The present research work introduces a novel combustion model formulation based on the flamelet concept suitable to investigate characteristic flow conditions in rocket engine thrust chambers. The new approach extends the validity domain of existing flamelet models to fully cover the observed flow conditions by incorporating kinetic rate effects into a state-of-the-art nonadiabatic flamelet manifold. This enables a transition between the fluid dynamic and reaction-rate-dominated combustion regimes, depending on the local Damköhler number of the numerical solution field. The approach is investigated based on a representative rocket engine test case focusing on highly nonadiabatic and de Laval nozzle flow conditions for the propellant combination methane-oxygen. Covering a range of propellant mixture ratios associated with significant variations of the underlying chemical effects, three flamelet combustion-modeling approaches of different manifold complexity are compared to high-fidelity laminar finite rate results based on the full chemical reaction mechanism. The obtained data show that the proposed model formulation is able to deliver a notably higher accuracy in the numerical prediction compared to the existing modeling approaches while retaining high computational efficiency to meet industrial design requirements.

本研究工作介绍了一种基于小火焰概念的新型燃烧模型公式，适用于研究火箭发动机推力室内的特征流动条件。新方法通过将动力学速率效应纳入最新的非绝热小火焰歧管中，扩展了现有小火焰模型的有效性域，以完全覆盖观察到的流动条件。这样就可以根据数值解域的局部Damköhler数，在流体动力学和反应速率主导的燃烧状态之间进行转换。该方法是基于代表性的火箭发动机测试案例进行研究的，该案例重点关注推进剂甲烷-氧气的高度非绝热和de Laval喷嘴流动条件。涵盖了与潜在化学作用的显着变化相关的一系列推进剂混合比，将基于不同化学反应机理的三种不同歧管复杂性的小火焰燃烧建模方法与高保真层流有限速率结果进行了比较。获得的数据表明，与现有的建模方法相比，所提出的模型公式能够在数值预测中提供显着更高的精度，同时保持较高的计算效率以满足工业设计要求。