Development Of Microthruster Research Articles

Micro-fibre based Porous Composite Propellants with High Regression Rates

Harnessing energy at micro-scale from high energy sources has gained significance in recent times for space propulsion and other applications. Conventional solid rocket propellants have advantages in terms of being efficient, compact and safe to handle, though with much lower regression rates as compared to solid explosives. An approach to high regression rates in composite propellants is demonstrated in the present work by the enhancement of fuel-oxidiser interaction, and by the incorporation of micro-scale porosity into the propellant grain. The porous polystyrene-ammonium perchlorate grain designed in this work, based on electrospun micro-fibres and aqueous impregnation, exhibits burning rates more than 25 times as compared to the non-porous grain. Such high regression rates using insensitive propellant compositions have practical implications in the development of micro-thrusters, and in gas generating devices such as MAV launch systems and turbine starters. Detailed preparatory procedure, characterisation techniques, and flame regression studies are included in this paper.

Development of a chemical microthruster based on pulsed detonation

The development of a microthruster based on gaseous pulsed detonation is presented in this study. The feasibility of cyclic valveless pulsed detonation at frequencies over 100 Hz is first experimentally investigated in a microchannel with 1 mm × 0.6 mm rectangular cross-section. Highly reactive ethylene/oxygen mixtures are utilized to reduce the time and distance required for the reaction wave to run up to detonation in a smooth channel. High-speed visualizations have shown that the reaction waves reach detonative state through highly repeatable flame acceleration and deflagration-to-detonation transition processes in the channel. The validated concepts are implemented for the development of an integrated pulsed detonation microthruster. The microthruster was fabricated using low temperature co-fired ceramic tape technology. The volume of the reaction channel in the microthruster was 58 mm3. Spark electrodes and ion probes were embedded in the ceramic microthruster. The channel and via holes were fabricated using laser cutting techniques. Ion probe measurements showed that the reaction wave propagated at velocities larger than 2000 m s−1 before reaching the channel exit. The pulsed detonation microthruster has been successfully operated at frequencies as high as 200 Hz.

Design, fabrication and performance evaluation of a vaporizing liquid microthruster

A recent application domain of MEMS technology is in the development of microthrusters for micro-/nanosatellites. Among the various types of MEMS microthruster developed so far, the vaporizing liquid microthruster (VLM) has been widely explored for its capability to produce continuously variable thrust in the micro-Newton (µN) to mili-Newton (mN) range. This paper reports the design and experimental validation of silicon MEMS VLM consisting of a microcavity, inlet channel and converging–diverging (C-D) in-plane exit nozzle integrated in two micromachined bonded chips and sandwiched between two p-diffused microheaters, located at the top and bottom surface of the device. Structural configuration was designed using simple analytical equations to achieve maximum thrust force by controlling the inlet propellant flow and heater power of VLM in an efficient way. In addition, a 3D model using a computational fluid dynamics technique was constructed to simulate the aft section of VLM for the investigation of its aerodynamic behavior. The device fabrication and testing have been briefly described. The fabricated VLM is capable to produce 1 mN thrust using maximum heater power of 3.6 W at a water flow rate of 2.04 mg s−1 using an in-plane C-D exit nozzle of throat area 130 µm × 100 µm. A detailed thrust force measurement was carried out with the variation of input heater power for different mass flow conditions and exit to throat area ratio of the exit nozzle, and the results were interpreted with the theoretical model. The model gives considerable physical insight in the operation of the VLM. Finally, a performance comparison with other published VLM results indicates that the present design can yield comparatively more thrust force with much less input power. A performance comparison with other published VLM results indicates that the present design can achieve improved performance by integrating two heaters with appropriate chamber volume in respect of propellant flow rate and input power for obtaining a supersaturated dry stream.

Problems of microthruster development

A description of a microthruster operating on the single-component propellant-hydrazine with the Shell-405 catalyst is presented. The hydrazine decomposition reaction at the specified parameters occurs on the surface. The gas temperature entering the nozzle actually depends on the catalyst temperature. It is established that if the temperature of hydrazine decomposition products at the catalyst outlet is lower than 400 K, ammonia condensation takes place in the supersonic nozzle part. The temperature distribution in the working chamber microchannel is calculated, the microthruster thrust and specific impulse are evaluated on the basis of a physico-mathematical model. A fundamental possibility of obtaining 10−3−10−2 N level thrust is shown.

Microthruster development in France under direction of the Centre National d'Etudes Spatiales /CNES/

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