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Nitrous oxide is suggested as a propellant for advanced small satellites.Due to its unique properties, nitrous oxide can be employed in cold-gas, monopropellant, bipropellant, and resistojet thrusters, thus covering all small satellite propulsion functions. This propellant does not require an expulsion system. The application of nitrous oxide is, therefore, considered beneficial for multi-mode small satellite propulsion system. To demonstrate nitrous oxide potential as a rocket propellant, catalytic decomposition is suggested for restartable monopropellant thruster concept. For this purpose nitrous oxide catalytic decomposition has been studied at the University of Surrey. The latest research results are presented in this paper. NOMENCLATUREE a -activation energy, kJ/mol AH°r -reaction enthalpy, kJ/mol m -mass flow rate, gm/s p -pressure, bar Q in -input energy, J Q h -input energy by heater, J Q dec -input energy released by nitrous oxide decomposition, J T-exhaust gas temperature, °C t -time, s

As the capability of small spacecraft (<500 kg) platforms increases, so do the complexities of their missions. The need for such functions as orbit raising and de-orbiting lead to an ever higher requirement for mission AV. Traditionally a cold gas nitrogen system would have been sufficient for a small satellite, however this is usually only suited to low AV missions, such as drag compensation. Propulsion technologies with higher specific impulse and higher storage densities are required.Hydrazine is the traditional step up, but its use has a high cost in terms of safety and infrastructure due to its toxic and flammable nature. This can negate the low cost advantages, which can be gained by having a small spacecraft. A green propulsion system with the equivalent performance of a hydrazine system would offer the advantages of safety, and handlability and hence lower overall propulsion system costs.This paper details the research being undertaken at the Surrey Space Centre to fill the gap between cold gas nitrogen and toxic, storable propellants. Research in the following areas will be described :-• A nitrous oxide monopropeilant thruster • A novel geometry hybrid rocket using common polymers as fuel with various oxidisers • A small kerosene / hydrogen peroxide bipropellant engine • A hydrogen peroxide storage programme • A hydrogen peroxide catalyst investigation

Small space vehicle propulsion is not only a technological challenge of scaling systems down, but also a combination of fundamental physical constraints. Several of these constraints related to small cold-gas, resistojet, and chemical thrusters are discussed in this paper. In particular, the trades of small size nozzle and reaction chamber designs are considered. NOMENCLATUREa -sonic velocity, m/s A c h s , A cs and A e -reaction chamber surface, crosssection, and nozzle exit areas, m 2 c -effective exhaust velocity, m/ŝand d th -chamber and nozzle throat diameters respectively, m F -thrust, N h -heat transfer coefficient, W/m 2 /K H °f-enthalpy of formation, J/mol I sp -specific impulse, s I ^t -minimum impulse bit, mN-s k( T) -heat transfer coefficient, J/m 2 /K 1 L -characteristic body dimension, m l ch and l noz -reaction chamber and nozzle lengths respectively, m m -propellant mass flow rate, kg/s n -number density, number of molecules per unit volume in the gas, #_of_molecules/m 3 p a , p ch and p e -ambient, chamber, and nozzle exit pressure respectively, Pa R = 8.31441 J/mol/K -universal gas constant Q g and £)/,/ -heat generation due to exothermic chemical reaction and heat losses out of the thruster, J Q and Q hl -.rates of heat generation inside and heat loss out of the reaction chamber respectively, W r , r c and r, -radius-vector, nozzle throat and inlet radii of curvature, respectively, m Re -Reynolds number T ch -chamber temperature, K t and t p -time and pulse duration respectively, s u and u e -flow and exit (or exhaust) velocities, m/s u -mean flow velocity, m/s U -maximum flow velocity, m/s V ch -reaction chamber volume, m 3 W-thruster mass, gm x -distance from beginning of boundary-layer, m a-nozzle divergence half-angle, deg. /?-nozzle convergence half-angle, deg. Y -specific heat ratio 1 American Institute of Aeronautics and Astronautics Downloaded by Duke University on September 24, 2012 | http://arc.aiaa.org |