机载导航卫星反射信号海面高度测量

对机载导航卫星反射信号海面高度测量模型、定点跟踪和模型拟合2种估计反射相对于直射信号时延的方法以及误差校正进行了研究。针对Z-V模型拟合法进行时延估计复杂度高的特点,提出了7-β模型和相应的时延估计方法;为消除定点跟踪法估计的时延中的海况偏差,提出了查表法和加权平均结合的校正方法。利用2015年12月3日CSIC-IEEC在波罗的海采集的机载数据进行了海面高度反演。结果表明通过已有时延估计方法和所提出的7-β模型拟合法对GPS L1和Galileo E1反射信号进行时延估计,并通过提出的误差校正算法对海况误差进行校正后可得到精度为亚米级海面高度。     

Relative navigation for multi-spacecraft system with GPS

A type of multi-spacecraft system with kinematical restraint but no structural restraint and force action is considered. Both the absolute and relative navigation information is required for this multi-spacecraft system, but the relative information is more critical and the accuracy requirements for relative information will be much higher than those for the absolute information. In this paper, the Global Positioning System (GPS)/Differential GPS (DGPS) are introduced and used for relative navigation. Relative motion of space vehicles is modeled. Relative position, relative velocity and relative attitude are represented and solved by GPS/DGPS measurements. Using a type of commercial GPS receiver onboard spacecraft and relative differential GPS technique, the relative navigation of space vehicles can be implemented in real-time     

Real-time navigation using the global positioning system

This paper presents the results of an investigation of the application of the Global Positioning System (GPS) to real-time integrated missile navigation. We present quantifiable measures of navigation accuracy as a function of GPS user segment parameters. These user segment parameters include antenna phase response accuracy, single versus dual frequency, and Kalman filter structure and size. We also formulate some new phase-locked loop (PLL) filter designs for application in GPS receivers, and demonstrate their superiority over more conventional filters     

Optimization of biased proportional navigation [guided projectiles]

An analytical treatment of the biased proportional navigation (BPN) is carried out with the aim of optimizing the bias parameter. It is shown that optimum biasing can lead to significantly more control-effort-efficient PN guidance in a wide variety of engagement situations, especially those involving higher target maneuvers. The performance of the BPN is compared with the standard (unbiased) PN system for the general case of a maneuvering target, and performance of the BPN is maximised to obtain the optimum bias value. The optimum bias is expressed through a simple algebraic equation, which can be readily solved. For the special (and very useful) case of the effective navigation constant being equal to three, the equation reduces to a quadratic, leading to an explicit expression for the optimum bias. Specific examples are provided to show the benefits of the BPN law. The higher control efficiency of the law is especially useful in extra-atmospheric interception, where the savings in control effort directly translates to a saving of propellent which forms part of the payload     

Inertial navigation system synthesis approach and gravity-inducederror sensitivity

A design criterion for improving the performance of the speed-damped inertial navigation system is presented. The single-axis speed-damped system is approached by optimizing the response of the system to a step-function disturbing signal. Butterworth, integral of time-multiplied absolute-value of error (ITAE), and solution-time standard forms are assumed to be the figures of merit for optimizing the system performance. The steady-state RMS (root-mean-square) gravity-induced navigation errors that are excited in the speed-damped system are determined for two gravity uncertainty models. The proposed figures of merit are compared. These comparisons reveal the sensitivity of predicted navigation errors to uncertainties in the gravity statistics, and simplify the choice of a suitable figure of merit for use in the design and error analysis of inertial navigation systems     

Multiposition alignment of strapdown inertial navigation system

The authors demonstrate that the stationary alignment of strapdown inertial navigation system (SDINS) can be improved by employing the multiposition/technique. Using an observability analysis, it is shown that an optimal two-position alignment not only satisfies complete observability conditions but also minimizes alignment errors. This is done by analytic rank testing of the stripped observability matrix and numerical calculation of the error covariance. It is also shown that an optimal three-position alignment accelerates the convergence of the alignment error compared with two-position alignment     

Target detection trials with a millimeter wave radar system

Unobstructed, large RCS targets, similar radar targets surrounded by moving foliage, and small targets in severe clutter have been used as test cases for two pre-processing algorithms and several threshold levels in an experimental millimeter wave radar system. The rather conventional "six-out-of-eight" pulse radar selection method with binary output has been compared to an algorithm that accepts a target if the pre-defined trigger level is crossed by the average of the eight consecutive pulses. In this case, however, the output is an analog value corresponding to the relative average video amplitude. In terms of plotted video, this process seems to give a slightly better combination of false alarm rate and detection probability. Large targets are easier to detect from foliage clutter with the conventional method.     

A strapdown inertial navigation system for the flat-Earth model

The development of a strapdown inertial navigation system (SINS) for aerodynamically controlled vehicles, which are limited to altitudes below 30 km (that is, a small distance compared with the Earth's radius of about 7000 km), or using the so-called flat-Earth model (FEM), is the principal objective of this work. In dealing with the FEM equations, the north, east, down (NED) frame on the surface of the Earth is taken as an inertial reference frame. Although, this frame is both accelerating and rotating, the accelerations associated with the Earth's rotation are negligible compared with the acceleration that can be produced by a maneuvering aircraft. Also, in this model, the gravity is taken as constant. In developing the SINS for the FEM, the aerodynamic force and moment have dominant roles, depending primarily on such variables as the angle of attack and sideslip, their derivatives, components of the angular velocity of the aircraft, and the control inputs. On the other hand, the SINS deals with such variables as the small-angle rotation vectors. Thus, it was necessary to link both set of variables as state variables of the strapdown FEM as is done in this work. The developed model is relevant for small (less than 200) angles of attack and sideslip.     

不依赖卫星的定位导航与授时体系研究

定位导航与授时(positioning, navigation and timing, PNT)体系是支撑人类社会发展的基础设施。当前以卫星为核心的天基PNT应用广泛,但在卫星拒止环境下面临严峻挑战。通过总结天基PNT在国民经济发展与国防建设中的重大作用,分析天基PNT的脆弱性与依赖风险,明确了对不依赖卫星的PNT体系的迫切需求。提出了不依赖卫星的PNT体系概念与特征,即以“惯导+时钟”为核心增强基础PNT能力,利用外源传感器多源融合提高精度保持能力,在卫星拒止条件下为用户提供可承载的高性能PNT服务。基于现有技术基础重点发展微小型、高精度、低成本惯导技术,微时钟技术以及智能融合技术是有效途径,能够实现不依赖卫星的PNT覆盖性、精确性与可承载性的全面提升。     

Fast alignment and calibration algorithms for inertial navigation system

Algorithms for fast estimating the azimuth misalignment angle and calibrating gyro drift rates are approached from the point of view of control theory. By introducing the Lyapunov transformation, the equivalence of Strapdown Inertial Navigation System (SINS) and Gimbaled Inertial Navigation System (GINS) is discussed, and it shows that the analysis results of GINS can be applied to SINS directly by using such kind of equivalence. A similar transformation that based on physical essence is introduced, so that the true states can be replaced by the so-called pseudo-states, and then the observable states of INS can be dynamically decoupled with the unobservable states. Consequently, the best completely observable subsystem model of INS can be obtained. Based on the simplified subsystem model of INS, the algorithms for fast estimating the azimuth misalignment angle and calibrating gyro drift rates are proposed. The proposed algorithms show that the azimuth misalignment angle and gyro drift rates can be estimated from the rates of leveling misalignment angles without using the gyro output signals.     

平台式惯性导航系统在线可靠性评估技术

平台式惯性导航系统是现代导航的重要设备,由于其所处环境和制造工艺各不相同,导致各个体之间的可靠性有着明显的差异。如果对收集到的性能误差数据进行统一处理,无法准确地反映个体之间的差异性,不利于掌握平台式惯性导航系统个体的可靠性规律。针对该问题,对平台式惯性导航系统的在线可靠性评估方法进行了重点研究。结合平台式惯性导航系统使用过程中的性能变化特点和失效机理,利用复合Poisson过程建立了其性能退化模型,并给出了性能参数评估方法。通过算例分析,说明建立的性能退化模型能够较好地描述平台式惯性导航系统的性能退化规律,有助于掌握平台式惯性导航系统的可靠性水平。     

惯性信息辅助的大视角目标快速精确定位

目标定位技术广泛应用于航空领域的侦察机、无人机等各类侦察打击任务中,目标定位精度的高低及效率对作战效果具有重要影响。针对仿射尺度不变的特征变换(ASIFT)算法对远距离大视角目标定位精度较低、速度较慢的问题,提出了一种基于惯性信息辅助的大视角目标快速精确定位方法。该方法首先对目标实测序列图像构造尺度空间,结合FAST特征检测与FREAK特征描述的方式进行匹配,实现对待定位目标的快速提取;然后利用机载惯性信息求解目标实测图与参考图之间的透视变换矩阵,利用该矩阵对实测图进行变换以减小图像间视角差异,克服了ASIFT算法盲目匹配计算的弊端,并通过FAST特征检测与FREAK特征描述相结合的方式提升了大视角图像的匹配速度;最后通过单应性矩阵映射关系实现对目标的精确定位。实验结果表明,大视角目标快速精确定位方法匹配耗时比ASIFT算法的减小了1个数量级,定位精度比目标平均值定位算法精度提高了1个数量级,有效提高了图像匹配定位在航空领域的应用效率。     

捷联惯导系统误差特点分析

对捷联惯导系统的误差源进行了深入分析,结果表明当陀螺仪刻度系数误差较大时,捷联惯导系统定位误差闭合现象较明显,即飞机返场时,误差有明显减小的现象;分析了舒拉调谐周期对惯导系统位置误差的影响并进行仿真,仿真结果验证了舒拉调谐对误差的调制作用,即在舒拉周期振荡分量的影响下,惯导系统的累积误差在某段时间内存在误差减小现象。     

Dynamics investigation of the two-channel automatic system controlling the mirror of aircraft-based surveillance-search system

The mathematical model of the control system for the mirror having a cardan suspension, installed aboard an aircraft is investigated. Control algorithms that provide the required dynamic properties in scanning, guidance and tracking modes were synthesized. We analyze dynamics of the control system for a mirror with regard to nonlinearities in controllers using the computer-aided simulation.     

平台惯导系统飞行试验性能误差分析

针对某型平台式惯导系统试飞中存在的性能超差现象,对惯导系统误差机理和飞行数据进行了分析,提出了影响飞行性能的误差因素。根据分析结果,从提高惯性元件精度和对相关误差进行标定和补偿两方面采取措施对惯导系统进行改进。改进后的惯导系统经过试飞验证,其性能满足设计指标,表明所采取的改进措施有效可行。     

图像传送及控制联络系统

本简要说明了利用电视广播技术，完成小区域音像传送及其他控制的方法，其特点是监控现场与控制室之间连线少，各种信号仅通过一根单芯信号电缆传送，使得系统安装，测试，维修简单。     

An advanced flight control and navigation system implementation fortactical helicopters

US Army helicopters that will allow single-crewmember operability in tactical environments require a high degree of automation. A tightly integrated flight control and navigation system has been studied. An architecture is discussed which meets anticipated flight control and navigation requirements, at an acceptable weight, using techniques of functional and physical integration. The architecture maximizes the sharing of sensor and computational resources in a flight critical environment and uses modular equipment packaging, dual fault-tolerant tetrad ring-laser-gyro inertial measurement units, and triplex self-checking processor pair-based flight critical processing channels     

基于惯性空间稳定的双轴惯导旋转调制方法

陀螺标度因数误差是影响长航时船用旋转调制惯导系统的关键误差源,其与地球自转和载体运动的耦合误差,可导致惯导系统误差发散。针对此问题,结合船用惯导使用特点,采用外航向、内俯仰的双轴旋转框架结构。在此基础上,提出了一种基于惯性系的双轴旋转惯导系统多位置转停调制方案,通过补偿地球自转和载体运动在双轴旋转惯导内外框架旋转轴上的投影分量,可显著降低陀螺标度因数误差对长航时导航精度的影响。数学仿真和船载试验结果表明,在载体航向角运动的场景下,该方法与传统的双轴旋转调制方案相比可有效抑制地球周期项振幅的增大,系统导航位置误差的发散也降低50%以上。     

Parametric navigation system synthesis on the basis of a magneto-inertial course transmitter

A technique for the parametric synthesis of the navigation system on the basis of the magneto-inertial course transmitter is proposed; the technique makes it possible to minimize the error scatter in determining the ground speed components, to analyze a nature of the error scatter variation in determining the location coordinates in time, to simulate and further choose regular intervals of using additional data obtained from the radionavigational satellite system.     

Control failure of the roll-isolated inertial navigation system under large pitch angle

Roll-isolation is an effective way for spinning vehicle to greatly reduce the roll gyro range of strapdown Inertial Navigation System(SINS) and increase the accuracy of inertial navigation.However, during a recent flight test, the roll-isolated control system failure was observed under a large pitch angle(706 h 6 85), which introduces a sharply increase in the roll angular velocity,the saturation of roll gyro and the inertial navigation failure. To address this issue, the governing equation of t...