Showing papers on "Burn rate (chemistry) published in 1997"

Cigarette sidestream smoke and free-burn rate control device

Abstract: A device for minimizing cigarette sidestream smoke and reducing free-burn rate of a burning cigarette, the device comprises: i) a non-combustible tubular element (12) encasing an effective length of a tobacco charge (22) of a cigarette (14) located in the tubular element; and ii) the tubular element having a means for both minimizing sidestream smoke emission from a burning tobacco charge and reducing free-burn rate of such burning tobacco charge to increase number of puffs from the burning tobacco charge.

Substantially smoke-free and particulate free inflator for inflatable safety restraint system

Abstract: Provided is an inflator for providing propellant gases to inflate a bag of an automotive safety system for restraining an occupant of a vehicle in which the inflator is a pure pyrotechnic inflator that generates propellant gases that are substantially free of metal-containing particulate and condensable materials without requiring storage of a pressurized gas. The propellant gases are generated from a solid gas-generating propellant composition including two components. A first component is fuel-rich and has a fast burn rate. The second component is a oxidizer that, although having a slow burn rate, operates in combination with the first component at an acceptably fast burn rate to oxidize carbon monoxide and/or hydrogen produced during combustion of the first component to carbon dioxide and/or water, respectively.

Use of the Nonlinear Dynamical System Theory to Study Cycle-to-Cycle Variations from Spark Ignition Engine Pressure Data

Abstract: Cycle-to-cycle variations in the pressure evolution within the cylinder of a spark ignition engine has long been recognized as a phenomenon of considerable importance. In this work, use of tools borrowed to the nonlinear dynamical system theory to investigate the time evolution of the cylinder pressure is explored. By computing a divergence rate between different pressure cycles versus crank angle, four phases during the combustion cycle are exhibited. These four phases may be identified with the four common phases evidenced by burn rate calculations [1]. Starting from phase portraits and using Poincare sections, we also study correlations between peak pressures, IMEP and the durations from ignition to appearance of a flame kernel. Accounting for the fact that, during the ignition phase of the combustion cycle, trajectories in a plane projection of the reconstructed phase portrait associated with cycles in the case of motored engine cannot be distinguished from trajectories corresponding to combustion cycles, we estimate the duration of the ignition phase without any prior assumption on the combustion processes. Fluctuations of ignition phase lengths have been found to be correlated with IMEP standard deviations.

Staged pyrotechnic air bag inflator

Abstract: An inflator (10) includes a housing (30) defining a chamber (50). An end-burning pyrotechnic charge (60) is in the chamber (50) in the housing (30) for, upon ignition, producing inflation fluid for inflating an inflatable device. The charge (60) has a first predetermined burn rate at a first location along the length of the chamber (50) and a second predetermined burn rate at a second location along the length of the chamber spaced apart from the first location. The second predetermined burn rate is different from the first predetermined burn rate. The inflator (10) includes an initiator (112) for igniting the charge (60), and portions (34, 90, 110) for directing inflation fluid from the chamber (50) to the inflatable device.

Combustion control requirements in high loading density, solid propellant ETC gun firings

Abstract: This report investigates and quantifies the requirements for increasing performance (muzzle kinetic energy) in existing high performance (tank) gun systems utilizing solid propellants. Factors studied include propelling charge mass or loading density, propellant specific energy, grain geometry, and the use of electrothermal-chemical concepts. Results indicate that significant increases in performance require not only increased system energy but, more importantly, propellant combustion control In terms of the mass generation rate to operate near ideal gun performance. Specific requirements and approaches (e.g., burn rate modification by plasma interaction) to obtain the "ideal" gun performance are discussed. Plasma-propellant interaction study results are also discussed.

The Application of Electrothermal-Chemical (ETC) Propulsion Concepts to Reduce Propelling Charge Temperature Sensitivity

Abstract: : The concept of the propelling charge temperature coefficient and its impact on gun performance is explored. Ballistic factors, in addition to the propellant burn rate dependence on initial temperature that often increases the magnitude of the propelling charge temperature coefficient, are identified and discussed. Techniques for moderating the effects of the propelling charge temperature coefficient are presented, with special emphasis on the utilization of electrothermal- chemical (ETC) concepts.

Investigation of the combustion process of gel propellants

Abstract: The trend today in high energy long storage rocket fuel and propellants research is to develop combustion systems which introduce the use of gel fuels and propellants. The idea stems from the fact that the addition of gelling agents and/or metallized additives to the liquid fuels enhances significantly the performance, density and specific impulse of liquid fuels. Understanding the evaporation and combustion process of single droplets of gel fuels is the first basic step to predict their behavior in the future combustion chambers. An experimental investigation of the evaporation and combustion phenomena of gelled fuel droplets is presented using a sophisticated state of the art computerized video system. The burning rate of the droplet was measured and the combustion process was viewed and investigated closely. A parametric investigation was conducted to evaluate the effect of the chamber pressure and the oxygen mass fraction on the droplet burning rate. The results indicate that gels burn at lower burning rates than liquids and ignition is more difficult.

Multiphase oxidation of metals

Abstract: The burning of metals in enriched oxygen atmospheres includes multiple phases and, therefore, satisfies the definition of being a heterogeneous system. For a reaction to be considered heterogeneous, however, the site of the chemical transformation must be at an interface. While the Wagner theory of metal oxidation gives satisfactory mechanisms for this type of reaction for solid metals, no comparable theory is available for metal oxidation when the temperature is above the melting point of the metal. Similitude theory is applied to metal combustion under conditions of the NASA/ASTM flammability test system in order to identify rate-controlling regimes and conditions of heterogeneous reaction. Above 4 MPa, the observed burn rate for iron is proportional to the sample dimension, L, as L−0.7. Analysis shows that the heat transfer rate between the reaction surface and the solid rod is proportional to L−0.5 and may, therefore, be the rate-determining process for the system.

Laser recoil response of HMX, RDS, N5, GAP/BTTN and others from 1 to 6 atmospheres

Abstract: HMX and RDX pellets were pressed and tested in the laser recoil device at pressures from 1 to 6 atm. The laser flux had an average power of 31 W/cm2, with sinusoidal oscillations from 15 to 43 W/cm2. A pressure insensitive force transducer obtained from Zarko was used in these tests. The force transducer has a resonant frequency at 160 Hz, but performs well at frequencies below 100 Hz. Samples inhibited with halocarbon grease, and with glass tubing were compared. HMX showed twice the response amplitude of RDX. Both HMX and RDX have a broad flat response. In contrast with the previously measured N5 double base propellant, the recoil response of neither HMX nor RDX changed significantly with pressure up to 60 psia. Burn rate vs. laser flux was measured for HMX, RDX and N5 up to 60 W/cm2 at 13.7 and 60 psia. At one atmosphere the rate vs. flux slope of HMX and RDX is about the same, but N5 is 3 times. Other ingredients tested include GAP, GAP/BTTN, NMMO gumstocks, BAMO pellets, and IRQ minimum smoke propellant. GAP and BAMO would not burn with a stable flame, but chuffed large puffs of nitrogen. NMMO ignited readily, but did not produce force oscillations. GAP/BTTN had a very high unstable laser recoil response, even bigger than N5 double base propellant, which did not decrease significantly with pressure.

Calculation study on burning rate characteristics of nitramine propellant

Abstract: A new steady combustion model of nitramine/AP propellant, the Two Combustion Zone model, was developed. In this model, the effect of the type and the numbers of the surplus oxidizer on the burning rate of the propellant was considered. The numerical results showed that the effect of the surplus oxidizer on the burning rate of the propellant is big enough, which must be considered in the combustion model.

Development of high burn rate azide polymer propellant

Abstract: The combustion characteristics of high burn rate azide polymer composite propellant were examined by using ⊘ 70 mm BMI composite rocket motor of L/D = 16. Azide polymer propellants have much higher specific impulse than HTPB propellant at below the AP content of 85%. AP/B/N propellants showed a plateau-mesa burning at a pressure range between 7 MPa and 15 MPa with a burn rate of approximately 28 mm/s. Very fine AP, however, diminished this favorable combustion characteristics. The AP/B/N also had excellent mechanical properties at an operational temperature range to enable the case bonded. The BMI motors burned stably and showed an excellent performance. The thickness of the insulation at the aft closure, 5.5 mm including a stress relief boot, was good enough.

A Preliminary Study of the Wet Towel Hybrid Rocket Motor.

Abstract: To overcome defects of conventional hybrid rocket motors such as poor mass ratio and low combustion efficiency, the authors propose a new idea of design. The point of this idea is that the motor uses ‘wet towel propellant’, in which liquid oxidizer fills gap space in the fuel bed consisting of plastic fibers or films. Specific impulses of the propellants are estimated theoretically with HAN and LOX as liquid oxidizer. To investigate pressure sensitivity of the burning rates of the propellants, burning rates of LOX-polyaramid propellants are measured in a pressure range of 0.1 to 2MPa with a specially designed strand burner. A striking feature in the experimental results is that the enhancement of the burning rate with ambient pressure is as big as the pressure exponent being larger than one. Further innovation is necessary for realization of the proposed idea.

Mechanical properties of azide polymer propellant at ignition stage

Abstract: The rates of pressure increase for the 6.8 grams of the igniter powder and the igniter pellets at an ignition stage are measured by using dummy rocket motors. The mechanical properties of high burn rate azide polymer propellant at low temperature and at simulated ignition condition are also studied here. Although the effect of the pressure increase on the deformation of the propellant surface is smaller than that caused by the thermal stress, it becomes important when the rate of pressure increase is very high. The rates of pressure increase are for powder, 1.9 × 107 kgf/cm2 /min, and for pellets, 2.6 to 7.5 × 105 kgf/cm2 /min. The high speed tensile test well simulates the ignition condition in a view point of the viscoelasticity. The propellant used here here has excellent mechanical properties even at a low temperature.

Laser recoil combustion response of hnf oxidizer from 1 to 6 atmospheres

Abstract: Solid oxidizer HNF, hydrazinium nitroformate, crystals were pressed into pellets and burned as a monopropellant in the laser recoil device at pressures from 1 to 6 atm. The laser flux was set at 2 different mean power levels with sinusoidal oscillations of 21 ± 11.5 W/cm2 or 52 ± 5.9 W/cm2. Oscillations were logarithmically swept over the range from 4 to 80 Hz. A Russian inductive force transducer was used in these tests. The force transducer has a resonant frequency at 140 Hz, but performs well at frequencies below 100 Hz. HNF pellets were inhibited with a glass ring on the sample pedestal. HNF burns nicely with a steep burn rate slope of 0.892 down to 5 psia. Burn rate vs. laser flux was measured up to 60 W/cm2 at 1,2, 3,4, and 6 atm pressure. The slope of the bum rate vs laser flux was nearly the same at each pressure. At one atmosphere the rate vs. flux slope of HNF is slightly lower than HMX and about the same rate as HMX. Ignition times decreased from 760 ms to 78 ms as the laser flux increased from 21 to 67 W/cm2. Ignition times also decreased as pressure increased. HNF thrust oscillations decreased from 65 to 20 mg/cm2 over the frequency range of 4 to 80 Hz. The thrust oscillations also decreased 50% in amplitude as the pressure increased to 4 atm. The thrust oscillations are about the same order of magnitude as for HMX, less than N5 and greater than RDX. A phase crossover frequency was not determined as it was below the 4 Hz frequency range of these tests. The phase lag decreased very slightly with pressure, but was nearly the same at low frequencies.

An Investigation to Improve Quality Evaluations of Primers and Propellant for 20mm Munitions

Abstract: To reduce the frequency of electrically initiated, 20mm munition hangfires (delayed ignitions), a joint Army/NASA investigation was conducted to recommend quality evaluation improvements for acceptance of both primers and gun propellant. This effort focused only on evaluating ignition and combustion performance as potential causes of hangfires: poor electrical initiation of the primer, low output performance of the primer, low ignition sensitivity of the gun propellant, and the effects of cold temperature. The goal was to determine the "best" of the Army and NASA test methods to assess the functional performance of primers and gun propellants. The approach was to evaluate the performance of both high-quality and deliberately defective primers to challenge the sensitivity of test methods. In addition, the ignition sensitivity of different manufacturing batches of gun propellants was evaluated. The results of the investigation revealed that improvements can be made in functional evaluations that can assist in identifying and reducing ignition and performance variations. The "best" functional evaluation of primers and propellant is achieved through a combination of both Army and NASA test methods. Incorporating the recommendations offered in this report may provide for considerable savings in reducing the number of cartridge firings, while significantly lowering the rejection rate of primer, propellant and cartridge lots. The most probable causes for ignition and combustion-related hangfires were the lack of calcium silicide in the primer mix, a low output performance of primers, and finally, poor ignition sensitivity of gun propellant. Cold temperatures further reduce propellant ignition sensitivity, as well as reducing burn rate and chamber pressures.

Use of the nonlinear dynamical system theory to study cycle-to-cycle variations from spark ignition engine pressure data

Abstract: Cycle-to-cycle variations in the pressure evolution within the cylinder of a spark ignition engine has long been recognized as a phenomenon of considerable importance. In this work, use of tools borrowed to the nonlinear dynamical system theory to investigate the time evolution of the cylinder pressure is explored. By computing a divergence rate between different pressure cycles versus crank angle, four phases during the combustion cycle are exhibited. These four phases may be identified with the four common phases evidenced by burn rate calculations [1]. Starting from phase portraits and using Poincare sections, we also study correlations between peak pressures, IMEP and the durations from ignition to appearance of a flame kernel. Accounting for the fact that, during the ignition phase of the combustion cycle, trajectories in a plane projection of the reconstructed phase portrait associated with cycles in the case of motored engine cannot be distinguished from trajectories corresponding to combustion cycles, we estimate the duration of the ignition phase without any prior assumption on the combustion processes. Fluctuations of ignition phase lengths have been found to be correlated with IMEP standard deviations.

Oscillations on Electrothermal-Chemical (ETC) Closed-Chamber JA2 Burn-Rate Reductions

Abstract: : Closed-chamber electrothermal-chemical firings of JA2 7-perf propellant revealed the presence of oscillations on its burn rates. This study was originated by questions raised about the probable causes for this phenomena. The discussion that follows analyses the fiber-optic links utilized for the data acquisition, the code used to deduce the propellant burn rate and the filtering for data smoothing, and finally, the effect of the plasma injection and its energy. The result of this analysis suggests that the plasma injection contributes to burn-rate oscillations, that their amplitude is proportional to the energy of the plasma injected, and that the oscillations are a function of the closed-chamber pressure.

Ultrasonic Measurement of the Regression Rate of the Melting Interface in Burning Metal Rods

Abstract: Results of tests in which metallic rods are burned in oxygen-enriched atmospheres often include the regression rate of the melting interface for the burning test specimen. This regression rate is used as an indication of a metallic materials' relative flammability when different metallic materials burn at the same test pressure and as an indication of a metallic materials' general ability to sustain burning under the test conditions. Most past methods for determining this regression rate have been based on visual interrogation which is costly, time consuming and often inaccurate. Inaccuracies typically result due to the specific method used (scale factors, optics, visualisation techniques, etc.), and often a limited access time with which to view the combustion event, or obscuration of the burning by condensed-phase products. An ultrasonic transducer has been fabricated to directly measure the regression rate of the melting interface established during the burning of metal rods, thus eliminating many of these problems. The transducer is described, typical results for an iron rod burning in pure oxygen are given along with a comparison of these results with regression rates obtained from visual interrogation. Excellent agreement is obtained between the regression rate obtained from the new ultrasonic transducer and regression rates obtained visually, thus demonstrating the transducer's usefulness.

Adiabatic Pyrolysis of Solid Propellants

Abstract: Several one-step, irreversible, zero-order pyrolysis models (Arrhcnius, KTSS, and Merzhanov Dubovitskii high activation energy pyrolysis), commonly used to study adiabatic burning of energetic materials with arbitrary pressure and initial temperature, are revisited. Motivated by experimental and theoretical work performed in 1984 by students of this laboratory, a relationship among the several interplaying parameters is found under steadystate conditions. This relationship corresponds to the Jacobian of the sensitivity parameters used in the Zeldovich Novozhilov approach. If the Vieille steady burn rate law is enforced, consistency requires an explicit pressure dependence for both Arrhenius and KTSS pyrolysis. But if the normal (or Zeldovich) steady burn rate is enforced, no explicit pressure dependence is required for both Arrhenius and KTSS pyrolysis. Other constraints arise for the Merzhanov Dubovitskii pyrolysis model. The unifying concept for these different trends is the Jacobian consistency between the implemented steady pyrolysis and ballistic laws. The dependence of the pre-exponential factor on surface activation energy (known as kinetic compensation) is shown to be linear (Arrhenius) or almost linear (Merzhanov Dubovitskii), for any given experimental data set under steady burning. Experimental results are reported for a variety of solid propellants of different nature. NOMENCLATURE * Copyright © 1997 by Luigi DeLuca. Published by the American Institute of Aeronautics arid Astronautics, Inc. with permission. ' Professor, Dipartimento di Energetica; Fax: (39-2) 2399-3940, e-mail: DeLuca@icil64.cilea.it. Associate Fellow. ' Professor, Dipartimento di Matematica ' PhD. Candidate, Dipartimento di Energetica ' Lab. Technician, Dipartimento di Energetica " MSc. Candidate, Dipartimento di Energetica a&, &(, = multiplicative factors denned by Eq. 3.3, Eq. 3.4 A, B = nondim. functions used in frequency response analysis Ac = multiplicative factor used in the Merzhanov-Dubovitskii pyrolysis, 1/s, in Eq. 2.3 AS, Bs = multiplicative factors defined by Eq. 2.1 (Arrhenius ), Eq. 2.2 (KTSS) c = specific heat, cal/g K £{...) = activation energy, cal/mole J5(...) = .Z?(...)/9J/T(...) , riondim. activation energy Is = external radiant flux intensity, cal/cmPs j = running counter, integer k = ZN steady sensitivity parameter defined in Eq. 6.3 ra = mass burn rate, g/crn s Ms = pre-exponential factor of Zeldovich (or normal) mass burn rate, g/cms, in Eq. 2.6 n = pressure exponent of ballistic steady burn rate defined by Eq. 2.4 n., = pressure exponent of pyrolysis law denned by Eq. 2.1 (Arrhenius ) and Eq. 2.2 (KTSS) IT-TS — pressure exponent of steady surface temperature defined by Eq. 3.1 p = pressure, atm Pref, Tref = reference pressure (68 atm), reference temperature (298 K) Q = heat release, cal/g (positive if exothermic) r =ZN steady sensitivity parameter denned in Eq. 6.4 rj, — burn rate, cm/s n,ref>TSiref = reference burn rate n(pref), reference surface temperature Ts(pTnf) 5R = universal gas constant; 1.987 cal/moleK T = temperature, K TI = initial propellant temperature, K ws = power of KTSS pyrolysis law, denned in Eq. 2.2 Greek Symbols a = thermal diffusivity, cm/s 8 = ZN Jacobian defined in Eq. 6.5 fj,, v = ZN steady sensitivity parameters defined in Eq. 6.2, Eq. 6.1 p = density, g/cm cTp = steady temperature sensitivity of burn rate, 1/K, denned by Eq. 2.4 T« = pressure exponent of surface temperature defined by Eq. 3.1 if} = variable defined in Eq. 4.19 Subscripts Arr = Arrhenius pyrolysis law bal = ballistic c = condensed g = gas hig = high KTSS = KTSS pyrolysis law low = low p = pressure pyr = pyrolysis ref = reference s = burn surface Vi = Vieille burn rate law Ze = Zeldovich burn rate law (...)1 = cold boundary value Superscripts (...) = steady-state value (...) = dimensional value (...) — average value

ontrol Requirements in High Loading Density, Solid Propellant ETC Gun Firings

Abstract: This report investigates and quantifies the requirements for increasing performance (muzzle kinetic energy) in existing high performance (tank) gun systems ropellants. Factors studied include propelling r loading density, propellant specific energy, etry, and the use of electrothermal-chemical concepts. Results indicate that significant increases in pe~o~ance require not only increased system energy but, more ~po~~tly, propellant combustion control in terms of the mass generation rate to operate near ideal gun erformanee. Specific requirements and approaches (e.g., burn rate modification by plasma interaction) to obtain the "ideal" gun pe~orma~ce are discussed. Plasma-propellant interaction study results are also discussed.

Ignition and unsteady combustion of ap-based composite propellants in subscale solid rocket motors with nozzle seal

Abstract: A mathematical model has been developed to predict the unsteady behavoir of AP-based composite propellants during the ignition transient period in solid rocket motors. Ignition occurs in such a very short delay that the unsteady behavior of the propellant has a dominant effect on ignition propagation. The motivation of this study was to eliminate the approximations used generally in the description of propellant combustion : - surface temperature criterion to determine the ignition delay, - steady state expression for the burning rate, - constant propellant flame temperature from ignition to steady state. The numerical package developed for the simulations comprises a multi-dimensional CFD code using a k-E turbulence model which is associated with a onedimensional unsteady heat conduction analytical procedure giving the propellant burning rate and an analytical description of the propellant flame. In addition, one considers that an ignition propagation occurs by conduction at the gas side propellant surface when the critical inflammation conditions to sustain such a propagation are satisfied. The effect of radiation is included in the computation process by adding at the propellant surface a constant radiation heat flux estimated a priori. The computer code has been applied to two different rocket motor configurations.

The study of nutation instability on spinning solid propellant rockets

Abstract: HMX and RDX pellets were pressed and tested in the laser recoil device at pressures from 1 to 6 atm. The laser flux had an average power of 31 W/cm, with sinusoidal oscillations from 15 to 43 W/cm. A pressure insensitive force transducer obtained from Zarko was used in these tests. The force transducer has a resonant frequency at 160 Hz, but performs well at frequencies below 100 Hz. Samples inhibited with halocarbon grease, and with glass tubing were compared. HMX showed twice the response amplitude of RDX. Both HMX and RDX have a broad flat response. In contrast with the previously measured N5 double base propellant, the recoil response of neither HMX nor RDX changed significantly with pressure up to 60 psia. Burn rate vs. laser flux was measured for HMX, RDX and N5 up to 60 W/cm at 13.7 and 60 psia. At one atmosphere the rate vs. flux slope of HMX and RDX is about the same, but N5 is 3 times. Other ingredients tested include GAP, GAP/BTTN, NMMO gumstocks, BAMO pellets, and IRQ minimum smoke propellant. GAP and BAMO would not burn with a stable flame, but chuffed large puffs of nitrogen. NMMO ignited readily, but did not produce force oscillations. GAP/BTTN had a very high unstable laser recoil response, even bigger than N5 double base propellant, which did not decrease significantly with pressure.