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Els sistemes de propulsió eficients són crucials per a l'exploració espacial, ja que determinen l'abast de les missions, les capacitats de càrrega útil i la durada dels viatges. A mesura que la humanitat té com a objectiu explorar més endins el cosmos, la necessitat d'innovacions de propulsió es fa cada cop més vital. La recerca de sistemes de propulsió més eficients implica un avenç significatiu de les tecnologies actuals. Els broquets reals utilitzats en missions pateixen deficiències a diferents altituds a causa de la seva pressió de sortida fixa, la qual cosa resulta en una expansió no òptima. En canvi, els aerospikes ajusten dinàmicament el seu flux d'escapament, mantenint una empenta òptima a través de diferents pressions externes, proporcionant així capacitats de compensació d'altitud. L'objectiu d'aquest estudi és dimensionar un aerospike utilitzant la teoria matemàtica actual, realitzar una anàlisi d'independència de la malla i realitzar simulacions CFD per extreure dades sobre les seves capacitats., Los sistemas de propulsión eficientes son cruciales para la exploración espacial, ya que determinan el alcance de las misiones, la capacidad de la carga útil y la duración de los viajes. A medida que la humanidad se adentra en el cosmos, la necesidad de innovaciones en materia de propulsión se hace cada vez más imperiosa. La búsqueda de sistemas de propulsión más eficientes supone un avance significativo respecto a las tecnologías actuales. Las toberas actuales empleadas en las misiones sufren deficiencias a distintas altitudes debido a su presión de salida fija, que provoca una expansión no óptima. Por el contrario, los aerospikes ajustan dinámicamente su flujo de escape, manteniendo un empuje óptimo a través de diferentes presiones externas, proporcionando así capacidades de compensación de altitud. El objetivo de este estudio es dimensionar un aerospike utilizando la teoría matemática actual, llevar a cabo un análisis de independencia de malla y realizar simulaciones CFD para extraer datos sobre sus capacidades., Efficient propulsion systems are crucial for space exploration as they determine the reach of missions, payload capabilities, and travel durations. As humanity aims to explore deeper into the cosmos, the need for propulsion innovations becomes increasingly vital. The quest for more efficient propulsion systems involves a significant advancement from current technologies. Actual nozzles employed in missions suffer deficiencies across different altitudes because of their fixed exit pressure, which results in non-optimal expansion. In contrast, aerospikes dynamically adjust their exhaust flow, maintaining optimal thrust through different external pressures, thus providing altitude compensation abilities. The purpose of this study is to dimension an aerospike using current mathematical theory, conduct a mesh independency analysis, and perform CFD simulations to extract data about its capabilities.

Aquest treball consta de l’estudi dels requisits, el disseny preliminar d’un motor híbrid de coet i la planificació i programació per tal de que l’associació Cosmic Research pugui participar a la competició europea de coeteria universitària EuRoC el 2023. L’objectiu és dissenyar un motor de propulsió química que permeti arribar a un apogeu de 9000 m, portant una càrrega de pagament d’1 kg, i que utilitzi un propergol no-tòxic i apte per ser manipulat per estudiants. El disseny s’inicia amb l’anàlisi de diferents combinacions de propergol per motors líquids i híbrids, que conclou amb la decisió de dissenyar un motor híbrid que utilitzi N2O com a oxidant i parafina com a combustible. Un cop determinat el propergol, es procedeix al dimensionat de la tovera, la cambra de combustió, el gra i el dipòsit d’oxidant. Utilitzant una tovera supersònica i un gra amb geometria anular per tal de mantenir una corba d’empenta neutre, s’obté que, per arribar a un apogeu de 9000m el motor ha de cremar 16,63 kg de propergol en 6,4 s, produint una empenta de 5745 N. Finalment, es desenvolupa un pla de fabricació, un pla d’assajos i una planificació i programació del disseny fabricació, assaig i implementació final del motor, per analitzar la viabilitat de que l’associació d’estudiants pugui desenvolupar el motor i operar-lo a la competició EuRoC l’octubre de 2023., This final degree project consists of the study of the requirements, the preliminary design of a hybrid rocket engine and the planning to ensure that the students association Cosmic Research can participate in the European rocket competition EuRoC in 2023. The aim of this study is to design a chemical propulsion engine capable of achieving an apogee of 9000 m, carrying a payload of 1 kg, using a non-toxic propellant that is suitable for handling by students. The design begins with the analysis of different propellant combinations for liquid and hybrid engines, which concludes with the decision to design a hybrid engine that uses N2O as the oxidizer and paraffin as the fuel. Once the propellant is determined, the nozzle, the combustion chamber, the grain and the oxidizer tank are dimensioned. Using a supersonic nozzle and a grain with annular geometry in order to maintain a neutral thrust curve, it is obtained that, to reach an altitude of 9000m the engine must burn 16,63 kg of propellant in 6,4 s, producing a thrust of 5745 N. Finally, a manufacturing plan, a test plan and a planning of the design, manufacture, testing and final implementation of the engine are developed, to analyse the feasibility of the student association being able to develop the engine and operate it in the EuRoC competition in October 2023.

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.

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

Rocket fuel injectors are in charge of propellant atomization and mixing processes, which determine the performance and stability of the liquid rocket engine itself. This small, yet significant components, are in charge of reducing the size and weight of the propellant droplets while mixing them in the shortest possible amount of time. The injector’s design and implementation must be suitable for performing their tasks in order to achieve maximum performance of the nozzle. Aiming to design and test a feasible rocket injector, a conceptual and physical design is presented, using cheap off-the-shelf stream jet nozzles, and its basic behaviour assessed. The numerical approach involves using computational fluid dynamics software (OpenFOAM) while the experimental approach involves lathe machining considerations. The numerical results indicate that the doublet injector design is feasible and mixture quality can be attained but still needs to be crafted and tested using a particular jet nozzle configuration.

The motivation of this thesis is to solve the two main problems of space exploration: the lack of fuel to perform space exploration beyond the Moon or Mars, and the limitations generated by the launching of large amounts of fuel from Earth to perform long distance missions. In this thesis we study the possibility of using lunar resources to produce fuel from them and not needing the raw material from Earth, the location of these resources on the Moon and the different methods to extract them. In order to favor a reliable study, a propulsive characterization test is carried out with a fuel composed only of resources available on the Moon: aluminum and oxygen. For this test, a hybrid rocket will be designed and manufactured to test the fuel. The results of this test will be used to study the effectiveness of this combination of lunar materials as a propellant. Obtaining feasible results in the test would be a milestone in space research, since it is estimated that the use of lunar resources for fuel production would drastically reduce the cost of space missions, thus allowing the possibility of solar system exploration and a future space economy. The fact of being able to obtain fuel from the Moon would expand our frontiers since it opens the door to the idea of being able to obtain it also from different bodies of the solar system. Thus, this thesis is a good starting point for future research on obtaining rocket fuel from the raw materials of other planets or asteroids, and consequently, to the long-term goal of establishing a sustainable presence on the Moon

This thesis compiles the process of upgrading two libraries (INIST and HGS) which provide to their users the thermodynamics properties and algorithms needed to analyse liquid rocket engine feed cycles. First, the Thermodynamic theory for real fluids, ideal fluids and mixtures is reviewed, as well as the algorithms and tools that the libraries implement. For ideal gases, the NASA 7 coefficient polynomials are used while for real fluids, data interpolated from NIST is applied. Then, five rocket engine feed cycles analyses are carried out using the libraries. Turbomachinery (pumps, turbines) as well as combustion chamber, nozzles, mixers and heat exchangers are studied with a thermodynamic analysis. The obtained results show good agrement with data obtained from the literature. Finally, some additional utilities of the potential of the library are given: (i) a chemical equilibrium algorithm based on the direct minimisation of Gibbs free energy that is used to obtain the combustion chamber mixture of dissociated species; (ii) a surrogate of RP1 for the simulation of RP1-LOX rocket cycles; (iii) a study on the effect of using ideal gas versus real fluid properties

State of the art rocket engines are reaching performance levels that have become limited by the available manufacturing techniques. With a great focus on these engines’ efficiency, rocket engine producers are focusing on increasing combustion chamber pressures and temperatures, which in turn produces increasingly greater structural and thermal loads. Complex cooling systems are usually used to tackle this issue, being the most predominant formula to use the stored cryogenic propellant as a coolant before being injected in the combustion pressure, more commonly known as regenerative cooling. However, these cooling systems are limited by the state of the art manufacturing techniques. To deal with this issue, efforts have been made to implement modern metal Additive Manufacturing (AM) techniques to rocket engine production. In order to show the extent of the benefits of such innovations, a rocket engine nozzle has been designed following the characteristics of SpaceX’s latest Raptor engines. This design has been tested through Computational Fluid Dynamics (CFD) and Finite Element Method (FEM) simulations so as to validate its performance parameters and study how the implementation of metal AM techniques would improve the engine’s performance. The economic feasibility of such implementation has been studied through the comparison to former techniques, as well as their impact to the environment

Most traditional simulators used in amateur experimental rocketry make use of basic equations and zero-dimensional models to recreate the performance of solid propellant rocket motors. The usefulness of these simulators is limited, and tend to fail with scale. A new simulator based around idealized and nonidealized phenomena is described in this work, along with the theory used to build it. This simulator considers the motor as a one-dimensional body and discretizes the propellant and chamber along the motor center axis in multiple finite volumes. The new simulator, dubbed PropSim yields results that, although not exactly equal to the real results from a static test, do present and take into account cases and circumstances overlooked by traditional simulators, correcting some of their shortcomings

This project consists in the definition of the technical requirements, the design and the development of a solid rocket engine. This rocket engine must fulfil a specific mission: to propel a 0.3 kg heavy payload to a height of 100 km, while following a parabolic trajectory, in a suborbital flight. The payload could make use of the very unique conditions experienced in a suborbital flight; namely, the microgravity conditions for a certain amount of time. The rocket would take-off from a launch platform lifted up to a height of 30 km, by means of an aerostatic balloon. The engine must fulfil 3 main properties: safety, cost and robustness. It shall be safe, for anyone working on any of its associated processes: the manufacturing of its structure, the processing of the propellant, the integration of all elements, the transportation, etc. The cost of the engine shall be as low as possible, in order to pave the way for students, universities or research centres with low funds to use this launch system. Besides, the engine must be robust enough; that is, it must be able to work as expected at nominal conditions but also at slightly adverse conditions. As has already been pointed out, the project is structured in 3 phases: definition, design and development. The first phase is based on research, and it must allow detailed acknowledgement on how a solid rocket engine works, thus allowing to define the technical requirements. The design phase consists on establishing technological solutions for the requirements, while fulfilling the 3 properties of the rocket engine that have already been mentioned. Finally, the development phase is about testing some critical elements of the engine, in order to validate, or either invalidate and thus improve, the design.

This thesis has the intention to conduct a study and design of a solid rocket motor and its propellant, for a later implementation, construction and testing of it in a test stand. Thestudy anddesign of the motor will be carried out using as outlines the requirements set either by the legislation or the ones imposed by the objectives of the study. These requirements, together with a bibliographic search for the state of the art and the expressions needed to compute the parameters defining the motor, and the use of adequate software will define the expected performance of the motor.The implementation and construction of the motor itself will be carried out by the thesis’ director, as he has the knowledge on how to operate the machinery to obtain the parts conforming the motor.Finally, the testing of the motor will have to be done by using a test stand, the design and construction of which arethe final master’s thesis of another student. During the test of the motor, the thrust and other parameters will be measured and compared with the calculated onesduring the design phase.

Els sistemes propulsius de combustible sòlid són sistemes simples en comparació a aquells de combustible líquid ja què l’emmagatzematge de combustible comparteix espai físic amb la cambra de combustió. Una ànima central en la càrrega de combustible genera el volum de la cambra de combustible i aquest crema en direcció perpendicular a la superfície de combustió. Donada una geometria de l’anima central es determinaria l’avanç de flama en diferents instants de temps. L’evolució en el temps de la superfície de combustió és determinant en la pressió a la cambra de combustió i en la llei d’empenta. En el treball es determinaria l’evolució de la pressió a la cambra i de la llei d’empenta. A més, la presència de bombolles o esquerdes en el gra provoquen alteracions durant la combustió que seria necessari simular.

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

Understand the different methods in use to feed both fuel and oxidiser into a rocket engine combustion chamber, as well as too refrigerate the nozzle. Understand the properties of the cryogenic fluids involved. Implement a model of an engine, The present Master Thesis discusses the efforts of the author to perform thethermodynamic analysis of two modern cryogenic rocket engines. The first one, withthe name SE21D, will be used as validation case to test the accuracy of the methodologydeveloped. Once it has been approved this simulation, the next step will consist in applythis scheme to the study case, called LE9.The special characteristic that distinguish these two models is the employment offeeding system type: Expander Bleed Cycle. They will be studied all the mechanicalcomponents that are involved in this feeding system, which includes pumps, turbines,mixer and valves. Secondary, the combustion process as well as the engine performanceare assessed. Finally, the heat exchange and regenerative cooling that occurs in theCombustion Chamber walls is modeled. The final results are accurate enough once thevalidity of this thermodynamic analysis model has been approved and this Master Thesisalso leaves room to future researchers to continue with this project.This simulation have been performed by means of an intensive bibliographic researchand the employment of MATLAB software and two additional packages, created in-housePolytechnic University of Catalonia and already validated, called HGS and INIST. Thislast one collects the fluid properties for the most common substances used as rocketpropellants and it was used to define the state of the fuel and oxidizer at each stage of thefeeding system. The latter, however, it is a chemical solver based on NASA’s polynomialsand it was utilized mainly in the combustion process that takes places in the CombustionChamber of both rocket engines.Nevertheless, the features of these two additional software packages were not enoughto accomplish the goal of this Master Thesis. For that reason, the author performedmaintenance and updating tasks of INIST package. The capacity of adding new data wasrestored as well as the inclusion of new fluid properties that complements the exist

Aiming to be the first step towards the design of a fully operational stabilization system for rockets, thebuilding of a model is done representing the attitudeof a rocket with anactive stability system. Several tests are done with the model both static and dynamic with different type of sensors that give us a view of the response of the system to different situations. Finally, it is described how the model could be upgradedand which steps follow to get a reliable stabilization system.

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.

This project presents the development of a numerical code aimed to perform the grain burnback analysis and predict the performance of a solid rocket motor. The Level Set Method (LSM) has been applied to calculate the grain burnback evolution of 2D and 3D grain configurations introduced by the user as an input. In addition, in order to implement the 3D grain burnback analysis, an innovative approach called Quasi-3D has been implemented towards reducing the computational time. The Quasi-3D approach consists of modeling the 3D grain surface through few 2D reference cross-sections, which are propagated using the LSM in 2D. Then, the 3D surface is reconstructed by interpolating the grain shape between the reference cross-sections. With respect to the internal ballistics simulation, 0D-unsteady and 1D quasisteady models found from the literature have been used. This way, the Matlab® code developed studies three different cases: 2D grain burnback and 0D unsteady flow, Quasi-3D grain burnback and 0D unsteady flow and Quasi-3D grain burnback and 1D quasi-steady flow. The results from 2D and Quasi-3D grain burnback and 0D unsteady flow tools have been compared against three reference cases, obtaining a high level of matching. The 1D quasi-steady flow model results are qualitative.

The accumulation of alumina is an inherent problem in Solid Rocket Motors (SRM). In the V on Karman Institute this problem have been studied using a scale model of Ariane 5 booster where alumina is modeled with water. The Stokes number is maintained. After a parametric test, the relevant parameter found to be responsible for the entrapment of particles, was the horizontal distance between the tip of the isolation ring (inhibitor) and the nozzle tip. This parameter is named as Obstacle tip to Nozzle tip (OT2NT). The problem found during Tóth investigation was that a large part of the water accumulation in the cavity was due to the dripping of the inhibitor. The rst objective of the experimental part have been to reduce the dripping modifying the inhibitor design. The design found to minimize the dripping is the one with no extra absorbing material and a wider top plate (24mm). A posteriori, a parametric tests was done. OT2NT has been con rmed to be the main responsible while the entrapment for the new system is around a half of Tóth results. The aim of the numerical part have been to apply an existent code, Open-Foam twoPhaseEulerFoam library, based on the nite volumes. Then this code have been modi ed in order to handle with ve dispersed phases with different properties, named as sixP haseEulerFoam. The codes have been verificated through a direct comparison with existing publications. Then a sensitivity study to numerical parameters has been done. For twoPhaseEulerFoam it has been found than the results are close to the existing ones while for the sixP haseEulerFoam although the behavior of the solution is the one expected, its values are overestimated giving only a qualitatively valid answer to the problem. Finally, twophaseEulerFoam is applied to the booster geometry studied in the experimental part. Due to the fact that OT2NT has been determined as the main responsible for the entrapment, three different geometries showing three different OT2NT (25:69, -4:31 and, Outgoing

According to the concept of a fluid-bipropellant rocket engine proposed in Design of a rocket engine fed with oxygen and methane in gaseous phase for educational purposes [TR01], a preliminary test stand design for the prediction and validation of the main performances of this engine has been developed. The main sensors and transducers are selected to measure the stated performances. After the proposed instrumentation, a virtual panel control has been also carried out.