Дисертації з теми "Mission design and analysi"
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Shastri, Bhardwaj. "Design and analysis of mission and system requirements for 'NetSat' mission with respect to structural and thermal limitations." Thesis, Luleå tekniska universitet, Rymdteknik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-76336.
Повний текст джерелаTanapura, Noravidhya. "Preliminary Mission Analysis and Design for a Small Satellite SWARM." Thesis, KTH, Rymd- och plasmafysik, 2012. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-104032.
Повний текст джерелаUnlusoy, Levent. "Structural Design And Analysis Of The Mission Adaptive Wings Of An Unmanned Aerial Vehicle." Master's thesis, METU, 2010. http://etd.lib.metu.edu.tr/upload/12611515/index.pdf.
Повний текст джерелаKim, Susan C. (Susan Cecilia). "Mission design and trajectory analysis for inspection of a host spacecraft by a microsatellite." Thesis, Massachusetts Institute of Technology, 2006. http://hdl.handle.net/1721.1/37566.
Повний текст джерелаIncludes 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.
Paris, Bethany L. "INSTITUTIONAL LENDING MODELS, MISSION DRIFT, AND MICROFINANCE INSTITUTIONS." UKnowledge, 2013. http://uknowledge.uky.edu/msppa_etds/9.
Повний текст джерелаInsuyu, Erdogan Tolga. "Aero-structural Design And Analysis Of An Unmanned Aerial Vehicle And Its Mission Adaptive Wing." Master's thesis, METU, 2010. http://etd.lib.metu.edu.tr/upload/12611657/index.pdf.
Повний текст джерелаZimmer, Aline [Verfasser]. "Mission Analysis and Conceptual Spacecraft Design for Human Exploration of Near-Earth Asteroids / Aline Zimmer." München : Verlag Dr. Hut, 2012. http://d-nb.info/1029400342/34.
Повний текст джерелаCucciarrè, Francesca. "Numerical and experimental methods for design and test of units and devices on BepiColombo Mission." Doctoral thesis, Università degli studi di Padova, 2013. http://hdl.handle.net/11577/3423379.
Повний текст джерелаL’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
Wertz, Julie (Julie Ann) 1978. "Expected productivity-based risk analysis in conceptual design : with application to the Terrestrial Planet Finder Interferometer mission." Thesis, Massachusetts Institute of Technology, 2005. http://hdl.handle.net/1721.1/35590.
Повний текст джерелаPage 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.
Gagliano, Joseph R. "Orbital Constellation Design and Analysis Using Spherical Trigonometry and Genetic Algorithms: A Mission Level Design Tool for Single Point Coverage on Any Planet." DigitalCommons@CalPoly, 2018. https://digitalcommons.calpoly.edu/theses/1877.
Повний текст джерелаSakarya, Evren. "Structural Design And Evaluation Of An Adaptive Camber Wing." Master's thesis, METU, 2010. http://etd.lib.metu.edu.tr/upload/12611514/index.pdf.
Повний текст джерелаCFD 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.
D'Anniballe, Alessandro. "Development of a sizing tool for preliminary mission analysis and design of propulsion systems for orbit control of small satellites in LEO -VLEO." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2017. http://amslaurea.unibo.it/14719/.
Повний текст джерелаBelguzhanov, Rustem. "Preliminary system design of the modular satellite." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2019.
Знайти повний текст джерелаKilian, Catherine A. (Catherine Anne). "Does "good design" add value? : a comparative analysis of two residential projects; the planned unit development of Mission Valley and the new town of Reston." Thesis, Massachusetts Institute of Technology, 1989. http://hdl.handle.net/1721.1/75539.
Повний текст джерелаAnthony, Niklas. "Prometheus Asteroid Redirection Mission : Mission Design, Spacecraft Design, Orbital Dynamics Code Development." Thesis, Luleå tekniska universitet, Institutionen för system- och rymdteknik, 2016. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-174.
Повний текст джерелаAlemany, Kristina. "Design space pruning heuristics and global optimization method for conceptual design of low-thrust asteroid tour missions." Diss., Atlanta, Ga. : Georgia Institute of Technology, 2009. http://hdl.handle.net/1853/31821.
Повний текст джерелаCommittee 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.
Moon, Jongki. "Mission-based guidance system design for autonomous UAVs." Diss., Atlanta, Ga. : Georgia Institute of Technology, 2009. http://hdl.handle.net/1853/31797.
Повний текст джерелаCommittee 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.
Cohanim, Babak 1980. "Mission design for safe traverse of planetary hoppers." Thesis, Massachusetts Institute of Technology, 2013. http://hdl.handle.net/1721.1/82476.
Повний текст джерелаThis 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.
Ogawa, Akira S. M. Massachusetts Institute of Technology. "Concurrent engineering for mission design in different cultures." Thesis, Massachusetts Institute of Technology, 2008. http://hdl.handle.net/1721.1/43175.
Повний текст джерелаIncludes 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.
Purville, Brian A. (Brian Alexander) 1978. "Design of information channels for mission-critical systems." Thesis, Massachusetts Institute of Technology, 2001. http://hdl.handle.net/1721.1/86680.
Повний текст джерелаIncludes bibliographical references (leaves 104-105).
by Brian A. Purville.
M.Eng.
Lin, Beldon Chi. "Integrated vehicle and mission design using convex optimization." Thesis, Massachusetts Institute of Technology, 2020. https://hdl.handle.net/1721.1/127070.
Повний текст джерелаCataloged 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
castello, brian. "CUBESAT MISSION PLANNING TOOLBOX." DigitalCommons@CalPoly, 2012. https://digitalcommons.calpoly.edu/theses/787.
Повний текст джерелаSPAGNUOLO, Gandolfo Alessandro. "INTEGRATED MULTI-PHYSICS DESIGN TOOL FOR FUSION BREEDING BLANKET SYSTEMS - DEVELOPMENT AND VALIDATION." Doctoral thesis, Università degli Studi di Palermo, 2020. http://hdl.handle.net/10447/395226.
Повний текст джерелаThe 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.
Curzi, Giacomo. "Trajectory design of a multiple flyby mission to asteroids." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2016.
Знайти повний текст джерелаPaskeviciute, Agne. "Preliminary Lander CubeSat Design for Small Asteroid Detumbling Mission." Thesis, KTH, Rymdteknik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-248427.
Повний текст джерелаGruvdrift 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.
Schumann, Benjamin. "Aeronautical life-cycle mission modelling framework for conceptual design." Thesis, University of Southampton, 2014. https://eprints.soton.ac.uk/366537/.
Повний текст джерелаMalakhoff, Lev A. "Combat aircraft mission tradeoff models for conceptual design evaluation." Diss., Virginia Polytechnic Institute and State University, 1988. http://hdl.handle.net/10919/53583.
Повний текст джерелаPh. D.
West, Jonathan David. "The Effects of Mission Statement Design on Behavioral Intention." Scholar Commons, 2016. http://scholarcommons.usf.edu/etd/6429.
Повний текст джерелаLocarini, Alfredo. "Design of a GNSS receiver for the ESEO mission." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2014. http://amslaurea.unibo.it/7617/.
Повний текст джерелаXiao, Size. "Pico-Satellite Design and Test for i-INSPIRE Mission." Thesis, The University of Sydney, 2013. http://hdl.handle.net/2123/9364.
Повний текст джерелаStill, Vincent. "Thermal Control Design and Simulation of a Space Mission." Thesis, Luleå tekniska universitet, Institutionen för system- och rymdteknik, 2018. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-71784.
Повний текст джерелаZhao, Wei 1966. "Multiple autonomous vehicle mission planning and management." Thesis, Massachusetts Institute of Technology, 1999. http://hdl.handle.net/1721.1/29165.
Повний текст джерелаIncludes 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.
Weber, A., S. Fasoulas, and K. Wolf. "Conceptual interplanetary space mission design using multi-objective evolutionary optimization and design grammars." Sage, 2011. https://publish.fid-move.qucosa.de/id/qucosa%3A38443.
Повний текст джерелаVisser, Benjamin Lee. "An application of linear covariance analysis to the design of responsive near-rendezvous missions." Thesis, Massachusetts Institute of Technology, 2007. http://hdl.handle.net/1721.1/59696.
Повний текст джерелаCataloged 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.
Waswa, M. B. Peter (Peter Moses Bweya). "Spacecraft design-for-demise strategy, analysis and impact on low earth orbit space missions." Thesis, Massachusetts Institute of Technology, 2008. http://hdl.handle.net/1721.1/46797.
Повний текст джерелаThis 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.
Rodgers, Lauren Rebecca. "Analysis of treatment effect in crossover designs with missing data." Thesis, University of Newcastle Upon Tyne, 2008. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.488660.
Повний текст джерелаCanalias, Vila Elisabet. "Contributions to Libration Orbit Mission Design using Hyperbolic Invariant Manifolds." Doctoral thesis, Universitat Politècnica de Catalunya, 2007. http://hdl.handle.net/10803/5927.
Повний текст джерелаEl 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).
Wickham, Mark E. "On-Board Spacecraft Time-Keeping Mission System Design and Verification." International Foundation for Telemetering, 1994. http://hdl.handle.net/10150/608549.
Повний текст джерелаSpacecraft 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.
Peloni, Alessandro. "Solar-sail mission design for multiple near-Earth asteroid rendezvous." Thesis, University of Glasgow, 2018. http://theses.gla.ac.uk/8901/.
Повний текст джерелаHinckley, David William. "Multi-Objective Optimization Mission Design for Small-Body Coverage Missions." ScholarWorks @ UVM, 2019. https://scholarworks.uvm.edu/graddis/1132.
Повний текст джерелаKozderka, Michal. "Parametric LCA approaches for efficient design." Thesis, Strasbourg, 2016. http://www.theses.fr/2016STRAD050/document.
Повний текст джерелаThis 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
Ward, Eric D. (Eric Daniel). "A socio-technical systems analysis of change processes in the design of flagship interplanetary missions." Thesis, Massachusetts Institute of Technology, 2016. http://hdl.handle.net/1721.1/107291.
Повний текст джерелаThis 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
Anisi, David A. "On Cooperative Surveillance, Online Trajectory Planning and Observer Based Control." Doctoral thesis, KTH, Optimeringslära och systemteori, 2009. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-9990.
Повний текст джерелаQC 20100622
TAIS, AURES
Enslin, Jason W. "An evolutionary algorithm approach to simultaneous multi-mission radar waveform design /." Online version of thesis, 2007. http://hdl.handle.net/1850/4770.
Повний текст джерелаMtshemla, Kanyisa Sipho. "Mission design of a CubeSat constellation for in-situ monitoring applications." Thesis, Cape Peninsula University of Technology, 2017. http://hdl.handle.net/20.500.11838/2633.
Повний текст джерелаReal-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)
Saranathan, Harish. "Algorithmic Advances to Increase the Fidelity of Conceptual Hypersonic Mission Design." Thesis, Purdue University, 2018. http://pqdtopen.proquest.com/#viewpdf?dispub=10792495.
Повний текст джерелаThe 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.
Weise, Peter Carl. "Mission-Integrated Synthesis/Design Optimization of Aerospace Subsystems under Transient Conditions." Thesis, Virginia Tech, 2012. http://hdl.handle.net/10919/76855.
Повний текст джерелаMaster of Science
Rivera, Francisco. "An object-oriented method of mission profile input for aircraft design." Thesis, This resource online, 1993. http://scholar.lib.vt.edu/theses/available/etd-09122009-040526/.
Повний текст джерелаYilmaz, Muhammed Yusuf. "Design And Analysis Of A High Voltage Exploding Foil Initiator For Missile Systems." Master's thesis, METU, 2013. http://etd.lib.metu.edu.tr/upload/12615437/index.pdf.
Повний текст джерелаIV 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.
Segato, Elisa. "Tecniche di taratura di stereocamere per missioni planetarie." Doctoral thesis, Università degli studi di Padova, 2010. http://hdl.handle.net/11577/3426956.
Повний текст джерелаGli 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.