117 results about "Spacecraft attitude control" patented technology

The invention relates to a rigid spacecraft attitude control method and system based on finite time constraint, and belongs to the technical field of spacecraft attitude control. Comprising the following steps: respectively constructing a spacecraft attitude control system model and a finite time convergence performance function containing a preset time constant; according to the finite time convergence performance function, dynamic attitude constraint is applied to the system state of the spacecraft attitude control system model; converting the attitude constraint into differential homeomorphic mapping of an unconstrained space; obtaining an error conversion variable according to differential homeomorphic mapping; and designing a virtual control law and constructing an error compensation module according to the error conversion variable, and designing an actual control input law according to the virtual control law and the error compensation module. According to the method, rapid convergence of the spacecraft attitude within the preset time is ensured, the dynamic response precision of the spacecraft in the rapid attitude maneuver process is remarkably improved, the control period delay is effectively avoided, and the power consumption is reduced.

The invention belongs to the technical field of spacecraft attitude control, and discloses a spacecraft attitude preset performance control method and equipment, and a computer storage medium. The method comprises the following steps: S1, establishing a spacecraft error attitude dynamics and kinematics model considering unknown bounded input time delay; s2, constructing a preset performance function, and performing preset performance constraint on the attitude error of the spacecraft through the preset performance function; introducing an error conversion function, and converting a constrained attitude error into an unconstrained error; performing state conversion on the spacecraft error attitude dynamics and kinematics model to obtain a spacecraft error attitude dynamics and kinematics model related to state variables; and S3, designing a preset performance control law based on the tracking error variable, so that the attitude error moves in a preset performance function envelope. According to the method, errors caused by unknown bounded input time delay can be compensated, and meanwhile, the steady-state and transient-state performance of spacecraft attitude control can be met.

The invention discloses a method for predicting the remaining service life of a control moment gyroscope based on parameter fusion, and relates to the technical field of health management of spacecraft attitude control systems. According to the method, optimal fusion of measurement parameters and implicit parameters is realized through adversarial learning feature extraction and a channel attention mechanism, and the method is particularly suitable for high-precision residual life prediction of the spacecraft control moment gyroscope under variable working conditions. The method comprises four main steps of data preparation and multi-source input, adversarial learning feature extraction, SE-Attention feature fusion and three-stage training optimization. In the data preparation step, original time sequences of voltage and current of a rotor motor are collected through a built-in electrical sensor of a CMG, and key physical parameters of a full-film lubrication friction coefficient, lubricating oil viscosity, a bearing clearance and contact stiffness are inverted based on an electromechanical coupling model; in the adversarial learning step, a game framework of a feature extractor and a working condition discriminator is constructed, and working condition-independent robust feature extraction is realized through a gradient inversion layer; in the SE-Attention feature fusion step, importance weights are adaptively distributed to different feature channels through an extrusion-excitation mechanism; and the three-stage training optimization step adopts a preheating-confrontation-fine tuning strategy to ensure the network convergence. The method solves the problem of domain adaptation, has the technical advantages of strong robustness, high precision, strong generalization, interpretability and self-adaptation, and can realize high-precision RUL prediction under variable working conditions.

The invention provides a spacecraft attitude control method, a spacecraft attitude control device and a spacecraft, and the spacecraft attitude control method comprises the following steps: calculating a pitch attitude angular acceleration at a starting moment of designing a pitch attitude angular velocity instruction by adopting a parabola attitude adjustment mode; determining a pitch attitude angular velocity instruction under each working condition according to the relationship between a pitch attitude angular velocity target value at the satellite-rocket separation moment and 0 and the magnitude relationship between the pitch attitude angular velocity target value and the pitch attitude angular velocity at the attitude adjustment starting moment of the pitch attitude angular velocity instruction; and tracking the pitch attitude angular velocity instruction of the current control period, and controlling the angular velocity of the yaw attitude and the angular velocity of the rolling attitude to be 0 degree / s so as to realize real-time control of the flight attitude of the carrier rocket. The reliability of the carrier rocket can be improved, meanwhile, the propellant consumption of the auxiliary power system is reduced, and the carrying capacity of the carrier rocket is improved.

The invention discloses a flexible agile spacecraft asymmetric attitude constraint control method based on event driving, and the method comprises the steps: building a flexible agile spacecraft attitude dynamic model considering multiple interferences according to the dynamic characteristics in a spacecraft attitude control task; comprising flexible vibration, unknown external disturbance torque and inertia parameter uncertainty; then designing a modal observer and a disturbance observer, wherein the designed modal observer and disturbance observer respectively obtain a vibration modal online observation value and an online observation value of composite disturbance composed of unknown external disturbance torque and inertia uncertainty; the method comprises the following steps: constructing an asymmetric TVIBLF for processing attitude angle constraint and a constant IBLF for processing angular velocity constraint; and finally, designing a flexible agile spacecraft attitude controller under the asymmetric full-state constraint when the condition is triggered, and proposing an event driving condition. The invention can effectively reduce the updating frequency and reduce the consumption of communication resources.

The invention belongs to the technical field of spacecraft control, and discloses a flexible spacecraft attitude control method and related equipment based on an all-drive system theory, and the method comprises the steps: building an attitude tracking error model containing flexible vibration, external interference and rotational inertia uncertainty, the model is converted into a second-order all-wheel-drive system model with a lumped disturbance term; converting the second-order all-drive system model into a linear steady-state closed-loop system containing a lumped disturbance term, and constructing a preset time sliding mode control law capable of enabling the state of the linear steady-state closed-loop system to converge within a preset time; constructing an RBF disturbance observer for performing approximation estimation on the lumped disturbance item; according to the second-order all-drive system model, the preset time sliding mode control law and the RBF disturbance observer, an RBF preset time pseudo-linear sliding mode controller is constructed; the flexible spacecraft is controlled by the pseudo-linear sliding mode controller according to the RBF preset time; therefore, the method has the advantages of realizing preset time convergence and improving attitude control precision and stability.

The application relates to the technical field of multi-spacecraft cooperative control, and discloses a multi-spacecraft grouping consistency restraint control method and device under optimized FLS, which comprises the following steps: acquiring the network topology of a heterogeneous multi-spacecraft system and constructing a multi-subgroup directed communication topology network, constructing a spacecraft attitude control system model based on modified Rodrigues parameters; approximating unknown nonlinear terms in the model through a fuzzy logic system, constructing a fuzzy logic system based on a gradient descent algorithm and Armijo line search criterion; constructing a restraint control term and using a backstepping method to construct a virtual control law, combining the output of the fuzzy logic system based on the gradient descent algorithm and the Armijo line search criterion to construct an actual control input law, so that grouping consistency restraint control of the heterogeneous multi-spacecraft system is realized. The application can improve real-time approximation precision and adaptive step length adjustment capability while guaranteeing grouping consistency, and can enhance topology applicability and task cooperation flexibility.

The present invention relates to the technical field of spacecraft attitude control, and in particular provides an adaptive non-singular sliding mode attitude control method for spacecraft unwinding problems. The method includes establishing a spacecraft attitude kinematics and dynamics model; constructing an attitude tracking motion model of the spacecraft based on the spacecraft attitude kinematics and dynamics model; establishing a non-singular sliding mode surface with anti-unwinding fixed-time convergence characteristics according to the attitude error model; designing an attitude tracking adaptive controller based on the spacecraft attitude tracking motion model and the non-singular sliding mode surface. In the case of external disturbances and uncertain moments of inertia, this method combines sliding mode control and adaptive algorithms to achieve the attitude tracking control task and improve the accuracy of attitude control.

The invention relates to a flexible connection load transfer function identification method based on active frequency sweep excitation, and relates to the field of spacecraft attitude control, and the method comprises the following steps: a plurality of three-super platform actuators sequentially output sine frequency sweep excitation moments of a load in three rotation directions, and a high-bandwidth fiber-optic gyroscope simultaneously measures the triaxial angular velocity of the load; the obtained time domain result is converted into a frequency domain result, and then load transfer function model identification is carried out; and adjusting controller parameters in the three-super platform control unit based on the identification model so as to increase the control bandwidth and improve the secondary control capability on the basis of ensuring the stability margin of the controller in the three-super platform control unit. And a technical basis is provided for ultrahigh-precision pointing, ultrahigh-stability control and high-quality imaging of optical loads of spacecrafts in the future.

The invention discloses a spacecraft attitude stability control method based on a liquid neural network. The method comprises the following steps: S1, establishing a spacecraft attitude dynamic model under multi-source complex disturbance; s2, designing a conventional spacecraft attitude stability controller, and constructing a data set in combination with the attitude dynamics model; s3, training and obtaining a liquid neural network interference observer by adopting a liquid time constant network and a neural circuit strategy architecture; s4, designing a spacecraft attitude stability controller based on a liquid neural network disturbance observer, and proving the stability of a closed-loop system by using a Lyapunov function; s5, verifying the estimation effect of the liquid neural network interference observer and the stability effect of the spacecraft attitude control system through simulation; results show that the method has the advantages that interference of the spacecraft attitude can be effectively estimated, stable control over the spacecraft attitude can be achieved, high real-time performance, low power consumption and high robustness are achieved, and the spacecraft attitude stable control efficiency can be improved.

1. Name of the design product: Spacecraft attitude control system (new 0.5 kg). 2. Use of the design product: The design product is used for high-precision attitude adjustment of spacecraft. 3. Design points of the design product: Combination of shape and pattern. 4. Picture or photo best indicating the design points: Perspective view.

The invention relates to the field of spacecraft attitude control, in particular to a load optical axis pointing control system and method for a truss type flexible spacecraft, and the system comprises a truss type flexible spacecraft structure platform, an attitude control module and a load optical axis pointing control module. The attitude control module is used for central star attitude angle fine adjustment and preliminary adjustment of optical load primary mirror optical axis direction. The load optical axis direction control module is used for optical axis direction fine adjustment and micro-vibration suppression. Two independent closed loops of central star attitude angle closed-loop control and load optical axis pointing closed-loop control are constructed, input signal sources, control objects and execution mechanisms of the two control loops are completely decoupled, and the problem of control effect divergence caused by mutual interference in traditional coupling control is solved in principle.

The application discloses a hotspot target tracking method and device based on load detection and star sensitivity reset, and belongs to the field of spacecraft attitude control. The method comprises the following steps: according to the time scale of load target detection information and historical load visual axis ground pointing point information, the latitude, longitude and altitude information of the target are estimated; according to the maximum attitude maneuvering ability of the system and the five-order polynomial guide model, the fastest time of target recapture is determined by taking time as an iteration amount, the target attitude of the target recapture process is autonomously planned, and closed-loop control is carried out; after the target recapture, the control system carries out closed-loop control according to the real-time detection information of the load, ensures that the target is located at the center of the visual field of the load, and the load continuously images; after the tracking imaging is completed, the guide process of returning to scanning imaging is autonomously planned; in the whole process, the current attitude is reset according to the double-vector attitude determination result of the high-precision star sensitive sensor with angular velocity as a constraint, so that the pointing accuracy is ensured.

This application belongs to the field of spacecraft attitude control technology and discloses a method, system, device and medium for attitude control of a rigid spacecraft. The method includes: S1, constructing a rigid spacecraft attitude dynamics model and kinematic equations; S2, constructing a T-S fuzzy switching model for the rigid spacecraft; S3, introducing a multi-node random communication protocol; S4, designing a fuzzy sliding mode controller based on the T-S fuzzy switching model; S5, forming a closed-loop control system for attitude control: Based on the T-S fuzzy switching model, multi-node random communication protocol and fuzzy sliding mode controller designed in S1-S4, a closed-loop control system is formed to achieve spacecraft attitude control and verify the global stability of the closed-loop control system. This application solves the technical difficulties of strong nonlinearity, large external disturbances, limited communication and high real-time performance in the attitude control of rigid spacecraft, effectively improving the attitude stability and control accuracy of spacecraft in complex space environments.

This invention relates to the field of spacecraft attitude control technology, and more particularly to an ultra-precise and ultra-stable on-orbit adaptive control method and system for pointing under additional stiffness disturbances. This on-orbit adaptive control method, by applying specific excitations to actuators and measuring their responses, combined with a recursive filtering algorithm, can accurately identify the total actuator stiffness, including the additional stiffness of the cable, in real-time on-orbit, overcoming problems caused by differences in ground testing environments and time-varying stiffness. Based on the real-time identified stiffness values, the parameters of the axial compensation controller and the actuator compensation controller are dynamically adjusted, enabling the control system to adapt to changes in the additional stiffness of the cable and maintain excellent control performance at all times. By employing dual compensation control at the axial load level and the actuator level, additional stiffness disturbances are synergistically suppressed at both the overall system and local actuator levels, improving the comprehensiveness and accuracy of control, and achieving ultra-precise and ultra-stable on-orbit adaptive control for pointing under cable additional stiffness disturbances.

The invention discloses a spacecraft attitude control single-machine model selection method and system and computer equipment. The spacecraft attitude control single-machine model selection method comprises the steps that disturbance torque is calculated based on a disturbance torque model; torque data of the disturbance torque and the suppression capability of the attitude control single machine on the disturbance torque are obtained, wherein the torque data comprise the torque distribution condition, the torque size and the torque direction; selecting a corresponding attitude control single machine combination and single machine control parameters by utilizing an optimization algorithm on the basis of spacecraft attitude control requirements, torque data and the suppression capability of attitude control single machines on disturbance torque; and based on the attitude control single-machine combination, the corresponding single-machine control parameters and numerical simulation, simulating the operation attitude of the spacecraft in different orbits and space environments, and selecting the attitude control single-machine combination according to a simulation result. According to the method, the precision and the stability of the satellite attitude control system can be remarkably improved, and the possible uncertainty in a traditional model selection method is reduced.

The application discloses a flexible spacecraft attitude control method and system, and the method comprises the following steps: establishing a flexible spacecraft attitude control model and a flexible accessory model; according to the homogeneous system theory, the flexible spacecraft attitude control model and the flexible accessory model are simplified for multiple times to obtain a simplified flexible spacecraft attitude control model and a simplified flexible accessory model; according to the coupling relationship between the simplified flexible spacecraft attitude control model and the simplified flexible accessory model, a full drive model is obtained; the output equation in the flexible accessory model is combined to design an observer, the modal variable in the flexible accessory model is estimated through the observer, the full drive model is combined, and a controller is designed based on a high-order full drive system method; the controller is substituted into the full drive model to obtain a linear closed-loop system; according to a parameterized design method, a parameterized controller is obtained, and a stable linear closed-loop system is obtained. The application can stabilize the flexible spacecraft attitude control system.

The utility model relates to a three-way control air valve for a satellite reaction flywheel, which relates to the technical field of spacecraft attitude control and comprises a cap, a threaded adapter, a three-way device and an ejector pin. The side face of the three-way device is connected with the threaded adapter, and the ejector pin is inserted into the three-way device in a penetrating mode. A connecting pipe is arranged at the end, away from the tip of the ejector pin, of the three-way device, and a first external thread is arranged on the side face of the connecting pipe. A first internal thread matched with the first external thread is arranged on the inner wall of the cap, and the cap is connected with the connecting pipe through the first internal thread; when the cap is rotated to move forwards on the connecting pipe, the cap pushes the tip of the ejector pin to move forwards. The device has the advantages that manual control is more convenient, and the air flow can be better controlled; the gas tightness is better, and gas circulation can be better guaranteed; most of the body parts are machined, so that the precision and the stability are ensured, and the stability is good.

The application discloses a normal vector-based follow-sighting attitude maneuver path planning method, and belongs to the technical field of spacecraft attitude control, and comprises the following steps: at any time during the follow-sighting maneuver, a camera optical axis is aligned with a target, and a satellite body axis perpendicular to the camera optical axis is set to point to a set normal vector direction, the set normal vector direction is set according to maneuver capability and pointing requirement, and the setting basis is that the body axis is in the set normal vector direction; the set normal vector direction is determined according to the follow-sighting task, and includes four cases, namely, keeping data transmission, keeping energy, keeping maneuver and keeping imaging. The normal vector of the plane formed by the target pointing and the ground pointing is applied, and the target attitude setting during the antenna data transmission to the ground while the follow-sighting is realized.

The application relates to a spacecraft attitude control method of a reaction wheel-chemical propulsion-electric propulsion hybrid actuator, which comprises the following steps: a dynamic model of a GEO satellite with flexible accessories is established; a spacecraft attitude control algorithm is designed by using a self-anti-interference control law; total control torques of three kinds of actuators, i.e., a reaction wheel, chemical propulsion and electric propulsion, are calculated; a dynamic distribution model of the hybrid actuator is designed to minimize energy consumption, and is used for torque distribution of the total control torques of the three kinds of actuators, so that the control torques of the reaction wheel, chemical propulsion and electric propulsion are obtained; and the particle swarm algorithm is used for iterative optimization of control parameters of the dynamic distribution model, so that the optimal control torques of the reaction wheel, chemical propulsion and electric propulsion are output, the autonomous distribution of the control torques of the actuators in the control process is realized, and the high-precision, short-time and low-energy consumption target of the rapid switching of various attitude control modes of the GEO satellite and between the various attitude control modes is realized.

1. Name of the design product: Spacecraft attitude control system (new 8.5 kg). 2. Use of the design product: The design product is used for high-precision attitude adjustment of spacecraft. 3. Design points of the design product: Combination of shape and pattern. 4. Picture or photo best indicating the design points: Perspective view.

The present invention discloses a method for spacecraft attitude control and RCS optimal allocation, which relates to the technical field of spacecraft control. Spacecraft parameters and spacecraft attitude control torque commands are obtained; a spacecraft attitude dynamics model is constructed; a state space equation is constructed according to the spacecraft attitude dynamics model; an estimated value of the external disturbance torque acting on the spacecraft is obtained according to the state space equation; an error conversion state space equation is constructed according to the finite-time performance function and the estimated value of the external disturbance torque acting on the spacecraft; a super-twisting sliding mode controller is constructed according to the error conversion state space equation; a spacecraft attitude dynamics model after chattering suppression is constructed according to the super-twisting sliding mode controller and the saturation function; an RCS optimal allocation result is obtained according to the spacecraft attitude control torque command and the spacecraft attitude dynamics model after chattering suppression. The present invention realizes the precise tracking control of the spacecraft attitude within a finite time, and optimizes the control allocation with the minimum command error and the overall optimal energy consumption.

The invention discloses a control moment gyroscope implicit parameter tracking method based on two-stage identification, and belongs to the technical field of spacecraft execution mechanism health monitoring and fault diagnosis. According to the method, through four core steps of electromechanical coupling model establishment, first-stage sliding window batch NLS coarse identification, Wiener process degradation modeling and second-stage adaptive EKF fine correction, a complete process from electrical telemetry data to implicit parameter accurate identification is realized. The method comprises the following steps: constructing a mapping relation between an electrical signal and an implicit parameter based on an SKF friction torque model and a direct current brushless motor voltage equation, and obtaining a preliminary implicit parameter value through nonlinear least square optimization by using electrical measurement data in a sliding time window; a nonlinear drift Wiener process is adopted to describe the evolution rule of hidden parameters along with time, and a first-stage result serves as an observation value and is subjected to smooth correction through extended Kalman filtering; according to the method, through a two-stage framework design of combining a sliding window batch NLS with a Wiener process driven adaptive EKF, the interpretability of a physical model and the sequence modeling capability of a random process are fully combined, working condition interference and parameter degradation information are effectively decoupled, noise interference of an identification result is remarkably reduced, and the degradation trend of equipment is accurately tracked. According to the method, the problems of serious electromechanical coupling interference, insufficient parameter identification precision, poor degradation trend tracking capability and the like in the existing CMG health monitoring technology are solved, reliable technical support is provided for health management and life prediction of the single-frame control moment gyroscope, and the reliability and safety of a spacecraft attitude control system are remarkably improved.

The application discloses a spacecraft magnetic moment attitude control method based on a non-inverse iteration algorithm, and the method establishes a linear periodic model for a spacecraft attitude control system which only relies on a magnetic moment; on the basis, according to a linear quadratic optimal control theory, an optimal controller design problem is converted into a solution problem of a discrete periodic Riccati equation associated with the optimal controller design problem; on the basis of a gradient descent principle, an error function is constructed, and a non-inverse iteration algorithm based on the gradient descent principle is proposed to solve the discrete periodic Riccati equation, so that a matrix inversion operation in a traditional solution method is avoided, and the calculation efficiency and the numerical stability are effectively improved; the obtained Riccati equation solution is used to obtain an optimal state feedback control law, so that the spacecraft attitude optimal control only using the magnetic moment is realized, and a new solution scheme is provided for an attitude control problem of a limited actuator in a space task.

The application relates to an on-orbit reconfiguration method of a software-defined thruster, and belongs to the field of spacecraft attitude control. The steps comprise: generating a propulsion subsystem software model according to a propulsion subsystem architecture, an installation layout and a configuration; determining a thruster reconfigurable boundary for the software model; in the case of partial thruster failure, generating a thruster reconfiguration scheme in combination with the thruster reconfigurable boundary; and outputting the thruster reconfiguration scheme for execution, and comparing whether the output performance meets the task requirements. The application makes up for the torque output under the condition of partial thruster invalidation, improves the effectiveness and reliability of on-orbit output of a propulsion system, and improves the on-orbit capability boundary of the propulsion system.

The present application relates to the technical field of spacecraft attitude control, and particularly relates to a load attitude determination method and device based on an included angle monitoring star sensor, wherein the method comprises: obtaining angle conversion relationships between a reference mirror coordinate system and a star sensor measurement coordinate system, between a central prism coordinate system and the reference mirror coordinate system, and between the central prism coordinate system and a star sensor monitoring coordinate system in a ground phase; obtaining angle conversion relationships between the central prism coordinate system and the star sensor monitoring coordinate system, and between the star sensor measurement coordinate system and a J2000 coordinate system in an on-orbit phase; and combining a corresponding relationship between the star sensor measurement coordinate system and the star sensor monitoring coordinate system to determine an angle conversion relationship between the central prism coordinate system and the J2000 coordinate system in the on-orbit phase. The present application can obtain high-precision load pointing.

The application discloses a kind of underactuated spacecraft attitude control method and system, method includes: establishing underactuated spacecraft attitude control model;According to the averaging theorem and homogeneous system theory, the underactuated spacecraft attitude control model is simplified, and the simplified attitude control model is obtained;The derivative of the simplified attitude control model is obtained, and the high-order full-drive model is obtained;Based on the full-drive characteristics of the high-order full-drive model, the controller of the simplified attitude control model is obtained, and the linear constant closed-loop system is obtained in combination with the simplified attitude control model;According to the eigenvalue of the linear constant closed-loop system, the complete parameterized controller is obtained, and the stable linear constant closed-loop system is obtained.The underactuated spacecraft attitude control method and system disclosed by the application ultimately obtain linear closed-loop average system, and nonlinearity is not retained, so that the system performance is easier to analyze.

The application discloses a method, system, device, medium and product for judging the instability degree of an electric power rope system, and relates to the technical field of spacecraft attitude control. The method comprises the following steps: acquiring the orbit inclination of the deployment orbit of a main star in the electric power rope system; determining the instantaneous equilibrium point distribution area of the electric power rope system based on the orbit inclination; positioning the optimal balance area of the electric power rope system in the instantaneous equilibrium point distribution area of the electric power rope system; acquiring the attitude angle of the electric power rope system at the current moment, and determining the instability degree of the electric power rope system at the current moment based on the attitude angle of the electric power rope system at the current moment, the instantaneous equilibrium point distribution area and the optimal balance area. The application improves the efficiency and reliability of the instability evaluation of the electric power rope system.

The invention relates to a spacecraft attitude combined control method based on a hybrid execution mechanism. The method comprises the following steps: establishing a GEO satellite dynamics model with flexible accessories; according to the established kinetic model, designing a spacecraft attitude control algorithm by using a golden section adaptive control law based on the feature model and a tracking maintenance adaptive control law based on the feature model, and generating a control instruction; according to the generated control instruction, establishing a dynamic allocation model of the hybrid actuating mechanism by taking an attitude maneuver angle error as a constraint condition, and switching a combined working mode of electric propulsion, reaction wheel and chemical propulsion in different attitude adjustment stages; and with minimum energy consumption as a target, solving an optimal error threshold used for switching a combined working mode in the dynamic distribution model of the hybrid execution mechanism by using a PSO-GA algorithm, and realizing autonomous distribution of the control torque of the execution mechanism in the spacecraft attitude combined control process. Therefore, the goals of high precision, short time and low energy consumption of the multiple attitude control modes of the GEO satellite and rapid switching among the multiple attitude control modes are achieved.

This invention relates to a method for identifying the transfer function of a flexible connection load based on active frequency sweep excitation, relating to the field of spacecraft attitude control. The method includes the following steps: several triple-platform actuators sequentially output sinusoidal frequency sweep excitation torques in three rotational directions of the load; a high-bandwidth fiber optic gyroscope simultaneously measures the three-axis angular velocities of the load; the obtained time-domain results are converted to frequency-domain results, followed by load transfer function model identification; based on the identified model, the controller parameters in the triple-platform control unit are adjusted to increase the control bandwidth and enhance secondary control capabilities while ensuring the stability margin of the controller in the triple-platform control unit. This invention has the advantage of maximizing secondary control capabilities, providing a technical foundation for ultra-high precision pointing, ultra-high stability control, and high-quality imaging of future spacecraft optical payloads.