Dissertations / Theses on the topic 'Mission design and analysi'

In the scope of this master thesis work, the proposed design for NetSat was analyzed for mission and system requirements with respect to structural and thermal limitations. Different load case scenarios for structural and thermal analysis were considered during the process which have been discussed. Based on results, the design is qualified and expected to satisfy all mission and system requirements with regards to structural and thermal limitations.

The thesis is a preliminary mission analysis and design of a small satellite swarm. The concept of the mission is to probe altitudes between 200 km and 6000 km to study the structures and dynamics of the magnetic field aligned currents. The mission lifetime is about 3 months. Aerodynamic drag at low altitudes is used for orbit and formation control. During the perigee passage, the satellite would decelerate due to drag, therefore, reducing its apogee. In addition, the attitude control of the spacecraft during the perigee passage could be used for formation control by changing its cross-sectional area. The simulations indicated that an appropriate insertion orbit should be at the perigee of 168 km and an apogee of 6000km. Moreover, from the orbital decay simulations, it was found that by maintaining a constant ram-facing area of 0.1 m2, it is possible for the satellite to decay in 90 days. The attitude simulations show that for at least one perigee passage at a perigee altitude of 168 km, the satellite is able to maintain its attitude and not tumble throughout the trajectory. In addition, investigation of the leader-follower satellite formation yielded that the relative translation of a circular orbit oscillates in all relative directions whereas in an elliptical orbit it only oscillates in the cross-track direction. Furthermore, the simulation has also shown that the relative translation of a leader-follower formation with a elliptical reference orbit, would spiral out of the radial-cross-track plane.

In this study, the structural design and analysis of a wing having mission-adaptive control surfaces were conducted. The wing structure was designed in order to withstand a maximum aerodynamic loading of 5 g due to maneuver. The structural model of the wing was developed by using MSC/PATRAN package program and that structural model was analyzed by using MSC/NASTRAN package program. The designed wing was then manufactured by Turkish Aerospace Industries Inc. (TUSAS-TAI). The finite element analysis results were verified by conducting ground vibration tests on the manufactured wing. The comparative results were used to tune the finite element model and the results obtained showed that the modeling was very successful.

Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2006.
Includes bibliographical references (p. 177-179).
The trajectory analysis and mission design for inspection of a host spacecraft by a microsatellite is motivated by the current developments in designing and building prototypes of a microsatellite inspector vehicle. Two different, mission scenarios are covered in this thesis - a host spacecraft in orbit about Earth and in deep space. Some of the key factors that affect the design of an inspection mission are presented and discussed. For the Earth orbiting case, the range of available trajectories - natural and forced - is analyzed using the solution to the Clohessy-Wiltshire (CW) differential equations. Utilizing the natural dynamics for inspection minimizes fuel costs, while still providing excellent opportunities to inspect and image the surface of the host spacecraft. The accessible natural motions are compiled to form a toolset, which may be employed in planning an inspection mission. A baseline mission concept for a microsatellite inspector is presented in this thesis. The mission is composed of four primary modes: deployment mode, global inspection mode, point inspection mode, and disposal mode. Some figures of merit that may be used to rate the success of the inspection mission are also presented.
(cont.) A simulation of the baseline mission concept for the Earth orbiting scenario is developed from the trajectory toolset. The hardware simulation is based on the current microinspector hardware developments at the Jet Propulsion Laboratory. Through the figures of merit, the quality of the inspection mission is shown to be excellent, when the natural dynamics are utilized for trajectory design. The baseline inspection mission is also extended to the deep space case.
by Susan C. Kim.
S.M.

Microfinance is a development tool used to reduce poverty among extremely poor households. Impoverished households can access lines of credit through microfinance institutions (MFIs), in order to create a new business, smooth household consumption, fund medical emergencies, etc. Many authors postulate that MFIs are drifting from a welfarist to an institutionalist approach to lending. Using MIXMarket data on specific MFIs in 118 countries between 1995 and 2011, the average loan balance of these organizations will be regressed against measure of outreach and sustainability of these institutions by charter type through a series of four, fixed effects models. The main research question is: given that a positive, overall shift in average loan balance indicates an institutionalist shift in mission, how does this impact microfinance institutions and the demographics they target on the intensive and extensive margins? These analyses will test the theory that MFIs with larger average loan balances serve households closer to the subsistence poverty level, a manifestation of mission drift toward the institutionalist philosophy of lending. The phenomenon of mission drift directly impacts the outcomes of microfinance institutions and the target demographic of the organization. The results of this study indicate that the mission of these organizations is drifting toward the institutionalist philosophy of lending. With this general result, mission drift can be observed within both the internal and external margins of the microfinance industry, which influences the chosen target market, profit generated, and structure of MFIs, as determined by the mission of the organization.

This thesis investigates the effects of camber change on the mission adaptive wing of a structurally designed unmanned aerial vehicle (UAV). The commercial computational fluid dynamics (CFD) software ANSYS/FLUENT is employed for the aerodynamic analyses. Several cambered airfoils are compared in terms of their aerodynamic coefficients and the effects of the camber change formed in specific sections of the wing on the spanwise pressure distribution are investigated. The mission adaptive wing is modeled structurally to observe the effect of spanwise pressure distribution on the wing structure. For the structural design and analysis of the UAV under this study, commercial software MSC/PATRAN and MSC/NASTRAN are used. The structural static and dynamic analyses of the unmanned aerial vehicle are also performed under specified flight conditions. The results of these analyses show that the designed structure is safe within the flight envelope. Having completed aero-structural design and analysis, the designed unmanned aerial vehicle is manufactured by TUSAS Aerospace Industries (TAI).

In this thesis work several numerical and experimental methods for design and test of units and devices onboard BepiColombo Mission are studied, implemented and described. BepiColombo Mission is the result of the joined efforts of European Space Agency and Japanese Space Agency: in 2015 two different orbiters (ESA Mercury Planetary Orbiter, MPO, which will support remote sensing and radio-science instrumentation, and JAXA Mercury Magnetospheric Orbiter, MMO) will be launched in the direction of Mercury to study the surface composition and morphology, the geology and the magnetosphere of the planet closest to the Sun. Italy plays an important role in the mission since it is involved in the design and development of the Spectrometer and Imagers for Mpo Bepicolombo Integrated Observatory SYStem (SIMBIO-SYS): this integrated package of instruments includes an imaging system with stereo (STC) and high spatial resolution (HRIC) capabilities along with a hyperspectral imager (VIHI) in the visible and near infrared range. Due to the proximity to the Sun, MPO will face an extremely harsh environment from a thermal point of view, therefore the orbiter, and in particular instrumentation exposed to the thermal fluxes, shall be equipped with sophisticated thermal control devices, such as baffling systems for heat rejection. Starting from the deep knowledge of the thermal scenario in which units and baffles will operate, thanks to the results obtained from detailed thermal and mathematical models, different innovative test-beds have been conceived and designed in order to simulate the environmental thermal fluxes in laboratory. At first, the Structural Thermal Models of SIMBIO-SYS baffles have been tested, subjecting the devices to the environmental infrared fluxes provided by infrared lamps and cold sources in vacuum conditions and assuring different temperature levels on the thermal interfaces of the units; after the test campaign, the thermal mathematical models of the baffles themselves have been validated thanks to the correlation with the experimental results, providing some useful information on the design of the Flight Models of the baffles. Afterwards an original set-up to test the Qualification Model of the Stavroudis baffle of HRIC unit has been designed: during tests, scheduled in January and February 2013, also solar fluxes will be simulated, thanks to CISAS solar simulator, with the aim to qualify the instrument reproducing in vacuum the maximum and minimum operative and non operative temperatures and the most critical heat fluxes (solar and infrared) in sequence. In parallel to this activity, from the need to calibrate and qualify the units in space-like environment simulating the operative conditions, two thermal vacuum chambers have been designed: calibration will be performed for HRIC and STC-VIHI units separately, with and without baffles. The activity started from the comprehension of the instruments calibration requirements and proceeded with the conceptual design of the units, the detailed thermal, structural and electrical design and concluded with the procurement, the assembling and the test activity, which has been performed in order to verify the initial requirements. Thanks to these activities, a series of methods, procedures and techniques, both numerical and experimental, have been developed and validated, with the aim to provide an original and useful contribution to the design and test of SIMBIO-SYS suite onboard BepiColombo mission
L’anno 2015 vedrà l’inizio della missione BepiColombo, promossa dall’Agenzia Spaziale Europea (ESA) in collaborazione con l’Agenzia Spaziale Giapponese (JAXA): la missione scientifica permetterà di approfondire la conoscenza di Mercurio, il pianeta più interno del Sistema Solare, studiandone la superficie, la composizione interna e il campo magnetico, consentendo inoltre di investigare sulle cause che hanno portato alla nascita dei pianeti e sulla loro evoluzione nel tempo. Il segmento di volo è costituito da 2 satelliti distinti: il Mercury Planet Orbiter (MPO), sotto la diretta responsabilità dell’ESA, che supporta la strumentazione per remote sensing e radioscienza, e il Mercury Magnetospheric Orbiter (MMO), che supporta la strumentazione per lo studio del campo magnetico e che è assegnato al controllo della JAXA. L’Italia riveste un ruolo fondamentale nell’ambito della missione dal momento che l’Agenzia Spaziale Italiana è coinvolta nella progettazione e nello sviluppo della suite SIMBIO-SYS (Spectrometer and Imagers for Mpo Bepicolombo Integrated Observatory SYStem), un pacchetto integrato di strumenti costituito da un sistema per imaging stereo (STC), da un sistema per imaging ad alta risoluzione (HRIC) e da uno spettrometro nel campo delle lunghezze d’onda del visibile e dell’infrarosso (VIHI). A causa della vicinanza del pianeta al Sole, MPO opererà in un ambiente ostile ed estremo dal punto di vista termico, di conseguenza il satellite e la strumentazione saranno dotati di sofisticati sistemi per il controllo termico attivo e passivo (ad esempio sistemi di baffling per la reiezione dei flussi). Partendo dalla comprensione e dalla conoscenza dello scenario termico in cui la strumentazione si troverà ad operare, grazie ai risultati dei modelli matematici previsionali, sono stati ideati e progettati diversi setup sperimentali innovativi al fine di simulare in laboratorio i flussi termici ambientali. Inizialmente è stata condotta una campagna di test sui modelli termo-strutturali (STM) dei baffles di SIMBIO-SYS, sottoponendo i dispositivi al flusso infrarosso planetario, simulato da lampade infrarosse e sorgenti fredde in condizioni di vuoto e assicurando diversi livelli di temperature alle interfacce termiche delle unità. In seguito alla campagna di test, i modelli matematici e termici dei baffles sono stati validati, mediante la procedura di correlazione con i risultati sperimentali; grazie alla validazione, è stato quindi possibile raffinare i modelli termici del modello da volo dei baffles. In secondo luogo è stato ideato e progettato un set-up per testare il Qualification Model del baffle Stavroudis di HRIC: durante i test, in programma per gennaio e febbraio 2013, saranno simulati anche i flussi solari, grazie all’innovativo simulatore solare progettato al CISAS, allo scopo di qualificare lo strumento riproducendo in vuoto le minime e massime temperature operative e non operative e i flussi termici (solare e infrarosso) più critici. All’attività precedentemente descritta è stato affiancato il design di due camere termovuoto che verranno utilizzate in fase di calibrazione e qualifica dei modelli da volo di STC, VIHI e HRIC, con e senza baffles. A partire dall’analisi delle prestazioni degli strumenti e da una serie di requisiti meccanici, termici, elettrici, di vuoto, di cleanliness e contamination, è stato effettuato uno studio di fattibilità, a cui sono seguiti il design preliminare delle camere, una serie di analisi strutturali e termiche di dettaglio (per simulare in camera da vuoto le interfacce meccaniche e termiche degli strumenti), la progettazione elettrica, il procurement dei componenti e l’attività di test sui sistemi progettati, al fine di verificare i requisiti iniziali imposti. Grazie a queste attività, sono stati sviluppati e validati una serie di metodi, procedure e tecniche, sia dal punto di vista numerico che sperimentale, al fine di fornire un contributo utile ed originale alla progettazione e alla verifica della strumentazione della suite SIMBIO-SYS a bordo della missione BepiColombo

Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, February 2006.
Page 238 blank.
Includes bibliographical references (p. 233-238).
During the design process, risk is mentioned often, but, due to the lack of a quantitative parameter that engineers can understand and trade, infrequently impacts major design decisions. The definition of risk includes two elements - probability and impact. As a result of heritage techniques used in the nuclear industry, risk assessment in the aerospace industry is usually purely reliability based, and is calculated as the probability of a failure occurring before the end of the design lifetime. While this definition of risk makes sense if all failures result in the same impact, for many non safety-critical systems, the impact of failures may vary, including variance by when a failure occurs. While current risk assessment techniques answer the question "What is the probability of failure?", the true question that needs to be answered for many missions is "How much return can be expected?" Depending on the question answered, the relative ranking of risk items may be different - leading to different risk mitigation investment decisions. Consequently, to complete an accurate risk assessment, it is important to combine system performance and reliability, and model the probabilistic nature of the expected value of the total system productivity.
(cont.) This expected value is defined as the expected productivity. While the expected productivity is easy to calculate for simple systems, it is more complex if a system has a path-dependant productivity function, as is the case with many aerospace systems. In these systems, the productivity in each state depends on the previous states of the system. An approach, called Expected Productivity Risk Analysis (EPRA), has been developed to model the systems described above in an efficient manner by finding the expected path, and then find the expected productivity given that path. EPRA has been tested against conventional Monte Carlo simulations with excellent results that consistently fall within the 95% confidence interval of the Monte Carlo results, while completing the simulation up to 275 times faster. The EPRA approach has been applied to two case-studies, to demonstrate the importance of using expected productivity in a trade study for a real mission, the Terrestrial Planet Finder Interferometer.
by Julie A. Wertz.
Ph.D.

Recent interest surrounding large scale satellite constellations has increased analysis efforts to create the most efficient designs. Multiple studies have successfully optimized constellation patterns using equations of motion propagation methods and genetic algorithms to arrive at optimal solutions. However, these approaches are computationally expensive for large scale constellations, making them impractical for quick iterative design analysis. Therefore, a minimalist algorithm and efficient computational method could be used to improve solution times. This thesis will provide a tool for single target constellation optimization using spherical trigonometry propagation, and an evolutionary genetic algorithm based on a multi-objective optimization function. Each constellation will be evaluated on a normalized fitness scale to determine optimization. The performance objective functions are based on average coverage time, average revisits, and a minimized number of satellites. To adhere to a wider audience, this design tool was written using traditional Matlab, and does not require any additional toolboxes. To create an efficient design tool, spherical trigonometry propagation will be utilized to evaluate constellations for both coverage time and revisits over a single target. This approach was chosen to avoid solving complex ordinary differential equations for each satellite over a long period of time. By converting the satellite and planetary target into vectors of latitude and longitude in a common celestial sphere (i.e. ECI), the angle can be calculated between each set of vectors in three-dimensional space. A comparison of angle against a maximum view angle, , controlled by the elevation angle of the target and the satellite’s altitude, will determine coverage time and number of revisits during a single orbital period. Traditional constellations are defined by an altitude (a), inclination (I), and Walker Delta Pattern notation: T/P/F. Where T represents the number of satellites, P is the number of orbital planes, and F indirectly defines the number of adjacent planes with satellite offsets. Assuming circular orbits, these five parameters outline any possible constellation design. The optimization algorithm will use these parameters as evolutionary traits to iterate through the solutions space. This process will pass down the best traits from one generation to the next, slowly evolving and converging the population towards an optimal solution. Utilizing tournament style selection, multi-parent recombination, and mutation techniques, each generation of children will improve on the last by evaluating the three performance objectives listed. The evolutionary algorithm will iterate through 100 generations (G) with a population (n) of 100. The results of this study explore optimal constellation designs for seven targets evenly spaced from 0° to 90° latitude on Earth, Mars and Jupiter. Each test case reports the top ten constellations found based on optimal fitness. Scatterplots of the constellation design solution space and the multi-objective fitness function breakdown are provided to showcase convergence of the evolutionary genetic algorithm. The results highlight the ratio between constellation altitude and planetary radius as the most influential aspects for achieving optimal constellations due to the increased field of view ratio achievable on smaller planetary bodies. The multi-objective fitness function however, influences constellation design the most because it is the main optimization driver. All future constellation optimization problems should critically determine the best multi-objective fitness function needed for a specific study or mission.

This study presents a camber morphing concept as an alternative to existing plain flap or aileron type hinged control surfaces used in wings. Structural aspects of the concept are investigated with static nonlinear finite element analyses by using MSC Nastran. In order to assess the aerodynamic characteristics
CFD based 2D solutions are obtained using ANSYS Fluent. The camber morphing concept is applied to the full scale hingeless control surface and implemented in the adaptive camber wing. Hingeless control surfaces and adaptive camber wing are manufactured and changes made in manufacture stages are incorporated into finite element models. Finite element analyses of the wing are conducted with static and dynamic loading and comparison with experimental dynamic analyses are performed.

The ever-growing necessity for faster and cheaper access to space, makes the building of artificial satellites to shift more and more towards small measures. The most exemplar case of small satellites is given by the so called CubeSats, i.e. modular satellites practically ‘built in blocks’ of approximately the same size and weight. Such an approach allows to fasten the design and decrease the overall project complexity, but at the same time has many limitations. The main one is equipping the satellites with a propulsion system for the control of their operative orbit. Such a task normally requires a consistent amount of propellant, and so a specific dedicated propulsion system, that weighs consistently (in terms of both mass and volume) on the budget of these small spacecrafts. The present thesis studies the feasibility of providing small satellites with a propulsion system that would enable them to perform orbit control maneuvers all along the mission duration. The concept is to create a computer tool able to carry out a rapid analysis of the satellite mission, for the determination of the needed Δv, and then a preliminary design of the main components of the required propulsion system. Different propulsion technologies can in this way be considered, being then able to do a trade-off to choose the best solution, in terms of mass and performances. Satellite models ranging from nano to mini-sat standard in LEO-VLEO missions of different durations (2, 5 and 7 years) have been used for feasibility simulations, and re-sults show that the use of some propulsion technology is possible to reach the fixed mission goals.

The main scope of the present thesis centres around two key points: modular block and binding interface of 4 types (mechanical/electrical/data/thermal). Throughout the development process, the preliminary design of the modular block and mechanical interface was continuously modified, that eventually evolved into the micro-satellite platform of different unit sizes for various mission applications. Similar to CubeSat platforms, it opens perspectives for a cheaper and faster satellite deployment by means of modularization and standardization of future satellite systems. The proposed design is fully analysed from different perspectives and later compared to the satellite built in accordance with the traditional approach. This gives the assurance of the validity of the comparison due to the similar mission. As a part of the validation process, the economic analysis was conducted for three different approaches in spacecraft design, which led to the determination of the feasibility of small-mass production and plans for the future satellite constellation.

This report will design a mission and spacecraft to redirect the first Near-Earth Object (NEO) to a stable orbit in the Earth-Moon system. The mission profile includes a soon-as-possible launch, spiral-out escape from the Earth-Moon system, rendezvous, ion beam redirection method, and decommissioning phases, each with accompanying orbital dynamics code written in Matlab. The spacecraft design will include power and mass budgets for each of the subsystems including power, thermal, communications, GNC, fuel, and thrusters. The orbital dynamics code is detailed in the final section of the report.

Thesis (Ph.D)--Aerospace Engineering, Georgia Institute of Technology, 2010.
Committee Chair: Braun, Robert; Committee Member: Clarke, John-Paul; Committee Member: Russell, Ryan; Committee Member: Sims, Jon; Committee Member: Tsiotras, Panagiotis. Part of the SMARTech Electronic Thesis and Dissertation Collection.

Thesis (Ph.D)--Aerospace Engineering, Georgia Institute of Technology, 2010.
Committee Chair: Prasad, JVR; Committee Member: Costello, Mark; Committee Member: Johnson, Eric; Committee Member: Schrage, Daniel; Committee Member: Vela, Patricio. Part of the SMARTech Electronic Thesis and Dissertation Collection.

Thesis (Sc. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2013.
This electronic version was submitted and approved by the author's academic department as part of an electronic thesis pilot project. The certified thesis is available in the Institute Archives and Special Collections.
Cataloged from department-submitted PDF version of thesis.
Includes bibliographical references (p. 116-125).
Planetary hoppers are a new class of vehicle being developed that will provide planetary surface mobility by reusing the landing platform and its actuators to propulsively ascend, translate, and descend to new landing points on the surface of a planetary body. Hoppers enhance regional exploration, with the capability of rapid traverse over hundreds to thousands of meters, traverse over hazardous terrain, and exploration of cliffs and craters. These planetary mobility vehicles are fuel limited and as a result are enabled by carrying sensor payloads that require low mass, low volume, and low onboard computational resources. This thesis describes methods for hoppers to traverse and land safely in this constrained environment. The key questions of this research are: - What types of missions will hoppers perform and how does a hopper traverse as part of these missions? - How does a hopper traverse from its current location to a new landing site safely? This thesis: - describes various hopper mission scenarios and considerations for their mission designs. - creates an operational concept for safe landing for the traverse hop mission scenario. - develops a method that can be used to rapidly and safely detect landing areas at long ranges and low path angles. - develops a method to do fine detection of hazards once at the landing area.
by Babak E. Cohanim.
Sc.D.

Thesis (S.M.)--Massachusetts Institute of Technology, System Design and Management Program, 2008.
Includes bibliographical references (p. 95-96).
The satellite is a highly complex system due to the tight physical constraints, high reliability requirements, and the scale of the product. Except for some commercial missions, most of the satellites are designed from concept to optimally achieve their missions. Historically, the multidisciplinary team spent several months or even a year to finish the concept design. As the information technology revolution occurred in 1990's, Integrated Concurrent Engineering (ICE) was invented to reduce cycle time and reduce resources but with higher quality. It is a new method of real-time team collaboration based on the quantitative computer-based calculations. It was introduced with significant success by JPL/NASA and The Aerospace Corporation. Some organizations followed in using ICE and also confirmed that the design period was reduced from months to weeks. Despite the remarkable successes of the ICE application in the United States and Europe, it is neither used nor well known in other parts of the world. The Japanese organizations, for instance, provide complex products and show their presence world wide, but there is no report of an organization utilizing the ICE approach. They applied the concurrent engineering in manufacturing long ago. It is unclear what brought this situation. The ICE approach has been well examined from the systems engineering perspective but not from the cultural aspect. This thesis analyzes the ICE approach to identify the key factors for successful implementation and operation from both systems engineering and cultural perspectives through the case studies of a implementation failure in a Japanese organization and some successes in Euro-American organizations. Then, the author proposes several ways for successful implementation in the Japanese organization and proposes how the ICE should be approached and be utilized to leverage the design capability of the organization.
by Akira Ogawa.
S.M.

Thesis (M.Eng.)--Massachusetts Institute of Technology, Dept. of Electrical Engineering and Computer Science, 2001.
Includes bibliographical references (leaves 104-105).
by Brian A. Purville.
M.Eng.

Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, May, 2020
Cataloged from the official PDF of thesis.
Includes bibliographical references (pages 177-183).
Convex optimization is used to solve the simultaneous vehicle and mission design problem. The objective of this work is to develop convex optimization architectures that allow both the vehicle and mission to be designed together. They allow the problem to be solved very quickly while maintaining similar fidelity to comparable methods. Multiple architectures are formulated, and the architectures are implemented and evaluated for a sounding rocket design problem and a hydrogen aircraft design problem. The methodology proves successful in designing the sounding rocket while taking into account the optimal trajectory and control strategy and extended to a multi-mission design case. The hydrogen aircraft was successfully designed, allowing for both the cryogenic tank design to be chosen in conjunction with the mission prole. For the rocket design problem, the integrated vehicle and mission problem can only be combined into alternating and integrated approach, and the integrated architecture for convergence to solution in 50% computation time while reaching similar solution. For the hydrogen aircraft case, a 50+% decrease in fuel burn was able to be achieved compared to regular kerosene with an integrated optimization approach. Future work includes studying the convergence properties as well as increasing the robustness of the architectures.
by Beldon Chi Lin.
S.M.
S.M. Massachusetts Institute of Technology, Department of Aeronautics and Astronautics

We are in an era of massive spending cuts in educational institutions, aerospace companies and governmental entities. Educational institutions are pursuing more training for less money, aerospace companies are reducing the cost of gaining ight heritage and the government is cutting budgets and their response times. Organizations are accomplishing this improved efficiency by moving away from large-scale satellite projects and developing pico and nanosatellites following the CubeSat specifications. One of the major challenges of developing satellites to the standard CubeSat mission requirements is meeting the exceedingly tight power, data and communication constraints. A MATLAB toolbox was created to assist the CubeSat community with understanding these restrictions, optimizing their systems, increasing mission success and decreasing the time building to these initial requirements. The Toolbox incorporated the lessons learned from the past nine years of CubeSats' successes and Analytical Graphics, Inc. (AGI)'s Satellite Tool Kit (STK). The CubeSat Mission Planning Toolbox (CMPT) provides graphical representations of the important requirements a systems engineer needs to plan their mission. This includes requirements for data storage, ground station facilities, orbital parameters, and power. CMPT also allows for a comparison of broadcast (BC) downlinking to Ground Station Initiated (GSI) downlinking for payload data using federated ground station networks. Ultimately, this tool saves time and money for the CubeSat systems engineer

Il Breeding Blanket (BB) del reattore DEMO rappresenta un sistema complesso in un ambiente pericoloso. Infatti, esso deve soddisfare diversi requisiti e vincoli ingegneristici sia di tipo nucleare, termo-strutturale che di sicurezza. Per questi motivi, è necessaria una progettazione omnicomprensiva che preveda l'applicazione di strumenti avanzati di simulazione basati su approcci multi-fisici. Questi strumenti devono eseguire simultaneamente diversi tipi di analisi. Tre di esse possono essere considerate prioritarie e propedeutiche per lo studio di tutti gli altri fenomeni riguardanti il BB, vale a dire l´analisi nucleare, termo-fluidodinamica e strutturale. In questa tesi, è proposto un innovativo approccio multi-fisico che copre i tre pilastri principali su cui è basato il progetto del BB (la neutronica, la termoidraulica e la termo-meccanica). Queste analisi devono essere condotte in maniera integrata, consentendo una valutazione olistica dei carichi volumetrici di potenza, delle prestazioni termiche sia del fluido di raffreddamento che delle strutture, nonché dei campi di tensione e deformazione. La strategia seguita per il conseguimento di questa sfida consiste nella creazione di una procedura “CAD-centric” e “loosely-coupled” (debolmente accoppiata) per la progettazione dei concetti di BB utilizzando una tecnica di analisi basata su sotto-modelli. Questa procedura prende il nome di Multi-physics Approach for Integrated Analysis (MAIA). Essa basa la sua architettura sull'uso di codici validati e sulla minimizzazione del loro numero. In particolare, MAIA è articolata in 10 fasi principali che vanno dalla creazione di un modello per le analisi nucleari generato dalla decomposizione in geometrie semplici di un generico CAD alla valutazione della potenza volumetrica, dal calcolo dei campi di temperatura e velocità nella struttura e nel refrigerante alla valutazione dei campi di spostamento, deformazione e stress, dalla stima dei tassi di produzione degli isotopi dell´azoto prodotti dall'attivazione dell'ossigeno presente nell'acqua al calcolo della loro distribuzione spaziale di concentrazione tenendo conto degli effetti del trasporto convettivo. Tutti i vari passaggi condividono gli stessi dettagli geometrici. In particolare, MAIA differisce dagli approcci convenzionali usati nell´accoppiamento multi-fisico su tre aspetti chiave. Innanzitutto, non introduce omogeneizzazioni dei modelli e dei carichi. In secondo luogo, MAIA permette di determinare, con un’alta risoluzione spaziale, i gradienti dei carichi per tutte le analisi coinvolte senza richiedere sforzi computazionali proibitivi. In terzo luogo, MAIA permette di mantenere la coerenza tra le tre analisi garantendo la congruenza tra gli input e gli output. Tuttavia, l´onere computazionale richiesto dall´approccio CAD-centric, su cui si basa la procedura MAIA, non permette di rappresentare il BB nel suo complesso ma solo alcune sue porzioni (una slice, per esempio). Ciò impone la definizione e, conseguentemente, la validazione di opportune condizioni al contorno per ogni sotto-modello utilizzato e per ogni analisi eseguita. A tal proposito, per quanto riguarda le analisi nucleari, le condizioni al contorno utilizzate nel modello locale della slice sono: definizione di una sorgente locale neutronica/fotonica per tener in conto l´effetto albedo dell´intero reattore, rappresentazione del Vacuum Vessel (VV) per simulare il back scattering verso il BB, e l´applicazione di condizioni di riflessione (“mirror”, specchio/simmetria, nella direzione poloidale e “white”, riflessione isotropica, in quella toroidale) per simulare la presenza delle slice adiacenti a quella analizzata. I risultati ottenuti mostrano una variazione della potenza depositata del -0.48 % tra il modello di riferimento DEMO e quello locale (slice). Inoltre, è stata eseguita un'analisi di sensibilità sulla distribuzione angolare della sorgente neutronica/fotonica locale determinando una discretizzazione ottimale in 10 suddivisioni poloidali. Questa suddivisione rappresenta un buon compromesso sia in termini di fedeltà dei risultati ottenuti, rispetto a quelli del modello di riferimento (DEMO), che di minimizzazione dell´onere computazionale. Per quanto riguarda l'analisi delle condizioni al contorno termo-idrauliche usate nel modello locale della slice, è stata applicata una condizione di simmetria termica poloidale. Assumendo una variazione delle portate comprese tra ~ -1.3% e ~ 0.6% e una fluttuazione della densità di potenza fino a ~ 6% tra slice vicine, è stata ottenuta una variazione della distribuzione delle temperature del ± 2.4% dimostrando, quindi, l'applicabilità di tali condizioni. Per quanto riguarda le analisi termo-meccaniche, le condizioni al contorno identificate per il modello locale della slice sono: simmetria sul piano inferiore della slice, Generalised Plane Strain su quello superiore e spostamenti radiali e toroidali impediti ai nodi che giacciono nella parte posteriore della back supporting structure lungo la direzione toroidale e poloidale. Queste condizioni, applicate al sotto-modello, producono una variazione compresa tra il -6% e il 4% tra gli spostamenti calcolati nella slice e quelli nel modello di riferimento DEMO, nonché una stima conservativa delle tensioni primarie e secondarie sia di membrana che di flessione. Inoltre, è stato anche studiato l'impatto della variazione (± 2.4%) di temperatura dimostrando che le fluttuazioni sulle deformazioni totale sono comprese tra il -0.3% e l’1.7%, fino a un massimo del 15% sulle tensioni equivalenti di membrana e tra il -7% e il 5% su quelle di flessione. Infine, la procedura MAIA è stata utilizzata per valutare l'impatto sul design del BB. La sua applicazione ha dimostrato la presenza di alcune criticità nel progetto. In particolare, i risultati fluidodinamici mostrano una violazione dei limiti di temperatura che non sono stati risolti introducendo soluzioni progettuali adeguate. Inoltre, queste violazioni producono, a loro volta, valori molto intensi delle tensioni equivalenti di Von Mises che potrebbero indicare un pericolo per l'integrità strutturale del BB. L´applicazione di MAIA al design del BB a permesso di dimostrare il valore aggiunto di questa procedura la quale potrebbe diventare uno strumento fondamentale e di riferimento per la progettazione del BB. Inoltre, la procedura MAIA ha permesso di mappare localmente variabili importanti come flussi neutronici e temperature, nonché le tensioni primarie e secondarie che sono utilizzate per la determinazione delle tensioni ammissibili applicate per la verifica dei criteri di progettazione. Al fine di dimostrare ulteriormente la versatilità e l'adattabilità della procedura MAIA, è stato studiato il problema di attivazione dell'acqua del sistema di trasferimento di calore primario (Primary Heat Transfer System, PHTS). Utilizzando la procedura MAIA, è stato possibile prendere in considerazione gli effetti dell´efflusso sulla concentrazione degli isotopi dell´azoto e fornire informazioni utili per lo sviluppo sia del design del BB che del suo PHTS.
The Breeding Blanket (BB) of the DEMO reactor represents a harsh system in a dangerous environment. It has to satisfy engineering requirements and constraints that are of nuclear, thermo-structural, material and safety kind. For these reasons, the application of advanced simulation tools, based on a multi-physics approach, is required for its comprehensive design. These tools have to simultaneously perform different kind of analyses among which three, and namely nuclear, thermofluid-dynamic and thermo-mechanical, can be prioritized and considered as propaedeutic for the investigation of all the other issues related to the BB. In this dissertation, a multi-physic approach, covering the three pillars of the BB design (the neutronics, thermal-hydraulics and thermo-mechanics), is proposed. These analyses have to be conducted in a strongly integrated way, allowing a holistic assessment of volumetric heat loads, thermal performances of coolant and structures as well as their stress and deformation states. The strategy, followed for the achievement of this challenge, consists of creating a CAD-centric and loosely-coupled procedure for the BB concepts design adopting a sub-modelling technique, named Multi-physics Approach for Integrated Analysis (MAIA). The MAIA procedure bases its architecture on the use of validated codes and on the minimisation of their number. It is articulated in 10 main steps that go from the decomposition of generic CAD in a format suitable for neutron/photon transport analysis to the nuclear analysis for the assessment of volumetric heating, from the assessment of temperature and velocity fields within coolant and structure to the evaluation of their displacement, deformation and stress fields, from the evaluation of nitrogen isotopes production rates from water oxygen activation to the calculation of their concentration spatial distribution taking into account the effects of passive convective transport. All the steps share the same geometry details and the consistency between input and output parameters. The new MAIA procedure differs from the conventional coupling approach with respect to three key aspects. First, it does not introduce homogenisations of models and loads. Second, MAIA can capture load gradients at high resolution in the three directions for all the analysis involved without requiring prohibitive computational efforts. And third, MAIA keeps the consistency between the three analyses maintaining the congruence between inputs and outputs. However, the computational effort required by the CAD-centric feature of MAIA procedure imposes the representation of BB portions and, therefore, the definition and validation of boundary conditions for each performed calculation. Regarding the nuclear analysis, it has been found that the set of reflecting and white conditions in the poloidal and toroidal directions, respectively, together with the presence of Vacuum Vessel (VV) and the definition of local neutron and photon source, produces a mismatch of -0.48 % in terms of power deposition between the DEMO and the local (e.g. slice) models. It has been demonstrated that the neutronic symmetry conditions are valid in the entire module up to the last slices nearby the caps. Furthermore, a sensitivity analysis on the angular distribution of local neutron and photon source has been performed indicating in 10 cosine bins the optimal discretisation choice in terms of compromise between the fidelity of the results obtained respect to those of the reference model and the relevant computational effort. Concerning the analysis of thermal-hydraulic boundary conditions, it has been found that the variation on mass flow rates (comprised between the ~-1.3 % and the ~0.6 %) as well as power density fluctuation (up to the ~6 % in the neighbouring domains) affect the temperature distribution for less than ±2.4 % demonstrating the applicability of poloidal symmetry conditions. As far as the thermo-mechanical analyses are concerned, it has been identified the set of boundary conditions (radial and toroidal displacements prevented to the nodes lying in the rear of the back supporting structure along the toroidal and poloidal direction, symmetry at the lower cut surface and Generalised Plane Strain to the top one) that produce a discrepancy in terms of displacement in the sub-model comprised between the -6 % and the 4 % as well as a conservative assessment of membrane and bending stresses both for primary and secondary stresses. The impact of the temperature variation has also been investigated showing that the fluctuations on total deformation are comprised between -0.3 % and the 1.7 %, on equivalent membrane stress up to 15 % while on equivalent bending stress between the -7 % and the 5 %. As a proof-of-concept, the MAIA procedure has been then used to evaluate the impact on the BB design, demonstrating that some criticalities are present in the design. In particular, the fluid-dynamic results show a violation of the temperature requirement limits that have not been solved introducing proper design solutions. Furthermore, these violations of thermal-hydraulic requirements produce very intense values of Von Mises equivalent stresses that could jeopardize the structural integrity of the segment box. This demonstrates that MAIA procedure can become the reference tool for the design of the BB. Moreover, the MAIA procedure has proven the possibility to locally map important variables such as the neutron flux and the temperature as well as the primary and secondary stress that are used for the determination of the allowable stress and applied for compering with design criteria. In order to further demonstrate the versatility and adaptability of the MAIA procedure, the water activation issue occurring within the blanket Primary Heat Transfer System (PHTS) has been studied. Using MAIA procedure, it has been possible to take into account the effects of the flow on the nitrogen concentration and to provide useful information for the development of both BB design and its PHTS.

In designing the trajectory for a multiple flyby mission to asteroids the choice of the targets is the most challenging problem. This dissertation faces this problem in the framework of the recently issued medium-size mission call (M5) from ESA: CASTAway. Starting from the preliminary work done in [6], this thesis develops a methodology for sequencing the potential targets in a multiple flyby mission. In order to reduce the computational time, the complete database of known small bodies is firstly pruned on the base of heuristic considerations. Using the assumption of small manoeuvres, a chief orbit concept could be used. Thus, two heuristic thresholds are defined in order to exclude non-promising targets given a chief orbit. The sequencing process takes chief orbit and promising targets as inputs and gives a set of candidate sequences. The results of such a process are analysed in the CASTAway framework and the best feasible sequence studied in details.

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.
Gruvdrift på asteroider förväntas att bli verklighet inom en snar framtid. Det första steget är att omdirigera en asteroid till en stabil omloppsbana runt jorden så att gruvteknik kan demonstreras. Bromsning av asteroidens tumlande är en av de viktigaste stegen i ett rymduppdrag där en asteroid ska omdirigeras. I detta examensarbete föreslås en preliminär asteroidlandare baserad på CubeSat-teknik för ett rymduppdrag där en asteroid ska omdirigeras. En asteroid av Arjuna-typ, 2014 UR, med en diameter på mellan 10.6 och 21.2 m är vald som kandidat för rymduppdraget. På grund av att asteroidens är relativt liten till storlek måste landningen utföras med en aktiv reglermetod och rymdfarkosten måste förankras till asteroiden. Med hjälp av en beslutsmetod utifrån flera mål, PROMETHEE, identifierades förankringsmetoden “mikro-ryggrads-gripare” som den mest lämpliga. Tre huvuduppgifter för rymduppdraget identifierades under designprocessen: dataflöde mellan landaren och moderfarkosten, Delta-V-budgeten och peknoggrannheten. Delta-V som krävs för landning på asteroiden uppskattas att vara högst 10 m/s. Bromsningen av tumlandet kostar högst 15 m/s. Osäkerheten med Delta-V för bromsning av tumlandet beror på olika uppskattningar av asteroidens storlek. Den nödvändiga minsta peknoggrannheten uppskattades vara 6 grader. Utformningen av landaren, baserad på CubeSat-teknik, använder till största delen komponenter som finns på hyllan, s.k. commercial-off-the-shelf. Det visas att en CubeSat-landare inte kan bromsa tumlandet för en asteroid som roterar snabbt kring flera axlar. Om den valda asteroiden roterar runt en axel med en rotationsperiod på 2.4 timmar, är det möjligt att bromsa tumlandet med endast 1.5 kg drivmedel. Den föreslagna landaren är en 12U CubeSat med en total massa på 15 kg och strömförbrukning på 65 W.​

This thesis introduces a novel framework for life cycle mission modelling during conceptual aeronautical design. The framework supports object-oriented mission definition using Geographical Information System technology. Design concepts are defined generically, enabling simulation of most aeronautical vessels and many non-aeronautical vehicles. Moreover, the framework enables modelling of entire vessel fleets, business competitors and dynamic opera-tional changes throughout a vessel life cycle. Vessels consist of components deteriorating over time. Vessels carry payload that operates within the vessel environment. An agent-based simulation model implements most framework features. It is the first use of an agent-based simulation utilising a Geographical Information System during conceptual aero-nautical design. Two case studies for unmanned aircraft design apply the simulation. The first case study explores how the simulation supports conceptual design phase decisions. It simulates four different unmanned aircraft concepts in a search-and-rescue scenario including lifeboats. The goal is to learn which design best improves life cycle search performance. It is shown how operational and geographical impacts influence design decision making by generating novel performance information. The second case study studies the simulation optimisation capability: an existing aircraft design is modified manually based on simulation outputs. First, increasing the fuel tank capacity has a negative effect on life cycle performance due to mission constraints. Therefore, mission definition becomes an optimisation parameter. Changing mission flight speeds during specific segments leads to an overall improved design.

A methodology is developed to address the analyses of combat aircraft attrition. The operations of an aircraft carrier task force are modeled using the systems dynamics simulation language DYNAMO. The three mission-roles include: surface attack, lighter escort, and carrier defense. The level of analysis is performed over the entire campaign, going beyond the traditional single·sortie analysis level. These analyses are performed by determining several measures of effectiveness (MOEs) for whatever constraints are applied to the model. The derived MOEs include: Campaign Survivability (CS), Fractlon of Force Lost (FFL), Exchange Ratio (ER), Relative Exchange Ratio (RER), Possible Crew Loss (PCL), and Replacement Cost (RC). RER is felt to be the most useful MOE since it considers the initial inventory levels of both friendly and enemy forces, and its magnitude is easy for the analyst to relate to (an RER greater than one is a prediction of a friendly force’s victory). The simulation model developed in this research is run for several experiments. The effects of force size on the MOEs ls studied, as well as a hypothetical multimission aircraft deployed to perform any of the three missions (albeit at lower effectiveness than the speciallzed aircraft for their given roles but nonetheless with a higher availability). Evaluation of specific technological improvements such as smaller radar cross section, higher thrust/weight, improved weapons ranges, is made using the MOEs. Also, a cost-effectiveness tradeoff methodology is developed by determining the acquisition cost ratio (ACR) for certain modified alternatives the baseline by determining the required initial inventory of modified aircraft to produce the same total effectiveness of the baseline aircraft.
Ph. D.

The purpose of this study was to determine whether the length and readability of a mission statement contribute to stakeholder behavior regarding the mission statement. The majority of studies in the mission statement literature have not attempted to find an empirical link between mission statement design and employee behavior. This study employed a 2 (length: long v. short) x 2 (readability: low v high) post-test only factorial design to test the relationship between message design and beliefs about the mission statement. Students at a large southeastern university (n=212) were shown the one of four treatments and asked to report their reactions on a brief questionnaire. Results indicated a significant link between readability and beliefs about the functionality of the mission statement. Using the theory of planned behavior, the effects of readability on beliefs about the mission statement were shown to be linked to behavioral intention. The results of this study partially support the relationship between message characteristics of mission statements and the behavioral intention of employees, as well as supporting the TPB model.

This thesis was carried out inside the ESA's ESEO mission and focus in the design of one of the secondary payloads carried on board the spacecraft: a GNSS receiver for orbit determination. The purpose of this project is to test the technology of the orbit determination in real time applications by using commercial components. The architecture of the receiver includes a custom part, the navigation computer, and a commercial part, the front-end, from Novatel, with COCOM limitation removed, and a GNSS antenna. This choice is motivated by the goal of demonstrating the correct operations in orbit, enabling a widespread use of this technology while lowering the cost and time of the device’s assembly. The commercial front-end performs GNSS signal acquisition, tracking and data demodulation and provides raw GNSS data to the custom computer. This computer processes this raw observables, that will be both transferred to the On-Board Computer and then transmitted to Earth and provided as input to the recursive estimation filter on-board, in order to obtain an accurate positioning of the spacecraft, using the dynamic model. The main purpose of this thesis, is the detailed design and development of the mentioned GNSS receiver up to the ESEO project Critical Design Review, including requirements definition, hardware design and breadboard preliminary test phase design.

The i-INSPIRE (initial-INtegrated SPectrograph, Imager and Radiation Explorer) satellite project involves the design and test of a tubular shape pico-satellite. This is a collaborative project of School of Aerospace, Mechanical and Mechatronic Engineering and the School of Physics within the University of Sydney. Expected for launch in 2013, i-INSPIRE will carry a photonics-based spectrograph, an imaging camera, and a miniaturized radiation detector. As the first pico-satellite to be fully constructed by an Australian university and to be launched into space, i-INSPIRE project will be a conceptual prototype for future more complex small satellite mission like the upcoming QB50. This thesis presents the i-INSPIRE pico-satellite subsystems design and final test by means of high altitude balloon launch. The current work and research is based on the concept of the TubeSat. The proposed satellite provides a robust bus to fulfil the scientific tasks. One onboard computer based on MSP430 microcontroller is specifically design for the mission. The AFSK (Audio Frequency-Shift Keying) based UHF (Ultra-High Frequency) communication implementation, cooperating with ground station, will accumulate experience and pave the way for future project. The main scientific aim of i-INSPIRE is the operation of a novel photonic spectrograph in space. Moreover, the unique radiation environment of space also enables us to probe the radiation environment in Low Earth orbit by one miniaturized Geiger counter. An imaging camera is added to undertake earth observation. During the test phase, subsystems and developed housekeeping software were first tested under laboratory conditions. The i-INSPIRE satellite was then launched with a high altitude balloon to verify its practical performance under simulated harsh space environment like low temperature.

The following document describes an example mission, which originated from a real life concept of an imaging satellite in a Sun Synchronous Orbit (SSO)around Earth. This report takes the reader through the thermal analysis and evaluation of space equipment performed with Airbus Space & Defence’s Systema/Thermica tool, at Space Structures GmbH. It details the full process of designing a thermal control system for a space project. The project started from a CAD file which was converted into a Geometric Mathematical Model (GMM) inside Thermica. This process requires an extensive knowledge of not only the software, but also the technical background behind what happens to a satellite in such an extreme environment. This thesis addresses this by showing a step-by-step approach of a full thermal evaluation, starting with the required theoretical background of the thermal environment and the different passive and active thermal design techniques. The next step involves gathering the required input information for the software; such as defining the conductance values between the components and calculating the per node power dissipation for each component considering each operational mode. The final step includes the designing, simulation, iteration and presentation of the temperature results across the spacecraft thermal model. The results of the initial simulation showed that some sensitive components were not within the specified temperature requirements, and therefore both radiators and heaters were sized and introduced to the model. After the third iteration of thermal control, the sensitive components’ temperatures were observed to be within the allowable margins of an ECSS Phase A study. This thesis can serve as a guide and complete document for future missions which plan the design of a Thermal Control System of a satellite in orbit around Earth.

Thesis (S.M.)--Massachusetts Institute of Technology, System Design & Management Program, 1999.
Includes bibliographical references (leaves 83-85).
This thesis investigates multiple autonomous vehicle mission planning and management. It begins by introducing the basic concepts and objectives of the multivehicle mission-planning problem. Then it formulates the problem mathematically and analyzes parameters in the objective function. The solution approach uses a hierarchical mission-planning scheme to take advantage of a scalable architecture. We develop a heuristic-based algorithm to solve the multiple-vehicle mission-planning problem. The algorithm has two phases: goal-point partitioning and routing. Goal-point partitioning uses a sweep procedure to group goal-points. Routing uses an implementation of simulated annealing combined with well-known TSP heuristics. Through the computational experiments conducted on both traveling salesman problem test cases, the TSPLIB library, and randomly generated test data, the routing algorithm performs quite well. It has been able to find TSP tours within one percent of optimality, and typically within one-half of one percent. The integration of the two-phase approach provides a solution to the multiple autonomous vehicle mission planning problem.
by Wei Zhao.
S.M.

Conceptual design optimization (CDO) is a technique proposed for the structured evaluation of different design concepts. Design grammars provide a flexible modular modelling architecture. The model is generated by a compiler based on predefined components and rules. The rules describe the composition of the model; thus, different models can be optimized by the CDO in one run. This allows considering a mission design including the mission analysis and the system design. The combination of a CDO approach with a model based on design grammars is shown for the concept study of a near-Earth asteroid mission. The mission objective is to investigate two asteroids of different kinds. The CDO reveals that a mission concept using two identical spacecrafts flying to one target each is better than a mission concept with one spacecraft flying to two asteroids consecutively.

Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2007.
Cataloged from PDF version of thesis.
Includes bibliographical references (p. 97).
This thesis investigates a new class of launch vehicles capable of being released from an aircraft which ultimately have the goal of achieving near-rendezvous conditions at orbital altitudes up to 800 km. These launch vehicles would be capable of carrying small payloads, on the order of two to six kilograms, and would be much more responsive to a customer's needs than the current space launch infrastructure, in which it may take months of preparation for a launch. To fully describe the mission in this thesis, it is broken up into three phases: atmospheric launch, orbit raising, and near-rendezvous operations. An analysis method known as Linear Covariance analysis is introduced to provide a platform of estimating the navigation covariance and dispersion of the spacecraft during the second and third phases, while the first phase, up to main-engine-cutoff, is examined using a three degree-of-freedom simulation. The goal of this thesis is to demonstrate the utility of Linear Covariance analysis to responsive space mission planning. This is accomplished by first explaining the mathematics that underlie the method. Next the software used for the analysis, Lincov Tools, is explained in detail, the mission is examined more closely, and the hardware for both the payload and launch vehicle are briefly discussed. Finally, the combination of the three degree-of-freedom simulation and Lincov Tools are employed to the space mission and the results are presented.
by Benjamin Lee Visser.
S.M.

Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2009.
This electronic version was submitted by the student author. The certified thesis is available in the Institute Archives and Special Collections.
Includes bibliographical references (p. 102-106) and index.
Uncontrolled reentry into the Earth atmosphere by LEO space missions whilst complying with stipulated NASA Earth atmospheric reentry requirements is a vital endeavor for the space community to pursue. An uncontrolled reentry mission that completely ablates does not require a provision for integrated controlled reentry capability. Consequently, not only will such a mission design be relatively simpler and cheaper, but also mission unavailability risk due to a controlled reentry subsystem failure is eliminated, which improves mission on-orbit reliability and robustness. Intentionally re-designing the mission such that the spacecraft components ablate (demise) during uncontrolled reentry post-mission disposal is referred to as Design-for-Demise (DfD). Re-designing spacecraft parts to demise guarantees adherence to NASA reentry requirements that dictate the risk of human casualty anywhere on Earth due to a reentering debris with KE =/> 15J be less than 1:10,000 (0.0001). NASA sanctioned missions have traditionally ad- dressed this requirement by integrating a controlled reentry provision. However, momentum is building for a new paradigm shift towards designing reentry missions to demise instead. Therefore, this thesis proposes a DfD decision making methodology; DfD implementation and execution strategy throughout the LEO mission life-cycle; scrutinizes reentry analysis software tools and uses NASA Debris Analysis Software (DAS) to demonstrate the reentry demisability analysis process; proposes methods to identify and redesign hardware parts for demise; and finally considers the HETE-2 mission as a DfD demisability case study. Reentry analysis show HETE-2 mission to be compliant with NASA uncontrolled atmospheric reentry requirements.
by Waswa M.B. Peter.
S.M.

In the analysis of clinical trials it is well-known that the omission of subjects randomized to treatments from the analysis can lead to bias in the final inference. A remedy that is widely adopted is to use the dictum of analysis by intention-to-treat (ITT), in which the groups as randomized are compared, even if the subjects have received treatments other than those prescribed in the protocol or have otherwise deviated from the prescribed regimen. A greater difficulty arises when subjects default and no response can be observed. This approach can often be rationalised in terms of comparing appropriate treatment policies. Crossover trials, in which subjects receive more than one of the trial treatments, present special problems. It is no longer clear that the use of ITT is appropriate. This work provides an overview of current approaches to missing or off-protocol responses in crossover trials and considers the underlying principle of what should be estimated from the data. Using simulation studies a rationale for the handling of missing data in crossover trials is developed. The model for the simple AB/BA crossover design is looked at in the simulation studies and analysed in the presence of ignorable or non-ignorable missing data and in the situation of off-protocol responses. This model is then extended to higher order crossover designs. From the outcome of these simulation studies we introduce a simple method for handling missing data in crossover trials. The motivation of this work lies in the data from Frank et al. (2008) where we have an AB/BA crossover trial with a high proportion of subjects who withdraw before the end of the trial due to side effects of their treatment.

Aquesta tesi doctoral està emmarcada en el camp de l'astrodinàmica. Presenta solucions a problemes identificats en el disseny de missions que utilitzen òrbites entorn dels punts de libració, fent servir la teoria de sistemes dinàmics.
El problema restringit de tres cossos és un model per estudiar el moviment d'un cos de massa infinitessimal sota l'atracció gravitatòria de dos cossos molt massius. Els cinc punts d'equilibri d'aquest model, en especial L1 i L2, han estat motiu de nombrosos estudis per aplicacions pràctiques en les últimes dècades (SOHO, Genesis...).
Genèricament, qualsevol missió en òrbita al voltant del punt L2 del sistema Terra-Sol es veu afectat per ocultacions degudes a l'ombra de la Terra. Si l'òrbita és al voltant de L1, els eclipsis són deguts a la forta influència electromagnètica del Sol. D'entre els diferents tipus d'òrbites de libració, les òrbites de Lissajous resulten de la combinació de dues oscil.lacions perpendiculars. El seu principal avantatge és que les amplituds de les oscil.lacions poden ser escollides independentment i això les fa adapatables als requeriments de cada missió. La necessitat d'estratègies per evitar eclipsis en òrbites de Lissajous entorn dels punts L1 i L2 motivaren la primera part de la tesi. En aquesta part es presenta una eina per la planificació de maniobres en òrbites de Lissajous que no només serveix per solucionar el problema d'evitar els eclipsis, sinó també per trobar trajectòries de transferència entre òrbites d'amplituds diferents i planificar rendez-vous.
Per altra banda, existeixen canals de baix cost que uneixen els punts L1 i L2 d'un sistema donat i representen una manera natural de transferir d'una regió de libració a l'altra. Gràcies al seu caràcter hiperbòlic, una òrbita de libració té uns objectes invariants associats: les varietats estable i inestable. Si tenim present que la varietat estable està formada per trajectòries que tendeixen cap a l'òrbita a la qual estan associades quan el temps avança, i que la varietat inestable fa el mateix però enrera en el temps, una intersecció entre una varietat estable i una d'inestable proporciona un camí asimptòtic entre les òrbites corresponents. Un mètode per trobar connexions d'aquest tipus entre òrbites planes entorn de L1 i L2 es presenta a la segona part de la tesi, i s'hi inclouen els resultats d'aplicar aquest mètode als casos dels problemes restringits Sol Terra i Terra-Lluna.
La idea d'intersecar varietats hiperbòliques es pot aplicar també en la cerca de camins de baix cost entre les regions de libració del sistema Sol-Terra i Terra-Lluna. Si existissin camins naturals de les òrbites de libració solars cap a les lunars, s'obtindria una manera barata d'anar a la Lluna fent servir varietats invariants, cosa que no es pot fer de manera directa. I a l'inversa, un camí de les regions de libració lunars cap a les solars permetria, per exemple, que una estació fos col.locada en òrbita entorn del punt L2 lunar i servís com a base per donar servei a les missions que operen en òrbites de libració del sistema Sol-Terra. A la tercera part de la tesi es presenten mètodes per trobar trajectòries de baix cost que uneixen la regió L2 del sistema Terra-Lluna amb la regió L2 del sistema Sol-Terra, primer per òrbites planes i més endavant per òrbites de Lissajous, fent servir dos problemes de tres cossos acoblats. Un cop trobades les trajectòries en aquest model simplificat, convé refinar-les per fer-les més realistes. Una metodologia per obtenir trajectòries en efemèrides reals JPL a partir de les trobades entre òrbites de Lissajous en el model acoblat es presenta a la part final de la tesi. Aquestes trajectòries necessiten una maniobra en el punt d'acoblament, que és reduïda en el procés de refinat, arribant a obtenir trajectòries de cost zero quan això és possible.
This PhD. thesis lies within the field of astrodynamics. It provides solutions to problems which have been identified in mission design near libration points, by using dynamical systems theory.
The restricted three body problem is a well known model to study the motion of an infinitesimal mass under the gravitational attraction of two massive bodies. Its five equilibrium points, specially L1 and L2, have been the object of several studies aimed at practical applications in the last decades (SOHO, Genesis...).
In general, any mission in orbit around L2 of the Sun-Earth system is affected by occultations due to the shadow of the Earth. When the orbit is around L1, the eclipses are caused by the strong electromagnetic influence of the Sun. Among all different types of libration orbits, Lissajous type ones are the combination of two perpendicular oscillations. Its main advantage is that the amplitudes of the oscillations can be chosen independently and this fact makes Lissajous orbits more adaptable to the requirements of each particular mission than other kinds of libration motions. The need for eclipse avoidance strategies in Lissajous orbits around L1 and L2 motivated the first part of the thesis. It is in this part where a tool for planning maneuvers in Lissajous orbits is presented, which not only solves the eclipse avoidance problem, but can also be used for transferring between orbits having different amplitudes and for planning rendez-vous strategies.
On the other hand, there exist low cost channels joining the L1 and L2 points of a given sistem, which represent a natural way of transferring from one libration region to the other one. Furthermore, there exist hyperbolic invariant objects, called stable and unstable manifolds, which are associated with libration orbits due to their hyperbolic character. If we bear in mind that the stable manifold of a libration orbit consists of trajectories which tend to the orbit as time goes by, and that the unstable manifold does so but backwards in time, any intersection between a stable and an unstable manifold will provide an asymptotic path between the corresponding libration orbits. A methodology for finding such asymptotic connecting paths between planar orbits around L1 and L2 is presented in the second part of the dissertation, including results for the particular cases of the Sun-Earth and Earth-Moon problems.
Moreover, the idea of intersecting hyperbolic manifolds can be applied in the search for low cost paths joining the libration regions of different problems, such as the Sun-Earth and the Earth-Moon ones. If natural paths from the solar libration regions to the lunar ones was found, it would provide a cheap way of transferring to the Moon from the vicinity of the Earth, which is not possible in a direct way using invariant manifolds. And the other way round, paths from the lunar libration regions to the solar ones would allow for the placement of a station in orbit around the lunar L2, providing services to solar libration missions, for instance. In the third part of the thesis, a methodology for finding low cost trajectories joining the lunar L2 region and the solar L2 region is presented. This methodology was developed in a first step for planar orbits and in a further step for Lissajous type orbits, using in both cases two coupled restricted three body problems to model the Sun-Earth-Moon spacecraft four body problem. Once trajectories have been found in this simplified model, it is convenient to refine them to more realistic models. A methodology for obtaining JPL real ephemeris trajectories from the initial ones found in the coupled models is presented in the last part of the dissertation. These trajectories need a maneuver at the coupling point, which can be reduced in the refinement process until low cost connecting trajectories in real ephemeris are obtained (even zero cost, when possible).

International Telemetering Conference Proceedings / October 17-20, 1994 / Town & Country Hotel and Conference Center, San Diego, California
Spacecraft on-board time keeping, to an accuracy better than 1 millisecond, is a requirement for many satellite missions. Scientific satellites must precisely "time tag" their data to allow it to be correlated with data produced by a network of ground and space based observatories. Multiple vehicle satellite missions, and satellite networks, sometimes require several spacecraft to execute tasks in time phased fashion with respect to absolute time. In all cases, mission systems designed to provide a high accuracy on-board clock must necessarily include mechanisms for the determination and correction of spacecraft clock error. In addition, an approach to on-orbit verification of these mechanisms may be required. Achieving this accuracy however need not introduce significant mission cost if the task of maintaining this accuracy is appropriately distributed across both the space and ground mission segments. This paper presents the mission systems approaches taken by two spacecraft programs to provide high accuracy on-board spacecraft clocks at minimum cost. The first, NASA Goddard Space Flight Center's (GSFC) Extreme Ultraviolet Explorer (EUVE) program demonstrated the ability to use the NASA Tracking and Data Relay Satellite System (TDRSS) mission environment to maintain an on-board spacecraft clock to within 100 microseconds of Naval Observatory Standard (NOS) Time. The second approach utilizes an on-board spacecraft Global Positioning System (GPS) receiver as a time reference for spacecraft clock tracking which is facilitated through the use of Fairchild's Telemetry and Command Processor (TCP) spacecraft Command & Data Handling Subsystem Unit. This approach was designed for a future Shuttle mission requiring the precise coordination of events among multiple space-vehicles.

Solar sailing is the use of a thin and lightweight membrane to reflect sunlight and obtain a thrust force on the spacecraft. That is, a sailcraft has a potentially-infinite specific impulse and, therefore, it is an attractive solution to reach mission goals otherwise not achievable, or very expensive in terms of propellant consumption. The recent scientific interest in near-Earth asteroids (NEAs) and the classification of some of those as potentially hazardous asteroids (PHAs) for the Earth stimulated the interest in their exploration. Specifically, a multiple NEA rendezvous mission is attractive for solar-sail technology demonstration as well as for improving our knowledge about NEAs. A preliminary result in a recent study showed the possibility to rendezvous three NEAs in less than ten years. According to the NASA’s NEA database, more than 12,000 asteroids are orbiting around the Earth and more than 1,000 of them are classified as PHA. Therefore, the selection of the candidates for a multiple-rendezvous mission is firstly a combinatorial problem, with more than a trillion of possible combinations with permutations of only three objects. Moreover, for each sequence, an optimal control problem should be solved to find a feasible solar-sail trajectory. This is a mixed combinatorial/optimisation problem, notoriously complex to tackle all at once. Considering the technology constraints of the DLR/ESA Gossamer roadmap, this thesis focuses on developing a methodology for the preliminary design of a mission to visit a number of NEAs through solar sailing. This is divided into three sequential steps. First, two methods to obtain a fast and reliable trajectory model for solar sailing are studied. In particular, a shape-based approach is developed which is specific to solar-sail trajectories. As such, the shape of the trajectory that connects two points in space is designed and the control needed by the sailcraft to follow it is analytically retrieved. The second method exploits the homotopy and continuation theory to find solar-sail trajectories starting from classical low-thrust ones. Subsequently, an algorithm to search through the possible sequences of asteroids is developed. Because of the combinatorial characteristic of the problem and the tree nature of the search space, two criteria are used to reduce the computational effort needed: (a) a reduced database of asteroids is used which contains objects interesting for planetary defence and human spaceflight; and (b) a local pruning is carried out at each branch of the tree search to discard those target asteroids that are less likely to be reached by the sailcraft considered. To reduce further the computational effort needed in this step, the shape-based approach for solar sailing is used to generate preliminary trajectories within the tree search. Lastly, two algorithms are developed which numerically optimise the resulting trajectories with a refined model and ephemerides. These are designed to work with minimum input required by the user. The shape-based approach developed in the first stage is used as an initial-guess solution for the optimisation. This study provides a set of feasible mission scenarios for informing the stakeholders on future mission options. In fact, it is shown that a large number of five-NEA rendezvous missions are feasible in a ten-year launch window, if a solar sail is used. Moreover, this study shows that the mission-related technology readiness level for the available solar-sail technology is larger than it was previously thought and that such a mission can be performed with current or at least near-term solar sail technology. Numerical examples are presented which show the ability of a solar sail both to perform challenging multiple NEA rendezvous and to change the mission en-route.

Missions concerning small-body celestial objects are of growing interest due to the resources and information they can provide. Such missions require detailed information about the surface of the body for interactions, such as landing on the surface, as well as predicting the gravity field of the object. This work provides a means of optimizing the mission elements of trajectory and imaging target schedules so that the level of knowledge of the surface can be increased. The information required to increase one's knowledge of the surface is described as a set of conditions placed on the collection of images taken of each facet of the surface; these requirements constitute the concept of "coverage" and were provided by NASA. Currently, no comparable optimization capability exists. The trajectory optimization task is done using an adapted form of the Non-dominated Sorting Genetic Algorithm-2 (NSGA-2) in which the genetic mutation and recombination operators are replaced with operators inspired by a different Evolutionary Algorithm, Differential Evolution. Since small-body objects have irregular distributions of mass, this optimization accounts for this by using a higher fidelity gravity model; the expense of the calculation causing a significant increase in fitness evaluation time. The trajectory optimization uses the maximization of possible coverage (the coverage achieved is every surface element were targeted for imaging at every opportunity) and minimization of a time quantity that typifies covering less but doing so quickly as the primary optimization objectives with an additional ancillary objective which rewards the fulfillment of the individual aspects of coverage so as to better condition improvement in the first objective. The optimization of imaging schedules is handled using a less adapted version of NSGA-2 in which the base operations were only tailored slightly. This differs from the previous task in that limitation are placed on the imaging process; namely that the camera may only target a single surface element at each opportunity and is thus only able to observe the faces caught within the narrow field-of-view. This optimization trades the minimization of time objective and the ancillary objective for the minimization of the required rotational effort of the imaging spacecraft. Both works result in sets of solutions to their respective problems that capture the trade-space between the considered objectives. The last work detailed here examines the consequences of how velocity domains are phrased in space trajectory optimization problems. Multiple means of framing the optimization domain are examined and the results detail the complications encountered by the more common formulations for a set of test problems.

Ces travaux de recherche portent sur la problématique de la mise en pratique de l'analyse de cycle de vie (ACV). La question principale est : comment faire une ACV plus rapide et plus facilement accessible pour la conception des produits ? Nous nous concentrons sur deux problématiques qui prolongent l'inventaire de Cycle de Vie (ICV) : • recherche des données manquantes : Comment ranger les données manquantes selon leur importance? Comment traiter l'intersection des aspects qualitatifs et les aspects quantitatifs des données manquantes? • Modélisation du cycle de vie : Comment réutiliser le cycle de vie existant pour un nouveau produit? Comment développer un modèle de référence? Pour la recherche des solutions nous avons utilisé l'approche "Case study" selon Robert Yin. Nos contributions font résultat de trois études de cas, dont la plus importante est l'ACV du High Impact Polypropylene (HIPP) recyclé. Nous avons publié les résultats de celle-ci dans la revue scientifique Journal of Cleaner Production. Suite aux études de cas nous proposons deux approches d'amélioration d'efficacité en ICV : nous proposons l'analyse de sensibilité préalable pour classifier les données manquantes selon leur impact sur les résultats d'ACV. L'approche combine les aspects quantitatifs avec les aspects quantitatifs en protégeant le respect des objectives d'étude. Nous appelons cette protection "LCA Poka-Yoké". La modélisation du cycle de vie peut être assistée grace à la méthode basée sur l'algorithme de King. Pour la continuation de la recherche nous proposons huit perspectives, dont six font l'objet d'intégration des nouvelles approches d'amélioration dans les concepts d'ACV basés sur la norme ISO 14025 ou dans le projet de la Commission Européenne PEF
This work addresses the different issues that put a brake to using Lifecycle assessment (LCA) in product design by answering the main question of the research: How to make Lifecycle assessment faster and easier accessible for manufactured product design? In the LCA methodology we have identified two issues to deal with and their consecutive scientific locks : • Research of missing data : How to organize missing data? How to respect quantitative and qualitative dimensions? • Modeling of the lifecycle scenario : How to translate methodological choices into the lifecycle scenario model? How to transform the reference scenario into a new one? We have dealt with these issues using the scientific approach Case study according toRobert Yin. Our contributions are based on three case studies, between which the most important is study of High Impact Polypropylene recycling in the automotive industry. We have published it in the Journal of Cleaner Production. As result of our research we present two methods to improve efficiency of the LifecycleInventory Analysis (LCI) : To organize the missing data: Preliminary sensitivity analysis with LCA Poka-Yoke ; To help with scenario modeling: Method of workflows factorization, based on Reverse engineering. For further research we propose eight perspectives, mostly based on integration of our methods into Product Category Rules (PCR)-based platforms like EPD International or the European PEF

Thesis: S.M. in Engineering and Management, Massachusetts Institute of Technology, School of Engineering, System Design and Management Program, Engineering and Management Program, 2016.
This electronic version was submitted by the student author. The certified thesis is available in the Institute Archives and Special Collections.
Cataloged from student-submitted PDF version of thesis.
Includes bibliographical references (pages 99-100).
In the engineering of complex systems, changes to flight hardware or software after initial release can have large impacts on project implementation. Even a comparatively small change on an assembly or subsystem can cascade into a significant amount of rework if it propagates through the system. This can happen when a change affects the interfaces with another subsystem, or if it alters the emergent behavior of the system in a significant way, and is especially critical when subsequent work has already been performed utilizing the initial version. These changes can be driven by new or modified requirements leading to changes in scope, design deficiencies discovered during analysis or test, failures during test, and other such mechanisms. In complex system development, changes are managed through engineering change requests (ECRs) that are communicated to affected elements. While the tracking of changes is critical for the ongoing engineering of a complex project, the ECRs can also reveal trends on the system level that could assist with the management of current and future projects. In an effort to identify systematic trends, this research has analyzed ECRs from two different JPL led space mission projects to classify the change activity and assess change propagation. It employs time analysis of ECR initiation throughout the lifecycle, correlates ECR generators with ECR absorbers, and considers the distribution of ECRs across subsystems. The analyzed projects are the planetary rover mission, Mars Science Laboratory (MSL), and the Earth-orbiting mission, Soil Moisture Active Passive (SMAP). This analysis has shown that there is some consistency across these projects with regard to which subsystems generate or absorb change. The relationship of the ECRSubsystem networks identifies subsystems that are absorbers of change and others that are generators of change. For the flight systems, the strongest absorbers of change were found to be avionics and the mechanical structure for the spacecraft bus, and the strongest generators of change were concentrated in the payloads. When this attribute is recognized, project management can attempt to close ECR networks by looking for ways to leverage absorbers and avoid multipliers. Alternatively, in cases where changes to a subsystem are undesirable, knowing whether it is an absorber can greatly assist with expectations and planning. This analysis identified some significant differences between the two projects as well. While SMAP followed a relatively well behaved blossom profile across the project, MSL had an avalanche of change leading to the drastic action of re-baselining the launch date. While the official reasoning for the slip of the launch date is based in technical difficulties, the avalanche profile implies that a snowballing of change may have had a significant impact as well. Furthermore, the complexity metrics applied show that MSL has a more complex nature than SMAP, with 269 ECRs in 65 Parent-Child clusters, opposed to 166 in 53 for SMAP, respectively. The Process Complexity metric confirms this, quantitatively measuring the complexity of MSL at 492, compared to 367 for SMAP. These tools and metrics confirm the intuition that MSL, as a planetary rover, is a more complex space mission than SMAP, an earth orbiter.
by Eric D. Ward.
S.M. in Engineering and Management

The main body of this thesis consists of six appended papers. In the  first two, different  cooperative surveillance problems are considered. The second two consider different aspects of the trajectory planning problem, while the last two deal with observer design for mobile robotic and Euler-Lagrange systems respectively.In Papers A and B,  a combinatorial optimization based framework to cooperative surveillance missions using multiple Unmanned Ground Vehicles (UGVs) is proposed. In particular, Paper A  considers the the Minimum Time UGV Surveillance Problem (MTUSP) while Paper B treats the Connectivity Constrained UGV Surveillance Problem (CUSP). The minimum time formulation is the following. Given a set of surveillance UGVs and a polyhedral area, find waypoint-paths for all UGVs such that every point of the area is visible from  a point on a waypoint-path and such that the time for executing the search in parallel is minimized.  The connectivity constrained formulation  extends the MTUSP by additionally requiring the induced information graph to be  kept recurrently connected  at the time instants when the UGVs  perform the surveillance mission.  In these two papers, the NP-hardness of  both these problems are shown and decomposition techniques are proposed that allow us to find an approximative solution efficiently in an algorithmic manner.Paper C addresses the problem of designing a real time, high performance trajectory planner for an aerial vehicle that uses information about terrain and enemy threats, to fly low and avoid radar exposure on the way to a given target. The high-level framework augments Receding Horizon Control (RHC) with a graph based terminal cost that captures the global characteristics of the environment.  An important issue with RHC is to make sure that the greedy, short term optimization does not lead to long term problems, which in our case boils down to two things: not getting into situations where a collision is unavoidable, and making sure that the destination is actually reached. Hence, the main contribution of this paper is to present a trajectory planner with provable safety and task completion properties. Direct methods for trajectory optimization are traditionally based on a priori temporal discretization and collocation methods. In Paper D, the problem of adaptive node distribution is formulated as a constrained optimization problem, which is to be included in the underlying nonlinear mathematical programming problem. The benefits of utilizing the suggested method for  online  trajectory optimization are illustrated by a missile guidance example.In Paper E, the problem of active observer design for an important class of non-uniformly observable systems, namely mobile robotic systems, is considered. The set of feasible configurations and the set of output flow equivalent states are defined. It is shown that the inter-relation between these two sets may serve as the basis for design of active observers. The proposed observer design methodology is illustrated by considering a  unicycle robot model, equipped with a set of range-measuring sensors. Finally, in Paper F, a geometrically intrinsic observer for Euler-Lagrange systems is defined and analyzed. This observer is a generalization of the observer proposed by Aghannan and Rouchon. Their contractivity result is reproduced and complemented  by  a proof  that the region of contraction is infinitely thin. Moreover, assuming a priori bounds on the velocities, convergence of the observer is shown by means of Lyapunov's direct method in the case of configuration manifolds with constant curvature.
QC 20100622
TAIS, AURES

Thesis (MEng (Electrical Engineering))--Cape Peninsula University of Technology, 2017.
Real-time remote monitoring of Africa’s resources, such as water quality, by using terrestrial sensors is impeded by the limited connectivity over the vast rural areas of the continent. Without such monitoring, the effective management of natural resources, and the response to associated disasters such as flooding, is almost impossible. A constellation of nanosatellites could provide near real-time connectivity with ground-based sensors that are distributed across the continent. This study evaluates the high level development of a mission design for a near real-time remote monitoring CubeSat constellation and ground segment for in-situ monitoring in regions of interest on the African continent. This would facilitate management of scarce resources using a low-cost constellation. To achieve this, the design concept and operation of a Walker constellation are examined as a means of providing connectivity to a low bit rate sensor network distributed across geographic areas of interest in South Africa, Algeria, Kenya and Nigeria. The mission requirements include the optimisation of the constellation to maintain short revisit times over South Africa and an investigation of the required communications link to perform the operations effectively. STK software is used in the design and evaluation of the constellations and the communications system. The temporal performance parameters investigated are access and revisit times of the constellations to the geographic areas mentioned. The types of constellation configurations examined, involved starting with a system level analysis of one satellite. This seed satellite has known orbital parameters. Then a gradual expansion of two to twelve satellites in one, two and three orbital planes follows. VHF, UHF and S-band communication links are considered for low data rate in-situ monitoring applications. RF link budgets and data budgets for typical applications are determined. For South Africa, in particular, a total of 12 satellites evenly distributed in a two-plane constellation at an inclination of 39° provide the optimal solution and offer an average daily revisit time of about 5 minutes. This constellation provides average daily access time of more than 16 hours per day. A case study is undertaken that decribes a constellation for the provision of maritime vessel tracking in the Southern African oceans using the Automated Information System (AIS). This service supports the Maritime Domain Awareness (MDA) initiative implemented by the South African Government, under its Operation Phakisa.
National Research Foundation (NRF) French South African Institute of Technology (F’SATI)

The contributions of this dissertation increase the fidelity of conceptual hypersonic mission design through the following innovations: 1) the introduction of coupling between the effects of ablation of the thermal protection system (TPS) and flight dynamics, 2) the introduction of rigid body dynamics into trajectory design, and 3) simplifying the design of hypersonic missions that involve multiple phases of flight. These contributions are combined into a unified conceptual mission design framework, which is in turn applicable to slender hypersonic vehicles with ablative TPS. Such vehicles are employed in military applications, wherein speed and terminal energy are of critical importance.

The fundamental observation that results from these contributions is the substantial reduction in the maximum terminal energy that is achievable when compared to the state-of-the art conceptual design process. Additionally, the control history that is required to follow the maximum terminal energy trajectory is also significantly altered, which will in turn bear consequence on the design of the control actuators.

The other important accomplishment of this dissertation is the demonstration of the ability to solve these class of problems using indirect methods. Despite being built on a strong foundation of the calculus of variations, the state-of-the-art entirely neglects indirect methods because of the challenge associated with solving the resulting boundary value problem (BVP) in a system of differential-algebraic equations (DAEs). Instead, it employs direct methods, wherein the optimality of the calculated trajectory is not guaranteed. The ability to employ indirect methods to solve for optimal trajectories that are comprised of multiple phases of flight while also accounting for the effects of ablation of the TPS and rigid body dynamics is a substantial advancement in the state-of-the-art.

The equations governing the thermodynamic behavior of a military aircraft have been implemented by the Air Force Research Lab (AFRL) and other Integrated Vehicle Energy Technology Demonstration (INVENT) contributors into a cohesive, adaptable, dynamic aircraft simulation program in Mathworks' Simulink®. The resulting model known as the "Tip-to-tail" model meets the design specifications set forth by the INVENT program. The system consists of six intimately linked subsystems that include a propulsion subsystem (PS), air vehicle subsystem (AVS), robust electrical power subsystem (REPS), high power electric actuation subsystem (HPEAS), advanced power and thermal management subsystem (APTMS), and a fuel thermal management subsystem (FTMS). The model's governing equations are augmented with experimental data and supported by defined physical parameters. In order to address the problems associated with the additional power and thermal loads for in more electric aircraft (MEA), this research utilizes exergy analysis and mission-integrated synthesis/design optimization to investigate the potential for improvement in tip-to-tail design/performance. Additionally, this thesis describes the development and integration of higher fidelity transient heat exchanger models for use in the tip-to-tail. Finally, the change in performance due to the integration of new heat exchanger models developed here is presented. Additionally, this thesis discusses the results obtained by performing mission-integrated synthesis/design optimization on the tip-to-tail using heat exchanger design parameters as decision variables. These results show that the performance of the tip-to-thermal management subsystems improves significantly due to the integration of the heat exchanger models. These results also show improvements in vehicle performance due to the mission-integrated optimization.
Master of Science

Increasing insensitivity demands on designing and producing munitions necessitates utilizing primarily insensitive initiation trains specifically in missile systems. Exploding Foil Initiator (EFI) is a high voltage detonator that is used as the initiation elementof rocket motor and warhead initiation trains of modern insensitive missile systems. In this thesis, EFI prototypes are designed and manufactured with the knowledge gained from detailed literature studies. An experimental setup is constructed including firing and testing means for EFI prototypes. That experimental setup is capable of firing EFI prototypes from 500 volts to 3000 volts voltage range. Besides, it allows measuring electrical characteristics like current and voltage traces and average velocity of the flyer plates of these prototypes.Using EFI prototypes,detonation tests of HNS &ndash
IV and PBXN &ndash
5 explosive pellets are carried out.Function times and detonation outputs of the prototypesare measured with the same experimental setup. A numerical study which predicts electrical performance of EFI prototypes and impact characteristics of flyer plates are carried out. Numerical code is validated with the experimental results.

Optical instruments for space missions work in hostile environment, it’s thus necessary to accurately study the effects of ambient parameters variations on the equipment. In particular, optical instruments are very sensitive to ambient conditions, especially temperature. This variable can cause dilatation and misalignment of the optical elements, and can also lead to rise of dangerous stresses in the optics. Their displacements degrade the quality of the sampled images. In this work the optics and mountings of a stereo-camera for the BepiColombo mission are modelled and processed by a thermo-mechanical FEM analysis, reproducing expected operative conditions. The output is elaborated into a MATLAB optimisation code, based on non-linear least square algorithm to determine the equation of the best fitting nth polynomial or spherical surfaces of the deformed lenses and mirrors; model accuracy is 10-8 m. The obtained mathematical surface representations are then directly imported into ZEMAX for sequential raytrace analysis. The results are spot diagrams, chief ray coordinates on the detector, MTF curves and Diffraction Encircled Energy variations due to simulated thermal loads. This helps also to design and compare different optical housing systems for a feasible solution. Different types of lenses and prisms constraints have been designed and analysed. The results show the preferable use of kinematic constraints instead of glue to correctly operate the instrument to maintaine in focus in orbit around Mercury considering an operative temperature between -20°C and +30°C.
Gli strumenti ottici che vengono utilizzati nelle missioni spaziali risentono delle variazione delle condizioni ambientali, per questo è necessario studiare l’effetto di queste ultime sull’equipaggiamento. In particolare gli strumenti ottici sono molto sensibili alle variazioni di temperatura, perché questa grandezza può causare la deformazione e il disalinneamento delle ottiche, inoltre può comportare l’insorgere di tensioni rilevanti che possono provocare la loro rottura. In questa tesi sono state effettuate delle analisi termo-elastiche utilizzando un software ad elementi finite (Nastran) riproducendo le condizioni operative in cui si troverà la stereocamera coinvolta nella missione BepiColombo. I risultati delle analisi sono stati elaborati in MATLAB per determinare le equazioni matematiche delle superfici degli elementi ottici deformati utilizzando un’ottimizzazione non lineare ai minimi quadrati, e considerando equazioni polinomiali, sferiche e planari. Le superfici matematiche sono state importate in un software raytrace (ZEMAX) per poter verificare la performance ottica dello strumento. I risultati mostrano come le variazioni di temperatura influenzino gli Spot Diagrams, la Diffraction Ensquare Energy e le curve MTF. Per migliorare la risposta del telescopio ai carichi termici sono stati ideati dei vincoli cinematici, il loro utilizzo compromette molto meno la performance della stereocamera rispetto a vincoli rigidi per le ottiche. È stata valutata l’influenza delle variazioni dei parametri ottici (focale, spostamento del centro ottico, spostamento degli Spot Diagrams sul piano immagine e distorsioni) sulla ricostruzione della profondità della superficie di Mercurio utilizzando la propagazione dell’incertezza secondo le metodologie GUM e Monte Carlo. Infine è stato ideato unsetup strumentale per determinare gli spostamenti e le rotazioni di alcuni elementi ottici della sterocamera in camera di vuoto riproducendo le condizioni operative dello strumento.