Effects of Propellant Gases on Thermal Response of Solid Rocket Nozzle Liners

Abstract

The propellant formulations, the gas properties in motor chamber, and the aluminum oxide particle sizes for two kinds of solid propellants with approximately 20% aluminum powder have been investigated. The scanning electron microscope photographs of aluminum oxide particles taken from a nozzle entrance show that the aluminized polycaprolactone polyol propellant with 47% volumetric fraction ammonium perchlorate/hexanitro hexaazaisowurtzitane and the bimodal oxidizers, 200=5 m, can offer greater possibility for increasing aluminum agglomeration than the aluminized hydroxy-terminated polybutadiene propellant with 64% volumetric fraction ammonium perchlorate and the trimodal oxidizers, 400=200=6 m. The aluminized polycaprolactone polyol propellant with energetic plasticizers results in locally greater mechanical erosion in four circumferential sections of the nozzle entrance in line with grain slots, due to the impingement of large particles. However, the hydroxyterminated polybutadiene propellant results in greater thermochemical ablation at the blast tube, the throat insert, and the exit cone of rocket nozzle, due to 2.3 times more water and carbon dioxide oxidizing species in exhaust gases than the polycaprolactone polyol propellant. Nomenclature A = cross-sectional flow area c p = specific heat of gases at constant pressure d p = diameter of aluminum oxide particle E 0 = activation energy of a chemical reaction h = convective heat transfer coefficient K 0 = preexponential factor of a chemical reaction k = mass fraction of the solid residue of material M 0 = molecular weight of gas mixture M c = molecular weight of carbon P w = pressure at nozzle wall R = ideal gas constant T w = temperature at nozzle wall CO 2 = mole fraction of CO 2 H 2 O = mole fraction of H 2 O 0 = initial density Subscripts t = conditions at nozzle throat w = wall conditions