44 results about "Control moment gyroscope" patented technology

The invention relates to the technical field of aircraft control, in particular to a control-surface-free aircraft control system based on synthetic jet and a momentum wheel, which comprises a lift force and attitude control subsystem, a steering control subsystem, a propulsion subsystem, a cooperative control unit and an attitude and navigation sensor. The lift force and attitude control subsystem independently controls different areas to generate local lift force difference through a synthetic jet generator matrix integrated on the surface, and rolling and pitching control is achieved. And the steering control subsystem changes the angular momentum direction of a flywheel through a control moment gyroscope arranged in the steering control subsystem, and generates reaction moment to control yawing. The propulsion subsystem provides a primary forward thrust. The cooperative control unit receives a sensor signal, generates an instruction according to a preset strategy, coordinates the subsystems to work cooperatively, and achieves whole-airspace flight control without an exposed control surface.

To enable a CMG to continuously maintain high-accuracy attitude control for a long time.SOLUTION: A saturation avoidance control device 6, which executes saturation avoidance control for avoiding control saturation of a control moment gyro (CMG) during attitude control of an object 1 whose attitude is controlled, comprises a torque control part that controls a torque generating device 10 that generates saturation avoidance torque τs that is coaxial with attitude control torque τg that is generated by the control moment gyro during the attitude control, without using gyro effect. The torque control part controls the saturation avoidance torque so that attitude variation of the object whose attitude is controlled by the saturation avoidance torque makes an inclination angle θg of a wheel shaft 21a of the control moment gyro gets close to an initial angle during the attitude control.SELECTED DRAWING: Figure 1

The invention relates to the technical field of control moment gyroscopes, in particular to an under-actuated brushless direct current motor pre-positioning method and device. The embodiment of the invention provides a pre-positioning method for an under-actuated brushless direct current motor. The pre-positioning method comprises the following steps: respectively judging whether a winding A, a winding B and a winding C in a three-phase winding are abnormal or not; if the at least two windings are not abnormal, selecting the two windings which are not abnormal, and determining a gating sequence of the two windings; according to the gating sequence, conducting the two non-abnormal windings with a current loop in sequence, and after the rotor is subjected to amplitude reduction oscillation swing under the action of preset current until the rotor is static, enabling the rotor to be located at a known pre-positioning position; and if at least two windings are abnormal, pre-positioning is stopped. The embodiment of the invention provides an underactuated brushless direct current motor pre-positioning method and device, electronic equipment and a storage medium. Pre-positioning can be achieved when a phase winding is disconnected.

The application discloses a kind of anti-swing assemblies based on double control moment gyroscopes, including shell, controller, attitude sensor, two double control moment gyroscopic components and two joints, shell is vertically arranged cylindrical piece, two joints are respectively arranged inside the both ends of shell, two double control moment gyroscopic components are arranged in two shells, and located between two joints, attitude sensor is arranged in the middle of lower joint, two joints are respectively used to be connected with outside setting, the gyroscopic moment direction of two double control moment gyroscopic components is mutually perpendicular in horizontal direction, controller is arranged in shell, and attitude sensor and two double control moment gyroscopic components are electrically connected with controller, two double control moment gyroscopic components are used to reduce the swing amplitude of shell, the anti-swing assembly based on double control moment gyroscopes can be applied by X direction and Y direction gyroscopic moment of two double control moment gyroscopic components in it respectively to make the swing amplitude of whole detection equipment reduce to keep stable.

Hubless spin motor-based control moment gyroscope provides full three-axis attitude control for small satellites using compact control moment gyroscopes based on hubless spin Motors, arranged in a pyramidal configuration within the satellite. The present invention comprises an a hubless spin motor, a flywheel, a gimbal (nutation) motor, a slip ring, an encoder, and essential mechanical supporting components. The said flywheel is directly attached to the hubless spin motor's rotor, generating angular momentum in both clockwise and counterclockwise directions. The slip ring and encoder are integrated with the gimbal motor in the center of the hubless spin motor, facilitating power transfer and measuring the gimbal angle, respectively. The said Hubless spin motor-based Control Moment Gyroscope is integrated into limited spaces without sacrificing performance, achieved by optimizing the shapes and mounting positions of each component.

The application provides a vibration isolation method, system and device of a control moment gyroscope system, which can be applied to the technical field of control. The method comprises the following steps: determining current kinematic parameters based on disturbance forces received by multiple components during operation; constructing a differential equation representing a dynamic relationship based on the current kinematic parameters and coupled dynamic characteristics; inputting a system state vector into a disturbance estimation network to output a disturbance estimation value; inputting the system state vector into the differential equation to obtain a disturbance theoretical value; inputting the disturbance estimation value and the disturbance theoretical value into a Hamilton function to obtain an initial control input quantity corresponding to a Hamilton function value satisfying a preset constraint condition; and updating the initial control input quantity in a gradient descent direction of a performance index function approximation value corresponding to the initial control input quantity until a difference between adjacent two performance index function approximation values is less than a preset threshold value, so as to obtain a target control input quantity.

This invention is an attitude control method for asymmetric spacecraft, comprising three steps: First, quantifying the disturbance torque of the asymmetric structure and establishing a multi-rigid-body dynamics model including a pyramid-configured control moment gyroscope group and the asymmetric spacecraft; second, designing the control law of the control moment gyroscope group, adjusting the rotation angle and angular velocity of the four frames to output control torque while avoiding singularities, for controlling the attitude of the asymmetric spacecraft; finally, designing the control law for the spacecraft base attitude, updating the inertial torque in real time by feeding back the current attitude error, and simultaneously providing nonlinear torque and compensation torque. The three torques are combined and provided as the desired control torque to the control law of the control moment gyroscope group.

The invention relates to a method for evaluating the operation stability of a high-speed shaft system of a high-torque control moment gyroscope, which relates to the field of testing and comprises the following steps of: recording the surface states of parts and bearings of the initial high-speed shaft system through a photographing microscope; the difference value between the inherent frequency of the debugging combination system and the target frequency is smaller than + / -5% of the target frequency; driving the high-speed shafting to rotate, driving the swing turntable to swing after the rotating speed is stable, monitoring the conditions of current, temperature, rotating speed curve and shafting noise of the high-speed shafting on the testing device in the running-in process in the running-in period, and replacing parts and re-testing if abnormity occurs; when the running-in process is qualified in the running-in period, the high-speed shaft system is disassembled, the surface of the part is observed through a photographing microscope, if abnormity occurs, the experiment is carried out again until the running-in process and the observation result both meet the requirements, and the running stability evaluation and assessment method has the advantage that the accuracy of running stability evaluation and assessment is improved.

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 invention relates to a multi-stage vibration isolation system for a space common-rack structure, belongs to the technical field of aerospace, and is applied to each inflatable structure of a central satellite. The multi-stage vibration isolation system comprises any two or three of a superstructure vibration absorption unit, a variable speed control moment gyroscope and a superconducting electromagnetic non-contact vibration isolation unit; through the superstructure vibration absorption unit, vibration caused by structure spinning and a high-order mode, vibration transmission near a driven propagation source and an equidistant point can be restrained, through a variable speed control moment gyroscope, an accurate active control moment is generated, and active damping and stabilization of the overall attitude and low-frequency rigid body vibration are achieved; non-linear rigidity is provided through the superconducting electromagnetic non-contact vibration isolation unit, micro-vibration before and after suppression of other two stages can be effectively isolated, function complementation, hierarchical progressive and multi-physical field coupling are achieved among unit structures, and system-level performance jump exceeding linear superposition is generated after integration of the unit structures, the superconductive electromagnetic non-contact vibration isolation unit and the superconductive electromagnetic non-contact vibration isolation unit and the superconductive electromagnetic non-contact vibration isolation unit.

The invention discloses a non-singular tandem pyramid configuration control moment gyroscope and a design method thereof, and belongs to the technical field of control moment gyroscope design, the design method comprises the following steps: step 1, establishing a tandem control moment gyroscope system of a pyramid configuration; step 2, establishing attitude dynamics models of the tandem control moment gyroscope and the spacecraft in the tandem control moment gyroscope system with pyramid configuration; 3, performing singular analysis on the torque output model of the tandem control torque gyro; 4, carrying out anti-saturation characteristic analysis on the torque output model of the tandem control torque gyro; 5, designing the control rate of the tandem control moment gyroscope with the pyramid configuration; according to the method, by increasing the degree of freedom, the designed non-singular tandem pyramid configuration control moment gyroscope can effectively avoid singularity under the condition of minimum energy consumption, and has the characteristics of simple system implementation, low power consumption, high efficiency and saturation resistance.

The present application relates to a kind of high-speed shafting running stability evaluation method of large torque control moment gyro, it is related to testing field, comprising the following steps: the initial high-speed shafting is recorded by photographing microscope, and the surface state of part and bearing is recorded, high-speed shafting is installed, and the difference between the natural frequency and target frequency of debugging combined system is less than ±5% between target frequency;Drive high-speed shafting rotation, after speed stabilizes, drive swing turntable swing, monitor the current, temperature, speed curve and the situation of shafting noise in the running-in period of high-speed shafting in the running-in process on testing device, replace part and retest when abnormal;When to run-in period is qualified, high-speed shafting is removed and disassembled, and the surface of part is observed by photographing microscope, and retest is replaced when abnormal, until running-in process and observation result all meet requirements, the present application has the advantages that the precision of running stability evaluation examination is improved.

The application discloses a sail spacecraft underactuated three-axis control method based on a planar control moment mechanism and belongs to the technical field of spacecraft attitude control and maintenance. The method is based on an attitude dynamics equation, a kinematics equation and a space geometric relation to establish a sail surface normal control law, a sail surface roll axis angular velocity control law and a sail surface roll angle control law, and then orderly combines the three control laws, so that the underactuated three-axis control is realized only by using the control moment on the sail surface. The application solves the problem that sail spacecraft represented by a solar sail and an atmospheric sail is difficult to autonomously generate a control moment in the roll axis direction to realize three-axis attitude control. The application can reduce the overall mass of a platform of the sail spacecraft due to the use of an additional roll axis moment generating mechanism, reduce the failure risk caused by the use of an additional control mechanism, is favorable for deep space exploration of the solar sail, and prolongs the service life of the sail spacecraft.

The present invention provides a composite anti-interference control method and system for a control moment gyroscope frame servo system, belonging to the technical field of mechatronic system control. According to the analysis of the working principle of the control moment gyroscope, considering the multi-source interference and uncertainty existing in the frame servo system under complex working conditions, a mathematical model of the frame servo system in the presence of multi-source interference is established, and an extended state observer is constructed to estimate the lumped interference with a bounded rate of change. Combining the interference estimation results, a fast terminal sliding mode controller forms a finite-time composite anti-interference control strategy to compensate for and suppress the influence of interference on the system control performance. The present invention effectively attenuates and suppresses the influence of multi-source unknown coupling interference on the servo accuracy of the control moment gyroscope frame servo system; has stronger robustness to unknown interference and can improve the system speed tracking accuracy within a finite time; enables the control moment gyroscope frame servo system to have good servo performance, and the actual angular velocity can accurately track the reference value with high precision.

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.

The embodiment of the present disclosure relates to a single-frame control moment gyro group on-orbit singularity analysis method. The method comprises the following steps: obtaining an angular momentum vector of an SGCMG based on each frame angle of the SGCMG, and establishing an angular momentum output equation of the SGCMG group; obtaining an output torque equation of the SGCMG group and a torque matrix of the SGCMG group according to the angular momentum output equation; obtaining a singular standard of the SGCMG by using the torque matrix and a singular frame angle of a singular state, and judging whether the frame angle is a singular point; if the frame angle is the singular point, judging the type and the singular direction of the singular point; drawing according to the singular direction and the angular momentum output equation; constructing a PID control model of the SGCMG, and simulating and analyzing a singular influence mechanism of the SGCMG group. According to the embodiment of the present disclosure, the visual drawing and the mechanism simulation running can directly show the singular point space distribution and changes of the SGCMG after the abnormality, and the spacecraft attitude maneuvering performance changes of the SGCMG group under external force interference can be accurately obtained.

The invention discloses a rigid-flexible coupling dynamics modeling method based on a rigid-flexible coupling spacecraft and application, and belongs to the technical field of spacecraft dynamics and control. The method comprises the following steps: S1, in a floating coordinate system, determining a deformation displacement vector of any point P on a neutral axis OBP of the flexible accessory; s2, determining a position vector and a deformation displacement vector of any point Q0 on the flexible accessory; and S3, constructing a rigid-flexible coupling dynamic model of the rigid-flexible coupling spacecraft. According to the method, a single-frame control moment gyroscope is used as a vibration suppression execution mechanism of a flexible accessory, and a rigid-flexible coupling dynamic model is discretized and then used for attitude control of a rigid-flexible coupling spacecraft and design of an active vibration control system. According to the method, an analysis basis and a dynamical model support are provided for efficiently suppressing the vibration of the large flexible structure of the spacecraft.

The invention provides a large control moment gyroscope group support combination in a satellite bearing cylinder, a support body combination comprises a plurality of groups of support bodies, and the plurality of groups of support bodies are installed to form a rotary symmetrical body; the bracket body combination is arranged on the bracket mounting plate; the bracket reinforcing frame is arranged on the bracket body combination; the reinforcing supporting rod is arranged between the support reinforcing frame and the satellite bearing cylinder and used for staggering the frequency of the support body combination in the satellite bearing cylinder. The support body is designed in a split mode and assembled integrally, the problem that a large control moment gyroscope group is difficult to install in a satellite force bearing cylinder is solved, and meanwhile light weight and modularization of the support combination are achieved on the premise that the design is guaranteed and the overall rigidity is improved. And the reinforcing supporting rods are designed on the bracket reinforcing frames, so that the rotating symmetric body bracket combination and the frequency in the satellite bearing cylinder are staggered, the vibration magnitude of the moment gyroscope body is reduced, the damage to a large control moment gyroscope product is avoided, and the requirement of an in-orbit use environment is met.

The present application relates to a method, apparatus, device, and storage medium for determining a single-frame control moment gyroscope group. By obtaining multiple angular momentum envelope sample points of each candidate single-frame control moment gyroscope group (SGCMGs), a target envelope value for each candidate SGCMG is determined based on the angular momentum envelope sample points of each candidate SGCMG. A comprehensive evaluation index for each candidate SGCMG is further determined based on the target envelope value. Finally, based on the comprehensive evaluation index of each candidate SGCMG, target SGCMGs are determined from each candidate SGCMG. The present application determines a comprehensive evaluation index for each candidate SGCMG based on the target envelope value. The comprehensive evaluation index is used to characterize the angular momentum output capability of the candidate SGCMGs. Furthermore, by using multiple angular momentum sample points, the problem of the inability to comprehensively and objectively judge the configuration capability caused by evaluating SGCMGs based on maximum or minimum angular momentum in the prior art is avoided, allowing the accurate selection of a single-frame control moment gyroscope group configuration that meets actual needs.

The present application relates to the field of aerospace technology, and particularly relates to a support table and a supporting method for a control moment gyro. The support table comprises a position moving device, a positioning locking device, a damping adjusting device and a mounting disc for mounting the control moment gyro. The position moving device comprises a support flexible beam, one end of the support flexible beam is connected to the mounting disc. When the positioning locking device is in a locking state, the positioning locking device locks and fixes the part of the support flexible beam in contact with the positioning locking device. When the positioning locking device is in a sliding state, the positioning locking device can slide on the support flexible beam, and the positioning locking device in the locking state on the support flexible beam can be changed. The support rigidity of the mounting disc can be changed. The damping adjusting device comprises a compression spring and a damping sheet. The damping sheet is connected to the mounting disc, and the compression spring is connected to the damping sheet. The elastic force of the compression spring is adjusted to adjust the damping of the mounting disc. The present application provides a support table and a supporting method for a control moment gyro, and the support rigidity and the damping can be adjusted.

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.

Embodiments provided herein include rotatable magnets (e.g., spherical dipole magnets) disposed within sets of coils that can be used to operate the magnets as reaction / momentum “spheres” and / or as control moment gyroscopes. The coils are able to exert three-dimensional torques onto the magnet in order to effect attitude control of a satellite or other system. The coils can also optionally exert translational forces onto the magnet in order to maintain the magnet in position and avoid contact with static components. Diamagnetic materials can also be included to provide stabilizing repulsive magnetic forces to maintain the magnet in position and / or to reduce the necessary performance of the coils with respect to applying stabilizing translational forces.

An embodiment of the present invention provides an ultra-small satellite attitude controller comprising: a casing forming an exterior and having an inner space; and multiple control moment gyro modules disposed in the inner space, each module having a gimbal motor and a spin motor for adjusting the attitude of a satellite. Each of the control moment gyro modules is equipped with an origin position detection unit for identifying the initial origin position of the spin motor through detection of a position guide bar protruding to the outside of the spin motor. The origin position detection unit is formed in a non-encoder structure so that the control moment gyro module can be manufactured in a small size.

The invention discloses a method for determining a target frame angle of a control moment gyroscope group under multiple constraints. The method can be applied to a control moment gyroscope group control system with a pentagonal pyramid structure. According to the method, the total angular momentum of the control moment gyroscope group reaches the target angular momentum, and meanwhile, the control moment gyroscope group can have high output capacity. According to the method, a complex optimization problem is simplified into a multi-constraint Newton iteration problem, the algorithm is simple and efficient, and the method is suitable for satellite on-orbit real-time calculation.

The embodiment of the invention relates to the technical field of spacecrafts, in particular to a discrete unloading method and device for a large-scale assembly control moment gyroscope. The method comprises the following steps: S1, calculating the angular momentum of a current control moment gyroscope needing to be unloaded; s2, judging whether the current angular momentum exceeds an unloading threshold value or not, and if yes, taking the current angular momentum as the angular momentum to be unloaded; s3, the jet system jets for a preset time in an on-satellite control period; s4, the unloading angular momentum of current injection unloading is calculated according to the preset injection time, the rotational inertia and the acceleration of the air injection system, and the angular momentum to be unloaded is updated according to the unloading angular momentum; and S5, judging whether the updated angular momentum to be unloaded is lower than a safety threshold value or not, if so, completing unloading, and if not, repeating the steps S3-S5 after a preset number of control periods. According to the scheme, the vibration of the flexible structure can be inhibited when the angular momentum of the control moment gyroscope is unloaded.

The invention discloses a control moment gyro system health monitoring method based on uncertainty, which comprises the following steps: S1, acquiring working condition data of a control moment gyro system, and detecting a plurality of physical quantity measurement values through a sensor; s2, taking the historical working condition data as neural network input, and taking the physical quantity data as neural network output; s3, inputting the working condition data into the trained neural network model to obtain N predicted values of each physical quantity, and calculating and obtaining a predicted mean value of each physical quantity and data uncertainty of each physical quantity; s4, obtaining an error between the measured value and the predicted mean value; s5, calculating a difference value between the error and the data uncertainty, judging whether the difference value exceeds a set threshold value or not, and if the difference value is smaller than or equal to the threshold value, judging that the health condition of the system is normal; and if the threshold value is exceeded, judging that the system health condition is abnormal. According to the method, health monitoring of the physical system of the control moment gyroscope is realized, further system reliability analysis is facilitated, and the method has a relatively good application prospect.

A non-linear double-end supported high-rigidity small control moment gyroscope belongs to the technical field of machinery, and comprises a frame assembly which is provided with a flange interface for mounting a main bracket and an auxiliary bracket and changes the direction of a rotor assembly connected to the main bracket and the auxiliary bracket by rotating the main bracket and the auxiliary bracket; the main bracket is fixedly mounted on the frame assembly and is provided with an interface for fixedly mounting the rotor assembly and supporting legs for increasing the anti-bending span; the rotor assembly is fixedly connected with the main support through a base, a rotor and a rotor motor are arranged in the rotor assembly, and the rotor motor receives the control signal transmitted through the frame assembly and drives the rotor to rotate so as to generate angular momentum required by attitude control; the auxiliary support is installed on the frame assembly and used for supporting the rotor assembly in an auxiliary mode; and the damping pad is fixedly connected to the auxiliary bracket and is pressed on the rotor assembly through a pressing force. The problems of insufficient mechanical resistance and large micro-vibration caused by insufficient rigidity in the prior art can be solved.

Systems and methods disclosed provide for attitude control of a body, the body including a plurality of thrusters and a plurality of gimballed rotors. Operations performed include: receiving a command vector representing a desired torque on the body, generating a Jacobian matrix that, when applied to a combined gimbal rate and thrust command vector, generates the command vector, generating components of a weighted pseudo inverse of the Jacobian matrix and using the components with the command vector in order to generate the combined gimbal rate and thrust command vector. Generating components of the weighted pseudo inverse of the Jacobian matrix can include generating a weighting matrix including a plurality of diagonal elements, the plurality of diagonal elements representing at least the plurality of gimballed rotors and the plurality of thrusters.

Approaches for balancing a two-wheeled vehicle, are described. In an example, balancing a two-wheeled vehicle is achieved using a control moment gyroscope (CMG) assembly which is controlled and operated based on a vehicle state information. The vehicle state information may be received from one or more sensors deployed in the two-wheeled vehicle. Once determined, the tilt corresponding to the vehicle state information may be compared a target position. The target position herein may refer to an equilibrium position wherein the two-wheeled vehicle is to remain balanced. In case the position of two-wheeled vehicle is at an offset with respect to the target position, i.e., is titled with respect to the target position, the control moment gyroscope (CMG) assembly installed in the two-wheeled vehicle may be actuated.

The application provides a distributed control moment gyro group mounting device for spacecraft configuration and layout design, which comprises micro-vibration suppression supports, control moment gyro elements and a moment gyro mounting plate; the micro-vibration suppression supports are provided in plurality, the moment gyro mounting plate provides mounting interfaces for the micro-vibration suppression supports, the plurality of micro-vibration suppression supports are arranged on the moment gyro mounting plate through the mounting interfaces, the control moment gyro elements are provided in plurality, the micro-vibration suppression supports provide interfaces for the control moment gyro elements, the micro-vibration suppression supports suppress the interference of the micro-vibration response of the control moment gyro elements on sensitive loads, and the plurality of control moment gyro elements are arranged on the plurality of micro-vibration suppression supports through the interfaces. The mounting mode has the advantages of good heat dissipation, convenient precision control and convenient operation, the interference of the micro-vibration response of the control moment gyro elements on sensitive loads is suppressed, and the stability, reliability and safety of high-precision sensitive loads on the satellite during on-orbit operation are ensured.