261 results about "Satellite attitude control" patented technology

The invention discloses a satellite posture all-round controlling method based on a magnetic moment device and a flywheel, relating to an all-round posture controlling method for completing a satellite orbit-injection phase by using the magnetic moment device and the flywheel. The invention solves the problems of low reliability and short service life of the traditional satellite posture all-round controlling technology. The satellite posture all-round controlling method comprises the following steps of: 1, setting controller parameters according to the requirement of a control system; 2, measuring a geomagnetic field intensity vector Bb, a satellite angular velocity vector Wb and a solar azimuth, and sending the measured data to a satellite controller; 3, calculating an expected control moment vector Tm and a control magnetic moment vector Mm, and sending the control magnetic moment vector Mm to the magnetic moment device; 4, acquiring an effective solar azimuth vector Alfa; 5, calculating a control input moment vector Tw and sending to the flywheel; and 6, jointly completing the satellite posture all-round control by the magnetic moment device according to the control magnetic moment vector Mm and the flywheel according to the control input moment vector Tw. The invention is suitable for the field of satellite posture control.

A ground simulation-test system of satellite attitude based on satellite network comprises simulation test unit of parts grade, digital simulation test unit of system grade and semiphysical simulation test unit of system grade. It is featured as using simulation test unit of parts grade and digital simulation test unit of system grade to provide physical data basis for semiphysical simulation test unit of system grade. The ground simulation testing-method of satellite attitude based on satellite network is also disclosed.

Simulation test equipment for validating the reliability of satellite attitude control and a test method thereof belong to the field of satellite attitude control. The test equipment comprises a ground simulation support module, a fault injection module and a satellite-born module, wherein, the ground simulation support module includes a monitoring terminal, a satellite model, a signal exciting source and a moment reverse-solution unit; the fault injection module includes an analog-to-digital converter, a signal transfer circuit, a fault injection processor, a man-machine interface unit and a digital-to-analog converter; and the satellite-born module includes a sensor, a satellite-born controller and an actuator. The test method comprises the following steps of: fault model setting; satellite initialization; orbit and attitude information renewal; the satellite-born controller collects signals of the sensor and outputs actuator command signals to the fault injection module; the actuator receives actuator pseudo command signals and outputs feedback signals to the fault injection module; the moment reverse-solution unit collects the pseudo feedback signals output by the fault injection module and outputs moment control signals to the satellite model. The equipment and the method of the invention have high feasibility and high performance-price ratio.

The invention relates to an independent orbital transfer of a satellite, comprising (1) star capture step, in which before the satellite establishes orbit-controlled ignition attitude, inertia attitude of the satellite is pre-estimated and satellite attitude is controlled; zero offset of gyro drift and an accelerometer is marked; (2) inertia attitude regulation step, used for establishing the orbit-controlled ignition attitude of the satellite, realizing maneuver of the satellite attitude, pre-estimating the inertia attitude of the satellite and a jet control is used for controlling the satellite attitude; (3) star orientation step, in which after the satellite establishes the orbit-controlled ignition attitude, a stable state control is carried out and a star sensor is used for filtering and correcting the satellite attitude; and (4) orbit-controlled orientation step, in which starting and shutdown control is carried out for a rail-controlled engine; the ignition attitude of the satellite is determined and attitude stable control is carried out in the period of the ignition. The invention solves the problem that spacecraft such as a deep space exploitation satellite transfers orbit, and ensures that the orbital transfer can be completed accurately and reliably.

The invention relates to a monoatomic dispersed platinum group metal based catalyst with a higher loading capacity and an application of the catalyst. Specifically, the invention relates to a bicomponent catalyst of titanium oxide and one selected from metals rhodium (Rh), ruthenium (Ru), platinum (Pt), iridium (Ir) and palladium (Pd) as well as preparation and an application of the catalyst. Themetal is highly dispersed in the form of a single atom on the carrier titanium oxide, and the metal content is up to 0.4-2%. The catalyst provided by the invention is suitable for the CO water gas shift reaction, and can improve the CO conversion rate and inhibit a methanation side reaction; and meanwhile, the catalyst also has very good low-temperature activity and use stability on catalytic decomposition of the liquid single-component propellant level anhydrous hydrazine used in a satellite attitude control engine in the aerospace industry.

The invention discloses a system of satellite attitude control integrated simulation and an implement method. The system comprises a simulation computer and an interface expansion board; simulation application under five working modes of complete mathematical simulation, component-level semi-physical simulation, component-level semi-physical simulation, whole-spacecraft semi-physical simulation, satellite-ground united simulation and in orbit attitude visualization can be realized through the combining connection of the simulation system and tested hardware; the system integrates a number of functions in one set to reduce the repeated simulation software development work to achieve the universality and generalization of the simulation system; the visualization and operability of a simulation process is enhanced; the whole flow comparison simulation analysis from scheme design to in orbit test of the satellite attitude control system is supported.

The invention discloses small satellite attitude control ground simulating device and method. The small satellite attitude control ground simulating device mainly comprises a simulation support module and a satellite tested module, wherein the simulation support module comprises a simulation computer, an embedded processor and a field programmable gate array; the satellite tested module comprises an attitude control processor, measuring parts and executing parts; the simulation computer, the embedded processor and the field programmable gate array are sequentially connected; and the attitude control processor, all measuring parts and all executing parts are respectively connected with the field programmable gate array. The invention is suitable for the simulation requirements of satellites in various models, simultaneously meets various ways of semi-physical simulation requirement, has simpler development and maintenance of simulation computer software, more flexible and convenient operation of the simulation process and the analysis of a simulation result, simple realization and low cost and greatly shortens the development time and development cost of the satellite attitude control ground simulating device.

The invention relates to an integrated executive mechanism of satellite attitude control and thermal control, which comprises a liquid accumulator, a circulating pump assembly and a first flow distribution valve which are connected in series to form a working medium circulating loop, heat exchange devices which are connected in parallel, a heat exchange loop formed by a first bypass branch, and a radiator/torquer integrated device, wherein a liquid working medium flows in the circulating loop; the circulating pump assembly is used for driving the liquid working medium to flow in the circulating loop at a certain flow velocity and flux; the radiator/torquer integrated device comprises at least two different pipelines which are connected in parallel and a device used for controlling the distribution of the liquid working medium in the different pipelines; and the circulating pump assembly is sequentially connected with the radiator/torquer integrated device. The integrated executive mechanism can realize the heat-dissipating function of a high heat-flux density device of a microsatellite and provides single-axis or multi-axis control moment for attitude control, integrates functions of attitude control and thermal control executive mechanisms, improves the function density of the microsatellite, and has simple structure and low cost.

The invention discloses a light high-accuracy digital control system for a miniature flywheel. The system belongs to a kind of posture control execution mechanism applied to miniature satellites which are lighter than 100Kg, and mainly comprises a field programmable gate array (FPGA) module, a three-phase full-bridge driving device, an RS-232 communication circuit, a Buck converter, an analogue-to-digital (AD) sampling circuit and a signal conditioning circuit, a Hall signal filter and the like, wherein the FPGA module comprises an instruction parser, a moment-current converter, a single-neuron adaptive proportion integration differentiation (PID) controller, a pulse width modulator (PWM), a commutation logic signal generator and the like. By the invention, a nonlinear transformation online correction method is adopted; the digital control system for the miniature flywheel is formed by taking the FPGA module as a control core based on a single-neuron adaptive PID control algorithm; and the high-accuracy moment output of the posture control execution mechanism for the miniature satellite is realized by adopting a moment mode. The invention has the characteristics that: the algorithm is simple and easy to implement; the system is quick in response, low in power consumption, high in reliability, flexible in design and particularly suitable for the miniature satellite; the power density of the miniature flywheel is increased; the mass and volume of the miniature flywheel are decreased; and cost is decreased.

The invention discloses a fault-tolerant control method for a discrete integral sliding mode of a satellite control system. A discrete integral sliding mode controller is designed based on the discrete linear system mode of a satellite, the virtual control law of the controller obtains a health factor W of an actuator through fault diagnosis and parameter identification; adopting a direct distribution method based on the reachable set of a moment to obtain the control moment uk of each actuator according to the controlling quantity vk and the health factor W of the actuator; and arranging the surfaces of reachable sets according to the sequence of angles between a surface normal vector and a target vector from small to large through the direct distribution method based on the reachable set of the moment. According to the method, the satellite attitude control problems of coexisting inaccurate health factors, multiple faults and multiple uncertainties can be effectively processed on line.

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.

The invention discloses an optical remote sensing satellite point target observation task dynamic imaging posture modeling method, which comprises the steps of obtaining the point target observation task center point imagining time of a satellite, camera station point position, posture pointing parameters and satellite orbit coordinate parameter sequences; calculating the camera pupil entrance radiant brightness and the camera integral time of the satellite on the observation task center point imagining time; on the basis of the camera integral time, processing the relevant parameters to obtain the camera station point scanning ground velocity of the satellite on the observation task center point imagining time, and using the scanning ground velocity as the constant scanning ground velocity; on the basis of the constant scanning ground velocity, solving the observation task path camera station point geographic position parameter sequence and the corresponding imaging time; calculating the posture rolling angle parameter, the posture pitch angle parameter, the posture yaw angle parameter and the corresponding angular speed parameter of each imaging time of the observation task, and completing the dynamic imaging posture model building on the basis. The effective combination of task planning and satellite posture control can be realized to a certain degree.

The invention discloses a low-orbit constellation GEO frequency interference avoidance method based on multiple coverage. In view of the problem that in a current low-orbit communication constellationGEO interference avoidance method, the attitude of a low-orbit satellite is periodically biased, other loads of the low-orbit satellite cannot work easily, and low-orbit satellite attitude control engineering cannot be achieved easily, GEO frequency interference avoidance is achieved through opening and closing of different low-orbit satellite beams in a low-orbit constellation, and continuity oflow-orbit constellation services is guaranteed. And meanwhile, in the whole working process of the low-orbit satellite, the satellite attitude can always keep a normal three-axis stable attitude to the ground, so that the work of other loads of the low-orbit satellite is facilitated, and the engineering realization of a satellite attitude control system is also facilitated.

The invention provides an attitude-orbit coupling control method for tracking the characteristic parts of non-cooperative targets. The steps include: establishing a relative attitude motion dynamics model based on relative quaternions and a relative orbital dynamics based on tracking the motion trajectory of the characteristic parts of non-cooperative targets Model; conduct attitude-orbit coupling analysis and design and track the trajectory of non-cooperative target feature parts; design a relative quaternion feedback attitude control law that does not depend on the dynamic information of the target star and an optimal real-time closed-loop feedback control law with attitude feedback information. The attitude-orbit coupling control method does not need the dynamic information of the target spacecraft, which is in line with the characteristics of non-cooperative targets; it is suitable for the target spacecraft in a three-axis stable state or a slow rotation state; it can enable the tracking spacecraft to track the characteristic parts of the target spacecraft Capability, good control accuracy, robustness and safety, and easy engineering implementation, so as to carry out reliable and effective capture tasks.

Provided is a flexible satellite attitude control method based on three-stage path planning, relating to the technical field of satellite attitude control. A flexible satellites with pyramidal CMG group as the actuator is selected as the study object, a satellite attitude control method combing the three-stage attitude path planning and rolling optimization tracking is proposed, a nonlinear state-space equation including satellite attitude dynamics, kinematics and flexible attachment vibration is established for the accurate prediction of satellite attitude future information. Based on the three-stage path planning strategy, the flexible satellite attitude maneuvering time is optimally designed, the vibration of the flexible attachments can be avoided when the satellite attitude maneuvering target is completed. By means of the nonlinear model prediction method, the attitude angular velocity curve and attitude angle position curve obtained by planning are taken as the desired tracking targets to design the rolling optimization tracking control law of the frame angular velocity of the pyramidal CMG group, therefore, the attitude wide angle rapid maneuvering control for suppressing the vibration of the flexible attachments can be realized.

The invention discloses an attitude control information joint estimation method under the satellite attitude angle deviation, and relates to the technical field of satellite attitude determination andcontrol. The problem that in the satellite attitude maneuvering process, attitude control information under the attitude small angle deviation is difficult to estimate precisely, so that the controlprecision is low is solved. The method comprises the steps that on the basis of measurement information of a spaceborne sensor and an actuating mechanism, under the attitude small angle deviation, joint precision estimation of variables such as the satellite attitude error quaternion, satellite attitude error angular velocity, reaction flywheel angular momentum, reaction flywheel friction moment,external disturbance moment and fiber-optic gyroscope constant drift is achieved in a loop iteration mode by means of satellite attitude dynamics, kinematics and a reaction flywheel dynamical model, and precise attitude control information is provided for high-precision satellite attitude control. The method is achieved through a software algorithm, and the method is simple and easy to implement.

The invention discloses a satellite attitude control method for space cooperation target tracking imaging. Firstly the conditions required to be met by an imaging time window are determined, and the expected attitude quaternion and angular speed are calculated through the given satellite orbit state and target orbit information so that the error amount between the attitude tracking error quaternion and the actual angular speed and the expected angular speed is calculated; then an attitude kinetic model is established; and finally control law parameters are selected, and the flywheel moment control amount is calculated. The method is based on the attitude quaternion representation so that the possible singularity problem caused by using euler angle in large-angle attitude motion can be avoided. Besides, the method is simple in algorithm, low in calculation amount and simple in project realization and has high pointing control accuracy and great pointing stability.

The invention discloses a flywheel motor adopting a no-cross one-range winding, relating to the technical field of the motor and solving the problems that the existing flywheel motor has complex structure, large power consumption and poor moment stability. A rotor in the invention is a wheel body of a flywheel; an outer rotor iron core and an inner rotor iron core of the rotor are fixedly connected on a torque output shaft; a no-iron core frame type stator is fixed between the inner rotor iron core and the outer rotor iron core; permanent magnets are fixed on the inner rotor iron core or the outer rotor iron core and the side wall corresponding to the no-iron core frame type stator; a winding of the no-iron core frame type stator consists of a plurality of groups of planar coils which areevenly distributed and fixed on the side wall of the frame; two adjacent groups of planar coils have an electric angle of 240 degrees, and have no cross; three position sensors are respectively fixedon the frames at the central positions of three adjacent groups of planar coils; and the ratio of the number Z of the planar coils and the number 2p of poles of the permanent magnets is 3/4. The flywheel motor is applicable to the technical fields of posture control of the satellites and control moment gyros and the like.

The invention discloses a device for measuring micro thrust of an engine, and relates to force measurement equipment. The device comprises a pedestal, a vertical beam with a knife edge, a propellant storage box, a cross beam, a balance weight, the engine, an electric control box, a force transducer, a signal processing unit and a calculator, and is characterized in that: the dead weight of the engine is balanced by using a knife edge structure and the balance weight; and the influence of a propellant pipeline and a control line on measurement accuracy is eliminated through relation between force and force arm according to a mechanical principle. When the range of the force to be measured is at the scale of hundreds of millinewtons, zero drift of the force is less than +/-1 percent F.S. every hour, the stable state measurement error is less than +/-1 percent of a full range, and high repeatability is guaranteed. The device can meet the accuracy requirement of small satellite attitude control engine thrust measurement and has broad application prospect in micro thrust measurement of the engine.