The combination of a space thermal baffle and radiator builds a passive cooling system for space thermal control. This solution was investigated during the Preliminary Design Review of a satellite optical instrument at the Mechanical Design Office of Airbus Defence and Space, Toulouse. However, the thermal analysis in the baffle sizing process had time-consuming steps that could be automated. This report presents the internship and final Master’s degree project that resulted in the development of a computational tool helping size thermal baffles by automating the generation of the numerical thermal model of the baffle and the radiator sink temperature computation. The tool was designed to provide the necessary inputs and outputs required to carry out an optimization on the baffle geometry. Operational results were obtained thanks to the tool such as the impact of the baffle’s sunshield inclination and of the specularity of its inner coating on the radiator sink temperature. A preliminary work on the baffle geometry optimization was carried out but remaining tasks have to be performed in order to make the tool more robust to input and output changes. Additional work is required to find the optimal baffle geometry in the framework of the instrument project and for prospective projects with similar baffle sizing needs.

In recent years, researchers realized the advantages of using low thrust propulsion system to navigate the spacecraft through the solar system. Designing a trajectory through space using low thrust propulsion system involves solving complex non-linear equations. Also, it often incorporates optimization of certain parameters like time of flight or propellant consumption, making it more complex. There are commercially available software like SEPSPOT and GMAT, which solves trajectory optimization problems. But these software are computationally tedious and requires a deeper understanding of concepts and software. This paper tries to bridge the gap by analyzing different simpler methods available for solving optimization problem. An Analytical method and the Indirect Optimization methods were examined in this investigation with the goal of assessing each approach, thereby determining the quicker yet accurate method. A few approximations like Directional Adaptive Guidance (DAG) and Orbital Averaging techniques were used to reduce the computational load. Results from both the methods were compared with each other and validated with results available from literature. Various missions like Geosynchronous Transfer Orbit (GTO) to Geostationary Earth Orbit (GEO), Low Earth Orbit (LEO) to Medium Earth Orbit (MEO) and pure orbital changes were taken and analyzed. It was concluded that by making proper approximations, one can get solutions quickly without compromising the accuracy of the solution. An error of 4% was observed between the two methods, which can be ignored for getting near estimates. After validating the methods,the analytical method was used to produce a non-dimensionalized contour graph which can be used as an estimate for getting time optimal solution for most of the missions within the sphere of influence of Earth.

Since the early 2000s the space industry has undergone significant changes such as the advent of reusable launch vehicles and an increase of commercial opportunities. This new space age is characterized by a dynamic entrepreneurial climate, lowered barriers to access space and the emergence of new markets. New business models are being developed by many actors and the merging of space and other sectors continues, facilitating innovative and disruptive opportunities. Already established companies are adapting in various ways as efforts to stay relevant are gaining attention. The previous pace of development that was exclusively determined by governmental programs are now largely set by private and commercial ventures. Relating to all trends, new technologies and driving forces in the space industry is no trivial matter. By analyzing and examining identified trends and technologies the author has attempted to discern those that will have a significant impact on the industrial environment during the next decade. Market assessments have been summarized and interviews have been carried out. Discussions and conclusions relating to the services provided by the Swedish Space Corporation are presented. This report is intended to update the reader on the current status of the space industry, introduce concepts and provide relevant commentary on many important trends.

This thesis presents a way to maximise the photographic coverage of Phobos, one of the two Martian moons, as part of the space mission Mars Moon eXplorer. This coverage is performed by the French hyperspectral imager MacrOmega, and two criteria are selected: the area covered and the resolution of the pictures. The approach considered is a greedy algorithm, and elements of basic theory are provided. This greedy approach is compared to a chronological algorithm, whose results were already approved for the mission.

This report details the development of a software model of an ADCS to form part of the *craft mission design tool, a software design tool used in conjunction with the hardware-in-the-loop testbed, NANOBED. Models for the ADCS have been developed, using Clyde Space Ltd systems as benchmarks, and include sub-models for actuators. The objective for these models is to provide data on power consumption and torques produced by the ADCS throughout an operational simulation, which along with models of other systems gives an indication of the feasibility of a satellite’s mission configuration. The control algorithms which define the logic behind the ADCSs have been discussed and implemented, taking into account effort limitations and distribution to the configuration of actuators. Simulations for several ADCS operational modes were performed and the results were as expected, with the ADCS following the requirements for each mode.

This report describes the feasibility of attitude determination and control of the student satellite MIST. It investigates the stability and controllability of the satellite system, it covers attitude determination based on magnetometer and sun sensor measurements available and finally compares two controllers and the resulting pointing accuracy of the satellite. The study shows that the pointing requirements can be met under nominal circumstances.

This document presents the results of the master thesis conducted from April 2019 to October 2019 under the direction of CNES engineer Eric Bourgeois, as part of the KTH Master of Science in Aerospace Engineering curriculum. Within the framework of development studies for the Callisto demonstrator, this master thesis aims at studying and developing possible guidance strategies for the boosted landing. Two main approaches are described in this document :

• Adaptive pseudo-spectral interpolation

• Convex optimization

The satisfying results yielded give strong arguments for choosing the latter as part of the Callisto GNC systems and describe possible implementation strategies as well as complementary analyses that could be conducted.

In this era of miniaturization and industrial needs for accessing space, nanosatellites are an appropriate and riskless answer. This paper examines the development and the progress of the mechanical assembly, integration and test (AIT) campaign on the CNES CubeSat demonstrator, EYESAT. This student  nanosatellite, designed to observe the zodiacal light in the solar system, is actually in integration and testing phase; and will be launched in 2019. This thesis led the project from the structural and thermal model to the end of the qualification model assembly as well as the beginning of the tests campaign.

Eye-Sat is a student-designed 3U-CubeSat, to be launched to a sun-synchronous orbit from where it will map the zodiacal light, a faint glare caused by the reflection of Sun on interplanetary dust. Such mission requires an accurate 3-axis attitude control, for which Eye-Sat is equipped with reactions wheels, magnetorquers, magnetometers and a star tracker. The star tracker can only be used for inertial pointing, which confines its use to shooting phases. A solution based on the remaining 3 equipment is proposed for the other mission phases, providing 3-axis pointing with high agility, for ground station tracking, at the cost of a slightly degraded accuracy. The magnetometers and magnetorquers work in closed-loop, while manoeuvres are performed in open-loop by the reaction wheels, which also ensure gyroscopic stabilisation of the spacecraft. Since this design relies on only one sensor, efforts have been put into making it robust to the imperfections of the magnetometers. Robustness to potential changes in the mission or the design has also been taken into consideration. Performance assessments carried out on a preliminary tuning have demonstrated the capacity of this magnetic-based mode to recover 3-axis pointing when exiting the survival mode, to provide a 3-axis pointing accuracy better than 8 deg  in the worst case, and to sustain slews up to 0.87 deg/s in download.

This master thesis report presents the modelling process of key elements of the Laser Interferometer Space Antenna mission (LISA mission) in a Modelbased systems engineering (MBSE) approach with SysML (Systems Modeling Language). The model implements a selected set of functions of the mission through executable graphical representations, called diagrams. It is shown how such diagrams can benefit the mission, by comparing this mean of information exchange to the traditional text- based systems engineering. The model represents the mission structure and behaviour through a system of nested layers. The deeper the layer is, the more it gives details on a system part. Each layer can be seen from different point of views, either focusing on the structure, the behaviour, or the performance of related system part.

In March 2019 the number of artificial objects bigger than 1 mm in orbit around the Earth is estimated to be more than 170 millions. Only a small fraction of them (0.03%) is catalogued. An impact of an operational satellite with one of these debris can damage the satellite and undermine its mission. So it is important to catalogue as many objects as possible in order to reduce the risk of a collisions. This is done by using the software tool Space Object Observations and Kalman Filtering (SPOOK), developed in Airbus Defence and Space in Friedrichshafen. The goal of this Master Thesis was to create newfunctionalities to this tool and improve the existing ones. In particular three main goals have been accomplished:

• a new model for the lighting ratio has been built to take into account the occultation of the Sun due to a covering body (for example the Earth or the Moon) and itsinfluence on the solar radiation pressure, necessary to have a good model for orbit propagation;

• a tracklet building algorithm has been built to distinguish different tracklets (consecutive observations of the same object along its orbit) as a starting point for the association of different measurements belonging to the same object at distant epochs, necessary to update a catalogue of space objects;

• a model to take into account the process noise has been improved giving some suggestion on how to tune the different parameters for different kinds of orbit.

The aim of this internship is to improve the test process, especially the data exchange and treatment, as well as the production traceability, of an aeronautical Diesel engine. New criteria are devised to enhance the verifications of the engine’s performance.The documentation used throughout the process is globally reworked. New documents are also created, to simplify and codify the process. The updates to forms, instructions and specifications are included. Search for data in the Production Records is simplified by sorting the documents composing them, both in paperand digital formats. This paper also presents the work accomplished with the test subcontractor.

This report deals with new technologies about control time saving and reliability within a manufacturing unit at Airbus Safran Launchers. Firstly, it analyses various kind of checked performances and describes how checking processes work in the different workshops and define what the main problems are. Then, it focuses on checking countersunk holes control and on three-dimensional control. It studies two kinds of new technologies and explains how it can decrease checking time relating to these two issues and how it can increase control reliability. To conclude, it also presents some other technologies which reach the same goal.

Many satellites share the need to transmit data to Earth. However, depending on their orbital parameters the total delivery time may be very long, starting from when a satellite acquires its data to ground station data delivery. In an era where the need for fast data acquisition grows ever larger, there is a need to optimise this issue. In this thesis, it is suggested that a communicating system of satellites that utilise a pathfinding algorithm offers a viable solution to decrease the total delivery time of the satellite data. This is examined by creating simulations of different satellite constellations in MATLAB and in each of these comparing two satellite systems, with only one of them utilising the pathfinding algorithm. The results from the simulations provided a clear indication that the pathfinding algorithm did in fact provide a significantly faster average delivery time and could thus serve as a viable solution to the problem. 

Transportation across the globe's oceans increases every year and is expected to keep increasing in the following decades. Consequently, there is a need to establish communication over the horizon through the Automatic Identification System (AIS) and the Very High Frequency (VHF) Data Exchange System (VDES), still in development, to track and communicate with vessels all over the globe regardless of the distance from shore.

In this Master thesis a feasibility study for the development of a system that fulfils that need is proposed consisting of a Low Earth Orbit (LEO) constellation providing VDES communication continuously all over the globe. A system engineering approach has been followed, identifying stakeholders and producing system requirements setting up a framework for the system. The key stakeholders were found to be the customers/users, the satellite provider, the satellite operator, the service provider and the payload provider. Furthermore, possible use-cases were presented and a system architecture was defined to outline the system, dividing the system into three segments: the space segment, the ground segment and the launch segment.

In addition, design proposals for a satellite constellation and a typical satellite in such a constellation were implemented. The satellite constellation was proposed to consist of 91 satellites at an orbit altitude of around 550 km in polar orbits of common inclination, this was regarding a minimum elevation angle of 10 degrees. The satellite is recommended to consist of a 6 U CubeSat using as payload the existing airborne transponder R5A from Saab TransponderTech, it builds on the Software Defined Radio (SDR) technology and is to be further developed for VDES applications.

Moreover, a link- and a data budget were implemented. Different launch options were addressed concluding that launching as secondary payload on a ride-share mission or as primary payload on a small satellite launch vehicle are the preferable options. A market analysis has been made providing details on how many AIS/VDES satellites that have been launched into LEO and by which service provider, as well as further details on small/nano satellites of extra interest to this work. A short risk evaluation was also done, identifying the most evident risks with developing, operating and disposing the system. In addition, Saab's potential role in the development of satellite VDES is discussed. 

In conclusion to this work it has been shown that it is possible to build a global continuous satellite constellation in LEO utilising as payload an SDR-platform to provide VDES services to vessels at open seas.

With the advent of private commercial suborbital spaceflight, a new demo-graphic of untrained individuals will begin to travel to space. These individuals are exposed to high levels of G-forces, resulting in medical considerations which are not a normal factor with high performance fighter pilots or astronauts.The acceleration profiles of the Virgin Galactic and Blue Origin spacecraft were obtained from publicly available data. Video analysis was performed on footage of spacecraft test launches and human centrifuge tests to obtain individual data sets. These data sets were used to develop the acceleration profiles for both spacecraft. Based on the spacecraft’s acceleration profiles and peak G-forces, medical conditions were investigated and considered to identify potential risks that may affect the passengers, particularly the elderly.

With a big picture view of increasing the accessibility of space, standardization is applied within a service-oriented space program. The development of standardized spacecraft interfaces for numerous and varied payloads is examined through the lens of the creation of an Interface Control Document (ICD) within the Peregrine Lunar Lander project of Astrobotic Technologies, Inc. The procedure is simple, transparent, and adaptable; its applicability to other similar projects is assessed.

The thesis at hand describes the design and analysis process of a deployment mechanism for large area composite structures which are z-folded and subsequently wrapped around a hub. The design process begins with the evaluation of previous attempts, deriving key factors for successful deployment such as maintaining tension in the deploying structure and reliable transmission of the torque. Subsequently new concepts for the key success factors are introduced. As a second step, a prototype of the new deployment mechanism design is manufactured and its characteristics are analyzed with a force gauge and Digital Image Correlation. The occurring malfunctions and their remedies are described. Recommendations for future, upscaled designs conclude the work.​

Optimal orbital trajectories are obtained through the solution of highly nonlinear large scale problems. In the case of low-thrust propulsion applications, the spacecraft benefits from high specific impulses and, hence, greater payload mass. However, these missions require a high count of orbital revolutions and, therefore, display augmented sensitivity to many disturbances. Solutions to such problems can be tackled via a discrete approach, using optimal feedback control laws. Historically, differential dynamic programming (DDP) has shown outstanding results in tackling these problems. A state of the art software that implements a variation of DDP has been developed by Whiffen and it is used by NASA’s DAWN mission [Mystic: Implementation of the Static Dynamic Optimal Control Algorithm for High-Fidelity, Low-Thrust Trajectory Design" , AAS/AIAA Astrodynamics Specialist Conference, (Keystone, Colorado), American Institute of Aeronautics and Astronautics, Aug. 21, 2006]. One of the latest techniques implemented to deal with these discrete constrained optimizations is the Hybrid Differential Dynamic Programming (HDDP) algorithm, introduced by Lantoine and Russell in [A Hybrid Differential Dynamic Programming Algorithm for Constrained Optimal Control Problems. Part 1: Theory", Journal of Optimization Theory and Applications, vol. 154, pp. 382-417, issue 2, Aug. 1, 2012]. This method complements the reliability and efficiency of classic nonlinear programming techniques with the robustness to poor initial guesses and the reduced computational effort of DDP. The key feature of the algorithm is the exploitation of a second order state transition matrix procedure to propagate the needed partials, decoupling the dynamics from the optimization. In doing so, it renders the integration of dynamical equations suitable for parallelization. Together with the possibility to treat constrained problems, this represents the greatest improvement of classic DDP. Nevertheless, the major limitation of this approach is the high computational cost to evaluate the required state transition matrices. Analytical derivatives, when available, have shown a significant reduction in the computational cost and time for HDDP application. This work applies differential algebra to HDDP to cope with this limitation. In particular, differential algebra is introduced to obtain state transition matrices as polynomial maps. These maps come directly from the integration of the dynamics of the system, removing the dedicated algorithmic step and reducing its computational cost. Moreover, by operating on polynomial maps, all the solutions of local optimization problems are treated through differential algebraic techniques. This approach allows us to deal with higher order expansions of the cost, without modifying the algorithm. The leading assumption of this work is that, treating higher than second order expansions, grants larger radii of convergence for the algorithm, improved robustness to initial guesses, hence faster rates of convergence. Examples are presented in this thesis to assess the performance of the newly constructed algorithm and to test the assumptions.​

Many space engineering and orbital mechanics applications seek for the usage of focused mathematical models, capable of providing useful insight onto particular phenomena or exploiting some theoretical and physical tools to reduce the computational costs and/or increase the level of accuracy reached. Orbital resonances are one of the phenomena that needs to be properly modelled, both for exploiting such features in the mission design phase and to predict possible resonant returns of threatening objects closely approaching a specified planet.

This work deals indeed with one of the possible models of orbital resonances, representing such a physical phenomenon in the B-plane reference frame with an analysis on the resonant trajectories performed at the moment of close encounter. Before this, flybys are an important source of uncertainty in the numerical simulations, which then need to be as accurate as possible to be used as benchmark. To this extent, a highly efficient method to account for general relativity effects in the N-body propagation is developed, tested and validated, to be then used as precise benchmark for the resonance analysis and application.

The B-plane resonance model is a strictly patched conics theory which does not account for perturbations. A semi-analytical extension of the current B-plane resonance model is proposed to account for perturbing effects inside the planet’s sphere of influence. Introducing a set of perturbing coefficients brings the model to match the simulation results at the B-plane point where such coefficients are computed, as well as to be a highly reliable approximation in its vicinity, performing a validation with Monte-Carlo simulated data.

An extension of the validation proposed would lead to a complete planetary protection or defence application, whereas in its final part the work will show the flexibility of the model by looking at it from a different perspective. A ballistic resonant flyby design application will be implemented by solving a multi-level optimisation problem, to modify an initial trajectory into a new one on the same Tisserand level.

Without dealing with the specific case of resonances, the B-plane reference frame embeds a smart geometrical framework where to express and design flyby deflections, whose power will be shown in terms of accuracy achieved and computational cost required.

Once completed by detaching from the patched conics approximation, such a model could bring remarkable simplifications in planetary protection applications, reducing the need of propagating a high number of Monte Carlo samples, and would increase the precision of the defence analyses against impacts from near-Earth threatening asteroids. About the application proposed here, internal and/or external integration could eventually lead to an enhanced efficiency of the current mission design strategies and could widen the internal proposed capabilities, providing high precision and almost optimal results with lowered computational costs.

Investigations on the Circular Restricted 3- Body Problem (CR3BP) and the motion about the Lagrangian points are not recent. Several past (upcoming) missions have used (plan to use) its dynamics. The existence of specific periodic orbits and their associated invariant manifolds is one property of the CR3BP that raises the interest. These periodic orbits are interesting for their great observation properties, eclipse avoidance, communication’s continuity with the Earth, etc. However, to reach them, optimized transfer trajectories have to be found. A numerical tool is developed to construct firstly these orbits, before using them as input parameters for the invariant manifolds. Indeed, when a spacecraft is inserted into one of these manifolds, it shall naturally reach the orbit without any additional cost. This numerical computation provides manifolds’s insertion points, which are used in return by a Nonlinear Programming (NLP) tool to eventually find the optimized trajectory. Families of periodic orbits and manifolds, together with optimized transfer trajectories, have been successfully computed, with a focus on the Halo Orbits of the Earth-Moon system. Some members of this family, the Near-Rectilinear Halo Orbits (NRHOs), are of a great interest both for their geometry characteristics (close approach of the secondary body) and stability properties. However, in the Earth-Moon system, the associated manifolds do not have points relatively close from the Earth. The thesis work hence does not ensure that using manifolds as transfer arcs is beneficial, compared to a direct transfer. Besides, the Time-Of-Flight (TOF) is significantly larger. Transfer strategies making use of the CR3BP dynamics still are interesting, radically different from the usual trajectories and offering a larger number of opportunities. They may be less expansive, and could particularly be used for uncrewed space missions.

Multiengine first stages are the new trend in recent rockets. Reusability and an oxygen/methane based engine complete this picture. ArianeGroup wants to develop its own rocket following these principles. This thesis presents the study of the feeding system for a seven Prometheus engine rocket. Several ways of connecting propellant tanks to engines were proposed and analyzed. Two configurations were selected and studied with more detail. One consists of a main feeding line which is then split in seven secondary lines. The other one adds one rank of pipes to reduce the number of feeding valves. Their performances were assessed according to classic space industry drivers. Furthermore, the impact of the two solutions on the efficiency of the tank was evaluated. CAD drawings and simulation models were made and could be a base for future work if one of the systems is chosen. The study shows that a falcon 9 like feeding system is performant in terms of mass and pressure losses but another cost-effective configuration is possible and gives good results.

Humanity has always wanted to explore the world we live in and answer different questions about our universe. After the International Space Station will end its service one possible next step could be a Moon Outpost: a convenient location for research, astronaut training and technological development that would enable long-duration space. This location can be inside one of the presumed lava tubes that should be present under the surface but would first need to be inspected, possibly by machine capable of capturing and relaying a map to a team on Earth.In this report the past and future Moon base missions will be summarized considering feasible outpost scenarios from the space companies or agencies. and their prospected manned budget. Potential mission profiles, objectives, requirements and constrains of the BATonomous Moon cave Explorer (BAT-ME) mission will be discussed and defined. Vehicle and mission concept will be addressed, comparing and presenting possible propulsion or locomotion approaches inside the lava tube.The Inkonova “Batonomous™” system is capable of providing Simultaneous Localization And Mapping (SLAM), relay the created maps, with the possibility to easily integrate the system on any kind of vehicle that would function in a real-life scenario.Although the system is not fully developed, it will be assessed from a technical perspective, and proper changes for a viable system transition for the space-Moon environment will be devised. The transition of the system from the Batonomous™ state to the BAT-ME required state will be presented from the requirement, hardware, software, electrical and operational point of view.The mission will be devised into operational phases, with key goals in mind. Two different vehicles will be presented and designed on a high engineering level. A risk analysis and management system will be made to understand the possible negative outcomes of different parts failure on the mission outcome.

This Master Thesis report presents the mechanical concept and design of a microgravity fluid science experiment that is planned to be launched to the International Space Station in approximately five years. Named AtmoFlow, this experiment aims to study the complex flow phenomena occurring in different planets’ atmosphere by implementing a scaled-down model of a planet-atmosphere system in microgravity. The design introduced here focuses on several of the experiment’s systems: the Turn Table, a rotating system involving a Fluid Cell Assembly, a Thermal Control System and two Rotation Drive Systems. Design methods, main choices and crucial conception steps are presented. Components’ dimensioning and choices are motivated. Models and analyses are derived from the selected design to ensure that the design meets the requirements on microgravity disturbances, structural and rotational characteristics stated during previous project’s phases. The Computer Assisted Design model of the experiment’s final concept and its integration sequence are eventually proposed and discussed.

This master thesis is a study of supervision tools in a context of high confidentiality. Indeed, the status “confidential defense” imposes governmental restrictions with national stakes and places the team project to higher security restrictions than a common internal project. First it introduces the context of the fusion between Airbus Defense and Space, and Safran, and then it gives the context of the ADELINE project. Then I proceed to the details of my activities at Airbus Safran Launchers through two main missions: data processing supervision on an avionic platform and supervision of several subcontractors via management of projects and compare both in a discussion part.

Asteroid mining is expected to become reality in the near future. The first step is to redirect an asteroid to a stable Earth orbit so that mining technologies can be demonstrated. Detumbling

of the asteroid is one of the important steps in asteroid redirection missions. In this thesis, a preliminary lander CubeSat design is suggested for a small asteroid detumbling mission.

The candidate asteroid for the detumbling mission is chosen to be 2014 UR, an Arjunatype asteroid with an estimated diameter ranging from 10.6 to 21.2 m. Due to the small size of the asteroid, the landing must be performed with an active control method after which the spacecraft must be firmly anchored to the asteroid. By using the multi-criteria decision making method PROMETHEE, the microspine gripper is chosen as the most suitable anchoring mechanism. Three main mission drivers are identified during the design process: data-flow between

the lander and the mothership, Delta-V budget and pointing accuracy. The Delta-V required for landing on the asteroid and despinning it is estimated to be 10 m/s and 0.15 m/s at most,

respectively. The uncertainty with the despinning Delta-V is due to varying estimates of the size of the asteroid. The required minimum pointing accuracy is estimated to be 6 degrees. The preliminary lander CubeSat design can be largely realised with commercial off-the-shelf components suggested in this work. Only some of the components have to be custom built or the technologies further developed. It is shown that a CubeSat lander is not able to detumble an asteroid that is rotating fast around multiple axes. However, if the considered asteroid is rotating around a single axis with a rotational period of 2.4 hours, it is be possible to despin it by spending just 1.5 kg of propellant. The suggested lander is a 12U CubeSat with an overall mass of 15 kg and power consumption of 65 W.

We propose using grain-of-sand-scale microscopy (1-micrometre resolution) to study the microstructure and composition of planetary material in a fast, in-situ, solid-state device. We have constructed a small and light prototype instrument (100 grams, 0.3 litres) from commercial-off-the-shelf components, targeted for applications in miniaturised robotic exploration, mounted to a robotic arm, or used as a hand-held tool. This microscope employs a programmable LCD aperture to virtually record multiple perspectives, and a dome studded with LEDs surrounding the sample to control illumination. With this prototype microscope we have captured rich and intuitive raw images for a human observer, and reconstructed 3D surfaces and photometric properties of the samples. The broad applicability of this method is demonstrated by integration into a novel exploration concept in which sensor projectiles are launched from a rover into inaccessible environments. Our microscope can there deliver 3D-maps of the surfaces they encounter and extract relevant morphological properties. Our prototype device is evaluated using a range of lunar and planetary simulants. We argue that this microscope delivers large scientific value on its own, and context for other instruments, with small resource requirements similar to those of a camera alone.

This thesis investigated different configurations of the attitude determination and control system (ADCS) for the MIST satellite, to find a satisfying trade-off between computational demand and estimation/pointing accuracy. A model of the satellite dynamics was developed and used in a simulation. The designed ADCS consists of a discrete extended Kalman filter (EKF) and a model predictive control (MPC) controller tunable in different ways. The filter works with a linearization of the spacecraft dynamics model which is performed about the last attitude estimate and it is also capable of estimating the residual magnetic moment of the spacecraft without any initial guess. Three different models were used with the MPC and compared: a linear-like, state-dependent model, a model linearized around a fixed equilibrium point, and a model linearized around the last attitude estimate. The simulation, developed with Simulink, served as a testbed for the different tunings. From the simulation results, the filter proved to be capable of estimating the residual magnetic moment of the satellite with satisfying accuracy. Estimation and pointing requirements were met on average with a mean absolute estimation error of 0.8 deg and a mean absolute pointing error of 3.5 deg. This performance was achieved in face of measurement and model uncertainty.

This work has as an objective to structure the content of a course on Attitude determination methods, part of an Aerospace Engineering Master program. A selection of books, papers, theses, web sites and films was reviewed to identify the most relevant topics within the areas of Static and Dynamic Attitude Determination and the ways to present them in a educational context. Theory is presented in a simplified way and examples were gathered to illustrate the theoretical part. 

Finally, a discussion is carried out on the main learning goals and challenges, required time for instruction and exercises and suggestion for a grading system.​

This report has been written during my internship/master thesis at Thales Alenia Space, Cannes, FRANCE. The subject of the thesis is ball bearing design, and is focused on the software RBSDyn. This software has been developed by CNES, the French Center for Space Studies, and is used to simulate bearings behaviors under various conditions. My mission was to verify, test and implement this software for the company. In order to do so, the first step was to understand the bearing theory, which is the first part of this report. The second step was to use the software and verify its results, which is presented in the second section. Eventually, the final goal of this internship was to create a sequence to help Thales

Alenia Space engineers to design and select bearings, using this software and an Excel tool that needed to be created. Note that for confidentiality reasons, the values and names used for internal TAS mechanisms have been removed of this document.

The FLARE Suit is a device that aims to protect astronauts from intense solar radiation when travelling out of the magnetosphere on future Deep Space missions. This suit is an emergency solution against solar particles that, due to their high density, can cause acute radiation sickness and lead to excessive destruction of internal organs (gastro-intestinal system, nervous system, blood forming bone marrow) and eventually to death. These symptoms will be a lot stronger out of the Earth’s magnetosphere, towards the Moon or Mars. For now, the FLARE Suit is being studied in the intra-vehicular environment as a supplement to already existing shielding provided by the spacecraft’s structure, but extra vehicular activities in space and on other planets can be considered. It consists of a bladder-suit that is to be filled with water when needed, the water being already present on any human carrying spacecraft. The suit can be deployed within a few minutes, be very lightweight at launch due to the resource utilization of on-board water, and does not use a lot of material compared to a fully shielded module since it is fitted to the individual human body. Furthermore, water has been shown to provide a decent shielding per mass capability, the third most shielding efficient material after hydrogenated boron nanotubes (H-BNNT) and high-density polyethylene (HDPE). Water could eventually be complemented with salt that shows high neutron blocking efficiency and could help shield from neutrons (present in secondary particles). Finally, it has a fully adaptive multilayered and shape changing design which allows for real-time scenario adaptation depending on the intensity of the solar radiation, the number of astronauts involved, the time frame and the water supplies. Preliminary design of the FLARE suit is show cased and radiation simulations are being performed in a 1D environment within the Columbus module, highlighting a fifty percent reduction in radiation exposure with 3.5 to 4 centimeters of water, depending on the salt content. Afterwards, a 3D environment is being built, however not tested. For this, selection of a relevant 3D human model, construction of the Columbus module and design choices on human morphology have been made. Matlab codes also were written to build and shape the 3D suit on top of the human model, following diverse radiation shielding strategies.

With a growing concern about space debris accumulation and the “French Space Operation Act” becoming strictly enforced in 2020, French space operators investigate new deorbiting methods or strategies. Among those, the Assisted Natural Reentry is an innovative deorbiting method developed by CNES to deorbit satellites with a low-thrust propulsive system. With correctly planned maneuvers, it becomes possible to minimize the reentry duration and the human casualty risk. This thesis aimed to further develop the Assisted Natural Reentry software as well as to compute the human casualty risk for several types of satellites. This study showed promising results as Assisted Natural Reentry could reduce the human casualty risk up to tenfold.

This thesis takes a closer look at the complex maneuver known as gravity assist, a popular method of interplanetary travel. The maneuver is used to gain or lose momentum by flying by planets, which induces a speed and direction change. A simulation model is created using the General Mission Analysis Tool (GMAT), which is intended to be easily reproduced and altered to match any desired gravity assist maneuver. The validity of its results is analyzed, comparing them to available data from real missions. Some parameters, including speed and trajectory, are found to be extremely reliable. The model is then used as a tool to investigate the way that different parameters impact this complex environment, and the advantages of performing thrusting burns at different points during the maneuver are explored. According to theory, thrusting at the point of closest approach to the planet is thought to be the most efficient method for changing speed and direction of flight. However, the results from this study show that thrusting before this point can have some major advantages, depending on the desired outcome. The reason behind this is concluded to be the high sensitivity of the gravity assist maneuver to the altitude and location of the point of closest approach.

In this report, have been treating the problem of validation of an orbit determination software. Orbit determination is one of the core element of satellite operations and is essential for any space mission. CNES is replacing its orbit determination tools and developed a generic software providing both routine operation and expert analysis tools. Several configurations for this software have been tested to validate the orbit determination tools. Optimal configurations have been set concerning the determination interval and the integrator and the software has been proven to give satisfying and expected results with another mission using real measurements. Optimal configurations are then to be used for routine operations while the new satellite configuration provide a new environment to compare the software results with its predecessor and serve as a reference configuration for the future studies.

The objective of this project is the development of a mission analysis tool for the nanosatellite company GomSpace Sweden. Although there are many existing software, they can be quite complicated and time consuming to use. The goal of this work is to build a simple app to be used at the earliest stages of space missions in order to obtain key figures of merit quickly and easily. By comparing results, assessing the feasibility of customer needs, analysing how various parameters affect each other, it enables immediate deeper understanding of the implications of the main design decisions that are taken at the very beginning of a mission. The tool shall aid the system engineering process of determining orbit manoeuvre capability specifically for CubeSat electric propulsion systems taking into account the most relevant factors for perturbation in Low Earth Orbit (LEO), i.e. atmospheric drag and Earth’s oblateness effects. The manoeuvres investigated are: orbit raising from an insert orbit to an operating orbit, orbit maintenance, deorbiting within the space debris mitigation guidelines and collision avoidance within the 12 to 24 hours that the system has to react. The manoeuvres cost is assessed in terms of Delta v requirements, propellant mass and transfer times. The tool was developed with MATLAB and packaged as a standalone Linux application.

The conception of constellations of satellites is now at the heart of the space industry programs. This review focuses on the development of a tool for the deployment of constellations composed of electric propulsion satellites, considering the perturbations due to the oblateness of the Earth. This Matlab tool named Constellation, is based on the software Mipelec from CNES (Centre National, d’Etudes Spatiales) and it is used for the first phases of space projects to select the best deployment among all the deployments considered for a given constellation. The low thrust transfers considering the Earth oblateness effects have been studied to develop a first Matlab tool. Then the tool for deployment of constellation has been developed. The different strategies for the RAAN rendezvous, the scenarios of deployments and the method to calculate the different transfers have been defined. The tool has been tested and its optimal configuration has been derived.

My Master Thesis takes place in the context of the MicroCarb mission. The goal of this mission is to identify the sinks and the sources of carbon dioxide on Earth in order to map them and to improve the knowledge of its cycle. To fulfill this mission, some particular guidance modes must be implemented in order to study their feasibility. My thesis consisted in defining and enriching the algorithms used to define the guidance laws, by implementing new tools and a new guidance law, and studying the induced performances in terms of data acquisition and with respect to the constraints related to the satellite. Alongside with this mission, the implementation of those elements support the development of the guidance library POLARIS, actual in its early phase, which is at first only dedicated to MicroCarb but which is intended to become multimissions. First, I describe the CNES as well as the guidance team I worked in. Then, the context of the Master Thesis is introduced. Once the context is established we will focus on the first elements I have been working on, as part of the Dazzling studies. Indeed, the spectrometer used in MicroCarb is very sensitive and has to be maintained at very low temperature. Thus the passive cooling mechanism must be protected from the Sunlight and from the light reflected by the Earth. I had to use a class of the Space mechanics library PATRIUS, called Assembly, in order to materialize the satellite and its numerous parts. Once implemented, I was able to perform some Dazzling Studies, highlighting some issues with the various strategies that were considered, and opening new perspectives. Moreover, a problem was detected on a crucial function of the guidance laws calculator. Once a new function was compiled, I had to made a cross validation using Scilab, and results were positive. This part will end with a Geometric Cape study, realized in order to quantify the influence of the satellite, and the MCV roll, over the Geometric Shifting. In the second part, we will introduce a guidance law which was not implemented initially, and on which I had to work during the last weeks of the thesis: The City mode. Although this mode is similar to an existing calibration mode, it has its own characteristics I had to take into account. The code for this acquisition mode worked well, but the results were not satisfying, considering the Dazzling problem and the kinematic constraints. Thus new strategies had to be considered, and more particularly the 2-scans mode. This mode brought a lot of satisfactions, but there is still more work to be done. This report ends with a general conclusion about my work and some perspectives which could be considered for future studies. I also present my personal contribution and some encountered difficulties I had to deal with.

During recent years, private funding and new technology have boosted a change of the space industry, resulting in smaller and cheaper satellites being available on the market. GomSpace Sweden, part of GomSpace Group, is located in this market and develops propulsion systems for small satellites, so called, CubeSats. One of these highly miniaturized and efficient products is the cold gas propulsion system, suitable for 3-unit CubeSats. Due to the current design, the propellant tank of the system can be filled through a port but only drained through the thrusters, which causes several issues. This thesis concerns the development of a solution to these issues and creating a complete fill and drain system.

A conceptual solution for a fill and drain tool was created and later designed as well as manufactured. The tool was tested and all initial requirements were successfully verified. However, through testing the functionality of the tool, several areas of improvement were found. A second version of the tool was digitally designed to meet these improvements. Also, a proposal of a more general fill and drain solution was digitally designed, which could be realized as a module in future development processes, including a fill and drain tank interface, fill and drain tool and a fill and drain procedure.

In conclusion of the thesis, a solution to the problem with the current design of the propulsion system was successfully developed and tested. An improved version of the solution was designed, which decreases the risk of damage to the propulsion system and simplifies the usage. Finally, a design of a general fill and drain system was created, which could possibly span over several future GomSpace products and lowering both the cost and time of development.