CN-121980667-A - On-orbit identification method for spacecraft system parameters under fault state of extension rod

CN121980667ACN 121980667 ACN121980667 ACN 121980667ACN-121980667-A

Abstract

An on-orbit identification method for spacecraft system parameters under a fault state of a stretching rod comprises the steps of constructing a gesture dynamics model of a rigid-flexible coupling configuration of the spacecraft system under a fault state of a coiled stretching rod, establishing an expression of the moment of the stretching rod to the spacecraft by vibration, bringing the expression of the moment into the gesture dynamics model to obtain an expression of a system control moment represented by a model parameter and a gesture parameter, giving the system control moment, applying the system control moment to a spacecraft rigid body through an executor of the spacecraft to enable the spacecraft to rotate and obtain corresponding gesture parameter measured values, sending the obtained gesture parameter measured values to the ground, expressing an estimated value of the system control moment by using the gesture parameter measured values and an estimated value of the model parameter, defining a system identification error as a function of the system control moment and the estimated value thereof, determining parameters to be identified based on the model parameter, and searching the parameters to be identified which can minimize the system identification error function as an optimized parameter identification result.

Inventors

FAN LIMING
ZHANG QIANG
HOU RUI
LUO YUFANG
LI QING
YIN JIANFENG
XING YANJUN
LIANG ZHONGJIAN
HE ZONGBO

Assignees

北京空间飞行器总体设计部

Dates

Publication Date: 20260505
Application Date: 20251209

Claims (9)

1. An on-orbit identification method for parameters of a spacecraft system in a fault state of a boom, wherein the boom is arranged outside the spacecraft, and the method is characterized by comprising the following steps: constructing a gesture dynamics model of a rigid-flexible coupling configuration of the spacecraft system under the fault state of expansion of the extension rod; establishing an expression of the moment of the boom vibration on the spacecraft, and bringing the moment expression into a gesture dynamics model to obtain an expression of the system control moment expressed by using model parameters and gesture parameters; giving a system control moment, applying the system control moment on a spacecraft rigid body through an actuator of the spacecraft, so that the spacecraft rotates and obtains a corresponding attitude parameter measurement value; The method comprises the steps of sending an obtained attitude parameter measured value to the ground, expressing an estimated value of a system control moment by using the attitude parameter measured value and the estimated value of a model parameter, and defining a system identification error as a function of the system control moment and the estimated value thereof; And estimating and positioning the fault state of the extension rod through the parameter identification result, and diagnosing and repairing the fault state of the extension rod.
2. The method for identifying the parameters of the spacecraft system in the fault state of the extension rod on orbit according to claim 1, wherein the attitude dynamics model of the rigid-flexible coupling configuration of the spacecraft in the fault state of the extension rod is as follows: Wherein omega b (t) is the attitude angular speed of the spacecraft at the moment t, Is the angular acceleration of the attitude of the spacecraft at the moment t, J k1 and J k are the moment of inertia of the rigid body of the spacecraft and the moment of inertia of the rigid body and the flexible accessory respectively, H k and M Jk are the coupling matrix between the rigid body and the flexible accessory and the equivalent moment of inertia of the flexible accessory respectively, and Z k1 (t) is the angular velocity omega b (t) of the attitude of the spacecraft and the vibration velocity of the extension rod Z k2 (t) is the intermediate state quantity of the boom expansion length L d and the attitude angular speed omega b (t), S (omega b (t)) is the cross matrix of the vector omega b (t), u r (t) is the system control moment, d (t) is the environmental disturbance, The flexible vibration state of the extension rod, the flexible vibration speed of the extension rod and the flexible vibration acceleration of the extension rod are respectively shown in the specification, D k and K k are respectively a damping matrix of the flexible accessory and a rigidity matrix of the flexible accessory, and D ηk (t) represents the disturbance moment of the vibration of the flexible accessory on the rigid body.
3. The method for identifying parameters of a spacecraft system in an on-orbit state in a fault state of a boom according to claim 2, wherein the disturbance moment d ηk (t) of vibration of a flexible accessory to a rigid body is: Where H k represents the rigid-flex coupling matrix between the rigid body and the flexible attachment, M Jk is the equivalent moment of inertia of the flexible attachment, and Z k1 is the intermediate state quantity.
4. The method for on-orbit identification of spacecraft system parameters in the event of a boom failure according to claim 3, wherein the expression of the system control moment expressed by the parameters to be identified and the attitude parameters is: j k1 ,J k ,H k ,M Jk ,D k ,K k is an unknown model parameter.
5. The method for on-orbit identification of spacecraft system parameters in the event of a boom failure according to claim 1, wherein a system control moment is designed, said system control moment being sinusoidal when applied to a spacecraft rigid body by an actuator of the spacecraft.
6. The method for on-orbit identification of a spacecraft system parameter in a fault state of a boom as claimed in claim 4, wherein the estimated value of the control moment of the system is expressed using the measured value of the attitude parameter and the estimated value of the parameter to be identified Wherein i is the number of samples measured by the sensor in the process of attitude maneuver, i is more than or equal to 1 and less than or equal to N and is an integer, and N is the total number of measured samples; The obtained attitude parameter measurement value comprises omega b (i) which is the attitude angular speed measured when the spacecraft rotates for the ith time under the action of the control moment, For the angular acceleration of the gesture measured when the spacecraft rotates under the action of the control moment for the ith time, eta (i) is the flexible vibration state of the extension rod measured when the spacecraft rotates under the action of the control moment for the ith time, The flexible vibration speed of the extension rod is measured when the spacecraft rotates under the action of the control moment for the ith time; The method is a spacecraft rigid body moment of inertia estimated value; Adding an estimate of moment of inertia for the rigid body to the flexible attachment; Is an estimate of the rigid-flex coupling matrix H k between the rigid body and the flexible attachment; an estimate of the equivalent moment of inertia M Jk for the flexible attachment; an estimate of the flexible attachment damping matrix D k ; an estimate of the stiffness matrix K k for the flexible attachment; Is an estimated value of the intermediate state quantity Z k1 ; is an estimated value of the intermediate state quantity Z k2 ; The estimated value of the control moment in the ith sampling measurement of the spacecraft is shown as a cross-multiplying matrix of omega b (i) in S (omega b (i)).
7. The method for identifying the parameters of the spacecraft system in the fault state of the extension rod according to claim 6, wherein the system identification error function J is as follows: u r (i) is the control moment measurement at the ith sample of the spacecraft.
8. The method for on-orbit identification of spacecraft system parameters in a fault state of a boom according to claim 1, wherein when determining parameters to be identified based on model parameters, the model parameters to be determined are: Extension arm extension length Moment of inertia matrix System damping matrix And stiffness matrix
9. The method for on-orbit identification of spacecraft system parameters under a fault state of a boom according to claim 1, wherein the method is characterized in that the variable weight particle swarm optimization identification method is used for searching the parameters to be identified which can minimize the system identification error function based on the range of the parameters to be identified, and specifically comprises the following steps: S1, defining each parameter X f (f is more than or equal to 1 is less than or equal to Size) to be identified as a particle, initializing the total number of particles as Size, defining the maximum iteration number as Gen, representing the iteration result of the particle in the lambda generation as X f (lambda), defining the individual optimum as the result of the f particle in the process of iterating to the lambda generation, which minimizes the system identification error function, representing as P f (lambda), defining the global optimum as the result of all particles in the process of iterating to the lambda generation, which minimizes the system identification error function, representing as BestS (lambda); S2, evaluating the particle fitness by using a system identification error function; S3, initializing individual particle history optimal P f (lambda) and overall particle optimal BestS (lambda) according to the fitness of the particles; S4, judging whether the identification precision of BestS (lambda) meets the requirement, outputting BestS (lambda) as an optimization result if the identification precision meets the requirement, and performing S5 if the identification precision does not meet the requirement; S5, solving the update speed and the update position of the particles: Wherein lambda represents the current cycle times, 1 is less than or equal to lambda and less than or equal to Gen, gen represents the maximum iteration times, X f (lambda) represents the value of a parameter X to be identified represented by the f-th particle in the lambda-th cycle, w (lambda) represents the inertia weight in the lambda-th cycle, V f (lambda) represents the update speed of the parameter X to be identified represented by the f-th particle in the lambda-th cycle, c 1 is a local learning factor, c 2 is a global learning factor, and rand (0, 1) represents a random number with a value between 0 and 1; S6, evaluating the adaptability of the updated particle X f (lambda+1) by using a system identification error J (X f ); S7, updating the individual particle history optimal P f (lambda) and the overall particle optimal BestS (lambda) according to the fitness of the particles; S8, judging whether the particle iteration algebra meets the requirement, if so, outputting BestS (lambda) as an optimization result, and if not, returning to S4 to continue iteration.

Description

On-orbit identification method for spacecraft system parameters under fault state of extension rod Technical Field The invention relates to an on-orbit identification method for spacecraft system parameters under a fault state of a boom, and belongs to the field of spacecraft system engineering and on-orbit application. Background With the continuous progress of space technology, the exploration and development of space resources has become a strategic research direction in many countries. In the fields of space science detection, on-orbit operation, on-orbit service and the like, a plurality of stretching mechanisms including mechanical arms and flexible connecting rods are widely researched to complete increasingly complex detection and operation tasks, so that the activity capability and the exploration range of the aircraft in a cosmic space are greatly expanded. However, due to the influence of microgravity, super vacuum, strong radiation, high temperature difference and other factors, problems of incomplete unfolding, inconsistent actions and the like occur in the rail working process of many stretching mechanisms, so that the parameters of the aircraft such as the pose, the size envelope, the inertia of the whole device, the rigidity, the damping coefficient and the like of the stretching rod are obviously different from the initial data, and finally the failure of the flight task is caused. In order to improve the adaptability and the mission success rate of the aircraft, a measuring means is added to the extension rod mechanism in engineering, but preliminary estimation is expected to be carried out through an on-orbit identification mode for the extension rod unfolding fault mode, the whole inertia and other non-measurable model parameters. In the existing research, firstly, an aircraft orbit and attitude dynamic model under the fault state of a boom is established, then an engine or a momentum wheel is used for driving an aircraft to finish orbit and attitude maneuver, the position and the attitude of the aircraft are measured to obtain original data, and then the on-orbit parameters of the aircraft are estimated by using algorithms such as a least square method, standard differential evolution, hopfield neural network and the like. However, the identification effect of the least square method in a nonlinear system is usually not ideal enough, the standard differential evolution method has too high requirement on model accuracy, and the Hopfield neural network algorithm has large calculated amount and unstable performance, so that the three algorithms are not enough in the model parameter identification process and are not widely applied to spacecraft engineering. Disclosure of Invention The invention solves the problems of overcoming the defects of the prior art, providing an on-orbit identification method for spacecraft system parameters in the fault state of the extension rod, comprising a complete identification process and an efficient identification algorithm, solving the problem of large system uncertainty in the fault state of extension rod expansion, providing a platform guarantee for completing the flight tasks such as scientific detection, on-orbit operation and the like, and enhancing the on-orbit adaptability and task completion capability of the spacecraft. The technical scheme of the invention is as follows: An on-orbit identification method for parameters of a spacecraft system in a fault state of a boom, wherein the boom is arranged outside the spacecraft, and the method comprises the following steps: Constructing a gesture dynamics model of a rigid-flexible coupling configuration of the spacecraft system under the state of a coiled extension rod unfolding fault; establishing an expression of the moment of the boom vibration on the spacecraft, and bringing the moment expression into a gesture dynamics model to obtain an expression of the system control moment expressed by using model parameters and gesture parameters; giving a system control moment, applying the system control moment on a spacecraft rigid body through an actuator of the spacecraft, so that the spacecraft rotates and obtains a corresponding attitude parameter measurement value; The method comprises the steps of sending an obtained attitude parameter measured value to the ground, expressing an estimated value of a system control moment by using the attitude parameter measured value and the estimated value of a model parameter, and defining a system identification error as a function of the system control moment and the estimated value thereof; And estimating and positioning the fault state of the extension rod through the parameter identification result, and diagnosing and repairing the fault state of the extension rod. Preferably, the attitude dynamics model of the rigid-flexible coupling configuration of the spacecraft in the fault state of the extension rod is as follows: Wherein omega b (t) is the attitude angu