Modelling of Hydrazine droplet evaporation and combustion in small rocket thrusters

Abstract

Hydrazine based bi-proppelant thrusters are currently widely applied and in use for standard satellite propulsion systems. Monomethylhydrazine (MMH) and nitrogen tetroxide (NTO) is a hypergolic engine fuel oxidizer combination which has been studied extensively in experiments[1]. However, numerical models of the combustion process in the combustion chamber are sparse and limited to multi-physics engineering models [2, 3] without validation. The available finite rate chemistry mechansims [1{5] are very stiff and not easily applied in computational fluid dynamics simulations. The challenges in these flows are the different length and time scales, the very complex chemistry, the pulsed operation with incomplete combustion and the multi-phase flow nature present in the combustion chamber. In order to allow prediction of thruster plume contamination and other operational parameters, a good numerical model of the thruster and its boundary conditions is necessary. This study aims at characaterizing the available reaction mechanisms for use in CFD application and combining enginering models of the droplet evaporation within the chamber with a gas phase based combustion model in order to obtain better prediction of the exit products of representative small thrusters. For this purpose the in house software TAU is used to couple[6] a Lagrangian droplet model including evaporation with finite rate chemistry in the Eulerian phase. Using this approach the respective evaporation and combustion characteristics of Hydrazine droplets in representative steady state combustion environments are investigated.