Abstract
The long-term development of ceramic rocket engine thrust chambers at the German Aerospace Center (DLR) culminates in compact designs of transpiration-cooled fibre-reinforced ceramic rocket engine chamber structures. Achievable benefits of the transpiration cooled ceramic thrust chamber are the reduction of engine mass and manufacturing cost, as well as an increased reliability and higher lifetime due to thermal cycle stability. The transpiration cooling principle however reduces the engine performance. Due to the transpiration cooling the characteristic velocity decreases with increasing coolant ratio. The goal of the chamber development is therefore to minimize the required coolant mass flow. The wall temperature can be calculated using known heat transfer correlations, for example given by Bartz, and employing a model given in literature for the reduction of the heat transfer coefficient based on coolant mass flow. By this method the required coolant mass flow ratio for different chamber diameters and pressure levels can be calculated. This paper discusses the application potential of DLR’s ceramic thrust chamber technology for high performance engines. Parametric variations of engine sizing (such as chamber pressure and diameter) are performed. For large diameters and high chamber pressures the required coolant ratio is below 1%.
