Numerically Efficient Non-Adiabatic Real Fluid Flamelet-Based Combustion Modeling Approach for Liquid Rocket Engine Thrust Chambers

In the present work, a consistent thermophysical modeling framework for the representation of real fluid effects via a single-phase dense gas approach is discussed and its prediction quality at transcritical conditions typical for liquid-propellant rocket engine injection is investigated. It is linked to a non-isobaric and non-adiabatic flamelet model in an efficient and robust manner. The inevitable accumulating error which is inherently associated with the applied energy correction procedure is quantified and evaluated as tolerable. All methods used in this context are independent of the propellant combination. The ability to adequately predict typical rocket engine performance characteristics, flow properties and the basic flame topology via the generated five-dimensional flamelet manifolds is demonstrated by its application in the numerical simulation of two respective supercritical load points of a subscale single flame combustor configuration operated with LOx/H2 and LOx/CH4 featuring a shear coaxial injector element, film cooling as well as nozzle expansion, which was extensively measured experimentally and provides validation data in terms of time-averaged OH∗ flame radiation images and wall pressure profiles. The influence of the selected thermal boundary conditions and the propellant injection states on the evolving flow field is analyzed in detail.