Validating ceramic-matrix composites (CMC) for energetic liquid bipropellant propulsion systems

This thesis presents the thermal analysis of a 2200 N (500 lbf) thruster made of Carbon/Silicon Carbide (C/SiC) composite, operating around 2 MPa (300 psi). Liquid oxygen and propylene are selected as propellants due to their high combustion temperature ≈3500 K. Convective, conductive and radiative heat transfers are modeled in the chamber and supersonic nozzle using COSMOS Floworks 2004. Implementation is verified for several adiabatic flow conditions amenable to 1-dimensional analytical solutions. Accounting for heat transfer, the analyses predict maximum chamber wall temperatures around 2200 K, ∼600 K below the SiC melting point; temperatures where C/SiC oxidation/ablation rates are expected to be small. Also, detailed, steady state thermal analysis of the interface between the aluminum injector and the C/SiC chamber shows the predicted temperatures of the supporting metal components are below their respective melting points. These analyses suggest that the thruster can operate without additional cooling for extended periods of time.