72 results about "Spacecraft attitude control" patented technology

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 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 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.

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 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 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 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 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.

The present application belongs to the technical field of spacecraft attitude control, and relates to a control moment gyro (CMG) on-orbit evaluation method and system and a terminal. The evaluation method acquires on-board telemetry data covering CMG core parameters, including mechanical system, electrical system, environmental adaptation and task performance parameters. Based on the on-board telemetry data, the CMG index values in the reliability, on-orbit performance, stability and health state dimensions are calculated through a multi-source method of collaborative basic index calculation, neural network prediction and data feature analysis. Finally, the comprehensive evaluation result is generated in combination with the index values in each dimension. The present application realizes the all-around accurate characterization of the on-orbit state of the CMG by means of rich data collection, fusion of multi-source algorithms and construction of a multi-dimensional system, improves the evaluation accuracy and robustness, and guarantees the landing of the evaluation.

This invention relates to the field of spacecraft attitude control technology, and more particularly to an active phase stabilization control method and system for flexible vibration during agile maneuvers of satellites. Based on a two-level collaborative architecture of primary control on the satellite platform and secondary control on the three-dimensional super-platform, this method addresses the problem of sustained oscillations easily triggered by large flexible appendages after agile maneuvers by designing an active phase stabilization controller. This controller is configured to specifically design the key flexible mode frequencies affecting maneuver stability. Through phase lead compensation, the damping ratio of the system is effectively increased, thereby rapidly suppressing flexible vibrations after maneuver cessation. By distributing the load control torque, after phase stabilization compensation, to each actuator on the three-dimensional super-platform, rapid stabilization of the load direction within seconds is achieved, overcoming the contradiction between rapid maneuvering and ultra-high stability that is difficult to achieve in existing technologies, and meeting the stability requirements of ultra-high resolution imaging.

The invention relates to the technical field of spacecraft attitude control, in particular to a hollow-core magnetic torquer for a spacecraft, which comprises a coil body, the coil body comprises an inner-layer coil, a middle-layer coil and an outer-layer coil, a base is mounted at the bottom of the outer-layer coil, and screw rods are mounted at the positions, close to the two ends, of the top of the base through shaft seats. The inner-layer coil is connected with the lead screw at one end of the base, the outer-layer coil is connected with the lead screw at the other end of the base, and a motor driving the lead screws to rotate is mounted at the top of the base and corresponds to the lead screws. According to the invention, the coils are arranged into three groups, and two groups of coils are arranged to be capable of relatively axially moving, so that the magnetic field intensity can be indirectly changed by changing the magnetic field superposition amount of the coils on the premise that the input current is fixed, and therefore, the magnitude of the output torque can be changed, and compared with the existing pulse width modulation control mode, the output torque can be changed. An additional control mode can be supplemented and replaced, and a standby control scheme is provided for the magnetic torquer.

The invention relates to the technical field of spacecraft attitude control, in particular to a non-cooperative target combined spacecraft robust attitude control method based on an all-drive system method. Comprising the steps that S1, after a controller is powered on or reset, preset combined spacecraft nominal parameters are loaded, and all gain parameters of the controller are configured; s2, in each control period, performing real-time measurement or estimation through an attitude sensor on a service satellite to obtain a current attitude quaternion and an angular velocity of the combined spacecraft; s3, calculating an attitude error; s4, substituting all gain parameters into a control law, and calculating to obtain a control torque required by the current control period; s5, torque distribution is carried out on the calculated control torque according to the characteristics of a service satellite execution mechanism, and a corresponding mechanism is driven to generate an actual control torque; and S6, judging whether the attitude of the combined spacecraft reaches and is stabilized in an expected state or not. The success rate and the control quality of the on-orbit service task are improved.

The invention discloses a control moment gyroscope remaining service life prediction method based on reconstruction errors, and belongs to the technical field of spacecraft health management. The method comprises the following steps: collecting a multi-dimensional sensor signal and carrying out wavelet denoising and normalization; performing singular value decomposition on the data to extract features; training an auto-encoder model by using the normal state features, and calculating a reconstruction error to quantify an abnormal degree; based on the reconstruction error, determining a threshold value and constructing a health index, and performing index moving average and adaptive Kalman filtering processing; and calculating an estimated degradation speed and combining theoretical design information to carry out weighted fusion correction to obtain a predicted remaining service life. According to the method, key information of multi-dimensional sensor data can be fully mined, accurate identification of an abnormal state and accurate prediction of the residual service life are realized, the accuracy and reliability of a control moment gyroscope health monitoring system are remarkably improved, and a powerful guarantee is provided for safe and stable operation of a spacecraft attitude control system.

The invention discloses an autonomous regulation and control method and device for a nominal frame angle of a variable topology CMG group, and belongs to the field of spacecraft attitude control. The method comprises the following steps: according to the axial symmetry of the mechanical installation configuration of the CMG group, summarizing all configurations of the CMG group in a full configuration state and an under-configuration state into a plurality of limited equivalent configuration classes; performing off-line optimization on each equivalent configuration class to sequentially obtain a nominal frame angle migration path corresponding to each equivalent configuration class; when the working state of the CMG group operating in orbit is changed, determining an optimal switching scheme from the migration path before the change to the migration path after the change according to the equivalent configuration classes before and after the change of the CMG group; and performing zero angular momentum switching on the CMG group according to the optimal switching scheme. According to the method, automatic switching and migration of nominal frame angles among CMG groups with different configurations can be realized, the on-satellite calculation and storage burden can be simplified to the greatest extent, and the system reliability is improved.

The invention relates to a flexible connection-based method for autonomously identifying additional rigidity of a cable between two stages of platforms, which relates to the field of spacecraft attitude control, and comprises the following steps of: in a cable-free connection state, sequentially outputting sine excitation by voice coil motors in all actuators, and measuring by an eddy current sensor to obtain motion displacement; calculating the stiffness of the actuators based on the two groups of data, and then calculating the equivalent stiffness of each actuator in the state of cable connection according to the previous step; the method has the advantages that the additional rigidity of the cable in the on-orbit state can be obtained, the influence of the cable is reduced by adopting control measures, the control capacity of a three-super platform is guaranteed, the stability of the three-super platform is improved, and the reliability of the three-super platform is improved. And a technical basis is provided for high-precision pointing, high-stability control and rapid and stable control of the load of a new-generation spacecraft.

The invention discloses an electromechanical integrated reaction momentum flywheel structure based on an air bearing, and belongs to the technical field of spacecraft attitude control equipment. Comprising a structure base, a rotating shaft, a porous air bearing, a motor rotor magnetic ring, a motor stator coil, an encoder magnetic ring, an encoder reading head, a bearing pressing ring, a flywheel and a shell, the structure base comprises a bottom base, a sleeve part and a bearing installation opening, and the rotating shaft comprises a flange, a magnetic ring installation shaft neck and a bearing installation shaft neck. The bearing installation shaft neck is arranged in a center hole of the porous air bearing in a penetrating mode, the motor rotor magnetic ring is tightly sleeved on the magnetic ring installation shaft neck, the motor stator coil is embedded between the sleeve portion and the motor rotor magnetic ring, a uniform gap is formed between the motor stator coil and the motor rotor magnetic ring, and the encoder magnetic ring is tightly sleeved on the lower end portion of the magnetic ring installation shaft neck. The encoder reading head is sleeved outside the encoder magnetic ring and is fixed with the bottom base; and a uniform gap is formed between the encoder reading head and the encoder magnetic ring. The device has the advantages of high reliability, long service life, low power consumption and high control precision.

The application discloses a space flight vehicle attitude control method based on a fractional order sliding mode considering an actuator, and belongs to the technical field of vehicle control. In view of the problem that the performance is reduced due to the neglect of the actuator dynamics by a traditional sliding mode attitude controller, the application firstly fuses an attitude dynamics, an aerodynamic moment and an actuator model to construct a second-order linearized attitude control model; subsequently, an extended state observer is designed to estimate and compensate the aggregated disturbance in the system in real time; further, a fractional order non-singular terminal sliding mode surface is designed, and an approaching law adopting a saturation function is combined to deduce an expected intermediate control quantity, and finally, a rudder deflection angle instruction is solved. Through the collaborative design of model construction, disturbance observation and advanced control law, the application realizes the finite time rapid convergence of the attitude angle tracking error under the premise of fully considering the dynamic characteristics of the actuator, effectively suppresses the control chattering, and significantly improves the control precision and robustness of the system under the conditions of parameter uncertainty and external disturbance.

The invention relates to a robust adaptive dynamic planning method for spacecraft attitude control, and belongs to the technical field of spacecraft control. According to the method, the neural network identifier is designed, online identification and reconstruction are carried out on unknown dynamic items caused by inertia parameter uncertainty in spacecraft attitude dynamics, the attitude tracking optimal control law based on adaptive dynamics is constructed on the basis, and under the condition that inertia parameter uncertainty exists, the attitude tracking optimal control law is optimized. The balance between tracking precision and control energy consumption is realized, and the applicability of the attitude control method in complex space tasks is improved.

The present application relates to a kind of ACMM composite control method and system for coping with interference torque uncertainty, belong to spacecraft attitude control field.The method of the present application includes: linearization modeling is carried out to spacecraft attitude, and the spacecraft dynamics model with periodic interference is formed;Interference torque w of periodic interference is modeled;For interference torque w, design disturbance rejection filter, output DRF control quantity u DRF ;Design extended state observer, state estimation and disturbance estimation are carried out to the spacecraft dynamics model of periodic interference, and after nonlinear feedback, output ESO control quantity u ESO ;DRF control quantity u DRF And ESO control quantity u ESO Superimposed, form composite control quantity u To the attitude control of spacecraft.The method of the present application makes the controlled spacecraft can converge to torque balance attitude faster, and can better cope with the influence of interference torque frequency uncertainty, effectively improve the attitude control precision and attitude stability of system.

1. Name of the design product: Spacecraft attitude control system (new type 5kg). 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.