299 results about "Control moment gyroscope" patented technology

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 relates to a double-frame magnetic-suspension CMG (Control Moment Gyroscope-CMG) control system, comprising a double-frame magnetic-suspension CMG body, a magnetic-suspension rotor control unit, an inner-frame control unit, an outer-frame control unit and a compound compensation control unit comprising a feed-forward compensator and a feed-back compensator. The feed-forward compensator carries out a feed-forward compensation for an inertia coupling disturbing moment of the rotor and the inertia moments of the inner and outer frames, and the feed-back compensator carries out the compensation for the impact of the gyroscope action disturbing moment; the outputs of the feed-forward compensator and feed-back compensator are respectively added to the outputs of the magnetic-suspension rotor control unit, inner and outer frames control units to obtain a total control value so as to realize the stable control to the double-frame magnetic-suspension CMG on the basis of compensating a dynamic frame effect. The double-frame magnetic-suspension CMG control system of the invention adopts the compensation to the output moments of magnetic bearings and frame motors to offset the coupling disturbing moment caused by the rotation of the inner and outer frames and to eliminate the displacement of the dynamic frame of the magnetic-suspension rotor, and enhances the response speed and accuracy of frames and the whole double-frame magnetic-suspension CMG.

The invention is an integrating double-framework magnetically suspended control moment gyroscope (MSCMG) magnetic bearing control system, comprising a control circuit, a collecting circuit and a communication interface; wherein, the control circuit comprises a FPGA module, a DSP module and the like; the collecting circuit consists of current, a displacement sensor interface circuit, a double-framework revolution / position interface circuit, an A / D module, and a D / A module; the communication interface comprises a USB interface, RS232 and an SPI interface; the control circuit in the system is used for receiving gyroscope rotor signal processed by the collecting circuit, controlling the stable suspension thereof, transmitting the controlled variable to an upper computer through the communication interface, and realizing on-line monitoring on the suspended gyroscope rotor, data collecting and on-line modification for the controlled variable under the control of the upper computer. The system provided in the invention has the advantages of effectively communicating an inner framework and an outer framework rotational signals under the control of high precision, realizing the integrating design of the control system and the measurement and control system, and satisfying the requirements of speed and precision for the data collecting and the on-line monitoring.

A modular control moment gyroscope (CMG) system for a spacecraft attitude control system (ACS) is formed by a plurality of CMG modules, wherein each CMG module has a modular enclosure design that is identical to that of the other CMG modules, such that the plurality of CMG modules are mountable in a spacecraft array bus structure in any desired one of multiple array configurations.

A moment distribution method for a rapid maneuvering satellite based on a mixed actuating mechanism relates to the technical field of spacecraft attitude control and solves the problems that the precision is low due to the fact that single gimbal control moment gyroscopes are in a dead zone and the actuating force is weakened after the satellite taking single gimbal control moment gyroscopes and flywheels as actuating mechanisms is maneuvered rapidly and maneuvered. The realizing process of the moment distribution is as follows: obtaining the gimbal angular speed distributed to the single gimbal control moment gyroscopes and the angular acceleration speed distributed to the flywheels, and assigning value to the optimized gimbal angular speed, meanwhile, judging whether each single gimbal control moment gyroscope is in the dead zone or not, if so, stopping using the single gimbal control moment gyroscope, otherwise, returning to the step two to obtain a new gimbal angular speed, comparing the new gimbal angular speed with the optimized gimbal angular speed, if the new gimbal angular speed and the optimized gimbal angular speed are different, storing the new gimbal angular speed into the optimized gimbal angular speed and returning to the step two; and if the new gimbal angular speed and the optimized gimbal angular speed are the same, assigning the new gimbal angular speed to the final gimbal angular speed of the single gimbal control moment gyroscope, and assigning the angular acceleration speed of the flywheels to the final angular acceleratioin speed of the flywheels. According to the moment distribution method, the satellite attitude can be adjusted.

A spacecraft posture fault tolerance control method of single gimbal control moment gyroscope groups includes the following steps that firstly, a kinetic equation and a kinematical equation are established when partial failures exist in the single gimbal control moment gyroscope groups (SGCMGs), wherein a coordinate system is defined, and a control system state equation is established; secondly, on the basis of in-orbit operation characteristics of a spacecraft, the spacecraft posture fault tolerance control method of the single gimbal control moment gyroscope groups is adopted; thirdly, the sliding mode control law is designed, wherein a sliding mode face is designed, the control law is primarily designed and the control law is improved. The method has the advantages that the designed control law is also applicable to fault tolerance control of aircrafts with flywheels as execution mechanisms; people do not need to learn about prior information of faults, real-time estimation of fault information and interference information is achieved through self-adaptation control, and fault time varying is allowed; the method is applicable to SGCMGs of any structure or partial failure modes of flywheels; the method is used in aircrafts of the SGCMGs, the rotation speed of a gyroscope frame serves as direct control amount, and the method conforms to engineering practice.

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 embodiment herein generally relates to a system to apply moments to a user, particularly during gait. The system is a portable gyroscopic-assisted system, mounted on a user's body, particularly the upper body, particularly to influence orientation of users having difficulty with balancing during gait. The system comprises a plurality of variable-speed control moment gyroscopes (VSCMGs). The VSCMGs generate moments, particularly to counteract a fall to any direction. The VSCMGs are placed close to the center of mass of the user. The couple moments of the VSCMGs are transmitted to the user through a support structure that is tightly attached to the user. Particularly, the system controls moments based on detecting pre-fall conditions.

The invention relates to a method for controlling the structural modal vibration of a dual-frame magnetic levitation control moment gyroscope (CMG). On the basis of establishing a nominal model of an electromagnet rotor system, a sine-swept excitation mode is adopted to perform the identification of frequency domains on an electromagnet rotor and measure the disturbance of system parameters, the properties of structural vibration modality and the high-frequency unmodeled dynamics of an electromagnet rotor model; and a comprehensive robust control method of a structural singular value mu is utilized to select a weighting function and inhibit the structural modal vibration. The method belongs to the technical field of aerospace control, and can be used for inhibiting the structural modal vibration of the dual-frame magnetic levitation CMG.

The servo control system of magnetically suspended control moment gyroscope frame with precise friction compensation includes mainly a magnetically suspended control moment gyroscope frame, a frame servo moment motor, a frame servo system controller and an adaptive friction compensator, which includes a frame servo system non-linear model identification unit and a frame servo system nominal model unit. The adaptive friction compensator is introduced for real-time acquiring the current signal and angular rate signal of the frame servo system and precise non-linear model identification, so as to compensate the friction of the frame servo system, regulate adaptively the friction moment compensation coefficient, raise the angular rate output precision and response speed of the frame servo system and improve the moment output precision and response speed of the whole magnetically suspended control moment gyroscope system.

A control moment gyroscope (CMG) framework control system and a CMG framework control method for restraining dynamic unbalance disturbance of a rotor are provided. A feed-forward compensation module and an algorithm switching module are embedded on the basis of a double-loop PI control system. In preset Ng control cycles, the whole framework control system is set to a PI control mode; in each control cycle from the (Ng+1)th control cycle, the speed precision of a CMG framework in the current control cycle is calculated through the algorithm switching module, and whether there is a need to switch the control mode of the current control cycle is determined; if the control mode is a feed-forward compensation mode after switching, subtraction operation is carried out on a moment reference value output by a speed loop PI controller and a fed-back torque of a framework motor to get the mechanical torque error e(j+1) of the current control cycle, and the feed-forward compensation module calculates the current feed-forward compensation value Delta i(qref, j+1) of the current control cycle according to e(j+1) and adds Delta i(qref, j+1) to the double-loop PI control system; and if the control mode is a PI control mode after switching, current feed-forward compensation is blocked, and control is performed according to the double-loop PI control system.

The invention relates to an integrated digital control system for a double gimbal magnetically-suspended variable-speed control moment gyroscope. The integrated digital control system mainly comprises a DSP and peripheral module thereof, an FPGA module, an interface circuit, a signal conditioning circuit, a power amplification module, an RS422 communication interface module. The control system obtains signals through the interface circuit and outputs the signals to the DSP; the DSP generates magnetic bearing, high-speed motor and frame system control quantity through a control algorithm and outputs the magnetic bearing, high-speed motor and frame system control quantity to the FPGA module; the FPGA module carries out PWM modulation on the control quantity and outputs PWM signals to the power amplification module, thereby realizing the integrated control of a magnetic bearing, a high-speed motor and a frame system. Besides, the real-time communication between the DSP and an upper computer is realized through the RS422 communication interface module, thereby realizing on-line adjustment of control parameters and on-line monitoring of the running status. The integrated digital control system helps to realize the high-precision high-reliability integrated control of the double gimbal magnetically-suspended variable-speed control moment gyroscope, and effectively helps to reduce the size, the weight and the cost of the control system.

According to an aspect of the present invention, there is provided an autonomous mobile apparatus including a movable body, a control moment gyro that generates a torque, a gyro unit that pivotably supports the control moment gyro about a first axis and a gimbal that pivotably supports the gyro unit about a second axis and is pivotable to the movable body about a third axis that is different from the second axis.

The present invention relates to a method for maneuvering an imaging satellite, and more particularly a method for commanding control moment gyroscopes on an imaging satellite to change the attitude of the satellite. In one embodiment, a method for generating a gimbal rate trajectory for maneuvering a satellite to point to a target includes providing a final attitude for pointing to the target and a final satellite rate for imaging the target, providing an initial attitude and an initial set of gimbal angles, and determining a gimbal rate trajectory from the initial set of gimbal angles to a final set of gimbal angles.

The invention relates to a parameter selection method adopting an onboard control moment gyroscope group vibration-isolating platform and belongs to the field of the attitude control and the vibration and shaking control of a spacecraft. (1) an idea of the cube vibration-isolating platform is adopted, the cube vibration-isolating platform is arranged between a configuration consisting of a plurality of control moment gyroscopes and a satellite body and the high frequency vibration caused by a plurality of control moment gyroscopes is isolated to the greatest extent; and (2) by reasonably selecting parameters of the onboard control moment gyroscope group vibration-isolating platform, the problem of effectiveness loss of other onboard systems, which is caused after the vibration-isolating platform is applied on a satellite, is avoided.

The invention relates to a satellite multiple attitude control mode test system based on a double gimbal control moment gyroscope (DGCMG) structure. Performance of three satellite attitude control systems which are based on a DGCMG, a single gimbal control moment gyroscope (CMG) and a flywheel is tested and verified through a DGCMG actuating mechanism. The satellite multiple attitude control mode test system based on the DGCMG structure comprises a platform system, the satellite attitude control system, a space environment simulation system and a ground station system. The platform system is composed of a tri-axial air bearing table, a satellite service comprehensive management system, a power source and a wireless bridge and used for simulating satellite dynamic characteristics and information management. The satellite attitude control system is composed of a jet propulsion system, the DGCMG, a fiber-optic gyroscope, a star sensor, a sun sensor and a global position system (GPS) receiver and used for determination of an attitude and an orbit and control of a satellite platform. The space environment simulation system is composed of a GPS simulator, a sliding block which is of a pyramidal structure, a sun simulator and a star simulator and used for simulating space interference torque, and part performance of a GPS satellite and a celestial body. The satellite multiple attitude control mode test system based on the DGCMG structure can provide ground testing and verification for multiple attitudes of a satellite.

A control moment gyroscope (CMG) is provided for deployment on a spacecraft. The CMG includes a stator housing, an inner gimbal assembly disposed in the stator housing, and a torque motor coupled to the stator housing and at least partially disposed around an intermediate portion of the inner gimbal assembly. The torque motor is configured to rotate the inner gimbal assembly about a gimbal axis.

The invention discloses a small-size speed change control moment gyroscope, and belongs to the technical field of control moment gyroscopes. The small-size speed change control moment gyroscope comprises a high-speed component, a connection bracket and a low-speed component, wherein the high-speed component supplies constant angular momentum in a control moment gyroscope mode and outputs fine moment required by whole satellite attitude control in a flywheel mode; the connection bracket is connected with the high-speed component and the low-speed component and guarantees the orthogonal perpendicularity between the angular momentum direction and the axial direction of a framework; the low-speed component supports the high-speed component and the connection bracket and supplies a mounting interface for a whole satellite; in the control moment gyroscope mode, large moment which is required by fastmoving of the whole satellite and is also orthogonal to the angular momentum direction and the axial direction of the framework is generated; in the flywheel mode, the fine moment output by the high-speed component is transmitted to the whole satellite for the attitude control. The speed change control moment gyroscope disclosed by the invention realizes the high-precision framework locking, the minimization, the light weight and the high rigidity, and is suitable for the application to small quick satellites.

Methods and systems are provided for orienting an agile vehicle using a control moment gyroscope array. A method comprises obtaining initial vehicle parameters for the vehicle and obtaining target vehicle parameters for the vehicle. The method further comprises determining command parameters based on a difference between the target vehicle parameters and the initial vehicle parameters, and simulating operation of the control moment gyroscope array using the command parameters and a torque value being at least equal to a maximum achievable torque for the control moment gyroscope array. When the simulated vehicle parameters are substantially equal to the target vehicle parameters, the method further comprises determining a torque profile for the control moment gyroscope array based on the simulated operation and operating the control moment gyroscope array using the torque profile.

The invention relates to an angular momentum management method for a variable-speed control moment gyroscope (VSCMG). A technical scheme is described as follow. An angular momentum management algorithm for the VSCMGs is designed to mainly solve the following problems: mixed mode instruction torque output based on singularity measurement, zero-motion rotor wheel speed balance, zero-motion frame configuration singularity avoidance, frame angular velocity dead-zone nonlinear processing and neglected item compensation, and gives corresponding theoretical analysis. The mixed mode of the VSCMGs gives full scope to the large torque output of the control moment gyroscope and large-angle fast attitude maneuver operation can be achieved by using the steering law. The present invention provides a novel VSCMGs angular momentum management algorithm which is improved significantly in torque output accuracy, angular velocity tracking, and frame singularity avoidance compared with a conventional angular momentum management algorithm, and which has a good engineering value.

The present invention discloses a liquid filling spacecraft posture dynamics physical simulation test system and method. The system comprises a set of liquid slosh moment simulation systems. The liquid slosh moment simulation system comprises: a liquid slosh moment generator formed by a plurality of single-frame control moment gyroscopes, wherein the liquid slosh moment generator is installed on the body of an air bearing table and performs angular momentum exchange with a triaxial air bearing table to simulate the interference of the liquid slosh on the attitude of a satellite; a liquid sloshcalculation communication module used for moment calculation and telemeasuring data forwarding; a ground monitoring module used for system state monitoring and parameter lettering; and a power moduleconfigured to control the control moment gyroscope set and calculate the power supply of the communication module. The system and the method are suitable for the physical simulation test of the triaxial air bearing table to analyze the influence of the liquid slosh on the attitude of the spacecraft so as to more really and effectively check and verify the control system, and the simulation methodcan avoid the downfall phenomenon of the air bearing table.

The invention relates to a rapid high-precision relative pointing control method of a noncoplanar rendezvous orbit, in particular to a feedforward PID (proportion, integration and differentiation) control based rapid high-precision relative pointing control method of the noncoplanar rendezvous orbit, and aims to solve the problem that rapid high-precision tracking-pointing control methods related to satellite attitudes under the noncoplanar rendezvous orbit do not exist in the prior art. The feedforward PID control based rapid high-precision relative pointing control method including adopting an Euler angle for attitude description of a spacecraft, creating dynamical and kinematical equations of the spacecraft, and acquiring an precise expected angle theta according to a noise-containing expected angle z by a Kalman filtering algorithm of a spaceborne computer; designing an attitude control law of each axis according to a formula of ; selecting two parallelly-placed single-frame control moment gyroscopes to control a yaw axis, and selecting two flywheels to control a rolling axis and a pitch axis respectively; calculating gyroscopic moment T and flywheel actual output moment uw to complete rapid high-precision relative pointing control of the noncoplanar rendezvous orbit. The feedforward PID control based rapid high-precision relative pointing control method is applicable to rapid high-precision relative pointing control of the noncoplanar rendezvous orbit.

The invention discloses a planet steady state controlling method replacing momentum wheels with a control moment gyroscope. When the number of the momentum wheels which work normally is three, the momentum wheels in a traditional way can only work in an angular momentum zero passing way to realize the state steady controlling, and the controlling precision is lowered. Aiming at the problem, the invention puts forward that a back-up control moment gyroscope high-speed rotor is used as an angular momentum fixing momentum wheel, a fixing bias angular momentum of a certain direction is provided by rotating a low-speed frame angle, so that the rest three momentum wheels can work in a bias state, the system can stay in a steady state controlling mode of a zero-momentum controlling system formed by the back-up control moment gyroscope high-speed rotor and the momentum wheels which work in a bias angular momentum state, and the problem that the precision when the momentum wheels are in zero passing use is damaged is solved. The planet steady state controlling method enables the cold back-up control moment gyroscope to play the back-up role of side-sway maneuver and steady state controlling, and improves the reliability of on-rail use of a planet activating mechanism.

The present invention provides a method for designing the parameter of the wave trap of the magnetic suspension flat high-speed rotor system. The magnetic suspension flat high-speed rotor system is considered with a kinetics model of the first order elastic mode and is translated into a complex-coefficient single-variate open-loop transfer function. The double-frequency Bode chart of the open-loop transfer function of the complex coefficient in different rotary speed is drawn. The phase angle allowance of elastic mode in each rotary speed is calculated, and the series-connected progression, central frequency and polar point amortization of the elastic self-excited vibration restraining used wave trap are designed according to the phase angle allowance curve. Basing on the double-frequency Bode chart the invention presents a method for designing the parameter of the wave trap which comprises a magnetic suspension flat high-speed rotor system with strong gyroscopic effect and is used for restraining the first order elastic self-excited vibration. Compared with the common multiple-variate system designing method such as state space analyzing method and the like, not only has excellent immediacy and robustness, but also can minimize the effect of the wave trap to the nutation stability, and thus the invention the invention is more suitable to be used in the actual magnetic suspension high-speed rotor system such as posture-controlling energy-accumulating fly wheel, the magnetic suspension counteracting wheel, the magnetic suspension control moment gyroscope and the like.

A high-precision magnetic suspension control torque gyro frame servo control system mainly includes a magnetic suspension control torque gyro frame, a frame servo torque motor, a frame servo system controller, a torque motor current detection circuit, a frame angular rate detection circuit, and an interference suppressor. In the present invention, the torque motor current and the frame angular rate are detected in real time by the current detection circuit and the frame angular rate detection circuit, and then sent to the interference suppressor to calculate the corresponding interference compensation amount, and the frame servo system controller adjusts the output current to realize the interference inhibition. The application of the invention overcomes the influence of the friction torque on the angular rate control accuracy of the frame servo system, suppresses the disturbance of the magnetic suspension high-speed rotor to the frame servo system and other unknown disturbances, improves the angular rate output accuracy and response speed of the frame servo system, and improves the performance of the frame servo system. Magnetic levitation controls the accuracy and response speed of torque gyro output torque.

The invention relates to a decoupling and disturbance-attenuation method for a rotor system of a magnetic suspension control moment gyroscope (CMG). The decoupling and disturbance-attenuation method includes the steps that a kinetic equation of the rotor system of the magnetic suspension CMG in a magnetic bearing coordinate system is built according to the Newton second law and a gyroscope technical equation, a radial four-way decoupling model is obtained based on the active-disturbance-rejection decoupling control principle, then active-disturbance-rejection controllers of channels are designed, and therefore radial four-way decoupling and disturbance attenuation of the rotor system are achieved. Compared with traditional disperse PID and intersection feedback decoupling control, the decoupling and disturbance-attenuation method has the advantages that the decoupling control accuracy is improved, and the robustness of the system for external disturbance and parameter changes is improved. The decoupling and disturbance-attenuation method belongs to the technical field of spaceflight control and can be applied to high-accuracy and high-robustness control over the magnetic suspension CMG.

A control moment gyroscope system for delivering a target torque to a spacecraft including a rotor assembly having a rotor and a motor to spin the rotor about a rotor axis. A gimbal assembly has a gimbal for supporting the rotor assembly and a gimbal torque motor to rotate the gimbal about a gimbal axis, which is normal to the rotor axis, to generate an output torque. A control system has a sensor for determining the output torque and a processor in communication with the rotor assembly, the gimbal assembly and the sensor. The processor requests the target torque and establishes a feedback control loop to generate a torque error signal based on the output torque for bringing the output torque within a predetermined range of the target torque.

A control moment gyroscope (CMG) is provided for deployment on a spacecraft. The CMG includes an inner gimbal assembly (IGA), which, in turn, includes an IGA housing, a rotor rotatably coupled to the IGA housing, and a spin motor coupled to the IGA housing and configured to rotate the rotor about a spin axis. The CMG further comprises a stator assembly, which includes: (i) a stator assembly housing rotatably coupled to the IGA housing, and (ii) a torque module assembly coupled to the stator assembly housing and configured to rotate the IGA about a gimbal axis. A gimbal bearing is disposed between the IGA housing and the stator assembly housing. The gimbal bearing resides between the spin axis and the torque module assembly such that the distance between the gimbal bearing and the spin axis is less than the distance between the gimbal bearing and the torque module assembly.

A gyro wave-activated power generator utilizing wave energy which has, as a component, a control moment gyro, and which comprises a floating body (1), a control moment gyro (5) supported by the floating body, and a power generator (10) connected to the gimbal shaft of the gyro through a gearing to increase speed(9), and is characterized by the gimbal (6) of the gyro being rotated by the movement of the floating body by wave energy to drive the power generator (10) for power generation.

A self-contained momentum control system (MCS) for a spacecraft is provided for small satellites. The MCS features a miniaturized gyroscopic rotor with a rotational speed in excess of 20,000 RPM. The MCS includes at least three control moment gyroscopic mechanical assemblies (CMAs) rigidly mounted within a single enclosure, where each CMA mounted in an orientation whereby the longitudinal axis of each CMA is either orthogonal to every other CMA or is parallel to another CMA but in the opposite orientation. In order to further reduce the size of the MCS, an electronics package that is configured to interface command and control signals with and to provide power to the CMAs is included within the MCS enclosure. A plurality of shock isolation devices are used to secure each of the CMAs to the enclosure in order to reduce the launch load upon the CMAs thereby allowing the use of smaller rotor spin bearings. The MCS enclosure surrounding the CMAs and support structure is hermetically sealed.