Protecting spacecraft against meteoroid/orbital debris impact damage: an overview

Meteoroids and orbital debris pose a serious damage threat to all spacecraft. The effects of a meteoroid/orbital debris (M/OD) impact depend on a variety of factors, including where the M/OD impact occurs, the size, composition, and speed of the impacting object, and the function of the impacted spacecraft system. These effects can be minimal, can degrade a functional spacecraft component, or can compromise spacecraft functionality, even to the point of mission loss or loss of life. To minimize the damage threat from the meteoroid/orbital debris environment, it is often necessary to install protective shielding around critical spacecraft systems. If a system cannot be shielded, operational constraints may need to be imposed to reduce the damage threat. This paper presents an overview of the research and development activities performed since the late 1950s with an aim of increasing the level of protection afforded satellites and spacecraft operating in the M/OD environment and ultimately mitigating the mechanical and structural effects of an M/OD impact.     

Vibration control of a flexible space manipulator during on orbit operations

Space manipulators are complex systems, composed by robotic arms accommodated on an orbiting platform. They can be used to perform a variety of tasks: launch of satellites, retrieval of spacecraft for inspection, maintenance and repair, movement of cargo and so on. All these missions require extreme precision. However, in order to respect the mass at launch requirements, manipulators arms are usually very light and flexible, and their motion involves significant structural vibrations, especially after a grasping maneuver. In order to fulfill the maneuvers of space robotic systems it is hence necessary to properly model the forces acting on the space robot, from the main terms, such as the orbital motion, to the second order perturbations, like the gravity gradient and the orbital perturbations; also flexible excitation of the links and of the joints can be of great importance in the manipulators dynamics. The case is furthermore complicated by the fact that the manipulator, together with its supporting spacecraft, is an unconstrained body. Therefore the motion of any of its parts affects the entire system configuration. The governing equations of the dynamics of such robotic systems are highly nonlinear and fully coupled. The present paper aims at designing and studying active damping strategies and relevant devices that could be used to reduce the structural vibrations of a space manipulator with flexible links during its on orbit operations. In particular an optimized adaptive vibration control via piezoelectric devices is proposed. The number of piezoelectric devices, their placement and operational mode should be correctly chosen in order to obtain maximum performance in terms of elastic oscillations reduction and power consumption. Even though an optimal placement cannot have a universal validity, since it depends on the type of maneuver and on the overall inertial and geometrical characteristics, an approach to solve the problem is proposed.     

单航天器无需机动对星座多星交会轨道全解

文献[1,2]已研究了单航天器无需变轨对Walker星座多星交会的轨道设计,在此基础上,依据交会点必是轨道交点的轨道特性,提出了轨道全解的解析法,可给出该轨道所能够交会卫星的最大数目,研究结果可为单航天器无需大机动变轨对星座多星接近的轨道应用提供理论参考。     

A space based radar on a micro-satellite for in-situ detection of small orbital debris

The increase in the number of satellites in the Near Earth Orbit is exponential. The consequent increase in pollution of the orbital environment is of growing concern to the international community. There are currently only two observation systems available for measurement of orbital debris. Ground based radar and telescopes can detect objects larger than about 7 cm. Passive space based systems provide an accurate statistical estimation of flux for debris smaller than about 0.1 mm in size. Consequently, there is no way of obtaining information about debris in the millimeter-size range. Considering that the relative speed between objects in space is commonly in the km/s range, millimeter sized debris carry enough energy to be deadly to astronauts or to totally destroy the functioning of any satellite. Then National space agencies have recommended launching orbital spacecraft carrying debris detection experiments for gaining a better understanding of small debris.CNES (the French Space Agency) is developing a new family of micro-satellites, that will make possible to put into orbit a totally new system of radar that could measure in-situ flux of debris. We present results of this system analysis, which would cumulate the advantages of both ground-based radar and in orbit passive experiments.The proposed method for detection is quite original and allows the radar to act like a band-pass filter with respect to the debris diameter. The optimum frequency is shown to be in the Ka-band. Two points are critical in the definition of the radar: the average power available and the false alarm probability in the detection criterion. Therefore, we present a special receiver chain in order to optimize the signal-to-noise ratio. The estimate of the radial velocity through Doppler frequency measurement may be used to discriminate orbital debris from meteoroids. This system could be built today using an existing Continuous Wave amplifier. Several hundreds of objects per year could be detected yielding an accurate statistical estimation.The orbital debris radar would be a major contribution to our knowledge of millimeter sized debris. This experiment would contribute to making the current models more accurate at all inclinations. The micro-satellite concept would make the orbital debris radar mission cheap enough for considering a constellation of such satellites.     

SAR卫星用能量功率兼顾型锂离子电池研究

未来大功率合成孔径雷达(SAR)卫星的发展对储能电池组提出了轻量化、倍率高及在轨寿命长等要求,而现有的SAR卫星用功率型锂离子电池产品已无法满足未来需求。通过优化正极材料和电池设计,加入功能电解液,在保证电池功率特性的同时,显著提升了电池的能量密度和循环稳定性。本文研制出的新一代能量功率兼顾型锂离子蓄电池,额定容量为25.0 Ah,初期放电比能量达到180.0 Wh·kg~(-1),2 C放电容量为0.2 C容量的91.0%,1 C-100%放电深度(DOD)和2 C-30%DOD循环性能优异,可以满足下一代大功率SAR卫星的供配电需求,且大幅降低了电源系统重量,提高了卫星储能系统的利用率和平台有效载荷能力。     

Super-capacitor energy storage for micro-satellites: Feasibility and potential mission applications

Small satellites, weighting between 100 and 200 kg, have witnessed increasing use for a variety of space applications including remote sensing constellations and technology demonstrations. The energy storage/stored power demands of most spacecraft, including small satellites, are currently accommodated by rechargeable batteries—typically nickel–cadmium cells (specific energy of 50 Wh kg⁻¹), or more recently lithium-ion cells (150 Wh kg⁻¹). High energy density is a primary concern for spacecraft energy storage design, and these batteries have been sufficient for most applications. However, constraints on the allowable on-board battery size have limited peak power performance such that the maximum power supply capability of small satellites currently ranges between only 70 and 200 W. This relatively low maximum power limits the capabilities of small satellites in terms of payload design and selection. In order to enhance these satellites' power performance, the research reported in this paper focused on the implementation of super-capacitors as practical rechargeable energy storage medium, and as an alternative to chemical batteries. Compared to batteries, some super-capacitors are able to supply high power at high energy-efficiency, but unfortunately they still have a very low energy density (5–30 Wh kg⁻¹). However, the provision of this high power capability would considerably widen the range of small satellite applications.     

Cost reduction trends in space communications by larger satellites/platforms

The paper analyses first the satellite growth trend in the past and the specific characteristics of communication satellites, as there are specific mass per channel and payload share vs. spacecraft mass.

With assistance of a cost model (derived from actual spacecraft cost) it is shown that larger satellites are more cost effective. The same applies to the launch cost, also showing a reduction in specific cost (per kg or per channel-year) for larger payloads.

Finally different types of communication satellites/platforms are compared (two smaller satellites, one large satellite, modular docked assemblies) for the same total communication capacity of 72 000 dual telephone channels. It is shown that for each orbital communication capacity a certain optimum spacecraft size exists which leads to minimum space segment cost.     

低功率水电弧加热发动机的初步研究

本文研究了以水为推进剂的电弧加热发动机。其中电弧加热发动机的流量为15mg/s、电流为8A,此时平均电压为80.5V、功率约为640W、平均推力为0.0975N,以此计算得到比冲为650s左右、效率为55%～60%。成功地验证了水做为电弧加热发动机的推进剂、产生推力的能力。     

Design and implementation of the GPS subsystem for the Radio Aurora eXplorer

This paper presents the design and implementation of the Global Positioning System (GPS) subsystem for the Radio Aurora eXplorer (RAX) CubeSat. The GPS subsystem provides accurate temporal and spatial information necessary to satisfy the science objectives of the RAX mission. There are many challenges in the successful design and implementation of a GPS subsystem for a CubeSat-based mission, including power, size, mass, and financial constraints. This paper presents an approach for selecting and testing the individual and integrated GPS subsystem components, including the receiver, antenna, low noise amplifier, and supporting circuitry. The procedures to numerically evaluate the GPS link budget and test the subsystem components at various stages of system integration are described. Performance results for simulated tests in the terrestrial and orbital environments are provided, including start-up times, carrier-to-noise ratios, and orbital position accuracy. Preliminary on-orbit GPS results from the RAX-1 and RAX-2 spacecraft are presented to validate the design process and pre-flight simulations. Overall, this paper provides a systematic approach to aid future satellite designers in implementing and verifying GPS subsystems for resource-constrained small satellites.     

Investigation of national policy shifts to impact orbital debris environments

Low earth orbit has become increasingly congested as the satellite population has grown over the past few decades, making orbital debris a major concern for the operational stability of space assets. This congestion was highlighted by the collision of the Iridium 33 and Cosmos 2251 satellites in 2009. This paper addresses the current state of orbital debris regulation in the United States and asks what might be done through policy change to mitigate risks in the orbital debris environment. A brief discussion of the nature of orbital debris addresses the major contributing factors including size classes, locations of population concentrations, projected satellite populations, and current challenges presented in using post-mission active debris removal to mitigate orbital debris. An overview of the current orbital debris regulatory structure of the United States reveals the fragmented nature of having six regulating bodies providing varying levels of oversight to their markets. A closer look into the regulatory policy of these agencies shows that, while they all take direction from The U.S. Government Orbital Debris Mitigation Standard Practices, this policy is a guideline with no real penalty for non-compliance. Various policy solutions to the orbital debris problem are presented, ranging from a business as usual approach to a consolidated regulation system which would encourage spacecraft operator compliance. The positive aspects of these options are presented as themes that would comprise an effective policy shift towards successful LEO conservation. Potential economic and physical limitations to this policy approach are also addressed.     

大型低轨航天器与星座卫星的碰撞风险研究

针对商业小卫星星座迅猛发展对航天器飞行安全造成潜在威胁的问题,以600 km高度星座和大型低轨航天器为研究对象,通过区域方法(BOX)和碰撞概率风险评估方法,分析了星座与大型低轨航天器的碰撞风险。根据星座轨道演化分析表明,整个星座卫星与大型低轨航天器可能发生碰撞的时间相对集中,持续时间约 1~ 2年。BOX方法计算结果表明,每颗卫星与大型低轨航天器交会,并进入红色预警门限的交会次数约10次左右。碰撞概率计算结果表明,约有5%的卫星进入红色预警门限,星座如果在寿命末期采取无控再入将对大型低轨航天器在1~2年内产生较大的威胁。     

Quality management for space systems in ISRO

In a little over four decades, the Indian Space Program has carved a niche for itself with the unique application driven program oriented towards National development. The end-to-end capability approach of the space projects in the country call for innovative practices and procedures in assuring the quality and reliability of space systems. The System Reliability (SR) efforts initiated at the start of the projects continue during the entire life cycle of the project encompassing design, development, realisation, assembly, testing and integration and during launch. Even after the launch, SR groups participate in the on-orbit evaluation of transponders in communication satellites and camera systems in remote sensing satellites. SR groups play a major role in identification, evaluation and inculcating quality practices in work centres involved in the fabrication of mechanical, electronics and propulsion systems required for Indian Space Research Organization's (ISRO's) launch vehicle and spacecraft projects. Also the reliability analysis activities like prediction, assessment and demonstration as well as de-rating analysis, Failure Mode Effects and Criticality Analysis (FMECA) and worst-case analysis are carried out by SR groups during various stages of project realisation. These activities provide the basis for project management to take appropriate techno-managerial decisions to ensure that the required reliability goals are met. Extensive test facilities catering to the needs of the space program has been set up. A system for consolidating the experience and expertise gained for issue of standards called product assurance specifications to be used in all ISRO centres has also been established.     

采用太阳电池阵和储能飞轮的电源系统设计与分析

飞轮储能装置具有比能量高、寿命长、任务期内无衰减等优点,可替代航天器中传统的化学储能装置。为论证太阳电池阵-储能飞轮电源系统的可行性,本文从航天器总体设计的角度分析了其关键设计要素,论述了其对航天器机、电、热等方面的影响,并给出提高系统可行性的合理化建议,以及针对低轨卫星的太阳电池阵-储能飞轮电源系统的设计举例。通过与传统电源系统的技术指标对比分析,表明太阳电池阵-储能飞轮电源系统具有较高的比功率,并在降低航天器质量、节约发射成本方面具有很大优势,在未来航天器的应用中具有很大的潜力。     

On optimal aerodynamic attitude control of spacecraft

Attitude control of spin-stabilized satellites by means of aerodynamic forces is investigated. A controller employing two rotatable control surfaces is proposed to control the roll-yaw motions of the spacecraft spin-axis. Optimal control theory is applied to synthesize a feedback control law for the control surface rotations which leads to asymptotically stable controller operation. The system response is interpreted in terms of performance criteria such as the maximum control surface excursions, the speed of response and the associated orbital energy loss due to drag. Even with a moderate size, the controller appears quite effective in maintaining the spacecraft attitude against external disturbances.     

System test results from the GNC experiments on the PRISMA in-orbit test bed

The PRISMA in-orbit test bed will demonstrate guidance, navigation, and control strategies for spacecraft formation flying and rendezvous. The project is funded by the Swedish National Space Board and the prime contractor is the Swedish Space Corporation. The project is further supported by the German Aerospace Center, the Technical University of Denmark, and the French Space Agency. PRISMA was launched on June 15, 2010 and after three weeks of operations, all on-board systems and units have passed an initial commissioning phase. Separation of the two PRISMA satellites from each other is expected by mid-August 2010.PRISMA consists of two spacecraft: MAIN and TARGET. The MAIN spacecraft has full orbit control capability while TARGET is attitude controlled only.The Swedish Space Corporation is responsible for three groups of guidance, navigation, and control experiments. These experiments include GPS- and vision-based formation flying during which the spacecraft will fly in passive as well as forced motion. The three experiments are: autonomous formation flying, proximity operations with final approach/recede maneuvers, and autonomous rendezvous. This paper presents system test results from two of these experiments as obtained with the flight-ready system. The system tests consist of a series of simulations performed on the flight model spacecraft with a large amount of hardware in the loop.     

Space light: space industrial enhancement of the solar option

The solar option can be enhanced significantly by space light technology. Reflectors in suitable orbits beam to Earth measured amounts of sunlight, the most versatile and bio-compatible energy source. The multitude of space light functions ranges from night illumination of rural and urban areas (by Lunetta systems) to photosynthetic production enhancement for the growth of food and of biomass for conversion to chemical fuels, local agricultural irradiation for crop drying and weather stabilization and to electric power generation by irradiating suitable photovoltaic or thermal ground receivers at night or by adding to the natural solar energy input during daytime (Soletta systems).

The Lunetta and Soletta concepts, developed by the author during the past ten years, building on the foundations laid by the great space pioneer Prof. H. Oberth (1928), are reviewed, along with their socio-economic merits. An assessment of terrestrial alternatives shows that many useful functions have no practical alternative, the major exception being electric power generation. Three systems are selected, bracketing the broad versatility of space light—Lunetta, Powersoletta and a large Biosoletta for large-scale seafood production in Antarctic and Artic waters. The systems, and several maintenance and supply requirements are described, sized and analyzed, along with suitable orbit selection for different applications. Models are developed for rural and urban area lighting, power generation at selected sites around the globe with photovoltaic and thermal ground stations and for the large-scale production of seafood at high southern and northern latitudes with ample nutrient upwell, but insufficient annual supply of solar energy. The economics of these systems is analyzed.     

Enhancing ground communication of distributed space systems

The functionality of a distributed system can be significantly enhanced by exploring non-traditional approaches that leverage on inherent aspects of distributed systems in space. Till now, the benefit of distributed systems in space has been limited to enhancing coverage, multipoint sensing, creating virtual baselines (e.g. interferometry) or to enhance redundancy. The list of benefits can be further expanded by understanding the nature of distributed systems and by productively incorporating it into mission and spacecraft design. For example, prior knowledge of the spatial evolution of such systems can lead to innovative communication architectures for these distributed systems. In this paper, different communication scenarios are investigated that can enhance the communication link between the distributed system and ground.The increasing trend towards highly miniaturized spacecraft (nano- to femto-satellites) and proposals to launch hundreds or even thousands of them in massively distributed space missions have expanded the interest in distributed systems with miniature spacecraft. It is important to understand how and which, functionalities and systems, scale with size and number. Scalable systems are defined and addressed at a basic level and the utility of scaling rules and trends in identifying optimal configurations of distributed systems is explored.In this paper we focus on the communication capability and identify methods to enhance the communication link between a distributed space segment, consisting of a number of simplistic, resource limited femto-satellites, and earth. As an example, the concept of forming a dynamic phased array in space with the elements of a distributed space system in low-earth orbit is investigated. At the ground receiver, the signals from different satellites forming the array should not differ in phase by more than one-third the transmission wavelength, to ensure constructive superposition. Realizing such a phased array places strict accuracy requirements on time synchronization and knowledge of relative separation between the satellites with respect to the ground receiver. These constraints are derived and discussed.     

Designing nonuniform satellite systems for continuous global coverage using equatorial and polar circular orbits

We present a method for designing nonuniform satellite systems for continuous global coverage using a combination of equatorial and near-polar satellite segments in circular orbits. Equations are derived to determine the basic design parameters of the satellite system itself and the conditions of its closure at the joint of near-polar and equatorial segments. We analyze specific features of near-polar and equatorial satellite systems and their advantages and disadvantages compared with existing classes of near-polar phased and kinematically correct satellite systems. We estimate the minimum required number of spacecrafts in satellite systems for a given fold of coverage and present calculated dependences for classes of near-polar phased and equatorial satellite systems with different types of closure. For the class of kinematically correct satellite systems, we analyze the characteristics of systems with a minimum spacecraft flight height and reveal that the number of satellites in the orbital plane depends on the flight height for different folds of coverage. We bring examples of the best near-polar equatorial satellite systems of global coverage for different folds and a class of satellite systems with a fixed number of spacecrafts and orbital planes in them.     

Dual attitude and parameter estimation of passively magnetically stabilized nano satellites

The attitude determination capability of a nano satellite is limited by a lack of traditional high performance attitude sensors, a result of having small budgets for mass and power. Attitude determination can still be performed on a nano satellite with low fidelity sensors, but an accurate model of the spacecraft attitude dynamics is required. The passive magnetic stabilization systems commonly employed in nano satellites are known to introduce uncertainties in the parameters of the attitude dynamics model that cannot easily be resolved prior to launch. In this paper, a batch estimation problem is formulated that simultaneously solves for the attitude of the spacecraft and performs parameter estimation on the magnetic properties of the magnetic materials using only a measurement of the solar vector. The estimation technique is applied to data from NASA Ames Research Center's O/OREOS nano satellite and the University of Michigan's RAX-1 nano satellite, where clear differences are detected between the magnetic properties as measured before launch and those that fit the observed data. To date this is the first known on-orbit verification of the attitude dynamics model of a passively magnetically stabilized spacecraft.     

“凤凰”计划关键技术及其启示

为了充分利用地球静止轨道(GEO)的轨道资源,美国于2012年启动了针对GEO卫星的重复利用验证计划——"凤凰"计划。文章介绍了"凤凰"计划的概况,重点对其系统组成、任务等方面进行了介绍,归纳了其中的关键技术,如模块卫星总体设计技术、无线能源传输技术和新型在轨捕获技术等。最后,给出了美国发展在轨服务的若干启示:卫星重用将带来航天器发展的重大变革,注重创新技术开发等。