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The microwave electrothermal thruster (MET) utilizes wave-exited microdischarges to heat gas flows, enhancing the specific impulse of the thruster. Our computational study investigates a 17.5 GHz helium-propellant MET, employing a two-dimensional, axisymmetric fluid model of plasma coupled with electromagnetic wave and gas flows. The discharges operate in the glow regime, remaining weakly ionized, and in thermal non-equilibrium. The plasma densities reach approximately 10 20 m − 3 , and the gas temperature is around 2000 K. Even a slight off-resonant frequency operation results in a significantly lower plasma density and gas temperature. Gas heating, primarily driven by electromagnetic Joule heating, plays a critical role in influencing the overall discharge behavior. The measured peak thrust and specific impulse are 8.24 mN and 292 s, respectively, at a mass flow rate of 3.2 mg/s with 30 W of power. Compared to a cold gas thruster, there is a significant increase in the specific impulse by a factor of approximately 1.7. The enhanced performance trades off with propulsive efficiency, which decreases by a factor of 1.5 from the peak 65% at 10 W. This is due to higher energy losses to cavity walls from heat conduction with increased power. These findings underscore the critical balance between the input power and mass flow rate to improve the MET performance, highlighting the importance of power management to maximize thrust and efficiency. Furthermore, the predicted thrust and specific impulse agree well with experimental values for nominally similar MET thruster studies in the literature. [ABSTRACT FROM AUTHOR]

The field of multifunctional energetics encompasses a range of materials including propellants, explosives, and pyrotechnics that possess the ability to be manipulated through various characteristics that can be switched between go and no-go, or those that have controllable energy release levels or have additional functions beyond energetic output. The development of multifunctional energetics harnessing electromechanical or piezoelectric properties of polymeric materials or binder systems has garnered increasing interest in recent years. Among polymers, fluoropolymers such as poly(vinylidene fluoride) (PVDF) and copolymers such as poly(vinylidene fluoride-trifluoroethylene) [P(VDF-TrFE)], which are used as the binder and oxidizer in the energetic formulations, have demonstrated the highest piezoelectric coefficients. In this study, we fabricated piezo-energetic composite films using aluminum nanopowders (10 wt. % active content) as fuel and P(VDF-TrFE) (70/30) as an oxidizer and investigated the piezoelectric response using a small-scale drop weight setup. Additionally, we employed a shaker setup to investigate the response of the films to vibrations. Our findings demonstrate that these piezo-energetic films can replicate the behavior of a commercial PVDF gauge at relatively low-pressure impacts, indicating their potential use as shock or pressure sensors in various fields, as well as an accelerometer gauge. Additionally, aging studies of up to one year indicated minimal loss in the energetic content of the created films, enabling the use of energetic gauges for an extended period. Our findings support the effectiveness of piezo-energetic composite films as pressure sensors or accelerometers and highlight their potential for energetic applications. [ABSTRACT FROM AUTHOR]

The reactivity of heterogeneous energetic materials (HEMs) intimately depends on the underlying microstructural effects. For reactive materials, key factors include the microstructure distribution, morphology, size scale of heterogeneities, reactant mixing, and chemical kinetics of the reactants. We report the development of a mesoscale model for simulating the evolutions of the hotspot field and associated reaction processes when such materials are exposed to external excitations. The model explicitly accounts for microstructure, interdiffusion between the reactant species, advection of the species mixture, and chemical kinetics of the reaction. An Arrhenius relation is used to capture the rate of reactive heat release. The particular material analyzed is a composite of poly(vinylidene fluoride-co-trifluoroethylene) and nanoaluminum [or P(VDF-TrFE)/nAl]. The excitation leading to the initial microstructural temperature increase that kicks off the exothermic reactive processes is the dissipative heating arising from dielectric breakdown under the electric field developed through piezoelectricity and flexoelectricity of P(VDF-TrFE). As such, the model resolves both the breakdown process and the diffusion, advection, and exothermic reaction processes. The evolutions of the temperature and species distribution fields under the combined effects of breakdown and chemistry are used to predict the effects of microstructure, diffusion, and kinetics on several key metrics characterizing the reactive responses of the material. This mesoscale framework admits the quantification of uncertainties in these predicted macroscopic behavior measures due to microstructure heterogeneity fluctuations through the use of multiple, random but statistically equivalent microstructure instantiations. Although the particular hotspot inducing mechanism considered is dielectric breakdown here, the framework can be adapted to analyze reaction initiation and propagation and establish microstructure–reaction behavior relations under other types of hotspot inducing mechanisms, such as thermomechanical inelastic dissipation, frictional heating, and laser or microwave excitation. [ABSTRACT FROM AUTHOR]

Direct thrust measurements are performed on a circular waveguide electron cyclotron resonance (ECR) thruster working at 5.8 GHz using a pendulum thrust balance with mechanically amplified displacement. Thrust levels between 1 and 3.5 mN are found for power levels in the range of 60–350 W and xenon flow rate between 2 and 8 SCCM. A maximum thrust efficiency of 3.5% is reached at 2 SCCM and 60 W. Plasma plume diagnostics are used to estimate the thruster partial efficiencies to understand the main losses, and to perform a comparative analysis between directly and indirectly measured thrust. Results suggest that the low energy conversion efficiency and propellant utilization efficiency (< 6.4 % and < 53%, respectively) are the key factors spoiling the ECR thruster performance. Finally, retarding potential analyzer measurements show the presence of energetic electrons with energy tail up to about 300 eV. [ABSTRACT FROM AUTHOR]

The topology of the applied magnetic field is an important design aspect of Hall thrusters. For modern Hall thrusters, the magnetic field topology most often features curved lines with a concave (negative) curvature upstream of the field's peak and a convex (positive) curvature downstream. Additionally, the advent of the magnetic shielding technique has resulted in Hall thruster designs with non-conventional field topologies that exhibit high degrees of concavity upstream of the field's peak. In this article, a detailed study is carried out on the effects that the magnetic field curvature has on the plasma discharge in a 2D configuration representative of a Hall thruster's radial–azimuthal cross section. The analyses are performed for discharges of three propellants of high applied interest: xenon, krypton, and argon. For each propellant, high-fidelity electrostatic reduced-order particle-in-cell (PIC) simulations are performed with various degrees of positive and negative curvatures of the magnetic field. Corresponding 1D radial PIC simulations are also performed for xenon to compare the observations against the 2D results. Most notably, it is observed that the instability spectra in the positive-curvature cases are mostly dominated by electron cyclotron drift instability, whereas the modified two stream instability is dominant in the negative-curvature cases. The distributions of electron and ion temperatures, in particular, as well as the contribution of various mechanisms to electrons' cross-field transport show notable variations between the positive and negative curvature values. Finally, the field curvature is observed to majorly influence the ion beam divergence along the radial and azimuthal coordinates. [ABSTRACT FROM AUTHOR]

The behavior of four algebraic closure models for anomalous electron transport is investigated using a fluid Hall thruster code. The models, which were selected because they have been previously described in the literature, are calibrated against a baseline experimental condition of a 9-kW-class magnetically shielded Hall thruster operating at 300 V and 15 A on xenon propellant. The extensibility of the models is then assessed by using this calibrated model to simulate three additional operating conditions—300 V and 30 A, 600 V and 15 A, and 300 V and 15 A operating on krypton propellant. The quality of the model prediction is quantified by comparing the model outputs to experimental measurements of discharge current, thrust, and ion velocity. It is found that while none of the models can predict the ion acceleration characteristics accurately, some compare favorably in terms of the scaling of thrust and discharge current across operating conditions. The limitations of the models are attributed to the coupling between the functional scaling of the closure models with respect to the local plasma properties and the fluid model. The role of the electron energy balance in this coupling is also highlighted. These results are discussed in the context of motivating improved closure models of the anomalous electron transport in Hall thrusters. [ABSTRACT FROM AUTHOR]

This Perspective paper deals with an overview of particle-in-cell/Monte Carlo collision models applied to different plasma-propulsion configurations and scenarios, from electrostatic (E × B and pulsed arc) devices to electromagnetic (RF inductive, helicon, electron cyclotron resonance) thrusters, as well as plasma plumes and their interaction with the satellite. The most important items related to the modeling of plasma–wall interaction are also presented. Finally, the paper reports new progress in the particle-in-cell computational methodology, in particular, regarding accelerating computational techniques for multi-dimensional simulations and plasma chemistry Monte Carlo modules for molecular and alternative propellants. [ABSTRACT FROM AUTHOR]

A model predicting the number of emission sites and total current from a porous conical electrospray emitter as functions of voltage is derived. A pressure balance between capillary and electric forces is used to determine an onset criterion for individual menisci, and an ionic emission scaling law is invoked to predict the current each meniscus emits. These submodels are integrated over a phenomenological meniscus size distribution and the area of the emitter to yield a model for emitter performance as a function of five free parameters, two for the ionic emission submodel and three for the meniscus size distribution. Bayesian inference is applied to determine these model parameters from an existing dataset [Dressler et al., J. Propul. Power 38, 809 (2022)]. The model predictions after training are compared to the experimental data, and it is found that the majority of the data are within a 90% credible interval. The ability of the model to capture key trends in the experimental data is attributed to the interplay of two effects: the distribution over meniscus size on the emitter and the position-dependent electric field. The calibrated model results also suggest that the emitter surface is wetted by a series of large but sparsely distributed pools of propellant. The performance and extensibility of the model are examined within the context of model-based design for porous electrospray array thrusters. [ABSTRACT FROM AUTHOR]

The radio frequency ion thruster (RIT) is an electric propulsion device that utilizes radio frequency (RF) power to ionize propellants; thus, its performance is influenced by the frequency and matching network of the RF. In this study, an L&T type wide-range matching network was employed to experimentally investigate the performance of a miniature RIT with 4 cm diameter at different frequencies. The experimental results show that changing the matching parameters of the L&T network at the same frequency leads to changes in the effective quality factor and varying performances of the thruster; under similar effective quality factors, changing the frequency does not significantly affect the performance of the thruster; however, the plasma resistance exhibits frequency dependence. As the frequency increases, the effective collision frequency increases, leading to an increase in plasma resistance, which is beneficial for improving the performance of the thruster at high frequencies. Finally, a set of methods for quickly comparing the performance of RITs was summarized based on the effective quality factor. These research results are also instructive for improving the ionization of induction-coupled discharge devices of other sizes. [ABSTRACT FROM AUTHOR]

Due to the orderly arrangement of tubular propellant, the permeability of combustion gases is improved, which is beneficial for enhancing the safety of the combustion system. However, current internal ballistic gas-solid flow calculation methods adopt a quasi-fluid assumption, which cannot accurately account for the characteristics of long tube shapes. Additionally, tubular propellants exhibit both overall movement and parameter distribution characteristics, necessitating the decoupling of gas and solid phases. These two deficiencies in previous studies have limited the effectiveness of gas-solid flow simulations for tubular propellant. This paper proposes a numerical calculation model suitable for tubular propellant charging based on the particle element method for internal ballistic two-phase flow. Firstly, considering the overall movement characteristics of tubular propellants, the concept of blank particle elements is introduced to represent pure gas phase regions. Then, based on computational requirements, the tubular propellants are divided to form the lumped element method and the multiple-element method. The moving boundary method is used to calculate the movement process of the propellant bed particle group and is compared with experimental results to verify the applicability of the two methods in tubular propellant beds. Analysis results show that the particle element method can effectively capture changes in the flow field inside the chamber and the position of tubular propellants. The lumped element method can quickly obtain the flow field distribution characteristics inside the chamber, while the multiple-element method can capture parameter distribution characteristics at different positions of the tubular propellants while ensuring overall movement. [ABSTRACT FROM AUTHOR]

Coaxial pulsed plasma thrusters are well suited for microsatellites, but the discharge therein is radially nonuniform, resulting in uneven propellant ablation that ultimately degrades the thruster performance. This study investigates how the propellant surface evolves during 50 000 discharges in a coaxial pulsed plasma thruster with different outer electrode inner diameters and initial voltages, encompassing the propellant surface profile, concave depth of the propellant, and carbon deposition on the propellant surface. The results show that the propellant surface changes in real time throughout the 50 000 discharges. Increasing the outer electrode inner diameter leads to a proportional increase in the concave depth of the propellant, resulting in unstable ablation, and the increased inward concave depth increases the likelihood of discharge failure. Furthermore, increasing the initial voltage increases the energy density in a unit area of propellant, further contributing to the increase in the concave depth and discharge failure. The energy density distribution on the propellant surface is obtained via simulation, and by comparing the results with experimental data, the critical ablation threshold for the propellant is estimated as being ca. 0.7 J/mm2. • The coaxial PPT discharged 50,000 times in 7 different working conditions. • The influence of initial voltage on the propellant evolution is analyzed. • The influence of electrode size on the propellant evolution is analyzed. • High initial voltage and large electrode size easily lead to discharge failure. [ABSTRACT FROM AUTHOR]

High Test Peroxide (HTP) is a well-known green and energetic material, widely used in propulsion systems in a variety of applications, including hypergolic systems. This experimental work mixed fumed silica with liquid HTP to synthesize Gelled HTP (GHTP). A high-speed camera and a time-resolved infrared camera were employed to analyze liquid/gel phase mixing, ignition, and flame spread through drop tests, adopting a promising low toxicity liquid fuel as a GHTP hypergolic pair. Liquid/gel impinging jets in ambient air were also performed, simulating a basic element of a hybrid liquid/gel propulsion system. Prior to ignition, GHTP revealed a lower mixing ratio compared to HTP, which ultimately led to an ignition delay time of 20.6 ±6.4 ms, while HTP reached 14.6 ±3.7 ms when 1 wt% of catalyst was used. The flame area, used as a parameter to quantify gas-phase reaction, also showed sensitivity to liquid/gel phase mixing. Using a less concentrated HTP solution, ignition of GHTP/liquid fuel seemed to occur around 6 s faster than liquid HTP/liquid fuel in impinging jets, probably due to differences in propellant residence time and the vapor formation process. A theoretical performance analysis considering GHTP as an oxidizer in a hybrid liquid/gel hypergolic system revealed that the gelled propellant, made with the addition of 6 wt% of fumed silica, can achieve 96% of liquid HTP's specific impulse with the selected fuel. Furthermore, the theoretical density specific impulse of the hybrid gel/liquid system would significantly surpass that of the liquid/liquid system if 10 wt% of aluminum were mixed with GHTP. To take advantage of the hybrid system, a new propulsion system composed of pistons coupled to propellant tanks is proposed to properly infuse the gelled and liquid propellants. The results from this experimental work, together with theoretical analysis, suggest that the hybrid liquid/gel system may be suitable for military applications. [Display omitted] • Use of gelled hydrogen peroxide as oxidizer in hypergolic propulsion systems. • Drop and impinging jets test reveals vapor and ignition delay times with green fuel. • Analysis before and after ignition using visible and infrared high-speed cameras. • Infrared data quantifies mixing prior to ignition and flame spread. • Gelled hydrogen peroxide may ignite faster than its liquid counterpart. [ABSTRACT FROM AUTHOR]

In the development of composite solid propellants, maintaining energy-release efficiency while enhancing the mechanical properties of the propellants presents a significant challenge. Optimizing the interfacial relationships among various propellant components can potentially improve both the mechanical performance and energy-release efficiency of these propellants. This study developed a straightforward method using tannic acid (TA) and Fe3+ ions to create an interfacial layer on ammonium perchlorate (AP) particles (referred to as AP@TA-Fe) for achieving enhanced AP energy-release efficiency and mechanical properties of the propellant. AP@TA-Fe composites with different TA-Fe mass ratios (2, 4, and 6 wt%) were prepared, tested, and characterized. The experimental results demonstrated the interface-engineered AP@TA-Fe displayed enhanced hydrophobicity and rapid energy release, and these effects intensified with increasing interfacial layer thickness. The TA-Fe interfacial layer significantly enhanced the tensile strength and elongation of the propellant under various temperature conditions. As the thickness of the TA-Fe layer increased, the damage mechanism of the modified propellant shifted from the dewetting of the AP particles to the tearing of the adhesive matrix. Consequently, the AP@TA-Fe composites are expected to serve as superior substitutes for traditional AP in propellant applications. [ABSTRACT FROM AUTHOR]

The document is a contents list of the journal "Dalton Transactions: An International Journal of Inorganic Chemistry," which covers various topics related to inorganic chemistry. It includes articles on metal complexes, electrocatalysts, nanocrystals, and organic synthesis. The journal aims to provide original research and perspectives to contribute to scientific progress in the field of chemistry. One article discussed in the document is about the synthesis and characterization of copper complexes that can be used in the degradation of organic dyes using visible light. Another article focuses on the synthesis and characterization of palladium complexes for use in focused electron beam induced deposition. [Extracted from the article]

Energetic Metal–Organic Framework (EMOF) compounds have gained significant attention in recent years as a hot research topic in the fields of explosives and propellants. This article provides an overview of the latest research progress of EMOFs in various areas, including heat-resistant explosives, burning rate catalysts and initiating explosives. It discusses the recent development trends of high-energy EMOFs, such as high-dimensional and solvent-free structural design, simplified and scalable synthesis conditions, environmentally friendly manufacturing processes with tunable structures, high-energy, low-sensitivity and multifunctional target products. The challenges and issues faced by EMOFs in heat-resistant explosives, burning rate catalysts and initiating explosives are presented. Furthermore, the key research directions for future applications of EMOFs in the fields of explosives and propellants are discussed, including solvent-free high-dimensional EMOFs design and synthesis, precise modulation of EMOFs molecular composition and pore structure, improvement of accurate prediction methods for physicochemical properties of high-energy EMOFs, low-cost large-scale production and development of multifunctional composite EMOFs as energetic materials, exploration of influencing factors, and comprehensive study on the application of novel and high-performance multifunctional EMOFs. [ABSTRACT FROM AUTHOR]

New graphene Schiff base iron nanocomplexes (S-792-Fe, S-550-Fe, and S-590-Fe) were designed and fabricated based on investigations of reactions mechanisms of AP pyrolysis, structural analysis of new materials and simulation results of quantum mechanics to deal with the long-standing contradiction between safety and performance for composite propellent combustion catalyst. By taking traditional catalyst Catocene as the benchmark, the S-550-Fe nanocomplexes exhibited competitive catalytic performances, much better anti-migration properties and significantly reduced impact and electrostatic sensitivities. After adding S-550-Fe, the burning rate of propellant enhanced from 13.87 to 19.77 mm s−1 at 15 MPa, and the impact and electrostatic sensitivities are improved to 50.2 cm and 172.12 mJ, respectively. The excellent catalytic performance of S-550-Fe is closely related to its molecular structure, which is manifested by an increase in charge density and a decrease in NH 3 oxidation energy barriers. The promotion effects of S-550-Fe on desorption and oxidation of NH 3 from the surface of AP is conducive to the rapid and complete pyrolysis of AP, which is reflected in the improvement of combustion rate and temperature gradient. The theory-driven design mode of this work may provide new pathways to tune combustion behaviors of composite propellants from the view of molecular-level insights. Effects and mechanisms of S-550-Fe on combustion and safety of AP-HTPB propellant were disclosed. [Display omitted] • Graphen Schiff base Fe nanocomplexes were first investigated in AP-HTPB propellant. • S-550-Fe is conducive to the combustion and safety performance of AP-HTPB propellant. • The structure-activity relationship of S-550-Fe at the molecular level was explored. • The influence of catalysts on NH 3 oxidation process was revealed. [ABSTRACT FROM AUTHOR]

The dielectric barrier discharge (DBD) plasma reforming of methane is ideal for fuel upgrading in rocket engine. However, the working conditions such as high flow rate, high pressure, and low temperature in the propellant supply system significantly affect the reforming performance. This research explores the performance of plasma methane reforming under extreme conditions, examining the characteristics of four kinds of structures, with aims to make reasonable engine design and settings based on actual needs. Simulation is conducted to investigate the effect of plasma reforming on methane ignition delay. The results indicate that prolonging the gas residence could promote methane decomposition. The methane conversion rate under Nanosecond pulsed (NP) excitation is 9.6% higher than that under AC excitation, and the energy utilization efficiency is 1.8% higher. However, the carbon deposition of NP excitation is more severe. As the pressure increases, the cracking reaction intensity decreases by 5%. NP excitation has ideal adaptability to high work pressure. Low temperature has little effect on methane plasma conversion but could affect product distribution, promoting the transformation towards H 2 and C 2 H 6. The ignition delay could be shortened by 71 ms with C 3 H 8 adding. The results of this study provide theoretical guidance for the ignition improving of the rocket engine. • Plasma reforming of methane in attitude and orbit control thrust. • 1.12s is ideal gas residence for plasma methane reforming and limited engine scale. • Working Pressure and Temperature has significant effect on production distribution. • Plasma reforming could shorten the ignition delay by up to 71 ms. • The order of ignition delay shortening effect is C 3 H 8 ＞C 2 H 6 ＞H 2. [ABSTRACT FROM AUTHOR]

Utilizing alternative propellants has been recognized as a strategy to reduce the total cost of propellants in electric propulsion-based missions. The aim of this study is to quantify the Hall effect thruster (HET) performance and operating characteristics using a Kr–Ar mixture to enable mission designers to evaluate the impact on mission and spacecraft design. We present the performance and plume plasma properties of the P5 5 kW-class HET operated with a Kr–Ar mixture with Ar volumetric flow rate fractions from 0 to 100%. The thruster is characterized at discharge power levels of 2.6 kW and 4.1 kW at constant discharge current and voltage over the range of Ar fractions. Despite higher ionization energy and lower mass of Ar, the thruster exhibited a similar level of thrust within 2% when comparing the pure Kr and 26%-Ar mixture cases. The derived ion energy distribution functions and analytical modeling suggest that the characteristic length for the ionization region is extended as the Ar fraction increases. The increased residence time of Kr at the extended ionization region and background energetic electrons from the ionized Ar neutrals are considered to cause this enhanced ionization of the injected Kr neutrals. This leads to a 6% higher Kr ion density at the 26%-Ar case even at the 16% less injected Kr neutral density than in the pure Kr case. The enhanced Kr ionization and generated Ar ions in the 26%-Ar case consequently led to a comparable thrust with that of the pure Kr case. The study indicates that mixing Ar with approximately 26% volumetrically with Kr can provide a similar or even higher thrust performance at the same discharge power. This will be particularly advantageous for various space missions that require high impulses by reducing the total cost of the propellant. [ABSTRACT FROM AUTHOR]

The use of fireworks as a sign of celebration is a common practice in many countries. During the last few years, this sector has been subjected to various social criticisms that have even led to a ban on their use in different places. One of the main causes is the environmental pollution they generate, due to emissions from combustion reactions associated with the use of pyrotechnic devices. Nitrate ester-based compounds as propellants are used in other solid rockets such as ammunition, but their use in pyrotechnics is currently limited to a small number of fireworks. This research analyses the substitution of black powder with nitrocellulose ([C6H7(NO2)3O5]n) as a propellant for aerial articles, as well as to the use of nitrocellulose as a perchlorate-free compound. In addition to achieving better clarity effects due to less smoke, the use of nitrocellulose reduces greenhouse gas emissions. However, this nitrogen-rich compound is more unstable, and the ignition sensitivity is higher. Therefore, the benefits and risks associated with this alternative compound need to be evaluated, but the evidence shown suggests a promising direction for the sustainable evolution of pyrotechnic compositions. [ABSTRACT FROM AUTHOR]

The undesired ignition of ammunitions during their different stages of life can lead to hazardous situations for goods and people. To tackle these potential issues, low-vulnerability ammunitions (LOVA) have been developed. However, this loss of sensitivity can lead to ignition issues. Therefore, the study of ignition phenomena like thermal decomposition is essential. This paper is about the thermal decomposition of three LOVA gun propellants based on RDX and nitrocellulose. Decomposition is studied by thermogravimetric analysis (TGA) between 303 and 573 K, under argon atmosphere and five heating rates (0.3, 0.5, 0.8, 1.0, 1.2 K min−1). A kinetic analysis method, considering ICTAC recommendations, is developed in this paper and applied to TG results. The latter allows to determine global kinetic parameters (pre-exponential factor, activation energy and reaction model based on the truncated Šesták-Berggren equation) to obtain a decomposition law depending on temperature and extent of conversion. Models with three independent parallel kinetic steps are obtained. Their integration provides mass loss curves in good agreement with the TG results. [ABSTRACT FROM AUTHOR]

Recently, because of the complex international situation and combat environment in the future, the development and application of new concept weapons have raised higher performance requirements for manufacturing technologies. However, at present, most weapons are still prepared using traditional charging methods (cast curing, pressure casting, and melt casting), which require subtractive manufacturing (SM) treatments before use. At present, the demand for weapon products is shifting towards reactive micro structures, high preparation efficiency, miniaturization, and controllable energy release. Besides, the modern "energetic-on-a-chip" trend was expected to reduce size and cost while increasing safety and maintaining performance. In this case, the traditional charging methods were not preferred due to their inherent drawbacks, such as being limited to the model, requiring long solvent drying times and recycling required and pores/cracks caused by the shrinkage of slurry, and so on. Therefore, it is necessary to innovate the processing and manufacturing technology of weapons and address the boundaries of existing charging methods, and this will enable the precise customization of high-quality energetic materials and avoid many defects. Additive manufacturing (AM), or 3D printing technology, has been booming recently. The application of additive manufacturing technology in the field of energetic materials (EMs) can promote the innovation of manufacturing technology for EMs and regulate the microstructure. Additionally, 3D printing technology can break through the existing design and development mode, expand explosive charging technology, and enable the distribution of different types of explosives and explosive density in a specific space area. Besides, 3D printing can fabricate "reactive microstructures" (RMS), which offer a deeper understanding of the EMs' combustion and detonation phenomena at the micro- and nanoscale. Thus, the explosive/propellant grains with multiple damage modes can be designed and manufactured. This paper aims to summarize the current progress in the 3D printing of EMs, analyze the corresponding mechanisms, and provide guidance for future research. [ABSTRACT FROM AUTHOR]

The impulse thruster exhibits the advantages of the simplistic structure and rapid response speed. It is often used for the attitude adjustment and orbit control of the aircraft and space vehicle. In order to study the internal ballistic variation law and the flow characteristics of the gunpowder gas, the interior working process of the L-type impulse thruster was investigated by the numerical simulation and experimental verification. First, an internal ballistic test was designed and conducted, and the change process of the thrust and pressure over time was measured, and the feasibility and rationality of the impulse thruster design were verified. Second, a three-dimensional internal ballistic model of the L-type impulse thruster including the ignition process was established. With the help of the dynamic mesh technology and secondary development capability of the user-defined function in the software Fluent, the coupling process of the propellant combustion and internal flow field change of the gunpowder gas was realized. The calculated results were in good agreement with the test data. Finally, the distribution of the pressure and velocity in the flow field was analyzed in detail, and the variation law of the thrust characteristics with the nozzle length and the expansion ratio was studied, which could provide the essential data for further optimization design research. The outcomes from this paper can offer technical means for advancing studies on the internal ballistic changing law of the L-type impulse thruster. [ABSTRACT FROM AUTHOR]

The response characteristics of the bi-swirl coaxial injector under oscillating backpressure were investigated numerically based on coupled level-set and volume of fluid model. The backpressure oscillation frequency was set ranging from 1000 to 5000 Hz. The frequency and amplitude response characteristics of liquid films were analyzed under different injection conditions. The amplitude of the pressure response inside the injector was determined by the acoustic damping of the system. Due to the central gas core inside the inner injector, the inner injector manifold was found to be more sensitive to the backpressure oscillation. The frequency characteristic of the liquid film response indicated the presence of the characteristic frequency where the response amplitude was maximum. This characteristic frequency was linearly positively correlated with the mass flow rate of the inner injector, while the flow in the outer injector could suppress the response amplitude of liquid films. Meanwhile, the nonlinearity of the response was dominated by the inner liquid film and was also the strongest at the frequency with the highest response amplitude. The simulated flow fields showed that the oscillation of the ambient airflow periodically hindered the liquid film, leading to Klystron effect. Finally, it was suggested that the characteristic frequency was related to the axial velocity of the inner liquid film at the injector outlet and the acoustic length of the gas core. In addition, the response amplitude of liquid films was found to be controlled by the spray cone angle and the total mass flow rate of propellants. [ABSTRACT FROM AUTHOR]

Aiming at the current groove blasting problems of additional damage, the feasibility of double base propellant for groove blasting was explored. Based on the propellant gas release behavior, the pressure change of the double base propellant in the closed hole was calculated. Combined with high-speed photography and digital image correlation (DIC) method, two groups of experiments were carried out with propellant loading density of 0.84 and 0.96 g/cm³ to investigate the dynmic destruction process of granite slabs under the action of propellant. The results show that the granite slabs in the two groups of experiments were cracked along the groove direction at 100 μs after ignition, and the cracks penetrated through the slabs at 200 μs; the specimen with a charge density of 0.96 g/cm³ had a larger separation speed between the upper and lower slabs after fracture, and the upper and lower slabs were cracked by the friction of the blocking rubber and the inertia of the specimen, and the cracks were in the vertical direction at 2 500 μs. The grooves around the blast hole provide space for the effect of the propellant gas, and the grooves can effectively guide the direction of crack propagation, no crushing zone formed around the hole wall. The quasi-static pressure generated by the combustion of double-base propellant is the main driving force for crack initiation and propagation. The experimental results have some implications for the use of double base propellant in controlled rock blasting projects. [ABSTRACT FROM AUTHOR]

In order to study the effect of pressure on the ignition and combustion process of CL-20/NEPE solid propellant, the ignition delay, burning rate, and maximum combustion temperature of different solid propellant formulations with an ambient pressure of 0.1~8.0 MPa were measured experimentally by a solid propellant laser ignition experiment system, and the agglomeration process and the characteristics of condensed phase combustion products were analyzed. The experimental results show that, with the increase of pressure, the ignition-delay time decreases, and the burning rate and the maximum combustion temperature increase. With the increase of pressure, the influence on propellant ignition and combustion characteristics becomes smaller. In the experiment, the dynamic agglomeration phenomenon of aluminum particles in the propellant was recorded by a high-speed camera combined with a microscopic camera lens, and the dynamic agglomeration phenomenon of the combustion surface of the propellant and the dynamic agglomeration phenomenon, after the initial agglomeration was separated from the surface, were analyzed and expounded. Based on the experiment and combined with the agglomeration phenomenon, a mathematical model capable of predicting the particle size of aluminum aggregates was proposed. [ABSTRACT FROM AUTHOR]

Incorporating silver wires into propellant has emerged as a highly effective strategy for enhancing propellant burning rates, a technique extensively deployed in the construction of numerous fielded sounding rockets and tactical missiles. Our research, employing a multi-faceted approach encompassing thermogravimetric-differential scanning calorimetry measurements (TG-DSC), combustion diagnoses, burning rate tests, and meticulous collection of condensed combustion products, sought to elucidate how variations in silver wire quantity and winding configuration impact the combustion properties of propellants. Our findings underscore the remarkable efficacy of double tightly twisted silver wire in significantly boosting propellant burning rates under ambient conditions. Moreover, at lower temperatures, the reduced gap between the propellant and silver wire further magnifies the influence of silver wire on burning rates. However, it is noteworthy that the relationship between burning speed and combustion efficiency is not deterministic. While a smaller cone angle of the burning surface contributes to heightened burning rates, it concurrently exacerbates the polymerization effect of vapor phase aluminum particles, consequently diminishing propellant combustion efficiency. Conversely, propellants configured with sparsely twinned silver wires exhibit notable enhancements in combustion efficiency, despite a less pronounced impact on the burning rate attributed to the larger cone angle of the burning surface. Remarkably, these trends persist at lower temperatures. Based on the principle of heat transfer balance, a theoretical model for the combustion of propellants with wire inserts is developed. The reliability of this theoretical model is validated through a comparison of calculated values with experimental data. Our research outcomes carry significant implications for guiding the application and advancement of the silver wire method in solid propellants for solid rocket motors, offering valuable insights to inform future research and development endeavors in this domain. [ABSTRACT FROM AUTHOR]

Laser ablation propulsion based on liquid propellants is a type of propulsion technology with a high specific impulse and good controllability that can be applied to space thrusters, gas metal arc welding, and extreme ultraviolet light. However, its basic mechanisms, such as flow evolution and thrust formation, have not yet been described in detail. In this study, the laser ablation of micro-scale fluid in the atmosphere was investigated. Flow evolution with different laser energy and fluid mass was observed using a schlieren system. According to the characteristic of flow evolution, a theoretical model of laser ablation propulsion in the atmosphere was established. For the first time, a theoretical hypothesis was proposed that the laser energy is divided into two parts, which act on fluid and air respectively. The model indicates that the impulses generated by fluids and air follow power laws with the laser energy, while the exponentials are 0.5 and 1, respectively. In the atmosphere, due to the shielding effect of a laser-maintained detonation wave on laser, the energy absorbed by the fluid is basically unchanged, while only the energy absorbed by the air changes. Significantly, the theoretical model is consistent with the impulse experiment and current studies. [ABSTRACT FROM AUTHOR]

The effect of initial grain temperature on the combustion response of composite solid propellants (AP/HTPB/RDX/Al) is experimentally investigated using a T-burner. Tests were conducted at different frequencies ranging from 200–400 Hz and at different pressures of 1–8 MPa at different initial propellant temperatures using a novel method (303 K, 243 K and 343 K) for two industrial-grade propellants, propellant-A and propellant-B. The test results revealed that the hot and cold temperatures increase the response. The combustion responses of the propellant-B are lower than propellant-A. Pulsed tests are conducted when the propellant showed low or negative response. This work highlights that the initial grain temperature is a crucial factor in determining the stability of solid rocket motors. The negative response compels us to further investigate toward a better understanding about the unsteady combustion behavior since there is no model available which could predict negative responses as of today. [ABSTRACT FROM AUTHOR]

Burning 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. [ABSTRACT FROM AUTHOR]

To address the issue of randomness in the mechanical properties of the hydroxyl‐terminated polybutadiene (HTPB) propellant, a stochastic constitutive model (SCM) with a lognormally distributed random parameter Λ was proposed to describe their mechanical behaviors, and the structural integrity of a HTPB propellant grain was analyzed based on it. The results indicate that the stress‐strain curves predicted by the SCM have a good agreement with the experimental curves, and the experimental curves fall within a 95 % probability interval predicted by the SCM. The mechanical response of HTPB propellant grain under ignition pressurization is associated with the random parameters Λ. The maximum equivalent stress and safety factor increase approximately linearly with the increase of random parameters Λ, while the maximum equivalent strain and maximum damage coefficient decrease approximately linearly with the increase of random parameters Λ. The error in the mechanical response of the grain obtained based on the SCM and the experimental constitutive model is basically not more than 2 %, the SCM can effectively characterize the randomness in the mechanical response of propellant grain caused by the dispersion of HTPB propellant mechanical properties. [ABSTRACT FROM AUTHOR]

In this paper, a launching system with a novel charge structure was proposed to improve the interior ballistic performance. According to the working characteristics of the novel launching system, a lumped‐parameter model was established. The parameters of the propellant and the interior ballistic characteristics of the launching system were obtained by experiments. According to the experimental results, the accuracy of the lumped‐parameter model was verified by code. The simulated results were in good agreement with the experimental results. Based on the lumped‐parameter model, the propellant parameters, such as the impetus, the burning rate, the web thickness, and the charge mass, were investigated to understand the interior ballistic performances of the launching system. An optimization method was proposed to design the propellant parameters of the launching system. The results show that the optimal scheme can increase the velocity of the projectile by 9.54 %. Compared with the traditional launching method, the velocity of the projectile is increased by 37.09 % while the peak pressure in the barrel has no change. [ABSTRACT FROM AUTHOR]

Artillery designers are always searching for means of improving a gun’s external and internal ballistic properties and this resulted in both new theories surrounding gun construction based on the properties of materials and new methods of building guns capable of firing heavier projectiles to greater ranges. The concept of the accelerator gun sought to achieve the same result without the need for larger and heavier guns and reducing recoil through the use of multiple propellant charges fired sequentially, thus accelerating the projectile without the need for a single high-pressure discharge. It was a quest initiated in France, pursued in the United States and then adopted by a desperate Germany during the latter years of World War 2. [ABSTRACT FROM AUTHOR]

Thermal response and chemical reaction of AP/HTPB in flame area of fire are different with that in near flame area of fire. AP/HTPB in flame area of fire is uniformly heated. This process is similar to that in the slow cook-off experiment. Thermal response and chemical reaction in AP/HTPB in near flame area of fire are not uniform, which cannot be characterized by slow cook-off experiment. In this study, a pool fire cook-off model has been established to examine hazard and critical condition of AP/HTPB explosion induced in near flame area of fire. The effects of fire temperatures on the AP/HTPB propellant explosion induced in near flame area of fire are studied using numerical simulation. Critical fire temperature (lowest fire temperature) for inducing explosion of AP/HTPB is 673K at that the initial explosion temperature of 576.4K at which explosion is initiated. At the fire temperature of lower than critical fire temperature, the AP/HTPB explosion cannot be initiated in the near flame area of fire. The ignition delay increases linearly with the decrease of fire temperature. At 673K fire temperature, the ignition delay is 1600s. The initial explosion temperature in the AP/HTPB propellant in the near flame area of fire increases from 736.5K to 790K with the decrease of fire temperature in range from 2000K to 1600K. However, initial explosion temperature of the AP/HTPB decreases from 790K to 576.4K with the decrease of fire temperature in range from 1600K to 673K. The initial explosion temperature in the AP/HTPB propellant in near flame area of fire reaches the maximum value of 790K at 1600K fire temperature. [ABSTRACT FROM AUTHOR]

The backflow of high-temperature products in an engine's combustion chamber is a key issue which can significantly reduce combustion efficiency. This is particularly problematic for hypergolic propellants, as the high-temperature products may still contain fuel or an oxidizer. If either the fuel or the oxidizer backflows into the manifold of the other, it can easily lead to micro-explosions, thereby creating a threat. To address this problem, this paper proposes a new design of an anti-backflow injector, aimed at effectively preventing the backflow of combustion products to the propellant manifold. A steady-state, non-reactive Computational Fluid Dynamics (CFD) model is employed to evaluate the steady internal flow characteristics of the proposed anti-backflow injector. Additionally, a complementary transient, multiphase fluid dynamics simulation is carried out to assess the response characteristics of the anti-backflow injector. Our analysis focuses on the response characteristics of the concave valve core. The study also explores the impact of different expansion port angles on the injection effect, finding that the vortex diameter at the injector outlet is positively related to the expansion port angle. It is also shown that the injection angle becomes more stable as the expansion angle increases. When the expansion port angles are 10° and 20°, the injection angles show similar trends. In terms of anti-backflow effect, taking the injection stiffness of 150% and a time range of 300 µs as examples, the response time of the anti-backflow models with expansion ports of 10°, 15°, and 20° is increased by 67 µs, 99 µs, and 213 µs, respectively, compared to the base models with the same expansion angles. Meanwhile, when the injection stiffness is 50%, the response time of the anti-backflow models with expansion ports of 10°, 15°, and 20° is increased by 207 µs, 210 µs, and 207 µs, respectively. When the injection stiffness is 20%, the response recovery speed of the anti-backflow models with expansion ports of 10° and 15° is increased by 41 µs and 96 µs, respectively. However, the performance of the anti-backflow model with a 20° expansion port is 216 µs of the base model. The optimized design of the anti-backflow injector has potential applications in solving the propellant backflow problem and contributes to the advancement of combustors. [ABSTRACT FROM AUTHOR]

In recent years, air-breathing electric propulsion emerged as a potential enabling technology for long-duration space missions in Very Low Earth Orbit (VLEO). In this work, we show how the complex relation between mission environment, spacecraft configuration, and propulsive performance can be associated to a single requirement for full air-breathing drag compensation, which becomes less stringent for higher VLEO orbits. The impact of the spacecraft shape and size on the performed analysis is then evaluated and the results compared with conventional electric propulsion solutions with stored propellant, showing how the adoption of air-breathing propulsion becomes more advantageous, from a volume and mass fraction perspective, when the spacecraft scale is reduced. • We introduce the total air-breathing electric rocket system efficiency. • A single requirement for full air-breathing drag compensation is identified. • We define the core spacecraft shape parameter to be optimized for air-breathing. • This feasibility requirement is not directly linked with platform size. • Air-breathing propulsion is more competitive, volume wise, at small scale. [ABSTRACT FROM AUTHOR]

Electrostatic–magnetic hybrid ion acceleration has been researched to realize high-thrust density and efficient thrust operations. In this study, a newly designed stagnant ring (SR) was installed in an electrostatic–magnetic hybrid thruster with the nominal anode inner radius of 40 mm to improve the thrust performance, and the effects of SR on the thruster operation of xenon propellant were analyzed. During constant-discharge-voltage operations, the discharge current increased with both electrically conductive (copper) and non-conductive (ceramic) SR materials, yet the thrust increased only with the copper SR, which was equipotential to the anode. At lower values of the radial width (H) of the copper SR, the thrust and discharge current were almost dominated by the propellant flow rate from the anode surface. However, the contribution of the propellant flow rate from the cathode tip on the thrust and discharge current became significant when H ≥ 9 mm. The thrust coefficient in electromagnetic acceleration form exhibited a maximum value at H = 9 mm, whereas that in electrostatic acceleration form monotonically increased when H ranged from 0 to 20 mm. The highest thrust efficiency was obtained at H = 12 mm. The study findings indicate that installing a conductive SR can effectively enhance the thrust performance in electrostatic–magnetic hybrid thrusters. • A stagnant ring was installed in an electrostatic–magnetic hybrid thruster. • Thrust performance was deteriorated with electrically non-conductive stagnant ring. • Thrust performance was improved with electrically conductive stagnant ring. • Propellant from the hollow cathode was utilized for thrust generation. [ABSTRACT FROM AUTHOR]

A detonation in multiphase air-kerosene-boron mixtures is studied experimentally and the initiation and propagation characteristics of multiphase pulse detonation are emphasized. The boron-rich kerosene gelled propellants are taken as the fuel to analyze its detonation characteristics in static premixed combustible gases. The factors affecting detonation initiation including component, atomization, and ignition method of propellant are elucidated respectively. Considering atomization, evaporation and combustion, a small molecular organic compound is employed to fabricate metallized gelled propellants enriched with substantial boron particles. The dynamic atomization of gelled propellant with varying boron content is imaged by means of visual observation devices, and the ejecting parameters are optimized properly from actual initiation conditions. A flame jet by hot turbulent jet is applied to amplify the ignition energy of deflagration to detonation transition in boron-rich gelled propellant, which has not been seen ever. Results show that the ignition energy is positively correlated with initial pressure formed by combustible gas products. The multiphase detonation wave of boron-rich gelled propellant is successfully induced by premixed H 2 /O 2 mixtures, and its detonation propagation characteristic is closely associated with the increase of initial pressure of premixed gas and mass mixing ratio of working medium. The pressure peak and propagation velocity of detonation wave show a considerable improvement with increasing initial pressure of premixed combustible gas and mass mixing ratio, and the highest values can reach up to 8.65 MPa and 2750 m/s respectively. • Detonation initiation and propagation was achieved by using gelled propellant with 40% boron content.. • The dynamic atomization mechanism and ignition method of boron-rich gelled propellant were reported. • The detonation initiation mechanism and implementation model of boron-rich gelled propellant were analyzed. • Factors influencing the detonation propagation characteristics of boron-rich gelled propellant were elucidated. [ABSTRACT FROM AUTHOR]

Advancing the precision of low-gravity liquid propellant mass gauging is a roadmap challenge for NASA and a significant obstacle to expanding human presence in space. We report flight test data for an implementation of a low-gravity liquid propellant mass gauging technology that is both non-invasive and propellant agnostic. The TRIO payload experiment flew twice on the Blue Origin New Shepard suborbital rocket and validated computational models of liquid distribution in microgravity at three different fill levels. The Propellant Refueling and On-orbit Transfer Operations (PROTO) payload experiment also flew twice on New Shepard and acquired mass gauging data before, during, and after a tank drain and fill that simulated in-space propellant transfers. Propellant volume estimates were derived at 1 Hz using the Modal Propellant Gauging (MPG) system. Analytical, semi-empirical, and finite element (FE) model results for predicted mode frequencies are compared to flight data for mode frequencies across a range of fill levels during the simulated propellant transfer. Flight (0-g), ground (1-g) and FEM mode frequencies for both 0- and 1-g are well-predicted by the semi-empirical models for frequency dependence on tank fill level. • Modal Propellant Gauging (MPG) is a non-invasive and propellant agnostic liquid mass gauging technology. • Demonstrated high-accuracy low-gravity mass gauge during liquid transfer between two tanks. • Added mass effect shifts modal frequencies in predictable ways. • Modal Peak Tracking is capable of identifying liquid mass effects in fluid-loaded tanks. [ABSTRACT FROM AUTHOR]

In order to determine the wetting and drying mechanism of Surelease®-Nitrocellulose aqueous coating solution during the coating of small-sized spherical propellant, and thus to provide theoretical guidance for the preparation of the coated spherical propellant. In this paper, combining with the surface wetting theory, the adhesion work, interfacial tension and the spreading coefficient of the coating system on the surface of the spherical propellant were calculated. The infrared spectrometer was used to explore the chemical interaction between the coating layer and the matrix propellant. In addition, based on the emulsion film-forming principle, the drying kinetics of the coating system was investigated in detail, and the influences of nitrocellulose content on the drying process were investigated. Meanwhile, the macro and micro morphology of the aqueous coating solution during the drying process were also investigated with the help of camera and scanning electron microscope. [ABSTRACT FROM AUTHOR]

In this paper, the creep experiments of Nitrate Ester Plasticized Polyether (NEPE) propellant under eight different stress levels are carried out. The creep experimental data is discussed with different constitutive models, and it is found that the four parameters fluid constitutive model and the composite time-hardening constitutive model can be used to fit the experiment data better. The composite time-hardening constitutive model is more suitable to predict the long-term creep behavior based on the short-term creep behavior. The accuracy of the constitutive model is proved by comparing the predicted results with the experimental results. [ABSTRACT FROM AUTHOR]

TKX-50, as a representative of a new nitrogen-rich energetic compound, has attracted extensive interests due to its high energy and low sensitivity properties. As an important energetic material, combustible agent has great influence on improving the energy level of explosives and propellants. Boron powder has become the most potential combustible agent in energetic fields because of its unique performances. In this work, nano-boron-based microspheres containing TKX-50 are directly fabricated by electrospray deposition. The prepared microspheres showed high monodispersity with particle size ranging from ~1–3 μm. Thermal behavior results show that the mass gain of the electrostatically sprayed sample is 163.3%, which is much higher than that of pure boron and physically mixed samples. Combustion behavior tests show that electrosprayed microspheres exhibit outstanding performance in comparison to physical mixed sample, the peak pressure improved by 85%, the burning time decreased by 21% and the pressurization rate improved by 136%. These excellent properties are probably related to their uniform microstructure. Therefore, electrospray preparation of microspheres with controllable structure has a good application prospect in energetic materials. [ABSTRACT FROM AUTHOR]

In order to improve the mechanical properties of HTPB four component and CL-20/CAP propellants and solve the problem of "dewetting" at the interface of the novel propellant filler matrix, a new Schiff base bonding agent was designed and synthesized. The bonding effect and mechanical properties of the propellant were characterized by Fourier infrared spectrosco-py, scanning electron microscopy and universal material tensioning machine. The results showed that the synthesized new Schiff base bonding agent contained the essential functional groups of the bonding agent and had obvious bonding on the AP surface. The new Schiff base bonding agent was expected to replace the toxic MAPO and significantly improve the mechanical properties of HTPB four component propellant, the strength of HTPB four component propellant reached 0. 7MPa, and the elongation was greater than 55% at room temperature. The elongation at high and low temperature are greater than 50% and 60% respectively. The processing performance of HTPB four component propellants can be significantly improved. In CL-20/CAP propellant, the new Schiff base bonding agent can improve the strength and elongation of the propellant at the same time. [ABSTRACT FROM AUTHOR]

In order to explore the thermal decomposition and combustion performances of zirconium powder (Zr) with high-energy oxidant composite, the thermal decomposition behavior, combustion heat and ignition combustion process of four kinds of energetic materials (EMs) of AP, CL-20, HMX, DAP-4 and Zr/EMs samples were compared and analyzed by using differential scanning calorimetry, automatic calorimetry and self-built combustion experimental system. Some combustion products were collected and characterized. The combustion mechanism of Zr/DAP-4 was analyzed. Results show that the thermal stability and combustion heat release of DAP-4 are better than that of the other three EMs samples. The addition of Zr leads to a higher combustion heat and a lower fa value of the sample, the thermal decomposition temperature remains at 373.8°C, and the combustion performance is better than that of pure DAP-4. Compared with the thermal decomposition of Zr/CL-20 and Zr/ HMX, Zr/DAP-4 has a higher initial decomposition temperature and peak temperature, a lower apparent activation energy fa (138. 7 kj/mol), and a higher combustion heat (17 134J/g). The entire combustion time of Zr/DAP-4 composite is only 170ms, indicating a rapid and intense combustion reaction and its flame intensity is also superior to the other three samples. The combustion products of Zr/DAP-4 include solid products such as ZrO2 and gas products such as H2O and HCI. Zr/DAP-4 composite energetic material has higher energy characteristics and better combustion performances, which make it has great potential for application in the formulation design of propellants. [ABSTRACT FROM AUTHOR]

This review paper explores the potential of oil palm biomass as a valuable cellulose source for the production of nitrocellulose-based propellants, contributing to the green revolution and sustainable energy solutions. It highlights the availability of the corresponding biomass in Malaysia and in line with global studies, the chemical compositions, as well as a brief description of current technologies for converting biomass of oil palm into value added products specifically cellulose. Steps to achieve maximum utilization of biomass from oil palm industry for cellulose production and prospective source for nitrocellulose-based propellant are also proposed. The methodology section outlines the pretreatment of lignocellulosic fibres, cellulose extraction, and nitrocellulose production processes. Overall, the review underscores the prospective of palm oil biomass as a sustainable cellulose source for propellant manufacturing, while acknowledging the need for further research and advancements in the field. [ABSTRACT FROM AUTHOR]

Electrospray thrusters are a promising electric micropropulsion technology which could be used to meet the propulsion needs of nanosatellites, or for fine attitude control of larger spacecraft. Multimodal propulsion is the integration of two or more propulsion modes into a system which utilizes a common propellant. Indeed, spacecraft mission simulations and models have shown that this type of multimode propulsion capacity is exciting because of the flexibility and adaptability it provides mission designers and planners. A single spacecraft would have potential to execute drastically different mission profiles, and the exact mission could even be determined post-launch. The current paper investigates a micro-propulsion system which combines a droplet and ion mode electrospray emitter into a unified multimodal system (using an ionic liquid as the common propellant for both systems). The high relative thrust droplet mode emitter was fabricated from P3 borosilicate glass while the high efficiency ion mode emitter, Carbon Xerogel dense porous substrate, was fabricated in-house. To characterize the multimodal thruster, a full beam and time-of-flight (ToF) experimental setup were developed at the RMC Advanced Propulsion and Plasma Exploration Laboratory (RAPPEL) and experiments were conducted using a custom vacuum chamber. The ion mode emitter, with a beam comprised purely of ions had an onset voltage around 1400 V with an estimated thrust performance of 0.14 μ N and specific impulse of 4040 s. For droplet mode, with a mixed beam comprised of around 17 % droplets and 83 % ions, an onset voltage of 1375 V with an estimated performance of thrust at 14 μ N and specific impulse of 140 s were measured. The prototype thruster demonstrates how various electrospray emitters could be combined into a multimodal system to provide flexibility and adaptability in providing effective thrust for small satellites. [ABSTRACT FROM AUTHOR]

This study investigates the evaporation and ignition characteristics of a single droplet of ammonium dinitramide (ADN)-based liquid propellant utilizing a waveguide resonant cavity device, in conjunction with a high-speed photographic imaging system and testing system. Experimental methods are employed to analyze the impact of microwave pulse width and duty cycle on the puffing and meicro-explosion phenomena of the droplet, as well as the delay time and duration of ignition. The experimental findings reveal that increasing the duty cycle enhances the ignition success rate and diminishes flame development time. Specifically, elevating the microwave duty cycle from 60% to 80% reduces the ignition delay time of the droplet from 132.8 ms to 88.1 ms, and the ignition duration from 23.1 ms to 19.9 ms. Furthermore, an increase in microwave energy pulse width expedites the combustion process of the flame and influences plasma generation. Increasing the pulse width of microwave energy from 20 µs to 40 µs prolongs the ignition delay time from 140.3 ms to 200.5 ms and extends the ignition duration from 56.7 ms to 77.8 ms. Additionally, it is observed that a higher duty cycle leads to a more pronounced puffing phenomenon that initiates earlier. In contrast, a higher pulse width results in a more pronounced puffing phenomenon that commences later. This study provides a thorough investigation into the microwave ignition mechanism of ADN-based liquid propellants, offering theoretical insights into the ignition and combustion stability of such propellants in microwave-assisted ignition systems. [ABSTRACT FROM AUTHOR]

The coupling action of gas flow with propellant burning surface regression is the critical process in solid rocket motor (SRM), which has an important effect on the unsteady combustion, erosive burning, dynamic combustion and pressure fluctuation of propellants. The immersed boundary method (IBM) is used to deal with the dynamic boundary of the flow field caused by the burning surface regression, and the face offsetting method (FOM) is applied to track the regression of the burning surface. The combination of these two methods can realize the coupling simulation of gas flow and burning surface regression. In this paper, the mathematical model of the immersed boundary method for all the flow field boundaries and the coupling method of gas flow with the propellant burning surface regression are given. And the implementation method of the immersed boundary method for all boundary conditions in the non-viscous flow field is given. The internal flow fields of the motor with the end-slotted and end-burning propellant and another motor with the perforated cylindrical propellant, considering the erosive burning, are simulated. The numerical results show that IBM, combined with FOM, has effectively realized the simulation of gas flow coupled with the burning surface regression. [ABSTRACT FROM AUTHOR]

Energetic materials, essential for rocket propulsion, play a crucial role in aerospace systems. These materials lack sufficient safety assessments and pose significant risks to life and infrastructure. The slow progress in this field is hindered by prohibitively high experimental costs and the inherently destructive nature of these experiments. This study aimed to evaluate modeling-based safety assessments, primarily utilizing the Lee–Tarver model. It specifically addressed the threat posed by shaped charges during transportation. Experimental analyses employed 50 mm JH-2 shaped charges, targeting RDX-based JH-2 formulation, since RDX is widely used in various rocket motor propellant formulations. TNT was also used to enhance the validation of the Lee–Tarver model findings. Various logistical configurations were explored, testing steel container walls ranging from 10 mm thickness to reinforced walls of 50 mm, 60 mm, and 65 mm steel. Comparative analysis with experiments reinforced the accuracy of the Lee–Tarver model's numerical predictions. Both simulations and experiments affirmed that a 65 mm steel protection suffices for the safe transportation of RDX-based JH-2 formulations and TNT within a 10 mm storage box in logistical setups. Furthermore, this study emphasizes the Lee–Tarver model's reliability for most energetic materials safety assessments, while acknowledging certain limitations. [ABSTRACT FROM AUTHOR]

The application of fluoropolymers in energetic materials has attracted increasing attention. The effect of poly(vinylidene difluoride) (PVDF) content on thermal reaction and combustion performance of MnO2-Al nanothermite was reported. Nanothermite samples were prepared by electrospray with the content of PVDF from 0 wt.% to 30 wt.%. Thermal reaction processes influenced by PVDF content were tested by simultaneous thermal analysis (TG-DSC). Finally, ignition threshold tests were designed, and combustion processes were recorded. The results showed that the components of nanothermite materials were evenly distributed and stable. The content of PVDF had a significant influence on thermal reaction process, which is divided into four stages. Ignition threshold was represented by value of instantaneous ignition current. With the increase of PVDF content, ignition threshold decreased first and then tended to a stable value. Besides, PVDF content had important effects on the combustion speed of the samples. This work has practical significance for designing thermite-based ignition propellant. [ABSTRACT FROM AUTHOR]

To enhance the comprehensive performance of artillery internal ballistics--encompassing power, accuracy, and service life--this study proposed a multi-stage multidisciplinary design optimization (MS-MDO) method. First, the comprehensive artillery internal ballistic dynamics (AIBD) model, based on propellant combustion, rotation band engraving, projectile axial motion, and rifling wear models, was established and validated. This model was systematically decomposed into subsystems from a system engineering perspective. The study then detailed the MS-MDO methodology, which included Stage I (MDO stage) employing an improved collaborative optimization method for consistent design variables, and Stage II (Performance Optimization) focusing on the independent optimization of local design variables and performance metrics. The methodology was applied to the AIBD problem. Results demonstrated that the MS-MDO method in Stage I effectively reduced iteration and evaluation counts, thereby accelerating system-level convergence. Meanwhile, Stage II optimization markedly enhanced overall performance. These comprehensive evaluation results affirmed the effectiveness of the MS-MDO method. [ABSTRACT FROM AUTHOR]