The hybrid rocket motor is a kind of chemical propulsion system that has been recently given serious attention by various industries and research centers. The relative simplicity, safety and low cost of this motor, in comparison with other chemical propulsion motors, are the most important reasons for such interest. Moreover, throttle-ability and thrust variability on demand are additional advantages of this type of motor. In this paper, the result of an internal ballistic simulation of hybrid rocket motor in a zero-dimensional form is presented. Further to validate the code, an experimental setup was designed and manufactured. The simulation results are compared with the experimental data and good agreement is achieved. The effect of various parameters on the motor performance and on the combustion products is also investigated. It is found that increasing the oxidizer flow rate, increases the pressure and specific impulse of the motor; however, the slope of the specific impulse for the high flow rate case reduces. In addition, by increasing the combustion chamber pressure, the specific impulse is increased considerably. The initial diameter of the fuel port does not have significant effect on the pressure chamber and on the specific impulse. Addition of a percentage of an oxidizer like ammonium perchlorate to the fuel increases the specific impulse linearly.
An Analysis and Reduction Design of Combustion Instability Generated in Hybrid Rocket Motor