Showing papers on "Burn rate (chemistry) published in 2009"

Size and Shape of Ammonium Perchlorate and their Influence on Properties of Composite Propellant

Abstract: Most of the composite propellant compositions contain solid loading up to 86 per cent. The main solid ingredients of composite propellant are ammonium perchlorate (AP) and aluminium powder. Therefore, it is a must to characterise these to improve processibility and quality of composite propellant. Effect of particle size on propellants slurry viscosity and ballistic parameters are well documented, however, the effect of oxidizer particle shape is not reported. In the present study, different methods for size and shape characterisation are discussed and effect of size and shape of AP on composite propellant properties are studied. The data indicate that as size of AP decreases, propellant slurry viscosity increases and burn rate increases. The particles having higher shape factor provides less endof mix (EOM) viscosity of propellant slurry and burn rate. Further, effect of size of ground AP on shape is also investigated. From the data thus obtained, it is inferred that as size of ground AP decreases, shape factor decreases, and particles become more irregular in shape.Defence Science Journal, 2009, 59(3), pp.294-299, DOI:http://dx.doi.org/10.14429/dsj.59.1523

Breeding a better stove : the use of computational fluid dynamics and genetic algorithms to optimise a wood burning stove for Eritrea

Abstract: Improved cooking stoves can bring significant benefits to women and children in rural African situations, due to reduced fuel consumption and improved indoor air quality. This investigation focuses on the use of Computational Fluid Dynamics (CFD) and Genetic Algorithms (GAs) to optimise a stove for Eritrea. Initial work focussed on developing a model of wood combustion in a fixed bed. An experimental investigation was carried out on regular wood cribs to determine the burn rate and temperature field above a wood fire. The experimental data was used to develop a numerical model using CFD software Fluent 6.2 and user-defined functions for the fixed bed of fuel. The model assumed that pyrolysis was limited by heat transfer through the fuel, and that char combustion was limited by oxygen diffusion to the fuel surface. Simulation results yielded a mean and maximum error of 16% and 42% respectively in fuel burn rate. In the second phase of the investigation, the numerical model of wood combustion was used as part of a larger CFD model to capture the behaviour of a complete stove. The model was compared with experimental data for rocket type stoves with different geometries. The model correctly identified the trends of fuel burn rate and heat transfer in the experimental data, though agreement with experimental values was poor and the model exhibited significant errors when altering stove height and diameter. In the final phase of the investigation, the stove model was used in conjunction with a genetic algorithm to optimise the stove shape. Two methods of genetic coding were investigated. The resulting stove is expected to half fuel consumption compared to the classic mogogo stove, though this remains to be experimentally verified.

Methods Of Controlling An Internal Combustion Engine Including Multiple Fuels And Multiple Injectors

Abstract: A method for controlling the start of combustion and peak burn rate of an internal combustion engine system on a cycle-to-cycle basis including multiple direct fuel injectors, multiple indirect fuel injectors, and multiple liquid fuels. The actual start of combustion and actual peak burn rate are calculated using in-cylinder pressure or ionization. A predetermined start of combustion and a predetermined peak burn rate are selected. The start of combustion and peak burn rate control algorithm includes controlling a fuel ratio, an injection ratio, and a residual gas recirculation ratio. A predetermined fuel ratio, a predetermined injection ratio, and a predetermined residual gas recirculation are calculated by combining feed forward ratios and closed loop control ratios corresponding to the three ratios. The predetermined ratios are employed so that the system achieves the predetermined start of combustion and predetermined peak burn rate.

Performance and Analysis of Vortex Oxidizer Injection in a Hybrid Rocket

Abstract: : A hybrid rocket motor is a type of rocket motor where fuel is placed in a combustion chamber as a solid, and then gaseous or liquid oxidizer is injected. When the two mix and are ignited, the surface of the fuel burns and the gases produced in the combustion develop thrust. Hybrid rocket motor performance is dictated by the rate at which the fuel burns. Fuel burn rate (or regression rate) can be increased by increasing oxidizer flow speed over the burning fuel surface. This is because flow over the burning surface creates shear stress which facilitates fuel and oxidizer mixing. One method for improving shear stress and thus regression rate is to induce an oxidizer vortex in the combustion chamber. The subject of this research is a method for inducing vortical flow that combines vortex and axial oxidizer injection within a cylindrical, interior burning fuel grain. A hybrid motor test stand has been developed to test both axial and vortex oxidizer flow configurations as well as any combination of the two. The apparatus is capable of measuring thrust, oxidizer flow rate, and chamber pressure. This, along with physical measurements of fuel grains, allows the determination of fuel regression rate, combustion efficiency, and specific impulse, all key rocket performance parameters. The apparatus is also equipped with millisecond scale combustion analyzers to measure the gases in the combustion products, to include CO, CO2, NOx, and unburned hydrocarbons. The high sample rate of these analyzers sheds light on vortex hybrid combustion processes as well as the phenomenon which could lead to combustion instability. Overall, this research is focused on identifying a possible way to increase hybrid rocket performance in order to bring this very safe and efficient type of propulsion to maturity.

Characterisation of emissions and combustion stability of a port fuelled spark ignition engine

Abstract: The chemical and physical limits of cycle-to-cycle combustion variability and engine out emissions of a gasoline port fuelled spark ignition engine have been investigated. The experimental investigations were carried out on a V8 engine with port fuel injection and variable intake valve timing. The chemical limits of stable combustion have been shown to be a function of burned gas, fuel and air mixture. The widest limit, gas fuel ratio of 9 was found at maximum brake torque spark timing. Retarding the spark timing by 10oCA caused a small reduction in the stable area, 20oCA retard reduced the stable combustion area significantly, whereby stable combustion occurred within an area of gas fuel ratio of 10. Burn rate analysis indicated increased variability in both the flame development and rapid burn period. The increase in variability in the rapid burn period is greater than that associated with the flame development. The variability is magnified from flame development through the rapid burn phase. This finding was consistent for unstable combustion caused by exceeding chemical and physical limits. Engine out emissions were investigated and characterised using engine global state parameters, for example AFR, burned gas fraction, for both stable and unstable combustion conditions. Carbon monoxide and oxides of nitrogen emissions correlations were unaffected by the presence of unstable combustion events whereas hydrocarbon emissions showed a significant increase. The incorporation of these findings were implemented into an engine simulation (Nu-SIM V8) investigating the impact for the New European Drive Cycle condition.

Regression fronts in random sphere packs: Application to composite solid propellant burning rate

Abstract: The burning rate of a composite solid propellant may be estimated by global modeling, such as the widely used BDP model. The backbone of such models is the “mixture law” that links the propellant burning rate rp with the burning rate of its own components, i.e., oxidizer rox and binder rb. However, different laws are available in literature which all read: 1/rp = q(ξ)/rox + (1 − q(ξ))/rb, with q(ξ) a function of oxidizer volume fraction ξ. This work attempts in analyzing numerically the validity of those empirical formulations by surface regression computation. Composite propellants are modeled by a random packing of monomodal spheres and the evolution of the regression front is computed via the resolution of Hamilton–Jacobi equations. It is shown that the popular choice q(ξ) = ξ is fairly valid but only provided that burn rate ratio Z = rox/rb is about 1. When Z > 1, combustion surface is no longer plane and global burning rate deviates from postulated laws. A special regime is also noticed for high rate ratio Z (typically Z > 5) because combustion then preferentially takes place through adjacent oxidizer particles. Computed results occur to be correctly modeled by percolation theory. This hints that percolation is a common feature of propellant combustion and a critical percolation threshold on volume fraction is numerically found to be about ξc ∼ 0.2. First validations show encouraging correlations with experimental data.

Predicting the Distribution of Combustion Timing Ensemble in an HCCI Engine

Abstract: Control of Homogeneous Charge Compression Ignition (HCCI) engines to obtain the desirable operation requires understanding of how different charge variables influence the cyclic variations in HCCI combustion. Combustion timing for consecutive cycles at each operating point makes an ensemble of combustion timing which can exhibit different shapes of probability distributions depending on the random and physical patterns existing in the data. In this paper, a combined physical-statistical control-oriented model is developed to predict the distribution of HCCI combustion timing (CA50, crank angle of 50% fuel mass fraction burnt) for a range of operating conditions. The statistical model is based on the Generalized Extreme Value (GEV) distribution and the physical model embodies a modified knock integral model, a fuel burn rate model, a semi-empirical model for the gas exchange process and an empirical model to estimate the combustion timing dispersion. The resulting model is parameterized for the combustion of Primary Reference Fuel (PRF) blends using over 5000 simulations from a detailed thermo-kinetic model. Empirical correlations in the model are parameterized using the experimental data obtained from a single-cylinder engine. Once the model is parameterized it only needs five inputs: intake pressure, intake temperature, Exhaust Gas Recirculation (EGR) rate, equivalence ratio and engine speed. The main outputs of the model are CA50 and the Probability Density Function (PDF) metrics of CA50 distribution. Experimental CA50 is compared with predicted CA50 from the model and the results show a total average error of less than 1.5 degrees of crank angle for 213 steady-state operating points with four different PRF fuels at diversified operating conditions. Predicted PDF of the CA50 ensemble is compared with that of the experiments for PRF fuels at different running conditions. The results indicate a good agreement between simulation and the experiment.Copyright © 2009 by ASME

Estimation of Pressure Index and Temperature Sensitivity Coefficient of Solid Rocket Propellants by Static Evaluation

Abstract: Burning rate of a solid rocket propellant depends on pressure and temperature. Conventional strand burner and Crawford bomb test on propellant strands was conducted to assess these dependent parameters. However, behaviour of propellant in rocket motor is different from its behaviour in strand form. To overcome this anomaly, data from static evaluation of rocket motor was directly used for assessment of these burningrate controlling parameters. The conventional empirical power law (r=aoexp[p{T-To}]Pn) was considered and a method was evolved for determination of pressure index (n) and temperature sensitivity coefficient (p) of burning rate for solid rocket propellants from static evaluation data. Effect of pressure index and temperature sensitivity coefficient on firing curve is also depicted. Propellant grain was fired in progressive mode to cover a very wide pressure range of 50 kg/cm2 to 250 kg/cm2 and propellant burning rate index was calculated to be 0.32 in the given pressure range. Propellant grain was fired at +35 °C and –20 °C temperatures and temperature sensitivity coefficient of burning rate was calculated to be 0.27 % per °C. Since both the values were evaluated from realised static evaluation curves, these are more realistic and accurate compared to data generated by conventional methods. Defence Science Journal, 2009, 59(6), pp.666-669 , DOI:http://dx.doi.org/10.14429/dsj.59.1573

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.

Solid propellant rocket motor having self-extinguishing propellant grain and systems therefrom

Abstract: A solid rocket motor includes a combustion chamber bounded by an outer casing, a propellant grain within the combustion chamber, and an igniter within the outer casing for igniting the propellant grain. A nozzle is coupled to the combustion chamber for releasing hot gasses evolved from burning the propellant grain to provide thrust for propelling the solid rocket motor. The propellant grain is a self-extinguishing propellant grain that includes at least one fuel, at least one oxidizing agent, at least one binder, and at least one surfactant that imparts the self-extinguishing property. The propellant grain provides a burning rate as a function of pressure that 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 propellant grain.

Uncertainty of Solid Propellant Burning Rate Measurements

Abstract: Laboratory measurements of a new propellant formulation provide critical screening of ballistic properties. The objective of this work is to quantify the uncertainty of laboratory measurements of propellant burning rates. The scope includes uncertainty analyses of the measured propellant burning rate from closed combustion bomb testing. The scope includes a Monte Carlo assessment of digital algorithms that propagate the uncertainties in the primary measurements throughout all the data analysis algorithms. The results are presented parametrically in terms of these variables to show the sensitivity of their sensitivity on the measurements. The results reveal that the burning rate can be determined to within 4.28% to 5.73% depending on the burn rate determination method used.

Flame Propagation of Bio-Ethanol in a Constant Volume Combustion Chamber

Abstract: This research presents an experimental study of the combustion of ethanol blends with gasoline in a constant volume combustion chamber. The test fuels are gasoline, E20 (ethanol20%), E85 (ethanol85%) and pure ethanol. The pressure in combustion chamber was measured. The images of flame propagation were recorded by schlieren technique with high speed video camera. The effects of equivalent ratio, temperature, pressure on the combustion characteristics were studied. The results show the greater percent of ethanol provides faster burn rate of combustion and higher peak of cylinder pressure. The peak combustion pressure E100 is 0.873 MPa which is higher than E85, E20 and E0. The ignition delay time decreases with higher percent of ethanol. As the results of flame propagation images, at stoichometric burn duration of pure ethanol is 13 ms which is shorter than E85, E20 and E0. The combustion duration of mixture at equivalence ratio of 1.0 is less than that of equivalent ratio 0.8, 1.2 and 1.4. The flame speed and peak pressure increase with percentage of ethanol. The faster flame speed of ethanol flame effects to shorter combustion duration and greater performance and efficiency.

Solid Rocket Motor Design for a High Altitude Composite Rocket

Abstract: The Rocket Team at The Ohio State University has initiated project Prometheus in an attempt to break an al titude of 38,000 meters. As part of the project, the team intends to design, construct, and test a large solid propellant rocket motor. Preliminary design efforts produced the size and performance constrains for the motor. Steel, titanium, and aluminum were examined for use as the motor casing. Aluminum 6061 alloy was chosen as the casing material. A finite element analysis was performed to ensure the structural integrity of the casing. A combination of hydroxyl -terminated polybutadiene, ammonium perc hlorate, and aluminum in an internal burning tube configuration will comprise the propellant grain. An analysis of the performance characteristics of the propellant was performed to optimize the grain design. The solid rocket motor will be tested to veri fy the theoretical performance. Nomenclature Case OD = outer diameter of the motor casing Case TH = thickness of the motor casing g = acceleration due to gravity Grain ID = inner diameter of propellant grain Grain L = length of propellant grain Grain OD = out er diameter of propellant grain Grain VOL = volume of propellant grain ISP = specific impulse p m = propellant mass flow rate B r = propellant burn rate B t = burn time � = th rust eff u = effective velocity

Heat release in gas phase near burning surface of Composite propellants -Ammonium perchlorate (AP) and Black powder (BP) propellant-

Abstract: To design the rocket motor, thrust and burning rate of propellant are important factors. The large range of burning rate makes the design criteria increase. The range of burning rate is expanded with catalyst or changing the size of oxidizers, however the range of burning rate of ammonium perchlorate composite propellants is small. Black powder has large range of burning rate. The burning rates of propellants are accelerated by conductive heat of reaction zone, therefore it is possible to increase burning rate by evaluating reaction zone. Reaction zone at burning surface of propellant is less than 1mm at 0.1MPa. It is difficult to observe the reaction zone with video camera and to evaluate gas reaction. We measured temperature histories in this reaction zone with small thermo-couple. The temperature fluctuation occurred near burning surface. We also analyzed the temperature vibration. The temperature fluctuation appeared by combustion of solid particles like aluminum or charcoal. Heat of reaction in gas phase is proportional to the partial derivative of the second order with respect to the distance (heat flow) at steady state. The heat flow fluctuated near the burning surface and heat of reaction generated, and the length, l l l l*, from burning surface to fluctuated heat flow, decreased with increasing the burning rate of black powder. The frequency of heat release vibration is 20Hz in the gas phase. The burning rate increases with decreasing the starting time of heat release and period of combustion in gas phase.

Performance of Nitramine Propellants with Different Phases of HMX

Abstract: In order to study the effects of phases of HMX on the nitramine propellants,the combustion performance and mechanical properties of nitramine propellants containing α-HMX and β-HMX were tested by closed combustion bomb and mechanical properties tester.Results show that the propellant with α-HMX is easy to be ignited and its gas generation brisance and combustion rate is lower than that with β-HMX,and the burning rate pressure exponent of the propellant with α-HMX is 0.1 higher than that with β-HMX and mechanical properties of the propellant with α-HMX is worse.In addition,when loading density is 0.12 g·cm-3,the two burning rate pressure exponents are larger than 1,and when loading density is 0.20 g·cm-3,the two burning rate pressure exponents reach the level of less than 1.

Experimental Study on Ignition Property of Composite Propellants on Different Condition

Abstract: The ignition characteristics of three typical composite propellants were studied,in order to promote the new high energy composite propellant's application.On different density of heat flow rate, environmental pressure and original temperature conditions, the ignition law of the HTPB propellant, two types of high energy propellant were obtained.The result shows that the new generation high energy propellant can be ignited easily, the ignition energy of it is less than the other two propellants ,its ignition security should be paid more attention.

Investigation on measuring gas generation rate for fuel-rich propellant during combustion process

Abstract: The equipment for measuring gas generation rate of fuel-rich propellant was established,and a experimental method was put forwardPickling asbestos and silica glass fiber cloth were used as filter in the testing equipment to realize reliable separation of gas-phase and condense phase combustion productsAnd,a reliable method was put forward by verifying the credible separation between gas and solid/liquidThe pressure change in the course of experiment was testedThe results show that the pressure in combustion chamber can be controlled by adjusting sample shape and consumption as well as the back end lid number,and the measurement of gas generation rate of sample under certain pressure can be realized

Numerical simulation for burning rate of GAP high-energy propellant

Abstract: The thermal and combustion characteristics of GAP were analyzed.The hypothesis about GAP decomposition in the initial stage of combustion was proposed.The burning rate equation of GAP and GAP high-energy propellant was derived based on the chemical structure.The calculated burning rates of GAP high-energy propellant agree well with most of the experimental data with the deviation of all data within±15% and that of a majority of data wihtin±10%.It proves that the model and program are reasonable.

DNS Simulation of Erosive Burning in Planar Periodic Rockets

Abstract: We examine the combustion response of a flame anchored to two 1/4-spaces of solid fuel and oxidizer, a configuration relevant to the combustion of heterogeneous solid propellants. A time-periodic shear flow is applied to model the shear that can be generated by the presence of acoustics or turbulence in a rocket chamber. To estimate the magnitude and frequency of the shear for the case of a turbulent flow, we present DNS results of a planar periodic rocket, a configuration that has its roots in a multiscale analysis. Such a configuration allows for the determination of the shear parameters as functions of motor geometry and downstream location. The response of the flame to this shear, the heat flux to the surface, and the burning rate are calculated numerically. Significant enhancement to the burn rate, commonly known as erosive burning, is found.

Development of Novel Polynitrogen-Based High-Performance Solid Propellants

Abstract: : In this STTR program, CFDRC, in collaboration with the University of Idaho, proposed to develop some novel high-performance polynitrogen energetic materials for propellant applications As a proof-of-concept, we synthesized a number of promising high energy amine substituted tetrazole based compounds These compounds were first theoretically investigated to obtain their heats of formation, densities, specific impulses, detonation pressures and velocities The synthesized compounds were characterized, and densities, melting points and impact sensitivities were measured in the laboratory Some of these compounds have remarkably low impact sensitivities We conclude that at least five compounds have excellent potential for use a propellant ingredient Impact sensitivities were further measured in a mixture of ammonium nitrate We observed that a small percentage of ammonium nitrate improves the impact sensitivities further while improving the oxygen balance In Phase II, these compounds will be formulated, and undergo a rigorous set of tests, such as burn rate, compatibility, mechanical strength, to evaluate their performance in practical situations

Linear Analysis of Erosive Burning in Solid Propellant Motors

Abstract: A theoretical investigation of the coupling between the core flow process and combustion sublayer process in a rocket chamber flow is presented. The focus of the investigation is on penetration of chamber disturbances into the sublayer and on the burn rate characteristic responses. Characteristic solutions to the local non parallel instability problem supported by gasifying propellant in a double slab geometry are sought. In the fully decoupled limit the only characteristic solution response is a pressure response, thus comparisons between coupled and decoupled pressure responses are presented. The propellant is typified by a set of burning characteristics, and their effect on the coupling is explored. The goal of the analysis is to clarify the extent strand burning characteristics (i.e., in zero cross flow conditions) can affect burn rate perturbations leading to the phenomenon of erosive burning. The findings of this research agree with experimental studies as it regards the effect of burning parameters on erosive behavior. Change in erosive sensitivity with propellant characteristics is thus correlated with a change in the core-sublayer interaction. n this paper we present an investigation of erosive burning in rocket motors by performing a linear analysis of the coupling between combustion sublayer and turbulent chamber flow. Erosive burning describes burn rate augmentation in solid rocket engines as a result of cross-flow conditions. Erosive burning is responsible for strong overpressure at the rocket firing stage. A better understanding of the influence of burning characteristics on erosive sensitivity can improve rocket design techniques and propellant choices. There have been many attempts to identify the mechanism responsible for the measured increase in burn rate over the zero-cross flow burn rate. “Flame bending”, “heat transfer from core to propellant surface”, “increase in heat flux due to turbulence” are classical examples. Most analyses have relied on empirical fits often based on turbulent profiles over a flat plate; none have been based on a rigorous analysis of the equations. Due to the large difference in length scales between combustion sublayer and chamber core flow, as discussed in more detail in §II of this paper, the development of the instability and the ensuing turbulent field are fairly independent of the burning characteristics. Therefore, the experimentally established link between erosive burning augmentation and propellant characteristics (e.g., King 3 ) cannot be resolved without hypothesizing an effect of the instability on the micron-size combustion profile; that is, hydro-dynamic considerations alone are not sufficient. Mukunda and Paul 2 argue that the effect of the burning characteristics can be taken into account by simply scaling the port mass flow rate by the strand value. Still, the proposed effect of the Reynolds number on erosive burning cannot be explained based on the relationship between the same parameter and the instability growth rate. The present paper attempts to provide a more rigorous analysis of the effect of burning parameters on the core flow-sublayer coupling by seeking characteristic solutions to the linearized and localized solid-gas fully-coupled problem. Although the linearized analysis does not provide a clear quantitative answer on the burning rate augmentation, burn rate eigenfunctions (fluctuation intensity) quantify the importance of

Studies on Starting Transient of Dual-thrust Motors

Abstract: In this paper parametric analytical studies have been carried out to examine the starting transient flow features in dual-thrust motors with horizontal and flip horizontal positions. Numerical computations have been carried out with the aid of a standard k-ω ω ω ω turbulence model. This model uses a control-volume based technique to convert the governing equations to algebraic equations. We conjectured from the numerical results that in dual-thrust motors with highly loaded propellants the mass flux of the hot gases moving past the burning surface is large. Under these conditions, the convective flux to the surface of the propellant will be enhanced, which in turn enhance the local Reynolds number. This amounts a reduction in heat transfer film thickness and enhanced heat transfer to the propellant with consequent enhancement in the dynamic burn rate and the starting pressure transient.