602 results about "Attitude control system" patented technology

The invention relates to a spacecraft attitude integral sliding mode fault tolerance control method taking consideration of a performer fault and provides a robustness attitude active fault tolerance control method based on an integral sliding mode surface for problems of the performer fault, external disturbance and control moment amplitude limits in a spacecraft attitude control process. The method comprises steps that firstly, a spacecraft attitude dynamics model taking consideration of the performer fault and containing external disturbance is established; secondly, on the condition that a performer is not in fault, a designed nominal controller can guarantee system stability, and input saturation amplitude limits can be easily satisfied through adjusting controller parameters; lastly, the fault information is introduced to design an integral sliding mode controller, robustness of external disturbance and the performer fault can be effectively improved, system stability is analyzed on the basis of an Lyapunov method. The method is advantaged in that stability of the attitude control system is guaranteed when a spacecraft operating on orbit generates the performer fault, and relatively strong fault tolerance capability and external disturbance robustness are realized.

Various embodiments of the present invention include an attitude control system for use with small satellites. According to various embodiments, the system allows rapid retargeting (e.g., high slew rates) and full three-axis attitude control of small satellites using a compact actuation system. In certain embodiments, the compact actuation system includes a plurality of single-gimbaled control moment gyroscopes (SGCMG) arranged in a pyramidal configuration that are disposed within a small satellite.

The invention discloses a reentry aircraft finite time control method based on a disturbance observer and relates to the reentry aircraft finite time control method. The reentry aircraft finite time control method belongs to the field of aircraft control technology. The reentry aircraft finite time control method comprises the following steps: a first step, establishing a reentry aircraft dynamic model, presenting a finite time gesture tracking task; a second step, performing feedback linear processing on a model which is established in the first step; a third step, providing a finite time control law; a fourth step, providing a finite time disturbance observer for estimating system uncertainty and outer disturbance; and a fifth step, providing a finite time global sliding mode control method based on the disturbance observer, and effectively improving tracking precision of a gesture control system through substituting an estimated disturbance value into the control law. The reentry aircraft finite time control method has functions of improving system error convergence speed, preventing a singular problem and improving uncertainty of the controlled system to the parameter, global robustness of outer disturbance and tracking precision.

The invention discloses a dual-redundancy attitude control system of a coaxial unmanned helicopter, which consists of a serial communication module, a PCM (Pulse-Code Modulation) decoding module, an AD acquisition module, a PWM (Pulse-Width Modulation) output module and a CPU control module. The serial communication module, the PCM decoding module and the AD acquisition module are used as system input, the CUP control module carries out improved digital PID control, and the PWM output module carries out output in the form of PWM pulses, and thus, the control of a backward-stage high power valve is realized. A debug method comprises the following fives steps of: firstly, connecting a defined I/O port to a high or low-level end according to a parameter debug object; secondly, adjusting a potentiometer to provide initial parameters for Kp and Kd and controlling the backward-stage high power valve; thirdly, keeping a neutral position of Futaba unchangeable, adjusting the potentiometer so that a steering engine achieves a critical state of self-excitation vibration; fourthly, operating the Futaba to output a signal exceeding a pulse threshold; and fifthly, operating the Futaba to return to a central position. The method realizes the debug and the storage of relevant parameters, decreases interference and improves the system stability and the response speed.

The invention discloses an inner system for controlling the gestures of an underwater robot, comprising a main controller connected with an adjustment mechanism, a gesture sensor, and an upper computer controller, wherein the adjustment mechanism is a balance adjustment mechanism arranged inside the underwater robot, and comprises a gravity slider, a double-shaft drive motor, a worm, a guiderail, a slide bar, a limit switch and a mechanism platform; the double-shaft drive motor is connected with the worm to drive the worm to rotate bidirectionally; the worm drives the slide bar to perform planar movement through the transmission of a lead screw; one end of the slide bar is connected with the lead screw of the worm and penetrates the gravity slider, while the other end thereof is connected with the guiderail by a slider and drives the gravity slider to move in a plane so as to realize the gravity adjustment of the underwater robot and control the transverse swing and longitudinal incline gestures of the underwater robot. The control system for gestures can be widely applied to the gesture control of the underwater robot.

The invention relates to a steering-law singularity avoiding aircraft attitude control system, comprising an attitude task managing unit, an attitude controller, an attitude measuring unit, a gyro group steering-law unit, a frame angle position measuring unit, a control moment gyro group and an aircraft. The gyro group steering-law unit calculates the frame angle rate of the control moment gyro group by cooperating control moment gyro angle momentum table, zero motion algorithm and pseudoinverse steering-law algorithm, and adopts the angle rate value as the set signal of the control moment gyro group to improve the control precision of the aircraft attitude. The steering-law singularity avoiding aircraft attitude control system introduces the steering-law design method that chooses frame angle rate and the initial value of frame angle position based on look-up table, realizes accurate moment output of the control moment gyro group, avoids the singularity problem produced in the process of calculating frame angle rate of the control moment gyro group by the pseudoinverse steering-law algorithm as well as avoids the deadlock problem caused by the control moment gyro steering law when adopting zero motion algorithm and robust pseudoinverse steering-law iterative computations.

The invention discloses an attitude control system and a control method of a four-rotor aircraft. Firstly, the information of attitude angles of the current aircraft is obtained, and the attitude angles include a roll angle, a pitch angle and a yaw angle; secondly, controllers for the roll angle, the pitch angle and the yaw angle are respectively designed, wherein the controller for the roll angle is the same as that for the pitch angle; and finally, after obtained control inputs of the three controllers are overlapped, a total driving signal is output, thereby driving a motor to operate, so as to control the attitude of the aircraft. A gyroscope and an accelerometer are combined, and a Kalman filter is designed in accordance with models of the attitude angles of the aircraft, so as to conjecture the attitude angles on line. Additionally, variables of the intermediate state of the system can be obtained, and a certain filtering effect is achieved.

Based on a neural network technology and an instruction filtering inversion method, the invention discloses a robust fault tolerance control system design structure based on instruction filtering inversion. Firstly, a mathematical model of an NSV attitude control system is given out, uncertainties and external disturbance caused by modeling errors are considered based on the mathematical model, a state equation of the NSV attitude control system under the fault of a control surface is also considered. The method mainly includes the two design parts including design of an auxiliary system and design of a controller based on the auxiliary system. The auxiliary system is introduced into the neural network to ensure robustness of the auxiliary system, and stability of a closed-loop system is strictly proved through Lyapunov theorem. Meanwhile, simulation is carried out on an attitude control system of a corresponding air vehicle, and a result shows that the method enables the uncertain flight control system with the external disturbance to have the ideal fault tolerance tracking performance under damage to the control surface.

The invention relates to a flexible spacecraft attitude control system and a flexible spacecraft attitude control method in allusion to flywheel low-speed friction. The system comprises six modules which are a flexible spacecraft dynamics real-time simulation module with actuating mechanism characteristics, a flexible spacecraft kinematics real-time simulation module, an attitude measurement module, an attitude determination module, an attitude control module and an executing mechanism module, wherein the flexible spacecraft dynamics real-time simulation module with the executing mechanism characteristics comprises a spacecraft body, a flywheel and flexible appendage dynamics; the flexible spacecraft kinematics real-time simulation module can select different types of attitude description modes according to task requirements; the attitude measurement module can select different real sensors and sensor simulators according to the task requirements; the attitude control module comprises a conventional PID control method, a robust control method and the flexible spacecraft attitude control method in allusion to flywheel low-speed friction, and switching can be carried out according to the task requirements of the system; and the executing mechanism module comprises a real flywheel and a thruster simulator.

The invention relates to a visual dynamic simulating device for deposition in a gas hydrate pipeline, which is used for simulating the generation, the transportation and the deposition of gas hydrate in high-pressure high-speed air flow and can be used for summarizing the high-speed motion characteristic, the transportation mechanism and the deposition rule of a pipeline gas-solid phase. The invention has the technical schemes that: the top end of a low-pressure gas storage tank is provided with a temperature sensor and a pressure sensor, the outlet end of the low-pressure gas storage tank is connected with a high-pressure gas pump and then is connected with a high-pressure gas storage tank, and the top end of the high-pressure gas storage tank is provided with a pressure sensor; the outlet end of a liquid storage tank is connected with a metering pump, and the upper end of a gas-liquid separator is connected with a vacuum pump and a tail gas treating device; the top end of a high-pressure reaction tube is connected with an outlet electrical heating insulation pipeline, and the bottom end of the high-pressure reaction tube is provided with a pressure sensor and an inlet electrical heating insulation pipeline; and a data acquisition system is formed by a data acquisition card, a computer, a cold light source and a high-speed camera, and a pipeline temperature attitude control system is additionally arranged. According to the visual dynamic simulating device disclosed by the invention, the generation process, the transportation process and the deposition process of the gas hydrate in the high-pressure high-speed air flow can be really simulated, and the visual dynamic simulating device is used for evaluating a gas hydrate inhibitor.

Various embodiments of the present invention include systems and methods for assessing the performance of an actuator of an attitude control system (ACS), such as a control moment gyroscope (CMG). In one embodiment, a system includes a support bracket assembly coupled to an actuator, wherein the actuator is configured to generate an output torque. The system also includes at least one sensor assembly that includes a sensor configured to measure the output torque about at least one axis of the support bracket assembly while the support bracket assembly remains substantially motionless.

The invention discloses a rigid spacecraft performer multi-fault diagnosis and fault tolerance control method. According to the method, the kinetic and dynamic model of a rigid spacecraft attitude control system is put forward, a fault model for existing of both the performance failure fault and the deviation fault of a rigid spacecraft is established, and then a fault detection observer adopting the adaptive threshold technology and a fault estimation observer based on the adaptive technology are respectively established so that online real-time detection and estimation of the fault time and the concrete situation of the fault can be realized, and finally a backstepping sliding mode fault tolerance controller is designed according to fault information established by the fault estimation observer. Attitude stabilization control of the rigid spacecraft under the condition of the performer efficiency damage and the deviation fault can be realized, and the influence of external disturbance on the system and the observers can also be considered in the design process. Besides, the fault detection observer and the fault estimation observer can be independently designed so that the engineering application can be more easily implemented.

The invention provides a method for performance evaluation of a spacecraft attitude control ground simulation system. According to the method, the credibility of a dynamic system, a kinematics system, a measurement system and a control system is studied based on the simulation system credibility theory, the overall credibility of the spacecraft attitude control ground simulation system is obtained through comprehensive calculation, and then the performance of the spacecraft attitude control ground simulation system is evaluated. According to the evaluation method, the credibility of the ground simulation system and an actual satellite attitude control system is considered from all aspects, simulation data are fully utilized, and the credibility of the ground simulation system is described objectively and quantitatively; the theoretical basis is reliable, the calculation process is simple, the application range is wide, and a new idea and reference are provided for performance evaluation study of other simulation systems.

Various embodiments of the present invention include assemblies and methods for utilizing antennas with high gain in small satellites. In one embodiment, a satellite comprising a payload configured for transmitting data is provided. The payload may include various components of the satellite, such as the attitude control system, electrical power system, and/or communication system. The satellite may be configured to communicate with one or more ground stations. The satellite includes a support structure comprising at least one deployable panel, wherein the support structure houses the payload. The satellite also includes at least one antenna coupled to the support structure, wherein the deployable panel is configured to cover the antenna in a non-deployed state and to expose the antenna in a deployed state.