Combustion Products Of Propellants Research Articles

COMBUSTION OF CONDENSED COMBUSTION PRODUCTS OF BORON-CONTAINING SOLID PROPELLANTS IN AIR

The analysis of combustion models of the main components of condensed combustion products (CCP) of boron-containing solid propellants (boron, boron carbide, boron nitride, and carbon) is carried out. Using the available experimental data, the parameters of the combustion model of boron carbide and boron nitride particles are determined. Using the combustion models of boron, B<sub>4</sub>C, BN, and carbon particles, parametric calculations of the combustion of both individual indicated particles and their mixtures, which simulates the composition of the CCP of boron-containing solid propellant, have been carried out. A comparative analysis of the combustion of particles of boron, B<sub>4</sub>C, BN, and carbon in proportions characteristic of the CCP of boron-containing solid propellants showed that combustion of BN has the least effect on the completeness of combustion of CCP. This is due to its low calorific value and low content in CCP. Ceteris paribus, carbon particles burn the fastest, but their contribution to the completeness of CCP combustion is less than that of boron and boron carbide due to their lower calorific value.

Theoretical analysis on applying steam reforming to the primary combustion of the boron-based fuel-rich propellant

The primary combustion products of boron-based fuel-rich propellant contain a significant amount of condensed boron, which poses challenges for complete burnout of the primary combustion products during the secondary combustion process. To enhance the ignition and combustion characteristics of these primary combustion products, a pioneering theoretical analysis on applying steam reforming to the primary combustion of the boron-based fuel-rich propellant was conducted. The results demonstrate that steam reforming improves the ignition and combustion characteristics of the primary combustion products through three main aspects: optimizing the composition of combustion products, improving energy distribution between the primary combustion and secondary combustion process, and elevating combustion temperature, respectively. It theoretically confirms the feasibility of this innovative combustion organization method for solid rocket scramjet. Furthermore, it has been found that there is an optimal steam-fuel ratio (the value is 1.0 for B35), which facilitates the conversion of combustible components in the primary combustion products from a gas–solid state to a pure gaseous phase. Simultaneously, the temperature of combustion products after steam reforming reaches the highest. Therefore, theoretically speaking, the steam-fuel ratio of solid rocket scramjet with steam reforming should not exceed the optimal steam-fuel ratio to ensure better ignition and combustion characteristics of the combustion products after steam reforming. This analysis provides valuable theoretical support for the design of solid rocket scramjets with steam reforming.

Improving the combustion efficiency and agglomeration of aluminum-water propellants via n-Al/CuO metastable intermolecular composites

In this work, we investigated the combustion and agglomeration characteristics of aluminum-water propellants by replacing the original Al with binary n-Al/CuO metastable intermolecular composites (MICs). Through laser ignition tests and thermogravimetric-differential scanning calorimetry, we evaluated the oxidation reactivity, ignition delay, burning rate, agglomeration properties, and condensed combustion products of aluminum-water propellants containing different CuO loadings. Compared with conventional aluminum-water propellants, the introduction of binary MICs is found to lower the initial temperature of Al and improve its oxidation activity. Both the burning and heating rates scale linearly with the CuO loading, increasing by a factor of 5 and 6, respectively, when the CuO loading reaches 5 wt.%. The combustion efficiency of the modified propellants is found to improve by 5−12 %. The mean size of the condensed combustion products drops from over 400 μm to around 200 μm due to MICs addition, indicating weakened agglomeration. The ignition delay time is slightly shortened, and the combustion intensity first increases but then decreases as the CuO loading increases. In summary, this work indicates that replacing Al with binary Al/CuO-MICs can significantly alter the combustion and agglomeration properties of aluminum-water propellants. The experimental data and insight from this work could help guide the development of advanced aluminum-water propellants for various propulsion and energy applications.

Elaborative collection of condensed combustion products of solid propellants: Towards a real Solid Rocket Motor (SRM) operational environment

A novel constant-pressure and constant-quenching distance Condensed Combustion Products (CCPs) collection system was developed, coupled with a timing control system, to collect the CCPs formed in the course of burning of aluminum-based composite propellants. The effects of adiabatic graphite plating, collection zone, quenching distance, time series of collection, and propellant burning rate on the microscopic morphology, particle size distribution and unburned aluminum content of CCPs were investigated. It was verified that the graphite plating can provide a high-fidelity high-temperature environment for propellant combustion. The combustion efficiency is improved by 2.44% compared to the bare propellant case. The time series of collection has a significant effect on the combustion efficiency of aluminum, and the combustion efficiency of aluminum in the thermal state (1.2–2.4 s) is 2.75% higher than that in the cold state (0–1.2 s). Similarly, the characteristics of the CCPs in different collection zones are different. At the quenching distance of 5 mm, the combustion efficiency of aluminum in the core zone (85.39%) is much lower than that in the outer zone (92.07%), while the particle size of the CCPs in the core zone (172 μm) is larger than that in the outer zone (41 μm). This indicates that the core zone is more likely to produce large-sized and incompletely burned agglomerates during the propellant combustion process. Different burning rates also lead to a significant difference in particle size distribution and combustion efficiency. High burning rates result in higher combustion efficiency. A detailed sequence of the elaborative collection process of CCPs is proposed, mainly including the setting of ignition delay time, burning rate, working pressure, plating length and time series of collection. The findings of this study are expected to provide a reliable tool for the evaluation of the combustion efficiency of solid propellants.

Morphology and Size Distribution of Condensed Combustion Products of Aluminum-Based Propellants

ABSTRACT Both glycidyl azide polymer (GAP) and hexanitrohexaazaisowurtzitane (CL-20) are high-energy compounds with very promising applications. However, there are relatively few relevant studies for GAP/CL-20 aluminum-based propellants. In this paper, the CCPs of aluminum-based GAP/CL-20 propellant under different operating conditions were collected by a gas collecting device, and the physicochemical properties of CCPs were carefully studied by various modern analytical and testing instruments. The results show that the CCPs consist mainly of smoke oxide particles (~1 μm), residual oxide particles (~20 μm) and agglomerates. Residual oxide particles are widely available, but have received less attention from researchers. The agglomerates can be divided into two categories: sphere-like agglomerates and non-spherical agglomerates. Sphere-like agglomerates include cracked spheres, solid spheres, hollow spheres, and spheres with oxide caps. Surprisingly, we found an agglomerate with two caps. Non-spherical agglomerates are larger in size than sphere-like agglomerates, and the non-spherical agglomerates include irregular agglomerates, coral-like aggregates, and filamentous agglomerates. As far as we know, coral-like aggregates and filamentous agglomerates in CCPs have not been reported in the literature. Pocket theory and inter-pocket bridge theory were used to explain the formation mechanism of fine filament agglomerates. In addition, the formation mechanisms of different types of agglomerates were analyzed and sorted out. The particle size distribution of CCPs at different pressures shows that the average particle size of CCPs decreases with increasing pressure. It is hoped that this paper can provide a reference for the design of high-energy propellant formulations.

Application Energetic Materials for Solid Composite Propellant to Support Defense Rocket Development

In its application in space technology, solid composite propellants are often used as fuel in rockets for military purposes. Increasing the energy of the propellant is carried out by observing two stages, the use of energetic materials and improvements to the process technology. The current development of propellant technology makes it possible to use new energetic materials, simple formulations, high energy, and smokeless. The purpose of this research is to find out developments related to the use of highly energetic materials as raw materials for composite propellants for defense rockets at the Rocket Technology Research Center, ORPA-BRIN. This study uses qualitative analysis methods with research designs in the literature studies and simulation results. In the context of mastering rocket propellant technology in Indonesia, the application of highly energetic materials is expected to be able to solve the problem of rocket propulsion performance. Currently, the Rocket Technology Research Center, ORPA-BRIN is developing a smokeless propellant composite with a composition based on the energetic materials AP/HTPB/Al and an oxidizing agent RDX. From the results of the combustion simulation software ProPEP and RPA, it shows that the composition of the resulting combustion gaseous (Al2O3 and HCl) shows a decrease when using RDX energetic material-based propellant. It's known that RDX can significantly reduce smoke in propellant combustion products. The application of the new highly energetic materials compound is expected to significantly solve the problem of solid rocket propulsion performance.

Drag and heat transfer of metal and oxide agglomerates in flow of combustion products of solid propellant

Modeling and simulation of the motion of agglomerates with a complex composition and shape in the flow of combustion products is of interest for problems related to design and optimization of solid propellant rocket motors. An approach is developed for modeling of unsteady processes that occur when a particle consisting of an aluminum droplet and a condensed oxide particle attached to it (aluminum droplet with oxide cap) is transported by a viscous incompressible fluid flow. Methods are proposed to take into account the sliding of an attached oxide particle over the surface of the aluminum droplet. The results of numerical computations are used to find the drag and heat transfer coefficients of a non-spherical particle at flow velocities corresponding to the formation of separation zones. The constructed model is one of the components of a more general mathematical model designed to calculate two-phase flows of combustion products based on a multi-scale approach to the simulation of two-phase flows in a complex environment.

Combustion and agglomeration characteristics of aluminized propellants containing Al/CuO/PVDF metastable intermolecular composites: A highly adjustable functional catalyst

In this experimental study, we explore how the combustion and agglomeration characteristics of solid propellants can be modified by replacing aluminum with ternary Al/CuO/PVDF(Polyvinylidene Fluoride) metastable intermolecular composites (MICs). Using thermogravimetric−differential scanning calorimetry, a laser ignition setup, and a high-pressure propellant combustion system, we examine the thermal reactivity, ignition behavior, burning rate, agglomeration, and condensed combustion products of various MICs-inclusion propellants. We find that the use of ternary Al/CuO/PVDF MICs can decrease the onset temperature and increase the oxidation efficiency of Al, as compared with common binary MICs. The ratio of CuO/PVDF is found to have a strong effect on the combustion intensity of the metallic powders, but only a weak effect on their ignition delay. Higher proportion of PVDF is found to have a negative effect on the combustion rate. Specifically, the combustion intensity is found to decrease first and then increase as the PVDF content increases. The exothermic Al/CuO reaction is found to alter the heat transfer of the condensed layer on the propellant surface, while PVDF plays a critical role in the thermal feedback of the gaseous reaction. When combined, these two factors lead to a ± 30% variation in the propellant burning rate. Furthermore, ternary MICs are found to reduce agglomeration during propellant combustion. The two optimal formulations are 2.5 wt.% CuO/2.5 wt.% PVDF and 3.5 wt.% CuO/1.5 wt.% PVDF. The use of these formulations is found to decrease the mean particle size of the condensed combustion products to only 11 μm, as compared with 76 μm for the original Al-inclusion propellant. High-speed microscope images reveal the existence of crack phenomena in the pockets surrounding the AP particles on the burning surface due to the Al/CuO reaction, as well as the floccule breakup mechanism of Al aggregates in the presence of PVDF, both of which reduce agglomeration. Overall, this study shows that replacing Al with ternary Al/CuO/PVDF MICs has significant effects on the agglomeration, combustion and ignition features of aluminized propellants. The findings of this study can be used to aid the development of highly adjustable functional catalysts for solid propellants.

Сравнительный анализ горения частиц бора, карбида бора, нитрида бора и углерода в воздухе

Combustion models of main components of condensed combustion products (CCP) of boron-containing solid propellants: boron, boron carbide, boron nitride, and carbon are analyzed. Using the known experimental data, the parameters of the combustion model of boron carbide and boron nitride particles are determined. Using the models of combustion of boron, B4C, BN and carbon particles, parametric calculations of the combustion of both individual abovementioned particles and their mixture, which simulates the composition of the CCP of boron-containing solid propellant, have been carried out. Calculations have shown that combustion of BN has the least effect on the completeness of CCP combustion. This is due to its low calorific value and low content in CCP. Ceteris paribus, carbon particles burn the fastest; however, their contribution to the completeness of CCP combustion is less than that of boron and boron carbide due to their lower calorific value.

Unique thermal and combustion behaviors of composite propellants containing a high-energy insensitive nitropyrimidine derivative

Owing to the novel insensitive high-energy-density material 2,4,6-triamino-5-nitropyrimidine-1,3-dioxide (ICM-102) has lower sensitivity and higher energy(see in Table S1, Supporting Information), it has great advantages in the application of low-vulnerability propellants. In this paper, the evaluation of ICM-102-containing composite propellants has been carried out, where the decomposition and combustion performances of these propellants have been comprehensively investigated. It has been shown that the main exothermic decomposition peak temperature can be increased from 235 ºC to 350 ºC, when 6 to 10 wt% of ICM-102 was used to replace RDX, but heat release rate was largely increased. Moreover, the burning rate of ICM-102 containing propellants at 3 MPa is in the range of 6.97–7.75 mm•s − 1, depending on its content. The pressure exponent (0.5–3 MPa) is between 0.34–0.44. Compared with the typical propellants with RDX, the burn rate and the flame temperature of ICM-102- containing propellants are higher. In addition, the particle size distributions of condensed combustion products of ICM-102 propellants are narrower and smaller than that of RDX-based one. It also shows that extra products such as aluminum carbide (Al20C) and aluminum oxide nitride (Al5O6N) were produced once ICM-102 was used.

Composition of solid and gaseous primary combustion products of boron-based fuel-rich propellant

The investigation of the primary combustion products of boron-based fuel-rich propellants is conducive to understanding the mechanism of the two-stage combustion process and provides important information for the prediction and organization of secondary combustion processes. The present investigation methods include numerical simulation and experimental analysis. The former deviates considerably from the actual conditions. The latter is currently a reliable and practical method, although it is difficult to quantify when the components are complex. In this work, SEM, LPSA, XRD, XPS, IC, ICP–MS, AAS, EDS, and chemical analysis methods were employed to qualitatively and quantitatively analyze the 3 solid products (aggregations, condensed-phase products, and flocculent products) at 1, 5, and 10 bar. GC, GC–MS, and IC combined with chemical tests were adopted to investigate the gaseous products (including chlorine-containing gas) for the first time. The solid products primarily included B2O3, B(HO)3, unburned boron, KB5O8(H2O)4, KCl, C, NH4Cl, NH4Mg(H2O)6Cl3, B13C2, Fe3O4, Mo, MgO, Mg(OH)2, MoO3, C3N4, BN, MgCO3(H2O)3 (or MgCO3), MoO2, Fe4(Fe(CN)6)3, and FeB (or BFe4), and the first 7 were dominant. The reaction degree of boron was less than 50 wt% (increasing with increasing pressure), indicating the significance of the secondary combustion process. The reacted boron was likely to react with O and C but not with N. The generated B2O3 vapor easily combined with H2O to form B(HO)3 or reacted with KCl (from the thermal decomposition of KP) and H2O to produce KB5O8(H2O)4 after cooling. There was less B13C2 (its content increased with increasing pressure, resulting in increases in particle size), since it was easily oxidized to generate B2O3. NH4Cl was produced by the oxidation-reduction reaction between the NH3 and HClO4 released from large-sized AP, and some of it crystallized with MgCl2 and H2O to form NH4Mg(H2O)6Cl3. The gaseous products primarily included N2, H2, CO, CH4, CO2, C2H4, C2H2, HCl, C4H6, C3H6, Cl2, C6H6, and other hydrocarbons. N2 and H2 had the highest content. The N2 content increased with increasing pressure, while the content of the combustible gases decreased, indicating that high pressure promotes the propellant reaction. The results also show that high pressure was conducive to the low-temperature decomposition of AP (producing more HCl and Cl2) and the reaction of boron and hydrocarbons (generating more B13C2), while it had an adverse effect on the C reaction (producing less COx and mostly CO). Nevertheless, the increasing pressure had little effect on the combustion product components but did affect the content within the pressure range.

Modelling of the formulated solid rocket propellant characteristics

This study is a mathematical model to obtain the characteristics performance of magnesium metal (powder) and carbon on a potassium nitrate-sucrose (KNSU) solid propellant formulation. Characterization of propellant is, as a general rule, important to determine its performance before it can be suitable for use for a rocket flight or any mission. Method of ballistic load cell evaluation was used to validate results and a mathematical model using the combustion exhaust products was solved to obtain the characteristics performance parameters of the propellant. The carbon constituent which acts as an opacifier and coolant was kept constant at 2% in order to arrest some of the heat during the combustion process and helped to lower the combustion temperature, because high combustion temperature could lead to combustion chamber rupture or failure. The effect of addition of magnesium which was optimized for 3% in the formulation contributed significantly in improving the overall performance of the propellant. The utilization of magnesium in KNSU propellant provided higher values parameters and better performance compared to when not included. This was confirmed with the model equations. The propellant combustion products equation was used to model and obtain the characteristics performance parameters. This gave propellant specific impulse (122.9s), combustion temperature (1821K), heat ratio (1.1592), molecular weight (36.89g/mole), propellant density (1912.5kg/m3) and characteristics velocity (1000m/s) result while maintaining the same chamber pressure.

Chemiluminescence of electronically excited species during the self-ignition of acetylene behind reflected shock waves

Acetylene is an effective additive to various rocket fuels. Acetam, an equilibrium highly concentrated solution of acetylene in liquid ammonia, is a promising fuel for use in liquid−propellant jet engines. To gain insights into the processes occurring in such systems and provide means for predicting their characteristics, the self-ignition of model acetylene−oxygen−argon mixtures is experimentally and theoretically studied. The experiments are performed behind reflected shock waves in a temperature range from 1100 to 1870 K at a total gas concentration of [M] ≈ 1.0 × 10−5 mol/cm3, with chemiluminescence signals from electronically excited C2* (λ ≅ 516 nm), CH* (429 nm), CO2*(363 nm), and OH* (308 nm) recorded. Numerical simulations within the framework of various detailed kinetic mechanisms closely reproduce the measured temperature dependences of the ignition delay time for the tested mixtures, but in some cases fail to satisfactorily describe the time profiles of the emission signals from C2*, CH*, CO2*, and OH*. The performed numerical simulations make it possible to explain the observed characteristic features of the emission profiles of electronically excited C2*, CH*, CO2*, and OH* during the self-ignition of C2H2–O2−Ar mixtures. Sensitivity analysis demonstrates that the existence and characteristics of the preignition stage in the CO2* and OH* emission signals are largely determined by the kinetic behavior of intermediate species, such as the ketene molecule and CH2 radical. Based on the results obtained, a new detailed kinetic mechanism has been proposed for describing processes in the combustion products of acetylene-containing propellants, in particular chemiluminescent glow, which are of considerable interest for the development of new technologies for astronautics and aeronautics.

Drag and heat transfer of metal and oxide agglomerates in flow of combustion products of solid propellant

Drag and heat transfer of metal and oxide agglomerates in flow of combustion products of solid propellant

Experimental investigation on the condensed combustion products of aluminized GAP-based propellants

The condensed combustion products (CCPs) of solid propellants significantly affect the combustion and internal flow inside the solid rocket motor. A constant-pressure quench vessel was used to collect the condensed combustion products of an aluminized glycidyl azide polymer (GAP)-based propellant. The effects of freezing medium, quench distance, chamber pressure (5∼9 MPa) and virgin aluminum size (13∼40 μm) on the physicochemical properties of the condensed combustion products were studied. The typical size distribution of the condensed combustion products is in three modes, 1∼2 μm, 20∼30 μm, ∼300 μm and their sizes vary from 0.3 to 600 μm, corresponding to smoke oxide particles and agglomerates. Freezing medium is found to have little impact on the particle size, while the agglomeration ratio in water is larger than nitrogen and argon. In contrast to the freezing medium, pressure does affect the size of CCPs, but has little effect on the agglomeration ratio. Quench distance has no significant effect on the particle size. Agglomeration is found to increase first and then decreases with the increase of virgin aluminum size. The influence mechanism of freezing medium and virgin aluminum size on CCPs is proposed. The combustion-agglomeration map is particularly obtained, which shows that freezing medium and virgin aluminum particle size have more profound influence on the agglomeration than quench distance and pressure. The low-pressure condition presents poorest performance with high fraction of agglomerates, large agglomerate size and low combustion efficiency. Results of this work are expected to provide better insight in the combustion of solid propellant and solid rocket motor.

Анализ существующих воспламенительных устройств и разработка перспективной конструкции воспламенителя стартового ускорителя современных ракетно-прямоточных двигателей на твердом ракетном топливе

The article briefly reviews the existing designs and technical solutions for solid propellant igniters in rocket engines. The technical and design solutions in the development of solid-propellant rocket engine igniters are analyzed. The results of the development of a promising igniter design with gunpowder filler axially located in the internal channel of a propellant grain of a launching free-flowing booster of a rocket-ramjet engine are presented. Ensuring the required level of structure reliability and durability in the launch mode of the solid propellant rocket-ramjet engine, the igniter will improve the engine traction characteristics due to the combustion of the igniter aluminum shell as an additional energy source and using its combustion products in the stream of solid propellant combustion products.

Dynamics of the gas-dust cloud observed in the upper atmosphere on October 26, 2017

The results of the study of the gas-dust cloud of solid fuel combustion products dynamics formed under a rocket stages separation are presented. The sequence of the digital images obtained in the twilight condition at a distance of ∼1500 km from the missile trajectory was analyzed. The height of the cloud center reaches up to ∼900 km, and its diameter is ∼2000 km. The glow of the cloud is determined by the scattering of sunlight on dispersed particles of the combustion products. The obtained data allow to specify the dispersed composition of the cloud and dynamic parameters of its development. The proposed methodology can be used to clarify the physical picture of the interaction of the combustion products of solid propellant with the environment, as well as environmental conditions and processes in the upper atmosphere.

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

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

Optimization of artillery projectiles base drag reduction using hot base flow

The CFD numerical simulations were carried out to investigate the base drag characteristics of a projectile with base bleed unit with a central jet. Different base bleed grain types with different combustion temperatures were used. The goal was to find a way to effectively control the base flow for base drag reduction and optimisate the latter using an adequate CFD software. Axisymmetric, compressible, mass-averaged Navier-Stokes equations are solved using the k-? SST, transition k-kl-?, and RSM turbulence models. The various base flow characteristics are obtained by the change in the non-dimensionalized injection impulse. The results obtained through the present study show that there is an optimum bleed condition for all base bleed grains tested. That optimum is dependent on the temperature of the grain combustion products. The optimum reduces the total drag for 6,9% in the case of air injection at temperature of 300 K and reaches up to 28% in the case of propellant combustion products injection at almost 2500 K. Besides, the increasing of molecular weight has a role no less important than temperature of the combustion products in terms of base drag reduction.

Application of the DRGEP and Engagement Methods to Reduce the Reaction Mechanisms for Calculation of Chemically Non-Equilibrium Flows in Aircraft Engine Nozzles

The procedure of the reaction mechanism reduction including the DRGEP and engagement methods is considered and applied to the problem of calculating the non-equilibrium flows of propellant combustion products (kerosene + oxygen) in nozzles.