全电推进GEO卫星的变轨策略研究

针对采用全电推进平台的GEO卫星的轨道转移问题,在给定电推进轨道控制模型的基础上,开展卫星由GTO向GEO的变轨策略研究。首先,在轨道远地点高度不变的基础上,给出了给定控制解析解的变轨策略,并给出了该种策略轨道控制律的解析解。其次,文章基于李雅普诺夫优化原理给出了优化的变轨策略,并结合仿真算例计算了该种策略控制律的数值解。在电推进优化转移策略研究的基础上,还分析了全电推进平台卫星在不同推力下的变轨时间需求。文章的变轨策略可为电推进系统在航天领域的应用提供参考。     

GEO卫星电推进与化学推进组合变轨方案研究

针对GEO卫星采用电推进和化学推进系统进行变轨的问题,在给定轨道控制模型的基础上,给出实现轨道半长轴、偏心率和倾角3个参数单独调整和联合控制的控制律。结合假设的运载火箭发射能力约束,给出航天器不同入轨高度对应的初始质量。以此为约束,给出化学推进和电推进不同组合的6种GTO向GEO转移方案,并对比分析各方案完成变轨所需的时间、推进剂、速度增量以及完成变轨后的质量剩余情况,研究方法可为电推进系统在航天领域的应用提供参考。     

一种GEO卫星平台电推进轨道转移技术

为解决传统化学推进方式比冲小、燃料携带量大等问题,将化学推进平台改为全电推进平台。全电推进可节省卫星燃料,增加卫星载重比,延长卫星使用寿命,并且支持"一箭双星"发射。通过调研国内外全电推进卫星平台的进展情况,设计出使用电推进的轨道转移方案,并与传统推进方式进行对比。分析结果表明:卫星发射质量为2 700 kg,使用2台300 mN的推力器并联工作,从地球同步转移轨道(GTO)至地球同步静止轨道(GEO)的转移时间约为6个月,消耗燃料约650 kg,可满足任务需求。     

中国数据中继卫星系统及其应用拓展

论述了中国天链一号数据中继卫星系统的建设历程和特点、应用实践和对中国航天事业发展的贡献。对中国未来数据中继卫星系统在大系统层面(包括优化顶层设计、提高工作效率和扩大覆盖区等)和卫星层面(包括相关的先进有效载荷和平台技术)的技术发展提出了建议,对今后应用领域的拓展(如地球同步转移轨道(GTO)/地球静止轨道(GEO)用户航天器、再入通信、深空通信和特殊卫星星座管理等)进行了讨论。     

卫星液体推进剂剩余量地面模拟试验装置

介绍了根据气体注入压力激励方法研制的卫星液体推进剂剩余量测量地面模拟装置。装置的主要组成包括卫星推进剂贮箱模拟系统,温度、压力测量系统,气体注入压力激励过程控制系统等。     

基于遥测数据的卫星在轨飞行温度仿真算法研究

研究一种基于遥测数据的卫星在轨飞行温度仿真计算方法,以卫星热控边界温度遥测参数作为仿真计算模型基准温度参数,挖掘星上设备温度与安装边界（热控边界）温度之间的数值定量关系,形成卫星温度关系数值矩阵。通过基准温度遥测数据与卫星温度关系数值矩阵之间数值运算,实现卫星在轨飞行温度仿真预计,计算误差小于2.5℃。     

实践四号卫星总体设计和运行结果初步分析

概述实践四号卫星是用于地球同步转移轨道(GTO)上的空间辐射环境探测与辐射效应试验的卫星总体方案,并对各项任务和飞行的初步结果进行初步分析。     

海洋卫星水色仪羽流污染DSMC模拟研究

在当前和今后很长一段时间内,我国卫星姿轨控系统都将液体小发动机作为主要推进系统。对于长寿命、高可靠卫星,羽流污染评估具有重要意义,而数值模拟则是有效方法。采用DSMC方法模拟了海洋卫星单组元轨控发动机羽流对水色仪表面可能造成的污染,通过这个算例验证了DSMC方法在羽流污染模拟中的重要作用。假设羽流组分为NH3、N2、H2,水色仪表面温度为100K,计算结果显示在水色仪所在位置发动机羽流造成污染影响最小,可以忽略不计。这表明推进系统选择合理,卫星结构设计合理。     

小卫星空间模拟器KM3B的研制

空间模拟器是用来模拟空间的真空、冷黑及太阳辐照环境,完成卫星的热平衡、热真空试验,满足卫星研制需要的关键地面大型试验设备。文章概要介绍了研制的一台有效试验空间为φ3 800 mm×4 800 mm的空间模拟器的组成、采用的模拟技术、调试及使用情况。使用结果表明:该设备的各系统运行稳定、可靠,模拟室的真空度、污染量、热沉温度、红外热流模拟系统、温度测量系统等均能满足卫星整星热平衡、热真空试验的要求。     

OTS卫星在轨道上的温控试验

本文主要介绍欧洲实验通信卫星(OTS)在轨道上运行近四年期间,在所有的分日和至日进行的似稳状态热试验情况。把观察到的卫星温度与利用“寿命初期”热涂层值预计的温度进行了比较。由于卫星的表面质量渐渐下降,导致卫星的温度逐步升高,而卫星温度与预计温度之间的差值也就不断扩大。目前可以利用最小二乘方法的指数曲线对取得的数据进行分析,从而能用外推法预测长时间内的温度;在比较各分日数据和至日数据之后,便可估算出两个散热器上的太阳光反射器的表面质量下降程度。     

小倾角GSO卫星的转移轨道优化设计

分析了采用小倾角地球同步轨道(GSO)策略后轨道转移变轨推进剂消耗增加的原因,对传统的地球同步转移轨道(GTO)提出了优化措施,根据月份发射的不同,选取不同的转移轨道近地点幅角ω,降低了转移轨道变轨所需的速度增量。速度增量平均降幅可达到55m/s,相当于延长了该卫星1年的推进剂消耗寿命。文章的研究结果可为优化卫星总体设计提供参考。     

Ariane 5 ECA's upper stage perigee reduction via passivation manoeuvre

When Ariane 5 ECA development has been decided by Europe to increase Ariane 5 performance, the rule of 25 years in GTO orbit for the upper stage has been anticipated. This was 14 years ago and this rule was known to be satisfied with a perigee lower than 250 km. Even when lowering slightly Ariane 5 ECA performance, this maximum perigee altitude has been held and the whole Launch System has been developed under CNES responsibility with this GTO perigee. In the meantime, more precise calculations demonstrated that such a GTO perigee was giving for the ESCA a mean lifetime higher than 25 years. So studies are in progress inside CNES to decrease the perigee and re-enter inside the 25 years lifetime domain. This paper presents a CNES study to reduce the orbital lifetime of Ariane 5's upper stage that last in GTO after each commercial mission. Usually the aimed orbit has a perigee altitude of 250 km, an apogee altitude near to the geostationary position and an inclination between 2° and 7°. These conditions make stage's mean lifetime superior to 90 years. The CNES study is to expose the possibility to decrease this lifetime by reducing the perigee altitude of the final upper stage orbit through a passivation process optimised to produce orbit modification. It is shown that taking into account material and functional stage constraints the optimised passivation process is able to decrease the perigee by a few tenths of kilometres.     

我国新一代中型高轨运载火箭发展研究

新一代火箭CZ-5、CZ-6和CZ-7陆续首飞成功,拉开了我国运载火箭更新换代的序幕。新一代中型运载火箭CZ-7于2016年6月和2017年4月圆满完成了两次飞行任务,为中型运载火箭的研制奠定了坚实的基础。在CZ-7火箭基础上,增加CZ-3A氢氧三子级,在海南文昌发射GTO轨道卫星,运载能力不低于7.0t,可快速形成更新换代能力,填补我国GTO轨道该吨位的运载能力的空白。为了进一步提升我国运载火箭的竞争力,对标国际先进水平,针对新一代中型高轨运载火箭开展构型优化研究,以提高火箭性能,降低火箭成本,提升火箭的使用维护性能,满足后续GTO发射任务需求。     

Atlas II and IIA analyses and environments validation

General Dynamics has now flown all four versions of the Atlas commercial launch vehicle, which cover a payload weight capability to geosynchronous transfer orbit (GTO) in the range of 5000–8000 lb. The key analyses to set design and environmental test parameters for the vehicle modifications and the ground and flight test data that validated them were prepared in paper IAF-91-170 for the first version, Atlas I.

This paper presents similar data for the next two versions, Atlas II and IIA. The Atlas II has propellant tanks lengthened by 12 ft and is boosted by MA-5A rocket engines uprated to 474,000 lb liftoff thrust. GTO payload capability is 6225 lb with the 11-ft fairing. The Atlas IIA is an Atlas II with uprated RL10A-4 engines on the lengthened Centaur II upper stage. The two 20,800 lb thrust, 449 s specific impulse engines with an optional extendible nozzle increase payload capability to GTO to 6635 lb. The paper describes design parameters and validated test results for many other improvements that have generally provided greater capability at less cost, weight and complexity and better reliability. Those described include: moving the MA-5A start system to the ground, replacing the vernier engines with a simple 50 lb thrust on-off hydrazine roll control system, addition of a POGO suppressor, replacement of Centaur jettisonable insulation panels with fixed foam, a new inertial navigation unit (INU) that combines in one package a ring-laser gyro based strapdown guidance system with two MIL-STD-1750A processors, redundant MIL-STD-1553 data bus interfaces, robust Ada-based software and a new Al-Li payload adapter. Payload environment is shown to be essentially unchanged from previous Atlas vehicles. Validation of load, stability, control and pressurization requirements for the larger vehicle is discussed.

All flights to date (five Atlas II, one Atlas IIA) have been successful in launching satellites for EUTELSAT, the U.S. Air Force and INTELSAT. Significant design parameters validated by these flights are presented. Particularly noteworthy has been the performance of the INU, which has provided average GTO insertion errors of only 10 miles apogee, 0.2 miles perigee and 0.004 degrees inclination. It is concluded that Atlas II/IIA have successfully demonstrated probably the largest number of current state-of-the-art components of any expendable launch vehicle flying today.     

北斗工程IGSO/MEO/GEO发射轨道设计

随着我国航天事业的蓬勃发展，运载火箭发射要求也呈现多样化。北斗卫星导航系统是我国自行研制的全球卫星导航系统，经历三步跨越式发展，目前已经全面建成。CZ-3A系列火箭承担了北斗工程全部发射任务，该工程对火箭倾斜同步转移轨道（IGTO）、中圆转移轨道（MTO）、地球同步转移轨道（GTO）新类型轨道要求。介绍了该类轨道特点，讨论了火箭发射方案、发射轨道设计及高空风双向补偿方法。实际飞行考核充分证明了发射轨道设计的正确性，设计方法确保了北斗工程全部发射任务取得圆满成功，为北斗工程顺利实施奠定了基础。     

Development status of the high-speed reentry system-DASH

Despite huge amount of data collected by the previous interplanetary spacecraft and probes, the origin and evolution of the solar system still remains unveiled due to limited information they brought back. Thus, the Institute of Space and Astronautical Science (ISAS) of Japan has been given a commitment to pave the way to an asteroid sample return mission: the MUSES-C project. A key to success is considered the reentry with hyperbolic velocity, which has not ever been demonstrated as yet. With this as background, a demonstrator of atmospheric reentry system, DASH, has been designed to demonstrate the high-speed reentry technology as a GTO piggyback mission. The capsule, identical to that of the sample return mission, can experience the targeted level of thermal environment even from the GTO by tracing a specially designed reentry trajectory. After the purpose of the mission was outlined at the last IAF symposium, the final fitting tests have been conducted in the ISAS Sagamihara Campus involving the flight model hardware. Furthermore, a series of rehearsals for recovery have been already executed. The paper describes the current mission status of the project.     

Single crystal mo solar thermal thruster for microsatellites

One potentially attractive propulsion concept offering significant payload gains for orbit transfer from LEO to higher orbits, station keeping and attitude control of spacecraft is thermal propulsion using light gas (typically hydrogen) as propellant and various kinds of heat energy. Solar Thermal Propulsion (STP) is a typical thermal propulsion with high Isp (500 – 1,000 s) in an appropriate thrust magnitude range and provides possibly much less space pollution than conventional chemical propulsion.

This paper presents the test results of a 30 mm dia. (medium-sized) windowless type of single crystal Mo thruster for orbit transfer of 50 kg class microsatellites. The cavity dia. is 20 mm, double the size of the previous model, and can apply to a primary solar reflector of up to 3.5 m dia., which is the maximum size containable in the H-II rocket fairing without segmentation. The performed mission analyses indicate that this size of STP is suitable to orbit transfer of 50 kg class microsatellites, such as LEO to GEO, or only multiple apogee kicks from GTO to GEO or deep space missions.     

北斗专列—CZ 3A系列火箭发展回顾和未来展望

CZ-3A系列运载火箭承担了北斗工程，包括北斗一号、北斗二号和北斗三号的全部发射任务，被称为“北斗专列”。在北斗工程历时26年的研制过程中，CZ-3A系列火箭突破了一系列关键技术，使火箭具备了从地球同步转移轨道(GTO)到倾斜同步转移轨道(IGTO)、中圆转移轨道(MTO)，从一箭一星发射至转移轨道，到一箭双星发射至转移轨道，再到一箭双星直接发射入轨的发射能力，实现了跨越发展。满足了北斗工程的发射任务需求，北斗工程共计44箭、59星，CZ-3A系列火箭发射均获得了圆满成功，成功率达到了100%。     

A preliminary study of pulse-laser powered orbital launcher

An air-breathing pulse-laser powered orbital launcher has been proposed as an alternative to conventional chemical launch systems. The aim of the present study is to assess its feasibility through the estimation of its achievable payload mass per unit beam power and launch cost. A transfer trajectory from the ground to a geosynchronous Earth orbit (GEO) is proposed, and the launch trajectory to its geosynchronous transfer orbit (GTO) is computed using the realistic performance modeled in the pulsejet, ramjet, and rocket flight modes of the launcher. Results show that the launcher can transfer 0.084 kg of payload per 1 MW beam power to a geosynchronous earth orbit. The cost becomes a quarter of existing systems if one can divide a single launch into 24,000 multiple launches.     

大型GEO通信卫星平台转移轨道段热分析

为了增强大型高价值卫星的抗风险能力,针对极端温度环境以及相较一般发射工作程序有所偏离的情况,进行转移轨道卫星平台的热性能分析。对抛整流罩、星箭分离、出地影、南板受照、北板受照5个典型状态下的卫星平台进行热分析,结果显示:卫星平台本体的温度包络为51.68～-16.35 ℃,未超出一般要求,表明卫星平台的热性能有一定保持能力,但平台本体温度与热控要求间的裕度仅3.3～3.6 ℃,须予以关注;同时,采用3D舱板的有限元分析方法揭示出服务舱板厚度方向热梯度在出地影时刻最大值达5 ℃,对结构精度和机构运动性能或带来潜在风险。