航天器解体模型研究的新进展

航天器解体模型用于描述航天器因爆炸或碰撞解体所产生碎片的数量、尺寸、面质比以及速度等特性的分布规律,对于空间碎片环境建模与演化、空间碎片撞击风险评估及空间解体事件分析具有重要意义。分析了目前广泛使用的NASA标准解体模型的特点,介绍了近年来国内外在航天器解体模型方面的研究进展。国外的研究主要介绍了日本九州大学和德国EMI实验室开展的卫星撞击试验以及NASA最近启动的新一轮卫星撞击研究项目。国内的研究主要介绍中国空气动力研究与发展中心（CARDC ）开展的卫星撞击试验和仿真研究,以及在此基础上发展的CARDC-SBM航天器碰撞解体模型,并比较分析了该模型与 NASA 标准解体模型的差别。对当前航天器解体模型研究存在的不足进行了分析,提出了下一步应关注的研究方向。     

ACE Spacecraft

The Johns Hopkins University Applied Physics Laboratory (JHU/APL) was responsible for the design and fabrication of the ACE spacecraft to accommodate the ACE Mission requirements and for the integration, test, and launch support for the entire ACE Observatory. The primary ACE Mission includes a significant number of science instruments - nine - whose diverse requirements had to be factored into the overall spacecraft bus design. Secondary missions for monitoring space weather and measuring launch vibration environments were also accommodated within the spacecraft design. Substantial coordination and cooperation were required between the spacecraft and instrument engineers, and all requirements were met. Overall, the spacecraft was kept as simple as possible in meeting requirements to achieve a highly reliable and low-cost design. This revised version was published online in June 2006 with corrections to the Cover Date.     

An Overview of the Fast Auroral SnapshoT (FAST) Satellite

The FAST satellite is a highly sophisticated scientific satellite designed to carry out in situ measurements of acceleration physics and related plasma processes associated with the Earth's aurora. Initiated and conceptualized by scientists at the University of California at Berkeley, this satellite is the second of NASA's Small Explorer Satellite program designed to carry out small, highly focused, scientific investigations. FAST was launched on August 21, 1996 into a high inclination (83°) elliptical orbit with apogee and perigee altitudes of 4175 km and 350 km, respectively. The spacecraft design was tailored to take high-resolution data samples (or `snapshots') only while it crosses the auroral zones, which are latitudinally narrow sectors that encircle the polar regions of the Earth. The scientific instruments include energetic electron and ion electrostatic analyzers, an energetic ion instrument that distinguishes ion mass, and vector DC and wave electric and magnetic field instruments. A state-of-the-art flight computer (or instrument data processing unit) includes programmable processors that trigger the burst data collection when interesting physical phenomena are encountered and stores these data in a 1 Gbit solid-state memory for telemetry to the Earth at later times. The spacecraft incorporates a light, efficient, and highly innovative design, which blends proven sub-system concepts with the overall scientific instrument and mission requirements. The result is a new breed of space physics mission that gathers unprecedented fields and particles observations that are continuous and uninterrupted by spin effects. In this and other ways, the FAST mission represents a dramatic advance over previous auroral satellites. This paper describes the overall FAST mission, including a discussion of the spacecraft design parameters and philosophy, the FAST orbit, instrument and data acquisition systems, and mission operations.     

The New Horizons Spacecraft

The New Horizons spacecraft was launched on 19 January 2006. The spacecraft was designed to provide a platform for seven instruments designated by the science team to collect and return data from Pluto in 2015. The design meets the requirements established by the National Aeronautics and Space Administration (NASA) Announcement of Opportunity AO-OSS-01. The design drew on heritage from previous missions developed at The Johns Hopkins University Applied Physics Laboratory (APL) and other missions such as Ulysses. The trajectory design imposed constraints on mass and structural strength to meet the high launch acceleration consistent with meeting the AO requirement of returning data prior to the year 2020. The spacecraft subsystems were designed to meet tight resource allocations (mass and power) yet provide the necessary control and data handling finesse to support data collection and return when the one-way light time during the Pluto fly-by is 4.5 hours. Missions to the outer regions of the solar system (where the solar irradiance is 1/1000 of the level near the Earth) require a radioisotope thermoelectric generator (RTG) to supply electrical power. One RTG was available for use by New Horizons. To accommodate this constraint, the spacecraft electronics were designed to operate on approximately 200 W. The travel time to Pluto put additional demands on system reliability. Only after a flight time of approximately 10 years would the desired data be collected and returned to Earth. This represents the longest flight duration prior to the return of primary science data for any mission by NASA. The spacecraft system architecture provides sufficient redundancy to meet this requirement with a probability of mission success of greater than 0.85. The spacecraft is now on its way to Pluto, with an arrival date of 14 July 2015. Initial in-flight tests have verified that the spacecraft will meet the design requirements.     

基于模型的载人航天器研制方法研究与实践

载人航天器具有系统规模大、技术难度高、单件小批量、无法通过多次飞行持续完善设计、可靠性要求高等特点。当前载人航天器研制中仍存在着参数化和模型化程度不高、基于模型的系统综合仿真验证不足、研制各环节缺乏数字化集成等问题，传统基于文本的系统工程方法已无法满足研制需求，亟需采用基于模型的系统工程方法。本文针对载人航天器的研制现状和应用需求，提出了面向载人航天器全生命周期的模型体系，定义了需求模型、功能模型、产品模型、工程模型、制造模型、实做模型等六类模型，提出了基于模型的研制流程，包含系统设计闭环验证、产品设计闭环验证、实做产品闭环验证3个验证环节，并深入探索了各研制环节中不同模型间的传递与关联关系。以某型号载人航天器为应用基础，系统地验证了提出的方法。     

Solid rocket propulsion technology for de-orbiting spacecraft

This paper presents the topic of using solid rocket propulsion for de-orbiting spacecraft,in order to fulfil space debris mitigation requirements. The benefits and disadvantages of using such means are discussed. A dedicated system can be implemented in the satellite design phase and shall be a key subsystem of platforms inserted into orbit. Uncontrolled, semi-controlled and controlled de-orbit can be completed using solid rocket motors. Their impact on the space debris environment is discussed....     

Reliability evaluation of avionics system with imperfect fault coverage and propagated failure mechanisms

Fault tolerance designs are essential techniques for systems that require high levels of reliability, such as aircraft or spacecraft control system. Imperfect Fault Coverage (IFC) may lead to the failure of a system or subsystem even with adequate redundancy. Previous studies of IFC mostly concentrated on evaluating Coverage Factor (CF), whereas the system failure behaviors with IFC have rarely been involved. Failures that occur in low-layer may be covered by high-layer. However, if the coverage is imperfect, uncovered failure will have functional and physical impact on the system behavior. In this thesis, the failure behavior and reliability of IFC of multi-layer systems are studied and a Binary Decision Diagram (BDD)-based modeling and simulation method are proposed to evaluate system reliability. As a case, the failure behavior of an aero engine electronic controller with IFC is studied. The results show that the IFC may impact system behavior without taking the IFC into account, the system maintenance intervals may reduce, and thus the maintenance costs will increase.     

A Comparison of Solar-Cell and Battery-Type Power System for Spacecraft

The electrical power systems of orbiting unmanned spacecraft generally consist of energy-conversion devices in combination with energy-storage and conditioning components. The development of efficient pulse-duration modulation regulators suggests configurations smaller, lighter, and cheaper than conventional dissipative voltage regulators. These savings may or may not be realizable, depending upon the system reliability design goal and amount of redundancy required to meet the goal. Three spacecraft electrical power systems, each containing a solar-cell energy converter and using different voltage regulation schemes, are compared for a common mission specification. Each system is made to meet a given reliability goal by a technique that adds redundant components in a manner that minimizes a system design characteristic such as weight. The reliability design goal is kept constant as mission length is increased, permitting system comparison in terms of weight and cost as a function of time.     

低轨卫星可控式冗余分包遥测技术

随着卫星功能日趋复杂,遥测参数越来越多,处理要求越来越快,遥测设计面临越来越大的压力,必须拓展新的设计思路.本文讨论了一种低轨卫星可控式冗余分包遥测技术,更好地缓解遥测信道“瓶颈”,满足用户多样化的遥测需求,提高卫星遥测能力.该设计采用PCM（Pulse Code Modulation,脉冲编码调制）遥测和分包遥测相结合的方案,在信道资源有限,遥测参数庞大、用户需求多样的约束条件下,通过星载软件模式控制的方法实现了动态可编程遥测处理和遥测数传通道的冗余处理,提高了遥测信道的资源利用率、遥测设计的灵活性和系统的可靠性,该技术已成功应用于某在轨卫星,并被多颗在研卫星所采用.     

实时高可用测控集群管理控制节点设计

为提高飞行器试验任务中的实时测控数据处理能力,满足高并发、大容量、强实时、高可靠的试验数据处理需求,基于实时操作系统建立了一个计算功能强大、实时性强、可靠性高的测控集群系统。为实现实时测控集群系统的原型开发,通过研究负责节点管理和任务分配的集群管理控制节点的功能,分别对系统监控软件、负载均衡软件和人机交互软件的详细设计方法开展了研究,重点解决了节点信息收集、任务分配、负载均衡、双工冗余热备份、人机交互、节点控制等技术难题,保证了实时测控集群系统的实时性和可用性,为系统软件原型开发奠定了技术基础。     

Distributed flight control system using fiber distributed datainterface (FDDI)

The design of fault-tolerant distributed control systems is discussed. The application described is a generic flight control system (FCS) for a fighter aircraft. The system is designed for high redundancy and expandability to meet various requirements. The communication network is based on the fiber distributed data interface (FDDI), which provides a high level of electromagnetic interference resistance and reliability. The choice of fiber optic media also provides an extremely high bandwidth and tremendous growth capability for future communications needs. This study is not specific to FDDI or flight control systems. The methods and configurations presented should be applicable to many real-time control implementations     

新一代多用途载人飞船概念研究

在"神舟"载人飞船进入成熟稳定期后,中国有必要尽早启动新一代多用途载人飞船的论证和研制。本文对国外新一代载人飞船的技术方案特点、新的设计理念及发展现状进行了分析,从适应多任务、降低运营成本、钝头体气动外形、更高安全可靠性以及新型轻质材料使用等多个方面总结了国外新一代载人飞船的技术发展趋势。初步分析了中国发展新一代载人飞船的近地轨道、载人登月、载人登小行星、载人登火星等任务需求,基本确定了新一代飞船的总体性能参数,并在此基础上梳理了新一代载人飞船技术途径,初步提出了两种方案设想,为中国新一代载人飞船的研制提供参考。     

我国空间站运行的测控通信工作模式初探

在分析我国载人空间站在轨运行的任务特点和对测控通信提出的新要求的基础上,对测控通信工作模式进行了探讨。针对任务中心,采用分布式模式进行任务的组织和实施,建立常态化工作体系;针对测控通信资源的使用,对资源进行简化、优化,发挥天、地基各自优势,并实现天地基资源统一调度;针对在轨故障诊断和应急处置,提出了“天地联合,以地为主”的载人航天故障诊断模式和实现途径;针对空间站遥操作需求,提出了一套工作模式和实施流程;针对测控通信长期执行任务的可靠性,提出了任务中心容灾备份工作模式和通信路由的备份模式。以期为后续任务工作模式设计提供参考。     

基于应力强度干涉理论和薄弱环节分析的光子晶体光纤陀螺寿命评估

在空间飞行环境中,航天器承受着各种环境的作用,而每种环境因素都在一定程度上影响着航天器的工作寿命。光子晶体光纤是一种新型光纤,其比传统保偏光纤更耐辐照,是长寿命光纤陀螺(Fiber Optical Gyroscope,FOG)的首选,可以满足长寿命卫星的应用需要。将光纤陀螺特征寿命定义为强度,将热、辐照和振动等环境因素定义为应力,运用应力强度分析理论,采用最坏情况分析方法,分析了在常见应力联合作用下光纤陀螺的薄弱环节,评估了光纤陀螺的在轨工作寿命,验证了光子晶体光纤陀螺在某长寿命通信卫星上的适应性。分析结果表明,热和辐照是影响光子晶体光纤陀螺的重要因素。研究结论可用于有针对性地进行改进设计,为长寿命高可靠卫星提供技术支撑,具有显著意义。     

A Kilowatt Rotary Power Transformer

Many future satellite configurations will require the transfer of signals and power across rotating interfaces. Satellite systems are particularly cost effective for both commercial and military communications applications if their useful lifetime can be demonstrated to be greater than five years. The rotary transformer has the desireable characteristics of high reliability and low noise which qualify it as a potential replacement for slip rings. This paper describes the development of a rotary transformer for typical spacecraft applications. The transformer is built in modular sections, each capable of transferring 500 watts across a gap at efficiencies greater than 88 percent, dc to dc. The design effort included a study of pertinent electrical characteristics required for typical spacecraft configurations, electrical design and analyses of the overall dc-dc converter, mechanical design of the transformer cores and their assembly, and a study of transformer core and winding characteristics. Breadboard test results have demonstrated the desired level of efficiency, satisfactory operation over temperature, and satisfactory electrical characteristics.     

航天智能控制技术让运载火箭“会学习”

高可靠和智能化是未来智能航天器的主要特点,本文聚焦航天器高可靠、智能化的发展需求。梳理了中国运载火箭从无到有、从有到全的发展历程,提出了航天智能技术从航天器的可靠性做起,航天器的可靠性从航天智能控制做起,航天智能控制从"会学习"的火箭做起。围绕航天智能控制技术如何使运载火箭"会学习"的发展架构,进一步探索了"边飞边学"和"终身学习"智能控制技术的理论研究和应用现状,支撑中国"会学习"运载火箭高可靠和智能化的发展。     

航天器控制高级语言的设计与实现

针对航天测控任务中存在的卫星控制人员手工操作多、任务准备周期长、自动化水平低等问题,在分析航天器控制流程的基础上,抽象出面向卫星控制人员的航天器控制高级语言,设计语言规范,描述Windows平台的编辑、编译、运行功能的设计与实现。该语言成功应用于某中心的卫星测控任务中,任务准备快速高效,大大减轻了卫星控制人员的工作量,任务过程自动执行,无需人工干预,提高了航天器控制过程的自动化水平。     

高超音速地面模拟设备的研究进展

首先介绍了航天器再入飞行中遇到的真实气体效应问题及其对航天器气动特性所产生的影响。然后对五六十年代以来高超音速地面模拟试验地位的转换进行了简要分析。以空气助力的轨道间运送器、美国国家空天飞机和火星探测器的再入飞行为例,探讨了再入流动对地面模拟设备的设计要求,并回顾了各国在此方面的主要研究进展,最后指出了该领域在未来的发展方向。     

下一代数据中继卫星系统发展思考

通过系统阐述中继卫星系统的发展过程,给出了主要国家和组织的中继卫星系统技术体制和现状.再结合卫星、载人航天器和深空探索的未来发展趋势,分析了下一代中继卫星系统的发展需求.在此基础上,从体系结构、卫星平台、链路调制体制、网络协议等方面,探讨并给出了下一代中继卫星系统的发展趋势和技术途径.为满足未来近地、深空航天任务,以及临近、低空快速移动用户的不同要求,节约系统成本,下一代中继卫星系统将向专业化和与其他系统融合的方向发展:星间链路将增加激光链路,数据速率可达到10 Gbit/s以上;多址业务成为主用,同时支持用户数能力将极大提高;对于链路调制体制,在采用CR(Cognitive Radio,认知无线电)和SDR(Software Defined Radio,软件定义无线电)技术的基础上,可实现实时自适应调整和根据需求加载配置;数据传输将采用网络化方式,天地间构成一体化DTN(Delay Tolerant Network,容延迟网络).     

NASA 二级轻气炮设备简介

随着人类航天活动日益频繁,地球轨道上空间碎片总数逐年增长。航天器表面空间碎片防护工作受到各航天大国的高度重视。航天器针对毫米级空间碎片主要采用被动防护方式。超高速撞击实验是防护方案设计工作的基础。NASA 在毫米级弹丸超高速撞击实验中采用的主要发射装置为二级轻气炮。本文对美国 NASA 和相关单位二级轻气炮设备及其未来发展趋势进行简要介绍,同时对我国相关单位超高速撞击实验设备进行分析,并对发展趋势进行讨论。