The successful incorporation of high performance nitrogen tetroxide oxidizer and the monomethylhydrazine fuel bipropellant rocket engines into integrated propulsion systems utilized on satellites and spacecraft is due in part to the use of high melting point material systems for the combustion chamber and exit nozzle. The 445 N attitude control rocket engine used on the Apollo service module and lunar module embodied molybdenum as the combustion chamber material. This material system was soon replaced by the niobium (columbium) alloy material system which is used in a majority of the small liquid rocket engines in service today. The material operating limit of 1600°C with the R-512 oxidation coating is more than adequate for the demonstrated performance levels of nearly 2900 N s/kg but the structural-mechanical differences between the coating and the base metal limits the operation time at temperature when the chamber is subjected to thermal cycling. New material systems such as the silicon carbide matrix system, rhenium and zirconium oxide have been evaluated and have demonstrated increased temperature capabilities and less sensitivity to the thermal cycling characteristic of rocket engine operation when subjected to station keeping and on-orbit attitude control. Tests conducted at Societe Europeenne de Propulsion with the silicon carbide material system and tests conducted at Marquardt with a layered rhenium and zirconium oxide matrix material system demonstrated increase durability at higher engine performance levels. Combustion chamber temperatures up to 2100°C were documented at specific impulse performance levels of 3000 N s/kg.