ACTIVE thrust termination or reversing of a separating motor are ways to facilitate the staging process and improve the performance of a multistage rocket [1–4]. This is done by opening a secondary nozzle located on the motor head and causing rapid depressurization of the motor chamber. During a rapid change in the chamber pressure the burning rate becomes a nonlinear function of pressure and its time rate of change. The strong dependence of the solid propellant burning rate on the chamber pressure results in a transient nonlinear behavior that may cause premature dynamic extinguishment of the motor. Solid propellants’ transient burning and extinguishment during rapid pressure change have been the subject of a number of studies and several transient burning models have been developed [5–8]. Zeldovich has proposed a comprehensive model based on the unsteady rate of heat transfer to the propellant surface and the time rate of change of the temperature distribution in the solid propellant [6]. This model requires specification of many experimentally determined parameters. The transient burning rate model proposed by Von Elbe and McHale is easy to implement and requires instantaneous pressure and its time rate of change for the regression rate prediction [5]. References [1, 2] studying motor transient behavior have used a control volume approach for the gas dynamic simulation of the motor’s internal ballistics. However during the active thrust termination (or reversing) process there is a considerable pressure variation along the motor port and there is a rapid propagation of expansion waves, through the thrust termination nozzles. These expansion waves play an important role in the transient thrust behavior and a more comprehensive flow model should be used for the internal ballistics’ simulation.