Design and Analysis of a CMC Radiation Cooled LOX/Propylene Thruster

The paper presents the design and performance analysis of an 800 lbf LOX/propylene rocket engine, designed to power the upper stage of a Nanosat Launch Vehicle (NLV). A precursor version of this engine with relaxed requirements was static fire tested twice and has undergone the first recorded flight test of a LOX/propylene engine by a team at California State University, Long Beach back in 2009. The revised engine has an expansion ratio of 70 with a targeted specific impulse of 342 s, and combustion and nozzle efficiencies of 95% or greater. The injector features a LOX-centered split triplet with a center-mounted TEA-TEB igniter. A single piece Cf/SiCm ceramic matrix composite is used for radiation cooling of the thrust chamber. Fuel film cooling (FFC) is added to lower the chamber wall temperature to the 2200-2400C range near the throat, maintaining the C/SiC structural integrity and minimizing throat erosion. Computational Fluid Dynamics analyses indicate that 10% FFC is sufficient to achieve this goal while almost meeting the targeted specific impulse. A proof-of-concept version of this engine with relaxed requirements designed for near sea-level testing is presented.

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