16 results about "Satellite attitude control" patented technology

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 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.

This invention provides a novel integrated attitude sensor for satellites, relating to the fields of onboard attitude control and electronics. It includes: a star sensor body; a solar sensor body mounted on the star sensor body; a magnetometer body within the star sensor body; and a gyroscope body within the star sensor body. By integrating traditional discrete structures into a single device, the size and weight are reduced, and costs are significantly lowered. Magnetometer and gyroscope data are fused in real-time within an FPGA. This device features rapid adaptability, modular design supporting plug-and-play functionality, and integrated design improving EMC performance. It solves the problems of existing traditional satellite attitude control systems, where various attitude sensors are independently configured and equipped with dedicated processors, resulting in high hardware redundancy, large overall size and weight, high power consumption and cost, hindering further improvements in attitude measurement accuracy and dynamic performance, and leading to high overall design complexity.

The present application relates to the field of aerospace technology, especially to a kind of earth observation satellite attitude controller parameter self-tuning method, the scheme includes: obtaining the desired direction and actual direction of satellite, and determining the control torque of flywheel, the satellite is earth observation satellite;Based on the control torque, obtain initial PID parameter, initial PID parameter is PID controller simulation parameter;Based on the PID controller simulation method of flywheel physical constraint, construct the fitness function of PID controller simulation parameter based on different satellite imaging mode;Satellite attitude control is simulated by PID controller simulation parameter, based on fitness function and initial PID parameter, the optimal fitness corresponding to target PID parameter under each satellite imaging mode is determined by iterative updating through particle swarm algorithm, the optimal solution of PID parameter is target PID parameter by using particle swarm algorithm combined with fitness function for iterative updating, the accuracy and efficiency of parameter setting are improved.

The present invention relates to a system and method for a ground-based simulation of an on-orbit operating environment of a laser communication terminal. The system comprises a plurality of subsystems, each subsystem comprising: a control unit, which performs data interaction with a timing and synchronization unit, a two-axis gimbal unit, a corresponding laser communication terminal, and a test unit; the timing and synchronization unit is configured for achieving time synchronization to a common reference between the respective subsystem and another subsystem; the two-axis gimbal unit is configured for simulating, on the basis of a satellite attitude control signal, an attitude of a satellite platform carrying a laser communication terminal, the laser communication terminal driving, on the basis of a pointing control signal of the control unit, a rotating mechanism to adjust a pose of the laser communication terminal; the test unit is configured for testing the laser communication terminal; and an optical path assembly unit configured for forming or disconnecting a free-space laser link with an optical path assembly unit of another subsystem. The present invention enables simulation of an on-orbit operating environment of a laser communication terminal, and performs comprehensive and dynamic functional testing for laser communication terminals.

The application discloses a satellite solar panel sun-pointing attitude regulation method and device and a storage medium, and relates to the technical field of satellite attitude control. The method comprises the following steps: determining a first output torque according to a roll axis feedforward target angle and a pitch axis feedforward target angle of a solar panel; determining a roll angle error and a pitch angle error of the solar panel according to the roll axis feedforward target angle and the pitch axis feedforward target angle respectively, and determining an error vector according to the roll angle error and the pitch angle error; determining a second output torque according to a recurrent neural network which is trained in advance based on the error vector; and determining a third output torque according to the first output torque and the second output torque, and driving the solar panel to complete sun-pointing orientation according to the third output torque. Thus, the application realizes high-precision, self-adaptive, non-accumulative error and low-vibration sun-pointing orientation control of a low-orbit satellite solar panel under dynamic atmospheric disturbance and mechanical load drift.

The application discloses a satellite attitude control demonstration system in a simulated weightless environment, which comprises a cubic robot and a three-axis rotating table; the three-axis rotating table comprises a fixed part, and outer, middle and inner rings are sequentially arranged in the fixed part; the inner ring is a cubic frame, and the cubic robot is embedded in the inner ring; the inner ring is rotationally connected with the middle ring, the middle ring is arranged in the inner part of the outer ring, and axial interference is generated between the inner ring and the middle ring and between the middle ring and the outer ring; the inner ring rotates around an X axis to realize rolling motion, the middle ring rotates around a Y axis to realize pitching motion, and the outer ring rotates around a Z axis to realize yawing motion, so that the cubic robot is freely rotated around the X, Y and Z axes, and three-dimensional attitude motion of the satellite in the weightless environment is simulated; during the free rotation of the cubic robot around the X, Y and Z axes, friction torque is generated due to the axial interference, the friction torque offsets eccentric torque caused by the gravity center deviation, and interference suppression is realized.

The present application relates to the technical field of satellite attitude control, and particularly relates to a maximum angular momentum determination method of a wheel system with non-zero initial angular momentum. The maximum angular momentum envelope surface of the momentum wheel combination can be obtained through analytical calculation, which can effectively improve the angular momentum output capacity of the momentum wheel combination, thereby improving the rapid maneuvering capability of the whole satellite and the momentum wheel anti-saturation capability, and there is no numerical iterative search process, compared with the search method, the calculation efficiency is high, the storage amount is small, and the on-board calculation is easy to realize; in addition, the obtained maximum angular momentum envelope surface has a clear boundary, and the momentum wheel output angular momentum with the maximum envelope characteristic can be accurately obtained according to the envelope surface, compared with the pseudo-inverse method, the angular momentum output capacity of the wheel system can be more fully utilized.

A high-precision satellite attitude control method controlled by only a magnetic torque device, first, an expected control torque is calculated according to an attitude angle and an attitude angular velocity, then an expected control magnetic torque is calculated according to geomagnetic field information and the expected control torque, then an inertia angular velocity information is combined with a gravity gradient torque estimation and a previous period theory control torque estimation, a satellite residual magnetic torque is filtered and estimated by using a formula provided by the application, finally, the expected control magnetic torque is deducted from the estimated residual magnetic torque and converted into a magnetic torque device control voltage for output.

The application discloses a kind of air float satellite attitude control demonstration system and its control method, the system includes main structure part and air float simulation structure part, main structure part is divided into three layers structure arrangement actuator, power supply module, attitude measurement module, control module, solar panel;Air float simulation structure part includes spherical shell, air float platform, counterweight and air pump.The application adopts full physical entity simulation, and the attitude deviation oscillation that satellite in-orbit weightlessness environment possibly appears is reproduced truly, can realize the unfolding and retraction of solar panel, while maintaining the stability of attitude.Structure is intuitive, clear and easy to understand, suitable as teaching demonstration device.Function modularization is easy to disassemble and replace.Layered structure divides each main function, and the division of labor is clear, with good universality, can be adapted to different accessories for testing.Beneficial to student understanding different accessory purposes, different standard accessories can be replaced according to target requirement.

This application discloses a method, apparatus, and storage medium for solar panel orientation control of a satellite, relating to the field of satellite attitude control technology. Specifically, based on a first attitude transformation matrix from the inertial coordinate system to the satellite body coordinate system and a second attitude transformation matrix from the solar panel's coordinate system to the satellite body coordinate system, the target angles for the solar panel's flip axis and pitch axis are determined. A first output torque is determined based on these target angles. The flip angle error and pitch angle error are determined based on the flip axis and pitch axis target angles, respectively, and a second output torque is determined based on these errors. A third output torque is determined based on the first and second output torques. The solar panel is then driven to achieve solar orientation based on the third output torque. Thus, this application achieves solar panel orientation control of a satellite solar panel under atmospheric disturbance conditions.

The application discloses a low-orbit satellite parameter adaptive control method with a solar sail driving device, belongs to the technical field of aerospace, and particularly relates to the technical field of low-orbit satellite parameter control. The method solves the technical problem of overcoming the influence of satellite moment of inertia and compensating the influence of SADA disturbance torque and atmospheric resistance torque on the satellite attitude control precision. The method comprises the following steps: constructing a satellite attitude kinematics and dynamics equation in the form of a quaternion, and on the basis, constructing a satellite attitude kinematics and dynamics equation for attitude controller design; defining the error attitude kinematics and dynamics equation of the satellite through an attitude control error and an angle control error; dividing and converting the parameters to be estimated and known state variables, and representing the negative influence of moment of inertia error on the attitude by matrix operation; designing an attitude controller control torque in a manner of combining nonlinear feedback control with adaptive control; and updating the adaptive estimation value in the attitude controller control parameter by using a parameter adaptive law.

This invention proposes a six-segment stacked integrated navigation satellite structure suitable for mass production, belonging to the field of spacecraft structure design and manufacturing technology. It solves the problem that existing technologies cannot simultaneously meet the requirements of mass production of navigation satellites while maintaining satellite attitude control and structural stability. It comprises, in sequence, an upper payload compartment, an integrated electronics compartment, an attitude control compartment, a two-dimensional solar panel drive compartment, a propulsion compartment, and a lower payload compartment. Each segment is connected by inter-segment locking blocks and slots. All segments are secured by a fixing device. The two-dimensional solar panel drive mechanism of the two-dimensional solar panel drive compartment is connected to the two-dimensional rotating solar panel via a drive shaft. The two-dimensional solar panel drive mechanism includes a drive structure and a locking structure. The drive structure drives the drive shaft to rotate the two-dimensional rotating solar panel in two dimensions. This invention employs a six-segment stacked integrated structure design, enabling satellite production to adapt to a parallel production line model, increasing production efficiency by approximately 70% compared to traditional methods.

The application provides a satellite attitude control method and device, electronic equipment and readable storage medium. The method comprises the following steps: establishing a center-of-mass offset calculation model of the system center-of-mass position changing with the residual fuel mass of different tanks according to the comprehensive center-of-mass position obtained by superimposing the dry mass center-of-mass position and the initial center-of-mass deviation; establishing a fuel uneven discharge model of different tanks according to different initial fuel masses, nominal flow rates and uneven discharge influence factors of different tanks; identifying a target working condition from a plurality of preset working conditions according to a target initial center-of-mass deviation and target uneven discharge influence factors of different tanks; establishing a coupling disturbance torque model of the 490N engine, and substituting the thrust vector detection data into the coupling disturbance torque model to obtain the current orbit disturbance torque in real time; and dynamically controlling at least one thruster group for attitude control during on-orbit control according to the target working condition and the current orbit disturbance torque.

A fault-tolerant satellite attitude control method based on magnetic torquers is proposed. First, in each control cycle, magnetic torquer and thruster fault detection is performed based on the control signal of the previous cycle and the attitude measurement signal of the current cycle, according to the control measurement consistency score. When a thruster fault is detected or a thruster fault signal is received, the backup thruster is switched if a backup exists. When a thruster fault is detected or a thruster fault signal is received and no backup is available, the attitude control mode is switched to pure magnetic torquer control. When a magnetic torquer fault is detected or a magnetic torquer fault signal is received from other methods or injected from the ground, the faulty magnetic torquer is removed, a backup magnetic torquer is introduced, and the magnetic torquer group configuration is reconstructed for magnetic torquer allocation calculation.

The application discloses a satellite attitude maneuvering sliding mode variable structure control method based on a segmented power approaching law, and belongs to the technical field of satellite attitude control. The application discloses a satellite attitude maneuvering sliding mode variable structure control method based on a segmented power approaching law, and belongs to the technical field of satellite attitude control. The application discloses a satellite attitude maneuvering sliding mode variable structure control method based on a segmented power approaching law, and belongs to the technical field of satellite attitude control. The application discloses a satellite attitude maneuvering sliding mode variable structure control method based on a segmented power approaching law, and belongs to the technical field of satellite attitude control. The application discloses a satellite attitude maneuvering sliding mode variable structure control method based on a segmented power approaching law, and belongs to the technical field of satellite attitude control. The application discloses a satellite attitude maneuvering sliding mode variable structure control method based on a segmented power approaching law, and belongs to the technical field of satellite attitude control. The application discloses a satellite attitude maneuvering sliding mode variable structure control method based on a segmented power approaching law, and belongs to the technical field of satellite attitude control. The application discloses a satellite attitude maneuvering sliding mode variable structure control method basedon a segmented power approaching law, and belongs to the technical field of satellite attitude control. The application can meet the application requirement of multi-target observation, shorten the convergence time, improve the Euler angle precision, has the characteristics of good robustness, short convergence time and high control precision, effectively solves the problems that the prior art cannot solve the multi-target observation attitude maneuvering control in the short arc segment of the low-orbit satellite and cannot meet the multi-target observation attitude maneuvering requirement of the microsatellite.