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 provides a single frame control moment gyro control method based on moment output capability optimization, relating to the prompt satellite attitude control field and solving the problems that the existing single frame control moment gyro control rule can not effectively avoid strange and big moment error can be caused when escaping strange. The method includes the concrete steps that: a satellite attitude control system obtains the current expected moment Tc output by CMG system according to the current attitude information; capability optimized frame angular velocity for the CMG system to output the expected moment Tc is calculated; and a frame rotating speed instruction of CMG system is obtained according to the expected moment Tc and the optimal frame angular velocity, so as to realize satellite attitude control. The method of the invention can effectively avoid strange problem, and meanwhile output moment error can be reduced, thus improving satellite attitude control accuracy.

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 relates to a satellite active fault-tolerant control method based on an observer and online control allocation. The active fault-tolerant control method based on the iterative learning observer is provided for the problems of actuator faults, limited input saturation and external disturbance in the satellite attitude control process. The method comprises the following steps that a satellite attitude control system model is established by considering the actuator faults and the external disturbance subjected by a satellite; the iterative learning observer is designed, and invalidation factors of an actuator are estimated; a virtual feedback controller is constructed based on fault estimation information, online control allocation is conducted simultaneously, and therefore precise estimation of the actuator faults and controller reconstruction can be conducted through the satellite. According to the method, the stability of the attitude control system is guaranteed when an on-orbit working satellite is subjected to actuator faults, and the method has the advantages that the control accuracy is high, and the fault-tolerant capability and robustness to the external disturbance are high.

The invention discloses a satellite attitude control semiphysical simulation system based on a multi-target machine. The satellite attitude control semiphysical simulation system comprises a real-time positioning sub system, an attitude measurement sub system, an attitude control comprehensive processing system, an execution mechanism, a signal conditioning box, a target machine, a working computer and a display system, wherein the real-time positioning sub system acquires real-time position information of a satellite; aiming at different working modes of the satellite during orbit flying, the attitude measurement sub system acquires real-time attitude information of the satellite; the attitude control comprehensive processing system manages all attitude measurement components and the execution mechanism; the target machine performs attitude calculation and torque calculation according to acquired real-time position information and the acquired attitude information and sends control information to the execution mechanism; and the execution mechanism sends actual execution information back to the target machine through the signal conditioning box, acquires the real-time attitude information of the satellite through dynamics and sends the attitude information to the attitude measurement components. The satellite attitude control semiphysical simulation system has high simulation instantaneity, high fault-tolerant capability and high integration degree and has great instructional significance for design and development of satellite attitude control systems.

The invention relates to a satellite fault diagnosis and fault-tolerant control method based on a T-S fuzzy model and a learning observer. The satellite fault diagnosis and fault-tolerant control method based on the T-S fuzzy model and the learning observer is used for solving the problems that a conventional fault diagnosis method cannot be used for effectively treating the influence caused by space disturbance torque and ensuring the robustness of a fault diagnosis method, and a conventional fault-tolerant control method has a poor fault-tolerant property. The satellite fault diagnosis and fault-tolerant control method based on the T-S fuzzy model and the learning observer comprises the following steps: step 1. establishing a mathematical model of a non-linear satellite attitude control system; step 2. establishing the T-S fuzzy model of the satellite attitude control system by using results obtained in the step 1; step 3. designing a T-S fuzzy learning observer by using the results obtained in the step 2 so as to realize the robust fault detection, isolation and fault reconstruction of a satellite attitude angular velocity estimation and executing mechanism; 4. designing a state feedback fault-tolerant controller by using the results obtained in the step 3 so as to ensure that a closed loop of a satellite attitude control T-S fuzzy system is stable. The satellite fault diagnosis and fault-tolerant control method based on the T-S fuzzy model and learning observer can be applied to an aerospace field.

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 processing method for upwards injecting a multi-parameter instruction by an agile satellite. According to the method, instructions generated when an on-satellite attitude control subsystem, a data transmission subsystem, a camera subsystem, a satellite affairs subsystem and other subsystems of the agile satellite execute mobile imaging tasks are formed into multiple instruction templates depending on typical working modes of various on-satellite subsystems and are stored in an on-satellite computer; the instruction templates corresponding to the types of the mobile imaging tasks are selected by a ground user, and a task number, an instruction template number and instruction parameters in the instruction templates are transmitted onto the satellite in accordance with a determined format of the satellite and the earth, and then an instruction set can be automatically generated by the on-satellite computer. Meanwhile, the on-satellite computer can modify contents in the instruction templates, add the instruction templates, and delete the instruction templates and the like at any time in accordance with an instruction on the ground. According to the method, not only can the complexity degree of instruction arrangement of the ground user be simplified, the upwards injecting data volume is greatly reduced, but also the reliability of instruction arrangement can be improved; particularly, the method provides a technique guarantee for task management and applications of the agile satellite.

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.

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 a ground simulation test system for satellite attitude control, comprising: control equipment, simulation equipment and acquisition simulation equipment; the control equipment, simulation equipment and acquisition simulation equipment are all equipped with optical fiber reflection memory cards; Optical fibers are interconnected in pairs to form a closed-loop data flow; control equipment, simulation equipment and acquisition simulation equipment realize multi-data interaction in the closed-loop data flow. The control device is used to design the simulation model according to the satellite to be simulated and tested; to remotely control the simulation device and to collect the simulation device; to save and monitor the interactive data in the closed-loop data stream. The simulation device is used to run the simulation model; the acquisition simulation device is used to simulate the input / output of the signal of the satellite to be simulated and tested or directly collect the signal of the satellite to be simulated and tested when the control device cannot design the simulation model. The invention has universality, super real-time performance, and wide application range, shortens the development cycle, and reduces the research and development cost.

The invention discloses a disturbance torque identification method suitable for satellite attitude control, and belongs to the technical field of spacecraft attitude control. The invention aims at solving the problem that a conventional filter algorithm cannot achieve the separated measurement of noise and disturbance torque in a high-precision attitude control task. The method comprises the following steps: 1, building a small quantized satellite attitude control mode with an unknown disturbance torque according to a to-be-identified satellite attitude control system; 2, employing an obtained disturbance torque, comprising noise, estimated by an unknown input observer according to the unknown input observer obtained according to the satellite attitude control system built at step 1; 3, employing the DFT (Discrete Fourier Transformation) and the IDFT (Inverse Discrete Fourier Transformation) to process the obtained disturbance torque, comprising noise, estimated at step 2 in an off-line manner, and obtaining the estimation results of the disturbance torque after noise removal; 4, carrying out the Fourier series fitting of estimation results, obtained at step 3 to obtain a mathematical expression of the disturbance torque. The method is used for the attitude control of satellites in orbit.

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 thermal control method of the non-attitude satellite, paste or spray the high-emissivity thermal control coating on the equipment in the star, fill the thermal conductive filler between the equipment installed on the partition and the installation surface, and spray the high-emissivity thermal control coating on the inner surface of the rest of the structure star Floor. The inner surface of the cabin board is covered with multi-layer heat insulation components to reduce the impact of drastic changes in heat flow outside the outer orbit of the star. Except for the installation surface of the instrument and equipment, the structural star module is pasted with third-order gallium arsenide solar cells. From the perspective of top-level system design, the present invention is based on the whole star isothermal design idea, fully utilizes the whole star structure heat storage, and utilizes solar cells as energy supply equipment and temperature control means at the same time, and adopts three-dimensional direction 3D isothermal heat pipe combination and integration The 3D isothermal device composed of a metal frame solves the problem of ensuring the temperature level of the entire satellite when there is no attitude control.

The invention discloses an optical imaging minisatellite attitude control system and an in-orbit working mode switching method, belongs to the field of satellite attitude control and aims at solving the problem that existing optical imaging satellites cannot be simply and effectively switched in the orbit. The optical imaging minisatellite attitude control system comprises an attitude measurement sensor, an actuating mechanism and an attitude controller, wherein the attitude measurement sensor comprises a sun sensor, a star sensor and a gyroscope; the actuating mechanism comprises a counteraction flywheel and a magnetorquer. According to an in-orbit working mode switching method of the optical imaging minisatellite attitude control system, the optical imaging minisatellite attitude control system has six working modes, namely a rate damping mode, a sun capture mode, a sun-orientated three-axis stabilization mode, an earth-orientated three-axis stabilization mode, a data transmission mode and a safety mode. The optical imaging minisatellite attitude control system and the in-orbit working mode switching method are suitable for all the optical imaging minisatellites.

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 configuration and control method for a satellite attitude control and large angular momentum compensation multiplexing flywheel. The configuration and control method comprises the following steps that a satellite pitch axis Y-axis is provided with a 25Nms angular momentum flywheel in a front mode, the positive angular momentum direction and the satellite body coordinate system+Y-axis direction are opposite, and offset momentum control is adopted in the satellite; the satellite yaw axis Z-axis is obliquely provided with two 50Nms angular momentum flywheels, the positive angular momentum direction is in the XOZ plane of the satellite body coordinate system and is separately biased towards the rolling axis+X-axis and rolling axis-X axis direction, and the angle between the positive angular momentum direction and the yaw axis Z-axis is 10 degrees; a satellite rolling axis X-axis is provided with a 25Nms momentum wheel in a front mode, the positive angular momentumdirection and satellite body coordinate system+X-axis direction are consistent; the flywheel rotation speed command is obtained by calculation, the control voltage command of the flywheels is obtained, the flywheels are driven, and the control angular momentum is output. The invention simultaneously completes the attitude control and the control output of the compensated angular momentum with a small number of flywheels.

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.