Controlling unsteady combustion induced gas flow fluctuations and the resultant motor vibrations is a very significant step in rocket motor design. It occurs when the unsteady heat release due to combustion happens to feed the acoustic oscillations of the closed duct forming a feed-back system. The resultant vibrations concerned may even lead to total failure of the rocket system unless analysed and tested thoroughly. This thesis aims developing a linear numerical analysis method for the growth rate of instabilities and possible mode shape of a liquid-propelled chamber geometry. In particular, A 3-D Helmholtz code, utilizing Culicks spatial averaging linear iterative method, is developed to find the form of deformed mode shapes iteratively to obtain possible effects of heat source and impedance boundary conditions. The natural mode shape phase is solved through finite volume discretization and the open-source eigenvalue extractor, ARPACK, and its parallel implementation PARPACK. The iterative method is particularly used for analyzing the geometries with complex shapes and essentially for disturbances of small magnitudes to natural mode shapes. The developed tools are tested via two simple cases, a duct with inactive flame and a Rijke tube, used as validation cases for the code particularly with only boundary contribution and heat contribution respectively. A sample 2-D and 3-D liquid-propelled combustion chamber is also analysed with heat sources. After comparing with the expected values, it is eventually proved that the method should be only used for determining the modes instability analysis, as to whether it keeps vibrating or decays. The methodology described can be used as a preliminary design tool for the design of liquid-propellant rocket engine combustors, rapidly revealing only the onset of instabilities.

To enable future environmentally friendly access to space by means of solid rocket propulsion a viable replacement to the hazardous ammonium perchlorate oxidizer is needed. Ammonium dinitramide (ADN) is one of few such compounds currently known. Unfortunately compatibility issues with many polymer binder systems and unexplained solid-state behavior have thus far hampered the development of ADN-based propellants. Chapters one, two and three offer a general introduction to the thesis, and into relevant aspects of quantum chemistry and polymer chemistry. Chapter four of this thesis presents extensive quantum chemical and spectroscopic studies that explain much of ADN’s anomalous reactivity, solid-state behavior and thermal stability. Polarization of surface dinitramide anions has been identified as the main reason for the decreased stability of solid ADN, and theoretical models have been developed to explain and predict the solid-state stability of general dinitramide salts. Experimental decomposition characteristics for ADN, such as activation energy and decomposition products, have been explained for different physical conditions. The reactivity of ADN towards many chemical groups is explained by ammonium-mediated conjugate addition reactions. It is predicted that ADN can be stabilized by changing the surface chemistry with additives, for example by using hydrogen bond donors, and by trapping radical intermediates using suitable amine-functionalities. Chapter five presents several conceptual green energetic materials (GEMs), including different pentazolate derivatives, which have been subjected to thorough theoretical studies. One of these, trinitramide (TNA), has been synthesized and characterized by vibrational and nuclear magnetic resonance spectroscopy. Finally, chapter six covers the synthesis of several polymeric materials based on polyoxetanes, which have been tested for compatibility with ADN. Successful formation of polymer matrices based on the ADN-compatible polyglycidyl azide polymer (GAP) has been demonstrated using a novel type of macromolecular curing agent. In light of these results further work towards ADN-propellants is strongly encouraged. / QC 20101103

Combustion instability due to feedback coupling between unsteady heat release and natural acoustic modes can cause catastrophic failure in liquid rocket engines and to predict and prevent these instabilities the mechanisms that drive them must be further elucidated. With this goal in mind, the objective of this thesis was to develop techniques that improve the understanding of the specific underlying physical processes involved in these driving mechanisms. In particular, this work sought to develop a small-scale, optically accessible liquid rocket engine simulator and to apply modern, high-speed diagnostic techniques to characterize the reacting flow and acoustic field within the simulator. Specifically, high-speed (10 kHz), simultaneous data were acquired while the simulator was experiencing a 170 Hz combustion instability using particle image velocimetry, OH planar laser induced fluorescence, CH* chemiluminescence, and dynamic pressure measurements. In addition, this work sought to develop approaches to reduce the large quantities of data acquired, extracting key physical phenomena involved in the driving mechanisms. The initial data reduction approach was chosen based on the fact that the combustion instability problem is often simplified to the point that it can be characterized by an approximately linear constant coefficient system of equations. Consistent with this simplification, the experimental data were analyzed by the dynamic mode decomposition method. The developed approach to apply the dynamic mode decomposition to simultaneously acquired data located a coupled hydrodynamic/combustion/acoustic mode at 1017 Hz. On the other hand, the dynamic mode decomposition's assumed constant operator approach failed to locate any modes of interest near 170 Hz. This led to the development of two new data analysis techniques based on the dynamic mode decomposition and Floquet theory that assume that the experiment is governed by a linear, periodic system of equations. The new periodic-operator data analysis techniques, the Floquet decomposition and the ensemble Floquet decomposition, approximate, from experimental data, the largest moduli Floquet multipliers, which determine the stability of the periodic solution trajectory of the system. The unstable experiment dataset was analyzed with these techniques and the ensemble Floquet decomposition analysis found a large modulus Floquet multiplier and associated mode with a frequency of 169.6 Hz. Furthermore, the approximate Rayleigh criterion indicated that this mode was unstable with respect to combustion instability. Overall, based on the positive finding that the ensemble Floquet decomposition was able to locate an unstable combustion mode at 170 Hz when the operator's time period was set to 1 ms, suggests that the dynamic mode decomposition based 1017 Hz mode parametrically forces the 170 Hz mode, resulting in what could be characterized as a parametric combustion instability.

La vectorisation de la poussée d'une tuyère propulsive supersonique axisymétrique est étudiée par le biais d'une injection fluidique transversale dans sa partie divergente. Cette étude menée dans le cadre du programme PERSEUS du CNES a été motivée par la recherche d'une solution alternative au pilotage conventionnel de la poussée par actionneurs mécaniques. Le travail de la thèse, tout en s'appuyant sur des approches expérimentale et numérique, comprend essentiellement une large étude paramétrique concernant principalement la position de l'injection, la forme de la tuyère, la nature et le débit du fluide injecté. L'analyse des résultats montre que pour certaines configurations optimales, des angles de déviation pertinents peuvent être obtenus pour des taux d'injections modérés. L'analyse numérique étendue aux écoulements chauds multi-espèces, plus proches des applications réelles, a montré que la vectorisation fluidique reste très performante lors de l'injection de produits de combustion dans le divergent. / Secondary injection into the divergent section of a supersonic rocket nozzle is investigated for the fluidic thrust vectoring effects. The study was conducted in the framework of CNES PERSEUS program and was motivated by the need for an alternative vectoring solution aimed for a small space launcher. The thesis work, based on the combined experimental and numerical approaches, essentially comprises of a wide parametric study mainly concerning the position of the injection, the shape of the primary and injection nozzles, flow regime, gas thermophysical properties and injected fluid mass-flow-rate. The analysis shows that for some optimal configurations, pertinent deflection angles can be obtained using the moderate injection rates. Furthermore, the study was extended to the hot flow multi-species investigation, simulating a case closer to the real applications. This numerical analysis indicated that the fluidic vectoring method remains effective with injection of combustion products into the divergent section of a propulsive rocket nozzle.

Cette thèse se focalise sur la compréhension et la prédiction des instabilités de combustion dans les systèmes à haute pression. Elle s'oriente autour de la simulation numérique d’un banc d'essai, opéré à l'université de Purdue, comprenant un injecteur caractéristique des moteurs-fusées et dont les propriétés acoustiques peuvent varier à l'aide d’un tube d'injection mobile. Une méthode d'initialisation et d'allumage pour les calculs LES de chambres de combustions terminées par une tuyère est présentée. Un point de fonctionnement instable est choisi pour étudier le mécanisme de l'instabilité. Les simulations sont comparées aux résultats expérimentaux en terme de fréquence et structure du mode instable. La fonction de transfert de flamme est calculée à l'aide du modèle n − τ puis implémentée dans un solveur acoustique (ne résolvant que les perturbations acoustiques à partir de l'équation de Helmholtz en écoulement réactif). Différents modèles d'impédance de tuyère, extraits de la littérature, sont comparés et leurs impacts sur les résultats de stabilité sont analysés. Le théorème d’impédance translatée est implémenté dans le solveur acoustique pour analyser, à faible coût de calcul, l’influence de la variation de la longueur du tube d'injection. Des écarts entre les fréquences prédites et celles trouvées expérimentalement subsistent mais la carte de stabilité de l’expérience est bien reproduite. / This thesis focuses on the understanding and the prediction of combustion instability in high-pressure devices. A model rocket combustor, tested experimentally at Purdue University, with continuously variable acoustic properties, thanks to a variable-length injector tube, is simulated. A method to initialize and ignite Large-Eddy-Simulation (LES) calculation of combustion chamber surrounded by nozzle is proposed. An unstable operating point is then chosen to investigate the mechanism of the instability. The simulations are compared to experimental results in terms of frequency and mode structure. The flame transfer function is calculated using the n − τ model to feed an acoustic solver which solves only the acoustic perturbation using a Helmholtz equation in reacting flows. The importance of the modeling of the nozzles impedance is studied through the main theories in the literature. The impedance translation theorem is implemented in the acoustic solver to analyze at low cost the influence of the variation of the injector tube. Despite differences in frequency of the instability, the stability map of the experiment is well reproduced.

Les oscillations de pression sont un enjeu majeur dans le design des moteurs à propergol solide car de faibles oscillations de pression (ODP) dans la chambre entraînent de fortes oscillations de poussée ce qui conduit à des vibrations néfastes pour les structures et les satellites embarqués. Les ODP sont encore aujourd'hui un vaste sujet de recherche et la simulation numérique est un outil indispensable dans leur analyse. De nombreux travaux ont permis de mettre en évidence divers mécanismes générateurs d'oscillations, mais la conception des nouveaux moteurs favorise la formation d'une instabilité hydrodynamique, appelée VSA et caractérisée par des détachements tourbillonnaire au niveau des discontinuités de la surface débitante. Etudiée dans les travaux sur le C1x [Vuillot 1995, Dupays 1996], il reste cependant divers points à aborder afin d'avoir une vision complète des mécanismes qui pilotent et modifient cette instabilité. Pour cela, il a été choisi dans ces travaux d'isoler le VSA dans une configuration académique et d'étudier dans un premier temps, l'impact du soufflage latéral, généré par un dégagement gazeux du à la combustion d'un bloc de propergol en aval de l'angle. Les deux approches utilisées, à savoir la simulation numérique et la stabilité linéaire, démontrent que le soufflage latéral possède un fort effet stabilisant sur le VSA. Dans un deuxième temps, l'impact de la combustion des particules d'aluminium et des résidus, présents dans un moteur à propergol solide, est analysé. Ces travaux montrent que les particules, via des mécanismes complexes, peuvent jouer à la fois un rôle stabilisant et déstabilisant sur le VSA. Pour finir, l'impact de la mise à l'échelle sur l'instabilité est étudié. Si en gaz seul, les résultats obtenus à échelle réduite sont directement transposables vers l'échelle réelle, la mise à l'échelle modifie le comportement des particules dans les structures tourbillonnaires et donc leur rôle sur l'instabilité. / Pressure oscillations (ODP) are a major issue in solid rocket motor design, as very small pressure oscillations induce strong thrust oscillations, involving vibrations detrimental to carrying load. ODP are still a vast and intense domain of research and the improvement of rocket motors mainly resorts to advanced numerical simulations. Extensive research have enabled to characterize several sources of instabilities, but new motor design promotes one hydrodynamic instability, called VSA and characterized by vortex shedding at geometry angles. VSA has be studied in the C1x configuration [Vuillot 1995, Dupays 1996] but several points still need to be studied in order to have a complete view of the phenomena driving and impacting this instability in a solid rocket motor flow. In this work, the VSA is isolated in an academic configuration and, in a first part, the impact of lateral blowing is studied. This blowing, never analysed so far, is due to burnt gases coming from the combustion of propellant block after the angle. This study has been performed following two approaches, numerical simulations and linear stability analysis. Both demonstrate the strong stabilizing effect of the lateral blowing. In a second part, the impact of aluminium particles combustion including the presence of residual particles, found in solid rocker motors, is analysed. This work shows that due to complex interaction mechanisms, particles can have a stabilizing or a destabilizing impact on the instability. Finally, the scaling impact is studied with and without particles. In purely gaseous configuration, the results obtained at reduced scale can be used directly at real scale as all the characteristics of the instability are preserved. However, with particles, the scaling modifies the particles behaviour and then the particles impact on the VSA.

Le nouveau moteur cryogénique de l'étage supérieur du nouveau lanceur Ariane présente la particularité d'être plusieurs fois réallumable, une fois la mise en orbite du lanceur. Le réallumage d'un moteur est particulièrement difficile durant les conditions de vol spatial. Ce moteur est composé d'un dôme LOX alimenté en oxygène liquide (LOX) qui est approvisionné par une vanne à boisseau positionnée en entrée d'une canne d'alimentation. Le mélange liquide / gaz formé dans le dôme LOX est injecté dans la chambre de combustion à travers des injecteurs reliant le dôme à la chambre. En conséquence, la distribution de l'écoulement diphasique en sortie des injecteurs revêt une importance particulière en terme d'allumage, de l'ouverture à la fermeture de cette vanne. La prise en compte de ces conditions de vol est primordiale pour qualifier le moteur. Cependant ces conditions ne peuvent pas être reproduites de façon représentative au cours d’essais au sol. Dans le cadre de ces études, un programme de recherche a été mis en place par le CNES (Centre National d'Études Spatiales) et SAFRAN Snecma pour étudier le remplissage du dôme LOX, via des études expérimentales et numériques. L'objectif est de connaître les conditions aux limites en sortie des injecteurs qui sont déterminantes pour appréhender la phase d'allumage dans la chambre de combustion. Des expériences ont été menées au LEGI (Laboratoire des Écoulements Géophysiques et Industriels) avec des fluides de substitution (de l'eau et de l'air), sans transfert de masse et de chaleur, sur la maquette du dôme d’alimentation d'un moteur de fusée. Les travaux présentés, menés à l'IMFT (Institut de Mécanique des Fluides de Toulouse), tentent de reproduire les expériences réalisées à l'aide de simulations numériques 3D incompressibles diphasiques. La géométrie du domaine de calcul est représentative de la maquette expérimentale, qui est composée d'une canne d'alimentation, d'un dôme, d'un allumeur et d'un grand nombre d'injecteurs. Le but de cette étude est de démontrer la faisabilité d'un calcul 3D instationnaire diphasique du remplissage du dôme oxygène du moteur-fusée avec le code industriel NEPTUNE_CFD, en prenant en compte la géométrie réelle et les phénomènes physiques prépondérants. La comparaison des prédictions avec les résultats expérimentaux est réalisée afin d'évaluer la capacité du code à prédire l'écoulement à l'aide des modèles de fermeture disponibles. Enfin, plusieurs études de sensibilité sur les modèles de fermeture sont menées pour estimer leur influence sur les résultats des simulations. Un travail important a été effectué pour imposer les mêmes conditions d'entrée que dans les expériences. Des études ont également été conduites sur un injecteur isolé. / New generation cryogenic upper-stage rocket engines are planned to be restartable during the orbit mission. The re-ignition of the engine is particularly difficult in space flight conditions. The engine contains a LOX dome fed with liquid oxygen (LOX) supplied by a bushel valve through a pipe. The gas / liquid mixture forming in the dome is injected into the combustion chamber through a number of injectors. Therefore the two-phase flow distribution at injectors outlet carries a real importance in terms of the ignition from the opening to the closing phases of the main valve. These flight conditions are of paramount importance, however, they are truly difficult to reproduce by experimental ground tests. In the framework of these studies, a research program set up by CNES (the French Space Agency) and SAFRAN Snecma, tries to study the LOX dome filling, through experiments and numerical studies. The aim is to identify the phenomena at sake to know the limit conditions at injectors, which will determine the ignition stage in the combustion chamber. Experiments are carried out at LEGI (Geophysical and Industrial Flows Laboratory in Grenoble) with substitution fluids (air and water), without heat and mass transfer on a rocket engine mockup. The work presented here, conducted at IMFT (Fluid Mechanics Institute in Toulouse), intends to reproduce the experimental results using incompressible two-phase flow simulations. The geometry used is representative of the experimental mockup composed of a feeding pipe, a dome, an igniter pipe and injectors. The aim of this study is to demonstrate the feasibility of a 3D unsteady two-phase flow calculation with the industrial code NEPTUNE_CFD, to simulate the LOX dome filling of the rocket engine, by taking into account the real geometry and the preponderant physical phenomena. The comparison of the predictions with the experimental results is carried out in order to estimate the code capability to predict the flow behavior, according to available closure laws. Finally, several sensitivity studies on the closure laws have been conducted to assess their influence on the numerical results. An important work has been carried out to obtain the proper inlet conditions to be imposed in the code in coherence with the experiments. Studies have equally been conducted on an isolated injector.

À ses débuts, la conquête spatiale a pu bénéficier des rivalités politiques de la Guerre Froide pour se développer rapidement sans réellement se soucier des efforts économiques à fournir. Aujourd’hui, de nombreux pays subissent le revers de la médaille de cette course effrénée : pour maintenir une flotte de lanceurs viable économiquement, les différentes agences spatiales doivent faire face à un dilemme opposant la minimisation des coûts de lancement à la maximisation de leur fiabilité. Dans cette logique d’optimisation, les industriels présents dans ce processus de réflexion se tournent vers la simulation numérique pour tenter d’améliorer leurs connaissances des technologies existantes, en particulier sur les zones d’ombres inaccessibles aux mesures expérimentales. Dans la lignée de plusieurs études théoriques et expérimentales, ces travaux visent à apporter un éclairage nouveau sur les phénomènes se produisant lors de l’allumage d’un moteur fusée cryotechnique. Ces recherches se tournent dans un premier temps vers l’amélioration de la modélisation de la flamme H2/O2. La validation d’une cinétique chimique réduite initialement destinée à la combustion H2/Air permet de justifier son utilisation lors de l’allumage. Puis, le développement d’un modèle de combustion turbulente pour le régime de flamme de diffusion est mené dans le but de palier aux limitations du modèle de flamme épaissie. Enfin, une analyse du cas où les régimes prémélangés et non-prémélangés sont présents tous les deux permet d’étudier un moyen simple de les distinguer même dans le cas où ils sont très proches. Dans un second temps, ces travaux se tournent vers l’étude de l’allumage dans un moteur fusée cryotechnique. Après avoir analysé de manière globale le calcul d’une séquence simplifiée, deux études plus approfondies sont menées pour investiguer, d’une part, les différents régimes de combustion, et d’autre part, les différents modes de propagation de la flamme propres à cette configuration. / The beginning of the conquest of space received benefits from the political competition of the Cold War and consequently grow quickly without considering the cost of these advances. The end of this unrestrained technological race brings to light the other side of the coin. In order to keep a fleet of launch vehicles up-to-date with the market, spatial agencies must answer a question : how can the cost of a launch be reduced without decreasing its efficiency. Through the use of numerical simulation, industrial partners may investigate this logic of optimisation. This solution might provide improvement in the knowledge of existing technologies, especially when experimental measurements are impossible. Following the path of theoretical and experimental results, this study aims to present a new view about the different processes occurring during the ignition of a space rocket engine. First, this research will present an improvement of the modelling of H2/O2 flame. The validation of a reduced chemical scheme basically developed for H2/Air will justify its use during the ignition sequence. Then, a turbulent combustion model for non-premixed flames will be developed in order to compensate the limits of the thickened flame model implemented in AVBP. Additionally, a study of both premixed and non-premixed regimes in a closed position will bring a simple method to distinguish them for a further active use. Secondly, this research will study the ignition process of a representative cryogenic space rocket chamber. The calculation of a simplified ignition sequence will be globally investigated. Finally, two-detailed analysis will lead to different combustion regimes and flame spreading processes

Submitted by PAULO ROBERTO SAKAI null (prsakai@yahoo.com.br) on 2016-01-15T12:06:37Z No. of bitstreams: 1 CARACTERIZAÇÃO DE JUNTAS SOLDADAS EM PAW E GTAW DE CHAPAS FINAS EM AÇO MARAGING 300 SUBMETIDAS A VÁRIOS REPAROS.pdf: 16627185 bytes, checksum: 4b08baddd60bec8be8acf6956cc8e2ee (MD5) / Approved for entry into archive by Ana Paula Grisoto (grisotoana@reitoria.unesp.br) on 2016-01-15T12:54:59Z (GMT) No. of bitstreams: 1 sakai_pr_dr_guara_int.pdf: 16627185 bytes, checksum: 4b08baddd60bec8be8acf6956cc8e2ee (MD5) / Made available in DSpace on 2016-01-15T12:54:59Z (GMT). No. of bitstreams: 1 sakai_pr_dr_guara_int.pdf: 16627185 bytes, checksum: 4b08baddd60bec8be8acf6956cc8e2ee (MD5) Previous issue date: 2015-12-18 / Este trabalho tem como objetivo caracterizar mecanica e metalograficamente, juntas soldadas de chapas finas em aço Maraging 300, submetidas a até três reparos, usadas na fabricação de envelopes motores foguete a propelente sólido desenvolvidos no Instituto de Aeronáutica e Espaço (IAE) em atendimento às necessidades de sua gama de lançadores. O envelope motor atua como elemento estrutural e também possui a função primária de suportar a pressão de trabalho durante a queima do propelente. Atualmente, o envelope motor é fabricado em aço 300M-ESR e o IAE tomou a decisão de substituí-lo pelo aço Maraging 300. Em função dos processos existentes no Instituto, neste trabalho utilizaram-se os processos de soldagem Plasma Arc Welding - PAW com a técnica keyhole e Gas Tungsten Arc Welding - GTAW, ambos em passe único, com metal de adição. Antes de serem submetidas aos ensaios, as juntas passaram por inspeção não destrutiva de acordo com os critérios da norma AWS D17.1. Os reparos foram feitos de forma manual e processo GTAW. Amostras da junta soldada e reparadas foram submetidas a ensaios de tração, dureza Vickers (HV) por microindentações, análises químicas, análises metalográficas e fractográficas. Corpos de prova dos cordões adjacentes aos reparos também foram avaliados. Os resultados mostram que após a solda e reparos e o tratamento térmico de solubilização e envelhecimento, a zona fundida e a região da linha de fusão da solda apresentam uma dureza abaixo das outras regiões afetadas termicamente. Para as condições da solda sem reparo e reparadas, o processo PAW apresentou um valor menor de dureza em todas estas regiões com relação ao processo GTAW. As análises da superfície dos corpos de prova soldados rompidos indicam o predomínio de um processo de ruptura iniciado próximo à linha de fusão da solda e que se propaga em direção ao interior do cordão. A natureza da fratura mostrou o domínio da formação de alvéolos (dimples). Os corpos de prova soldados GTAW apresentaram uma resistência mecânica mais alta do que os do processo PAW. Igualmente, os corpos de prova soldados PAW obtidos dos cordões das regiões adjacentes aos reparos tiveram valores de resistência inferiores. Embora os valores de resistência mecânica das juntas soldadas submetidas a até três reparos no mesmo ponto tenham apresentado grande variabilidade, não há indicativo de diminuição da resistência com relação a junta sem reparo. / This work aims at mechanic and metallographic characterization of Maraging 300 welded joints sheets, submitted to up to three repairs, used for the fabrication of solid propellant rocket motors at the Institute of Aeronautics and Space – IAE as to comply with its range of launchers. The rocket motor is a structural part and also has the primary function of supporting the nominal pressure during the propellant burning. At present, the rocket motor is fabricated in 300M-ESR steel and IAE has decided to replace such a steel for the Maraging 300 one. Due to IAE’s existing processes, Plasma Arc Welding – PAW with the keyhole technique and the Gas Tungsten Arc Welding – GTAW have been used, both single-pass welding with filler. Before they have been submitted to the tests, the joints went through non-destructive inspection according to AWS D17.1 Standard. Manual repairs and GTAW process have been made. Samples of the welded and repaired joints were submitted to tensile testing, Vickers hardness, chemical analysis, fractrographic and metallographic analysis. Body tests of the beads adjacent to the repairs have also been assessed. Results show that after welding, repairs and solubilization and aging heating treatment, the melted zone as well as the weld joins lines zone present hardness below other heat affected zones. As for the conditions of the non-repaired and repaired welds, the PAW process has demonstrated lower hardness values in all zones in what regards the GTAW process. The welded and fractured body tests surfaces analysis indicate the predominance of a fracture process started next to the weld joins lines which goes towards the bead interior. The nature of the fracture has shown the predominance of dimples. The GTAW welded body tests presented higher mechanical strength than that of the PAW process. Similarly, the PAW welded body tests obtained from the beads of the zones adjacent to repairs presented lower strength values. Although the mechanical strength values of the welded joints submitted to up to three repairs in the same point have shown great variability, there is no indication of strength decrease regarding the non-repaired joint.

Sous la direction de SNECMA, un nouveau moteur cryotechnique pour Ariane 5 est en phase de développement. Ce moteur, VINCI, fonctionne à l'oxygène et à l'hydrogène liquides et devra être rallumable en cours de vol. Le transitoire de remplissage d'une cavité intermédiaire par laquelle transite l'oxygène liquide avant d'arriver dans la chambre de combustion s'avère être une étape critique qu'il faut s'efforcer de maîtriser. Cette cavité, appelée dôme LOX, est directement reliée à la chambre de combustion par les injecteurs à oxygène. Des outils numériques sont actuellement en cours de conception : ils permettront à terme de simuler le remplissage de cette cavité dans l'espace. Afin de valider ces outils numériques, un programme expérimental a vu le jour, impliquant SNECMA, le CNES (Centre National d'Etudes Spatiales) et le LEGI (Laboratoire des Ecoulements Géophysiques et Industriels). Il s'agit de mener à bien des expériences « simples » et reproductibles afin de disposer d'une base de données expérimentales qui servira de cas tests pour les simulations. Un banc d'essais expérimental a été progressivement mis en place et instrumenté au LEGI. Le fluide utilisé en substitution de l'oxygène liquide est de l'eau. On a choisi de respecter une similitude du nombre de Weber entre le cas « réel » et les conditions expérimentales. Ce dernier équivaut au rapport des forces d'inertie sur les forces de tension de surface. Deux campagnes d'essais ont été réalisées, dans lesquelles on s'est centré sur l'étude des aspects hydrodynamique du transitoire de remplissage : variation du débit total et des pressions au cours d'un essai, évaluation du taux de vide dans la cavité, visualisation de l'écoulement dans la cavité et en sortie des injecteurs. L'instrumentation à disposition est constituée d'un débitmètre à effet Coriolis, de capteurs de pression, d'une sonde optique, de caméras rapides et d'un laser pour l'imagerie. La première campagne d'essais a visé le remplissage de la cavité en eau seule. Le paramètre de contrôle principal était la pression génératrice de l'écoulement liquide. Dans la deuxième campagne d'essais on injecte simultanément dans la cavité un écoulement d'eau et un écoulement d'air. Cela se rapproche plus des conditions réelles du transitoire de remplissage, au cours duquel la cavité est balayée par un écoulement d'hélium. Le paramètre de contrôle supplémentaire est le débit de gaz initial. Ces campagnes ont également souligné l'importance du profil d'ouverture de la vanne de l'écoulement liquide sur le transitoire de remplissage de la cavité. Ces campagnes d'essais constituent une première étape dans la compréhension du transitoire de remplissage du dôme LOX. Elles ont permis de visualiser la forme de l'écoulement dans la cavité et en sortie des injecteurs et d'identifier certains phénomènes intervenant dans le remplissage de la cavité. En particulier, nous avons mis en évidence l'existence d'un délai de mise en place de l'écoulement par les injecteurs, qui peut être responsable d'un pic de pression dans la cavité au cours du transitoire. L'influence de la fraction gazeuse sur l'écoulement dans les injecteurs a été soulignée mais reste à quantifier de façon précise. La prochaine étape de l'étude concerne les aspects énergétiques du transitoire de remplissage, notamment les transferts thermiques ayant lieu, dans la réalité, entre l'hélium, l'oxygène et les parois du dôme LOX. / Under the supervision of SNECMA, a new cryotechnic engine is being developed for Ariane 5. This engine, named VINCI, uses liquid oxygen and liquid hydrogen as propellant. It must be re-ignitable in flight. The filling transitory phase of an intermediate tank where the liquid oxygen passes through before entering the combustion chamber, has proved to be a very important stage that must be handled. This tank, called LOX dome, is directly linked to the combustion chamber through the oxygen injectors. Numerical tools are currently under development. They will allow to simulate the filling of this tank in the space. In order to validate those numerical tools, an experimental program has been launched. It involves SNECMA, the CNES (Centre National d'Etudes Spatiales : National Centre for Spatial Studies) and the LEGI (Laboratoire des Ecoulements Géophysiques et Industriels : Laboratory of Geophysical and Industrial Flows). Simple and repeatable experiments must be run. They will allow to gather experimental data that will further be used as test cases for the simulations. A test bench has been brought into service step by step at the LEGI, as well as scientific instruments. Water is used in place of liquid oxygen. A similarity of flows based on the Weber number has been chosen between the real case and the experiment. The Weber number measures the relative importance of the fluid inertia compared to its surface tension. Two experimental campaigns have been realized, that have focused on the dynamic aspects of the filling transitory phase : variations of the total flow and of the pressures measured during an experiment, evaluation of the void fraction in the tank, flow visualization in the tank and at the outlet of the injectors. The scientific instrumentation used is made of a Coriolis flow-meter, pressure probes, an optical probe, and high speed cameras with a laser for the flow visualization. The first experimental campaign has studied the tank filling with water only. The main control parameter is the reference pressure of the liquid flow. In the second campaign, both liquid and air flows are simultaneously injected in the tank. It aims at reproducing the real conditions of the filling transitory phase, where helium is injected in the tank with the liquid oxygen. The additional control parameter is the initial gas flow. Those campaigns have shown as well the importance of the valve opening that controls the liquid flow. Those campaigns are a first step in the understanding of the filling transitory phase of the LOX dome. They have permitted to visualize the flow in the tank and at the outlet of the injectors and to point out some important phenomena occurring during the tank filling. In particular, they have highlighted the existence of a delay before the flow can develop through the injectors. This delay can be responsible for a pressure peak in the tank during the transitory phase. The influence of the gas fraction on the flow through the injectors has been underlined as well but still must be accurately quantified. The next step of the study concerns the energetics of the filling transitory phase, especially the thermal transfers that occur between the helium, the oxygen and the walls of the tank.