123 results about "Satellite attitude control" patented technology

The invention provides a satellite attitude control system fault diagnosis and fault-tolerant reconstruction method and system, and relates to the technical field of satellites, and the method comprises the steps: obtaining attitude data through a Kalman filtering state observer, building a fault diagnosis model, generating a residual sequence, and carrying out the adaptive dynamic threshold calculation to obtain a fault feature vector; and inputting the feature vector into a pre-trained deep belief network for fault diagnosis, and determining the fault type and degree. According to a diagnosis result, a mixed integer programming model considering attitude control precision and actuator power constraint is established, an improved genetic algorithm is adopted for solving, an actuator reconstruction scheme and controller parameters are obtained, and fault-tolerant control is achieved. According to the invention, faults can be diagnosed rapidly and accurately, fault-tolerant reconstruction is carried out according to fault types and degrees, the reliability and stability of a satellite attitude control system are effectively improved, and the satellite attitude control precision is ensured.

The invention discloses a multi-satellite cooperative distributed game control method and device, a computer program and a storage medium, belongs to the technical field of spaceflight, and particularly relates to a satellite attitude control and game control method. The problem that in the prior art, resolving of the overall control quantity depends on a center unit, and constraint processing on the size of the microsatellite control torque is not accurate enough is solved. The method comprises the following steps: S1, according to attitude dynamics modeling and pseudo-linearization processing of the combined spacecraft, obtaining a state dependence coefficient model; s2, obtaining a local optimal control strategy of each satellite in the combined spacecraft through a model prediction control method according to the state dependence coefficient model in combination with a dynamic model in a prediction time domain and flywheel constraint conditions; and S3, obtaining a distributed Nash equilibrium solution through a distributed iteration and global coordination method according to the local optimal control strategy of each satellite in combination with communication topology information. The method is suitable for efficient cooperation tasks of large-scale satellite groups and microsatellite systems.

The invention discloses an energy scheduling method, system and device for a low-earth-orbit satellite and a medium. The method comprises the steps that the energy storage level is calculated according to the current generation power of the low-earth-orbit satellite, the current remaining electric quantity of an energy storage unit and the health degree of a battery; generating a future illumination time sequence according to the orbit parameters of the low-orbit satellite and the satellite attitude control data; according to the current operation state data of each device loaded on the low earth orbit satellite and a preset task list, generating corresponding task prediction energy consumption data arranged in a descending order according to the importance level through a historical energy consumption model; according to the energy storage level, the future illumination time sequence and the task prediction energy consumption data, calculating the dominant power supply amount, matching the energy consumption demand in the task prediction energy consumption data, and carrying out the preferential power supply of the equipment of the task item with the high importance level. And carrying out power supply adjustment on the other equipment by adopting at least one strategy of delaying to the illumination period, reducing the operation power and shortening the operation duration so as to realize efficient scheduling on the energy of the low-orbit satellite.

The invention relates to the technical field of satellite attitude control, in particular to a satellite-borne laser semi-autonomous pointing control method, electronic equipment and a storage medium. The method is completed by cooperation of a ground system and an on-satellite system, the ground system calculates a top target passing moment T * in advance, generates a unified laser pointing control instruction and injects the instruction into a satellite; and the satellite-borne system receives and analyzes the instruction, obtains parameters such as T * and beam selection, and then autonomously controls a specified load for work preparation. And the satellite-borne system autonomously calculates the optimal moment T1 and the target attitude when passing through the calibration field through high-precision orbit extrapolation based on the latest navigation data, finally autonomously executes attitude maneuver and laser pointing tasks, and autonomously returns after the tasks are completed. Instructions in a uniform format are generated and injected from the ground, ground operation is greatly simplified, efficient cooperation of ground decision triggering and on-satellite autonomous execution is achieved, pointing errors caused by orbit extrapolation are greatly reduced, the pointing precision and the task success rate are improved, and universality is good.

An attitude control method for a spin detection satellite, which method belongs to the field of satellite attitude control. The method comprises: S10, establishing a satellite dynamic equation and converting same into a state-space equation; S20, designing a two-layer linear extended state observer, wherein the first-layer linear state observer is used for observing and acquiring a current angular velocity estimate and a current disturbance estimate of a satellite, and the second-layer state observer observes a tracking error of the first-layer linear state observer on the basis of the observation result of the first-layer linear state observer; and S30, then performing calculation on the basis of the current disturbance estimate and the tracking error thereof, so as to obtain the total internal and external disturbance torque estimate of the satellite, using attitude parameters representing the direction of a spin axis to design a PD control law, and compensating for the total internal and external disturbance torque estimate of the satellite, so as to calculate a control torque, which is used for satellite attitude control within a current cycle. The method has the advantages of fewer tuning parameters, ease of adjustment and stable performance, thereby greatly improving the observation precision.

The present invention provides a satellite flexible vibration suppression attitude control method. By adopting the idea of cascading input-to-state stability design, an attitude angle closed-loop control and an angular velocity tracking control form a cascaded system. The attitude controller is designed in a hierarchical manner. A second-order command filter is used to assign saturation constraints for tracking the virtual angular velocity and the virtual control quantity. The control method has the characteristics of small computational load and easy implementation in engineering. The flexible vibration suppression attitude controller realizes the smoothing processing of the angular velocity tracking command and the actuator control command by configuring the constraint amplitude, damping coefficient, and natural oscillation frequency of the command filter, so that the maximum angular velocity and angular acceleration of the satellite are constrained, and the angular acceleration of the excitation source of the satellite flexible attachment is limited. During the large-angle attitude maneuver, the satellite is constrained by the second-order saturation command filter, and the angular acceleration becomes zero after reaching the maximum angular velocity, and the vibration of the flexible attachment is suppressed.

The invention relates to the technical field of satellite attitude control, in particular to a satellite-borne laser full-autonomous pointing control method, electronic equipment and a storage medium. The method comprises the following steps: firstly, presetting the position of a ground fixed calibration field, a laser light emitting phase and pointing information on a satellite; the on-satellite system autonomously carries out long-term orbit extrapolation, judges whether a satellite is about to pass through a calibration field, calculates to obtain a rough target moment T * and a day and night state thereof, and autonomously notifies a ground user. And the satellite autonomously controls work preparation of laser and camera loads according to T *. And at the moment before T *, the satellite-borne system autonomously solves the optimal moment T1 and the target attitude of passing through the calibration field through iterative extrapolation calculation and preferential selection based on high-precision navigation data. And finally, the satellite autonomously maneuvers to a target attitude, laser is controlled to accurately point to the calibration field at the T1 moment, and the satellite autonomously returns after the task is completed. According to the method, on-satellite autonomy from task planning to execution of the whole process is realized, and the autonomy, the working efficiency and the pointing precision of the satellite are greatly improved.

The invention relates to a satellite-borne SAR (synthetic aperture radar) beam pointing calibration method based on antenna installation deviation and difference beam image characteristics. The method comprises the following steps: calculating beam pointing deviation caused by satellite installation of an antenna before a satellite is launched; loading the beam pointing deviation caused by satellite loading of the antenna to a satellite attitude control subsystem, and roughly compensating the beam pointing deviation of the SAR antenna; after a satellite is launched, distance direction beam pointing deviation introduced by in-orbit space environment changes is obtained according to the distance direction difference beam antenna directional diagram null depth of an imaging result, and the distance direction beam pointing deviation of the SAR antenna is precisely compensated; after the satellite is launched, according to an imaging result and difference beam image accumulation characteristics of azimuth beam pointing deviation, azimuth beam pointing deviation introduced by in-orbit space environment change is obtained, and the azimuth beam pointing deviation of the SAR antenna is precisely compensated; and the satellite performs fine compensation according to the calibration period. The method can effectively solve the problems that the ground target aiming precision is poor and the SAR image quality is reduced.

The invention discloses a fault-tolerant momentum gear train control method and device, and belongs to the field of satellite attitude control. The method comprises the following steps: determining a momentum wheel configuration scheme and a configuration scheme meeting the envelope capability based on the envelope capability of a satellite platform; if it is determined that a failure angular momentum wheel exists in the large-angular momentum momentum wheel train, target angular momentum feedforward compensation load platform rotation angular momentum is used for forming whole-satellite zero momentum, and the large-angular momentum momentum wheel train is used for compensating non-rotation-axis angular momentum generated by the large-angular momentum momentum wheel train; if it is determined that a failure torque wheel exists in the large-torque momentum wheel train, new target angular momentum is established for the large-angular momentum momentum wheel train, and the large-torque momentum wheel train absorbs angular momentum changes caused by rotating speed control of the large-angular momentum momentum wheel train so as to achieve whole-satellite zero momentum control. According to the invention, when a failed wheel appears in one wheel train, the interference caused by the failed wheel is compensated by using the other wheel train, so that the stable transition of the whole-satellite attitude is realized.

The invention discloses a parameter self-adaptive control method for a low-orbit satellite with a solar panel driving device. Belongs to the technical field of aerospace, and particularly relates to the technical field of low-orbit satellite parameter control. The satellite attitude control method solves the technical problems of overcoming the influence of the rotational inertia of the satellite and compensating the influence of the SADA vibration disturbance torque and the atmospheric resistance torque on the satellite attitude control precision. The method comprises the following steps: constructing a satellite attitude kinematics and dynamics equation in a quaternion form, and constructing a satellite attitude kinematics and dynamics equation for designing an attitude controller on the basis of the satellite attitude kinematics and dynamics equation; defining an error attitude kinematics and dynamics equation of the satellite through the attitude control error and the angle control error; carrying out segmentation conversion on a to-be-estimated parameter and a known state variable, and expressing the negative influence of a rotational inertia error on the attitude by matrix operation; designing the control torque of the attitude controller by adopting a mode of combining nonlinear feedback control with self-adaptive control; and updating the self-adaptive estimated value in the control parameters of the attitude controller by adopting a parameter self-adaptive law.

The invention discloses a satellite attitude control method and system when gyro data cannot be used and a satellite, and relates to the technical field of satellite attitude control, the method comprises the following steps: when it is determined that the gyro data of the satellite cannot be used, judging whether star sensing data collected by a star sensor configured by the satellite in a current control period is valid, if yes, executing the step 2, and if not, executing the step 3; determining the attitude data of the satellite at the current moment according to the star sensor data collected by the star sensor in the current control period; and when a preset condition is met, performing attitude control on the satellite according to the attitude data of the satellite at the current moment. In the invention, when the gyro data of the satellite cannot be used, the attitude data of the satellite at the current moment can be determined by using the star sensor data in the current control period acquired by the star sensor configured on the satellite, so that the attitude control can be effectively and timely performed on the satellite, the task implementation can be guaranteed to the greatest extent, and the satellite attitude control accuracy is improved. The engineering implementation operability is high.

The invention discloses a high-precision attitude determination method and device for high-frequency estimation and timing correction of a rotating load, and belongs to the field of satellite attitude control. The method comprises the following steps: determining the attitude of a satellite by utilizing short-period attitude estimation of a gyroscope in a star-sensitive invisible sky area; when a satellite enters a star sensor visible sky area, resetting an initial filtering estimation attitude of a gyroscope by using a star sensor attitude quaternion, and performing short-period attitude estimation based on the reset initial filtering estimation attitude of the gyroscope so as to perform fusion attitude determination by using the short-period attitude estimation of the gyroscope and star sensor correction; wherein the attitude estimation period of the gyroscope is smaller than the star sensor correction period. According to the method, the attitude quaternion estimation period of the gyroscope is reduced, the attitude estimation frequency of the gyroscope is improved, and the attitude determination strategy is fused through the long and short periods in the visible sky area of the star sensor, so that the deviation in the attitude determination process caused by the motion characteristics of the high-angular-speed rotating load is effectively reduced, and the high-precision attitude determination of the high-speed rotating load is realized.

The invention discloses a dynamic modeling and verification method and system for sailboard driving and flexible coupling, and the method comprises the steps: measuring a driving torque in a sailboard driving process under different open-loop rotation angular velocity commands of a sailboard driving mechanism, and then carrying out the calculation to obtain the angular acceleration of the rotation of a sailboard; then, obtaining a spectrogram of the sailboard driving torque and the sailboard rotation angular acceleration through fast Fourier transform, extracting the main frequency of the spectrogram, and fitting the relationship between the sailboard driving torque and the sailboard rotation angular acceleration and the open-loop rotation angular velocity command of the driving mechanism; and finally, the fitted sailboard driving torque is added to satellite rigid body attitude dynamics, the sailboard rotation angular acceleration is added to a sailboard flexible vibration equation, and the satellite attitude control angular velocity stability and the coupling vibration influence on flexible accessories are analyzed.

The invention discloses an optical satellite drift angle accurate calculation and attitude correction method and system, and relates to the technical field of satellite attitude control, and the method comprises the steps: calculating a prime unitary circle curvature radius according to a WGS84 ellipsoid parameter, and converting the geodetic coordinate of a ground target point into an ECEF rectangular coordinate; calculating Julian days based on UTC time, determining an age difference angle and a nutation angle, and converting an ECEF coordinate into a J2000 coordinate; calculating the projection of the earth rotation angular velocity in a J2000 coordinate system, and determining a relative velocity vector; determining an orthogonalized body coordinate system in combination with a unit vector from the target satellite to the ground target point; and projecting the relative velocity vector to a body coordinate system, calculating a drift angle and a double-rotation-sequence Euler angle, and correcting a real-time execution attitude by rotating a matrix around a Z axis. According to the method, high-precision image motion compensation can be realized, the imaging quality of a high-resolution optical satellite in a complex task scene is improved, and a technical foundation is laid for wide application of sub-meter remote sensing images.

The invention discloses a high dynamic load multi-information fusion attitude determination method and device, and belongs to the field of satellite attitude control. The method comprises the following steps: a satellite performs load platform attitude determination by using attitude extrapolation of a gyroscope in a star-sensitive invisible sky area; when a satellite enters a star sensor visible sky area, attitude quaternion resetting of a gyroscope is carried out by utilizing a star sensor measurement attitude, gyroscope attitude extrapolation is carried out based on the reset attitude quaternion, fusion attitude determination is carried out by utilizing the attitude extrapolation of the gyroscope and star sensor correction, and fusion attitude determination is carried out by using a first double-star sensor group at a front semi-arc section of the visible sky area; carrying out fusion attitude determination on the rear semi-arc section of the visible sky area by using a second double-star-sensor group; the two star sensors in each double-star-sensor group are integrally mounted, and the projections of the two star sensors in each double-star-sensor group on the YOZ plane coincide. According to the invention, the star-sensitive measurement precision of the high dynamic load can be ensured, and the attitude determination precision requirement of the optical satellite under the high-speed motion characteristic of the in-orbit load platform can be met.

The invention provides a method and device for determining a satellite attitude control time sequence and a readable medium. The method comprises the following steps: establishing a dynamics simulation model of multi-satellite stacking release; establishing a satellite attitude racemization control system simulation model; a delayer is arranged in the satellite attitude racemization control system simulation model, and the initial racemization time of each satellite is simulated through the delayer; carrying out joint simulation on the multi-satellite stacked release dynamics simulation model and the satellite attitude racemization control system simulation model, and for each satellite, when the angular velocity of the satellite is zero, recording as racemization consumption time; according to the initial racemization time, the racemization consumption time and the dynamic simulation model, whether collision exists in the satellite release and racemization process or not is judged, and if no collision exists, the initial racemization time serves as a satellite attitude control time sequence.

The invention discloses a satellite attitude control method and device for propulsion system faults, and belongs to the field of satellite attitude control. The method comprises the following steps: initializing a software available state of a satellite control propulsion system according to satellite control computer software; initializing a hardware available state of the satellite control propulsion system according to a hardware interface of the system; and determining whether target data is successfully recovered when the satellite control computer is reset or cut off, if so, performing satellite attitude control according to the available state of the software, and otherwise, performing satellite attitude control according to the available state of the hardware. The method can effectively improve the safety and stability of satellite operation when the propulsion system breaks down.

The application relates to a high-dynamic on-line interference estimation method for a ring-scan load on a satellite. The method designs an on-line high-dynamic interference prediction method based on an observer aiming at the periodic characteristics of the interference, and realizes high-precision estimation of fast-changing interference. The high-precision estimated value of the interference torque is used as a feedforward item of the satellite attitude control torque, so that the load interference is effectively inhibited, and the pointing accuracy and stability of the satellite are ensured.

The present application provides a method for calculating the satellite inertia matrix under abnormal conditions of satellite solar panel deployment. The method comprises: monitoring the deployment status of the satellite's solar panels in real time based on a preset sensor, obtaining deployment status data of the solar panels, and determining the deployment status of the solar panels based on the deployment status data; in response to the abnormality in deployment of the solar panels, obtaining a first inertia matrix of the satellite body and a second inertia matrix of the solar panels, and calculating the total inertia matrix of the satellite based on the first inertia matrix and the second inertia matrix. The present application achieves a more real-time and accurate calculation of the satellite's inertia matrix under abnormal conditions of solar panel deployment by obtaining the first inertia matrix of the satellite body and the second inertia matrix of the solar panels when an abnormality in deployment of the solar panels is detected in real time, and calculating the total inertia matrix of the satellite based on the first inertia matrix and the second inertia matrix. This improves the stability of the satellite's attitude control.

The present application relates to a kind of satellite load observation attitude control and data transmission, orbit control autonomous intelligent planning system and method, the system includes mission guide module, attitude control module, orbit control module and data transmission module;The attitude control module includes several thrusters for satellite attitude control, the orbit control module is used for satellite autonomous orbit control calculation, and carries out satellite orbit control, the data transmission module is used for satellite autonomous data transmission calculation, and satellite data is transmitted to ground;The mission guide module is in system control core, executes load observation task implementation strategy, according to priority reasonably arrange satellite's working time sequence, so that the satellite is intelligently switched between attitude control and data transmission to ground, orbit control, controls satellite autonomous intelligent operation in orbit.Autonomous intelligent planning, autonomous data transmission calculation, autonomous orbit control calculation are beneficial.

The invention relates to the technical field of satellite attitude and orbit control, in particular to a non-inductive coordination control method for quick zero passage of a momentum wheel under a redundant momentum wheel train. The method comprises the following steps: firstly, interpreting the change trend of the rotating speed of the momentum wheel according to the change condition of the rotating speed of the momentum wheel; secondly, judging the zero-crossing state of the momentum wheel according to the rotating speed change trend and the actually measured rotating speed of the momentum wheel; and finally, based on the characteristic of zero motion of the momentum wheel set, applying zero-crossing compensation torque to the momentum wheel in a zero-crossing state to enable the momentum wheel to quickly pass through a rotating speed zero-crossing interval, and applying reverse compensation torque to other momentum wheels to enable the resultant torque of the compensation torque of the momentum wheel set to be zero. The rotating speed of the momentum wheel quickly passes through a zero-crossing interval through a method of increasing zero-crossing compensation torque, and the influence of zero-crossing of the rotating speed of the momentum wheel on the satellite attitude control precision is reduced. The method can be popularized and applied to all satellites with zero-crossing momentum wheel rotating speed and high attitude control precision requirements.

The application belongs to the technical field of satellite attitude control, and relates to a large flexible spacecraft distributed fault-tolerant attitude and vibration control method, a large flexible spacecraft rigid-flexible coupling dynamics model with distributed actuators is established by using the Lagrange method, a hierarchical Stackelberg game control framework is established by taking an attitude control subsystem as a leader layer and a vibration suppression subsystem and a fault-tolerant control subsystem as follower layers, performance functions of the three subsystems are defined, and through game of spacecraft distributed fault-tolerant attitude and vibration control, strategy cooperation and dynamic balance between at least two control targets are realized; the application realizes three-axis attitude stabilization and low-frequency vibration suppression of a target spacecraft, and shows stronger fault-tolerant capability under node failure.

The application discloses a mode switching method applied to a variable rotating speed control moment gyro, which clearly defines multiple working modes of the control moment gyro, switching logic between behaviors of the control moment gyro in the modes and the modes, can realize non-control, moment control and backup reaction wheel functions of the control moment gyro, and can make the control moment gyro effectively applied in satellite attitude control.

The application discloses a satellite attitude control method and device based on mode conversion and a satellite, relates to the field of satellite control, and is characterized in that: a fine control mode, a coarse control mode and a sun capture mode used for controlling the attitude of the satellite and having different attitude determination accuracies are set in advance, the fine control mode is determined as a current target attitude control mode when a first remote control signal is received, the stop control mode is determined to be entered when it is determined that the satellite failure this time is a system out-of-control failure, the attitude of the satellite is regulated according to a second remote control signal when the second remote control signal sent by a ground remote control module is received, and the target attitude control mode is determined again according to the type of the failure carried component when it is determined that the satellite failure this time is a failure of the carried component used for attitude control, so that the satellite is switched to a new target attitude control mode. According to the scheme, the satellite can be switched between different modes according to different states of the satellite, reliable control of the attitude of the satellite is realized, the scheme has high flexibility, strong self-adaptability and is beneficial to practical application.

The utility model relates to a satellite three -axis integration reaction flywheel module relates to satellite attitude control technical field, including base module, non -magnetic motor, harmonic reducer, wheel disc, regular quadrangular prism and reaction flywheel, the middle part of the top of base module is equipped with round recess, the wheel disc is rotatably connected in the round recess, is equipped with the tooth on the circumference of wheel disc, the side of base module is connected with non -magnetic motor, the output shaft of non -magnetic motor is connected through harmonic reducer and the tooth of wheel disc, is used for driving wheel disc rotation, the top of wheel disc is connected with regular quadrangular prism, all are equipped with quick detachable interface on the four sides of regular quadrangular prism, the reaction flywheel is connected on the side of regular quadrangular prism through quick detachable interface. Advantages lie in: compared with traditional three motor systems more energy -conserving, reduce control coupling, be convenient for on -orbit maintenance or replacement, improve the heat dissipation efficiency, prolong the flywheel life.

The present invention discloses a variable dual-vector attitude control method for a large elliptical orbit satellite. The method comprises: determining a satellite orbit mission and segmenting the satellite orbit according to different satellite orbit missions to obtain satellite orbit segments for different missions; constructing a reference coordinate system corresponding to each mission satellite orbit segment and a satellite body attitude coordinate system; and tracking the reference coordinate system corresponding to each mission satellite orbit segment via the satellite body attitude coordinate system to perform satellite attitude control. The present invention can plan satellite attitude based on the mission characteristics of satellites at different altitude segments of a large elliptical orbit, lighting conditions, and satellite angular momentum.

The invention relates to the technical field of satellite attitude control, in particular to a compensation coefficient calculation method and device for a satellite reverse flywheel, equipment and a medium. The compensation coefficient range corresponding to the target angular momentum needing to be compensated for for the reverse flywheel is determined based on the target angular momentum needing to be compensated for for the reverse flywheel of the satellite, the preset variation corresponding to the target angular momentum, the inertia of the reverse flywheel and the maximum angular velocity of the rotating motor of the reverse flywheel; and calculating a residual angular momentum obtained after the compensation coefficient compensates the target angular momentum, and determining the compensation coefficient when the residual angular momentum is smaller than a preset minimum residual momentum as a target compensation coefficient of the reverse flywheel, so that accurate adjustment and calculation of the compensation coefficient of the satellite reverse flywheel can be realized. The accuracy of the compensation coefficient of the satellite reverse flywheel is improved, and then the shooting precision and stability of the rotary shooting device on the satellite are improved.

The invention relates to an attitude control parameter calculation method for a satellite with a large-flexibility solar panel. The attitude control parameter calculation method comprises the following steps: determining a target angle of rotation of a B axis of a two-dimensional solar panel driving mechanism relative to a zero position during sun tracking rotation of the solar panel; calculating a target rotational inertia by using the target angle and the satellite rotational inertia of the B axis of the two-dimensional solar panel driving mechanism at the zero position; and calculating PID control parameters by using the target rotational inertia. According to the attitude control parameter calculation method, for a satellite with a large-flexibility solar panel, self-adaptive calculation is carried out on the inertia change of the satellite when the B axis of the two-dimensional solar panel driving mechanism rotates to obtain the target rotational inertia, then the PID control parameters are calculated according to the target rotational inertia, and the robustness requirement of attitude control can be met.

A satellite control system fault detection method based on cooperative monitoring of energy and angular momentum comprises the following steps: firstly, setting a detection statistical period value, calculating a satellite average energy value, a theoretical angular momentum change value and a measurement angular momentum change value of each statistical period, then comparing the satellite average energy value, the theoretical angular momentum change value and the measurement angular momentum change value with a satellite average energy value of a previous statistical period, and scoring; comparing and grading the consistency of the theoretical angular momentum change value and the measured angular momentum change value, comprehensively judging the grading value, giving a satellite attitude control abnormity mark, and when the satellite attitude control abnormity is confirmed, performing subsequent processing to enable the satellite to autonomously switch into an attitude control safety capture mode.

The application relates to a flywheel speed control method and device based on a polynomial trajectory planning, belongs to the flywheel speed control method and device based on the polynomial trajectory planning, and solves the problem of how to reduce the impact disturbance introduced when a flywheel tracks an expected speed instruction of attitude control and improve the attitude stability of a satellite. The method comprises the following steps: receiving an expected speed instruction sent by an attitude control system; for each attitude control period of the attitude control system, a speed response curve between two adjacent expected speed instructions is constructed by adopting an expected speed sequence based on the polynomial trajectory planning, so that the acceleration is smooth and the jerk is continuous. The flywheel speed control method and device based on the polynomial trajectory planning are suitable for flywheel speed control in satellite attitude control.