Abstract
This study is a mathematical model to obtain the characteristics performance of magnesium metal (powder) and carbon on a potassium nitrate-sucrose (KNSU) solid propellant formulation. Characterization of propellant is, as a general rule, important to determine its performance before it can be suitable for use for a rocket flight or any mission. Method of ballistic load cell evaluation was used to validate results and a mathematical model using the combustion exhaust products was solved to obtain the characteristics performance parameters of the propellant. The carbon constituent which acts as an opacifier and coolant was kept constant at 2% in order to arrest some of the heat during the combustion process and helped to lower the combustion temperature, because high combustion temperature could lead to combustion chamber rupture or failure. The effect of addition of magnesium which was optimized for 3% in the formulation contributed significantly in improving the overall performance of the propellant. The utilization of magnesium in KNSU propellant provided higher values parameters and better performance compared to when not included. This was confirmed with the model equations. The propellant combustion products equation was used to model and obtain the characteristics performance parameters. This gave propellant specific impulse (122.9s), combustion temperature (1821K), heat ratio (1.1592), molecular weight (36.89g/mole), propellant density (1912.5kg/m3) and characteristics velocity (1000m/s) result while maintaining the same chamber pressure.
Highlights
All Solid Rocket Motors have the same form of principle but there is universal design method to utilize for design of various subsystems of Solid Rocket Motors such as the propellant
The aim of this work is to improve the performance of a KNSU rocket propellant fuel using magnesium metal and carbon
A rocket motor operates on the basic principle of converting heat energy, from chemical reactions, to kinetic energy
Summary
All Solid Rocket Motors have the same form of principle but there is universal design method to utilize for design of various subsystems of Solid Rocket Motors such as the propellant. For applications of rocket either for unguided military mission for upper atmosphere or any launch vehicle for space missions, the optimum Solid Rocket Motors design has to satisfy optimum total impulse, an optimium thrust-time profile, an optimum nozzle configuration an optimum chamber pressure and a preferred solid propellant grain configuration [1,2]. Different formulations of various combinations of chemical constituents would give propellants of different physical and chemical properties, as well as combustion characteristics and performance [4,5].
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