New propulsion systems play a key role for future space transportation systems (STS). While high efficiency with respect to specific impulse is mandatory aspects concerning system design (complexity, manufacturing, weight), operating aspects (reusability, maintenance, high reliability and safety) and cost reduction attract increasing attention. A significant step forward seems to be the application of high temperature usable ceramic matrix composites (CMC) within a cryogenic combustion chamber design. The approach is using a porous fibre ceramic inner combustion chamber liner. Operating temperatures of the CMC material lie far upper than 1800 • C. Nevertheless an appropriate cooling method has to be applied accompanied by temperatures of about 2700 • C running the combustion process. While the porous CMC is responsible for both thermal load absorption and the effusion cooling the mechanical load carrying structure consists on a carbon fibre reinforced plastic (CFRP) outer shell. Using both ring shaped ceramic combustion chamber segments on 1-10 MPa pressure levels and a complete ceramic thrust chamber design on 1 MPa pressure level samples demonstrated already the physical principles. The paper will give an overview of the current development status, whereby the experience achieved so far is very promising out looking to the general goals.
A thermal barrier coating (TBC) system for rocket chambers made of Cu-based high strength alloys has been developed in a pilot project in line with EB-PVD (electron-beam physical vapor deposition) technology aiming at TBC application on Cu-based walls of real rocket combustion chambers. The TBC system consists of a metallic bond coating compatible with Cu-based material and an yttria partially stabilized zirconia TBC. The TBC overlayer is a distinctive ceramic structure designed for an exceptionally low Young's modulus to withstand the extreme mismatch stresses between the internally LN-cooled high thermal expansion Cu metal base and the low thermal expansion hot ceramic shell. The TBC system has been qualified under close-to-service conditions on cylindrical LH 2 -cooled combustion chamber segments, where they have performed superior.As EB-PVD technology is a line-of-sight process that is rather able to coat internal cavities, a transient liquid phase (TLP) joining technique for fully coated parts has been developed, that allows to assemble complete components out of vapor-accessible fully coated parts. It is capable, e.g. to incorporate sinuous cooling passages in the throat areas of combustion chambers, and/or to assemble oversized parts out of smaller components by maintaining parent metal properties. A manufacturing process is outlined for making internal TBC armored combustion chambers. q
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