427 results about "Kinetic equations" patented technology

The invention provides a vehicle mass identification method and a device. The vehicle mass identification method comprises the following steps of: acquiring wheel rotating speed and driving torque of a vehicle during running; acquiring a longitudinal acceleration value of the running vehicle; constructing a longitudinal kinetic equation of the vehicle; and on the basis of the longitudinal kinetic equation, setting a roughly estimated vehicle mass by taking the wheel rotating speed, the driving torque and the longitudinal acceleration value of the running vehicle as measurement input parameters and taking a sum of rolling resistance and air resistance as a variable, and identifying the mass of the vehicle through a recursive least square method with a plurality of forgetting factors.

The invention relates to a method of taking over and controlling attitude after a space tethered robot captures a flexible target satellite. Taking regard of nondeterminacy of inertia, the coupling effect, external interference and other factors, the method of taking over and controlling attitude after space tethered robot captures flexible target satellite establishes a complex attitude and orbit coupling kinetic equation, designs inner and outer ring terminal sliding mode controllers, and gives consideration on the saturation characteristics of a thruster and a tether so as to stably control the attitude and angular velocity of the complex. The method of taking over and controlling attitude after a space tethered robot captures a flexible target satellite includes the following steps: establishing an attitude kinetic equation of a complex after a space tethered robot captures a target satellite; designing inner and outer ring terminal sliding mode controllers and the corresponding adaptive law; taking the inner ring control law and the outer ring control law as the input of a control system to take over and control the attitude after the flexible target satellite is captured; and proving Lyapunov stability. The method of taking over and controlling attitude after a space tethered robot captures a flexible target satellite can solve the problem that the flexible complex parameters are not determined and the self-supplied thruster of the space tethered robot is saturated.

The invention discloses an iterative learning trajectory tracking control and robust optimization method for a two-dimensional motion mobile robot. The method includes the steps that firstly, a kinetic equation of a two-dimensional motion mobile robot discrete non-linear motion system model is established; a discrete non-linear state space expression is established; a P type open-closed loop iterative learning controller based on the iterative learning control technology is established; then the robust convergence of the established discrete non-linear control system is theoretically analyzed; then parameter item splitting is conducted on control gains of the P type controller, meanwhile, a quadratic performance index function based on controller parameters is designed, and the purpose is to optimize the control parameters; finally, monotone convergence characteristics of output errors and parameter selection conditions generated when a control algorithm acts on a controlled system are analyzed and optimized, and the two-dimensional motion mobile robot can rapidly track an expected motion trajectory at high precision. The method has the advantages that the robust optimization iterative learning controller is suitable for tracking control in an ideal state and suitable for trajectory tracking tasks under the condition that interference exists outside. A designed iterative algorithm is simple and efficient, introduction of a large number of additional parameter variables is not needed, and engineering realization is easy.

The invention provides an impedance control method for a flexibility joint mechanical arm based on a connection and damping configuration, and belongs to the technical field of robot controlling. The problem in a traditional mechanical arm control method that during the flexibility joint mechanical arm control process, due to the fact that the residual oscillation is large, the aim of stable control cannot be achieved is solved. The technical points are that according to a CAD three-dimensional model, a kinetic parameter and a kinematic parameter of the flexibility joint mechanical arm are obtained; according to the parameter identification, a key parameter of a flexibility joint is obtained; a kinetic equation of the flexibility joint mechanical arm is established; a gravity compensation algorithm and an external force compensation algorithm based on the motor position are established; a gravity compensation value and an external force compensation value of the mechanical arm are calculated and obtained; a minimum Hamiltonian function value on the expectation balance position of motor position information is calculated and obtained; an expectation connection matrix and a damping matrix in the impedance control are calculated; an impedance control rule of the flexibility joint based on the connection matrix and the damping matrix are obtained. The impedance control method for the flexibility joint mechanical arm based on the connection and the damping configuration can be used for controlling service robots, medical robots and space robots.

The invention discloses a heavy vehicle weight real-time identification method based on CAN information and a function principle. According to the method, when a vehicle runs, vehicle wheel speed, engine rotating speed, engine practical torque percentage, a gear position, accelerating pedal opening degree, a clutch pedal signal and a braking pedal signal are acquired through a vehicle CAN bus; engine torque, accelerated speed and a rotating weight conversion coefficient are obtained by computing; a vehicle longitudinal kinetic equation based on the function principle is established, on the basis of the kinetic equation, the work done by equivalent outer force generated by driving force and air resistance is regarded as system input quantity, the work done by the equivalent accelerated speed generated by the sum of rolling resistance, slope resistance and acceleration resistance is regarded as an observable data vector, and heavy vehicle weight is estimated in real time through a recursive least-squares method with which a forgetting factor is selected according to the feature of the vehicle weight.

The invention relates to an industrial robot positioning precision calibration method used for improving robot absolute positioning precision. The method comprises the steps that: a robot kinetics model is established; a rotary joint equation is established by using a circumferential method; joint torsion angle parameter values are calculated; the actual pose of a robot end is measured accurately by utilizing a laser tracker; D-H algorithm inversion kinetics equation is improved to obtain geometrical structural parameters; calibration of joint distance parameter values and calibration of joint offset are realized, and thereby first calibration is completed. Error correction is carried out on D-H parameters based on first calibration results, and if positioning precision can not meet requirements, second calibration can be carried out by substituting parameter correction values, instead of theoretical parameter values, into the kinetics equation. Absolute positioning precision of the robot is improved with the calibration method, which is proved in experiments. According to the invention, the calibration method is advantaged in that the method is simple and practical and introduces small external errors.

The invention discloses a safety analysis calculating device for transient nuclear heat coupling of a supercritical water reactor, and belongs to the field of nuclear power plant design and reactor safety analysis. The safety analysis calculating device for transient nuclear heat coupling of the supercritical water reactor is applicable to the supercritical water reactor. According to the safety analysis calculating device, a transient physical analysis program is built by using the space-time disperse solution of a double-group space-time kinetic equation, a transient heat engineering analysis program is built through calculation of flow distribution among channels and the solution of flowing heat exchange inside the channels, on this basis, a physical calculation program is embedded into a transient system analysis program to build a supercritical transient coupling analysis program so as to analyze safety of a system. Particularly, due to the fact that a space-time kinetics quasi-static solution model built through the double-group neutron space-time kinetic equation is combined with a multi-channel analysis method, compared with a sub-channel model, the calculation time is shortened on the basis that precision is ensured, then economy of the supercritical water reactor design is improved, and the safety analysis calculating device accords with the development trend of safety analysis.

The invention relates to a motion coordination control method for a space dual-manipulator system, and belongs to the technical field of space manipulators. In order to overcome the defects that no space manipulator system specific to dual manipulators exists at present, the coordinated relation between a manipulator and a satellite body is not considered in an existing space manipulator system and tracking errors are high, the motion coordination control method for the space dual-manipulator system is provided. The motion coordination control method for the space dual-manipulator system includes the steps that a kinematical equation and kinetic equation of the space dual-manipulator system are established; the space dual-manipulator system is subjected to trajectory planning according to the initial postures and tail-end postures of the manipulators; and trajectories of the space dual-manipulator system are subjected to tracking control through a PD controller. The motion coordination control method for the space dual-manipulator system is suitable for a space manipulator control system.

Provided is a fault tolerant control method for a spacecraft equipped with a flywheel based on robust adaptation. The method comprises five steps of: 1, establishing a spacecraft kinematical equation; 2, establishing a spacecraft kinetic equation; 3, performing desired angular velocity locus design; 4, performing controller output torque design; and 5, performing numerical simulation. Under the circumstance that a FDD device is not used, the method online estimates a triaxial torque fault factor and designs a robust adaptive fault tolerant controller. The research aims at enriching spacecraft PFTCS methods and providing technical support for spacecraft attitude control in the future.

The invention discloses a tri-axial platform servo motor control method based on the combination of a BP neural network and an active disturbance rejection controller. The tri-axial platform servo motor control method includes the following steps that 1, the kinetic equation of a tri-axial platform arbitrary-frame servo system containing a friction model is built; 2,according to the kinetic equation in the formula (I),the active disturbance rejection controller is designed; 3,a BP neural network parameter on-line setting module is designed and combined with the active disturbance rejection controller, and active disturbance rejection controller on-line parameter setting is achieved. By the adoption of the tri-axial platform servo motor control method based on the combination of the BP neural network and the active disturbance rejection controller, a flat top phenomenon and a dead zone phenomenon of a servo system can be basically eliminated, wherein the flat top phenomenon occurs during position signal tracking, and the dead zone phenomenon occurs during speed signal tracking; therefore, the position tracking precision and the speed tracking precision of the servo system can be improved.

The invention provides a spacecraft attitude tracking control method based on discontinuous adaptive control, belongs to the technical field of spacecraft attitude tracking control, and solves the problem that a spacecraft attitude tracking control system is poor in robustness so that a spacecraft attitude tracking control effect is poor in the case of modeling uncertainty, external disturbances and input saturation effects. The spacecraft attitude tracking control method comprises the following specific steps of step 1, establishing an earth centered inertial oIxIyIzI, a spacecraft body coordinate system oBxByBzB and an expected reference coordinate system oRxRyRzR; step 2, according to the coordinate systems established in step 1, obtaining the spacecraft attitude kinematics and a kinetic equation described by using an attitude quaternion, and a spacecraft error attitude kinematics equation and the kinetic equation, namely the attitude tracking control system; and step 3, based on the step 2, based on a sliding mode surface of an integral terminal, designing an attitude tracking controller considering an unknown external disturbance torque and the rotational inertia uncertainty.The spacecraft attitude tracking control method based on the discontinuous adaptive control provided by the invention can be applied to spacecraft attitude tracking control.

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.

The invention discloses a rail vertical irregularity estimation method and system based on extended Kalman filtering. The system is characterized in that vibration acceleration sensors and gyroscopes are installed on a train body and a frame, the output ends of the vibration acceleration sensors, the gyroscopes and train speed sensors are connected to a central processing unit, and spatial sampling of train body vertical vibration, train body nodding angular speed, frame vertical vibration, frame nodding angular speed and the like is realized through train speed signals. A vehicle rail coupling system state equation is established according to a kinetic equation for a vehicle rail coupling model; a sensor observation equation is established according to signals acquired by all sensors; a filtering iterative equation is configured to obtain time updating equations and measurement updating equations of state estimation and estimation error covariance; an optimal state at each moment is obtained according to extended Kalman filtering, and rail irregularity estimation is finally obtained after inversion by using the time updating equation of the state estimation. The rail vertical irregularity estimation method and system based on extended Kalman filtering have the advantages of low cost, good engineering feasibility, simplicity in operation and online real-time detection.

The invention discloses a several-for-one collaborative guidance method based on target probability distribution information, and belongs to the field of aircraft guidance. The problems that due to the fact that an existing guidance method is adopted in a near space hypersonic target, the target detection precision is low, the trajectory prediction accuracy is low, terminal guidance time is too short, so that the requirement for the overload response speed of an interceptor is too high are solved. The method is characterized by comprising the steps that relative motion measuring information between the interceptor and the target is acquired; Kalman filtering is used for processing the relative motion measuring information by considering the state value of the interceptor, the estimated value of the target state at the current moment is obtained, and a target terminal moment location probability density function is obtained by considering the kinetic equation of target motion; the integrals of the target terminal moment location probability density function in a plurality of interceptor reachable sets are constructed as a collaborative guidance cost function; several-for-one collaborative guidance optimizes problem solving is carried out. The method is applied to a plurality of aircrafts to intercept the hypersonic target in a collaborative mode, and has wide application prospects.

The invention discloses an efficient multi-objective optimization method for a satellite constellation and belongs to the field of constellation system of a spacecraft. The method comprises the stepsof determining an initial condition based on a Walker-delta constellation configuration, and establishing a constellation orbit kinetic equation, an earth coverage analysis module and an earth observation resolution model; optimizing an earth orbit height, an orbital inclination and an ascending node right ascension with a coverage percentage and a ground pixel resolution by use of a sequential radial basis function multi-objective optimization strategy; and constructing an objective function based on l2 weighting and an improved Pareto fitness function, replacing a high time-consuming constellation performance simulation model with an RBF (Radial Basis Function) agent model to optimize design, and updating and managing the RBF agent model through sequence sampling in an interest interval.Therefore, a Pareto non-inferior solution set satisfying engineering requirements is acquired as a satellite constellation design scheme, the coverage percentage of a constellation for a target observation region is realized as high as possible, the pixel resolution of an effective load is realized as low as possible, the calculation cost and the design cost of the satellite constellation are reduced, and the Pareto leading edge search capability is improved.

The invention provides a tracking control method of plane paths of a self-managing airship. The method comprises the following steps of: (1) giving an expected track value; giving an expected plane path; giving an expected pitch angle Theta c and an expected rolling angle Phi c; and giving an expected speed; (2) calculating navigation: calculating and removing an expected yaw angle Psi c needed by an error between an expected position and an actual position; (3) controlling and calculating a posture kinesiology: calculating and removing an expected angle speed Omega c needed by the error between an expected posture and an actual posture; (4) horizontally and longitudinally disintegrating a kinesiology: horizontally and longitudinally disintegrating a kinetic equation and an expected speedvalue; (5) longitudinally controlling and calculating the kinesiology: calculating and removing a control quantity mu 1on needed by the error between an expected longitudinal speed and an actual longitudinal speed; and (6) horizontally controlling and calculating the kinesiology: calculating and removing the control quantity mu 1at needed by the error between an expected horizontal speed and an actual horizontal speed. By the method of the invention, any parameter plane paths can be tracked for ensuring the stable property of a closed-ring system and simplifying the control calculation.

The invention discloses a combined navigation method of joint entropy extended Kalman filter based on strong tracking. The combined navigation method comprises the following steps of: establishing a state equation of an aircraft navigation system; establishing a kinetic equation of an SINS / GNSS combined navigation system; establishing a measurement equation of the SINS / GNSS combined navigation system; dispersing the kinetic equation and measurement equation of the SINS / GNSS combined navigation system; linearizing a nonlinear dispersion kinetic equation; by adopting a joint entropy extended Kalman filter algorithm based on strong tracking, outputting information of SINS / GNSS combined navigation according to the linearized kinetic equation. As fading factors in the strong tracking theory are introduced into the joint entropy extended Kalman filter algorithm, model uncertainty of a non-linear SINS / GNSS combined navigation system and non-gaussian problem of noise statistical characteristics are effectively solved, the navigation precision can be improved, and the stability of the navigation process can be improved.

The invention aims to provides a measuring device and method suitable for a non-orthogonal angle between saddle-shaped coil X and Y axes of an SERF atom magnetometer, and belongs to the technical fields of optical detection, weak magnetic detection and uniform magnetic field coils. Aimed at the problem that non-orthogonality of three-dimensional magnetic coils influences the sensitivity of the SERF atom spin magnetometer, the measuring method and device of the non-orthogonal angle between the saddle-shaped coil X and Y axes, which are based on Bloch kinetic equations, are provided. The invention fills in the blanks of the measuring method and device of the non-orthogonal angle between the saddle-shaped coil X and Y axes of the atom magnetometer, effective references are provided for the atom magnetometer to estimate and compensate a magnetic field deviation value, and the improvement of the sensitivity of the SERF atom magnetometer is ensured.

The invention discloses a multisource disturbance actuator saturation integrated spacecraft relative attitude control method. The method includes: starting from spacecraft relative attitude kinematicand kinetic equations to perform characteristic analysis and classification of multisource disturbances in a system; then, estimating disturbances with different characteristics by taking advantages of estimation performances of a disturbance observer, a radial basis function neural network and an extended state observer; next, designing an anti-saturation compensator to avoid damages caused by saturation upper limit exceeding of a spacecraft execution mechanism; finally, designing a composite controller according to a backstepping method on the basis of multisource disturbance observation values and anti-saturation compensation items. Therefore, adverse effects of multisource disturbances on the system are avoided, system robustness is improved, accurate spacecraft relative attitude control in a saturation range is guaranteed by the execution mechanism, and smoothness in space operation task completion is guaranteed.

The invention discloses a nonlinear modeling method and a linear trimming method for a composite rotor aircraft, and belongs to the technical field of control of composite rotor aircrafts. According to the method, the aerodynamic characteristics of all parts are analyzed, and a split modeling method is adopted to carry out dynamic modeling; on the basis of the model, combined with a quasi-Newton iteration method and a sequence quadratic programming method, a quasi-Newton-sequence quadratic programming method is proposed to dynamically balance the dynamic model of the composite rotor aircraft and make small disturbance linearity at the equilibrium point, so that an approximate linearized kinetic equation for the equilibrium point can be obtained. The dynamic model established by using the modeling method disclosed by the invention is high in confidence coefficient and simple in calculation, in addition, the new trimming method provided by the invention can avoid solving the inverse matrix of the objective function Hessian matrix, the calculation is simplified, the relatively high convergence speed is ensured, and the method can be used for researching the control method of the composite rotor aircraft.

The invention discloses a two-wheeled self-balance robot self-adaptive sliding mode changing structure control method and system. According to a classical mechanics analytic method and based on a Lagrange algorithm based on energy analysis, molding of a kinetic equation of a two-wheeled self-balance robot of is realized, a sliding mode changing structure controller is designed according to the kinetic equation. The sliding mode changing structure controller comprises a speed sliding mold changing structure controller and an angle sliding mold changing structure controller; the speed sliding mold changing structure controller and the angle sliding mold changing structure controller give each other feedbacks; a feedback equation is that theta r=beta V; and self-adaptive controlling is carried out on the system on the basis of a function approximation mode. Through adoption of the technical scheme of the present invention, the modeling process is enabled to be more simplified and comprehensive; the robustness and the respond speed of the system are raised; simultaneously, since a mutual feedback relation exists between the speed and the angle of the system, when an inclination angle of the system is overlarge, the system will automatically decelerate; while the speed is reduced, the system will return to the balance position; in a condition of facing different road surface conditions, the system can adapt to external environment and large scope load changes, thereby guaranteeing the safety and stability of the system.

The invention provides a gravity-gradient-invariant-based method for estimating an orbit element. The method comprises: step one, carrying out preparation work; step two, carrying out decomposition of an ideal gravity gradient tensor under an east-noth-up (ENU) coordinate system and obtaining feature values; step three, solving decomposition of a measuring epoch J2 model gravity field tensor under the ENU coordinate system and obtaining feature values; step four, solving all measuring epochs r and phi by using feature values of a J2 gravity gradient matrix; step five, calculating an initial orbit element by using the r and phi; and step six, carrying out orbit element smoothing. Therefore, with utilization of the gravity gradient matrix invariant information measured during the satellite operating process, a geometrical distance between the satellite and the earth's core and the latitude of the earth's core are obtained by iterative solution; a measurement equation that uses a semi-major axis, an eccentricity ratio, an orbit inclination angle, a perigee depression angle, and a true anomaly in an orbital element as state parameters is provided by using an obtained data as an observed quantity; and a perturbation kinetic equation of the orbital element is introduced innovatively and five orbital elements are estimated by using a batching least square method. The gravity-gradient-invariant-based method has advantages of high autonomous degree, high anti-interference capability, and low building cost and the like and has advantages that the traditional method does not have in the deep space exploration field.

The invention discloses a remote sensing quantitative inversion method for soil moisture supply amount. The method comprises the integration of dynamic simulation of crop growth, remote sensing quantitative inversion, soil moisture balance and kinetic equation so as to obtain the soil moisture supply amount from a certain time interval to the entire growth period of a crop growth period. The characteristics of dynamic simulation technology and remote sensing quantitative inversion technology for crop moisture stress information are integrated, so that real-time and dynamic monitoring of the soil moisture supply amount in the crop growth period is realized. The largest advantage of the method is that the monitoring of the soil moisture supply amount is dynamic in the aspect of time and continuous in the aspect of space.

The invention relates to a method for controlling a flexible satellite based on a feature model, which is characterized in that the time dimension, the sampling time and the parameters M and m are determined according to the kinetic equation of the flexible satellite; the coefficient range of the feature model is determined according to each obtained variable; the parameters of the feature model are identified by utilizing a gradient method; and a control law is designed according to the coefficients of the feature model obtained through the identification, and the attitude angle of the flexible satellite is controlled through the kinetic equation that the control law is fed back to the flexible satellite. The method has the advantages that the time dimension and the sampling period of the flexible satellite are introduced; the change rate of the flexible satellite is depicted; the bottleneck problem of the feature modeling of the flexible satellite is solved; the expression of the parameter range of the feature model of the flexible satellite is provided; the parameter property of the feature model is qualitatively researched; the boundary of the parameters of the feature model is relative to the sampling period, the modeling error, the system order and the change rate of the system from the given parameter range; and the theoretical foundation of the self-adaptive control based on the feature model is laid for the flexible satellite. The method is suitable for the feature model of the attitude kinetics of an aircraft so as to lay the foundation for the aircraft based on the attitude control of the feature model.

The invention relates to the technical field of vehicles and provides a measuring method for the load of a hybrid electric vehicle. The measuring method comprises the steps that S1, the vehicle total mass m and the vehicle rotational inertia coefficient sigma are measured; 2, the rolling resistance Ff, the air resistance Fw and the acceleration resistance Fj are calculated according to the vehicle total mass m and the vehicle rotational inertia coefficient sigma; S3, the motor output driving force Ftn in the vehicle running process is measured; and S4, the rolling resistance Ff, the air resistance Fw, the acceleration resistance Fj and the motor output driving force Ftn are substituted into a kinetic equation Ft=Ff+Fw+Fi+Fj, and the gradient is calculated according to the acquired gradient resistance Fi. According to the measuring method for the load of the hybrid electric vehicle, the gradient of the vehicle in the running process is calculated by adopting the formula of the Newton second law, and a gradient sensor does not need to be installed on the vehicle; accordingly, not only is the cost reduced, but also the power performance and the economic performance of the whole vehicle can be optimized.

The invention relates to a Mars complex terrain region safe landing trajectory generation method with minimum fuel consumption, and belongs to the technical field of planet landing. The method comprises the steps that a Mars powered lowering kinetic model is built at first, a navigation function is built according to the target landing terrain, obstacle avoidance control force obtained through the navigation function can effectively avoid collision between a lander and an obstacle, and the lander is safely landed to a target landing point. According to the Mars complex terrain region safe landing trajectory generation method with minimum fuel consumption, the obtained obstacle avoidance control force is introduced into a Mars powered lowering kinetic equation, the control force is partially used for achieving obstacle avoidance, and landing trajectory optimization is carried out on the improved kinetic equation, so that obstacle avoidance is achieved, and meanwhile fuel consumed by the powered lowering trajectory is saved. The Mars complex terrain region safe landing trajectory generation method can not only take the target landing area terrain into consideration, but also effectively reduce fuel consumption, and overcomes the defects that obstacle avoidance cannot be achieved and fuel consumption is large in traditional optimization trajectory and obstacle avoidance.

The invention provides a global sliding-mode control method for a bridge crane system. The method comprises the following steps: determining a kinetic model; determining a control target; designing adamping signal and transforming a kinetic equation; designing a control rule; and implementing the control method. According to the control method, the bridge crane system has good robustness in the whole operating process while positioning control and anti-sway control of a trolley of the bridge crane system are implemented, so that the control efficiency of the bridge crane is improved, and theprobability of safety problems of the bridge crane system in actual operation caused by external disturbance and uncertain parameters. The global sliding-mode control method has good application prospect and economical benefit.

The invention relates to a geostationary orbit spacecraft electrical propulsion transfer track control method. The method comprises the following steps: determining initial conditions, a carrier rocket transmitting a geostationary orbit spacecraft to a first initial orbit, and determining a Kepler radical of the first initial orbit; establishing a first initial orbit kinetic equation by means of the Kepler radical, reducing the orbit inclination angle of the first initial orbit through a thrust azimuth alpha and performing rounding to obtain a second initial orbit, and obtaining transfer time tf1 and a propellant consumption amount mfuel1; fixing a thrust accelerated speed in a second preset plane, transferring the second initial orbit through a thrust azimuth beta to a geostationary orbit, and obtaining transfer time tf2 and a propellant consumption amount mfuel2; calculating total time tf=tf1+tf2 of a geostationary orbit transfer process and a propellant consumption amount of the transfer process; and by taking minimization of the geostationary orbit total time tf as a design object, optimizing the thrust azimuth alpha at a first initial orbit phase to obtain an optimal geostationary orbit transferring scheme.

The invention discloses a method for determining current rigidity and displacement rigidity of a permanent magnet biased hybrid magnetic bearing. The method comprises the following steps of: establishing relation among the current rigidity, the displacement rigidity, controller proportional gain and oscillation frequency of simple harmonic motion of a magnetic suspension rotor according to a kinetic equation of the simple harmonic motion of the magnetic suspension rotor; obtaining related data about the controller gain and the oscillation frequency of the simple harmonic motion of the magnetic suspension rotor by adjusting the controller proportional gain; and determining the current rigidity and the displacement rigidity of the magnetic suspension rotor by fitting the coefficient of a unitary linear regression equation. The method has the advantages of simplicity and practicability and can be applied to the analysis and design of a magnetic suspension rotor control system.

The invention discloses a method for estimating a disturbance moment in a power decreasing process. In the process of decreasing soft landing power in the moon, the disturbance moment generated by centroid shift is possible to endanger landing. An important measure for early warning and answering is that the magnitude of the disturbance moment is estimated online. The method comprises the following steps of firstly, recursively calculating the residual mass of a lander by utilizing motor power and a pulse width command given out by a guidance and control system, and utilizing a curve (function) bound on the ground to calculate the height of center of mass of the lander and inertia; secondly, calculating a control moment according to the height of center of mass and a pulse width command; thirdly, carrying out difference on a gyro output to obtain an angular acceleration calculation value; fourthly, inversely calculating the disturbance moment by utilizing an attitude kinetic equation; finally, utilizing a filter to reduce the influence of noise on a disturbance moment estimated value. The invention provides a method for monitoring the change of the disturbance moment of the lander, and the method is beneficial to improve the robustness of the control system in the landing process and reduce the risk of landing.