7 results about "Control moment gyroscope" patented technology

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.

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 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 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.

A satellite attitude maneuvering method after specific failure reconstruction of a pentagonal pyramidal SGCMG configuration is disclosed, belonging to the field of spacecraft attitude control technology. This includes a single-axis pointing maneuvering strategy for satellites equipped with a pentagonal pyramidal configuration and requiring single-axis spatial pointing. The designed method ensures that the angular momentum of the control moment gyroscope combination is in an optimal state under the target observation attitude by designing the initial angular momentum configuration and target attitude. The target maneuver is divided into two stages: singularity avoidance and target maneuver, ensuring that the angular momentum of the control moment gyroscope combination does not become singular during the maneuver. The maneuvering strategy proposed in this invention guarantees that the satellite does not experience angular momentum singularity during the target maneuver and that the angular momentum is in an optimal state during target observation. For satellites that lose three-axis maneuvering capability due to partial failure of the control moment gyroscope, this method ensures their mission execution capability and improves satellite availability.