Burning Surface Of Propellant Research Articles

Effect of Organic Fluoride on Combustion Performance of HTPB Propellants with Different Aluminum Content

ABSTRACT Aluminum is an important contributor to energy of solid propellants. Completeness of aluminum combustion has key effect on the release of energy. In this paper, hydroxyl-terminated polybutadiene (HTPB) based propellants with aluminum content from 16% to 24% were prepared with a kind of organic fluoride (OF). Meanwhile, propellants without OF were also prepared as a comparison. The theoretical specific impulse and the heat of explosion of these propellants were calculated and measured, respectively. The particle size, morphology and elemental composition of the condensed combustion products for these propellants were characterized by laser measurement of particle size, scanning electron microscope (SEM), X-ray photoelectron spectroscopy (XPS) and X-ray diffraction (XRD). The burning rates of propellants were tested by CCD line-scanning camera and a high-pressure chamber measurement system. The high-speed photographic system was used to study the agglomeration process of the melting aluminum particles on the propellant burning surface. The results indicate that with increasing aluminum content, both the theoretical specific impulse and the measured heat of explosion show a trend of increasing until an aluminum content of 20%. The large particle size and residual active aluminum in the condensed combustion products increase with increasing aluminum content in the propellants, but decrease significantly after addition of OF. Under the same pressure, the burning rates of the blank and OF propellants decrease with increasing aluminum content, but the burning rates of OF propellants are slightly higher than those of the blank propellants. For the propellants with OF, the melting aluminum particle on the burning surface is less agglomerated and brighter, meaning more complete aluminum combustion.

Formation of solid residues in combustion of boron-containing solid propellants

Boron-containing solid propellants are promising propellants for ducted rockets. The experimental data show that combustion of boron-containing propellants is accompanied by an intense conglomeration of the boron particles near the burning surface. As a result, the boron particles conglomerates which are the branched coral structures composed of sintered boron particles are formed in combustion of boron-containing solid propellants. This results in decrease in combustion efficiency of ducted rockets. To calculate the ignition and combustion of a conglomerate in the secondary combustor of a ducted rocket, it is necessary to know the size, structure and shape of the boron particle conglomerates that can significantly affect these processes. In this paper, we consider a model and a method for calculating the boron particles conglomeration in combustion of the boron-containing solid propellants. The process of the boron particles conglomeration is considered as a result of competition between two main processes: the formation of adhesive bonds between the contacting boron particles and rupture of these bonds under the action of the aerodynamic detached force from the side the flow of gaseous combustion products of solid propellant. The criterion of formation of residues (slag) on the burning surface of boron-containing propellants is considered. It is shown that depending on the propellant burning rate law, two different types of propellants with different conglomeration behavior can exist. The results of calculations of the boron particles conglomeration which demonstrate the slag formation are shown.

Microwave plasma enhancement of multiphase flames: On-demand control of solid propellant burning rate

This effort explores the microwave-supported plasma enhancement of an aluminized, ammonium perchlorate composite solid propellant flame. The technique is enabled through novel alkali metal doping, which provides increased levels of ionization and allows efficient microwave energy deposition to the flame structure and subsequent perturbation of the steady-state propellant burning rate. Three potential modes of energy deposition are identified in composite propellants: (1) plasma-enhancement promoted by the presence of sodium, (2) strong absorption by high temperature metallic oxide products, and (3) direct absorption of energy by the propellant burning surface condensed phase. Equilibrium calculations are used to identify propellant compositions with high free-electron populations, and solution of the Boltzmann equation using the BOLSIG + code is used to show the significant effects of sodium doping on microwave absorption of the flame. Experimental emission spectroscopy and imaging of the microwave-enhanced propellant flame structure show evidence of these mechanisms. Propellant burning rates can be increased by up to 60% through the application of 1 kW, continuous 2.46 GHz microwave radiation. Propellant doping with 3.5% by weight of sodium nitrate shows significant improvement in the effectiveness of microwave application in modifying the propellant burning rate. This technique provides high levels of dynamic propellant combustion control from a solid-state system. The on-demand control of energetic material burning rates through low level doping and microwave irradiation is a promising technique which could lead to the development of a new class of ‘smart’ dynamically controllable energetic materials.

Ignition and combustion of a single aluminum particle in hot gas flow

For simulating the aggregated aluminum bulks on the burning surface of solid propellants, large aluminum particles (40–160 µm) are used in this work. The isolated aluminum particles are ignited in hot oxidizing gas. Based on the bright-spot diameter profiles and the known respective reaction mechanisms, the total ignition and combustion process of aluminum particle can be divided into three stages, namely, pre-heating, ignition and combustion. The initial and bright-spot diameters of the aluminum particle are measured directly from the images by using the in-house automated data processing routines. Ignition delay time, ti, and combustion time, tc, are also obtained by post-processing the sequential images and can be associated with the particle diameters, D, in the form of ti = aD + b and tc = αD, respectively. The changing trends of ignition delay time and combustion time with the effective oxidizer mole fraction in the range of 22.8%–49.1% are distinctly different. The oxidizing environments with a high effective oxidizer mole fraction can result in short combustion time but long ignition delay time. For small particles (40–110 µm), the environmental effective oxidizer mole fraction exerts a limited effect on the sum of ignition delay time and combustion time, which indicates total time. By considering the effects of particle sizes and effective oxidizer mole fractions of environments, the percentages of ignition delay time in the total time are analyzed. These results suggest that with the goal of decreasing the total time, suitable methods can be employed for different conditions. Furthermore, we observe and discuss the phenomenon of aluminum particle microexplosion in an environment with high effective oxidizer mole fraction, which decreases particle combustion time by a large margin.

Aluminum agglomeration involving the second mergence of agglomerates on the solid propellants burning surface: Experiments and modeling

The agglomeration of aluminum particles usually occurs on the burning surface of aluminized composite propellants. It leads to low propellant combustion efficiency and high two-phase flow losses. To reach a thorough understanding of aluminum agglomeration behaviors, agglomeration processes, and particles size distribution of Al/AP/RDX/GAP propellants were studied by using a cinephotomicrography experimental technique, under 5MPa. Accumulation, aggregation, and agglomeration phenomena of aluminum particles have been inspected, as well as the flame asymmetry of burning agglomerates. Results reveals that the dependency of the mean and the maximum agglomeration diameter to the burning rate and the virgin aluminum size have the same trend. A second-time mergence of multiple agglomerates on the burning surface is unveiled. Two typical modes of second mergence are concluded, based upon vertical and level movement of agglomerates, respectively. The latter mode is found to be dominant and sometimes a combination of the two modes may occur. A new model of aluminum agglomeration on the burning surface of composite propellants is derived to predict the particulates size distribution with a low computational amount. The basic idea is inspired from the well-known pocket models. The pocket size of the region formed by adjacent AP particles is obtained through scanning electron microscopy of the propellant cross-section coupled to an image processing method. The second mergence mechanism, as well as the effect of the burning rate on the agglomeration processes, are included in the present model. The mergence of two agglomerates is prescribed to occur only if their separation distance is less than a critical value. The agglomerates size distribution resulting from this original model match reasonably with the experimental data. Moreover, the present model gives superior results for mean agglomeration diameter compared to common empirical and pocket models. The average prediction error is lower than 5% for the four propellants tested. Results of this study are expected to provide better insight and enrich in the theoretical frame of aluminum agglomeration.

Microexplosions and ignition dynamics in engineered aluminum/polymer fuel particles

Aluminum particles are widely used as a metal fuel in solid propellants. However, poor combustion efficiencies and two-phase flow losses result due in part to particle agglomeration. Recently, engineered composite particles of aluminum (Al) with inclusions of polytetrafluoroethylene (PTFE) or low-density polyethylene (LDPE) have been shown to improve ignition and yield smaller agglomerates in solid propellants. Reductions in agglomeration were attributed to internal pressurization and fragmentation (microexplosions) of the composite particles at the propellant surface. Here, we explore the mechanisms responsible for microexplosions in order to better understand the combustion characteristics of composite fuel particles. Single composite particles of Al/PTFE and Al/LDPE with diameters between 100 and 1200µm are ignited on a substrate to mimic a burning propellant surface in a controlled environment using a CO2 laser in the irradiance range of 78–7700W/cm2. The effects of particle size, milling time, and inclusion content on the resulting ignition delay, product particle size distributions, and microexplosion tendencies are reported. For example particles with higher PTFE content (30wt%) had laser flux ignition thresholds as low as 77W/cm2, exhibiting more burning particle dispersion due to microexplosions compared to the other materials considered. Composite Al/LDPE particles exhibit relatively high ignition thresholds compared to Al/PTFE particles, and microexplosions were observed only with laser fluxes above 5500W/cm2 due to low LDPE reactivity with Al resulting in negligible particle self-heating. However, results show that microexplosions can occur for Al containing both low and high reactivity inclusions (LDPE and PTFE, respectively) and that polymer inclusions can be used to tailor the ignition threshold. This class of modified metal particles shows significant promise for application in many different energetic materials that use metal fuels.

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.

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.

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.

Numerical Study of the Dynamic Characteristics of Pintle Nozzles for Variable Thrust

Unsteady numerical simulations of pintle nozzles were implemented to investigate dynamic characteristics of various pintle configurations. To take into account the effects of the pintle shape and movement, a sliding mesh method was applied. The physical nozzle throat based on pintle location was analytically investigated and found to compare well with numerical results. The static and dynamic results are verified with the experimental results. The flow separation shock trains as the pintle strokes are analyzed according to the three pintle models. The response lag and sensitivity of the chamber pressure and nozzle performance were evaluated for pintle reciprocation, insertion, and extraction processes to better understand the dynamic performance of the pintle nozzle. The pressure coupling effects of the propellant burning surface during the pintle reciprocation are conducted, which are compared with the cold-flow cases.

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.

Numerical analysis of AP/HTPB composite propellant combustion under rapid depressurization

A two-dimensional steady/unsteady combustion model has been developed for ammonium perchlorate/hydroxyl-terminated polybutadiene (AP/HTPB) composite propellant. An equivalent source term method is applied to describe the decomposition at the burning surface, and the combustion field in the gas phase is solved coupling with the thermal field in the solid phase. Firstly, a series of numerical simulations were carried out for steady combustion for pressure ranging from 0.2 MPa to 5.0 MPa. It is shown that the flame develops from a premixed structure at low pressure to a diffused one at high pressure. And the calculated burning rate agrees well with the experimental result. Then the unsteady combustion under rapid depressurization was predicted for the initial pressure of 5.0 MPa. The simulation results indicate that during the depressurization, the dominant factor of combustion turns from the diffusion flame to the AP decomposition flame. As the pressure decays, the heat release rate of the overall flame reduces and the flow rate of the combustion gas increases, with the out-of-phase blowing effect enhanced. And thus the heat feedback from the flame to the propellant burning surface is reduced. The higher the depressurization rate is, the faster the heat feedback and the surface temperature decrease.

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.

Coupled simulation of fluid flow and propellant burning surface regression in a solid rocket motor

Propellant burning surface regression is one of the most distinguishing characteristics of solid rocket motors. Simulation of this phenomenon, especially when coupled with a fluid flow, requires accurate mathematical models and robust and efficient numerical techniques to determine the evolution of the propellant burning surface as it regresses with speed as determined by solid propellant combustion. In this paper, an integrated framework is presented for the coupled simulation of propellant burning surface regression and internal fluid flow in a solid rocket motor. The arbitrary-Lagrangian–Eulerian scheme is employed to formulate the compressible viscous fluid flow on moving meshes. Face-offsetting method is used to model the propellant burning surface regression of solid rocket motor. Automatic mesh smoothing and remeshing techniques are further utilized to address the deformation and distortion of fluid domain meshes. We present the theoretical foundation of our method and finally demonstrate its accuracy, efficiency, and flexibility for a laboratory-scale solid rocket motor.

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.

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.

Studies on Gravity Influence on Solid Propellant Burn Rate

Experimental studies have been carried out using the in-house developed propellant samples at the atmospheric conditions to examine the influence of propellant surface orientation / attitude on burn rate. A series of burning tests are conducted with different grain orientations, viz., vertical, inverted and horizontal. We have observed 5 % burn rate augmentation on end-burning grains when the burning surface evolution was against the earth gravity compared to the normal vertical candle burning condition. We conjectured that the coupled effects of the instantaneous variations of the propellant burning surface attitude and the flight acceleration during the mission could alter the flame structure due to the local gravitational influence, which in turn alter the burn rate. This paper throws light for developing a suitable gravitational force dependant burn rate model for improving the performance prediction of solid rocket motors for aerospace applications.

Influence of the composition and ionizing radiation on the stability of combustion of composite propellants

The results of an experimental study of the acoustic admittance of the burning surface of composite propellants performed with the use of a two-end combustion chamber (T-chamber) are presented. The effects of the composition of the composite propellant (type of fuel-binder, content of aluminum powder, burning rate catalysts) and of ionizing γ-radiation on the acoustic admittance, which characterizes the tendency of the combustion chamber to high-frequency instability, are analyzed.

Vortex Flow Analysis of a Large Segmented Solid Rocket Motor

It is not uncommon to find pressure oscillations in large size segmented solid rocket motors. During the static test of segmented solid rocket motor (SRM-3), unanticipated pressure oscillations were seen after some of the propellant has burnt and the oscillations sustained for certain duration of time. The purpose of this paper is to present the results of the analysis carried out on the experimental data using commercial CFD software 'FLUENT' and compare the results with experimental data to find out the cause of the pressure oscillations. Inhibitors are provided for the full web thickness of propellant at the segment joint interfaces to prevent end-burning of propellant at the joints. The char and erosion rate of these inhibitions are lower than the burning rate of the propellant. As burning of the propellant progresses, annular inhibition wall starts projecting above the burning surface of the propellant. These obstacles (inhibition) are acting as wall and sheared gas flow occurs in the rocket motor. CFD analysis was carried out on a quadrilateral mesh for the geometry of SRM-3 at the time of 32 s after ignition when the pressure oscillations peaked. It has been found that the vortex shedding frequency obtained by CFD analysis closely matches with the frequency of pressure oscillations occurring during static testing. Thus, it has been revealed that the obstacle vortex shedding generated by the compartmentalization of the combustion chamber by the protrusion of inhibition at the segment joints was the cause of the pressure oscillations in SRM-3.

Effect of Nitrate Ester on the Combustion Characteristics of PET/HMX -based Propellants

The effect of nitrate ester NG/TEGDN on the combustion characteristics of PET/HMX-based propellants has been experimentally investigated using of high-speed photography technique and scanning electron microscopy. It is indicated that the increase of NG/TEGDN content has little impact on the propellant burning rates at the same pressure. Furthermore, propellant can not be self-sustaining combustion at low pressure (£1 MPa). The increase of NG/TEGDN content does not affect the flame structure of propellant, but it plays an important role in condensed phase reaction zone. The flame structure of propellant is estimated. The thermal decomposition products in different combustion zones are also discussed. Scanning electron microscopy examination of quenched sample indicates that a liquified layer forms during combustion of these propellants. Numerous gas bubbles are present. Especially, the burning surface of propellant with low NG/TEGDN content shows signs of crystallization. The thickness of condensed phase reaction zone, by cross-section examination of propellant burning surface, has also been investigated. The results show that the thickness of condensed phase reaction zone increases with NG/TEGDN content increasing. These observations suggest that the condensed phase zone plays significant role in propellant combustion. Defence Science Journal, 2011, 61(3), pp.206-213 , DOI:http://dx.doi.org/10.14429/dsj.61.567