Progress in Space Debris Research

During recent years, A de-orbit disposal of SinoSat 2 satellite and the depletion of the residual propellant after SC/LV separation for all LM-4 series launch vehicles were carried out. Stuffed Whipple Shields based on hypervelocity impact particles were developed. Routine observation and collision avoidance were performed. The main progress in space debris research will be introduced from three aspects: mitigation, spacecraft protection, observation and collision avoidance.      

基于Adaboost的填充式防护结构超高速撞击损伤预测

填充式防护结构的显式弹道极限方程在对弹丸进行超高速撞击损伤预测时,由于填充材料、填充方式的不同,会导致预测结果与实测数据存在一定偏差。对此,采用机器学习方式将该问题转化为二分类问题,以碰撞过程中的弹丸撞击参数、防护结构参数作为分类特征,构建了基于Adaboost的填充式防护结构超高速撞击损伤预测模型。该模型以分类回归树（CART）作为弱分类器,通过对一系列弱分类器的加权组合生成强分类器,并通过对训练样本的循环使用,实现了小样本集下的撞击损伤预测。实验结果表明,建立的Adaboost预测模型对填充式防护结构的超高速撞击损伤具有良好的预测效果,总体预测率与安全预测率相比于NASA的弹道极限方程均提高了14.3%,具有更强的通用性。通过不同训练样本规模下的交叉检验,证明了该模型具有良好的鲁棒性与准确性。      

The Giotto mission to Halley''s comet

ESA's Giotto mission to Halley's comet is a fast flyby in March 1986, about four weeks after the comet's perihelion passage when it is most active. The scientific payload comprises 10 experiments with a total mass of about 60 kg: a camera for imaging the comet nucleus, three mass spectrometers for analysis of the elemental and isotopic composition of the cometary gas and dust environment, various dust impact detectors, a photopolarimeter for measurements of the coma brightness, and a set of plasma instruments for studies of the solar wind/comet interaction. In view of the high flyby velocity of 68 km/s the experiment active time is very short (only 4 hours) and all data are transmitted back to Earth in real time at a rate of 40 kbps. The Giotto spacecraft is spin-stabilised with a despun high gain parabolic dish antenna inclined at 44.3° to point at the Earth during the encounter while a specially designed dual-sheet bumper shield at the other end protects the spacecraft from being destroyed by hypervelocity dust impacts. The mission will probably end near the point of closest approach to the nucleus when the spacecraft attitude will be severely perturbed by impacting dust particles leading to a loss of the telecommunications link.     

基于缩放实验的Whipple结构撞击特性

针对航天器空间碎片防护问题,基于缩放实验方法,开展了7 km/s以上超高速碰撞仿真研究.建立了单板和Whipple防护结构的仿真模型,并对铝-铝撞击问题和镉-镉撞击问题进行了多工况仿真.通过实验结果与数值仿真的对比,表明了数值仿真技术的正确性,并从仿真角度验证了缩放实验方法的有效性.对缩放实验方法的适用性进行了仿真验证,结果表明该方法对弹丸形状适用性较好,对3～4 km/s以上撞击速度的适用性较好,但对Whipple防护结构后板存在一定误差.分析了Whipple结构后板的失效模式,提出了失效模式的不连续性导致了缩放实验方法的误差.最后通过数值仿真计算了Whipple结构7 km/s以上弹道极限特性,提出了失效模式的不连续性造成了在该速度段弹道极限曲线的分叉现象.     

CubeSail: A low cost CubeSat based solar sail demonstration mission

CubeSail is a nano-solar sail mission based on the 3U CubeSat standard, which is currently being designed and built at the Surrey Space Centre, University of Surrey. CubeSail will have a total mass of around 3 kg and will deploy a 5 × 5 m sail in low Earth orbit. The primary aim of the mission is to demonstrate the concept of solar sailing and end-of-life de-orbiting using the sail membrane as a drag-sail. The spacecraft will have a compact 3-axis stabilised attitude control system, which uses three magnetic torquers aligned with the spacecraft principle axis as well as a novel two-dimensional translation stage separating the spacecraft bus from the sail. CubeSail’s deployment mechanism consists of four novel booms and four-quadrant sail membranes. The proposed booms are made from tape-spring blades and will deploy the sail membrane from a 2U CubeSat standard structure. This paper presents a systems level overview of the CubeSat mission, focusing on the mission orbit and de-orbiting, in addition to the deployment, attitude control and the satellite bus.     

基于电学成像的蜂窝夹层结构超高速撞击损伤监测

针对航天器遭受空间碎片和微流星体撞击的问题,对蜂窝夹层结构的超高速撞击损伤监测进行研究。提出将碳纳米管薄膜共固化在蜂窝夹层结构面板表面使之具有自感应能力,结合电学成像技术对超高速撞击造成的损伤进行监测和识别。采用二级轻气炮对自感应蜂窝夹层结构进行了超高速撞击,在撞击前后分别向感应层注入微小的激励电流,根据边界电压变化重建损伤引起的电导率变化图像,从而提供有关撞击和损伤的信息。试验结果表明,基于碳纳米管薄膜的感应层性能良好,重建的电导率变化图像能够较好地反映损伤个数、位置和近似尺寸,验证了所提出技术方法的有效性,为航天器结构超高速撞击监测提供了一种新的技术手段。     

Impact effects from small size meteoroids and space debris

When the impact risk from meteoroids and orbital debris is assessed the main concern is usually structural damage. With their high impact velocities of typically 10–20 km/s millimeter or centimeter sized objects can puncture pressure vessels and other walls or lead to destruction of complete subsystems or even whole spacecraft. Fortunately chances of collisions with such larger objects are small (at least at present). However, particles in the size range 1–100 μm are far more abundant than larger objects and every orbiting spacecraft will encounter them with certainty. Every solar cell (8 cm² area) of the Hubble Space Telescope encountered on average 12 impacts during its 8.25 years of space exposure. Most were from micron sized particles.     

Improved kHz-SLR tracking techniques and orbit quality analysis for LEO missions

Satellite gravity field missions such as CHAMP, GRACE and GOCE are designed as low Earth orbiting spacecraft (LEO) with orbit heights of about 250–500 km. The challenging mission objectives require a very precise knowledge of the satellite orbit position in space. For these missions precise orbit information is typically provided by GPS satellite-to-satellite tracking (SST) observations supported by satellite laser ranging (SLR).     

On the validity of the aluminum equivalent approximation in space radiation shielding applications

The origin of the aluminum equivalent shield approximation in space radiation analysis can be traced back to its roots in the early years of the NASA space programs (Mercury, Gemini and Apollo) wherein the primary radiobiological concern was the intense sources of ionizing radiation causing short term effects which was thought to jeopardize the safety of the crew and hence the mission. Herein, it is shown that the aluminum equivalent shield approximation, although reasonably well suited for that time period and to the application for which it was developed, is of questionable usefulness to the radiobiological concerns of routine space operations of the 21st century which will include long stays onboard the International Space Station (ISS) and perhaps the moon. This is especially true for a risk based protection system, as appears imminent for deep space exploration where the long-term effects of Galactic Cosmic Ray (GCR) exposure is of primary concern. The present analysis demonstrates that sufficiently large errors in the interior particle environment of a spacecraft result from the use of the aluminum equivalent approximation, and such approximations should be avoided in future astronaut risk estimates. In this study, the aluminum equivalent approximation is evaluated as a means for estimating the particle environment within a spacecraft structure induced by the GCR radiation field. For comparison, the two extremes of the GCR environment, the 1977 solar minimum and the 2001 solar maximum, are considered. These environments are coupled to the Langley Research Center (LaRC) deterministic ionized particle transport code High charge (Z) and Energy TRaNsport (HZETRN), which propagates the GCR spectra for elements with charges (Z) in the range 1 ? Z ? 28 (H–Ni) and secondary neutrons through selected target materials. The coupling of the GCR extremes to HZETRN allows for the examination of the induced environment within the interior of an idealized spacecraft as approximated by a spherical shell shield, and the effects of the aluminum equivalent approximation for a good polymeric shield material such as generic polyethylene (PE). The shield thickness is represented by a 25 g/cm² spherical shell. Although, one could imagine the progression to greater thickness, the current range will be sufficient to evaluate the qualitative usefulness of the aluminum equivalent approximation. Upon establishing the inaccuracies of the aluminum equivalent approximation through numerical simulations of the GCR radiation field attenuation for PE and aluminum equivalent PE spherical shells, we further present results for a limited set of commercially available, hydrogen rich, multifunctional polymeric constituents to assess the effect of the aluminum equivalent approximation on their radiation attenuation response as compared to the generic PE.     

Synergistic approach of asteroid exploitation and planetary protection

The asteroid and cometary impact hazard has long been recognised as an important issue requiring risk assessment and contingency planning. At the same time asteroids have also been acknowledged as possible sources of raw materials for future large-scale space engineering ventures. This paper explores possible synergies between these two apparently opposed views; planetary protection and space resource exploitation. In particular, the paper assumes a 5 tonne low-thrust spacecraft as a baseline for asteroid deflection and capture (or resource transport) missions. The system is assumed to land on the asteroid and provide a continuous thrust able to modify the orbit of the asteroid according to the mission objective. The paper analyses the capability of such a near-term system to provide both planetary protection and asteroid resources to Earth. Results show that a 5 tonne spacecraft could provide a high level of protection for modest impact hazards: airburst and local damage events (caused by 15–170 m diameter objects). At the same time, the same spacecraft could also be used to transport to bound Earth orbits significant quantities of material through judicious use of orbital dynamics and passively safe aero-capture manoeuvres or low energy ballistic capture. As will be shown, a 5 tonne low-thrust spacecraft could potentially transport between 12 and 350 times its own mass of asteroid resources by means of ballistic capture or aero-capture trajectories that pose very low dynamical pressures on the object.     

Assessing attitude error of FORMOSAT-3/COSMIC satellites and its impact on orbit determination

An attitude determination and control system (ADCS) is critical to satellite attitude maneuvers and to the coordinate transformation from the inertial frame to the spacecraft frame. This paper shows specific sensors in the ADCS of the satellite mission FORMOSAT-3/COSMIC (F3/C) and the impact of the ADCS quality on orbit accuracy. The selection of main POD antenna depends on the beta angles of the different F3/C satellites (for FM2 and FM4) during the inflight phase. In particular, under the eclipse, alternative attitude sensors are activated to replace the Sun sensors, and such a sensor change leads to anomalous GPS phase residuals and a degraded orbit accuracy. Since the nominal attitude serves as a reference for ADCS, the 3-dimensional attitude-induced errors in reduced dynamic orbits over selected days in 2010 show 9.35, 10.78, 4.97, 5.48, 7.18, and 6.89 cm for FM1–FM6. Besides, the 3-dimensional velocity errors induced by the attitude effect are 0.10, 0.10, 0.07, 0.08, 0.09, and 0.10 for FM1–FM6. We analyze the quality of the observed attitude transformation matrix of F3/C and its impact on kinematic orbit determination. With 249 days of GPS in 2008, the analysis leads to the following averaged 3-dimensional attitude-induced orbit errors: 2.72, 2.62, 2.37, 1.90, 1.70, and 1.99 cm for satellites FM1–FM6. Critical suggestions of geodetic payloads for the follow-on mission of F3/C are presented based on the current result.     

空间飞行器碎片防护结构的简化设计

给出一种基于实验和理论分析的航天器碎片防护结构简化设计方法 ,该方法可用于进行大型空间飞行器碎片防护结构的方案选择和初步结构设计。利用空间碎片的工程环境模型和防护结构几何经验公式 ,采用“设计碎片”的概念 ,对防护结构进行几何结构设计和质量估算 ,并采用改进的防护性能验证算法进行空间碎片的风险评估。通过对惠式防护结构的计算 ,得到的计算结果基本符合实际要求。     

Laboratory calibration measurements of a piezoelectric lead zirconate titanate cosmic dust detector at low velocities

A cosmic dust monitor for use onboard a spacecraft is currently being developed using a piezoelectric lead zirconate titanate element (PZT). Its characteristics of the PZT sensor is studied by ground-based laboratory impact experiments using hypervelocity particles supplied by a Van de Graaff accelerator. The output signals obtained from the sensor just after the impact appeared to have a waveform that was explicitly related to the particle’s impact velocity. For velocities less than ∼6 km/s, the signal showed an oscillation pattern and the amplitude was proportional to the momentum of the impacting particle. For higher velocities, the signal gradually changed to a single waveform. The rise time of this single waveform was proportional to the particle’s velocity for velocities above ∼6 km/s. The present paper reports on results for the low velocity case and especially discusses the effect of an outer coating of the sensor with a paint, which is used to reduce heating by solar radiation.     

Large constellations assessment and optimization in LEO space debris environment

Recent plans for large constellations in Low-Earth Orbit have opened the debate on both their vulnerability and their influence on the already hazardous space debris environment. In fact, given that large constellations normally employ satellites of small size, there might be situations in which cm-size debris could have enough energy to cause fragmentation of a significant part of these spacecraft upon impact, while smaller debris could affect the functionalities of critical subsystems, even compromising the success of disposal operations planned at end-of-life. In this context, this paper investigates: (1) collisions with large objects that could initiate the fragmentation of a significant part of the satellite, and (2) impacts with small debris that might perforate the spacecraft hull thus causing relevant performance/functionality degradation. These two points are merged in a simple statistical tool for risk assessment, which analyses the effects of the main parameters of the constellations on its vulnerability (i.e. operational life, number of satellites, spacecraft cross section, satellites reliability). In more details, the tool relates impact probability (for both small and large debris) to the ballistic response of spacecraft structures and protections, defining the critical configurations that might compromise the expected disposal operations. This method requires a limited knowledge of the spacecraft internal layout, as it is based on a statistical analysis of impact damage instead of a complete evaluation of the vulnerability of each subsystem. In parallel, non-debris related failures are also investigated and statistic models of spacecraft reliability characteristic are proposed. Among the results, it is shown that reducing the lifetime of individual satellites in a constellation might improve the success rate of post-mission disposal, thanks to the reduction of the spacecraft exposure to the space environment with the consequential degradation of its performance. On the other hand, reducing the lifetime would seriously affect the debris environment: the increase in traffic in the most crowded altitudes would be not counterbalanced by the higher post mission disposal success rate, causing an overall increase of the total number of uncontrolled resident objects.     

R3D-B2 – Measurement of ionizing and solar radiation in open space in the BIOPAN 5 facility outside the FOTON M2 satellite

Solar and space radiation have been monitored using the R3D-B2 radiation risks radiometer-dosimeter on board a recent space flight on the Russian satellite Foton M2 within the ESA Biopan 5 facility mounted on the outside of the satellite exposed to space conditions. The solar radiation has been assayed in four wavelength bands (UV-C, 170–280 nm, UV-B, 280–315 nm), UV-A (315–400 nm) and PAR (photosynthetic active radiation, 400–700 nm). The data show an increasing tumbling rotation of the satellite during the mission. The photodiodes do not show a cosine response to the incident light which has been corrected. After calibration of the signals using the extraterrestrial spectrum, doses have been calculated for each orbit, for each day and for the total mission as basic data for the biological material which has been exposed in parallel in the Biopan facility. Cosmic ionizing radiation has been monitored and separated in 256 deposited energy spectra, which were further used for determination of the absorbed dose rate and flux. Basic data tables were prepared to be used by other Biopan 5 experiments. The paper summarizes the results for the Earth radiation environment at the altitude (262–304 km) of the Foton M2 spacecraft. Comparisons with the predictions of NASA Earth radiation environment experimental models AE-8 and AP-8, and the PSB97 model are also presented, which calculate the fluxes of ionizing radiation from a simulation. AP-8 is a model for trapped radiation.     

Hypervelocity impact on the GIOTTO Halley Mission dust shield: Momentum exchange and measurement

The two layer dust shield on the GIOTTO Halley Mission is constructed in a meteoroid bumper configuration. The dust shield is instrumented so that parameters associated with the hypervelocity collision of cometary particles on the exposed surface can be determined. A multisensor detector array provides simultaneous sensing of the momentum exchange of particles impacting and subsequently penetrating the outer layer of the dust shield. Current knowledge of momentum exchange during hypervelocity impact relative to the GIOTTO Halley Mission and the dust shield experiment is reviewed. The sensors used for determination of momentum exchange exhibit a functional dependence on projectile velocity leading to an enhancement of the sensor signal as the relative impact velocity increases. The GIOTTO Mission provides a very unique opportunity to obtain hypervelocity momentum exchange information at a known impact velocity. Therefore, with the dust experiment, a determination of the velocity index for both momentum and multilayered penetration sensor is possible. Results of analysis of analytical and laboratory studies indicate that the velocity index for hypervelocity impact is approximately 2.0 at the 68 km/sec encounter impact velocity of the GIOTTO Mission. A clear determination of the size and mass distribution of the cometary dust near the comet will be possible from the in-situ measurement of the DIDSY GIOTTO experiment.     

Magnetic torquer induced disturbing signals within GRACE accelerometer data

The GRACE (Gravity Recovery And Climate Experiment) gravity field satellite mission was launched in 2002. Although many investigations have been carried out, not all disturbances and perturbations upon satellite instruments and sensors are resolved yet. In this work the issue of acceleration disturbances onboard of GRACE due to magnetic torquers is investigated and discussed. Each of the GRACE satellites is equipped with a three-axes capacitive accelerometer to measure non-gravitational forces acting on the spacecraft. We used 10 Hz Level 1a raw accelerometer data in order to determine the impact of electric current changes on the accelerometer. After reducing signals which are induced by highly dominating processes in the low frequency range, such as thermospheric drag and solar radiation pressure, which can easily be done by applying a high-pass filter, disturbing signals from onboard instruments such as thruster firing events or heater switch events need to be removed from the previously filtered data. Afterwards the spikes which are induced by the torquers can be very well observed. Spikes vary in amplitude with respect to an increasing or decreasing current used for magnetic torquers, and can be as large as 20 nm/s². Furthermore, we were able to set up a model for the spikes of each scenario with which we were able to compute model spike time series. With these time series the spikes can successfully be removed from the 10 Hz raw accelerometer data. Spectral analysis of the time series reveal that an influence onto gravity field determination due to these effects is very unlikely, but can theoretically not be excluded.     

Development of in-situ Space Debris Detector

Due to high relative velocities, collisions of spacecraft in orbit with Space Debris (SD) or Micrometeoroids (MM) can lead to payload degradation, anomalies as well as failures in spacecraft operation, or even loss of mission. Flux models and impact risk assessment tools, such as MASTER (Meteoroid and Space Debris Terrestrial Environment Reference) or ORDEM (Orbital Debris Engineering Model), and ESABASE2 or BUMPER II are used to analyse mission risk associated with these hazards. Validation of flux models is based on measured data. Currently, as most of the SD and MM objects are too small (millimeter down to micron sized) for ground-based observations (e.g. radar, optical), the only available data for model validation is based upon retrieved hardware investigations e.g. Long Duration Exposure Facility (LDEF), Hubble Space Telescope (HST), European Retrievable Carrier (EURECA). Since existing data sets are insufficient, further in-situ experimental investigation of the SD and MM populations are required. This paper provides an overview and assessment of existing and planned SD and MM impact detectors. The detection area of the described detectors is too small to adequately provide the missing data sets. Therefore an innovative detection concept is proposed that utilises existing spacecraft components for detection purposes. In general, solar panels of a spacecraft provide a large area that can be utilised for in-situ impact detection. By using this method on several spacecraft in different orbits the detection area can be increased significantly and allow the detection of SD and MM objects with diameters as low as 100 μm. The design of the detector is based on damage equations from HST and EURECA solar panels. An extensive investigation of those panels was performed by ESA and is summarized within this paper. Furthermore, an estimate of the expected sensitivity of the patented detector concept as well as examples for its implementation into large and small spacecraft are presented.     

Validation of a Compression Mass Gauge using ground tests for liquid propellant mass measurements

To properly estimate orbital lifetimes and predict the maneuverability of spacecraft, the remaining liquid propellant mass must be accurately known at every moment of a space mission. This paper studies the Compression Mass Gauge (CMG) method to determine the mass of liquid contained in a tank in a low-gravity environment with high accuracy. CMG is a thermodynamic method used to determine the quantity of liquid by measuring the gas pressure change when the tank volume changes, and has been previously theoretically and experimentally studied by researchers. The primary objective of this investigation is to explore the effects of attitude disturbance and the spacecraft thermal environment on the accuracy of the method. A ground test system, consisting of several test apparatuses, was fabricated and described as part of this study. The test results and analyses indicate that the CMG performs well and has an accuracy of ±1%. Additionally, demonstrations were performed to show that measurement errors do not increase drastically or exceed ±1% when the test system is vibrated to simulate the tank being perturbed as a result of an attitude disturbance. Liquid sloshing resonance was found to have a significant effect on the gauging accuracy. Measurements in a real thermal environment in which heat transfers into and out of the propellant tank were also conducted. The results show that the gauging accuracy is acceptable for normal liquid propellant. Furthermore, theoretical research shows that heat leakage has a significant influence on cryogenic propellant mass gauging.