Search Results for « Rocket Propellants »

Abstract Solid rocket propellants containing nitramine are considerably superior to doublebase propellants, both as regards their performance and mechanical properties. The pressure exponent of non‐modified nitramine propellants is n ≥ 0.9. The possibility of changing the burning rate or, respectively, the pressure exponent has, however, only been realisable to a restricted extent up till now in propellants with an increased nitramine content. In the following propellant system containing nitramine, the effects of ammonium perchlorate on burning behaviour are studied: ammonium perchlorate (AP) hexogen (RDX) nitroglycerin (NG), or trimethylolethane trinitrate (TMETN) polyurethane binder (PU) AP concentrations already as great as 10% produce considerable changes in the burning behaviour of the propellants described above. It is possible to reduce the pressure exponent by the addition of ammonium perchlorate from n ≥ 0.9 to n ≤ 0.65. The burning rates may also be influenced with AP concentrations ≤ 20% and by varying the AP particle size by the factor of 2. All the propellants prepared were easily castable and showed exceptionally good viscoelastic properties (strain at break ε R > 200%) in the temperature range between −40 °C to + 50 °C. The thermal chemical stability is not influenced negatively in any way by the combination of nitric acid esters and ammonium perchlorate.

Abstract The cookoff behaviour of several plasticized HTPB/AP and HTPB/RDX propellants has been evaluated using the standard Small‐Scale Cookoff Bomb (based on the NWC design) and a modified version of the standard equipment (whereby the reaction violence can be measured quantitatively) to investigate the effects of (1) plasticizer level on the thermomechanical properties and its role in moderating the cookoff response, and (2) heating rate. A change from slow to fast cookoff behaviour was observed for the plasticized HTPB/RDX propellants; the measured pressures/energy output at the slow heating rate were much lower than those at the fast heating rate for the same propellants. The thermal distribution/gradient at various positions and depths in the cylindrical propellant specimen, during fast and slow cookoff, were measured in order to understand the reversal in the severity of the response shown by the plasticized RDX propellants when the heating rate was changed. Thermomechanical properties of the propellants were measured and correlated with cookoff behaviour, to provide further insight into the cookoff mechanisms.

Abstract Rocket propellants with reduced smoke and high burning rates recommend themselves for use in a rocket motor for high accelerating tactical missiles. They serve for an improved camouflage on the battle field and may enable guidance control due to the higher transmission of their rocket plume compared to traditional aluminized composite propellants. In this contribution the material based ranges of performance and properties of three non aluminized rocket propellants will be introduced and compared to each other. The selected formulations based on AP/HTPB; AP/PU/TMETN and AP/HMX/GAP/TMETN have roughly the same specific impulse of I SP = 2430 Ns/kg at 70:1 expansion ratio. The burning rates in the pressure range from 10–18 MPa vary from to 26–33 mm/s for the AP/HTPB propellant, 52–68 mm/s for the formulation based on AP/PU/TMETN and 28–39 mm/s for the propellant based on AP/HMX/GAP. With 58% and 20% AP‐contents the propellants with nitrate ester plasticizers create a much smaller secondary signature than the AP/HTPB representative containing 86% AP. Their disadvantage, however, is the connection of high performance to a high level of energetic plasticizer. For this reason, the very fast burning propellant based on AP/PU/TMETN is endowed with a low elastic modulus and is limited to a grain configuration which isn't exposed too much to the fast and turbulent airstream. The mechanical properties of the AP/HMX/GAP‐propellant are as good or better as those of the AP/HTPB propellant. The first one exhibits the same performance and burn rates as the composite representative but produces only one fifth of HCl exhaust. For this reason it is recommended for missile applications, which must have high accelerating power together with a significantly reduced plume signature and smoke production.

Abstract Experiments to determine the flow pattern in an indoor rocket test stand during static rocket motor firing are described. After a theoretical treatment of the flow properties, a description is given of the experimental set‐up. Tests were performed with short burning (1.8 s), low thrust (2,8 kN) rockets as well as with a long burning (40 s) rocket motor with higher thrust (9.5 kN). With the short burning rocket motors flow velocities obtained were between 3.3 m/s and 7.9 m/s. The stationary flow velocities of the long burning rocket motor amounted up to 24 m/s, while the increase in temperature of the flow was appr. 200 K. Experimental flow velocity appeared to be lower while the experimental flow temperature appeared to be higher than the theoretical values, mainly due to after burning of the fuel‐rich combustion gases.