Burn rate (chemistry)

About: Burn rate (chemistry) is a research topic. Over the lifetime, 847 publications have been published within this topic receiving 8908 citations. The topic is also known as: Burning rate.

ALE3D Model Predictions and Experimental Analysis of the Cookoff Response of Comp B

Abstract: ALE3D simulations are presented for the thermal explosion of Comp B (RDX,TNT) in a Scaled Thermal Explosion Experiment (STEX). Candidate models and numerical strategies are being tested using the ALE3D code which simulates the coupled thermal, mechanical, and chemical behavior during heating, ignition, and explosion. The mechanical behavior of the solid constituents is represented by a Steinberg-Guinan model while polynomial and gamma-law expressions are used for the equation of state of the solid and gas species, respectively. A gamma-law model is employed for the air in gaps, and a mixed material model is used for the interface between air and explosive. A three-step chemical kinetics model is used for each of the RDX and TNT reaction sequences during the heating and ignition phases, and a pressure-dependent deflagration model is employed during the rapid expansion. Parameters for the three-step kinetics model are specified using measurements of the One-Dimensional-Time-to-Explosion (ODTX), while measurements for burn rate are employed to determine parameters in the burn front model. We compare model predictions to measurements for temperature fields, ignition temperature, and tube wall strain during the heating, ignition, and explosive phases.

Oxidizer Enhanced Hybrid Rocket Engine: Regression Rates and Performance

Abstract: Hybrid rocket motors have a significant advantage over solid and liquid rockets in that they are intrinsically safer. Keeping the liquid oxidizer separate from the solid fuel grain inhibits the two propellants from mixing rapidly and eliminates the potential of a serious explosion. However, the inability of the oxidizer and fuel to mix quickly in a typical hybrid rocket motor results in low propellant burning rates causing a reduction in performance. This is a result of a fundamental difference in the combustion process. A classic solid rocket motor produces a premixed (faster burning) flame, whereas a hybrid rocket engine produces a diffusion (slower burning) flame due to the oxidizer and fuel mixing mainly through diffusion processes. The addition of solid oxidizers into the hybrid fuel grain has the potential to increase the burning rate of the fuel grain. This causes the combustion zone to act more like that of a solid rocket motor. Also, if the solid oxidizer levels are kept below stoichiometric values, a rapid combustion reaction cannot take place in the solid fuel, preserving the safe characteristics of hybrid. This paper will explore the performance and safety implications associated with the oxidizer enhanced fuel grain by the use of a laboratory scale hybrid rocket engine test stand. The test stand has the ability to measure thrust, chamber pressure, line pressure and gaseous oxidizer flow rates. The thrust is measured via a strain gauge mounted on a spring steel thrust stand. Laboratory scale Hydroxyl-Terminated Polybutadiene (HTPB) /Ammonium Nitrate (NH4NO3) /Oxygen (O2) hybrid rocket engine performance is compared verses their solid oxidizer content. Analysis of the data produces empirical regression rates. To compare the relative safety of the hybrid propellants, samples of HTPB hybrid propellant were mixed with various percentages of ammonium nitrate. The samples were exposed to an ignition source and their ability to self sustain combustion was recorded.

Experimental Analysis of the Influence of Length to Diameter Ratio on Erosive Burning in a Solid Tubular Propellant Grain

Abstract: Erosive burning usually refers to the increase in the propellant burning rate caused by high velocity combustion gasses flowing over the propellant surface. It may seriously affect the performance of solid-propellant rocket motors [1]. A series of experiments had been made to study the effects of length to the diameter ratio in a single tubular propellant grain on the erosive burning phenomenon. In the same combustion pressure and different grain geometries, the burning pattern of AP1based propellant were recorded. Furthermore, pressure-time curve for each condition was obtained. The mean velocity gradient is obtained by some thermo-gas-dynamical analysis on experimental data. The results can be used for preliminary design of AP based tubular propellant rocket motors. This method may be used for other types of tubular solid propellants which defer in chemical formulation.

Self-extinguishable solid propellant

Abstract: Solid composite propellant compositions include at least one oxidizing agent, at least one binder, and at least one surfactant. The surfactant provides the solid propellant the property of being “self-extinguishing”, where the burning rate of the solid composite propellant as a function of pressure includes a negative pressure dependence portion, wherein the burning rate in the negative pressure dependence portion decreases with increasing pressure until a cutoff pressure is reached which results in extinguishment of the solid composite propellant. The solid composite propellant can also include at least one catalyst that modifies the burning rate of the solid composite propellant. Solid composite propellants can be extinguished without the need for depressurization by reaching a cutoff pressure, and with a tailored burning rate.