55results about How to "Improve attitude measurement accuracy" patented technology

A miniature combined attitude measuring system for small satellite is composed of a combined inertial-magnetic measuring unit, a signal processing circuit and a satellite carried computer for storing data and program. Said measuring unit has 3 sensor sets based on microelectromachinal system. Each sensor set has microgyroscope, microaccelerometer and micromagnetometer. Its advantages are high real-time performance, high reliability and high anti-interference power.

The invention discloses a method for improving precision of a spaceborne linear array static infrared horizon sensor. The method comprises the following steps of initial judgment of an earth-space boundary position, selection of a reference signal, normalization of a detector signal, selection of a boundary judgment element, high-precision detection of the earth-space boundary position, and high-precision altitude calculation. Combined with the characteristics of the horizon sensor and based on the characteristic that a detection element signal corresponding to a transition zone changes regularly with the movement of the earth-space boundary in a field of view of the detector, the invention creatively provides a method for locating the earth-space boundary position highly precisely by the boundary judgment element to improve the measurement precision of the horizon sensor. By the invention, the problem that an instantaneous field of view angle of a detection element, the measurement precision and the altitude measurement range are restricted mutually in the design of the conventional horizon sensor is solved. The method has the characteristics of small calculation quantity and high generality, and is easy to implement, so the method can be applied to spaceborne linear array static infrared horizon sensors with different orbit heights.

The invention discloses a gyroscope assisted positioning method. The method comprises the following steps: step one, measurement of attitude angle: individually subjecting a measured value of an accelerometer and a measured value of a magnetic sensor to the Gauss-Newton iteration so as to obtain an attitude error speed, obtaining an attitude quaternion speed by using a gyroscope through a quaternion differential equation, adding the attitude error speed and the attitude quaternion speed together so as to obtain an attitude quaternion; step two, obtaining an initial point: carrying out track calculation according to the data of the accelerometer and the gyroscope chip so as to improve the track period to the data period of an inert measuring unit; step three, calculating the speed and the displacement distance; step four, calculating the position point, and step five, correcting the position. The method effectively solves the problems of gyroscope drift, carrier linear acceleration and interference of the surrounding local magnets, then carries out project correction on the positioning result according to the road data, and calculates the next position point by utilizing the correction result, and thus effectively reduces the accumulated error.

The invention discloses a capsule endoscope space attitude determination method which is characterized by comprising detecting the current co-acceleration vector of the capsule endoscope; detecting a magnetic field vector generated by an external preset directional magnetic field at a current position of the capsule endoscope; calculating the current attitude angle of the capsule endoscope by substituting the co-acceleration vector and the magnetic field vector into a preset formula. The present invention calculates the attitude angle of the capsule endoscope through the detected co-acceleration vector and magnetic field vector, wherein the magnetic field vector is generated by an externally preset horizontal directional magnetic field; the magnetic field intensity is high and the direction is fixed , which can avoid the measurement inaccuracy caused by strong magnetic interference when geomagnetic sensors are used to detect the magnetic field vector of geomagnetic field. The integral accumulated error caused by the angular velocity meter in the prior art is avoided and the attitude measurement accuracy is increased. The invention also discloses a capsule endoscope space attitude determination system and device which have beneficial effects as described above.

The invention belongs to the technical field of ship navigation guidance and control, and particularly relates to a ship horizontal attitude measurement method based on fusion complementary filtering and Kalman filtering. According to the invention, strapdown inertial navigation calculation is carried out by utilizing gyroscope output signals and accelerometer output signals of a micro-electromechanical inertial measurement unit, the signals are acquired in real time, and the advantages of each sensor are comprehensively utilized, so that high-precision horizontal attitude measurement of the system is realized. The linear acceleration and the Coriolis acceleration of the carrier in the maneuvering state are compensated through the navigation parameters subjected to Kalman filtering, a gyro integral error is compensated through complementary filtering, the horizontal attitude is kept at high-precision output, and even if the system has the motion linear acceleration, the attitude of the carrier can be accurately adjusted. The complementary filtering effect and the optimal calculation of the misalignment angle are still ensured, the system attitude measurement precision is effectively improved, and the method has a certain engineering application value.

The invention relates to a calibrating system and method of a multi-view-field star sensor based on a three-axis turntable. The calibrating system comprises the multi-view-field star sensor, the three-axis turntable, a single-star simulator and a data collecting device. The calibrating method comprises the following steps: step one, collecting data; step two, calibrating a pinhole model; step three, calibrating a complete model; step four, solving a structure parameter; and step five, evaluating a calibrating result. The calibrating method uses a generating angle of a rolling-rotating axis ofthe three-axis turntable to switch a data collecting view field, and uses the other two axes to generate a calibrating track for each view field. Data collection is completed by one time, so an outerparameter model in each view field calibrating model is kept unchanged, and the structure parameter between the view fields can be solved without being affected by an installing error. At the same time, according to an unchanged outer parameter, calibration of each view field can be joint-optimized, and the effect on parameter estimation precision caused by parameter coupling is reduced.

The invention discloses an inertia / odometer / RTK multi-information fusion method. Aiming at the problem of relatively poor observability of inertia / odometer combined heading, a positive and reverse RTSsmoothing method is used to perform weighting fusion on inertia / RTK data processing results, and an optimal heading estimation result obtained by fusion replaces a Kalman filtering estimation resultof inertia / odometer combined navigation at the corresponding moment to perform attitude correction, so that the heading precision of the inertia / odometer combined navigation can be improved and the position and attitude measurement precision of a combined navigation system is improved.

The invention discloses a roller shutter exposure star sensor star point position correction method based on average speed, which comprises the following steps of 1) according to two continuous framesof star maps (a qth frame and a (q + 1)th frame), obtaining the star point centroid position and the average speed of the star point moving along the horizontal and vertical directions in the (q + 1)th frame of star map, 2) taking the kth row of exposure time of the (q + 1)th frame of image as a reference, and obtaining the corrected star point position in the (q + 1)th frame of star map, 3) resolving the angular distance measured value between the star points after correction, comparing the angular distance measured value with the angular distance theoretical value under a celestial coordinate system, and verifying the correctness of the star point position after correction in the step 2) according to the angular distance error, and 4) according to the corrected star point centroid coordinates solved in the step 2) and the verification conclusion in the step 3), resolving to obtain the star sensor attitude corresponding to the kth row exposure moment of the (q + 1)th frame of image.According to the method, the position correction relation is established based on the line exposure moment and the star point average movement speed in the roller shutter exposure star map, and the attitude measurement precision of the star sensor in the roller shutter exposure mode under the star point movement condition is improved.

The invention discloses a star sensor demarcation method based on region division, comprising steps of (1) placing a star sensor on a two-shaft rotation table and performing imaging observation on a single-star simulator, (2) rotating the rotation table, performing measurement using the approximately uniform angle and choosing a standard grid point, (3) dividing the star sensor vision filed into n mutually independent correction areas according to the grid points, and generating a set of correction functions for each correction area, (4) choosing test points within the range not beyond the distribution boundary of the grid points, and (5) calculating the residual errors of all test points after demarcation. The invention can obtain high demarcation accuracy and is applicable to demarcation test and data processing of a large vision field optical sensor.

The invention relates to a method for improving the attitude measurement accuracy of a star sensor, belonging to the field of satellite analysis. The method comprises the following steps: establishinga satellite attitude determination system model so as to acquire the estimated state value of satellite attitude; subtracting an estimated gyroscopic drift value from the estimated state value so asto obtain a sample sequence; subjecting the sample sequence to discrete Fourier transform so as to obtain the frequency spectrum of the sample sequence and reading the frequency of the errors of the estimated gyroscopic drift value form the frequency spectrum of the sample sequence; estimating the amplitude of the periodic errors of gyroscopic drift so as to calculate the estimated amplitude valueof the low-frequency errors of the star sensor; establishing a reference model for the low-frequency error of the star sensor based on the frequency of the periodic errors of gyroscopic drift and theestimated amplitude value of the low-frequency errors of the star sensor; and updating the estimated state value of the satellite attitude according to the reference model for the low-frequency errors of the star sensor. The method provided by the invention fully takes the influence of the low-frequency errors of the star sensor on the attitude of the star sensor into consideration, so the attitude measurement accuracy of the star sensor is improved.

The invention provides an asynchronous multiframe star atlas fusion method of a star sensor. The method comprises the following steps: (1) exposure time of an image sensor of the star sensor is dynamically and periodically adjusted, in order to consecutively record star atlas data acquired by the star sensor in a period of time; (2) star point information in the acquired star atlas is extracted, positioning precision of all star points is determined, and star points with low positioning precision are removed; (3) relative displacement between star points in star atlas at different time is determined; (4) according to positioning precision of star points with higher positioning precision extracted in the step 2, and relative displacement obtained in the step 3, the information is converted to a star atlas directional coordinate system with the newest frame, and mean value of the same star point data in different star atlases with positioning precision as weight is taken; (5) in the star atlas directional coordinate system of the newest frame, an attitude determination algorithm is used, and high precision attitude measurement based on asynchronous multiframe star atlas fusion is completed. The measurement accuracy of the single probe star sensor is improved.

The invention discloses an area array static horizon sensor provided with an uncooled infrared focal plane, which is used for measuring the satellite attitude. The horizon sensor consists of four subsystems, namely a large-view field infrared optical system, a detector, a front electronic system and a rear DSP image information processing system. The product obtains infrared radiation images of carbon dioxide absorption bands on a radiation wave band between 13.5 and 16.25 micrometers of the earth by means of staring imaging, and calculates the rolling and pitching attitude angle information of a satellite relative to the local vertical of the earth through processing of the infrared images, wherein the measuring precision can reach 0.2 DEG. The area array static horizon sensor provided with the uncooled infrared focal plane has the characteristics of high precision, low power consumption, light weight, long service life and so on, and is a novel-generation infrared horizon sensor with innovation and wide application prospect.

The invention discloses a motion attitude sensing method based on electromagnetic wave polarization three-dimensional electric / magnetic signal stationary characteristics. The method comprises the steps of receiving electromagnetic wave signals through a sensor at a motion platform reference attitude, obtaining a signal cross correlation matrix, computing a CRB (Cramer-Rao bound) value estimated through attitude parameters, and judging whether the electromagnetic wave signals can be used for motion platform attitude measurement in the environment; computing the signal cross correlation matrix at the motion platform reference attitude, and judging whether the electromagnetic wave signals are stationary at the environment or not according to characteristic values of the reference attitude and a real-time attitude cross correlation matrix; converting the motion platform attitude under a real-time attitude coordinate system to be at a geographic coordinate system, and accomplishing motion platform attitude sensing. The method has the advantages that attitude / course sensing can be realized only by one receiving point on the motion platform, the system can be miniaturized, the measurement accuracy is high, the system can be miniaturized, signals are not interfered by multipath effect, the method can not only be used for actively transmitting electromagnetic wave signals, can also passively receive space distributed broadcast electromagnetic wave signals to measure attitude, and improves the system invisibility. The method is applied to aerospace vehicles.

The invention discloses a real-time posture measuring method for a motion capture system oriented to a four-rotor aircraft. According to the real-time posture measuring method, the motion capture system is utilized to obtain position information of five mark points mounted on the four-rotor aircraft, a body axis system is established by combining with kinetic characteristics of the four-rotor aircraft, and a posture can be resolved through transforming relationships between coordinate systems. Compared with a traditional MEMS (Micro-Electro-Mechanical System)-IMU posture measuring method, the method disclosed by the invention has the advantages of high posture measuring accuracy, and capabilities of achieving dynamic real-time posture measurement, providing correction information for an airborne low-cost MEMS inertial navigation system, and improving posture estimation accuracy, and has the characteristics of high accuracy, strong real timeliness, implementation easiness and the like.

A miniature combined attitude measuring system for small satellite is composed of a combined inertial-magnetic measuring unit, a signal processing circuit and a satellite carried computer for storing data and program. Said measuring unit has 3 sensor sets based on microelectromachinal system. Each sensor set has microgyroscope, microaccelerometer and micromagnetometer. Its advantages are high real-time performance, high reliability and high anti-interference power.

The invention discloses a method for temperature field measurement and directional temperature drift compensation of star sensors. The method comprises the following steps that 1, optical axis drift distances and hot defocus distances of multiple star sensors under distribution of a set working temperature field are calculated; 2, according to calculation results in step 1, the number and distribution of least temperature measurement points which can sense the optical axis drift distances and the hot defocus distances are determined; 3, multiple heaters are arranged in advance according to a model established in step 1 and a temperature measurement point layout determined in step 2, and the number and distribution of the temperature measurement points are calculated and analyzed; 4, a mathematic fitting formula of the optical axis drift distances and the hot defocus distances of multiple similar temperature field sample areas is obtained, and the optical axis drift distance and the hotdefocus distance of a current value of the changed temperature field are calculated; 5, through testing calibration of the temperature field, the mathematic fitting formula of the optical axis driftdistances and the hot defocus distances in step 4 is corrected; 6, compensation quantities of directional measurement data of optical axes of the star sensors are calculated. According to the method,the low-frequency errors caused by hot drift of the optical axes of the star sensors are reduced.

The invention discloses an MEMS gyroscope in-orbit constant drift calibration method and system for a microsatellite. The method comprises the following steps: determining a calibration mode of a gyroscope in an attitude determination system; determining a time interval of calibration and a group number k of data volumes of the star sensor and the MEMS gyroscope required by calibration according to a task condition; obtaining attitude quaternion and gyro data of a star body relative to an inertial system output by the star sensor at the current moment, judging the data validity, and if the data validity, adding 1 to the data group number; if not, abandoning the group of data, and waiting for the next group of data until the data volume meeting the condition reaches k groups; and carrying out calibration calculation based on the k groups of data to obtain constant drift of a measured angular velocity of the gyroscope under the relative inertial system of the satellite body, judging a calculation result, and if the calibration is successful, returning the calibration result to the attitude determination system for carrying out zero compensation on the gyroscope data.

The invention relates to a method for computing position deviation of pixels of a CCD (charge coupled device). The method includes generating a group of longitudinal dynamic interference fringes on the surface of the CCD, and exposing the CCD at fixed frame frequencies within a period of time to generate a plurality of groups of longitudinal dynamic interference fringe images; processing the longitudinal dynamic interference fringe images to obtain phases of output of the pixels at the same moment; computing phase difference between the pixels and first pixels at the same moment by the aid of combinations of the phases and movement directions of the longitudinal dynamic interference fringes; acquiring wave vectors of the longitudinal dynamic interference fringes and transverse position deviation of all the pixels by means of computation by the aid of combinations of the phase difference and the sizes of the pixels of the CCD; solving the longitudinal position deviation of all the pixels. The method has the advantages that dynamic interference fringe data are effectively utilized, accordingly, the phases of all the pixels at the same moment can be accurately computed, and the high-precision position deviation of the pixels can be obtained.

The invention relates to an optical system for a star sensor. The optical system is applied to high-precision star sensors. The optical system is of an improved peaceful structure and consists of six lenses, wherein the six lenses are all spherical lenses; an aperture diaphragm is arranged on first glass; the first glass is made of fused quartz, can be used for correcting aberration and can be used as protective glass of the optical system. The optical system has the characteristics of wide spectrum, large visual field, relative large aperture and the like; the distortion is very low in relatively large spectral range and visual fields, concentration ratios of discs of confusion of each visual field are uniformly distributed, and the imaging quality is good. The optical system can work in a temperature range from -40 DEG C to 50 DEG C, has good image quality and defocusing amount at each temperature point and can meet attitude measurement requirements of the high-precision star sensors working at severe space temperature environments.

The invention provides an inertial navigation system and a high-rotating-speed rotating body attitude measurement method. The method comprises the following steps: measuring normal inertia informationand geomagnetic information of a rotating body in a present resolving period; and establishing a geomagnetic attitude calculation coordinate system, performing attitude calculation in the geomagneticattitude calculation coordinate system by utilizing the geomagnetic information and the normal inertial information, and performing coordinate transformation on the calculated attitude to obtain attitude information of a projectile body coordinate system relative to the geographic coordinate system. According to the invention, the sensor cost can be reduced, and the problems of insufficient rollangle range and rapid divergence of measurement precision along with time in the existing inertia / terrestrial magnetism combined attitude measurement system are solved.

The invention discloses a measuring method for disturbance acceleration in attitude estimation of an aircraft. The method comprises the following steps: acquiring the attitude quaternion of the aircraft and acquiring measured acceleration in a body-axis coordinate system; acquiring disturbance acceleration in a reference coordinate system according to the attitude quaternion and the measured acceleration in the body-axis coordinate system and then acquiring discrete disturbance acceleration; and subjecting the discrete disturbance acceleration to segmented processing, then carrying out first-order low-pass filtering, and then transferring the processed discrete disturbance acceleration to the body-axis coordinate system for subtraction by the measured acceleration in the body-axis coordinate system so as to obtain accurate aircraft acceleration without the disturbance acceleration. Compared with the prior art, the measuring method provided by the invention is low in cost, needs no GPS sensor for measurement and can improve the attitude measurement precision of subsequent multi-sensor fusion technology.

The invention discloses a double-refraction sun sensor. The double-refraction sun sensor comprises a light filtering module, a single-axis crystal camera lens, an image sensor and a data processing module, and is characterized in that the light filtering module is used for bundling light to form relatively fine incident light; the single-axis crystal camera lens is used for carrying out double refraction on the incident light to form two beams of refraction light; the image sensor is used for imaging the two beams of refraction light; the data processing module is used for extracting an image point center, calculating vector information of the incident light according to the two beams of refraction light and calculating three-axis attitude information of a carrier. When the carrier is equipped with the double-refraction sun sensor, the invention further discloses a measurement method of carrier three-axis attitudes; three-axis attitude measurement can be realized by utilizing a spacecraft / unmanned aerial vehicle through independent sun measurement. The double-refraction sun sensor disclosed by the invention can realize a relatively large view field and can guarantees that attitude measurement precision is not lost.

The invention provides miniaturized radar dynamic attitude measurement equipment, which comprises a box body 1, and a fiber-optic gyroscope combination 2, an accelerometer 3, an IF plate 4, a main control board 5 and a power supply module 6 which are arranged in the box body 1, wherein the fiber-optic gyroscope combination 2 is used for measuring triaxial angular rate data in real time; the accelerometer 3 is used for measuring acceleration data in a carrier movement process, converting an accelerometer signal into a current signal, outputting the current signal to the IF board 4, converting the current signal output by the accelerometer into a pulse signal, and outputting the pulse signal to the main control board 5; the main control board 5 is used for receiving and collecting triaxial angular rate data measured by the fiber-optic gyroscope combination 2 and acceleration data measured by the accelerometer 3, and carrying out dynamic attitude measurement; the power supply module 6 is used for converting an input direct-current power supply into a power supply required by each component in the product; and the main control board 5 is specifically used for carrying out attitude correction on triaxial angular rate data measured by the fiber-optic gyroscope combination 2 and acceleration data measured by the accelerometer 3 through a complementary filter to obtain attitude information.

The invention relates to a method for improving attitude measurement precision of a multi-probe star sensor, and belongs to the technical field of satellite attitude control. The method comprises thefollowing steps of deducing a relationship between the position of a star point in the field of view and the installation array variation between probes, and establishing a measurement model and a state model based on the star point position as the observed quantity; and analyzing star point error characteristics in the field of view of the product by a ground test result, and correcting filter estimation parameters based on star point noise characteristics, and the product installation array estimation accuracy is improved. The method can further improve the product attitude measurement precision, and according to a ground simulation and test result, the method can improve the product attitude measurement precision from 1" (3sigma) to 0.6" (3sigma).

The invention provides a high-precision carrier dynamic attitude measurement method and system, and the system comprises two micro-electro-mechanical inertial measurement units, one unit employs a double-position reciprocating transposition mode, and the other unit employs a mode of being fixedly connected with a carrier. Relative attitude, position, speed and the like between the two micro-electro-mechanical inertial measurement units are introduced as new observed quantities, and zero offset of gyroscopes and accelerometers of the two micro-electro-mechanical inertial measurement units is effectively estimated, so that the two sets of systems can be better compensated. The micro-electromechanical inertial measurement unit fixedly connected with the carrier is used as attitude output inertial navigation, so that the phenomenon of peak jump of an attitude calculation result caused by reciprocating rotation of the first micro-electromechanical inertial measurement unit is successfully eliminated, and the precision of the whole system is improved. In the using process, accurate position information does not need to be introduced, so that satellite signals do not need to be introduced, and the system has better anti-interference performance, stability and concealment.

The invention discloses a self-adaptive horizontal attitude measurement method based on motion state monitoring, which comprises the following steps of: based on a newly established state space model, taking a horizontal attitude angle as a state variable, taking an angular velocity increment delta omega b for compensating random constant zero offset as control input of a system equation, and taking a specific force fb for compensating random constant zero offset as quantity measurement; meanwhile, the judgment condition of the maneuvering state of the carrier is improved, and the maneuvering information of the carrier is judged by comprehensively utilizing the acceleration information and the angular velocity information output by the MEMS-IMU, so that the measurement noise matrix of the filter is automatically adjusted, and the influence of the maneuvering state of the carrier on the calculation of the horizontal attitude is effectively reduced. The method has no special requirement on the maneuvering state of the carrier, and can ensure that the system has high attitude measurement precision in different motion states without external information assistance.

The invention discloses a star map joint non-uniformity correction method and device and a storage medium, and the method comprises the steps: receiving a plurality of calibration images which are collected under the condition of no stray light interference; calculating a correction parameter according to the plurality of calibration images; receiving a star map to be corrected; processing the correction star map according to the correction parameters to obtain a two-point correction star map; introducing a sobel edge detection operator, and constructing a sobel weight factor; substituting thesobel weight factor into the guided filtering model for improvement to obtain an improved guided filtering model; and performing guided filtering on the two-point correction star map according to theimproved guided filtering model, and completing non-uniformity correction to obtain a non-uniformity corrected star map. According to the method, the purpose of completing star map non-uniformity correction while protecting weak star point signals in daytime can be achieved, and the attitude measurement precision of the star sensor can be improved.

The invention discloses an adaptive horizontal attitude measurement method based on motion state monitoring. Based on a newly established state space model, the horizontal attitude angle is used as a state variable, and the angular velocity increment Δω of random constant value zero offset is compensated. b As the control input of the system equations, the specific force f to compensate for the random constant zero bias b As a quantity measurement, at the same time, the judgment conditions of the carrier maneuvering state are improved, and the acceleration information and angular velocity information output by the MEMS-IMU are comprehensively used to judge the carrier maneuvering information, and the automatic adjustment of the filter measurement noise array is realized, which effectively reduces the The effect of vehicle maneuvers on the horizontal attitude solution. This method has no special requirements on the maneuvering state of the carrier, and can ensure that the system has a high attitude measurement accuracy in different motion states without the assistance of external information.

The invention discloses a measuring method for disturbance acceleration in attitude estimation of an aircraft. The method comprises the following steps: acquiring the attitude quaternion of the aircraft and acquiring measured acceleration in a body-axis coordinate system; acquiring disturbance acceleration in a reference coordinate system according to the attitude quaternion and the measured acceleration in the body-axis coordinate system and then acquiring discrete disturbance acceleration; and subjecting the discrete disturbance acceleration to segmented processing, then carrying out first-order low-pass filtering, and then transferring the processed discrete disturbance acceleration to the body-axis coordinate system for subtraction by the measured acceleration in the body-axis coordinate system so as to obtain accurate aircraft acceleration without the disturbance acceleration. Compared with the prior art, the measuring method provided by the invention is low in cost, needs no GPS sensor for measurement and can improve the attitude measurement precision of subsequent multi-sensor fusion technology.