21 results about "Right ascension" patented technology

The invention discloses a strict regression frozen orbit design method. The method comprises the following steps: calculating an initial estimated value of a target strict regression frozen orbit mean element number based on an analytic formula of a regression orbit and a frozen orbit under a third-order gravity field; under the high-order gravity field dynamic model, iteratively correcting the orbit semi-major axis and the orbit inclination angle based on the right ascension change rate of the ascending node and the deviation value of the longitude of the revisit ascending node; taking the geocentric vector difference of the revisit ascending node as an optimization target, taking the eccentricity rate as an optimization variable, and adopting an optimization algorithm to obtain a current optimal frozen eccentricity rate; iteratively correcting the semi-major axis of the orbit and the inclination angle of the orbit, and optimizing the eccentricity ratio to obtain the orbital number of the strictly frozen regression orbit.

Disclosed are a method and an astrophotographic apparatus for acquiring images of targets in a sky area. The method includes: step 1, an image acquisition device is driven through a controlled rotatable component to point to the position near a target in a sky area, and images of the sky area are acquired through the image acquisition device; step 2, the images of the sky area are analyzed to obtain the right ascension coordinates and the declination coordinates of the center point of images for synchronizing the coordinates to the controlled rotatable component; step 3, the image acquisition device is driven by the controlled rotatable component to point to a target position of the target right ascension coordinates and the target declination coordinates corresponding to the target celestial body to acquire the images; and step 4, an image processing is performed on the images to obtain the processed images.

The present application relates to the technical field of altazimuth telescope, in particular to a pointing model calibration method of altazimuth telescope, the error function of the telescope is obtained by two-dimensional surface fitting, the star image is shot every 10 degrees or so in the azimuth and elevation directions, the right ascension and declination of the image center are obtained by comparing with the standard star map, the right ascension and declination are converted into the real azimuth and real elevation in the geodetic horizontal coordinate according to the shooting time, the geographical longitude and latitude of the telescope, the altitude, the air temperature and air pressure; the real azimuth and real elevation are subtracted by the azimuth and elevation read by the angle sensor of the telescope respectively, the azimuth error matrix and the elevation error matrix are obtained respectively; the error matrix is fitted to obtain the azimuth error function, the real azimuth and real elevation of any target are subtracted by the azimuth pointing error and the elevation pointing error respectively, and the accurate pointing of the telescope can be obtained.

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

The application discloses an astronomical cross-identification method for a heterogeneous many-core processor, and comprises the following steps: S1, obtaining the strip id of two star catalogs based on declination and the multi-node data distribution in strip units; S2, determining the number of heterogeneous many-core processors of each computing node and the maximum number of threads that can be simultaneously run by each heterogeneous many-core processor; S3, rearranging the star catalog data in each strip in strip units, and dividing the star catalog data into grids based on right ascension; S4, judging whether the right ascension span of each strip is too large during internal identification, so as to select the identification calculation of the CPU end in step S5 in a multi-thread parallel mode or the identification calculation on the heterogeneous many-core processor in step S6; S7, determining the identification selected result; and S8, obtaining the ID of a star body in two star catalogs respectively. The method can guarantee the calculation accuracy and integrity, improve the parallelism of the cross-identification algorithm and the use efficiency of the heterogeneous many-core processor, and realize the improvement of the overall calculation performance of the astronomical cross-identification.

The utility model relates to a kind of hybrid reduction drive equatorial telescope based on cross roller bearing, including shell, right ascension shaft hybrid reduction drive assembly, declination shaft reduction drive assembly, cross roller bearing, laser beam parallelism adjusting mechanism, pedestal;Right ascension shaft hybrid reduction drive assembly includes right ascension shaft motor, right ascension primary reducer and right ascension secondary reducer, and right ascension primary reducer is harmonic reducer, and right ascension secondary reducer is worm gear transmission pair;Declination shaft reduction drive assembly includes declination shaft motor, declination shaft reducer;The outer ring of cross roller bearing is fixed on rear shell, the end surface of inner ring is fixed with the output end of right ascension shaft hybrid reduction drive assembly, and the other end surface of inner ring is fixed with right ascension output seat;Declination shaft reduction drive assembly is fixed on right ascension output seat.The utility model, with high degree of freedom, small, light weight, the characteristics of strong carrying capacity, and the reduction ratio of this equatorial telescope is high, tracking is stable, tracking precision is high, and power consumption is low.

The invention provides an observation constellation configuration control method and device based on a maximum allowable drift domain, and belongs to the technical field of universe navigation. According to the method, based on the observation performance of a target constellation on the surface of a central celestial body, a critical equivalent view field is obtained through geometric equivalence, and a maximum allowable drift domain is obtained through analysis and derivation. And determining a first maximum permissible drift distance (namely, a phase maximum permissible drift distance) and a second maximum permissible drift distance (namely, an ascending node right ascension maximum permissible drift distance) from the maximum permissible drift domain in an inscribed interception mode, and taking the first maximum permissible drift distance and the second maximum permissible drift distance as target constellation configuration control threshold values, so as to realize configuration control on the target constellation. According to the method, the calculation amount required for determining the maximum allowable drift distance is reduced, so that the precision of the maximum allowable drift distance is improved under the same calculation power, and the configuration control effect of the target constellation is improved.

The invention provides an ultra-low orbit communication constellation design method based on a fixed sub-satellite point trajectory constellation, which comprises the following steps: (1) task demand description, namely designing an ultra-low orbit constellation, so that a target location T can realize continuous communication with duration of T every day; if the view field half cone angle of the satellite sensor is that the target location is in the sensor coverage range of any satellite, the communication is successful; (2) deriving a sub-satellite point regression condition; (3) acquiring the orbit semi-major axis and the inclination angle of the constellation, and further determining the satellite orbit height; (4) determining the constellation scale to ensure that the target location is within the sensor duration range of the satellite at any moment; and (5) determining the constellation phase, namely determining the right ascension and the latitude argument of the ascending node of each satellite. The method is reasonable in conception, can ensure that a target site can realize continuous uninterrupted communication with fixed duration every day, adopts a sub-satellite point sharing design to facilitate installation of ground equipment, and can flexibly adjust constellation configuration.

The application relates to a steady-state gain correction GNSS autonomous navigation system based on normalized orbit parameters, which comprises a GNSS information conversion module, an orbit extrapolation updating module, a steady-state gain Kalman filter correction module and a steady-state gain offline calculation module. The application adopts a normalized orbit semi-major axis, an orbit inclination, an ascending node right ascension, an eccentricity vector and a flat latitude amplitude angle as state variables of navigation filter estimation, adopts a Cowell method to perform one-step prediction of the above-mentioned orbit element information, converts position in a WGS84 system output by a GNSS receiver into orbit element information in combination with one-step predicted speed information, and adopts a steady-state gain Kalman filter algorithm to complete filter correction of the above-mentioned orbit element information.

The current disclosure provides systems and methods of directly measuring angular velocity and angular acceleration of space objects and using the measured angular velocity and angular acceleration as inputs into new and unique algorithms for initial orbit determination. Sensors measuring locations and times of light events may be used to generate a virtual rate track image for identification of space objects. Right ascension and declination of the space object events versus time may be fit to polynomials or splines to determine associated angle, angular rate, and angular acceleration of the space objects. New and unique initial orbit algorithms may then be applied to estimate orbital elements of the space objects.

The application discloses a spacecraft trajectory planning method for constellation multi-target traversal flyby, and belongs to the technical field of aerospace. The method comprises the following steps: introducing a flyby time constraint condition and an ideal flyby azimuth, constructing a task star flyby state optimization problem, and generating a relative motion trajectory for flying by a single target. A tangential velocity pulse is applied to an apogee of the task star to adjust a semi-major axis, and a difference in right ascension drift rate of the task star and target stars under J2 perturbation at different orbit semi-major axes is utilized to realize transfer of the task star among orbit planes of the target constellation. A correction maneuver is applied during orbit plane drift of the task star to realize matching of an amplitude angle of the task star apogee and a target star phase, so that the task star enters a resonance orbit for satellite traversal flyby in the constellation. The above-mentioned transfer trajectory optimization process of the task star among adjacent orbit planes is repeated to realize flyby traversal of a single task star to all satellites of the target constellation, and further realize spacecraft trajectory planning.

The application discloses a GNSS-based extrapolation method for a large-ellipse frozen orbit, which comprises the following steps: step one: acquiring the position and speed of a navigation receiver orbit according to the effective state of navigation data; step two: calculating satellite orbit elements (distance and speed values, moment of momentum, semi-major axis, orbital angular velocity, eccentricity, orbital inclination, ascending node right ascension, latitude amplitude angle, and perigee amplitude angle) by using t GNSS moment of momentum, semi-major axis, orbital angular velocity, eccentricity, orbital inclination, ascending node right ascension, latitude amplitude angle, and perigee amplitude angle); step three: judging the effectiveness of navigation data according to the calculation results of the satellite orbit elements; step four: extrapolating and calculating the orbit parameter true anomaly and latitude amplitude angle at the current satellite time; and step five: calculating the instantaneous orbital angular velocity, thereby completing the GNSS-based orbit extrapolation design for a large-ellipse frozen orbit. The application only uses the orbiting information of the GNSS time to extrapolate the orbit elements of the satellite, thereby achieving satellite autonomous positioning and autonomous attitude control, reducing the positioning requirements of the satellite to the ground station, and improving the autonomous ability of the satellite.

The invention relates to a right ascension axis visual polar axis calibration method based on laser beam indication and an equatorial telescope, and the method comprises the following steps: adjusting a laser beam through a laser beam parallelism adjusting mechanism, and enabling the laser beam to be parallel to the right ascension axis of the equatorial telescope; the direction of an optical imaging unit is adjusted, the optical imaging unit comprises an optical lens and an imaging camera used in cooperation with the optical lens, and the tail end of the laser beam is made to be located in the view field of the optical imaging unit; a latitude adjusting seat of the equatorial telescope is adjusted, so that the laser beam points to the vicinity of the polar axis; performing exposure by using an optical imaging unit to obtain recognizable star points; the polar axis position is positioned based on the relative position of the recognizable star points; and the latitude adjusting seat is adjusted, so that the imaging point of the tail end of the laser beam on the display screen coincides with the polar axis position. According to the invention, laser beam indication is combined with an optical imaging technology, accurate polar axis alignment of the equatorial telescope before use is realized, and the equatorial telescope is suitable for astronomical observation scenes of southern and northern hemispheres.

The invention provides a heterogeneous orbit constellation configuration regression period determination method and system, and the method comprises the steps: finding a group of coprime operation turns and regression number for each satellite, and enabling the operation turns and the regression number to meet the condition that the ratio of the operation turns to a first set parameter is equal to the ratio of the regression number to a second set parameter, the regression period of the satellite can be obtained; wherein the first set parameter is the sum of the average change rate of the perigee argument and the average change rate of the mean anomaly; the second set parameter is the difference between the earth rotation angular velocity and the right ascension precession rate of the ascending node; and calculating the least common multiple of the time periods of all satellites to obtain the regression period of the constellation. The method has the advantages that compared with a traditional method, the efficiency is improved by 13.3% when compatibility analysis is carried out on the regression cycle calculated through the method, and the maximum interference level in the regression cycle is consistent with that of annual simulation.

The present application relates to the technical field of spacecraft attitude orbit control, and particularly relates to a kind of coplanar installation electric thruster autonomous position keeping control method and device.Method includes: when receiving autonomous position keeping control instruction, based on current star time and pre-determined position keeping control period, the orbit inclination vector variation in position keeping control period is determined;Based on the orbit inclination vector variation in position keeping control period, the total firing duration in position keeping control period is determined to determine the firing duration corresponding to each operating day;Based on the orbit inclination vector variation in position keeping control period, the firing point right ascension is determined;Based on the firing duration corresponding to each operating day and the firing point right ascension, the coplanar installation electric thruster is used to autonomously carry out position keeping control in each operating day.This scheme can predict the position keeping parameters of multiple days, and the calculation is simple and easy to implement.

The invention discloses an astronomical telescope off-axis star guiding device with a focus reducer. The astronomical telescope off-axis star guiding device comprises a telescope, a reflector, a main detector system, a focus reducing lens group, a star guiding detector and the like. The off-axis large-view-field star guide of the telescope is realized, the star guide system shares the main light path of the telescope, the light collecting capability and the stability are higher than those of an external star guide mirror, and the focus compressor device improves the star guide detection view field and is beneficial to finding a proper reference star. The invention also discloses a method for calculating the offset of the guide star, the guide star transmits an image shot in a certain period of time to a program for processing, and the program calculates the WCS coordinate (world coordinate system) of the image in a star map matching manner, so as to extract right ascension and declination information of an image reference point. And the offset of the central point is obtained through comparison with historical data and is fed back to a telescope motion system for correction, so that long-time high-precision observation of the target celestial body is realized.

The application discloses a kind of space target observation right ascension and declination data three-dimensional simulation method and system, wherein the method comprises: adding observation equipment and observed target, mark J2000 position;Calculate the J2000 position of observation equipment and observed target in the visible time period;Convert the J2000 position of observed target in the visible time period to the required J2000 position;Convert the J2000 position data after conversion into right ascension and declination;According to the calculated right ascension and declination, set to the same height, show in three-dimensional situation.This application effectively avoids the defects of two-dimensional projection through three-dimensional visualization technology, can accurately present the motion trajectory and spatial relationship of space observed target, and provides reliable support for related research, task planning and the like.

The present disclosure provides a sun synchronous satellite orbit point root acquisition method and device, which relates to the technical field of spacecraft and can solve the problem that the satellite orbit point root cannot be acquired based on the launch site position and the descending node local time in the related art. The specific technical solution is: calculating initial orbit parameters according to the sun synchronous satellite epoch time, the preset orbit height and the preset descending node local time; calculating the time of the satellite orbit point; adjusting the ascending node right ascension according to the position of the launch site; adjusting the perigee amplitude angle; acquiring the calculated value of the descending node local time according to the time of the satellite orbit point, the adjusted ascending node right ascension, the adjusted perigee amplitude angle and other parameters in the initial orbit parameters; when the calculated value of the descending node local time matches the preset descending node local time, determining that the orbit parameters corresponding to the calculated value of the descending node local time are the sun synchronous satellite orbit point root; when they do not match, re-executing the above adjustment steps until they match.

This invention belongs to the field of remote sensing satellite ground data processing technology, and discloses a method and system for high-precision star centroid extraction and restoration in dynamic imaging mode. Based on the image-side coordinates of stars in the star image and the right ascension and declination of the stars, this invention calculates the satellite attitude and angular velocity, constructs a model of star centroid error and angular velocity, and determines whether the stars are in a low-dynamic or high-dynamic trailing state. For low-dynamic trailing star images, sub-pixel centroids are extracted using two-dimensional elliptical Gaussian fitting; for high-dynamic trailing star images, BM3D denoising, an improved EVSSM model restoration, and two-dimensional elliptical Gaussian fitting are used sequentially to extract the centroids. The extracted centroids are used for attitude calculation, and the centroid extraction accuracy is evaluated by the mean square error of the optical axis angle. This invention achieves full-process optimization from star image restoration to sub-pixel-level centroid extraction under dynamic imaging conditions through angular velocity threshold determination, noise suppression, model restoration, and fitting collaborative processing, improving the stability and reliability of attitude calculation.

The application discloses a kind of near-circular orbit space target initial orbit determination method based on single optical satellite, belong to space-based situation awareness technical field.The method includes: S1, using two groups of observation data successively, according to the right ascension and declination calculation space target true anomaly angle variation, based on circular orbit assumption through cosine theorem and search iteration estimation space target distance to geocenter within two groups of observation period;S2, repeat S1 for all observation data within period, using elliptical orbit equation and least square method calculates semi-major axis, eccentricity, and iteratively optimizes initial true anomaly;S3, according to semi-major axis, eccentricity and initial true anomaly calculation space target position and certain point velocity vector, utilize initial orbit normal vector and two-body orbit model restores initial orbit;If eccentricity is larger, then using Lambert specific orbit calculation velocity.The method can efficiently generate initial orbit using short arc angle data, avoid iteration not convergent, improve near-circular orbit space target orbit determination accuracy.

This invention discloses an optimization method for low-thrust transfer between circular orbits based on piecewise constant yaw angle guidance, belonging to the field of spacecraft orbit control technology. The implementation method of this invention is as follows: the low-thrust transfer process is divided into a transfer segment, a drift segment, and an adjustment segment; a piecewise constant yaw angle guidance strategy is proposed, using latitude argument angles α, β, γ as switching points, dividing each orbital period into four arc segments, maintaining a constant yaw angle and thrust switching state within each arc segment; based on the Gaussian perturbation equation, analytical solutions are derived for the long-term changes of the orbital semi-major axis, orbital inclination, and right ascension of the ascending node under this strategy and perturbation influence; the parameters of the optimal drift orbit are estimated based on the impulse assumption; a mixed-integer nonlinear programming model is established with the objective of minimizing the total velocity increment, and based on the dynamic relationship that "inclination changes are only generated by the thrust of arc segment 1," the model is transformed into several nonlinear programming sub-problems for solution, obtaining near-optimal guidance parameters. This invention avoids tedious numerical integration, significantly improves the efficiency of orbit design and optimization, and the resulting transfer scheme has fuel consumption close to the theoretical optimum. Moreover, the algorithm is robust and suitable for rapid orbit design tasks such as large-scale constellation deployment.