Burning Rate Of Solid Propellant Research Articles

Evaluation of solid propellant burning rate by using strand burner and static firing tests

Abstract The main objective of this paper is to compare the different methods used to measure the burning rate of hydroxyl-terminated polybutadiene (HTPB) propellant. The burning rate measurement is carried out under varying crosslinking density, which is determined by the curing ratio, i.e., the equivalent ratio of diisocyanate to total hydroxyl (NCO/OH) ratio. Four batches with different curing ratios were produced using fixed production processes. To measure the burning rate, cured solid propellant strands were tested using the acoustic wave emission method at different pressures ranging from 4 to 10 MPa and also by using static firing test of subscale motors. The study showed that there was little difference between the results obtained from the acoustic wave emission method and subscale motors.

Closed vessel burning rate measurements of composite propellants using microwave interferometry

AbstractBurning rate as a function of pressure is one of the primary evaluation metrics of solid propellants. Most solid propellant burning rate measurements are made at a nearly constant pressure using a variety of measurement approaches. This type of burning rate data is highly discretized and requires many tests to accurately determine the burning rate response to pressure. It would be more efficient to measure burning rate dynamically as pressures are varied. Techniques used to make transient burning rate measurements are reviewed briefly and initial results using a microwave interferometry (MI) technique are presented. The MI method used in tandem with a closed bomb enables nearly continuous measurement of burning rates for self‐pressurizing burns, capturing burning rate data over a wide range of pressures. This approach is especially useful for characterization of propellants with complex burning behaviors (e. g., slope breaks or mesa burning). The burning rates of three research propellants were characterized over a pressure range of 0.101–24.14 MPa (14–3500 psi). One research propellant exhibited a slope break at a pressure of 6.63 MPa (960 psi). Using MI in a closed pressure vessel, 14 propellant strand burns resulted in a nearly continuous burning rate curve over a pressure range of 0.41–24.13 MPa (60–3500 psi) that reasonably matched conventional burning rate measurements. The development of this technique provides an opportunity to quickly characterize the burning rate curve of solid propellants with greater fidelity and efficiency than traditional quasi‐static pressure testing techniques.

Methodology for Studying Combustion of Solid Rocket Propellants using Artificial Neural Networks

The combustion properties of energetic materials have been extensively studied in the scientific literature. With the rapid advancement of data science and artificial intelligence techniques, predicting the performance of solid rocket propellants (SRPs) has become a key focus for researchers globally. Understanding and forecasting the characteristics of SRPs are crucial for analyzing and modeling combustion mechanisms, leading to the development of cutting-edge energetic materials. This study presents a methodology utilizing artificial neural networks (ANN) to create multifactor computational models (MCM) for predicting the burning rate of solid propellants. These models, based on existing burning rate data, can solve direct and inverse tasks, as well as conduct virtual experiments. The objective functions of the models focus on burning rate (direct tasks) and pressure (inverse tasks). This research lays the foundation for developing generalized combustion models to forecast the effects of various catalysts on a range of SRPs. Furthermore, this work represents a new direction in combustion science, contributing to the creation of a High-Energetic Materials Genome that accelerates the development of advanced propellants.

Исследование газодинамики горения смесевого твердого топлива при колебаниях давления

The non-stationary burning rate of a solid rocket propellant under harmonic variations in pressure over the combustion surface is studied. The physical and mathematical model is based on the equations of heat transfer and oxidizer decomposition in the solid phase, and on the flow model for the reacting products of solid propellant gasification. Calculations are carried out for the unsteady burning rate of mixed solid propellants with harmonic pressure variations over the combustion surface. The dependence of the amplitude of the burning rate fluctuations on the frequency of the pressure variations are obtained. The amplitude of the burning rate varies non-monotonically with frequency. As the frequency increases, the amplitude first increases and then decreases. At a frequency of pressure variation with a semi-period greater than or equal to the characteristic time of relaxation for heat transfer in solid propellants, the instantaneous burning rate at the lowest point of the pressure curve is lower than the equilibrium value and is higher at the highest point. At high frequencies, the burning rate exceeds the corresponding equilibrium value at minimum pressure and is lower than the equilibrium value at maximum pressure.

Effect of Acoustic Pressure Oscillations on Burning Rate Augmentation of Composite Solid Propellants at Different Initial Grain Temperatures

ABSTRACT The effect of acoustic pressure oscillations, besides initial grain temperature on the burning rate of composite solid propellants (AP/HTPB/RDX/Al), is experimentally investigated using a T-burner. Mean burning rate is a by-product of the T-burner experiment while screening different propellants for its stability behavior by evaluating its combustion response. Experiments at different frequencies ranging from 200 to 400 Hz and at different pressures of 2–8 MPa and at different initial grain temperatures (room temperature (303 K), hot (343 K), and cold (243 K)) by soaking the propellant samples by a novel method for two industrial-grade propellants (propellant-A and propellant-B) were conducted. Many of the tests were self-excited for propellant-A, whereas propellant-B needed to be pulsed in most of the tests. Various levels of acoustic amplitudes were imposed up to 300 kPa, which is not otherwise possible to impose it externally using any other acoustic drivers, for example, a rotary valve (up to 20 kPa max. is possible). The test results revealed that the acoustic pressure oscillations augment the burning rate tremendously. It was found that the burning rate augmentation factor of propellant-B is lower than propellant-A in most of the cases. The investigations revealed the following new findings: (a) Increasing the initial grain temperature in the presence of acoustic pressure oscillations tends to decrease the augmentation; (b) Many of the results of the augmentation showed a rising and falling trend with acoustic amplitude; (c) The burning rate augmentation factor from self-excited and pulsed tests surprisingly form a continuous trend and shown together. The positive and negative augmentation compels us to further investigate toward a better understanding about the steady combustion behavior of solid propellants since there is no model available as on today to predict the burning rate augmentation precisely due to the combined effect of acoustic pressure oscillations and initial grain temperature.

Self-assembly Hybridization of CL-20 and Hydrazone-based Polymeric Carbon Nitrides for High Performance Explosive

A strategy for desensitization of the highly sensitive explosive CL-20 with polymeric carbon nitrides is described. The hybrid energetic material (MHGP-CL-20) was synthesized via a self-assembly approach, where CL-20 molecules are immobilized on the hydrazone-based polymeric carbon nitrides. The good detonation performances (detonation velocity 9608 m·s−1, detonation pressure 42.8 GPa) show that MHGP-CL-20 retains the excellent performance of CL-20 to the greatest extent. More importantly, its sensitivities were effectively reduced with the impact sensitivity (IS) 10 J and the friction sensitivity (FS) 160 N, which is much lower compared with raw ε-CL-20 (IS: 4 J; FS: 94 N). In addition, its specific impulse was improved to 276.6 s, higher than that of ε-CL-20 (272.4 s). By preparing solid composite propellants, the burning rate of the propellant grain PMC20 based on MHGP-CL-20 is higher than PC20 based on ε-CL-20. The addition of hybrid energetic complex could keep the stability of combustion and improve the burning rate of solid propellant. The combination of PCNs material and CL-20 sustains the excellent performance of CL-20 and reduces the sensitivity effectively. The preparation of hybrid energetic material MHGP-CL-20 finds a new balance in the contradictory system of high energy and low sensitivity.

Effect of Carbon Nanotubes (CNTs) on the performance of Solid Rocket Propellants (SRPs): A short review

Burning rate catalysts (BRCs) are usually utilized to adjust the burning rates of solid rocket propellants (SRPs). Carbon nanotubes (CNTs), for their unique structure and properties, have been studied by several researchers worldwide with a wide range of applications. In this paper, several worthy CNTs-based materials and effects are described. Specifically, the influence of CNTs on the thermal behavior of energetic materials (EMs) are discussed, and we investigated the impact of CNTs/metal (metal oxide, MO) composites on the thermal degradation of EMs. Applications for SRPs and solid rocket nozzle motors might benefit from a focus on the examination of the burning rate, pressure exponent characteristics, and hazardous aspects. It was discovered that CNTs, as opposed to the comparable micro-sized additives, can modify the combustion behavior and speed up the burning of SRPs. Lastly, the difficulties encountered in implementing some of these applications are also examined in terms of manufacturing, processing, pricing, and potential future uses.

Insight into the precise catalytic mechanism of CuO on the decomposition and combustion of core–shell Al@AP particles

Burning rate modulation is a crucial but challenging technology for designing solid propellants. The core–shell structured Al@AP and Al@AP/CuO were prepared and their thermal reaction properties were investigated at different pressures. The precise catalytic mechanism of doped CuO in oxidizer and the key reason for its burning rate modulating effect on solid propellants were elucidated by combining DFT calculations. The *NH4ClO4 on the CuO (111) surface is more likely to undergo a dehydrogenation reaction first. The large difference in activation energy between *NH4ClO4→*NH3+*ClO4+*H and *NH4ClO4 → NH3 + HClO4 explains the reasons that CuO can reduce the ignition delay of the Al@AP. CuO promotes the decomposition of the AP@Al at the low- and high-temperature decomposition (LTD and HTD) stage. For the Al@AP-CuO, the temperature of the high-temperature decomposition peak (THTD) does not shift to the lower temperature range with increasing pressure. These results are responsible for a reduction of pressure exponent in the presence of minor CuO as reported in the literature. CuO can facilitate the transformation process *NH3→*NH2 on the CuO (111) surface at low-temperature and low-pressure conditions. However, at high-temperature and high-pressure conditions, it can inhibit the reaction *NH2→*NH→*N, so the burning rate of solid propellants can be controlled.

Prediction of Solid Propellant Burning Rate Characteristics Using Machine Learning Techniques

AbstractWhen formulating a new solid propellant, one of the most important aspects of its performance is the burning rate's response to a change in pressure. For this reason, it is useful to be able to predict the burning rate response of a given propellant before the propellant formulation is created such that experimental trade studies are minimized or reduced in scale. While many theoretical and phenomenological models have been proposed to explain various aspects of energetic material combustion, little work has been made publicly available in the application of machine learning models to predicting solid propellant burning rates. To facilitate model creation, the material formulation and burning rate parameters for over 600 publicly available propellant formulations have been collected into a coherent data set. This work utilizes the large amount of publicly available data to inform a random forest machine learning (ML) model in the prediction of solid propellant burning rate parameters. This ML model operates over a large parameter space including ammonium perchlorate composite, plastic bonded high explosive, and double‐base propellants. The model's accuracy, adaptability, and prediction capabilities are presented and discussed. The effects of different materials on a propellant's expected burning rate are examined.

Design, preparation, and combustion performance of energetic catalysts based on transition metal ions (Cu2+, Co2+, Fe2+) and 3-aminofurazan-4-carboxylic acid.

Development of energetic catalysts with high energy density and strong catalytic activity has become the focus and frontier of research, which is expected to improve the combustion performance and ballistic properties of solid propellants. In this work, three energetic catalysts, M(H2O)4(AFCA)2·H2O (AFCA = 3-aminofurazan-4-carboxylic acid, M = Cu, Co, Fe), are designed and synthesized based on the coordination reaction of transition metal ions and the energetic ligand. The target products are characterized by single crystal X-ray diffraction, Fourier transform infrared spectroscopy, differential thermal analysis, optical microscopy, and scanning electron microscopy. The results reveal that Cu(H2O)4(AFCA)2·H2O crystallizes in the monoclinic space group, Dc = 1.918 g cm-3. Co(H2O)4(AFCA)2·H2O, and Fe(H2O)4(AFCA)2·H2O belong to orthorhombic space groups, their density is 1.886 g cm-3 and 1.856 g cm-3, respectively. In addition, the designed catalysts show higher catalytic activity than some reported catalysts such as Co(en)(H2BTI)2]2·en (H3BTI = 4,5-bis(1H-tetrazol-5-yl)-1H-imida-zole), Co-AzT (H2AzT = 5,5'-azotetrazole-1,1'-diol), and [Pb(BTF)(H2O)2]n (BTF = 4,4'-oxybis [3,3'-(1-hydroxy-tetrazolyl)]furazan) for the thermal decomposition of ammonium perchlorate (AP). The high-temperature decomposition peak temperatures of AP/Cu(H2O)4(AFCA)2·H2O, AP/Co(H2O)4(AFCA)2·H2O, and AP/Fe(H2O)4 (AFCA)2·H2O are decreased by 120.3 °C, 151.8 °C and 89.5 °C compared to the case of pure AP, while the heat release of them are increased by 768.8 J g-1, 780.5 J g-1, 750.9 J g-1, respectively. Moreover, the burning rates of solid propellants composed of AP/Cu(AFCA)2(H2O)4·H2O, AP/Co(AFCA)2(H2O)4·H2O and AP/Fe(AFCA)2(H2O)4·H2O are increased by 2.16 mm s-1, 2.53 mm s-1, and 1.57 mm s-1 compared with the case of pure AP. This research shows considerable application prospects in improving the combustion and energy performance of solid propellants, it is also a reference for the design and preparation of other novel energetic catalysts.

INFLUENCE OF BURNING RATE INHIBITORS ON THE THERMAL BEHAVIOR AND DECOMPOSITION KINETICS OF AMMONIUM PERCHLORATE

The addition of burning rate inhibitors (BRIs) to propellant formulations can be effective in reducing the burning rate of composite solid propellants (CSPs). Ammonium perchlorate (AP) is commonly used as an oxidizer for CSPs and is a major component of propellants. Therefore, it is of great significance to study the thermolysis performance of AP in the presence of BRIs. In this work, differential scanning calorimetry and thermogravimetric analysis were utilized to study the thermolysis of AP. Oxamide (OXA), ammonium sulfate (AS), ammonium oxalate (AO), and carbamide (CA) were selected as the BRIs. The results showed that as the BRI content increased, the endothermic peak remained essentially unchanged. The exothermic peak in the low-temperature decomposition (LTD) stage shifted to a hightemperature field, while the exothermic peak in the high-temperature decomposition (HTD) stage shifted to a low-temperature field. Furthermore, associated kinetic studies showed that the decomposition activation energies of AP with inhibitors during the LTD stage were higher than that of pure AP, proving that the selected BRIs could effectively inhibit the thermolysis of AP. However, during the HTD stage, the values of the activation energies for decomposition gradually increased in the order of AP + AO, AP + OXA, AP + CA, AP, and AP + AS, indicating that only AS had the effect of inhibition at this time.

Tunning Combustion Behaviors of Composite Modified Double-Based Propellants with a Novel Energetic Burn Rate Modifier

ABSTRACT The flexible control in burning rate and stable combustion of solid propellants over a broad range of pressures essentially determine the performance of solid rocket motors. The nanoscale metals and their oxides are usually utilized as the traditional burn rate modifiers in adjusting the combustion behaviors of solid propellants, but the energy characteristics of propellants are greatly compromised owing to the inert nature of additives used. In this paper, a series of metal complexes of triaminoguanidine-glyoxal polymer (TAGP-Ms) with high nitrogen content were synthesized and employed as energetic burn rate inhibitors in tunning the combustion properties of composite modified double-based (CMDB) propellants. The influences of prepared TAGP-Ms on the thermal decomposition and combustion of CMDB propellants were comprehensively investigated based on thermal analysis technique and combustion characterization methods. It was found that the peak temperatures of the second exotherm displayed for the propellants containing TAGP-Ms inhibitors were significantly increased by more than 30°C (except for TAGP-Cu), indicating the TAGP-Ms is capable of suppressing the thermal decomposition of CMDB propellants. More importantly, the heat releases of CMDB propellants were substantially enhanced by 16%~34% determined by DSC analysis. In comparison to the blank reference, the energetic TAGP-Ms inhibitors could promote the heat of combustion for the involved CMDB propellants by 5%~9%. The burning rates of CMDB propellants were notably reduced over a wide range of pressures by adding TAGP-Ms, whereas the calculated pressure exponents at pressures ranging from 10 to ~ 22 MPa appear to be relatively high due to the accelerated HMX decomposition. The measured combustion wave temperatures suggest that the CMDB propellants with TAGP-Ms undergo a three-stage combustion process with the highest temperatures in the range of 2160 ~ 2230°C.

In Situ Cutting of Ammonium Perchlorate Particles by Co-Bipy "scalpel" for High Efficiency Thermal Decomposition.

Burning rate of solid propellants can be effectively improved by adding catalysts and using smaller size ammonium perchlorate (AP). Although few reports, the exploration of changing the size of AP primary particles by catalysts is of great significance for improving combustion performance. Here, taking Co-bipy as an example, the potential advantages of such materials as AP decomposition catalysts are reported. Due to the existence of NO3 - combined with oxygen rich environment provided by AP, the structural self-transformation from micronrods to nanoparticles can be quickly realized during the heating process. More importantly, when Co-bipy decomposes, it can play the role of "scalpel" and in situ cut AP particles. Results show that high-temperature decomposition of Co-bipy/AP occurs at 305.8 °C, which is 137.5 °C lower than that of pure AP. Catalytic mechanism is discussed by in situ IR and TG-IR, CoO can effectively increase the content of reactive oxygen species and weaken the N-H bond, realizing the rapid oxidation of NH3 . Eventually, the behavior of Co-bipy cutting AP particles is tested. This interesting catalyst structure self-transformation behavior can not only realize the influence on AP, but also perform a positive function in the combustion process of solid propellants, such as opening the adhesive AP interface.

ADN Solid Propellants with High Burning Rates as Booster Material for Hypersonic Applications

AbstractRAM‐ and SCRAM‐Jet engines can not generate static thrust and only work at high speeds. This starting velocity of the motor v0 is usually generated by carrier rockets or boosters, which mostly use solid propellants based on ammonium perchlorate (AP) and hydroxyl‐terminated polybutadiene (HTPB). Modern solid propellants based on ammonium dinitramide (ADN) and Glycidyl Azide Polymer (GAP) have significant advantages in terms of environmental compatibility and thermodynamic performance. In the present work, it is shown that the already high burning rate of ADN solid propellants can be more than doubled by means of metallic fibers, which significantly increases the thrust of such engines. In order to achieve such results, 3 wt% of Al−Mg alloy fibers are added to the propellant. A formulation with 3 mm long fibres shows the highest burning rate with 84.0 mm/s at 26.5 MPa. Small scale motor tests confirm these investigations. The burning mechanism is studied by SEM/EDX investigations and by measuring the thermal conductivity. These show that the fibres remain on the burning surface and thus transfer the heat from the flame to the depth of the propellant. The thermal conductivity of samples with AlMg5 fibers is about 22 % higher than the reference.

On the Use of Fluorine‐Containing Nano‐Aluminum Composite Particles to Tailor Composite Solid Rocket Propellants

AbstractThe burning rate of solid propellants is an important factor for optimizing rocket motors and improving performance. The enhanced burning rate can increase thrust and reduce a propulsion system‘s overall size and weight. In this study, a novel nano‐aluminum/THV composite additive was prepared and introduced into a solid ammonium perchlorate/polybutadiene composite solid rocket propellant to enhance its burning rate. The morphology of the composite particle additive and its effects on combustion were characterized. The use of small quantities (<15 wt.%) of the additive resulted in a burning rate enhancement of up to 2.1 times that of the conventional coarse aluminized propellant with a specific impulse loss of only 3 seconds, and as much as 4.7 times enhancement with a predicted loss of 22 seconds in theoretical specific impulse. Some of this loss may be recovered by the improved combustion efficiency in smaller rocket motors because the additive was shown to significantly reduce the aluminum agglomeration at the propellant burning surface and reduce the size of reaction products which may reduce two‐phase flow losses. The additive also provides wide burning rate tailorability, favorable for motor, grain, and thrust curve design. The burning rate enhancement mechanism is thought to be a physical cratering mechanism governed by the burning rate disparity between the binder/oxidizer system and the nano‐aluminum/fluoropolymer additive and not a chemical catalytic effect.

Combustion Characteristics of Nanoaluminium-Based Composite Solid Propellants: An Overview

The nanosized powders have gained attention to produce materials exhibiting novel properties and for developing advanced technologies as well. Nanosized materials exhibit substantially favourable qualities such as improved catalytic activity, augmentation in reactivity, and reduction in melting temperature. Several researchers have pointed out the influence of ultrafine aluminium (∼100 nm) and nanoaluminium (<100 nm) on burning rates of the composite solid propellants comprising AP as the oxidizer. The inclusion of ultrafine aluminium augments the burning rate of the composite propellants by means of aluminium particle’s ignition through the leading edge flames (LEFs) anchoring above the interfaces of coarse AP/binder and the binder/fine AP matrix flames as well. The sandwiches containing 15% of nanoaluminium solid loading in the binder lamina exhibit the burning rate increment of about 20–30%. It was noticed that the burning rate increment with nanoaluminium is around 1.6–2 times with respect to the propellant compositions without aluminium for various pressure ranges and also for different micron-sized aluminium particles in the composition. The addition of nano-Al in the composite propellants washes out the plateaus in burning rate trends that are perceived from non-Al and microaluminized propellants; however, the burning rates of nanoaluminized propellants demonstrate low-pressure exponents at the higher pressure level. The contribution of catalysts towards the burning rate in the nanoaluminized propellants is reduced and is apparent only with nanosized catalysts. The near-surface nanoaluminium ignition and diffusion-limited nano-Al particle combustion contribute heat to the propellant-regressing surface that dominates the burning rate. Quench-collected nanoaluminized propellant residues display notable agglomeration, although a minor percentage of the agglomerates are in the 1–3 µm range; however, these are within 5 µm in size. Percentage of elongation and initial modulus of the propellant are decreased when the coarse AP particles are replaced by aluminium in the propellant composition.

Nanodiamonds characterization and application as a burning rate modifier for solid propellants

A nanodiamond powder (n-D) synthesized by detonation method was characterized in detail, and the combustion of solid propellants contained 10 wt. % of n-D as a burning rate modifier was investigated. n-D characterization was implemented by XRD, SEM, DSC, and size distribution analyses. Specific surface area analysis was performed using BET-method. The size of primary nanodiamond particles has not changed after acid treatment and drying, but the particles' agglomeration decreased the value of a specific surface area. The effect of n-D modification on the burning rate of solid propellants with an active and inert binder was investigated. The addition of the n-D to propellants formulation instead of carbon black powder has significantly increased their burning rate (by at least 50 %) in the studied pressure range. The observed trend was associated with the high chemical reactivity of n-D and its extremely high value of specific surface area.

Statistical Evaluation of Burning Rate Data of Composite Propellants Obtained from Acoustic Emission Technique

The acoustic emission technique has been considered to be one of the most reliable and robust methods for the measurement of the steady burning rate of composite propellants. In this work, attempts were made to quantify the measurement variability of the burning rate of composite solid propellants by acoustic emission method using statistical tools. A total of 1100 individual measurements were subjected to statistical treatment. The combination of confidence interval and repeatability limit delineated the extent of natural dispersion in the burning rate measurement data. The very high coefficient of variation values for the propellant compositions, having a burning rate of more than 25 mm s–1 raised concerns about the suitability of the acoustic emission method for high burning rate compositions. The Reliability interval approach was employed to determine the statistically significant sample size for different composite propellants having a burning rate range of 5–31 mm s–1. The entire set of data was screened for identification of outlying observation using the Dixon Q test, and the extent of contamination was quantified. Moreover, the application of statistical techniques could have far-reaching implications for quality control perspectives of burning rate measurement by acoustic emission and could be implemented as reference tolerance limits and preventive measures for ensuring the good health of the instrument as well as propellant processing.

Investigation of Electrically Controlled Ammonium Nitrate – Epoxy Solid Propellant at High Pressures

AbstractThe effect of electric voltage/power application on the burning rate of solid propellants composed of ammonium nitrate (AN) and epoxy at high pressures was investigated experimentally. Such an effect can yield an effective means for real‐time managing of the thrust of a solid rocket motor. A burning rate increase up to about 4 folds when increasing voltage application from 70 V to 200 V, was demonstrated. It was attributed to energy (heat) release due to an electric current passing through the conductive melt layer formed at the burning surface. The relative increase in burning rate due to electric effect was higher at the lower pressure range, diminishing for the higher‐pressure end because of the lower relative electric energy contribution. The common pressure dependence correlation of the burning rate could be applied only for limited ranges. However, because of the coupling between the pressure and specific electric power (power per unit surface area) effects, a more complex correlation which includes a variable pressure exponent, was revealed and applied successfully to fit the data over a range of specific electric power and pressure.

ЧИСЛЕННОЕ МОДЕЛИРОВАНИЕ ГОРЕНИЯ СМЕСЕВОГО ТВЕРДОГО ТОПЛИВА, СОДЕРЖАЩЕГО БИДИСПЕРСНЫЙ ПОРОШОК БОРА

A problem of combustion of the composite solid propellants containing various powders of metals and non-metals is relevant in terms of studying the effect of various compositions of powders on the linear rate of propellant combustion. One of the lines of research is to determine the effect of the addition of a boron powder on the burning rate of a composite solid propellant. This work presents the results of numerical simulation of combustion of the composite solid propellant containing bidispersed boron powder. Physical and mathematical formulation of the problem is based on the approaches of the mechanics of two-phase reactive media. To determine the linear burning rate, the Hermance model of combustion of composite solid propellants is used, based on the assumption that the burning rate is determined by mass fluxes of the components outgoing from the propellant surface. The solution is performed numerically using the breakdown of an arbitrary discontinuity algorithm. The dependences of the linear burning rate of the composite solid propellant on the dispersion of the boron particles and gas pressure above the propellant surface are obtained. It is shown that the burning rate of the composite solid propellant with bidispersed boron powder changes in contrast to that of the composite solid propellant with monodispersed powder. This fact proves that the powder dispersion should be taken into account when solving the problems of combustion of the composite solid propellants containing reactive particles.