雷达波束驻留模式下空间碎片数量的置信区间估计

碎片数量估计是空间碎片环境统计特征描述的重要内容之一,对于空间碎片环境模型验证、航天器碰撞风险分析以及碎片数量增长趋势预测有重要意义.针对波束指向正东、正南任意仰角的雷达波束驻留(Beam-park)模式(天顶指向是波束指向仰角为90°时的特例),给出了一种估计碎片数量置信区间的方法.对于给定轨道高度范围内一个具有穿越雷达波束可能性(即雷达散射截面足够大,且轨道倾角相对测站纬度足够大)的碎片,将其是否真正穿越波束这一随机事件用(0-1)分布来建模,根据所采集的轨道高度和倾角数据,计算出该轨道高度范围内碎片穿越波束的平均概率,进而采用中心极限定理来估计碎片数量的置信区间.仿真结果表明了方法的有效性.      

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.     

A debris image tracking using optical flow algorithm

This study proposes a motion detection and object tracking technique for GEO debris in a sequence of images. A couple of techniques (called the “stacking method” and “line-identifying technique”) were recently proposed to address the same problem. Although these techniques are effective at detecting the debris position and motion in the image sequences, there are some issues concerned with computational load and assumed debris motion. This study derives a method to estimate motion vectors of objects in image sequence and finally detect the debris locations by using a computer vision technique called an optical flow algorithm. The new method detects these parameters in low computational time in a serial manner, which implies that it has an advantage to track not only linear but also nonlinear motion of GEO debris more easily than the previous methods. The feasibility of the proposed methods is validated using real and synthesized image sequences which contain some typical debris motions.     

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.     

Assessment of space debris collisions against spacecraft with deorbit devices

Deorbit methods have been employed to remove space debris from orbit. One of these methods is to utilize atmospheric drag. In this method, a membrane loaded into the spacecraft is expanded to increase atmospheric drag. Although this method works without requiring fuel, it has the disadvantage of a high risk of collision with other debris owing to its larger area. Area-time product and energy-to-mass ratio have been used as indices to evaluate the risk of collisions between spacecraft and debris. However, the evaluation criteria were uncertain because these two indices are independent. In this paper, we propose a new evaluation index, single-sheet collision factor (SSCF), that comprehensively evaluates the collision risk based on experiments simulating debris collisions. As a result of the hypervelocity collision experiment, we found that the penetration-area mass of the spacecraft affects the severity of debris collisions. In this paper, the product of the exterior-wall thickness, the exterior-wall density, and the space debris cross-sectional area defines the penetration-area mass of the spacecraft. Furthermore, we compare and evaluate various deorbit methods using SSCF. The comparison showed that the penetration-area mass of the SSCF could be quantitatively determined for the debris-collision severity due to difference in structural materials of spacecraft. SSCF will be used to create rules for space-environment conservation with the expansion of the space-development market.     

Ariane 5 GTO debris mitigation using natural perturbations

GTO objects can potentially collide with operative satellites in LEO and GEO protected regions. Internationally accepted debris mitigation guidelines require that these objects exit these protected regions within 25?years, e.g. by re-entering and burning up in Earth’s atmosphere. In this paper, an inventory of the GTO debris generated from Ariane 5 launches in the period 2012–2017 is provided, and it is expected that none of these objects will re-enter within 25?years. For future launches, natural perturbations can be exploited to increase compliance with mitigation guidelines without the use of extra propellant or complex de-orbiting systems, which is attractive from an economic point of view. The lifetime of GTO objects is very sensitive to initial conditions and some environmental and body-related parameters, mainly due to the effect of solar gravity on the perigee altitude. As a consequence, the lifetime of a specific GTO object cannot be predicted accurately, but its probability of re-entering in less than 25?years can be estimated with proper accuracy by following a statistical approach. By propagating the orbits of over 800,000 simulated Ariane 5 GTO objects, it was found that the launch time leading to the highest probability of compliance with debris mitigation guidelines for GEO launches from Kourou corresponds to about 2 PM local time, regardless of the date of launch, which leads to compliance rates ranging from 60 to 100%. Current practice is to launch at around 5–9?PM, so a change in procedures would be required in order to reach a higher degree of compliance with debris mitigation guidelines, which was predicted to be on average below 20% for the objects generated in the period 2012–2017.     

Micro-meteoroids and space debris impact risk assessment for the ConeXpress satellite using ESABASE2/Debris

Micro-meteoroid and space debris impact risk assessments are performed to investigate the risk from hypervelocity impacts to sensitive spacecraft sub-systems. For these analyses, ESA’s impact risk assessment tool ESABASE2/Debris is used. This software tool combines micro-particle environment models, damage equations for different shielding designs and satellite geometry models to perform a detailed 3D micro-particle impact risk assessment. This paper concentrates on the impact risk for exposed pressurized tanks. Pressure vessels are especially susceptible to hypervelocity impacts when no protection is available from the satellite itself. Even small particles in the mm size range can lead to a fatal burst or rupture of a tank when impacting with a typical collision velocity of 10–20 km/s. For any space mission it has to be assured that the impact risk is properly considered and kept within acceptable limits. The ConeXpress satellite mission is analysed as example. ConeXpress is a planned service spacecraft, intended to extend the lifetime of telecommunication spacecraft in the geostationary orbit. The unprotected tanks of ConeXpress are identified as having a high failure risk from hypervelocity impacts, mainly caused by micro-meteoroids. Options are studied to enhance the impact protection. It is demonstrated that even a thin additional protective layer spaced several cm from the tank would act as part of a double wall (Whipple) shield and greatly reduce the impact risk. In case of ConeXpress with 12 years mission duration the risk of impact related failure of a tank can be reduced from almost 39% for an unprotected tank facing in flight direction to below 0.1% for a tank protected by a properly designed Whipple shield.     

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.     

Atmospheric drag modelling for orbital micro-debris at LEO altitudes

Data from satellite impact experiments and the scanning of recovered spacecraft offers an extended timebase to examine, using a consistent methodology, the microparticle fluxes. New penetration data from the TiCCE experiment on Eureca /1, 2/ adds to this database and shows that - despite an expected growth in the micro-debris flux - the observed flux is not greater than either LDEF or SMM. The question arises: “is this consistent with the micro particle flux being dominated by space debris or by meteoroids”.

To assist this assessment, numerical modelling using the Gear method /3/ of explicit time integration of the atmospheric drag lifetime of micron dimensioned orbital debris in both circular (LEO) and eccentric (GTO) orbits has been performed for the relevant space exposures. Results are applied to the data to examine whether the recent variations in flux can be attributed to varying levels of, orbital micro-debris caused by atmospheric drag and its changes during the solar cycle.     

A Partially Orthogonal EnKF approach to atmospheric density estimation using orbital debris

A key requirement for accurate trajectory prediction and space situational awareness is knowledge of how non-conservative forces affect space object motion. These forces vary temporally and spatially, and are driven by the underlying behavior of space weather particularly in Low Earth Orbit (LEO). Existing trajectory prediction algorithms adjust space weather models based on calibration satellite observations. However, lack of sufficient data and mismodeling of non-conservative forces cause inaccuracies in space object motion prediction, especially for uncontrolled debris objects. The uncontrolled nature of debris objects makes them particularly sensitive to the variations in space weather. Our research takes advantage of this behavior by utilizing observations of debris objects to infer the space environment parameters influencing their motion.The hypothesis of this research is that it is possible to utilize debris objects as passive, indirect sensors of the space environment. We focus on estimating atmospheric density and its spatial variability to allow for more precise prediction of LEO object motion. The estimated density is parameterized as a grid of values, distributed by latitude and local sidereal time over a spherical shell encompassing Earth at a fixed altitude of 400 km. The position and velocity of each debris object are also estimated. A Partially Orthogonal Ensemble Kalman Filter (POEnKF) is used for assimilation of space object measurements to estimate density.For performance comparison, the scenario characteristics (number of objects, measurement cadence, etc.) are based on a sensor tasking campaign executed for the High Accuracy Satellite Drag Model project. The POEnKF analysis details spatial comparisons between the true and estimated density fields, and quantifies the improved accuracy in debris object motion predictions due to more accurate drag force models from density estimates. It is shown that there is an advantage to utilizing multiple debris objects instead of just one object. Although the work presented here explores the POEnKF performance when using information from only 16 debris objects, the research vision is to utilize information from all routinely observed debris objects. Overall, the filter demonstrates the ability to estimate density to within a threshold of accuracy dependent on measurement/sensor error. In the case of a geomagnetic storm, the filter is able to track the storm and provide more accurate density estimates than would be achieved using a simple exponential atmospheric density model or MSIS Atmospheric Model (when calm conditions are assumed).     

DebrisWatch I: A survey of faint geosynchronous debris

Recent anomalies exhibited by satellites and rocket bodies have highlighted that a population of faint debris exists at geosynchronous (GEO) altitudes, where there are no natural removal mechanisms. Despite previous optical surveys probing to around 10–20 cm in size, regular monitoring of faint sources at GEO is challenging, thus our knowledge remains sparse. It is essential that we continue to explore the faint debris population using large telescopes to better understand the risk posed to active GEO satellites. To this end, we present photometric results from a survey of the GEO region carried out with the 2.54 m Isaac Newton Telescope in La Palma, Canary Islands. We probe to 21st visual magnitude (around 10 cm, assuming Lambertian spheres with an albedo of 0.1), uncovering 129 orbital tracks with GEO-like motion across the eight nights of dark-grey time comprising the survey. The faint end of our brightness distribution continues to rise until the sensitivity limit of the sensor is reached, suggesting that the modal brightness could be even fainter. We uncover a number of faint, uncatalogued objects that show photometric signatures of rapid tumbling, many of which straddle the limiting magnitude of our survey over the course of a single exposure, posing a complex issue when estimating object size. This work presents the first instalment of DebrisWatch, an ongoing collaboration between the University of Warwick and the Defence Science and Technology Laboratory (UK) investigating the faint population of GEO debris.     

Removing orbital debris with lasers

Orbital debris in low Earth orbit (LEO) are now sufficiently dense that the use of LEO space is threatened by runaway collision cascading. A problem predicted more than thirty years ago, the threat from debris larger than about 1 cm demands serious attention. A promising proposed solution uses a high power pulsed laser system on the Earth to make plasma jets on the objects, slowing them slightly, and causing them to re-enter and burn up in the atmosphere. In this paper, we reassess this approach in light of recent advances in low-cost, light-weight modular design for large mirrors, calculations of laser-induced orbit changes and in design of repetitive, multi-kilojoules lasers, that build on inertial fusion research. These advances now suggest that laser orbital debris removal (LODR) is the most cost-effective way to mitigate the debris problem. No other solutions have been proposed that address the whole problem of large and small debris. A LODR system will have multiple uses beyond debris removal. International cooperation will be essential for building and operating such a system.     

On the effect of considering more realistic particle shape and mass parameters in MMOD risk assessments

One of the primary mission risks tracked in the development of all spacecraft is that due to micro-meteoroids and orbital debris (MMOD). Both types of particles, especially those larger than 0.1 mm in diameter, contain sufficient kinetic energy due to their combined mass and velocities to cause serious damage to crew members and spacecraft. The process used to assess MMOD risk consists of three elements: environment, damage prediction, and damage tolerance. Orbital debris risk assessments for the Orion vehicle, as well as the Shuttle, Space Station and other satellites use ballistic limit equations (BLEs) that have been developed using high speed impact test data and results from numerical simulations that have used spherical projectiles. However, spheres are not expected to be a common shape for orbital debris; rather, orbital debris fragments might be better represented by other regular or irregular solids. In this paper we examine the general construction of NASA’s current orbital debris (OD) model, explore the potential variations in orbital debris mass and shape that are possible when using particle characteristic length to define particle size (instead of assuming spherical particles), and, considering specifically the Orion vehicle, perform an orbital debris risk sensitivity study taking into account variations in particle mass and shape as noted above. While the results of the work performed for this study are preliminary, they do show that continuing to use aluminum spheres in spacecraft risk assessments could result in an over-design of its MMOD protection systems. In such a case, the spacecraft could be heavier than needed, could cost more than needed, and could cost more to put into orbit than needed. The results obtained in this study also show the need to incorporate effects of mass and shape in mission risk assessment prior to first flight of any spacecraft as well as the need to continue to develop/refine BLEs so that they more accurately reflect the shape and material density variations inherent to the actual debris environment.     

Space Debris Mitigation in France,Germany, Italy and United Kingdom

The Space systems today provide growing benefits to enhance the quality of humankind. However, as a by-product, the orbiting objects inevitably leaves some debris which after 50 years of space activities represent a concern for all space agencies and manufacturers and operators. Since last year no international agreement was in place to mitigate the growing population of space debris objects. The successful result obtained at UN-COPUOS in 2007 and available in the OOSA web site, now gives to the public, a set of voluntary international guidelines that could, if adopted by each space fairing Country, help in maintaining the present space environment. More further steps are necessary in the future to define a legal and normative framework. The paper will present the seven established UN Space Debris guidelines as well as examples of the minimum steps to be carried out at national level to enable the UN-COPUOS to start the discussion of the legal aspect associated with the space debris issue.     

Material selection for a CubeSat structural bus complying with debris mitigation

The increasing number of commercial, technological and scientific missions for CubeSats poses several concerns about the topic of space junk and debris mitigation. As no regulation is currently in place, innovative solutions are needed to mitigate the impact that Low Earth Orbit objects can have during uncontrolled re-entry and the associated potential events of surface collision. We investigated the requirements, in terms of materials selection, for the development of a 3D-printed structural bus able to withstand loads during launch and in-orbit operations, with the objectives to be as light as possible and requiring the least amount of heat for demise during atmospheric re-entry. The selection indicated magnesium alloys as the best candidates to improve the reference material, aluminium 6061 T6, resulting in both mass-reduction and improved demisability. We also analysed how the relative importance of these two objectives can modify the selection of materials: if minimizing the heat to disintegration were valued more highly than lightness, for example, the new best candidates would become tin alloys. Our analysis, furthermore, suggested the importance of Liquid Crystal Polymer as the sole plastic material approaching the performance of the best metal choices. This contribution, thus, provides novel insight in the field of 3D-printed materials for the fast-growing CubeSat segment, complying with the debris mitigation initiatives promoted by space agencies and institutions.     

Protecting and manoeuvring of spacecraft in space debris environment

The paper gives an overview on the fields of debris research performed at the TUBS. The orbital debris flux of all objects larger than 1cm has been established and simulated by a mathematical model in the past mainly on the basis of simulating explosion fragments. However the flux in the millimeter and submillimeter size range seems to be largely influenced by collisions and their ejecta on high circular or on eccentric orbits. The angular distribution of the impact flux on targets at various altitudes and on various inclinations are presented. This angular distribution has also influence on the surface impact flux on a space station, where also the self shielding has to be considered. Results for the ISS are presented. The risk of impacts of larger not shieldable objects on a space station may become too high, so that collision avoidance manoeuvres must be envisaged, the feasibility of which using onboard detectors is discussed.     

Optimising filtering of two-line element sets to increase re-entry prediction accuracy for GTO objects

Predicting re-entry epoch of space objects enables managing the risk to ground population. Predictions are particularly difficult for objects in highly-elliptical orbits, and important for objects with components that can survive re-entry, e.g. rocket bodies (R/Bs). This paper presents a methodology to filter two-line element sets (TLEs) to facilitate accurate re-entry prediction of such objects. Difficulties in using TLEs for precise analyses are highlighted and a set of filters that identifies erroneous element sets is developed. The filter settings are optimised using an artificially generated TLE time series. Optimisation results are verified on real TLEs by analysing the automatically found outliers for exemplar R/Bs. Based on a study of 96 historical re-entries, it is shown that TLE filtering is necessary on all orbital elements that are being used in a given analysis in order to avoid considerably inaccurate results.     

防护结构在斜撞击下的行为研究

以惠普尔（Ｗｈｉｐｐｌｅ）防护屏结构为例,通过实验研究了铝合金柱状弹体以第一门槛值附近速度斜撞击ＬＹ１２合金防护屏的特点,研究了不同速度下,不同撞击角度下,防护屏损伤的变化规律,对于速度在第一门槛值附近的斜撞击,以斜撞击存在一个临界入射角度,当入射角度大于临界入射角度时,发生滑弹现象,入射角度越大,产生的滑弹碎片数量就越多。