<p>A series of experimental tests were performed to study the transient response of gas/liquid injectors exposed to transverse detonation waves. A total of four acrylic injectors were tested to compare the response between gas/liquid and liquid only injectors, as well as compare the role of various geometric features of the notional injector design. Detonation waves are produced through the combustion of ethylene and oxygen, at conditions to produce average wave pressures between 128 and 199 psi. The injectors utilize water and nitrogen to simulate the injection of liquid and gaseous propellants respectively. Quantification of injector refill times was possible through the use of a high-speed camera recording at a frame rate of 460,000 frames per second. High frequency pressure measurements in both the gaseous and liquid manifolds allow for quantification of the temporal pressure response of the injectors. Variations in simulant mass flow rates, measured through the use of sonic nozzles and cavitating venturis, produce pressure drops up to 262 psi across the injector. Injector refill times are found to be a strong function of the impulse delivered across the injector. Manifold acoustics were found to play a large role in injector response as manifolds that promote manifold over-pressurizations during the injector recovery period recover quicker than designs that limit this response.</p>

Electrically controllable solid propellants are an area of interest as a viable solution to the lack of throttle-ability in solid propellant rocket motors. Existing studies have focused on propellants compositions using hydroxyl-ammonium nitrate, ammonium nitrate, or lithium perchlorate as oxidizers. Additionally, the thermochemical and electrochemical reaction mechanisms have not yet been fully defined. The research in this thesis explores the nitrate and perchlorate oxidizer families to compare their cation-anion relationships. Using these oxidizers, pseudo electrically controllable solid propellant compositions were created with the addition of multi-wall carbon nanotubes to enhance ohmic heating capabilities. These additives were selected based on theory that with a non-complexing polymer, an oxidizer melt layer is required for ions to dissociate and electrically controlled ignition to occur. Using an applied voltage, ignition delay and current draw experiments were performed to expand on prior findings that ignition delay follows oxidizer melt temperature while mobility is associated with the size of the ionic radii. Additionally, neat oxidizer pellets were electrically decomposed to determine their linear regression rate. These results help to characterize the mechanism of reaction. This advances the knowledge of oxidizers in electrically controllable applications. / Master of Science / Solid propellant rocket motors have been extensively studied and used in both space and military applications because they do not use air as the source of oxygen. Their main limitation is the lack of throttle-ability, or inability to control propellant burning. This is because solid propellants, which are generally composed of an ionic oxidizer salt, a polymer fuel, and additives, are pre-combined and stored within the rocket motor. An emerging viable solution to this limitation is electrically controllable solid propellants. With an applied voltage, the oxidizer is heated and melts, allowing ions to dissociate and current to flow between electrodes. This reaction can then be controlled by turning the power supply on and off. Cations, or ions which have a net positive charge, move to the negatively charged cathode while anions, which have a net negative charge, move to the negatively charged anode. The research in this thesis explores different cation-anion oxidizer pairings using both a propellant composition and as a pure oxidizer under an applied voltage. The results help to characterize the mechanism of reaction of each oxidizer in an electrically controllable context and determine their effectiveness in these propellant applications.

Increasing interest in the development of nontoxic monopropellants for the replacement of hydrazine and its derivatives stems from the desire for safer and thus more cost-effective alternatives. Ionic liquid monopropellants based on the hydroxylammonium nitrate and ammonium dinitramide ionic oxidizer salts have received the majority of attention over the last two decades and present a promising alternative with higher performance and more attractive handling qualities than hydrazine. These monopropellants are employed using catalytic methods which lead to their decomposition and ignition. However, the development of compatible catalysts remains a limiting step in the technological readiness of these alternative monopropellants. Due to their ionic nature, the development of ionic liquid monopropellants has led to many investigations on the utilization of electrolysis to achieve combustion. Separately, there has been a longtime interest in the use of gelled propellants for enhanced handling and operating safety. Atomization and combustion inefficiencies associated with gels have continued to limit their use. Monopropellants composed of gel polymer electrolytes present a unique opportunity which combines the safety features of gelled propellants as well as the ionic conductivity seen in ionic liquids, allowing them to decompose and ignite electrolytically. In this research, a family of electrically controlled monopropellants that utilize electrolysis in this fashion was developed from a gel polymer electrolyte. Their fundamental properties, including those pertaining to rheology, conductivity, thermal stability, and combustion, are explored as the composition of the oxidizer salt is varied. / Master of Science / Current advancements in rocket propulsion include interests in developing alternative green propellants for use in spacecraft propulsion systems with the hope of replacing current options which may be toxic to handle and present a serious safety hazard. Alternative propellants are generally thought of as not requiring special safety equipment or protocols in their handling, thereby reducing costs. Several promising options belonging to a category of propellants known as ionic liquids have made significant progress in development since the 1990s and have the potential to be used alongside a novel electrical combustion method known as electrolysis. Gelled propellants are another possible alternative which have been researched for their appealing safety qualities for some time. While not researched for their use as rocket propellants until very recently, gel polymer electrolytes have received interest in this application due to their composition which includes a polymer, commonly used as rocket fuel, and an oxidizer salt. Due to their inherent electrical conductivity, their potential to use electrolysis in a similar manner to ionic liquids to achieve combustion is of interest. The research detailed in this thesis was completed to characterize fundamental material and combustion properties of a gel polymer electrolyte propellant as its oxidizer constituents are varied.

The objective of this master thesis was to investigate an ablative liner for the T-Minus DART booster that will accelerate a dart to Mach 5.2 within five seconds. An oxyacetylene torch test was used to sort out the obviously bad materials. Glass fiber/epoxy, with and without alumina as fire retardant, and carbon fiber/epoxy were selected for further investigation. A sub-scale motor was built to expose the materials for conditions similar to the booster conditions in regard to temperature, chemistry, flow velocity and pressure. The target pressure could not be reached in the sub-scale motor but a polynomial function was fitted to the data in order to extrapolate the data and estimate the ablation rate at 7 MPa. The final design is always based on measurements on full scale motors. This could not be done within this report. Recommendation for future work is to use an insulator of 1.8 mm of carbon fiber/epoxy or 1.3 mm of glass fiber/epoxy/alumina for the sub-scale firings to come.

A technical project alongside the University courses can deepen the understanding and increase the motivation for the subject of choice. As a student, there is often a hurdle to start such a project because of a lack of inspiration. And even after overcoming this, the costs associated with such a project may put students off. With my project I show how a 3rd semester Mechanical Engineering student can design and manufacture a rocket engine with all testing components on a student budget. Cost structure and resource planning are explained in detail. I launched the project in December 2020 and in September 2021 it was presented at the StuFoExpo21. A general curiosity for the topic and a basic understanding of mechanical engineering was sufficient for starting the project. Importantly, I gained the most valuable knowledge during the implementation of the project, through active failure-iteration and reading specialised literature. The project is focussed on the design and manufacturing of a rocket engine and its testing components. A special feature is the cooling jacket of the combustion chamber. It has been 3D printed in the SLUB Makerspace, a facility at TU Dresden. Further work packages of the project were the programming of sensors and control systems, first open-air combustion tests of the injector head, safety checks and a Risk & Safety analysis. The first testing and other preliminary work were performed in collaboration with fellow students. During the entire design and manufacturing process I was in continuous exchange with the research group “Space Transportation” of the Institute of Aerospace Engineering at TU Dresden. Special thanks go to Dipl.-Ing. Jan Sieder-Katzmann and Dipl.-Ing. Maximilian Buchholz for their help during this process. For 2022 I plan a test campaign of the rocket engine to collect sensor data and to perform engine thrust measurements. / Ein selbständiges, fachbezogenes Projekt neben dem eigentlichen Studium kann die Kenntnisse vertiefen und die Motivation für das Studienfach fördern. Oft jedoch fehlt Studenten die Inspiration für ein solches Projekt. Und selbst wenn diese Hürde genommen ist, schrecken die Kosten davon ab zu beginnen. Am Beispiel meines Projektes im Bereich Luft- und Raumfahrt zeige ich, wie man als Maschinenbaustudent im 3. Semester einen Raketenantrieb und alle Testkomponenten mit einem Studentenbudget entwerfen und herstellen kann. Kostenstruktur und Ressourcenplanung werden im Detail erläutert. Das Projekt startete ich im Dezember 2020. Im September 2021 wurde es auf der StuFoExpo21 vorgestellt. Für den Beginn reichten Neugier und ein Grundlagenverständnis im Maschinenbau aus. Die meisten und wichtigsten Wissensbausteine erlernte ich während des Projektes durch aktive Fehleriteration und aus der Fachliteratur. Das Projekt umfasst das Design und die Fertigung eines Raketenantriebs und der dazugehörigen Testkomponenten. Eine Besonderheit, die Kühlkammer-Ummantelung des Triebwerks, wurde unter Nutzung der Ressourcen an der TU Dresden, SLUB, mit einem 3D-Drucker hergestellt. Weitere wichtige Schritte waren die Programmierung der Sensorik und der Steuerungseinheiten für den Test, ein erster offener Injektor Test mit Treibstoffverbrennung, Sicherheitstests und eine „Risk & Safety Analysis“. Tests und Vorbereitungsarbeiten erfolgten in Zusammenarbeit mit Kommilitonen - der gesamte Entwicklungsprozess fand in ständigem Austausch mit der Forschungsgruppe „Raumtransportsysteme“ des Instituts für Luft- und Raumfahrt in Dresden statt. Besonderer Dank gilt Dipl.-Ing. Jan Sieder-Katzmann und Dipl.-Ing. Maximilian Buchholz für ihre wertvollen Hinweise. Für 2022 ist eine Testkampagne des Raketentriebwerks geplant, um Sensorwerte aufzuzeichnen und den Schub des Triebwerks zu messen.

<p> Lifted non-premixed turbulent jet flames in the Transverse Instability Combustor (TIC) have been analyzed using qualitative and quantitative methods. Lifted flames in the TIC have been observed to stabilize about zero to five injector exit diameters downstream of the dump plane into the chamber and exhibit pulsating, unsteady burning. Anchored flames immediately begin reacting in the injector recess and burn evenly in a uniform jet from the injector exit through the entire optically accessible region. Statistically significant, repeatable behavior lifted flames are observed. It is shown that the occurrence of lifted flames is most likely for an injector configuration with close wall-spacing, second greatest for a configuration with close middle-element spacing, and lowest for a configuration with even element-spacing. For all configurations, of those elements that have been observed to lift, the center element is most likely to lift while the second element from the wall was likely. Flames at the wall elements were never observed to lift. Evidence is shown to support that close injector element spacing and stronger transverse pressure waves aid lateral heat transfer which supports flame stability in the lifted position. It is hypothesized that the stability of lifted flames is influenced by neighboring ignition sources, often a neighboring anchored flame. It is also shown that instances of lifted flames increase with the root-mean-squared magnitude of pressure fluctuation about its mean (P’ RMS) up to a threshold, after which flames stabilize in the anchored recess position.</p> <p>Dynamic mode decomposition (DMD) and proper orthogonal decomposition (POD) analyses of CH* chemiluminescence data is performed. It is found that lateral ignition of the most upstream portion of lifted flames is dominated by the 1W mode. Furthermore, it is shown that low-frequency high energy modes with spatial layers resemble intensity-pulses, possibly attributable to ignition. These modes are trademarks of CH* chemiluminescent intensity data of lifted flames. It was also shown that the residence time in the chamber may be closely associated with those low-frequency modes around 200 Hz. DMD and POD were repeated for a downstream region on the center element, as well as a near-wall element, highlighting differences between the lifted flame dynamics in all three regions. </p> <p>It is shown that lifted flames are best characterized by their burning behavior and in rare cases may stabilize in the recess, while still being “lifted”. Furthermore, it is shown that flame position differentiation can extend into an initial period of highly stable combustor operation. Dynamic mode decomposition is explored as potential method to understand physical building blocks of proper orthogonal spatial layers. Non-visual indicators of lifted flames within the high-frequency (HF) pressure signal are sought to seek a method that allows for observation of lifted flames in optically inaccessible combustors, such as those in industry. Some attributes of power-spectral diagrams and cross-correlations of pressure signals are provided as potential indicators. </p>

Earlier research has demonstrated that downstream of combustion in a rotating detonation engine, exhaust flow periodically reverses circumferential direction. For small periods, the circumferential flow reaches velocity magnitudes rivaling the bulk flow of exhaust, manifesting as a swirl. The minimization of this swirl is critical to maximizing thrust and engine performance for rocket propulsion. During this study, numerous nozzle contours were iteratively designed and analyzed for losses analytically. Once a nozzle was chosen, further losses were validated through computational fluid dynamics simulations and then tested experimentally. Three different configurations were run with the RDRE: no nozzle, a nozzle without a spike, and a nozzle with a spike. Images of the exhaust quality were recorded using OH* chemiluminescence in high-speed cameras. One camera was used to confirm the existence of a detonation and the frequency of detonation. The second camera is pointed perpendicular to the exhaust flow to capture the quality of exhaust. Quantitative results of the turbulent velocity fluctuations were obtained through particle image velocimetry of the side-imaging frames. All frames in each case were exported and converted to several time-averaged frames whereupon the time-averaged turbulent velocity fluctuation profiles could be compared between cases for swirl attenuation.

Combustion in cryogenic engines is a complex phenomenon, involving either liquid or supercritical fluids at high pressure, strong and fast oxidation chemistry, and high turbulence intensity. Due to extreme operating conditions, a particularly critical issue in rocket engine is wall heat transfer which requires efficient cooling of the combustor walls. The concern goes beyond material resistance: heat fluxes extracted through the chamber walls may be reused to reduce ergol mass or increase the power of the engine. In expander-type engine cycle, this is even more important since the heat extracted by the cooling system is used to drive the turbo-pumps that feed the chamber in fuel and oxidizer. The design of rocket combustors requires therefore an accurate prediction of wall heat flux. To understand and control the physics at play in such combustor, the Large Eddy Simulation (LES) approach is an efficient and reliable numerical tool. In this thesis work, the objective is to predict wall fluxes in a subcritical rocket engine configuration by means of LES. In such condition, ergols may be in their liquid state and it is necessary to model liquid jet atomization, dispersion and evaporation.The physics that have to be treated in such engine are: highly turbulent reactive flow, liquid jet atomization, fast and strong kinetic chemistry and finally important wall heat fluxes. This work first focuses on several modeling aspects that are needed to perform the target simulations. H2/O2 flames are driven by a very fast chemistry, modeled with a reduced mechanism validated on academic configurations for a large range of operating conditions in laminar pre- mixed and non-premixed flames. To form the spray issued from the atomization of liquid oxygen (LOx) an injection model is proposed based on empirical correlations. Finally, a wall law is employed to recover the wall fluxes without resolving directly the boundary layer. It has been specifically developed for important temperature gradients at the wall and validated on turbulent channel configurations by comparison with wall resolved LES. The above models are then applied first to the simulation of the CONFORTH sub-scale thrust chamber. This configuration studied on the MASCOTTE test facility (ONERA) has been measured in terms of wall temperature and heat flux. The LES shows a good agreement compared to experiment, which demonstrates the capability of LES to predict heat fluxes in rocket combustion chambers. Finally, the JAXA experiment conducted at JAXA/Kakuda space center to observe heat transfer enhancement brought by longitudinal ribs along the chamber inner walls is also simulated with the same methodology. Temperature and wall fluxes measured with smooth walls and ribbed walls are well recovered by LES. This confirms that the LES methodology proposed in this work is able to handle wall fluxes in complex geometries for rocket operating conditions.

Concerns about the use of certain chemical species within the aerospace field are growing in recent years. A European regulation, REACh, now makes the use of hydrazine uncertain in – among others- attitude control thrusters. Green monopropellants, which are alternatives for this species already exist, but they all require a catalyst to react. Catalysts constitute the limiting factor for the lifespan of satellites because of the number of thermal cycles they endure. A joint project between ONERA, the French aerospace research center and CNES, the French space agency, was born to develop a high-performance green monopropellant thruster operating without any catalyst. Sizing the thruster and particularly its combustion chamber is not an easy task because of the explosive properties and the lack of knowledge regarding the monopropellant reaction process. The thesis aims at simulating the flow in a combustion chamber using CNES05, a new promising green monopropellant. This monopropellant has a very low vapor pressure and is an energetic liquid. As such, its reaction above a certain temperature -which is called decompositionis not well understood and must be observed closely. For this matter, a test bench was created, and it paved the way for the development of a specific model of decomposition. Indeed, even if the CNES05 decomposition cannot be modeled with the classical theory of isolated droplets, the setup showed us the order of magnitude of the reaction kinetics and the presence of a break up phenomenon. Using this model, the simulations of the flow inside the combustion chamber give us the heat flux profile through its walls, a sizing parameter for the thruster. Large recirculation zones are observed and the influence of the angle of injection seems to be the major injection parameter of influence. The sensitivity of the parameters used in the model is also studied.

La combustion cryogénique dans les moteurs de fusée dits à propulsion liquide utilise généralement un couple d'ergols, le plus couramment composé d'hydrogène/oxygène (H2/O2). Privilégiée pour le fort pouvoir calorifique du dihydrogène, cette combustion à haute pression, induit des températures de fonctionnement très élevées et nécessite l'intégration d'un système de refroidissement. La prédiction des flux thermiques aux parois est donc un élément essentiel de la conception d'une chambre de combustion de moteur fusée. Ces flux sont le résultat d'écoulements fortement turbulents, compressibles, avec une cinétique chimique violente induisant de forts gradients d'espèces et de température. La simulation de ces phénomènes nécessite des approches spécifiques telles que la Simulation aux Grandes Echelles (SGE) qui réalise un très bon compromis entre précision et coût de calcul. Cette thèse a ainsi pour objectif la simulation par SGE des transferts de chaleur aux parois dans les chambres de combustion de moteurs fusée opérant en régime sous-critique. Le régime sous-critique implique un état liquide pour un des ergols, dont il faut traiter l'injection et l'atomisation. Dans un premier temps ce travail s'intéresse à plusieurs éléments de modélisation nécessaire pour réaliser les simulations visées. Le comportement des flammes H2/O2 est décrit par un schéma cinétique réduit et validé sur des configurations académiques. La prédictivité de ce schéma est évaluée sur une large gamme de fonctionnement dans des conditions représentatives des moteurs fusée. La simulation de l'injection de l'oxygène liquide (LOx) est un autre point critique qui nécessite de décrire l'atomisation et la phase dispersée ainsi que son couplage avec la phase gazeuse. La déstabilisation et l'atomisation primaire du jet liquide, trop complexe à simuler en SGE 3D, sont omises ici pour injecter directement un spray paramétré grâce à des corrélations empiriques. Enfin, la prédiction des flux thermiques utilise un modèle de loi de paroi spécifiquement dédiée aux écoulements à fort gradient de température. Cette loi de paroi est validée sur des configurations de canaux turbulents par comparaison avec des simulations avec résolution directe de la couche limite. La méthodologie basée sur les modèles développés est ensuite employée pour la simulation d'une chambre de combustion représentative du fonctionnement des moteurs cryogéniques. Il s'agit de la configuration CONFORTH testée sur le banc MASCOTTE (ONERA) et pour laquelle des mesures de température de paroi et de flux thermiques sont disponibles. Les résultats des SGE montrent un bon accord avec l'expérience et démontrent la capacité de la SGE à prédire les flux thermiques dans une chambre de combustion de moteur fusée. Enfin, dans un dernier chapitre ce travail s'intéresse à une méthode d'augmentation des transferts thermiques via une expérience de JAXA utilisant des parois rainurées dans la direction axiale. Par comparaison avec une chambre à parois lisses, les résultats démontrent la bonne prédiction par la SGE de l'augmentation du flux de chaleur grâce aux rainures et confirment la validité de la méthode développée pour des géométries de paroi complexes. / Combustion in cryogenic engines is a complex phenomenon, involving either liquid or supercritical fluids at high pressure, strong and fast oxidation chemistry, and high turbulence intensity. Due to extreme operating conditions, a particularly critical issue in rocket engine is wall heat transfer which requires efficient cooling of the combustor walls. The concern goes beyond material resistance: heat fluxes extracted through the chamber walls may be reused to reduce ergol mass or increase the power of the engine. In expander-type engine cycle, this is even more important since the heat extracted by the cooling system is used to drive the turbo-pumps that feed the chamber in fuel and oxidizer. The design of rocket combustors requires therefore an accurate prediction of wall heat flux. To understand and control the physics at play in such combustor, the Large Eddy Simulation (LES) approach is an efficient and reliable numerical tool. In this thesis work, the objective is to predict wall fluxes in a subcritical rocket engine configuration by means of LES. In such condition, ergols may be in their liquid state and it is necessary to model liquid jet atomization, dispersion and evaporation.The physics that have to be treated in such engine are: highly turbulent reactive flow, liquid jet atomization, fast and strong kinetic chemistry and finally important wall heat fluxes. This work first focuses on several modeling aspects that are needed to perform the target simulations. H2/O2 flames are driven by a very fast chemistry, modeled with a reduced mechanism validated on academic configurations for a large range of operating conditions in laminar pre- mixed and non-premixed flames. To form the spray issued from the atomization of liquid oxygen (LOx) an injection model is proposed based on empirical correlations. Finally, a wall law is employed to recover the wall fluxes without resolving directly the boundary layer. It has been specifically developed for important temperature gradients at the wall and validated on turbulent channel configurations by comparison with wall resolved LES. The above models are then applied first to the simulation of the CONFORTH sub-scale thrust chamber. This configuration studied on the MASCOTTE test facility (ONERA) has been measured in terms of wall temperature and heat flux. The LES shows a good agreement compared to experiment, which demonstrates the capability of LES to predict heat fluxes in rocket combustion chambers. Finally, the JAXA experiment conducted at JAXA/Kakuda space center to observe heat transfer enhancement brought by longitudinal ribs along the chamber inner walls is also simulated with the same methodology. Temperature and wall fluxes measured with smooth walls and ribbed walls are well recovered by LES. This confirms that the LES methodology proposed in this work is able to handle wall fluxes in complex geometries for rocket operating conditions.