Propellant Surface Research Articles

A mechanism for shattering microexplosions and dispersive boiling phenomena in aluminum–lithium alloy based solid propellant

The microexplosive nature of multicomponent liquid fuels has been both studied and fielded to decrease droplet residence times and increase completeness of combustion. However, little work has focused on investigating microexplosive metal fuels to enhance the metal fuel combustion efficiency in traditional energetic material formulations. Microscopic surface videography was performed on two solid propellant formulations, one using aluminum (baseline) and the other with 80/20wt% Al–Li alloy as fuel additives. It was observed that the propellant combustion with neat aluminum formed large molten droplets at the surface as aluminum particles agglomerate, which is a well-known problem with aluminized propellants. In contrast, the Al–Li propellant formed an Al–Li melt-layer on the propellant surface during combustion. Droplets were ejected from the surface melt-layer through dispersive boiling. Above the surface, further dispersive boiling is observed from the ejected droplets and droplet-shattering microexplosions are also observed. These dynamics are thought to be a result of a large disparity in volatility (i.e., boiling points) between the metals in the molten alloy and the large Lewis number in the droplet, so that superheating occurs before the more volatile component (here Li) can diffuse to the surface. A Lewis number of 7440 was estimated for molten 80/20wt% Al–Li alloy, which is nearly three orders of magnitude larger than typical multicomponent liquid hydrocarbon droplets that microexplode, suggesting a higher propensity for molten droplet microexplosions. This would also indicate that a smaller amount of the volatile component might be necessary for microexplosions and dispersive boiling than observed for liquid hydrocarbon fuels. These dynamics are important for metal fuel applications, because injectors cannot be used to decrease droplet size in a metallized energetic material formulation.

Temperature measurements in metalized propellant combustion using hybrid fs/ps coherent anti-Stokes Raman scattering.

We apply ultrafast pure-rotational coherent anti-Stokes Raman scattering (CARS) for temperature and relative oxygen concentration measurements in the plume emanating from a burning, aluminized ammonium-perchlorate propellant strand. Combustion of these metal-based propellants is a particularly hostile environment for laser-based diagnostics, with intense background luminosity and scattering from hot metal particles as large as several hundred micrometers in diameter. CARS spectra that were previously obtained using nanosecond pulsed lasers in an aluminum-particle-seeded flame are examined and are determined to be severely impacted by nonresonant background, presumably as a result of the plasma formed by particulate-enhanced laser-induced breakdown. Introduction of femtosecond/picosecond (fs/ps) laser pulses improves CARS detection by providing time-gated elimination of strong nonresonant background interference. Single-laser-shot fs/ps CARS spectra were acquired from the burning propellant plume, with picosecond probe-pulse delays of 0 and 16 ps from the femtosecond pump and Stokes pulses. At zero delay, nonresonant background overwhelms the Raman-resonant spectroscopic features. Time-delayed probing results in the acquisition of background-free spectra that were successfully fit for temperature and relative oxygen content. Temperature probability densities and temperature/oxygen correlations were constructed from ensembles of several thousand single-laser-shot measurements with the CARS measurement volume positioned within 3 mm or less of the burning propellant surface. The results show that ultrafast CARS is a potentially enabling technology for probing harsh, particle-laden flame environments.

Study on Combustion Characteristics and Propelling Projectile Motion Process of Bulk-Loaded Liquid Propellant

ABSTRACTData are presented showing that the problem of gas–liquid interaction instability is an important subject in the combustion and the propellant projectile motion process of a bulk-loaded liquid propellant gun (BLPG). The instabilities themselves arise from the sources, including fluid motion, to form a combustion gas cavity called Taylor cavity, fluid turbulence and breakup caused by liquid motion relative to the combustion chamber walls, and liquid surface breakup arising from a velocity mismatch on the gas–liquid interface. Typically, small disturbances that arise early in the BLPG combustion interior ballistic cycle can become amplified in the absence of burn rate limiting characteristics. Herein, significant attention has been given to developing and emphasizing the need for better combustion repeatability in the BLPG. Based on this goal, the concept of using different geometries of the combustion chamber is introduced and the concept of using a stepped-wall structure on the combustion chamber itself as a useful means of exerting boundary control on the combustion evolution to thus restrain the combustion instability has been verified experimentally in this work. Moreover, based on this background, the numerical simulation is devoted to a special combustion issue under transient high-pressure and high-temperature conditions, namely, studying the combustion mechanism in a stepped-wall combustion chamber with full monopropellant on one end that is stationary and the other end can move at high speed. The numerical results also show that the burning surface of the liquid propellant can be defined geometrically and combustion is well behaved as ignition and combustion progressivity are in a suitable range during each stage in this combustion chamber with a stepped-wall structure.

Development of a Direct Drive Vacuum Arc Thruster Passively Ignited for Nanosatellite

Active attitude control systems require a non-negligible amount of space to be mounted on board satellites. For the nanosatellite case, it becomes even more difficult to accommodate an active attitude control system due to size and mass constraints, and choices of active attitude control system are therefore limited. In this paper, small size and lightweight are realized by direct drive and passive ignition of a thruster. The direct drive is a power supply system without a booster circuit. The passive ignition system is an ignition system using microdischarge generated by interaction with a plasma environment and the propellant surface. Availability of these two systems was verified. As a result, when the thruster was placed into a high-density plasma environment ( $10^{12}~\text{m}^{-3})$ , it was confirmed that the direct drive and passive ignition can operate normally. Moreover, expected performance of this vacuum arc thruster was measured. The results show that the impulse bit was of micronewtons order, and the ion velocity was measured to be approximately 12 km/s. Moreover, it was found that the impulse bit changes with the applied voltage and the capacitance value of the discharge circuit.

Combustion of Aluminum Particles near the Burning Surface in AP/AN Composite Propellants

AbstractAluminum (Al) particles are commonly used in ammonium perchlorate (AP) composite propellants of solid rockets for increasing performance. When propellants including Al particles burn, Al particles easily agglomerate on the burning surface of the propellant. The diameters of agglomerated Al particles are greater than those of mixed particles. The combustion efficiency of the propellant decreases with increasing burning time of the agglomerated Al particles. Therefore, it is important to observe how the agglomerated Al particles burn on the burning surface of AP composite propellant. A lot of researchers have studied Al agglomerate characteristics. Previous studies clarified the relation between the agglomerated Al particle diameter and luminous flame diameter around Al particles near the burning surface. The shapes of luminous flames around agglomerated Al particles are spherical or elliptical. This study evaluates the shapes of the luminous flame around agglomerated Al particles at a constant diameter or a different diameter. When the proportion of the luminous flame diameter (Df) to the diameter of agglomerated Al particles (D0) is 1.54–1.71 at a constant D0, the luminous flames are almost perfectly spherical. Otherwise, the luminous flames are elliptical at a constant D0. Furthermore, when Df/D0 is close to the mean value, the luminous flame is more spherical than elliptical at different D0. The evaporation rate and the burning rate of Al vapor are inversely proportional to D0. The oxidation gas temperatures were changed and the activation energy of Al vapor was obtained as 39.2 kJ mol−1.

Ignition of a composite propellant by a hot particle under conditions of a nonideal thermal contact

A mathematical modeling of the solid-phase ignition of a condensed material by a single diskshaped hot metal particle under conditions of a nonideal thermal contact between the particle and the propellant due to the natural surface roughness of the latter is performed. The ranges of values of the initial temperature of the local energy source and of the parameter characterizing the propellant surface roughness at which radiative heat transfer across the gas gap between the propellant and particle significantly (up to 25%) influences the ignition delay time, the main integrated characteristic of the process, are determined from numerical simulations.

Simple Surface-Tracking Methods for Grain Burnback Analysis

Practical surface-tracking methods are successfully developed to simulate the evolution of the grain geometry, which is dynamically regressing during the combustion of the solid propellant. Three methods, namely the front-tracking, emanating-ray, and least-distance methods, are proposed. The front-tracking method is based on the Lagrangian approach, whereas both the emanating-ray and the least-distance methods are formulated from the Eulerian viewpoint. Three two-dimensional test cases have been examined to compare the programming complexity, simulation accuracy, and computational efficiency of the proposed methods. It is found that the least-distance method performs superior to the other two methods in many numerical respects. The least-distance method is implemented on the tetrahedron grid system to track the outwardly regressing surface of a three-dimensional cubical propellant. Comparison between the predicted erosive volume and burning surface area with the theoretical results yields satisfactory agreement. A grain design similar to one of the Falcon Launch V cases has been employed to demonstrate the usefulness of the proposed methodology in application to a practical three-dimensional grain burnback analysis.

Application of Nano‐Aluminum/Nitrocellulose Mesoparticles in Composite Solid Rocket Propellants

AbstractIn this work we investigate the potential application of nano‐aluminum/nitrocellulose mesoparticles as an ingredient for solid composite rocket propellants. The basic strategy is to incorporate nanoaluminum in the form of a micrometer scale particle containing a gas‐generator, to enable easier processing, and potential benefits resulting from reduced sintering prior to combustion. The mesoparticles were made by electrospray and comprised aluminum nanoparticles (50 nm) and nitrocellulose to form micrometer scale particles. In this study, 80 % solids loaded composite propellants (AP/HTPB based) were made with the addition of micrometer sized (2–3 μm) aluminum (10 wt‐%), and compared directly to propellants made by directly substituting aluminum mesoparticles for traditional micrometer sized particles. Propellant burning rate was relatively insensitive for mesoparticles containing between 5–15 wt‐% nitrocellulose. However, direct comparison between a mesoparticle based propellant, to a propellant containing micrometer scale aluminum particles showed burning rates approximately 35 % higher while having a nearly identical burning rate exponent. High speed imaging indicate that propellants using mesoparticles have less agglomeration of particles on the propellant surface.

Response of TEGDN Propellants to Plasma Ignition with the Same Magnitude of Ignition Energy as Conventional Igniters in an Interrupted Burning Simulator

AbstractIn order to consider the potential influence of ignition energy factors on the response of double base propellants plasticized with triethyleneglycol dinitrate (TEGDN propellants), the influence of different ignition methods at the same magnitude of ignition energy level on the response of TEGDN propellants was investigated in an interrupted burning simulator. Compared to conventional ignition methods, plasma ignition exhibits a significantly shorter ignition delay and lower ignition energy. There are stronger ablation and impact interactions of plasma flow with the surface of propellants. For TEGDN propellants coated with titanium dioxide, a greater amount of melted white layer is deposited on the surface of propellants. The content of copper on the surface of recovered plasma‐ignited samples observed by X‐ray Fluorescence (XRF) spectroscopy is much larger than that on the surface of recovered conventionally ignited samples, indicating more deposition of copper wire discharge on the surface of the samples. The test results will benefit the design of plasma generator and electrical parameters of pulse power to satisfy certain propellant compositions.

Numerical Evaluation of the Use of Aluminum Particles for Enhancing Solid Rocket Motor Combustion Stability

The ability to predict the expected internal behaviour of a given solid-propellant rocket motor under transient conditions is important. Research towards predicting and quantifying undesirable transient axial combustion instability symptoms typically necessitates a comprehensive numerical model for internal ballistic simulation under dynamic flow and combustion conditions. On the mitigation side, one in practice sees the use of inert or reactive particles for the suppression of pressure wave development in the motor chamber flow. With the focus of the present study placed on reactive particles, a numerical internal ballistic model incorporating relevant elements, such as a transient, frequency-dependent combustion response to axial pressure wave activity above the burning propellant surface, is applied to the investigation of using aluminum particles within the central internal flow (particles whose surfaces nominally regress with time, as a function of current particle size, as they move downstream) as a means of suppressing instability-related symptoms in a cylindrical-grain motor. The results of this investigation reveal that the loading percentage and starting size of the aluminum particles have a significant influence on reducing the resulting transient pressure wave magnitude.

Iron Nanoparticle Additives as Burning Rate Enhancers in AP/HTPB Composite Propellants

AbstractBurning rate measurements were carried out for ammonium perchlorate/hydroxyl‐terminated polybutadiene (AP/HTPB) composite propellants with iron (Fe) nanoparticles as additives. Experiments were performed in a strand burner at pressures from 0.2 to 10 MPa for propellants containing approximately 80 % AP and Fe nanoparticles (60–80 nm) at concentration from 0 to 3 % by weight. It was found that the addition of 1 % Fe nanoparticles increased burning rate by factors of 1.2–1.6. Because Fe nanoparticles are oxidized on the surface and have high surface‐to‐volume ratio, they provide a large surface area of Fe2O3 for AP thermal decomposition catalysis at the burning propellant surface, while also providing added energy release due to the oxidation of nanoparticle sub‐shell Fe. The increase in burning rate due to Fe nanoparticle content is similar to the increase in burning rate caused by the addition of iron oxide (Fe2O3) particles observed in prior literature.

Gas dynamics and heat transfer inside a solid propellant crack during ignition transient

To study the gas dynamic and heat transfer phenomena inside a single isolated longitudinal solid propellant surface crack, two 3-D geometric models with different crack shapes were constructed. Concerning the influence of propagation of jet from the igniter on the flame spreading phenomena in the crack, flow region around the opening of the crack was also included in the above geometric models. A theoretical framework was then adopted to model the conjugate heat transfer in the combustion channel and the crack cavity. Numerical simulation results indicate that the ignition shock wave can spread into the crack cavity. Extremely high overpressure and pressurization rate were observed along the crack front. It is possible that the crack may propagate before the flame front reaches it. An ignited region located at the crack front near to the channel surface in downstream direction was generated long before the flame front reached the crack opening in both models.

Laboratory method for measurement of the specific impulse of solid propellants

A new laboratory express-method of determining the specific impulse of solid propellants based on the measurement of the reactive force of gasification products escaping from the burning propellant surface is presented in this work. The values of the specific impulse for a model composite solid propellant by varying the pressure in the combustion chamber are determined.

Influences of particle size and content of RDX on burning characteristics of RDX-based propellant

Research on the burning characteristics of propellants prepared with cyclotrimethylenetrinitramine (RDX) has not yet provided sufficient systematic experimental data. In this study, the thermal decomposition behaviors and the burning characteristics of RDX/hydroxyl-terminated polybutadiene (HTPB) propellants prepared with five series of RDX with weight mean diameters of 41 μm, 80 μm, 145 μm, 300 μm, and 515 μm and at various RDX contents of 50–80% were investigated.The thermal decomposition behavior of the RDX propellants was not affected by the particle size of RDX. RDX and HTPB decomposed almost separately in propellant matrices. The flame structure and the burning surface of propellants became more heterogeneous with increasing particle diameters of RDX. The burning rates of RDX propellants increased with increasing RDX content. The increasing ratio of the burning rate with increasing RDX content was not dependent on the mean particle diameter of RDX, but incremented with higher combustion pressure. The pressure exponents increased at higher RDX content and did not have particle size dependence below 145 μm of mean particle diameter of RDX. The burning rates of propellants increased with decreasing mean diameter of RDX particles. The relations between the burning rate and the mean particle diameter of RDX were expressed by a linear regression line on a double logarithmic plot. The RDX particle size dependence of the burning rates became smaller with higher RDX content. The interparticle distances of RDX particles in the propellant matrix played a key role in determining the particle size dependence on the burning rate.

Numerical modeling for the combustion of simulated solid rocket motor propellant

The numerical procedure for the burning of Ammonium Perchlorate (AP) with a Fuel-Binder (Hydroxyl Terminated Polybutadience HTPB) heterogeneous propellant is presented. This model accounts for the two-step reaction mechanism for the primary diffusion flame at the interface between Binder (B) and the oxidizer AP and the AP premixed flame. The complete coupling between the gas-phase, the condensed-phase, and the unsteady non-uniform regression of the propellant surface is considered. The parameters used in this model are chosen to fit experimental data for the burning of the composite AP/Binder. The propagation of the unsteady non-planer regression surface is described, using the Essentially-Non-Oscillatory (ENO) scheme with the aid of the level set strategy. The Alternating-Direction-Implicit (ADI) solver is employed to solve the full Navier–Stokes equations in the gas phase for the variable density model. The results show the effect of various parameters on the surface propagation speed, flame structure, and the burning surface geometry. A comparison between the computational and experimental results is presented.

Modeling of ammonium dinitramide (ADN) monopropellant combustion with coupled condensed and gas phase kinetics

A comprehensive multi-phase combustion model has been developed to study the physiochemical processes involved in the combustion of ammonium dinitramide (ADN). The numerical model is based on the conservation equations of mass, species concentration, and energy, and takes into account finite-rate chemical kinetics in both condensed and gas phases. Based on an extensive review of the literature on ADN thermal decomposition, three global decomposition reactions in the condensed phase of ADN are included. A detailed chemical kinetics scheme involving 34 species and 165 reactions is employed in the gas phase. Detailed combustion-wave structures and burning rate characteristics of ADN are described. The optimized gas-phase kinetics mechanism was able to predict the multi-stage flame structure. Good agreements between the predicted and measured profiles of temperature and species mole fractions were obtained at different pressures. Reasonable agreements between calculated and measured values of propellant burning rates and surface temperatures were obtained over a broad range of pressure from 0.7 to 350atm. The burning rate increases with pressure, except in the mid range of ∼60–100atm. The coupled condensed- and gas-phase analysis employed in the current model is able to capture this irregular/unstable combustion behavior in the mid range, where the burning rate decreases with the increase in pressure.

Mechanism study of stabilization of double-base propellants by using zeolite stabilizers (nano- and micro-clinoptilolite)

The mechanism of stabilization of double-base rocket propellants (DBPs) using inorganic stabilizer (nano- and micro-clinoptilolite) was investigated. The surface structures of the stabilizers, the double-base propellants containing the new stabilizers and the effect of the stabilizers on the surface behavior of propellants and vice versa were checked using XRD and Electron Microscope (AFM and TEM techniques). The results obtained from XRD suggested that the crystalline structure of the new inorganic stabilizers was completely changed when it was introduced into the propellants which may be attributed to the pressing processes of DBPs with stabilizers under very high pressure during their mixture preparation. The results obtained from Atomic Force Electron Microscope (AFM) and TEM indicated that nano-clinoptilolite particles become more regularly arranged on the propellant surface than micro-clinoptilolite which gives the stabilizer a higher ability to absorb more nitrogen oxide. The work aimed chiefly to use zeolite stabilizers for DBPs instead of classically used organic compounds; in order to avoid the harmful and carcinogenic organic products coming from the reaction of NOx gases with these organic stabilizers. This is achieved by studying the thermal behavior of these zeolites via investigation of their surface interaction with NOx gases obtained during stabilization process and suggesting possible interaction mechanism.

Thermites for Solid Rocket Motor Ignition

Fast and repeatable ignition transients for small solid rocket combustors can be difficult to achieve. This work sets out to characterize a thermite mixture to fill this need. Igniter function was determined using high-speed imaging, allowing an examination of product droplet size and combustion time as a function of packaging technique. Safety testing (electrostatic discharge, drop weight impact, and friction) indicated that this material is far safer than existing ignition compounds. A simple dual-criteria ignition model is applied to igniter sizing, and our modeling results were successfully evaluated using hot fire tests of motors with various exposed propellant surface conditions. The end result is a safe, inexpensive, reliable, and readily available method of igniting small (500 g–50 kg propellant mass) solid rocket combustors.

Effect of Specific Surface Area of Aluminum on Composite Solid Propellant Burning

This paper proposes the use of micron-sized aluminum (pyral, average particle size ) with a higher specific surface area as a good candidate to enhance the burn rate of the composite propellant. Experiments were performed in the pressure range of 10 to 70 bar in a window bomb for measuring the burn rate. Comparison of these burn rate results with those obtained using micron- and nanosized aluminum found that the performance of pyral was in between that of micron- and nanosized aluminum. The reason for the high burn rates observed with pyral is due to the flake like appearance of pyral with a large specific surface area. It is argued that, if the specific surface area is large, then the thickness becomes the characteristic length scale. This ensures the heat release from the aluminum combustion to occur closer to the propellant surface as the thickness of pyral is in nanometers. Both the x-ray diffraction and heat of formation analyses indicated that pyral had higher purity than nanoaluminum, which has an implication to the specific impulse of the propellant. In addition to this, it shows that the micron-sized catalyst is effective with pyral, whereas it is ineffective with nanoaluminum. This study also reports the mechanical properties of the propellant containing pyral.

Simulation Research of Vaporization and Pressure Variation in a Cryogenic Propellant Tank at the Launch Site

In order to improve depiction of pressure variation and investigate the interrelation among the physical processes in propellant tanks, a 2D axial symmetry Volume-of-Fluid (VOF) CFD model is established to simulate a large-sized liquid propellant tank when the rocket is preparing for launch with propellant loaded at the launch site. The numerical model is considered with propellant free convection, heat transfer between the tank and the external environment, thermal exchange between propellant and inner tank wall surfaces, gas compressibility, and phase change modeled under the assumption of thermodynamic equilibrium. Vaporization rate of the vented LH2 tank and prediction of pressure change in the tank pressurized with GHe are obtained through simulation. We analysis the distributions of phase, temperature, and velocity vectors to reveal interactions among the propellant’s own convection motion, heat transfer and phase change. The results show that the vaporization rate is mainly affected by heat leaks though the tank wall when the tank is vented, but it does not completely accord with the trend of the leakage because of convection motion and temperature nonuniformity of the liquid propellant in the tank. We also find that the main factors on pressure variation in the pressurized tank are the heat transfer on the tank wall surface bonding the ullage and propellant vaporization which has comparatively less influence.