Design of Propellant Composite Thermodynamic Properties Using Rocket Propulsion Analysis (RPA) Software

Abstract

Rocket Propulsion Analysis (RPA) is software for predicting the performance of a rocket engine. It is usually used in conceptual and preliminary design. Heat capacity and specific impulse are two properties related to the performance of a propellant. This work aimed to design AP/HTPB-based solid propellant composite with various compositions and predict the heat capacity and specific impulse using the RPA software. The materials used were ammonium perchlorate (AP) as the oxidizer, Hydroxy-Terminated Polybutadiene (HTPB) as the fuel binder, Al powder as the metal fuel, and other additives. Four propellants with different formulations were prepared and tested for heat capacity and specific impulse. The experimental heat capacity was obtained using a differential scanning calorimeter (DSC), while the specific impulse was obtained using a bomb calorimeter. The same propellant formulations were used as the input to the RPS to predict the heat capacity and specific impulse. The results show that the experimental heat capacity of the propellant ranges from 1.576 to 4.08 J g–1 K–1, and the simulation result ranges from 1.78 to 3.48 J g–1 K–1. The overall average deviation is 16.3%. The predicted specific impulse at vacuum and sea level ranges from 231.3 to 234.0 s and from 219.8 to 220.9 s, respectively. Meanwhile, the experimental specific impulse at vacuum and sea level varies from 236.2 to 240.3 s and from 228.5 to 232.9 s, respectively. The overall average deviation is 3.7%. Therefore, the RPA is reliable for predicting specific impulse of propellant, but it is not accurate enough for predicting the heat capacity of propellant composite.