Dottorato in INGEGNERIA AERONAUTICA E SPAZIALE

Titolo della tesi: Heat Transfer Modelling and Analysis of Oxygen/Methane Uncooled and Film-Cooled Liquid Rocket Engines

The high hot gas temperatures and chamber pressures typical of modern liquid rocket engines (LREs) yield high local wall heat fluxes which have to be inevitably managed by the engine active cooling system. Nowadays, regenerative cooling alone might not be enough to counteract such high thermal loads in high-pressure high--performance engines, so cooling capabilities are typically enhanced with the addition of further strategies, such as film cooling or mixture ratio biased peripheral injectors. The prediction of wall heat flux in LREs is of paramount importance during the design phase both for sizing and safety purposes, especially at the nozzle throat where the maximum thermal load occurs. Numerical simulations can help in the prediction, provided that they can be effectively used during the design phase and that suitable modeling is employed. In this framework, this thesis aims at evaluating the suitability of different modeling solutions to predict in affordable times the wall heat flux of LREs employing the oxygen-methane propellant combination, which is nowadays attracting the attention of many developers as a possible cheaper and denser replacement to hydrogen. In particular, the first part of the results presented in this thesis is devoted to the analysis of the throat heat flux, whereas the second part addresses the cylindrical region of the combustion chamber. Computational fluid dynamics (CFD) simplified approaches are presented initially in case of absence of an active cooling system, and validated against experimental data as well as more accurate yet simplified numerical simulations carried out with a higher level of model completeness. Then, the cooling strategy is introduced focusing on gaseous film cooling and mixture ratio bias techniques. Due to the extreme lack of experimental data in the literature regarding the oxygen--methane propellant combination, attention is focused on a second thrust chamber representative of a possible methane--fueled upper stage. The simplified approaches validated for the uncooled case are employed and further improved to perform parametric analyses aimed at investigating how the main design parameters, such as the secondary flow mass flow rate and mixture ratio, affect the throat heat flux and the engine performances. Eventually, the second and final part of the results addresses the low--order modeling of gaseous and liquid film cooling in the cylindrical part of the combustion chamber. An extensive literature review allowed to select the most appropriate formulations to be implemented in the EcosimPro/ESPSS (European Space Propulsion System Simulation) framework, eventually providing a new component to be included in the multi--physics platform. The reliability of the predictions is assessed by analyzing transients and comparing the steady--state results against the selected experimental test cases and CFD numerical simulations performed employing the approaches above.