THRUST CHAMBER MATERIALS AND DESIGN CONCEPTS EVALUATION

A program was conducted to investigate and evaluate new and unique designs of non-regenerative combustion chambers and nozzles using new materials or new techniques for applying the best materials. Analysis of information obtained through literature search, contact with material suppliers, and laboratory testing indicates that materials which will have the best performance, i.e. lowest reactivity with high temperature fluorinated propellant combustion products, are essentially carbon base materials--graphites, carbides, pyrolyzed composites. Rocket motor tests were conducted using LF2/hydrazine blend propellant composition which is equivalent to a theoretical combustion temperature of over 7000 F to evaluate candidate materials and to demonstrate design concepts and scaling relations. Eight 100 pound and three 3750 pound thrust chambers were tested at chamber pressures of 150 and 200 psia. Material and design concepts included prestressed tantalum carbide, arc cast hypereutectic hafnium carbide, pyrolyzed composites of Carb-I-Tex 7000 and PTB, heat sink design of pyrolytic graphite, tungsten, high-density graphite, hot- bonded Grafoil, and Thompsine Tape. The experimental findings were supplemented with technical and analytical support in an effort to establish the limitations that the propellant environment imposes on the selected materials and design concepts. The erosion or chemical reaction of liner materials appear to be diffusion controlled in most instances and a method for prediction of erosion based upon surface temperature and chamber pressure is reported. Conclusions and recommendations for further work are offered.