Low-cost access to space demands durable, cost-effective, efficient, and low-weight propulsion systems. Key components include boost and upper stage rocket engine nozzles and extensions. Nozzle material options include ablatives, actively cooled alloys, and radiation-cooled composites and metals, each of which has known limitations. Metallic nozzles have high density and limited temperature capability. Carbon/carbon (C/C) is an attractive alternative, but has manufacturability, oxidation resistance, and joining ability concerns.

In this work, an innovative, integrated ceramic matrix composite (CMC)-C/C composite flange or spool transition element was demonstrated that could enable the attachment of a domestically sourced 2D C/C nozzle extension to a metallic manifold. Thermal structural design and analysis showed that this flange mitigates high stresses at the joint between the high-thermal-expansion metal manifold and the low-thermal-expansion C/C extension.

Melt infiltrated (MI) carbon fiber-reinforced zirconium carbide, hafnium carbide, zirconium-silicon carbide, and hafnium-silicon carbide CMCs were previously demonstrated for use in liquid and solid propellant rocket engine thrust chamber and hypersonic leading edge and nose-tip applications. C/C primary structures were then combined with integral MI CMC liners to provide high-temperature-capable structures bridging the gap between the low-density C/C and the more durable, but higher- density CMCs.

In current work, a variation of this process was applied in which the CMC matrix was infiltrated into a partially densified C/C preform, resulting in an integrated CMC-C/C structure.

This MI processing approach is much faster and more cost-effective than chemical vapor infiltration (CVI) or preceramic polymer infiltration and pyrolysis (PIP) processing commonly used for production of CMCs. Relative to other MI processes, this approach relies on wicking of the metal into the preform rather than on pressure infiltration. Also, this MI process is focused primarily on more refractory matrices (e.g., zirconium-and hafnium-based ceramics) than on the silicon carbide-based matrices typically produced by other suppliers. In addition to the MI process itself, the success of processing and the effectiveness of the resultant MI CMCs rely on the use of fiber interface coatings produced by ultraviolet activated chemical vapor deposition (UVCVD), which protects the fibers from degradation during processing, and provides mechanical slip between the fibers and the ceramic matrix during operation.

The fully developed integrated CMCC/ C material system can support a range of potential NASA missions. This work directly targeted future launch and exploration vehicle propulsion systems, including the Space Launch System, as potential end-use applications of this innovative technology. Of particular interest are large boost-scale nozzles and upper stage nozzle extensions similar in scale to the J-2X nozzle extension, the RL10B-2 upper stage nozzle, the niobium alloy nozzles used on the SpaceX Merlin engine, and future large nuclear propulsion nozzles. More generally, the versatility of this concept makes it relevant to a variety of structures including combustion chambers, leading edges, thermal protection systems, airframes, and other propulsion components that require mating of C/C to dissimilar materials.

This work was done by Timothy Stewart of Ultramet for Marshall Space Flight Center. NASA is seeking partners to further develop this technology through joint cooperative research and development. For more information about this technology and to explore opportunities, please contact Ronald C. Darty at This email address is being protected from spambots. You need JavaScript enabled to view it.. MFS-32993-1