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

A need to find better achievement has always been required in the space industrythrough time. Advanced technologies are provided to accomplish goals for humanityfor space explorer and space missions, to apprehend answers and widen knowledges. These are the goals of improvement, and in this thesis, is to strive and demandto understand and improve the mass of a space nozzle, utilized in an upperstage of space mission, with an expander cycle engine. The study is carried out by creating design of experiment using Latin HypercubeSampling (LHS) with a consideration to number of design and simulation expense.A surrogate model based optimization with Multidisciplinary Design Optimization(MDO) method for two different approaches, Analytical Target Cascading (ATC) and Multidisciplinary Feasible (MDF) are used for comparison and emend the conclusion. In the optimization, three different limitations are being investigated, designspace limit, industrial limit and industrial limit with tolerance. Optimized results have shown an incompatibility between two optimization approaches, ATC and MDF which are expected to be similar, but for the two limitations, design space limit and industrial limit appear to be less agreeable. The ATC formalist in this case dictates by the main objective, where the children/subproblems only focus to find a solution that satisfies the main objective and its constraint. For the MDF, the main objective function is described as a single function and solved subject to all the constraints. Furthermore, the problem is not divided into subproblems as in the ATC. Surrogate model based optimization, its solution influences by the accuracy ofthe model, and this is being investigated with another DoE. A DoE of the full factorial analysis is created and selected to study in a region near the optimal solution.In such region, the result has evidently shown to be quite accurate for almost allthe surrogate models, except for max temperature, damage and strain at the hottestregion, with the largest common impact on inner wall thickness of the space nozzle. Results of the new structure of the space nozzle have shown an improvement of mass by ≈ 50%, ≈ 15% and ≈ -4%, for the three different limitations, design spacelimit, industrial limit and industrial limit with tolerance, relative to a reference value,and ≈ 10%, ≈ 35% and ≈ 25% cheaper to manufacture accordingly to the defined producibility model.