GenCorp Aerojet made extensive use of FLUENT in designing a novel gaseous hydrogen-oxygen injector element for rocket propulsion. The rocket application requires that the combustion process be stable and complete, and take place in as short a distance as possible without compromising the structural integrity of the injector itself. The rocket industry has only limited experience with gaseous propellant systems and even less with hydrogen-oxygen systems. CFD models of the combustion process substituted for design experience by providing detailed understanding of how and where mixing and combustion would occur.

The rocket-based combined cycle (RBCC) engine concept integrates small rocket thrusters into the rear-facing base area of struts placed in the flow path of a scramjet (supersonic combusting ramjet) engine. At vehicle takeoff, the rockets provide the primary thrust with additional thrust coming from an ejector effect as air is drawn into the engine inlet, entrained and accelerated by the rocket exhaust. Because the exhaust in this application is mixing with ambient air, the potential exists for additional combustion in the engine flow path and thermal choking of the flow. For these reasons, the engine is required to run at a very nearly stoichiometric oxygen-to-fuel-mass mixture ratio of 7 to minimize free hydrogen in the exhaust.

The simulations predicted a comfortable temperature margin at the injector face, which was confirmed by subsequent test firings.

Initial CFD simulations used non-reacting propellant streams, with the idea that they would give a reasonable representation of mixing and combustion. Later work made it clear that heat released from combustion changes the entire character of flow and that any non-reacting model would be grossly inadequate. FLUENT was used to simulate the reacting flow using the PDF chemistry model. After an injector design was selected based on predicted gas temperatures near the injector face and combustion efficiency, a more complete model employing conjugate heat transfer and including the solid injector face and propellant passages was created. This model's prediction that a comfortable temperature margin existed at the face was confirmed by subsequent test firings.

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