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978-3-8439-2144-2, Reihe Strömungsmechanik
Edder José Rabadán Santana
Numerical Investigation of a Generic Supersonic Combustion Chamber under Realistic Flight Conditions
188 Seiten, Dissertation Universität Stuttgart (2015), Softcover, A5
The main objective of this thesis is to investigate the proposed combustor configuration under realistic hypersonic flight conditions in order to determine its characteristics and operability limits as well as to evaluate its performance and establish references for further conceptual designs. To achieve that the supersonic combustion chamber was numerically investigated at the prescribed design point given by a flight Mach number of 8 at an altitude of 30 kilometers. The investigation was conducted in sequential steps, first the combustor was investigated for a fuel-off case, which was aimed to study the flow field properties prior the fuel injection. In this study it was found that the thermodynamic properties, particularly temperature, were within acceptable levels for H2 self-ignition. Accordingly, the second step was to investigate the mixing process. For this purpose numerical simulations for non-reacting fuel-on cases were conducted. The results showed that the mixing concept, a central lobed strut injector, provided a satisfactory mixing efficiency up to 80% and combustion was very likely to occur. The injection system and the mixing process increased the pressure losses by less than 6% which was considered an acceptable value.
After the mixing case was evaluated, the biggest part of this research was focused on the reacting cases. The combustor was investigated at the flight design point and for two additional off-design conditions given by 1) a flight Mach number of 8 at an altitude of 28 kilometers and 2) a flight Mach number of 6 at an altitude of 26 kilometers. Along these off-design conditions different angles of attack were also investigated. The results showed that combustion was successfully achieved for all cases and the combustion chamber was able to operate properly at the investigate flight conditions. There was no evidence of engine failure neither by boundary layer separation nor by thermal choking.