154 results about "Inertial coordinate system" patented technology

The invention relates to an indoor airship-oriented finite time trajectory tracking control system, platform and method, the system comprises an airship dynamic model, a disturbance estimation module, a sliding mode filter and a self-adaptive controller, the airship dynamic model is formed by constructing a motion stipulation inertial coordinate system and an airship body coordinate system of an airship; the disturbance estimation module estimates unknown disturbance based on a double-layer neural network according to an airship dynamic model and updates the network in real time; the sliding mode filter estimates the influence of noise on the motion speed of the airship according to the unknown disturbance condition of the airship and updates the network in real time; the adaptive controller adjusts the virtual control input of the airship motion trail by controlling the convergence condition of the tracking error within the finite time; according to the method, the disturbance problem caused by uncertainty and dead zones in the airship and the deviation problem of speed signal estimation caused by noise are solved.

The invention relates to a satellite-borne microwave radiometer cold air calibration attitude planning method based on multi-objective optimization, and belongs to the field of satellite microwave remote sensing, and the method comprises the steps: building a satellite attitude kinematics model and an attitude dynamics model through a geocentric equator inertial coordinate system, an orbital coordinate system and a satellite body coordinate system; establishing a brightness temperature sum model received by the microwave radiometer in the cold air calibration process; according to the method, a microwave radiometer receiving brightness temperature sum, brightness temperature fluctuation, platform maneuvering capability, energy and a satellite sensor field of view are comprehensively considered, a multi-target optimization model of microwave radiometer cold air calibration consideration constraints is established, solving is performed, constraint conditions are established, and a solution or a solution set is obtained. According to the method, the brightness temperature sum and the brightness temperature fluctuation in the observation area of the microwave radiometer in the cold air calibration process are optimized by adopting multi-objective optimization, and the on-orbit cold air calibration effect and the system reliability of the satellite-borne microwave radiometer are improved.

The invention discloses an inertial navigation system periodic oscillation suppression method based on time-frequency transformation, and relates to the technical field of underwater navigation, and the method comprises the following steps: S1, modeling an error propagation model of an INS, and defining coordinate systems required by INS navigation, including an inertial coordinate system, a navigation coordinate system and a body coordinate system; based on the INS navigation equation, obtaining an INS error model, and carrying out time domain analysis to obtain an analytical solution of time domain error divergence; s2, designing a time-frequency transformation NTFT, proposing a normal phase time-frequency transformation method, and constructing a complete observation equation; s3, substituting a null method into a calculated longitude and latitude result, and accurately extracting frequency, phase and instantaneous amplitude in inertial navigation longitude and latitude; and S4, establishing a resolving framework of the inertial navigation system periodic oscillation suppression method based on time-frequency transformation, and constructing a negative feedback loop through input gyroscope and acceleration information to suppress periodic errors.

The invention belongs to the technical field of satellite communication, and discloses a satellite communication beam switching method. The positions of a satellite and an aircraft are determined under a geocentric inertial coordinate system, an elevation angle is calculated, and visible wave beams are screened to form a neighbor cell list. And screening out a candidate beam list from the neighbor cell list based on the received power threshold. And if the list is not empty, constructing a decision matrix containing the receiving power, the signal-to-noise ratio, the number of available channels and the satellite residual service time for each candidate beam. And objectively calculating the weight of each performance index based on a matrix by using a CRITIC method. And by using the weight, a TOPSIS model based on Tanimoto similarity optimization is adopted to calculate the relative close degree of each candidate wave beam. And finally, selecting the candidate beam with the maximum relative closeness as the optimal beam to trigger switching. According to the method, through the objective weight and the optimization model, unnecessary frequent switching is reduced, and the communication stability and the system resource utilization rate are improved.

The invention discloses an orchard crawler differential robot trajectory tracking method based on adaptive incremental control, and relates to the technical field of agricultural robot control. The method comprises the following steps: receiving environmental data collected by a robot sensor, constructing a probability grid map of an orchard environment according to the environmental data, and aligning the probability grid map with a pre-constructed map inertial coordinate system; the course angle error and the distance error between the current position and a target point are calculated based on the pose of the robot, an error vector is formed, and the target point is dynamically planned according to the fruit maturity and reachability priority detected by a visual system and a laser radar point cloud. According to the method, coordinate transformation is not needed, and different from traditional tracking control, a control instruction is generated directly based on pose increment, the course and distance error of a target point is calculated directly based on an inertial coordinate system, and real-time matrix projection operation is avoided.

The invention relates to an ephemeris orbit calculation method suitable for orbit extrapolation on an elliptical orbit satellite. The ephemeris orbit calculation method comprises the following steps: calculating ephemeris orbit base point parameters on the ground; ephemeris orbit base point parameters which are injected from the ground are provided; and performing orbit extrapolation by using the ephemeris orbit base point parameters to calculate the position and the speed of the satellite in the inertial coordinate system. According to the method, orbit extrapolation is carried out on the basis of ephemeris orbit base point parameters injected on the ground, the calculation amount is low, the method is suitable for satellite-borne calculation, the calculation precision is high, the high-precision requirement of real-time orbit forecasting in the elliptic orbit satellite inter-satellite communication period can be met, the calculation capacity constraint of a satellite service system can be met, and the completion degree of an inter-satellite communication task is improved.

The invention relates to a GNSS (Global Navigation Satellite System) auxiliary measurement-based autonomous navigation method for a lunar satellite formation release process, which comprises the following steps of: establishing a kinetic model of a primary satellite and a secondary satellite under a lunar center inertial coordinate system J2000; establishing and fusing an inter-satellite measurement model and a distance measurement model of a main satellite fused GNSS signal to obtain a collaborative observation model; and acquiring orbit prediction information based on the kinetic model, acquiring collaborative theory observation information based on the collaborative observation model, performing fusion processing on the orbit prediction information and the collaborative theory observation information by adopting an extended Kalman filtering algorithm, and estimating orbit states of the primary satellite and the secondary satellites. According to the autonomous navigation method, full-autonomous orbit determination can be realized in orbit without depending on the support of a ground station.

The application provides a method for analyzing a pointing angle of an inter-satellite link, comprising: obtaining orbital elements of a satellite and orbital parameters of a target satellite matched with an SGP4 model in a J2000.0 inertial coordinate system; determining a target recursion time t; obtaining positions of the satellite and the target satellite at the time t in the J2000.0 inertial coordinate system; determining a pointing vector from the satellite to the target satellite at the time t in the J2000.0 inertial coordinate system; determining the pointing vector in a satellite orbital coordinate system; determining the pointing vector in a satellite body coordinate system; determining the pointing vector in a satellite antenna coordinate system; determining a pointing angle of the satellite antenna coordinate system at the target recursion time; and determining whether a cycle number reaches a set cycle number, if the cycle number does not reach the set cycle number, continuously obtaining the pointing angle of the satellite antenna coordinate system at a next target recursion time until the pointing angle of the satellite antenna coordinate system at the set cycle number is obtained.

The invention belongs to the technical field of electromagnetic field calculation, and discloses a method and a system for calculating an electromagnetic scattering field of a moving target coated with a lossy dispersive medium. The method comprises the following steps: firstly, carrying out three-dimensional modeling on a target surface by adopting triangular surface elements to form a model file, and importing the model file into a main program to obtain geometric feature data of a target; secondly, carrying out relativistic coordinate transformation on the space-time component by utilizing Lorentz transformation, and mapping a dynamic problem to an inertial coordinate system synchronous with a target; then establishing an equivalent reflectivity model for the lossy dispersive medium coated area by adopting a transmission matrix method; then efficiently solving a transient scattering field in the motion coordinate system by adopting a time domain physical optical method; and finally, reconstructing a calculation result to a laboratory coordinate system through inverse Lorentz transformation, and outputting electromagnetic field distribution in a dynamic scene. Through the method provided by the invention, the electromagnetic scattering field of the moving target coated with the lossy dispersive medium can be quickly calculated.

This invention discloses a control method for intelligent traffic cones, relating to the field of intelligent traffic cones. It establishes a communication connection between the intelligent traffic cones and a central control system based on a fixed base. Simultaneously, the algorithm module analyzes the current position of the moving cone group in the inertial coordinate system based on image data collected by a camera and feeds this position information back to the central control system. When deployment is required, the central control system generates corresponding deployment path information based on the current position and pre-arranged formation of each moving cone group and issues a deployment command and corresponding path information to the control module. After receiving the deployment command, the control module, combining attitude calculation from the inertial navigation module, angular velocity measurement from the wheel speed sensor, calculation from the algorithm module, and the directional error adjustment mechanism of the PID controller, accurately executes the path movement task, effectively solving the path conflict or overlapping coverage problems that may occur during the collaborative deployment of multiple intelligent traffic cones.

The invention discloses a method and a device for predetermining a communication time period between a remote sensing satellite and a high-orbit communication satellite. The specific implementation scheme is as follows: determining at least one visual time period and time period parameters of the remote sensing satellite and the high-orbit communication satellite; for each visual time period in the at least one visual time period, calculating a satellite attitude angle of the remote sensing satellite in the current area working mode based on the visual time period, the time period parameters and the time intervals of different areas of the remote sensing satellite, the earth and the sun; acquiring a geocentric inertial system pointing vector of the remote sensing satellite and the high-orbit communication satellite under the geocentric inertial coordinate system corresponding to the visual time period; calculating an azimuth angle and a pitch angle of the antenna based on the satellite attitude angle, the terrestrial inertial system pointing vector, the at least one visual time period and the time period parameters; and in response to the detection that the azimuth angle and the pitch angle of the antenna meet the angle constraint condition and the detection that the visual time period does not meet the telemetering and data transmission work constraint conditions, taking the visual time period as the communication time period of the remote sensing satellite and the high-orbit communication satellite.

The multi-detection-view-angle target matching method based on the error direction consistency comprises the steps that multiple targets are detected at different angles through multiple aircrafts, and images of the multiple targets at different detection view angles are obtained; detecting each image by adopting a target recognition algorithm, and calculating a sight angle of a target in a camera coordinate system; acquiring a sight angle under an inertial coordinate system, and acquiring position and attitude angle information of multiple aircrafts; calculating direction vectors of the target under different detection view angles and direction vectors among the multiple aircrafts; constructing a target matching model under multiple detection view angles by using the characteristic that the direction vector of the same target under different detection view angles and the corresponding aircraft are in the same plane and have a unique intersection point; according to the characteristic that the error direction of the current multi-target data frame has consistency, target rough matching under multiple detection view angles is completed, and a matching threshold value is calculated; and performing differential matching with the multi-target rough matching value under each detection view angle based on the matching threshold, and coding the successfully matched target.

The application discloses a kind of based on liquid rocket's interstage separation gap calculation method and system, including establishing the inertial coordinate system of liquid rocket launch, rocket body coordinate system and the coordinate system of storage tank;Obtain the dynamics equation of liquid rocket mass center motion in inertial coordinate system and the dynamics equation of liquid rocket rotation around mass center in rocket body coordinate system;Obtain the displacement H of liquid rocket core first stage and liquid rocket core second stage and attitude angle;Obtain the displacement L1 of liquid rocket core second stage mass center deviating from nozzle center line, the displacement L2 of liquid rocket core second stage mass center deviating from liquid rocket core first stage axis and the distance L3 of nozzle outer edge to nozzle center line;Obtain interstage separation gap as L4, L4=R-L3-(L1-L2), solve the problem that core second stage engine nozzle and interstage section sleeve may collide in the process of large-diameter liquid rocket separation, to improve the precision of separation scheme design, attitude control system design and storage tank anti-slosh structure design.

The invention relates to the field of satellite orbit calculation, in particular to a low-orbit satellite ephemeris analysis method based on a satellite orbit plane coordinate system, which comprises the following steps: acquiring TLE two-line orbit data of a target satellite, geodetic coordinates of a terminal, UTC time and earth rotation average angular velocity; calculating the state of the satellite in the orbital plane; calculating a transformation matrix from a geocentric inertial coordinate system ECI to a satellite orbit plane coordinate system; the LBH coordinates of the terminal are converted into ECEF rectangular coordinates; a rotation matrix from the ECEF to the ECI is constructed, and the ECI system speed of the terminal is calculated; according to the transformation matrix, transforming the ECI coordinate of the terminal into a final satellite orbit plane coordinate system; performing simplified calculation based on a unified calculation framework of a satellite orbit plane coordinate system; and outputting the distance, the relative speed, the Doppler frequency offset and the propagation time delay result which are calculated under the satellite orbit plane coordinate system. And the calculation complexity and the realization difficulty of high-precision orbit forecasting are obviously reduced.

A method for error self-calibration and compensation of liquid floated gyro platform inertial navigation system is provided. In the condition of static base, the platform is locked at 13 frame angle positions, then the platform is controlled to track inertial coordinate system through the stable loop, and the measurement values of accelerometer and frame angle sensor are collected. The platform attitude angle in the earth coordinate system is calculated by using the platform attitude solution method of inertial navigation system, and is compared with the platform attitude angle measured by the platform frame angle sensor to obtain the attitude misalignment angle observation. The state misalignment angle sequence of each tracking inertial coordinate system process is linearly fitted to obtain the comprehensive angular velocity error of the inertial navigation system at each calibration position. The residual of the zero order, first order and cross term drift error coefficients is solved by using the total least squares algorithm, the estimated value of the measured parameter is corrected, and the self-calibration result of each drift error coefficient is obtained by iterative calculation until the iterative convergence criterion is met.

The embodiment of the invention provides a three-dimensional maneuvering target interception method, a storage medium and equipment. The method comprises the following steps: determining a constraint relation between the speed and the maximum field angle of an interception aircraft and the speed and the maximum field angle of an intercepted aircraft; calculating a collision angle control gain; obtaining a relative speed expected value in the inertial coordinate system according to the expected collision angle in the speed coordinate system; constructing a sliding mode vector as an auxiliary variable according to the relative speed expected value, the sight line direction unit vector and the sight line conversion rate information; setting a gain value of auxiliary variable control; a guidance instruction to intercept the aircraft is calculated. The embodiment of the invention provides a three-dimensional space guidance law design method considering collision angle and field angle constraints at the same time, the problem of three-dimensional space interception under multiple complex constraint conditions is solved, and an air defense and anti-missile combat system is assisted to obtain higher combat performance.

A relative navigation sensor field-of-view obstacle identification and avoidance method comprises the following steps: firstly, setting initial attitude control parameters of a tracking aircraft, and calculating a unit vector of a relative navigation sensor field-of-view optical axis in an inertial coordinate system; then, respectively calculating the included angles between the unit vector relative to the visual field optical axis of the navigation sensor and the unit vectors of the earth, the moon and the non-target aircraft; and finally, sequentially judging whether interference elimination operation of the earth, the moon and the non-target aircraft needs to be carried out or not. Through the above steps, the method effectively eliminates the influence of the earth, the moon and the non-target aircraft on the target recognition of the relative navigation sensor, and improves the precision of the target recognition of the relative navigation sensor.

The invention relates to a satellite-borne autonomous high-precision inertial coordinate system and earth-fixed coordinate system conversion method. The method comprises the following steps: calculating a polar shift parameter in earth orientation parameters, and a difference UT1-UTC of world time UT1 and coordinated world time UTC; calculating the Julian century number based on solar system centroid force class hours and the Julian century number based on world time UT1; calculating an age difference parameter by using the Julian century number based on solar system centroid force class hours; calculating nutation parameters by using the Julian century number based on solar system centroid force hour; calculating earth rotation parameters by using the Julian century number based on solar system centroid force class hours; respectively calculating a polar shift matrix, an age difference matrix, a nutation matrix and an earth rotation matrix by using the polar shift parameter, the age difference parameter, the nutation parameter and the earth rotation parameter; and calculating a conversion matrix of the earth-fixed coordinate system and the inertial coordinate system based on the polar shift matrix, the age difference matrix, the nutation matrix and the earth rotation matrix so as to perform conversion between the earth-fixed coordinate system and the inertial coordinate system.

The present application relates to a kind of inertial measurement device platform body axis unbalance moment estimation and offset method, belong to inertial measurement technical field;Inertial coordinate system OXYZ is established based on inertial measurement device;Establish the control model of platform body axis stable loop in inertial measurement device under the influence of unbalance moment;The model of unbalance moment acting on platform body axis is established;According to unbalance moment model, establish unbalance moment parameter observer, estimate the unbalance moment parameter that platform body axis is subjected to by unbalance moment parameter observer, obtain unbalance moment parameter estimated value;According to unbalance moment parameter estimated value, calculate platform body axis compound control moment;Through platform body axis compound control moment, improve the control precision of platform body axis stable loop;The present application has the characteristics of low cost, high precision, strong engineering practicability, is applicable to the high-precision control of platform system platform body axis stable loop under the influence of unbalance moment.

The present application provides a method for planning and controlling a large-curvature curve for an autonomous driving vehicle, comprising the following steps: obtaining path information in a current scene, collecting a reference path, and planning a target speed; obtaining vehicle status information; obtaining a reference path point in the reference path based on the position of the vehicle's preview point, calculating a lateral position error and a heading angle error from the horizontal and vertical coordinates and heading angle of the reference path point and the horizontal and vertical coordinates and heading angle of the preview point, and inputting the calculated lateral position error and heading angle error into a lateral controller; the lateral controller is established based on a variable parameter state feedback control algorithm using a table lookup method; the difference between the reference speed generated by a tracking differentiator for the target speed and the longitudinal speed of the vehicle in the inertial coordinate system XOY is input into the longitudinal controller; the longitudinal controller is established based on a terminalless model predictive control algorithm; TD is added to the model predictive control algorithm to arrange a transition process for a specified speed signal, so that the longitudinal speed tracking is more stable and smooth.

The invention discloses a multi-quad-rotor unmanned aerial vehicle formation collision avoidance fault-tolerant control method, and relates to the technical field of unmanned aerial vehicle control, and the method comprises the steps: constructing a kinetic model and a dual-state constraint model of a multi-quad-rotor unmanned aerial vehicle, and building an inertial coordinate system and a vehicle body coordinate system; according to the invention, the dynamic model and the dual-state constraint model are constructed, unified modeling processing is carried out on the motion behavior and the space safety boundary of the multi-quad-rotor unmanned aerial vehicle, and real-time acquisition and dynamic compensation are carried out on the state and fault information of the unmanned aerial vehicle in combination with an online estimation mechanism of a distributed fault diagnosis observer. The system can still maintain stable operation under the conditions of actuator faults and coupling disturbance; the position error and the formation error of the unmanned aerial vehicle are constrained, the mapping from virtual control to actual control is realized in combination with an attitude coupling compensation mechanism, the collision risk in the formation flight process is effectively avoided, and the trajectory tracking precision is improved.

The invention discloses a step-by-step marginalized point-line visual inertia SLAM (Simultaneous Localization and Mapping) method and system. Firstly, an environment image and IMU data are obtained and preprocessed; point-line features in the image are extracted, line features with the length larger than the minimum line segment length threshold value in all layers of an image pyramid are reserved, and the direction included angle and the center point Euclidean distance of any two reserved line features are calculated; if the direction included angle of the two line features is smaller than or equal to an angle threshold value and the center point Euclidean distance is smaller than or equal to a distance threshold value, combining the two line features into a new line feature; then, aligning the visual inertial coordinate system, and estimating an initial state; and finally, performing local visual inertia optimization and mapping by using a sliding window based on a step-by-step marginalization strategy to obtain a prior error. And meanwhile, loopback detection is carried out, and loopback errors and prior errors are added into the next round of sliding window optimization to carry out tight coupling state estimation. The method has relatively high robustness and consistency, and meanwhile, the real-time performance is ensured.

This invention discloses a method for precise airdrop and rapid in-flight alignment based on inertial navigation / satellite integrated navigation. The method includes: after the navigation device is powered on, establishing a carrier and navigation inertial coordinate system through inertial solidification; periodically acquiring angular rate and specific force data from the inertial measurement unit and position and velocity data from the satellite receiver; sequentially calculating carrier system tracking, specific force integration, navigation system tracking, and acceleration integration; and recursively solving the problem using a Rodrigues parameter multi-vector attitude determination algorithm to calculate the real-time attitude matrix, repeating this process until alignment is complete. This invention simplifies the initial alignment problem without initial values ​​for arbitrary attitudes into an unconstrained linear optimal estimation problem for Rodrigues parameters, utilizing specific force and acceleration integration as measurement information, and fully utilizing all effective data during the alignment process. It requires no initial attitude information and can rapidly complete alignment under complex dynamic base conditions such as airdrop swaying and rotation, outputting attitude results in real time, significantly improving alignment accuracy and speed.

A virtual object display device comprises a display and a display control device configured to perform display control of the display. The display control device includes: a coordinate system calculation unit configured to detect movement and rotation of the virtual object display device in a real world, and use an inertial coordinate system in which a coordinate origin thereof follows the movement of the virtual object display device and an effective field of view of the display is rotated therein in accordance with the rotation of the virtual object display device so as to define an arrangement position of an inertial coordinate system virtual object; and a display control unit configured to, when the inertial coordinate system virtual object is included in the effective field of view of the display, display the inertial coordinate system virtual object within the effective field of view.

The application discloses a spacecraft single-pulse absolute reachable set solving method based on differential algebra technology and belongs to the technical field of on-orbit motion of space spacecrafts; firstly, a geocentric inertial coordinate system, a geocentric geodetic rotating coordinate system and a cylindrical coordinate system are established, and a dynamics model of on-orbit motion of the spacecraft in the cylindrical coordinate system is established; then, the earth shape perturbation, the solar pressure perturbation and the third body gravity perturbation of the sun and the moon are converted into the cylindrical coordinate system for representation, and a dynamics model considering multiple perturbation forces is obtained; finally, the absolute reachable set is solved in the dynamics model considering multiple perturbation forces based on Jet Transport technology; the dynamics model considering multiple perturbation forces is established in the cylindrical coordinate system, the absolute reachable set is solved by using Jet Transport technology based on the dynamics equation, and the absolute reachable set is simulated by using C++ and MATLAB; compared with the traditional Monte Carlo method, the solving method based on Jet Transport technology is high in accuracy and calculation efficiency.

The invention relates to the technical field of integrated navigation, in particular to a positioning method and system based on optical angle measurement and astronomical-inertial integrated navigation. According to the method, firstly, a star image and a navigational satellite image are obtained at the same time by shooting a star map, after recognition and processing, the attitude of a camera relative to a geocentric inertial coordinate system is calculated by utilizing the star, and then the coordinate of the navigational satellite in the coordinate system is determined. And obtaining angle information of the navigation satellite relative to the observer under the earth coordinate system through coordinate transformation. Optical angle measurement information is fused with attitude, speed and position information output by an inertial navigation system, a Kalman filtering state equation and an observation equation are constructed, and finally real-time high-precision calculation and updating of the position of the observer are achieved. According to the method, the accuracy of optical angle measurement, the limited distance reference of celestial navigation and the autonomous continuity of inertial navigation are fused, and the navigation positioning accuracy and reliability under the complex environment and the satellite denial condition are effectively improved.

A low-orbit remote sensing satellite three-axis rapid maneuvering attitude trajectory planning and tracking method relates to the technical field of aerospace, in particular to the technical field of low-orbit remote sensing satellite three-axis rapid maneuvering attitude trajectory planning and tracking, which effectively solves the balance problem of maneuverability and stability in the process of attitude transformation and improves the superiority of anti-interference capability. The method comprises the following steps: establishing a satellite kinematics model and a satellite dynamics model; calculating the quaternion of the desired coordinate system relative to the satellite initial coordinate system at the maneuvering starting moment; completing attitude trajectory planning through an improved tracking differentiator; calculating the total control rate; obtaining the real control rate through an actuator; calculating the quaternion of the satellite body coordinate system relative to the inertial coordinate system and the angular velocity of the satellite body coordinate system relative to the inertial coordinate system; obtaining the real body quaternion and the real body angular velocity through a measuring mechanism; and cyclically planning the trajectory and tracking.

The invention provides a quad-rotor unmanned aerial vehicle time sliding mode control method and device based on actuator faults, and relates to the technical field of unmanned aerial vehicle flight control. The method comprises the following steps: establishing a quadrotor unmanned aerial vehicle body coordinate system and an inertial coordinate system based on the structural characteristics of a quadrotor unmanned aerial vehicle with an X-shaped layout; based on the influence of the partial failure fault and the bias fault of the actuator, constructing a four-rotor unmanned aerial vehicle kinetic model which takes the uncertain item external disturbance and the actuator fault as lumped disturbance in a unified manner; introducing an auxiliary equation to design a predefined time disturbance observer; respectively designing a non-singular predefined time sliding mode controller based on a position subsystem and an attitude subsystem to carry out segmentation processing on a sliding mode surface; and based on a sliding mode controller and in combination with a disturbance observer, inner and outer ring coordination control of the four-rotor unmanned aerial vehicle is realized. According to the method, unknown disturbance and actuator faults can be effectively handled, the response speed, robustness and fault-tolerant capability of the system are remarkably improved, and the method is suitable for high-precision flight tasks in a complex environment.

The invention discloses an inter-satellite link fault avoidance method and system based on time quantitative calculation under the sun-sun effect, and mainly solves the problems that the calculation complexity of inter-satellite link faults of a satellite network is high and the link fault duration cannot be directly calculated under the influence of the sun-sun effect in the prior art. According to the implementation scheme, the method comprises the following steps: initializing constellation parameters, establishing a J2000.0 geocentric inertial coordinate system, and determining a satellite position at any moment; the right ascension and declination of the sun are calculated, the included angle between the sunlight and each satellite orbit plane is calculated according to the parameters, the link fault position existing in the orbit is determined, and the link fault range of each orbit is calculated; calculating positions and fault duration of all fault links in each orbit according to the link fault range and the positions of satellites; and the satellites autonomously perform link avoidance according to the link fault range and the fault duration. According to the method, the calculated amount can be effectively reduced, the response speed is improved, the duration of the link fault is directly obtained, and the method can be used for analyzing the inter-satellite link fault condition of the satellite network under the sun-sun effect.

The application discloses a satellite solar cell array rotation angle estimation method, which comprises the following steps: step 1, establishing an earth inertial coordinate system, a satellite body coordinate system and a solar cell array coordinate system, and defining a satellite solar cell array rotation angle; step 2, establishing a state equation of an augmented state composed of a sun vector, the solar cell array rotation angle and a gyro constant drift noise; step 3, calculating an augmented state quantity estimation value by using an extended Kalman filter; and step 4, calculating a satellite solar cell array rotation angle estimation value. The method fuses gyro angular velocity and sun sensor vector data, constructs a nonlinear dynamic model containing a sun vector, a rotation angle and a gyro bias, and designs an EKF algorithm to estimate the solar cell array rotation angle in real time, thereby solving the strong dependence of a traditional scheme on a star sensor and a mechanical sensor, and significantly improving the autonomous recovery capability and energy guarantee reliability of a satellite in an extreme fault scene.