Showing papers on "Burn rate (chemistry) published in 2002"

Hydrogen Addition For Improved Lean Burn Capability of Slow and Fast Burning Natural Gas Combustion Chambers

Abstract: One way to extend the lean burn limit of a natural gas engine is by addition of hydrogen to the primary fuel. This paper presents measurements made on a one cylinder 1.6 liter natural gas engine. Two combustion chambers, one slow and one fast burning, were tested with various amounts of hydrogen (0, 5, 10 and 15 %vol) added to natural gas. Three operating points were investigated for each combustion chamber and each hydrogen content level; idle, part load (5 bar IMEP) and 13 bar IMEP (simulated turbocharging). Air/fuel ratio was varied between stoichiometric and the lean limit. For each operating point, a range of ignition timings were tested to find maximum brake torque (MBT) and/or knock. Heat-release rate calculations were made in order to assess the influence of hydrogen addition on burn rate. Addition of hydrogen showed an increase in burn rate for both combustion chambers, resulting in more stable combustion close to the lean limit. This effect was most pronounced for lean operation with the slow combustion chamber.

Multiphase cfd simulations of solid propellant combustion in gun systems

Abstract: The Army Research Laboratory has developed a scaleable, 3D, multiphase, computational fluid dynamics (CFD) code with application to gun propulsion (interior ballistics) modeling. The NGEN3 code, which incorporates general continuum equations along with auxiliary relations into a modular code structure, is readily transportable between computer architectures and is applicable to a wide variety of gun propulsion systems. Two such systems are the Army's Modular Artillery Charge System (MACS) and the Future Combat System (FCS). The MACS is being developed for indirect fire cannon on both current and developing (e.g., Crusader) systems. The efficiency of the MACS charge is dependent on proper flamespreading through the propellant modules; a process that has been repeatedly demonstrated in gun firings, successfully photographed using the ARL ballistics simulator, and numerically modeled using the NGEN3 code. The FCS requires weapons systems exhibiting increased range and accuracy. One of the technologies under investigation to achieve these goals is the electrothermal-chemical (ETC) propulsion concept, in which electrically generated plasma is injected into the gun chamber igniting the high-loadingdensity (HLD) solid propellant charge. NGEN3 code development and application to the MACS and FCS is currently a DoD HPC Challenge Project (No. 112) and is being greatly advanced by access to the DoD high performance computers (HPCs). Associate Fellow AIAA. Propulsion Physics Team Leader, Ballistics and Weapons Concepts Division, Weapons and Materials Research Directorate. Mechanical Engineer, Propulsion Physics Team, Ballistics and Weapons Concepts Division, Weapons and Materials Research Directorate. This paper is declared a work of the U.S. Government and is not subject to copyright protection in the United States. INTRODUCTION A solid propellant gun system consists of a reaction chamber connected to a gun tube through which a projectile is guided once propelled by pressurization of the chamber. Chamber pressurization is accomplished by placing a solid propellant (SP) charge in the chamber and igniting it by various means. Current SP charges are generally complex structures consisting of hundreds or even thousands of distinct regularly formed (e.g., spherical, cylindrical) grains, which may be loaded in either regular or random arrangements. In addition to small-scale voidage between grains (i.e., porosity) many charges also contain large-scale voidage (i.e., ullage), which surrounds the entire charge (such as when the charge does not fill the entire chamber volume) or separates distinct subcharges (i.e., increments or modules) that together comprise the whole charge. The addition of energy to the chamber, usually near the gun breech, or rearmost end of the chamber, and in some cases through a tube extending along the centerline of the chamber, ignites the SP. In general, all of the grains are not ignited simultaneously, but an ignition flame spreads from the breech to the projectile base. The burning of the SP transforms chemical energy into heat as hot gases evolve from the surface of each grain of propellant. Initially the projectile resists movement allowing the pressure in the chamber to climb rapidly. Since the burn rate of the propellant is proportional to the pressure, hot gases are produced at an accelerated rate until peak pressure is reached in the chamber. Movement of the projectile down the gun tube, usually slight before peak pressure and much more significant afterwards, causes the chamber volume to increase, and generates rarefaction waves, which lower the pressure and thus the burn rate of the propellant. Upon ignition and burning, the gas dynamic flowfield in the gun chamber takes on a highly complex structure that includes the dynamics of propellant motion and combustion and various gas dynamic flow phenomena such as turbulent mixing, highly transient pressure waves, steep gradients in porosity and temperature, nonideal thermodynamics, and gas generation.

Investigation of the Burning Behavior of Cryogenic Solid Propellants

Abstract: Cryogenic Solid Propellant (CSP)-technology is a new approach to develop more powerful rocket motors. CSPs include the advantages of classical solid propellants to save weight as well as those of a high energy content and safety of modern liquid propellants. The charges consist of liquid and/or gaseous fuels and oxidizers, both frozen. Two main versions of CSP-technology can be realised: 1. Mono-CSPs show the burning behavior of solid propellants. Experiments with mono-CSPs have been carried out under inert pressure conditions in a window bomb. Mono-CSPs have a stable burning behavior with a constant regression rate which follows the Vieille's law under varying pressure conditions. 2. The advantage of high safety is obtained by assembling oxidizer and fuel in sandwich configurations. The grain geometry governs the burning behavior. Such systems can be externally controlled, e.g. by the heat from a gas generator or they can work self-sustained. A Rod-in-Matrix burner shows self-sustained combustion in an inert pressure atmosphere with overall burning rates in a similar range as solid rocket propellants which obey also a Vieille-like pressure law. Disc stack burners have also been investigated, the combustion of which is strongly dependent on the disc thickness. For a short time Mach's nodes have been observed in the exhaust plume of a disc stack burner. Currently, the temperature ranges are limited to the boiling temperature of liquid nitrogen. Therefore, liquid oxidizers like H2O2 have been used. However, for the first time a propellant strand of polymer rods embedded in solid oxygen was prepared and burnt. The experiments with CSPs end in the combustion of a small rocket motor showing no serious technical obstacles. Simplified models based on the heat flow equation can simulate the burning characteristics of the frozen energetic materials including phase transitions.

Direct-fired heater

Abstract: A combustion system for a portable forced air heater having two frusta-conical sections attached to a circular burner tube, wherein each frusta-conical section has pre-determined vent hole patterns that allow the gas heater to have a variable burn rate.

Solid propellant formulations and methods and devices employing the same for the destruction of airborne biological and/or chemical agents

Abstract: High temperature incendiary (HTI) devices and methods destroy biological and/or chemical agents. Preferably, such HTI devices include dual modal propellant compositions having low burn rate propellant particles dispersed in a matrix of a high burn rate propellant. Most preferably, the HTI device includes a casing which contains the dual modal propellant and a nozzle through which combustion gases generated by the ignited high burn rate propellant may be discharged thereby entraining ignited particles of the low burn rate propellant. In use, therefore, the high burn rate propellant will be ignited using a conventional igniter thereby generating combustion gases which are expelled through the nozzle of the HTI device. As the ignition face of the propellant composition regresses, the low burn rate particles will similarly become ignited. Since the low burn rate particles burn at a lesser rate as compared to the high burn rate propellant in which such particles are dispersed, the ignited particles per se will be expelled through the nozzle and will therefore continue to burn in the ambient environment. Such continued burning of the particles will thereby be sufficient to destroy chemical and/or biological agents that may be present in the ambient environment.

Pyrotechnic Burn Rate Measurements: Strand Testing

Abstract: Burn rate is one of the most fundamentally important properties of pyrotechnic materials. While burn rate may be measured as a mass burn rate (mass of pyrotechnic composition consumed per unit time, e.g., g/s), linear burn rate is most commonly used. Linear burn rate can be defined as the distance the burning surface of a pyrotechnic composition advances inwardly (perpendicular to the burning surface) per unit time, and typically would be reported as inches per second (or mm/s). Even for a specific pyrotechnic material with a defined composition (including prescribed particle size and shape) there are a number of factors that will affect its burn rate. Generally the most important factors, ranked roughly in order of importance, are: ambient pressure, loading pressure (composition density), temperature, and burning surface area. Accordingly, for burn rate measurements to be most useful, they must take each of these additional factors into consideration.

Solid rocket motor fire tests: Phases 1 and 2

Abstract: JHU/APL conducted a series of open-air burns of small blocks (3 to 10 kg) of solid rocket motor (SRM) propellant at the Thiokol Elkton MD facility to elucidate the thermal environment under burning propellant. The propellant was TP-H-3340A for the STAR 48 motor, with a weight ratio of 71/18/11 for the ammonium perchlorate, aluminum, and HTPB binder. Combustion inhibitor applied on the blocks allowed burning on the bottom and/or sides only. Burns were conducted on sand and concrete to simulate near-launch pad surfaces, and on graphite to simulate a low-recession surface. Unique test fixturing allowed propellant self-levitation while constraining lateral motion. Optics instrumentation consisted of a longwave infrared imaging pyrometer, a midwave spectroradiometer, and a UV/visible spectroradiometer. In-situ instrumentation consisted of rod calorimeters, Gardon gauges, elevated thermocouples, flush thermocouples, a two-color pyrometer, and Knudsen cells. Witness materials consisted of yttria, ceria, alumina, tungsten, iridium, and platinum/rhodium. Objectives of the tests were to determine propellant burn characteristics such as burn rate and self-levitation, to determine heat fluxes and temperatures, and to carry out materials analyses. A summary of qualitative results: alumina coated almost all surfaces, the concrete spalled, sand moisture content matters, the propellant self-levitated, the test fixtures worked as designed, and bottom-burning propellant does not self-extinguish. A summary of quantitative results: burn rate averaged 1.15 mm/s, thermocouples peaked at 2070 C, pyrometer readings matched MWIR data at about 2400 C, the volume-averaged plume temperatures were 2300–2400 C with peaks of 2400–2600 C, and the heat fluxes peaked at 125 W/cm2. These results are higher than other researchers’ measurements of top-burning propellant in chimneys, and will be used, along with Phase 3 test results, to analyze hardware response to these environments, including General Purpose Heat Sources (GPHS) and Radioisotope Heater Units (RHU). Follow-on Phase 3 tests burning propellant blocks up to 90 kg will be briefly described.

Computation of combustion in hybrid rocket motor using a boundary layer theory

Abstract: A boundary layer numerical solution was used for the computation of combustion in hybrid rocket motor according to the boundary burning theory. Various value of mass transfer number and the oxidizer mass flux were considered in the calculation. The regression rate of the solid fuel was formulized based on the calculation. The results show good agreement with the related references and useful for further study.

One-dimensional combustion instability studies with moving boundaries in an end-burning test motor

Abstract: A quasi-steady one-dimensional arbitrary moving boundary code is progressively developed for the objective of solving; full transient, moving boundary, rocket motor internal flow field problem. In this analysis, the propellant surface is modeled as moving, contrary to the common approach that assumes a stationary propellantburning surface. A one-dimensional Godunov type exact Riemann Solver is developed that can handle the boundary conditions for advancing and retarding walls. Moving propellant boundaries are handled with fixed grid approach by clipping the boundary inside grid points. The classical time-dependent inflow/outflow characteristic boundary conditions are implemented for exhaust. The mass injected from the propellant surface is modeled as the mass generated inside the control volume, represented in the source terms of the Euler equations. The resulting code is verified for moving boundary test cases, nozzle type geometries, and small test motors. Finally solid propellant combustion instability prediction capability is searched for an end-burning test motor configuration by comparing different burn rate models. Both steady state burning rate and a transient burning rate law is applied as a boundary condition. Acoustic oscillations at different frequencies are observed during the operation time. The pulsed oscillations occurred in higher modes. Results are discussed for fast and slow regression rates, long motors and steady state pressure levels. Frequency analysis of the flow variables and burning rate variations are done real time during the computations using non-stationary data analysis techniques. Pressure response function of the model propellant is assumed known from experiments available from the literature.

An approach to decreasing the burning rate of NEPE propellant

Abstract: Three factors,including temperature profile of fizz zone,heat of reaction of condensed phase and temperature of burning surface have been found to affect the burning rate of NEPE propellant by the light of analysis of the heat balance on the burning surface.The possible methods to decrease burning rate are presented,the action of burning rate depressants and the effect on energy of propellant are researched.The results show that the approach for decreasing the burning rate of propellant is feasible.