166results 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 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 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 discloses a star point motion parameter estimation method for a high-dynamic star sensor. The method specifically comprises the following steps: (1) acquiring a star map under a dynamic condition through a fixed star sensor; (2) carrying out filtering preprocessing on the fixed star sensor image obtained in the step (1); (3) carrying out threshold segmentation on the image processed in the step (2) and carrying out connected domain extraction to obtain rough star points; (4) selecting the brightest star point in the star points obtained in the step (3); (5) obtaining a sub-pixel star point track by using the brightest star point adjacent area obtained in the step (4) and adopting a Zernike moment sub-pixel edge extraction method; and (6) utilizing the sub-pixel star point trajectory obtained in the step (5), and adopting least square fitting to obtain star point motion parameters. The method has the beneficial effects that the precision of star point trailing length and angle estimation is improved, so that the extraction precision of star points in a dynamic environment is improved, and the precision of a fixed star sensor is improved.

The application discloses a kind of combined navigation system online calibration method, system, storage medium and terminal, comprising: establishing the state space model of combined navigation system online calibration, the state space model includes the space installation error of auxiliary sensor, the constant drift of gyroscope, scale factor error and the bias error of accelerometer, while introducing the reference information provided by multiple auxiliary sensors;Error correction is carried out to the calibration parameter and navigation parameter of the online calibration model;The output information of different navigation sensors is information fusion;The state space model is carried out observability analysis and design maneuver trajectory.The present application can solve the problem that part of calibration parameter cannot be estimated and the observability of IMU scale factor error is weak in the existing combined navigation system online calibration method, while improving the calibration efficiency and precision of combined navigation system.

The embodiment of the invention provides a data processing method and device and electronic equipment, and relates to the technical field of satellites.The method comprises the steps that star sensor data obtained by collecting spacecraft attitude information through a star sensor and gyroscope data obtained by collecting spacecraft attitude information through a gyroscope are obtained and comprise data from the k-n moment to the k moment; carrying out fusion and linear interference filtering processing on the data of the star sensor and the gyroscope at each moment by using Kalman filtering to obtain preliminary prediction data of attitude information, and calculating a first residual value of the data of the star sensor; the star sensor data, the gyroscope data and the first residual value from the k-n moment to the k moment are input into the trained LSTM network, a second residual value of the preliminary prediction data at the k moment is predicted, and the second residual value represents an error of nonlinear interference; and compensating the second residual value to the preliminary prediction data to obtain attitude information of the spacecraft at the moment k. According to the embodiment of the invention, the error of spacecraft attitude information acquisition is reduced.

A navigation method of a vehicle equipped with a star tracking device, a first inertial measurement device and a satellite positioning device, the star tracking device and the first inertial measurement device being connected to a single support. The method includes controlling movements of the support; simultaneously with these movements, calculating, in a horizontal plane, first positions of the vehicle, and second positions of the vehicle; calculating first deviations between the first positions and the second positions corresponding to each tracking heading and depicting them in a polar reference frame as a function of the corresponding heading and the values of the first deviations; carrying out a circular regression on said first deviations to determine, in this polar reference frame, a first circle; and determining a radius of the circle and a second deviation between a centre of the first circle and an origin of the polar reference frame.

The embodiment of the invention provides an interactive multi-model factor graph integrated navigation method based on tight coupling attitude constraint. In a specific embodiment, the method comprises the following steps: constructing an integrated navigation system; constructing an inertial measurement unit factor pre-integrated by the inertial navigation system; constructing a direct celestial navigation factor and a refraction celestial navigation factor which are respectively and tightly coupled with an inertial measurement unit factor based on the star sensor observation condition of the celestial navigation system; constructing a horizontal radio navigation factor and a vertical radio navigation factor based on the position information of the land-based radio navigation system; acquiring height measurement data of the barometric altimeter to construct barometric altimeter factors; dynamically adjusting the weight ratio of the height measurement data of the vertical radio navigation factor to the height measurement data of the barometric altimeter factor through an interactive multi-model algorithm; and carrying out global state estimation of the integrated navigation system on a factor graph framework formed by the factors.

A kind of star sensor star detection capability analysis method based on polarization degree image belongs to star sensor detection technical field.For solving the problem that less detection means and low utilization rate of sky light radiation information when near-earth space all-time star sensor observes star, the present application includes the star sensor of integrated focal plane micro-polaroid array is built, four gray star maps are collected by short exposure using the star sensor of integrated focal plane micro-polaroid array;Calculate the sky Stokes vector diagram;Calculate the polarization degree distribution of starry sky to obtain the polarization degree star map;Utilize the physical process of radiation energy to polarization degree conversion to construct the polarization signal-to-noise ratio calculation formula of star in polarization degree star map;The polarization signal-to-noise ratio of star in field of view is calculated;The polarization signal-to-noise ratio of star is compared with the signal-to-noise ratio threshold value of near-earth space all-time star sensor detection, and the detectable ability of star in polarization degree star map is judged.The present application improves the robustness of near-earth space all-time star sensor technology to complex environment.

Examples of the present disclosure include a method for approximating a non-carth imaging camera model for a non-earth sensor includes obtaining a physical camera model, receiving a target range value, determining a plurality of space-based coordinates of a plurality of points about a line of sight of the sensor in proximity to a distance corresponding to the received target range value, for each of the plurality of locations, determining line and sample coordinates on an image thereof via the physical camera model based on the space-based coordinates of the plurality of points, and based on the determined line and sample coordinates for each of the plurality of points, fitting an approximate camera model thereto.

This invention discloses a polarization-oriented light method based on Rayleigh point information, comprising: acquiring a sky polarization image and calculating the measured polarization angle of each measurement point in the sky polarization image; calculating the ideal polarization angle of each measurement point in the sky polarization image, and obtaining the polarization angle error of each measurement point based on the measured polarization angle and the ideal polarization angle; selecting Rayleigh points from all measurement points based on the polarization angle error, and performing polarization-oriented light orientation based on all Rayleigh points to obtain the heading angle of the carrier. This invention is applied to the field of polarization-oriented light navigation. By comparing the measured polarization angle with the theoretical polarization angle of each measurement point before polarization-oriented light orientation, and selecting Rayleigh points in the sky polarization image based on the polarization angle error between the measured and theoretical polarization angles, and then performing polarization-oriented light orientation only using the selected Rayleigh points, the accuracy of polarization-oriented light orientation is significantly improved.

The star sensor acquires a star point image through a detector, a data processor scans the obtained star point image using a 3*3 template; judges whether the gray value of the pixel point in the center of the template is greater than an image threshold T1 and whether the gray values of the 8 pixels around the pixel point are all less than the image threshold T1, and obtains a star point image after removing isolated points; the data processor divides the star point image after removing isolated points into blocks, divides the size of a sub-block into 32*32 pixels, and then calculates the image mean value in each sub-block; judges the mean value of each sub-block, and filters out the background and noise with high gray value; performs a pixel traversal operation on the star point image after removing isolated noise using the determined image threshold T1, and sends the pixels not less than the threshold T1 to a target extraction module. The star point image processing method under a complex background can judge and filter the star point image under a complex background, and can stably output the attitude information of the star sensor.

The invention discloses a method and a system for determining an optimal attitude of a high-dynamic star sensor, and belongs to the technical field of spaceflight measurement and control. Aiming at anisotropic distribution of centroid positioning errors caused by'trailing 'of a star point image under high dynamic condition, the method comprises the following steps: acquiring a trailing star point image and extracting an observation star vector; calculating a prior attitude matrix; analyzing the anisotropy error and trajectory orientation of the center of mass in the motion trajectory direction and the vertical direction; calculating an error component in an image plane coordinate axis direction; and constructing a weighted least square problem based on the error component, and updating a prior attitude to obtain an optimal attitude matrix. By accurately matching the weight of anisotropy error distribution, noise in a large error direction is effectively suppressed, and information in a high-confidence direction is fully utilized, so that the limitation of a traditional method is overcome, attitude estimation under a high-dynamic condition is ensured to meet the minimum variance unbiased estimation requirement, and the accuracy of attitude estimation is improved. And the precision and reliability of attitude determination are obviously improved.

The application provides a spacecraft autonomous navigation method and electronic equipment. The spacecraft autonomous navigation method comprises the following steps: determining technical indexes of an optical sensor according to navigation requirements, and determining a configuration of the optical sensor according to orbit information of the spacecraft; collecting a star map by using the optical sensor with the determined technical indexes and configuration, processing the star map, identifying a refraction star and extracting a refraction angle; and using a Kalman filtering algorithm to perform information fusion on information output by an inertial navigation system and a refraction apparent height calculated based on the refraction angle, so as to correct navigation errors and obtain corrected navigation information. The application corrects drift errors of the inertial navigation by astronomical navigation, maintains continuous navigation information by using the inertial navigation when the astronomical system is unavailable, and achieves the purpose of improving navigation accuracy.

The application discloses a kind of non-cooperative target centroid intelligent positioning method and system fusing multi-dimensional physical constraint.The method is directed to sparse, noisy point cloud data, and constructs a comprehensive evaluation function containing four-dimensional information of geometry, orbital dynamics, time sequence continuity and surface physical properties;Using a two-stage solution framework of surrogate model and hybrid optimization, first, a lightweight neural network surrogate model is used with a differential evolution algorithm for global coarse search, and then switching to a high-fidelity orbit model combined with an adaptive particle swarm optimization algorithm for local fine search;Finally, the optimal state estimation and its covariance matrix, evaluation decomposition and other decision support information are output.The application improves the accuracy, robustness and computational efficiency of centroid positioning, and enhances the interpretability of the results, suitable for on-orbit servicing, space debris removal and other high-risk space missions.

The invention discloses a field-of-view expansion type sun sensor based on diffraction order multiplexing, which comprises a dual-band band-pass optical filter, a diffraction type light introducer and a color CMOS (Complementary Metal Oxide Semiconductor) image sensor, and is characterized in that the dual-band band-pass optical filter is used for selecting two pieces of separated sunlight with a preset wave band for incidence; the diffractive light introducer diffracts transmitted sunlight into multi-level light beams; the color CMOS image sensor is used for acquiring a diffraction spot image containing spectral information of two wave bands; the current effective diffraction order can be determined through the mass center positions of the light spots in the two corresponding wave band channels and the mapping relation between the light spot distance and the diffraction order, and finally the sun vector is obtained through calculation in combination with a diffraction optical model. According to the invention, the inherent contradiction that a large view field and high precision of a traditional digital sun sensor cannot be considered is solved, the problem of level ambiguity in a diffraction extended view field is effectively overcome, the lightweight design of the sun sensor adapts to the high-density integration requirement, and technical support is provided for spacecraft attitude measurement.

This invention provides a method and system for correcting the polarization of fragmented sky light for UAV navigation, relating to the fields of UAV navigation and multi-sensor fusion technology. The method includes: during UAV flight, acquiring a raw polarization image containing sky regions in real time and performing distortion correction to obtain a distortion-free polarization image; using a lightweight sky segmentation model to identify the distortion-free polarization image and extract effective sky fragment regions; extracting the polarization information of the effective sky fragment regions based on the distortion-free polarization image covered by the effective sky fragment regions; calculating the absolute geographic heading angle of the UAV at the current moment using a fragmented polarization heading solution model; obtaining the predicted heading angle at the current moment and correcting it using an unscented Kalman filter method; calculating the deviation between the corrected heading angle and the desired heading angle to control the UAV to fly along a predetermined route. This invention effectively achieves stable and reliable control of the UAV's positioning trajectory.

The present application relates to the technical field of photoelectric detection and navigation, and discloses a solar vector orientation method, device and computer equipment based on thermal sensitive vision, to solve the problem of signal attenuation and low reliability of existing visible light or polarized light solar detection technology under complex weather. The method comprises: acquiring a thermal infrared image of a sky area by using a near-infrared thermal sensitive vision sensor; performing image preprocessing to enhance the solar target information; detecting and segmenting the solar target area in the image, and accurately positioning the solar center coordinates by using a centroid estimation or ellipse fitting method; converting the image coordinates into the solar position observation vector in the carrier coordinate system of the bearing sensor according to the sensor calibration parameters and installation parameters; and combining the theoretical solar vector calculated by an astronomical algorithm to solve the orientation information of the carrier. The present application is suitable for complex lighting conditions, has the advantages of strong anti-interference and high precision, and can be used for attitude determination in aerospace and marine navigation.

This invention discloses a method for orientation using atmospheric polarized light in the solar azimuth region, provides a visual modulation model of polarization angle in pixel coordinates, and presents an orientation strategy utilizing the solar azimuth region. This invention incorporates the horizontal attitude angle into the polarization angle in the pixel coordinates and achieves orientation using the separation point and the image principal point. Specifically, a seed line extraction algorithm is provided to update the solar azimuth in real time, fusing the morphological representation and polarization angle representation of the solar azimuth for optimal estimation. Finally, the horizontal attitude angle is input from an external sensor, such as inertial navigation, to complete the orientation. This invention can directly calculate the polarization angle in the pixel coordinates under tilted conditions, effectively improving the orientation accuracy and speed of atmospheric polarized light navigation.

The invention provides a spacecraft autonomous navigation method and electronic equipment. The autonomous navigation method for the spacecraft comprises the following steps: determining a technical index of an optical sensor according to a navigation requirement, and determining a configuration of the optical sensor according to orbit information of the spacecraft; acquiring a star map by using an optical sensor with determined technical indexes and configuration, processing the star map, identifying a refraction star and extracting a refraction angle; and carrying out information fusion on the information output by the inertial navigation system and the refraction apparent height calculated based on the refraction angle by using a Kalman filtering algorithm so as to correct the navigation error and obtain the corrected navigation information. The drifting error of the inertial navigation is corrected through the astronomical navigation, and the navigation information is kept continuous through the inertial navigation when the astronomical system is unavailable, so that the purpose of improving the navigation precision is achieved.

The application discloses a kind of star group between heterogeneous measurement data association and high-precision fusion orbit determination method, belong to space relative navigation technical field.The method includes grouping to the member star in star group, constructs the measurement set of the angle measurement and ranging of all member stars in time t, then calculates the position assumption set of member star, calculates the distance assumption value between member stars under each association assumption, constructs association degree evaluation index by distance assumption value and actual distance measurement, constructs cost matrix according to association degree evaluation index, and then obtains optimal association result, obtains the ranging and angle measurement data corresponding to all member stars at current time relative orbit determination, including using optimal estimation method to carry out state estimation, and including determining system state, constructing system measurement equation and system equation;Finally, the above steps are repeatedly executed for each measurement time, until filter convergence, obtain the high-precision orbit determination result of member star.

This invention relates to star sensors, specifically to lenses for shortwave imaging detectors in all-weather star sensors and their design methods. It addresses the problem that shortwave imaging detectors are not yet widely used in all-weather star sensors, resulting in a lack of suitable optical lenses. The invention comprises a first lens, a second lens, a third lens, a fourth lens, a fifth lens, a sixth lens group, and a seventh lens arranged sequentially along the optical path. The sixth lens group is a cemented lens, including a first cemented lens and a second cemented lens arranged sequentially along the optical path. This six-lens-plus-one-cemented-lens configuration results in low lens distortion, uniform dispersion under a telecentric optical path, which is beneficial for centroid extraction and improves the lens's response in the near-infrared band, enabling all-weather detection.

The application discloses a polarization differential reference-based navigation orientation method and device, equipment and a medium, which are applied to a polarization differential reference station and a user station, and the user station is located within the effective radius of the reference station. The polarization angle non-perpendicular error of each pixel point observation area is taken as a state vector to be estimated, a state equation of a polarization angle non-perpendicular error time domain interference observer is established; a polarization vector is constructed based on a polarization angle measurement value; a measurement equation and a system model of the interference observer are established; a corresponding relationship between the polarization angle non-perpendicular error to be compensated and a sensor observation vector is established; the polarization angle non-perpendicular error value is determined according to the corresponding relationship and a real-time sensor observation vector obtained through a user station navigation system; the polarization angle is compensated, and polarization heading calculation is carried out by using the compensated polarization angle, so that real-time heading information of a carrier is obtained. According to the embodiment of the application, the navigation precision and environmental adaptability of the user end combined navigation system can be improved.

This invention relates to the field of astronomical navigation technology, and in particular to a high-precision method for synchronizing satellite measurement integration with horizontal reference information. The method includes a calculation system and a calculation method. The calculation system comprises an astronomical navigation calculation unit, a satellite measurement sensor, and an inertial measurement unit. The inertial measurement unit is communicatively connected to both the astronomical navigation calculation unit and the satellite measurement sensor, and the satellite measurement sensor is also communicatively connected to the astronomical navigation calculation unit. This invention achieves high-precision navigation calculation through cross-clock domain synchronization and timestamp synchronization, and can be widely applied in astronomical navigation systems.

The invention relates to a star sensor, in particular to a lens for a short-wave imaging detector in an all-day star sensor and a design method of the lens, and solves the problem that the short-wave imaging detector is not popularized on the all-day star sensor at present, so that almost no applicable optical lens exists. The optical system comprises a first lens, a second lens, a third lens, a fourth lens, a fifth lens, a sixth lens group and a seventh lens which are sequentially arranged along an optical path, the sixth lens group is a glued lens and comprises a first glued lens and a second glued lens which are sequentially arranged along the optical path, the mode that the six lenses and one glued lens group are adopted, the lens distortion is small, and the imaging quality is high. The defocused spots are uniform in a quasi telecentric light path, so that the mass center extraction of the light spots is facilitated, the response of the lens in a near-infrared band is improved, and all-day detection can be carried out.

The application discloses a method for improving star detection capability by fusing energy and polarization information, and belongs to the technical field of star sensor detection. In order to solve the problem of insufficient star detection capability of a near-earth space all-time star sensor in a strong atmospheric background light interference environment, the star sensor integrated with a micro-polarization plate array is used to obtain four gray star images through one-time exposure of the sky; a polarization degree calculation value is calculated to obtain a polarization degree star image and determine a sky background polarization degree; a polarization degree segmentation threshold is set; the polarization degree star image is subjected to threshold segmentation to obtain a polarization degree star image segmentation result; a weight coefficient is determined according to the ratio of the radiant energy of the star and the sky background; the sky background polarization degree is compared, and the gray star image or the accumulated star image with the highest signal-to-noise ratio is subjected to weighted processing based on different conditions to obtain a final high signal-to-noise ratio star image after processing. The star detection capability is effectively improved.

This invention discloses a star map matching method based on the statistical features of star pairs, comprising the following steps: constructing star pairs; searching for similar stars from an image star pair list and a star catalog star pair list; determining the rotation angles of the image and the star catalog; initially determining matching parameters; and precisely determining the matching parameters. This invention's method can achieve a relatively high correct matching rate even with star position errors, false star images, missing star images, and high-density star fields.

This invention relates to the field of astronomical navigation technology, specifically a star identification method applied to deep-space, large-field-of-view, high-distortion images. The method includes: Step 1, applying precession and nutation corrections to a standard star catalog based on the image imaging time, removing noise and dark current from the image itself, calculating the brightness index of observed stars in the image, and selecting the brightest star closest to the center region as the reference star; Step 2, filtering surrounding neighboring stars based on the central bright star, constructing a brightness sequence ring model including star spacing and brightness order, and quickly matching to filter out possible corresponding candidate star pairs; Step 3, extending the candidate star pairs into the longest possible continuous matching sequence by gradually expanding the most matching star points, prioritizing the retention of the longest matching chain with smooth extension; then checking whether these sequences can close end-to-end. If they can close, they are considered a correct match; if they cannot close, the most reasonable matching result is selected based on the sequence length.