flow field numerical study within thrusrt chamber of aerospace small liquid propellant rocket engine

The spray combustion flowfield within thrust chamber of a 10N MMH/NTO thruster were investigated by numerical simulation used various droplet vaporization models, and various chemical reaction models. Several measured data from hot gas test of the thruster, such as chamber pressure, thrust, etc, were used to evaluate the simulation results. The results revealed that using all the models involved in this paper could reasonably predict the flow field within thrust chamber of small hypergolic bi-propellant rocket engine. In theory, the hypergolic propellants droplet high pressure vaporization theory maybe accords with the characteristics of hypergolic propellants droplet vaporization in high pressure condition even more. Whereas, droplet residence time in small combustion chamber was very short, the droplet size was small, and the chamber pressure was not too high. Because of all above reasons, it is feasible to use general droplet vaporization model. For gas turbulent combustion process, using finite-rate/Eddy-Dissipation Model and 4-step finite-rate reaction model both could reasonably reflect the main thrust chamber characteristics of a small hypergolic bi-propellant rocket engine. The former can't predict intermediate products. Whereas, the latter considered the decomposition process of MMH and NTO, so it can reflect actual combustion process to a certain degree. On the other hand, it may take more computational time.