448 results about "Inertial coordinate system" patented technology

The invention discloses an autocollimation method of a carrier aircraft rotating type strapdown inertial navigation system under a shaking base, and belongs to the field of navigation. The autocollimation method comprises the following steps of: firstly obtaining the longitude and latitude of the position of a carrier, secondly collecting output signals of a fiber-optic gyro and a quartz flexible accelerometer in an inertia measurement unit, and then carrying out automatic compensation on a constant error of an inertia device by adopting a double-shaft rotating table rotatory IMU (inertial measurement unit) to form a rotating type strapdown inertial navigation system; and then calculating a rough initial posture array by utilizing gravity acceleration in an inertial coordinate system as a reference vector, establishing a state equation and a measuring equation again of the system, designing a vanishing adaptation Kalman filter to accurately estimate a misalignment angle of the carrier, and correcting a strapdown posture matrix by utilizing the misalignment angle to finish initial alignment and enter a navigation state. The method isolates the influence of shake of a vessel to initial alignment of a carrier aircraft, and dynamic random disturbance in measured noise is restrained through estimating the initial posture array of the system via the vanishing adaptation Kalman filter method, so that rapid autocollimation of the carrier aircraft rotating type strapdown inertial navigation system is realized.

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.

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 relates to a method for acquiring an initial course attitude of a fiber optic gyro strapdown inertial navigation system. The method is characterized by: extracting gravitational acceleration information from an acceleration meter through filtering processing and compensation, projecting an earth gravitational vector to an inertial coordinate system, with rotational-angular velocity information of the earth included in a projecting component, taking the earth gravitational vector in the inertial coordinate system as reference information, and making use of the output of a gyro and the acceleration meter to estimate an initial attitude angle. On the basis of ensuring that an initial horizontal angle satisfies a precision requirement, the method of the invention substantially improves the precision of an initial course angle of the fiber optic gyro strapdown inertial navigation system.

The invention provides a coarse alignment method for a fiber optic gyro strapdown inertial navigation system based on single axis rotation. The coarse alignment method comprises the following steps: (1) determining initial position parameters of a carrier by a GPS; (2) collecting the data output by a fiber optic gyroscope and a quartz accelerometer and processing the data; (3) determining a transfer matrix Ti<n> of a navigation coordinate system and an inertial coordinate system according to the mutual position relation of the coordinate systems; (4) allowing a single axis of an inertial measurement unit to continuously rotate, setting superposition of an initial time IMU coordinate system s and a carrier coordinates b, then allowing the inertial measurement unit to continuously rotate around a positive direction of an azimuth axis ozb of the carrier coordinate system with an angular velocity Omega=6 degrees/s; (5) determining the relative position relation of the inertial coordinate system and a base inertial coordinate system; and (6) determining an expression of a strapdown matrix after coarse alignment is over according to the calculated relative conversion relation of the coordinate systems calculated in step (3), (4) and (5). The coarse alignment method can help obtain higher coarse alignment precision under a swinging interference condition.

The invention relates to a precise orbit determination method of a navigation satellite for assisting a clock error between stations, which is characterized by comprising the following steps: configuring an atomic clock and a nanosecond stage time synchronization system at each station of an area network, receiving a distance measurement signal of a navigation satellite by each station of the area network to obtain the carrier phase data of each station, detecting the periodic trip generation moment of the carrier phase data by a Blewitt method and carrying out time synchronization between the stations by a two-way satellite time frequency transmission method TWSTFT to obtain a clock error between the stations; resolving a differential equation under an inertial coordinate system to obtain an orbit of the navigation satellite. Before orbit determination data are processed, the method firstly realizes the time synchronization between the stations with high accuracy (i.e.: obtaining the clock error between the stations). During the processing of the orbit determination data, an improved non-error method is adopted, only the orbit needs to be resolved, and the clock error between the stations does not need to be resolved, which not only can improve geometrical factors but also is beneficial to separating a system error (the clock error between the stations). Thus, the orbit determination accuracy can be enhanced.

The invention relates to the field of radio navigation and, more specifically, it relates to methods of determining the current values of the orbital position of the artificial satellite of the Earth which is used in a satellite radio navigation system (SRNS), by the ephemeredes and can be used when performing the radio navigation measurements in the equipment of the consumers of the SRNS signals. The variants of a fast method for determining the coordinates of the satellite are proposed, which consist in that the satellite coordinates are determined at preset points of the orbit spaced from each other by a selected time interval, and further determining the velocities of the satellite at the nodal points, while at the points between the nodal points the satellite coordinate are determined in the geocentric inertial system of coordinates are in a system the Greenwich geocentric coordinates using special formulas or the Taylor polynomial of the second or third order, and determining these at the nodal points using the six-order Taylor polynomial.

The invention provides a spacecraft attitude tracking control method based on discontinuous adaptive control, belongs to the technical field of spacecraft attitude tracking control, and solves the problem that a spacecraft attitude tracking control system is poor in robustness so that a spacecraft attitude tracking control effect is poor in the case of modeling uncertainty, external disturbances and input saturation effects. The spacecraft attitude tracking control method comprises the following specific steps of step 1, establishing an earth centered inertial oIxIyIzI, a spacecraft body coordinate system oBxByBzB and an expected reference coordinate system oRxRyRzR; step 2, according to the coordinate systems established in step 1, obtaining the spacecraft attitude kinematics and a kinetic equation described by using an attitude quaternion, and a spacecraft error attitude kinematics equation and the kinetic equation, namely the attitude tracking control system; and step 3, based on the step 2, based on a sliding mode surface of an integral terminal, designing an attitude tracking controller considering an unknown external disturbance torque and the rotational inertia uncertainty.The spacecraft attitude tracking control method based on the discontinuous adaptive control provided by the invention can be applied to spacecraft attitude tracking control.

The invention relates to a satellite starlight refraction navigation error determination and compensation method. The method comprises the following steps: firstly, generating data of a satellite orbit by virtue of STK software; establishing a satellite attitude planning model; determining an actual observation visual field, and simulating an observation star map comprising refraction stars and non-refraction stars; calculating to obtain the tangential height h of the refraction stars; calculating to obtain the tangential height h'' of the refraction stars with errors of refraction angles and errors of an atmospheric model, wherein tangential height errors are mainly caused by the measurement precision errors of the refraction angles and the errors of the atmospheric model; calculating to obtain the position of a satellite under a geocentric inertial coordinate system; and performing filtration by an extended Kalman filtration method, and outputting a starlight refraction navigation estimated position and position errors. The method provided by the invention can realize accurate prediction of navigation precision of a satellite starlight refraction navigation system, and is comprehensive in error analysis and accurate and reliable in results.

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 relates to the quadrotor unmanned aerial vehicle fault-tolerant control, and aims to realize attitude stability under the condition of an actuator fault of a quadrotor unmanned aerial vehicle. The technical scheme adopted by the invention is that an observer-based quadrotor unmanned aerial vehicle fault-tolerant control method includes the following steps: first defining an inertia coordinate system {I}, a body coordinate system {B} and a target coordinate system {Bd}, and through analysis of an action principle of the actuator on the quadrotor unmanned aerial vehicle, an unknown diagonal matrix is utilized to represent influence of an actuator fault on kinetic characteristics of the quadrotor unmanned aerial vehicle, and a nonlinear kinetic model when a fault occurs in the quadrotor unmanned aerial vehicle actuator is obtained; a unit quaternion-based attitude representation method is adopted; and an observer technology based on an immersion-invariant set method is adopted to observe the actuator. The observer-based quadrotor unmanned aerial vehicle fault-tolerant control method provided by the invention is mainly applied to quadrotor unmanned aerial vehicle design and control occasions.

The invention provides a method for tracking a ground station through a low earth orbit space-borne antenna. The method comprises the following steps: 1) calculating UTC at the next pointing moment of the antenna; 2) calculating the number of transient satellite orbit at the next pointing moment of the antenna; 3) calculating the position of the satellite under a J2000 inertial coordinate system; 4) calculating the position of the ground station under the J2000 inertial coordinate system at the next pointing moment of the antenna; 5) calculating the opposite position vectors of the satellite and the ground station under the J2000 inertial coordinate system; 6) calculating the vector of the opposite position vectors of the satellite and the ground station under geo-center camber line; 7) calculating the vector of the opposite position vectors of the satellite and the ground station under the satellite orbit coordinate system; 8) calculating the vector of the opposite position vectors of the satellite and the ground station under a satellite body system; 9) calculating the vector of the opposite position vectors of the satellite and the ground station under an antenna coordinate system; 10) obtaining an antenna pointing angle, and driving the antenna to perform pointing tracking for the ground station.

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 provides a camera optical axis direction calculation method based on high-precision posture information, and the method comprises the steps: calculating the orientation (shown in the description) of a camera optical axis in a rail coordinate system according to the orientation (shown in the description) of a scanning camera optical axis in a satellite system {b} and the conversion matrix Abo of the satellite system {b} relative to the rail coordinate system {o}; calculating the orientation (shown in the description) of the camera optical axis in an inertial coordinate system {i} according to the orientation (shown in the description) of the camera optical axis in the rail coordinate system {o} and a conversion matrix Aoi of the rail coordinate system {o} relative to the inertial coordinate system {i}; calculating the orientation (shown in the description) of the camera optical axis in an earth-fixed reference coordinate system according to the orientation (shown in the description) of the camera optical axis in the inertial coordinate system {i} and a conversion matrix (HG) of the inertial coordinate system {i} relative to the earth-fixed reference coordinate system. The method can achieve the real-time calculation on the ground and correct the orientation of the optical axis, also can enable the information of a rail and the earth-fixed reference coordinate system on the ground to be stored on a satellite in advance so as to autonomously carry out real-time calculation and correct the orientation of the optical axis, and can improve the imaging capability of the satellite in a simple mode.

The invention provides a ground testing method of a star sensor. The method comprises the following steps of: adjusting the star sensor by using a theodolite so that the Z axis of the star sensor points to true north and the Y axis of the star sensor vertically faces towards zenith; receiving a quaternion of the star sensor under an inertial coordinate system; converting the posture of the star sensor into a quaternion under a WGS84 coordinate system; converting the quaternion under the WGS84 coordinate system of the star sensor into a three-axis Euler angle; continuously operating the star sensor for 30 minutes, storing the difference between a roll angle and the local longitude and the difference between a drift angle and the local latitude in real time, and counting precisions of the roll angle and the drift angle; and storing the difference between a pitch angle and the local longitude in real time, and counting the precision of the pitch angle. According to the method disclosed by the invention, the testing process is simple; any special equipment is unnecessary; the method is not only used for testing the precision of the three-axis Euler angle of the star sensor but also used for completely testing the three-axis polarity of the star sensor; and in addition, only when the time precision satisfies requirements, the method is also used for testing the absolute longitude of the star sensor.

The invention relates to an inertial space gyro calibration test method based on a three-shaft turntable; while a gyro assembly attitude rotation is provided by the three-shaft turntable, the compensation dosage of the earth autorotation angular velocity is calculated according to the real-time rotating angle output of three rotating shafts of the three-shaft turntable, so that the influence of earth autorotation on the gyro measurement output is eliminated, and the inertial space constant angular velocity required for a gyro calibration algorithm is ensured. At the same time, based on attached rotation characteristics of the three rotation shafts of the three-shaft turntable, the method defines a rotation sequence rule of the three rotation shafts and a three-shaft turntable body coordinate system and a reference inertial coordinate system having the initial zero coinciding with the ground horizontal coordinate system, and provides a reference benchmark for determining the inertial system angle increment according to the gyro output and for determining the inertial system attitude according to the rotation angle output of the three rotation shafts of the three-shaft turntable. The method can significantly improve the precision of a ground gyro calibration test, can guarantee the ground to perform effective calibration and test result calibration on a gyro assembly, can establish a good foundation for performing relevant calibration tests of satellites in orbit, and improves the determination precision of the gyro attitude.

The invention relates to a method for correcting in-orbit real-time interaction of double star sensors. The method comprises the steps of obtaining quaternion information of a satellite body coordinate system based on a first star sensor and a second star sensor relative to an earth core inertial coordinate system according to satellite attitude quaternion information output by the first star sensor and the second star sensor and an installation relationship between a satellite body and the first star sensor and the second star sensor, and performing synchronous calculation with current time calculated by an on-star track; calculating corrected deviations of real-time interaction between the first star sensor and the second star sensor according to the quaternion information of the satellite body coordinate system based on the two star sensors having undergone time synchronization relative to the earth core inertial coordinate system, and finally performing real-time interaction correction on the first star sensor or the second star sensor. According to the method, attitude deviations resulting from slow variation and shape change between the star sensors and a satellite reference can be reduced, and high-precision triaxial attitude determination precision of the satellite is guaranteed.

The invention provides a star-sensor-based method for determining the attitude of a carrier relative to a geographical coordinate system. The invention belongs to the technical field of attitude measuring. First, according to a chronometer and a Greenwich sidereal time formula, or through GPS timing, a spring equinox Greenwich time angle of a measured time is obtained; a transformation matrix Cwi from an equatorial inertial coordinate system i to a WGS-84 coordinate system w is obtained by calculation; the geographical coordinate of the carrier is obtained by using external information or a locating apparatus, and a transformation matrix Ctw from the system w to a geographical coordinate system t is obtained by calculation; then a transformation matrix Csi from the system i to a star sensor coordinate system s is obtained by the star sensor; with the three attitude matrices, the attitude matrix Csb of the carrier geographical coordinate system b relative to the system t is obtained by calculation. The attitude matrix Csb is the installation constant matrix of the star sensor. According to the invention, the spring equinox Greenwich time angle is obtained based on the timing apparatus, and real-time calculation of the attitude information of the carrier relative to the geographical coordinate system is realized by computerized programs.

The invention discloses a UAV formation control algorithm based on the Lagrange's equation. An improved manual potential energy field is used in the aspect of formation UAV path programming, and motion of formation UAVs is seen as a constraint dynamic system in the formation maintaining aspect. When the constraint dynamic system is established, a spatial formation constraint equation is established according to a needed formation shape by utilizing spatial geometric relations of the UAVs relative to an inertia coordinate system, a Lagrange multiplier is introduced to establish a constraint dynamic equation set including the formation constraint formation UAVs, and certain stability is combined to eliminate the Lagrange multiplier directly and solve the Lagrange multiplier. The spatial geometric relation of the UAVs is established via the constraint equation, and the formation dynamic equation including the formation constraints is solved to obtain a flight track of the formation UAV to maintain an expected formation, and formation maintenance and transformation are relatively flexible and effective during flight of the UAV formation.