Research On Error Calibration And Separation For Inertial Measurement Systems

The primary goal of the dissertation is to improve measurement accuracy of inertial measurement system. The problem is studed from two aspects which include the high accuracy self-calibration technology of the inertial system and separation method of guidance instrumentation systematic error for flight test of the missile.For multi-position calibration of inertial measurement system on static base, the calibration method of gimbaled inertial system is analyzed particularly. A thorough study is put into that how the fixed error of platform frame-axis and angle sensors error affect the error model and calibration accuracy. The result shows that two kinds of error have remarkable effect on the fixed error of the gyro and all errors of accelerometer, but has little effect on bias error, scaled factor error, quadratic coefficient error of the gyro. The integrated physical mechanism and mathematical model of the continuous calibration method on static base are brought forward, and the observability, filter method, and sensitive analysis are discussed thorouthly, which provides a foundation for engineering practice. The continuous calibration has a shorter process and higher accuracy than the multi-position calibration. Besides, the coefficients of instrumentation systematic error will drift when working temperature changes. The cross validation technology is introduced to select the temperature drift model of error coefficients, and the specific modeling criterion is deducted. The new criterion has a higher precision and is more applicable in small samples than the familiar AIC criterion and MDL criterion.The configuration fixed error of the gyro-free inertial system affects seriously the navigation accuracy. The thesis presents respectively the definitions of the configuration fixed error for a six-accelerometer configuration and a nine-accelerometer configuration, and a new calibration scheme and compensation scheme are put forward. The simulation results show that the relative calibration accuracy of orientation fixed error is less than 3 percent, and the relative accuracy of location fixed error is less than 8 percent with a large angular velocity although the accuracy of location fixed error is seriously affected by the angular velocity of calibration.The spacecraft gyro system is an important kind of attitude sensor. The on-line calibration methods are researched for the drift error of gyros, and the performance of the explicit and the implicit calibration scheme is compared. The result shows that two calibration methods have a same calibration precision, but the precision of the explicit calibration method is affected by the modeling precision of spacecraft attitude dynamic. For the redundant gyro measurement unit in a spacecraft, the implicit calibration with the model replacement technique is used to analyze the drift. The simulation result shows that the model replacement technique can improve the calibration precision in the implicit calibration for redundant gyro system. The separation model of guidance instrumentation systematic error is a foundation to implementing error separation. How to calculate precisely the circumstance function matrix for the gimbaled system is lucubrated. The method of using exterior data and the method of calculating iteratively are put forward. The simulation result shows that two methods can increase the calculating accuracy of the circumstance function matrix. For the separation of guidance instrumentation systematic error of the strapdown system, the thesis set up the separation model based on the circumstance function matrix. The six-freedom trajectory simulation verified the correctness of the model. The separation of the initial launched parameters error of the maneuvering launched missile is a new problem, the mechanism how the initial launched parameters error acts on the telemetry and exterior data is lucubrated, and the separation model of is deducted. The result shows that the majority of coefficients including the guidance instrumentation systematic error and the initial launched parameter error are detachable except several coefficients. The six-freedom trajectory simulator has verified the correctness of the model.Because the circumstance function matrix is correlative approximately, present methods can not completely resolve the separation problem. After analyzing traditional methods, the derived characteristic root method, the partial least square method and the support vectors machine method are presented to separating and converting the guidance instrumentation systematic error. The derived characteristic root method revises traditional principal components method, and the partial least square method and the support vectors machine method can avoid to calculating inversion of the diseased matrix. The simulation result shows the new methods have a better performance than the traditional methods.In summing up it may be stated that the research work has contributed to improve using accuracy of the inertial system, and enhance fighting efficiency for strategic missile.