1414 results about "Star sensor" patented technology

The invention relates to a multi-sensor autonomous navigation method for the low-orbiting satellite with fault-tolerance function and based on federated UKF algorithm, belonging to satellite autonomous navigation method. The method comprises the following steps of: constructing an orbital dynamics equation of earth satellite in a rectangular coordinate system; constructing a subsystem measurement equation with the output values of a star sensor and an infrared earth sensor as measurement quantities; constructing a subsystem measurement equation with the output values of magnetometer and a radar altimeter as measurement quantities; constructing a subsystem measurement equation with the output value of an ultraviolet sensor as measurement quantity; selecting a Sigma sampling point; constructing a predictive equation and an update equation of discrete UKF algorithm; respectively and independently performing Sigma sampling point calculation of each subsystem, and performing predictive update and measurement update; determining whether the output of each sub-filter is valid according to the predicted filter residual, isolating in case of malfunction, otherwise, inputting the filter result to a main filter for information fusion; constructing a non-reset federated UKF filter equation based on the UKF algorithm; and outputting earth satellite state estimated value X and variance matrix P thereof according to the steps.

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.

The invention discloses a SINS / CNS deep integrated navigation system and a realization method thereof. Wherein the navigation system comprises a strapdown inertial navigation system (SINS), a celestial navigation system (CNS), an integrated navigation filter, a inertial navigation posture measurement information structure unit; the realization unit comprises the following steps: 1. a large viewing field star sensor assists the strapdown inertial navigation system to obtain a high-precision mathematic horizontal reference; 2. CNS positioning is carried out based on the mathematic horizontal reference; 3. SINS / CNS deep integrated system model and measurement mode are established; 4. integrated navigation system information is fused; 5. the SINS and the CNS assists each other to realize high-precision positioning. In the invention, the star sensor high-precision posture information is employed to assist SINS to obtain high-precision SINS strapdown matrix which serves as the mathematic horizontal reference for CNS positioning, and on the basis, positions and postures of the CNS are employed to comprehensively calibrate the SINS, thus realizing SINS / CNS deep integration and finally achieving high-precision positioning and navigation.

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 relates to a method for calibrating a star sensor, which can effectively solve the problems of low calibrating speed, low efficiency and high hardware requirements of the star sensor. The method comprises the following steps of: calibrating internal parameters of the star sensor to obtain a coordinate of a mark point in a coordinate system of a photogrammetric system; carrying out image processing on a picture and calculating internal parameters of cameras; establishing a transit surveying coordinate system; carrying out measurement on a manual surveying mark of a calibrating field by using a transit surveying system to obtain a three-dimensional coordinate of the manual surveying mark in the coordinate system of the transit surveying system; carrying out common point conversion by utilizing a three-dimensional coordinate value to obtain a conversion relation of the coordinate system of the calibrating field and the coordinate system of the transit surveying system and converting the coordinate system of the transit surveying system and a cubic prism coordinate system into the coordinate system of the calibrating field; and taking a picture of the calibrating field again to obtain a relation of the two cameras and the cubic prism coordinate system. The method for calibrating the star sensor of the invention has the advantages of simple calibrating process, low required precision for the hardware, high calibrating precision and high speed, greatly improves the working efficiency and is an innovation on the calibration of the 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.

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 a celestial autonomous navigation method based on star sensors, which comprises the following steps: calculating attitude information based on a geocentric inertial coordinate system, which is output by a star sensor; calculating the optical axis direction based on the geocentric inertial coordinate system; converting the optical axis direction based on the geocentric inertial coordinate system into optical axis direction based on a WGS84 coordinate system; reading the included angles alpha 0 and beta 0 between the X and Y directions of the star sensor and the horizontal direction from a laser level meter; calculating the direction in the WGS84 coordinate system when the optical axis direction is perpendicular to the horizontal level; calculating the longitude alpha and latitude beta of the underground point S of the carrier; and outputting the attitude q and the longitude alpha and latitude beta of the underground point of the carrier in the geocentric inertial coordinate system. The invention avoids measurement and control errors caused by horizontal reference platforms, enhances the measuring accuracy, and simultaneously outputs the attitude of three axes and the longitude and latitude of the carrier in the geographic coordinate system in real time, thereby completely realizing celestial autonomous navigation.

A reflective star sensor light shield comprises three thin-wall cone barrels, wherein the three thin-wall cone barrels are successively connected. Inner walls of two cone barrels which are far from a lens are smooth clean surfaces and are sprayed with black paint or black films with high reflectivity and a low scattering rate. The inner wall of the cone barrel which is close to the lens is a baffle vane or a barrel type structure of eliminating veiling glare threads and is sprayed with black paint or black films with low reflectivity and the low scattering rate. The whole light shield can possess a optimal length through setting an algebraic relation of cone angles of each cone barrel. The reflective star sensor light shield of the invention has a short length and light weight, and is convenient for processing and manufacturing. The reflective star sensor light shield possesses strong inhibition capability for stray light, such as sunlight and ground gas light, which is a strong light source out of field of view. The invention can be used for the star sensor to shade light and be used for optical remote sensors to explore deep space, such as a star camera and the like.

The invention discloses a near-space missile-borne strap-down inertial navigation system transfer alignment method based on a star sensor. The method comprises the following steps: 1) establishing a missile-borne strap-down inertial navigation system transfer alignment state equation by taking an inertial coordinate system (launching point coordinate system for short) on a carrier launching point as a navigation coordinate system and a strap-down inertial navigation system (SINS) on a missile to be launched as sub-inertial navigation; 2) calculating navigation information and observed quantity of the missile-borne strap-down inertial navigation system; 3) establishing a measurement equation; 4) by depending on the state equation and the measurement equation established, estimating a mathematics platform misalignment angle, a speed error, a position error and an installation error of the missile as well as flexural deflection of the carrier through a sparse grid integral kalman filter, and correcting a sub-inertial navigation system, thus finishing a transfer alignment process.

The invention provides a simulator of a star sensor. The simulator comprises a collimating objective lens, wherein a focal plane of the collimating objective lens is provided with a star point target; the star point target is imaged at infinity through the collimating objective lens; a light splitting component is arranged between the collimating objective lens and the star point target; the focal plane of a light splitting optical path on the side face of the light splitting component is also provided with a detection target; the star point target is fixed on a focal plane adjusting mechanism; and a target lighting device is used for lighting the star point target. In the simulator, a technology which is different from the scheme of the conventional star simulator is adopted, so that the problem of poor accuracy of simulation of angular distance between stars of the conventional star simulator is solved; and the simulator can be used for extremely-high-accuracy simulation of a star and asteroid array of an asteroid autonomous navigation extremely-high-accuracy star sensor, verification of measurement accuracy and autonomous navigation and positioning accuracy of the star sensor, and has broad application prospect in research of a deep space autonomous navigation sensor.

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 a dynamic measuring device and method for plumb line deviation kept on the basis of an astronomical attitude reference. An INS / GPS attitude measurement subsystem and an LGU / GPS attitude measurement subsystem are built, wherein an initial value adopted for attitude updating by the LGU / GPS attitude measurement subsystem is provided by attitude information output by a star sensor. The difference between the attitude output by the LGU / GPS attitude measurement subsystem and the attitude output by the INS / GPS attitude measurement subsystem is obtained, and further the plumb line deviation is calculated. Finally, jump errors in the measuring values of the plumb line deviation are finally corrected, and low frequency errors in the measuring values of the plumb line deviation are corrected by utilizing global gravity model data. The systems in the invention are simple; the method provided by the invention is strong in robustness, and the measuring results are stable; requirements to GPS accuracy are lowered; compared with a traditional vector airborne gravitometer which relies on high precision difference GPS, the dynamic measuring device provided by the invention only requires common GPS single point positioning to meet accuracy requirements. Therefore, the application scope of the measuring method is expanded.

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.

The invention discloses a star sensor function test method based on electric injection star map, which comprises simulating the track parameter and attitude of satellite motion and simulating a star map shot by a star sensor, using the star point center to refresh and transmit part of the star map, laminating the part refreshes in the star map background of the star sensor, positioning star point, recognizing star map and calculating altitude, checking the results of the star map recognition and altitude calculation of the star sensor. The invention discloses a relative device. The invention has high refresh rate of star map, which can simulate the track altitude of star sensor and can process ground simulation test with whole star system, without additional external devices and special application condition demand, but with strong versatility.

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 invention provides an integrated navigation system of a strapdown inertial navigation system (SINS) / a central nervous system (CNS) / a global navigation satellite system (GNSS) of a geostationary earth orbit (GEO) transfer vehicle. The SINS serves as a core of the integrated navigation system. Navigation information of the GEO transfer vehicle is calculated and output by the SINS in real time, fault detection and isolation are conducted to data output by a GNSS receiver, a globe sensor and a star sensor by utilizing a residual error chi 2 detecting method improved by a kalman filter. Information fusion is conducted to output information of the globe sensor and the star sensor in a celestial navigation system, pseudorange measuring information output by the GNSS receiver and the navigation information output by the SINS. Navigation errors, inertial device errors, globe sensor errors and GNSS receiver time errors of the GEO transfer vehicle are evaluated in real time, and the navigation errors of the GEO transfer vehicle are corrected through a manner of closed loop feedback correction in real time so that in-orbit self navigation with high precision and high reliability of the GEO transfer vehicle is achieved, and therefore meaningful effects of good subjectivity, high precision, good robustness and high reliability are obtained.

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 relates to a star sensor installation error matrix and navigation system star-earth combined calibration and correction method, and aims at solving the problem that a conventional star sensor calibration method cannot preferably complete the calibration of an installation error matrix of a star sensor and cannot calibrate and correct the star sensor installation error matrix and navigation system deviation at regular intervals. The method is realized according to the following steps: (1) acquiring information theta<t,m> and X<t,m>; (2) establishing an attitude information and orbital parameter information measurement model; (3) determining values shown in the specification; (4) calculating values shown in the specification; (5) solving an arithmetic mean value; (6) establishing a practical attitude installation matrix and orbital parameter information correction model of the star sensor; (7) determining the direction of delta theta; (8) correcting the (6); (9) determining the attitude and the orbital parameter information; and (10) rerunning every N attitudes. The star sensor installation error matrix and navigation system star-earth combined calibration and correction method is applied to the field of satellite attitude determination technology and satellite navigation technology.

The invention discloses a method for dynamically measuring attitude of a star sensor based on top accurate angle relevance and belongs to the field of navigation and positioning. Based on the prior art that dynamic compensation is performed on each measurement exposure frame of the star sensor and a vector matrix matched with a fixed star containing dynamic error and noise effect is obtained, transformation matrixes between every two adjacent measurement frames of the star sensor are accurately measured through an assembly of three tops fixedly linked with the star sensor, the vector matrixes matched with adjacent measurement frames of the star sensor are correlated via the transformation matrixes, and finally, a correlated measurement equation is established by using a series of correlated measurement frames, that is a series of measurement frames are processed as a single measurement frame, and an attitude matrix is solved by using the least square method. With the adoption of the method, dynamic error and noise effect can be effectively restrained, high-accuracy dynamic attitude and change information of a carrier in motion can be obtained.

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 present invention belongs to the technical field of deep space navigation sensor, specifically discloses an imaging structure of a fixed star sensor, wherein a secondary aspheric surface is applied to an optical surface of an aberration correction lens, which not only can correct the aberration to make system imaging reach a level of near diffraction limit, but can also reduce difficulties ofsystem processing and resetting, and save cost; simultaneously the aberration correction lens can play a part in protecting a window, so as to prevent the system from the damage of space environment,such as irradiation, temperature difference, etc.