• combustion
• injection
• LPR
• propulsion

The objective of the current thesis work could be basically summarized as the realization of the flight version designs for the injector assembly and the combustion chamber head of a 500 N liquid bipropellant rocket engine. The work is embedded in the German Space Administration(DLR)-STERN project “SMART ROCKETS” consisting in the design, manufacture and flight campaign of a Liquid Oxygen-Ethanol propelled rocket. Both design problems were carried on in parallel, with the aim of not only realize a functional design capable of satisfying the driving requirements, but also amenable of being incorporated in the evolving engine concept, whose components interface features represent crucial aspects for the whole assembly realization and validation.The path toward the definition of the flight version coaxial swirl injector started with the characterization of the existing test injectors built within the project, based on existing flow models and experimental data gained through a series of cold flow and combustion tests. Those tests evidenced that the available fluid dynamic models, based on the strong hypothesis of inviscid flow, were only capable of accurately describe the behavior of the oxidizer injector while gave unacceptable prediction errors for the fuel injector, leading to the difficulty in find a reliable massflow-pressure correlation for the injection assembly. Therefore the focus shifted in enhancing the flow model including the viscosity effects believed to be the reason of the fuel injector discharge capability drop. The characterization of the Boundary Layer structure inside the swirl injector chamber under some simplifying hypotheses brought to a close form analytical solution of the viscous swirler problem. Thus, further cold flow tests were planned and conducted in order to verify the validity of the obtained solution: the test results matched with great accuracy the predictions of the built viscous model over a wide range of measurements, leading to the assessment of the enhanced injector flow model and the building of the operative map for both injectors. This gave the possibility to design the flight version injectors meeting the baselined requirements.The definition of the flight version combustion chamber head also started considering the analysis of the current project process: the interface nature of this element forced its design to be subject to several constraints coming from the other subsystems development. The optimization process of the flight version focused on the choice of the combination of geometric and material features of the designed concept, facing the compromise between weight saving and thermal-structural integrity. The use of lightweight materials such Aluminum alloys seemed to be the most promising solution for a lightweight design and an active cooling strategy has been implemented in order for the lightweight structure to withstand the thermal loads coming from combustion.A thermal model has been built for the convective heat transfer of burning gases and coolant beside the injector plate and coupled with a structural analysis, through which the optimum “leading thickness” of the structure has been obtained as the outcome of a thermal/stress analytical optimization, then validated through the use of FEM codes. Tuned the interface features together with the other mattering subsystems, a final lightweight design has been assessed with the choice of a High Temperature Aluminum Alloy, ready to be manufactured and capable of satisfying both thermal-structural and interface requirements.