The objective of this thesis is the development of a software application facilitating a quick and extensive mission analysis including all subsystems of the satellite. Furthermore mission analysis for the Sonate mission is done, utilizing the developed program. Usually mission analysis is done by applying separately different tools which are specific for each subsystem. The drawback of this method is that the system is not analyzed as a whole; rather all subsystems are investigated separately from each other. In case one or more system parameters should change, the effect of this variation onto other parts of the system will not directly be accessible and visible. A simulation tool considering all these subsystems overcomes this weaknesses and helps to develop a balanced system. Different system configurations and operation scenarios can be evaluated quickly and compared to each other in order to find an optimal solution. The simulation system developed during this thesis exhibits a client-server structure, where the subsystems are separated modules acting as clients. The program has a highly modular structure which renders it possible to have an arbitrary extension in the future. Aside from the server, connecting all modules, the following modules were implemented: orbital dynamics module containing orbit propagation and attitude determination power subsystem consisting of the subparts power production, power consumption and power storage archiving module saving all results generated for further evaluation simulation clock providing the simulation time configuration module with a GUI to facilitate the configuration of all simulation parameters simple thermal module providing temperatures for the satellite panels Mission analysis was done for the Sonate satellite utilizing the developed program in order to find a configuration for the solar panels of the satellite such that all the subsystems of the satellite can be supplied with sufficient power during the mission.

Le but de ce travail est de développer un outil d'aide à l'analyse de mission pour un satellite artificiel autour de la Lune. Nous développons tout d'abord une théorie analytique qui décrit suffisamment bien la dynamique du satellite lunaire : nous considérons les quatre perturbations majeures de natures différentes qui l'influencent, ainsi que leurs différents couplages. Les résultats sont obtenus sous forme fermée, sans aucun développement en série de l'excentricité ni de l'inclinaison de l'orbite du satellite : la solution s'applique donc à une large gamme de valeurs. Nous utilisons la méthode des Transformées de Lie pour moyenniser deux fois l'Hamiltonien du problème, dans des variables canoniques, ce qui permet d'intégrer des orbites avec un temps de calcul réduit d'un facteur environ 200 000. Grâce à cela, nous produisons des cartes inédites d'espaces de phase (a,i) qui permettent de sélectionner les paramètres orbitaux selon les besoins de la mission lunaire. De nombreuses vérifications analytiques par rapport à la littérature ont été réalisées, et se sont avérées concluantes; la qualité des deux moyennisations a également été vérifiée. Le logiciel développé est souple et permet un traitement automatisé; les intégrations sont automatiquement vérifiées. Nous avons aussi apporté quelques améliorations significatives au manipulateur algébrique des FUNDP, comme l'ajout de fractions symboliques. Par ailleurs, nous résolvons le problème zonal complet du satellite artificiel, étudions l'effet de C22 sur l'inclinaison critique ainsi que l'effet de la Terre sur les durées de vie orbitales limitées de certains satellites lunaires. The aim of this work is to develop a tool helpful to mission analysis of a lunar artificial satellite. We first develop an analytical theory which sufficiently well describes the dynamics of the lunar satellite : we consider the four main perturbations of various kind which influence it, together with their several coupling. The results are obtained in closed form, without any series expansion in eccentricity nor inclination of the orbit of the satellite : so the solution applies for a wide range of values. We use the Lie Transform method for averaging twice the Hamiltonian of the problem, in canonical variables, which allows to integrate orbits with a CPU time reduced by a factor of about 200 000. Thanks to that, we produce unpublished (a,i) phase space maps from which the orbital parameters can be selected on the basis of the needs of the lunar mission. Many conclusive analytical checks with the literature have been performed, and both averaging processes have been checked. The software developed is flexible and allows an automated treatment; the integrations are automatically checked. We also improved significantly the algebraic manipulator of the FUNDP, like the inclusion of symbolic fractions. Moreover, we solve the complete zonal problem of the artificial satellite, we study the effect of C22 on the critical inclination, and also the effect of the Earth on the limited orbital lifetimes of some lunar satellites.

A conceptual dust detection mission, KnightSat III, using pico-scale satellites is analyzed. The purpose of the proposed KnightSat III mission is to aid in the determination of the size, mass, distribution, and number of dust particles in low earth orbits through a low cost and flexible satellite or a formation of satellites equipped with a new dust detector. The analysis of a single satellite mission with an on-board dust detector is described; though this analysis can easily be extended to a formation of pico-scale satellites. Many design aspects of the mission are discussed, including orbit analysis, power management, attitude determination and control, and mass and power budgets. Two of them are emphasized. The first is a new attitude guidance and control method, and the second is the online optimal power scheduling. It is expected that the measurements obtained from this possible future mission will provide insight into the dynamical processes of inner solar system dust, as well as aid in designing proper micro-meteoroid impact mitigation strategies for future man-made spacecraft.

MATS Satellite

The 2010 National Space Policy of the United State of America introduced by President Obama directed NASA to set far reaching exploration milestones that included a crewed mission to a Near Earth Asteroid by 2025 and a crewed mission to Martian orbit by the mid-2030s. The policy was directly influenced by the recommendations of the 2009 Review of United States Human Space Flight Plans Committee, which called for an evolutionary approach to human space exploration and emphasized the criticality of budgetary, programmatic, and program sustainability. One potential method of improving the sustainability of exploration architectures is the utilization of orbital propellant depots with commercial launch services. In any exploration architecture, upwards of seventy percent of the mass required in orbit is propellant. A propellant depot based architecture allows propellant to be delivered in small increments using existing commercial launch vehicles, but will require three to five times the number of launches as compared to the using the NASA planned 70 to 130 metric ton heavy lift launch system. Past studies have shown that the utilization of propellant depots in exploration architectures have the potential of providing the sustainability that the Review of United States Human Space Flight Plans Committee emphasized. However, there is a lack of comprehensive analysis to determine the feasibility of propellant depots within the framework of human space exploration. The objective of this research is to measure the feasibility of a propellant depot and commercial launch based exploration architecture by stochastic assessment of technical, reliability, and economic risks. A propellant depot thermal model was developed to analyze the effectiveness of various thermal management systems, determine their optimal configuration, quantify the uncertainties in the system models, and stochastically compute the performance feasibility of the propellant depot system. Probabilistic cost analysis captured the uncertainty in the development cost of propellant depots and the fluctuation of commercial launch prices, and, along with the cost of launch failures, provided a metric for determining economic feasibility. Probabilistic reliability assessments using the launch schedule, launch reliability, and architecture requirements of each phase of the mission established launch success feasibility. Finally, an integrated stochastic optimization was performed to determine the feasibility of the exploration architecture. The final product of this research is an evaluation of propellant depots and commercial launch services as a practical method to achieving economic sustainability for human space exploration. A method for architecture feasibility assessment is demonstrated using stochastic system metrics and applied in the evaluation of technical, economic, and reliability feasibility of orbital propellant depots and commercial launch based exploration architectures. The results of the analysis showed the propellant depots based architectures to be technically feasible using current commercial launch vehicles, economically feasible for having a program budget less than $4 billion per year, and have launch reliability approaching the best single launch vehicle, Delta IV, with the use of redundant vehicles. These results serve to provide recommendations on the use of propellant depots in exploration architectures to the Moon, Near Earth Objects, Mars, and beyond.The 2010 National Space Policy of the United State of America introduced by President Obama directed NASA to set far reaching exploration milestones that included a crewed mission to a Near Earth Asteroid by 2025 and a crewed mission to Martian orbit by the mid-2030s. The policy was directly influenced by the recommendations of the 2009 Review of United States Human Space Flight Plans Committee, which called for an evolutionary approach to human space exploration and emphasized the criticality of budgetary, programmatic, and program sustainability. One potential method of improving the sustainability of exploration architectures is the utilization of orbital propellant depots with commercial launch services. In any exploration architecture, upwards of seventy percent of the mass required in orbit is propellant. A propellant depot based architecture allows propellant to be delivered in small increments using existing commercial launch vehicles, but will require three to five times the number of launches as compared to the using the NASA planned 70 to 130 metric ton heavy lift launch system. Past studies have shown that the utilization of propellant depots in exploration architectures have the potential of providing the sustainability that the Review of United States Human Space Flight Plans Committee emphasized. However, there is a lack of comprehensive analysis to determine the feasibility of propellant depots within the framework of human space exploration. The objective of this research is to measure the feasibility of a propellant depot and commercial launch based exploration architecture by stochastic assessment of technical, reliability, and economic risks. A propellant depot thermal model was developed to analyze the effectiveness of various thermal management systems, determine their optimal configuration, quantify the uncertainties in the system models, and stochastically compute the performance feasibility of the propellant depot system. Probabilistic cost analysis captured the uncertainty in the development cost of propellant depots and the fluctuation of commercial launch prices, and, along with the cost of launch failures, provided a metric for determining economic feasibility. Probabilistic reliability assessments using the launch schedule, launch reliability, and architecture requirements of each phase of the mission established launch success feasibility. Finally, an integrated stochastic optimization was performed to determine the feasibility of the exploration architecture. The final product of this research is an evaluation of propellant depots and commercial launch services as a practical method to achieving economic sustainability for human space exploration. A method for architecture feasibility assessment is demonstrated using stochastic system metrics and applied in the evaluation of technical, economic, and reliability feasibility of orbital propellant depots and commercial launch based exploration architectures. The results of the analysis showed the propellant depots based architectures to be technically feasible using current commercial launch vehicles, economically feasible for having a program budget less than $4 billion per year, and have launch reliability approaching the best single launch vehicle, Delta IV, with the use of redundant vehicles. These results serve to provide recommendations on the use of propellant depots in exploration architectures to the Moon, Near Earth Objects, Mars, and beyond.

The master's thesis treat of procedure of a small business entity's developement. It deals with analysis of firm's enviroment and with proposal of vision, mission, goals and strategy, which should help the small business entity be successful during its transition from a starting stage to a groving stage of the life cycle. At the end a few words about implementation are added as well as possible risks which could appear.

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. / Målet med detta examensarbete var att utveckla ett verktyg för missionsanalys för nanosatellitföretaget GomSpace Sweden. Det finns många andra mjukvaror för att nå samma mål men de är ofta komplicerade och tidskrävande. Det specifika målet var således att skapa en enkel applikation som kan användas i de tidiga stegen av utformning av rymduppdrag för att snabbt och enkelt få fram viktiga parametrar. Genom att jämföra resultat, uppskatta genomförbarheten av kundbehov och analysera hur olika parametrar påverkar varandra kan omedelbar förståelse erhållas rörande påverkan av designbeslut som tas i början av rymduppdragen. Verktyget ska stödja systemingenjörsprocessen genom att uppskatta banförflyttningskapacitet för elektriska framdrivningssystem för CubeSats och ta i beaktande de mest relevanta faktorerna gällande störningar i låg jordbana (LEO), i.e. atmosfäriskt motstånd och effekterna av Jordens form. De undersökta manövrarna är: banhöjning från injektionsbana till operationell bana, banunderhåll, bansänkning som följer riktlinjerna för rymdskrot och kollisionsundvikande inom de 12 till 24 timmar som systemet har på sig att reagera. Kostnaden för manövrarna är uppskattade genom DeltaV-krav, massan av bränslet och förflyttningstider. Verktyget utvecklades med MATLAB och paketerades som en fristående applikation i Linux.

Optimizing launcher trajectories is essential for effective mission planning, and specialized software like ASTOS provide an initial, precise overview. However, as launcher development progresses, there is a growing need for the creation of an autonomous flight trajectory software that offers greaterflexibility in adjusting simulation parameters and better represents actual, real-life trajectories. Thisreport introduces an initial version of a comprehensive six-degree-of-freedom launcher trajectory calculation software developed using MATLAB and Simulink. The emphasis is on the development strategy, encompassing discussions on dynamics equations, essential features, and crucial models necessary for accurate simulations. Real-world scenarios often deviate from optimized trajectories, and the software addresses these deviations using sensitivity analysis through Monte Carlosimulations, enabling a thorough examination of uncertainties in input parameters and their impact on trajectories. The article delves into the establishment of the dispersion analysis tool and offers suggestions for further enhancements for both the Simulink model and this dispersion analysis tool. / Optimering av flygbanor är avgörande för effektiv uppdragsplanering, och specialiserad programvara som ASTOS ger en initial, exakt översikt. Men när flygbanans utveckling fortskrider finns det ett växande behov av att skapa en autonom flygbana som erbjuder större flexibilitet när det gäller attjustera simuleringsparametrar och bättre representerar faktiska, verkliga banor. Den här rapporten introducerar en initial version av en omfattande beräkningsprogramvara utvecklad med MATLAB och Simulink för sex frihetsgraders lanseringsbana. Tyngdpunkten ligger på utvecklingsstrategin, som omfattar diskussioner om dynamikekvationer, väsentliga funktioner och avgörande modeller som är nödvändiga för exakta simuleringar. Scenarier i verkligheten avviker ofta från optimerade banor, och programvaran adresserar dessa avvikelser med känslighetsanalys genom Monte Carlo-simuleringar,vilket möjliggör en grundlig undersökning av osäkerheter i inmatningsparametrar och deras påverkan på banor. Rapporten går in i skapandet av spridningsanalysverktyget och erbjuder förslag till ytterligare förbättringar för både Simulink-modellen och detta dispersionsanalysverktyg.

International Telemetering Conference Proceedings / October 27-30, 1997 / Riviera Hotel and Convention Center, Las Vegas, Nevada / Lunar Prospector’s project engineering staff selected a Windows PC platform as the RF test console for the Lunar Prospector spacecraft. The spacecraft test team chose the PCbased RF test console because the PC provides a low-cost, common platform with a graphical user interface. The PC provides point-and-click, menu-driven windows that are common throughout the satellite factory. The PC RF test console is being used to exercise the Lunar Prospector spacecraft RF link for RF commanding, telemetry, and ranging signals during factory test, including thermal vacuum chamber testing. For spacecraft command and control at the factory, the PC-based RF test console is networked to a UNIX workstation over RS-422. The PC RF test console and spacecraft interface are controlled through a coax switch residing in a test rack next to the workstation. The PC RF test console is connected directly to the Lunar Prospector spacecraft using coax cable through the spacecraft Telemetry, Commanding, & Ranging (TC&R) RF antenna hat for both transmit and receive functions. The PC RF test console is also connected hard-line to the spacecraft transponder through the transponder RS-422 connection. This connection provides the ability for spacecraft telemetry to be received at the PC at RF or baseband. The same hard-line spacecraft telemetry data is provided to the UNIX workstation for comparison. NASA’s Lunar Prospector project is the first of the Discovery series of “faster, better, cheaper” missions to be competitively awarded. Lunar Prospector project funding was capped by NASA to ensure that no overruns would occur. The mission was funded to support the scientific community’s desire to verify the presence of ice on the moon and collect environmental data to understand the dynamics that may have led to polar ice deposits. The Lunar Prospector mission received funding in 1996 with a launch planned for September 1997

The Auroral Acceleration Region (AAR) is a key region in understanding the interactionbetween the Magnetosphere and Ionosphere. To understand the physical, spatial, and temporal features of the region, multi-point measurements are required. Distributed small-satellite missions such as constellations of multiple nano satellites (for example multi-unit CubeSats) would enable such type of measurements. The capabilities of such a mission will highly depend on the number of satellites - one reason that makes low-cost platforms like CubeSats a very promising choice. In a previous study, the state-of-the-art of miniaturized payloads for AAR measurements was analyzed and evaluated on the capabilities of different multi-CubeSat configurations equipped with such payloads in addressing different open questions in AAR. This thesis will provide the mission analysis of such a multi-CubeSat mission to the AAR and possible mission design. This includes defining the mission scenario and associated requirements, developing a mathematical description of AAR that allows for specific regions in space to be targeted, an optimisation process for designing orbits targeting these regions, conversion of a satellite formation to appropriate orbits, verifying the scientific performance of this formation and the various costs associated with entering, maintaining, and exiting these orbits.