A continuing demand exists to develop the capabilities of nanosatellites. A key element limiting the range of nanosatellite applications is the accommodation of a propulsion system. This research investigated this need and considered the miniaturisation of a monopropellant thruster. A literature review considered all aspects of micropropulsion together with enabling technologies. Assessment of present micropropulsion developments revealed that few would conform to the nanosatellite constraints. In addition the complexities associated with the miniaturisation of a propulsion system such as the modification of fluid flow, were highlighted. A review of the possible applications of a propulsion enhanced nanosatellite resulted in the creation of an inspection mission scenario. Assessment of present micropropulsion developments revealed none could fulfil the mission requirements, but a miniaturised chemical propulsion could. This led to the initiation of research to miniaturise a monopropellant thruster that would meet the mission requirements within the platform constraints. Hydrogen peroxide was selected as the propellant as it is considered to be a Green, non-toxic and non-carcinogenic propellant. The effect of scaling on the thermal characteristics of the thruster was evaluated using numerical models, which considered the effect of chamber wall thickness. It was concluded that a thin walled chamber should be combined with a heat-shield to allow radiated heat to be reflected back towards the decomposition chamber. The options available for the manufacture of a micropropulsion system were considered with respect to machining accuracy, materials and cost. There are two main options: Micro-Electro-Mechanical Systems (MEMS) technologies or micro conventional precision machining methods. It was concluded that at present the use of the latter was preferred as the level of machining accuracy is higher and conventional materials can be used. Following these analyses the detailed miniaturisation of the monopropellant thruster began, with a focus upon two major components: the decomposition chamber and the exhaust nozzle. The use of hydrogen peroxide as a rocket monopropellant was prevalent in the 1960's. Since then its use has waned in favour of other monopropellants such as hydrazine, which exhibit higher performance and improved storage characteristics. At that time significant research was conducted into the performance of hydrogen peroxide, but its use for low thrust applications was not considered. An analysis of available empirical data was conducted to determine the optimal configuration of a decomposition chamber in terms of the geometry of the decomposition chamber as well as the morphology of the catalyst bed. Two different catalyst morphologies were considered: a monolithic catalyst bed and a compressed powder catalyst bed. The monolithic morphology was based upon a ceramic foam substrate coated with a manganese oxide catalyst. Overall it generated good decomposition characteristics, but suffered from severe internal structural degradation. A compressed silver powder catalyst generated excellent decomposition characteristics and enabled the effect of catalyst bed length to be investigated as a function of decomposition chamber diameter. The results from these tests indicate that a compressed silver powder catalyst bed is a suitable alternative to silver gauzes for use in small diameter decomposition chambers, in addition the results showed that an optimal mass flow rate exists for each length of catalyst bed and a shorter bed is preferred due to thermal characteristics.