524results about "Navigation by astronomical means" patented technology

The present invention relates to a solar azimuth estimation method and system based on multi-channel feature enhancement and region-aware attention, belonging to the technical field of intelligent navigation for low-altitude economy unmanned systems. Aiming at the problem of decreased accuracy in solar azimuth estimation based on polarization images under complex cloud cover conditions, the present invention proposes a deep learning framework integrating multi-channel features and direction-aware attention. First, based on polarization light field information acquired by a division-of-focal-plane polarization camera, a three-channel composite input feature composed of a polarization intensity map, an adaptive threshold gradient map, and high-frequency residual edge information is constructed. Second, a ResNet backbone network embedded with a squeeze-and-excitation mechanism is adopted, and a direction-aware polarization attention module is introduced to achieve adaptive fusion of multi-scale features through luminance guidance, deep feature enhancement, and a gradient edge branch.

The invention discloses an agile satellite attitude fusion method and system considering multi-source reference difference, and belongs to the field of remote sensing satellite ground data processing, and the method comprises the steps: calculating a thermal environment factor based on satellite measurement data and satellite sensor installation parameters; extracting a space-time low-frequency error based on observation data of the star camera and the star sensor; performing normalization processing on the thermal environment factor, and constructing a relation model of the space-time low-frequency error and the thermal environment factor by adopting a multi-layer perceptron network to obtain compensated attitude data of the star sensor and the star camera; establishing a multi-source absolute attitude sensor data fusion model based on the compensated attitude data of the star sensor and the star camera; and carrying out absolute and relative attitude data fusion filtering processing on the processed gyroscope data and the fusion result of the star sensor and the star camera by adopting a bidirectional dynamic unscented Kalman filter to obtain a high-precision attitude data result. According to the method, the attitude determination precision of the agile satellite is improved, and the requirement of a high-precision remote sensing task is met.

It discloses an atmospheric polarized light orientation method using a solar azimuth region, provides an angle of polarization (AOP) vision-modulated model in a pixel frame, and gives an orientation strategy utilizing the solar azimuth region; a horizontal attitude angle is introduced into the AOP in the pixel frame, and orientation is implemented by utilizing a separation point and a principal point of photograph; specifically, a seed line extraction algorithm is given to update the solar azimuth in real time, a morphological expression and an AOP expression of the solar azimuth are fused for an optimum estimation to the solar azimuth, and ultimately the orientation is achieved by using an external sensor such as an inertial navigation system to input the horizontal attitude angle.

The invention relates to the field of data processing, in particular to an on-orbit calibration method and device for a scale factor, and the method comprises the steps: obtaining attitude data of a spacecraft after maneuvering through a star sensor, converting the attitude data into an angular acceleration signal under a test mass coordinate system, constructing a Kalman filter in combination with an excitation voltage read by an inertial sensor, and carrying out the on-orbit calibration of the scale factor; on-orbit calibration of the angle scale factor is completed, and on-orbit calibration of the line scale factor is completed through the conversion relation between the angle scale factor and the line scale factor. By combining and considering related factors such as linear acceleration and angular acceleration, the calibration process can more comprehensively consider the motion state of the spacecraft, and more accurate calibration is provided. In the maneuvering process of calibrating the scale factor, other parameters such as centroid offset and rigidity of the inertial sensor of the spacecraft can be calibrated at the same time. Through the technical scheme of the invention, the additional mass increase can be reduced, the unnecessary control complexity is avoided, the reliability of the system is improved, and the cost is reduced.

The invention provides a satellite inertial installation error calibration method and system based on iterative filtering, and aims to solve the problem that the attitude measurement precision of a star sensor is influenced by an installation error between inertial navigation and the star sensor. The method comprises the following steps: constructing a state space model and a measurement model of iterative extended Kalman filtering through a star vector under a star sensor body coordinate system, an ephemeris vector under an inertial coordinate system and inertial information, and deducing and calculating to obtain a measurement Jacobian matrix; and substituting the measured Jacobian matrix into the measurement model of the iterative extended Kalman filter to iteratively estimate the satellite inertial installation error, namely the installation error of the inertial astronomical integrated navigation system, so that the function of calibrating the installation error of the inertial astronomical integrated navigation system is realized. The method eliminates the influence of the installation error between the inertial navigation system and the star sensor on the attitude measurement of the star sensor, is suitable for the application occasion of combined navigation, and is good in adaptability and high in precision.

The invention relates to a sun meridian detection and course angle real-time calculation method based on polarized light field vector characteristics, and aims to solve the problems that a traditional sun meridian detection method is easily interfered in low signal-to-noise ratio scenes such as haze and cloud layer shielding, lens distortion causes polarization angle errors, the false detection rate is high, and the calculation accuracy is high. In order to solve the problems that the real-time performance and robustness of course calculation are affected due to the fact that single polarization angle information is difficult to fully reflect the sun direction, polarization parameter calculation is carried out by utilizing a multi-channel polarization image, and the sun meridian is detected by adopting Hough transform through fisheye lens correction, bilateral filtering and candidate region screening, so that the real-time performance and robustness of course calculation are improved. And a multi-point inversion method is combined to accurately obtain a sun vector and a zenith angle and an azimuth angle thereof, and finally a course angle is calculated in real time. The problems of imaging distortion and insufficient real-time processing in the traditional sun positioning technology can be effectively solved, and a powerful navigation positioning technical guarantee is provided for the fields of unmanned aerial vehicles, autonomous navigation, polarized light navigation and the like.

The present invention relates to a real-time heading angle calculation method based on the vector characteristics of the polarized light field. Aiming at the problems of traditional solar meridian detection methods, such as being easily disturbed in low signal-to-noise ratio scenarios like haze and cloud cover, the occurrence of polarization angle errors and high false detection rates caused by lens distortion, and the difficulty of using single polarization angle information to fully reflect the solar direction, which affects the real-time performance and robustness of heading calculation. This method calculates polarization parameters by using multi-channel polarization images.

The invention discloses a polarized light orientation error compensation method, which comprises the following steps: acquiring an atmospheric polarization image, and constructing a polarization vector non-orthogonal error angle model; solving the polarization vector non-orthogonal error angle model to obtain a non-orthogonal error angle of each observation point in the atmospheric polarization image; and calculating to obtain an E vector of each observation point based on the non-orthogonal error angle, and obtaining a course angle of the carrier. The method is applied to the field of polarized light navigation, the non-orthogonal error is quantitatively introduced into the construction process of the polarization vector non-orthogonal error angle model, and the polarization vector non-orthogonal error is compensated based on the prior information according to the characteristic that the non-orthogonal error angle changes slowly in a short time, so that the polarization vector non-orthogonal error is compensated. Therefore, the non-orthogonal error angle of the observation point in the atmospheric polarization image is estimated, the influence of the non-orthogonal error angle on the orientation result can be better compensated, and the method can be effectively suitable for cloudy weather orientation with the large non-orthogonal error angle.

Methods and systems are provided for determining navigational information for an entity such as a vehicle using celestial measurements. One method involves determining a Kalman gain based at least in part on inertial measurement data at or before a current point in time associated with the celestial measurement data, determining a celestial measurement error based at least in part on a relationship between the celestial measurement data and a prior estimate for the celestial measurement data, determining a current estimated error associated with the location of the entity based at least in part on a prior estimated error associated with the location, the Kalman gain and the celestial measurement error, and determining an updated estimate of the location of the entity based at least in part on the current estimated error and a prior estimate of the location of the entity.

The invention provides an integrated navigation method combining starlight refraction positioning with inertial navigation in a near space, and belongs to the technical field of starlight refraction navigation positioning. The method comprises the following steps: firstly, determining a state model of starlight refraction positioning in a near space by taking an error equation of an inertial navigation system of a near space aircraft as a state equation; designing a near space starlight refraction positioning algorithm, identifying a refraction star by using a star sensor on the aircraft, acquiring a refraction angle as a measurement, and establishing a measurement model of starlight refraction positioning in a near space; and obtaining the optimal estimation value of the system state variable by using an extended Kalman filter (EKF) method. The method has the advantages of being suitable for near space, high in reliability, high in autonomy and the like, and can be finally used for pure astronomical positioning and astronomical / inertial integrated navigation.

The invention discloses a lightweight high-precision spaceflight electronic system based on a self-calibration inertial measurement unit, and the system is characterized in that a comprehensive information processing module provides an operation environment for flight control, multi-information comprehensive processing, guidance, attitude control, radar height measurement and terrain matching, and is responsible for inertial measurement unit self-calibration, inertial measurement error compensation, signal collection and logic judgment, and external communication; and performing satellite navigation solution and Beidou short message communication. The three-self-inertial-unit module is used for measuring apparent acceleration and angular velocity in three directions of a coordinate system of a spaceflight electronic system; the double-shaft indexing mechanism controls the frame shaft system to drive the inner table body to rotate and position; the satellite guide radio frequency module is responsible for down-conversion, filtering amplification and anti-interference processing of radio frequency signals and outputting digital intermediate frequency signals; the power supply electric control module is responsible for secondary power supply conversion, primary power supply and distribution, ignition time sequence control, ignition circuit self-inspection and time sequence voltage telemetering acquisition; the electrical interconnection assembly is responsible for the electrical interconnection function in the aerospace electronic system.

The invention discloses a star point coordinate extraction and star sensor attitude optimization method based on deep learning. The whole process of pseudo star map generation, star point coordinate extraction and attitude calculation is optimized. Multi-factor interference such as spacecraft jitter and optical distortion is simulated through dynamic star point modeling, and a high-fidelity pseudo star map is generated; a noise adaptive detection network is designed to fuse denoising and key point detection, and star point positioning precision in a low signal-to-noise ratio environment is improved; and optimizing attitude solution, constructing a loss function based on star point coordinates, and performing back propagation to optimize attitude parameters to realize an algorithm closed loop. According to the method, real data can be expanded to realize effective training of the model, and the positioning precision is improved; the star point coordinate positioning precision is improved to a sub-pixel level (RMSElt; 0.3 pixel); an end-to-end optimization framework reduces error accumulation in a traditional assembly line, and the robustness of the system is improved by 40%.

The invention discloses an X-ray pulsar period and pulse time delay estimation method, and belongs to the technical field of pulsar navigation. The method comprises the following steps: through H test, according to a principle that an average integral square error between a Fourier series estimator and a real contour is minimum, searching an optimal harmonic wave truncation order under each observation; according to prior information of an observation event, adaptive selection of harmonic wave truncation orders under different data qualities is realized; re-estimating the phase of the packing data according to the equal photon intensity segmentation principle in the packing, so that the folded contour is closer to the real contour; and quantitative analysis is carried out in combination with simulation and actual measurement data, the optimal packing quantity is selected, and a folding contour with the optimal distinction degree is provided for statistical magnitude inspection. The method has the advantages of being suitable for a navigation pulsar source, high in reliability, high in environment adaptability, high in estimation precision, high in estimation stability and the like, and the X-ray pulsar navigation performance can be further improved.

The invention discloses a modular multi-size cubesat attitude control integrated system suitable for batch production, and relates to the technical field of satellites. Comprising a structural unit, an attitude monitoring subsystem, an attitude control subsystem, a data processing subsystem, a data transmission communication subsystem and a power supply subsystem, and the five subsystems are installed on the structural unit. The structural unit is a hollow hexahedron composed of six mounting plates; the attitude monitoring subsystem is electrically connected with the power supply subsystem and is also connected with the data processing subsystem through a signal line; the attitude control subsystem is electrically connected with the power supply subsystem and is also connected with the data processing subsystem through a signal line; the data processing subsystem is electrically connected with the power supply subsystem and is also connected with the data transmission communication subsystem and the power supply subsystem through signal lines; the data transmission communication subsystem is electrically connected with the power supply subsystem. The system has the advantages of being short in development and final assembly period, light in structure, unified in interface, multiple in function, good in expandability, capable of being used in a multi-size compatible mode and suitable for batch production.

The invention relates to the technical field of image processing and sensors, in particular to a star point detection algorithm based on an event camera, and the algorithm comprises the following steps: S1, carrying out the preprocessing and noise filtering of an event stream; carrying out noise filtering on an original event stream of an event camera, and removing noise events which do not conform to space-time correlation through space-time density analysis and adaptive time window adjustment; s2, star point centroid extraction and trajectory analysis; and carrying out time slice interception and clustering processing on the event stream after noise reduction, calculating a weighted centroid of a star point event cluster, and fitting a motion trail of a star point. High precision of star point detection is ensured through calculation and dynamic adjustment, and noise is removed by setting a threshold according to camera parameters; the angular velocity is utilized to determine time slices, and the centroid is calculated and the trajectory is fitted through clustering and weighting, so that the detection precision is far higher than that of a traditional method, and star points can be accurately identified in a complex environment.

The invention discloses a polarization vector non-orthogonal error identification method. The method comprises the following steps: acquiring an atmospheric polarization image; constructing a polarization vector non-orthogonal error angle model; and substituting the prior information and the observation information of each observation point in the atmospheric polarization image into the polarization vector non-orthogonal error angle model, and estimating to obtain a non-orthogonal error angle of each observation point in the atmospheric polarization image. The method is applied to the field of polarized light navigation, polarization vector non-orthogonal errors caused by multiple scattering in cloudy weather are quantitatively introduced into a model construction process, and the polarization vector non-orthogonal errors are estimated based on prior information and observation information of observation points according to the characteristic that a non-orthogonal error angle changes slowly in a short time. Therefore, the influence of the non-orthogonal error angle on the orientation result can be better compensated, and the method can be effectively applied to a polarized light orientation method in cloudy weather.

The invention provides a low-altitude airway flight height acquisition method and device, a storage medium and electronic equipment, and the method comprises the steps: determining the current air pressure height of an aircraft relative to a reference plane according to the current air pressure of the aircraft; according to the current geodetic height of the aircraft, the current satellite-based height of the aircraft relative to the datum plane is determined, and the current geodetic height is the height, measured through a satellite-based positioner at the current moment, between the aircraft and the geodetic ellipsoid; and performing fusion processing according to the current air pressure height and the current satellite-based height of the reference surface to determine the current fusion height of the aircraft relative to the reference surface. By uniformly setting the reference surface, acquiring the current air pressure height and the current satellite-based height of the aircraft relative to the reference surface, and performing fusion processing on the current air pressure height and the current satellite-based height, the height measurement error is effectively reduced, and the current fusion height with uniform measurement reference and higher accuracy is obtained.

The invention relates to the technical field of space target detection, in particular to a dynamic space target detection method based on event information, and the method comprises the steps: reading a target event flow according to a certain time length, and obtaining an initial pseudo image frame; eliminating noise events based on the corresponding event density information by using a pre-established circular local sliding window, and judging whether a certain noise reduction condition is met or not; if not, eliminating noise events based on the corresponding event density information by using the self-adaptive filtering characteristic of a pre-established multi-neighborhood radius local sliding window, and calculating the centroid information of the detection space target; judging whether the centroid information meets a certain precision condition or not; and if not, correcting the motion track by using the data of the inertial measurement unit to obtain centroid information meeting a certain precision condition, and obtaining motion information of the detection space target. Therefore, the problems of low space target extraction precision, difficult tracking and the like in a high-dynamic maneuvering scene are solved.

The present application provides an underwater point source polarization light detection device and a method for measuring a heading angle thereof, which are applied to the field of polarization navigation technology, wherein the polarization light detection device comprises a base, a rotating detection platform arranged on the base, and three groups of polarization light detection modules arranged on the top of the rotating detection platform, wherein the three groups of polarization light detection modules adopt the same structure and each group of detection modules is composed of a pair of polarization analysis channels whose transmission directions are perpendicular to each other, the other two groups of detection modules are positioned so that the polarization light detection directions of the two are left-right symmetrical about the zenith polarization light detection direction in the same plane and the angles between the two and the zenith polarization light detection direction do not exceed 5°, the rotating detection platform is configured to be able to rotate 360° around the axis where the zenith polarization light detection direction is located, and the heading angle measurement method using the detection device has the advantages of good adaptability to underwater environments and a small number and type of polarization analysis channels used.

The invention discloses a cage cable-based FAST cable-driven pulley high-altitude inspection and maintenance device, and the device is characterized in that a FAST telescope comprises a plurality of process systems, such as an active reflection surface system and a feed source supporting system; in the feed source supporting system, six steel cables are led out from six supporting towers to pull and control a feed source cabin to move on a focal plane, a plurality of pulleys used for hanging cables and optical cables are arranged on each steel cable in cooperation with astronomical observation, and the inspection and maintenance device comprises a cableway, a cable crane, a cableway hanging basket and a lifting pulley platform. The cableway is erected between the cable crane and the lifting pulley platform, and transverse movement of the cableway is achieved through walking of the cable crane. The cableway hanging basket is used for moving between the upper anchoring end and the lower anchoring end along the cableway to be close to the position of the cable driving tackle needing to be inspected and maintained, and then the cableway hanging basket is controlled by carrying maintenance personnel to vertically ascend and descend or arrives at the position of the tackle needing to be inspected and maintained along the cableway ascending / descending mechanism for inspection and maintenance.

The present application discloses an inertial orientation method for a navigation constellation based on stargazing measurement, the method comprising: obtaining constellation orbit parameters in an inertial coordinate system determined by intersatellite ranging and an inertial space pointing of an intersatellite relative vector in a star map background, calculating an intersatellite rotation matrix of a single observation baseline at any sampling time, and calculating intersatellite observation values ​​and relative vectors of intersatellite links obtained by a stargazing camera based on the inertial space pointing; constructing an intersatellite orientation observation equation according to the intersatellite observation values, the relative vectors of the intersatellite links and the rotation matrix, and constructing a navigation constellation orientation observation equation group based on the intersatellite orientation observation equation; correcting the rotation matrix based on the navigation constellation orientation observation equation group to obtain a corrected rotation matrix, and correcting the constellation orbit parameters based on the corrected rotation matrix to obtain the corrected constellation orbit parameters.

The invention belongs to the technical field of attitude change detection, and particularly relates to a real-time monitoring system and method for attitude change of a front-rear view surveying and mapping camera. The method comprises the following steps: firstly carrying out ground calibration, then determining a reference moment as a moment when the real-time monitoring system starts to start, then constructing a monitoring link at the reference moment, measuring a corresponding matrix obtained at a working moment, finally confirming the monitoring link at the working moment, and deducing a transformation matrix of the front-view camera and the rear-view camera under a J2000 coordinate system. According to the method, the real-time performance of the measurement process and the comprehensiveness of system monitoring can be considered while the geometric accuracy of the surveying and mapping camera is ensured.

The present invention discloses an optical navigation feature planning optimization method for autonomous operation on the surface of an extraterrestrial body, which belongs to the field of autonomous navigation. The implementation method of the present invention is as follows: at a camera sampling time interval, according to the planned detector trajectory and attitude, combined with the camera parameters, the visible field of view range of the optical camera at the corresponding moment is calculated; considering the uncertainty of the detector position and attitude, combined with the camera field of view range formula, the uncertainty area of ​​the field of view range is obtained, and the intersection area of ​​the uncertainty area is taken to obtain the estimated field of view range for navigation feature selection at the current moment; before the start of the detection task, the Cramér-Rao lower bound is used as the performance indicator to optimize the selection of optical navigation features within the sequence estimated field of view range, and obtain optical navigation features that meet the requirements of autonomous operation on the surface of an extraterrestrial body. According to the present invention, the preferred navigation feature points are planned for navigation, the influence of the uncertainty of the detector position and attitude on the navigation performance is eliminated, and the robustness and efficiency of autonomous operation navigation are improved.

Navigation with light refraction or dispersion and landmark data is provided. A system can include a data processing system. The data processing system can receive a first image of a surface of a planet from a first camera. The data processing system can generate a first position dataset based on the first image and data representing landmarks of the surface of the planet. The data processing system can receive, by a second camera oriented towards an atmosphere of the planet, a second image. The data processing can generate, via a celestial body catalog, a second position dataset based at least in part on an amount of refraction or dispersion of light of a celestial body in a second image. The data processing system can determine, based on a filter applied to the first position dataset and the second position dataset, a position and attitude of a vehicle.

The invention relates to a low-orbit star network-oriented random target tracking estimation method, which comprises the following steps of: establishing a dynamic filter AUKF of multiple targets in a geocentric inertial coordinate system by taking position, speed and thrust acceleration vectors of the targets j (j = 1, 2,..., Nt) in the geocentric inertial coordinate system as state vectors; establishing a maneuvering target space-based observation model based on the measurement equation of the target under the geocentric inertial coordinate system; based on the calculation result of the maneuvering target space-based observation model at the current moment, constructing a maneuvering starting and ending detector, when it is judged that the target starts maneuvering, using a maneuvering dynamics filter AUKF to track and estimate the position, speed and maneuvering acceleration information of the target, and when it is judged that the maneuvering of the target is ended, using the maneuvering dynamics filter AUKF to track and estimate the maneuvering speed of the target. And a standard Kalman filter UKF is switched to carry out state updating on the position and the speed of the target. According to the method, the relay observation and whole-course tracking performance estimation problems of the low-orbit star network on the uncertain maneuvering target are solved, and the method can be used for verification and evaluation of a large-scale star network system detection simulation test.

The invention discloses a near-earth space all-day star sensing system based on a sub-wavelength metal grating and a working method thereof, and belongs to the technical field of star sensors. In order to increase the number of fixed stars which can be detected by the near-earth space all-time star-sensitive system, the near-earth space all-time star-sensitive system comprises an optical system, an InGaAs detector, a spectrum optical filter, a micro-polaroid array, a memory, a processor, a control unit and a time unit, the InGaAs detector is respectively connected with the micro-polaroid array, the memory and the control unit; the micro-polaroid array is connected with the optical system, and the optical system is connected with the spectrum optical filter; the memory is respectively connected with the processor and the time unit; and the processor is connected with the control unit. According to the invention, the polarization filtering work of the sky background is realized, and the number of observable fixed stars is effectively increased on the basis of ensuring the miniaturization of a near-earth space all-day star-sensitive system.

The invention relates to the technical field of digital twinning, and particularly discloses a space debris collision detection digital twinning method and system, and the method comprises the steps: carrying out the grid processing of space debris through a space grid between the earth and the moon; performing space debris collision detection on the space debris by using a grid collision screening algorithm; and according to a space debris collision detection result, screening out space debris which may be collided, and sending out an early warning. According to the method, the space grid collision screening algorithm based on the space grid between the earth and the moon is adopted to carry out collision screening and secondary screening on the space debris, a large number of space debris which cannot be collided can be efficiently rejected, and therefore resources are concentrated to carry out more accurate analysis on high-risk collision.

The invention relates to an attitude prediction type star tracking method, and relates to the technical field of star sensors. The method comprises the following steps: obtaining historical angular velocity data; predicting the angular velocity of the next frame; predicting the attitude quaternion of the next frame; projecting a sky area fixed star to an image surface; performing window matching on the projection star and the observation star of the next frame; calculating the accurate attitude quaternion of the next frame; the target with the window matched successfully is used for attitude calculation, and the accurate attitude of the next frame is obtained; and repeating the steps to realize autonomous star tracking. According to the attitude prediction type star tracking method disclosed by the invention, the angular velocity of the next frame can be quickly and accurately predicted through polynomial fitting of historical angular velocity data, and the real-time performance and the accuracy are remarkably improved; the method can well adapt to high-angular-speed maneuver, and the tracking capability of the star sensor in a complex dynamic environment is enhanced; the method reduces the dependence on an external sensor, improves the independence and anti-interference capability of the system, simplifies the calculation process, and improves the overall calculation efficiency.

The invention provides an inertial navigation aided navigation positioning method and system, and the method comprises the steps: calculating the coordinate of each navigation satellite in an ECEF coordinate system according to the satellite ephemeris of a satellite navigation system; according to the navigation satellite coordinate and the inertial navigation system positioning result, the azimuth angle and the pitch angle of the navigation satellite relative to the optical detection equipment under the ENU coordinate system are solved, and the navigation satellite is tracked; acquiring star maps collected by the optical detection equipment, matching the star maps through a star map recognition algorithm, resolving a coordinate transformation matrix between two continuous frames of star maps, and recognizing fixed stars and satellites in the star maps based on the transformation matrix; converting the identified fixed star in the field of view into an ECEF coordinate system, calculating an observation angular distance between the fixed star and the satellite, and calculating a satellite observation vector in the ECEF coordinate system according to the observation angular distance; and performing error correction on the positioning result of the inertial navigation system based on the satellite observation vector. Through the scheme, the positioning precision of the inertial navigation system can be improved, and the influence of interference deception signals is reduced.

The invention provides a fast pulse period estimation method based on phase difference correction. The method comprises the following steps: step 1, determining a mapping relation between a folded waveform phase difference and a period estimation error; 2, folding the waveform in different time periods, calculating the phase difference of the folded waveform, and calculating a period estimation error according to the mapping relation; and step 3, setting a precision threshold value, and when the folded waveform phase difference meets the precision threshold value, calculating a pulse period estimation value according to the period estimation error. Compared with a traditional epoch folding period estimation method, the method has the advantages that the calculation speed is greatly increased, meanwhile, the precision is equivalent to that of the prior art, the method is not sensitive to observation duration, and the method is of great significance to on-orbit real-time pulse receiving period estimation.