DLR (CC BY-NC-ND 3.0).
The thrust chamber (consisting of the injection head, combustion chamber and expansion nozzle) as part of a rocket engine is one of the most thermally and mechanically highly stressed components. Therefore, it has a great influence on the performance, reliability and cost of a launcher system. The combination of modern composite materials with innovative design concepts and cooling methods within the framework of Black Engine technology promises advantages in terms of simplified production, reliability and reusability.
The name ‘Black Engine’ is derived from the predominantly black carbon fibres used in the hybrid construction of the composite components. Following on from the previous sub-scale developments on lox/LH2 thrust chambers, Technology Readiness Level 5 of this technology for a 60 kN lox/LCH4 thrust chamber is to be demonstrated through ground tests in the current Black Engine project. Within the project, the system (consisting of injection head, combustion chamber and nozzle) is being developed for a combustion chamber pressure of 70 bar and an ambient pressure of 1 bar.
Taking advantage of the favourable material properties of ceramic matrix composite (CMC), the combustion chamber is cooled by transpiration of the fuel through the combustion chamber wall. For this purpose, part of the fuel is used for cooling instead of combustion. This highly efficient cooling method with its low pressure requirement in contrast to conventional regenerative cooling, combined with the high temperature resistance of the material, potentially results in a very efficient overall system. The modular design of the CMC combustion chamber also enables additional simplifications and redundancies in production. The very small thermal expansion of the materials is advantageous in respect to durability and reusability. In addition, the load-decoupled structural design – the mechanical loads are absorbed by the cold CFRP housing – promises a long service life and high reliability.
The multi-shell CMC expansion nozzle is cooled by the fuel film transpiring through the porous inner wall and no longer requires additional cooling due to its high temperature resistance in the supersonic part. This increases the efficiency of the entire rocket engine system by reducing the demands on the cooling system and the general complexity of the design.
The aim of the project is initially to demonstrate the technology for a micro-launcher with a relevant fuel combination and thrust class. Nonetheless, due to its robustness and scalability, the innovative technology seems to be suitable for extremely powerful first-stage propulsion systems of large launchers as well.