RT Conference Proceedings
T1 Test of a turbo-pump fed miniature rocket engine
A1 Scharlemann, C.
A1 Schiebl, M.
A1 Marhold, K.
A1 Tajmar, M.
A1 Miotti, P.
A1 Guraya, C.
A1 Seco, F.
A1 Kappenstein, C.
A1 Batonneau, Y.
A1 Brahmi, R.
A1 Lang, M.
AB The increasing application of microsatellites (from 10 kg up to 100 kg) for a rising number of various missions requires the development of suitable propulsion systems. Microsatellites have special requirements for a propulsion system such as small mass, reduced volume, und very stringent electrical power constraints. Existing propulsion systems often can not satisfy these requirements. The present paper discusses the development and test of a bipropellant thruster complying with these requirements. The main development goal of this effort was the utilization of ethanol in combination with hydrogen peroxide (H2O2) as a non-toxic propellant combination. The Turbo-Pump Fed Miniature Rocket Engine (TPF-MRE) is a bipropellant thruster consisting of four subsystems: the propellant pumps, a decomposition chamber with a monolithic catalyst, a turbine, and the thruster itself. The turbine is driven by the decomposed hydrogen peroxide and magnetically coupled with a power generator. The power produced is then used to generate a pressure head in order to deliver the propellant into the combustion chamber. This system therefore constitutes a self-sustaining system and does not rely on the limited power supply of a micro-satellite. Previous test have shown that although the thruster can be operated with ethanol and oxygen, it was not possible to ignite the thruster when utilizing hydrogen peroxide in a 70% concentration by weight. A minor redesign of the thruster and the test facility was therefore initiated. This redesign together with the use of hydrogen peroxide in higher concentration was speculated to improve this behavior. However, even though the monolithic catalysts were able to decompose hydrogen peroxide in a concentration of 87.5% with nearly 100% efficiency, it was not possible to ignite or operate the thruster. Subsequently, a thorough investigation of the baseline design and operational conditions of the thruster was conduced. It was found that the failure of the thruster to ignite is due to a combination of reasons. The combustion chamber length is too short to facilitate sufficient mixing of the propellants, making an ignition impossible or very difficult at least. Additionally, the combustion chamber pressure which was chosen such that it accommodates the performance of commercially available mircopumps is considered too low. This further deteriorates the conditions for which an ignition is feasible.
PB American Institute of Aeronautics and Astronautics Inc.
SN 1563478188
SN 9781563478185
YR 2006
FD 2006
LK https://hdl.handle.net/11556/2352
UL https://hdl.handle.net/11556/2352
LA eng
NO Scharlemann , C , Schiebl , M , Marhold , K , Tajmar , M , Miotti , P , Guraya , C , Seco , F , Kappenstein , C , Batonneau , Y , Brahmi , R & Lang , M 2006 , Test of a turbo-pump fed miniature rocket engine . in Collection of Technical Papers - AIAA/ASME/SAE/ASEE 42nd Joint Propulsion Conference . Collection of Technical Papers - AIAA/ASME/SAE/ASEE 42nd Joint Propulsion Conference , vol. 3 , American Institute of Aeronautics and Astronautics Inc. , pp. 2446-2455 , AIAA/ASME/SAE/ASEE 42nd Joint Propulsion Conference , Sacramento, CA , United States , 9/07/06 . https://doi.org/10.2514/6.2006-4551
NO conference
DS TECNALIA Publications
RD 30 jul 2024