78results about How to "Improve attitude accuracy" patented technology

The invention discloses a state estimating and interfusion method based on strap down inertial navigation system that includes the following steps: using the sensor in IMU to induce the kinetic characteristic; taking strap down inertial computing; using the state signal of acceleration estimating system; judging the reliability of state estimating value; taking state information interfusion; outputting navigation parameter. The invention has the following advantages: no additional hardware cost, fully autonomy; effectively improving the navigation accuracy of the system; and supplying 50 navigation signals per second to control display device.

The invention relates to a three-dimensional gesture determining and local positioning method of a lunar surface rover, which comprises the following steps: (1) ascertaining the rolling and pitching angles by use of a triaxial accelerometer with sensitivity while the rover is still; (2) determining the drift angle gesture by means of a sun sensor; (3) using the axial gesture and the gyro deviation as the state quantity, the rolling and pitching angles established by the triaxial accelerometer, the drift angle determined by the sun sensor as well as three gyro outputs as the measuring information, building a state equation and a measuring equation, and estimating the triaxial and gyro deviations by means of extended Kalman filter; (4) after compensating the gyro outputs by virtue of the estimated gyro deviations while the rover is in motion, calculating the gesture changes of the rover, finishing the preestimation of the gyro gesture, and fulfilling gesture update; (5) acquiring the information about the rotation speed of the driving wheels of the rover, the rotating angle of the steering wheel, the rotating angle of the left and right rocker arms, and getting the position increment of the rover in the body coordinate system by use of the forward kinematics relationship. The invention has the advantages of high precision of gesture determining and positioning, simple calculation and easy implementation of the engineering.

The invention discloses a method for improving attitude determination precision of an airborne laser radar by solving three attitude angle errors through two-dimensional coordinate errors of feature points on a geodetic coordinate system under the condition of selecting an INS with low cost and low precision. The method is characterized in that the method combines three-dimensional initial correction coordinates of GPS positioning data, INS attitude determination data, laser distance measurement data, scanning angle data and other feature points on the geodetic coordinate system, solves the attitude angle errors by differences between the initial correction coordinates of the feature points in the two-dimensional direction of the geodetic coordinate system and real coordinates according to the mapping relation between the coordinate errors of laser footprints on the geodetic coordinate system and the attitude angle errors, and utilizes the solved attitude angle errors to amend the INS attitude data with low precision, thereby improving the attitude determination precision of the airborne laser radar. The method can greatly reduce system cost and amend the attitude angle errors with high frequency, thereby being applicable to the situations of platform attitude frequent change in light and small aerial remote sensing applications.

The invention belongs to the technical field of navigation methods, and concretely relates to an inertia and mileometer combination high-precision attitude measurement method. The method can be applied in fields of vehicle long-endurance combination navigation and complex landform plotting. A technical scheme adopted in the invention is characterized in that combination Kalman filtering calculation is carried out with a dead reckoning result as a reference, and a horizontal misalignment angle and a horizontal gyro drift are estimated and corrected. A Kalman filtering state equation adopts 19 state variables, a system error model is established, the horizontal misalignment angle and the horizontal gyro drift are established in a real-time and online manner by using the dead reckoning result, and are fed back and corrected to improve the attitude precision, so the system combination navigation precision is effectively improved.

The invention discloses a method for improving resolving precision and anti-jamming capability of an attitude heading reference system. The attitude heading reference system mainly comprises a micro mechanical gyro, a micro mechanical accelerometer and a geomagnetic sensor, and is used for measuring attitude information of a movement carrier in real time. The method specifically comprises the following steps of: firstly, initializing the system according to output of the accelerometer and the geomagnetic sensor for completing a coarse alignment process; secondly, according to output of the gyro, determining an updating equation of a quaternion, calculating a quaternion of an updating system in real time, and realizing an updating process of the system; thirdly, with the output of the accelerometer and the geomagnetic sensor as the reference, realizing a correction process of the system by applying a kalman filtering technology; and fourthly, by combining with a fuzzy control theory, adding a fuzzy control module in the system, estimating the measurement noise of the system in real time, and realizing self-adaptive kalman filtration. According to the invention, the calculating complexity of the system is greatly lowered while the precision of the attitude heading reference system is improved; and especially, when the system is interfered by larger noise, the method can be used for rightly estimating and measuring the characteristic of the noise, thereby improving the attitude precision and enhancing the stability and reliability of the system.

The invention provides a multi-sensor-based attitude determination system, which comprises a sensitive subsystem, a signal acquisition and preprocessing subsystem and an integrated processing subsystem, wherein the sensitive subsystem integrates four classes of sensitive parts such as star sensitive probes, sun sensitive probes, fiber optic gyros and magnetic field probes; the signal acquisition and preprocessing subsystem is realized by an FPGA chip and a peripheral circuit thereof, parallelly receives output data of the four classes of sensitive probes and completes the smoothing and preprocessing of the data; and finally, the integrated processing subsystem taking a DSP (or ARM) chip as a core is utilized to realize the high-precision attitude determination of the system by an ant colony particle filter-based combined attitude determination method according to sunlight information, starlight information, inertial information and geomagnetic information after smoothing and preprocessing. The multi-sensor-based attitude determination system with high precision, integration level and reliability is realized, and has important practical significance for high-resolution earth observation of new-generation earth observation satellites.

The invention discloses an online calibration method of a star sensor assisted gyroscope for ships, and relates to the technical field of inertial sensor error calibration in inertial navigation. Theonline calibration method comprises following steps: (1) initializing a star sensor / inertia combined navigation system; (2) collecting the output data of an inertia device and a star sensor; (3) calculating the inertia device to obtain attitude quaternion and navigation information; (4) obtaining the real value of the carrier attitude quaternion; (5) utilizing a fuzzy logic control method to determine the forgetting factor of a simplified Sage-Husa adaptive filtering method, and then subjecting the carrier attitude quaternion information to filter calculation; (6) compensating the gyroscope output angular velocity information based on the gyroscope output error, and carrying out navigation calculation; and (7) storing and outputting navigation information. The applicability of online calibration on the star sensor assisted gyroscope in a complicated environment is improved. The problem that the estimate on noise error is not sufficient in measurement is solved. The attitude precision of a star sensor is effectively improved.

The invention relates to a satellite attitude determination method and an attitude determination error analytical method, in particular to a satellite attitude determination method and an attitude determination error analytical method based on two star sensors, and aims at solving the problems that the existing measuring error in the star sensor optical axis direction is large and presently no attitude determination error analytical method exists. The first scheme is that the component of the optical axis vector of each star sensor in an inertial system is calculated according to measuring output of the two star sensors, the component of the optical axis vector of each star sensor in a satellite body system is calculated according to an installation matrix of each star sensor, the gesture of the satellite body system relative to the inertial system is calculated, and a gesture matrix is subjected to unit orthogonalization processing. The second scheme is that an attitude determination error matrix C* in the satellite gesture matrix and not subjected to unit orthogonalization is separated; the expression of the attitude determination error matrix C* is evaluated according to the star sensor installation matrixes; and a satellite attitude determination error matrix delta C subjected to unit orthogonalization is gained. The satellite attitude determination method and the attitude determination error analytical method are used for satellite attitude determination and attitude determination error analyzing.

The invention relates to a multi-mode attitude determination method for a remote sensing micro nano satellite, and belongs to the technical field of spacecraft attitude control. The method comprises the steps that a magnetometer, a sun sensor and three star sensors serve as original attitude measurement data sources, according to a telemetry command received by the satellite and by combining withthe working state of the sensor, two attitude determination modes are designed, and the two modes are the sun pointing attitude determination mode and the remote sensing imaging posture determinationmode. For the sun pointing attitude determination mode, the magnetometer and the sun sensor serve as measurement input, in combination with the Calman filtering method, the sun pointing attitude determination mode has the advantages of being low in power consumption and high in reliability, and certain precision attitude determination can be provided to meet the control requirements of the micro nano satellite sailboard sun pointing. The remote sensing imaging attitude determination mode uses the three star sensors as the measurement input, in combination with the three-star sensitive attitudefusion method, high-precision micro nano satellite attitude and angular velocity determination can be achieved, and the high-precision attitude control requirements of micro nano satellite remote sensing imaging are met.

Belonging to inertial navigation calibration methods, the invention in particular relates to a calibration method for installation error angles between a dual-shaft rotation inertial navigation system's rotating shafts and sensitive shafts. According to the method, four installation error angles delta theta M1, delta gamma M1, delta psi M2, and delta gamma M2 are calibrated, and then compensation is conducted to complete demodulation work of dual-shaft rotation inertial navigation. The specific steps include: 1. calibration and compensation of the inner ring error angles delta theta M1 and delta gamma M1; and 2. calibration and compensation of the inner ring error angles delta psi M2 and delta gamma M2. With the invention, by calibration of the installation error angles between the dual-shaft rotation inertial navigation system's rotating shafts and sensitive shafts, navigation attitude demodulation error caused by the installation error angles can be eliminated, and the navigation attitude precision of the system is improved.

The invention provides a rotating member disturbance compensation method based on a momentum wheel. The rotating member disturbance compensation method includes (1) performing disturbance analysis to the rotating member; (2) determining disturbance compensation functions of the rotating member; (3) designing a realizing method of invoked disturbance compensation on the basis of time slice interruption; (4) performing fusion of a compensation torque and an attitude control torque; and (5) determining a manner of outputting control voltage to an I / O (input / output) port of the momentum wheel so as to realize the disturbance compensation to the rotating member. By the rotating member disturbance compensation method, attitude control accuracy is improved while fuel consumption of jet propulsion system compensation is reduced; since attitude control periods and control computer time slice interrupt laws are combined in the rotating member disturbance compensation method, hardware sources are fully utilized; and the problem of inconsistence of control periods and compensation periods is solved through a manner of combination of hardware and software, and compensation control in orbit in real time is realized.

The invention discloses a high-precision attitude data calculation method for an AHRS (Attitude and Heading Reference System), mainly solving the problems that as an accelerometer and a gyroscope are finite in precisions and have accumulative errors, the obtained attitude data errors are large and cannot meet technical requirements in the prior art. The high-precision attitude data calculation method comprises the following steps of initializing a sensor and a normalized coordinate system transfer matrix; calculating the attitude data and the coordinate system transfer matrix R1 of a carrier at the moment; calculating a coordinate system transfer matrix R2 at the moment; determining a final coordinate system transfer matrix at the moment, and normalizing the final coordinate system transfer matrix, so as to obtain the normalized coordinate system transfer matrix at the moment; solving the attitude data at the moment according to the normalized coordinate system transfer matrix obtained in the step (4); and returning to the step (2), and calculating the attitude data of the next moment. By utilizing the scheme, the high-precision attitude data calculation method is convenient to implement, is high in precision and has very high practical and promotional values.

The invention discloses a communication-in-motion antenna multi-platform heading and attitude determination method based on transfer alignment. According to the method, an inertial measurement unit isinstalled on a satellite communication-in-motion antenna, multi-platform navigation calculation is carried out, and error compensation is carried out on a strapdown navigation algorithm result of theinertial measurement unit by using output information of navigation equipment on an antenna installation carrier and a transfer alignment algorithm; and finally, a multi-platform heading and attitudedata validity evaluation method is used to obtain heading and attitude information beneficial to improving the tracking performance of the antenna. The multi-platform heading and attitude determination method is established by combining the azimuth angle and signal intensity information of the antenna, the tracking performance of the antenna is improved, and the method has the advantages of no need of manual operation, simple implementation and wide application range.

The invention discloses a method for correcting a long-period attitude error of a satellite based on ground control. The method mainly comprises three steps: step 1, calculating the initial values of the frequency and amplitude of the long-period attitude error based on the included angle of optical axes of star sensors; step 2, calculating an attitude correction amount by using a ground control point and an exterior orientation observed value; and step 3, fitting the long-period attitude error by using the initial values calculated in the step 1 and the attitude correction amount calculated in the step 2. The method can further improve attitude determination precision of star sensor-gyro combined treatment through detection and correction of the long-period attitude error of the satellite, which is beneficial for improving the global control-free mapping capacity of a mapping satellite.

The invention proposes a calibration method for an attitude angle error of a biaxial rotary inertial navigation system. By constructing a plurality of coordinate systems, an installation deviation angle, that is, a roll misalignment angle existing between the shaft system of the IMU mounted on the inner frame and an inner ring shaft and an outer ring shaft is calibrated, which reduces the influence of a cone error angle on the system attitude; meanwhile, the Kalman filter method is used to calibrate the non-orthogonal angle between the rotating axes and the swing angle of the shaft itself, andthe 16-position rotation method is used to obtain the observation information and navigation data in the process of Kalman filtering, which further improves the attitude accuracy of the system and the navigation accuracy of the system during long-term navigation.

The invention provides a method for restraining attitude errors of a modulation type strapdown inertial navigation system. The method comprises the following steps that: a rotary mechanism drives an inertial component to rotate to a position in which an inertial measurement unit (IMU) coordinate system is coincided with a carrier coordinate system; the inertial component performs single-shaft forward and backward rotation and stop motion; navigational parameters are designed according to a navigational parameter design principle under rotation modulation; the angular velocity and the linear acceleration of the motion of measurement carriers of an optical fiber gyro and a quartz accelerometer are acquired in real time; the control angular rate is corrected; carrier attitude information is obtained by performing navigation solution through using IMU measurement information and the corrected control angular rate; solved velocity errors are obtained by updating the carrier movement velocity; and the obtained carrier attitude information and velocity errors are used as navigation information which is finally output by the system. By adopting the method, oscillation errors related to the rotation modulation in the solved attitude information are eliminated, the navigation accuracy is improved, and the applicability of the solved attitude information of the system is enhanced.

The invention provides an SINS / DVL / ES combined navigation method based on information pre-evaluation and navigation result compensation correction. DVL data validity is judged before SINS / DVL combination, and when the DVL data information is invalid, DVL navigation information does not participate in combined navigation; at the same time, a residual error x2 testing method is introduced for evaluating whether combined navigation results are stable, and a compensation correction method is designed for real-time evaluation and correction compensation of Kalman output results when the combined navigation results are oscillated. The combined navigation method can effectively suppress the divergence speed of the SINS / DVL combined navigation attitude and speed, improves the combined system attitude and speed accuracy, improves the position accuracy in a certain extent, and also can solve the oscillation problem of attitude errors brought by declining of the DVL data information quality.

The invention relates to a attitude control feedback loop based on a combined fixed attitude of a multi-rate sensor, which comprises an attitude sensor subsystem, an attitude determining subsystem based on multi-scale modeling, and an attitude control subsystem, wherein the attitude sensor subsystem comprises a gyro, a solar sensor and a magnetometer and outputs attitude measurement information with multi-rate characteristics; the attitude determining subsystem is used for describing measurement information of each sensor from different scales by using a multi-scale analyzing method and estimating the satellite attitude angle on the basis of the established multi-scale model; and the attitude control subsystem is used for controlling the satellite attitude and feeding the controlled attitude information back to the attitude sensor subsystem so as to form the attitude control loop. The attitude control feedback loop is used for carrying out multi-scale modeling by using the multi-scaleanalyzing method, thereby improving the attitude determining precision while reducing the complexity of an attitude combining and fixing process.

The invention discloses a self-calibration and compensation method for a framework nonorthogonal angle of a triaxial rotating inertial navigation system. The method includes the following steps that the triaxial rotating inertial navigation system is powered on to be preheated, and coarse alignment is completed; the triaxial rotating inertial navigation system controls a framework to rotate according to a preset selection strategy and carries out navigation solution at the same time; posture output and speed output in the navigation process are recorded, and a recursive least-square method is used for calibrating the framework nonorthogonal angle of the triaxial rotating inertial navigation system; influence on the posture output of the triaxial rotating inertial navigation system is compensated through a compensation model. The calibration method for the framework nonorthogonal angle of the triaxial rotating inertial navigation system is simple in process, high in calibration precision and capable of achieving self-calibration of the framework nonorthogonal angle without machine disassembling and facilitates engineering implementation. The compensation algorithm of the framework nonorthogonal angle is easy to implement, and the precision of posture output of the triaxial rotating inertial navigation system is remarkably improved after compensation.

The invention relates to a UAV (unmanned aerial vehicle) group mounted polarization array attitude measurement and target detection method, and the invention reveals the change rule between the attitude position of a multi-loss electromagnetic vector sensor and the receiving of signals, and achieves the attitude measurement of a platform according to the signals of a plurality of earth satellite navigation systems which achieves the cooperative navigation or the spectrum of a UAV base station signal and maximization. A feature value algorithm is employed for calculating a general electromagnetic wave generalized spatial spectrum, and the polarization signal of a blind signal, the position of a plane, and the frequency shift information are obtained. The installation attitude and position information of a sensor in a plane body is maintained, and the method achieves the multi-target DOA under the conditions that the position of a UAV is not accurate and there is the frequency shift problem.

The invention discloses an attitude error compensation method of a laser gyroscope dual-axis rotary inertial navigation system. The attitude error compensation method is implemented by sequentially executing the following steps: constructing a six-degree-of-freedom kinetic model of the laser gyroscope dual-axis rotary inertial navigation system, obtaining a kinetic equation when an inner frame axis and an outer frame axis rotate, solving a deviation angle caused by the deformation of a rubber shock absorber when the inner / outer frame axis rotates by adopting a fourth-order Runge-Kutta method,and fitting the corresponding relationship between the angular rate of the frame axes obtained by dynamic analysis and the deviation angle caused by the deformation of the rubber shock absorber to obtain attitude error compensation information; and compensating the attitude output according to the attitude error compensation information. According to the attitude error compensation method, the attitude accuracy of the laser gyroscope dual-axis rotary inertial navigation system can be effectively improved; compared with the original uncompensated attitude output, the output accuracy of a pitching angle and a heading angle of the error-compensated attitude can be improved by about 8 times, and thus the effectiveness of the method can be fully proved.

The invention discloses an SDRE parameter adjustment method suitable for attitude maneuver and tracking control of wheel-controlled satellites, relates to an adjustment method of attitude maneuver and tracking control of wheel-controlled satellites, and aims at solving the problems that the prior art needs to ensure that the maximum torque computed by a control system does not exceed the maximum torque of a satellite execution mechanism, needs to ensure the attitude precision of the control system, cannot ensure the problem of emphasizing the different aspects of the control, cannot ensure the problem of emphasizing the different directions at different control stages and does not select the parameters of weighting matrixes as functions of states. The SDRE parameter adjustment method comprises the following steps: 1, abbreviating a satellite attitude dynamic equation and a satellite attitude kinematical equation; 2, writing the expression of an optimizing index J; 3, designing R(x) and Q(x); 4, solving uc; 5, computing Tc and sending the Tc to an execution mechanism to control the satellite attitude. The SDRE parameter adjustment method suitable for attitude maneuver and tracking control of wheel-controlled satellites is applied to the field of attitude maneuver and tracking control of satellites.

The present disclosure relates to a method (200) for estimating information related to a vehicle pitch and / or roll angle. The method comprises a step of obtaining (220) a first estimate of the information related to the pitch and / or roll angle. The method is characterized by the steps of capturing (210) an image of an area covering at least a part of the horizon using a camera mounted on the airborne vehicle, and determining (240) an improved estimate of the information related to the pitch and / or roll angle based on the first estimate of the information related to the pitch and / or roll angle, and a digital elevation model.

The invention discloses a satellite communication in motion attitude detection device and method. The device comprises satellite communication in motion, a microprocessor, an attitude detection moduleand a complementary filter, the attitude detection module comprises an X-axis attitude detection module, a Y-axis attitude detection module and a Z-axis attitude detection module; the method comprises the following steps of three-axis coordinate system building and attitude detection module layout, real-time information collection and uploading, satellite communication in motion three-axis weighted fusion angular velocity value obtaining, satellite communication in motion attitude angle obtaining and satellite communication in motion attitude angle correction. Accordingly, design is reasonable, by means of fusion of multiple micromechanical gyroscope homologous sensors and fusion of micromechanical gyroscope and accelerometer heterogenous sensors, the satellite communication in motion attitude can be accurately obtained, the cost is low, the reliability is high, the precision is high, and the practicability is high.

The invention discloses an integrated electric actuator for an unmanned aerial vehicle, which is specifically applied to the field of control surface control for unmanned aerial vehicles. According to the scheme, a brushless direct-current motor is used as a power source, and due to the scheme of serially connecting the brushless direct-current motor, a planetary reducer and a hollow angular displacement sensor, the actuator is compact in structure, and the spatial dimensions thereof are reduced, the working life thereof is prolonged, and the reliability thereof is improved; a shell of the actuator adopts a sealed structure design, and is high in dustproof and waterproof capacities, and adaptive to bad working environments; a control driver adopts the control strategy of a position and two-stage current limiting, a current-limiting ring is mainly used for limiting an excessive current occurring in a system under abnormal conditions, and protecting the safety of the electric actuator and the unmanned aerial vehicle; and with regard to a square-wave position angle instruction, real-time acquisition for the angle instruction is realized by means of interrupt capture, and the accuracy of acquisition for the angle instruction is improved by adopting the control strategy of angle amplitude-limiting filtering and instruction compensation, therefore, the accuracy of the flight attitude of the unmanned aerial vehicle can be greatly improved, and can achieve 0.1 degrees.

The invention relates to a dual-mode inertia / vision / astronomical navigation method for a lunar surface patrol detector, and the method comprises the steps: employing an inertia / vision / astronomical integrated navigation mode when the lunar surface patrol detector moves, and employing a starlight vector provided by astronomical navigation and the relative position and attitude errors provided by aninertia navigation system and a visual navigation system to correct the attitude error and the inertial device error of the lunar surface patrol detector. When the lunar surface patrol detector is static, the navigation system works in an inertia / astronomy combined navigation mode, and the position and attitude errors of the navigation system can be corrected according to starlight vectors and starlight heights provided by the astronomy navigation system and horizontal speed errors provided by the inertia navigation system. According to the invention, high-precision position, speed and attitude information can be provided for the lunar surface patrol detector, and the method has important practical significance for autonomous navigation of the deep space detector.

The invention relates to a star map recognition method, a star map recognition apparatus and a star map recognition system based on triangle matching. According to the present invention, part of starpoints are selected from bright stars, and are recognized, and re-selection is performed if the recognition fails, such that the interference caused by pseudo stars in the star map can be eliminated;the matching is performed with the multiple star points, such that the redundant matching probability can be substantially reduced in the case of the high error margin; the traditional absolute errorthreshold is changed into the relative error threshold, such that the small angular distance is corresponding to the small error margin so as to increase the calculation speed; and the primary recognition on the star points is completed by comparing the star number, and the primary recognition result is verified by comparing the theoretical angular distance value calculated with right ascension and declination to the actual measurement value, such that the number of the searching of the star number according to the angular distance and the number of the star number comparing are reduced, and the total recognition speed is improved.

The invention discloses an attitude determination method, a device, a computer readable storage medium and electronic equipment. The method comprises the steps of obtaining an initial pixel coordinateof a target object in an image; determining an attitude angle between the target object and an optical axis of the camera according to the initial pixel coordinates and internal parameters of the camera; obtaining corrected pixel coordinates corresponding to the target object according to the initial pixel coordinates, the attitude angle and the camera internal parameters; and determining the attitude of the target object according to the corrected pixel coordinates, a preset model and the attitude angle. According to the method, the initial pixel coordinates are corrected by utilizing the attitude angle between the target object and the optical axis of the camera, the corrected pixel coordinates are obtained, and the attitude of the target object is obtained according to the corrected pixel coordinates, so that the determined attitude of the target object is relatively high in accuracy.

The invention discloses a 6R industrial robot inverse kinematics solving method. The method comprises the following steps that a 6R robot geometric simplification model and a D-H table are established; according to the geometric simplification model, a function relational expression between a joint angle of a first joint and spatial coordinates of the origin of a tail end connecting rod coordinatesystem are established; under the condition that the joint angle of the first joint is determined, the spatial position relation of a second connecting rod and a third connecting rod is simplified into a triangle, and a function relational expression of joint angles of the second joint and the third joint is established; under the condition that the joint angles of the first joint, the second joint, the third joint and the fourth joint are determined, z-axis direction vectors of a fourth connecting rod coordinate system, a fifth connecting rod coordinate system and a sixth connecting rod coordinate system are used for establishing a space geometric model with the joint axis of the fifth joint as the center axis, and a joint angle function relational expression of the fifth joint is established firstly through the space geometric model; and then a function relational expression of the joint angles of the fourth joint and the sixth joint is established. According to the 6R industrial robot inverse kinematics solving method, the efficiency and attitude precision of the robot inverse solution method are improved.

The invention discloses an SLAM method based on a multi-fisheye camera and a double-pinhole projection model. The SLAM method comprises the following steps: (1) calibrating internal and external parameters of the camera; (2) inputting a plurality of fisheye images, if the system is not initialized yet, correcting the input images into appropriate double-pinhole perspective images through a double-pinhole projection model, and extracting features from the corrected images of each view; (3) converting map points to each view for feature tracking and matching in combination with external parameters of a camera; (4) enabling map points of the fixed local map to be subjected to multi-view attitude optimization, and updating the estimated main body attitude; (5) combining all poses and map points of the local map to carry out multi-view bundle adjustment (BA) optimization; and (6) adopting a bidirectional symmetrical query mode to ensure that loopback can be correctly detected when the samescene is passed forward and backward, and carrying out loopback correction after loopback detection succeeds.