Design and Analysis of Combustion Chamber for HAN Based Mono Propulsion System Thruster for Spacecraft Application
This paper presents a preliminary dimensional study of combustion chamber using Hydroxyl Ammonium Nitrate (HAN) propellants for spacecraft application. The combustion chamber consists of two parts namely thrust chamber and Convergent-Divergent (C-D) nozzle. The design for combustion chamber is very much important because the chemical energy in the propellant released within this closed volume i.e., thrust chamber and gets expanded through the C-D nozzle part. So the chamber must be designed to provide a necessary space for the propellants to react and release maximum available energy and also it should prevent the loss of energy in the form of heat. The C-D nozzle should be optimally designed to allow the maximum conversion of enthalpy into kinetic energy. So, the thrust chamber and C-D nozzle are designed in an optimum size for releasing the heat to convert maximum available heat energy from the combustion of HAN propellant into exhaust velocity for HAN based monopropellant thruster. In this work the combustion chamber i.e. thrust chamber and C-D nozzle are designed at 16 bar pressure to generate a thrust of 11 N. CFD analysis is done to show the pressure and temperature variation in the combustion chamber modeled for 11 N thrust and chamber pressure of 16 bar for spacecraft application. From the analysis result it is found that monopropellant engine with the propellant combination of HAN+ Methanol+ Ammonium Nitrate + Water is suitable for design of Attitude & Orbit Control System (AOCS) thrusters.
Hydroxyl Ammonium Nitrate.Computational Fluid Dynamics.Attitude and Orbit Control System.Mono Propellant Thrusters.Combustion
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