Study And Design Of A Solid Rocket Motor's Igniter

Solid rocket motors have been widely used and developed for the both tactical missions and the civilian applications due to their high performances,simplicity and reliability.For the past few decades,different types of solid propellants rocket motors are used for different missions.The shortcomings of solid rocket motors are that the thrust development is pre-defined by the primary shape of the propellant grain and the free burning surface.Hence,the concept of a dual pulse solid rocket motor has been developed,where more than one pulse is used in the motor burning time.The dual pulse solid rocket motor design contains of two burning chambers(first pulse and second pulse),separated by a bulkhead,designated as pulse separation device(PSD)and a nozzle.The ignition transient of solid rocket motors(SRM)represents the ignition transient includes the entire time from initiation of the ignition signal to the attainment of steady state conditions of the rocket motor operation Generally,the period of the ignition transient varies of tens of milliseconds for SRM.A better understanding of the ignition transient of a solid propellant rocket motor becomes increasingly more important in the aim to get more sophisticated solid propellant rocket motors.In the design and development of solid propellant rocket motors,the use of numerical simulation able to know and predict the ignition transient phenomenon of a given solid rocket motor.Numerical simulations have been carried out with the application of computational fluid dynamics(CFD)methods to simulate the ignition transient of SRM.CFD is one of fluid mechanics disciplines that uses numerical analysis and data structures to simulate and analyze the physical phenomena that can happen to the fluid flow.This master's degree research focus on the dual pulse solid-propellant rockets-motor(DPSRM)ignition transient prediction theory with computer simulation for a practical configuration.The 2D unsteady axisymmetric Navier-Stokes equation systems were numerically solved by using the Advection Upstream Splitting Method(AUSM)scheme in this master research based on ANSYS FLUENT.A numerical calculation model of the dual pulse solid rocket motor ignition transient process,including the physical model and mathematical equation model through the use of the user-defined function(UDF)code on the solid propellant mass flow rate boundary condition or so-called the adding mass source zone has been done.The numerical results demonstrate the flow field in the combustion chambers,the temperature distribution and the pressure evolution in the solid rocket motor combustion chambers.The rupture of the pulse separation device related to the high pressure generated during the second pulse ignition.The igniter's mass flow rate has a significant effect on the rupture of the pulse separation device.