Showing papers on "Burn rate (chemistry) published in 2007"

Reaction Propagation of Four Nanoscale Energetic Composites (Al/MoO3, Al/WO3, Al/CuO, and B12O3)

Abstract: Nanoscale composite energetics (also known as metastable intermolecular composites) represent an exciting new class of energetic materials. Nanoscale thermites are examples of these materials. The nanoscale thermites studied consist of a metal and metal oxide with particle sizes in the 30-200 nm range. They have potential for use in a wide range of applications. The modes of combustion and reaction behavior of these materials are not yet well understood. This investigation considers four different nanoaluminum/metal-oxide composites. The same nanoscale aluminum was used for each composite. The metal oxides used were molybdenum oxide (MoO 3 ), tungsten oxide (WO 3 ), copper oxide (CuO), and bismuth oxide (Bi 2 O 3 ). The reaction performance was quantified by the pressure output and propagation velocity using unconfined (or open burn) and confined (burn tube) experiments. We examine the optimization of each composite in terms of pressure output and propagaton speed (or burn rate) for the open burn experiment. We find that there is a correlation between the maximum pressure output and optimum propagation speed (or burn rate). Equilibrium calculations are used to interpret these results. We find that the propagation speed depends on the gas production and also on the thermodynamic state of the products. This suggests that condensing gases or solidifying liquids could greatly enhance heat transfer. We also vary the density of these composites and examine the change in performance. Although the propagation wave is likely supersonic with respect to the mixture sound speed, the propagation speed decreases with density. This behavior is opposite of classical detonation in which propagation (detonation) speed increases with density. This result indicates that the propagation mechanism may differ fundamentally from classical detonations.

Development and characterization of high performance solid propellants containing nano-sized energetic ingredients

Abstract: This paper addresses the development of a pair of layered solid propellants suitable for use in a fast-core gun-propellant charge application. A baseline propellant combination was formulated using RDX particles and thermoplastic-elastomer binder as the major ingredients and CL-20 and nitroguanadine as separate additives for high- and low-energy propellants. The propellant’s burning rate was characterized and insufficient burning-rate ratio between the fast and slow baseline propellants was found. Impetus obtained from the combustion of the combined baseline propellants was also found to be far from the demanded value of 1300 J/g. Several modifications were made by introducing nano-sized aluminum particles and ultra-fine boron particles as well as high-energy oxidizer HNF into the propellant formulation. It was found that the addition of nano-sized aluminum particles can enhance the propellant burning rate only when the propellant contains oxidizers with a positive oxygen balance. Without the presence of positive oxygen balance oxidizer, the exothermic reaction of aluminum and boron particles occurs at a large distance from the burning surface introducing an energy-sink effect. The results obtained from the combustion of the advanced propellants show that an average impetus of 1299 J/g, a flame temperature of 3380 K with a burn rate ratio around 3 between the fast- and the slow-burning layers can be achieved. These conditions are desired for fast-core layered propellant applications. The impact sensitivities of the baseline, intermediate and advanced propellants were measured. The results show that addition of HNF and nano-sized aluminum exhibited improved impact sensitivity at levels that can be considered acceptable for deployment.

Solid Rocket Motor Internal Ballistics Simulation Using Three-Dimensional Grain Burnback

Abstract: Internal ballistics simulations of solid rocket motors have been conducted with the propellant grain's 3-D burning surface geometry described by a new minimum distance function approach and the internal flowfield represented by 1-D, time-dependent, single-phase compressible flow equations. The combustion model includes erosive burning and unsteady, dynamic burning corresponding to transient energy storage in the heated surface layer of the propellant. The integrated internal ballistics code (Rocballist) is used to investigate the role of these two burning rate augmenting mechanisms in solid rocket motor performance. Two tactical motors are used as test cases. Results indicate that dynamic burning can be the dominant factor in producing a short-duration ignition pressure spike in low-L* motors, particularly if the L/D ratio is not too large and the port cross section is nonrestrictive (e.g., center perforated grain). However, when L/D is large and the port cross section is noncircular in the aft section (aft fins/slots), erosive burning can take over in dominating the burning rate to the extent that an otherwise progressive pressure-time trace becomes regressive/neutral. That is, erosive burning can effectively prolong the initial pressure spike in some star-aft motors. The results also show that with sufficiently accurate models of dynamic burning and erosive burning, it is reasonable to expect reliable internal ballistics predictions with suitable simplified flowfield models, thereby realizing significant reductions in computation time compared with 3-D, multiphase reacting flow simulations.

An Advanced GAP/AN/TAGN Propellant. Part I: Ballistic Properties

Abstract: A solid rocket propellant based on glycidyl azide polymer (GAP) binder plasticized with nitrate esters and oxidized with a mixture of ammonium nitrate (AN) and triaminoguanidine nitrate (TAGN) was formulated and characterized. Non-lead ballistic modifiers were also included in order to obtain a propellant with non-acidic and non-toxic exhaust. This propellant was found to exhibit a burning rate approximately twice that of standard GAP/AN propellants. The exponent of the propellant is high compared to commonly used composite propellants but is still in the useable range at pressures below 13.8 MPa. This propellant may present a good compromise for applications requiring intermediate burn rate and impulse combined with low-smoke and non-toxic exhaust.

Sb6O13 and Bi2O3 as Oxidants for Si in Pyrotechnic Time Delay Compositions

Abstract: Lead and its compounds in detonator time delays are being phased out owing to environmental and health concerns. It was found that changing from the conventional rolled lead elements to rigid aluminium tubes caused a significant decrease in the burn rate. It also impaired ignitability, especially of slow burning compositions such as SiBaSO4. Consequently, potential alternatives for the latter and also the fast burning SiPb3O4 system were sought. Bi2O3, prepared by thermal decomposition of bismuth subcarbonate, gave fast burning compositions with silicon as fuel (155 mm s−1 with 20% Si). This system was ignitable by the spit of a shock tube. The SiSb6O13 system required an initiating composition and yielded slow burning compositions. The lowest sustainable and reproducible burn rate in lead tubes, in the absence of additives, was 4.8 mm s−1. In lead tubes, it was possible to reduce the burn rate further by adding fumed silica: A composition obtained by adding 10% fumed silica (add-on basis) to a 10% Si–90% Sb6O13 composition still burned reliably at a burn rate of 2.3 mm s−1.

Solid propellant burn rate, propellant gas flow rate, and propellant gas pressure pulse generation control system and method

Abstract: Systems and methods of controlling solid propellant burn rate, propellant gas pressure, propellant gas pressure pulse shape, and propellant gas flow rate, rely on pulse width modulation of a control valve duty cycle. A control valve that is movable between a closed position and a full-open position is disposed downstream of, and in fluid communication with, a solid propellant gas generator. The solid propellant in the solid propellant gas generator is ignited, to thereby generate propellant gas. The control valve is moved between the closed position and the full-open position at an operating frequency and with a valve duty cycle. The valve duty cycle is the ratio of a time the control valve is in the full-open position to a time it takes the valve to complete one movement cycle at the operating frequency. The valve duty cycle is controlled to attain a desired solid propellant burn rate, propellant gas pressure, propellant gas pressure pulse shape, and/or propellant gas flow rate.

Bi-propellant rocket motor having controlled thermal management

Abstract: A bi-propellant rocket motor having controlled thermal management is disclosed. The rocket motor produces thrust using a solid or gel-phase primary propellant that can be either fuel- or oxidizer-rich, with a complementary self-pressurizing secondary propellant selected to balance the primary propellant in terms of the equivalence ratio. The motor houses multiple propellant grains arranged in such a configuration that each chamber containing a primary grain serves as both propellant storage and the main combustion chamber for that propellant grain as it burns with the secondary propellant. The secondary propellant is stored separately, and the flow routed past the primary propellant chamber to provide cooling for adjacent primary propellant chambers limiting the temperature rise in the motor structure.

Non-Steady Burning Effect on Solid Rocket Motor Performance

Abstract: Solid rocket performance during rapid pressure excursions differs greatly from predictions based on steady state burning rate data. Rapid pressurization following a chamber filling interval produces indicated burning rate overshoots. Transient internal ballistics of a solid propellant rocket motor during rapid pressurization part of chamber filling phase has been considered in this work. Quasi one-dimensional unsteady Euler equations with a transient propellant burning model that accounts for the effects of time rate of change of the chamber pressure on the burning rate have been used to simulate the internal ballistics of rocket motors. The compressible convective flow solver used in this study is based on Roe’s scheme. The effects of rapid chamber pressure change on the propellant and motor transient behavior have been studied. Furthermore, prediction of the level of pressurization rate has been achieved.

Mechanism and laws of combustion of composite solid propellants with a coolant

Abstract: The model of combustion of a composition consisting of a quasi-homogeneous composite propellant (matrix) and coolant particles is considered. The model is based on the leading role of exothermal decomposition of the matrix and on the cooling effect of the second component by virtue of transverse heat transfer between the components in the condensed and gas phases. Formulas for combustion characteristics (temperature, burning rate, its sensitivity to pressure, and initial temperature) are derived and analyzed. The calculated dependences of these characteristics on pressure, particle size, concentration, and thermal effects of decomposition of the components show that transitional regimes with a stronger dependence of the burning rate on pressure than that of the initial propellant are reached in a certain range of parameters. An algorithm is proposed, and a parametric identification of the model on the basis of experimental data is performed.

Experimental Analysis on the Measurement of Ballistic Properties of Solid Propellants

Abstract: Ballistic properties of solid propellants play an important role in the performance of the solid propellant rocket motors. Therefore, ballistic properties of a likely propellant should be known and provided to the design engineers. In this study, a specific AP/HTPB composite solid propellant (SCP) was examined to obtain steady-state linear burning rates as a function of pressure and propellant initial temperature, temperature sensitivity, and pressure deflagration limit (PDL). In some tests micro-thermocouples were embedded into the propellant samples to measure the temperature profiles in the solid, condensed, and gas-phase flame regions. From the measured profiles, burning surface temperatures were obtained and the activation energies were extracted from these data. A high-pressure strand burner was developed to perform ballistic property measurements of the SCP. Propellant samples were ignited by Ignition Wires. Experiments were performed at different pressures and temperatures to determine the ballistic properties of the propellant at different environmental conditions. Auxiliary systems such as Cryogenic Cooling System and Electrical Heating System were built to provide different environmental conditions in the strand burner. An inert gas (N2) was selected for both conditioning and pressurizing the test chamber. Experiments revealed that the PDL limit for the SCP is 1 atm, which means it can burn even at atmospheric pressure. Results showed that temperature sensitivity of the SCP decreased with the increase of chamber pressure. Experiments aimed to measure burning surface temperature of the SCP did not represent reliable results due to the fact that the exact location of the burning surface could not be obtained even with fine-wire (25 mum) thermocouples.

Theoretical Analysis of Plasma Enhancement of Propellant Burning Rate

Abstract: A theoretical model of the combustion of a nitramine solid propellant in the presence of a plasma jet is proposed. Unlike standard double-base compositions, nitramine propellants exhibit experimental evidence that plasma induces a burning rate enhancement. The model is based on heat transfer considerations and proposes a closed-form solution of the enhancement of the propellant burning rate as a function of the thermophysical parameters of the system. The model provides a good qualitative agreement with experimental results.

Production of synthetic gas from organic waste

Abstract: A pyrolysis device has a pyrolyzer (retort) and a thermal oxidizer, each with its own burner, but coupled so that the burner of the pyrolyzer can be down-regulated as heat from the thermal oxidizer is used to hear the pyrolyzer. One or more automatic controllers preferably adjust a burn rate of the extrinsic gas by the first burner as a function of a rate at which heat is transferred from the thermal oxidizer to the pyrolyzer. In preferred embodiments, the burn rate of the extrinsic gas can be adjusted to ≤ 50%, ≤ 30% or even lower relative to a base rate that would be required to pyrolyze the waste without the heat thermal oxidizer.

Laser-assisted combustion of solid propellants

Abstract: This paper describes experimental work on laser-assisted combustion and its control of solid propellants. Burning rates of HTPB/AP composite solid propellants were measured at pressures up to 90 kPa under laser irradiation coupled with the propellant preheat temperature distribution measured by thermography. With custom made (fuel-rich) non-self combustible solid propellants, the combustion control, including ignition and interrupting by the laser irradiation, was feasible having the burning rate coincident with Vieille’s law and a pressure exponent of approximately 0.5, and therefore, this combustion control concept seems to be applicable to a thruster such as one for microsatellites. The heat balance on the burning surface evaluated with a laser power density of 1.4 W/mm 2 at 50 kPa was 1)the chemical reaction was endothermic at a produced heat flux of -0.7 W/mm 2 , 2)the heat was supplied to the solid phase of the propellant at a heat flux of 3.7 W/mm 2 , and 3)the heat feedback rate to the burning surface was 3.0 W/mm 2 , which suggest that approximately 70% of the energy to the burning surface was provided by the combustion itself, and the laser supplied the supplementary heat flux to compensate for the heat insufficiency for combustion sustenance.

Combustion Characteristics of GIBR Layered Square Flake Gun Propellant

Abstract: In order to study the progressivity of GIBR(gradual increasing burning rate) layered propellant with square flake shape,the physical model about the combustion process was put forward,and the shape function was calculated according to the physical model and the parallel layer burning law.The change of Ψ,σ with Z was calculated and analyzed at different parameters such as X1,K and β.The result by theoretic study was validated by interrupted burning method both in gun chamber and closed bomb.The results show that the gas generating rate of the layered propellant can be controlled by adjusting the burning rate,dimension and thickness ratio of outer-layer to innerlayer.The results also show that the factual combustion law and theoretic model are consistent and the combustion process approximately reflects the parallel layer burning law.

Numerical Analysis of the Effects of Solid Metal Fuel Particle Characteristics on the Combustion Process

Abstract: Combustion using aluminum particles as fuel is an attractive energy source where high energy densities are desired. Very little experimental literature or computational results are available for metal combustion in high-pressure chambers, as most experimental and computational work has been done on chamber operating at near atmospheric pressures. This paper attempts to improve our understanding of metal-fueled combustion chambers at pressures above atmospheric. A numerical model of solid Aluminum fuel particle combustion is developed to investigate the effects of radiation and fuel particle size on the combustion process. Of specific interest are particle specific burn rate, residence time, combustion efficiency, coupled radiation effects, and flame characteristics. This computational model is applied to a linear-type dump combustor. The effects of a range of particle sizes are investigated using mono-dispersed and poly-dispersed particle distributions. Combustion efficiency and characterization of the combustion process are addressed by studying particle ignition delay, surface combustion time, and particle flame radiation intensity as a function of particle diameter and mass fraction. The computational results of this detailed theoretical combustor reveal fundamental physics relating particle sizes and distributions to the variables commonly used to define the effectiveness and performance of the combustion process. The computational models include nonisotropic turbulence models, empirically derived ignition criteria and reaction rates, as well as convective and radiant heat transfer. The numerical results were compared with test data with reasonable agreement.

Influence of Burning Rate Pressure Exponent of Propellant on Gun Interior Ballistics

Abstract: In order to study the interior ballistic behavior of new type high-energy propellant,influence of burning rate pressure exponent of propellant on gun interior ballistic behavior was theoretically simulated by classical interior ballistic mathematical model.The results show that the value of burning rate pressure exponent of propellant and its change law can bring about a great change of interior ballistic behavior.The new type high-energy propellant whose burning rate pressure exponent is decrescent as pressure increasing can obtain higher interior ballistic piezometric efficiency,and reduce the sensitivity of bore pressure and muzzle velocity on loaded conditions of propellant.It will appear an evident error when interior ballistic behaviors of this kind of high-energy propellant are evaluated by traditional average burning rate pressure exponent.

Influence of Environment Temperature on Thermal Self-Ignition of Propellant

Abstract: Taking the single base 9/7 propellant as an example,the mathematical model of the self-ignition of propellant is established through testing the self-ignition time under different conditions of the environment temperature.In this experiment,the heater heats the hot piece,and the temperature control equipment controls its temperature.The reactor which is filled with propellant was placed at the middle of hot piece cavity,using vapor to heat.Then,after putting the temperature detector into the propellant,the temperature is tested and time is recorded.According to the experiment,the numerical simulation result is in accordance with that of the experiment.

Effect of Grain Size on Progressive Combustion of the Variable Burning Rate Propellant

Abstract: The relation of grain size with progressive combustion of the propellant was studied by closed bomb test. The characteristics of p-t and L-B curves for different prescription and grain size of the variable burning rate propellant were analyzed,and the influence law of grain size on burning properties were obtained.Results show that,for certain prescription and thickness for the inside and outside layers of the variable burning rate propellant,the ratio of length to diameter(1.5/1-2.0/1) affects the progressive combustion more obviously.The ratio of length is higher,the progressive combustion is better.