541 results about "Star sensor" patented technology

The present invention discloses a high-precision satellite attitude determination method based on star sensor and gyroscope, which comprises the following steps: step 1. establishing a status equation of a satellite attitude determination system; step 2. establishing a measurement equation of the satellite attitude determination system; step 3. performing an online real-time model error estimation through predictive filtering; and step 4. performing a status estimation on a compensated model through 2-order interpolation filtering to obtain the attitude of a satellite. By applying predictive filtering to performing an online real-time model error estimation and correcting the system model, the invention overcomes the shortcoming existing in the conventional estimation process that error is processed into zero-mean white noise; and in addition, the invention can process any nonlinear system and noise conditions to obtain an estimation result of higher precision and is applicable to thefield of high-precision attitude determination.

A system for, and method of recovering a solar-powered spacecraft from an anomaly that renders the attitude of the spacecraft unknown includes maintaining a power-safe attitude by switching between two orthogonal axes using solar panel current sensors. The system and method may also include simultaneously determining spacecraft attitude using a star sensor. The system is applicable to spacecraft operating in a solar wing-stowed configuration.

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.

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.

The invention relates to an independent orbital transfer of a satellite, comprising (1) star capture step, in which before the satellite establishes orbit-controlled ignition attitude, inertia attitude of the satellite is pre-estimated and satellite attitude is controlled; zero offset of gyro drift and an accelerometer is marked; (2) inertia attitude regulation step, used for establishing the orbit-controlled ignition attitude of the satellite, realizing maneuver of the satellite attitude, pre-estimating the inertia attitude of the satellite and a jet control is used for controlling the satellite attitude; (3) star orientation step, in which after the satellite establishes the orbit-controlled ignition attitude, a stable state control is carried out and a star sensor is used for filtering and correcting the satellite attitude; and (4) orbit-controlled orientation step, in which starting and shutdown control is carried out for a rail-controlled engine; the ignition attitude of the satellite is determined and attitude stable control is carried out in the period of the ignition. The invention solves the problem that spacecraft such as a deep space exploitation satellite transfers orbit, and ensures that the orbital transfer can be completed accurately and reliably.

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.

A star sensor attitude-determination method during the autonomous orbital control fault recovery includes: (1) predicting the satellite inertia attitude according to data measured by a gyro; (2) calculating the filtering modified innovation amount according to the inertia attitude of the satellite and the optical axis vector and the lateral axis vector under an inertial coordinate system which are measured and output by the star sensor, and calculating the error of the innovation amount between the front and the back periodicals, which is used for judging the consistency of star sensor data; (3) judging the consistency of the star sensor data; (4) fixing the attitude of the star sensor according to the two vectors; (5) introducing the star sensor which is combined with the gyro for the correction of the satellite attitude under the condition that the star sensor data arranges the initial value of the attitude estimation. The method can improves the reliability of the orbital control fault recovery, saves the time for the fault recovery and ensures that the orbital control is accurately recovered in time.

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.

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.

The invention is a SINS/CNS/GPS combined-navigating half-entity simulating system, comprising SINS, CNS and GPS subsystems, combined navigating computer and track generator, using SINS and GPS entity systems, star map simulator and star sensor simulator entity as physical models, combined navigating computer, where the combined navigating computer adopts PC104 embedded microcomputer, the track generator software produces nominal track data, and dynamic characteristics of the system adopt digital model and it performs SINS/CNS/GPS combined navigating operation in the combined navigating computer to complete SINS/CNS/GPS combined navigation. And it effectively reduces testing cost and shortens development cycle.

A spacecraft attitude fixing method of DD2 (Divided Difierence 2) filter and a counting method based on Euler-q (Euler-quaternion) are provided, which relates to a state estimation and attitude determination method based on vector observation. The invention is characterized in that: according to the reference ephemeris and the reference vector and the observation vector observed by a star sensor, the spacecraft attitude quaternion can be gotten through Euler-q counting method; through DD2 filter, on one hand, feedback compensation the attitude matrix obtained through resolving by gyro; on the other hand, revise the palstance output of three-axle gyro and conduct compensation to the constant drift of gyroscope. As for that, the current spacecraft attitude information when getting high precision can be achieved, the constant drift of gyroscope can be scaled on line. The invention is an unattended attitude determining method, which is characterized in high precision and good timeliness, being convenient and feasible, and can be used in the attitude fixing system of various spacecrafts.

The invention discloses an attitude standard deviation estimation and correction method of a star sensor and a payload. The attitude standard deviation estimation and correction method comprises the following steps of 1, according to payload characteristics, determining a location formula of an incident light vector of a payload and building a payload error model with payload standard deviation, 2, according to star sensor characteristics, building a star sensor error model with star sensor standard deviation, 3, through the payload, determining measured deviation by observation data of a known target and corresponding star sensor measured data, and building a payload and star sensor standard deviation measurement equation according to the measured deviation, the payload standard deviation and the star sensor standard deviation, 4, carrying out estimation of the star sensor or payload standard deviation according to the payload and star sensor standard deviation measurement equation, and 5, correcting the star sensor or payload measured data by the estimated standard deviation. The attitude standard deviation estimation and correction method can reduce star sensor and imaging payload attitude standard deviation and improve imaging quality and an image positioning precision.

A star sensor includes (a) a scan mirror for scanning at least one star; (b) a detector array, coupled to the scan mirror, for detecting the one star; and (c) a processor, coupled to the detector array. The processor includes a first filter configured to reduce noise spikes in the detected one star, and provide a detection mask of filtered data. Also included is a second filter configured to reduce non-contiguous samples in the detection mask. A centroid calculator is included to determine a location of the one star, after the first and second filtering. The first filter includes a median filter, followed by an averaging filter, both configured to filter the one star in an along-scan direction of the scan mirror. The first filter includes another median filter, which is configured to filter the detected one star in the cross-scan direction of the scan mirror. An adder is included to subtract (a) output data from the other median filter from (b) output data from the averaging filter and provide filtered star data to the second filter.

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.

The invention discloses a star light refraction satellite autonomous navigation method based on a single star sensor. The star light refraction satellite autonomous navigation method comprises the following steps: 1, installing the star sensor on a satellite according to an optimal installation angle; 2, after the star sensor shoots a star map, identifying normal stars in the star map by using a triangle algorithm; 3, calculating an optical axis direction and a satellite gesture of the star sensor by using the identified star sensors; 4, selecting a star from the star map according to the optical axis direction of the star sensor to generate a stimulation refraction star map; 5, identifying a refraction star by using the stimulation refraction star map, and calculating a star light refraction angle according to an identification result; and 6, substituting the star light refraction angle into a system model, and obtaining navigation information of the satellite by using an optimum estimation method by a spaceborne computer. According to the star light refraction satellite autonomous navigation method, the precision of autonomous navigation of a star light refraction satellite is improved, and the design cost is lowered.

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.

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.

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.

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.

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.

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.

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.

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.

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.

The invention provides a ship's inertial navigation and astronomical positioning method based on attitude measurement, comprising the following steps of: (1) collecting the output data of an optical fiber gyroscope and a quartz flexible accelerometer after the initial alignment of an inertial navigation system is finished; (2) collecting the output of a CCD (Charge Coupled Device) star sensor, namely the attitude information of the coordinate system of the CCD star sensor relative to an inertial coordinate system i; (3) collecting the attitude matrixes continuously output by the inertial navigation system; (4) resolving the conversion matrix of an earth based coordinate system e relative to the system i; and (5) calculating out the position matrix through the information in steps (1), (2), (3) and (4), and calculating out the position information according to the position matrix. The method is an accumulation-free navigation positioning algorithm and has high positioning precision.

The invention discloses a method for improving data processing precision of star sensors. Star-sensor measurement errors are divided into three categories: A, low-frequency errors caused by star-sensor internal deformation, star-sensor installation position changes and other factors; B, intermediate-frequency errors caused by satellite-orbit parameter changes, star map transformation, jitter and vibration of satellites in orbit movement and other factors; C, high-frequency errors of star sensors, namely random errors obeying zero-mean Gaussian white noise distribution. After the star-sensor measurement errors are solved, the deviation of attitude estimation is basically eliminated, and attitude determination accuracy is increased by over 15 percent.

A calibration method based on star field for the star-sensitive sensor includes such steps as creating a posture conversion array of star-sensitive sensor, creating the distortion model of star-sensitive sensor, and eliminating the least square parameters.

The invention relates to strapping inertia/astronomical combination navigation semi-physical simulation system. It includes strapping inertia subsystem, astronomical subsystem, combination navigation terminal, trace generating terminal, demo verification and evaluation terminal. The trace generation terminal is used to form nominal trace data which introduces into astronomical subsystem and combination navigation terminal as information processing reference source. The astronomical subsystem is made up of ARM +liquid crystal light valve star map simulator and CMOS sensor +DSP star sensor. The combination navigation terminal is realized by DSP and used to receive the output data. The control terminal is used to realize demo verification and evaluation. The invention can effectively reduce experimentation cost and lead time. Thus, it has great theory and practice meanings for the combination navigation system dynamic performance study and engineering application.

A method of estimating the alignment of a star sensor (20) for a vehicle (12) includes generating star tracker data. A vehicle attitude and a star sensor attitude are determined in response to the star tracker data. A current alignment sample is generated in response to the vehicle attitude and the star sensor attitude. A current refined estimate alignment signal is generated in response to the current alignment sample and a previously refined estimate alignment signal via a vehicle on-board filter (38).

The invention discloses a layout design method of a multi-star sensor configuration layout. The layout design method comprises the steps of (1) defining minimum included angles between an optical axis of a star sensor and sunlight, ground gas light and a star object; (2) creating a layout design model; (3) creating a sunlight inhibition pyramid, a ground gas light inhibition pyramid and a star object inhibition pyramid of each star sensor in a satellite stereo model; (4) adjusting the layout of each star sensor in real time on the satellite model; (5) adjusting the included angle between the optical axes of two star sensors from 20 theta s to 180 degrees, so that the included angle is greater than twice of a sunlight inhibition angle; and (6) rotating the star sensors, so that the relative movement of fixed stars is uniformly distributed on two coordinate axes which are perpendicular to the optical axis of each star sensor. By using the layout design method, the best attitude measuring precision of multi-star sensor attitude combined determination can be obtained, and the layout design efficiency of multiple star sensors is improved.