as the configuration of control mechanism1,θ2,...,θmIs the actuator failure factor matrix, θi∈[0,1],i＝1,2,...,m，θiSmaller means lower residual efficiency of the respective actuator; i is3×3Is an identity matrix; 03×3Is a zero matrix.
4. A spacecraft reconfigurability evaluation method according to claim 1, wherein the spacecraft reconfigurability evaluation model in S2 is:
wherein rho is a quantitative evaluation index, kappa is an allowable performance threshold value when the spacecraft is reconfigurable, Q and R are symmetric matrixes, and Q-Q is satisfiedT≥0，R＝RT＞0；x∈Rn、u∈RmRespectively a state vector and a control input vector of a nominal system; t is tmisIs the task completion time.
5. A spacecraft reconfigurability evaluation method according to any one of claims 1 to 4, wherein the time-limited condition means that a fault occurrence time is unknown and a task completion time is determined.
6. A spacecraft reconfigurability evaluation method according to any one of claims 1 to 4, wherein a spacecraft reconfigurability evaluation model is established according to system control accuracy requirements, actuator health states and allowable performance thresholds.
7. A computer-readable storage medium, on which a computer program is stored which, when being executed by a processor, carries out the steps of the method as claimed in claim 1.
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| CN105955299A (en) * | 2016-06-08 | 2016-09-21 | 北京宇航系统工程研究所 | Reconfigurable integrated measurement-control, navigation, flight control system and reconstruction method thereof |
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| CN105955299A (en) * | 2016-06-08 | 2016-09-21 | 北京宇航系统工程研究所 | Reconfigurable integrated measurement-control, navigation, flight control system and reconstruction method thereof |
| EP3293639A1 (en) * | 2016-09-07 | 2018-03-14 | LYNX Technik AG | Reconfigurable digital signal processing device |
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