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En este trabajo se presenta un estudio sobre el lanzamiento de cohetes orbitales desde una plataforma aérea, centrándose en el análisis del comportamiento y la trayectoria del cohete cuando se lanza bajo estas condiciones. El estudio ha sido supervisado por el profesor Josep Oriol Lizandra Dalmases y la profesora Blanca Tejedor Herrán, de la Universitat Politècnica de Catalunya. El estudio inicia con la presentación de los principios fundamentales que definen el comportamiento del motor y las ecuaciones gobernantes de la trayectoria del cohete. También se añaden detalles esenciales sobre la caracterización de la atmósfera, las órbitas, el combustible y la división por etapas del cohete. Posteriormente, se desarrolla el código de Matlab para la simulación del comportamiento del motor y el lanzamiento desde la superficie. El motor y la trayectoria están basados en los del cohete Electron de Rocket Lab, cuyo perfil será usado para validar el modelo. Finalmente, se optimiza el programa para el lanzamiento aéreo. Se determina la importancia de los parámetros dados por la plataforma aérea de velocidad, altura y ángulo de inclinación iniciales, valorando el incremento de velocidad que proporcionan. Con los beneficios obtenidos en cuanto a velocidad, se proponen mejoras como el aumento de la capacidad de carga de pago, disminución de la cantidad de combustible o reducción de potencia mediante la extracción de motores y combustible. Se determina la viabilidad técnica a efectos cualitativos del sistema de lanzamiento aéreo. Sería necesario realizar un estudio futuro sobre la plataforma aérea a utilizar y el método de separación para valorar el sistema de lanzamiento desde el aire completamente.

En els pròxims anys es preveu un gran creixement de l’activitat espacial i, sobretot, en l’àmbit dels satèl·lits i els coets llançadors. La necessitat de posar en òrbita satèl·lits al menor preu possible serà la tendència a llarg termini i calen nous coets que compleixin aquestes demandes. La competitivitat de la indústria aeroespacial creixerà encara més del que és actualment i serà vital trobar noves fórmules per abaratir al màxim el cost de cada missió. En aquest treball es dona una possible solució dissenyant un coet optimitzat per transportar petits satèl·lits a una òrbita baixa LEO. Es farà un estudi del mercat actual per trobar el balanç entre rendiment i cost del model, a m'és de definir la configuració òptima, i se simularà la seva trajectòria amb MATLAB.

El canvi climàtic i els seus efectes és un problema latent a la societat actual. L’efecte hivernacle i l’escalfament global pot portar conseqüències devastadores a la forma de vida actual. No només els humans, sinó totes les espècies presents al planeta patiran aquests efectes si no es troben solucions de forma ràpida. El sector del transport aeri és un dels grans emissors de gasos d’efecte hivernacle, a causa de la quantitat de gent que mou i el seu ràpid desenvolupament. Les tecnologies actuals no son capaces de fer front als reptes que suposa reduir la contaminació en aquest sector. Es per aquest motiu que des de l’any 1997, amb el Protocol de Kyoto, s’ha intentat fixar uns límits per intentar reduir les emissions del món aeronàutic. L’objectiu d’aquest treball és desenvolupar l’estudi dels diferents paràmetres d’un motor turbofan amb l'esperança de reduir les emissions ambientals d’aquest. Per fer-ho, s’elabora un anàlisi dels primers anys de l'aviació i els primers combustibles, passant pels diferents tipus de motors i estudiant de forma exhaustiva el turbofan. A la mateixa vegada, s’investiga sobre els combustibles fòssils actuals utilitzats en aviació, així com altres tipus de combustibles alternatius més ecològics per trobar la solució més adient. En especial, l’hidrogen serà estudiat en profunditat. A partir dels resultats obtinguts, s'intentarà optimitzar els paràmetres del motor amb bescanviadors de calor per descobrir la millor solució possible al problema presentat. Per dur a terme l'anàlisi, s’utilitza l’entorn MATLAB. Aquest projecte acaba amb un anàlisi de l’impacte ambiental que pot tenir aquest tipus de motor tenint en compte tot el procés des de l'extracció de l’hidrogen fins a la seva combustió, fins a arribar a la conclusió que si no es desenvolupen mètodes més ecològics i menys costosos, la implementació d’aquest turbofan no seria possible des d’un punt de vista del medi ambient.

The present study explores the complex dynamics of spacecraft in Earth’s orbit at altitudes ranging up to a few hundred kilometres. Spacecraft have the challenge of navigating through an atmosphere that consists of gases with extremely low densities, leading to the generation of substantial aerodynamic forces, particularly drag. The long-term impacts of these forces give rise to challenges that cannot be analysed with the assumption of the classical continuous gas model used in fluid dynamics. The primary focus of this study revolves around the thermosphere, a critical atmospheric layer that plays a significant role in the operation of objects in Low Earth Orbit (LEO) as well as in the forecast of re-entering objects. The thermosphere is characterised by its relatively low densities and high kinetic temperatures. Fluid dynamics in the low Earth orbit (LEO) altitudes is governed by the (largely unknown) laws of free molecular flow. The investigation of the dynamic consequences of thermospheric interaction is of the highest importance, and under constant efforts for development, as it improves the accuracy of orbital trajectories of space objects. Furthermore, the analysis of these dynamic effects plays a crucial role in the optimisation of spacecraft design, the mitigation of drag forces, and ultimately the improvement of total vehicle efficiency. Historically, computational fluid dynamics (CFD) simulations have been widely utilised as a major approach for the modelling of spaceship drag. Computational Fluid Dynamics (CFD), which involves the numerical solution of governing equations for fluid flow, offers a comprehensive analysis of the flow properties surrounding a spacecraft. However, this methodology frequently requires significant computational resources and time, and yet can not be employed above 100 km above sea level. Therefore, the present thesis aims to investigate alternative approaches based on Statistical Mechanics, specifically focusing on the Direct Si

In this report an extensive explanation of the Final Degree’s Thesis Design of a test bench for solid rocket motors and complete flight rocket vehicles can be found. This project starts by studying different types and configurations of solid rocket motors and test benches in order to to fulfill weight, dimensions and force requirements for the posterior testing. Later, the main parameters of a solid rocket motor and its derivatives are determined so as to set out which of them can be measured. Once all the research segment is completed, the final design of the test bench is presented along with its previous iterations. A structural static analysis through the ANSYS software is conducted to verify and validate the overall test bench. Last but not least, the sensing configuration of the chamber pressure, exterior temperature and solid rocket motor thrust is presented. Finally, a set of conclusive thoughts are extracted from the test bench design, as well as a budget summary and an environmental impact assessment.

Este proyecto tiene como objetivo el diseño de un planeador a escala de juguete. Dicho planeador pretende ser un producto infantil comercializable que pueda acercar el mundo de la aviación al público más infantil. Contará con capacidad de control automático del ángulo de cabeceo, de forma que pueda contrarrestar posibles ráfagas de viento y regular la trayectoria del vuelo. La envergadura máxima posible es de 40 cm, de forma que el usuario pueda lanzarlo con la mano con facilidad. Se ha buscado maximizar el tiempo de vuelo, de forma que sea más atractivo para el cliente. El planeador se ha diseñado mediante métodos semi-empíricos y numéricos, y posteriormente se ha fabricado un prototipo que ha permitido verificar el diseño y detectar posibles correcciones.

This thesis presents a study on super heavy two-stage rockets targeting high Earth orbit, focusing on geostationary and lunar transfer orbits. The work has been done under the supervision of UPC professor Josep Oriol Lizandra since December 2021 to June 2022. The first part, presents the most important theoretical concepts necessary to develop the study. This includes nozzle ideal theory, rocket 2D flight equations, as well as the most common ascent flight strategies. In addition, some essential rocket design guidelines such as the optimum number of stages and propellant types are introduced. This part also explains the different existing engine cycles in terms of turbopump arranging and propellant combinations. In the second part, first of all, the mission is defined and characterised by considering two particular destinations (GEO and the Moon) and setting a general ascent strategy. This strategy consists of using the gravity turn to reach a low circular parking orbit and then perform a second burn to reach HEO. Thereafter, rocket configuration is studied, thus the oxidiser and fuel are selected, the number of engines is set and the general vehicle shape is defined. Then, the engines are preliminary designed using a perfect gas model and based on the SpaceX Raptor engine. A MatLab code has been developed to simulate engine performance and rocket flight, allowing to set the most advantageous nozzle area ratio. The code has allowed to simulate rocket ascent and study the optimal trajectory using the gravity turn. For this purpose, the indicated magnitude of the pitching manoeuvre is sought, which allows gravity to begin to curve the trajectory and the parking orbit reached to be as circular as possible. Afterwards the code has been modified to optimise rocket characteristics. Most of the study is based on SpaceX Starship rocket, which has been taken as reference, nevertheless the thrust to weight ratio and the propellant ratio between the two stages have been varied

En el present estudi, es tracta un anàlisi d’un motor de tipus Turbofan, prenent com a referència el motor GEnX de General Electric, utilitzant dos tipus de model, el de gas ideal caloríficament perfecte i el de gas ideal caloríficament no perfecte, caracteritzat aquest últim per l’ús del coeficient de calor específica a pressió constant que és depenent de la temperatura. Al llarg del projecte es presenten les hipòtesis considerades en cadascun dels models, posant èmfasi en les adaptacions fetes per tal de garantir que les dades inicials siguin les que es disposen de la informació del fabricant, com per exemple, el càlcul realitzat per a trobar la temperatura d’entrada a turbina que coincideix amb el que disposa la informació del Type Certificate Data Sheet del model en qüestió. L’anàlisi realitzat ha estat elaborat mitjançant la programació d’un codi numèric amb el software de Matlab per a cadascun dels models on s’han obtingut els resultats dels paràmetres més importants com ara l’impuls i el consum específic així com les pressions i temperatures de cada etapa del turbofan. D’aquest anàlisi i de la comparació dels resultats dels dos models s’ha pogut extreure que els valors de les primeres etapes del turbofan fins a la turbina coincideixen amb un percentatge d’error gairebé inexistent. Per altra banda a partir de la turbina els resultats comencen a diferir en un percentatge d’error més notable. S’ha programat també els diferents rendiments que caracteritzen un motor d’aquest tipus, el rendiment tèrmic, propulsiu i global validant el percentatge d’error entre els dos models així com els gràfics del cicle termodinàmic per a cada cas. Amb tota la informació generada s’ha pogut determinar la similitud entre ambdós models i la utilitat en les hipòtesis realitzades per al cas de gas ideal caloríficament perfecte. S’ha disposat també d’un estudi per a l’optimització del motor variant quatre paràmetres característics, decidint els valors òptims per millorar les prestacion

Aquest treball tracta principalment sobre l’estudi del llançament d’un coet de combustible sòlid incloent l’ascens a través de l’atmosfera fins a la inserció orbital. Primer de tot, conté un petit estudi sobre coets del mateix tipus de combustible, a més a més d’algunes trajectòries utilitzades i també sobre els diversos sistemes de refrigeració que pot tenir una tovera. A continuació, es comença amb el desenvolupament del treball, passant per la definició del coet que es farà servir, les característiques tant de l’estructura i massa, com del motor, el seu ascens a través de l’atmosfera i finalment la seva inserció a òrbita. Tot això, és estudiat en una simulació creada amb MATLAB, la qual presenta resultats numèrics, gràfics de l’evolució de diferents variables i representacions de trajectòries i òrbites en 3D. Finalment, amb tots els resultats, s’extreuen conclusions sobre la viabilitat de l’ús d’aquests tipus de coets més petits i barats per a la inserció a òrbita baixa terrestre de petites càrregues útils.

El present estudi es centra en realitzar una simulació d’un compressor centrífug instal·lat en un microturbojet Merlin M160G, seguint els passos generalment emprats per aquest tipus d’anàlisis: modelat, mallat, introducció dels paràmetres, simulació i validació dels resultats. El modelat de la geometria ha estat elaborat fent enginyeria inversa des del component fabricat, utilitzant la peça escanejada amb un escàner 3D i reconstruïda utilitzant SolidWorks. La resta dels passos van ser realitzats dins de l’entorn Ansys Student. El resultat de la simulació, malgrat les limitacions marcades per l’ús de versions de programari per estudiant, ha estat un model capaç de predir els paràmetres del flux de sortida, sota diferents règims de funcionament, amb un error mitjà de 12.7% respecte els valors experimentals obtinguts utilitzant un banc de proves disponibles en la universitat. Un valor raonable considerant les aproximacions superficials, de simulació i els errors propis dels instruments de mesura. Addicionalment, aquests resultats han estat utilitzats per introduir altres simulacions derivades, com la simulació estructural, i mostrar la versatilitat de l’ús de mètodes computacionals aplicats a problemes de mecànica de fluids., This study is centered in performing the simulation of centrifugal compressor installed in a micro turbojet Merlin M160G, following the generally accepted procedure of: modeling, meshing, solver setup, simulation and validation of the results. The modeling of the geometry was done using reverse engineering techniques with the 3D scans of the components which served as templates for their recreating using SolidWorks. The rest of the simulation was carried out using Ansys Student. The results of the simulation, given the limitations imposed by the educational versions of the software, were able to predict the output flow parameters, under different power settings, with a mean error of 12.7% in respect to the experimental values obtained using a jet engine test bed available at the university. This value is reasonable, considering the simplifications in the geometry, simulation and the errors in experimental measurement itself. Additionally, the results of this simulation were used to introduce to other simulations, like structural, showing the versatility of computational methods applied to fluid mechanics problems.

En aquest projecte s'estudia i s'avalua el rendiment d'un coet de diverses etapes de propel·lent sòlid durant tot el seu ascens atmosfèric i fins a la seva inserció orbital. Per tant, en primer lloc s'analitza la composició química del fuel i l'oxidant i s'elucida una proposta preliminar d'acord amb els resultats obtinguts del programa CEARUN de la NASA, que empra la termodinàmica química d'equilibri per analitzar problemes de combustions i coets. Posteriorment, es resolen numèricament les equacions diferencials que caracteritzen el moviment del vehicle al llarg de la trajectòria. Per fer l'estudi, s'ha dut a terme diferents simulacions modificant paràmetres de rellevància, com ara la relació d'àrees entre la gola i la sortida de la tovera, el temps d'encesa de cada etapa, la distribució de propel·lent entre fases o la resposta al fet d'incloure una tercera etapa. Finalment, es proposa una alternativa de coet optimitzat, tenint en compte els resultats obtinguts i d'acord amb la inserció orbital proposada., This project studies and evaluates the performance of a multi-stage solid propellant rocket throughout its atmospheric ascent and until its orbital insertion. Therefore, the chemical composition of the fuel and oxidant is first analyzed and, then, a preliminary proposal is elucidated under the results obtained from NASA's CEARUN program, which employs equilibrium chemical thermodynamics to analyze combustion and rocket problems. Subsequently, the differential equations that characterize the movement of the vehicle along the trajectory are solved numerically. To conduct the research, different simulations were executed by modifying parameters of relevance, such as the ratio of areas between the throat and the exit of the nozzle, the ignition time of each stage, the distribution of propellant between phases, or the response of including the third stage. Finally, an optimized rocket alternative is proposed, taking into account the results obtained and following the suggested orbital insertion.

The future plans of the space agencies all across the globe, mainly focus on the advance of the Extravehicular Activities (EVAs) that are orientated in the repair and maintenance of the space vehicles. This thesis presents a brief history of the space suits (EMU, Orlan-M) utilized in space applications and the propulsion units (Hand-Held Maneuvering Unit, Manned Maneuvering Unit and the Simplified Aid for Extravehicular Activity Rescue-SAFER which is currently being used for performing self-rescue maneuvers, in the case the EVA crew member becomes untethered from the orbiter) that have been used during the space activities. After introducing the space environment and the critical conditions that surround the astronaut during these activities, the fundamentals of the Cold Gas Propulsion Systems on which the function of the SAFER propulsion unit is based, are presented. The aim of this thesis is to perform a performance analysis of the SAFER propulsion unit, with the extraction of the basic space propulsion concepts, i.e. thrust, specific impulse, etc. and the verification of the produced results with the technical characteristics of the unit. The orbital dynamics applying on the astronaut during an EVA are also being studied in this thesis, starting from the scenario where the crew member becomes untethered and floats freely in the vacuum of space and at the same time, examining the case where the SAFER propulsion unit is being activated for the performance of the self-rescue maneuver, and the return of the astronaut to the orbiter.

Colloid electrospray thrusters are unique amongst electric thrusters due to their inherent scalability and the possibility of covering a broad range of Isp and thrust. The main research fronts are finding a suitable propellant, improving their mi- crofabrication, and understanding the underlying physics. In this work, a series of experiments have been carried out with a single emitter capillary electrospray with the ionic liquid EMI-BF4 and how the operational variables affect its per- formance has been studied. A breakdown of all the loss mechanisms that make up for the total efficiency of a colloid thruster is also presented. The past work on colloid thrusters was reviewed, along with the basic physics of electrosprays and the experimental techniques used to characterize them. It is observed that for the broad ranges of flow rates covered in the experiments, EMI-BF4 with a capillary emitter always emits a large proportion of ions along with some droplets. At minimum flow rate, the droplets are still present, causing the efficiency to be low at all the flow rates studied. The effect of varying the emit- ter temperature and voltage is studied, concluding that raising their value would increase slightly the current but have little impact on the efficiency and Isp. Des- pite it became clear that EMI-BF4 is not suitable for its use in colloid thrusters, it lead to identifying what are the desired characteristics of a favorable propellant for these kinds of thrusters. These include having electrochemical reaction products that can be drained from the emitter and a very high ion evaporation energy to prevent ion emission. The efficiency of a colloid thruster was analyzed and each loss mechanism was iden- tified. A method to calculate the losses from electrochemical reactions is presented. Identifying each factor of the overall efficiency assists the optimization of the ef- ficiency and the assessment of the suitability of a thruster by studying the main loss mechanisms

The present Bachelor thesis studies, develops and validates the Vortex Cloud Method in a 2-dimensional domain. This method is able to model the flow over a body accounting with flow separation and therefore, overcomes the main limitation of the classical potential models, where flow detachment is not allowed. The Vortex Cloud Method is a meshless model that computes the solution by solving the Navier Stokes Equations through several steps. First, the surface of the evaluated body is discretized in panels and the vorticity of each panel is obtained through potential methods. The next step consists on concentrating the vorticity in discrete vortices and shed them at a certain distance of the surface. After that, the free discrete vortices are convected under the influence of other vortices, the surface vorticity and the free stream. Moreover, viscous diffusion is simulated by a applying a random displacement to the vortices. Finally, the pressure coefficient distribution along the body surface is computed as well as the resulting force coefficients that determine the aerodynamic behaviour of a given body. The theoretical bases of the method and the implementation of each step are extensively explained in this work. Moreover, the implemented model is validated and discussed through the comparison of the results with experimental data.

This study focuses on the aerodynamics of airfoils by matching the inviscid external flow (computed with a panel method) and the viscous flow within the boundary layer next to the wall. For the computational design of the airfoil, MATLAB software is employed., Incoming

Following the near future environmental targets set by the aviation industry, there now exists an increasing amount of research and wish on changing the propulsion system of aircraft. The purpose of this study is to analyze the feasibility of implementing a Hybrid Propulsion System in a light commercial aircraft, taking into account some reasonable requirements to fulfill in a 2030 scenario. Those are especially focused on the available payload, range, weight distribution, and power distribution. To this end, the established preliminary sizing methodology is based on modifying the traditional conventional propulsion approach. Furthermore, the methodology takes into consideration each hybrid aircraft configuration, but this paper is based only on a series architecture in which the full power provided by the fuel source is used to charge the batteries that are the main power source of the engines. The final results provide information about the total power and energy required for a traditional flight performance and, despite they overcome the range expectations, they still do not bring much benefits in comparison with conventional propulsion in terms of range. However, the results show the possibility to perform 2,000 km range saving up to 63 % of the fuel consumed. That if, batteries specific energy density turn out to be a key factor for hybrid propulsion. The study shows the technological shortfall that needs to be developed and looked into make hybrid aircraft feasible by 2030.

Outgoing

The study focuses on the gas expansion through a two-dimensional, axisymmetric nozzle using the method of characteristics for the (assumed inviscid) bulk flow, and matching it to the viscous boundary layer. Nozzles are truncated using the Thrust ideal contour nozzle concept, aiming at estimating the coefficient of thrust, heat losses to obtain an optimised nozzle length. For the computational design of the nozzle, Matlab software is employed., Incoming

In the present time, the study and analysis of the physical phenomena occurring inside a Hall-effect thruster is strictly limited by the problem of the flow of plasma in the enclosed constant-area portion of its annular channel. Several approximations,each of them with their own drawbacks, have been proposed to successfully simulatethe plasma physics inside these ion thrusters. In this thesis, in order to simulate the Hall-effect thruster plasma flow, a one-dimensional hybrid Particle-In-Cell plasma model will be developed. For the validation of the developed model, the already well known plasma ordinary differential equations solution will be used. Finally, in order to improve the results obtained, as a future work, a 2D or even a 3D extrapolation of this validated model could be done and thus, obtaining a better approach of the physics inside the Hall-effect thruster

Dawn mission represents a revolutionary achievement in space research due to the fact that it was the first spacecraft to orbit two different extraterrestrial bodies, the asteroids Vesta and Ceres. It was managed by NASA’s JPL since its design, its launch in 2007 until its end in November 2018. Its objective was accomplished thanks to its Ion Propulsion System (IPS), which made possible the required manoeuvres and trajectories required to fulfil the mission’s objectives. This thesis focuses on the study of this innovative IPS, specifically on how it works during the mission and trying to understand why it operates in one way or another. Specially, the thesis centres in the most characteristic manoeuvre of the entire Dawn mission, the escape trajectory from Vesta to begin cruise to Ceres. In order to do this, the JPL’s SPICE library and toolkit is employed during the whole computations due to its high reliable source of data. From this information provided by SPICE, some calculations are performed regarding the Dawn IPS such as its trajectory, thrust force and even though its input power given by the spacecraft’s solar arrays. One of this thesis’s goals is to prove the reliability of this SPICE given data as well as its programming tools, which is confirmed after analysing the extracted results and comparing them to corroborated information. Furthermore, together with the SPICE computations, the same IPS parameters are estimated employing some theoretical models related to space manoeuvres similar to the ones performed by Dawn. The feasibility of the application of these established procedures is also wanted to be demonstrated in the case of each determined value. After the calculations, it is found out that these methodologies are validated due to the absolute similarity to the results obtained with SPICE. The estimated low thrust climb trajectory results in a truly approximate escape trajectory, and the thrust force value provided by a low thrust escape is incredib

This report outlines the pertinent theory and methods for supersonic nozzle design using the method of characteristics. The programmes developed design ideal nozzle contours, using numerical methods based on the gas properties, for desired exit velocities. Ideal contours include the most rapidexpansion without inducing shock waves followed by the straigh tening sec-tion to provide uniform exit conditions. Then based on the ideal results, thecontours can be truncated and the shape modified to induce minor shocksand only sacrifice minimal thrust to save length and thereby weight.

Aquest treball pretén desenvolupar l'estudi preliminar d'un dispositiu propulsor elèctric de tipus iònic per a un vehicle espacial d'òrbita baixa (a uns 200 km d'altitud), mitjançant el model d'ordre zero de Brophy. A més, com a alternativa als gasos nobles d'alt pes molecular, típicament utilitzats en aquests tipus de propulsors, es valorarà la possibilitat de fer servir com a propulsant una mescla de nitrogen i oxigen, en la proporció en què aquests es troben en aquelles altituds, deixant així (per un altre estudi posterior) oberta la possibilitat de prendre aquests gasos directament de l'exterior.

Present and investigate a methodology for the optimisation of control systems and structural parameters of aeroelastic systems, which enables a rapid acquisition of design criteria with low computational costs., This project presents and investigates a methodology for the optimisation of control systemsand structural parameters of aeroelastic systems, which enables a rapid acquisition of designcriteria with low computational costs.The aerolastic system is modelled via a set of non-linear intrinsic beam equations for thestructure together with a two-dimensional model for the unsteady aerodynamics. The sys-tem’s number of degrees of freedom is truncated by a modal reduction while maintainingthe relevant physical information for the optimisation process. The aerolastic system hasbeen verified through a comparison of the flutter dynamics with benchmark results. Anadjoint sensitivity analysis is then performed to extract the derivatives of a cost function,for instance the strain energy during a gust perturbation, with respect to some paramet-ers of interest such as the weights of a PID-type controller or the wing’s mass distribution,to be applied to a gradient-based optimisation and find a minimum of this objective function.This methodology, which has been implemented with MATLAB, is very flexible and canbe easily applied to different cost functions as well as optimisation paramete

Dissenyar, implementar i testar una tecnologia de boosters, modular, robusta i versàtil, per a coets de modelisme.

Development of a preliminary study of Hall Effect thruster for a spacecraft., Space propulsion, and more specifically electric propulsion, has been growing widely strong for the last 20 years thanks to the increasingly frequent technological advances in this field. That’s why there are plenty of studies arising seeking for new ways of computing simulations of propulsion systems. This is the case of the Particle-in-Cell method. This method is nowhere near new, for it started being used in the 1950s, but the computational advances in the recent years have opened new doors for this numerical method that makes it one of the best options. So, this thesis has the main purpose of proving that Particle-in-Cell simulations of a onedimensional Hall-effect thruster channel give similar results to experimental data obtained via analytical models. For it, first an introduction to different electric propulsion systems and plasma physics will be done in order to have the basis of Hall thrusters functioning so the analysis can be performed. Once the basis are settled, the analytical model will be performed comparing the results with those obtained by professor E. Ahedo and then, using the results obtained in the analytical model as reference, the numerical model using the PIC method will be computed. Finally, by discussing and comparing the two models, we will see that the results obtained in both models resemble a lot those obtained by professor E. Ahedo and the experimental data from real life thrusters, with the only difference that some discrepancies appear. In the case of the analytical model, values in electrons temperature will tend to have a sudden decrease due to not taking into account wall losses and plume divergence, while in the PIC numerical model a little oscillation appears at the beginning due to the finite difference solver used. Although the appearance of such discrepancies, the results obtained are exactly the ones that were expected, hence proving the viability of the PIC model for plasma computation.

Aquest treball consisteix en estudiar la viabilitat tècnica d'un sistema propulsor basat en una presa d'aire, una hèlix multi-pala (o ventilador) impulsada per un motor elèctric, i una tovera convergent, esperant obtenir així un major empenyiment a altes velocitats de vol, respecte al que s'aconseguiria amb una hèlix convencional.S'estudiarà també la possibilitat de disposar una cambra de combustió a continuació del ventilador, utilitzant hidrogen com a combustible, amb la qual cosa s'incrementaria de forma notable l'empenyiment durant determinades fases del vol, tals com l'enlairament i l'ascens.

A Venturi Unmanned Technologies s’està començant la fase de disseny d’un nou producte consistint en un UAV d’ala fixa d’altes prestacions. Per tal de dur a terme aquest procés és essencial treballar amb dissenys preliminars realistes, que permetin optimitzar amb precisió el disseny final. Davant d’aquesta necessitat, en aquest treball es desenvolupa una eina complementària al programari de disseny ja utilitzat per l’empresa, basada en mètodes experimentals, que permet obtenir les característiques aerodinàmiques i de consum d’un avió durant el vol. Per realitzar aquest sistema, en el projecte es construeix un avió radio-control amb mètodes de prototipatge ràpid per tal d’obtenir geometries precises facilitant-ne la possibilitat de variacions, s’implementa una controladora de vol per fer el mostratge de dades a altituds i velocitats constants i tota l’electrònica necessària per mesurar els diferents coeficients aerodinàmics i consums durant l’operació de l’avió.

Desenvolupar l'estudi preliminar d'un dispositiu propulsor elèctric tipus Magnetoplasmadinàmic (MPD) per un vehicle espacial, mitjançant un model quasi unidimensional, basat en les equacions de la dinàmica de plasmes magnetitzats., Magnetoplasmadynamic (MPD) thrusters are an electrical propulsion system capable of accelerating plasma to high exhaust velocities using self-magnetic field. The characteristics and technology of these thrusters make them a suitable candidate for deep space of far-term human exploration of Mars for example. However, they are still not yet flight operational because of their electric power consumption, which are of the order of kW and MW. The purpose of this study is to perform a preliminary analysis of a MPD thruster for a spacecraft, through a quasi one-dimensional, self-field accelerated plasma flow model, with zero axial current, base on the magnetized plasma dynamic equations. To do so, first it would be required and justify why the MPD is a capable propulsion system. Next, using analytical and numerical models it would be studied the magnetized plasma dynamic and validated by comparing it to contrasted values from experimental using. Throughout this report, good results have been obtained with the MPD one-dimensional model and the possible space mission. Results show that one-dimensional models can be solved by applying Matlab code. Furthermore, it confirms that MPD thrusters perform as expected and are worthy candidate of future space propulsion. However, some conditions and concepts need improvement. The study has achieved its goal to develop a preliminary approach to MPD performance. With a hope that this will allow the reader to undersant the importance of this type of propulsion in a near future.

The world of space market evolves towards low-cost solutions, therefore solar sail propulsion is intended to supersede the chemical and electric propulsion. If in addition we combine the solar sail with the suitable orbit, the overall cost can be reduced.What is more, solar sails are propelled only by solar energy, and the Sun is an inexhaustible source of energy, enabling this type of propulsion to perform unique missions.On the other hand,the use of solar sails in space missions has just started a few years ago, being an innovative system with a lot of benefits.This study provides a detailed explanation of solar sailsand a model that describes and demonstrates its effect.The present study also describes the motion of a spacecraft around the L2 Lagrange point of the Earth-Moon system with its corresponding simulation. The interest of orbiting the Lagrange point has increased in the last years because of their natural properties. Furthermore, models for the dynamics of a solar sail in the Earth-Moon system are not as well studied as the Sun-Earth system. In addition, the majority of the studies done,do not take into account Sun’s gravity, when, in fact, it seems that the Sun’s gravity has a significant impact in the Earth-Moon system.This study starts from the circular restricted Three-body problem Earth, Moon and spacecraft, in particular, the periodical Haloorbits around the L2 Lagrange point, for subsequently analyse the effect of the perturbations of a third body (the Sun) and the solar radiation on the spacecraft movement.With the model described, the study investigates the use of a solar sail system with variable orientation in order that the spacecraft could remain closely to the nominal Halo orbit, and even to close the orbit.

Aquest treball aborda l'estudi preliminar, incloent la viabilitat tècnica i econòmica, sobre el disseny d'un mòdul per a l'Estació Espacial Internacional, així com l'anàlisi de la missió corresponent per al seu acoplament.

Desenvolupar l'estudi preliminar d'un dispositiu propulsor elèctric de tipus iònic per a un vehicle espacial., 1. Introducció.2. Anàlisi de missió, incloent càlcul de trajectòries i/o maniobres orbitals típiques, amb un cert grau de detall.3. Justificació del motor iònic com a sistema propulsor adequat per a la missió en particular, a partir d'una anàlisi comparativa de diferents tipus de propulsors per a vehicles espacials, tot assenyalant els avantatges i inconvenients de cadascun d'ells.4. Breu introducció a la física de plasmes magnetitzats.5. Anàlisi preliminar d'un propulsor iònic, mitjançant mètodes analítics simplificats.6. Anàlisi numèrica mitjançant un model quasi-unidimensional, fent ús de la teoria de Brophy.7. Validació dels resultats, per comparació amb altres models numèrics i/o resultats d'assaigs experimentals.8. Anàlisi d'actuacions per diferents condicions d'operació.9. Conclusions.

Desenvolupar l'estudi preliminar d'un dispositiu propulsor elèctric de tipus Hall per a un vehicle espacial., 1. Introducció.2. Anàlisi de missió, incloent càlcul de trajectòries i/o maniobres orbitals típiques, amb un cert grau de detall.3. Justificació del motor d'efecte Hall com a sistema propulsor adequat per a la missió en particular, a partir d'una anàlisi comparativa de diferents tipus de propulsors per a vehicles espacials, tot assenyalant els avantatges i inconvenients de cadascun d'ells.4. Breu introducció a la física de plasmes magnetitzats.5. Anàlisi preliminar d'un propulsor d'efecte Hall, mitjançant mètodes analítics simplificats.6. Anàlisi numèrica mitjançant un model quasi-unidimensional.7. Validació dels resultats, per comparació amb altres models numèrics i/o resultats d'assaigs experimentals.8. Anàlisi d'actuacions per diferents condicions d'operació.9. Conclusions.

Scramjet engines are the fastest form of propulsion that has been conceived, being capable offlying in a flight regime above Mach number 5. The concept of scramjet engine is not recent,as its development started more than half a century ago. However, a scramjet vehicle has onlyflown three times (by NASA), being the last one in 2004 and reaching a Mach number of 10.The working principle of a scramjet engine is the same as the subsonic engines: compressing,combusting and expanding air to produce thrust. Nevertheless, scramjet engines do not havemoving parts such a compressor or a turbine, as the flow at hypersonic speed is compresseddirectly by the inlet and the combustion is held in supersonic regime.In this study, a parametric model for the combustor and nozzle of a scramjet engine has beendeveloped, focusing in the combustion process. Under quasi-one-dimensional steady flowconditions, the properties at the exit of the combustor and nozzle are computed by calculatingthe composition of the gas at the combustor for different conditions of temperature, fuel ratio,pressure and velocity.Also, in order to assess the performance of the engine, a simplified inlet model is assumed.Finally, the results of this document are compared with those of other studies. Some possibleimprovements of the study are also included.

This Final Degree Project has the aim of developing a preliminary study of the design and performance of a NuclearThermal rocket for a Spacecraf

This study aims to develop a preliminary analysis of a magnetoplasmadynamic (MPD)thruster, a device which represents the next generation of space propulsion technologyfor new interstellar missions, for instance cargo and piloted missions to Mars. In order tounderstand the principles that govern this plasma accelerator, a quasi-1D model and aperformance analysis have been carried out. Quantitative results are shown for differentthruster configurations. The constant area geometry has been found to be limited by theeffects of finite Magnetic Reynolds numbers, which generally determine whether the flowthat exits the thruster is sonic or not. The variable area geometry, contrarily, alwaysensures that the flow finishes being supersonic, forRmranging from 4 to 10. With thistheoretical model, the performance has resulted in nearly twice the common values forMPD thrusters tested in laboratories.

Es planteja estudiar la viabilitat d’un planador no tripulat (“drone”) capaç de retornar càrregues útils des de l’estratosfera (20-25km d’alçaria), partint des d’un globus d’heli no recuperable. El vehicle ha poder caure i retornar a la base, o a un dels waypoints pre-programats, de forma autònoma i sense propulsió de cap mena.En el treball, s’estudiarà la configuració més adient i es faran simulacions del seu comportament, per saber en quines condicions podria operar. S’estudiaran diferents opcions per estabilitzar l’actitud del vehicle mentre estigui creuant zones de baixa densitat atmosfèrica, i es farà un disseny previ de l’estructura i del sistema de control., Estudi de l'estat de l'art i aplicacions-Disseny conceptual-Desenvolupament d'un model-Simulacions-Estudi de viabilitat tècnica

Aquest treball consisteix en elaborar un programa informàtic eficient i robust que calculi la trajectòria de vol que ha de seguir un avió comercial en la fase creuer, per tal que aquesta trajectòria sigui òptima, d'acord amb un dels dos criteris alternatius següents: 1) mínim temps de vol, 2) mínim consum de combustible.L'optimització de la trajectòria tindrà en compte les següents consideracions:1. Moviment bidimensional (altitud de vol constant).2. Curvatura de la terra apreciable.3. Intensitat i direcció del vent conegudes al llarg de zona (geogràfica) considerada.4. Algoritme robust, fiable, i que requereixi un temps de càlcul baix-moderat (per possible aplicació a navegació en temps real).

1. Introducció.2. Anàlisi de missió.3. Elecció d'un sistema propulsor per la missió en particular, a partir d'una anàlisi comparativa de diferents tipus de motor coet (avantatges i inconvenients de cadascun). 4. Anàlisi preliminar del sistema propulsor elegit, incloent càlculs numèrics, basats en models matemàtics.5. Conclusions., Desenvolupar l'estudi preliminar d'un coet per un llançador de micro-satèl·lits.

Aquest treball pren com a punt de partida el problema circular restringit de tres cossos Terra, Lluna i vehicle espacial, en particular, les òrbites periòdiques (Halo) al voltant del punt de Lagrange L2, per, posteriorment, analitzar l'efecte de les pertorbacions d'un tercer cos (el Sol) i la radiació solar sobre el moviment del vehicle espacial. que fa que, amb rigor, les òrbites Halo al voltant de L2 no isguin possibles.Mitjançant l'anàlisi de sistemes dinàmics i teoria de control, el treball investiga l'ús de sistemes propulsius de baix empenyiment, combinats possiblement amb un sistema de vela solar per tal que el vehicle pugui romandre el més a prop possible de l'òrbita Halo nominal., 1. Introducció del problema de tres cossos circular restringit: sistema sinòdic, punts d'equilibri (o punts de Lagrange) i estabilitat d'aquests, i trajectòries periòdiques.2. Obtenció de les trajectòries periòdiques al voltant de L2 en el sistema Terra-Lluna.3. Anàlisi dels sistemes propulsius de baix empenyiment i alt impuls específic.4. Anàlisi d'una vela solar: pressió solar, forces sobre la vela en funció de l'orientació de la vela.5. Estudi de la dinàmica del vehicle espacial sotmès a les pertorbacions degudes al Sol (tercer cos) i a la radiació solar.6. Anàlisi de lleis de control òptimes (en termes de minimitzar el consum de combustible) que mantenen el vehicle en les proximitats de l'òrbita Halo nominal.

A computationally efficient, quasi-one-dimensional, supersonic combustion model has been developed to simulate high-speed engine flowfields such as in a scramjet. The model can be used in design and optimization studies, where a wide range of engine inlet conditions may be encountered. The model solves a series of ordinary differential equations that come from the equations of fluid motion coupled with combustion. The effects of area change, wall skin friction, mass injection, fuel mixing, and heat transfer are included. The equations are derived assuming an open thermodynamic system with equilibrium or with finite-rate chemistry. Results will show the differences on combustor performance between considering equilibrium or not. The inclusion of chemical kinetics allows for a prediction of fuel ignition, a finite rate process that inherently cannot be predicted using equilibrium methods. Therefore, incorporating the timescales of fuel mixing and ignition is crucial for an accurate prediction of combustor performance in most scramjet operating conditions. Furthermore, the effects of dissociation, fuel mixing and friction are discussed. Eventually, a simple analysis of the air inlet and the nozzle is developed in order to have an overall vision of the scramjet and its feasibility