For enhanced range, higher payload capacities and for miniaturized propulsion systems, today's strategic and tactical weapon system designers demand for higher density and specific impulse of the propellant. In order to enhance the density impulse of HTPB/DOA/RDX/AP/Al based composite propellant, studies have been carried to replace conventional HTPB/DOA binder system with hydroxyl terminated block copolymer of polybutadiene and ϵ‐caprolactone with NG as plasticizer. Total eight numbers of compositions were formulated with varying content of RDX. Both binder systems were compared in propellant compositions by evaluating various physical, thermal and ballistic properties. Various rocket performance parameters of each formulation were theoretically predicted by NASA CEC‐71 program and burning rate was measured in pressure ranges of 3‐7 and 7‐11 MPa by the acoustic emission technique. In addition, density, viscosity build up, calorimetric values, thermal decomposition and sensitivity parameters of each composition were also assessed and compared. In an outcome, it was concluded that HTBCP25/NG based propellant compositions enhance the density by 4.4–5 % and calorimetric values by 12–15 % as compared to HTPB/DOA based compositions. Strand burning rate data show enhancement of burning rate by 40–70 % at 7 MPa pressure in HTBCP25/NG based compositions. Impact and friction sensitivity data also revealed their utility in propellant compositions for future applications.
In the present study, the mathematical prediction with the Paul‐Mukunda model is carried out for maximum pressure rise with erosive burning in multi‐grain solid rocket propellant. For this study, a cluster of 7 tubular solid double‐base propellant grains is selected. The erosive burning model has given a fair idea of the maximum pressure rise in rocket motors. The maximum pressure rise due to the erosive burning effect is quite a lot higher than without the erosive burning effect. The erosive burning model helps in studying maximum pressure rise for various configurations of propellant grains. It is found that lowering the outer diameter (OD) of propellant grains is giving low maximum pressure in rocket motor in comparison to increasing the inner diameter (ID) of propellant grains. Ap/At (port area to throat area) ratio is maintained same for both the cases. Although in both the cases predicted maximum flow velocity of propellant gases is almost same. It shows that keeping the same Ap/At ratio and erosive burning effect, the maximum pressure is reduced significantly by lowering OD of propellant grains than increasing the ID of the propellant grains in rocket motor. This study will help to reduce the maximum pressure rise in rocket motors for safe working.
In a systematic study to compare the effects of the values of burning rate and pressure exponent in RDX‐AP based composite propellant, various compositions with varying percentages of zirconium carbide (ZrC) and zirconium silicate (ZrSiO4) were formulated to select a suitable candidate. Various rocket parameters of each formulation were theoretically predicted by the NASA CEC‐71 program and the burning rate was evaluated in pressure range of 3–11 MPa. In addition, density, sensitivity, and thermal properties of compositions having maximum effects on pressure exponent’s values were also evaluated. It was concluded that ZrSiO4 enhances the pressure exponent “n” value substantially, whereas ZrC doesn’t have significant effects on it as compared to base composition and also provides higher density values of composite propellant formulated.
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