两种滤波方法在SINS/GPS组合导航中的分析比较

本文针对SINS/GPS组合导航系统中卡尔曼滤波发散的情况:介绍并分析了自适应滤波和H∞滤波算法原理,并进行仿真计算,验证了这两种滤波用于SINS/GPS组合导航系统的可行性和有效性,对实际应用中组合导航系统滤波器的设计具有一定的指导意义。     

两种滤波方法在SINS/GPS组合导航中的应用与比较

本文针对SINS/GPS组合导航系统中卡尔曼滤波发散的情况,引入了自适应滤波和H∞滤波,分析了它们各自的特性,最后进行仿真计算,验证了这两种滤波用于SINS/GPS组合导航系统的可行性和有效性,对实际应用中组合导航系统滤波器的设计具有一定的指导意义.     

基于GPS软件接收机的SINS／GPS紧组合系统仿真

为了系统验证SINS／GPS紧组合系统的性能,基于GPS软件接收机,进行了仿真系统构建。仿真系统由轨迹发生器、GPS中频信号模拟器、IMU信号模拟器、GPS软件接收机、SINS导航解算模块、组合滤波算法和导航性能分析模块等部分构成,其中详细设计了GPS软件接收机中的捕获和跟踪算法、SINS解算以及基于伪距和伪距率的组合滤波算法。仿真结果表明：紧组合导航系统收敛性较好,能够一定程度上抑制惯导系统误差的积累,有较好的导航性能。设计的该系统满足紧组合导航系统性能验证的需要,也为后续的超紧组合研究奠定了良好的基础。     

本期导读

利用掩星和地基GPS研究日食电离层效应(王泽民,等)一文利用COSMIC掩星GPS数据和地基CORS网GPS数据来共同分析2009-07-22发生在亚洲东部和太平洋西部的日全食,进而分析了日食期间电离层效应的物理过程。基于多尺度预处理的GPS/SINS组合导航系统(李廷军,等)一文基于多尺度分析的思想,以离散小波变换为工具,利用小波对惯性元件输出的信息进行并行阈值消噪以削弱惯性元件误差对SINS及组合系统性能的影响;对GPS输出的信息进行并行多尺度预处理,并结合传统的Kal man滤波方法,对系统进行综合滤波;将上述方法引入到GPS/SINS组合导航系统中,利用实测数据进行验证,并给出了基于不     

面向城市道路的多传感器融合定位导航技术

方便地获取高精度、高可靠的轨迹数据是交通、旅游等行业智能化发展的关键。鉴于此,本文设计一款集GPS、SINS和OBD于一体的多源车载组合导航系统来收集轨迹数据,提出数据融合的方法来弥补因传感器噪声导致的位置累积误差,并在GPS失锁时有效预测轨迹位置信息。此方法通过梯度提升与决策树相结合建立INS误差补偿模型,并引用粒子群算法优化模型的回归参数,可有效避免误差积累;再利用联邦滤波器实现GPS、SINS和OBD数据融合,提高了轨迹信息的准确性。实际道路测试证明,基于此方法的组合导航系统,在多种路况下可收集连续精准的轨迹数据。     

两种滤波方法在SINS／GPS组合导航中的应用与比较

针对SINS／GPS组合导航系统中卡尔曼滤波发散的情况,引入了自适应滤波和H∞滤波,分析了它们各自的特性,最后进行仿真计算,验证了这两种滤波用于SINS／GPS组合导航系统的可行性和有效性,对实际应用中组合导航系统滤波器的设计具有一定的指导意义。     

CNS/SINS组合导航系统仿真分析

天文导航和捷联惯性导航(CNS/SINS)相结合构成的组合导航系统,实现导航优势的互补,提高导航的精度和可靠性。文中阐述在惯性系下,以弹道导弹为研究对象而建立CNS/SINS组合导航的数学模型,并利用MATLAB编程设计基于离散型卡尔曼滤波的组合导航系统仿真,对比SINS和CNS/SINS对载体位置误差、速度误差以及姿态角估计的影响,进而验证CNS/SINS组合导航系统的可靠性。     

基于双差伪距/伪距率的GPS/SINS紧组合导航

GPS/SINS组合导航系统的观测模型直接关系着导航系统的精度。提出了一种基于双差伪距/伪距率的GPS/SINS紧组合观测模型。分析了采用双差伪距和采用双差伪距/伪距率两种观测模型对组合导航输出参数精度的影响。实测数据结果表明,采用双差伪距和采用双差伪距/伪距率作为观测值均能实现组合导航系统的收敛。引入双差伪距率观测值明显改善了系统的可观测性,不仅提高了组合导航中速度和姿态角的估计精度,也加快了速度误差和姿态角误差估计的收敛速度。     

浅谈GPS/SINS组合导航系统的设计

SINS/GPS组合导航系统是一种性能较好的导航系统,它结合了GPS的高精度定位,误差无积累及INS的自主性、实时性等优点。两者的结合可使导航系统的成本下降,可靠性增加,精度提高。本文概要地介绍了SINS/GPS的研究背景、结构组成、建立了系统的总体设计方案,给出详细的软件设计框图.并介绍实现系统各个功能的软件算法。实际应用结果表明:该系统的导航精度、成本、体积等指标均达到了设计要求。     

改进Sage-Husa算法在飞机组合导航中的应用

针对机载组合导航系统,考虑不同飞行阶段的气压高度,提出一种改进的Sage-Husa自适应滤波算法,以提高组合导航系统定位精度.该算法通过引入气压高度,实时计算并修正滤波异常判定的调节因子,以满足飞机不同飞行阶段的滤波需求.通过捷联式惯性导航系统(SINS)、全球卫星导航系统(GNSS)定位误差特性仿真、卡尔曼滤波组合算...     

Performance of GPS and GPS/SINS navigation in the CE-5T1 skip re-entry mission

A CE-5T1 spacecraft completed a high-speed skip re-entry to the earth after a circumlunar flight on October 31, 2014. In addition to the strapdown inertial navigation system (SINS), a lightweight GPS receiver with rapid acquisition was developed as a navigation sensor in the re-entry capsule. The GPS receiver effectively solved the poor accuracy problem of long-term navigation using only the SINS. In contrast to ground users and low-earth-orbit spacecraft, numerous factors, including high altitude and kinetic characteristics in high-speed skip re-entry, are important for GPS positioning feasibility and were presented in accordance with the flight data. GPS solutions started at nearly 4900 km orbital altitude during the phases of re-entry process. These solutions were combined by an inertial measurement unit in a loosely coupled integrated navigation method and SINS navigation initialization. A simplified GPS/SINS navigation filter for limited resources was effectively developed and implemented on board for spacecraft application. Flight data estimation analyses, including trajectory, attitude, position distribution of GPS satellite, and navigation accuracy, were presented. The estimated accuracy of position was better than 42 m, and the accuracy of velocity was better than 0.1 m/s.     

高精度、微小型惯性/GPS组合导航系统研究

针对INS/GPS组合导航系统航向角可观测性差的问题,提出了把航向信息耦合到Kalman滤波器量测方程中的方法,以及用航位推算校正航向的方法来抑制航向角发散的趋势。为满足低成本、小型化的要求,以MEMS陀螺和加速度计为核心,构建了高精度组合导航系统。仿真试验表明,两种算法均达到了较高的定位/定向精度,对导航系统在小型化领域的推广应用具有实际意义。     

基于多尺度预处理的GPS/SINS组合导航系统

基于多尺度分析的思想,以离散小波变换为工具,利用小波对惯性元件输出的信息进行并行阈值消噪以削弱惯性元件误差对SINS及组合系统性能的影响;然后,对GPS输出的信息进行并行多尺度预处理,并结合传统的Kalman滤波方法,对系统进行综合滤波;将上述方法引入到GPS/SINS组合导航系统中,利用实测数据进行验证,并给出了基于不同方法的大量实验曲线。实验结果表明,该方法可以有效削弱惯性元件以及GPS误差对系统的影响,提高了GPS/SINS组合导航系的精度和可靠性。     

State transformation extended Kalman filter for GPS/SINS tightly coupled integration

Extended Kalman filter (EKF) is a widely used estimator for integrated navigation systems, and it works well in general situations. However, in adverse conditions such as partially observable environments and highly dynamic maneuvers, the performance of the traditional EKF-based strap-down inertial navigation system (SINS)/GPS integrated navigation system is easily to be affected by the dynamic changes of the specific force, thus leading to the problem of error covariance inconsistency. Though the inconsistency problem can be overcome to some extent if the system matrix, the states and the error covariance matrix are propagated as fast as possible in the SINS calculation rate, the problem cannot be fully solved. State transformation extended Kalman filter (ST-EKF) mechanization, with a new converted velocity error model for the SINS, is proposed, which can also be used to solve the inconsistency problem. In the ST-EKF, the specific force vector in the system error model is replaced by the nearly constant gravity vector for local navigation. Since the propagation and the updating of the ST-EKF can be executed simultaneously in the updating interval, the computation cost is greatly reduced compared with the traditional EKF. Experiments for the GPS/SINS tightly coupled navigation, including linear vibration Monte Carlo test and an unmanned aerial vehicle flight test, are implemented to evaluate the performance of the proposed ST-EKF. The results show that the proposed ST-EKF has superior performance to the traditional EKF, especially in partially observable situations.     

Design of a low-cost attitude determination GPS/INS integrated navigation system

Due to their complementary features of GPS and INS, the GPS/INS integrated navigation system is increasingly being used for a variety of commercial and military applications. An attitude determination GPS (ADGPS) receiver, with multiple antennas, can be more effectively integrated with a low-cost IMU since the receiver gives not only position and velocity data but also attitude data. This paper proposes a low-cost attitude determination GPS/INS integrated navigation system. The proposed navigation system comprises an ADGPS receiver, a navigation computer unit (NCU), and a low-cost commercial MEMS IMU. The navigation software includes a fault detection and isolation (FDI) algorithm for integrity. In order to evaluate the performance of the proposed navigation system, two flight tests have been performed using a small aircraft. The first flight test confirmed the fundamental operation of the proposed navigation system and the effectiveness of the FDI algorithm. The second flight test evaluated the performance of the proposed navigation system and demonstrated the benefit of GPS attitude information in a high dynamic environment. The flight test results show that the proposed ADGPS/INS integrated navigation unit gives reliable navigation performance even when anomalous GPS data is provided and gives better navigation performance than a conventional GPS/INS unit.