Combustion Of Composite Propellants Research Articles

Microflower-like Fe-Co-MOF with enhanced catalytic performance for decomposition of ammonium perchlorate and combustion of ammonium perchlorate-based composite propellants

Ammonium perchlorate (AP) serves as a crucial component in solid propellants. Enhancing its thermal decomposition with catalysts improves the combustion performance of solid propellants significantly. To enhance the catalytic performance of metal-organic frameworks (MOFs) on AP, this study controlled the morphology of MOFs by adding metal atoms, resulting in a Fe-Co-MOF catalyst with higher specific surface area and dual-metal synergistic effect. The catalytic effect of the catalyst on AP was investigated using DSC and TG-IR, followed by studying its influence on the combustion performance of AP-based composite propellants. The results show that the introduction of 5% Fe-Co-MOF reduced the decomposition temperature of AP to 298.6 °C, decreased the activation energy to 151.6 kJ mol-1, and increased the heat release by 110.6%. Additionally, the ignition delay of the propellant decreased by 71 ms, and the combustion rate increased by 43.8%. Mechanistic studies demonstrate that the abundant catalytic sites and oxygen vacancies of Fe-Co-MOF facilitate the charge transfer rate during the AP thermal decomposition process and promote the increase in AP heat release by suppressing the high-temperature conversion of N2O. This research paves the way for enhancing the application of MOF materials in AP-based solid propellants.

Understanding AP/HTPB composite propellant combustion from new perspectives

Combustion of the AP/HTPB composite propellant at the mesoscale is simulated by a partitioned numerical framework, in which the unsteady solid-phase conduction, gas-solid interface regression, and gaseous combustion are sequentially calculated on two sets of overlapped grids. The gas-solid interface is traced by the level-set method. The mass and species exchange conditions at the gas-solid interface are converted into source terms by the immersed-boundary method. Effectiveness of the entire framework is justified by applying the integrated solver to a Miller pack example. The predicted burning rates show good agreement with experimental data over a wide range of pressure conditions, and the gas-phase temperature field clearly depicts the final diffusion flame. The flame zone closely above the regressing interface can be identified by the fluctuation of the projected velocity components perpendicular to the main stream direction. Statistical results conditional on Lagrangian tracing of fluid parcels released from the AP and binder parts show distinction between the oxidizer-rich and fuel-rich flames in terms of the reaction intensity and the flame structure. Statement of SignificanceFrom the algorithm/simulation perspective: a partitioned numerical solution framework is built upon overlapped structured (for the gas phase) and unstructured (for the solid phase) meshes; the level-set approach is used on the unstructured mesh to identify solid-gas interface with complex geometry. From the analysis perspective: the flame front is newly defined in terms of the tangential velocity components; the fuel/oxidizer evolution pattern is analyzed from conditional statistics based on Lagrangian tracing; new insights in heterogeneous combustion include, for instance, effects of pressure on the tangential velocity fluctuation and flame properties conditional on different ingredients.

Modeling of aluminum agglomeration at solid propellant burning surface by combining aluminum ignition with LDP-SC cluster analysis

Aluminum particles at burning surface may agglomerate during the combustion process of composite solid propellant, involving the aggregation and ignition of aluminum particles as well as the secondary coalescence of agglomerates. So far it is still a huge challenge to capture the agglomeration process and obtain the size distribution of agglomerates accurately. Modeling of aluminum agglomeration was carried out by constructing several submodels, including propellant combustion, cluster analysis of aluminum particles, aluminum ignition and coalescence based on the packed three-dimensional propellant microstructure unit. Dynamic combustion process of propellant microstructure unit was simulated firstly, which characterizes the temperature distribution above the burning surface. Next, aluminum clusters were grouped using a new local density peaks-spectral clustering (LDP-SC) algorithm for initial aluminum particles in microstructure unit. The ignition of aluminum particles in clusters exposed from burning surface was then examined by considering their distributions in the flame of propellant. These exposed aluminum particles in a cluster yield an agglomerate once an aluminum particle is successfully ignited. Secondary mergence of multiple agglomerates was also considered when the distance between two particles in different agglomerates was less than a critical separation distance. With these submodels, the agglomeration process of a typical propellant and properties of aluminum clusters and agglomerates were estimated. The predicted size distribution of agglomerates shows reasonable agreement with experimental data, indicating that the present model is capable of predicting aluminum particle agglomeration during composite propellant combustion.

Influence of pre-ignition technology on combustion and explosion characteristics of composite propellants

The slow cook-off test is widely used for analyzing the safety of propellants under accidental thermal stimulation. To examine the influence of the pre-ignition technology of a composite propellant on thermal safety and combustion characteristics, a quantitative evaluation experiment of slow cook-off reaction intensity of a typical composite propellant charge was designed and executed, considering hydroxyl-terminated polybutadiene (HTPB) and hydroxyl‑terminated block copolyether (HTPE) propellants containing ammonium perchlorate (AP) and aluminum (Al). The experimental method for quantitative evaluation of the slow cook-off reaction intensity is highly essential to investigate the safety of ammunition. The influence of pre-ignition on reaction intensity was analyzed at different temperatures, the combustion behavior of the heated propellant was studied using a closed bomb experiment, and the influence of pre-ignition on combustion evolution was analyzed. Finally, the process was characterized by theoretical modeling. The results show that the slider speed was 169.3 m/s when the HTPB/AP/Al propellant was ignited at 234 °C. The slider speed was reduced to 43.4 m/s and 61.8 m/s by pre-ignition at 120 °C and 214 °C, respectively. When the spontaneous ignition temperature of HTPE/AP/Al propellant was 212 °C, the slider speed was 110.5 m/s. Pre-ignition at 120 °C and 192 °C reduced the slider speed to 39.4 m/s and 43.7 m/s, respectively. Pre-ignition can effectively reduce the reaction intensity. Under the same conditions, the reaction strength of HTPE/AP/Al propellant is marginally lower than that of HTPB/AP/Al propellant. In the slow heating process before ignition, with the increase in ignition temperature, when HTPB/AP/Al propellant was pre-ignition at 120 °C and 214 °C, the pressurization time was 9.8 ms and 6.4 ms, respectively, while before ignition, the pressurization time decreased to 5.5 ms. At 120 °C and 192 °C, the pressurization time of HTPE/AP/Al propellant was 50.9 ms and 36.6 ms, respectively, while before ignition, the pressurization time decreased to 26.9 ms. The reduction in ignition temperature decreased the pressure growth rate of propellant during combustion and reduced the pressurization time. A calculation model of the propellant combustion process was established to characterize the growth process of the combustion pressure at different ignition temperatures. This model can effectively predict the combustion evolution of composite propellants after the slow cook-off ignition and aid in ensuring propellant safety.

Combustion of Composite Propellants with Titanium

Combustion of Composite Propellants with Titanium

Controlling the combustion and agglomeration characteristics of a solid composite propellant via a DC electric field

This study explores the use of a DC electric field to control the combustion and agglomeration characteristics of a solid composite propellant as typically deployed for rocket propulsion. Experiments were performed in a custom-built reacting flow facility that supports high-speed photography and the collection of condensed combustion products (CCPs). The propellant burning rate, aluminum agglomeration process, CCP size distribution, and combustion efficiency were systematically examined under a range of electric-field voltages (−5 to 5 kV) in both the forward and backward directions. It was found that as the electric-field voltage increased, both the propellant burning rate and combustion efficiency increased, while the degree of aluminum agglomeration decreased, reducing the particle size of the CCPs. The introduction of an electric field can bring about profound effects via particle charge acceleration, Rayleigh crushing, and ion excitation. For example, aluminum near the burning surface can split into smaller particles, the diffusion flames around droplets can exhibit a higher stand-off ratio, and more hollow agglomerates can form, as evidenced by an inspection of scanning electron microscopy images of the CCPs. The electric field was found to have a more pronounced effect in the forward direction than in the backward direction. A physical mechanism was proposed for how an electric field can alter propellant combustion. Through this, the combustion and agglomeration characteristics of solid composite propellants can be precisely tuned with an electric field without resorting to changes in the propellant formulation. The findings of this study contribute to a better understanding of how an electric field can be used to control propellant combustion and thrust in solid rocket motors.

Decomposition and combustion of HTPB-based composite propellants containing intercalated HMX crystals with desired high energy but low burn rate

Two types of qy-HMX (insensitive HMX with improved energy density, numbered 1 and 2) have been used and studied as an ingredient of HTPB-based composite propellants (CPs), and the decomposition and combustion performances have been investigated. It has been shown that the main exothermic peak temperature of blank CPs containing 21 wt% of qy-HMX-1 is 258.9 °C, which is 4.7 °C higher than that of HMX-containing one, indicating improved thermostability. The corresponding heat of explosion for blank qy-HMX-2-based CPs was measured to be 5830 J∙g−1, which is 300 J∙g−1 higher than that of HMX-based one. Interestingly, the qy-HMX-2-based CPs has higher energy density than corresponding CL-20-based ones, but it has much lower burn rate under the same pressure, e.g. 4.38 mm∙s−1 at 3 MPa for the former vs. 9.03 mm∙s−1 for the latter. Under the effect of typical conventional catalysts such as nano-CuO and nano-Fe2O3, the qy-HMX-1-based CPs showed increased heat of explosion and higher burn rate. In addition, the particle sizes of condensed combustion products (CCPs) of CPs containing qy-HMX-2 are much smaller than that of HMX-based one, where the Al4C3 was produced as its unique CCPs ingredient.

GASEOUS EMISSIONS CHARACTERIZATION FROM POTASSIUM NITRATE-BASED ROCKET COMPOSITE PROPELLANT COMBUSTION

GASEOUS EMISSIONS CHARACTERIZATION FROM POTASSIUM NITRATE-BASED ROCKET COMPOSITE PROPELLANT COMBUSTION

Modifying the ignition, combustion and agglomeration characteristics of composite propellants via Al-Mg alloy additives

The performance of propellants, explosives and pyrotechnics is known to benefit from the use of metallic alloys as fuel additives. Previous studies have shown that switching from pure Al to Al-Mg alloy can improve the agglomeration and combustion features of composite propellants. However, the optimal ratio of Al and Mg has yet to be explored. We experimentally examine the effects on ignition, combustion and agglomeration arising from substituting pure Al particles with four different Al-Mg alloys whose Mg content ranges from 10 to 60%. Using thermogravimetric−differential scanning calorimetry, we find that switching to Al-Mg alloy can enable oxidation to occur earlier and can increase the oxidation rate and the enthalpy of thermal reaction, which is found to be positively correlated with the Mg content. Using laser ignition measurements, we find that the addition of Mg can promote the ignition and combustion of Al, with (Al/Mg)0.5 exhibiting the most balanced combination of a high combustion intensity, a short ignition delay, and a high mass burning rate. Through high-pressure measurements taken at 1 to 7 MPa, we find that switching to Al-Mg alloy can increase the propellant burning rate (by nearly a factor of two) and the pressure exponent (from 0.23 to 0.43). We also find that the burning rate and pressure exponent are both highest for the propellant containing (Al/Mg)0.1, but that both quantities decrease with further increases in Mg content. High-speed microscopic surface imaging at 1 MPa reveals that increasing the Mg content can increase the overall combustion intensity of the propellant and decrease the diameter of spherical agglomerates, consistent with the observed size distribution of the condensed combustion products. From these measurements, it can be concluded that (Al/Mg)0.5 provides the optimal balance of ignition, combustion and agglomeration characteristics in the formulation of composite propellants containing Al-Mg alloy.

Combustion promotion and agglomeration reduction of the composite propellant using graphene

Graphene and its composites are promising catalysts for energetic materials. In order to realize the application of graphene in solid propellant, the effects of graphene content on the ignition, combustion and agglomeration characteristics of composite propellant were thoroughly conducted based on thermogravimetric-differential scanning calorimetry measurements (TG-DSC), laser ignition, combustion diagnoses, burning rate test and condensed combustion products collection. Results show that graphene effectively promotes the thermal decomposition of AP. 3 wt.% graphene advances the thermal decomposition temperature of AP by 68.3 °C, and 1 wt.% graphene increases the apparent decomposition heat of AP by 2.6 times. Laser ignition experiments show that the combustion intensity and mass burning rate increase, while the ignition delay time decreases with increasing graphene content. The addition of graphene is found to increase the burning rate at 0.1 MPa, but the effect is not obvious at 1-7 MPa, thus reducing the pressure exponent. 0.5 wt.% graphene can reduce the pressure exponent from 0.43 to 0.34. Compared with the baseline propellant, more aluminum particles are ignited on the surface of the propellant containing graphene and the large size agglomerates in the condensed combustion products decrease significantly. Overall, graphene plays a comprehensive effect on the ignition, combustion and agglomeration characteristics of aluminized propellant. Results of this study can help guide the application and development of graphene or graphene-inclusion materials in solid propellant for solid rocket motors.

Very-High-Pressure Burning Rates of Aluminized and Nonaluminized AP/HTPB-Composite Propellants

To further characterize the roles of aluminum and ammonium perchlorate (AP) characteristics in ammonium perchlorate/R45-M hydroxyl-terminated polybutadiene (AP/HTPB)–composite propellants, aluminized and nonaluminized formulations were tested at very high pressures, up to 68.9 MPa (10,000 psi). A total of nine formulations were examined: seven nonaluminized with varying AP distributions, concentrations, and average particle sizes, and two aluminized with varying aluminum concentrations. All nine propellants showed an exponent or “slope” break above 20 MPa (2900 psi) and a postbreak pressure exponent greater than one. Decreasing the AP particle size decreased both the characteristic pressure where the exponent break occurred (or ) and the pressure exponent after the break. AP distribution also affected , whereas changes in AP concentration did not. The inclusion of aluminum lowered compared with the nonaluminized formulations. Additionally, increasing the aluminum concentration appeared to lower the postbreak pressure exponent. The present study provides one of the first fundamental, systematic studies on the exponent break feature of AP/HTPB-composite propellants and adds new data to the limited database available in the open literature on the AP-based, composite propellant burning rate exponent break at very high pressures, with emphasis on the effect of mixture variables and Al on the characteristic break pressure and the postbreak pressure exponent.

Isolating the effects of oxidizer characteristics and catalytic additives on the high-pressure exponent break of AP/HTPB-Composite propellants

AP/HTPB-composite propellants are widely used in the field of energetics due to their burning rate controllability using a variety of formulation variables. At higher pressures, the burning rate becomes less controllable as it experiences a transition regime often referred to as an “exponent break” typically between 20.7 and 34.5 MPa. The pressure exponent drastically increases to values greater than 1, making the burning rate extremely sensitive to pressure fluctuations. This study systematically evaluated the effects of catalysts on the exponent break and high-pressure burning rates of AP/HTPB-composite propellants containing a variety of AP particle sizes, distributions, and concentrations. The burning rates of seven formulations with varying AP characteristics and either Mach I iron oxide or titania nanoparticles or in-situ titania as a catalyst were evaluated between pressures of 6.89 MPa and 68.9 MPa. All formulations with the exception of two, 46.0-µm AP with Mach I titania and 138.9-µm AP with 0.50% in-situ titania, showed an exponent break. For the others, the characteristic pressure where the exponent break occurred changed with the inclusion of the catalysts, increasing it to 34.5 MPa and higher. The characteristic pressure was found to be dependent on the corresponding baseline burning rates, occurring where the additive burning rate curve intersected with its respective baseline burning rate curve. Additionally, all the burning rates fell above the AP deflagration rate curve, corroborating the existence of an AP barrier and the general theory in the literature that AP decomposition dominates in the very-high-pressure regime. This study adds new data to the severely limited database for very-high-pressure, AP-based, composite propellant burning rates in the open literature and provides one of the first fundamental studies on the exponent break feature with emphasis on the role of mixture variables and catalytic additives on the characteristic break pressure and the post-break pressure exponent.

Three-dimensional spatial distributions of agglomerated particles on and near the burning surface of aluminized solid propellant using morphological digital in-line holography

The particle size distribution of aluminum particles during combustion plays an important role in predicting the combustion performance of aluminized solid propellants. However, conventional models face difficulties in estimating the distributions of the agglomerated particles in propellants with novel or complex compositions. This study investigated the accuracies of the experimental techniques applied to propellant combustion to obtain the three-dimensional (3D) size distributions of agglomerated particles on the burning surface and particles in the plume. The 3D spatial distributions of aluminum particles on the burning surface and in the plume during solid aluminum composite propellant combustion were obtained using digital in-line holography (DIH) and morphological watershed image segmentation. Overall, 68,321 individual particles at distances of up to approximately 20 mm from the burning surface of the propellant were analyzed under 1 MPa pressure. The field of view results suggest that approximately 7% of the particles were detected on the burning surface, displaying a trimodal distribution in the particle size range of 30–1,200 μm, and almost all the particles (>96%) were agglomerated. The mean diameters and agglomeration fractions in the height range of 0–1 mm were larger than those in the overall combustion field; the mean diameters in this height range were similar to those of the particles on the burning surface. We found that the process of “second mergence” occurred at and up to 1 mm above the burning surface. As the height from the burning surface increased, the mean particle diameter and agglomeration fraction decreased rapidly; however, when the height exceeded 11 mm, the mean diameter did not change significantly, indicating that the particles had almost burned completely and formed condensed combustion products. Significant differences were found between the results obtained from the empirical agglomeration models and those obtained from the 3D DIH because these models were developed based on 2D data. The results presented herein further provided an insight into the agglomeration characteristics of aluminum particle combustion in composite propellants.

Investigation of Liquid Breakup Process Solid Rocket Motor Part A: Horizontal Converging–Diverging Nozzle

Abstract This paper presents a solid rocket motor (SRM), both experiment and simulation of alumina molten flow patterns using the cold flow case. The combustion of aluminum composite propellants in SRM chambers causes high-temperature and pressure conditions resulting in the liquid alumina (Al2O3) as a combustion product, and it tends to agglomerate into molten droplets, which impinge on the propulsion chamber walls and then flow along the nozzle wall. This liquid alumina in the flow creates problems such as chemical erosion of the propellant and mechanical erosion of the nozzle. Thus, particle size and droplet distribution are considered to affect the erosive behavior. Furthermore, for the rocket motor, converging–diverging (C–D) of the nozzle is used because of its high performance in terms of the rate of change of momentum. In this study, to investigate the relationship between air velocity and molten particle size, the study was mainly focused on the horizontal arrangement of the combustion chamber with the cold flow with the liquid.

Effect of Long-Chain Bonding Agent on the Combustion of Composite Propellant and Modification of Combustion Performance Using Nano Additives

ABSTRACT Novel long-chain bonding agent MST was synthesised and introduced into the bonding agent matrix in order to enhance the mechanical properties of composite propellant by improving the interfacial interactions between the bonding agent and solid particles, in order to withstand stresses produced due to combustion loading conditions, changes in environmental condition, transportation and handling without effecting on the ballistic performance. In this work, the replacement of the long-chain MST bonding agent instead of the most famous reference MAPO bonding agent has improved the mechanical properties especially strain value corresponding to stress value which doubling the maximum strain value with an acceptable value of maximum stress with a good value with Young’s modulus in propellant formulations. Different compositions of composite propellant were carried, in order to measure the linear burning rate and to study the effect of the replacing MST bonding agent instead of tradition bonding agent MAPO. Due to the presence of a bonding agent, the burning process of the composite propellant should become more regular and stable burning but it reduces slightly the burning rate as in formulation F2 in which the pressure exponent (n) value reached .067. It should be modifying the reduction in burning rate due to the presence of MST in order to achieve promising burning properties. By using nano-aluminum it was secured improving performance and achieved a high burning rate of the propellant formulations based on MST bonding agent as well as the composite propellant formulations based on MAPO bonding agent. And by using nano-copper chromite the values of the burning rate of composite propellant formulations became greater than those values of formulations based on nano-aluminum with the same weight ratio and particle size, due to the nature of copper chromite and its effect on the binder and the ammonium perchlorate.

Combustion of Aluminum and Boron Agglomerates Free Falling in Air. I. Experimental Approach

This paper presents a review of studies of the combustion of composite propellants containing a combined fuel based on aluminum and boron. A method for studying the combustion of large particles of the combined Al + B fuel in air is described. Burning agglomerated Al/B particles 300–700 μm in diameter were obtained by ignition of miniature pieces of a composition containing 32% binder and 68% micron-sized aluminum and boron powders in the ratio Al/B = 81/19 placed in a burning metal-free sample. Agglomerates formed by the merger of many small particles burned in free fall in air. Procedures are described that were used to process video records of the combustion process and study condensed combustion products (combustion residues of agglomerates) in order to determine the burning time and analyze the transformation of the combined fuel into oxide.

Initial Studies on Development of High-Performance Nano-structured Fe2O3 Catalysts for Solid Rocket Propellants

Space launch vehicles and strategic military vehicles commonly employ composite solid rocket propellants containing Fe2O3 as ballistic modifier. Nanoscale catalysts, including nanoscale Fe2O3, have been reported to exhibit superior activity in the thermal decomposition and combustion of composite rocket propellants. However, scalable methods to prepare such nano-structured catalysts with high performance as ballistic modifiers and systematic studies relating the synthesis parameters to the catalyst characteristics and consequently to the thermal and combustion properties of the composite propellant are scarce. In this paper, we report a novel and facile route to prepare nano-structured Fe2O3 with enhanced catalytic activity in the ballistic modification of ammonium perchlorate (AP)-based composite solid rocket propellant. A submerged spray precipitation method using air-assisted liquid-centered coaxial atomization has been developed to prepare these nano-structured Fe2O3 catalysts. The prepared Fe2O3 catalysts possess higher surface area and exhibit superior activity in the thermal sensitization of AP, leading to an 88% increase in the burning rate of AP-based composite solid rocket propellants, than the Fe2O3 catalyst prepared via traditional precipitation method. Merits of the developed preparatory route, influence of the atomization process on the nano-structure morphology and subsequent benefits of these nano-structured catalysts on the ballistic properties of AP-based propellants are demonstrated and discussed in this paper.

Aging Effects on the Burning Rates of Composite Solid Propellants with Nano-Additives

The aging behavior of composite propellant is a key parameter in determining the usable service life of an energetic system. The effect of aging on composite propellant burning rates is an understudied topic in the open literature. Recent research in the authors’ laboratory has led to the development of high-burning rate propellants containing titania nanoparticles and fluorescing propellants containing quantum dots. The current study assessed the aging behavior of composite ammonium perchlorate (AP)/hydroxyl terminated polybutadiene propellants, with specific emphasis on fluorescent properties and burning rates. Propellant formulations were based on either 80% monomodal or 85% bimodal AP distributions with and without titania nanoparticles and quantum dots. Propellant samples were thermally aged at temperatures of 63 and 85°C and humidity levels of 25 and 75% for up to 195 days to yield approximate service lives in the range of 0–25 years. Fluorescent imaging experiments were conducted with an ultraviolet light source to measure the uniformity of propellant samples. Ballistic testing was conducted in a constant-volume strand burner over the pressure range of 500–2250 psia (3.5–15.5 MPa). Spatial uniformity of the propellant fluorescence was not significantly altered with propellant age but did decrease in intensity to some degree. In general, the burning rate and pressure index of propellants decreased with increasing service life. Propellant formulations containing 85% bimodal AP were found to be more stable than formulations containing 80% monomodal AP. Inclusion of quantum dots or titania nanoparticles in the propellant formulation increased aging stability.

Aluminized composite propellant combustion modeling with Heterogeneous Quasi-One dimensional (HeQu1-D) approach

The Heterogeneous Quasi One-dimensional (HeQu1-D) model for AP/HTPB composite propellant combustion is extended to aluminized propellants. Following the serial burning approach of HeQu1-D, a statistical particle path through an aluminized propellant is taken to consist of aluminized binder-matrix coated AP particles of various sizes. Extending the idea of homogenization of fine-AP particles, aluminum particles are assumed to be homogenized with the binder-matrix. Large Al particles (nominal size ≥ 15 µm) in the binder-matrix are assumed to get ejected into the gas phase and hence do not contribute to heat feedback to the burning surface. With these modifications, the model is shown to accurately predict the burn rate variation with pressure and initial temperature of propellants using Al of nominal sizes in the range of 15–50 µm (termed conventional aluminum, CAl). The experimentally observed reduction in propellant burn rate with substitution of AP by CAl is shown to be due to the increase in fuel richness and energy sink effect of melting of Al. Catalytic effects due to addition of burn rate modifiers (Fe2O3 in space shuttle booster propellant, for instance) are also accounted for by modifications to gas phase rate parameters. Increase in burn rate up to 25% with Al particles of a few µm (3–6 µm) in comparison with non-aluminized propellant is explained by recognizing that there will be non-negligible fraction of sub-micron Al due to the associated particle size distribution. This leads to heat release by sub-micron Al combustion close to the surface and an enhancement in the heat feedback, by a combination of convective and radiative mechanisms. Dramatic increase in burn rate with sub-micron Al (a factor of 4–5 as compared to CAl) is captured by invoking radiation from fine-Al/Al2O3 particles to propellant surface. Agglomeration of sub-micron Al particles is invoked to explain the saturation of burn rate enhancement with reduction in Al particle size and the effectiveness of sub-micron Al substitution in smaller fractions (bi-modal Al). Predictions for over fifteen different propellants with varying fractions of fine and coarse aluminum from earlier literature have been presented. Comparisons of the predictions from the model with experimental results from different sources is shown to be excellent-to-good for most cases. The hitherto unknown radiation dominated ablation (r˙>50mm/s) of coarse AP particles and the effect of Al particle size on propellant temperature sensitivity are brought out. The MATLAB® code based on the model offers opportunity for design of AP-HTPB composite propellants with combination of fine and coarse aluminum with confidence.

A novel approach to composite propellant combustion modeling with a new Heterogeneous Quasi One-dimensional (HeQu1-D) framework

This paper is concerned with a new and novel approach to modeling composite propellant burn rate behavior. It is founded on the fact that composite propellant combustion is largely boxed between the premixed limits – of pure AP and fine AP-binder (HTPB, here) whose burn behaviors are taken as known. The current strategy accounts for particle size distribution using the burn time averaging approach. The diffusional effects are accounted for through a calibrated heterogeneous quasi-one-dimensional model (HeQu1-D for short) that allows for the flame temperature dependence on the local AP size-binder thickness geometry. Fine AP-binder homogenization is adopted as in recent models with refinement on the particle size as a function of pressure. The specialty of the present approach is that it invokes local extinction for fuel rich conditions for specific particle sizes when the heat balance causes the surface temperature to drop below the low pressure deflagration limit of AP; this feature allows for the prediction of extinction of propellant combustion. Combining these ideas into a MATLAB® calculation framework that uses a single dataset on properties of AP and binder consistent with burn rate vs. pressure of pure AP and fine AP-binder system allows for making the predictions of propellants with multiple particle sizes and different fractions. Comparisons of burn rate data over nearly thirty compositions from different sources appear excellent to good. It is found that it is important to treat the full particle size distribution to achieve better predictions. Low burn rate index (∼0.25) observed with addition of SrCO3 is captured by extending the model to include the effect of binder melt; the gas phase effect is accounted for by calibration against catalytic effect on the fine AP-binder propellant. An interesting deduction from the model is that the temperature sensitivity of propellants should not exceed that of AP. The robustness of the current model and speed of determining the burn rate behavior allow for the possibility of determining the particle size distribution required to meet the burn rate specifications of a specific propellant for practical applications before actually embarking on making the propellant.