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The design of an efficient and reliable ignition system represents a challenge for propulsion engineers especially for hybrid rocket motors where the fuel grain and oxidizer are separated. In the current work, a simple design model is used to size the igniter, and to select the solid propellant type and mass. The analysis is based on the heat transfer model of the main rocket motor grain, thermochemical model of the ignition process and an empirical regression rate model of the grain. Following this analysis, a design point of the igniter including the chamber pressure, igniter dimensions and a propellant type are selected to perform the required task and to ensure the ignition of the main engine. Series of experiments are performed for different propellant types, different propellant mixture ratios and different grain size to select the optimum igniter. The flame length, temperature and ignition time are observed to assess the performance of the igniter. Finally, the selected igniter design is tested successively in the hybrid rocket motor.

DESIGN AND TESTING OF THE IGNITION SYSTEM FOR HYBRID ROCKET MOTOR