Flow field numerical study within thrust chamber of aerospace small liquid propellant rocket engine

Paper number

IAC-06-C4.P.1.06

Author

Mrs. Kunmei Xu, Beijing University of Aeronautics and Astronautics, China

Coauthor

Prof. Guobiao Cai, Beijing University of Aeronautics and Astronautics, China

Coauthor

Mrs. Yuntao Zheng, Beijing University of Aeronautics and Astronautics, China

Year

2006

Abstract

Hypergolic bi-propellants are adopted in many small aerospace rocket engines, such as MMH/NTO. The combustion process of hypergolic propellants was complex extremely. So far, the reaction mechanism was known little. Nevertheless, several programs had been developed to simulate the flow field within the combustion chamber of hypergolic bi-propellants rocket engine, such as the PHEDRE program in France, the ROCFLAM software in Germany, and the HPRECSA code in China. In these programs, droplets vaporization models and chemical reaction models were different each other.
In this paper, the spray combustion and flow field within combustion chamber of a 10N thruster were studied by numerical simulation with various droplets vaporization models, and various chemical reaction models according to the characteristics of spray combustion and flow within small hypergolic bi-propellant rocket engine. 
Available test data, such as chamber pressure, thrust, et al, were used to evaluate the simulation results. The result revealed that all the models used in this paper can reasonably predict the spray combustion and flow field within combustion chamber of small hypergolic bi-propellant rocket engine. In theory, the hypergolic propellant droplets high pressure vaporization theory accord with the characteristics of hypergolic propellants vaporization. Whereas, droplets retain time in small combustion chamber was very short, the droplets size was small, and the chamber pressure was not too high. Because of all above reasons, it is feasible to use general droplets vaporization model. For gas turbulent combustion process, using finite rate/Eddy-Dissipation Model and 4-step finite rate model all can reasonably reflect the performance of combustion chamber. The former can’t predict intermediate product. Whereas, the latter considered the decompose process of MMH and NTO, so it can reflect actual combustion process to a certain degree.
Based on the study of theoretical models of spray combustion, the author made some application development in the CFD software FLUENT to develop a software for flow field prediction of the aerospace small rocket engine. The software can predict spray combustion two-phase flow field within the whole thrust chamber of the small hypergolic bi-propellant thruster, and gas flow field within nozzles of cold gas thruster, single-propellant thruster, and hypergolic bi-propellant thruster.

Abstract document

IAC-06-C4.P.1.06.pdf

Manuscript document

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