惯性测量系统误差标定及分离技术研究
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摘要
论文以提高惯性测量系统使用精度为目的,从惯性测量系统高精度标定与飞行试验误差分离两个方面开展了深入研究,得到许多有意义的结果。
对于惯性测量系统地面标定试验,分析了惯性系统特别是平台式惯性系统的标定方法。对于多位置标定试验,深入讨论了平台框架轴误差与角度传感器误差对误差系数标定精度的影响,这两项误差主要影响陀螺仪安装误差以及所有加速度计误差系数的估计。提出了平台式系统连续翻滚自标定方法,详细分析了其物理机理并推导了数学模型,并对模型的可观测性、滤波方法、参数敏感性等进行了详细的讨论,为工程实现提供了理论基础。该方法标定时间短,估计精度高。针对惯性测量器件误差系数随温度变化发生漂移的情况,将交叉验证技术引入到误差系数温度漂移建模中,详细推导了基于交叉验证的温度漂移建模的准则,与传统的AIC准则、MDL准则相比,交叉验证技术建模的准确度更高,而且更适合小样本下的模型选择。
无陀螺惯性系统是一种全新的惯性测量系统,其使用精度受构型安装误差影响较大。论文给出了六加速度计与九加速度计两种构型下构型安装误差的定义,同时提出了一种构型安装误差的标定方案与补偿方案,仿真结果表明方向安装误差标定相对误差优于3%,位置安装误差虽然受标定中转台角速度影响较大,但在较大的角速度下相对误差可小于8%。补偿方案计算过程简单,可以补偿85%以上的安装误差。
空间陀螺仪系统是空间飞行器一种重要的姿态测量敏感器,论文研究了空间陀螺仪误差漂移模型的在轨标定方法,对比分析了显式标定方法与隐式标定方法的性能,结果表明,二者的精度基本相当,但显式标定方法受航天器姿态动力学建模精度的影响比较明显。针对常见的空间陀螺仪冗余配置,利用模型置换技术应用隐式标定方法分析了陀螺仪冗余配置情况下的标定方法,提高了标定精度。
弹道导弹制导工具误差分离模型是实现工具误差分离的基础。讨论了平台式系统环境函数矩阵的精确计算问题,提出了利用外测数据与应用迭代技术精确计算环境函数矩阵的方法,仿真计算表明,这两种方法能够有效提高环境函数的计算精度。针对捷联惯性系统工具误差分离问题,建立了基于环境函数矩阵的捷联系统工具误差分离模型,并利用六自由度弹道仿真验证了模型的正确性。初始发射参数误差分离是机动发射导弹工具误差分离中出现的新问题。论文深入分析了初始发射参数误差对弹道遥测、外测数据的影响,推导了初始发射参数误差分离模型,结果表明,除少部分误差系数无法分离外,大部分工具误差系数与初始发射参数误差是可分离的,且分离模型是线性的。六自由度弹道仿真软件验证了模型的正确性。
在制导工具误差分离中,由于环境函数矩阵存在严重的复共线性,该问题始终没有得到很好的解决。在分析传统方法的基础上,提出了衍生特征根方法、偏最小二乘方法以及支持向量机方法,衍生特征根方法是对传统主成份方法的改进,偏最小二乘方法与支持向量机方法则完全避免了病态矩阵的求逆问题。仿真算例和工程应用表明,这三种方法能够在一定程度上克服环境函数矩阵严重复共线性所造成的估值偏差较大的问题。
论文的工作对于提高惯性测量系统使用精度、提高导弹作战效能具有重要意义。
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.
对于惯性测量系统地面标定试验,分析了惯性系统特别是平台式惯性系统的标定方法。对于多位置标定试验,深入讨论了平台框架轴误差与角度传感器误差对误差系数标定精度的影响,这两项误差主要影响陀螺仪安装误差以及所有加速度计误差系数的估计。提出了平台式系统连续翻滚自标定方法,详细分析了其物理机理并推导了数学模型,并对模型的可观测性、滤波方法、参数敏感性等进行了详细的讨论,为工程实现提供了理论基础。该方法标定时间短,估计精度高。针对惯性测量器件误差系数随温度变化发生漂移的情况,将交叉验证技术引入到误差系数温度漂移建模中,详细推导了基于交叉验证的温度漂移建模的准则,与传统的AIC准则、MDL准则相比,交叉验证技术建模的准确度更高,而且更适合小样本下的模型选择。
无陀螺惯性系统是一种全新的惯性测量系统,其使用精度受构型安装误差影响较大。论文给出了六加速度计与九加速度计两种构型下构型安装误差的定义,同时提出了一种构型安装误差的标定方案与补偿方案,仿真结果表明方向安装误差标定相对误差优于3%,位置安装误差虽然受标定中转台角速度影响较大,但在较大的角速度下相对误差可小于8%。补偿方案计算过程简单,可以补偿85%以上的安装误差。
空间陀螺仪系统是空间飞行器一种重要的姿态测量敏感器,论文研究了空间陀螺仪误差漂移模型的在轨标定方法,对比分析了显式标定方法与隐式标定方法的性能,结果表明,二者的精度基本相当,但显式标定方法受航天器姿态动力学建模精度的影响比较明显。针对常见的空间陀螺仪冗余配置,利用模型置换技术应用隐式标定方法分析了陀螺仪冗余配置情况下的标定方法,提高了标定精度。
弹道导弹制导工具误差分离模型是实现工具误差分离的基础。讨论了平台式系统环境函数矩阵的精确计算问题,提出了利用外测数据与应用迭代技术精确计算环境函数矩阵的方法,仿真计算表明,这两种方法能够有效提高环境函数的计算精度。针对捷联惯性系统工具误差分离问题,建立了基于环境函数矩阵的捷联系统工具误差分离模型,并利用六自由度弹道仿真验证了模型的正确性。初始发射参数误差分离是机动发射导弹工具误差分离中出现的新问题。论文深入分析了初始发射参数误差对弹道遥测、外测数据的影响,推导了初始发射参数误差分离模型,结果表明,除少部分误差系数无法分离外,大部分工具误差系数与初始发射参数误差是可分离的,且分离模型是线性的。六自由度弹道仿真软件验证了模型的正确性。
在制导工具误差分离中,由于环境函数矩阵存在严重的复共线性,该问题始终没有得到很好的解决。在分析传统方法的基础上,提出了衍生特征根方法、偏最小二乘方法以及支持向量机方法,衍生特征根方法是对传统主成份方法的改进,偏最小二乘方法与支持向量机方法则完全避免了病态矩阵的求逆问题。仿真算例和工程应用表明,这三种方法能够在一定程度上克服环境函数矩阵严重复共线性所造成的估值偏差较大的问题。
论文的工作对于提高惯性测量系统使用精度、提高导弹作战效能具有重要意义。
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.
引文
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