The development of high aspect-ratio, high precision micromachining in silicon or silicon carbide suggests the feasibility of rnicrofabricated, high chamber pressure chemical rocket engines. Such an engine, approximately 20x15x3 mm in size, would produce about three pounds of thrust using 300 sec I sp propellants. As part of the present work, the feasibility of these engines has been investigated and a liquid-cooled, pressure-fed thrust chamber has been designed, fabricated, and tested to evaluate the feasibility of the concept. The results of the tests to date using oxygen and methane as propellants support the feasibility of the concept, producing a maximum thrust of 1 N at a chamber pressure of 12 atm. Given the 1.2 gram mass of the thrust chamber, this corresponds to a thrust-to-weight ratio of 85:1. The characteristic exhaust velocity, c*, a measure of combustion effectiveness, appears to be nearly independent of chamber pressure, indicating that chemical reaction rates are not limiting the combustion. Additionally, when effects of chamber heat loss are included, c* appears to approach its predicted ideal value, indicating that the transport and mixing of propellants in the combustion chamber is of the right order to provide for complete combustion. The thrust chamber was fabricated by etching the required patterns into each side of six 0.5 mm thick silicon wafers, and then diffusion bonding the six wafers together to create the one-piece thrust chamber. A packaging technique is presented to interface high pressure and high temperature fluids to the silicon rocket engine chip. Additionally, initial modelling work has lead to the development of a methodology for mapping the feasible design space of microrocket engines, and for optimizing the performance of such systems given current limitations in microfabrication technology.
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