This thesis introduces the analysis of a novel device which, capitalizing on recent advances in gossamer solar sail technology, offers the possibility of propellantless satellite deorbiting and attitude control. By taking advantage of aerodynamic drag effects, a lightweight sail can rapidly deorbit a satellite. At the same time, the sail provides an ideal substrate for a large area magnetic torque coil for attitude control. Through the use of orbit propagation software, the performance of an implementation of this "MagSail" on a Low Earth Orbit (LEO) small satellite is simulated. The analysis is set forth in three parts. First the orbit decay profile of the satellite under the effects of atmospheric drag is presented. The results are interpreted for various initial orbits. Next, the actual torque generation of the MagSail is analyzed. Emphasis is placed on how various design parameters change the magnetic moment of the sail. Finally, a six degree of freedom simulation, combining both orbit propagation and PD attitude control demonstrates a possible implementation of the sail's attitude control capabilities. The work presented in this thesis provides an in-depth look at the deorbiting performance of large-area, low-mass LEO satellites. This research provides a theoretical framework for the development of compact, cost-effective propellantless propulsion and space debris mitigation systems.

<div>With the increasing frequency of satellite launches and proposed constellations for Earth observation and communications, the number of objects in orbit is expected to increase significantly in the coming decades. This trend brings to question the current orbital debris problem, which will continue to worsen if proper deorbit guidelines are not met for future satellites. Currently, the primary system that is utilized for deorbiting is a propulsion system that performs a maneuver at the end of mission and delivers the satellite on an orbit that satisfies deorbit guidelines. However, this method suffers from reliability concerns since it requires that the host spacecraft be operational to perform the maneuver and it also significantly increases the cost and complexity of the satellite if a propulsion system is not already being used for the mission. These issues with complexity, cost, and robustness could potentially be solved by a different class of deorbiting system, a drag sail.</div><div><br></div><div>In this thesis, the effects of perturbations on the orbital and attitude motion of drag sails are analyzed in order to extend the current understanding of drag sails as a deorbiting system. The perturbations considered in the orbit-attitude propagator are aerodynamic drag, solar radiation pressure, Earth oblateness effects, and gravity gradient torques. Different drag sail types, sail materials, deployment epochs, and deployment locations in Earth orbit are also considered in order to extend the analysis to different types of missions and to understand how these parameters influence the effects of the perturbations. Additionally, a self-shadowing algorithm and efficient implementation approach is developed in order to improve the fidelity of the aerodynamic drag and solar radiation pressure perturbations.</div>

Once a satellite has completed its operational period, it must be removed responsibly in order to reduce the risk of impacting other missions. Geostationary Transfer Orbits (GTOs) offer unique challenges when considering disposal of spacecraft, as high eccentricity and orbital energy give rise to unique challenges for spacecraft designers. By leveraging small satellite research and integration techniques, a deployable drag sail module was analyzed that can shorten the expected orbit time of launch vehicle stages in GTO. A tool was developed to efficiently model spacecraft trajectories over long periods of time, which allowed for analysis of an object’s expected lifetime after its operational period had concluded. Material limitations on drag sail sizing and performance were also analyzed in order to conclude whether or not a system with the required orbital performance is feasible. It was determined that the sail materials and configuration is capable of surviving the expected GTO environment, and that a 49 m2 drag sail is capable of sufficiently shortening the amount of time that the space vehicles will remain in space.

Orbital debris poses a serious threat to ongoing operations in space.  Recognising this threat, the European Commission has funded the three-year Technology for Self Removal (TeSeR) project with the goal of developing a standard scalable Post Mission Disposal (PMD) module to remove satellites from orbit following the completion of their mission.  As the project coordinator and key member of the TeSeR Project, Airbus Defence and Space Germany will invest significant resources in achieving this goal over the course of the project. This thesis details the initial analysis of potential PMD module designs conducted by the author during an internship within the AOCS/GNC department of Airbus Defence and Space Friedrichshafen between 1 April 2016 and 31 August 2016.  Three main concepts, drag sails, drag balloons and Electrodynamic Tethers (EDTs), were evaluated during this time with an emphasis on determining the ability of each design to permit passive attitude stabilisation of the satellite during PMD.  Following the required modification of a pre-existing MATLAB/Simulink model, several key findings were made for each device concept.  It was found that no drag sail designs investigated permitted passive aerodynamic attitude stabilisation at orbit heights above 550 km.  When deorbiting from 800 km, however, the lack of the desired and stable attitude was not found to have a significant increase on the deorbit time or the area‑time product. Drag balloon designs were predicted to be comparatively unstable and less mass efficient for deorbiting purposes, with area‑time products up to approximately 50 per cent higher than the equivalent mass drag sail designs.  In spite of this, unstable drag balloons were found to provide shorter deorbit times than stable balloons due to the contribution of the satellite body and solar array to the total frontal area of the satellite.  This indicated that attitude stabilisation is not required for satellites equipped with drag balloon devices. Modelling of bare EDTs suggested that tethers with lengths of 1000 metres or more would not permit passive attitude stabilisation at an orbit height of 800 km.  Simulation of a 500 metre EDT, however, indicated that passive attitude stabilisation can be achieved with EDT devices and proved that EDTs can generate significantly higher drag forces than aerodynamic devices while possessing a significantly lower device mass.  Following the analysis of these results, a recommendation was made for future work to be aimed at improving the EDT model used in this investigation.

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.

Due to the unsustainable space debris environment in Low Earth Orbit, debris objects must be removed to ensure future safe satellite operations. One proposed concept for deorbiting larger space debris objects, such as decommissioned satellites or spent upper rocket stages, is to use a chaser spacecraft connected to the debris object by an elastic tether, but the required technology is immature and there is a lack of flight experience. The inoperable satellite, Envisat, has been chosen as a representative object for controlled re-entry by performing several high thrust burns. The aim of this paper is to develop a control system for the deorbit phase of such a mission. Models of the spacecraft dynamics, the tether, and sensors are developed to create a simulator. Two different tether models are considered: the massless model and the lumped mass model. A switched linear-quadratic-Gaussian (LQG) controller is designed to control the relative position of the debris object, and a switched proportional-integral-derivative (PID) controller is designed for attitude control. Feedforward compensation is used to counteract the couplings between relative position and attitude dynamics. An analysis of the system suggests that the tether should be designed in regard to the control system and it is found that the lumped mass model comes with higher cost than reward compared to the massless tether model in this case. Simulations show that the control system is able to control the system under the influence of modeling errors during a multi-burn deorbit strategy and even though more extensive models are suggested to enable assessment of the feasibility to perform this mission in reality, this study has resulted in extensive knowledge and valuable progress in the technical development. / En ökande mängd rymdskrot har lett till en ohållbar miljö i låga omloppsbanor och föremål måste nu tas bort för att säkerställa framtida satellitverksamhet. En föreslagen metod för att avlägsna större skrotföremål, såsom avvecklade satelliter och använda övre raketsteg, är att koppla en jagande rymdfarkost till föremålet med en elastisk lina. Dock är den teknik som behövs inte mogen och det finns en brist på praktisk erfarenhet. Den obrukbara satelliten Envisat har valts som representativt objekt för kontrollerat återinträde genom flera perigeumsänkande raketmanövrar. Syftet med detta arbete är att utveckla ett system för att kontrollera de två sammankopplade rymdfarkosterna under avlägsningsfasen under ett sådant uppdrag. Modeller för farkosternas dynamik, den sammankopplande linan och sensorer byggs för att utveckla en simulator. Två olika modeller för linan undersöks: den masslösa modellen och den klumpade nodmassmodellen. En omkopplande regulator designas genom minimering av kvadratiska kriterier för att kontrollera skrotföremålets relativa position till den jagande farkosten. Vidare designas en omkopplande proportionerlig-integrerande-deriverande (PID) regulator för att reglera pekningen hos den jagande farkosten. Kompensering genom framkoppling används för att motverka de korskopplingar som förekommer mellan translations- och rotationsdynamiken. En analys av systemet visar att linan bör designas med reglersystemet i åtanke och det framkommer att nackdelarna överväger fördelarna för den klumpade nodmassmodellen jämfört med den masslösa modellen. Simuleringar visar att reglersystemet klarar att kontrollera systemet under ett scenario med flera manövrar och under inverkan av modellfel. Även om mer omfattande modeller föreslås för att möjliggöra en fullständig bedömning av genomförbarheten för detta uppdrag så har denna studie resulterat i en omfattande kunskapsvinst och värdefulla framgångar i det tekniska utvecklingsarbetet.