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541 results about "Star sensor" patented technology

Star sensor calibrator and method for calibrating high-precision star sensor

The invention discloses a star sensor calibrator and a method for calibrating a high-precision star sensor. The star sensor calibrator has the structure that a two-dimensional adjustable plane mirror is arranged on the optical path between a single-star simulator and a star sensor to be calibrated; the position forming an included angle of 90+/-15 degrees with the two-dimensional adjustable plane mirror is provided with a laser device of a laser angle measuring device; the position 50-200 cm away from the center of the two-dimensional adjustable plane mirror is provided with a high-precision two-dimensional guide rail which can vertically and horizontally move on an optical hover platform; a laser detector of the laser angle measuring device is installed on the high-precision two-dimensional guide rail; after being reflected by the two-dimensional adjustable plane mirror, the laser emitted by the laser device is incident into the laser detector; and a data processing computer is respectively communicated with the star sensor to be calibrated and the laser detector. The invention does not need to rotate the star sensor, increases the tangential distortion and the deflection angle of the installation error, and can meet the requirements for very-high-precision (sub-arcsec) star sensors. The invention has the advantages of simple method and small amount of calculation.
Owner:NAT UNIV OF DEFENSE TECH

Decoupling control method for relative orbits and attitudes of formation satellites

The invention discloses a decoupling control method for relative orbits and attitudes of formation satellites, relates to the technical field of the control of the orbits and attitudes of a spacecraft formation, and solves the problems of large satellite calculation amount and low orbit solving efficiency caused by high control dimension of the formation satellites due to serious coupling of the relative orbits and the attitudes of the formation satellites. The method gives two decoupling conditions at first, so that the control of the relative orbits and the attitudes can be designed independently; a thrust vector mobility decoupling constraint condition is introduced to the initialization control of the relative orbits according to satellite attitude mobility constraints indirectly; and during the optimal thrust vector attitude tracking, the possible orientation in space, which meets solar avoidance constraints, of a star sensor optical axis (1) is sought by using a geometric method, and the optimal attitude quaternion and attitude angular velocity are calculated finally by using a double-vector attitude determination algorithm. The method provides important reference value for the control of the orbits and the attitudes of the spacecraft formation.
Owner:HARBIN INST OF TECH

Correction method for on-track aberration of star sensor

The invention provides a correction method for on-track aberration of a star sensor. The method comprises the following steps of: calculating the annual aberration constant of a satellite according to a formula; measuring the linear velocity of the satellite in an inertial coordinate system by using a satellite borne device; calculating the altitude of the satellite in the inertial coordinate system; calculating the optical axis direction of the star sensor in the inertial coordinate system; calculating the linear velocity vertical to the optical axis direction of the star sensor in the inertial coordinate system; calculating the diurnal aberration constant vertical to the optical axis direction of the star sensor; calculating the included angle between a fixed star direction and the optical axis direction in the view filed of the star sensor; calculating aberration synthesis caused by all factors; and calculating an altitude quaternion. In the invention, a mathematical model for eliminating the diurnal aberration and annual aberration of the star sensor and the peculiar motion aberration of the sun is induced; after the model is used to eliminate aberration, altitude information with high accuracy can be further provided for an aircraft, the angular speed of a non-gyro aircraft can be calculated by using altitude, and the accuracy of angular speed calculation can be further improved.
Owner:HARBIN INST OF TECH

SINS/CNS integrated navigation system based on comprehensive optimal correction and navigation method thereof

The invention provides an SINS/CNS integrated navigation system based on comprehensive optimal correction and a navigation method thereof, and belongs to the technical field of integrated navigation. The integrated navigation system comprises an astronavigation subsystem, an inertia navigation subsystem and an information fusion subsystem. The navigation method comprises the following steps: analyzing celestial fix based on starlight refraction, building up navigation system state equations, building up navigation system measuring equations and performing information fusion of an integrated navigation system based on Kalman filtering. According to the invention, by utilizing the basic principle of starlight refraction indirection sensitive horizon and a large viewing field star sensor, the characteristics of a plurality of fixed stars can be observed at the same time, and the starlight refraction indirection sensitive horizon method is applied to aircrafts that do not fulfill track kinetics, so that the problem of high-precision autonomous horizon of the celestial navigation system is solved. According to the invention, position and posture information of the celestial navigation system are fully utilized to perform comprehensive optimal correction on the SINS deviation, so that the integrated navigation accuracy is significantly improved.
Owner:BEIHANG UNIV

Inertia/astronomy/satellite high-precision integrated navigation system and navigation method thereof

The invention provides an inertia/astronomy/satellite high-precision integrated navigation system and a navigation method thereof. The inertia/astronomy/satellite high-precision integrated navigation system comprises a sensor module, an integrated navigation resolving module and a real-time display module, wherein the sensor module is sequentially connected in series with the integrated navigation resolving module and the real-time display module and comprises an inertia sensor, a star sensor and a satellite receiver; the integrated navigation resolving module comprises a four-step Runge-Kutta strapdown navigation resolving module, an astronomy attitude determination resolving module, an unequal-interval Kalman filtering module and a position and speed compensation algorithm module; and the real-time display module comprises a three-dimensional real-time display module. The navigation method realizes the real-time acquisition of sensor data, utilizes the inertia sensor data to carry out strapdown navigation resolving based on a four-step Runge-Kutta algorithm and can carry out the integrated navigation resolving on the inertia sensor data, star sensor data and satellite receiver data and carry out the real-time three-dimensional display on an integrated navigation result by utilizing a projection algorithm.
Owner:NANJING UNIV OF AERONAUTICS & ASTRONAUTICS

Star sensor reference cube-prism installation error calibration apparatus

ActiveCN105318891AHigh precisionQuick installation error calibrationMeasurement devicesSingle starTheodolite
The present invention belongs to the technical field of optoelectronic equipment calibration, and particularly relates to a star sensor reference cube-prism installation error calibration apparatus. According to the present invention, a photoelectric autocollimator and a single star simulator are respectively placed on two orthogonal axes of a reference plane, a detected star senor is placed at the intersection point position of the two axes so as to make the normal lines of the two orthogonal light reflection surfaces of the detected star senor reference prism be respectively parallel to the two orthogonal axes, the optical axes of the photoelectric autocollimator and the single star simulator are respectively adjusted to parallel to the reference plane through a theodolite, the star sensor is installed on a star sensor three-dimensional adjustment reference base, the input optical axis of the star sensor and the output optical axis of the single star simulator are adjusted to achieve a parallel state through the star sensor three-dimensional adjustment reference base, a detected reference cube-prism is arranged on the upper surface of the housing of the detected star senor, the installation angle error of the reference cube-prism round the X-axis and the Y-axis is measured by using photoelectric autocollimation, the star sensor three-dimensional adjustment reference base rotates 90 DEG, and the installation angle error of the reference cube-prism round the Z-axis is measured.
Owner:BEIJING AEROSPACE INST FOR METROLOGY & MEASUREMENT TECH +1

Star map data based method for measurement of in-orbit precision of star sensor

A star map data based method for measurement of in-orbit precision of a star sensor. The single-frame static star map data obtained by the star sensor is utilized for measure the key performances such as in-orbit pointing accuracy and viewing angle of the star sensor; a quaternion estimation algorithm (QUEST), which converts the eigenvalue solution into the solution of a root of a fourth-order equation to speed up, is employed to obtain an optimal attitude matrix Aq; then according to the attitude matrix Aq and the identified n navigational stars, conducting back calculation to obtain the theoretical position of each navigational star on the image sensor, wherein the theoretical position is the theoretical point of intersection (xit, yit) of the primary optical axis of the star sensor and the interface; calculating the errors between the theoretical positions of the star points (xit, yit) and the measured actual positions of star points, wherein the actual star point position is the actual point of intersection (xi, yi) of the primary optical axis of the star sensor and the interface; and calculating a mean value of the errors and converting the mean value to an equivalent angle value, so as to obtain the precision of the star sensor. Compared to the commonly used in-orbit comparison method, the method provided by the invention eliminates the influence of the errors of time alignment and installation matrix between the star sensors, and reaches higher measurement accuracy.
Owner:BEIJING INST OF SPACECRAFT SYST ENG

Semi-physical simulation testing system and method for deep space autonomous navigation star sensor

ActiveCN102116641AVerify measurement accuracyNot affected by atmospheric environmentInstruments for comonautical navigationInformation processingPilot system
The invention provides a semi-physical simulation testing system for a deep space autonomous navigation star sensor and a method thereof. A star sensor simulator is placed on a similar support frame, and an autonomous navigation star sensor is fixed on a three-axis rate rotating table; the center of an entrance pupil of the autonomous navigation star sensor is arranged on a horizontal rotation central axis of the three-axis rate rotating table; an emergent pupil of the star sensor simulator is in butt joint with the entrance pupil of the autonomous navigation star sensor, and an optical axis of the star sensor simulator is superposed with the optical axis of the autonomous navigation star sensor; and the autonomous navigation star sensor, the three-axis rate rotating table and the star sensor simulator are all connected with a control and information processing computer. By adopting the system and the method, star simulation for the very high-precision star sensor for autonomous navigation of a minor planet can be solved, and the measurement precision of the star sensor and the autonomous navigation positioning precision are verified, thereby having broad application prospects for developing the deep space autonomous navigation sensors.
Owner:BEIJING INST OF CONTROL ENG

Integrated navigation system and method based on SINS (Strapdown Inertial Navigation System) and star sensor

The invention discloses an integrated navigation system and method based on an SINS and a star sensor. The integrated navigation system comprises the SINS, the star sensor and a filter, wherein the SINS is used for detecting the attitude information of a carrier, and amending the attitude information according to the optimum estimate of a state error term; the star sensor is used for acquiring the longitude and latitude, in a star sensor coordinate system, of an imaged fixed star, the direction unit vector, in a geocentric inertial coordinate system, of a reference fixed star matched with the imaged fixed star, and the longitude and latitude, in the star sensor coordinate system, of the reference fixed star; when the number of fixed stars observed by the star sensor is one or two, the filter is used for acquiring the optimum estimate of the state error term of the SINS according to an observation equation, wherein the observation equation is established by taking a longitude and latitude difference as a state quantity and by taking the pre-established error equation of the SINS as a state equation, and the longitude and latitude difference is composed of the longitude difference and latitude difference, in the star sensor coordinate system, between the reference fixed star and the imaged fixed star. According to the invention, the application range of the integrated navigation system can be widened.
Owner:HARBIN INST OF TECH

Measuring method for attitude angular velocity of spacecraft based on star sensor

The invention discloses a measuring method for an attitude angular velocity of a spacecraft based on a star sensor. The invention aims to solve the problem that the precision of a traditional method is seriously influenced by a measurement error of the star sensor. According to the technical scheme, the measuring method comprises the following steps of: reading a star map at an initial moment t0,thereby obtaining a measuring vector of the star sensor at t0 and navigating star information; supposing t=t0+deltat; reading the star map at the moment t; performing star point extraction and sequence star map identification, and numbering navigating stars simultaneously appearing in a previous frame map and a present frame map, thereby obtaining the total number n of the navigating stars simultaneously appearing at the moment of t minus deltat and t and a corresponding set omega1 of measuring vector pairs of the star sensor; initializing a Kalman filter, estimating the attitude angular velocity of the spacecraft and resetting an initial value of the filter; and supposing t=t+deltat, and repeating the steps of reading the star map at the moment t, performing star point extraction and sequence star map identification, numbering navigating stars and obtaining n and omega1. According to the measuring method, the influence of random noise of the measurement of the star sensor on the estimation of the attitude angular velocity can be eliminated and the measuring precision can be increased.
Owner:NAT UNIV OF DEFENSE TECH

DSP and FPGA parallel multi-mode star image processing method for star sensor

The invention discloses a DSP and FPGA parallel multi-mode star image processing method for a star sensor, the processing algorithm of a star image is classified and optimized according to different working modes of the star sensor, and star image processing can be completed jointly by the DSP and the FPGA. The method comprises the following steps: under a capture mode, the FPGA completes image filtering processing to obtain a quasi star point pixel, and the DSP performs capture mode processing on the quasi star point pixel; under non-high-dynamic tracking mode, the FPGA completes image window intercepting processing, and the DSP performs non-high-dynamic tracking mode processing on a window pixel; under a high-dynamic mode, the FPGA completes the image window intercepting processing while performing pixel binning, and the SP performs non-high-dynamic tracking mode processing on the window pixel; star image data preprocessed by the FPGA can be directly stored in the chip of the FPGA, and an external image memory is not needed. According to the method, the star image processing speed is effectively increased on the premise that the accuracy is not affected, and the system configuration is simplified.
Owner:BEIJING INST OF CONTROL ENG

Optical remote sensing satellite rigorous imaging geometrical model building method

The invention provides an optical remote sensing satellite rigorous imaging geometrical model building method. The optical remote sensing satellite rigorous imaging geometrical model building method comprises the following steps that the geometrical relationship between the image point coordinates of an optical remote sensing satellite and the satellite is determined according to design parameters and on-orbit calibration parameters of an optical remote sensing satellite camera and the installation relation of the camera and the satellite; the shooting position of a satellite image is determined according to the GPS carried by the optical remote sensing satellite, the observation data of a laser corner reflector and the installation relation of the laser corner reflector and the satellite; the shooting angle of the satellite image is determined according to a star sensor carried by the optical remote sensing satellite and the observation data of a gyroscope and the installation relation of the gyroscope and the optical remote sensing satellite, a collinearity equation of all image points of the optical remote sensing satellite is built, and a rigorous imaging geometrical model of optical remote sensing satellite images is formed. The optical remote sensing satellite rigorous imaging geometrical model building method is the basis of optical remote sensing satellite follow-up geometrical imaging processing and application.
Owner:SATELLITE SURVEYING & MAPPING APPL CENTSASMAC NAT ADMINISTATION OF SURVEYING MAPPING & GEOINFORMATION OF CHINANASG

Outfield precision testing method for high-precision star sensor

ActiveCN104280049AAchieving Frequency Domain StrippingAccurate calculationNavigation by astronomical meansStar patternObservational error
The invention discloses an outfield precision testing method for a high-precision star sensor. Accuracy in calculation of a star pattern posture of the star sensor and frequency domain stripping of short-period error terms of the posture are realized by utilizing a procession correction formula, an earth rotation model, an atmosphere correction model and a power spectral density formula, and the posture measuring error of the star sensor can be rapidly and effectively analyzed and evaluated. According to the method, the hardware resource is not occupied, the method is realized by utilizing software and does not need the ground intervention. According to the method, by combining the posture identification principle of the star sensor and adopting a data smoothing method, the posture truth value of each frame star pattern moment is calculated, so that a foundation is established for the subsequent analysis on the posture measuring error; by adopting the method for combining the time domain and the frequency domain, the short-period error term is decomposed into an airspace low-frequency error, a high-frequency error and a time-domain error by virtue of the power spectral density analysis, so that reasonable data support is provided for further evaluation of the precision indexes of the star sensor, and the index performance of the star sensor can be improved.
Owner:BEIJING INST OF CONTROL ENG

Self-adaptive controlled-array star sensor

InactiveCN101592490AImprove update rateHigh precision attitude measurementInstruments for comonautical navigationPicture interpretationSkySelf adaptive
The invention discloses a self-adaptive controlled-array star sensor, which comprises a plurality of observing view fields and a central signal processing unit comprising a central time sequence controller, a imaging driving unit, a mass center extracting unit, a start atlas recognition unit and an attitude calculation unit, wherein the central sequence controller performs the time sequence control of the observing view fields; the observing view fields shot the images of the star atlases according to the time sequence control and the imaging drive provided by the imaging driving unit; the mass center extracting unit extracts the mass center of star points according to the images of the start atlases; the star atlas recognition unit performs star atlas recognition according to the extracted mass centers of the star atlases; and the attitude calculation unit calculates the attitude of a spacecraft according to the star atlas recognition results. The sensor can improve the attitude measurement precision and data update rate in both a synchronous mode and an asynchronous mode. In addition, with the plurality of observing view fields, the sensor avoids attitude loss caused by influences such as uneven distribution of stars in sky coverage, has excellent dynamic performance, and can reduce weight and save power consumption and cost.
Owner:BEIHANG UNIV

Magnetic sensor and star sensor-based full attitude capture method and device thereof

The invention provides a magnetic sensor and star sensor-based full attitude capture method which comprises the steps: (1) obtaining an angular velocity of satellite attitudes through a gyro; (2) judging whether the angular velocity is within the range of a first threshold value, if not, carrying out rate damping through a magnetic torquer, so that the angular velocity is adjusted to be within the range of the first threshold value; (3) obtaining a sun vector by a magnetic sensor; (4) judging whether the inclined angle between the normal of a satellite sailboard and the sun vector is within the range of a second threshold value, if not, enabling the satellite to rotate by a momentum wheel, so that the inclined angle is adjusted to be within the range of the second threshold value; (5) measuring the earth magnetic field intensity vector by the magnetic sensor to obtain an attitude angle of the satellite relative to the earth; (6) judging whether the attitude angle is within the range of a third threshold value, if not, enabling the satellite to rotate by the momentum wheel, so that the attitude angle is adjusted to be within the range of the third threshold value; (7) obtaining an attitude quaternion of the satellite by a star sensor to determine attitude information of the satellite.
Owner:SHANGHAI ENG CENT FOR MICROSATELLITES
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