| With the advantages of comprehensive information,complete autonomy,high concealment and high precision,INS has been widely used in many fields such as aviation,navigation and aerospace.However,the error of inertial devices will change with the change of time and environmental conditions and other factors,which will lead to the decline of accuracy of navigation system after a long time of operation,so the imu needs to be calibrated again.As for the calibration of the Strapdown Inertial Navigation System(SINS),it is traditionally required to disassemble the imu from the carrier on a regular basis and complete the laboratory calibration by returning to the factory for re-installation on the turntable,which has high cost and long cycle and also affects the service efficiency of the carrier.In order to ensure the accuracy of imu,improve its maintainability,extend the calibration period and even achieve the goal of no calibration,the on-line calibration technology of airborne imu is studied in this paper.Using the external observation information provided by the Global Positioning System(GPS)and the maneuvering of the vehicle itself,various errors of the airborne strapdown inertial navigation were estimated by the Kalman filter,so as to achieve online calibration.This paper mainly includes the following research contents:Firstly,the working principle and error propagation characteristics of SINS and GPS were studied,and the error model of SINS/GPS integrated navigation system was established.At the same time,the trajectory generator is designed to simulate the flight parameter information(including attitude,position,speed,etc.)of the aircraft in various flight states(including takeoff,landing,turning,uniform speed,uniform speed flight,etc.),and then the specific force and Angle increment information of the inertial device are calculated.The accuracy of the error model is verified by Matlab simulation.Secondly,the observability of SINS/GPS integrated navigation System was studied.The piecewise Constant System(PWCS)observability analysis theory and method were used,and Singular Value Decomposition(SVD)method was combined to analyze the simulation data of SINS and GPS in different maneuver states obtained by trajectory generator.The effects of different maneuver states on the observability and observability of SINS/GPS integrated navigation system were obtained.Then,Matlab simulation experiment was carried out to verify the estimation effect of the filter on the amount of different states under different maneuvering states.The combination of different maneuvering states was used to obtain a relatively optimal calibration trajectory,namely: uniformly accelerated linear motion +S maneuvering + uniform linear motion.Use this path,to the precision of the gyroscope estimate the 0.001°/ h,the accelerometer estimate 1e-04 m/s^2.In addition,it is concluded that the more complex maneuver combination is,the better the error estimation effect will be.Finally,the actual inertial assembly was placed on the platform,some inertial device errors were artificially added,the turning movement of the aircraft was simulated by rotating the inertial assembly,and the estimated results of the filter were compared with the Matlab simulation results,further verifying the relevant conclusions about the observability of two horizontal tables and three gyros in each state.Imu rotation experiment results also got a simulation experiment didn't get the conclusion,namely: can on the basis of rotation to change course,then roll,stopped the original days to gyro,this can effectively improve the convergence speed and precision of the two horizontal gyro,at the same time also can speed up the day to the convergence speed of gyro. |
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