نوع مقاله : مقاله پژوهشی
1 دانشکده مهندسی هوافضا دانشگاه صنعتی امیرکبیر
2 دانشگاه صنعتی شریف
3 پژوهشکده سامانه های حمل و نقل فضایی
عنوان مقاله [English]
Combustion of hypergolic propellants increases the specific impulse in the thrusters due to high temperature products. In this paper, the combustion process will be investigated through the axis of a bi-propellant thruster by an in-house code with chemical reaction mechanism. This code includes several models for injection, droplet evaporation, liquid film, combustion and heat transfer through liquid and gas films. The Astrium thruster with MMH as fuel and NTO as oxidizer has been simulated. By implementing a detail mechanism with 1619 steps, the thruster has been simulated at different total mass flow rates and results have been validated by experimental data. Then, injector dimension effects on the droplet evaporation and combustion have been investigated. Results show that by increasing the injector dimension, the droplet evaporation length increases, so the flame structure changes in the combustion chamber. Therefore, the combustion products enter the nozzle with higher temperature and as a result, the thruster specific impulse increases.