Showing papers on "Burn rate (chemistry) published in 2011"

Analysis of RDX-TAGzT pseudo-propellant combustion with detailed chemical kinetics

Abstract: A detailed model of steady-state combustion of a pseudo-propellant containing cyclotrimethylene trinitramine (RDX) and triaminoguanidinium azotetrazolate (TAGzT) is presented. The physicochemical processes occurring within the foam layer, comprised of a liquid and gas bubbles, and a gas-phase region above the burning surface are considered. The chemical kinetics is represented by a global thermal decomposition mechanism within the liquid by considering 18 species and eight chemical reactions. The reactions governing decomposition of TAGzT were deduced from separate confined rapid thermolysis experiments using Fourier transform infrared spectroscopy and time-of-flight mass spectrometry. Within the gas bubbles and gas-phase region, a detailed chemical kinetics mechanism was used by considering up to 93 species and 504 reactions. The pseudo-propellant burn rate was found to be highly sensitive to the global decomposition reactions of TAGzT. The predicted results of burn rate agree well with experimental burn...

Effect of application of various ignition conditions in closed-vessel tests on burning rate calculation of a fine-grained propellant

Abstract: The most widely used experimental method for determination of dynamic characteristics of gun propellant combustion is burning a specific amount of a propellant in a closed chamber (vessel) and measuring the resulting pressure versus time profile. The aim of this work is to conduct experimental investigations at different conditions of propellant ignition, which permit to verify the view on legitimacy of standard conditions of closed-vessel tests. This paper considers first of all the influence of conditions of propellant ignition during closed-vessel tests on possible deviations in determination of the burning rate. For this purpose, experimental tests of a fine-grained, single-base propellant are carried out in a conventional closed vessel and a micro-closed vessel, applying different ignition methods for the same loading density. The results of experimental tests and calculations show significant influence of the ignition system type used on determined values of the burning rate and other dynamic characteristics of propellant combustion, such as dynamic vivacity, absolute quickness, and relative quickness.

Numerical Study of Erosive Burning in Multidimensional Solid Propellant Modeling

Abstract: Using numerical simulations, the combustion response of heterogeneous solid propellants to an imposed crossflow velocity field are examined. It is shown that this model flow is useful to capture the influence of shear flow in the so-called erosive burning phenomenon observed in actual experiments. Previous numerical studies on a model quarter-plane problem and on homogeneous solid propellants have shown that the presence of shear in the crossflow plays a role in increasing the heat transfer to the propellant surface, thus enhancing the burn rate. In the current work, the response of twoand three-dimensional packs to an imposed velocity field are first examined for different propellant morphologies and at different pressures. The imposed velocity field has its root in a separate nonreactive multiscale analysis. It is shown that, with the model flow, the variations of the erosive burning rate with shear parameters are captured, the trends being in line with experimental results. Furthermore, a comparison with experimental results present in the literature is examined, where it is shown that the results compare qualitatively well with the experiments using the estimated shear rates from the experiment. These results suggest that the influence of the shear flow on the primary diffusion flame may be a leading factor in the erosive burning effect.

Effect of Protrusion in the Enhancement of Regression Rate in Hybrid Rocket

Abstract: It has been reported in literature that the use of protrusion in the combustion chamber of a hybrid rocket motor enhances the regression rate. This study reports careful experiments conducted to show that the improvement in burn rate with protrusions is only up to a certain fraction of the overall burn time. From the results obtained, it is seen that an X/L of 0.5 is the best location for a graphite protrusion. The protrusions are also shown to increase the combustion eciency by as much as 45 % when placed at an X/L of 0.5. This, more than the improvement in the regression rate is very useful, especially for small scale motors whose combustion eciencies are otherwise very low.

Enhancement of regression rate in hybrid rockets

Abstract: This study has identified a potential HTPB based propellant that gives a very good density specific impulse but yet does not burn like a solid propellant. Due to low solid loading in this propellant, pre-curing of the HTPB and IPDI mixture is a necessity to get uniform density across the length of the grain. Nea rly 7 hours of pre-curing gave the best results in terms of uniformity of the propellant co mposition. The paper also discusses the enhancement of regression rate of this propellant w ith known burn rate modifiers like iron oxide and copper chromite.

Modeling the Energy Release and Burn Rate Characteristics of ZPP Based Initiators

Abstract: Bridge-wire initiators using zirconium potassium perchlorate (ZPP) as pyrotechnic are commonly found in the aerospace, defense, and automotive industries. To measure pyrotechnic output, it is common practice to discharge initiators into small isochoric test vessels and monitor the transient pressure response. The maximum pressures obtained in these tests are usually compared to values determined theoretically based upon a thermochemical analysis. However, an analytical prediction for the rate of pressure increase is problematic because a complete kinetic description of ZPP combustion is impractical; furthermore, the burn rate parameters for ZPP are not reported in the literature. In this research, a previous analytical model describing the energy release characteristic for THPP based initiators is applied to ZPP initiators. ZPP based initiators are discharged into 100 cm 3 test vessels filled with argon at initial pressures ranging between 1.72 and 22.4 MPa. From these experiments the total energy content of the ZPP formulation used in this work is determined to be 5600 J/g. Estimates for the burn rate parameters for ZPP are based upon an earlier model, termed the TPPM, originally developed to describe the transient pressure response of high pressure argon gas environments heated by burning THPP particles. To estimate the burn rate parameters for ZPP a quasi-analytical approach using the TPPM is presented. This approach assumes the initial pressure rise determined from experimental data is proportional to the initial burn frequency which is defined based upon Vielle’s Law. Results from this analysis yield a burn rate exponent (n) and burn rate parameter (A) of approximately 0.47 and 1.94 x 10 -3 cm/ms/MPa n , respectively. These values for the burn rate parameters are used as input in a previously developed model, termed the MTPPM, describing the transient pressure response of argon gas heated by burning pyrotechnic over a wide range of initial pressures. Predictions from the MTPPM using the burn rate parameters of ZPP determined in this research are shown to accurately replicate the measured transient pressures over all initial densities tested. Therefore, the method for predicting burn rate parameters developed in this work can potentially be applied to other pyrotechnic formulations.

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.

Role of Moisture in Activated Charcoal in Composite Solid Propellant

Abstract: Experiments were conducted to examine the role of moisture in activated charcoal on the performance parameter like burn rate and burn rate pressure index of the aluminized composite solid propellant. Propellants with moistu re present in the activated charcoal were cured and shown to be free from blow holes and voids. They were also shown to have good mechanical properties. It was noticed that burn rat e and burn rate pressure index is higher for propellant composition containing very small pe rcentage (0.16 % and 0.6 % of the weight of the propellant) of moisture trapped insid e the pores of activated charcoal as compared to the composition without moisture. Propellants were also prepared to examine the role of moisture in activated charcoal when mix ed along with other known burn rate modifiers like copper chromite and iron oxide. Acti vated charcoal with moisture along with known burn rate modifiers like copper chromite or iron oxide futher increased the burn rate as compared to the propellant containing only copper chromite or iron oxide.

Effect of mixer size on density and burn rate of composite solid propellant

Abstract: Although composite propellant has been studied for a long time, the role of density of the composite propellant has not been addressed adequately in the literature. This paper examines the role of density of the propellant with regards to burn rate and burn rate pressure index. Propellants were prepared in the four different size (15, 70, 200 and 1000 g) mixers with identical input of ingredients to examine the effect of mixer size on density and burn rate of the composite solid propellant. It is noticed that density of the propellant changes significantly with change in the mixer size from 15 g to 1000 g. This is because for the smaller mixers the surface area to volume ratio is large and the actual percentage of aluminum and ammonium perchlorate (especially coarse) that goes into the propellant reduces. High burn rate with decrease in mixer size is also accompanied with increase in burn rate pressure index. These high burn rate and burn rate pressure index with decrease in mixer size is because of larger fraction of coarse AP particles as compared to fine AP particles remaining attached to the mixer and decreases the coarse to fine AP ratio in the cured propellant.

Modeling the Combustion of a Sub-micron Aluminum Particle (PREPRINT)

Abstract: : We examine the shrinking-core model of aluminum combustion, one that has been proposed for sub-micron drop diameters In this model a core of liquid aluminum is surrounded by alumina through which O atoms diffuse to the aluminum surface It is shown that the volume changes intrinsic to this model necessarily lead to fracturing of the alumina and the creation of cracks and voids A simple mathematical model is described which, because of the length and time scales, is quasi-steady, permitting analytical solutions When the drop is sufficiently small this leads to a simple formula for the burn time as a function of the atmospheric pressure, temperature, and oxygen concentration, which is tested against experimental data By introducing a fractal ingredient into the description, motivated by the fracturing, it is possible to generate agreement with experimental data on the variations of burn time with drop diameter, a d025-t law in contrast with the familiar d sq-t law of classical fuel-drop combustion For larger drop diameters a non-linear differential equation for the burn rate is derived whose integration yields the burn time as a function of drop diameter and shows a transition from the d025 law to a d sq law

Burn-back Equations for High Volumetric Loading Single-grain Dual-thrust Rocket Propellant Configuration (Review Paper)

Abstract: Dual-thrust mode is adopted in solid propellant rocket propulsion through tailoring of burning area, nozzle, rocket motor chamber, propellant type, multiple propellant blocks. In the present study, mathematical formulation has been evolved for generation of burning surface area with web burnt for a simple central blind hole in a solid cylindrical propellant geometry with proper partial inhibition on external and lateral surfaces. The burn-back equation has been validated by static firing and parametric study was conducted to understand effect of various control geometrical parameters. The system is utilised for high volumetric loading, single propellant, single composition, single-chamber, single nozzle dual-thrust mode of burning profiles in rocket application. Defence Science Journal, 2011, 61(2), pp.165-170 , DOI:http://dx.doi.org/10.14429/dsj.61.41

Influence Factors of Slow Cook-off Characteristic for Solid Propellant

Abstract: The influence factors of slow cook-off characteristic of solid propellant were studied with the object of HTPE propellant and GAP propellant by means of slow cook-off experimental device and thermocouple and pressure transducer,including composition,burning rate,heating rate,restriction condition,and free volume.HTPE propellant has a very excellent slow cook-off characteristic,hut the addition of HMX increases the response extent of slow cook-off,and free volume has not obvious effect on the response extent of slow cook-off.The burning rate from 10.1 mm·s-1 to 32.2 mm·s-1 and the heating rate from 3.3 ℃·h-1 to 1 ℃·min-1 have scarcely effect on the response extent of slow cook-off characteristic of GAP propellant.The response extent of slow cook-off becomes more violent after increasing restriction condition for HTPE propellant and GAP propellant.

The role and importance of porosity in the deflagration rates of HMX-based materials

Abstract: The deflagration behavior of thermally damaged HMX-based materials will be discussed. Strands of material were burned at pressures ranging from 10-300 MPa using the LLNL high pressure strand burner. Strands were heated in-situ and burned while still hot; temperatures range from 90-200 C and were chosen in order to allow for thermal damage of the material without significant decomposition of the HMX. The results indicate that multiple variables affect the burn rate but the most important are the polymorph of HMX and the nature and thermal stability of the non-HE portion of the material. Characterization of the strands indicate that the thermal soak produces significant porosity and permeability in the sample allowing for significantly faster burning due to the increased surface area and new pathways for flame spread into the material. Specifically, the deflagration rates of heated PBXN-9, LX-10, and PBX-9501 will be discussed and compared.

Star Grain Regression under Spin Induced Acceleration Effect

Abstract: Spinning is used in some of solid rocket motors to increase the flight trajectory precision or for stability requirements. The angular acceleration due to the spin effect increases the burning rate of solid propellant and changes the motor performance by increasing the operating pressure and decreasing the burning time. So it is important to know the grain regression taken place in the solid propellant rocket motor in the acceleration field. In this study, we represent the grain regression analysis of two-dimensional axis-symmetric star grain configuration of the solid propellant rocket motor under spin induced acceleration effect to study how the spin affects on the internal ballistics of the solid rocket motor. Grain regression is done by two methods - geometrical approach and numerical approach. The burning rates on the propellant surface are different with its radial distance, acceleration vector angle and surface slope when the rocket is spinning. With the different burn rates on the propellant surface, the propellant surface perimeter and port area are computed by using the numerical method, and the results are compared with that of constant burn rate.

Principle and Method of Measuring Actual Burning Rate of Propellant by Closed Bomb

Abstract: A principle and a method of measuring actual burning rate of propellant by closed bomb were presented.Through designing standard structure propellant and revising the form function,it can eliminate the influence of burning surface area change on burning performance test.Using the methods of mathematics transform to dispose the testing data,it can eliminate the influence of pressure change on burning performance test.Through tests to validate this method,the results testify that the principle and method can get actual burning rate of propellant.

Heat Transfer Induced Irregular Burning

Abstract: In the analysis of a ballistic evaluation motor (BEM) for extruded double base motors, an irregular burning phenomenon was encountered. The irregular burning manifested as a drop in pressure. This was not combustion instability as acoustic instability is associated with high frequency pressure oscillations 2 , usually a mean pressure rise and steep fronted shock waves. L* instability is also eliminated as a possibility as the pressure is much higher than the pressure associated with this phenomena (>5MPa), and the pressure did not oscillate at the frequencies associated with L*-instability 3 . The reason for this phenomenon lies in the unique design of the motor. An internal/external burning cylindrical grain, inhibited on one end, was used. The external grain initially vented onto the cold steel wall, resulting in a significant heat loss. Experimental measurements of the steel wall temperature indicated that the heat losses to the wall were in the order of 14-15kJ or 4-6% of the heat released by combustion. This heat loss is also only experienced by the external burning surface of the propellant, thus amplifying the heat loss effect for the outside wall of the propellant. Inserting a silica-phenolic sleeve eliminated the irregular burning. This indicated that the heat transfer at the external wall surface is the most likely source of the irregular burning. This was also confirmed by the increase in severity of the irregular burning for cold firings, where the total heat loss was increased due to the lower temperature of the motor wall and more energy was absorbed by the casing to heat the surface. The burn rate was measured from the pressure/time curve because pressure was the only practical parameter that could be measured. It is not, however, the mechanism that drives the burn rate but is directly related to it. The propellant decomposes at the surface when the surface temperature reaches the ignition temperature. This mass flux from the surface moves into the flame zone and then combusts releasing heat. This heat is then absorbed by the surface, speeding up the decomposition (usually, some form of the Arrhenius equation applies to the surface decomposition), increasing the mass flux and finally increasing the heat released by the combustion. The increased mass addition and flame temperature resulted in the increase in pressure. In most cases the propellant is an internally burning grain that is axis symmetric. In the case of this BEM motor it is also burning from the outside. This has a significant effect on the heat transfer within the motor. In an internally burning axis symmetric motor the radiative heat transfer away from the surface is the same as the incoming energy from the surface directly opposing it. Thus, both surfaces transfer energy to each other. This may not be significant in terms of conductive as most of the conductive heat transfer to the surface comes from the combustion nearest the surface, but can be significant in terms of radiation. Basic propellant combustion theory states that radiation can account for up to 25-30% of the heat transfer to the surface. The energy radiated from the propellant surface should be the same as the energy radiated from the opposing surface. In the case of the motor burning inwards the opposing surface (in this case a steel wall) is not another source of radiant heat but will only be a heat sink. This, combined with the cooling of the combustion gasses, causes the propellant surface not to heat up at the same rate as the propellant surface in the internally burning grain. Thus, its burn rate is lower, resulting in a drop in pressure. This phenomenon warranted a more fundamental approach in the burn rate calculation and heat transfer modeling to confirm the above assertions. To accomplish this, a complete heat transfer calculation was performed. In parallel a CFD simulation with radiation and conduction effects was performed. The results of the two approaches showed remarkable similarity and both pointed to the heat losses as the origin of the irregular burning.

Effects of Ammonia Borane on the Combustion of an Ethanol Droplet at Atmospheric Pressure

Abstract: Adding hydrogen containing energetics to liquid fuels has the potential to change the combustion behavior of the fuels, thereby enhancing performance. One potential energetic additive is ammonia borane (AB), which contains 19.6 wt.% hydrogen and can be dissolved in anhydrous ethanol (up to 6.5 wt.%). Single droplet combustion experiments are performed that vary the AB concentration in ethanol from 0 to 6 wt.%. These experiments indicate that the addition of 6 wt.% of AB increases the burn rate of the fuel by 16% during normal regression. As the ethanol droplet containing AB burns, gas formation within the droplet causes it to expand slightly. Gas bubbles generated during burn reach the surface of the droplet and rupture, jettisoning liquid and increasing the burn rate. High speed (5 kHz) planar laser-induced uorescence (PLIF) measurements show an increase in OH production during combustion, indicating that the reaction rate of the fuel increases with the addition of AB. PLIF is also used to capture the jetting dynamics. The eects of hydrogen gas and other decomposition products of AB in the fuel vapor surrounding the droplet are evident in these dynamic measurements that aect the burn rate. AB addition to the fuel causes the droplet to shatter towards the end of the droplet lifetime causing atomization and rapid combustion of the remaining fuel. This droplet shattering is similar to some other systems, but has not previously been reported for ammonia borane systems.

In-Cylinder Combustion Analysis of a Yamaha R6 FSAE Engine: Achieving Improved Engine Performance and Efficiency through Burn Rate Combustion Diagnostics

Abstract: .................................................................................................................................... ii Acknowledgements .................................................................................................................. iii List of Tables ........................................................................................................................... vi List of Figures ......................................................................................................................... vii List of Abbreviations .................................................................................................................x Chapter 1: Background ..............................................................................................................1 Chapter 2: Working Principles of the System............................................................................5 2.1 Spark Ignition Engine Pre-Mixed Flame Development ...................................................5 2.2 Calculating Performance and Efficiency Indicators from the P-V Diagram ...................6 2.3 Quantifying Combustion Variability ...............................................................................8 Chapter 3: Modeling Techniques and Equations .....................................................................10 3.1 Combustion Modeling Approaches ...............................................................................10 3.2 Derivation of the Mass Fraction Burn Rate ...................................................................13 Chapter 4: Instrumentation & Programming ...........................................................................20 4.1 Piezoelectric Pressure Sensors .......................................................................................20 4.2 Incremental Angular Encoders ......................................................................................27 4.3 Data Acquisition ............................................................................................................28 4.4 Data Integrity and Analysis ...........................................................................................30 Chapter 5: Component Machining & Dynamometer Setup .....................................................34

A Model of Burning Rate Characteristics for Heterogeneous Propellants over a Wide Range of Pressure

Abstract: The variation of steady burning rate for heterogeneous propellants over a wide range of pressures is studied by experimental methods and theoretical analysis. Two kinds of heterogeneous propellants, GATo-1 composite modified double-base (CMDB) propellant and 86 series composite propellant are selected. An explanation that two inflection points observed in burning rate-pressure curve is presented. It is concluded that condensed phase reaction dominates the burning process in low pressure region which presents a small pressure exponent, while heat feedback from gas phase predominates in high pressure region with a large pressure exponent. In medium pressure region, these two factors influence the burning process cooperatively and pressure exponent exhibits the trend of reducing and then rising. Results calculated from the model show good agreement with experimental data. Thermal parameters analysis indicates that burning rate is more sensitive to the thermal conductivity of condensed phase, activation energy of condensed phase, specific heat capacity and delay time of heat release of condensed phase reaction in high pressures. Insights and guideline of burning rate characteristics in high pressures and interior ballistic properties for solid propellant impulsive microthruster are provided by our model.

Regarding the evaluation of the solid rocket propellant response function to pressure coupling

Abstract: High frequency combustion instabilities imply a major risk for the solid rocket motor stable working and they are directly linked to the propellant response to chamber pressure coupling. This article discusses a laboratory testing method for the measurement and evaluation of the pressure coupled response for non-metalized propellants in a first stage. Experimental researches were done with an adequate setup, built and improved in our lab, able to evaluate the propellant response by interpreting the pressure oscillations damping in terms of propellant response. Our paper aims at defining a linearized one-dimensional flow study model to analyze the disturbed operation of the solid propellant rocket motors. Based on the applied model we can assert that the real part of propellant response is a function of the oscillations damping, acoustic energy in the motor chamber and various losses in the burning chamber. The imaginary part of propellant response mainly depends on the normalized pulsation, on the burning chamber gas column and on the pressure oscillations frequency. Our research purpose was obviously to minimize the risk of the combustion instabilities effects on the rocket motors working, by experimental investigations using jet modulating techniques and sustained by an interesting study model based on the perturbation method.

Development of a Novel Thermoelectric Propellant Temperature Controller For Strand Burning Studies

Abstract: Development of a device capable of precise temperature control and rapid thermal equilibration of propellant strands is discussed. The device is comprised of three dual-stage thermoelectrics controlled by a pulse width modulating, proportional, integral, derivative hardware controller and a set of gaseous nitrogen cooling blocks for conveyance of heat out of a pressurized combustion vessel. The device is capable of operating over temperatures ranging from -25 to 70°C and at pressures from vacuum to 11.7 MPa (1700 psi). Operability testing shows the device is capable of thermal equilibration to any defined temperature with an error of ±3.75°C in less than ten minutes. The device will not only be useful in investigating temperature dependence of solid propellant burning rate but will be particularly useful in determining the burning rate effects of phase change in propellants such as aluminum and ice (ALICE) and hydrogen peroxide, aluminum, and water (PALICE) frozen propellants.

Combustion Characteristics of Pool Fires of Biofuels and Their Blends with Diesel Fuel

Abstract: The combustion characteristics of pool fires of biofuels, canola methyl ester (CME), soy methyl ester (SME) and their blends with No. 2 diesel were studied. The fuels were burned in cups of two sizes (4.2 cm and 5.7 cm in diameter and 3.8 cm height) that simulated pools of liquids during spills. The mass burn rate, the radiation emission from the flames, the flame temperature field, and the emission index of CO and NO were recorded. Blends of CME and SME with diesel fuel were tested with biofuel concentrations of 25, 50 and 75% volume. The fuel mass burning rate in the smaller container was less than that in the large container for all flames. The radiation fraction of heat release and the temperature profiles were comparable for all the flames. The CO emission index decreased with the biofuel content in the fuel, and the NO emission index, albeit being much smaller than in the spray flames of corresponding fuels, did not show a systematic and significant dependence on the fuel type.

Specific Design Features of Solid Propellant Rocket Motors for Shoulder-Launched Weapon Systems *

Abstract: Solid propellant rocket motors for Shoulder Launch ed Infantry Weapon Systems (SLWS) are characterized with a very short burning time, high-pressure combustion and a wide spectrum of design solutions for rocket motor structure. Interior ballistic behaviour of such rocket motors depends o n many factors such as design structure, propellant grain shape, propellant grain joint to the rocket motor case, type and location of the igniter, spinning mode and nozz le design. Erosive burning also plays important role due to high combustion gases mass fl ow rate. Numerical simulation of the igniter combustion gases flow through the hollo w of the propellant grain tubes with gas temperature distribution was carried out in thi s paper. Results confirmed assumptions that igniter interior gas flow affected duration of the pressure rise. A mathematical model approach for prediction of the curve p = f( t) which was included in a model of the corrected propellant grain burnin g surface for two types of short-time rocket motors has been presented. A good agreement with measured curves was achieved.