201 results about "Inertial coordinate system" patented technology

The invention discloses a four-rotor unmanned aerial vehicle model prediction control method based on a deep Q-network (DQN). The method comprises the following steps: firstly, establishing an inertial coordinate system and a body coordinate system, and establishing a kinematic equation and a kinetic equation of the four-rotor unmanned aerial vehicle based on a Newton second law and an Euler-Lagrange equation; according to the method, an unmanned aerial vehicle robust controller based on model predictive control (MPC) is designed, optimal control input is generated through dynamic optimization, and the trajectory tracking performance of the unmanned aerial vehicle is improved. For various disturbances existing in the whole tracking process, a DQN reinforcement learning algorithm and an MPC method are combined to design a flight control system, control rate errors caused by the disturbances are compensated, the stability of attitude control is enhanced, and the tracking precision and the anti-interference capability are improved. And finally, verifying the robust performance of the flight control system through a simulation experiment.

The invention, which relates to the technical field of the radar system, discloses a swash plate machine phase-sweeping radan installation angle adaptive correction method comprising the following steps: acquiring three-axis attitude angle data of an installation reference surface, and synchronously acquiring azimuth angle rotation data of a radar antenna; based on the three-axis attitude angle data of the installation reference surface, establishing a coordinate transformation matrix from a radar antenna coordinate system to an inertial coordinate system, and mapping the three-axis attitude angle data of the installation reference surface to the inertial coordinate system; obtaining a beam scanning conical surface deviation angle according to the coordinate transformation matrix and azimuth angle rotation data of the radar antenna; generating antenna attitude correction according to the beam scanning conical surface deviation angle; and inputting the antenna attitude correction quantity into an antenna driving controller, adjusting the attitude of two axes of pitching and orientation of the antenna, and compensating the beam scanning conical surface deviation. According to the method, the precise coordinate transformation matrix and the angle difference compensation model are established, and the inertial measurement unit and a closed-loop feedback control mechanism are combined, so that the self-adaptive real-time correction of the mounting angle is realized.

The invention is suitable for the technical field of multi-unmanned aerial vehicle cooperative control, and provides a multi-quadrotor formation maneuvering flight event triggering cooperative control method and system, and the method comprises the steps: firstly building a kinetic model of a quadrotor unmanned aerial vehicle in a distributed formation system under an inertial coordinate system, designing a multi-quadrotor unmanned aerial vehicle cooperative formation synchronization error, and carrying out the cooperative formation of the quadrotor unmanned aerial vehicle; establishing a preset performance control function to constrain the formation synchronization error, and performing corresponding equivalent error transformation; designing a dynamic event trigger track subsystem controller; designing a robust attitude subsystem controller; and constructing a Lyapunov function, and verifying whether a closed-loop system is stable or not and whether an expected formation effect can be achieved or not. The method can effectively restrain the synchronization error of the multi-quadrotor unmanned aerial vehicle formation system, reduces the error overshoot, improves the coping capability for uncertainty, environmental interference and abrupt trajectory change, and guarantees the stable formation flight of the multi-quadrotor unmanned aerial vehicle system.

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 discloses a two-stage installation error outfield calibration method for a space target fine-following camera and a two-dimensional rotary table, and belongs to the technical field of error calibration. The method comprises the following steps: extracting shooting information of each angle in an inertial coordinate system according to a plurality of starry sky images at different shooting angles; according to the inertial sight vector truth value of the equator coordinate under the inertial coordinate system, calculating a horizon sight vector truth value under the east-north-sky coordinate system by combining a transfer matrix from the east-north-sky coordinate system to the inertial coordinate system; calculating a horizon sight vector measurement value of the corresponding angle according to the rotary table rotation angle of each angle and the equipment installation matrix; and inputting the errors of the truth values and the measured values of the horizon sight line vectors under all the shooting angles into a least square iterative algorithm for optimization calibration, and outputting an optimal rotary table installation matrix and an optimal camera installation matrix. Calibration of the fine-following camera of the ground test two-dimensional rotary table can be realized, and the pointing precision of the fine-following camera is improved.

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 provides a non-singular sliding mode control method, device and equipment for a quad-rotor unmanned aerial vehicle, and a medium, and relates to the technical field of unmanned aerial vehicle flight control. According to the method, a quadrotor unmanned aerial vehicle body coordinate system and an inertial coordinate system are established, and an attitude model of the quadrotor unmanned aerial vehicle is established by considering air resistance and external unknown disturbance conditions according to a rigid body motion law; introducing an auxiliary equation to construct a predefined time disturbance observer, and rapidly estimating external disturbance by adjusting predefined time in combination with a Lyapunov function; then designing a non-singular predefined time sliding mode controller; and controlling the quadrotor unmanned aerial vehicle to perform attitude tracking according to the nonsingular predefined time sliding mode controller, and controlling the attitude tracking error of the quadrotor unmanned aerial vehicle to converge within the predefined time by adjusting predefined time parameters. According to the method, the external disturbance can be quickly estimated within the preset time, and the convergence rate of the four-rotor unmanned aerial vehicle is flexibly controlled.

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 provides an inter-satellite link sun-sun avoidance method for a giant satellite base satellite network. The method comprises the following steps: initializing the giant satellite base satellite network; calculating the date of the inter-satellite link influenced by the sun; calculating the included angle between each inter-satellite link and the sun position vector in the date influenced by the sun; and avoiding the sun. According to the method, on the basis of the uniform distribution characteristic of a Walker constellation, the dates of the same-orbit link and the different-orbit link influenced by the sun-sun are calculated under a J2000 geocentric inertial coordinate system, the dates of the inter-satellite links influenced by the sun-sun are calculated according to the motion law of the sun and the characteristics of the different types of inter-satellite links, and the data of the inter-satellite links influenced by the sun-sun are calculated within the dates influenced by the sun-sun. According to the relative position relation of the satellites, the included angle between the inter-satellite link and the sun is calculated in real time, whether the inter-satellite link is affected by the sun or not is judged, the optical axis direction of the laser terminal does not need to be obtained in real time, the included angle between the inter-satellite link and the sun does not need to be calculated in real time on the date without the sun influence, and invalid calculation expenses are reduced.

The invention provides a centimeter-level broadcast ephemeris method and device based on a GNSS precise ephemeris, and the method comprises the steps: converting a classical Kepler orbit element into a nonsingular orbit element for the near-zero eccentricity rate of a middle earth orbit satellite; performing fast Fourier transform on a satellite orbit, selecting a certain number of additional parameters from the optional additional parameter set according to a spectral analysis result of the Fourier transform, and combining the additional parameters with an original broadcast ephemeris model to construct a centimeter-level broadcast ephemeris model; interpolating the precise ephemeris to obtain dense ephemeris data points, and determining centimeter-level broadcast ephemeris model parameters; calculating an error in a geocentric inertial coordinate system, converting the error into an orbital coordinate system, and determining the precision of a centimeter-level broadcast ephemeris model by using a space signal ranging error; and broadcasting the centimeter-level broadcast ephemeris model for satellite orbit determination and user high-precision positioning. The method provides support for real-time high-precision satellite positioning application.

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 a flight path tracking control method for an unmanned autonomous submersible vehicle, and belongs to the technical field of unmanned submersible vehicle control. The method comprises the following steps: firstly, establishing a nonlinear dynamic model of the submersible vehicle, analyzing kinematics and dynamic characteristics of the nonlinear dynamic model, considering various underwater forces and moments, and defining an inertial coordinate system and a submersible vehicle body coordinate system to obtain a kinematics equation and a dynamic equation so as to provide a basis for control design; secondly, a non-singular integral terminal sliding mode control method is adopted, a sliding mode surface is designed, the singular problem of traditional sliding mode control is avoided, a sliding mode control law is deduced, system finite time convergence to a balance point is achieved, and an expected state is rapidly tracked. And finally, a brand new and simple adaptive control strategy is provided, an adaptive law is designed according to system errors and sliding mode surface information, a gain matrix is adjusted in real time, and system uncertainty and external interference are compensated. The method has the advantages of being high in robustness, rapid in convergence, simple and easy to implement.

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 provides a correlation method and system for space debris observation data and cataloguing orbit data. The method comprises the following steps: firstly, within a time range of currently processed batch observation data, recurring to an orbit matrix distributed at equal time intervals through an orbit of an orbit recursion cataloguing target, and generating a position interpolation function based on a time sequence and the orbit matrix; and for observation data in a processing time range, firstly unifying the observation data and an observation station position to a J2000.0 inertial coordinate system, and obtaining a multi-target position prediction value corresponding to observation time through an interpolation function. Calculating the difference between the predicted value and the measured value of the observation data; and calculating an associated object and confidence according to the difference. The method is easy to implement, high in efficiency during batch observation data processing and suitable for association of observation data of different observation devices and different types of orbit space debris and catalogued orbit data.

The invention provides a satellite flexible attitude dynamics modeling method considering the change of the mass center of a whole satellite, and the model is used for simulating the on-orbit flexible attitude behavior of the satellite. The method comprises the following steps: establishing an inertial coordinate system Oi, an orbital coordinate system Oo, a star coordinate system Ob and a sailboard coordinate system Of; establishing an attitude motion equation Eq.1 of a whole satellite system, an equation Eq.2 of flexible vibration of the sailboard and an attitude motion equation Eq.3 of the sailboard relative to a central body, wherein the Eq.1, Eq.2 and Eq.3 consider the influence of the mass center change of the whole satellite and the flexible vibration of the large sailboard; and calculating a direction cosine matrix RoI of the centroid orbital coordinate system, a direction cosine matrix RsI of the star coordinate system, an attitude conversion matrix Rso of the star coordinate system relative to the centroid orbital coordinate system and an Euler angle, and taking the Euler angle as attitude control input. According to the satellite attitude dynamics calculation method considering the satellite centroid change and the sailboard elastic deformation in real time, the problem that the instantaneous centroid change cannot be simulated in an existing attitude dynamics algorithm is solved.

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 invention provides a state estimation method and device, electronic equipment and a storage medium. According to the specific implementation scheme, the method comprises the following steps: acquiring a state quantity, first data acquired by a first inertial measurement unit arranged on first equipment, and second data acquired by a second inertial measurement unit arranged on second equipment; the state quantity comprises a target parameter, a relative position and a relative attitude between the second equipment and the first equipment, an angular velocity bias and an acceleration bias of the first inertial measurement unit, and an angular velocity bias and an acceleration bias of the second inertial measurement unit; the target parameter comprises at least one of the following two: the relative speed between the second equipment and the first equipment, and a projection result obtained by projecting the relative speed between the second equipment and the first equipment under an inertial coordinate system to a reference coordinate system; obtaining a state estimation equation meeting a predetermined lever arm constraint condition; an estimated state quantity is determined based on the state quantity, the first data, the second data, and a state estimation equation.

The invention relates to an error-free strapdown inertial navigation method based on a launching inertial coordinate system, and relates to the technical field of strapdown inertial navigation. According to the error-free strapdown inertial navigation method based on the launching inertial coordinate system, under the condition that the output angular velocity of a gyroscope and the output specific force of an accelerometer meet a polynomial hypothesis, smooth angular motion or linear motion can always be infinitely approximated by a polynomial according to mathematical knowledge, and the linear motion can be infinitely approximated by the polynomial according to a navigation equation of strapdown inertial navigation under the launching inertial coordinate system. And a strapdown inertial navigation value updating algorithm for the attitude, the speed and the position under the engine inertial system is exported by adopting a Taylor series expansion method. The algorithm does not have any principle error and comprises perfect compensation for non-exchangeable errors of attitude, speed and position updating, and the algorithm precision is far higher than that of a traditional two-subsample algorithm.

Provided in the present disclosure are a method and apparatus for state estimation, and an electronic device and a storage medium. A specific implementation comprises: acquiring state variables, first data which is collected by a first inertial measurement unit arranged in a first device, and second data which is collected by a second inertial measurement unit arranged in a second device, wherein the state variables comprise: a target parameter, relative positions and relative poses of the second device and the first device, an angular velocity bias and an acceleration bias of the first inertial measurement unit, and an angular velocity bias and an acceleration bias of the second inertial measurement unit, and the target parameter comprises at least one of the following two pieces of information: the relative velocity between the second device and the first device, and a projection result obtained by projecting the relative velocity between the second device and the first device in an inertial coordinate system to a reference coordinate system; acquiring a state estimation equation meeting a predetermined lever arm constraint condition; and on the basis of the state variables, the first data, the second data and the state estimation equation, determining estimated state variables.

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.

The invention provides a space debris radar observation segmental arc initial orbit determination method and system. Radar observation segmental arc data of space debris are acquired. And obtaining a distance measurement error and an angle measurement error of radar observation equipment. And calculating the sight line vector of the space debris relative to the observation station in the geocentric inertial coordinate system and the position coordinate of the observation station in the geocentric inertial coordinate system. And calculating an orbit position speed initial value. And calculating a target function corresponding to the current track position. And calculating partial derivatives of the relative position and speed of the target function. And comparing the current target function with the minimum value of the historical target function, and recording according to a comparison result. And judging whether a convergence condition is met or not, and if yes, outputting the position and speed corresponding to the minimum value of the historical target function. Otherwise, updating the track position and the speed value. The method is easy to implement and can be effectively applied to determining the initial orbit of the space debris radar observation arc section.

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.