ESA/ESTEC的空间环境试验能力研究

这里的空间环境主要指电子、质子、离子、太阳紫外、原子氧、碎片、极端温度、污染等环境,这些环境在航天器中产生总剂量效应、单粒子效应、充放电效应等各种有害效应,甚至会引发航天器故障与异常。鉴于空间环境不利影响的严重性和复杂性,欧洲空间局（ESA）在欧洲空间技术研究中心（ESTEC）的产品保证与安全部门建立了空间环境试验室,目的为ESA航天器的空间环境防护提供先进的试验验证手段。文章介绍ESA/ESTEC的空间环境地面试验能力,包括空间环境模拟设备、测试仪器及其试验相关的标准;介绍ESTEC航天器研制组织体系及其空间环境试验室所在的产品保证与安全部门的职能和作用,分析研究了这些部门及空间环境试验室对ESA航天器质量、可靠性、安全性的基础保证作用;最后就完善我国空间环境试验能力提出建议。     

航天器力学试验回顾、启发与展望

中国航天器的研制与火箭的发展密切相关,航天器力学试验也是借鉴了火箭研制中力学试验的经验。振动试验至今仍沿用了x、y、z 3个方向分别进行振动试验的方法;1997年ESTEC建立了HYDRA 6自由度振动试验系统,开始了以x向为主的6自由度振动试验的新方法,这对我国航天器振动试验技术是一个重要的启迪。     

KM6太阳模拟器冷却系统

太阳模拟器是在冷黑和真空环境中模拟太阳辐照的一种重要设备,其冷却系统主要用于太阳模拟器氙灯的冷却和光学组件的冷却。冷却系统主要由高压水冷却、低压水冷却、低压氮气冷却、氟里昂制冷和去离子水等5部分组成,文章分别介绍了它们的流程和特点。文章参考了ESTEC冷却系统的经验,对KM6太阳模拟器冷却系统进行了优化设计,特别采用了去离子水、密闭循环、模块化机组、PCL控制系统等多种技术,提高了系统的可靠性。此系统设计方法可以应用到其他类似太阳模拟器冷却系统和大型常温冷却系统的设计。     

Family of deployable/retractable structures for space application

New trends in the frame of space applications lead to the necessity of using deployable/retractable structures, working either as beams (with payloads all along their length) or masts (loaded at their tip). SENER—under ESA/ESTEC and Spanish Space Program contracts—are developing a family of structures with deployment and retraction capabilities (LTS, SENERMAST, CTM) so as to cover all ranges of potential necessities in the space community (antennas, experiment support, solar arrays, heat rejection systems …). This paper consists of a summary of the performances and range of applications of LTS, SENERMAST and CTM, and pays special attention to the large truss structure (LTS) development and verification.     

Space optical payloads, new application area for high temperature composites

High temperature composites have been extensively developed in order to produce thermal protection systems of reusable re-entry vehicles and launchers. This development effort covers all aspects including sizing, design, manufacturing processes characterization, non destructive inspection, and all industrial facilities which have also been installed. Strong interest recently appeared for these materials to meet requirements for different space applications. In particularly, for more stringent optical payloads, new materials with high performance requirements have appeared. In the field of high dimensionally stable structures for telescopes, materials have to meet severe requirements, such as low coefficients of thermal expansion, good specific modulus, long-term stability (moisture and chemical insensitivity), etc. Carbon/carbon (C/C) composites can meet these specifications. To demonstrate this capability a structure has been designed, manufactured and will be submitted for complete testing (work supported by ESA/ESTEC). The main available results (part feasibility, characterizations, analysis and stability performance budgets) are presented. For future telescope mirrors, silicon carbide is already known as a good candidate. However, an innovative concept based on silicon carbide sandwich honeycomb technology, which allows optimized design, has been developed. The first characterization results and manufacturing capabilities are presented.     

Thermomechanical behaviour of CFRP tubes for space structures

Carbon fiber reinforced plastic (CFRP) tubes, with the increasing dimensions and performances requested for space structures, are becoming a basic building element of boom-type structures for large precision reflectors, towers and payload support structures such as the Modular Payload Support Structure, the Shuttle Pallet Satellite or the European Retrievable Carrier. It is very important for such applications that the CFRP tubes have minimum thermal distortions and very high stiffness.An extensive test program was performed to characterise the CFRP tubes that are used for such applications. Measurements of coefficient of thermal expansions, thermal conductivity, thermal cycling, microstructure behaviour, as well as mechanical tests and outgassing tests were performed. The main purpose was to correlate the microcracking with the thermal cycling and the coefficient of thermal expansion and thermal conductivity.These types of activities for the CFRP tubes were performed for the first time in Europe and important results were found, especially in the area of microcracking generation and correlation with engineering parameters. The influence of the thermal cycling speed on the microcracking was also studied. Most of the tests were conducted at ESTEC (European Space Research & Technology Centre, Holland) by the European Space Agency in the frame of the technology research activities.     

Microgravity measurement assembly (MMA) for spacelab module missions

The microgravity measurement assembly (MMA) is a precision measurement facility for ground and on-orbit disturbance accelerations on board Spacelab, being currently under development by MBB/ERNO under DFVLR contract. MMA is using a new generation of micromechanical acceleration detectors developed by CSEM under ESTEC contract. Small dimensions of the triaxial sensor packages allow for installation very close to scientific experiments; mass is significantly reduced compared to conventional systems. Six or more of these mini-sensor packages are installed at the most g-sensitive experiments of Spacelab Module Missions. Acceleration and housekeeping data are processed in real time by a dedicated microcomputer and transmitted to the ground. Thus, for the first time, synchronized and comparable precision acceleration data are available in real time on ground for on-line judgement of the microgravity environment desired for experiment success, offering the possibility, for example of experiment repetition in case of excessive g-disturbances. Furthermore, MMA allows for immediate feedback to the crew concerning the microgravity effects of their dynamic behavior, with the aim of crew training towards lower disturbances. An additional mobile sensor package can be installed at vibration sources, e.g. pumps, centrifuges etc. or any arbitrary location inside the Spacelab Module. An impact hammer can be used together with MMA in order to measure in-flight structural transfer functions. The MMA on-board system and ground station and its planned utilization for the German Spacelab Mission D-2 is described.     

[Interior] Configuration options,habitability and architectural aspects of the transfer habitat module (THM) and the surface habitat on Mars (SHM)/ESA's AURORA human mission to Mars (HMM) study

This paper discusses the findings for [Interior] configuration options, habitability and architectural aspects of a first human spacecraft to Mars.In 2003 the space architecture office LIQUIFER was invited by the European Space Agency's (ESA) AURORA Program committee to consult the scientists and engineers from the European Space and Technology Center (ESTEC) and other European industrial communities with developing the first human mission to Mars, which will take place in 2030, regarding the architectural issues of crewed habitats.The task was to develop an interior configuration for a transfer vehicle (TV) to Mars, especially a transfer habitation module (THM) and a surface habitat module (SHM) on Mars. The total travel time Earth—Mars and back for a crew of six amounts to approximately 900 days. After a 200-day-flight three crewmembers will land on Mars in the Mars excursion vehicle (MEV) and will live and work in the SHM for 30 days. For 500 days before the 200-day journey back the spacecraft continues to circle the Martian orbit for further exploration. The entire mission program is based on our present knowledge of technology. The project was compiled during a constant feedback-design process and trans-disciplinary collaboration sessions in the ESA-ESTEC concurrent design facility.Long-term human space flight sets new spatial conditions and requirements to the design concept. The guidelines were developed from relevant numbers and facts of recognized standards, interviews with astronauts/cosmonauts and from analyses about habitability, sociology, psychology and configuration concepts of earlier space stations in combination with the topics of the individual's perception and relation of space.Result of this study is the development of a prototype concept for the THM and SHM with detailed information and complete plans of the interior configuration, including mass calculations. In addition the study contains a detailed explanation of the development of the Design process including all suggested design and configuration options.     

European Sail Tower SPS concept

Based on a DLR-study in 1998/99 on behalf of ESA/ESTEC called “System Concepts, Architectures and Technologies for Space Exploration and Utilization (SE&U)” a new design for an Earth-orbiting Solar Power Satellite (SPS) has been developed. The design is called “European Sail Tower SPS” and consists mainly of deployable sail-like structures derived from the ongoing DLR/ESA solar sail technology development activity. Such a SPS satellite features an extremely light-weight and large tower-like orbital system and could supply Europe with significant amounts of electrical power generated by photovoltaic cells and subsequently transmitted to Earth via microwaves. In order to build up the sail tower, 60 units - each consisting of a pair of square-shaped sails - are moved from LEO to GEO with electric propulsion and successively assembled in GEO robotically on a central strut. Each single sail has dimensions of 150m × 150 m and is automatically deployed, using four diagonal light-weight carbon fiber (CFRP) booms which are initially rolled up on a central hub. The electric thrusters for the transport to GEO could also be used for orbit and attitude control of the assembled tower which has a total length of about 15 km and would be mainly gravity gradient stabilized. Employing thin film solar cell technology, each sail is used as a solar array and produces an electric power in orbit of about 3.7 MW_e. A microwave antenna with a diameter of 1 km transmits the power to a 10 km rectenna on the ground. The total mass of this 450 MW SPS is about 2100 tons. First estimates indicate that the costs for one kWh delivered in this way could compete with present day energy costs, if launch costs would decrease by two orders of magnitude. Furthermore, mass production and large numbers of installed SPS systems must be assumed in order to lower significantly the production costs and to reduce the influence of the expensive technology development. The paper presents the technical concept and an economic assessment as well as results of a recent solar sail deployment ground demonstration at DLR's facilities in Cologne.     

基于dSPACE的挠性卫星姿轨控实时仿真系统

以某在研挠性卫星星载计算机在回路中的仿真试验为背景 ,利用dSPACE多处理器系统构建了卫星姿轨控实时仿真系统。介绍了实现该系统的软、硬件构成。重点介绍了实现该实时系统的关键———如何用dSPACEDS4 2 0 1s板上的RS2 32 / 42 2串口模拟星载计算机与各路姿态敏感器、反作用飞轮串口之间的复杂通信及串口通信Simulink仿真模型的设计。本文用Simulink ,Stateflow和S函数混合建模方法 ,设计了该实时系统的串口通信接口模型。并以星上自主模式为例进行了实时仿真。仿真结果证明了本文设计的卫星姿轨控实时仿真系统的可行性和串口通信接口模型设计的有效性。     

智能仿真软件设计平台的实现

以运载火箭控制系统仿真软件为对象,通过分析仿真模型的共性,建立通用模型数据库和函数库,完成仿真软件各组成部分通用模块设计,搭建了一个智能化的专用仿真软件设计平台。平台的实现突破传统的软件设计思路,用仿真模型描述代替人工代码输入,在缩短仿真软件研制周期的同时也提高了软件的设计质量,为仿真软件设计创造了一个良好的开发环境。     

分布式SAR相位同步误差的影响分析与试验验证

相位同步是分布式SAR系统必须考虑的突出问题之一。建立了双向脉冲交换相位同步全数字仿真模型,仿真得到相位同步误差信号。基于全数字仿真系统分析了不同类型相位同步误差对InSAR测高性能的影响。最后利用地面半实物仿真试验得到实际分布式SAR系统相位同步误差及其对InSAR测高的影响,实测相位同步误差标准差为2.6度,引起的InSAR相对测高精度损失在厘米量级;并且半实物试验结果与全数字试验结果一致,表明了相位同步方案及其影响分析的正确性。本文工作对分布式SAR相位同步分系统的设计与实现具有重要指导意义。     

基于Perl和OCEAN的集成电路仿真分析方法

提出一种集成电路仿真分析方法。将Perl语言和OCEAN语言应用于集成电路(IC)仿真中,自动生成一组完备、准确的测试文件,以批处理和远程非图形用户界面的方式完成大量长时间的电路仿真,并利用所得大量仿真结果进行曲线绘制、拟合、信息提取。将方法运用于某双系统导航接收机中某衬底偏置低压射频混频器的仿真分析,结果表明方法高效、有效、准确、便捷,可用于航空航天领域芯片设计的仿真分析。     

基于CAD三维模型几何特征简化的卫星桁架结构快速建模仿真与试验验证

为提高卫星桁架结构设计过程的仿真建模质量,优化仿真过程对三维模型信息的利用方式和效率,文章提出一种基于CAD三维模型几何特征简化的卫星桁架结构快速建模仿真方法,并给出分别采用梁单元和壳单元建模的路径及软件界面。将该方法应用于某卫星结构的仿真建模,得到了三维有限元模型模态分析数据,其与力学试验数据的对比结果表明:横向模态分析与试验误差最大,为6.07%;各方向误差均满足指标要求。基于CAD三维模型几何特征简化处理的卫星桁架结构快速建模仿真方法合理可行,建模过程便捷、高效,数据可信。     

航天训练模拟视景的光照仿真方法研究

针对航天员太空观察感知的视景仿真,提出了一种大规模场景下航天器的光照仿真实现方法。通过对航天器的真实光照环境进行建模分析,在此基础上实现了太空场景的光照仿真,并结合太空光照环境特点采用色调映射算法进行了光照亮度调节。光照一致性分析结果表明,航天训练模拟器的视景仿真实现在满足实时性指标的前提下,有效地提高了太空场景仿真的逼真度,所采用的亮度调节方法也切实可行。这种光照仿真方法不仅适用于各种空间活动的视景仿真,对于大场景中的天文景观的仿真也同样可以达到较好效果。     

基于VERICUT虚拟加工的应用

VERICUT是美国CGTECH公司开发的世界上先进的数控加工仿真软件。本文以VERICUT6．1．2软件为平台，构建了VMC10数控加工仿真系统。通过一个典型实例展现了仿真、优化的全过程，这将为构建其它数控加工仿真系统提供参考。     

空间电子、质子和紫外综合辐照模拟试验研究

文章介绍了地面模拟地球同步轨道15年电子、质子和紫外环境的综合辐照模拟试验技术,为长寿命卫星热设计及热控涂层选型提供可靠依据。试验采用空间低能综合环境试验设备、太阳吸收率原位测试系统,针对卫星各种表面材料如S781白漆、SR107-ZK白漆、F46镀银和OSR二次表面镜等进行了空间低能综合辐照试验,与已有的飞行试验数据进行对比研究,结果表明：本次试验能够较准确地反映航天器在轨道上材料的退化现象。     

液体火箭发动机动态特性仿真技术研究进展

仿真技术是研究液体火箭发动机动态特性的主要手段之一。总结了国内外液体火箭发动机动态特性仿真技术的研究进展,将仿真技术的研究进展分为三个阶段:专用仿真程序阶段、通用仿真软件阶段和多学科联合仿真阶段。对各个阶段的进展进行了总结和评述,分析了各阶段仿真技术的主要特点。在总结研究进展的基础上,对今后发动机动态特性仿真技术的发展方向提出了设想。     

一体化建模和仿真环境（IMASE）

一体化建模／仿真环境能为用户提供友好的仿真建模环境，交互功能强的仿真运行环境和功能充分的结果分析环境，从而使仿真试验全过程都能方便地进行。ＩＭＡＳＥ是在ＭＰＳＥ或ＹＦＳＩＭ／ＹＦＲＩＯ的基础上，扩建仿真数据库和模型单元库，并且通过它们的生成源程序和增强结果分析功能，从而使运载火箭控制系统仿真无论在ＹＨ—Ｆ１上，还是在ＹＦ—Ｆ２上，都有一个一体化建模／仿真环境。     

基于Vega的实时场景渲染技术研究

视景仿真因其能够帮助人们建立一个具有身临其境的沉浸感 ,能与复杂系统进行交互 ,并能促进构想与创造环境 ,而成为社会各个应用领域发展中不可或缺的高科技手段。Vega用于虚拟现实和实时视景仿真软件的开发 ,提供了快捷和很容易就可构造出复杂应用程序的手段 ,可以很好地实现视景仿真的实时场景渲染 ,使虚拟场景更加逼真 ,在视景仿真领域有着广泛的应用