Link-16战术数据链抗干扰性能评估与仿真

分析了JTIDS系统的结构与特点,重点探讨了其抗干扰性能,在Simulink仿真平台实现了对JTIDS系统的仿真,给出了具体的仿真模型并对模型中的每个模块进行了较为详细的说明,对仿真结果进行了简要分析.     

星载导航软件产品线设计方法与应用

在分析星载导航软件存在问题的基础上,从配置项级、架构级和构件级三个维度阐述了软件产品线的设计方法,结合典型的航天器产品应用实例,介绍了产品线的复用过程,并借助软件一体化平台工具,详细描述了星载导航软件产品线的建立、应用和管理方法。通过实践摸索,基本解决了软件研发中存在的问题,实现了产品线开发和项目开发分离,研制能力得到了提升。星载导航软件产品线经过不断迭代,逐步提高了复用率和成熟度,确保了软件产品质量。     

影响航天器陀螺仪装调合格率的因素研究

根据速率积分陀螺仪在装调过程中出现的问题,在详细分析影响装调精度的原因基础上,通过工艺试验实施了验证,确认了电机自检频率问题出现的原因,并提出了解决问题的工艺措施,改进了工艺实施方案。经过总装生产实践,验证了该改进方案的可行性、可靠性,表明了该技术方案合理,有效地满足了航天器用速率积分陀螺仪装调合格率的要求,大幅度提高了生产效率。     

CCD太阳敏感器误差分析与补偿

介绍了星载CCD太阳敏感器的工作原理,用空间几何模型阐述了线阵CCD太阳敏感器测量两轴姿态信息的方法,进行了误差分析,建立了系统的误差补偿模型。通过标定试验,确定了模型参数,进行了物理补偿,同时针对系统的残差,提出了数学补偿的方法。测试结果表明敏感器实现了较高的测量精度,验证了误差补偿方法的有效性。     

论航天企业质量文化建设

回顾了航天科工集团公司质量文化建设历程,介绍了航天企业预防型质量文化建设的3个阶段,分析了航天企业质量文化精神、行为、物质3个层次的结构特点,并阐述了其内涵,叙述了不同阶段的质量文化建设任务,分析了质量文化与质量经营思想的动态契合性,在此基础上对当前航天企业质量文化建设提出了一些建议。     

SiP陶瓷芯片水汽含量控制工艺方法研究

结合宇航级陶瓷封装芯片的特点和水汽控制要求,对SiP陶瓷封装工艺进行了研究,分析了陶瓷封装过程中产生水汽的各个因素,优化了封装生产工艺流程,增加了预烘转运过程下厂监制,保证了转运时间的工艺过程管控,通过试验验证了管控措施的有效性.     

让青春奏响绚丽的乐章—东北地区空管局党委引导团员青年成才纪实

近年来,东北地区空管局党委在和谐空管建设中,着力引导团员青年成长成才,围绕安全生产和服务保障,充分发挥了团员青年的生力军突击队作用;围绕和谐空管和企业文化建设,充分发挥了团员青年的参与和促进作用,让团员青年在本职岗位上充分展露了才华,实现了人生价值,用青春奏响了一曲曲绚丽的乐章。     

SFMEA方法在飞行控制软件中的应用

为提高飞行控制软件的安全性,分析了飞行控制软件的安全性薄弱环节,探讨了影响飞行控制软件安全性的隐患。简要阐述了软件失效模式、影响(SFMEA)方法,给出软件失效模式分类的方法以及飞行控制软件常见的失效模式,给出适用于航天飞行控制软件的软件危害性等级,结合某型号飞行控制软件进行了SFMEA方法的尝试性应用,分析了软件失效的局部影响和最终影响,形成了52个SFMEA的分析表格,总结了302个飞行控制软件的失效模式,发现了部分安全性薄弱环节,并提出了相应的改进措施,结果表明,SFMEA方法对提高飞行控制软件的安全性有一定的工程价值。     

信息高速公路及其标准化

简述了信息高速公路的提出及其含义,介绍了国内外信息高速公路的发展动态,分析了信息高速公路可能产生的效益,论述了开展信息高速公路标准化的意义,并提出了其标准化的设想。     

压制式干扰对SAR成像质量的影响初步分析

研究了几种压制干扰样式,建立了SAR雷达干扰方程,分析了SAR雷达回波信号仿真方法,计算了干扰信号对SAR成像的干扰效果.     

镁铝贫氧推进剂的能量分析

利用冲压发动机热力计算程序，对镁铝中能贫氧推进剂的能量特性进行了系统研究，并探讨了高能级分对该类贫氧推进剂能量性能的影响，研究结果表明，在镁铝中能贫氧推进剂中增加镁粉含量（或减少铝粉含量）贫氧推进剂的比冲下降，在一定空燃比范围内，增加空燃比有助于提高冲压发动机的比冲。提高贫氧推进剂中ＣＬ－２０和硼粉含量，可以显著提高共能量，而采用叠氮类含能粘合剂取代惰性粘合剂时贫氧推进剂的比冲降低。     

Analysis of the Accuracy of Ballistic Descent from a Circular Circumterrestrial Orbit

The problem of the transportation of the results of experiments and observations to Earth every so often appears in space research. Its simplest and low-cost solution is the employment of a small ballistic reentry spacecraft. Such a spacecraft has no system of control of the descent trajectory in the atmosphere. This can result in a large spread of landing points, which make it difficult to search for the spacecraft and very often a safe landing. In this work, a choice of a compromise scheme of the flight is considered, which includes the optimum braking maneuver, adequate conditions of the entry into the atmosphere with limited heating and overload, and also the possibility of landing within the limits of a circle with a radius of 12.5 km. The following disturbing factors were taken into account in the analysis of the accuracy of landing: the errors of the braking impulse execution, the variations of the atmosphere density and the wind, the error of the specification of the ballistic coefficient of the reentry spacecraft, and a displacement of its center of mass from the symmetry axis. It is demonstrated that the optimum maneuver assures the maximum absolute value of the reentry angle and the insensitivity of the trajectory of descent with respect to small errors of orientation of the braking engine in the plane of the orbit. It is also demonstrated that the possible error of the landing point due to the error of specification of the ballistic coefficient does not depend (in the linear approximation) upon its value and depends only upon the reentry angle and the accuracy of specification of this coefficient. A guided parachute with an aerodynamic efficiency of about two should be used at the last leg of the reentry trajectory. This will allow one to land in a prescribed range and to produce adequate conditions for the interception of the reentry spacecraft by a helicopter in order to prevent a rough landing.     

Optimization of the Flight of a Spacecraft with a Solar Sail from the Earth to Mars with a Perturbation Maneuver near Venus

The trajectories of the fastest flight of a spacecraft (SC) with a solar sail from the Earth's sphere of activity to the Martian sphere of activity including the section of a perturbation maneuver near Venus are investigated. The planetary spheres of activity are assumed to be point-like; i.e., the maneuver section and the initial and final positions of the SC coincide with the corresponding positions of the planets. The initial velocity of the SC is assumed to be equal to the Earth's velocity, so that no leveling of the velocities of the SC and Mars in the final point of the flight is required. The perturbation maneuver is considered as a jump of the heliocentric velocity of the SC at the point of its contact with Venus, which does not change the magnitude of its Venus-centric velocity. The orbits of planets are assumed to be circular and coplanar; the SC trajectory lies at the plane of these orbits. The sail is planar with a specularly reflecting surface. The trajectories of optimum flights are determined as a result of solving the boundary value problem of the Pontryagin maximum principle. The families of solutions to this problem depending on the initial angular positions of Venus and Mars are constructed by the method of continuation over a parameter.     

Spacecraft dynamics with regard to elastic gravitational stabilizer deployment

The present paper deals with the study the dynamics of the spacecraft with gyro-gravitational system of stabilization. The deployment of the boom of the gravitational stabilizer commences after placing the spacecraft into the orbit and completion of the preliminary damping, when the gyroscopes are uncaged. Primarily the boom is the pre-stressed tape wound on the special drum. When the drum starts deploying the tape, it turns into the elastic cylindrical rod with the mass at its tip. The objective of the study is the creation of the generalized mathematical model and the conducting of the computer modelling of the spacecraft dynamics. The equations of motion are worked out with the use of the Lagrangian formalism. The numerical simulation of typical modes of system functioning is conducted. It is shown that the folding and the following deployment of the boom result in the turn of the spacecraft by 180° about the axis of the pitch. The results illustrate the behaviour of the main system variables.     

The Use of Quaternions in the Optimal Control Problems of Motion of the Center of Mass of a Spacecraft in a Newtonian Gravitational Field: I

The problem of optimal control is considered for the motion of the center of mass of a spacecraft in a central Newtonian gravitational field. For solving the problem, two variants of the equations of motion for the spacecraft center of mass are used, written in rotating coordinate systems. Both the variants have a quaternion variable among the phase variables. In the first variant this variable characterizes the orientation of an instantaneous orbit of the spacecraft and (simultaneously) the spacecraft location in this orbit, while in the second variant only the instantaneous orbit orientation is specified by it. The suggested equations are convenient in the respect that they allow the general three-dimensional problem of optimal control by the motion of the spacecraft center of mass to be considered as a composition of two interrelated problems. In the first variant these problems are (1) the problem of control of the shape and size of the spacecraft orbit and (2) the problem of control of the orientation of a spacecraft orbit and the spacecraft location in this orbit. The second variant treats (1) the problem of control of the shape and size of the spacecraft orbit and the orbit location of the spacecraft and (2) the problem of control of the orientation of the spacecraft orbit. The use of quaternion variables makes this consideration most efficient. The problem of optimal control is solved on the basis of the maximum principle. Several first integrals of the systems of equations of the boundary value problems of the maximum principle are found. Transformations are suggested that reduce the dimensions of the systems of differential equations of boundary value problems (without complicating them). Geometrical interpretations are given to the transformations and first integrals. The relation of the vectorial first integral of one of the derived systems of equations (which is an analog of the well-known vectorial first integral of the studied problem of optimal control) with the found quaternion first integral is considered. In this paper, which is the first part of the work, we consider the models of motion of the spacecraft center of mass that employ quaternion variables. The problem of optimal control by the motion of the spacecraft center of mass is investigated on the basis of the first variant of equations of motion. An example of a numerical solution of the problem is given.     

固体火箭超燃冲压发动机补燃室构型的影响分析

针对不同补燃室结构参数对固体火箭超燃冲压发动机补燃室掺混燃烧性能的影响进行研究,分析各级燃烧室的长度与扩张角度对补燃室性能的影响。采用基于密度的二阶迎风格式对补燃室掺混燃烧进行模拟,湍流模型和燃烧模型分别采用SST k-ω模型和涡团耗散模型。结果表明,提高燃烧效率与降低总压损失是相互矛盾的;燃烧效率随燃烧室长度的增加而增大,随燃烧室扩张角度的增加而减小;总压恢复系数随燃烧室长度的增加而减小,随燃烧室扩张角度的增加而增大;一级燃烧室的结构参数对燃烧效率与总压恢复系数的影响最大。当补燃室的总长与出口面积一定时,以发动机的总体性能参数作为补燃室构型的优化目标,对一、二级燃烧室长度与一、三级燃烧室扩张角度进行优化。     

Determination of kinematic state of an orbiting multibody using GNSS signals

Precise attitude determination of the members of a free-flying multibody system is a not so immediate task, due essentially to the large motion of its appendages coupled with their relevant flexibility effects. In fact, sensors used to this aim in current projects, such as optical encoders usually positioned near the joints of each arm, are almost blind to these effects, and clusters of specific redundant sensors should, therefore, be required in order to reconstruct both elastic deformations and rigid motion.Satellite navigation systems (GNSS) offer a suitable and reliable solution to this problem. To exploit the phase of the signal, instead of the traditional pseudo random code, ensures a very high accuracy of the order of magnitude of centimeter. Such a process requires the solution of an initial ambiguity problem, related to the number of integer wavelength included in the length of the member.The aim of the paper is to investigate the capability of this GNSS based technique to reconstruct the kinematics of a flexible multibody system orbiting around the Earth. This analysis requires a simulation including both the multibody dynamics and the navigation system constellation to define the satellites lines-of-sight at each time step.Concerning multibody equations of motion, a Newtonian formulation is adopted in this work. A special attention is required about the choice of the state variables. As the internal forces are associated to the relative displacements between the bodies, which are small fractions of the distance of the multibody spacecraft from the center of the Earth, the task of obtaining these forces from inertial coordinates could be impossible from a numerical point of view. So, the problem is reformulated in such a way that the equation of motion of the system contains global equations, with no internal forces, and local equations, with internal forces. In the latter, only quantities of the same order of the spacecraft dimensions are present.Accuracies achievable in LEO orbit with current GPS and upcoming Galileo systems are evaluated to show the interest of the proposed technique.     

Generalized restricted circular three-body problem as a model for dynamics of binary asteroids

Exploration of the Solar System has recently revealed the existence of a large number of asteroids with satellites, which has stimulated interest in studying the dynamics of such systems. This paper is dedicated to the analysis of the relative motion of a binary asteroid. The conditions of existence of such a system (i.e., when its components do not run away) are derived in the Introduction. Then it is assumed that the satellite has no significant effect on the motion of the main asteroid, the latter being modeled as a dumbbell-like precessing solid body. The equations of motion of this system are a two-parameter generalization of the corresponding equations of the restricted circular three-body problem. It is demonstrated that in the system under consideration there exist steady-state motions in which the small asteroid is equidistant from attracting centers at the ends of the dumbbell (an analog to triangle libration points). The conditions of existence of such motions are derived, and the positions with respect to the dumbbell are analyzed in detail. Examination of the stability of the triangle libration points is reduced to investigation of a characteristic equation of the sixth degree. The stability conditions are derived in the case when the main asteroid executes near-planar motion.     

Uncontrolled Attitude Motion of the Foton-12 Satellite and Quasi-Steady Microaccelerations onboard It

The results of determination of the uncontrolled attitude motion of the Foton-12 satellite (placed in orbit on September 9, 1999, terminated its flight on September 24, 1999) are presented. The determination was carried out by the onboard measurement data of the Earth's magnetic field strength vector. Intervals with a duration of several hours were selected from data covering almost the entire flight. On each such interval the data were processed simultaneously using the least squares method by integrating the satellite's equations of motion with respect to the center of mass. The initial conditions of motion and the parameters of the mathematical model employed were estimated in processing. The results obtained provided for a complete representation of the satellite's motion during the flight. This motion, beginning with a small angular velocity, gradually sped up. The growth of the component of the angular velocity with respect to the longitudinal axis of the satellite was particularly strong. During the first several days of the flight this component increased virtually after every passage through the orbit's perigee. As the satellite's angular velocity increased, its motion became more and more similar to the regular Euler precession of an axisymmetric rigid body. In the last several days of flight the satellite's angular velocity with respect to its longitudinal axis was about 1 deg/s and the projection of the angular velocity onto the plane perpendicular to this axis had a magnitude of approximately 0.15 deg/s. The deviation of the longitudinal axis from the normal to the orbit plane did not exceed 60°. The knowledge of the attitude motion of the satellite allowed us to determine the quasi-steady microacceleration component onboard it at the locations of the technological and scientific equipment.     

空间太阳能电站姿态运动-结构振动耦合建模与分析

针对多旋转关节空间太阳能电站构型，利用基于能量等效原理的连续体等效方法将其等效为柔性梁模型，并考虑重力梯度影响，建立了姿态运动与结构振动的耦合动力学模型；结合Runge-Kutta 法和Newmark法的优点，提出了适用于求解姿态运动与结构振动耦合动力学方程的改进算法，相比于经典Runge-Kutta 法大幅提高了效率；利用改进算法得到了不同参数下的动力学响应。在此基础上，推导了结构振动量级随结构尺寸的六次方量级增加的规律，仿真结果表明尺寸过大引发不稳定现象；分析了姿态运动和重力梯度对结构振动频率和振幅的影响；发现了姿态运动周期受结构柔性影响而增大的现象，这种现象在低轨以及大初始姿态角下影响更为明显。