Showing papers on "Burn rate (chemistry) published in 2003"

Enhanced Propellant Combustion with Nanoparticles

Abstract: Enhanced burn rate results are presented for ammonium perchlorate/Al nanoparticle strand burners at atmospheric (and higher) pressure and for the comparative combustion in a high pressure closed vessel of a solid propellant containing 15% of either conventional micrometer-scale Al or nanometric Al The burn rate at the smallest nanometric Al particle size appears to be asymptotically approaching an inverse particle-diameter-squared dependence

Synthesis and characterization of iron carbonyl‐modified hydroxyl‐terminated polybutadiene: A catalyst‐bound propellant binder for burn‐rate augmentation

Abstract: Iron was coordinately linked to the hydroxyl-terminated polybutadiene (HTPB) backbone using iron carbonyl via a ligand displacement reaction. The modified HTPB thus obtained was reddish brown in color and was characterized by GPC, FTIR, NMR, thermal, and propellant studies. No significant changes in the rheology, molecular weight, and molecular weight distribution were seen in the modified resin when the bonded Fe content was >0.8.0 wt %. However, the hydroxyl value of the resin decreased by 3–7 irrespective of the weight percent of the bonded Fe, and this was more likely due to the Fe-catalyzed oxidation of the CH2OH moiety, mostly to the CHO group. Apparently, this has not affected the cure characteristics of the binder, as demonstrated by the good mechanical properties of the gum stock and the propellant. The catalytic efficiency of the bonded Fe on the burn rate of the propellant was more efficient than was the free Fe added to the propellant. The aging characteristics of the resin for the bound iron content of ≤0.8 wt % was apparently good, as its viscosity and molecular weight did not undergo any drastic changes even after 18 months' storage under ambient conditions. © 2003 Wiley Periodicals, Inc. J Appl Polym Sci 90: 2813–2823, 2003

Electrically controlled extinguishable solid propellant motors

Abstract: Motor designs that enable the control of ignition, combustion burn rate, extinguishment, and reignition of solid propellants by precise application of electrical power are provided. Design of such motors, including choice of electrode materials, the form of electric power, and exemplary facial-extent electrode and axial-extent electrode configurations are also provided.

Erosive Burning Study Utilizing Ultrasonic Measurement Techniques

Abstract: A 6-segment subscale motor was developed to generate a range of internal environments from which multiple propellants could be characterized for erosive burning. The motor test bed was designed to provide a high Mach number, high mass flux environment. Propellant regression rates were monitored for each segment utilizing ultrasonic measurement techniques. These data were obtained for three propellants RSRM, ETM- 03, and Castor@ IVA, which span two propellant types, PBAN (polybutadiene acrylonitrile) and HTPB (hydroxyl terminated polybutadiene). The characterization of these propellants indicates a remarkably similar erosive burning response to the induced flow environment. Propellant burnrates for each type had a conventional response with respect to pressure up to a bulk flow velocity threshold. Each propellant, however, had a unique threshold at which it would experience an increase in observed propellant burn rate. Above the observed threshold each propellant again demonstrated a similar enhanced burn rate response corresponding to the local flow environment.

Methods of controlling solid propellant ignition, combustion, and extinguishment

Abstract: An apparatus and method to control ignition, combustion burn rate, extinguishment, and reignition of solid propellant is provided.

Development of erosive burning models for cfd predictions of solid rocket motor internal environments

Abstract: Four erosive burning models, equations (11) to (14). are developed in this work by using a power law relationship to correlate (1) the erosive burning ratio and the local velocity gradient at propellant surfaces; (2) the erosive burning ratio and the velocity gradient divided by centerline velocity; (3) the erosive burning difference and the local velocity gradient at propellant surfaces; and (4) the erosive burning difference and the velocity gradient divided by centerline velocity. These models depend on the local velocity gradient at the propellant surface (or the velocity gradient divided by centerline velocity) only and, unlike other empirical models, are independent of the motor size. It was argued that, since the erosive burning is a local phenomenon occurring near the surface of the solid propellant, the erosive burning ratio should be independent of the bore diameter if it is correlated with some local flow parameters such as the velocity gradient at the propellant surface. This seems to be true considering the good results obtained by applying these models, which are developed from the small size 5 inch CP tandem motor testing, to CFD simulations of much bigger motors.

Method of improving the burn rate and ignitability of aluminium fuel particles and aluminium fuel so modified

Abstract: A method of improving the burn rate and ignitability of aluminium fuel particles, and a thus modified aluminium fuel for use in propellant and explosive compositions and pyrotechnic charges Aluminium fuel particles are treated with an aqueous solution of hydrofluoric acid and a fluoride and/or complex fluoride salt of an alkali metal and/or alkaline earth metal to form a surface layer of a fluoride complex bound to the aluminium fuel particle

Composite propellant compositions

Abstract: The invention relates to propellant compositions comprising a solid inorganic perchlorate oxidizing agent, a nitrogen-containing fuel, and a burn rate catalyst. Such compositions may be used as a propellant material, (e.g., in rocketry), a pyrotechnie material, an explosive material, a light generating material, a heat generating material, or a sound generating material.

ALE3D Model Predictions and Experimental Analysis of the Cookoff Response of Comp B

Abstract: ALE3D simulations are presented for the thermal explosion of Comp B (RDX,TNT) in a Scaled Thermal Explosion Experiment (STEX). Candidate models and numerical strategies are being tested using the ALE3D code which simulates the coupled thermal, mechanical, and chemical behavior during heating, ignition, and explosion. The mechanical behavior of the solid constituents is represented by a Steinberg-Guinan model while polynomial and gamma-law expressions are used for the equation of state of the solid and gas species, respectively. A gamma-law model is employed for the air in gaps, and a mixed material model is used for the interface between air and explosive. A three-step chemical kinetics model is used for each of the RDX and TNT reaction sequences during the heating and ignition phases, and a pressure-dependent deflagration model is employed during the rapid expansion. Parameters for the three-step kinetics model are specified using measurements of the One-Dimensional-Time-to-Explosion (ODTX), while measurements for burn rate are employed to determine parameters in the burn front model. We compare model predictions to measurements for temperature fields, ignition temperature, and tube wall strain during the heating, ignition, and explosive phases.

Improved Strand Burn Rate Reproducibility Using a New Preparation Methodology for Paint-Based Inhibitors

Abstract: : Small-scale ballistic characterisation of composite rocket propellants is a critical step in the development of rocket motors and is achieved by measuring linear strand burning rates determined using the Crawford strand burning technique. To utilise this technique, the propellant strands must burn with an end-burning configuration, which is achieved through the application of a water-based acrylic paint as an inhibitor. To be effective, the inhibitor needs to be of such a quality that the resultant coating on the propellant strands will be defect-free. If the inhibitor coating contains defects, erroneous and irreproducible burn rate data can result, a problem that has been evident at this facility in the past To improve the strand burn rate reproducibility, a series of paint/water compositions, ranging between 65 and 78%(v/v) paint/water, were prepared using a modified preparation methodology. Of the inhibitor compositions investigated, a 70%(v/v) paint/water ratio was found to be optimal. Using this inhibitor composition, coupled with the introduction of a more rigorous inhibitor preparation methodology, a 60% reduction in erroneous burn rate data and a 32% improvement in burn rate reproducibility was achieved.

Effect of Porosity on the Burn Rate of a Gas-Generating Composition

Abstract: In the course of the development of an oxidizer-based extruded composite propellant for use as gas generant, a simple model to describe and evaluate the actual porosity using simple methods was developed. It is briefly described and used to assess the influence of porosity on the burn rate of propellants produced. For the formulation studied the presence of microporosities does not seem to significantly influence the burn rate. On the other hand, volume fraction of porosities in extruded composition grains can exert a large influence and can increase considerably under thermal treatment with a resulting augmentation in burn rate.

ETC closed-vessel experiments with alkali-doped propellants

Abstract: Experiments with the objective to stimulate combustion of solid propellant through addition of electric energy are reported. In this paper, the idea of stimulating combustion by conducting an electric current through the flame is tested. The current leads to resistive heating of the flame, which adds to the reaction heat, and thus amplifies the heat feedback to the unburnt propellant. The electric conductivity of the flame must be sufficiently high in order to reduce random delay and to increase controllability of the electric discharge. The propellant composition has to be modified to meet this requirement. Thermochemical calculations give the equilibrium state and composition, which are used to estimate the flame conductivity. Calculations are carried out for several propellant compositions: some being considered for preparation, while others are intended solely for comparison. The calculation results indicate that, given a high flame temperature, a substantial increase in conductivity can be obtained by doping the propellants with alkali compounds. A double-base propellant doped with potassium nitrate has been used in the experiments so far. The experiments are performed in a closed vessel where a propellant slab is pinched between two copper plates connected to a pulsed power supply. Results presented include current and voltage recordings, yielding the added electric energy, and pressure measurements from which the burn rate can be determined.

Radiative Heat Loss Measurements During Microgravity Droplet Combustion in a Slow Convective Flow

Abstract: Radiative heat loss from burning droplets in a slow convective flow under microgravity conditions is measured using a broad-band (0.6 to 40 microns) radiometer. In addition, backlit images of the droplet as well as color images of the flame were obtained using CCD cameras to estimate the burning rates and the flame dimensions, respectively. Tests were carried out in air at atmospheric pressure using n-heptane and methanol fuels with imposed forced flow velocities varied from 0 to 10 centimeters per second and initial droplet diameters varied from 1 to 3 millimeters. Slow convective flows were generated using three different experimental configurations in three different facilities in preparation for the proposed International Space Station droplet experiments. In the 2.2 Second Drop-Tower Facility a droplet supported on the leading edge of a quartz fiber is placed within a flow tunnel supplied by compressed air. In the Zero-Gravity Facility (five-second drop tower) a tethered droplet is translated in a quiescent ambient atmosphere to establish a uniform flow field around the droplet. In the KC 135 aircraft an electric fan was used to draw a uniform flow past a tethered droplet. Experimental results show that the burn rate increases and the overall flame size decreases with increases in forced-flow velocities over the range of flow velocities and droplet sizes tested. The total radiative heat loss rate, Q(sub r), decreases as the imposed flow velocity increases with the spherically symmetric combustion having the highest values. These observations are in contrast to the trends observed for gas-jet flames in microgravity, but consistent with the observations during flame spread over solid fuels where the burning rate is coupled to the forced flow as here.

Destroying airborne biological and/or chemical agents with solid propellants

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

Pyrotechnic Burn Rate Measurements: Interstitial Flame Spread Rate Testing

Abstract: There are two general types of burning, sometimes described as parallel burning and propagative burning. In the previous article in this series, a method for determining parallel burn rate was discussed (strand testing), and a simplified method of testing under ambient conditions was suggested. In the current article, a method for determining one type of propagative burn rate (interstitial flame spread rate) is discussed.

A nitroguanidine based gas generant containing mica

Abstract: Gas generants in an inflator having nitroguanidine as a fuel and mica as a slagging agent. A gas generant with nitroguanidine has many desirable properties such as little hygroscopicity, a high conversion efficiency, and a suitable burn rate. Mica is beneficial ingredient to the gas generant because it reduces the amount of undesirable gases as well as reduces the amount of solid combustion particles from escaping the inflator.

Solid Fuel Rockets

Abstract: A solid fuel rocket is distinguished from a liquid fuel rocket by the type of fuel that it uses. It is more accurate to refer to the two basic types of rockets as solid propellant and liquid propellant rockets. Both types of rockets generate thrust by exhausting combustion products through a supersonic nozzle. Rocket propellants contain both a fuel and an oxidizer to produce energy. Both carry their own oxygen, so they can operate above the atmosphere and can be used in space. A solid fuel rocket, commonly called a solid rocket motor, is a completely self-contained device that converts chemical energy into kinetic energy in a controlled way. Although there is much to know about the science, engineering, and manufacture of solid rockets, they are simple devices. Solid rockets consist of four main components: (1) propellant grain, (2) a case that is the thrust chamber that contains the pressurized combustion gases, (3) a nozzle for directing and accelerating the gases away from the motor, and (4) an igniter. The propellant grain consists of the propellant charge shaped to deliver the desired thrust profile. The case contains the pressure of propellant combustion and is frequently a major portion of the vehicle airframe. Insulation is necessary to protect the case from the high-temperature combustion products. The igniter provides the heat necessary to initiate combustion at the propellant surface. The nozzle directs and accelerates the propellant exhaust gases. The propellant is typically a solid rubberlike material, similar to a pencil eraser, that contains fuel and oxidizer particles. The propellant grain can have a wide variety of shapes, but it is generally a hollow cylinder designed to burn from the inside out, thereby not exposing the case to the extreme temperatures until near the end of the burn. The propellant is bonded to the case. The rate at which the propellant burns and thereby generates thrust is designed into each propellant formulation. For instance, propellants that have fine oxidizer particles burn at higher rates than formulations that contain coarse oxidizer particles. Burn rate also varies with combustion chamber pressure, initial temperature of the propellant grain, and other factors. The internal shape of the grain, which can vary the amount of exposed surface area, hence the burn rate, is typically established when it is cast (poured) and cured inside the case. An igniter, generally located in the head end and fires down the bore or grain center perforation, initiates the motor. Hot combustion gases are ducted from the motor through a supersonic nozzle, which accelerates the gas and converts the pressure and temperature of propellant combustion to kinetic energy. Directional control is attained through a number of different means, but commonly the nozzle is jointed, allowing it to be vectored by mechanical actuators. Keywords: history; propellants; rockets; solid fuel; grain design; motor case; nozzle; propellant material

An experimental study on the ignition and combustion characteristics of pasty propellant

Abstract: The ignition and combustion characteristics of pasty propellant are important parameters for the motor design. An experimental study on the pasty propellant named PEPA/AP was carried out. The results show that the autoignition temperature of this propellant is about 150℃, so it is easy to ignite. As the chamber pressure is greater than 0.6MPa, the critical combustion diameter is less than 1mm. As the test motor is in operation, its propellant burning rate accords with the static burning rate. It is predicted that the propellant will be suitable for multiplestart motor.

Deflagration Behavior of HMX-Based Explosives at High Temperatures and Pressures

Abstract: We report the deflagration behavior of several HMX-based explosives at pressure from 10-600 MPa and temperatures from 20-180 C. We have made laminar burn rate measurements with the LLNL High Pressure Strand Burner, in which burn wires are used to record the time-of-arrival of the burn front in the cylindrical sample as a function of pressure. The explosive samples are 6.4 mm in diameter and 63 mm long, with ten burn wires embedded at different positions in the sample. Burning on the cylindrical surface is inhibited with an epoxy layer. With this direct measurement we do not have to account for product gas equation of state or heat losses in the system, and the burn wires allow detection of irregular burning. We find that formulation details are very important to overall deflagration behavior - the presence of 10% or less by weight of binder leads to physical deconsolidation and rapid deflagration at high pressures, and a larger particle size distribution leads to slower deflagration. High temperatures have a relatively minor effect on the deflagration rate until the beta-to-delta phase transition temperature is reached, beyond which the deflagration rate increases approximately 40-fold.

Convolution integral methods for the computation of linear burn rate of solid propellant subjected to arbitrary pressure excitation

Abstract: The computational methods for the time dependent burn rate of solid propellant with frequency and time domain convolution integral methods are presented. It is observed that time domain convolution integral method is faster compared to frequency domain convolution integral method. The accuracy of the time domain convolution integral method is checked by performing harmonic excitation analysis. The maximum oscillation amplitudes found from the response function and time domain convolution integral method are compared. It is found out that sampling period and the number of sampling points used in the inverse Fourier transform has a profound impact on the accuracy of the transient burn rate computations.

정적연소기에서의 메탄-공기 혼합기의 연소특성(2) : 비균질급기

Abstract: A cylindrical constant volume combustion chamber was used to investigate the flow characteristics at spark plug and the combustion characteristics of inhomogeneous charge methane-air mixture under several parameters. The flow characteristics such as mean velocity and turbulence intensity was analyzed by hot wire anemometer. Combustion pressure development measured by piezoelectric pressure transducer was used to investigate the effect of initial charge pressure, excess air ratio and ignition times on combustion pressure and combustion duration. Mean velocity and turbulence intensity had the maximum value at 200 or 300ms and then decreased to beneath 0.05m/s gradually at 3 seconds. Second mixture is accompanied by an increase in the combustion rate, and that the higher the mass which is added in the second stage injection, the faster the burn rate.