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Composite propellants are tested using the quench particle collection bomb (QPCB) for the pressure ranging from 2 to 8 MPa to estimate the particle size distribution of aluminum agglomerates from quenched combustion residues emerged out from the burning surface. The major ingredients included in the propellants are ammonium perchlorate (AP), aluminum (Al), hydroxyl-terminated polybutadiene (HTPB), and toluene diisocyanate (TDI). Five propellant compositions are considered in this study; two of them are mixed with catalysts. Propellant formulation variables like the coarse AP/fine AP ratio, total solid loading, catalyst percentage, and aluminum content are varied to assess their effects on the aluminum agglomeration process at different pressures. Unburnt aluminum in agglomerates is continuously getting combusted as they move out from the propellant burning surface. Large agglomerates comprise both Al2O3 and unburnt aluminum. The majority of agglomerates are spherical in shape, and the sizes vary from 31 to 115 $$\mu$$ m for non-catalyzed propellants and from 28 to 136 $$\mu$$ m for catalyzed propellants over the tested pressure conditions. These results can give further insight into the aluminum agglomeration process of catalyzed and non-catalyzed propellants and also affect the choice of the propellant ingredient percentage aimed at reducing aluminum agglomeration, which causes two-phase flow losses of thrust and slag accumulation in full-scale solid rocket motors.

A flake aluminum powder (f-Al) is prepared by a bi-directional rotation mill method using a spherical aluminum powder (s-Al) as a raw material. Then, these two kinds of aluminum powders are used to prepare ammonium perchlorate-based composite modified double-base (AP-based CMDB) propellants. The properties of two kinds of AP-based CMDB propellants, such as the thermal decomposition performance, mechanical sensitivity, and combustion performance, are comprehensively researched. The results show that the exothermic peak temperature of the f-Al/AP-based CMDB propellant is lower than that of the s-Al/AP-based CMDB propellant. The impact and friction sensitivities of the propellant decrease by 12% and 187% after using f-Al to replace s-Al. In addition, the burning rate of the f-Al/AP-based CMDB propellant is 5.5% higher.

The heat of combustion and the enthalpy of formation of 7,7′-bis(terfurazan[3,4-b:3,4′-d:3″,4″-f]azepine (I) and 1,1′-dioxide-7,7′-bis(terfurazan[3,4-b:3,4′-d:3″,4″-f] azepine (II), were experimentally determined. Product I was studied by X-ray diffraction analysis, and its crystallographic characteristics were determined. The efficiency of using compounds I and II as components of solid composite propellants was analyzed, and the type of compositions in which they are more effective than HMX was determined.

This study experimentally confirms the intensification of kerosene combustion in the case where a mixture simulating steam reformation products in a model combustion chamber with a supersonic velocity of the flow at the inlet. It is shown that the used mixture has a higher chemical activity than ethylene. The use of steam reformation products of hydrocarbon or synthetic propellants in schemes with pulse-periodic combustion control increases combustion efficiency without the use of special design solutions for organizing initiation and stable combustion.

This paper presents the results of an experimental study of the influence of the strength of the thin-walled case of a PETN charge on the performance of small-sized pulsating explosive devices for various applications.

The standard enthalpies of formation of 1,1-bis(methoxy-NNO-azoxy)-3-nitro-3-azabutane and 1,1,8,8-tetrakis (methoxy-NNO-azoxy)-3,6-dinitro-3,6-diazaoctane were experimentally determined to be 87.7 $$\pm$$ 3.9 and 283.8 $$\pm$$ 6.2 kJ/mol, respectively. Calculations have shown that solid composite propellants containing these two compounds as gasifying components in metal-free compositions based on an active binder and ammonium perchlorate are inferior in the maximum achievable effective impulse at the third stage of the rocket system $$I_{ef}$$ (3) to compositions based on HMX, but in designing special compositions with a limited content of organic explosive (not higher than 30–35%), these two compounds provide 5–10 s higher values of $$I_{ef}$$ (3) than when using HMX.

This paper describes an experimental study pertaining to laser ignition of igniters of liquid rocket engines operating on such propellants as O $$_{2g}$$ + kerosene, O $$_{2g}$$ + ethanol, and O $$_{2g}$$ + H $$_{2g}$$ , and to how they ignite the propellant of a model combustion chamber. A quartz fiber based optical device is described, which allows for multiple ignition of the igniter and the combustion chamber without replacing the device. This study also touches upon possible types of pulsed lasers used in the laser ignition system, a fiber-optic laser igniter that remain hermetic and operable at a combustion chamber pressure of up to 10 MPa and more than 100 ignition cycles.

This paper presents a comparative analysis of the maximum possible solid rocket motor pressures obtained using the exponential and real laws of the propellant burning rate as a function of pressure. It is shown that at high exponents in the burning rate law, there is a large (tens of percent) difference between the maximum motor pressure obtained by the above two methods. In the case where the exponent is in the range 0.3–0.4, the exponential burning rate law overestimates the motor pressure by only 2–3%.

A comparative evaluation of the use of various metals and nonmetals as fuel components of a solid propellant containing ammonium perchlorate as an oxidizer and rubber as a binder was performed using the flight range criterion of a rocket-ramjet aircraft taking into account the expected completeness of combustion of individual components. The influence of the energetic properties of the propellant on the technical and economic characteristics and flight performance of the aircraft was also taken into account. The aircraft flight range was calculated by numerical integration of flight dynamics equations. Based on the results, it is recommended that a number of fuel components should be studied in more detail, in particular, for the purpose of the full or partial replacement of boron currently used in solid propellants for rocket-ramjet engines.

A method for calculating the optimal ratio of the fuel and solid propellant in the gas generator of a high-velocity flying vehicle is proposed. A thermodynamic approach is used to describe the processes in the gas generator. The ratio of the solid propellant and fuel in the gas generator is determined by their physical and chemical properties. An increase in the density and heat of combustion of the solid propellant and in the fuel porosity leads to reduction of the volume fraction of the solid propellant in the gas generator. An increase in the density, specific heat, specific heat of sublimation, and temperature of sublimation of the fuel leads to an increase in the volume fraction of the solid propellant in the gas generator.

The conditions of occurrence, nature, size, and shape of combustion hotspots of energetic condensed systems and the spatial dynamics of the hotspot combustion wave are discussed. A model for the propagation of hotspot combustion waves in condensed systems is briefly described based on the cellular automata method. A possible explanation for the occurrence of unsteady effects in the combustion of solid rocket propellants is proposed.

This paper describes the experimental study of amplitude (with a frequency of up to 50 kHz) and integral characteristics of the self-magnetic field of high-temperature combustion products of hydrocarbon fuel, which flow out of the nozzle of a standard liquid-propellant engine (LPE) with simulation of the heat of a combustion chamber with injection of aluminum-magnesium alloy particles (c-phase) into the combustion chamber. It is determined that the amplitude of the magnetic field intensity generated by high-temperature (up to 3500 K) combustion products depends on the LPE operation modes and the presence of k phase particles in the particle flow. The magnetic field amplitude increases by 20% during the LPE burnout ≈0.2 s earlier than the pressure drop in the combustion chamber. The total volumetric electric charge generated by the combustion product flow with the c-phase is estimated.

The achievement of stable colloidal suspensions of reactive metal powders in liquid propellants is crucial for obtaining enhanced thrust per unit mass. Aluminium is of interest due to its availability, stability, and high combustion enthalpy (32 000 J/g). In this manuscript, ultrafine spherical aluminium particles with the average size of 15 μm are produced by wet milling. Aluminium particles are effectively surface-modified with a polymeric surfactant and sterically stabilized in an organic solvent (toluene). Organically modified aluminium demonstrates a drastic change in surface properties from hydrophilic to hydrophobic, with effective transfer from the aqueous to organic phase. The stabilized particles are effectively dispersed in a liquid rocket pro-pellant (hydrazine). The impact of aluminium particles on hydrazine combustion characteristics is evaluated by using a thermodynamic code named ICT (Institute of Chemical Technology in Germany, 2008). Aluminium particles offer an increase in the combustion temperature, oxygen balance, characteristic exhaust velocity, and specific impulse. The optimum solid loading level of aluminium in the hydrazine fuel is found to be 6 wt.%.

The development of novel energetic materials with highest possible performance is of current interest. Synthesis of such materials is performed at various stages of pilot plant production all over the world. However, their synthesis involves hazardous production processes. This paper discusses relatively safe and eco-friendly approaches and techniques such as microwave technology and the use of ionic liquids for the synthesis of high-performance energetic materials that can be used as explosives and propellants. In addition, the use of dinitrogen pentoxide as an efficient nitrating agent for the synthesis of energetic materials is considered.

The standard enthalpies of formation of the compounds 1-(2,2-bis(methoxy-NNO-azoxy)ethyl)pyrazole, 1-(2,2-bis(methoxy-NNO-azoxy)ethyl)-3-nitropyrazole, and 1-(2,2-bis(methoxy-NNO-azoxy)ethyl)-4-nitropyrazole were measured experimentally to be 273.6 ± 6.7, 231.0 ± 3.3, and 213.8 ± 7.9 kJ/mol, respectively. These enthalpy values were used to determine the contribution of the replacement of the H atoms at the N atoms in the heterocycles by CH2CH(N2O2Me)2 groups (151.9 kJ/mol). Calculations have shown that 1-(2,2-bis(methoxy-NNO-azoxy)ethyl derivatives of pyrazole, 3- and 4-nitropyrazole, 3,4-dinitropyrazole, and 3,4,5-trinitropyrazole and the bis-derivative of bis-furazano[3,4-b;3′,4′-e]piperazine are inferior to HMX as gasifying components of solid composite propellants in metal-free compositions with an active binder. Only the derivative of 3,4-dinitropyrazole, which added in a small amount together with ammonium Perchlorate to aluminum-free propellant compositions provides a specific impulse of 249 s.

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.

The energy potential of compositions based on a number of hypothetical tetrazole derivatives with a very high mass content of nitrogen (73-75%) have been studied by thermochemical calculations. Quantitative dependences of the energy parameters of solid composite propellants on the nature of the high-enthalpy polynitrogen oxidizer and the presence of metal in the composition have been determined.

Owing to its high mass and volume heats of combustion, boron is a promising component of solid propellants for air-breathing engines. Its application is limited by difficulties of organizing high-efficiency combustion. Experimental investigations of combustion of individual boron particles demonstrate a large number of unique features, which are not typical for other materials: variable ignition temperature, two stages of combustion, and drastic reduction of the burning rate for particles with sizes of several micrometers or smaller. Models that cover the entire range of temperatures, concentrations, and particle sizes are physically non-obvious, can be hardly reproduced, and do not provide the accuracy needed for solving practical problems. In this paper, we propose a simple diffusion model of combustion, which ensures an adequate description of combustion of boron particles 34.5 and 44.2 μm in size at temperatures above 2240 K.

The principles of computational determination of the characteristics of both coarse and fine fractions of condensed combustion products formed at the propellant surface and evolving as part of the multiphase flow in the combustion chamber. The characteristics of condensed products at the propellant surface are determined by modeling the processes involved in their formation using a decision-making system based on an expert system approach. An evolution model for non-one-dimensional flow with the mutual influence of the characteristics of condensed and gaseous combustion products is developed to determine the characteristics of condensed products in the multiphase flow of combustion products. The developed tools allow the characteristics of condensed products to be determined depending on the propellant composition, characteristics of the propellant grain, and combustion chamber parameters. Parametric analysis of the developed models was performed for combustion chamber conditions.

Bis(4-nitrofurazan-3-yl)furazan (DNTF) and 3,4-bis(4-nitrofurazan-3-yl)furoxan (DNFF) have been studied as potential components of composite solid rocket propellants. Their heats of combustion and enthalpies of formation have been determined experimentally. X-ray diffraction analysis has shown that the DNTF and DNFF crystals are orthogonal with the same space group P212121. It has been found DNTF and DNFF are ineffective in rocket propellants with a hydrocarbon binder; however, when using DNTF for compositions without aluminum and with an active binder, it is easy to provide a specific impulse of 254.5 s at combustor and nozzle exit pressures of 40 and 1 atm, respectively, and at a density of 1.77 g/cm3, and when using DNFF, a specific impulse of 258 s at a density of 1.79 g/cm3.

This paper presents an approach to the generalization of the features of solid propellant combustion in gas generators of propulsion systems based on a one-dimensional differential model. The heat release rate function represented by mutually orthogonal Laguerre polynomials in the differential model is identified by solving the inverse problem of combustion theory and using measured gas parameters in the gas generator. This makes it possible to establish the relationship between the solid propellant combustion features and the gas generator operating modes and formulate boundary conditions for numerical investigation of the solid propellant gas generator.

A solution is obtained to the problem of determining the conditions of hydrodynamic stability in the presence of flow of gaseous combustion products over the propellant gasification surface. The cross-flow has a small gradient along the direction of motion. Analysis of the obtained dispersion equation shows that hydrodynamic instability with oscillations develops. The transfer coefficients of natural turbulence are represented as the sum of two terms: the first is responsible for the transfer in the absence of cross-flow, and the second takes into account the enhancement of transport processes in the case of cross-flow. Their dependence on the initial temperature of solid propellant combustion predicts a reduction in the negative erosion effect in accordance with experimental data. Accounting for the relaxation time of the evaporation process has a stabilizing effect. In the limit of strong relaxation, this leads to oscillatory stability (in the absence of crossflow) where the perturbations neither grow nor decay. However, arbitrarily weak cross-flow leads to instability.

Five variants of calculating the burning rate of a solid propellant as a function of the pressure in a solid-propellant rocket motor are considered. Two variants of analytical expressions are proposed for approximating real dependences. In all variants, the pressure in the rocket motor can be presented by simple analytical expressions as a function of solid propellant parameters, charging conditions, and structural factors of the charge and motor.

A model for solid propellant gasification is proposed which contains a two-phase medium in an intermediate stage. The formation of the gas phase proceeds in two ways: chemical reactions result in gaseous products, which, in turn, initiate the formation of bubbles in which vapor forms from the liquid phase of the propellant. Gaseous products play an important role only in the very early stage of bubble development; their critical pressure is used to determine the minimum size of gas-phase nuclei. The bubble volume grows primarily by evaporation of the liquid phase. A kinetic equation for the bubble concentration and the necessary boundary conditions are formulated. Arguments are given suggesting that a temperature maximum cannot occur in the gasification zone and that natural turbulence can be generated by collapsing bubbles. The sound produced by solid propellant combustion is explained by the collapse of a huge number of microscopic bubbles. If the processes in the two-phase zone are neglected, the formulated system of equations is transformed into the Belyaev–Zel’dovich model equations.

Solid propellant combustion products at a pressure of 40 atm have been studied by molecular beam mass spectrometry. A facility with a four-stage molecular beam formation system has been designed. The possibility of quantifying the composition of the combustion products is shown using as an example a composite fuel based on ammonium dinitramide and polycaprolactone.

Energetic polymers containing nitro, nitrato, and azido groups release high energy during combustion and thereby increase the performance of the systems. A number of energetic polymers have been found suitable for use as binders in high-performance propellant and explosive formulations. This review describes the synthetic and application aspects of various modern energetic polymers for explosive formulations and propellants.

A performance analysis and experimental study of a hybrid gas generator to be used in a ducted rocket are presented. Such a system exhibits potential advantages with regard to safety, performance, costs, availability of the fuel components, storability, and thrust control. A combination of a paraffin wax fuel and oxygen in the gas generator ensures a high regression rate and reveals oxidizer-to-fuel ratios as low as 0.14 in the gas generator (compared to the stoichiometric ratio of 3.4). A fuel regression rate correlation versus the oxidizer mass flux is derived, presenting a major advantage for the fuel flow rate management in comparison to control of the solid propellant gas generator burning rate through the pressure exponent, which requires mechanical interference with the hot nozzle flow to ensure a change in the combustor pressure and a corresponding change in the burning rate. Evaluation of the ducted rocket (with different oxidizers) versus pure ramjet performance shows a higher specific thrust for the former, though the latter exhibits a higher specific impulse.

This paper presents the result of a study of the chemical composition, physicotechnical properties, and structure of various types of granulated ammonium nitrate (high-density, porous, and aerated) made in Russia and abroad. It is shown that thermal treatment of high-density ammonium nitrate granules (Russian Standard (GOST) No. 2-2013) leads to changes in the crystal structure (porization) that increase the retention capacity with respect fuel oil. The detonation velocity of ammonium nitrate/fuel oil compositions based on aerated ammonium nitrate was measured.

This paper presents the new way of the occurrence of “natural” turbulence in the Gusachenko–Zarko mechanism of negative erosion effect during the propellant burning. It is shown that the propellant gasification region can generate hydrodynamic instability if its burning rate at a constant temperature depends on the pressure. The hydrodynamic instability of the propellant combustion that decompose according to the solid phase → liquid phase → gas and solid phase → gas scheme occurs under quite different conditions. The gasification region in propellants of the first type is more inclined to instability generation than that in propellants of the second type. The hydrodynamic instability occurs if the critical value of the Reynolds number, which depends on the properties of the propellant and environmental conditions, is exceeded.

The solid-state ignition of a metallized composite propellant (ammonium perchlorate + 14% butyl rubber +5% aluminum powder + 6% plasticizer) under local heating by several sources of limited power capacity (dimensions of the hot particle x p = 4 mm and y p = 2 mm) was studied by mathematical modeling. For the temperature of the heated steel particles and the distance between them varied in the ranges 700 < T p < 1500 K and 0.1x p < Δx < 1.5xp, respectively, the values of T p and Δx were determined for which the ignition delay corresponds to the initiation of combustion of the composite propellant by a single particle, by a plate at a constant temperature or by several particles. In the region of low initial temperatures of the local sources (T p < 1100 K), the limiting values Δx → 0.1x p and Δx > 1.5x p, were identified for which the characteristics and mechanism of ignition of the propellant by a group of heated particles can be studied using the “plate–propellant–gas” model and the “single particle–propellant–gas” model, respectively. Decreasing the distance Δx at T p < 1100 K decreases the induction period to 50% and reduces the minimum initial temperature of the source required to initiate propellant combustion from 830 to 700 K. At T p > 1100 K, the ignition of the metallized composite solid propellant by a single or several particles can be studied using relatively simple one-dimensional models of condensed material ignition by a plate at constant temperature. The variation in the ignition delay in this case is less than 5%.

Combustion of solid propellants in rocket propulsion systems usually occurs in an intense cross-flow of combustion products (solid rocket motor), gaseous oxidizer (hybrid rocket motor) or air (ramjet and air-breathing engines). This leads to the so-called erosive burning effects, resulting in a change in the burning law under the influence of the gas flow. The main approaches to modeling the erosive burning of solid propellants in a high-velocity cross-flow of gases are considered. Methods for the criterial description of the results of experimental studies of the erosive burning of solid propellants under transonic and supersonic flow conditions are analyzed.

Experimental data demonstrating the correlation of parameters in the power-law dependence of the burning rate of composite solid propellants on pressure are reported. The reasons for changes in the burning rate due to changes in propellant mixing conditions are discussed. The deviation of the pressure in the combustor of a solid-propellant rocket motor is analyzed with due allowance for the correlation of parameters in the burning rate law. It is shown that the relative deviation of the burning rate depends on pressure at which propellant combustion occurs. Moreover, for each propellant, there exists a pressure level at which the burning rate deviation is theoretically equal to zero, regardless of the differences in propellant compositions and properties.

A mathematical model of combustion of a composite solid propellant called ALICE (frozen suspension of nanosized aluminum in water) is presented. The model takes into account the combustion of aluminum nanoparticles in water vapor, the motion of combustion products, and the smaller velocity of particles as compared to the gas. The calculated burning rate is consistent with available experimental data on the burning rate of ALICE as a function of pressure.

Multiphase reactive systems can exhibit highly dynamic combustion phenomena that could be better understood by using recently developed high-repetition-rate optical diagnostic and imaging approaches. Here, we present an overview of recent activities using high-speed (5 kHz) OH planar laser-induced fluorescence to visualize and make measurements in several multiphase reactive systems. This technique is used to visualize the dynamically changing OH concentration in the gas phase near the surface of solids, liquids, and gels. In addition to gas-phase OH imaging, condensed phases of various solid propellants, gels, and liquids are found to fluoresce when exposed to the laser radiation centered at 283.2 nm. Simultaneous imaging of condensed phases and gasphase OH radical fluorescence has proven to be particularly useful for various measurements, and several examples are presented.

The problem of calculating the characteristics of the agglomerates formed during combustion of high-energy composite solid propellants is considered. It is shown that the mathematical models developed by the authors can be used for different propellant formulations to evaluate not only the dispersion of agglomerates, but also their quantity, chemical composition, and structure. The rules (algorithm) of using the developed models for a wide range of propellant formulations are determined. Modeling results for a number of propellant formulations based on various components are analyzed.

A method of increasing the efficiency of mixing in an air-breathing rocket engine is discussed. Three-dimensional simulations confirm the advantages of the proposed method based on using normal supply of products of incomplete combustion of a pasty propellant into the wake zone behind a bluff body in an air flow.

A possibility of ballistic efficiency of space rockets with main rocket propulsion engines by means of developing a system of gasification of unburned residues in propellant tanks with the use of gas-generating compositions with self-sustained combustion is considered. The criterion for choosing gas-generating compositions is the increment of the characteristic velocity of the rocket stage acquired due to combustion of gasified propellant residues. The results of the present study show that the proposed method of gasification of liquid propellant residues increases the energy characteristics of the rocket.

The propellant performance and the velocity of expansion of detonation products of explosives based on RDX and HMX and their mixtures with particulate Al were investigated. Measurements were performed in air at atmospheric pressure and in rarefied air. Comparison of the results shows that the detonation products of explosives, including those containing no Al additives, react with the surrounding air. The completeness of the reaction increases with increasing velocity of detonation products in air.

This paper presents a physicomathematical model and a method for calculating the parameters of the transition of a SRM to steady-state operation based on the design features of the propellant charge, the initial conditions in the motor, and the interrelated processes such as the formation and motion of compression waves and shock waves in the channel of the charge; the motion and mixing of the igniter combustion products containing a condensed phase with the initial gas and the combustion products of the charge; gradual heating and ignition of the surface of the charge. The developed programis used to explain the difference between the experimental pressure-time curves in the region of the transition of a large-size motor model to steady-state operation, obtained in bench tests under atmospheric conditions and on a high-altitude test facility at an initial motor pressure of 0.25 · 105 Pa. It is shown that ignition of the charge in vacuum produces more favorable conditions for heat transfer than ignition at atmospheric pressure. In the channel, this is due to an increase in the velocity of the igniter combustion products behind the transmitted shock wave, and in the submerged region and the charge slots, due to the minimum absorption of the radiation from the combustion products by the rarefied initial gas (air) present there.

A numerical simulation of solid-phase ignition of a composite propellant by a single small disk-shaped metal particle heated to a high temperature is performed. In the “heat flux amplitude-ignition delay ” coordinates, an region of stable initiation of combustion of a typical composite solid propellant under local heating by a source of limited energy capacity is selected. The limiting amplitudes of heat fluxes during ignition of the condensed substance under conductive and radiative heating are compared.

A thermal explosion synthesis method for production of magnesium diboride, MgB2, has been investigated, revealing the formation of a high-quality material via an exothermic reaction after heating a mixture of magnesium and boron powders to a temperature of ≈600°C. The reaction product has been characterized by different thermal and x-ray analyses. It has been also revealed that aluminum diboride, AlB2, is not formed from aluminum and boron powders in the range of temperatures applied.

The dependences of the detonation velocity and the propellant performance measured using the M-40 technique on the charge density for aluminized explosives with different mass fraction of Al were studied. The fractions of the energy of Al combustion utilized during the chemical reactions and during the acceleration of the flyer plate were estimated. Regression dependences of the detonation velocity and the propellant performance on the charge density were obtained. The effect of the addition of particulate Al, Ti, Zr, and W in an amount of 5–30% on the detonation velocity of high-density explosive charges based on plasticized RDX was investigated. It is found that the reduction in the detonation velocity with the addition of various metallic additives is determined by the longitudinal sound velocity of the additive, and not by its density. Simple formulas for calculating the detonation parameters of high-density metallized explosives were obtained.

This paper presents an analysis of the possibility of improving the ballistic performance of solid rocket propellants by using the α-modification of hexanitrohexaazaisowurtzitane (CL-20) as a component of solid crystal hydrates or crystal solvates. If small molecules (water, hydrazine, and methanol) occluded in the main material are enclosed in the structural cavity of the crystal of the main component and, consequently, the crystal lattice parameters do not change relative to the original compound, then, there is a real opportunity to improve the ballistic performance of the rocket propellants developed on the basis of the original energetic compound by increasing the density of the component.

Combustion of mixtures of a narrow fraction of ammonium perchlorate (AP) with hydrocarbon binders and combustion catalysts diethylferrocene and 1,1′-bis(dimethyloctyloxysilyl)ferrocene, as well as nano-sized Fe2O3 is studied. It is shown that the efficiency of ferrocene compounds from the viewpoint of increasing the burning rate depends on the oxidizer/fuel ratio in the propellant and on the place of the leading reaction of combustion. In composites with a high oxidizer/fuel ratio whose combustion follows the gas-phase model, the catalyst efficiency is rather low. In systems with a low oxidizer/fuel ratio where the contribution of condensed-phase reactions to the burning rate of the system is rather large, the catalyst efficiency is noticeably greater, and it is directly related to the possibility of formation of a soot skeleton during combustion. The close values of the catalytic activity of ferrocenes and Fe2O3 in the case of their small concentrations in such compositions testify that the main contribution to the increase in the propellant burning rate is made by Fe2O3 formed due to rapid oxidation of ferrocene on the AP surface and accumulated on the soot skeleton. Thermocouple measurements of propellants with a low oxidizer/fuel ratio are performed, and it is shown that the temperature of their surface is determined by plasticizer evaporation. A phenomenological model of combustion of the examined propellants is proposed.

Optical methods in infrared (IR) and visible (VIS) ranges for detection of combustion zone propagation in solid high-energy materials, such as pyrotechnic compositions inserted into pyrolytic graphite (pyrographite) tubes and ignited at one end by a CO2 laser beam are presented. The pyrographite tube is used as a thermal management transducer enabling detection of combustion zone movement because of unique thermal conductivity anisotropy of pyrographite resulting in low thermal conductivity of the tube along its axis and high thermal conductivity along the tube radius. In the first method, an IR thermal camera is applied for detection of heat zone movement induced on the external side surface of the pyrographite tube by the combustion zone travelling inside the tube. According to the second method, a VIS camera and a thermochromic layer covering the external side surface of the pyrographite tube are used for visualization of heat zone movement registered as the color change boundary traveling along the thermochromic layer. The change in color of this layer is caused by its thermochromic substance response to heat delivered by the heat zone. As thermochromic substances, leuko dyes or chiral-nematic liquid crystals are used. These methods seem to be particularly promising for continuous measurements of burning rates of solid high-energy materials, such as solid rocket propellants and pyrotechnic compositions.

This paper discusses the results of experiments on the underwater combustion of ballistite propellant in a centrifugal force field where the gasification front moves in the direction of acceleration. The necessary conditions for the combustion were provided using a mobile localizer of the combustion zone, which was a shell of a heat-resistant material put on the test sample of the propellant.

The brightness temperature and pressure profiles of the detonation products of pressed charges of benzotrifuroxane were determined by a pyrometric method, and the heat of explosion and propellant performance were experimentally determined. The temperature of the detonation products (4100±150 K) was significantly lower than the calculated values reported in most theoretical papers. Compared to HMX, benzotrifuroxane has a higher heat of explosion but lower expansion velocity of the shell (T-20) method and Gurney energy.