Qualification and in-flight demonstration of a European tether deployment system on YES2

This paper highlights the design, qualification and mission performance of the tether deployer system on the second Young Engineers’ Satellite (YES2), that featured a tethered momentum transfer. The deployer is designed with a broad range of near-term tether applications in mind. The system contains the tether, including features to enhance safety and wound up in controlled manner onto a spool core, optical deployment sensors, a “barberpole” friction brake controlled by a stepper motor and a triple tether cutter system. To initiate the deployment a spring-based ejection system was developed, and to apply accurate momentum transfer a timer and release system is present on the subsatellite side. A small, 6 kg re-entry capsule was developed as subsatellite. On September 25th, 2007, YES2 deployed a 32 km tether in orbit and gathered a wealth of data. Confidence is gained from the mission results for use of the deployer in future missions.     

YES2 optimal trajectories in presence of eccentricity and aerodynamic drag

YES2 (launching 2007) aims to demonstrate a tether-assisted re-entry concept, whereby payload will be returned to Earth using momentum provided from a swinging tether. Deployment takes place in two phases: (1) deployment of 3.5 km of tether to the local vertical and hold, and (2) deployment to 30 km for a swinging cut. Optimal trajectories are determined for both phases after comparing the effect of different cost functions on the deployment dynamics. Closed-loop control is provided by linearizing the dynamics around the optimal trajectories and solving a receding horizon control problem for a set of linear feedback gains. The controllers are tested in a flexible tether model with large disturbances to the hardware model and environmental variables. Closed-loop simulations show that the system can be controlled quite well using only feedback of length and length rate.     

Low cost planetary exploration: surrey lunar minisatellite and interplanetary platform missions

In order to meet the growing global requirement for affordable missions beyond Low Earth Orbit, two types of platform are under design at the Surrey Space Centre. The first platform is a derivative of Surrey's UoSAT-12 minisatellite, launched in April 1999 and operating successfully in-orbit. The minisatellite has been modified to accommodate a propulsion system capable of delivering up to 1700 m/s delta-V, enabling it to support a wide range of very low cost missions to LaGrange points, Near-Earth Objects, and the Moon. A mission to the Moon - dubbed “MoonShine” - is proposed as the first demonstration of the modified minisatellite beyond LEO. The second platform - Surrey's Interplanetary Platform - has been designed to support missions with delta-V requirements up to 3200 m/s, making it ideal for low cost missions to Mars and Venus, as well as Near Earth Objects (NEOs) and other interplanetary trajectories. Analysis has proved mission feasibility, identifying key challenges in both missions for developing cost-effective techniques for: spacecraft propulsion; navigation; autonomous operations; and a reliable safe mode strategy. To reduce mission risk, inherently failure resistant lunar and interplanetary trajectories are under study. In order to significantly reduce cost and increase reliability, both platforms can communicate with low-cost ground stations and exploit Surrey's experience in autonomous operations. The lunar minisatellite can provide up to 70 kg payload margin in lunar orbit for a total mission cost US$16–25 M. The interplanetary platform can deliver 20 kg of scientific payload to Mars or Venus orbit for a mission cost US$25–50 M. Together, the platforms will enable regular flight of payloads to the Moon and interplanetary space at unprecedented low cost. This paper outlines key systems engineering issues for the proposed Lunar Minisatellite and interplanetary Platform Missions, and describes the accommodation and performance offered to planetary payloads.     

COLUMBUS Orbital Facility and Automated Transfer Vehicle: a challenge for agency & industry

Long term continuous operation of the COLUMBUS Orbital Facility (COF) flight- and ground segment requires continuous mission control and operations support capability to ensure proper operation and configuration of the COF systems in support of ongoing science and technology payloads. The ISS logistics scenario will be supported by the Automated Transfer Vehicle (ATV). These operational needs require the built-up of a new ground infrastructure in Europe and USA, enabling an efficient operations for preparation, planning and mission execution. The challenge for the European space community consists in the development and operation of a user friendly operational environment but keeping costs within budgetary constraints. Results of detailed definition studies performed by both agency and industry for the ground infrastructure indicate solutions to those technical and programmatic requirements by using of existing centers and facilities, re-use of C/D phase products (Hardware, Software) and COTS equipment to avoid costly new developments, using engineering expertise of the industrial personnel from flight element phase C/D. The concept for operations execution defines the task sharing between Operations Control Facilities (OCF), Operations Support Facilities and User Operations Sites. Operations support consists of on-line engineering support, off-line engineering support, payload integration, logistics support and crew training support performed by industry. DASA RI has made internal investments in organizational concepts for mission operations as well as in mission technologies and tools based on the standard COLUMBUS Ground Software (CGS) toolset and on knowledge based systems to enable an efficient industrial operations support. These tools are available as prototypes being evaluated in a simulated operational environment.     

载人飞船有效载荷搭载技术及试验验证

载人飞船飞行任务中搭载的有效载荷设备质量、体积大并且安装要求苛刻,而返回舱内空间狭小且舱内环境复杂。针对这个问题,文章提出了通过改装返回舱座椅作为有效载荷搭载安放的方案。经过试验验证表明,该方案对飞船的影响很小,并且满足有效载荷设备搭载要求。     

System design of the Hayabusa 2—Asteroid sample return mission to 1999 JU3

The Japan Aerospace Exploration Agency is currently developing the second asteroid sample return mission, designated as Hayabusa 2. Following the successful return of Hayabusa from the asteroid “Itokawa”, Hayabusa 2 is designed as a round-trip mission to the asteroid “1999 JU3”. The 1999 JU3 is a C-type asteroid, which is believed to contain organic matter and hydrated minerals. Thus, it is expected that successful sample collection will provide additional knowledge on the origin and evolution of the planets and, in particular, the origin of water and organic matter. The current mission scenario will enable the spacecraft to reach 1999 JU3 in the middle of 2018 and perform an asteroid proximity operation for 1.5 years. Three touch downs for sampling and one 2-m-class crater generation by means of a high-speed impact operation are planned during the asteroid proximity operation. The samples are to be brought back to the Earth by a re-entry capsule. The present paper describes the system design of Hayabusa 2, some key technical challenges of the mission, and the development status.     

再论空间系绳在回收容器返回中的应用

在已有的研究结果的基础上，对空间系绳在回收容器回中的应用，作了进一步分析，包括从回收容器离开空间站至回收容器从系绳上脱落的系绳伸展运动，（２）从系绳脱落的回收容器的返回运动，（３）释放容器后的系绳的收回运动。对回收容器的动态释放，除了两阶段指数型伸展程序外，还提出了另外两种导致系绳后摆的方式。对静态释放与动态释放，不仅对系绳长度，而且对系绳张力作了计算与分析。为了系绳收回过程的稳定，不仅要施加张力     

Ground assisted rendezvous with geosynchronous satellites for the disposal of space debris by means of Earth-oriented tethers

Previous studies have shown that extended length Earth-oriented tethers in the geosynchronous (GEO) region can be used to re-orbit satellites to disposal orbits. One such approach involves the extension of a GEO based tether, collection of a debris object, and retraction of the tether, which transfers the retracted configuration to a higher energy orbit for debris disposal. The re-extension of the tether after debris disposal returns the configuration to the near-GEO altitude. The practical feasibility of such a system depends on the ability to collect GEO debris objects, attach them to a deployed tether system, and retract the tethers for transfer to the disposal orbits.This study addresses the collection and delivery of debris objects to the deployed tether system in GEO. The investigation considers the number, type and the characteristics of the debris objects as well as the collection tug that can be ground controlled to detect, rendezvous and dock with the debris objects for their delivery to the tethers system.A total of more than 400 objects are in drift orbits crossing all longitudes either below or above the geostationary radius. More than 130 objects are also known to librate around the stable points in GEO with periods of libration up to five or more years. A characterization of the position and velocity of the debris objects relative to the collection tug is investigated. Typical rendezvous performance requirements for uncooperative GEO satellites are examined, and the similarities with other approaches such as the ESA's CX-OLEV commercial mission proposal to extend the life of geostationary telecommunication satellites are noted.     

Mars Express spacecraft: design and development solutions for affordable planetary missions

The spacecraft designed to support the ESA Mars Express mission and its science payloads is customized around an existing avionics well suited to environmental and operational constraints of deep-space interplanetary missions. The reuse of the avionics initially developed for the Rosetta cometary program thanks to an adequate ESA cornerstone program budget paves the way for affordable planetary missions.The costs and schedule benefits inherited from reuse of up-to-date avionics solutions validated in the frame of other programs allows to focus design and development efforts of a new mission over the specific areas which requires customization, such as spacecraft configuration and payload resources. This design approach, combined with the implementation of innovative development and management solutions have enabled to provide the Mars Express mission with an highly capable spacecraft for a remarkably low cost. The different spacecraft subsystems are all based on adequate design solutions. The development plan ensures an exhaustive spacecraft verification in order to perform the mission at minimum risk. New management schemes contribute to maintain the mission within its limited funding.Experience and heritage gained on this program will allow industry to propose to Scientists and Agencies high performance, low-cost solutions for the ambitious Mars Exploration Program of the forthcoming decade.     

80 mN霍尔推力器空心阴极寿命试验

我国的多个GEO卫星平台即将采用电推进系统完成轨道保持任务,其中比冲为1 600s的80 mN霍尔推力器是国际公认的最适合完成该项任务的推力器,也是目前国外卫星和深空探测器应用最广的电推力器.为满足15年GEO卫星寿命要求,80 mN霍尔推力器必须达到7500h和8 000次点火的寿命指标.空心阴极作为霍尔推力器的重要组件,其寿命和点火次数必须达到相应的指标.为此,上海空间推进研究所开展了80 mN霍尔推力器空心阴极的寿命试验,试验采用模拟推力器阳极的三极管工作方式进行.截止2013年8月上旬,试验件1完成10 322 h寿命试验（含4 549次点火）,试验件2完成24 248次加热器热循环试验.空心阴极的寿命已经达到任务要求,两个试验件的放电电压、触持极电压和点火时间等性能指标变化很小,目前试验还在持续进行中.     

多体回收小型返回舱

把有些航天器(如空间站和侦察卫星等)上的物品返回地面,有两种方法可选用:一种是搭载天地往返运输系统返回,另一种是利用航天器上设置的专用返回舱返回。TAURUS和FAST是德国在80年代末设想的两种多体回收小型返回舱。它们预先装在空间站的贮存库中,需要时即可携带待返物品离开空间站、再入大气层并返回地面。文中主要介绍TAURUS返回舱的运行程序及主要构件(有效载荷舱、弹射装置及辅助设施等)。     

Sounding rocket experiment of bare electrodynamic tether system

An overview of a sounding rocket, S-520-25th, project on space tether technology experiment is presented. The project is prepared by an international research group consisting of Japanese, European, American, and Australian researchers. The sounding rocket will be assembled by the ISAS/JAXA and will be launched in the summer of 2009. The sounding rocket mission includes two engineering experiments and two scientific experiments. These experiments consist of the deployment of bare electrodynamic tape tether in space, a quick ignition test of hollow cathode system in space, the demonstration of bare electrodynamic tether system in space, and the test of the OML (orbital-motion-limit) current collection theory.     

多用途飞船返回舱虚拟振动试验研究

利用虚拟振动试验软件系统、多用途飞船返回舱的有限元模型和"神舟一号"返回舱的历史试验数据,对多用途飞船返回舱进行虚拟正弦和随机振动试验,考核了返回舱经受动力学环境的能力,研究发现设计方案在一阶共振区响应过大,存在较大的设计缺陷,经过多次迭代优化设计后的模型满足设计要求。研究结果为多用途飞船返回舱的设计和优化提供了依据。     

Ground and on-orbit command and data handling architectures for the Active Rack Isolation System microgravity flight experiment

The Active Rack Isolation System [ARIS] International Space Station [ISS] Characterization Experiment, or ARIS-ICE for short, is a long duration microgravity characterization experiment aboard the ISS. The objective of the experiment is to fully characterize active microgravity performance of the first ARIS rack deployed on the ISS. Efficient ground and on-orbit command and data handling [C&DH] segments are the crux in achieving the challenging objectives of the mission. The objective of the paper is to provide an overview of the C&DH architectures developed for ARIS-ICE, with the view that these architectures may serve as a model for future ISS microgravity payloads. Both ground and on-orbit segments, and their interaction with corresponding ISS C&DH systems are presented. The heart of the on-orbit segment is the ARIS-ICE Payload On-orbit Processor, ARIS-ICE POP for short. The POP manages communication with the ISS C&DH system and other ISS subsystems and payloads, enables automation of test/data collection sequences, and provides a wide range of utilities such as efficient file downlinks/uplinks, data post-processing, data compression and data storage. The hardware and software architecture of the POP is presented and it is shown that the built-in functionality helps to dramatically streamline the efficiency of on-orbit operations. The ground segment has at its heart special ARIS-ICE Ground Support Equipment [GSE] software developed for the experiment. The software enables efficient command and file uplinks, and reconstruction and display of science telemetry packets. The GSE software architecture is discussed along with its interactions with ISS ground C&DH elements. A test sequence example is used to demonstrate the interplay between the ground and on-orbit segments.     

NASA'S Robotic Lunar Lander Development Project

Over the last 5 years, NASA has invested in development and risk-reduction activities for a new generation of planetary landers capable of carrying instruments and technology demonstrations to the lunar surface and other airless bodies. The Robotic Lunar Lander Development Project (RLLDP) is jointly implemented by NASA Marshall Space Flight Center (MSFC) and the Johns Hopkins University Applied Physics Laboratory (APL). The RLLDP team has produced mission architecture designs for multiple airless body missions to meet both science and human precursor mission needs. The mission architecture concept studies encompass small, medium, and large landers, with payloads from a few tens of kilograms to over 1000 kg, to the Moon and other airless bodies. To mature these concepts, the project has made significant investments in technology risk reduction in focused subsystems. In addition, many lander technologies and algorithms have been tested and demonstrated in an integrated systems environment using free-flying test articles. These design and testing investments have significantly reduced development risk for airless body landers, thereby reducing overall risk and associated costs for future missions.     

Launch cost reduction by the reuse of the core stage of Ariane 5

The high cost of launching payloads into Earth orbit is a main limiting factor on the development of space. In order to reduce the high cost of launch, reuse of (parts of) the launch vehicle is needed. This study analyses the possibilities of recovering and reusing the core stage of Ariane 5. Recovery of the core stage sets demands on re-entry trajectory, attitude, stability, thermal protection, structural strength, terminal deceleration, salt water protection, recovery and refurbishment. All these subject areas require solutions to their individual problems. Added subsystems to the stage are defined and their mass is determined. These masses are used to determine the financial feasibility of the recovery concept, by weighing the payload demise and operational cost against the gains of reduced production cost. It is concluded that the recovery is technologically feasible, using a detachable ablative heat shield on the nose of the stage and a stabilisation device (an inflatable drag cone), a parachute system and an engine enclosure device. Total mass of these systems is 1320 kg, with financial savings amounting to $8.5 million per flight.     

Tethers and debris mitigation

In recent years, the use of tethers has been proposed for reduction of space debris either through momentum transfer or use of electrodynamic effects. Tethers have been shown to at least theoretically allow for quick, elegant and cost-effective deorbit of defunct satellites or spent stages. On the other hand, the large risk that tethers themselves may pose to other satellites in orbit has been recognized as well. The large collision area of tethers, combined with operational hazards and meteoroid risk may result in a large orbital exposure. For example, in 1997, the ESA/Dutch 35-km tether deployment of YES from TEAMSAT was inhibited after an analysis of the collision risk for the case the tether operation would fail. The question rises how these two points of view compare to eachother. This paper intends to highlight a representative selection of the proposed tether applications while taking into account the added risks caused by the tethers themselves.Typical applications from recent literature will be briefly described, such as an Ariane 502 spent stage re-entry from GTO and the concept of deboost of defunct satellites by interaction of a conductive tether with the Earth magnetic field.Mass savings of the tethered sytems versus conventional equivalents will be evaluated.Based on a crude risk analysis, involving elements such as mission complexity, dynamic stability, meteoroid risk and orbital life time, a general outline of limiting factors can be given for the various applications. Special attention is reserved for implementation of mechanisms that help reduce this tether risk, such as the DUtether (Tether Degradable by Ultraviolet), utilization of airdrag and solar pressure, the effect of residual current in bare tethers, tether retrieval etc.It is proposed how a net tether-induced mitigation can be compared to that of conventional alternatives, i.e. deboost by rocket engine or a completely passive approach.This comparison is put in the perspective of an ever-increasing occupation of the space environment.It is concluded that tethers can in fact help mitigate the debris risk and that for each application a useful niche can be defined. It is argued that eliminating pollution directly after use of the precious resource of space is not only good custom, but also an important way to make the risk of debris controllable and independent of future trends. Although tethers may have large exposure in terms of area-time product, they deliver a quick cleaning service that may be appreciated by the future users of space.     

A cost-engineered launch vehicle for space tourism

The paper starts with a set of major requirements for a space tourism vehicle and discusses major vehicle options proposed for this purpose. It seems that the requirements can be met best with a Ballistic SSTO Vehicle which has the additional advantage of lowest development cost compared to other launch vehicle options — important for a commercial development venture.

The BETA Ballistic Reusable Vehicle Concept is characterized by the plug nozzle cluster engine configuration where the plug nozzle serves also as base plate and re-entry heat shield. In this case no athmospheric turn maneuver is required (as in case-of the front-entry Delta-Clipper DC-Y concept). In our specific case for space tourism this mode has the avantage that the forces at launch and reentry are in exactly the same direction, easing passenger seating arrangements. The second basic advantage is the large available volume on top of the vehicle providing ample space for passenger accomodation, visibility and volume for zero-g experience (free floating), one of the major passenger mission requirements. An adequate passenger cabin design for 100 passengers is presented, as well as the modern BETA-STV Concept with its mass allocations.