A deorbiting drag device is being designed and built by the University of Toronto Institute for Aerospace Studies/Space Flight Laboratory (UTIAS/SFL) to be demonstrated on the Canadian Advanced Nanospace eXperiment 7 (CanX-7) satellite. CanX-7 will address the growing issue of space debris by designing a drag sail device that will be demonstrated for cubesat-sized satellites. Mission analysis done to ensure the drag device functions properly and deorbits within the required lifetime is performed while varying different properties such as drag coefficient, effective drag area, and solar cycle variations. The design evolution of the device is documented and the chosen design, along with several stages of prototyping, is described. The individual components that make up the device are described as are preliminary numerical analyzes. Finally, the test plan required for the device is described with several deployment experiments and risk reduction tests documented.

A deorbiting drag device is being designed and built by the University of Toronto Institute for Aerospace Studies/Space Flight Laboratory (UTIAS/SFL) to be demonstrated on the Canadian Advanced Nanospace eXperiment 7 (CanX-7) satellite. CanX-7 will address the growing issue of space debris by designing a drag sail device that will be demonstrated for cubesat-sized satellites. Mission analysis done to ensure the drag device functions properly and deorbits within the required lifetime is performed while varying different properties such as drag coefficient, effective drag area, and solar cycle variations. The design evolution of the device is documented and the chosen design, along with several stages of prototyping, is described. The individual components that make up the device are described as are preliminary numerical analyzes. Finally, the test plan required for the device is described with several deployment experiments and risk reduction tests documented.

Outer Space Activities & Environmental Protection Abstract This thesis aims at addressing various connecting factors between outer space activities and environmental protection. It examines the present legal regime of the outer space with particular regard to the environmental protection. The analysis of the most relevant space treaties is provided, reviewing the provisions concerning the protection of the environment, while addressing the emerging environmental issues. The significant role of international bodies is being emphasized, as well as the current challenges the regime of outer space is facing. Although the space-related activities offer an extensive tool to enhance environmental protection of Earth, the environment of space is nowadays being exposed to an imminent threat called space debris, which could prevent all possible future space activities if left unattended. The pollution of the space environment is at a critical level, and the international community is required to respond within a short time in the future. The introduction of mitigating efforts will no longer itself be effective without the addition of active employment of remediating measures, such as the active debris removal. On the other hand, space activities already widely contribute to preserving the environment of Earth...

The constant growth of space debris and the associated risks force the space community to find solutions to mitigate them. Today the most advanced solutions to dispose of satellites and rocket stages after the end of mission consists of moving them either into a graveyard orbit or towards an atmospheric re-entry ending in the demise of both spacecraft and its materials. Alternative solutions should be considered, such as providing a sustainable solution by reusing materials in space. However, it is crucial to understand better the ageing process of the materials present in currently active spacecraft and space debris. The space environment causes degradation and damage over time, making the state of those materials uncertain for potential re-use. Degradation effects have been studied as a source mechanism to result in paint flakes, ejecta particles, or delaminated insulation foils released into the space environment and sustaining a positive feedback loop through potential impacts into spacecraft. A better understanding of degradation effects would also help to better characterize the small debris environment and its evolution. The current materials databases used by the space industry could be useful tools to select materials for satellite missions with respect to their reusability, but they often do not include the evolution of material properties in space after the end of mission.This study will investigate the impact of the damage effects of radiation and meteoroid impact on solar panels. During this research, the methodology used to analyse these effects was explained. The results showed that radiation caused the most damage and could cause solar panels to lose more than a third of their performance over a period of 50 years. It was therefore possible to estimate the quantity of solar panels available for re-use. It was concluded that the results were valid, but that to obtain more accurate data, all the different types of deterioration would also have to be considered. / Creaternity

Approved for public release; distribution in unlimited. / The International Space Station (ISS) must be able to withstand the hypervelocity impacts of micrometeoroids and orbital debris that strike its many surfaces. In order to design and implement shielding which will prevent hull penetration or other operational losses, NASA must first model the orbital debris and micrometeoroid environment. Based upon this environment, special multi-stage shields called Whipple and Enhanced Stuffed Whipple Shields are developed and implemented to protect ISS surfaces. Ballistic limit curves that establish shield failure criteria are determined via ground testing. These curves are functions of material strength, shield spacing, projectile size, shape and density, as well as a number of other variables. The combination of debris model and ballistic limit equations allows NASA to model risk to ISS using a hydro-code called BUMPER. This thesis modifies and refines existing ballistic limit equations for U.S. Laboratory Module shields to account for the effects of projectile (debris/ micro-meteoroid) densities. Using these refined ballistic limit equations this thesis also examines alternative shielding materials and configurations to optimize shield design for minimum mass and maximum stopping potential, proposing alternate shield designs for future NASA ground testing. A final goal of this thesis is to provide the Department of Defense a background in satellite shield theory and design in order to improve protection against micrometeoroid and orbital debris impacts on future spacebased national systems. / Lieutenant, United States Navy

L’analyse de cycle de vie environnementale (ACV), d’après l’ISO 14040/44, a été identifiée par l’agence spatiale Européenne, ArianeGroup et plusieurs acteurs de l’industrie spatiale européenne comme la méthodologie la mieux adaptée pour réaliser l’évaluation environnementale des missions spatiales. Cependant, le secteur spatial est un domaine d’activité dont la particularité est de s’étendre au-delà des limites de l’environnement terrestre. Ainsi, s’il est possible d’évaluer les impacts sur l’environnement classique via la méthodologie ACV, les impacts générés sur l’environnement orbital ne sont aujourd’hui pas considérés dans son cadre de référence.Par conséquent, il s’agit ici d’étendre le champ des études ACV dans le domaine du spatial afin de caractériser les étapes du cycle de vie ayant lieu en orbite, c’est-à-dire la phase d’usage pendant la mission spatiale et la phase de fin de vie (ou Post-Mission Disposal). L’anticipation de cette dernière est devenue une étape cruciale dans la conception des missions spatiales du fait des règlementations visant à limiter la prolifération des débris spatiaux. En effet, seulement 6% de la population en orbite autour de la Terre sont des satellites actifs, le reste étant considéré comme des débris, résultat des missions et activités spatiales passées. Etant donné cet enjeu de durabilité majeur pour l'industrie spatiale, les études environnementales se doivent de mettre en évidence les transferts de pollution potentiels non seulement entre les impacts environnementaux classiques tels que Changement Climatique ou la Toxicité mais aussi ceux relatifs à l’environnement orbital, au premier rang desquels figurent les débris spatiaux.Afin de caractériser cet impact sur l’environnement orbital, nous proposons de définir une nouvelle catégorie d’impact en ACV dénommée ‘Orbital Space Use’. La conformité avec le cadre de référence de l’ACV est assurée au travers de la construction de chaines de cause-à-effet (ou impact pathways) reliant les flux élémentaires du système d’étude aux mécanismes environnementaux (indicateur midpoint) et au dommage final (indicateur endpoint) au sein de l’Aire de Protection ‘Ressources Naturelles’. En effet, les orbites proches de la Terre qui supportent les activités spatiales sont considérées à ce titre comme une ressource pouvant être impactée par des ‘stresseurs’. Les débris spatiaux sont aujourd’hui identifiés comme les principaux ‘stresseurs’ vis-à-vis de l’accès et de la pleine utilisation de la ressource orbitale. Ainsi, le développement d’un modèle de caractérisation a été entrepris. Il permet d’évaluer l’impact potentiel d’une mission spatiale sur la population des débris déjà présente en orbite. Les facteurs de caractérisation obtenus sont appliqués à un cas d’étude comparant trois scénarios de fin de vie différents dans le but de prouver l’applicabilité du modèle. En outre, une première approche s’intéressant à la quantification des externalités économiques négatives engendrées par la prolifération des débris est développée. Elle constitue une étape préliminaire en vue d’un développement d’une catégorie de dommage dite ‘endpoint’. Enfin, les challenges méthodologiques restants et les potentiels développements complémentaires sont abordés. / Several actors of or related to the European space industry, such as ArianeGroup and the European Space Agency (ESA), have identified life cycle assessment (according to ISO14040/44) as the most appropriate methodology to measure and minimise their environmental impact. Nevertheless, space systems deal with a strong particularity which adds new aspects considering the scope of the LCA framework. Space missions are the only human activity that crosses all segments of the atmosphere and stays “out” of the natural environment and ecosystems. Regarding space systems with a holistic approach, environmental impacts could occur not only in the conventional ecosphere but also in the outer space (i.e. the orbital environment).Consequently, the current scope of LCA studies should be broadened to take into account the on-orbit lifetime as well as the end-of-life disposal of the spacecraft. Yet, it is becoming a crucial point of the space mission design due to the future increase of the orbital population composed in a major part by space debris. In this way, LCA studies of space missions could indicate trade-offs not only between typical impact categories (toxicity and climate change for example) but also with regard to impacts generated in the orbital environment with a particular focus on space debris related impacts.Hence, the priority has been given to the integration of a new impact category called orbital space use in the life cycle impact assessment framework. To address the environmental burdens comprehensively in this new category, impact pathways linking elementary flows to environmental mechanisms (midpoint) and damages (endpoint) are developed within the Area-of-Protection ‘Natural resources’. Space debris is identified as the main stressor of the orbital environment. Thus, ‘characterisation factors’ are defined and calculated at midpoint level to assess the potential impact of a space mission on the orbital environment. The methodology is applied to a case study to prove its applicability: the potential impact of a theoretical space mission is addressed through the comparison of three disposal scenarios. Also, a first attempt regarding the characterisation of the endpoint damage is provided taking into account the economic externalities caused by space debris. Finally, remaining methodological challenges and perspectives for future work are provided.

Vikram Sarabhai Space Centre,Trivandrum / In the space surrounding the earth there are two major regions where orbital debris causes concern. They are the Low Earth Orbits (LEO) up to about 2000 km, and Geosynchronous Orbits (GEO) at an altitude of around 36000 km. The impact of the debris accumulations are in principle the same in the two regions; nevertheless they require different approaches and solutions, due to the fact that the perturbations in the orbital decay due to atmospheric drag effects predominates in LEO, gravitational forces including earth’s oblateness and luni solar effects dominating in GEO are different in these two regions. In LEO it is generally known that the debris population dominates even the natural meteoroid population for object sizes 1 mm and larger. This thesis focuses the study mainly in the LEO region. Since the first satellite breakup in 1961 up to 01 January 2003 more than 180 spacecraft and rocket bodies have been known to fragment in orbit. The resulting debris fragments constitute nearly 40% of the 9000 or more of the presently tracked and catalogued objects by USSPACECOM. The catalogued fragment count does not include the much more numerous fragments, which are too small to be detected from ground. Hence in order to describe the trackable orbital debris environment, it is important to develop mathematical models to simulate the trackable fragments and later expand it to untrackable objects. Apart from the need to better characterize the orbital debris environment down to sub millimeter particles, there is also a pressing necessity of simulation tools able to model in a realistic way the long term evolution of space debris, to highlight areas, which require further investigations, and to study the actual mitigation effects of space policy measures. The present thesis has provided newer perspectives for five major issues in space debris modeling studies. The issues are (i) breakup modeling, (ii) environment modeling, (iii) evolution of the debris environment, (iv) collision probability analysis and (v) reentry prediction. The Chapter 1 briefly describes an overview of space debris environment and the issues associated with the growing space debris populations. A literature survey of important earlier work carried out regarding the above mentioned five issues are provided in the Chapter 2. The new contributions of the thesis commence from Chapter 3. The Chapter 3 proposes a new breakup model to simulate the creation of debris objects by explosion in LEO named “A Semi Stochastic Environment Modeling for Breakup in LEO” (ASSEMBLE). This model is based on a study of the characteristics of the fragments from on orbit breakups as provided in the TLE sets for the INDIAN PSLV-TES mission spent upper stage breakup. It turned out that based on the physical mechanisms in the breakup process the apogee, perigee heights (limited by the breakup altitude) closely fit suitable Laplace distributions and the eccentricity follows a lognormal distribution. The location parameters of these depend on the orbit of the parent body at the time of breakup and their scale parameters on the intensity of explosion. The distribution of the ballistic coefficient in the catalogue was also found to follow a lognormal distribution. These observations were used to arrive at the proper physical, aerodynamic, and orbital characteristics of the fragments. Subsequently it has been applied as an inverse problem to simulate and further validate it based on some more typical well known historical on orbit fragmentation events. All the simulated results compare quite well with the observations both at the time of breakup and at a later epoch. This model is called semi stochastic in nature since the size and mass characteristics have to be obtained from empirical relations and is capable of simulating the complete scenario of the breakup. A new stochastic environment model of the debris scenario in LEO that is simple and impressionistic in nature named SIMPLE is proposed in Chapter 4. Firstly among the orbital debris, the distribution of the orbital elements namely altitude, perigee height, eccentricity and the ballistic coefficient values for TLE sets of data in each of the years were analyzed to arrive at their characteristic probability distributions. It is observed that the altitude distribution for the number of fragments exhibits peaks and it turned out that such a feature can be best modeled with a tertiary mixture of Laplace distributions with eight parameters. It was noticed that no statistically significant variations could be observed for the parameters across the years. Hence it is concluded that the probability density function of the altitude distribution of the debris objects has some kind of equilibrium and it follows a three component mixture of Laplace distributions. For the eccentricity ‘e’ and the ballistic parameter ‘B’ values the present analysis showed that they could be acceptably quite well fitted by Lognormal distributions with two parameters. In the case of eccentricity also the describing parameter values do not vary much across the years. But for the parameters of the B distribution there is some trend across the years which perhaps may be attributed to causes such as decay effect, miniaturization of space systems and even the uncertainty in the measurement data of B. However in the absence of definitive cause that can be attributed for the variation across the years, it turns out to be best to have the most recent value as the model value. Lastly the same kind of analysis has also been carried out with respect to the various inclination bands. Here the orbital parameters are analyzed with respect to the inclination bands as is done in ORDEM (Kessler et al 1997, Liou et al 2001) for near circular orbits in LEO. The five inclination bands considered here are 0-36 deg (in ORDEM they consider 19-36 deg, and did not consider 0-19 deg), 36-61 deg, 61-73 deg, 73-91 deg and 91- 180 deg, and corresponding to each band, the altitude, eccentricity and B values were modeled. It is found that the third band shows the models with single Laplace distribution for altitude and Lognormal for eccentricity and B fit quite well. The altitude of other bands is modeled using tertiary mixture of Laplace distributions, with the ‘e’ and ‘B’ following once again a Lognormal distribution. The number of parameter values in SIMPLE is, in general, just 8 for each description of altitude or perigee distributions whereas in ORDEM96 it is more. The present SIMPLE model captures closely all the peak densities without losing the accuracy at other altitudes. The Chapter 5 treats the evolution of the debris objects generated by on orbit breakup. A novel innovative approach based on the propagation of an equivalent fragment in a three dimensional bin of semi major axis, eccentricity, and the ballistic coefficient (a, e, B) together with a constant gain Kalman filter technique is described in this chapter. This new approach propagates the number density in a bin of ‘a’ and ‘e’ rapidly and accurately without propagating each and every of the space debris objects in the above bin. It is able to assimilate the information from other breakups as well with the passage of time. Further this approach expands the scenario to provide suitable equivalent ballistic coefficient values for the conglomeration of the fragments in the various bins. The heart of the technique is to use a constant Kalman gain filter, which is optimal to track the dynamically evolving fragment scenario and further expand the scenario to provide time varying equivalent ballistic coefficients for the various bins. In the next chapter 6 a new approach for the collision probability assessment utilizing the closed form solution of Wiesel (1989) by the way of a three dimensional look up table, which takes only air drag effect and an exponential model of the atmosphere, is presented. This approach can serve as a reference collision probability assessment tool for LEO debris cloud environment. This approach takes into account the dynamical behavior of the debris objects propagation and the model utilizes a simple propagation for quick assessment of collision probability. This chapter also brings out a comparison of presently available collision probability assessment algorithms based on their complexities, application areas and sample space on which they operate. Further the quantitative assessment of the collision probability estimates between different presently available methods is carried out and the obtained collision probabilities are match qualitatively. The Chapter 7 utilizes once again the efficient and robust constant Kalman gain filter approach that is able to handle the many uncertain, variable, and complex features existing in the scenario to predict the reentry time of the risk objects. The constant gain obtained by using only a simple orbit propagator by considering drag alone is capable of handling the other modeling errors in a real life situation. A detailed validation of the approach was carried out based on a few recently reentered objects and comparison of the results with the predictions of other agencies during IADC reentry campaigns are also presented. The final Chapter 8 provides the conclusions based on the present work carried together with suggestions for future efforts needed in the study of space debris. Also the application of the techniques evolved in the present work to other areas such as atmospheric data assimilation and forecasting have also been suggested.

Thesis (MEng)--Stellenbosch University, 2013. / ENGLISH ABSTRACT: Throughout its lifetime in space, a spacecraft is exposed to risk of collision with orbital debris or operational satellites. This risk is especially high within the Low Earth Orbit (LEO) region where the highest density of space debris is accumulated. This study investigates orbital decay of some LEO micro-satellites and accelerating orbit decay by using a deorbitsail. The Semi-Analytical Liu Theory (SALT) and the Satellite Toolkit was employed to determine the mean elements and expressions for the time rates of change. Test cases of observed decayed satellites (Iridium-85 and Starshine-1) are used to evaluate the predicted theory. Results for the test cases indicated that the theory tted observational data well within acceptable limits. Orbit decay progress of the SUNSAT micro-satellite was analysed using relevant orbital parameters derived from historic Two Line Element (TLE) sets and comparing with decay and lifetime prediction models. The study also explored the deorbit date and time for a 1U CubeSat (ZACUBE-01). A proposed orbital debris solution or technology known as deorbitsail was also investigated to gain insight in sail technology to reduce the orbit life of spacecraft with regards to de- orbiting using aerodynamic drag. The deorbitsail technique signi cantly increases the e ective cross-sectional area of a satellite, subsequently increasing atmospheric drag and accelerating orbit decay. The concept proposed in this work introduces a very useful technique of orbit decay as well as deorbiting of spacecraft. / AFRIKAANSE OPSOMMING: Gedurende sy leeftyd in die ruimte word 'n ruimtetuig blootgestel aan die risiko van 'n botsing met ruimterommel of met funksionele satelliete. Hierdie risiko is veral hoog in die lae-aardbaan gebied waar die hoogste digtheid ruimterommel voorkom. Hierdie studie ondersoek die wentelbaanverval van sommige Lae-aardbaan mikrosatelliete asook die versnelde baanverval wanneer van 'n deorbitaal meganisme gebruik gemaak word. Die Semi-Analitiese Liu Teorie en die Satellite Toolkit sagtewarepakket is gebruik om die gemiddelde baan-elemente en uitdrukkings vir hul tyd-afhanlike tempo van verandering te bepaal. Toetsgevalle van waargenome vervalde satelliete (Iridium-85 en Starshine-1) is gebruik om die verloop van die voorspelde teoretiese verval te evalueer. Resultate vir die toetsgevalle toon dat die teorie binne aanvaarbare perke met die waarnemings ooreenstem. Die verloop van die SUNSAT mikrosatelliet se wentelbaanverval is ook ontleed deur gebruik te maak van historiese Tweelyn Elemente datastelle en dit te vergelyk met voorspelde baan- elemente. Die studie het ook ondersoek ingestel na die voorspelde baan-verbyval van 'n 1-eenheid cubesat (ZACUBE-01). Die impak op wentelbaanverval deur 'n voorgestelde oplossing vir die beperking van ruimterommel, 'n deorbitaalseil, is ook ondersoek. So seil verkort 'n satelliet se ruimte- leeftyd deur sy e ektiewe deursnee-area te vergroot en dan van verhoogde atmosferiese sleur en sonstralingsdruk gebruik te maak om die vervalproses te versnel. Hierdie voorgestelde konsep is 'n moontlike nuttige tegniek vir versnelde baanverval en beheerde deorbitalering van ruimtetuie om ruimterommel te verminder.

Made available in DSpace on 2014-06-11T19:25:31Z (GMT). No. of bitstreams: 0 Previous issue date: 2009-12-22Bitstream added on 2014-06-13T18:53:36Z : No. of bitstreams: 1 merguizosanchez_d_me_rcla.pdf: 2937684 bytes, checksum: 8831a2baec961b9a95e300124d0b3a90 (MD5) / Coordenação de Aperfeiçoamento de Pessoal de Nível Superior (CAPES) / A estabilidade dos membros das constelações Galileo e GPS é investigada. Devido à ressonância 2:1 entre w e W, ocorre um aumento significativo da excentricidade. Este aumento causa riscos de colisão entre os satélites descartáveis e os ativos. Como a ressonância não depende do semi–eixo do satélite, estratégias usuais de aumentar a altitude não resolvem o problema. Então, condições iniciais especiais são achadas tais que os satélites descartáveis permanecem estáveis, com baixa excentricidade por pelo menos 250 anos. Outra estratégia de atacar o problema é mover o objeto descartável para uma órbita particular, acelerando o crescimento da excentricidade. Este estudo é brevemente apresentado. A dinâmica de satélites heliossíncronos é também estudada. Devido o arrasto atmosférico, a altitude do satélite sempre decai e portanto ele cruza o valor ressonante do semi-eixo. Sempre que isso ocorre, um salto na inclinação é observado e em alguns casos, há alguns cruzamentos tais que a inclinação permanece aprisionada (durante algum tempo) no centro de libração. Este evento é importante, pois isso pode ser explorado para realizar manobras de baixo custo para controlar o satélite numa determinada inclinação. Através do sistema exato, investigamos estas quasecapturas e seu aproveitamento em manobras de manutenção de inclinação. / The stability of the disposed members of the Galileo and GPS constellations is investigated. Due to the 2:1 resonance between w and W, a significant increase of the eccentricity occurs. These increase cause risk of collisions between the operational and disposed satellites. As the resonance does not depend on the semi-major axis of the satellite, usual strategies of raising the altitude do not solve the problem. Therefore, special initial conditions are found such that the disposed satellites remain stable with small eccentricity, for at least 250 years. Another strategy to attack the problem is to move the disposed object to a particular orbit, accelerating the growth of the eccentricity. This study is briefly presented. The dynamics of the sun-synchronous satellite is also studied. Due to the atmospheric drag, the altitude of the satellite always decays and therefore it crosses the resonant value of the semi-major axis. Whenever this happens, a jump in the inclination is observed and in some cases, there are some crossing such that the inclination remains locked (during some time) in the center of the libration. This event is interesting since it can be exploited to perform inexpensive maneuvers to control the satellite at desired inclination.

Shape and attitude of near-Earth objects directly affect the orbit propagation via drag and solar radiation pressure. Obtaining information beyond the object states (position and velocity) is integral to identifying an object. It also enables tracing origin and can improve the orbit accuracy. For objects that have a significant distance to the observer, only non-resolved imaging is available, which does not show any details of the object. So-called non-resolved light curve measurements, i.e. photometric measurements over time can be used to determined the shape of space objects using a two step inversion scheme. It follows the procedure to first determine the Extended Gaussian Image and then going through the shape reconstruction process to retrieve the closed shape even while measurement noise is present. Furthermore, it is also possible to generate high confidence candidates when follow-up observations are provided through a multi-hypotheses process.