Abstract For near-Earth spacecraft, many environmental forces are utilized to unload reaction wheels. However, effects like magnetic force or atmospheric friction hardly exist in deep-space environments where solar sails operate. Thus, this paper proposes an unloading method that uses reflectivity control devices to generate Solar Radiation Pressure (SRP) torque to conduct the unloading process of solar sails operating in deep space. Firstly, the basic principles of the Reflectivity Control Device (RCD), satellite kinetic model, and Pulse Width Pulse Frequency (PWPF) modulator are given. Secondly, based on the torque envelope obtained by traversing all switch combinations of RCD, a novel strategy using a PWPF modulator is specially designed to achieve efficient unloading. In the end, we carry out numerical simulations and analyze the influence of different parameters, through which the validity of the proposed torque resolution scheme is demonstrated.
Compared with the traditional anomaly detection methods, machine learning algorithms do not rely on manual and have the ability to extract advanced features of data. However, anomaly detection of spacecraft telemetry data by supervised machine learning is a challenging problem due to the lack of priori knowledge. This paper presents a signal anomaly detection algorithm based on attention mechanism. First, the long-distance characteristics of spacecraft telemetry data are captured by attention mechanism. Then, the stacked autoencoder compresses the data dimension and reconstructs the input signal to obtain the error reconstruction sequences. Furthermore, the anomaly indexes of the error reconstruction sequences are marked by the window threshold method to realize the anomaly detection of the spacecraft telemetry signal. Finally, the effectiveness of the algorithm is verified by a group of examples based on multi-channel spacecraft telemetry signals.
In view of the need to keep secret the transmitted control commands and the coordinates of unmanned aerial vehicles（UAVs） when they are used for reconnaissance or fire attack, a new physical layer encryption method is proposed in this paper. By considering that UAVs and ground stations have a pair of finite dimensional predictive bidirectional channel vectors（ERCVs）, the continuous encryption function（CEF） is used firstly in this method. The two ERCVs of UAVS and ground stations are converted into two quasi-continuous pseudo-random number sequences（QCPRNs） of arbitrary dimensions, and then the QCPRNs are further converted into two uniformly distributed（UD） QCPRNs sequences, and the UD-QCPRNs are superimposed on the information transmitted by UAVS in a modular manner. UD-QCPRNs is generated at the ground station to complete the decryption on the ground, and the information and coordinates are hidden for the safe ground communication of the UAV. The results of simulations show that the proposed method can counter any possible detection methods of the attacker and is suitable for the safe flight of UAV or the completion of war tasks.
The limitations of common unloading methods is analyzed such as water flotation, suspension and air floating, and a magnetic unloading mechanism is designed. A radial parallel dual structure is applied to the new mechanism which overcomes the difficulty of requiring external cooling equipment when the large bearing mechanism is suspended. The use of double-sided and seamless splicing permanent magnets realizes the consistency of magnetic field strength in space work and reduces the introduction of interference force; At the same time, the non-magnetic and non-conductive non-metallic coil winding is selected, and there is no additional hysteresis resistance and damping torque, and the requirement of zero gravity unloading is realized. The result of application test shows that the unloading mechanism meets the requirements of ground research and verification of space magnetic levitation rotary platform.
The research of air-breathing high speed aircraft is focused. Aiming at the high speed aircraft in the flight process of the great uncertainty and stamping engine dynamic pressure control method based on backstepping sliding mode is proposed to solve the problem of strict constraint on dynamic pressure during operation.Firstly,the precise linearization of the dynamics model of high speed aircraft is implemented,then based on the exact linearization model,the backstepping sliding mode controller is designed and the stability of the closed-loop system is proved by Lyapunov stability theory.Finally,the simulation verified that the proposed control method can accurately track the flight pressure in the case of large disturbance and large uncertainty and the stable operation of the ramjet is ensured.
In order to solve the problem of low tracking precision of a single observer for a near space vehicle over long distance, a multi-observer cooperative tracking method is applied. In order to accurately fit the motion state information of high maneuvering target in near space, an interactive multi-model quadratic filtering method is proposed. Based on the target tracking information, a differential auto regressive integrated moving average（ARIMA） model is used to predict the target trajectory. The simulation results show that the tracking and prediction precision of the target motion information is effectively improved by using the quadratic filtering trajectory tracking and prediction method, compared with the traditional tracking and prediction methods. The angular velocity tracking error is reduced from 30% to less than 5%, and the velocity tracking error is reduced from 20% to less than 2%. Moreover, it can effectively predict the trajectory of the target in the future 100s without a lot of prior information, and the error is less than 0.5km.
In flight test for lifting body aircraft, debris in failed mission can be spread along the flight path in a vast area. In order to accurately assess the safety risk of flight test, a method based on reliability assessment on intersected flight profile is proposed in this paper. The influence of the current system state and past flight profile can be reflected by using this method. In this paper, the calculation of debris spreading is converted into reliability assessment on intersected flight profile and flight simulation, and the approach is given with a simulation example. This method can be used in a variety of lifting body aircraft flight test safety risk assessment.
Regarding the difficulty of multi-view and multi-angle airborne image matching, an image matching method based on mutual attention mechanism and fine-grained feature segmentation and alignment is proposed in this paper. Based on the usage of convolutional neural network to extract local fine-grained features, self attention mechanism is used to excavate global information and fine-grained features are incorporated into the expression by introducing this method, and then similarity fine-grained features among matched images are enhanced by the mutual attention mechanism. Thereafter, the features are segmented and aligned according to the attention score of each fine-grained feature. Finally, the improved triplet loss is used to constrain the entire model, and the model is established more robust on multi-view and multi-perspective data. The experimental results on the University-1652 data set show that the matching performance of the proposed method is better than that of current state-of-the-art methods.
In recent years, quadrotor is attracted more and more attention in scientific research and industrial fields. The design and implementation of the quadrotor flight control system are always one of the core problems, and attitude control is the basis of whole flight control. An Explicit Model Predictive Control（EMPC） method around the attitude control problem of the quadrotor is presented is this paper. Firstly, the force analysis of the quadrotor is implemented, and the mathematical model is established, the attitude model of the quadrotor is obtained by decoupling. Then, explicit forecast attitude controller is designed, including the linearization of the system model and the establishment of the objective function, the offline and online optimization online control only needs to find the appropriate control area, which can be controlled by simply calculating the amount, reducing the online computation and still pursuing the optimal solution in some degree. A complete quadrotor nonlinear simulation model is established in MATLAB/Simulink simulation environment, and the feasibility and effectiveness of the design algorithm are verified.
Based on manifold theory, a method of feature learning is proposed, which is based on the local preserving projection of manifold learning to the operating data of the attitude system and fault detection based on the statistical characteristics. Firstly, the local manifold structure of the statistic is found by the local preserving projection method. The purpose of the projection method is used to map the closure points in the original space to the closure points in the low dimensional space. Secondly, T~2 and square prediction error（SPE） are established, normal data training results are used for statistics, and kernel density estimation（KDE） is used to determine the fault control limit, and fault detection is realized. Through data simulation verification, the proposed method can effectively realize fault detection by a typical fault detection rate of 100%.
Aiming at the problem of optimal control strategy for spacecraft on orbit by using tangential monopulse jet to avoid space collision risk, a method is proposed.Firstly, a relative motion equation of the satellite with small eccentricity orbit after tangential control related to the original orbit is derived. Then, the spatial proximity geometric relationship is analyzed. There are distance constraints both from the perpendicular direction and space proximity, under which the calculation method of the minimum control quantity corresponding to control time is developed and the function curves at different control time are drawn. Besides, the monotone function relationship between control quantity and control effect is proven, and the sensitivity of the control time to the control effect is discussed. As long as the control time is obtained by integrating multiple constraints such as station tracking, satellite conditions and fuel saving, the minimum tangential control quantity and the control direction can be determined through the above control time and control function curve.
In order to determine the initial orbit of the pointed target after separation in space test, the impulse thrust maneuver model is proposed for approximating the target separating process in orbit, and the Gauss perturbation motion equation whose singularity has been eliminated by using the vernal equinox orbital elements, is established to quantitatively describe the relationship between the increment of instantaneous velocity caused by the thrust and the change of orbital elements. Under the main perturbation forces having a marked impact on the dynamic differential equation of the target and the consideration of the obit prediction precision and modeling complexity, the method of high-precision orbit prediction for the target is developed by each selected perturbation acceleration formula. Simulation results show that the relative error of the initial values of the target orbit compared with the data of a certain spacecraft simulation software is not more than 1.576%. The decay characteristics of the target orbit and the variation of the visible distance of the target relative to the satellite are analyzed, which have important value for reference to orbit design of the accompanying satellite, separation window scheme and optimal initial orbit determination for the target in space test.
In order to solve the problems of the existing sea-launched rockets in the process of operation, such as heavily relying on center scheduling, low fault tolerance rate of wireless measurement and control data and complex interference of sea wave motion on Marine wireless communication, A wireless test and fire control implementation method of sea launch rocket is presented in the Blockchain Scenario. Based on the simulation design of the system antenna by Ansoft HFSS three-dimensional electromagnetic field simulation software, the hardware design and algorithm control of the system wireless transceiver module, main control module, acquisition module and distributed block storage of telemetry data are studied in this paper. At last, the experiment of wireless data transmission and surveillance video transmission at three distances of 100m, 3km and 5km is implemented. The test results show that the upstream and downstream throughputs of TCP mode data of the system is not less than 71.8Mbps under the condition of 5km, the monitoring video transmission of the unmanned launch platform is smooth, and the error data can be traced throughout the whole process, which set a new benchmark for the establishment of the block chain communication management system in the sea launch operation of carrier rocket.
针对卫星陀螺仪故障检测中存在的冗余依赖、微小故障覆盖问题,提出一种基于长短时神经网络(LSTM)的故障检测方法.首先对卫星陀螺仪建模,考虑到卫星姿态控制回路对陀螺仪微小故障覆盖影响,利用半物理仿真平台采集陀螺仪正常与故障数据;然后使用部分正常数据训练LSTM神经网络,使得网络具有预测陀螺仪输出的能力,并将另一部分正常数据输入到训练好的网络模型,得到预测误差,进一步设定故障阈值;最后,将测试数据输入提出的故障检测模型,仿真验证其时效性和准确性.结果表明,在采样频率为10Hz时,对于陀螺仪的卡死、噪声以及偏差故障,基于LSTM神经网络的故障检测模型能在故障发生2s内检测出故障,并达到了 98.9％的准确率.
针对高低温偏差下固体发动机推力、耗尽时间和秒耗量散布大的问题,提出了一种以速度模值增量为自变量的固体运载火箭序列凸优化制导算法.该算法以速度增量替代传统的飞行时间为自变量建立动力学模型,采用序列凸优化算法求解多约束耗尽关机制导问题,并基于速度模值增量在线辨识高低温工况,修正内弹道模型.该方法相比以飞行时间为自变量的序列凸优化算法具有更高的制导精度和鲁棒性.最后,通过某型固体运载火箭高空飞行段制导数值仿真,验证了以速度增量为变量的序列凸优化算法和内弹道修正模型的有效性.
综述了液体火箭发动机的故障模式,总结了液体火箭发动机故障诊断技术的最新成果,包括基于物理模型、信号分析和人工智能的故障诊断方法;将不同故障诊断方法的应用进展及其诊断效果进行了对比分析;并对液体火箭发动机故障诊断方法的发展趋势进行了展望.
基于对飞行器飞行控制软件的需求(包括功能需求以及性能需求)的分析结果,提出一种基于战星嵌入式实时操作系统的多核分布式飞行控制软件架构,架构共分为5层,从下至上依次为:操作系统层、硬件接口层、框架层、应用层和重用构件层.该架构目前已经应用于多个项目,具有可移植性好、可靠性高、实时性强等特点,能够降低项目研制时间与成本,提高飞行控制软件的可靠性与安全性.
针对超低轨道升力式航天器对地观察的优势及其高机动特性,设计了一种近地点位于临近空间的太阳同步冻结回归轨道,并对气动力辅助与发动机推力相结合的轨道保持策略进行了研究.策略将轨道保持过程分为3个阶段:第1阶段自远地点飞向大气层,不施加控制;第2阶段在大气层内飞行,通过控制攻角和倾侧角调整航天器所受气动力,小幅改变轨道的升交点赤经;第3阶段自跃出大气层到远地点,利用轨控发动机调整轨道参数,回到远地点时除升交点赤经其他轨道参数不变.以燃料最省为性能指标,对轨道保持策略进行了仿真分析,结果表明可以实现14.7天太阳同步冻结回归轨道的在轨运行.
针对多四旋翼无人机编队队形保持不变情况下的航迹优化问题,在飞行过程中以图论方法为基础将四旋翼无人机编队看作一个虚拟刚体,将编队中的跟随无人机看作领航四旋翼无人机的约束来处理,优化过程中仅考虑领航四旋翼无人机的最大转弯半径约束、队形变换时间约束及转换而来的约束,并将编队飞行所需能量最小作为优化目标,求解多四旋翼无人机编队定时到达目标的航迹优化问题,然后基于Gauss伪谱法对领航四旋翼无人机进行航迹优化,将其转化成一个非线性规划问题.仿真结果表明,利用本文提出的算法可以获得一条平滑飞行航迹,满足定时到达目的地约束、多无人机性能约束和环境约束要求;所获得的航迹能安全绕过障碍物,提高了无人机编队的生存能力,有一定的工程应用价值.
基于TDOA/FDOA对地面辐射源目标的定位原理,建立了多星TDOA/FDOA联合定位模型和定位精度分析模型.针对三星和四星系统不同星下点构型,仿真分析了星下点构型对目标定位精度的影响.三星星下点构型包括星下点直线型、等腰锐角三角形、等边三角形、等腰直角三角形、等腰钝角三角形等5种情况.四星星下点构型包括直线型、三角形、凹四边形和凸四边形等4种情况.分析结果表明,增加卫星数量能够提高探测区域内 目标的定位因子和定位精度;星下点非直线型的三星及四星构型,定位精度优于直线型构型,各卫星之间的基线长度尽可能长.
