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.

Mitigation of Ignition-Induced, Two-Phase Flow Dynamics in Guns through the Use of Stick Propellants

Abstract: : Pressure waves arising in gun chambers from ignition-induced flow dynamics can be deleterious to a weapon system, either catastrophically through the failure of the gun or projectile, or more subtly through degraded ballistic reproducibility or projectile reliability. One way to improve the flow dynamics during the ignition phase of the interior ballistic cycle, and thus to mitigate pressure-wave development, is to increase the permeability of the propellant bed to ignition and combustion gases. A method by which this can be accomplished is through the use of stick propellants, which produce natural flow channels when bundled into a charge. We describe herein an investigation into the effects of stick propellant grain geometry on the development of pressure waves in guns. Specifically, several slotted- and unslotted-stick M30A1 propellants are considered. A series of preliminary studies of these propellants is briefly described, including closed-bomb testing and computer simulations of one- dimensional charges using a two-phase flow interior ballistic model. We present a detailed description of firing tests at ambient, reduced, and elevated temperatures using these propellants in full-bore, base-ignited, 155-mm bagged charges, specifically designed to promote the formation of pressure waves. by comparison with a previous study, the results indicate improved performance, as evidenced by decreased pressure-wave levels, in progressing from granular to stick propellants. It is also shown, for the lots tested, that the temperature coefficient of pressure, Delta P/Delta T, is dependent on the geometry, such that the ambient-to-hot coefficient for the slotted-stick propellant is twice that for the unslotted-stick propellant.

Continuous measurement of the burning rates of solid rocket propellants.

Abstract: : An experimental system for the direct and continuous measurement of solid rocket propellant burning rates was developed. A sample of solid propellant is bonded to a holder with an attached shaft which moves the propellant sample within a two-dimensional rocket motor. The flat burning surface of the propellant sample recedes normally as a servomechanism moves the propellant in a direction opposite to the receding propellant surface. The servomechanism operates in such a manner that the burning propellant surface is maintained at a fixed position within the rocket motor. Since the burning surface of the propellant remains fixed, the direct measurement of the velocity of the shaft yields the burning rate. The servomechanism incorporates a radioactive isotope feedback transducer system for detecting the position of the burning surface of the propellant sample. A collimated beam of gamma rays, provided by a 50 millicurie source of Cesium137, is transmitted through the rocket motor walls and is transformed into an electrical feedback signal by a scintillation probe coupled with a linear ratemeter. (Author)

Thrust Characterization of Very High Burning Rate Propellants

Abstract: : A series of closed bomb combustion tests were conducted on several very high burning rate (VHBR) propellants in which, in addition to the standard pressure measurements, the thrust produced by a burning propellant column was measured using a dynamic force gage located at the end of the propellant column. The results indicate that in several instances the total force imparted to the force gage substantially exceeded that due to the gas pressure alone. The difference is ascribed to the thrust or impulse resulting from very high apparent burning rates. In fact, an analysis of the data for two VHBR propellants indicates that effective burning rates from 100 to 300 m/s are required to produce the observed thrust levels. Such burning rates are not yet explainable by conventional combustion models. Additional physical processes such as convective flamespread or stress-induced surface fracture must be invoked to explain such high apparent burning rates.

Development of Ionic Liquid Monopropellants for In-Space Propulsion

Abstract: A family of new, low toxicity, high energy monopropellants is currently being evaluated at NASA Marshall Space Flight Center for in-space rocket engine applications such as reaction control engines. These ionic liquid monopropellants, developed in recent years by the Air Force Research Laboratory, could offer system simplification, less in-flight thermal management, and reduced handling precautions, while increasing propellant energy density as compared to traditional storable in-space propellants such as hydrazine and nitrogen tetroxide. However, challenges exist in identifying ignition schemes for these ionic liquid monopropellants, which are known to burn at much hotter combustion temperatures compared to traditional monopropellants such as hydrazine. The high temperature combustion of these new monopropellants make the use of typical ignition catalyst beds prohibitive since the catalyst cannot withstand the elevated temperatures. Current research efforts are focused on monopropellant ignition and burn rate characterization, parameters that are important in the fundamental understanding of the monopropellant behavior and the eventual design of a thruster. Laboratory studies will be conducted using alternative ignition techniques such as laser-induced spark ignition and hot wire ignition. Ignition delay, defined as the time between the introduction of the ignition source and the first sign of light emission from a developing flame kernel, will be measured using Schlieren visualization. An optically-accessible liquid monopropellant burner, shown schematically in Figure 1 and similar in design to apparatuses used by other researchers to study solid and liquid monopropellants, will be used to determine propellant burn rate as a function of pressure and initial propellant temperature. The burn rate will be measured via high speed imaging through the chamber s windows.